* & 

MCR-86-1329 

NAS8-36609 

DPO-664 

m-6 

Volume I 

Final 

Technical 

Report June 1986 

Executive Summary, Concept Definition 

Study for Recovery Study Results 

of Tu m b I i n g - 

Sat el I i tes 

__ 

(NASA-CR-179228) COBCEPT DEFINITION STUDY N88-14118 

POR BECOVEBY OF TULIBLXNG SATELLITES. ..VOLUHE 

Technical 1: EXECUTIVE Report SU8UAEY, {Hartin STUDY Barietta RESULTS Pioal Unclas 

Aerospace) 212 p CSCL 228 G3/18 0114849 
MCR-86-1329 

Contract NAS8-36609 

DR-6 

DPD-654 

. Volume I 

Final 

Technical 

Report June 1986 

Executive Summaw/ CONCEPT DEFINITION STUDY 

Study Results FOR RECOVERY OF 

TUMBLING SATELLITES 

Prepared for: 

National Amronwrtict and Space Administration 

Marshall Space Flight Center 

Alabama 35812 

Prepared by: 

0. A. Cable 

W. L. DeRocher, Jr. 

J. A. Cathcrrt 

M. G. Keelw 

L. Madayav 

T. K. Nguyen 

J. A. Preesa 

Approved by: 

D. A. Cable 

Program MI-r 

1-A - 

MARTIN MARIETTA 

DENVER AEROSPACE 

P.O. Box 179 

Denver, Colorado 80201 
. 

FOREWORD 

This document was prepared by t4arti n Narietta Corporation under contract 

NAS8-36609, Concept Def i ni ti on Study for Recovery of Tumbl i ng Sate1 1 i tes, to 

fullfil the requirements of Data Procurement Document 654, Data Requirement 

DR-6, final Technical Report, Volume I, Executive Summary and Study Results. 

This effort was accomplished for the George C. Marshall Space Flight Center of 

the National Aeronautics and Space Admi ni strati on under the technical 

direction of Mr. Herbert Lenox and Mr. Stephen B. Hall, as Contract Technical 

Idanage rs. 

ii 
. 

a 

CONTENTS 

Page 

1 . 0 

1.1 

1.2 

1.2.1 

1.2.2 

1.2.3 

1.2.4 

1.2.5 

1.3 

1.4 

1.4.1 

1.4.2 

1.4.3 

1.4.4 

1.4.5 

1.5 

1.5.1 

1.5.2 

1.5.3 

1.5.4 

Title Page ............................................. Foreword ............................................... i 

ii 

Contents ............................................... ii 1 

Figures ................................................ vii 

................................................. xi Tables 

EXECUTIVE SUMlARY ...................................... 1-1 

1-3 

1-4 

1-5 

Recovery System Definition ............................. 1-8 

1-10 

System Hardware Requirements ........................... 1-11 

Concept Definition - Task 2.0 .......................... 1-15 

Subsystem Evaluation Approach .......................... 1-16 

Evaluation of Extendible Booms ......................... 1-16 

Spin/Despin Mechanism Definition ....................... 1-19 

Grapple Mechani sm Interface Device ..................... 1-19 

Grapple Mechani sm Eval uation ........................... 1-21 

Conceptual Tumbl i ng Sate1 1 i te Recovery System 

Design - Task 2.2 .................................... 1-26 

TSR Conceptual Design Drivers .......................... 1-26 

Modular TSR System Design .............................. 1-27 

MMC Envel oper Grapple Mechanism ........................ 1-29 

Conceptual Recovery Systems - Summary .................. 1-30 

Study Objectives and Approach .......................... Requirements Analyses and Trades - Task 1 

Concept Identification and Evaluation .................. 

Satellite Recovery System Mission Model 

Design Reference Missions .............................. 1-13 

Evaluatlon of Recovery System Mechanisms - Task 2.1 .... 1-16 

.............. 

................ 

iii 
CONTEkTS (continued) 

Page 

1.6 Supporti ng Devel opment P1 an ............................. 1-32 

1.6.2 STS Cargo Bay/Proximity Operations ..................... 1-33 

1.6.3 F1 ight Hardware Program ................................. 1-35 

1.7 Supporting Research and Technology Report .............. 1-36 

1 .a Cost Estimate and Work Breakdown Structure ............. 1-36 

1.6.1 Ground Demonstration Activities ........................ 1-32 

2.0 STUDY RESULTS . INTRODUCTION AND BACKGROUND ............ 2-1 

2.1 Study Objectives ....................................... 2-2 

2.2 Study Approach ......................................... 2-4 

2 . 2 . 1 Study Schedul e ......................................... 2-9 

2.1.1 Ground Rules ........................................... 2-4 

3.0 

3.1 

4.0 

4.1 

4.2 

4.2.1 

4.2.2 

4.2.3 

4.2.4 

4.2.5 

4.3 

4.3.1 

4.3.2 

4.3.3 

4.4 

4.4.1 

4.4.2 

REQUIREMENTS ANALYSES AND TRADES . TASK 1 .............. 3-1 

Introduction and Approach .............................. 3-1 

CONCEPT IDENTIFICATION AND EVALUATION . TASK 1.1 ....... Introduction and Approach .............................. Concept Identification ................................. Debris Capture Device ................................... Tel eoperator Retrieval Mani pul ator ..................... Experimental Materials Hand1 i ng Device ................. Docking and Retrieval Mechanism ........................ Space Bola ............................................. Initial Concept Evaluation ............................. Introduction ........................................... Evaluation Criteria .................................... Concept Evaluation ..................................... A1 ternative Concept Identification Approach ............ Failure Mode Analysis .................................. Failed Satellite Motion Analysis ....................... 

4-1 

4-1 

4-4 

4-4 

4-5 

4-6 

4-0 

4-9 

4-10 

4-10 

4-11 

4-13 

4-14 

4-17 

4-18 

iv 
CONTENTS (continued 

Page 

4.4.3 Spacecraft Dynamics Analysis ........................... 4-21 

4.4.4 Recovery System Differentiators ........................ 4-23 

4.4.5 Recovery System Definition ............................. 4-26 

5.0 

5.1 

5.2 

6.0 

6.1 

6.2 

6.2.1 

6.2.2 

6.2.3 

6.2.4 

6.2.5 

6.3 

6.4 

6.4.1 

6.4.2 

6.4.3 

7.0 

7.1 

7.2 

7.2.1 

7.2.2 

7.2.3 

7.2.4 

7.2.5 

7.2.6 

7.3 

SATELLITE RECOVERY SYSTEM MISSION MOGEL . TASK 1.2 ..... 5-1 

Introduction and Approach .............................. 5-1 

Description of Mission Model ........................... 5-2 

SYSTEM HARDWARE REQUIREMENTS . TASK 1.3 ................ Introduction and Approach .............................. Tumbl i ng Sate1 1 i te Recovery System Operations Concept . Introduction ........................................... Operations ............................................. Maintenance/Refurbishment .............................. Operations Control ..................................... Space Station Operations ............................... Functional Analysis .................................... Mission Model Analysis ................................. Introduction ........................................... Derived Base1 ine Definition ............................ Derived Baseline Application ........................... 

DESIGN REFERENCE MISSIONS . TASK 1.4 ................... Introduction ........................................... DRM Operations Analyses ................................ DRM 1 . System A. Case 1 ............................... DRM 2 . System A. Case 2 ............................... DRM 3 . System B. Case 1 ............................... DRM 4 . System B. Case 2 ............................... DRM 5 . System C. Case 1 ............................... DRM 6 . System C. Case 2 

Requirements Allocation ................................ 

............................... 

6-1 

6-1 

6-5 

6-5 

6-5 

6-8 

6-8 

6-8 

6-8 

6-11 

6-11 

6-12 

6-17 

7-1 

7-1 

7-1 

7-1 

7-3 

7-7 

7-12 

7-14 

7-21 

7-28 

V 
CONTENTS (conti nued 

8.0 

8.1 

9.0 

9.1 

9.2 

9.2.1 

9.2.2 

9.3 

9.4 

9.5 

9.5.1 

9.5.2 

9.5.3 

9.5.4 

10.0 

10.1 

10.2 

10.3 

10.4 

10.5 

10.6 

10.7 

Page 

CONCEPT DEFINITION . TASK 2 ............................ Concept Definition Approach ............................ 8-1 

8-1 

EVALUATION OF RECOVERY SYSTEM MECHANISMS . TASK 2.1 .... Subsystem Evaluation Approach .......................... Evaluation of Extendible fiooms ....................... A . Introduction ........................................... Evaluations ............................................ 

Definition of Grapple Mechan'ism Interface Devices ...... 

Introduction ........................................... 

Stinger-Type Grapple Mechanism ......................... Eval uation of Large Envel opment-Type 

Spin/Cespin Mechanism Defini tfon ....................... 

Grapple Mechanism Evaluation ........................... 

Small Gripper Definition ............................... 

Grapple Mechanisms ................................... 

TUMBLING SATELLITE RECOVERY SYSTEM CONCEPTUAL 

DESIGN . TASK 2.2 .................................... TSR Conceptual Design Drivers .......................... Modular TSR System Design .............................. TSR System . Compact Design ............................ MMC Enveloper Grapple Mechanism ........................ 

OMV/TSR Kit Interfaces ................................ Conceptual Recovery Systems . Sumnary ................. 

TSR Capture Envelope Flexibility ....................... 

9-1 

9-1 

9-2 

9-2 

9-2 

9-11 

9-11 

9-1 3 . 

9-13 

9-14 

9-15 

9-16 

10-1 

10-1 

10-2 

10-5 

10-7 

10-1 0 

10-12 

10-1 3 

11.0 SUPPORTING DEVELOPMENT PLAN ............................ 11-1 

11.1 Objectives and Sumnary ................................. 11-1 

11.3 STS Cargo Bay/Proxjmi ty Operations ..................... 11-4 

11.4 Flight Hardware Program ................................ 11-5 

11.2 Ground Demonstration Activities ........................ 11-1 

vi 
FIGURES 

Figure 

1 00-1 

1.1-1 

1.2-1 

1.2-2 

1.2-3 

1.2-4 

1.2-5 

1.2-6 

1.2-7 

1.2-8 

1.2-9 

1.2-10 

1.4-1 

1.4-2 

1 04-3 

1.4-4 

1.4-5 

1.44 

1 04-7 

1.4-8 

1.4-9 

1.4-10 

1.5-1 

1.5-2 

1.5-3 

1.5-4 

1.5-5 

1.6-1 

1.8-1 

Conceptual Modular Recovery System Oesign .......... Study Task Flow ................................... 

Concepts Attached to OMV .......................... Task 1 Clpproach ................................... 

Recovery Scenario Differentiators ................. Recovery Systems A & B Definition ................. 

Mission Model Data Base ........................... TSR Operations Concept ............................ Functional Analysis of System C Recovery .......... 

Requi rements A1 1 ocation Exemplar .................. Extendible Booms .................................. 4-Bar Linkage Evaluation .......................... Manipulator Evaluation ............................ Spin/Despin Mechanism ............................. Ground Assembled Interface fl ange ................. Small Gripper Definition . PFMA End Effectors ..... Stinger-Type Grappler ............................. Debris Capture Evaluation ......................... MSFC Envel oper Eval uation ......................... MMC Envel oper Evaluation .......................... Key Design Drivers ................................ Tumbling Satellite Recovery Kit ................... TSR Grapple Envelope .............................. Conceptual TSR . System B ......................... Conceptual TSR . System C ......................... Supporting Development Plan Schedule .............. Total TSR Kit Funding Profile ..................... 

Recovery System C Definition ...................... 

DRM 6. System C. Case 2 ........................... 

Page 

1-2 

1-4 

1-5 

1-6 

1-9 

1-9 

1-11 

1-12 

1-12 

1-13 

1-14 

1-15 

1-17 

I-la 

1-19 

1-20 

1-20 

1-21 

1-22 

1-23 

1-24 

1-25 

1-26 

1-28 

1-30 

1-31 

1-32 

1-35 

1-37 

vi i 
FIGURES (continued) 

Figure 

2.1-1 

2.2-1 

3.1-1 

.4. 1.1 

4.1 -2 

4.2.1-1 

4.2.2-1 

4.2.3-1 

4.2.4-1 

4.2 . 5-1 

4.4-1 ~ 

4.4.1 -1 

4.4.2 -1 

4.4 . 4-1 

4 . 4 . 5-1 

4.4.5-2 

5.1-1 

5.2-1 

5.2-2 

5.2-3 

6.2.2-1 

6.3-1 

6 . 4.2-1 

6.4 . 2-2 

6.4.2-3 

6 . 4 -3-1 

6.4.3-2 

6.4.3-3 

6.4.3-4 

6.4.3-5 

6.4.3-6 

TSR Study Objectives .............................. Study Task Flow ................................... Task 1 Approach ................................... 

Concepts Deployable from OW ...................... 

Tel eoperator Retrieval Manipul ator ................ Experimental Materi a1 s Hand1 i ng Device ............ Docki ng Retrieval Mechanism ....................... Space Bola ........................................ 

Concepts Attached to OMV .......................... 

Debris Capture Device ............................. 

Concept Identification and Eva1 uation 

Assessment ...................................... Spacecraft Failure Anomaly Analysis ............... Defense Meteorological Support Program ............ Recovery System Di fferentiators ................... Recovery Systems A & B Definition ................. Recovery System C Definition ...................... Mission Model Development ......................... Recovery Candidate Diversity ...................... Mission Model Data Base ........................... Maintenance of Mission Model ...................... TSR Operations Concept ............................ Functional Analysis of System C Recovery .......... Mission Model Configuration Derivation ............ Location of Major Principal Axis .................. Angular Momentum and Spin Rate .................... Mi sal ignment of Principal and 

Geometric Axes .................................. Misalignment of Spin Axes ......................... OMV Controllability Deadband ...................... Extendible Boom Issues ............................ Geometric Coning .................................. Surface Irregularities ............................ 

Page 

2-2 

2-5 

3-1 

4-2 

4-3 

4-5 

4-6 

4-7 

4-8 

4-9 

4-16 

4-18 

4-21 

4-25 

4-28 

4-30 

5-1 . 

5-7 

5-8 

5-9 

6-6 

6-9 

6-13 

6-15 

6-16 

6-18 

6-20 

6-20 

6-21 

6-22 

6-23 

. viii 
FIGURES (continued) 

'a Figure 

7.2.1-1 

7.2.2-1 

7 . 2 . 3 -2-1 

7 . 2 . 4 . 2 -1 

7.2.5.2-1 

7 . 2 . 6-1 

7.3-1 

7.3-2 

7.3-3 

8.1-1 

8.1 .2 . 

8.1-3 

9.1 -1 

9.2-1 

9.2.2 -1 

9 . 2 . 2 -2 0 9.2.2-3 

9.2.2-4 

9.2.2-5 

9.2.2-6 

9.2.2-7 

9.2.2-8 

9.4-1 

9.4-2 

9.5.1-1 

9.5.2-1 

9.5.3-1 

9.5.4 -1 

9.5.4-2 

9.5.4-3 

9.5.4-4 

9 . 5 . 4-5 

Design Reference Mission 1. System A. Case 1 ...... DRM 2. System A. Case 2 ........................... DRM 3. System B. Case 1 ........................... DRM 4. System B. Case 1 ........................... DRM 5. System C. Case 1 ........................... DRM 6. System C. Case 2 ........................... 

Requirements . Allocation (Concluded) ............... 

Concept Definition . Task 2 ....................... Full -Up Recovery System ........................... 

Requirements Allocation ........................... 

DRM Requi rements A1 1 ocation Summary ............... 

Family of Recovery Systems ........................ Recovery Subsystems Evaluation .................... Extendible Booms .................................. Scissors Evaluation ............................... Tubular Extension Eval uation ...................... Tripan Evaluation ................................. 4-Bar Linkage Evaluation .......................... Tel escopi ng Boom Eval uation ....................... Fixed Shaft Evaluation ............................ Manipulator Evaluation ............................ ESAM Evaluation ................................... Ground Assembl ed Interface Fl ange ................. Robotically Operated Interface Device ............. Grapple Mechanism Eval uation ...................... Small Gripper Definition . PFMA End Effectors ..... Sti nger-Type Grappler ............................. Enveloper Mechanisms .............................. Mu1 ti segmented Arm Evaluation ..................... C-Clamp Evaluation ................................ Space Bola Evaluation ............................. Debris Capture Evaluation ......................... 

Page 

7-2 

7-5 

7-11 

7-14 

7-19 

7-23 

7-29 

7-30 

7-32 

8-2 

8-3 

8-5 

9-1 

9-2 

9-3 

9-4 

9-5 

9-6 

9-7 

9-8 

9-9 

9-10 

9-12 

9-13 

9-14 

9-15 

9-16 

9-17 

9-17 

9-19 

9-19 

9-21 

ix 
FIGURES (continued) 

e Figure 

9.5.4-6 

9.5.4-7 

9.5.4-8 

10.2-1 

10.3-1 

10.3-2 

10.3-3 

10.4-1 

10.5-1 

10.5-2 

10.7-1 

10.7-2 

10.7-3 

11.2-1 

11.3-1 

11 -4-1 

MSFC Envel oper Evaluation ......................... MMC Enveloper Evaluation .......................... Envel opment-Type Grappl er Evaluations . Summary ... Tumbl ing Sate1 1 ite Recovery Kit ................... 

MMC Enveloper Grapple Mechanism ................... 

TSR Grapple Envelope .............................. 

TSR . A Compact Design ............................ 

TSR . OMV Risk Reduction .......................... 

MMC Enveloper Target Capture Envelope 

Flexibility ..................................... 

Conceptual TSR . System B ......................... Conceptual TSR . System C ......................... Final Recovery System Eval uation .................. Ground Demonstration Schedule ..................... Cargo Bay/Proximi ty Operations 

Conceptual TSR with MK Envel oper ................. 

Experiments Schedule ............................ Development Program Schedul e ...................... 

Page 

9-22 

9-23 

9-24 

10-3 

10-5 

10-6 

10-7 

10-8 

10-1 0 

10-1 1 

10-1 4 

10-1 6 

10-1 8 

11-3 

11-6 

11-7 

X 

. 
TABLES 

Tab1 e 

1.2-1 

1.6-1 

1.6-2 

2 02 . 1 -1 

4.3.2-1 

4.4-1 

4.4.4-1 

6.1-1 

6.1 -2 

6.1-3 

6.4.1-1 

7.2.1 -1-1 

7.2.1 -2-1 

7 -2.2.1 -1 

7.2.2.2-1 

7 . 2 . 3 . 2 -1 

7 . 2.4.1 -1 

7 . 2 . 4 . 2-1 

7.2.5.1 -1 

7 . 2 . 5 . 2 -1 

7.2 . 6.1 -1 

7 . 2 . 6.2 -1 

10.1-1 

10.6-1 

11 02-1 

11.3-1 

. Page 

Recovery System Evaluation ........................ 1-7 

Ground Demonstration Approach ..................... 1-33 

cargo Bay/Proximi ty Operations Experiments ........ TSR Study Schedule ................................ 2-10 

Concept Selection Criteria ........................ 4-11 

Recovery System Evaluation ........................ 4-15 

Fail ed Sate1 1 i te Motion Orientation 4-24 

Mission Operations Requirements ................... 6-2 

Mi ssi on Functional Requi rements ................... 6-3 

Mission Performance Requirements .................. 6-4 

Mission Model Derived Base1 ine .................... 6-12 

Pre-Mission Tasks ................................. 7-3 

DRM 1 Mission Event Sequence ...................... 7-4 

DRM 2 Pre-Mission Activities ...................... 7-6 

DRM 2 Mission Activities .......................... 7-7 

DRM 3 Mission Activities ........................... 7-10 

DRM 4 Pre-Mission Activities ...................... 7-13 

DRM 4 Mission Activities .......................... 7-15 

DRM 5 Pre-Mi ssion Activities ............. .'. ....... 7-18 

DRM 5 Mission Event Sequence ...................... '7-20 

DRM 6 Pre-Mission Activities ...................... 7-23 

DRM 6 Mission Activities .......................... 7-25 

Key Design Drivers ................................ 10-1 

OMV/TSR Kit Interfaces/kcomnodatfons ............. 10-12 

Ground Demonstration Activities ................... 11-2 

Cargo Bay/Proximi ty Operations Experiments ........ 11-4 

1-34 

............... 

x i 
I 

1 .O EXECUTIVE SUMRY 

This Tumbling Satellite Recovery (TSR) study report documents the first 

assessment of the design requirements and the conceptual definition of a 

."front end kit" to be transported on the currently defined Orbital Maneuvering 

Vehicle (OMV) and the Space Transportation System (STS) Shuttle Orbiter, to 

conduct remote, tel eoperated recovery of disabled and noncontrol 1 ab1 e, 

tumbl i ng satel 1 i tes. Studies re1 ated to recovery of disabled satel 1 i tes or 

space debris have been conducted by NASA and DCD with academic and industry 

assistance for over 20 years. None of the literature data quantified the 

dynamic characteristics of a tumbling sate11 ite, nor did they appear to 

address the full spectrum of TSR system requirements. This study investigated 

both aspects with useful results. 

The study group conducted a thorough review of the prior recovery-related 

studies and actual hardware development efforts to identify a concept for the 

design effort. Five candidate concepts were selected and evaluated as 

potential recovery concepts, using a set of preliminary evaluation criteria. 

The assessment revealed that none of the candidates truly had all the elements 

@ needed to do the task. 

The principal reasons behind this were: (1) typical "complex" satellite 

motion was not defined; and (2) a remote recovery system as a front-end kit 

for a well defined Orbital Maneuvering Vehicle was not conceptualized. This 

study conducted an explicit analysis to determine the most likely tumble mode 

presented by a non-controll ab1 e, disabled tumbl ing satel 1 i te. 

"complex" motion was estimated to most likely be flat, single axis spin about 

the major principal inertial axis. This implies that a recovery system could 

have to deal with a fifteen-foot diameter, 30-foot long spacecraft, with 

33,000 pounds mass that would be spinning end over ena about a transverse 

axis, a quantified challenge that can be answered successfully. 

Sate1 1 i te 

An operations concept was developed and provided a solid frame of reference 

for the description and operations analysis of a set of six Design Reference 

Missions (DRMs). 

requirements were developed and allocated, on a top level basis, to a group of 

In addition, a set of functional and operational 

@ recovery "subsystem" accommodations. 

1-1 
Rather than design a wide range of individual recovery systems to satisfy the 

six DRMs, the study team selected a modular system readily configured into 

recovery systems capable of being tailored for specific mission scenarios. As 

shown in Figure 1.0-1, the selected design concept provides a modular system 

architecture that is composed of a number of subsystem mechanisms. Thus, a 

large disabled satellite in a flat spin beyond the range of the Orbiter would 

require a full-up recovery system, as shown in Figure 1.0-1, that has an 

extendible boom, a spin table to match rates, and a large envelopment-type 

grapple mechanism to allow capture and rigiditation for transport back to the 

Orbiter. 

with an extendible boom and a small gripper and thus save launch weight. The 

recommended conceptual recovery system can serve as a well-founded focal point 

for future study efforts. 

However, a controll ab1 e, non-spinni ng sate1 1 i te could be recovered 

In addition, a Supportfng Research and Technology Report recommended four 

specific areas for studies and hardware development to'address recovery system 

technology issues. 

program cost estimate of 18 million in constant fiscal year (FYI 84 dollars. 

Finally, a TSR kit cost analysis established a total 

I 

OM" 7 

Exmndibh Bean 

'-\ Tv-3 

\ I 

/ 

Interface Maehmian \ 

\ I 

\ 

' Figure 1 .O-1 Conceptual Mc Idul ar Recovery System Design 

1-2 
1.1 Study Objectives and Approach 

There were two major Objectives out1 ined for the tmbl ing satellite recovery 

study. The first of these objectives was to develop realistic candidate 

recovery systems to support the NASA and MSFC decision making process in 

preparing to develop an operational capability for remote recovery of disabled 

satellites, using the OMV as a transport vehicle. A second major objective 

was to define the full range of remote recovery capability required. The 

first objective may be restated in this form: 

related work in this area, to evaluate new concepts and to focus these efforts 

into a channeled conceptual framework that would lead to cost effective 

development of the remote tumbling satellite recovery kit. An alternative 

form of the second objective is to examine a broad perspective of potentipl 

recovery scenarios and to define the full range of required remote, disabled 

sate1 1 i te recovery capabil i ty. 

to review all known previous, 

The MSFC had identified three generally defined levels of capability and 

requested MMC to establish the rationale for an increasing level of capability 

and then to provide designs for the requisite systems. These are: (1) System 

A, the basic OMV, using the NASA baseline configuration. By determining the 

limits of inherent OMV recovery capability, it was anticipated that the 

boundary between that level of capability and the next would be more easily 

determined; (2) System 8, the basic OMV with some minimal hardware addition or 

changeout, such as end effectors, batteries, or special avionics hardware; (3) 

System C, a "full-up" capacity to recovery satellftes with "complex" motion, 

which was undefined at that time. 

@ 

The study team proceeded to develop the rationale for differentiating these 

levels of capability, to define requirements for each level, and to provide 

conceptual designs for MSFC's recovery Systems 8 and C. 

The approach used for the TSR kit study is outlined in Figure 1.1-1. 

approach was centered on conducting the four major tasks outlined in the 

contract statement of work. The interrelations of these tasks are also shown. 

The 

1-3 
Figure 1.1-1 Study Task Flow 

1.2 Requirements Anal’yses and Trades - Task 1 

The overall objective of Task 1 was to perform the type of analyses and trades 

that -- woul d enable i denti ficati on of recovery system requi rements for the broad 

range of recovery systems. An associated objective was to develop the 

supporti ng rational e and the system di fferentiators that would clearly 

delineate the level of capability required for each of Systems B and C. 

approach used in conducting Task 1 is shown in Figure 1.2-1. The major 

objectives of each of the four. subtasks are shown, together with the 

The 

interaction between each of them. 

1-4 
1 I 

RECOVERY CONCEm 

IDENTIFICATION 81 e EVALUATION - TASK 1.1 

RECOVERY CANDIDATES 

(TSR MISSION MODEL) - TASK 1.2 

I I 

RECOVERY SYSTEM REQUIREMENTS - TASK 1.3 

- OPERATIONS CONCEPT 

- FUNCTIONAL ANALYSIS 

- MISSION MODEL ANALYSIS 

F 

DESIGN REFERENCE MISSIONS - TASK 1.4 

- SYSTEMS A, B, & C 

- OPERATIONS ANALYSIS 

- REFINED REQUIREMENTS 

J 

Figure 1.2-1 Task 1 Approach 

1.2.1 

was to conduct a thorough survey of prior related recovery studies and 

hardware development efforts to obtain an understanding of remote recovery 

requirements. It was expected that i d e n t i f i c a t i o n o f a number of viable 

concepts would evolve from this process, and evaluation of tnee concspts 

would provide a set of concepts for concept definition in Task 2. An example 

of some concepts identified in the survey effort is snown in Figure 1.2-2. 

Concept Identification and Evaluation - The approach used in this task 

From the large number of initial concepts, five were selected for further 

evaluation. Evaluation criteria vere selected based on insights and 

background provided by pre-proposal efforts, and new perspectives gai ned from 

tk concept identification surwy and preliminary assessments. The rssults of 

the concept evaluation are presented in Table 1.2-1. The selection critzria, 

wighting factors and total scores for the concept evaluation are provided in 

the table. The results were not surprising, as the three top rankgd 

"concepts" all appeared to possess more of what the study team was beginning 

to understand as requirements for a "full-up" System C recovery device. 

1-5 
However, by simple addition, one could note that the maximum score any concept 

could receive was 710 points. 

only 80% of that score, and it became clear that more effort should be 

expended to clearly define the overall recovery problem. 

The highest rated recovery concept received 

A new and successful approach to addressing these issues was initiated and 

completed. Thi s approach included an analysi s of hi storical fai 1 ure mode 

data, collection of tumble mode data from actual failed satellites, and a 

dynamics analysis to define and validate expected tumble motion. 

Figure 1.2-2 Concepts Attached to OMV 

1-6 
Table 1.2-1 Recovery System Evaluation 

Experimental 

Teleoperator Nbtwiais Docking 81 

Weighting Grapple Despin Hurdling Retrieval Space 

Factor Device Owice Mechanism Bola 

EVALUATION 

CRITERIA 

Debris 

Capture 

Device 

Minimum Risk to OMV & Recovery System 

An excellent source of satel 1 i te fai lure data was obtained at Goddard Space 

Flight Center (GSFC). 

the cognizance of the GSFC and the Jet Propulsion Laboratory (JPL) during the 

1977-1984 time period. 

facts. First, tMt a typical anomaly would result in major mission loss only 

1% of the time. (However, major on-orbit failums have occurred recently and 

are to oe expected.) A second major finding was that power and attitude 

control and stabilization (ACS) subsystem failures accounted for 30% of the 

satellite subsystem anomalies. This fact implies that approximately 30% of 

remote, disabled satellites could be non-controllable from the ground and 

tumbling or spinning in some undefined manner. A third major consideration 

was that nearly 35% of satellite anomalies wee related to the payload or 

experiment package. 

controllable from the ground and completely stable. 

It examined anomalies of 44 unmanned spacecraft under 

An assessment of this data base established three 

Tnus, a remote, disabled satellite could be totally 

Thus, a full scale recovery system should be capable of recovering: 

non-spinning satel lites; stable, spin-stabilized satel lites; and completely 

non-control led, tumbl i ng/spi nni ng satel 1 i tes. 

stable, * 
Next, it became necessary to define typical non-controllable, tumbling or 

"complex" satellite motion to qualify the upper range of recovery System C 

capability. A search was made for satellites that had failed in a noncontrollable mode, and for which satellite orientation mode and rate data were 

available. 

orientation motion was revealed to be similar in all cases. An external 

torque created angular momentum levels that resulted in a spin-up of the 

satellite in a general tumble mode, with spin, precession and nutation 

components. This general tumble motion quickly converged to flat, single axis 

spin about the major principal inertia axis. 

this scenario of satellite motion after a major disturbance was confirmed by a 

dynamics analysis and discussions with dynamics experts outside of Martin 

Mari etta . 

A set of six such cases were identified and the resultant 

The representative nature of 

The sate1 1 i te failure envi ronment analysis di rectly supported a final 

definition of a set of differentiators (Figure 1.2-3) that led to an enhanced 

definition of MSFC's Systems A, 6, and C. 

listed and those applicable to each system are check marked. 

System A scenario is defined as dealing with a target that is beyond Orbiter 

The scenario differentiators are 

For example, the 

range, the target's attitude is stable, and the target has .properly situated 

recovery support elements (RSE), as indicated on the figure. This implies 

that the target is not spinning, or is spinning so slowly that OMV can match 

the spin rate. The RSE must be fully accessible to a head-on approach with 

the OW RMS end-effector, and not obstructed by arrays or antennas. 

conditions exist, the System A scenario exists and the satellite could be 

recovered by the basic OMV. 

B and System C. 

If these 

Similar scenarios were developed for both System 

1.2.2 Recovery System Definition - The overall intent of this series of 

analyses was to develop a rationale for the development of a logical family of 

satellite recovery systems, with well defined levels of capability. 

paragraph 1.2.1, the recovery system differentiators provided a framework for 

defining the scenarios. Figure 1.2-4 highlights the scenarios and 

descriptions for Systems A and 9. 

In 

a 

1 -8 
1. Within Reach of OMV Grapple Devices 

2. STS RMS Grapple Fixtures or FSS Berthing Latches 

3. Beyond Grapple OMV Rate-Matching Capability 

Figure 1.2-3 Recovery Scenario Differenti ators 

Scenario 

Beyond Orbiter Altltude 

Target. Controllable 

Accessible OMV Recovery Support Elements 

TSR Kit Confiaurations 

- Basic OMV 

-dud RMS ' 

End Effector 

KIT WITH: - Extendible Boom 

- Grapple Mechanism 

Interface. Device 

- Small Gri er 

- Grapple fl 

Figure 1.2-4 Recovery Systems A B Definition 
System A is defined as a basic reference configuration OMV, with its 

capabil i ty to dock to unobstructed RMS grapple fixtures and to flight support 

structure (FSS) latch pins. The capability of the OMV for capture operations 

is restricted so it is apparent that to accomplish a System A recovery 

I successfully, the target will have to be relatively stable. 

For System 8, the typical recovery scenario is shown in Figure 1.2-4. 

mu1 tiple degree-of-freedom manipulator arm is viewed as essential to a1 1 ow the 

OMV access to obstructed grapple fixtures and to align the captured target's 

center of mass with the OMV orbit transfer thrust vector. A second element is 

a grapple mechanism interface device to allow the ready exchange of various 

grapple mechanisms. A third element is a small gripper mechanism to enable 

grapple of small "hard points" on satellites that have no recovery support 

el ements . 

A 

The definition of a full-up System C is provided in Figure 1.2-5. 

the recovery target is beyond Shuttle range and is spin-stabilized between 5 

and 50 RPM. 

an extendible boom, a spin/despin mechanism to match rates with the target, a 

grapple mechanism interface device, and a "stinger" type grapple mechanism. 

For Case 1, 

In this situation, the recovery system configuration will incluae 

The more difficult System C scenario is Case 2, where the satellite is not 

controllable. 

of Case 1. A spin/despin mechanism is required to match the spin rate(s) of 

the target. The tumble or spinning configuration of the satellite, however, 

will require a large gripper or large envelopment type grapple mechanism to 

provide a firm, smooth grapple of the target, to rigidize the attachment and 

then maintain the grip to enable despin and stabilization of the target. A 

representative system configuration is shown in Figure 7.2-5. 

I The system capability for System C, Case 2, is similar to that 

1.2.3 Satellite Recovery System Mission Model - The mission model, developed 

specifically for this study effort, was composed of specific spacecraft data 

re1 ated to potential recovery missions, for sate1 1 i tes expected to be on-orbi t 

in the mid-1990s. The model provided the basis from which to design a 

recovery system capable of accomnodating the wide diversity of target size, 

shape, mass and inertia distributions. 

1-10 
1 Systam C-Full Up Capability i 

- Extandiblr Boom 

- Spin/Drpin Mechanism 

- Grapple Mechanism Interface Device 

- Large Envelopment Grapple Mechanism 

- Bamsiaht TW , - Stingor Type Grapple Mechanism 

- Cw l-Controllable Targets 

- Beyond Orbiter Range 

- StablsGround Controllable 

- Spin-Stabilitad-10-50 rpm 

- Case 2-Noncontrollable Targets 

- Within OMV Range 

- Noncontrollablr 

- Target Tumbling/Spinning 

Grapple Fix! 

Spin TIDIE, &Q 

<,' .* * i. _I \* E IS 

OE POOR QUALIT% 

Spin-12 rpm 

OMV 

Extendiblo- '-Stinger Grapple Mechanism 

Probe 

Figure 1.2-5 Recovery System C Definition 

The specific satellite data collected in developing the model reflected a wide 

diversity in target size, mass and configuration. This is illustrated in the 

mission model data base, shown in Figure 1.2-6. The early stages of 

development of many of these programs is reflected in the absence of technical 

information in portions of the matrix. However, the data base was considered 

to be representative of what is to be expected in the mid-l990s, and provided 

the basis from which to develop a multi-purpose recovery system. 

@ 

1.2.4 System Hardware Requirements - Recovery system requirements were 

generated and refined through a series of independent analyses that included 

an operations concept, functional analysis, and mission model assessment. 

The generation of a preliminary concept of operations supports the development 

of a broad range of mi ssion operations requi rements, i ncl udi ng those re1 ated 

to ground and aerospace support equipment. The operations concept is outlined 

in Figure 1.2-7. 

storage facility and assembled, in modular form, for each specific mission. 

It will be mated with the OMV in the Orbiter cargo bay during launch 

processing, deployed into a standby orbit by the Orbiter and transferred to 

the disabled satellite by the OMV. 

In general terms, the TSR kit will be stored in the OMV 

a 

1 -1 1 
ORIGINAL PAGE IS 

OB XIOR QUALITY 

Ixy 2 lyt 2 Irr 2 ARRAY 

HST 25.027 590 174 174 22.386 56.273 

INTELSAT 

VI 4,961 304 151 151 6.197 5,103 

' Irr 2 ANENN 

lsLdm 

125,000 

5,091 

- 

- 0 

- 71,750 

135 72.5 

0 30 

08 

54 

- 57,384 

20 

-58 

- 5,953 

-3,150 -35 330 216 

527 -13 

165 172 

2.91 5 

4.175 

TSR KIT MISSION .LALMcHTo OMVTTSR KIT 

STORAGE * PREPARAllON ' ' NOMINAL ' CARGOBAY 

RETRIEVAL ACTIVITIES STS ORBIT DEpLoyMpcT 

- 25.369 

* 

t . OMVTTSR MISSION FwxlvERy MISSION TRANSIT KIT 

4 TRANSITTO + CoNlROL- MISSION -@ cONlRoL-@ TO 

TRANSFER TRANSFER ORBITER TARGET 

-om UU' 

-206 1 -159 1 325 1 1 206 1 

-r 

TSR KIT 

OMVTTSR KIT ORBITER b DISASSEMBLY, 

ECOVERED 

-2719 I -471 I 1121 I I 

' CARGOBAY . mRN . REFURBISHMENT 

4 SATELLITE 

CARGO BAY 

STOWAGE STOWAGE - TO EARTH &=ORAGE ~ 

Figure 1.2-6 Mission Model Data Base 

+ 

I 

Figure 1.2-7 TSR Operations Concept 

ORIGINAL PAGE IS 

OE POOR QUALITY 

1-12 
A functional analysis of a full-up System C recovery operation was conducted 

to support development of mission functional requirements. This analysis is 

high1 ighted in Figure 1.2-8. The recovery operation commences with the 

disabled satellite in visual contact. The OMV/TSR kit operators conduct the 

sequence of operations shown, under tel eoperated control from the ground 

station. Once the target has been rigidized, or firmly grappled for transport 

back to the Orbiter, the OMY operator will execute short firings of the lower 

thrust RCS engines to ensure the new mass is aligned with the OMV thrust 

vector, and the OMV/TSR kit/disabled satellite will be transported to the 

Orbiter. 

and mission requirements. 

These analyses produced a detailed set of functional , operational 

STABILIZE 

OMVITSAI 

-Allow "Sys1.m" to 

Roach Ylnlmum Enorgy 

Stabllb "Syatam" 

-Apply OW Thmatm to 

Figure 1.2-8 Functional Analysis of System C Recovery 

1.2.5 Design Reference Missions - A set of six Design Reference Missions 

(DRMs) were selected and approved by MSFC. Two DRMs were chosen for each of 

the three designated levels of capability, i.e., Systems A, B, and C. An 

operations analysis was completed for each of the DRMs, to expand and refine 

the growing requirements data base. The DRM operations analyses included a 

breakout of: (1) required "pre-mission" activities; (2) specific or direct 

mission activities; and (3) post mission activities. The last DRM, DRM 6, was 

1-13 
defined and developed to describe the most challenging recovery mission. 

DRM 6 involves recovery of the Upper Atmospheric Research Satellite (UARS), a 

10,000 pound disaoled satellite in a 324-nautical-mile circular orbit, at an 

inclination of 57 degrees. 

uncontrolled, flat, single axis spin and requires a full-up System C recovery 

system. 

The UARS, as shown in Figure 1.2-9, is in an 

The identification of DRMs that encompass tne full range of required wovery 

system capabi 1 i ty, and the subsequent analysis of recovery preparation and 

execution acti vi ties, greatly expanded the evolving set of recovery system 

requi rements. 

Grapple Mechanism ,/ ( S pace% I a) 

W 

Figure 1.2-9 DRM 6, System C, Case 2 

A final step in the requirements analysis phase of this study was the 

allocation of requi rements to recovery system design accommodations. 

level requirements, generated from all the Task 1 analyses and trades, wre 

grouped into categories that related to specific recovery subsystem design 

accommodations. This 

allocation process clearly suggested the need for a specific set of subsystem 

mechanisms and elements i ncl udi ng an extendible boom (perhaps with multiple 

The top 

Representative a1 locations are shown in Figure 1.2-1 0. 

1-1.4 
degree-of-freedom articulation capability) , a spin/despin mechanism, a grapple 

mechanism interface device, a variety of grapple mechanisms, a boresight or 

proximity television system and adequate 1 i ghti ng. fhi s a1 1 ocation supported 

the development of a system architecture and design concept for the second 

major study task, Concept Definition. 

@ 

Shall Recover Sltellltor with Recesredl0brtruct.d Extondible Boom 

RSFr (Due to Deployed BoOmr, AntonMS) (MuilCOOF Manlpulator Arm) 

Shall Rocover Sateilltor with Protuberances 

Ertondlng out from Satellit. Envelope 

Shall Mlnimlre Risk to OW 

Shall Mlnlmire Rlrk to TSR 

ShaU Recow Sltrlllter In Not+SL.bh, Tumbk Mod. 

Extmdlblo Boom 

Extmdlble Boom 

Extmdlbb Boom 

Shall k Clprbl. Of Remote Re.llgmmt Of Target CG Extendlblo Boom 

Prlor to Trurrporl (MuitCOOF M8nlpUiltOr Arm) 

shrli Wtch Spin R8U Of COfttrdlrbk Spln 

shall Accommodate SlUlIlU Spin R8tn from 0 to 55 RPM 

' Shall Wtch SpinlTurnble R81.r d WorrCOftWollabk 

mil Match Spin Rat.. with TBD Aecuny tu Accomplish 

5.telllt.r with TED Levd 

Spin T8bk 

Spln Tabk 

sprn Tabk 

Spin Tabk Armmy 

spinnlng 18rg.t 

EmrlOpnnn(. Gnpw 

ShaU k Caprbk of 0.rpl SplnlD.rpbl M.ch8nlm 

d Angukr Momonturn mthout ""8 am8ga to TSR, OMV, md 

Tugat 

Figure 1.2-10 Requirements Allocation Exemplar 

1.3 Concept Definition - Task 2.0 

The objective of the second major study task, Task 2, Concept Definition, was 

to produce a set of conceptual designs for MSFC Systems B and C that would 

serve as a focal point for continuing design and development efforts aimed at 

creating a front-end kit for the OMV to remotely recover disabled satellites. 

The initial approach for this task was to use the results of the initial 

concept survey to select the best of the concepts, or any other new concepts, 

and refine these concepts .into meaningful , effective conceptual system 

designs. 

became clear that none of them could satisfy the widely variant requirements 

During the evaluation of those previously developed concepts, it 

c 

that were evolving out of parallel analysis tasks. e 

1-1 5 
Thus, the study team, with the approval of the MSFC study technical director, 

restructured Task 2.1 as an identification and evaluation of candidate 

recovery subsystem elements. These subsystems included an extendfbl e boom, a 

spin/despin table, a grapple mechanism interface device, and a set of grapple 

mechanisms for different recovery scenarios. 

For Task 2.2, Conceptual Design, the preferred subsystem components were 

selected and a set of recovery systems were designed for Systems B and C, each 

with the capability to conduct both of their respective design reference 

missions. In addition, the MMC study team developed a new design for an 

envel opment type grappl e mechanism, the MMC envel oper. 

1.4 Evaluation of Recovery System Mechanisms - Task 2.1 

1.4.1 

maJor subsystems was conducted. 

discussions with mechanisms experts at MSFC and MMC produced the initial 

candidates. 

design requirements, all were evaluated qualitatively. By looking at the 

apparent advantages and disadvantages of each candidate, more evaluation 

considerations were derived. The study team did not consider it productive to 

develop new mechanism concepts, with the major exception of the envelopment 

type grappler, as many generic mechanisms for each component were available. 

The evaluations are presented in a format which describes the candidate and 

high1 ights the advantages and disadvantages of each. 

Subsystem Evaluation Approach - A survey of candidates for each of the 

Industrial brochures, 1 ibrary inqui ries, and 

Though most were viewed quickly as falling short of perceived 

1.4.2 Evaluation of Extendible Booms 

Shown in Figure 1.4-1 are examples of some of the extendible boom concepts 

identified and evaluated during this phase of the study. 

considered a wide range of alternatives to avoid unsupportable elimination of 

questionable candidates. 

evaluation included: the need to be compact when not deployed (to minimize 

Shuttl e cargo bay del i very space/cost) , articul ati on capabi 1 i ty to provide 

access to obstructed recovery support elements (RSE) and to realign the target 

center of mass with the OW orbit transfer thrust vector prior to orbit 

transfer, and capacity to accommodate grapple/despin force and torque loads. 

The study team 

Some of the driving design requirements in this 

1-16 
Manipulator Arm Fixed Shaft 

Scissors Mechanism 

Tripan &Bar Closed Tubular 

Linkage Extension Boom 

Figure 1.4-1 Extendible Booms 

A total of eight extendible boom concepts were identified and evaluated for 

application to the TSR kit. For this Executive Sumnary, evaluations of the 

two best concepts are presented. 

The 4-bar linkage illustrated in Figure 1.4-2 was previously developed by MMC 

for space application. The linkage was designed so that when extended, the 

end to be attached to a grapple mechanism will travel in a linear path 

perpendicular to the base plate. 

The 4-bar linkage offers a non-complex accommodation to the recovery system 

extension requirement. 

capability is limited in bending and torsion. The tradeoff of increased 

weight for increased load carrying capability will not, however, create a 

significant increase in compacted volume. The I-bar linkage, like the Tripan, 

could be considered for the TSR kit in the absence of the articulation 

requirement for access to obstructed grapple fixtures and a1 ignment of target 

center of mass prior to orbit transfer. 

As a light weight mechanism, its load carrying 

0 
ADVANTAGE 

* Compact 

Light Weight 

Can Support Grapple, Derpin Torque Loads 

Non-Complex Control 

DISADVANTAGE ~- 

lnsufflclent Articulation for Target Realignment 

Insufficient Articulatlon .for Access to 

Obstructed Recovery Support Elements 

Figure 1.4-2 4-Bar Linkage Evaluation 

The robotic manipulator is the most complex of the extendible boom candidates 

evaluated, but provides the most capability. 

Figure 1.4-3 was previously developed by MMC, and it is typical of robotic arm 

design. A manipulator arm enjoys the advantage of being a multiple 

degree-of-freedom (DOF) device, capable of providing target realignment and 

access to obstructed grapple mechanisms. The MMC design includes four DOF, as 

a result of joints at the base, elbow, and two at the spin/despin mechanism or 

grapple mechanism interface. A motor, gearbox, and control sensor are part of 

each joint, and together create control complexity and increased re1 iab.il ity 

and maintai nabi 1 i ty requi rements. Like the 4-bar mec hani sm, the mani pul ator 

has a very compact stowed configuration. 

weight and complexity of a robotic arm are offset by its articulation, 

compacted volume, and load bearing capability. 

a TSR kit extendible boom concept, for both Systems B and C. 

The manipulator depicted in 

The study team felt that the higher 

The manipulator was favored as 

1-18 
U 

ADVANTAGE 

Compact 

Can Support Grapple, Despin Torque Loads 

Existlng Designs 

DISADVANTAGE 

Hervy 

Increased Control Complexity 

Enables Realignment of Target Mass for I Transport 

Enables Access to Obstructed Recovery Support Elements ‘b 

Figure 1.4-3 Manipulator Evaluation 

. 1.4.3 Spin/Despin Mechanism Definition - The concept selected for the 

Spin/Despin Mechanism subsystem is an existing design, functioning as part of 

an existing three axis gimbal within the MK Space Operations Simulation (SOS) 

Laboratory. An assessment of this design was accomplished in order to provide 

a conceptual level definition for follow-on efforts. 

of a spin platform, DC torque motor, servo control amplifier, tachometer, and 

electrical brake, as shown in Figure 1.4-4. 

The mechanism consists 

1.4.4 Grapple Mechanism Interface Device - The grapple mechanism interface 

device definition process identified the requi remnt for unique interface 

devices for ground assembly of a recovery system to be deployed mated with OMV 

and delivered into a working orbit by STS. 

flange, conceptually illustrated in Figure 1.4-5, provides an interface 

connection between recovery subsystems, such as that between the grapple 

mechanism and the spin/despin mechanism. A bolted assembly is required for 

each subsystem interface. 

device, it will be used as a readily applied interface between several 

components of a modular, interchangeable recovery system. 

The ground assembled interface 

Initially intended as a grapple mechanism interface 

0 

1-1 9 
BIDIRECTIONAL SERVO CONTROL SYSTEM 

- REVERSIBLE CURRENT DC TORQUE MOTOR 

- TACHOMETER TO SENSE SPEED 

- SERVO CONTROL AMPLIFIER TO ADJUST & DISSIPATE CURRENT 

- ELECTRICAL BRAKE TO SECURE TABLE 

SPIN TABLE '--- TACHoMmR 

DC TORQUE 

BRAKE 0 'MOTOR 

Figure 1.4-4 Spin/Despin Mechanism 

Flange 

Extendible 

Figure 1.4-5 Ground Assembled Interface Flange 

1-20 
1.4.5 Grapple Mechanism Evaluation - Three categories of grapple mechanisms 

were i d e n t i f i e d t h a t addressed the grapple mechanism requirements of Systems B 

and C. These included: a small gripper type mechanfsm for System B; a 

"stinger" type mechanism for recovery of disabled, spi n-stabil ized satel 1 i tes, 

and large grippers or envelopment type grapplers for recovery of noncontrollable, tumbling or spinning satellites. 

alternatives were recognized and evaluated and selections were made. The 

criteria identified and applied in the evaluations included: compactness, 

accommodation of a range of satellite size and hardpoint size, strength of 

grapple, target damage potential, positioning flexibility, and accommodation 

of grapple, despin, and transport loads. 

Within these categories, 

1.4.5.1 

(PM) End Effector and the JPL PFMA Smart Hand were selected by the study 

team as two possible accommodations of the System B, small gripper mechanism 

requirement, and are illustrated in Figure 1.4-6. Both provide parallel jaw 

motion, minimizing reaction away from the mechanism during grapple. 

intermeshing parallel plate designs and square recessed shapes of the jaws 

Small Gripper Definition - The MSFC Proto-Flight Manipulator Arm 

The 

enable grapple of hardpoints of varied size and shape. 

AwANrAGE 

EMbios Grappio of a Vark 

' Intormhlng Jaw flatos Enablo Plckup of 

Smallor Hardpoints 

Accompllrhor Firm Grip 

Ad]ustablo Grip Form and Cloruro Rat. 

* Grrpplo Unaffected by Powor Loss 

Dorign Avallablo. Provon 

of Round, Flat, Or Irrogul8rly-Sh8p.d H&OlnU 

USADVANTAGE 

Si10 of Hardpdnll, Qrrpplod Llmltod by Goornotry of Yoehanlrm 

Figure 1.4-6 Small Gripper Definition - PFMA End Effectors 

1-21 
These existing mechanism designs reflect the state-of-the-art for 

non-dexterous hands and provide capable accommodation of the small gripper 

requi rement. 

mission model hardpoi nt configurations will enable complete assessments of the 

applicability of these mechanisms to the System B requirements. 

1.4.5.2 Stinger-Type Grapple Mechanism - The obvious selection for the 

grapple mechanism of System C, involving recovery of spin stabilized 

spacecraft, is the Apogee Kick Motor (AKM) Attachment Device, commonly called 

the "Stinger". 

mission, where the recovery of two spin stabilized spacecraft was safely 

accomplished without damage to either vehicle. The stinger is depicted in 

Figure 1.4-7. 

Fol 1 ow on efforts i nvol vi ng analyses of transport 1 oads and 

I 

Its capability was proven in the Westar-Palapa B recovery 

ADVANTAGE 

En8bl.r Gr8pple of Spln Stablllxod Target 

DISADVANTAGE 

Figure 1 . 4-7 Sti nger-Type Grappler 

1.4.5.3 Evaluation of Large Envel opment-Type Grapple Mechanisms - 

Accommodation of the System C, large envelopment-type grapple mechanism 

requirement was provided to the study team by a large number of varied 

concepts. The six most promising were identified as a possible focus for 

follow on efforts, and were evaluated in detail. 

evaluations of the better concepts is provided in this Executive Summary. 

A review of three 

1-22 
The first of the large envelopers, the Debris Capture Device developed by LTV, 

is shown in Figure 1.4-8. 

mounted on adjustable arms, with which it accompltshes a two point grapple. 

The grapple force is applied through the ann assembly, by a hydraulic system 

at the base of the mechanism. The toroids minimize damage to recovery 

candidates by confoni ng to i rregul ar sate1 1 i te surfaces and spread4 ng the 

forces due to contact dynamics. 

It consists of a pair of low pressure toroids 

@ 

AwANTAGE 

Accommodates a Wldo lung. ot 5.tolllto SIzos 

aft Contact. Ylnlmltr Target D8mrgo 

The Debris Capture Device is one of the preferred concepts, yet it has several 

disadvantages. The mechanism structure is large and heavy. Added control 

complexity and reliability/maintainability requirements result from the 

hydraulic systems, toroid control, and the many linkages involved. 

grapple and despin, the mechanism's capability to provide sufficient 

rigidiration dud ng transport is questionable, as contact with the recovery 

candidate is sustained by a compressible device. Finally, as presently 

configured, it occupies a significant volume in the Orbiter cargo bay during 

launch and return. 

Following 

DEADVANTAGE 

Not CompCt 

* HeIvy 

Complex Comrol 

' aUOStlOnr#. Transport Rlgldlutlon C8p8bllky 

Bu1)7 Llmltd Flexlblllty In Tlght Targot 

InCr0aS.d RoH8bllltylUalntalnablllty 

Con fguRoquirermnf8 

Figure 1.4-8 Debris Capture €valuation 

1-23 
The MSFC Enveloper, depicted in Figure 1.4-9, has undergone preliminary 

design/development by MSFC. 

been incorporated in the mechanism. The base links of each arm are connected 

and operated by an actuation system between the two, which effects an 

inward/outward rotation of the arms. The remaining two arm links on each side 

operate in conjunction with one another, and with the similar links of the 

opposing arm, as a result of a motor within the base and a connecting 

linkage. 

f n posi ti oni ng for envel opment, and an i ncreased capabi 1 i ty in recoveri ng 

smaller satellites as compared to a system without the outboard arm segment 

articulation. 

Two methods of actuating the grappler's arms have 

The combined effect of the two systems is an increase in flexibility 

AWANTAGE 

Llght Wolght . 

Rolatlvoly Compact 

Accommod.too 0 Rang0 of !htoltIto Slzos 

Florlblllty In Posltlonlng for Envetopnwnt 

USADVANTAGE 

Incroarod Roll~blllty/Yalntalnablllty Roqulronwnta 

Figure 1.4-9 MSFC Enveloper Evaluation 

The final mechanism evaluation, shown in Figure 1.4-10, involves the MMC 

Enveloper, which was conceptually designed by the study team to provide a 

System C grapple mechanism that was designed from the requirements allocated 

in Task 1 to the envelopment grappler. 

1-24 
The mechanism incorporates a two arm, six member str'ucture. Driven by DC 

torque motors in the joints, each of the links can be operated independently. 

To maintain symnetry, control software is employed so that opposing links, 

e.g., the base links of each arm, operate in conjunction. The geometry of the 

members allows a folded, and extremely small compacted volume, with sufficient 

surface area at the end of each arm to enable a two point capture. Like the 

MSFC enveloper, appropriate structural member desfgn and composition will 

produce a re1 at4 vely 1 i ght weight mechani sm. 

U 

The substantial increase in flexibility of the grapple operation, as a result 

of independent actuation of the links, provides a number of advantages. 

Flexibility in positioning for envelopment and around protuberances is 

optimized. The design enables envelopment before contact, thereby minimizi ng 

target reactions and negating target motion away from the OMV. Target damage 

is minimized, as contact dynamic forces will be reduced with the grapple 

elements closer to the recovery target when first contact occurs. The 

mechanism's i ncreased fl exi bil i ty though, creates a proportional increase i n 

control complexity and re1 iability/maintainability demands. These 

disadvantages are outweighed by the number of advantages shown in Figure 

1.4-10, and this enveloper mechanism was selected for inclusion in the System 

C design configuration. 

~~ ~ ~~ 

* Ught Wolght 

Compact 

Optlmum RexIblUty In Porltlonlng for 

Ennlopmnt 

Accommodator I Wlde Rang. of 5.t.11110 Skn 

ontrol of Mulll Io Llnkr uallros 

sontact Forcr/&nlmlros !&ollllo D1ma90 

Accompllrher Enwlopment Boforo Contact 

' Accompllrher Rlgld Gnpplo 

WANTAGE 

Addod CotWoi Comploxlty wllh 

ConWoUI#. Arm Sognwnts Yultlpl. 

Figure 1.4.09 WC Enveloper Evaluation 

1-25 
1.5 Conceptual Tumbling Satellite Recovery System Design - Task 2.2 

The objective of the second concept definition task, Task 2.2, was to develop 

select the preferred set of subsystem mechanisms and to integrate them into a 

basic OW TSR kit system that met the study team's objective of system 

modul ari ty and ready interchange of subsystem components. 

l a set of conceptual designs for TSR Systems B and C. The approach used was to I 

1.5.1 

architecture for the recovery systems was influenced heavily by a number of 

key factors driven out by the requirements analysis. These are shown in 

Figure 1.5-1. The first of these was the inherently broad range of recovery 

scenarios identified during Task 1. This fundamental reality caused the study 

team to select from two apparent options: (1) operate from a design concept 

that would provide an equally wide range of recovery systems; or (2) develop a 

design concept with a modular design as a framework that could be configured 

readily into recovery systems tailored for specific missions. 

approach was selected. 

TSR Conceptual Design Drivers - The formulation of a design 

The latter 

Broad Range of Recovery Scenarios Dictates 

- Wide Range of Recovery Systems, or 

- Modular System Easily Configured into Recovery System Tailored 

for Specific Missions 

Recovery Kit Must Be Compact - Efficient STS Operations 

Minimum Risk to OMV 

Recovery Operations Bounded by OMV Controllability 

Target Rigidized for Return to STS or Space Station 

Figure 1.5-1 Key Design Drivers 

1-26 
Another key design criterion was the need for a compact design. The TSR 

system is being designed as an OMV front end kit and carried into orbit in the 

STS cargo bay. The OMV has been configured for compactness to minimize cargo 

bay space necessary for delivery into orbit. Though no other study reflected 

this requirement, MMC believed the architectural design should be as compact 

as is possible and selected subsystem options that supported this design 

factor. 

@ 

. 

Another prominent design factor was related to minimizing risk to the OMV 

during all phases of recovery operations. 

to maintain a proper distance from the OMV during operations, a design 

requirement readily accomnodated by the selection of an effective extension 

device. A related secondary design driver was the perceived necessity to 

retain control of the OMV during recovery operations. The primary concern 

here was that with contact dynamics forces in excess of OMV control authority, 

ground controllers could lose the capability of controlling OM and the TSR 

kit, with potential resulting damage to both. 

One such element involved the need 

Finally, no other study had considered the requirement to have the target 

satellite firmly grappled for transport back to the Shuttle or to Space 

Station. 

different recovery systems. 

This requirement influences the choice of grapple mechanisms for the 

1.5.2 Modular TSR System Design - A representation of a modular design for a 

family of recovery systems is provided in Figure 1.5-2. 

displays the major components of the system and the inherent capacity to 

interchange, add, or subtract individual components to tailor the kit for 

speci f ic mi ssi ons . 

This expanded view 

One of the primary features of the design is a structural interface element 

that is readily attached to the OW docking latches in manual mating 

operations in the STS cargo bay during launch processing. The electric power 

and communications and data management (OM) interface will be through the 

OMV payload umbilical mounted on the front face of the OMV. 

1 -27 
Modulu 

- Tailored to Spcific 

Miuiom 

r 2. \ I 

Small Grippr 

Figure 1 .5-2 Tumbling Sate1 1 i te Recovery Kit 

The next component is the extendible boom. 

manipulator arm with pitch and yaw positioning at the grapple mechanism 

interface was selected. 

stowed position. 

between the OMV and a spinning target. 

reach recovery support elements that are obstructed to an approach by OMV due 

to deployed target solar arrays and antennas. 

alignment of a target's center of mass with the OMV orbit transfer thrust 

vector following capture and prior to orbit transfer. 

A four degree of freedom 

This mechanism is capable of folding into a compact 

Its most important function is to provide safe clearance 

It also provides the capability to 

Additionally, it enables 

A third major component is the spin table. 

efficiently to the interface flange on the end of the extendible boom. 

shown configured within the spin table is a boresight, wide angle view 

television camera. 

or configured to spin at the same rate as the spin table. 

The spin table can be mounted 

Also 

On this base, it can be mounted in a fixed configuration, 

In this representation of the recovery system, both the System 6 and System C 

recovery kits are shown. 

mechani sm attached to the flanged grapple mechani sm interface device. 

The full-up System C has the MMC enveloper grapple 

Thi s 

1 -28 
system will be capable of recovering satellites with what was previously 

defined as the more complex tumble motion. 

evolve to single axis spin about the satellite's major principal axis. 

The System B configuration is also represented in Figure 1.5-2. This system 

incl udes the structural and electrical /C&DM interface with OW, the extendi bl e 

boom and a small gripper, connected to the system with the grapple mechanism 

interface flange. 

attached to the smaller gripper to provide locallzed viewing of the attachment 

to the hard point of a remote, controllable satellite. This figure portrays 

how amenable this design is to subsystem interchange and its capacity for 

accommodati ng a wi de vari ety of recovery mi ssi ons . 

This tumble motion is expected to 

In addition, a close proximity television camera is 

1.5.3 MK Enveloper Grapple Mechanism - The conceptual design of the MMC 

enveloper, which was selected as the grapple mechanism element for one of the 

conceptual System C recovery configurations, was influenced by an increasing 

concern on the part of the study team regarding the potential impact of 

contact dynamics between the TSR kit and the target during grapple operations. 

@ Even with perfect conditions during recovery, with no major target 

protuberances/appendages and given a reasonably sylnrnetric target for recovery, 

as the operator begins to grapple and rigidize, the grapple mechanism will 

begin a series of contacts with the target. 

relative position changes between the target and recovery system that are 

expected to be complex, and have not yet been modeled. When initiating this 

grapple mechanism closure operation with a two- or even three-point gripping 

device, such as a Clclamp, in which the target is not enveloped, it is 

possible that the target position will change in such a manner that a new 

approach and grapple positioning setup will be required following each 

contact. If this target reaction was found to be dominant, a "two point" 

grapple would be untenable. 

appeared to be accompl i shed more feasibly using an envel opment approach. 

Thus, the MK enveloper was designed from the operational concept that 

envelopment of a spinning target with a spinning grapple mechanism would 

provide a higher probability of successful grapple and rigidization. 

enveloper is shown in a grapple configuration in Figure 1.5-3. 

These contacts will produce 

For this reason, grapple of a spinning target 

The MMC 

1 -29 
- Three Pairs of Grapple Elements Independently Controlled 

- Grapple Mechanism Deployed to Optimum Envelopment Configuration 

- Target Enveloped-Elements Closed Slowly to Minimize Contact Dynamics 

- Grapple Mechanism Rigidized for Dmpin and Transfer to STS or's 

- dc Torque Motors-Harmonic. Planetary 

- Then Spun Up-Minimize Deployment Dynamics 

Figure 1.5-3 TSR Grapple Envelope 

1.5.4 Conceptual Recovery Systems - Sumnary - The recommended recovery system 

architecture and conceptual system designs are presented in a format that 

illustrates the efficacy of the MMC modular, inteKhangeable element approach. 

Shown in Figue 1.5-4 are the system configurations for System 6, for both of 

the recovery scenarios described in DRM 3 and DRM 4. For Case 1, the scenario 

is a controllable, stable target, with a recovery support elemnt, an RMS 

grapple fixture, that is obstructed from a direct OMV approach by a deployed 

solar panel. 

of the structural /mechani an interface element, a multiple degree-of-freedom 

manipul ator arm (to gain access to the grapple fixture), a grapple mechani sm 

interface flange and an RMS end effector. 

For this recovery candidate, the conceptual TSR system consists 

For the second System 6 scenario, a controllable satellite with no grapple 

fixtures or flight support system latch pins, the recovery system is shown, 

also, in Figure 1.5-4. The mechanical interface provides the ready attachment 

to the OMV, and in this configuration, a small gripper, for attachment to 

target hard points is attacned to the interface flange and the extendible boom. 

1 -30 
System B-Target Controllable, Recovery Support Element (RSE) Obstructad 

a 

RMS End Effector 

System B- Target Controllable. No RSE Available 

Camera 

Small Gripper 

Figure 1.5-4 Conceptual TSR - Sysbm 3 

The recovery system configurations for both System C recovery scenarios a= 

illustrated in Figuro 1.5-5. In the first case, with a controllable, spin 

stabilized target such as INTELSAT-6, the mechanical interfaces (both the 

structural and umbilical 1, the extendible boom and the spin table are 

included. The "stinger" type grapple mechanism, attached to the grapple 

mechanism interface flange, will be used to secure a solid grip on tile 

spi nni ng INTELSAT kick motor. 

* 

The most difficult recovery scenario is the full-up System C scenario, which 

is most likely a tumbling/spinning satellite. 

the enti re modular system, including an enveloper-type grappler. 

This recovery scenario requires 

Thus, the conceptual modular design contains all of the "fundamental" 

accomnodations to enable recovery of tne full range of identified and defined 

System B and System C mission requirements. 

1-31 
System C-Target Controllable, Spin-Stabilized 

1..6 

Stinger 

System C-Target Noncontrollable, Tumbling/Spinning 

Large Envelowr 

7-m 

Grapple Mechanism 

Figure 1.5-5 Conceptual TSR - System C 

Supporti ng Devel opment P1 an 

The purpose of the supporting development plan is to: outline research and 

technol ogy development incl udi ng ground-based testi ng and simulation, and 

Orbiter cargo bay or proximity operations activities; and to structure the 

fl i ght hardware development program needed to establish the technical 

readiness of an OMV tumbling satell i te recovery front-end kit. 

These activities must be integrated into a comprehensive program by recovery 

kit planners, and must be coordinated with concurrent OMV development 

activities. 

Research and Technology (SR&T) Report are addressed in either ground or 

fl ight-oriented experiments. 

. 

The technol ogy development issues identified in the Supporting 

1.6.1 

demonstrati on approach is hi ghl i ghted in Tab1 e 1.6-1. 

of the ground-based program is to design, develop and exploit a set of 

software and hardware demonstration units to examine the issues identified in 

Ground Demonstration Activities - The recommended ground-based 

The pri nci pal ‘el ement 

1 -32 
the Supporting Research and Technology Report, presented in Volume I1 of this 

Final Technical Report. 

the demonstration units exist at MSFC and at MMC. The approach recommends 

continued use of ground demonstration devices as laboratory tool s to evaluate 

evolving mechanism and system concepts. 

It is clear that the simulation facilities to develop 

Tab1 e '1.6-1 

Design, Develop, & Exploit Recovery Kit Ground Demonstration Unit(s) 

- Evaluate Concepts Feasibility 

- Recovery System Deployment Characteristics 

- Contact Dynamics in Recovery Operations 

- Recovery System OperationsIOperator Assessment 

Utilize Existing MSFC/Martin Marietta Simulation Capabilities to Address 

Identified Technology Issues 

- Contact Dynamics Concerns 

- Force & Moment Measurements, Resulting Position/Motion States 

- Computer Simulations Using Varying Configurations, Evaluate Human Factors 

Limitations 

Demonstrate Use of Recovery Demonstration Unit as Laboratory Tool 

. - Evaluate Alternative Concepts 

- Evaluate Subsystem Mechanisms - Grapple Devices 

- Eventual Use as Astronaut Trainer for Flight Experiment 

- Identify Logical Flight Experiment Candidates 

1.6.2 STS Cargo Bay/Proximi ty Operations - The definition of on-orbi t fl ight 

experiments to support technology development of the TSR kit will evolve and 

be refined through experience with the ground demonstration units. It appears 

that an on-orbit experiment will be required to validate the recovery concept 

agreed upon for development, and to verify contact dynamics forces and torques 

and the impact of relative movement between the target and recovery system 

during the recovery operation. The cargo bay/proximi ty operations experiments 

are outlined in Table 1.6-2. 

1-33 

. . ... 
Table 1.6-2 Cargo Bay/Proximi ty Operations Experiments 

Deflne an STS Cargo BaylProxImity Operations Equlpment Set 

- Scaled Satellite Recovery System 

- Extendible Boom, Spin Table, Envelopment Grappler 

- Equipped with Interface to STS & RMS 

- Scaled Composite Recovery Target 

Conduct Remote Recovery Experiments in Zero-G 

- Remote Recovery Operations 

- Spin Axis Alignment, Spin Rate MatchinglPhasing 

- Operations, Operator Limitations 

- Recovery System Deployment Dynamics 

- Target-Recovery System Contact Dynamics 

Cargo Bay Experiments Should Be Phased to 

Support Flight Hardware Phase C/O CDR. 

The on-orbit remote satellite recovery experiments will be conducted with high 

fidelity equipment to validate the system concept. Thus, definition of the 

requirements and conceptual design of the scaled down experiment equipment 

should begin prior to the start of flight hardware Phase C/D for the TSR 

sy s tem. I 

l The experimental recovery equipment would be an extendible boom, a spin table 

and an envelopment grappler. The system would be designed to interface with 

the STS RMS end-effector and equipped with an operating interface in the STS. 

The spacecraft target would be a modification of a current rented bus, 

designed to be controllable to produce mu1 tiple tumble and spin modes and 

rates. 

zero-gravity remote recovery operations to enable validations of ground-based 

experiments. Operators will gain actual experience in conducting remote 

recovery operations, such as spin axis alignment, spin rate matching and 

phasi ng with target, recovery system deployment dynamics and act1 vi ties 

conducted in reaction to targethecovery system dynamics. 

This experiment would enable operators to conduct the first 

1-34 
1.6.3 Flight Hardware Program - The objective of the actual OMV tumbling 

satelli-te recovery kit flight hardware program it to be prepared to conduct 

free flight operations fn 1993 with actual or sfnulated targets. This flight 

hardware will be developed on a schedule consistent with development plans for 

OMV and other OMV front-end kits. It will be conducted using the generally 

accepted NASA/MSFC approach of conceptual , definition and devel opment phases 

(Phases A, B, and C/D). 

The Supporting Development Plan (SOP) schedule is provided in Figure 1.6-1. 

The schedule outlines an integrated TSR kit development program that includes 

ground-based and on-orbit STS flight experiments and a flight hardware program 

that provides for free flight operations in 1993. 

I I 1 986 I 7987 I 1988 I 1989 1990 I lesl I 1992 I 1993 

- Continued Flight Hardware Evaluation 

Definition of Ground Demonstration Unit 

- Devolopmmt. Use of Ground D.mo Unit 

AT? PDR CDR CIR FRR 

vv vvv 

Free Flight Phase C/O 

Verif ication 

- Cargo Bay/Proximity 

Operations Schedule 

Figure 1 .6-1 Supporting Devel opment P1 an Sc hedul e 

1 -35 
1.7 Supporting Research and Technology Report 

The Supporting Research and Technology (SRBT) Report for the Tumbling 

Satel 1 i te Recovery (TSR ) conceptual def i ni ti on study i s presented in Vol ume I 1 

of this Final Technical Report. 

The FlMC study team did not identify any problem areas requiring new state-ofthe-art techno1 ogy devel opment initiatives. 

areas where research and 1 aboratory experiments coul d support resol ution of 

technology issues that could lead to development of a cost efficient remote, 

di sabl ed sate1 1 i te recovery system. These tec hnol ogy i ssues i ncl uded: 

provision of terrestrial estimates of disabled satel 1 i te motion, depl oyment 

dynamics of the capture device, contact dynamics between the target and 

recovery system and assessments of operator control capabilities for thi s 

remote teleoperations mission. 

They did identi fy a number of 

1.8 Cost Estimate and Work Breakdown Structure 

A cost estimate of the TSR kit program was prepared, based on a Work Breakdown 

Structure (WBS) and WBS Dictionary prepared by MMC. These study results are 

presented in Volume I11 of this Final Technical report. The initial WBS was 

presented to MSFC early in the study, and a coordinated effort between MSFC 

planners and the MMC study team resulted in creation of the final version. 

The total program cost for the TRS kit program, leading to a flight ready 

system was estimated at $18 million, based on the cost elements shown in 

Figure 1 -8-1 . 

1-36 
Phase 

*7 

C/D ($16.9M) 

I I 

! I 1.3.1- 

i 

1.3.1 Prog Mgmt 

1.3.2 SE&I 

Fiscal Year 

I 

i 

\ I 

1.3.3 1 1.3.3 Dgn & Anal. 

1.3.4 HW Proc 

1.3.6 Launch Ops 

1.3.7 Miss. Ops 

1.3.5 sys Assy 

7- 1.3.l+ 

6- 1.3.2 

5- 

4- 

3- 

2- 

"($0.23M) ($-a? Phase B Phase 8 

(S0.20M) ($0.30M) 1.3.3 

.3.4 

i 

I 

I 

I 

I 

I 

- B 1993 

Figure 1.8-1 Total TSR Kit Funding Profile 

c 

1-37 
2.0 STUDY RESULTS - INTRODUCTION AND BACKGROUND 

The study results that tnis Final Technical Report documents are the first 

assessment of the design requi rements and the conceptual definition of a 

"front end kit" to be transported on the currently defined Orbital Maneuvering 

Vehicle (OMV) and the Space Transportation System (STS) Shuttle Orbiter, to 

conduct remote, teleoperated recovery of disabled and possibly noncontrollable, tumbling satellites. Studies related to recovery of disabled 

satellites or space debris have been conducted by the NASA and 000, academia, 

and industry for over 20 years. 

considerad in previous studies, including Gemini, Apollo, and earlier versions 

of the ONV, including the Teleoperator Retrieval System . (TRS), that ?Jas under 

contract for actual design and development for retrieval of Skylab in the 

mi d-1970s. 

Many different transport vehicles were 

Tne unique aspect of tnis study is the fact tnat the conceptual tumbling 

satellite recovery (TSR) system was groundruled as a front-end kit for a well 

defined OMV transport vehicle, that is, in turn, being delivered into a near 

earth orbit by the oDerationa1 Shuttle. 

Where many previous studies examined only partial aspects of the remote 

recovery problem, such as how to grapple tumbling satellites (in various 

tumble modes), or how to despin satellites, tnis study encompasses J full 

range of issues related to the recovery problem. As the study progressed, an 

operations concept WS developed, and this concept provided a solid frame of 

reference for description and operations analyses of a group of selected 

Design Reference Missions (DRMs). 

analyses, a set of systems requirements and recovery concept requirements were 

developed and allocated on a top level basis, to a set of recovery Subsystem 

acconodations. From this overall reaui remnts analysis, a recovery system 

"design concept'' was defined and served as a foundation for tne total concept 

definition phase, a major study task. 

nature of "complex" satellite tumble motion surfaced, an explicit analysis was 

conducted to determine the most likely tumble mode presented by a noncontrolled, di sabled satall i te. Thi s study effort was designed to provide the 

broadest possible contextual framework in which to determine tumbling 

satel 1 i te recovery system requi r2ments. 

Using a series of systems requirements 

When a significant concern over the 

* 

2 -1 
2.1 Study Objecti ves 

The NASA Marshall Space Flight Center 

SUPPORTING OBJECTNB 

- IDENTIN REQUIRED NEW 

YSFC) expressed two major objectives in 

outlining the expected result of the tumbling satellite recovery study, as 

shown on Figure 2.1-1. The first of these objectives was to develop realistic 

candidate recovery systems to support the decision making process in preparing 

to develop an operational capability for remote recovery of disabled 

satellits in the mid-1990s timeframe. A second major objective was to define 

the full range of remote recovery capability reauired in this same period. 

s2ries of supporting objectives were also outlined including: 

A 

the 

identification of required new technology initiatives inandated by development 

of this capability; a supporting development plan to include requisite ground 

demonstrations and flight experiments; and cost analyses to support estimation 

of eventual total development costs. 

PREVIOUS STUDIES ALTERNATlVE SYSTEM 

ASSESSMENT MECHANISMS ASSESSMENT 

-CONCEmAL 

DESIGN A 

”/*“ /-.-*- 

--’ 

. “, ‘I * 

CONCEPT -/,,,/+”- -, SYSTEM A-BASIC OMV, EVALUATION 1*-1 NASA GENERIC / NASA CONFlGURATlON 

)”,,.””” _,...--s-.- SYSTEM 8-MINIMUM CAPABILITY 

~~~0e-y SY~MS- ADD-ON KIT .a 

<” 

SYSTEM C-FULL UP REMOTE 

RECOVERY KTT 

.I ,~ CIS 

*”I *’ 

DEFINE FULL 

RANGE OF REQUIRED 

REMOTE RECOVERY 

MOTION ASSESSMENT b 

TARGET ORIENTATION 

CAPABILITY 

TARGET FAILURE 

TECHNOLOGY 

- PROGRAM COST ESTlMATES 

- FLIGHT HARDWARE DEVELOPMENT 

/ 

POTENTIAL TARGETSEARLY 1990’s 

MODES 

PLAN 

. Figure 2.1-1 TSR Study Objectives 

2-2 
, 

The first and principal objective, that of defining viable recovery concepts, 

came from a MSFC perspective that in over 20 years of general, and in some 

cases very specific, examinations of concepts related to remote recovery, a 

large number of widely variant approaches had been considered and suggested to 

the NASA and to DOD. 

technical managers, Mr. Herbert Lenox, was to review all known previous 

related work in this area, to evaluate new concepts and to focus these efforts 

into a more channeled conceptual framework that would lead to cost effective 

development of the required capability. The study team believed tnat review 

of previous concepts and identification of system and concept level 

requirements would provide the framework on which to develop new alternative 

candidate concepts, for conpari son with the older concepts. 

assessment would lead to selection of viable candidate recovery systems, and 

provide the desired focus for the continuing development effort. 

0 

The specific objectives expressed by one of the MSFC 

This general 

The second principal objective was to examine a broad perspective of potential 

recovery scenarios and to define the full range of required remote, disabled 

sate1 1 i te recovery capabi 1 i ty. The MSFC requested Martin F4arietta to study 

and evaluate three generically defined levels of capability. 

0 - 

The first of these expressed levels of capability was called System A and is 

the basic OMV, using the NASA baseline configuration at study start time. 3Y 

determining the limits of inherent OPlV recovery capability, it was anticipated 

that the boundary between that level of capability and the next would be more 

edsi ly determined. 

A second level of capability, termed System 8, was described as the basic OMV 

with some mi nimal hardwar2 additions or changeouts, such as end effectors, 

batteries, or special avionics hardware. This level of capability was 

expressed rather generally to challenge the imagination of the study group to 

examine intermediate, but realistic, levels of capability with some rationale. 

The final desired level of recovery capability, labeled System Cy was 

described as a "full -up" capacity to recover satellites with ''complex'' 

motion. This "system" was to be a distinctive OMV front-end kit, capable of 

recovering some defined maximum level of remote, disabled satellite, in som 

@ degree of complex satellite motion. 

2-3 
Thus, 'it was the responsibility of the MMC 

for diffetentiating these levels of capabi 

study team to provide the rationale 

i ty. defining the requi rements for -- - 

I 

I C. 

each level, and providing conceptual designs for f4SFC's recovery Systems B and I 

2.1.1 Ground Rules - The study ground rules .ere explicit. The first of 

these was to make maximum use of completed satellite recovery studies and 

research efforts on actual hardware generation programs. Also, illMC was 

directed to review related, ongoing OMV satellib recovery related efforts. , 

I A second study guideline was to use MSFC's OMV baseline reference 

configuration for all study tasks, including Characterization of the OblV's 

sate1 1 i te recovery capabi 1 i ty and def i ni ti on of the interfaces , be tween 

Systems 8 and C, and the OMV. 

MSFC requested that any simulation work or ground demonstrati on acti vi ties, 

recommended by the study team as requisite Supporting Research and Technology 

(SRT) program activities, include consideration of MSFC simulation 

facilities. 

zero-gravity demonstration concepts, where it was considered cost efficient to 

verify preliminary conceptual mechanisms or techniques. 

A final guideline bdas to propose early, low cost Shuttle 

2.2 Study Approach 

The general approach used by Martin Marietta in the conduct of our Concept 

Definition Study for Recovery of Tumbling Satellites was to view the study as 

a preliminary Phase A activity. Some of the benefits derived from this 

differentiation were: (1) the development of a preliminary operations 

concept , out1 i ni ng ground and space-based operations ( for both Shuttle and 

Space Station bases); (2) requirements detailed at a lower level of 

decomposition, including some level of effort in attempting to quantify the 

requi rements; (3) development of a fundamental design concept that supports 

clarification of varying levels of recovery capability and improved 

delineation of a family of conceptual recovery systems that were efficiently 

configured to enable .recovery over a broad range of recovery scenarios. 

2 -4 
The approach used is outlined in Figure 2.2-1. 

major tasks in the contract statement of work. 

tasks are also shown. 

It was centered on the four 

The interrelations of these 

I 1 STUDY INPUT DATA TASK 1- REOUIREMENTS ANALYSES & TRADES 

1.1 CONCEPT IDENIIFEATION & EVALUATION 

1.2 MISSION MODEL (RECOVERY CANDIDATE 

1.3 RECOVERY SYsllEy REOUIREMENTS 

ANALYSES) 

DERIVATION 

MISSION MODEL 0 OSSA,OMV 

SATELLITE CONflGURATIONS 

SATELLITE FAILURES 

PREVIOUS RECOVERY SYSTEM 

-NASA CENTERS 

- MODES, MOTION ORIENTATION 

STUDIES, CONCEPTS 

II I 

lDEMlFlCATION & EVALUATION 

2.2 DEFINITION OF ALTERNATIVE 

TASK 3 - PROGRAMMATICS & PLANNING 

3.1 SUPPORTING DEVELOPMENT PROGRAM 

3.2 SUPPORTING RESEARCH & TECHNOLOGY PUN 

3.3 PROGRAM COSTS 

TASK 4 - STUDY MANAGEMENT 

4.1 STUDY PLANNING, ORGANlZATlOW & MANAGEMENT 

4.2 PERFORMANCE REVIEWS 

4.3 REPORTS 

Figure 2.2-1 Study Task Flow 

The principal data inputs to the satellite recovery study were: (1) mission 

model information, provided by the Office of Space Science Applications, OMV 

and Space Station programs, and specific satellite configuration data provided 

by NASA, Canada, and comnercial satellite program offices; (2) satellite 

failure mode and tumble mode/motion orientation; and (3) a large number of 

previously completed and ongoing study efforts related to remote satell ite 

recovery. 

Task 1 was one of the two principal study tasks and the general approach was 

to use a series of systems requirements analysis (SRA) processes that were 

primarily oriented to deriving system and concept requirements for the family 

of recovery systems to be defined. 

2-5 
1 

This task had four subtasks, and they were all inter-related and mutually 

supportive. 

previously known study and hardware efforts related to satellite recovery were 

reviewed and a group of potential concepts evaluated against selected system 

requirements-oriented criteria. Our assessmnt of this effort was that none 

of tnese concepts were worthy of continued detailed design. 

critical decision as the initial contract guidance was to complete furtner 

definition of some of these preferred concepts. 

diversions in the overall study effort. 

The first of these was a survey-type effort in which all 

This was a 

This decision resulted in two 

The first diversion was to comprehend more completely what to expect in regard 

to probable or "typical" satellite failure nodes, and the resulting tumble 

mode and motion orientation created by these failure modes. 

interesting to note that none of the previously conducted studies addressed 

this issue in a definitive manner. The result of this series of analyses, 

explained in detail in Section 4.4, supported definition of a broad range of 

recovery scenarios and enhanced definition of MSFC's conceptual Systems B 

and C. 

It was very 

A second study diversion occurred later when, with the realization that none 

of the previous concepts could fully meet top level system requi rements and 

design accommodations, it became apparent that the approach to the second 

major task, Task 2, Concept Definition, would have to be altered. This change 

is detailed in Sections 9.0 and 1.0. 

A second Task 1 subtask was the development of a TSR mission model. 

the study team's view that a clear understanding of the types of missions and 

dc tual si ze , shape and confi gurati on of potenti a1 recovery candi dates expected 

to be on-orbit in the mid-1990s would enhance definition of the requirements. 

This data was obtained and an assessment of tne actual satellite 

configurations enabled development of a set of composite target model 

characteristics and an improved comprehensi on of potenti a1 , future fai 1 ure 

modes and resulting target motion. 

It was 

2 -6 
In the third Task 1 subtask, Recovery System Requirements Derivation, a series 

of activities were conducted to blend with results of the initial two subtasks 

to enable requirements identification and evaluation. First, a concept of 

operations and a functional analysis of the actual recovery of a disabled 

satellite in complex tumble motion (see Section 6.01, were both completed. 

Next, a composite target model was developed as a "worst case" recovery 

target, and was used to conduct a top level quantification of some of the 

identified system and concept requirements. 

@ 

The last of the four Task 1 subtasks was to define design reference missions 

(DRMs) for satellite recovery. 

missions identified previously in the prior study task, it was a relatively 

straightforward effort to define and describe a set of six recovery DRMs. 

DRMs were selected for each of the MSFC defined levels of capability, Systems 

A, B, and C. 

the previously deri ved preliminary operations concept. 

further identification of recovery system requirelaents for both System B and 

System C, particularly in the area of system requirements, such as ground 

support equipment (GSE) and aerospace support equipment (ASE), for operations 

based at both the STS and Space Station. 

the STS as a base for OMV/TSR kit operations, and others designed using the 

Space Station as an operations base. This was done to highlight the potential 

for, and to identify the requi rements related to a growth option of eventually 

stationing sate1 1 i te recovery equipment at the Space Station. 

With the broad range of satellite recovery 

Two 

A top level operations analysis was conducted on each DRM, using 

Thi s acti vi ty provi Jed 

Some of the DRMs were defined with 

As a final requirements related activity, the study team conducted an 

allocation of all identified requirements to a group of candidate recovery 

system components. These components, including a structural interface, an 

extendiDle boom, a spin/despin table, a grapple mechanism interface device, a 

boresight blevision camera, and a family of grapple mechanisms for varying 

recovery scenarios, were viewed as the essential subsystems required for 

conceptual definition of a family of remote, disabled satellite recovery 

systems. This a1 1 ocati on of requi rements and definition of requi si te recovery 

subsystems provided tk principal input to the second major study task. 

2-7 
L 

The Concept Defi ni ti on task, Task 2, was designed to provide a1 ternati ve 

conceptual designs for satellite recovery systems and an assessment of no 

fewer than three grapple mechanisms. At the study Midterm Review in December 

1955, the MSFC contract technical managers agreed with the lJv;1C study team that 

Task 2 should be dedicated to an identification and assessment of alternative 

I conceptual designs for each of the agreed-upon recovery subsystem elements/ 

components. 

selection of preferred subsystem components and integration of these into 

conceptual designs for Systems B and C. 

Conceptual design of recovery systems would then be based on the 

At this point in the study effort, it became apparent to the study team that 

the overall recovery system could be designed in a modular format and composed 

of a number of subsystem mechanisms that would be readily interchangeable. 

Thus, if a particular recovery scenario required only an extendible boom and a 

small 

structural interface unit and mated with the OMV in this form. If a "complex" 

motion scenario were to surface, this might require a spin table and, perhaps, 

a large envelope type grapple mechanism. In this case, the design of the 

modular, expandable recovery system would enable the removal of the mall 

gripper on the end of the extendible boom .and with appropriate interface 

mecnanisms, the addition of a spin table and a large envelope grapple 

mechanism to accomplish the "full up'' System c type mission. 

I gripper (a System B mission), these elements could be attached to a 

I 

In the Subsystem Component Eva1 uati on task, Task 2.1, a number of a1 ternati ve 

candidates for each of the subsystems were identified and evaluated. 

of current concepts, including those deve 1 oped previ ously by NASA, Martin 

llarietta and other aerospace contractors was conducted. 

providing a wide range of applicability to each subsystem was identified and 

e va 1 uate d . 

A search 

A set of alternati ves 

In the second Task 2 subtask, Definition of Alternative Conceptual Recovery 

Systems (Task 2.21, the preferred subsystem alternatives wrp integrated into 

a set of System B and System C recovery devices. 

systems is comprised of a subset of the total recovery system mechanisms that 

make up a complete recovery system. These recovery system configurations are 

described in Section 9.0. 

Each of the four specified 

2 -8 
A major portion of the effort expended in the definition of conceptual systems 

tasks was the conceptual design of a large, envelopment-type grapple mechanism 

designed for compactibility for transport in the Shuttle cargo bay. This 

"Martin Marietta enveloper" mechanism was designed to meet all of the grapple 

requirements identified in the study effort. It is a conceptual design that 

will enable recovery of the most likely "complex motion" cases expected in a 

"full -up" System C scenario. 

0 

Programmatics and Planning, Task 3, was comprised of three general supporting 

study efforts. The fi'rst of these was the Supporting Research and Technology 

Report (SR&T). The Supporting Research and Technology Report was preparea 

from a compilation of those technical issues requiring further study and 

clarification to enable eventual definition and development of an OMV remote 

satellite recovery kit. 

new technology development initiatives is that there are no critical 

technology %how stoppers" related to the development effort. 

ground-based studies and demonstrations were identified, as well as a number 

of Shuttle cargo bay or proximity operation type on-orbit experiments. 

1 and 2 provided inputs to Task 3, the Supporting Development Plan and TSR 

Program Cost Estimates. 

The study team's assessment of the requirement for 

A number of 

Tasks 

@ 

Task 4, Study Management, as shown in Figure 2.24, is supported by the three 

other tasks and was comprised of study planning and management, the conduct of 

three directed reviews, 5 .e. , an Orientation Meeting, a Midterm Review, and 

the Fjnal Review, and preparation of a final report in three volumes. In 

addition, the study team conducted two additional Technical Interchange 

Meetings at MSFC to present interim progress reports and to seek advice and 

recommendations for changes in approach where deemed appropriate by the 

contract technical manager. 

2.2.1 

recovery study is shown in Table 2.2.1-1. The requirements oriented 

activities (Task 1) were scheduled and completed with the exception of the 

Design Reference Missions, prior to the Midterm Review. A Technical 

Interchange Meeting was conducted in October 1985. The Task 2 and 3 

activities were conducted in early 1986, and a Technical Interchange Meeting 

Study Schedule - The schedule used to conduct the tumbling satellite 

e 

2 -9 
was conducted in 1 ate March 1986, to Secure acceptance of the recovery system 

modular design concept and the design of the Martin Marietta Enveloper grapple 

mechanism. 

the final two months of the study period. 

The final review and final report material were generated during 

Table 2.2.1-1 TSR Study Schedule 

TimcPhased Schedule 

Months after Go-Ahead 

Tnckc 

Contract Award 

1.0 Requirements Analyses & Trades 

1.1 Concept Identification & Evaluation 

1.2 Satellite Requirements by Category 

1.3 System & Hardware Design Requirements 

1.4 Design Reference Missions 

2.0 Concept Definition 

2.1 Recovery System Conceptual Design 

2.2 End-Effector Concept Evaluation 

3.0 Programmatics & Planning 

3.1 Supporting Development Program 

3.2 Programmatic Planning - 

4.0 Study Management 

4.1 Planning Organization & Management 

4.2 Performance Reviews 

4.3 Reports 

Meetings 

Documentation 

DR-1 

DR-2 

DR-3 

DR-4 

DR-5 

DR-6 

DR-7 

Legend: 

P Preliminary 

F Final 

U Update 

C COR Comment 

7 

F 

F 

F 

2-10 
3.0 REQUIREMENTS ANALYSES AND TRADES - TASK 1 

IS 

3.1 Introduction and Approacn 

The objective of Task 1 was to perform the type of analyses' and trades that 

would identify recovery system requirements for the MSFC designated range of 

recovery systems, i.e., Systems A, By and C. The identification of system 

requirements was designed to support identification of satellite recovery 

concepts that muld show promise of satisfying recovery requirements. Tne 

approach used in conducting Task 1 is shown in Figure 3.1 -1. 

objectives of each of the four tasks are showny together with the interaction 

between each of them. Task 1.1 was designed to secure a thorough survey of 

prior related satellite recovery studies and hardware development efforts. 

This was expected to produce a basic conceptual understanding of the problems 

related to remote recovery and serve as an ini tialization of the requi rements 

definition process. 

Section 4.0. . 

The major 

The Task 1.1 analysis and results are presented in 

RECOVERY CONCEPT 

IDENTIFICATION & 

EVALUATION - TASK 1.1 

RECOVERY CANDIDATES 

(TSR MISSION MODEL) - TASK 1.2 

f + 

RECOVERY SYSTEM REQUIREMENTS - TASK 1.3 

- OPERATIONS CONCEPT 

- FUNCTIONAL ANALYSIS 

- MISSION MODEL ANALYSIS 

! 

DESIGN REFERENCE MISSIONS - TASK 1.4 

- SYSTEMS A, B, & C 

- OPERATIONS ANALYSIS 

- REFINED REQUIREMENTS 

Figure 3.1-1 Task 1 Approach 

3-1 
The second requirements oriented task, Task 1.2--TSR Mission Model 

Derivation--was intended to high1 ight the types of candidates available for 

recovery and was used by the study team to illustrate how potential recovery 

targets will drive recovery system requirements. 

given in Section 5.0. 

The Task 1.2 infomation is 

I In Task 1.3, the results of concept identification and recovery target 

assessment were used to develop an operations concept for an OMV/TSR kit, and 

to conduct a functional analysis of a full-up System C recovery operation. 

Additionally, the study team then conducted a detailed analysis of the mission 

model to build a "worst case" set of composite targets and to begin to 

quantify the recovery requirements as they are impacted by recovery target 

candidates. 

Section 6.0. 

The approach to and results from these analyses are given in 

The final requirements analysis task was selection and approval of design 

reference missions (DRMs). 

levels of capability (Systems A, By and C) were defined by Martin Marietta and 

approved by the MSFC contract technical monitors. A top level operations 

analysis was conducted for each of the DRMs to gain additional insight on 

requirements, particularly system requirements, such as GSE and ASE for the 

TSR kit. The design reference mission information is presented in Section 7.0. 

A set of six DRMs, two for each of the three MSFC 

The analyses conducted in Task 1 produced a broad base of system 

requirements. These were then a1 1 ocated to recovery system accommodations and 

provided a sound basis for conceptual definition, the second major study task. 

I 
4.0 CONCEPT IDENTIFICATION AND EVALUATION - TASK 1.1 

e 4.1 Introduction and Approach 

Concept Identification and Evaluation, Task 1.1, was the first of the 

requirements analysis tasks. The study team conducted an extensive search for 

any documentation di rectly or i ndi wtly related to sate1 li te recovery. 

Johnson Space Center (JSC) and MSFC libraries wee queried, as well as a large 

number of knowledgeable individuals at JSC, MSFC, the Jet Propulsion 

Laboratory (JPL), and Goddard Space F1 ight Center (GSFC). 

provided source leads on other aerospace contractors and tnese produced 

additional sources. Martin Marietta had completed a number of recovery 

related studies, experiments, and actual hardware demonstrati on uni ts and 

these were all incorporated into the survey data base. 

The 

The NASA centers 

As was to be expected, tnere was a great deal of variety involved in this 

collection of data. Some of the studies addressed the overall question of how 

to remotely recovery "tumbling" satellites. 

on the mechanics of physically attaching a grapple mechanism to satellites in 

various states of motion. 

OMV" and some of these are shown in Figure 4.1-1. Another group of recovery 

system analysts apparently presumed that tumble motion would typically be so 

complex that attachmnt to the target would be impossible or inappropriate and 

chose to deploy rocket propelled nets to encircle the target, or to fire 

harpoons or adhesive grapples at the target to achieve a fim connection. 

These concepts were categorized by the study team as "deployable from the OMV" 

and are shown in Figure 4.1-2. As can be seen, some of these concepts 

addressed only partial aspects of the problem, such as simply despinning the 

disabled satellite using plum impingement, yo-yo despin packages, self 

contained thruster stabilization packages, or electromagnetic despin. None of 

these recovery concepts considered the necessity of providing a firm 

attachment to the target satellite once despin was completed. 

Most of these focused primarily 

@ These were categorized as ''concepts attached to 

4 -I 
I 

f 

1 

I 

i 

I 

I 

f 

4-2 
. 

i 

i 

I 

c 

i 

n 

t 

1 1 

B 

0 

4-3 
blost importantly, none of these conceptual study and hardware developmnt 

efforts *re conducted from the perspective of transporting a recovery system 

front end kit using a well defined OMV and an operational Shuttle. 

These concepts represent a myriad of satellite recovery/retrieval and space 

debris collection concepts that have evolved over many years. 

vehicles for the recovery devices have varied from Apollo to Gemini, and to 

early versions of the OMV, such as Earth Orbital Teleoperator System (EOTS) 

and Teleoperator Retrieval System (TRS). 

concepts that were sufficiently "systems"-oriented to be considered as 

"concepts" for overall evaluation under this analysis task. 

selected five of the best of these concepts for further definition and 

evaluation against a set of evaluation criteria. 

definition and evaluation process are presented herein. 

The transport 

There were only a few recovery 

The study team 

The results of this 

4.2 Concept Identi f icati on 

4.2.1 

concept developed by Vought, a Ling-Tempco Vought (LN) company, as a space 

debris collection, front-end system for the Teleoperator Mamuveri ng System 

(TMS) (now OMV) in 1982. 

of inflatable, low pressure toroids, that are mounted on adjustable ams for 

varying the reach of the system. 

mecnanism of the system. 

beam that is capable of spin about an axle. A television camera mounted on 

the spin axis rotates witn the arms, and the capture area is illuminated by 

lights situated on the beam. This viewing capacity is supported by the TMS 

pan/ti 1 t/zoom camera. 

Debris Capture Device - The Debris Capture Device (DCD) is a recovery 

The DCD, shown in Figure 4.2.1-1, consists of a pair 

This portion of the DCD is the grapple 

This grapple mechanism is mounted on a rotatable 

The DCD, mated to the TMS or OMV, would operate by observing the tumbling 

debris dnd approach the target along its spin axis. 

portion is ''spun up" until the TV image of the debris is no longer spinning. 

This will reflect the fact that the spin rates have been matched. The arms 

would then be closed to allow the soft, compliant toroids to grip the 

irregularly shaped debris and damp out oscillations. 

envelopment of a debris element by the DCO. 

The grapple meCnJniSm 

Figure 4.2.1-1 shows the 

4-4 
Vought Grapple Mechanism Includes Inflatable, Low-Ptwsure Toroids 

Adjustable Dual Arms for Varying Reach 

Rotatable Axis for Spin Matching 

Television Cameras on Arms for Proximity Opedons 

Initial Assessment 

Addressed Most Identified TSR Functional Requirements 

Has Applicability for Capture of Tumbling AcW Satellites & Space Debris 

Figure 4.2.1-1 Debris Capture Device 

The general assessment of the DCD was that is could satisfy most of the TSR 

requirements identified at this early point in the study effort. Later, it 

would be devalued for its lack of compactness for transport as an OMV front 

end k i t i n the Shuttle and for lack of an extendible boom to provide safety 

clearance between OMV and a target satellite. 

@ 

4.2.2 Teleoperator Retrieval Manipulator - The Teleoperator Retrieval 

Mani pul ator was a NASA/MSFC concept modi f i ed by tkrti n Marietta for 

applications related to recovery of satellites with "complex" motion. It is 

illustrated in Figure 4.2.2-1. The recovery device consists of a spin table 

to match target spacecraft spin rates, an extendible manipulator ann to natcn 

coning angles created by free precession of the target, and a rotatable 

grapple wrist with a small RMS snare-type grappler to grasp an RMS end 

effector on the tumbling satellite. 

complex coning motion of the satellite and grasp the grapple fixture, then 

apply forces to reduce the spinning and coning rates to zero. 

Such a manipulator am would follow the 

4-5 
NASA Concept Modified by Martin Marietta 

Major katures 

Device Designed for Multiaxis Motion 

Spin Table to Match Spacecraft Spin Rates 

Extendible Manipulator Arm to Match ConinglFree Precession Angles 

Rotatable Grapple Arm for Longitudinal, Body-Axis Spin 

initial Assessment 

Enables Grapple of Satellites with multiple Spin Axes 

Possesses Most Subsystem Elements Required for Recovery 

Fi gu re 4.2.2 -1 Te leope rat0 r Ret ri e val Mani p ul ato r 

The initial assessment of this device was that it would enable grapple and 

despin of sate1 lites with this type of motion assuming that a grapple fixture 

was properly located on the tumbling satellite. The system could not deal 

effectively with targets that had no grapple fixtures or had complex motion 

that requi red envelopment. (The envelopment requirement was developed in 

subsequent analysis efforts. ) 

to be effective for this type of recovery scenario; however, the study did not 

addr2ss the likelihood of this or any other type of tumble motion. 

In general, the recovery device was considered 

4.2.3 

Handling Device (EMHD) recovery concept evolved out of a fdarshall Space Flignt 

Center/Martin Marietta Aerospace study conducted duri ng 1970. 

shown in Figure 4.2.3-1, was developed for use as a front end kit for the 

Apollo command module. The recovery system included a flexible interface 

between the kit and the command nodule to snable ready adaptation of a variety 

Experimental Materials Handling Device - The Experimental Materials 

The system, 

I 

I of potential grapple mechanisms to the recovery system. The study considered 

4-6 
several alternative grapple mechanisms, such as the "C-clamp", shown in Figure 

4.2.3-1, clamps with three and four arms and even some adhesive grapples. In 

fact, pnysical representations of all these grapple mechanisms were built and @ tests conducted to validate them. 

Thi s device included a spi nni ng interface to enable matchi ng of the spi n rate 

about one potential spin axis of the tumbling satellite. 

The assessment of the EMHD concept was that it contained most of the 

mecnanisms assumed as essential elements during this preliminary phase of the 

TSR study. However, there were no provisions for dealing with multi-axis spin 

or tumble, and the EMHD study provided no assessmnt of expected or probable 

"complex" motion. As subsequent analysis suggested that single axis spin is 

the most probable state of a non-controllable tumbling satellite, the initial 

assessment of EMHD meeting most requirements prosd to be quite accurate. In 

addition, it turned out to have potential for applications in recovery 

scenarios that -ere later categorized as System B and System C cases. 

concept included a grapple mechanism interface device to enable ready 

application of special purpose grapple fixtures, sucn as have already been 

developed for recovery of the Solar Maximum satellite, and WESTAR and Palapa-B. 

Tnis 

Martin Marietta Concept 

Major Features 

Developed for Use with Apoilo Command Module 

Device Includes Spinning Interface with Variable Grapple Fixtures 

- C-Clamps, Two & Four Arms, Adhesive Grapple 

Equipment Built & Tests Conducted 

Initial Assessment 

Simple, Can Accommodate Target with Single-Axis Spin Grapple Mechanism Interface Concept Support8 Use of Simple and Complex 

Grapple Mechanisms 

0 

0. Figure 4.2.3-1 Experimental Materials Handling Device 

4-7 
4.2.4 Docking and Retrieval lilechanism - The Docking and Retrieval Mechanism 

(DRM) concept evolved from another Marshal 1 Space F1 ignt Center/Marti n 

Marietta study on an earlier version of the Ol4V that was known as Earth 

Orbital Teleoperator System (EOTS). 

MSFC directed Martin Marietta to design and fabricate an engineering prototype 

design of the D@l adaptable for installation in the MSFC mobility Unit 

simulator facility, now known as the Teleoperation and Robotics Engineering 

Facility. 

At the completion of a definition study, 

The DRM concept is shown at Figure 4.2.4-1. 

The principal mechani sins of the DRM included: 

spin-up to 100 revolutions per minute (RPM) and despin; (2) an extendible boom 

to enable tne EOTS to station keep with the payload while extending the 

docking mechanism, in lieu of using the EOTS translational capability to 

effect the insert, capture, latch and dock sequence; and (3) a section 

containing a latcn mechanism that could be converted readily to a grapple 

mecnani sin section. Thi s potential recovery concept contained an extendible 

docking probe for stationkceping with the EOTS, but the boom would, in fact, 

also provide a safe clearance tone between OMV and a tumbling recovery target. 

(1 ) a spin table capable of 

Martin Marietta Design from Earth Orbital Teleoperator Systems (ECTTS) Study 

(1 976) 

Major Features 

Extendible Docking Probe Mechanism 

Payload Docking Receptacle 

Control Electronics & Displays Prototype Hardware Delivered to MSFC 

Initial Assessment 

Spin Table Capable of Spin & Despin to 100 rpm 

System Includes Extendible Probe Mechanism 

Grapple Mechanisms Could Replace Latch Mechanism 

Figure 4.2.4-1 Docking Retrieval Mechanisin 

4-8 
The deliverable DRM hardware included the EOTS extendible docking probe 

mecnanim, a 

controls and e 

This concept 

include many 

sate1 1 i tes. 

payload docking receptacle, the DFUd control electronics and the 

displays shown in Figure 4.2.4-1. 

was chosen for extended evaluation because it did appear to 

of the features seen as requisite to remote recqvery of disaolsd 

The primary exception was a grapple mechanism and its related 

interface device, i.e., that element required to accomnodate ready changeout 

of grapple devices for varying recovery scenarios. 

that a grapple mechanism element could replace the DRM latcning mechanism, 

making this a feasible concept for this evaluation task. 

The study team believed 

4.2.5 Space Bola - The last of the five recovery concepts selected for the 

expanded evaluation was the Space Bola, a recovery concept introduced in the 

early 1960s. The space bola concept is illustrated in Figure 4.2.5-1. 

Martin Marietta-Mid-1960s 

Major Features 

Inflatable, Extendible Grappling Arms To Envelop Satellite 

Rockets in Arm Tips Fired To Achieve Grapple 

Initial Assessment 

Flexible, Extendible Grapple Mechanism, To Accommodate Varying Target 

Sizes 

Contact Sensor & Envelopment by Arms Pddea Rigidization for Transport 

Spin Table, Extendible B&m Readily Added 

Figure 4.2.5-1 Space Bola 

4-9 
The Space Bola concept was patterned after the bolas used by the gauchos of 

South America. The gauchos make their Dolas &y attaching balls of stone or 

iron to the ends of a cord. 

bisecting point. By hurling the bola at the legs of animals, the animals 

become entangled and are captured. 

grappling unit, which is a set of three inflatable arms that are inflated to a 

ful ly extended position to provide a capturing volume, and a contact sensor. 

On approaching the target the grappling phase would commence at contact with 

the target. The extended arms are partially deflated, and small solid 

propellant charges in the end of the arms drive the arms around tne target. 

Velcro pads on tne ends of the arms lock the arms together as they overlap. 

The arms are then reeled back into the grappling unit until a firm attachment 

is achieved. 

satellite would be despun and prepared for return to the STS. 

A third ball is attached to the main cord at the 

The Space Bola concept consists of a 

Witn the spin table, added by the study team, the disabled 

I The major redeeming value of this concept was the potential provided by the 

Space Bola's enveloping grapple mecnani sm, a feature that'became increasi ngly 

more relevant as the study effort progressed. 

satellite) and a spin table could be readily added to make tnis a viable 

concept. With these additions, the Space Bola was included in the group of 

recovery concepts for expanded evaluation. 

In addition, the study team 

I believed an extendible boom (for clearance between OMV.and the tumbling 

4.3 Ini ti a1 Concept Eva1 uati on 

4.3.1 Introduction - The survey of prior remote mcovery related study and 

hardwar% act1 v i t i e s provided the study requi rements analyst s task some i ni ti a1 

insights on fundamental recovery requi rements. As one would expect, as the 

study progressed and other analysis tasks were completed, the breadth and 

depth of both recovery system and recovery concept requirements was expanded. 

Thus, the initial concept evaluation was conducted from a relatively limited 

View of total rvquiements and is presented herein from this perspective. In 

actuality, the resulting evaluation, and its perceived shortcomings, provided 

tne study group with an alternative analysis approach that produced highly 

rewarding results. 

I 

4-1 0 
4.3.2 Evaluation Criteria - The evaluation criteria eventually selected for 

the concept comparisons were derived from: 

provided by pre-proposal efforts; (2) new perspectives gained from the concept 

identification and assessments; and (3) discussions with various MSFC and MTK, 

personnel. Each 

of the evaluation criteria was weighted to illustrate its relative importance 

to the criterion considered to be of highest importance. 

(1) insights and background 

0 

The selected evaluation criteria are shown on Table 4.3.2-1. 

The "capability to recovery a broad spectrum of satellite configurations" was 

selected as the highest vale evaluation criterion. 

gathering data on the projected satellite mission model for the mid-1990s and 

it was clear that there was to be a wide diversity in satellite size, Shape, 

mass and potential failure mode for recovery candidates in this period. 

The study team was 

Table 4.3.2-1 Concept Selection Criteria 

Capability to Recover Broad Spectrum of Satellite Configurations 

Minimum Risk to OMV & Recovery System during Recovery 

Capability to Accommodate High Single-Axis Satellite Spin Rates 

Minimum Risk to Target Vehicle 

Compatibility with OMV; Minimum Impact on OMV Design 

Dependence on Recovery Vehicle Support Elements 

Modularity of Subsystems to Enable TSR System Growth for 

Flexible Mission Capability 

Capability to Deal with Wide Range of Tumble Mode Complexity 

Weight to Orbit (Mass 8, Volume) 

Development Risk/Cost 

Weight 

10 

9 

9 

8 

7 

7 

6 

5 

5 

5 

Another highly ranked criterion was the necessity for the eventual recovery 

system to conduct operations with minimm risk to the OMV and the TSR kit 

during recovery operations. It was not known at that time what the potential 

complexity of recovery operations might be, but safety of recovery equipment 

appeared to be high in importance. 

'4-1 1 
This evaluation was actually conducted twice, before and after the satellite 

failure modes analysis described in Section 4.4.1. The next criterion, 

capability to accomodate high, single axis spin rates, was derived from that 

analysis, and added to the evaluation criteria, with the high designated 

weight shown in Table 4.3.2-1. 

The sole purpose of a remote disabled satellite mission is recovery of a 

valuable sateliite for refit and return to functional operation. Thus, 

minimum risk’to the target vehicle was also considered high in potential 

ranking for concept evaluation. 

that certain deployed target elements such as antennas and non-retractable 

solar panels might have to be sacrificed during recovery; in some cases to 

enable access to onboard recovery support elements, such as RMS grapple 

fixtures, and in others to preclude damage to OMV and the TSR kit during orbit 

transfer maneuvers. 

Orbiter could also result in the sacrifice of its on-orbit deployable 

appendages. 

It was clear and becam increasingly apparent 

Return of a recovered satellite to the ground in the I 

The conceptual design of a recovery vehicle Has considered to have to be 

basically compatible to that of the OMV and to have minimum impact on the 

design of OMV, and, thus, this criteria was rated relatively high. 

be designed and developed well in advance of the TSR kit and the TSR kit 

should be designed to be cost efficiently integrated into the OMV program. 

Continuing analyses proved it would be equally important to design the TSR kit 

for compatibility with STS Shuttle operations. 

OMV will 

Dependence on target recovery support elements (RSE) was actually a negative 

criterion and those concepts requi ring this feature bere downgraded. It was 

felt that many potential targets would not have RMS grapple fixtures or flight 

support structure berthing rings, or given they were present, would not likely 

be in a position where they would support attachment to a TSR kit. 

target were spinning or tumbling, the recovery system would have to grapple on 

an axis of spin through the target center of gravity, and the RCE was not 

likely to be there. 

If the 

4-1 2 
The next criterion was titled "Modularity of Subsystems to Enable TSR System 

Growth for Flexible Mission Capability." 

inclusion of this evaluation criterion. 

scenarios described in Section 7.0 led the study team to believe that anoverall recovery system built with a number of readily interchangeable 

mechanisms would be a recovery system design concept candidate. 

A parallel analysis resulted in the 

The broad range of potential recovery 0 

Tumble mode complexity was another evaluation criteriorr that was downgraded 

due to otner ongoing Task 1 analyses. Many of the "previous" concepts had 

been designed to accomnodate complex satellite motion comprised of spin on 

multiple axes, while others were designed for recovery of single axis spin 

cases. 

motion. 

multi-axis spin or tumble, so the weight of this evaluation criterion was 

lowered. 

None of the studies actually addressed "complex" or worst case tumble 

Subsequent analysis revealed that complex motion was not likely to oe 

The last two criteria =re mass and volume required for the recovery kit and 

development risk and cost. 

considered as manageable development criteria and rated accordingly. 

Though both are important criteria, they were 

0 4.3.3 Concept Evaluation - Shown in Table 4.3.3-1 is the result of the 

evaluation of the five pre-selected concepts. The selection criteria, 

weighting factors and total scores are provided in the table. 

provided by five.experienced study team members and their associates. 

the score values was then normalized for each evaluation criteria and edch 

recovery concept. 

The scoring was 

Each of 

As was expected, the Experinental Materials Handling Device, the Debris 

Capture Device, and the Docking Retrieval Mechanism all received higher 

scores, as tnese devices appeared to possess, at that point in the study 

effort, more of what the study team was beginning to understand as 

requirements for a "full-up" recovery device. 

.' . 

4-1 3 
recovery concept received only 80% of that score, and therefore it became 

clear that more effort was needed to understand clearly what was required to 

4.4 Alternative Concept Identification Approach 

The total assessment of the concept identification and evaluation task is 

shown on Figure 4.4-1. 

and tasks, it was clear that Task 1.1, Concept Identification and Evaluation, 

had not contributed to those important issues shown on the top of Figure 

4.4-1. This task had not provided differentiators that would support a 

refinement of the definition of MSFC's three generic recovery system levels of 

capabi li ty, i .e., Systems A, B, and C. 

understanding of what the OMV operator would expect to see, in terms of 

typical satellite motion at the site of the "tumbling satellite." 

the study efforts and hardware activities addressed a broad range of perceived 

tumble motion. 

satellites in multi-axis, general tumble motion, some configured more for 

specific single axis spin, and others, like net and harpoon concepts, were 

designed apparently for tumble motion dth no discernable steady state motion 

configuration. MSFC had referenced '@complex motion" as that required for the 

full -up system, but no defi ni ti on of complex motion was readily avai 1 able. 

In looking at some of the specified study objectives 

The survey added no clarity to the 

In fact, 

Some concepts were configured for recovery of disabled 

The study team developed and conducted a new and ultimately successful 

approach to the clarification and comprehension of these issues. 

the bottom half of Figure 4.4-1, this approach included a search for 

historical failure mode data, the resulting tumble mode data from actual 

failed satellites, and a dynamics analysis to enable full comprehension of any 

real world satellite "complex motion'' findings. The results of this series of 

analyses are highlighted in Figure 4.4-1. 

satellite failure inode analysis. 

As shown on 

The first of these analyses ;das 

\ 

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4-1 6 
4.4.1 Failure Mode Analysis - It was assumed that a seach for satellite 

failure data wuld be rewarded quickly and with highly useful data. An 

extensive series of telecons to offices in NASA headquarters and centers, and 

to similar Department of Defense focal points produced limited results. 

Fortunately, one excellent source of satellite failure data was secured. Mr. 

Edward Shockey at Goddard Space Flight Center (GSFC) had contracted with 

Planning Research Corporation (PRC) to conduct a study of spacecraft on-orbit 

anomalies and lifetimes. 

Orbit Anomalies and Lifetimes", PRC R-3579, February 10, 1984. 

examined the orbital performance records of 44 umanned spacecraft under tne 

The report is titled "Analysis of Spacecraft OnThe study 

cognizance of GSFC and the Jet Propulsion Laboratory (JPL), during the 

1977-1984 time period. 

anomalous spacecraft behavior, ranging from momentary mal feasance to comp 

spacecraft failure. Tnis is an extensive collection of United States 

satellites and presented a highly representative data base for use in ma& 

assumpti ons regarding satel 1 i te fa5 1 ure cause and resulting tumble moti on 

synopsis of relevant failue data is presented in Figure 4.4.1-1. The re 

The PRC study examined each recorded incident of 

e te 

ng 

A 

o rt 

presented two overall statistical measures of merit. The first of these was 

an attempt to measure the mission effect of the anomaly. As shown in Figure 

4.4.1-1, with a total of 602 reported anomalies, a satell i te anomaly wi 11 

result in greater than 67% mission loss in only 6 out of over 600 cases. That 

equates to major on-orbit satellite failure, in only 1% of the satellites, or 

very rakly. 

occurred recently and are to be expected. 

.a 

But, as will be shown later, nrajor oworbit failures have 

The other measure of merit used in the GSFC report was highly relevant to this 

study effort. The report included a breakout of anomalies ~y major spacecraft 

subsystem. These eight subsystems categories are snow in Figure 4.4.1-1. 

When assessing this satellite subsystem failure data, two highly relevant 

observations can be made. First, the power and attitude control and 

stabi lization (ACS) subsystem anomalies account for nearly 30% of satel 1 i te 

subsystem failures. 

spacecraft stability, the potential for loss of satellite control from major 

failure could be projected at nearly one in three. The impact of this is 

expanded in subsequent portions of this report. 

Thus, with power and ACS the major contributors to 

4-1 7 
Mission Effect 

Nonnegligible, but Small 117 

113 to 2/3 Mission Lou 32 

213 to Nearly Total Mission Lou 5 

Essentially 1 Total Mission Lou . - 

602 

Negligible 447 

Spacecraft Subsystem Anomaly 

Timing, Control, & Command 

Telemetry 81 Data Handling 

power supply 

Attitude Control & Stabilization 

Propulsion 

Environmental Control 

Structure 

Payload/ExparimmtaI 

Number 

55 

112 

56 

123 

26 

16 

6 

- 208 

602 

Percent 

2.6 

1 .o 

100% 

- Few Major Failures 

- 35% Could Fail in Controllable Mode, Not Tumbling 

- 30% Could Fail in Noncontrollable, "Tumbling" Mode 

Figure 4.4.1-1 Spacecraft Failure Anomaly Analysis 

A second major relevant recovery-oriented statistic is presented by an 

assessment of the pay1 oad/experimental subsystems. The GSFC/PRC report noted, 

as shown again in Figure 4.4.1-1, that nearly 35% of satellite anomalies were 

related specifically to the payload or instrument package. Thus, one could 

just as readily project that one-third of major satellite failures could 

result in a useless mission, a disabled payload package, with a totally 

control 1 ab1 e satel 1 i te. 

In summary, a remote, disabled satellite is equally likely to be completely 

stable and controllable, with power and ACS subsystems intact, or totally out 

Of control, in some form of "complex" tumble motion. 

the study team was to deternine just what that tumble motion would most likely 

be. 

The next step taken by 

4.4.2 

data was conducted. The study team began to query NASA and DOD sources for 

examples of satellites that had failed in a non-controllable mode and, more 

importantly had, for some reason, had tumble or satellite motion orientation 

Failed Satellite Motion Analysis - Another search for satellite failure 
data collected on them. The search was difficult, again, for obvious 

reasons. 

care to discuss failures. 

was collected on six different missions. 

Primarily, it was determined that program/project offices did not @ However, satellite failure and related motion data 

The first of these were two failures of the Defense Meteorological Satellite 

Program (DMSP), Block 5D, flights 1 and 2. Data on these failures was 

obtained from Roger Hogan, Radio Corporation of America (RCA) in Princeton, 

New Jersey, and from Ray Skrynska, Aerospace Corporation, Los Angeles, 

California. 

The f i rst reported non-control 1 ab1 e satel 1 i te failure, wi th re1 ated tumbl e 

motion data, was DMSP Block 5D1, Flight 1. 

a result of a massive failure in nitrogen supply to the reaction control 

system (RCS) thrusters and some hydrazine leakage. 

to about 3.1 revolutions per minute (RPM) in a general three axis tumble mode, 

resul ti ng in complete 1 oss of atti tude control. 

was lost and batteries were soon depleted because of the disorientation to the 

sun and the satellite eventually lost all power. As DoD was extremely 

interested in recovering use of this scarce resource, tumble motion data was 

acquired prior to loss of power through reading attitude control sensors 

onboard the satellite. The initial motion was general tumble, with spin, 

precession and nutation components. 

converged to flat, single axis spin about the major principal axis. 

case, after several months, the spinning solar panels realigned with the sun, 

the batteries became recharged and with power again available, the 

reprogrammi ng of onboard software enabled ground control 1 ers to stabi 1 i ze and 

reestabl ish atti tude control of the satel 1 i te. 

This Do0 payload failed in 1976 as 

The satellite was spun UP 

Spacecraft atti tude control 

0 

This general tumble motion quickly 

In this 

The second non-control1 ab1 e satel 1 i te failure case occurred in the second 

flight of DMSP Block 5D1 program. DMSP is a DoD weather satellite and a 

vitally needed military support system. 

thought to be the result of a ground controller error. 

checkout of the satellite, the solar boom did not deploy. 

attempting to "shake out" the boom instead loaded the satellite with enough 

"excess" angular momentum to exceed the spacecraft's atti tude control 

capability. Once again control was lost, placing the spacecraft initially in 

The flight two failure was generally 

During operational 

A ground controller 

4-1 9 
a general, multi-axis spin, with a spin rate of 7 RPM. Again, the satellite 

motion, due to system flexibilities including the solar boom, antennas and 

fuel sloshing in the propellant supply tanks, quickly converged to what the 

study team would later find is the "classic dynamic solution'' - flat, single 

axis spin about the principal axis of maximum moment of inertia. Fortunately, 

as with the flight one case, this rotational mode did force deployment of the 

solar panel , and it realigned with the sun. The batteries recharged and, 

again the spacecraft was recovered for full operational use. The DMSP flight 

1 and 2 configuration is shown at Figure 4.4.2-1. The spin axis, single axis 

rotation, is illustrated there, and one can see how the solar panel could 

deploy with forces provided by this spin. 

. 

Further examples of satel lite non-controllable failure motion were provided by 

short failures in TIRCS-N and NOM-6. 

low levels of angular momentum, slightly beyond the capacity of.the angular 

momentum wheels to retain control of the satellites, and tumble motion 

developed. 

as in the DMSP cases. 

In both cases, hydrazine leaks induced 

The final tumble, or single axis spin rate mode, evolved quickly 

The last two cases were both classified military satellites. One was a three 

axis stabilized spacecraft, the other a dual spinner type satellite. In both 

cases, subsystem failures produced a tumbling, non-controllable satellite that 

entered general, mu1 ti-axis tumble motion immediately, and very quickly, the 

motion converged to single axis spin about the major principal axis. 

Thus, the process of collecting and assessing motion orientation data on 

actual non-controll able satel 1 i tes produced resul ts that appeared to show a 

somewhat consistent pattern. Different failure causes resulted in noncontrollable motion with an apparently predictable pattern. 

defining "complex", non-controllable motion was to fully comprehend the 

characteristics of what practical spacecraft dynamics experience was 

demonstrating. 

The next step in 

n -7n 
Block 5D Configuration-Flights 1 &. 2 

Array 

Figure 4.4.2-1 Defense Meteorological Support Program 

4.4.3 Spacecraft Dynamics Analysis - An independently conducted analysis 

provided an analytic solution to the question of why non-controllable 

satel 1 i te general tumble motion rather quickly coalesces into si ngl e axi S 

spin. 

0 

The result of that independent analysis is sumnarized herein. 

Satellite tumble motion is usually initiated by some source of torque on the 

satel 1 i te system, either internally or externally induced. 

failure, the NOAA-8 satellite, a sun-synchronous satellite in a 450 nautical 

mile, polar orbit suffered a battery explosion that disabled the satellite and 

induced a 1.5 RPM spin on the satel 1 ite. The study team intends to study this 

failure independently to: 

tumble motion; and (2) assess the nation's capability to determine tumble 

motion orientation and rates, using terrestrial sources, for mission planning 

support. 

In another recent 

(1 1 improve the general understanding of complex 

4-21 
The induced torque from the failure creates a level of angular momentum beyond 

the control laws and the control capability of the satellite's attitude 

control and stabilization (ACS) system. The torque also induces a high level 

of additional kinetic energy into the spacecraft system, a system that is a 

quasi-rigid body. This excess energy will create an initial state of general 

mu1 ti-axis tumble, with spin, precession and possibly nutation. 

As satellites are truly quasi-rigid bodies, i.e., not solid, inflexible rigid 

bodies, this kinetic energy will be dissipated quickly by various damping 

Sources, such as flexible appendages, viscous fluid flow in propellant tanks, 

and even friction in structural fasteners. 

When the additional kinetic energy introduced into the satellite system by the 

external torque source reaches a "steady state" minimum energy level, basic 

dynamic theory supports the thesis that the angular velocity vector coalesces 

with the angular momentum vector. 

"dynamic coning'' and the satellite motion becomes flat, single axis spin about 

the major principal axis, which is the axis of maximum moment of inertia. 

These results are in agreement with the precept that a spinning system will, 

in the presence of damping, seek the lowest possible energy level consistent 

with conservation of angular momentum. 

This eliminates free precession and 

The thesis that satellite "complex motion" is single axis spin, at varying 

rates, about the spacecraft's major principal inertia axis is supported: 

analytically; (2) by observed actual satellite failure cases; and (3) by 

further val Sdation through personal telephone discussions with the dynamics 

experts 1 i sted bel ow. 

(1 ) 

a. Dr. Farrenkopf, TRW 

b. Dr. Cochran, Auburn University 

c. Dr. Hubert, RCA 

d. Dr. Likins, President, Lehigh University 

e. Numerous Martin Marietta spacecraft dynamics experts. 

With this assessment of what non-controllable satellite complex motion is 

likely to be, the study team determined, and was supported by MSFC during 

study reviews, that the requirements and conceptual design of the tumbling 

satel 1 i te recovery "System C" should proceed from this recognized complex, 

uncontrol 1 ab1 e satel 1 i te motion mode. 

4-22 
4.4.4 Recovery System Differentiators - The definition of complex, or worst 

case, non-controllable motion provided a framework for refining the 

description of MSFC's hypothetical Systems B and C. 

analysis supported this objective by clarifying what the logical stateh) of 

motion orientation for disabled satellites could be. These potential states 

of motion are outlined in Table 4.4.4-1. 

@ The spacecraft failure 

A disabled satellite could be, in fact, attitude stabilized and under control 

from the ground. This would be the case when a payload or instrument package 

were to fail while the power and ACS subsystems are not affected. The 

recovery environment in the case of a controllable satellite will be dependent 

upon whether a disabled satellite has recovery support elements attached to it 

or not. Recovery support elements (RSE) include: (1 ) RMS grapple fixtures, 

which can be grappled by the OMV RMS grapple mechanism/end effector; or (2) 

STS flight support structure (FSS) berthing "latch pins." Some targets will 

be equipped with RCE elements, others will not. 

Another recovery environmental state, as shown in Table 4.4.4-1, will be the 

case of a controllable, spin-stabilized satellite!, such as a comnercial, 

comnunications satellite that experienced a kick motor failure after departi ng 

the Orbiter cargo bay. The satellite will be spinning at some rate, probably 

varying from 5 to 50 RPM, though stable, and must be recovered from that 

motion orientation. 

A final motion orientation, or recovery environment, is the case where the 

disabled satellite is non-controllable, and the fallure mode has induced a 

tumbling or spinning action. Again, from the failure mode analysis, the 

tumbling case is as likely to occur as the controllable recovery case. 

In any event, continuing analysi s demonstrated that the recovery requi rements 

and the recovery system capability differs in each of these prospective 

di sabl ed satel 1 i te envi ronments. 

4-2 3 
Table 4.4.4-1 Failed Satellite Motion Orientation 

Disabled Satellite Can Be: 

- Controllable with Recovery Support Elements Available 

- Controllable without Recovery Support Elements Available 

- Stable, Spin Stabilized, Spinning 

- Non-Controllable, TumblingSpinning 

Recovery System Requirements/Capabilities Different in 

Each Case 

The failed satellite environment analysis/sumnary directly supported a final 

definition of a set of differentiators that led to an enhanced definition of 

MSFC's Systems A, 6, and C. These differentiating elements are shown in the 

matrix of Figure 4.4.4-1 . The selected recovery scenario differentiators are 

all listed and those applicable to each system are checkmarked. 

System A scenario is defined as dealing with a target that is beyond Orbiter 

range, the target's attitude is controllable (either automatically or from the 

ground) and the target has properly si tuated recovery support el ements 

(RCE). This implies that the target is not spinning, or is spinning so slowly 

that OMV could match the recovery target's spin rate. 

exercises have shown these limitations to be single axis spin at rates no more 

than two to three degrees per second. The RCE must be fully accessible to a 

head-on approach with the OMV RMS end-effector, and the RCE cannot be 

obstructed by antennas or solar arrays. 

recovery environment, or recovery workspace, then the System A scenario exists 

and the satellite could be recovered by the basic OMV. 

Thus, the 

Computer simulation 

If all these conditions exist in the 

4-24 
Target Is beyond Orbiter Range 

ABC 

xx 

1. Within Reach of OMV Grapple Devices 

2. STS RMS Grapple Fixtures or FSS Berthing Latches 

3. Beyond Grapple OMV Rate-Matching Capability 

Target's Attitude Can Be Controlled from Ground 

Target Has Properly Situated' Recovery Support 

Target Is Controllable, Spin Stabilized at High Rate@ 

Target Cannot Be Controlled from Ground and in 

Elements (RSE)* 

Minimum-Energy Flat-Spin Mode. 

Target Is Prevented from Reaching Minimum-Energy 

State (Due to Internal or External Torques, or Bizarre 

3-a Failures) & Exhibits General Motion at High Rates 

(Low Probabilitv of Occurrence) 

Figure 4.4.4-1 Recovery System Differentiators 

xx 

X 

X 

X 

The System B scenario description varies only slightly from this scenario. 

Again, the recovery target is beyond the Shuttle's range, and the target is 

controllable and not spinning beyond the.capability for OMV to match the 

target's spin rate. For System 8, however, the target is defined as having no 

RCE to grapple with OMV retrieval equipment, or the RCE element is obstructed 

by deployed antennas or solar arrays for normal OMV approaches. As will be 

shown subsequently, the capability requirements for System B are limited in 

nature and substance as requested by MSFC. 

The System C scenarios are two in number. 

fundamental scenarios that will be accommodated with a basic System C, and the 

recovery system to accommodate these scenarios will vary slightly for each Of 

the two scenarios. 

Shuttle range, controllable, but spin stabilized at speeds ranging from 5 to 

This implies that there are two 

In the first System C scenario, the target is beyond 

50 RPM, beyond OMV spin rate matching capability. 

e 

4-25 
The second System C scenario, as shown in Figure 4.4.4-1, involves a case 

where the disabled target is not controllable, through loss of power or ACS 

subsystem failures, and is in a minimum energy, flat, single axis spin. Note 

that, in this case, the target does not necessarily have to be out of Shuttle 

range. The spinning, tumbling satellite would present a large danger to the 

Shuttle and crew if an extravehicular activity (EVA) recovery were to be 

attempted in this scenario. 

This series of trades and analyses provided the rationale required to define 

the MSFC family of recovery system capabilities and provide the foundation 

upon which to define the requirements for the three, actually two, levels of 

capability required in the OMV satellite recovery kit. 

will be discussed subsequently. 

one should note that none of the systems included the final recovery scenario 

differentiator shown. 

energy state, or. retains general motion characteristics (spin, precession and 

nutation) at high rates because of excessive atmospheric drag or other 

external torques, the recovery requirements can become excessi ve. 

System C does not include these requirements, primarily because of the low 

probabi 1 i ty of occurrence of this scenari 0. 

This system capability 

However, referring back to Figure 4.4.4-1 , 

When the target is prevented from reaching a minimum 

The full -Up 

4.4.5 Recovery System Definition - The overall intent of this series of 

analyses was to develop a rationale for the development of a logical family Of 

satellite recovery systems, with well defined levels of capability. MSFC, in 

the contract statement of work, described three 1 eve1 s of capabi 1 i ty , Systems 

A, B, and C, but directed the study team to evaluate the recovery problem 

areas and to definitively describe the boundaries between the levels Of 

capability. In paragraph 4.4.4, the recovery system differentiators provided 

a framework for defining the scenarios that could logically apply to an 

expanding level of recovery system capability. Figure 4.4.5-1 high1 ights the 

scenario descriptions and system descriptions for Systems A and 6. 

4-26 
The typical System A scenario was described as a situation in which the 

disabled satellite is beyond Orbiter range, the target is controllable and not 

spinning beyond OMV's spin rate matching capability, and has ClMV recovery 

support elements that are accessible to the OMV. 

antennas or solar arrays will not interfere with a normal OMV approach. 

System A is defined as a basic reference configuration OMV, with its 

capability to dock to unobstructed RMS grapple fixtures and to flight support 

structure (FSS) latch pins. 

0 

In this case, the deployed 

The capability of the MSFC reference configuration OMV for capture operations 

is restricted primarily by pilot workload, based on computer graphic 

simulations conducted recently at both MSFC and MMC. Simulations conducted at 

MSFC have indicated that ground controllers can control OMV operations against 

targets with from two to three degrees per second of target roll, one degree 

per second of target pitch or yaw, and less than 0.5 degree per second Of 

multi-axis motion. The OMV capture capability is lfmited by excessive pilot 

workload impacted by three primary factors. 

delay of signals relayed through the Tracking and Data Relay Satellite System 

(TDRSS). A second factor limiting OMV capture operations is the small 

envelope of the capture environment, both for an FSS latching pin capture and 

the capture of an RMS grapple fixture. The third and most limiting factor is 

the complexity of attempting to dock (and capture), while having to translate 

and rotate to maintain a position for the eventual capture. 

that to accomplish a successful System A recovery, the target will have to be 

re1 ati vely stab1 e. 

The first of these is the time 

@ 

It is apparent 

For System B, the typical recovery scenario, is shown on Figure 4.4.5-1. The 

disabled target is: 

stable; and (3) the target either (a) does not have recovery support elements 

(RSE), or (b) the RSE(s) are inaccessible to the approach of a basic OMV, and 

the target is not spinning or is spinning within OMV rate matching capability. 

(1) beyond orbiter range; (2) it is controllable and 

4-27 
Scenario TSR Kit Confiaurations 

- Basic OMV 

Beyond Orbiter Altltude 

Target Controllable 

Accessible OMV Recovery 

\ Support Elements 1 

Beyond Orblter Range Target Controllable 

Recovery Sup ort Elements 

InaccessiblePLacking Tar et not S innin 

wittin OMV !pin &te 

or 

KIT WITH: - Extendible Boom - Grap le Mechanism 

Interface Device 

- mall Gri er - grapple fl 

Figure 4.4.5-1 .Recovery Systems A & B Definition 

The required System B capability is an OMV front-end kit with a configuration 

including four subsystem elements. Some form of "extendible boom" is required 

to allow the OMV access to obstructed grapple fixtures or other recovery 

support elements. A mu1 tiple degree-of-freedom manipulator arm is viewed as 

essential to align the captured target's center of mass with the OW orbit 

transfer thrust vector, once a firm grapple is achieved. A second element is 

a grapple mechanism interface device. This recovery element will allow the 

ready interchangeability of various grapple mechanisms required for other 

scenarios. 

small "hard points" on satellites that have no recovery support elements. 

television camera located in close proximity to the small gripper (or to an 

RMS end-effector to be used for grapple of obstructed RCEs in some System B 

recoveries) is also seen as a required element. 

A third element is a small gripper mechanism to enable grapple of 

A 

This description of a System B capability appears to meet the MSFC definition 

of limited capability, minimum impact on OtiV. 

kit will be structurally interfaced with the OMV docking latches, and the. 

electrical and comnand and data management (C8DM) interface will be with the 

OMV pay1 oad accommodations umbi 1 ical . 

As will be shown later, this 

4-28 
The definition of a full-up System C is provided in Figure 4.4.5-2. The 

full-up recovery system was designed to deal with two complex motion cases. 

Case 1, shown in Figure 4.4.5-2, deals with the scenario in which the recovery 

target lis beyond Shuttle range, and is spin-stabilized between 5 and 50 RPM. 

In this situation, the recovery system configuration will include an 

extendible boom, a spin/despin mechanism to match rates with the target, a 

grapple mechanism interface device, and a "stinger" type grapple mechanism, 

similar to that used on WESTAR and Palapa-B recoveries. 

require a boresight camera to support alignment of the spin axis of the 

OMV/TSR kit and the target's spin axis. 

0 

This system will also 

The most difficult System C scenario is Case 2, the situation in which the 

satellite is not controllable, due to some major malfunction, and has 

excessive angul ar momentum. An excessi ve torque created a tumbl i ng sate1 1 i te 

that quickly assumed a state of single axis spin about a single, major 

principal inertia axis. The satellite could actually be within Shuttle range, 

but would have a tumble motion considered dangerous to close approach and 

recovery by EVA, MU, or other Shuttle dependent techniques. Thus, the 

recovery would have to be a remote retrieval using OW and a TSR kit. 0 

The system capability for System C, Case 2, is similar to that of Case 1. A 

spin/despin mechanism is required to match the spin rate(s) of the target. 

The tumble or spinning configuration of the satellite, however, will require a 

large gripper or large envelopment type grapple mechanism to provide a firm, 

smooth grapple of the target, to rigidize the attachment and then maintain the 

grip to enable despin and stabilization of the target. 

system configuration is shown in Figure 4.4.5-2. 

A representative 

In sumnary, these Task 1.1 analyses provided failure data that supported the 

identification of a broad range of remote, disabled satellite scenarios. 

derived scenario differentiators provided a rationale for a clarification of 

MSFC's generic levels of recovery system Capability, Systems, A, 8, and C. 

With these levels of capability more refined, it was possible to provide a 

The 

. general description of the specific recovery systems. 

4-29 
I System C-Full Up Capability 1 

- Case 1-Controllable Target, 

- Beyond Orbiter Range 

- Stable-Ground Controllable 

- Spin-Stabilized-10.50 rpm 

- Case 2-Noncontrollable Target, 

- Within OMV Range - Noncontrollable 

- Target Tumbling/Spinning 

GnppIe Fixture 

Spun la 

OMV 

... :-

..-... ' s& AXIS Grapple InterfaceExtendibleProbe 

-Stinger Grapple Mechanism 

- Extlndiblr Boom 

- Spin/Despin Mechanism 

- Grapple Mechanism Interface Device 

- Large Envelopment Grapple Mechanism 

- Stinger Type Grapple Mechanism 

Figure 4.4.5-2 Recovery System C Definition 

Spin-12 rpm 

OMV 

pln Table-12 rpm ' 

ORIGINAL PAGE IS 

OB EOOR QUAJJIX 

a -7n 
d 

5.0 SATELLITE RECOVERY SYSTEM MISSION MODEL - TASK 1.2 

5.1 Introduction and Approach a 

Task 1.2 involved tne development of a TSR mission inodel, a data base of 

mi ssi on and spacecraft speci f ic i nformati on regard1 ng sate1 1 i tes expected to 

be either on-orbit during the mid to late 199Os, were scheduled for launch 

during that period, or were representative of expected future follow-on 

systems. The model provides the basis from which to design a recovery system 

capable of accomnodating the diversity of target sizes, shapes, mass 

distributions, and configurations characteristic of operational satellites in 

the 1990s. 

missions and the acquisition of detailed information regarding those missions 

and the spacecraft involved. 

developing the model. 

Development of the model included the selection of appropriate 

Figure 5.1-1 illustrabs the approach used in 

t 

DETERMINATION MISSION ACQUlSmON OF 

OF b M508 

APPROPRIATE MISSION 

MISSION SPECIFIC - 

- RN=oMRY 

MISSIONS SPEClRCATlON RELATED DATA 

MISSION 

MODEL 

Figure 5.1-1 Mission Model Development 

5 -1 
c 

Through telephone conversations with the Office of Space Science Applications 

(OSSA) Division Chiefs, mission specific infonation regarding a number of 

satellite missions of varied maturity was obtained. Based on their 

oierational schedules, the study team selected 16 missions as the most 

appropriate for inclusion in the initial data base. Although the nature of 

the study effort precluded consultation witn DOD planning offices, the 

information provided by the OSSA offered a data base of sufficient size and 

di versity. 

To gather specific recovery elated data on the selected missions, a survey 

letter was sent to all NASA program/project managers. 

information regarding the satellite's physical configuration, 

shape, mass di stri buti on, etc, the questi onnai pe requested information 

concerning satellite orbital parameters, atti tude control system, hazards, 

plum sensitivity, safety, safing modes, and the number and cost of the 

satellites. The inquiry was addressed to the NASA managers, as opposed to 

contacting contractors directly, in order to avoid limitations imposed by 

A-109 restrictions. The study team received information from all programs; 

however, most of the data is of a preliminary nature due to the early stages 

of some of the programs. 

characteristics provided by several of the programs were, as yet, insufficient 

for analysis, and the mission model was limited to 13 of the 16 initially 

se 1 ec ted mi s si on s . 

In addition to detailed 

.e., its size, 

As a result, it was felt that the satellite 

5.2 Description of Mission Model 

The mission model development process provided a collection of diverse mission 

objectives and spacecraft configurations for 13 missions of varied maturity. 

As a summary, a brief description of each is outlined below. 

The Advanced X-Ray Astrophysics Facility (AXAF), designed for launch and 

servicing by the Shuttle, will incorporate a 1.2 meter, grazing incidence, 

X-ray telescope to accomnodate instruments collecting high spatial resolution 

and spectral data on quasars, galaxies, clusters of galaxies, and the 

intergalactic medi um, Evolving from the second High Energy Astronomy 

Observatory, it will have four times the spatial resolution and at least 100 

times the sensitivity. AXAF is one of the largest satellitss in the model, 

5-2 
having a cylindrical geometry, roughly 15 feet in diameter and 49 feet in 

length and a wight of 19,000 pounds. 

32 feet, extend from the spacecraft along a transverse axis, while two 

antennas extend 6 feet along an axis perpendicular to that of the arrays. 

AXAF is a potential development start for 1987. 

Two large solar arrays, measuring 10 by 

The Cosmic Background Explorer (COBE) mission, scheduled for launch by the STS 

in 1988, is being designed to explore and study diffuse radiation beteen tne 

one micrometer and 9.6 mi 11 imeter wavelengths. 

subsystem will be used to achieve the final circular, sun-synchronous, 900- 

kilometer orbit from a 300-kilometer STS parking orbit. Twelve solar arrays 

form the perimeter of the spacecraft, from which an omni antenna and the 

propulsion subsystem extend. 

approximately 18 feet and 13 feet respectively. 

10,000 pounds. 

An on-board propul si on 

The Explorer length and diameter are 

Its weight is approximately 

From an altitude of 400 kilaneters, the X-Ray Timing Explorer (EXP) will 

conduct intensive studies of the changing luminosity of x-ray sources, over' 

times ranging from milliseconds to years. 

energies from 2,000 to 100,000 electron volts will study known sources and 

transient events. The EXP leased platform and payload form a boxy structure 

with no more than a fifteen-foot width, an eight-foot height, and a fiVe-fOOt 

length. 

platform. After the scheduled 1992 launch, the payloads are planned to be 

replaced every 2.5 years. The platform and its initial payload, the X-Ray 

Timing Experiment, will weigh approximately 6,600 pounds. 

Instruments sensitive to x-ray 

Two rectangular solar arrays, seven by eight feet, extend from the 

The Gravity Probe-B mission involves one of the smaller spacecraft of the 

mission model. It will weigh only 2900 pounds and has a conical geometry 

which tapers from a six-foot diameter to less than one foot over J fifteenfoot length. 

perimeter, increasing the six-foot diameter to fifteen feet. The spacecraft 

will enable the testing of a fundamental concept of general relativity, by 

measuring the precession of orbiting gyroscopes as they move through a 

gravitational field twisted by the earth's rotation. 

planned to follow a functioning prototype that will be tested on a 1989 

Four solar arrays are symmetrically attached around the 

The GPB mission is 

Shuttle fl ignt. a 

5-3 
A new start in 1985 and scneduled for launch in 1989, the Geopotential 

Research Hission (GRM) will provide detailed global mapping of the Earth's 

gravitational and magnetic fields to an accuracy that is an order of magnitude 

of improvement over all previous global models. 

similar spacecraft, launched to a 275-kilometer orbit, self-deboosted to a 

160-kilometer altitude, and then separated a distance varied from 150 to 550 

kilometers. The gravitational field mapping will be accomplished by measuring 

changes in the relative velocity of the vehicles. Magnetometers isolated-on 

the fourteen-foot boom of tne leading vehicle will provide the geomagnetic 

measurements. The spacecraft cy1 indrical geometries, of 21 -foot length and 

3.5-fOOt diameter, consist mostly of propel lant, =qui red to naintain thei r 

peculiar low-earth orbit. 

Configuration to eacn vehicle. 

pounds. 

The mission will involve two 

Two solar arrays are firmly attached in a wing-like 

The spacecraft weights are 6200 and 5700 

A significantly larger spacecraft, the Gama Ray Observatory, will collect 

data on gamma rays, by observing known sources and by making the first full 

sky gama ray survey. Three individual instruments will measure specific 

gamma ray ranges from one-tenth of a million to 30 billion electron volts. 

The observatory is being designed to be launched, serviced and retrieved by 

the STS. Four 100-pound thrusters will boost it to the 450-kilometer circular 

operating orbit and 28.5-degree inclination. The three instruments help to 

create a very irregular satellite surface. 

approximately 70 feet from tip to tip and a high gain antenna extends 21 feet 

from the surface. 

dimensions are roughly 15 feet in diameter and 24 feet in length. 

I 

I 

Two solar arrays measure 

GRO has a weight of approximately 33,000 pounds. Its 

The principal mission in the NASA astronomy program is the eight foot diameter 

Hubble Space Telescope (HST). 

of wavelengths from the infrared to the ultraviolet, providing fine angular 

resolution, will enable extragalactic astronomy and observational cosmology 

for tasks such as investigation of stars in other galaxies to determine their 

rotation, age, mass, and chemical composition. 

operating orbit, the Shuttle and Space Station will serve as a base from which 

to service and replace the HST science instruments, as tecnnological advances 

and scientific prior'ities evolve. 

spacecraft, measuring 14 feet in diameter and 43 feet in length, and wighing 

roughly 25,000 pounds. 

The telescope's ability to cover a wide range 

8 

At the 320 nautical mile HST 

The telescope's size mandates a large 

5-4 
INTELSAT VI is representative of STS-launcned, spin stabilized commercial 

comnunicati on satellites intended for geosynchronous operating orbits. 

inclusion in the model is based on an assumed failuFe of the apogee kick motor 

after deployment from the Shuttle. Typical of these spacecraft, INTELSAT VI 

has several different antennas extending from one end of its cylindrical 

envelope. 

feet in diameter and 21 feet in length. 

satellites will be typically spin stabilized in the nominal Shuttle operating 

orbit at rates ranging from 30 to 55 revolutions per minute. 

Its 

@ 

Solar drums form the envelope shape, which measures close to 13 

For the assumed failure, the 

The LANDSAT program had its origin in conceptual studies and planning 

performed in the late 196Os, culminating with the launch of LANDSAT 1 in July 

of 1972. The program has focused on the development and application of space 

remote sensing technology to assist man in his understanding and management of 

the earth's resources. 

of LANDSAT spacecraft is LANDSAT D, launched in March of 1984. Like its 

predecessor LANDSAT 4, LANDSAT D carries the Thematic Mapper sensor that 

enaoles vastly superior measurement capabi 1 i ties than the sensors of previ ous 

spacecraft. An on-board propul si on subsystem was designed to accompl ish 

altitude changes between its 709-kilometer operating. omit and the nominal 

Shuttle orbit. The spacecraft can be approximated as a cylinder, over seven 

feet in diameter and 13 feet in length, weighing 4400 pounds. 

large antenna extending more than thirteen feet from its surface and a solar 

array measuring eight by fourteen feet. 

The most sophisticated and most recent in the family 

It employs a 

The Long Duration Exposure Faci lity (LDEF) was placed into its 250 nautical 

mile orbit by the STS during the Solar Maximum repair mission. 

to support a diverse set of international experiments in each of a series of 

planned flights. Close to two-thirds of its initial 21,400-pound wight 

resulted from the hardware of 57 experiments, ranging from fiber optic data 

transmission tests to packaged tomato seeds. The spacecraft measures 14 feet 

in diameter and 30 feet in length, and provided valuable experience in 

deploying a Shuttle payload nearly as large as the HST. 

totally passive and was gravity gradient stabi 1 ired during its one-year 

exposure. 

It is designed 

The facility is 
The RADARSAT, a Canadian satellite program, evolved out of Canadian 

participation and interest in the U.S. SEASAT and LANDSAT programs. 

produced the ground-based image processor to process data channeled down from 

SEASAT. The RADARSAT project is an earth resource oriented satellite system. 

It's synthetic aperture radars are designed for surveillance of North American 

shipping lanes extending to the arctic region. Three other Sensor systems are 

included in its payload which provide crop and forestry monitoring, ocean/Sea 

State monitoring for ocean meteorology, and ice flow measurement in the far 

north. The spacecraft body, similar to LANDSAT in shape, measures rougnly 15 

feet in diameter and 21 feet in length and has a weight of 13,400 pounds. Its 

protuberances include a large antenna, six feet wide and 49 feet long, and two 

large solar arrays, both ten feet wide and 49 feet long. 

Canada 

The Space Station Spartan (3s) program is intended to provide the scientific 

community with a short-duration free-flying carrier for modular single 

instrument payloads based at the Space Station. 

and of fer wick-turnaround schedules and frequent flights. 

propul sion subsystem wi 11 place the carrier into its mi ssion position, 

nominally 250 nautical miles from the Space Station. 

dimensions do not exceed that of a cylinder, fifteen feet in diameter and five 

feet in length, from which a large solar array measuring sixteen by fourteen 

feet extends. 

pointing sounding rocket program and ttw STS/Spartan program. A full system 

flight demonstration is scheduled prior to the Space Station Initial 

Operati onal Capabi 1 i ty date. 

It is desigwd to be reusable 

An on-board 

The spacecraft 

The 3s program is an evolutionary step from NASA's fineThe goal of the Upper A~sphers Research Satellite (WRS) is to extend 

scient1 fic understanding of the cnemi cal and physi'cal processes occurring i n 

the Earth's stratosphere, microsphere, and lower thermosphere. 

objective is to understand the mechanisms that control the Structure and 

variablility of the upper atmosphere, the response of the upper atmosphere to 

natural and human-related perturbations, and the role of the upper atmosphere 

in climate and climate variability. It will use remote sensing instruments, 

currently in development, to measure trace molecule species, temperature, 

winds, and radiative energy input from and lost to the upper atmosphere. 

Presently in a conceptual development stage, the UARS is scheduled for a 1989 

STS launch to a 600-kilometer operating orbit. 

Its primary 

Its weight will be 

5-6 
b 

approximately 13,000 pounds, and its dimensions can be roughly estimated as 

15 feet in diameter and 30 feet in length. 

spacecraft is very irregular and a number of protuberances extend a 

The surface configuration of the 

significant distance froin it. 

Figure 5.2-1 depicts the diversity of configuration between COBE and GRO, 

which is representative of the model. 

WEIGHT: 10,000 LBS WEIGHT: 33,000 LBS 

Figure 5.2-1 Recovery Candidate Di versi ty 

Figure 5.2-2 provides a portion of the mission model data base obtained to 

date. As previously mentioned, the early stages of development of sane of the 

programs are responsible for the lack of technical information within the 

matrix, such as the moments and products of inertia. The mission model 

presents a wide range of satellite size, mass distribution, and configuration; 

and its evolution provides the basis from which to develop and refine a multipurpose recovery system. 

5-7 
U 

=LL 0 

za x 

52 a 

5 

5 -8 
Since the TSR system is being designed to recover satellites in the mid-1990s 

time frame, it can be affected by cnanges to tne mission model. Figure 5.2-3 

depicts a mission model maintsnance strategy based on monitoring changes to 

NASA mission planning and design. Space Station evolution, NASA launcning 

strategy, and the development of individual mission designs. were identified as 

having tne potential to indirectly cause significant changes in the model. It 

is recommended that these areas then be monitored in follow-on efforts, in 

order to maintain the inode1 and assess impacts on the TSR system design. 

a 

SPACE ~ MISSION 

MODEL STAflON 

Ev~ulloN 

1 

The exi stence of tne Space Station would allow mi ssi ons planned as free-flyers 

to be collocated on a Space Station platform. Spacecraft subsystems common to 

free-flyers, i.e., electrical and propulsion, could be combined as part of a 

platform. Where the number of missions may be unaffected, the number of 

satellites would decrease. The type of satellite could also change; the 

mission model may include mostly large spacecraft for very unique missions or 

the number of Space Station Spartan spacecraft, as they are dedicated to the 

Space Station, could be increased. 

MISSION 

PROGRAW 

PROJECT 

MATURATlON 

1 

AS THE TSR SYSTEM MAY BE AFFECTED BY CHANGES TO ME MISSION 

MODEL, A MAINTENANCE STRATEGY IS RECOMMENDED TO 

MAINTAIN ME MODEL AND ASSESS IMPACTS TO THE SYSTEM DESIGN. 

Figure 5.2-3 Maintenance of Mission Model 

5 -9 
It is expected that the NASA launching strategy will be impacted as a result 

of the recent Challenger tragedy. 

the model could ~e affected as missions are delayed and decisions with regard 

to mission importance are made. 

influenced, snould expendaDle launch vehicles be considered to augment launch 

capacity. 

The quantity and types of satellites within 

Relative satellite sizes may also be 

The fact that many of the programs are in an early stage of development can 

result in revisions of the data collected to date. 

provided for the AXAF satellite, for example, have increased 25 percent over 

the course of the contract. 

as their designs mature is amplified by a revised launching strategy and/or 

the evolution of the Space Station. 

being &signed for a dedicated STS launch to its operating orbit, could be 

impacted by a revised launching strategy, under which a dedicatsd launch may 

not be appropriate. 

The moments of inertia 

The importance of monitoring individual missions 

Specifically, the GRO mission, presently 

5-1 0 
6.0 SYSTEM HARDWARE REQUIREMENTS - TASK 1 .3 

0 6.1 Introduction and Approach 

Several areas were developed and analyzed which contributkd to the initial 

definition of conceptual recovery system requi rements. They included a 

preliminary concept of operations, a functional analysis, and an analysis of 

the mission model. 

conceptual requirements and form the basis from which to create an operations 

concept and system specification and initiate follow-on efforts. 

Together, the analyses resulted in a detailed set of 

The recovery system requirements are presented in the tables on the following 

pages. 

developed as a result of the preliminary operations concept. 

requirements are based on OW, Space Station, Shuttle, and ground operations 

imposed constraints. The functional analysis was responsible for tne mission 

functional requi rements li sted in Table 6.1-2. 

mi ssi on performance requi rements; developed tnrough an analysi s of the mi ssi on 

model, they provide a refinement of tne functional requirements of Table 6.1-2. 

The mission operations requirements, provided in Table 6.1-1, 'ere 

Many of the 

Table 6.1-3 li sts a set of 

6-1 
Table 6.1-1 Mission Operations Requirements 

o The TSRS shall be OMV compatible with little 

impact on the OMV reference configuration. 

o The TSRS electrical and communication 

requirements interface will be accomplished 

through the OMV payl oad accomnodations umbil icalm 

o The nominal TSRS mission will be accomplished 

within OMV mission time limitations. 

o The nominal TSRS mission will be accomplished 

within OMV mission range 1 imitations. 

0 The nominal TSRS mission will be accomplished 

within OMV mission propellant limitations. 

0 The nominal TSRS mission will be accomplished within the OMV attitude control deadband. 

o The nominal TSRS mission will be accomplished 

within the tel eoperation control 1 imi ts induced 

by communication delays and thruster operation. 

o The nominal TSRS mission shall be accomplished 

with minimum risk of damage to the OhV. 

o The TSRS stowed diameter shall not exceed the 

Shuttle cargo bay envelope diameter. 

o Stowage of the TSRS in the Shuttle cargo bay 

shall require a minimum distance parallel to the 

centerline of the cargo bay. 

o TSRS mission operations shall be accomplished 

without exceeding Shuttle and Space Station 

proximity operations procedures. 

o TSRS mission operations shall be accompl ished 

wi thout exceeding Shuttle work1 oad 1 imitations. 

o The TSRS shall accommodate operations testing at 

a ground assembly area. 

o The TSRS shall accommodate a mating to the ON in 

either the vertical or horizontal Shuttle payl oad 

processi ng sequence. 
Table 6.1 -2 Mission Functional Requi tgments 

0 

0 

0 

0 

0 

0 

0 

0 

0 

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0 

The satellite envelopment, rigidization, stabilization, and transport loads shall not exceed OMV 

ACS limitations. 

Control of the TSRS mission operations shall be 

accompli shed from the OMV mi ssi on operati ons 

control center. 

The TSRS snall recover ground controllable and 

noncontroll able di sabled satel 1 i tes. 

The TSRS shall accommodate on-orbit EVA and 

automated subsystem reconfi gurati on at Space 

Stati on. 

Tne nominal TSRS mission shall be accomplished 

with minimum risk of damage to the TSRS. 

The TSRS shail be capable of enveloping the 

specified configurations of target satel lites. 

The TSRS shall oe capable of rigidizing contact 

with the target satellite. 

The TSRS shall be capable of stabilizing the 

target satellite relative to the TSKS and the OMV. 

The TSRS shall match the satellite spin rate and 

phase angle with sufficient accuracy to 

accompli Sh attachment. 

TW TSRS shall align its spin axis to the spin 

axis of the satellite w 

accompli sh attachment. 

Tne TSRS mall ~e capab 

of the satellite center 

tne OMV major thrust ax 

tn- sufficient accuracy to 

e of remote realignment 

of mass, with respect to 

s, after stabilization. 

6-3 
. Table 6.1-3 Mi ssi on Performance Requi rements 

o The TSRS shall recover satellites with or without 

Rec o ve ry Support El eme n t s . o The TSRS shall recover satellites with recessed 

Recovery Support Elements due to the satellite's 

de p 1 oyed conf i gu ra ti on. 

o The TSRS shall accommodate a range of satellite 

Sizes not to exceed a cylindrical envelope of 

180-i nch diameter and 590-i nch length. 

o The TSRS shall accommodate the range of grapple 

diameters from mal 1 nard points to 780 i nches. 

o The TSRS shall recover satellites in stabilized 

and nontorqued, nonstabi 1 i zed conf i gurati ons. 

o The TSRS shall accommodate 11 degrees of geometric 

coni ng in a nontorqued, nonstabi 1 i zed conf i gurati on. 

o The TSRS snall recover satellites having irregular 

surfaces ranging from 50% to 100% of the saellite 

e nvel ope. 

o The TSRS shall accommodate the range of spin rates 

from zero to 55 revolutions per minute. 

o The TSRS snall recover satellites with protuberances extending from the satellite envelope with 

lengths ranging from 51 inches to 757 inches. 

cnangeabi 1 i ty to speci f ical ly accomnodate the 

wide range of recovery scenarios. 

o The TSRS shall provide maximum subsystem intcro Tne TSRS shall accommodate timely and economic 

o 

subsystem reconfi gurati on. 

Tne TSRS and the TSRS/OMV interface shall accommodate torques associated with attacnment, 

ri g i d i tati on, stabi 1 i zati on, and transport. 

o The nomi nal TSRS mission shall be accompli shed 

with minimum damage to the recovered vehicle. 

o The contact forces between the satellite and the 

TSRS afte r envel opment and ri gi di zati on sha 11 be 

sufficient for stabi li zation and transport. 

6-4 
6.2 Tumbling Satellite Recovery System Operations Concept 

0 6.2.1 

for a modularly designed TSR system at this early stage in the conceptual 

definition phase was considered essential to the development of a broad range 

of mission operations requirements, including those related to ground and 

aerospace support equipment. 

on deployment from the Shuttle and from the Space Station. 

TSR concept of operations supported both the completion of a functional 

analysi s of on-orbi t operations, and the operational analysi s conducted on the 

six selected design reference ni ssi ons. 

Introduction - The generation of a preliminary concept of operations 

The operations concept included operations based 

Tnis preliminary 

The tumbling satellite recovery system will be one of a family of kits to be 

developed to conduct operations remotely, as front ends for a ground 

controlled, teleoperated OMV. 

spacecraft services avail able through the OMV structural and electrical 

payload accommodations interfaces. The TSR kit equipment and supplies will be 

collocated with OMV at ground storage facilities. The TSR ground operations 

console(s) will be situated at the OMV Operations Support Center to enable 

coordination of training and mission operations. 

elements of the TSR kit will be situated at the OMV storage facility. 

The kits will be designed to utilize fully the 

At the Space Station, the 

As described previously, the TSR kit is conceived to be a modular system, 

comprised of a number of subsystem mechanisms thgt can be readily integrated 

in varying combinations. This will enable the user to quickly configure a 

specific remote recovery system to meet a broad spectrum of unique mission 

requi reinents. 

6.2.2 Operations - The TSR operations concept is outlined in Figurx 6.2.2-1. 

The TSR kit elements would be stored at the OlyV ground storage facility. For 

each mission the requi red subsystem elements would be assembled in the OMV 

front end kit assembly area and ground tested, using recovery kit ground 

support equipment. 

facility, the recovery kit would be transported to the STS payload processing 

facilities at the STS launch site. 

Following assembly and checkout at the ground support 

6-5 
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At the launcn facility, like the OMV, the TSR kit would undergo further test 

and checkout prior to a mating witn the OMV in the norizontal or vertical 

payload processing sequence, as selected by OMV program directors. 

assembly and checkout approach recommended for the OMV recovery kit is to 

emphasize ground testing and verification, with necessary adjustments and 

replacements done on the ground. If recovery kit subsystems were to fail 

during on-orbit checkout, it will ix difficult to replace them at tne 

Orbiter. 

mated TSR kit will be deployed from tne cargo bay with the Orbiter RMS. 

Orbiter will then be maneuvered away from the mated OMV/TSR kit to a safe 

distance for the OMV orbit transfer. The OMV will then transport the attached 

recovery kit to a rendezvous with the disabled satellite. 

The 

Following tne launch into an operating/standby orbit, the OMV and 

The 

The actual recovery operation will commence with visual sighting of the 

disabled satellfte. Tne OMV will maneuver to witnin visual range of the 

satellite and commence actual recovery operations. These operations are 

described as part of tne functional analyses of paragraph 6.2.3. 

satellite recovery operations will be controlled from the ground-based OblV 

Operations Support Center (OSC), so mission control is transferred to the OSC 

The on-orbit 

.. 0 at this time. 

Upon completion of a full-up System C envelopment recovery operation, the 

satellite will be stabilized. It may be necessary to release the enveloped 

target and regrapple it to align the new center of mass with the OMV omit 

transfer thrust vector. The OMV ACS tnrusters will be fired in small 

translation maneuvers to determine the proper positioning of tne payload 

center of mass, prior to orbit transfer to the Orbiter. 

The recovered sate1 li te wi 11 be posi ti oned in the Orbiter usi ng tire RblS end 

effector, in the Orbiter trunnions or in a pre-configured cradle arrangement. 

The OMV and TSR kit will be repositioned in the cargo bay, and the Orbiter 

will return its new cargo to the appropriate STS launch/landing facility. 

6-7 
6.2.3 Mai ntenance/Refurbi shment - When the recovery mi ssion has been 

completed and the Orbiter has returned to the launch site, the OMV satellite 

recovery kit will be detached from the OMV and returned to the OMV front end 

kit assembly and checkout area. 

refurbished and/or repaired, and returned to storage for follow-on mi %ions. 

Here, the kit will be disassembled, 

6.2.4 Operations Control - The OMV satellite recovery kit will be operated by 

a ground controller and will require adequate ground consoles and ground 

support equipment. This equipment will be located with the OMV ground control 

console(s) at the OMV Operations Support Center. 

kit command and data management formats wi 11 be exactly like those used by 

OW, as the recovery kit will be linked for C&DM through the OMV payload 

umbilical(s). 

The OMV satellite recovery 

6.2.5 Space Station Operations - When the Ot4V satellite recovery kit is based 

at the Space Station, the kit equipment will be transported to and stored in 

an OMV kit storage aEa. As is true in ecovery operations that are initiated 

from the ground and conducted from the Oroiter, the mission recovery kit used 

on any mission will be tailored for tnat specific mission. The kit will Dt! 

assembled and attached to the OMV by astronauts on EVA or by robotic or 

teleoperated manipulators. 

be deployed from the Space Station by the mobile RMS. 

mission will be controlled from the ground at the OMV Operations Support 

Center. 

refurbi shment wi 11 be conducted at the Space Station, either by EVA, 

te 1 e ope rati on or automat i c a I 1 y . 

Once the OMV and recovery kit are mated, they wi II 

The remainder of the 

Upon return to the Space Station, maintenance, repair, or 

6.3 Functi mal Analysi s 

A functional analysis of a full-up System C recovery operation was conducted 

to support the development of mi ssion functional requi rements. 

is shown in hignlighted "flowchart" form in Figure 6.3-1. The actual recovery 

operation commences with the rendezvous to within visual range of the disabled 

satellite. The OMV will circumnavigate the satellite to determine target 

motion orientation and rates. The target is expected to be spinning in a 

This analysi s 

6-8 
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It 

0 

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flat, single axis spin about the mdor principal axis of maximum moment of 

inertia. The OMV and TSR ground control lers, usi ng OMV and TSR TV cameras 

wi 11 evaluate protuberances, i ,e., solar arrays, antennas, instrument packages, 

and develop an approacn strategy. 

The TSR operator will determine the spin axis, as the approach to the target 

will take that path, but the OMV operator will approach from that side of the 

spinning target that offers the least interference. 

The OMV operator will then align the OMV translation axis with the target's 

spin axis and translate to a distance short of deployment and target 

envel optnent di stance. 

The next phase of recovery operations involves TSR system equipment 

deployment. 

stability and will be conducted in a sequence designed to minimize dynamic 

interaction between tne TSR kit and the OMV. First, the extendible boom Mill 

be fully deployed. Next, the large envelopment-type grapple mechanim,wilI AX 

deployed to an envelopment diameter, more than adequate for the target's 

grapple position configuration. Following this, the spin mechanism will be 

activated to spin up the recovery system to match tne target spin rate and 

phase relationship of the target. The OMV operator will avoid actions that 

mignt result in plum impingement on the target. 

Deployment of the system has the potential of impacting OMV 

The OMV and TSR kit operators now conduct envelopment operations, including 

adjustment of spin rate and phase during translation to an envelopment grapple 

range, vis-a-vis the target and the recovery system grapple mechanism. 

the recovery kit grapple mechanism is in position, centered as closely as 

possible to the target's center of mass, the ground controller will slowly 

close the jaws of the grapple mechanism. 

associated contact dynamics, will be avoided until envelopment is achieved. 

When 

Inadvertent contact, and the 

The grapple mechanism will then be closed further to provide a rigid grapple 

of the target so tnat the ground controller can then "despin" the captured 

satellite, using a reverse braking action of the direct current (DC) torque 

motor in the spin table. OMV reaction control thrusters wi 11 provide reacti ve 

impulse to maintain a stable system during despin operations. 

6-1 0 
When the OMV/TSR kit and captured satellite "system" is stabilized, the TSR 

kit controller may need to release the target and regrapple it to rpfine the 

alignment of the total system's center of mass with the OMV orbit trdnsfer 

thrust vector. 

release and regrapple may be required to realign the new system center of 

mass, to ensure it is within the OMV center of mass offset capabilities. 

will enable OMV to translate and control the system during orbit transfer to 

tne Orbiter. 

0' An iterative series of OMV translational maneuvers and target 

This 

6.4 Mission Model Analysis 

6.4.1 Introduction - Analysis of the mission model was accomplished in order 

to define the boundaries of expected or possible recovery scenarios, in tzrms 

of recovery candidate characteristics. 

the model, the analysis was maintained at a top level of complexity, producing 

quantities appropriate for this early phase of TSR development. 

Given the nature of the data within 

A portion of the analysis resulted in the Mission Model Derived Baseline, a 

composite of worst case recovery candidate characteristics, shown in Table 

6.4.1-1. 

was derived independently and together do not apply to any single recovery 

candidate. 

subset of mi ssi on performance requi rements. 

Each of the characteristics, althougn assembled to form a composite, 

The baseline enabled the development and quantification of this 

To further define the recovery scenario an effort was ma& to bound the 

diameter necessary to surround a recovery candidate and to verify an 

extendible boom requirement. 

were thereby applied to a recovery candidate with the characteristics of the 

Mission Model Derived Baseline. Issues werp identified and analyzed, 

clarifying the recovery scenario and providing additional requirement 

definition and refi nement. 

The previously defi ned functions and operations 

6-1 1 
Table 6.4.1-1 Mission Model krived Baseline 

ENVELOPESHAPE: CYLINDRICAL, IRREGULAR SURFACE (50-100% OF ENVELOPE) 

EWELO PE DlAM€fER 180 INCHES . 

LENGTH: -50 FEET 

MASS: 34,000 POUNDS 

IN Y-Z PLANE, 

MAXIMUM 11 DEGREES OFF PLANE 

MOMENT OF INERTIA ABOUT PRINC IPAL AXIS; 125,000 SLUG-FOOT 

mm M 68,000 FOOT-POUND-SECOND 

SPIN RATE; 55 REVOLUTIONS / MINUTE 

6.4.2 Derived Baseline Definition - The first step in deriving ttx composite 

of worst case characteristics involved the definition of satellite 

configurations, i .e., shape, dimensions’and mass, and was accompli shed through 

simplification and categorization of the mission model. 

tne satellites was graphically represented by an envelope shape within which 

all surfaces of the satel lite, excl’udi ng protuberances, would lie. 

Protuberances, strictly solar arrays and antennas, we analyzed separately as 

their locations and dimensions varied significantly. It was then possible to 

neatly categorize the envelope shapes, based on sire, conformance to the 

envelope or surface irregularity, and the satellite mass. As a result of the 

nature of launch vehicle payload accommodations, the envelope shapes were all 

cylinders or modified cylinders. 

di scri mi nators are provi ded in Figure 6.4.2-1. 

Initially, each of 

The resultant categorizations and their 

The largest and smallest volume or envelope size, largest and smallest mass, 

and smooth versus irregular surfaces weere the obvious extremes within the 

discriminators. 

characteristics in o.rder to satisfy boundary definition, although it was 

understood that the smallest volume and mass, previously mentioned in the 

The largest volum and mass were chosen as worst case 

6-1 2 
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6-1 3 .I 
sumnary for the GPB satellite, produced additional, unique recovery 

requi rements. 

were included in the composite along with the GRO weight of 34,000 pounds. 

irregular satellite surface was considered to provide a more difficult 

recovery than a smooth surface and was also selected as a worst case 

characteristic. The fiftn category of Figure 6.4.2-1 included the most 

irregular surface. 

envelope diameter, it provided the level of conformance of the actual surface 

to the envelope diameter for the composite. 

sdtel li te configurations are representati ve of this level of surface 

i rregul ari ty. 

Tne HST dimensions of 15 feet in diameter and 50 feet in length 

An 

Defined as varying between 50 and 100 percent of the 

Both the GRO and the UARS 

An analysis of tne mass distributions within the mission node1 was 

accomplished in an effort to define the worst case values for moments of 

inertia and the location of the principal axes of inertia. 

The various satellites in the model ware unexceptional in being designed for 

stabiiity, as evidenced by the relatively mall products of inertia. 

Therefore, for all cases the principal axes were assumed to be not far off the 

geometric axes. 

calculations were used to determine the location of the principal axes and the 

magnitude of tbe moments of inertia about the principal axes. 

satellites in the mission model with sufficient data to make the calculations 

provided a range of locations of the principal axes with respect to the 

geometric axes. 

In taking advantage. of this, simplified two-dimensi onal - 

Those 

As 'was described in Section 4.4.3, an initially torqued, disabled satellite 

having damping provisions will reach a steady state spin about that principal 

axis with the largest moment of inertia. 

moments of inertia about the three principal axes were compared for each of 

the individual configurations. 

Z transverse axis, with the Y transverse axis being very close in magnitude., 

The calculated magnitudes of the 

In each case the major principal axis was the 

The accuracy of the analysis was considered sufficient to generally define one 

of the transverse principal axes as the major principal axis, and to bound the 

location of the principal axis off the transverse geometric axis. 

locating the principal axis, with respect to the previously defined envelope 

The region 

shape, is shown in Figue 6.4.2-2. 

6-1 4 
\ / 

Y 

Figure 6.4.2-2 Location of Major Principal Axis 

The magnitudes of the calculated moments of inertia about the principal axes 

approximated the values of the moments of inertia about the geometric axes. 

For those satellites for which products of inertia were not provided, the 

value of the moment of inertia about the major principal axis \vas estimated as 

the largest moment of inertia about the geometric axes. The largest value of 

moment of inertia about the principal axis is 125,000 slug-feet squared, for 

the AXAF satellite, and was included in Figure 6.4.1-1. 

a 

The final step in the derivation of the composite baseline involved the 

determination of worst case satellite motion. Two selected disabled satellite 

fai1uE modes were assumed in order to determine a maximum angular momentum 

and spin rate for three axis stabilized spacecraft, as a mference of 

comparison to the worst case motion characteristics of spin stab1 lized 

spacecraft. Failure mode A assumed the failure of the reaction wheels of an 

atti tude control system, effecti vely transferring thei r momentum to the 

spacecraft. FailuE mode 3 involved the fai1ul.e open of an attitude control 

system thruster exhausting the available propellant. The location of the 

thrusters rclative to the center of gravity of the spacecraft, and the type 

and amount of available propellant *re required to calculate the angular 

6-1 5 
momentum produced by failure mode 8. In the cases where sufficient data was 

available, it was assumed that the available propellant was restricted to the 

propellant of one tank, even in the presence of tank cross-strapping, due to 

cross-strapping isolation valves. 

detenni ned simply as the maximum momentum transfer capabi 1 ity of the wheel 5. 

Conversations with Hughes and Ford Aerospace defined tne maximum spin rate 

occurring in low earth orbit for disabled spin stabilized spacecraft as 35 to 

55 revolutions per minute, for the Hughes 393 and 376 DUS respectively. 

angular momentum for the larger and more recent 393 bus bias calculated as the 

product of the moment of inertia about the spin axis and the spin rate. 

For failure mode A the angular momentum was 

Tne 

MOTION 

CHARACTERISTICS 

I 

I 

The maximun angular momentum and spin rate of spin stabilized spacecraft and 

of thme axis stabilized spacecraft, as a result of failure modes A and 6, are 

snown in Figure 6.4.2-3. The angular momentum value and resultant spin rate 

for failure mode 6 did not include the Mission Model Derived Baseline value of 

moment of inertia about the major principal axis, but was calculated uniquely 

for each satellite of the model. The calculation for GRO produced the maximum 

Val ue of angular momentum of 68,000 foot-pound-second. 

FAIlUHt -A FAllURt mt B 

MOMENllJM THRUSTER SPIN 

STAB'L1ZED TRANSFER FROM OPEN 

SPACECRAFT REACTION WHEEL FAILURE 

ANGULAR 

MOMENNM 

( FT-LB-SEC ) 

SPIN 

RATE 

( REV/MIN ) 

16,800 1260 68,000 

55 2.4 9.0 

Figurp 6.4.2-3 Angular Momentum and Spin Rate . 
6.4.3 Derived Base1 ine Application - The diameter necessary to completely 

surround a recovery candidate, normal to its major principal axis, has been 

termed "envel opment diameter. " The effort to bound envel opnent diameter and 

further define the recovery scenario assumed the envelope dimensions provided 

by the Mission Model Derived Baseline as an initial reference. Three 

independent issues were identified which affect the envelopment diameter 

magnitude: misalignment of the major principal axis and the geometric axis of 

the recovery candidate (geometric coning), misalignment of the spin axes of 

the recovery candidate and the tumbling satellite recovery system, and the 

deadband of the OMV attitude control system. 

on envelopment diameter is depicted in Figure 6.4.3-1. 

Geometric coning and its affect 

In the presence of geometric coning, the effect on the magnitude of the 

envelopment diameter is dependent on the location of the major principal axis, 

or the type of geometric coning. As previously shown, the misalignment of the 

principal axes and the geometric axes is slight, so that the principal axes 

can be described as longitudinal and transverse. Geometric coning is produced 

by misalignment of the longitudinal principal axis with respect to the 

1 ongi tudinal geometric axes, and mi sal i gnment of one of the transverse 

Principal axes with respect to its transverse geometric axis. The 

misalignment of the transverse principal axes will be within the geometric 

transverse axis plane and/or off the transverse geometric axis pl ane towards 

the longitudinal geometric axis. 

therefore does not affect envelopment diameter. The magnitude of the change 

in envelopment diameter varies considerably for the remaining two cases Of 

geometric coning. 

a 

The former produces no geometric coning, and 

The two cases, 1 ongi tudinal and transverse, are respectively depicted in 

Figure 6.4.3-1. It can be seen that rotation about the principal axis creates 

an increase in envel opment diameter. For the previously defined maximum angl e 

of misalignment between the principal and geometric axes, the magnitude of the 

change in envelopment diameter of the longitudinal case is significantly 

greater than that of the transverse case. 

angle of misalignment for spin about the longitudinal principal axis is much 

smaller than that for spin about the transverse principal axis. 

Therefore, the tolerance of the 

6-1 7 
% 

P 

tu 

- E 

c 

c tu 

0 I -?e- --/ c 

Q 

n c 

i, 

E. 

6-1 a 
Although geometric coning produced by misalignment of the longitudinal 

principal axis was examined, it was previously shown that the longitudinal 

principal axis will generally not be the major principal axis. 

of spin about the longitudinal axis cannot be ruled out, however, due to spin 

stabilized spacecraft and exceptions within the mission model for which the 

diffeence between the moments of inertia about the longitudinal and 

transverse principal axes becomes difficult to distinguish. 

stabilized spacecraft, angles of misalignment are normally in the range of 

fractions of a degree, which represents approximately a one-inch increase in 

envelopment diameter for Intelsat VI. for the exceptional case, a mall 

spacecraft, the increase in envelopment diameter will be significant relative 

to its dimensions, but in terms of the more than 180 inches required to 

envelope the larger spacecraft, it will not effect an overall increase in the 

TSRS envelopment diameter capability. 

principal axis, the driving case, produces an increase in envelopment diameter 

of approximately four inches, based on the envelopment dimensions of the 

Mission Model Deri ved Baseli ne. 

The occurrence 

For spin 

Misalignment of the transverse 

Tne second issue affecting envelopment diameter, the misalignment of the spin 

axes of the recovery candidate and the TSR system, is illustrated in Figure 

6.4.3-2. The misalignment of the spin axis increases the envelopment diameter 

in much the same way as geometric coning. 

been bounded, although the esultant maximum increase in envelopment diameter 

is not expected to be severe. 

The angle of misalignment has not 

The OMV as used in a TSR mission will have two controllability deadbands. 

fi rst is associated with the automatic pi lot attitude control and produces 

rotational deadbands for roll, pitch, and yaw. The second is effected by 

te leope rator capabi I i ty, dependent on feedback time delay and the thruster 

impul se magnitudes. 

The 

The effect of tne attitude control deadband was addressed and is depicted in 

Figure 6.4.3-3. The desi red initial contact of the TSR system with the 

recovery candidate was designated as 15 degrees off the diameter normal to the 

approach direction, to influence a reaction in the direction of the the TSR 

system and the OMV along the approach diameter. 

deadband of one degree results in a six-incn translational deadband at the 

point of contact, equi valently i ncreasi ng the envelopment di meter. 

An assumed rotational 

6-1 9 
Transverse 

Principal 

Axis 

Figure 6.4.3-2 Misaligment of Spin Axes 

Dirtma from 

Tmt to OMV 

Initial Point of 

Attitud. Control 

l’, OMV Attitud. Control Dudbnd r 

ow 

Figure 6.4.3-3 OMV Controllabi lity Deadband 

6-20 
Several issues were identified which provide justification for an extendible 

boom requirement. Of these, thrte addess the proximity of the OMV to the 

recovery candidate, where an extendi ble boom dimi ni shes the possi bi 1 i ty of 

contact between the vehicles; and two are issues of access and aiigrment, 

accommodated by a boom having multiple degrees of freedom. 

is shown in Figure 6.4.3-4. 

Their relationship 

MlSALlGNME 

OF TARGET 

n 

DISTANCE ISSUES 

tI SURFACES I 

OBSTRUCTED 

RECOVERY . 

SUPPORT 1 ELEMENCS b1 

n n 

REALIGNMENT 

OF TARGET 

CENTER OF 

MULTIPLE 

DEGREEOFFREEWM 

WES 

Figure 6.4.3-4 Extendible Boom Issues 

The first issue, geometric coning about a transverse axis, is depicted in 

Figure 6.4.3-5. The condition establishes a distinct barrier between the 

recovery candidate and the OMV. To avoid vehicle damage, as the recovery 

candidate is spinning relative to tne stable OMV, a distance greater than the 

''barrier" distance must be preserved. For the angle of misalignment of the 

Mission Model Derived aaseline, the "barrier" distance is 18 inches. 

6-21 
"Safe" Distance Limits Proximity of OMV so That 

Spinning Surface of Satellite Doer Not Contact 

Stable Sllrfrce of OMV 

Figure 6.4.3-5 Geometric Coni ng 

Figure 6.4.3-6 shows sate1 lite surface i rregulari ties typical of those of the 

Mission Model Derived Baseline. As can be seen, a distance is equired so 

tnat after envelopment, rigidization will not cause the surface of the 

recovery candidate away from the point of envelopment to contact the OMV. 

surface configuration at and near the point of envelopment varies with each 

recovery candidate. The "safe" distance, however, will not exceed that for a 

surface with a 50 percent to 100 percent diameter at and near the point of 

envelopment. The 50 percent to 100 percent ratio results in a 45-inch ''safe" 

di stance. 

The 

Tne renai ni ng issues, obstructed RSEs, CG realigrment, and protuberances, were 

identified but not completely analyzed. RSEs accessiale to the Shuttle RMS 

can ~e obstructed by deployed protuberances, requiring a multiple degree of 

freedom extendible boom, in order to provide the proper gripper orientation 

for recovery. 

CG location re1 ati ve. to the TSR system and the @lV, exceeds the CG off set 

limitations of the OMV. As was discussed in the functional analysis, the ealignment may be accommodated by articulation of a multiple degree of freedom 

extendi Dle boom. 

Realigrment of the recovery candidate CG is required, where the 

6-22 
Actual Surface 

I 

I 

X 

I 

When Actual Diameter Is Lm than Envelopa 

Diameter, OMV "Safe" Distance Required to Keep 

Spinning Surface of Satellite from Contacting Stabla 

Surfitl of OMV. 

Figure 6.4.3-6 Surface Irregularities 

@ Protuberances, in general, were reviewed for each of the satellites in the 

mission model. Dimensions were sumnarized for both solar arrays and antennas, 

but no effort was made to provide a boundary dimension to include in the 

Mission Model Derived Baseline. 

surface are 0 to 63 feet and 4 to 17 feet for solar arrays and antennas 

respectively. 

based on protuberances, was considered to be too severe to accommodate vJith an 

extendible boom. 

The ranges of extension from the satellite 

The proximity limitation of the OMV to the recovery candidate, 

6-23 
7.0 DESIGN REFERENCE MISSIGNS (DRM) - TASK 1.4 

e 7.1 Introduction 

The Design Reference Mission (DRM) task, contract statement of work Task 1.4, 

was interpreted by MhC to include: 

group of DRMs covering a broad range of potential remote recovery scenarios; 

and (2) operations analyses of these DRMs to support refinement of 

requirements. 

(1) the selection (and MSFC approval) of a 

A set of six DRMs was chosen early in the study and included two scenario 

"cases" for each of the three MSFC defined recovery systems, Systems A, B and 

C. Each of the DRMs was analyzed in detail and a sequence of events for each 

was outlined to display the results of functional and operational analyses. 

The DRM operations analysis included a breakout of: 1) required "pre-mission" 

activities; 2) "specific" or direct mission activities and 3) post-mission 

activities. The "specific" mission events were those activities directly 

included in the actual conduct of the recovery mission. 

were those activities, required upon mission completion, that would ensure 

continued orderly Space Transportation System or Space Station operations, 

such as cleanup operations, refurbishment and storage of equipment and tools. 

The DRM event sequencing included a detailed description of recovery 

activities. 

Post-mission events @ 

7.2 DkM Operations Analyses 

7.2.1 

remote recovery missions in which the basic OMV, without any TSR elements, 

would conduct recovery operations. 

designed with two types of retrieval /recovery support mechanisms. 

is the standard STS RMS end effector, a grapple mechanism designed to enable 

grapple and rigidization with an RMS grapple fixture situated on the 

spacecraft, and that is accessible to a large transport vehicle like the OMV. 

The other OMV recovery support mechanism is a set of STS flight support 

structure (FSS) docking latch pins to enable retrieval or recovery of 

spacecraft where these FSS berthing pins are accessible to the GIliV. 

DRM 1 - System A, Case 1 - This DRM and DRM 2 were designea to describe 

The MSFC reference configuration OMV is 

The first 

7-1 
An onorbit recovery scenario for DRM 1 is illustrated on Figure 7.2.1-1. This 

System A (basic OMV) recovery DRM describes a realistic mission in which the 

OMV captures the Hubble Space Telescope (HST), using the OMV FSS berthing 

latches to mate with HST's FSS latch pins. 

mandated by a set of instrument failures that occur prior to the first planned 

HST maintenance mission, and has left HST with an effective capability of only 

20%. 

spacecraft 'S atti tude control systems are operable and the satel 1 i te is. stab1 e. 

This "recovery" mission is 

For this DRCl, the HST is significantly disabled; however, the 

Baseline OMV h 

Space Telescope r with FSS Berthing 

OMV with FSS 

Berthing Latches 

Pins 

Case 1: Stable Target-FSS Capture 

Figure 7.2.1-1 Design Reference Mission 1, System A, Case 1 

This recovery DRM is initiated from the Space Station, as the Space Station 

ana OMV are both assumed fully operational and the reusable OMV has been 

tested and fully integrated into the Space Station, located at 28.5 degrees in 

inclination. 

7.2.1.1 

DRM 1 are presented in Table 7.2.1.1-1. 

fully into Space Station operations prior to this mission. The OMV procedure 

for retrieval ana for recovery in this case should be very similar to a normal 

OMV recovery mission. 

prior to conduct of the DRM. 

DRM 1 Pre-Mission Activities - The primary pre-mission activities for 

The OMV will have been integrated 

However, they must be developed, tested and exercised 

7-2 
Table 7.2.1 -1-1 Pre-Mission Tasks 

e Ground 

- 

Devel op Operati onal Procedures 

- OMV RendezvouslTarget C1 ose 

- OMV Grapple and Rfgidiratfon - OMV Orbit Transfer to Space Station 

- OMV/HST Separation 

- Conduct recovery operations using OMV Operations Support Center 

Space Station 

- Conduct local exercises on routine recovery operations, and OMV 

dep 1 oymen t /recove ry 

7.2.1.2 DRM 1 Mi ssi on Acti vi ties - Thi s missi on i s assumed to commence when 

the OMV is deployed from the Space Station, probably done with a mobile RMS. 

This is shown in Table 7.2.1.2-1. The OMV will be controlled to transfer to 

the immediate vicinity of the HST and achieve rendezvous. 

operated by OMV ground control operators throughout this mission. There are 

no TSR elements on the OMV for this DRM and thus no TSR requirements will be 

derived from this mission. OMV will navigate around the target, HST, and the 

teleoperator will align the OMV for its translational maneuvers toward the FSS 

berthing pins. 

HST and prepare to return to Space Station. 

recovery of an HST that is remote from the Space Station. 

The ONV will be 

The operator will achieve a firm three-latch hookup with the 

These activities will enable 

7.2.1.3 DRM 1 Post-Mission Activities - Following this mission, the OMV will 

be refurbished and stored in storage depots at Space Station. These 

activities are strictly OMV-related and will not be detailed here as they do 

not i nfl uence TSR requi rements. 

7.2.2 DRM 2 - System A, Case 2 - The second System A DRM is designed to 

outline an OMV remote retrieval/recovery operation using the OMV RMS end 

effector/grapple mechanism as a recovery support tool. 

DRM, the Cosmic Background Explorer (COBE) instrument package was considered 

to be totally failed. 

attitude control system (ACS) is operable and the satellite is in a 

semi-control 1 able mode. The COBE pointing control reaction wheel system has 

failed, but its magnetic torquers and momentum wheels prevent rotation above 

In the second recovery 

However, as in the HST failure, the satellite's 

7 -3 
Table 7.2.1.2-1 DWI 1. Mission Event Sequence 

I - Event 

Man and Activate OMV Operations 

Checkout OMV Ground Control 

Support Center 

Consol e/Equi pment 

Prepare GMV at Space Station for 

Recovery Mi ssi on 

- Remove from Storage 

- Fuel 

- Conduct Checkout 

Geploy from Space Station, Transit 

to Safe Standoff Distance, with 

Cold Gas Engines to Main Engine 

Ignition Position 

Grbi t Transfer to Rendezvous with HST 

I Rendezvous and Vi sua1 ly Acquire HST 

OMV Ground Control1 er Conducts Inspection 

Maneuvers Around HST 

HST POCC Operator C1 oses Contamination 

Shields, Retracts Solar Arrays and 

Antenna, Inerts Atti tude Control 

Sys tem 

OMV Ground Controller Aligns OMV for 

Transit to HST Berthing Pins 

OblV Ground Control 1 er Grapples HST 

Berthing Pins with OMV FSS Docking 

Latches, Ac uires Firm Grapple for 

Orbit Trans 4 er 

OClV Ground Control 1 er Conducts Transfer 

of OCiV/HST to Space Station Vicinity 

Space Station Mission Control Assumes 

Mission Control and Completes Recovery 

of OMV/hST 

I 

I 

7-4 
one degree per second on the roll axis and close to zero on the other two 

axes. 

matching this target roll rate. 

The OMV reference configuration vehicle is expected to be capable of 

This DRM was formatted to be a dedicated STS mission, as the COBE satell ite 

will be in a 99 degree inclination polar orbit. 

also to demonstrate the widest possible range of System A remote recovery 

scenarios. 

configuration OMV is shown prepared to match the one degree per second spin 

rate of the COBE satellite and attach the OMV RMS end effector to the COBE's 

MtS grapple fixture. 

The STS scenario was chosen 

DRM 2 is illustrated in Figure 7.2.2-1. The MSFC reference 

Baseline OMV 

Standard RMS 

Grapple Fixture (GF) 

Standard RMS 

End Effector 

Cosmic Background 

Explorer (COB E) 

Case 2: Stable Target-RMS GF Capture 

Figure 7.2.2-1 DRM 2, System A, Case 2 

7.2.2.1 

highlighted in Table 7.2.2.1-1. 

requirements, so they are presented at a top level. 

mission is baselined to begin at rendezvous with COBE. 

DRM 2 Pre-Mission Activities - For DRM 2, the pre-mission events are 

None of these activities impact TSR 

The actual recovery 

7 -5 
Table 7.2.2.1-1 DRM 2 Pre-Mission Activities . 

- Event 

Prepare OMV for Launch in STS 

I Devel op Recovery Operating Procedures 

- Rendezvous 

- Approach & Survey Target 

- Grapple & Despin - Rigidize for Transport to STS 

I - Match Target Roll Rate 

I Perform Simulations, Exercises 

Launch CMV to Standard STS Operating 

A1 ti tude 

I Deploy OMV from STS with RMS 

Maneuver STS to Safe Observation 

Transfer OMV to Rendezvous with COBE 

Distance From OW 

7.2.2.2 DRM 2 Mission Activities - Control of the recovery mission is 

transferred to the OMV ground control center during OMV transit to the 

operational orbit of COBE, currently planned for a 486 nautical mile circular 

orbit. The DRM 2 recovery mission events are outlined on Table 7.2.2.2-1. 

The OMV ground controller maneuvers OMV to a position to rendezvous with the 

target including visual acquisition and then proceeds to circle the failed 

satellite to gather data to develop an approach strategy. 

The ground controller will maneuver OW to align with the target's axis of 

spin, match the spin rate of the disabled satellite, extend the OMV RMS end 

effector to snare the grapple fixture on COBE, despin the satellite to a 

stable condition, and retract the snare wires into the end effector to provide 

a rigid OMV/COBE mate for transfer back to the Orbiter. As in the case of DRM 

1 

requirements, and thus are described in general terms. 

the events of this DRM do not impact on tumbling satellite recovery 
Table 7.2.2.2-1 DRM 2 Mission Activities 

Event 

Ground Controller Maneuvers OMV to 

- 

Vi sua1 ly Acqui re COBE 

OMV Maneuvered to Ci rcumnavi gate 

COBE and Select Approach Strategy 

OMV Maneuvered to Close Proximity of COBE to Match Spin Rate 

With OMV Translating in Circular 

Manner to Match Rates, OMV MS 

is Extended and Grapple of COBE 

Achieved 

OMV Despins COBE to Stabilize Satellite 

OW Ground Controller Uses RMS End 

Effector to Rfgidize COBE to OMV 

for Transport to STS 

OMV Comnanded to Return COB€ to STS 

Mission Control Transferred Back to 

STS . 

7.2.2.3 DRM 2 Post-Mission Activities - Following completion of the recovery 

mission, the OMV and the COBE satellite'will be positioned in the Orbiter 

cargo bay for return to earth. 

satellite into transfer support cradles and/or STS trunnion mounts, and 

similarly will be used to position OMV on the STS longeron and keel attach 

fittings in the cargo bay. 

OMV will be refurbished and prepared for a future mission and COB€ will be 

returned for repair and/or upgrade and a return launch to its operational 

orbit. 

The Orbiter RMS wi'll be used to place the COBE 

The Orbiter will then be returned to earth where 

7.2.3 DRM 3 - System 8, Case 1 - DRMs 3 and 4 were developed to describe 

examples of the type of missions a satellite recovery "System B" would be 

designed to conduct. System B is the recovery system envisioned to deal with 

situations where the disabled satellite is: 

a. Outside of Orbiter range (either higher or lower than efficient STS 

operating a1 ti tudes 1; 

7 -7 
b. 

C. 

d. 

Controllable from the ground and stable, i .e., within OMV 

retrieval /recovery operating capabil f ties; 

Devoid of accessible recovery support devices (RMS grapple fixture or 

FSS latch pins); 

Not spin-stabilized or spinning at speeds not in excess of OMV spin 

matching capabilities, which are projected to be very low. 

The recovery target for DRM 3 is the LANDSAT-D satellite that has, without 

prior warning, suddenly 1 ost the remaining power of a marginally operating 

solar panel. The LANDSAT-D i s i n a 380 nautical mile circular polar orbit. 

The satellite has not as yet developed significant rotation and is generally 

stable. The LANDSAT-D failure mode did not introduce torque into the 

satellite system and LANDSAT'S altitude of 380 nautical miles will minimize 

short term atmospheric and gravitational torques. As this is a polar orbit 

recovery, the mission base of the recovery operation is the Orbiter, and not 

the Space Station. 

7.2.3.1 Dh 3 Pre-Mission Activities - The activities requisite to initiation 

of DRM 3 are similar to those for DRM 2, shown on Table 7.2.2.1-1. 

additional preparation activity relates to the use of the System B recovery 

kit. 

recovery system will consist of an extendible bootii, a grapple mechanism 

interface device, and, for DRM 3, a standard RMS end effector, to serve as a 

grapple mechanism. These recovery system elements will be collected from the 

OMV and OMV kit storage area and assembled and checked out, using ground 

support equipment. The System 8 kit is then transported to the STS launch 

processing area and tested again to ensure operability. 

then mated to the OMV in the STS cargo bay, and continuity checks are 

conducted to validate a functional OMV/TSR kit interface. 

testing is done on the ground, where repair or replacement activities can be 

accomplished readily. 

The 

This is the "cost effective, minimum OMV impact" recovery kit. The 

The System B kit is 

All operational 

7.2.3.2 DRM 3 Mission Activities - As this Dm is detailed solely for the 

purpose of refining TSR requirements, the recovery mission is presumed to 

commence at the point of rendezvous with.the target vehicle, the disabled 

7-8 
LANDSAT satellite. 

acquisi tion range of the ground control lab1 e LANDSAT sate1 1 i te. The recovery 

mission activity sequenc'e for DRM 3 is presented in Table 7.2.3.2-1, and 

illustrated in Figure 7.2.3.2-1. The actual recovery mission is initiated at 

rendezvous with LANDSAT. The OMV and attached TSR kit are maneuvered into 

close proximity of LANDSAT by the OMV ground controller at the OMV Operations 

Support Center. OMV is then flown around the disabled satellite to determine 

its motion characteristics and select an approach and grapple strategy. This 

motion may be known in advance as a result of analysis of radar returns; 

however, this data cannot routinely be assumed to be available. 

The mated OMV/TSR is initially positioned within visible 

0 

The System 6 recovery elements will be deployed when the OkV/TSR kit assembly 

is aligned to track the target's RMS grapple fixture. 

possibly a several degree of freedom (DOF) manipulator arm, will be extended 

to provide clearance between the OMV and the target. 

access to the LANDSAT grapple fixture that is obstructed by the satellite's 

array of solar panels from a straight-in approach. 

LANDSAT'S small roll rate by CW, the TSR RMS end effector/grapple mechanism 

is prepared for grapple and the snare mechanism grips the fixture, retracts it 

into the end effector mechanism and rigidizes it for transport. The final 

recovery mechanism action is to use the multiple DOF extension arm to align 

the LANDSAT center of gravity with the OMV orbit transfer thrust vector to 

enable forward trans1 ation maneuvers required to return OMV/TSR/LANDSAT to the 

Orbiter. 

The extendible boom, 

It will also provide 

Once aligned with the 

0 

7.2.3.3 DRM 3 Post Mission Activities - When the recovery mission is 

completed and the LANDSAT satellite has been secured back in the Orbiter, the 

OMV and attached TSR kit will be positioned in the longeron and sill trunnion 

latches in the Orbiter and prepared for deorbit, if LANDSAT repairs cannot be 

completed onorbi t. 

Back on the ground, the TSR kit will be detached from the OMV and 

disassembled, refurbished, and stored for future missions. 

7-9 
Table 7.2.3.2-1 DRM 3 Mission Activities 

Event 

OMV/TSR Kit Del ivered to Visual 

- 

Contact with LANDSAT 

- Approximately 2000 Feet from LANDSAT 

Operations Support Center I OMV/TSR Control Transferred to OMV 

OMV/TSR Maneuvered by Ground 

Controller to Close Vicinity 

of Target - 50-100 ft. 

- Using OMV TV/Lighting System 

OMV/TSR Circumnavigate Target to 

Determi ne Spin Rate/Spi n Axes, 

Locate Grappl e Fixture, Observe 

Docking Obstruction( s 1 

Spin Axis, Positions OMV to Match 

Target Spin Rates 

Full Length, TSR Kit Support Camera 

Deployed and Activated 

TSR RMS Grapple Mechanism Prepared 

to Grapple. 

Ground Control 1 er Maneuvers TSR Kit 

to Engage LANDSAT Grapple Mechanism - 

Taking Care to Avoid Contact with 

Obstructing Solar Panel, & Maintain 

Position with Slowly Rotating LANDSAT. 

OMV and TSR Controllers Use Respective 

TV honitors to Conduct Mission 

I 

OMV Ground Control 1 er Determines Target 

TSR Extendible Eoom is Deployed to 

I 

Ground Controller Activates OMV RMS 

End Effector to Snare LANDSAT and 

Retract Fixture to Achieve Rigid Grapple 

to StabJe Position 

I OMV/TSR Maneuvered to Despi n LANDSAT 

Ground Controller Uses Mu1 tiple DOF 

Extendible Boom Arm to Move LANDSAT 

and Align Its Center of Mass with 

OMV to Limit C.G.. Offset and Permit 

Transfer of Mated OMV/TSR/LANDSAT 

Back to Orbiter 

7-1 0 
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7-1 1 
7.2.4 DRM 4 - System B, Case 2 - This is the second of the two recovery 

scenarios designated for recovery System B. 

a scenario in which a disabled satellite is: 

This DRM was selected to describe 

a. 

b. 

c. Not spin-stabilized or spinning; 

Disabled, all scientific instrumentation has failed; 

Controllable from ground and stable; 

d. 

e. 

Beyond Orbiter's nominal operating range; and 

Does not have standard grapple fixtures, i.e., RMS grapple fixtures 

I or FSS latch pins. 

I The specific mission Snvolves recovery of one of the Geopotential Research 

Mission (GRM) satellites that will fly in pairs in a circular, low earth orbit 

at 100 nautical miles with an orbital inclination of 90 degrees. The mission 

of this satellite is refined measurement of the earth's gravitational field. 

The failure mode for this satellite is postulated as a total instrument 

failure, though the spacecraft power and attitude control subsystems are not 

affected and the spacecraft remains fully control 1 ab1 e and stab1 e. 

7.2.4.1 

recovery candidate, i s i n a near-polar orbit, the mission will be conducted 

from the Orbiter. 

The only difference in the recovery system used for this DRM, as opposed to 

DRM 3, is the type of grapple mechanism used. As the target satellite does 

- not have any standard grapple mechanisms on it, a small, special purpose 

grapple mechanism must be used to attach to a "hard point" on the GRM 

satellite. The GRM satellite mission is actually composed of two satellites 

when deployed, one positioned about 80 miles behind the other. 

Will be carried into earth orbit in an STS "cradle" mechanism. Each of the 

GRM satellites has an extension on it to enable structural interface with the 

cradle, and it is this element that will be grappled by the small gripper 

device of the System B TSR kit. 

DRM 4 Pre-Mission Activities - As GRM, the disabled satellite 

The pre-mission tasks are described in Table 7.2.4.1-1. 

The satellites 

7-1 2 
Table 7.2.4.1-1 DRM 4 Pre-Mission Activities 

Event 

Assemble TSR Kit System Elements 

for GRM Mission (Extendible boom, 

Small Gri pper 1 

Conduct Assembly and Checkout Tests 

Using Ground Support Equipment at 

OMV Storage Area 

Transport TSR Kit to Launch 

Processi ng Area 

Conduct Final Ground Testing in 

Vertical or Horizontal Launch 

Processi ng F1 ow 

Attach TSR Kit to OMV in Cargo Bay 

Conduct Power and C&DM Continuity 

Checks Between OMV and TSR Kit 

through OMV Payload Unbilical 

OMV/TSR Kit Del ivered to Nominal 

Orbiter Operational Orbit 

OMV/TSR Kit Deployed From Cargo 

Bay with RMS 

OMV Transits to Rendezvous with 

GRM 

7.2.4.2 DRM 4 Mission Activities - Again, the DRM 4 mission is similar in 

most regards to DRM 3. 

illustrated in Figure 7.2.4.2-1 . The mission is described in Table 7.2.4.2-1 and 

The recovery mission commences when the OHV/TSR rendezvous with the GRM i n i t s 

low earth orbit. 

atmosphere and geopotential forces in the lower altitude and necessary 

recovery actions will be expedited to minimize the effects of these forces 

during retrieval operations. As GRM will be stable, the tasks of the OKV and 

TSR ground controllers will be simplified. The OMV ground controller can 

translate di rectly to the target sate1 1 i te and the TSR operator will extend 

the extendible boom, position the small gripper device on the GRM cradle 

The OMV will be more strongly affected by the earth's 
interface, and effect a rigid grapple on GRM. Finally, the multiple degree of 

freedom extension arm will be used to reposition the center of mass of GRM, to 

align it with the OMV thrust vector, to support transfer of the mated 

OMV/TSR/GRM package back to the Orbiter for repair or return to earth. 

I Minimum OMV ImpactILow-Cost Increase in Capability 

I Grapple Mechanism 

Geopotential Research 

Case 2: No Grapple Fixture 

Figure 7.2.4.2-1 DRM 4, System B, Case 1 

7.2.4.3 DRM 4 Post-Mission Activities - The GRM satellite will be repaired at 

the Orbiter if at all possible, and returned to earth if not. The System 8 

recovery kit will be retained on the OW while positioned in the. Orbiter for 

transport to earth. 

OW, either before or after OMV is removed, and the kit will be refurbished 

and stored for future use. 

When back on earth, the TSR kit will be removed from the 

, 7,2.5 DRM 5 System C, Case 1 - The final two DRMs, DRM 5 and DRM 6, are 

missions designed to demonstrate the most difficult of the remote, disabled 

satellite recovery scenarios. System C, as defined by MSFC, was intended to 

represent a "full up" recovery system, capable of dealing with the most 

"complex" type of satell ite motion and these two scenarios are representative 

cases requi ri ng this 1 eve1 of capabil i ty. \ 
Table 7.2.4.2-1 DRM 4 Mission Activities 

Event 

OMV/TSR Kit Delivered to Visual 

- 

Contact with GRM 

- Approximately 2000 Feet from GRM 

OMV/TSR Control Transferred to OMV 

Operations Support Center 

OMV/TSR Maneuvered by Ground 

Controller to Close Vicinity of Target - 50-100 ft. 

- Using OMV TV/Lighting 

Systems 

OMV/TSR Ci rcumnavi gates Target to 

Determi ne Spin Rate/Spi n Axes, 

(if any), Locate GRM Cradle Interface Device 

Ground Controller Determines Approach Strategy, Positions TSR C1 ose 

to GRM Cradle Interface Device 

TSR Extendible Boom is Deployed to 

Full Length, TSR Kit Support Camera 

Depl oyed and Acti vated 

Prepared for Grapple 

TSR Small Gri pper-Type Grappl e Mec hani sm 

Ground Control 1 er Maneuvers TSR Kit 

to Engage GRM Cradle Interface Device 

Ground Control 1 er Activates OMV TRS 

Gripper to Capture GRM, Engage and 

Achieve Rigid Grapple 

Ground Control 1 er Uses Mu1 ti pl e DOF 

Extendible Boom Arm to Reposition 

WI and Align Its Center of Mass 

with OMV to Limit C.G. Offset and 

Permit Transfer of Mated OMV/TSR/ 

GRM Back to Orbiter 

AS has been shown in preceding portions of this report, the more difficult 

recovery scenarios invol ve situations where the disabled satel 1 i te iS 

spinning, either in an attitude controllable mode or the complex case where 

the satellite has failed in a mode where attitude control has been lost due to 

any of a wide variety of potential causes. 

7-1 5 
Again, as previously shown, when a satellite has failed in a non-controllable 

mode wi th angul ar momentum 1 eve1 s exceedi ng the atti tude control capabi 11 ty Of 

the spacecraft, it will develop an initial multi-axis tumble or spin mode at 

some tumble rate. 

satellite will dissipate rapidly due to friction in joints, viscous fluid flow 

in propellant tanks, and flexible appendages on the satellite. When the 

system reaches a new minimum energy, steady state condition, the initial 

"tumble mode" coalesces into flat, single axis spin about the satellite's 

major principal axis of maximum moment of inertia. Thus, the complex motion 

for which a "full up" remote, disabled satellite recovery system should be 

designed is, in actuality, flat, single axis spin at some relatively steady 

rate. 

seven revolutions per minute (rpm) and spin rates for potential, 

non-controllable satellites were projected to be in the neighborhood of 

between 3 rpm and 10 rpm. 

disabled satellite is expected to range from 10 to 50 rpm and possibly higher, 

though the trend in spin stabilizing of spacecraft appears to be headed toward 

the lower ranges of spin rate. 

The high level of kinetic energy initially imparted to the 

1 

I 

The study team discovered cases where the spin rate was as high as 

The spin rate for a controllable, spin stabilized, 

I 

DRM 5 describes the recovery of the INTELSAT-6 satellite. This satellite was 

delivered into low earth orbit on an Orbiter flight and deployed for transfer 

into geostationary orbit. The perigee kick motor failed enroute to its higher 

orbit and the INTELSAT-6 satellite was left in a spin-stabilized, controllable 

mode in a totally useless orbit within the range of the OMV/Tumbling Satellite 

Recovery kit. 

I The recovery kit subsystems required to conduct this mission include the 

extendible boom, a spin table, a grapple mechanism interface device and a 

"stinger" type grapple mechanism to support grapple of the spinning perigee 

kick motor. 

integrated recovery kit and prepared for a mate with the OMV. The satell i te 

is in a low inclination orbit, as a result of its 28 degree Orbiter launch 

injection and this recovery mission can and will be conducted from a Space 

Station frame of reference. 

These subsystems will be assembled into a complete totally 

, 
The development and depl oyment concept for the recovery systems considered for 

use in recovering all remote, disabled satellites is to create a versatile Set 

of component subsystems that can be assembled and integrated into specific 

recovery kits, tailored for the unique recovery scenario pre'sented to the 

user. 

assumed to be spinning, and in DRM 5, the INTELSAT-6 spacecraft is spinning, 

as it is spin-stabilized. Thus, unlike the System B remote recovery scenarios 

previously analyzed, this recovery scenario requires a "spin table". This 

recovery subsystem, the spin table, will enable the total TSR kit to be spun 

up to match the spin rate of INTELSAT-6. With both the kit and the target 

spinning at similar rates, the grappling element of the recovery system can be 

maneuvered to a position wherein the grapple mechanism subsystem can be 

employed to achieve a firm grapple. The significant point to remember here is 

that the overall recovery system will be developed in such a manner that the 

spin table can easily be included as a part of the total system when required 

by the scenario, or deleted to reduce the total weight to be transferred to 

the remote, disabled satellite. This design philosophy will produce a 

resultant savings in operations costs, by minimizing propellant usage in going 

out to and coming back from the target. 

0 

In the "full-up" System C scenarios, the disabled satellites are 

a 

Similarly, the grapple mechanism required for this scenario will be a 

relatively special purpose mechanism, rather than a general purpose tool and 

will be attached to the grapple mechanism interface device, another subsystem 

element of the TSR. This grapple mechanism will be very similar to the 

mechanism used in the recent Westar and Palapa-B retrieval missions, in which 

an astronauthanned maneuvering unit (MMU) served as the basic recovery 

system. 

System because of the implied requirement and resultant design accomnodation 

inclusion of the grapple mechanism interface device into the basic recovery 

system. The concept of a modular design, allowing use of only the subsystem 

elements specifically needed for the given recovery mission will: 

the breadth and scope of potential remote recovery opportunities that can be 

conducted with this system and 2) enable efficient and cost effective 

deployment of the recovery system(s1. 

Again, this grapple mechanism can be incorporated easily into the 

1) increase 

7-1 7 
I' 

7.2.5.1 

are highlighted in Table 7.2.5.1-1. 

elements will be stored in the servicing storage area on the Space Station. 

The recovery system has been tested thoroughly in all possible assembly 

configurations. 

configuration for DRM 5 will be selected and assembled. 

will consist of an extendible boom, a spin/despin table, a grapple mechanism 

interface device and a sti nger-1 i ke grapple mechanism. 

components will be assembled in the servicing area, either manually using 

astronaut EVA or robotically using a Space Station "smart front-end" kit. 

DRM 5 Pre-Mission Activities - The pre-mission activities for DRM 5 

All of the remote recovery subsystem 

In preparation and planning for this mission, a TSR kit 

This recovery kit 

These modul ar 

Once assembled, the TSR kit will be mated to the OMV, again using either 

astronauts on EVA, or robotically. 

will be deployed from Space Station using the lriobile RMS, transit out to a 

safe launch distance from Space Station (using cold gas jets) and perform an 

orbit adjust to a rendezvous position close to the disabled INTELSAT-6 

sate1 1 i te. 

With assembly of the OMV/TSR kit, the OMV 

Table 7.2.5.1-1 DRM 5 Pre-Mission Activities 

Event 

Assemble TSR Kit System Elements for 

INTELSAT-6 Recovery Mission in OMV 

Servicing Area at Space Station - 

EVA or Robotic Assembly, (Kit includes Extendible BOOP, Spin Table, 

Grapple Mechani sm Interface Device, 

Stinger-type Grapple Mechanism). 

Mate TSR Kit to OMV in OMV Service 

Area (EVA or Robotic Mating) 

Deploy OMV/TSR Kit From Space 

Station with Mobile RMS (MRMS) 

Conduct Operational Checkout of OW/ 

TSR Kit in Proximity Operations Under 

Space Station Mission Control 

Transfer ObN/TSR Kit to Rendezvous 

with INTELSAT-6 

7-1 8 
7.2.5.2 DRM 5 Mission Activities - The satellite recovery mission commences 

when the OMV/TSR kit is maneuvered to a close proximity position with 

INTELSAT-6, within visual range of the target. The OW and TSR kit, will be 

controlled, at this point, by the OW Operations Support Center, with separate 

control’ consoles for the OMV and the TSR. 

in Figure 7.2.5.2-1 and outlined in Table 7.2.5.2-1. 

a 

The mission events are illustrated 

The OMV control 1 er will maneuver OMV/TSR to a1 ign the TSR transl ation/grappl e 

axis with the spin axis of INTELSAT-6. 

i.e., the extendible boom, spin table, and grapple mechanism, will be 

deployed, 

to that of the target. 

OMV transl ation maneuvers to position the stinger-type grapple mechanism for a 

firm grapple of the target. The OMV and TSR controllers will maneuver both 

systems to achieve a firm grapple of the target, while minimizing torque 

producing forces caused by incidental contact with the target prior to the 

completion of a firm grapple. 

The TSR capture support mechanisms, 

The spin table will be activated to generate a TSR spin rate equal 

These activities will prepare the recovery system for 

Fullup Recovery System 

Spin-12 rpm 

Spin Table-12 rpm 

\-Perigee Kick Motor 

ExtendibleJ l- Stinger Grapple Mechanism 

Probe 

Case 1: Spin Stabilized Target 

Figure 7.2.5.2-1 DRM 5, System C, Case 1 

7-1 9 
Table 7.2.5.2-1 DRM 5 Mission Event Sequence 

Event 

OMV/TSR Maneuvered by OMV Ground 

- 

Control to Close Visual Contact 

With Target 

OMV Operator Maneuvers OMV/TSR 

to Align OMV/TSR Axis With Spin 

Axis of Spin-Stabilized Target 

TSR Extendible Boom is Fully Extended by Ground Control 1 er 

and Sti nger-type Mechani sm is 

Deployed. OMV Maintains Stable 

Atti tude duri ng TSR Element 

Deployment 

TSR Ground Controller Spins Up TSR 

to Match INTELSAT-6 Spin Rate 

OMV and TSR Controllers Maneuver OMV 

and Grapple Mechanism to Proper Grapple Position on Target Spin 

Axis and Center of Mass 

OMV and TSR Controllers Maneuver 

OMV and TSR Grapple Mechanism 

to Achieve Firm Grapple of 

INTELSAT While Minimizing 

Contact Dynamics Effects 

TSR Controller Despins INTELSAT-6 to 

Perform Satell ite Stabil ization 

OMV/TSR Operators Take Actions to 

Align INTELSAT-6 For Transit to 

Space Station 

OMV Small Thrusters Ignited to Test 

Grappl e Rigidity and Correct Target Mass Alignment for Transit. 

OMV Orbit Adjust Engines Fired to 

Return OMV/TSR/INTELSAT to Space 

Station 

Following grapple actions, the TSR controller will despin the satellite and 

OMV and TSR controllers will take necessary follow-on actions to align the 

mass of INTELSAT-6 .with the OMV thrust vector for transfer back to Space 

Station. 

These activities describe the basic mission activities for DRM 5. 

7 -20 
7.2.5.3 DRM 5 Post Mission Activities - When OMV is transferred back to the 

Space Station and delivered back to the servicing area by the Mobile RMS, the 

TSR kit will be dentated from OW, either by robotic operations or EVA. The 

subsystem el ements wi 11 be di sassembl ed and i ndi vidual ly refurbished and 

tested. They will each be positioned in individual storage locations at Space 

Station, ready for assembly into the next required TSR kit configuration. 

7.2.6 DRM 6 - System C, Case 2 - This DRM was defined and developed to 

describe the most chall engi ng remote, disabled satel 1 i te recovery mission 

envisioned by the study team. 

attitude control of the system is lost, and it accumulates significant angular 

momentum, presents .potentially the most difficul t recovery scenario. 

A satellite that fails in a mode in which 

An orbiting satellite can be subjected to external torque of some form, such 

as reaction control system (RCS) engine failure, or propellant or pressurant 

1 eaks that produce angul ar momentum 1 eve1 s exceedi ng the satel 1 i te 's atti tude 

control system (ACS) capability. 

When this occurs, the satellite will assume some initial condition of general 

tumble motion, more than likely a multi-axis tumble or spin configuration. As 

shown previously in the condition where additional torque is not added, this 

motion will very quickly, within minutes or hours, stabilize and become single 

axis spin about the satellite's major principal axis; i.e., its principal axis 

of maximum moment of inertia. Thus, it is realistic to assume that a typical 

complex motion satel 1 f te recovery scenario might involve a 10-1 5,000 pound 

satellite, 15 feet in diameter, spinning at 7-10 revolutions per minute about 

an axis perpendicular to a 20-30 foot long satellite! This is a somewhat 

accurate description of the DRM 6 scenario chosen by the study team for 

definition, and, ultimately, for refinement of recovery system requirements. 

DRM 6 involves recovery of the Upper Atmospheric Research Satellite (UARS), a 

10,000 pound satellite, estimated at this time for launch in 1990 into a 324 

nautical mile circular orbit at an inclination of 57 degrees. The satellite 

failure mode is a reaction control motor that fails open, depleting its 

hydrazine fuel supply. The resultant torque generated by the RCS engine 

7-21 
throttle-open failure produces sufficient angular momentum to exceed LIARS 

attitude control authority and produce a multi-axis tumble mode of around 

7 rpm. 

operation will be conducted using an OMV launched in the Orbiter, because of 

the 57 degree inclination of the recovery target. 

This scenario is illustrated on Figure 7.2.6-1. The recovery 

The DRM 6 recovery mission is the most technically challenging of the DRMs and 

the TSR kit for this mission contains all of the subsystems being considered 

for the total system, including the most technically challenging of the family 

of grapple mechanisms being considered for the total recovery system, This 

kit will include the extendible boom, the spin/despin table, the grapple 

rnec hani sm interface device and a 1 arge envel opment-type grapple rnec hani sm, 

similar to that shown in Figure 7.2.6-1. This grapple mechanism must be 

capable of being deployed and spun up at speeds ranging from 5-100 rpm. It 

must have grapple arms that can be extended to grapple, circle or envelop a 

satellite that is 15 feet ih diameter. The structural components of the total 

system must be strong enough to accomnodate the forces and torques generated 

by 1) the contact dynamics of the grapple action, i.e., interactions between 

the grappler of the recovery system and the target and, 2) similar forces 

(though smaller) introduced into the recovery system by the despin motor while 

despinning the grappled satellite. This recovery kit will be assembled on the 

ground during the pre-mission, or planning, phase. 

7.2.6.1 

missions, the planning and preparation activities will be conducted on the 

ground as shown in Table 7.2.6.1-1. 

DRM 6 Pre-Mission Activities - As in most Orbiter oriented recovery 

Those specific recovery system components to be used for DRM 6 will be 

assembled in an OMV/TSR kit assembly area and thoroughly tested. 

delivery to the launch area, the TSR kit will be tested in the Payload 

Vertical (or Horizontal ) Processing area using special TSR ground support 

equipment (GSE). 

onorbit deployment from the cargo bay and prior to the OMV flight out to a 

UARS rendezvous. 

replace components at that time (though not impossible), so the launch 

processing tests will be thorough. 

Following 

It will be relatively inefficient to test it following 

It is assumed that it would be difficult to repair or 

7-22 

a 
Fullup Recovery Systems 

- Event 

Grapple Mechanism 

/(Space Bola) 

/ n 

Case 2: Uncontrolled Spin 

Figure 7.2.6-1 DRM 6, System C, Case 2 

Table 7.2.6.1-1 DRM 6 Pre-Mfssfon Activities 

Ground 

Assembly and checkout of DRM 6 TSR 

Kf t at OMV/TSR Storage/Assembly 

Facil i ty - (Ki t i ncl udes Extend4 bl e 

Boom, .Spin Table, Large Enveloper 

Grapple Mechanism) 

Transport of TSR Kit to Launch 

Processing Payl oad Assembly Area 

Checkout in Payl oad Processi ng Area, 

Mate to OMV, and Installation of 

OMV/TSR in Vertfcal or Horizontal 

Processing F1 ow 

Onorbi t 

Following Launch to Orbiter Operational Orbit, the Mated OMV/TSR 

Kit fs Deployed from Cargo Bay 

Using Orbiter RMS 

Orbf ter is Maneuvered to Provide 

Separatf on from OMV Prior to OMV 

Ignition 

OW is Launched to Rendezvous with 

the UARS Satellite 

7-23 
The OMV/TSR kit will then be transported into a properly phased 57 degree 

orbit at nominal Orbiter altitude by the STS. 

physically deployed from the Orbiter, with the use of the RMS, and the Orbiter 

will maneuver out to an appropriate separation distance from the OMV. The OMV 

will receive launch and orbit transfer signals from the OMV Ground Operations 

Support Center to deliver OMV/TSR to a rendezvous with the disabled UARS 

sate1 1 ite. 

The mated OMV/TSR.will be 

7.2.6.2 DRM 6 Mission Activities - The recovery mission for DRM 6 commences 

when the OMV and TSR operators obtain visual contact with the Upper 

Atmospheric Research Satellite (UARS). The specific mission activities for 

DRM 6 are outlined in Table 7.2.6.2-1. 

The OMV/TSR operators will obtain the first view of a 25 foot long satellite 

spinning end-over-end in, a fl at spin about an .axis perpendicul ar to the 

longitudinal axis of the spacecraft. 

The OMV and TSR cameras will be used to estimate the spin rate and determine 

the spin axis, as the OMV operator circumnavigates the spinning UARS. 

operator will use care to assure adequate clearance between OMV and the 

numerous protuberances on the UARS. 

to locate the spin axis (or axes if motion is multi-axis spin; Le., general 

motion, including precession and nutation - which should occur only if the 

satellite is experiencing some form of external torque) on the spinning UARS. 

This axis should be the colinear major principal axis, angular momentum vector 

and angular velocity vector. The OMV and TSR operators will identify this 

axis through visual techniques including determining the spot where the target 

appears as a single point and everything around that point prescribes apparent 

perfect circles, rather than elliptic contours. 

The OMV 

The primary objective of this activity is 

The study team examined a number of alternatives for providing support to the 

OMV and TSR operators in finding the central spin axis, including radar and 

laser ranging, but found no feasible solutions. 

corner reflectors properly positioned on the target were required to enable 

electronic or visual ranging for motion orientation, and proper preposi tioni ng 

of the ref1 ectors appears infeasible. 

In most cases, a set of 

7-24 
Table 7.2.6.2-1 DIU4 6 Mission Activities 

- Event 

a OMV/TSR Rendezvous to Within 

Visual Range of UARS (Under 

OMV Operations Support Center 

(0%) Control 1 

OMV Ci rcumnavi gates Target to 

Determine Sate1 1 i te Motion: 

Spin Axes; Rates 

OMV OSC Determines Target Spin Axes and Develops Approach Strategy 

TSR Ground Control 1 er (GC) Deploys TSR Kit, Incl udi ng Boom Extension, Envel opment-Type Grapple Mechanism Deployment 

Mechanism to Match Spin Rate(s) of UARS. OMV Maintains Attitude 

Control During TSR Deployment 

OMV GC, with TSR GC Support, Conducts 

Trans1 ation Maneuvers to Position 

Spinning Grapple Mechanism at or Near 

Target Center of Mass 

Mechanism to Surround Target Wi thout 

Touching It 

TSR GC Conducts Grapple Closure Operations 

to Achieve Firm Grapple and Rigidization 

with Target 

TSR GC Operates TSR Spin Table 

TSR GC Mani pul ates Envel oper Grapple 

TSR GC Performs Despin Operations to 

Stabi 1 i ze Target 

TSR Operator Releases UARS, Regrapples Target as Required to Provide Proper Mass Distribution for OMV Transfer 

OMV GC Operates Small Attitude Control 

Thrusters to Test for Rigidity and 

Correct Orientation of Center of Mass 

OMV GC Fires Orbit Adjust Thrusters to 

Return OMV/TSR/UARS to Orb1 ter 

7-25 
Once the spin axis is identified, an approach strategy is selected and the 

OMV/TSR kit is aligned with the target spin axis. The TSR ground controller 

will deploy the recovery system elements. The extendible boom will be 

actuated first to provide target/OMV clearance. Next, the large, 

envelopment-type grapple mechanism will be unfolded and extended to a open 

configuration adequate to envelop the target. 

"spun up" by the TSR ground controller to match the spin rate and phase angle 

of the target. The TSR ground controller should be able to refine spin rate 

matching errors with occasional small adjustments during the recovery 

operation. 

Finally, the TSR system will be 

Following deployment of the recovery subsystems, the OMV operator, in 

coordination with the TSR operator, will translate the mated systems into a 

position wherein grapple can be effected. The extensive rotating mass of the 

TSR system may introduce complications preventing a nominal OMV transition to 

the target grapple, but simulations will be required to validate this 

assumption and to develop an appropriate safe and efficient approach technique. 

When the OMV/TSR systems are positioned such that the spinning, enveloper 

grapple mechanism has spin rate matched with the target and is surrounding the 

target at or near the target's center of mass, the grapple mechanism closure 

can be initiated. 

precision in this phase of conceptual study. The principal uncertainty, of 

course, relates to the potential for unpredictable contact with the target and 

resultant contact dynamics, which can possibly affect the target, the TSR kit 

and the OMV itself. It will be difficult to develop a highly reliable contact 

dynamics model for recovery operations, certainly until a specific recovery 

system is selected. It is highly likely that the "contact dynamics" issue can 

be addressed only in space, and will mandate a cargo bay (or Orbiter proximity 

operations) experiment eventually. 

"envelopment" of the target represents i ts capture and that once envel oped the 

grapple mechanism can slowly accomplish a firm grapple and rigidization with 

the target, minimizing damage to the target, the TSR mechanisms and the OMV. 

The grapple mechanism concept selected and designed for this mission is 

intended to minimize these potential disturbing factors. 

This is a phase of operations difficult to define with 

The study team used the assumption that 

7-26 
One of the more apparent sources of inappropriate contact is the fact that all 

of the grapple mechanism arms or fingers must close at the same rate to 

*minimize dynamic stability problems for the TSR itself. Thus, as the 

mechanism is closing and one arm makes contact with one target surface first, 

without all arms contacting the target at the same time, a force is imposed on 

the target. This could disturb its motion, causing target dynamic 

instability, that could result in unexpected motion creating damage to 

recovery elements. Another area of uncertainty relates to the natural 

limitations in the control of OMV. There will be positioning "deadzones" in 

controlling OMV pitch and yaw motions, and also in OMV translation, that will 

cause inherent inaccuracies in correctly posi tfoning a grapple mechanism, 

extended from between 10-15 feet from the OMV, with respect to the UARS 

target. These positioning uncertainties, coupled with 2-3 second time del ays 

in controlling both OW and TSR mechanisms will create possibilities for 

undesired levels of contact with the target prior to achieving a firm grapple. 

The TSR operator will achieve a firm grapple with all fingers and/or arms of 

the grapple mechanism and apply additional gripping pressure to rigidize the 

contact between TSR and UARS. a 

The TSR operator will then activate the TSR spin motor to slowly despin the 

firmly grappled UARS recovery target. 

attitude control system wlll automatically retain attitude control of the 

mated OMV/TSR/UARS package during despin. This same approach was used 

successfully using EVA in the Manned Maneuvering Unit (MMU) retrieval 

operations involving both Westar and Palapa-B in 1985. 

This despin will be conducted at very 

- low rates to minimize the torques on the TSR subsystems and on OMV. The OMV 

When the angular momentum of the UARS has been eliminated by TSR despin, the 

TSR ground controller will release UARS and regrip the satellite appropriately 

to properly distribute the target's mass distribution for OMV transit back to 

the Orbiter. 

Finally, the OMV operator will fire the small thrusters to test the rigidity 

of the OMV/TSR/UARS attachments and then fire the OMV main thrusters to return 

the recovery elements to the Orbiter. 

These activities describe the basic mission activities for DRM 6. 

7-27 
7.2.6.3 DRM 6 Post Mission Activities - Following rendezvous of the returned 

OMV/TSR/UARS at the Orbiter, the UARS will be grappled using the Orbiter RMS 

and positioned in a flight support system (FSS) cradle for repair or in 

appropriate trunnion fittings for return to Earth. 

orbit, the TSR kit will be removed from the OMV by astronaut EVA and 

temporarily stored in the cargo bay. 

by remote RMS operations or by astronaut EVA, and UARS returned to its 

operational orbit by OMV. 

remated to OMV and returned to earth for refurbishment and storage. 

If the UARS is repaired on 

The UARS and OMV will be mated, either 

When OMV returns to the Orbiter the TSR kit will be 

7.3 Requirements Allocation 

The overall objective of Task 1, Requirements Analyses and Trades, was to 

conduct a series of systems requirements analyses to identify the requi rements 

related to development of a system capable of remote recovery of disabled 

satellites. 

illustrate the recovery scenarios and capability inherent in each of the 

recovery systems A, B, and C; (2) further decompose the level of systemoriented and concept-oriented requirements; and (3) support the selection of 

candidate hardware concepts. These six DRMs described operations re1 ated to 

the six recovery scenarios considered most likely to occur in the mid-1990s. 

The recovery capability required for each of the DRMs was described in 

general. The operational analysis did generate addi tional system level 

requi rements, complementing the other requi rements analyses in Task 1. 

in Figures 7.3-1 and -2 is the study team's approach to using the DRMs to 

support selection of hardware concepts. 

from the DRM definition and analyses and from other Task 1 analyses, were 

grouped into categories that related to speci fic recovery subsystem 

accommodations. 

strongly implied the need for an extendible boom. 

requi rement to recover sate1 15 tes with obstructed recovery support elements 

(DRM 3) suggested an extension device capable of articulation to enable an OMV 

to close and attach an RMS end-effector to a grapple fixture. The requirement 

to minimize risk to the OMV, especially in the presence of jutting appendages 

on a spinning target, also appeared allocatable to an extendible boom. 

The six recovery kit DRMs were developed and analyzed to: (1) i 

Shown 

The top level requirements, generated 

For example, a number of requirements identified in the DRMs 

One of these, the 

7 -28 
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Also shown on Figure 7.3-1 are the spin-related requirements that would appear 

to be accommodated by the addition of a spin table or mechanism that will 

enable the recovery device to match target spin rates with the accuracy 

required to accomplish envelopment and firm grapple, while minimizing damage 

to the satellite. Similarly, the spin mechanism must be capable of reverse, 

or braking operation, to despin recovery targets safely. 

be possible in either direction so the grapple can be effected on the most 

optimum si de of the sate1 1 i te. 

Spin and despin must 

Figure 7.3-2 illustrates the remainder of this top level allocation of 

recovery requirements to potential subsystem components of a TSR kit. The DRM 

analysis demonstrated the need for a variety of grapple mechanisms. 

grapple mechanisms could be readily accommodated on the recovery system with a 

grapple- mechanism interface device, configured with electrical and 

communications and data management (C&DM) interface hardware. 

Figure 7.3-2 is the broad range of grapple mechanisms needed to accomplish the 

DRMs. 

match spin rates, and to secure a firm grapple with minimum impact on the 

target appears to dictate a boresight TV camera, perhaps on the spin table, 

possibly a TV camera near grapple contact points and perhaps contact sensors 

on grapple mechanism contact points. 

These 

A1 so shown on 

Finally, the requirement to observe target motion characteristics, to 

Each of the DRMs was evaluated to determine how many of' the candidate 

"recovery system accomodations" would apply to each scenario. The result of 

this sumnary correlation effort is presented on Figure 7.3-3. For DRMs 1 and 

2, there are no recovery subsystem accomnodations, as these DRMs were designed 

for System A, the baseline OMV. Several of the candidate TSR subsystem 

components were required, in some form, in all four of the remaining DRMs. 

These were an extendible boom, a grapple mechanism interface device, an OMV 

structural interface device, and an OMV umbilical connect for power and C&DM. 

In addition, the wide variety of grapple mechanisms outlined in this summary 

appear to (1) support the need for an efficient method of configuring the 

overall recovery system for changeover of grapple mechanisms, and (2) 

reinforces the contractual requirement to evaluate no fewer than three grapple 

mechanisms in the Concept Definition phase, Task 2. 

7-31 
7 -32 
8.0 CONCEPT DEFINITION - TASK 2 

8.1 Concept Definition Approach 

The objective of the second major study task, Task 2, Concept Definition, was 

to conduct design trades that would enable development of conceptual system 

designs for a set of satellite recovery systems. 

produce a set of conceptual designs for MSFC Systems 8 and C that would wve 

as a focal point for continuing design and development efforts aimed at 

creating a front-end kit for tne OMV to remotely recovery disabled satellites. 

The primary goal was to 

The initial approach for this task was to use the results of the concept 

survey to select the best of the concepts and any other new, original concepts 

and refine these concepts into meaningful, effective conceptual system 

designs. As previously noted, during the evaluation of those prior concepts, 

it became clear that none of them could satisfy the widely variant 

requirements that wee evolving out of parallel analysis tasks. The study 

team developed an alternative approach to tne conduct of Task 2 and it is 

illustrated in Figure 8.1-1. 

0 The primary inputs to the development of this approach are shown on t& left 

side of Figure 8.1-1. 

suggested that the satellites could fail in either a stable and controllable 

mode, or be non-stable and tmbling/spinning about a single principal axis, in 

the non-torqued case. This fact introduced several implications for the 

conceptual design phase. 

The analysis of the actual recovery environment 

First, and most importantly, it validated the requi rement for a family of 

varying capability recovery systems. This, of course, led to the recovery 

scenari o analysis that produced the enhanced defi ni ti on of the capabi 1 i ties 

required for MSFC's Systems 8 and C. Second, it highlighted some of the 

recovery system requi rements and provided a foundation upon which to conduct 

analyses that led to identification of a number of the key requirements. For 

example, a spinning target with whirling appendages, such as solar arrays and 

antennas, signified the need to protect the OMV and the recovery system from 

damage during recovery qerations. This insight, in turn, suggested the need 

for an extension device to provide clearance during recovery operations. This 

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8-2 
is a requirement that was not included in any of the previous recovery system 

studies. As another example, the realization that the targets could very 

likely be spinning, some at high rates, led to the establishment of a spin a table as a potenti a1 vstem accomnodati on. 

Finally, the satellite failure environment assessment supported clarification 

of the scope of conceivable satellib recovery activities, and enabled 

definition and approval of the set of six DRMs for the OMV TSR system. 

operati ons analyses of these DRMs out1 i ned a wi de range of different grapple 

mechanisms requi red for potential recovery operations. To enable ready 

applicaticm of these grapple mechanisms to a "recovery system", the 

requirement for a grapple mechanism interface device also became apparent. 

Thus, a full-up recovery system was seen to contain the major subsystem 

components graphical ly represented i n Figure 8.1 -2. 

The 

In addition, the operations analyses of the DRMs led to further expected 

decomposi ti an of requi rements, i ncl udi ng the need for 1 ight weight subsystems , 

and compact systems for cost efficient STS operations while mated with OMV 

during transport into nominal satellite orbits. 

ModulU 

- Tailored to Spuific 

MiuioM 

OMV 7 

Figure 8.1-2 Full-Up Recovery System 

8- 3 
The general assessment of the Task 1 requirements analyses was that the 

overall architecture of the recovery system should be a modular design, 

including several major subsystem components. These elements could be readily 

interchanged to allow tailoring of the "overall system" for the unique 

requirements of each specific recovery mission, The wide range of potential 

recovery missions was a major factor in the slection of tnis architecture. 

In addition, as shown in Figure 8.1-3, the study team believed that, using 

this des gn architecture, the full system could be developed incrementally and 

cost eff ciently, starting with a System B capability and growing to a full-up 

System C 

recovery system to conduct missions requiring only an extendible boom and a 

small gripper will save in fuel usage and operating costs. 

In addition, the capability to use only a portion of a full-up 

Thus, referring back to Figure 8.1-1, the study barn, with the approval of the 

MSFC Contract Technical Manager, dedicatsd Task 2.1 to an extended evaluation 

of candidate recovery subsystem elements. These subsystems include an 

extendible boom, a spin/despin table, a grapple mechanism interface device, 

and a set of grapple mechanisms for different recovery scenarios. 

0 For Task 2.2, Conceptual Design, the preferred subsystem components ere 

selected and a set of recovery systems were designed for Systems B and C, each 

with the capabi li ty to conduct both of the related design reference missions. 

In addition, the MMC study team developed a new design for an envelopment type 

grapple mechanism, which we termed the "MMC Enveloper." 

8-4 
System A System B System C 

- Extends Recovery - Expands Capture - Full-Up Capability 

Range Geometry of OMV - Recovers Controllable 

- Limited Capability - Increased Grappling Spinners 

for Controllable Capability for - Noncontrollable . Targets Controllable Satellites TumblelSpinners 

Capability Increases through Modular, Incremental Growth 

Development lmpllcatlonr 

Can Be Developed & Tested Incrementally 

System Elements Are Modular 

- Elements Added as Needed 

- Common Interfaces Support Element Addition 8 Deletion 

Cost Efficient 

- System Elements Common 

- Use Only Those Needed-Weight to Orbit 

Figure 8.1-3 Family of Recovery Systems 

8- 5 
9.0 EVALUATION OF RECCVERY SYSTEM MECHANISMS - TASK 2.1 

9.1 Subsystem Evaluation Approach 

A survey of candidates for each of the major subsystems listed in Figure 9.1-1 

was conducted. Industrial brochures, 1 ibrary inquiries, and discussions with 

mechanisms experts at MSFC and MMC produced the initial candidates. Though 

most were viewed quickly as falling short of previously perceived design 

requirements, all were evaluated qualitatively. By looking at the apparent 

advantages and disadvantages of each candidate, more eval uation considerations 

were derived. The study team did not consider it productive to develop new 

mechanism concepts, with the major exception of the envelopment type grappler, 

as many generic mechanisms for each component were available. The evaluations 

are presented in a format which describes the candidate and highlights the 

advantages and disadvantages of each. 

- 

velopment Grappler 

Figure 9.1-1 Recovery Subsystems Eva1 uation 

9 -1 
9.2 Eva1 uation of Extend1 bl e Booms 

9.2.1 Introduction - Shown in Figure 9.2-1 are examples of some of the 

extendible boom concepts identified and evaluated during this phase of the 

study. The study team considered a wide range of alternatives to avoid 

unsupportable elimination of questionable candidates. A total of eight 

extendible boom concepts were identified and evaluated for applications in the 

TSR Kit. Some of the driving design requirements in this evaluation 

included: the need to be compact when not deployed (to minimize Shuttle cargo 

bay delivery spacekost), articulation capability to provide access to 

obstructed recovery support elements (RSE) and to align target center of mass 

with the CMV orbit transfer thrust vector prior to orbit transfer, relative 

weight, and capacity to accommodate grappleldespi n force and torque 1 oads. 

Manipulator Arm 

Scissors Mechanism 

Tripan &Bar 

Linkage 

Figure 9.2-1 Extendible Booms 

Fixed Shaft 

Closed Tubular 

Extension Boom 

9.2.2 Evaluations - The first of the extendible boom concepts selectea for 

evaluation is the scissors mechanism depicted in Figure 9.2.2-1. 

structure, a collection of links, each rotatively pinned at the center and 

ends, will most likely be actuated by a motor driven screw at the base of the 

mechanism. The ends of the two links attachea at the base will be forced 

The scissors 

9-2 
ADVANTAGE 

Llght Wolght 

Non-Complox Control 

* No Arclculrtlon for Aceoat to Obatructod 

Not Compact 

R.comy Support Elomontt 

DISADVANTAGE 

Llmltod Capaclty for Grapplo, Doapln 

No Artlculalon !or T8rg.t Roallgnmont 

Torqw Lo8ds 

Figure 9.2.2-1 Scissors Evaluation 

together by travel on the screw. Each of the links will thereby be caused to 

move from a horizontal position, where the mechanism is in a stowed configuration, to a vertical position, or the fully deployed mechanism configuration. 4 

Of the disadvantages listed in the figure, the most significant is the 

inability of the mechanism to support grapple and despin torque loads, 

specifically those in bending and torsion. 

compression could be accommodated by a nonbackdrivable actuation system. 

increase in the strength of the links in order to increase the load bearing 

capability, would only add to an already relatively large stowed volume; as 

the width of each of the links contributes to the perpendicular distance from 

the mechanism base, the significant stowed volume dimension. 

Load support in tension and 

Any 

Although articulation of the extendible boom, listed among the scissors 

mechanism disadvantages, is a mission specific requirment, it is recomended 

that the TSHS extendible boom selection be capable of satisfying all mission 

requi rements. Chosen for evaluation because of its simp1 ici ty , the scissors 

mechanism provides a light weight, variable method of satisfying the grapple 

mechanism extension requirement, but its disadvantages make it unwarranted as 

an extendible boom choice for the TSRS. 

9 -3 
The tubular extension mechanism, illustrated in Figure 9.2.2-2, provides an 

extremely compact and 1 i ght wei ght method of extension. The mechanism empl oys 

two long, tape-like plates, joined at the edges by a connection which allows a 

limited amount of sliding to occur between the plates. 

so that when constricted, their surfaces are flush, but when freed, the plates 

will curve, and together form an oval. 

operation, the plates may be constricted then spooled. When freed they extend 

from the base. 

system within the base. 

antenna, and it has been used in this capacity for space application. 

The plates are bowed, 

Similar to a tape measure in 

Actuation can be accomplished by a nonbackdriveable motor 

Existing designs employ the device as an extendible 

Load bearing capability is not inherent in the design of the tubular extension 

mechanism and its capacity for grapple and despin torque loads is very 

limited. 

capabil i ty to provide target realignment and access to obstructed RSEs. 

in the conceptual design process the mechanism was considered as a method of 

extension,' however, the lack of articulation and load bearing capability make 

it an unsuitable choice for the TSR kit extendible boom. 

Also listed below as a disadvantage is the mechanism's lack of 

Early 

AWMAGE 

Compact 

Llght Welght 

Existing Peslgn 

Non-Complex Control 

DISADVANTAGE 

* Limited Capacity tor Grapple. Oarpin 

No Articulation for Target Realignment 

No Artlculatlon for Access to Obstructed 

Torque Loads 

Recovery Support Elements 

Figure 9.2.2-2 Tubular Extension Evaluation 

9-4 
Depicted in Figure 9.2.2-3, the Tripan mechanism structure, like that of the 

scissors mechanism, is a collection of links rotatively hinged at the center 

and ends. Where the scissors mechanism structure emanates from two 1 nks and 

two base attachment points, the Tripan structure extends from six links and 

three base attachment points which collectively form a triangle. Each of the 

two links extending from a base attachment point is pinned at the center and 

at the far end to another of the links extending from a separate base 

attachment point. 

additional 1 inks then begins the Tripan Is triangular structure progression. 

Again, like the scissors mechanism, the Tripan could incorporate motor drive 

screws as the methoa of actuation, where the base attachment points are forced 

to the center of the triangular base. 

structure is used today in varied industrial applications. 

The pinning of the other ends of these base links to 

Not recently designed, the Tripan 

The Tripan mechanism differs significantly from the scissors mechanism in that 

its structure provides the added load bearing capability in both bending and 

torsion; yet while heavier than the scissors mechanism, it still offers a 

light weight method of extension. 

provide access to obstructed grapple fixtures, decrease in significance, the 

Tripan could be considered as an extendible boom selection. 

Should the articulation requirement, to 

' Llght Wolght 

Can Support Grrpplo, Dospln loquo Load8 

Non-Complor Control 

Not Compact 

Hlghw Rdlrblllty. U~lntrlMblllty Domrndr 

No Artlculrtlon for Trrgot Rrllgnmont 

No Artlculrtlon for Accorr to Obrtructod 

Recovery Support Elemontr 

Figure 9.2.2-3 Tripan Evaluation 

9-5 
The 4-bar linkage illustrated in Figure 9.2.2-4 was previously developed by 

NIX for space appl ication. The 1 inkage was designed so that when extended, 

the end to be attached to a grapple mechanism will travel in a linear path 

perpendicular to the base plate. 

arms is hinged at the elbow plate; the remaining two links have attached gears 

which mesh at the elbow. A rotation of the geared link secured at the base iS 

effected by a nonbackdriveable torque motor. This rotational actuation and a 

two to one gear ratio at the elbow are responsible for the linear extension. 

The linkage is very compact in terms of the critical volume dimension; 

although the length of the links, constrained by the radius of tne Orbiter 

cargo bay, limits the extension to roughly 15 feet. 

One of the two links of each of the linkage 

ADVANTAGE 

' Compact 

Like the previous concepts, the 4-bar offers a non-complex accommodation to 

the recovery system extension requirement. As a light weight mechanism, its 

load carrying capability is limited in bending and torsion. The tradeoff of 

increasea weight for increased load carrying capability will not, however, 

create a significant increase in compacted volume. 

the Tripan, could be considered for the TSRS in the absence of the 

articulation requirement for access to obstructed grapple fixtures and 

alignment of target center of mass prior to orbit transfer. 

The 4-bar linkage, like 

DISADVANTAGE 

Insufflclent Artlcuktlon for Target Realignment 

Insufflclent Arnculatlon for Access to 

Obstructed Recovery support Elements I Llght Welght 

Can Support Grapple, Oespln Torque Lo&# 

Non-Complex Control 

Figure 9.2.2-4 4-Bar Linkage Evaluation 

9 -6 
One possible design for the telescoping extendi bl e boom concept, previously 

developed by MMC and shown in Figure 9.2.2-5, involves a series of tubes, 

threaded to allow extension of one with respect to another. 

mechanism is accomplished by reversible torque motors surrounding each tubular 

section. 

motor, will extend that section and those sections of smaller diameter. 

tel escopi ng mechanism wi 11 provide a vari ab1 e method of extension, capabl e of 

supporting the grapple and despin torque loads of a recovery mission. 

a Actuation of the 

Rotation of any of the sections, effected by its corresponding 

The 

Although the telescoping mechanism is incapable of articulation, the other 

disadvantages identified in the figure, make it unwarranted as a TSRS boom 

selection for even limited recovery missions. 

which the tubular sections are made must be capable of maintaining a threaded 

interface, precludes the boom from being a light weight structure. 

Additionally, the structure is difficult to design to be compact in terms Of 

the volume dimension perpendicular to the base. The grounded or initial 

tubular section, into which the other sections are compacted is proportional 

to the overall extension of the mechanism. A decrease in the length of this 

section, in an effort to decrease the compacted volume, will produce a 

decrease in the overall extension. 

requirements higher than the previous concepts are also introduced, as a 

result of torque motors associated with each tubular section. 

The fact that the material of 

Re1 iabil i ty and maintainabil i ty 

W 

I I AWANTAGE MSADVANTAGE I 

Cn Support Grapple, Despln Torque Loads * He8q 

' Not Comp8cffDlfflcult to Deslgn to be Compact 

No Artlcuktlon for Target Reallgnment 

No Artlculatlon for Access to Obrtructed 

Recovery Support Elements 

' Hlgher ReIIablllty/YalntalnablIlty Requlrementr 

. Figure 9.2..2-5 Telescoping Boom Evaluation 
The simplest of the extendible boom concepts is a fixed shaft. 

relatively heavy, there are no constraints regarding its composition, and a 

number of materials could be used to reduce its weight. 

greatest capability to support grapple and despin torque loads, it also 

includes the most significant disadvantage: inflexibility in requiring 

valuable Orbiter cargo bay volume. A fixed shaft, though obviously not 

recommended as a TSRS selection, can be used for comparative evaluation, and 

was included for completeness. A summary of its advantages and disadvantages 

is provided in Figure 9.2.2-6. 

Although 

But by providing the 

ADVANTAGE 

C8n Support Gmpplo, Dorpln torquo Lo8dS 

Non-Complox Control 

UWVANTAGE 

' HO8Vy 

* Not Compact 

No ArtlCul8tlOn lor Trrgot Rorlignmont 

No Arllculallon lor Accorr to Obrtructrd 

Rocovrry Support Elomnts 

Figure 9.2.2-6 Fixed Shaft Evaluation 

Contrary to the fixed shaft, of the extendible boom concepts selected for 

evaluation, the robotic manipulator is the most complex. The manipulator 

depicted in Fisure 9.2.2-7 was previously developed by MMC, and it is typical 

of robotic arm design. A manipulator arm enjoys the advantage of being a 

mu1 ti pl e degree-of -f reedom (DOF) device, capabl e of provi di ng target 

realignment and access to obstructed RSEs. 

as a result of joints at the base, elbow, and two at the spin/despin mechanism 

or grapple mechanism interface. A motor, gearbox, and control sensor are part 

The MMC design includes four DOF, 
of each joint, and together create significant control complexity and 

increased re1 iabil i ty and maintainability requirements. 

of aluminum alloy are incorporated, in an effort to minimize weight, yet the 

mechanism weight remains re1 ati vely high. 

manipulator has a very compact stowed configuration. 

Tubular arms composed a 

Like the 4-bar mechanism, the 

The study team felt that the higher weight and complexity of a robotic arm are 

offset by its articulation, compacted volume, and load bearing characteristics. 

The manipulator was favored as the preferred TSR kit extendible boom concept. 

ADVWAGE 

Cornput 

Cn Support Gnppio, Dospin Torquo Loda 

Exlaling Dnlgm 

En8bles Reallgnmont of Tug01 u888 for 

En8blos Aecoaa to Obatructod Rwovory 

Transport 

Suppyt Elemonts 

DISADVANTAGE 

HO8Vy 

Incroaaod Control Comploxlty 

Figure 9.2.2-7 Manipulator Evaluation 

The Extendible Stiff Arm Manipulator (ESAM), a more specific design of robotic 

man3 pul ator devel oped at MSFC, was i ncl uded in the extendi bl e boom subsystem 

evaluation. 

extendible structural section deployed from within a grounded section, both of 

square cross section. 

Shown in Figure 9.2.2-8, the ESAM structure consists of an 

Roll, pitch, and yaw end effector positioning are 

9-9 
incorporated in the wrist assembly of the extendible section. 

positioning, coupled with the azimuth, elevation, and extendhetract 

capability of the grounded section, make ESAM a six DGF mechanism. 

the ESAM joints includes a reversible DC motor, for which a control system 

effects position and rate control. 

27 inches. 

This 

Each of 

The ESAM extension range is approximately 

Most of the advantages and disadvantages of the manipulator can also be 

attributed to ESAM, as both are robotic manipulators. A significant 

difference though, exists between the methoas of articulation of the 

mechanisms. 

obstructed RSEs nor target realignment, as its articulation is accomplished at 

either the base or the end effector. Modification of the ESAM design could 

provide a more appropriate extension range and adequate load bearing 

capability, but a complete redesign would be necessary to meet the 

articulation requirement. 

ESAM, while a six DOF mechanism, does not enable access to 

ADVANTAGE 

Compact 

Can Support Grrppk, Dmpln Torquo Lorda 

Modl!lcrtlon to Erlatlng Dnlgn 

USADVANTAGE 

* Hervy 

Increased Control Comploxlty 

Figure 9.2.2-8 ESAM Evaluation 

9-1 0 
9.3 Spi n/Despi n Mec hani sm Def i ni ti on 

The concept selected for the Spin/Despin Mechanism subsystem is an existing 

design, functioning as part of a three axis gimbal within the WMC Space 

Operations Simulation Laboratory. An assessment of this design was 

accompl ished in order to provide conceptual level definition for foll ow-on 

efforts. The mechanism consists of a spin platform, DC torque motor, servo 

control ampl ifier, tachometer, and electrical brake. 

Rotation of the spin platform in either direction is effected by the 

reversible backdrivable DC torque motor, through a direct interface between 

the two. The backdrivable system necessitates an electrical brake to hold the 

grapple mechanism stationary during stowage, before spin up, and after 

despin. Speed control is accomplished by coupling the tachometer signal with 

current adjustment of the amplifier. 

increases the current to effect a proportional torque increase of the motor, 

until the tachometer senses the correct speed. In the despin mode, the DC 

torque motor functions as a generator. The current generated is dissipated 

across the ampl ifier, producing a proportional decrease in torque. 

In the spin up mode, the amplifier 

9.4 Definition of Grapple Mechanism Interface Devices 

The grapple mechanism interface device definition process identified the 

requirement for unique interface devices for both ground assembly, for Orbiter 

based operations, and on-orbi t recovery kit assembly, for potential Space 

Station operations. 

are discussed bel ow. 

An interface device was defined for each case and they 

The ground assembled interface flange, conceptually illustrated in Figure 

9.4-1, provides an interface connection between subsystems, such as that 

between the grapple mechanism and the spin/despin mechanism. 

identified as part of ground assembled recovery kits which require no on-orbit 

changeout of subsystems. 

interface. 

The flange was 

A bolted assembly is required for each subsystem 

Initially intended as a grapple mechanism interface device, it can be used as 

a readily applied interface between several components of a modular, interchangeable recovery system. 

9-1 1 
9 

I F IanI I 

Extendible I Boom 

I 

I 

Figure 9.4-1 Ground Assembled Interface Flange 0 

For recovery kits based at the Space Station, it was assumed that assembly of 

recovery kit elements would be accomplished robotically. 

requires an interface mechanism capable of providing a simply accomplished, 

firm connection during assembly. The Robotically Operated Interface Device, 

illustrated in Figure 9.4-2, was identified and assessed as a possible 

on-orbit interface accommodation. 

This, in turn, 

The device includes both a passive and an active element. The latter consists 

of an alignment cone and electrically operated latches. 

includes a plate with lips on opposite edges, on which the latches secure, and 

a cavity hollowed to the shape of the cone. 

opposing subsystems at the interface. As the mechanisms are brought together, 

the mating of the cone and cavity provides an accurate alignment. 

of the latches is accomplished by a small motor, effecting a simultaneous 

rotation of the latches through an open/lock sequence. The travel of the 

The passive element 

The elements will be attached to 

Actuation 

9-1 2 
latches, in degrees, is small and the rotation is accomplished quickly. A 

secure connection is maintained as. a result of incorporating a nonbackdriveable system. Like the ground assembled interface flange, the device is 

required at each subsystem interface. 

The device has been previously developed, which provides the additional 

advantage of being an existing design. As with other robotic devices, it has 

higher re1 iabil i ty and maintainabil i ty requirements and additional control 

compl exi ty . 

USADVANTAGE 

Addltlon8.l Control Complrxlty 

Hlghr Rell8blllty. M8lnt8in8blllty Requlremrnta 

Figure 9.4-2 Robotically Operated Interface Device 

9.5 Grapple Mechanism Evaluation 

9.5.1 Introouction - Three categories of grapple mechanisms were identified 

which address the grapple mechanism requirements of Systems B and C. 

mechanisn types and their corresponding systems are shown in Figure 9.5.1 -1. 

Within the categories, a1 ternatives were recognized and evaluated, from which 

the selections for the more detailed evaluation of follow-on efforts may be 

The 

9-1 3 
Small Gripper - System B 

Stinger - Spin Stabilized, Controllable - System C 

Large Enveloper - System C 

Figure 9.5.1 -1 Grappl e Mec hani sm Eva1 uati on 

accomplished. The criteria identified and applied in the evaluations 

i ncl uded: weight, compactness, accomnodati on of a range of sate1 1 i te and 

hardpoint sizes, strength of grapple, target damage potential, positioning 

flexibility, accomnodation of grapple, despin, and transport loads, and 

compl exi ty . 

9.5.2 Small Gripper Definition - The MSFC Proto-Flight Manipulator Arm (PFMA) 

End Effector and the JPL PFMA Smart Hand were selected by the study team as 

two possible accommodations of the System B small gripper mechanism 

requirement and are illustrated in Figure 9.5.2-1. 

motion, minimizing reaction away from the mechanism during grapple. 

intermeshing parallel plate designs and square recessed shapes of the jaws 

enable grapple of hardpoints of varied size and shape; although in both cases, 

the maximum jaw opening is limited to 3.5 inches. Variable closure rates and 

grapple forces are incorporated, with maximum grapple forces of 90 and 120 

pounds. 

grapple force in the event of power loss. 

Both provide parallel jaw 

The 

Each of the.mechanisms includes a provision which maintains the 

n in 
ADVANTAGE 

Enables Grapple of a Varle of Round, Flat. or Irregularly-Shaped Har$olnts 

Figure 9.5.2-1 Small Gripper Definition - PFMA End Effectors 

DISADVANTAGE 

Slre of Hardpoints Grappled Llmlted by Geometry of Mechanlsm 

These existing mechanism designs reflect the state-of-the-art for 

non-dexterous hands and provide capabl e accommodations of the small gripper 

requirement. 

mission model hardpoint configurations will enable complete assessments of the 

appl icabil i ty of these mechanisms to the System B requirements. 

Follow on efforts involving analyses of transport 1 oads and 

9.5.3 

the grapple mechanism of System C, involving recovery of spin stabilized 

spacecraft, is the Apogee Kick Motor (AKM) Attachment Device (Stinger). 

capability has been proven in the Westar-Palapa recovery mission, where the 

recovery of two spin stabilized spacecraft was safely accompli shed without 

damage to either vehicle. 

Stinger-Type Grapple Mechanism - The obvious and unique selection for 

Its 

The stinger is depicted in Figure 9.5.3-1. 

9-1 5 
ADVANTAGE I DISADVANTAGE 

Enables Grapple of Spin Stabiilzd Targot 

Design Available, Proven 

I 

Figure 9.5.3-1 Stinger-Type Grappler 

9.5.4 Evaluation of Large Envelopment-Type Grapple Mechanisms - Accommodation 

of the System C, 1 arge envel opment-type grapple mechanism requi rement was 

offered to the study team by a large number of varied concepts. The five most 

promising were identified as a possible focus for follow on efforts, and are 

illustrated in Figure 9.5.4-1. An additional MSFC concept was later 

identified and evaluated. A summary of each of the concept evaluations is 

provided bel ow. 

Shown in Figure 9.5.4-2, the Mu1 tisegmented Arm proposes envelopment-type 

capture of recovery candidates. 

its two arms are actuated by a combination of spring force and cable tension. 

Cables are attached to the links at the tips of each arm and wound through the 

remaining links in such a way, that when tightened by a motor and reel at the 

Composed of a hinged series of small links, 

base, cause an inward curling motion of the arms. Springs located in each 

segment, oppose the cable tension and with the cables released, effect the 

fully deployed, straightened configuration of the arms. A high cable tens 

is required to overcome the force of the springs and produce sufficient 

contact force at the satellite surface. Inherent in the linkage structure 

mechanical stops, which 1 imit the minimum grapple diameter and compacted 

on 

a re 

vol ume. 

a-1 6 
Space 

C-Clamp 

Martin Marietta Enveloper 

Multisegmented Arm 

Figure 9.5.4-1' Enveloper Mechanisms 

ADVANTAGE 

Accompllshes Rlgld Grapple 

MSADVANTAGE 

Not Compact 

Llmlted Capaclty for Target with Spln on 

Longitudlnal Axis 

Relatively Heavy 

Llmlted Flexlblllty In Posltlonlng of Grapple 

S stem Not Applicable to Wlde Range of 

Mechanlsm 

4atellite Sires 

Figure 9.5.4-2 Mu1 tisegmented Arm Evaluation 

9-1 7 
The mechanism provides a consistent but una1 terabl e envelopment motion, and 

its recovery capability is 1 imited in tight target configurations. 

of hinges within the arms prevents the Multisegmented An from being a light 

weight device. Additionally, the negligible surface area at the tips of the 

arms, contributes to an already limited capability of the mechanism to recover 

satellites with spin on the longitudinal axis. 

significant disadvantages of not having a compact stowed volume and a limited 

The number 

These, and the more 

envelopment diameter range, withhold the 

for foll ow on efforts. 

The C-C1 amp mec tiani sm and its evaluation 

Although designed for recovery of satell 

Mu1 tisegmented An from consideration 

summary are shown in Figure 9.5.4-3. 

tes with spin on the longitudinal 

axis, it also accommodates transverse axis spin for satellites within a small 

range of diameters. 

through a variety of possible actuation methods, are caused to rotate. 

Satellite capture is then accomplished between the large surface areas of the 

arm tips. For satellites with large diameters and. transverse axis spin, 

actuation of the mechanism is accomplished in the same way, but the arms 

envelop the satel 1 ites, with satel 1 i te surface contact possible over the 

length of the arms. 

compacted volume and severely restricts the mechanism's capability in dealing 

with tight satellite configurations. 

The mechanisms two solid arms are hinged at the base, and 

The solid composition of the arms prevents a small 

A1 though 1 imi ted as an overall System C grapple mechanism, the C-Clamp 

provides a viable method of capturing a specialized case of disabled 

satellites. 

I 

The Space Bola, depicted in Figure 9.5.4-4, provides the most unique recovery 

method of the concepts selected for evaluation. The mechanism consists of 

three inflatable arms of light weight, flexible material, with a Velcro 

covering on the an tip surfaces and small rockets attached near the ends. 

Initially, the arms are inflated to deploy them from a stowed configuration. 

They are then deflated, simultaneous to the firing of the rockets. 

the arms are forceci to wrap around a recovery candidate, their ends make 

contact, and a Velcro connection is secured. 

is applied to the arms by motor driven spools. 

Together, 

To rigidize the grapple, tension 

9-1 8 
ADVANTAGE 

Llght Weight 

A pllcabk to a Wide Ran e of Satelllto Slres 

(eongitudlnal Axis Spin 8niy) 

Figure 9 5.4-3 

DISADVANTAGE 

Not Compact 

Llmlted Capaclty for Target wlth Spin on Transverso Axis 

Limited Flexibility, to Grapple In Tight Target Configurations, in Positioning of Mechanism 

ADVANTAGE 

Llght Weight 

Compact 

C-C1 amp Eva1 uation 

DISADVANTAGE 1 

Limited Flexibility In Positioning for 

Envelopment 

Gra le 0 eration not Repeatable - &Lo Pastenar closed - Rocket Reload Required 

Strength of Arms Questionable for Despin, Target Transport to STS a Space Station 

Rigld Grapple Is Questlonable 

Figure 9.5.4-4 Space Bola Evaluation 

9-1 9 
In concept, the Space Bola provides an extremely compact and light weight 

accommodation to the System C grapple mechanisM requirement, but many aspects 

of the envelopment operation are questionable or impractical. Of these, the 

most si gni f icant are the strength of the vel cro connect1 on and the possi bi 1 i ty 

of an inaccurate envelopment by the arms, requiring a reload of the rockets. 

Additionally, the i nfl exi bili ty of the envelopment method makes recovery of 

sate1 1 ite configurations involving protuberances impractical . 

Seemingly a simple concept, yet a significant development effort would be 

required to produce an operational mechanism in which confidence was high. 

The Debris Capture Device developed by LTY, is shown in Figure 9.5.4-5. It 

consists of a pair of low pressure toroids mounted on adjustable arms, with 

which it accomplishes a two point grapple similar to that of the C-Clamp. 

grapple force is applied through the arm assembly, by a hydraulic system at 

the base of the mechanism. 

by conforming to irregular satellite surfaces and aissipating the forces of 

contact dynamics. 

surface area of the toroids securely accommodates a wide range of satellite 

sizes. 

The 

The toroids minimize damage to recovery candidates 

Although providing only a two point capture, the large 

The Debris Capture Device is one of the preferred concepts of those selected 

for evaluation, yet it has several disadvantages. 

large, mandating a voluminous mechanism of relatively high weight. 

control complexity and re1 iabil i ty/maintainabil i ty requirements resul t from 

the hydraulics, toroids, and the many linkages involved. Following grapple 

and despin, the mechanism's capability in providing sufficient rigiditation 

during transport is questionable, as contact wfth the recovery candidate is 

sustained by a compressible device. 

The mechanism structure is 

Added 

Finally, as presently configured it 

offers 1 imi ted compactness. 

9-20 
WANTAGE 

Accommodates a Wlde Range of Satelllte Slres 

Soft Contact, Mlnlmlzes Target Damage 

DISADVANTAGE 

Not Compact 

Heavy 

Complex Control 

Quertlonable Transport Rlgldlzatlon Capability 

Bulk 

Conri’gurat ions 

Increased RellabllltylMalntalnabIlIty Requirements 

Llmlted Flexlblllty In Tlght Target 

Figure 9.5.4-5 Debris Capture Evaluation 

The MSFC Enveloper, depicted in Figure 9.5.4-6, has undergone preliminary 

design/development by MSFC. 

combination of the Multisegmented Arm, in which all of the arm links operate 

in conjunction, and the MMC Enveloper, evaluated below. 

The mechanism is similar in operation to a 

Two methods of actuating the grappler’s arms have been incorporated in the 

mechanism. 

actuation system between the two, which effects an inward/outward rotation of 

the arms. 

another, and the similar links of the opposing arm, as a result of a motor 

within the base and a corresponding linkage. The linkage causes an equal 

rotation of all four arms, producing a curling motion like that of the 

Multisegmented Arm. The combined effect of the two systems is an increase in 

flexibility in positioning for envelopment, and an increased capability in 

recovering small er sate1 1 i tes. 

The base links of each arm are connected and operated by an 

The remaining two ann links operate in conjunction with one 

9-21 
t ADVANTAGE DISADVANTAGE 

Llght Welght Increased RellabllltylMalntalnablllty 

Relatively Compact 

Accommodates a Range of Satelllte Sizes 

Flexlblllty In Posltlonlng for Envelopment 

Requlrements 

The device is designed for envelopment-type capture, yet the width of the end 

links also enables a two point capture, thereby accommodating longitudinal 

axis spin. 

actuation system, in an effort to minimize target damage. 

light weight mechanism but its compacted volume remains large. 

increase in control complexity and reliabilityhaintainability requirements, 

relative to some of the other concepts, results from its increased flexibility. 

Hard rubber pads are provided on the links and the cross member 

It is a relatively 

A slight 

The final mechanism evaluation, shown in Figure 9.5.4-7, involves the MNC 

Enveloper, which was conceptually designed by the study team to provide a 

System C grapple mechanism which was designed from the requirements allocated 

to the envelopment grappler, done in Task 1. 

mechanism is presented in the Task 2.2 accomplishments (see Section 10.0); 

therefore only a brief overview and evaluation are provided here. 

A detailed description of the 

9-22 
The mechanism incorporates a two arm, six member structure. 

torque motors in the joints, each of the links can be operated independently. 

To maintain symmetry, control software is employed so that opposing links, 

i.e., the base links of each am, operate in conjunction. The geometry of the 

members allows a folded, and extremely small compacted volume, with sufficient 

surface area at the end of each arm to enable a two point capture. 

MSFC enveloper, appropriate structural member design and composition will 

produce a re1 ati vely 1 i ght weight mechanism. 

Driven by DC 

Like the 

The substantial increase in flexibility of the grapple operation as a result 

of the independent activation of the links, provides a number pf advantages. 

The flexibili ty in positioning for envelopment, around varied protuberance 

configurations is optimized. The design enables envelopment before contact, 

thereby dissipating target reactions and negating target reaction away from 

the mechanism. 

reduced with the grapple elements closer to the recovery target when first 

contact occurs. 

proportional increase in control complexity and re1 iabil ity/maintainabil i ty 

demands. 

in Figure 9.5.4-7, and this enveloper mechanism was selected for inclusion in 

the System C design configuration. 

Target damage is minimized, as contact dynamic forces will be 

The mechanism's increased flexibility though, creates a 

These disadvantages are outweighed by the number of advantages shown 

AOVMAGE 

Llght Welght 

Compact 

Opllmum Flexlblllty In Poslllonlng for 

Envelopment 

Accommodates r Wlde Range of S8teIllte Slz.8 

onlrol of Multl le Llnks E uallres 

%ontact Force/&nlmltes SAelllie 08nuge 

Accompllshes Envelopment Before Contact 

Accomplishes Rlgld Grapple 

DISADVANTAGE 

Added Control Complexity wlth Mulllple Conlrollrble Arm !3gmenls 

Figure 9.5.4-7 MMC Envel oper Eva1 uation 

9-23 
The matrix of Figure 9.5.4-8 provides a summary of the System C envelopmenttype grapple mechanism evaluation. 

evaluations are shown appropriately applied to each of the mechanism concepts. 

The advantages identified in the 

MULTISEGMENTED 

ARM 

C-CLAMP 

SPACE BOLA 

DEBRIS 

CAPTURE 

MSFC 

ENELOPER 

MMC 

ENVELOPER 

Figure 9.5.4-8 Envelopment-Type Grappler Evaluations - Summary 

9-24 
10.0 TUMBLING SATELLITE RECOVERY SYSTEM CONCEPTUAL DESIGN - TASK 2.2 

e The objective of the second concept definition task, Task 2.2, was to develop 

a Set of conceptual designs for TSR systems. The approach used was to select 

the preferred set of subsystem mechanisms and to fntegrate them into a basic 

interchange of subsystem components. This, of course, enables the system to 

be assembled readily into a variety of different recovery systems tailored to 

the unique recovery scenario presented to the user. 

TSR kit that met the study team's objective of system modularity and ready 

10.1 TSR ConceDtual Desi an Dri vers 

The formulation of a design architecture for the recovery systems was 

influenced heavily by a number of key factors driven out by the requirements 

analysis. These are shown in Table 10.1-1. The first of these was the 

inherently broad range of recovery scenarios identified during Task 1. This 

fundamental reality caused the study team to select from two apparent 

options: 

range of recovery systems; or (2) develop a design concept with a modular 

design as a framework thatcould be configured readily into recovery systems 

tailored for specific missions. The latter approach was selected. 

(1) operate from a design concept that would provide an equally wide 

@ 

Table 10.1-1 Key Design Drivers 

Broad Range of Recovery Scenarios Dictates 

- Wide Range of Recovery Systems, or 

- Modular System Easily Configured into Recovery System Tailored for Specific Missions 

Recovery Kit Must Be Compact - Efficient STS Operations 

Minimum Risk to OMV 

Recovery Operations Bounded by OMV Controllability 

Target Rigidized for Return to STS or Space Station 

10-1 
Another prominent design factor was related to minimizing risk to the OMV 

during all phases of recovery operations. 

to maintain a proper distance from the OMV during operations, a design 

requirement readily accommodated by the selection of an effective extension 

device. 

retain control of the OMV during recovery operations. 

here was the concern that with contact dynamics forces in excess of OMV 

control authority, ground controllers coul d 1 ose the capabil i ty of control 1 i ng 

OMV and the TSR kit, with potential resulting damage to both. 

One risk element involved the need 

A related secondary design driver was the perceived necessity to 

The primary concern 

Finally, no other study had considered the requirement to have the target 

satellite firmly grappled for transport back to the Shuttle or to Space 

Station. 

different recovery systems. 

This requirement influences the choice of grapple mechanisms for the 

10.2 Modular TSR System Design 

The representation of a modular design for a family of recovery systems is 

provided in Figure 10.2-1. 

of the system and the inherent capacity to interchange, add, or subtract 

individual components to tailor the kit for specific missions. 

This expanded view displays the major components 

As shown on Figure 10.2-1, one of the primary features of the design is a 

structural interface element that is readily attached to the OMV docking 

latches in manual mating operations in the STS cargo bay during launch 

processing. The electrical power and communications and data management 

(C&DM) interfaces wil-1 be through the OMV payload umbilical mounted on the 

front face of the OMV. 

Another key design criterion turned out to be the need for a compact design. 

The TSR system is being designed as an OMV front end kit and carried into 

orbit in the Orbiter cargo bay. The OMV has been configured for compactness 

to minimize cargo bay space necessary for delivery into orbit. Though no 

other study ref1 ected this requirement, MMC believed the structural design 

should be as compact as possible and selected subsystem options that supported 

this design driver. 

10-2 
0 

c 

10-3 
1 

The next component, referring to Figure 10.2-1, is the extendible boom. 

four degree of freedom manipulator arm with pitch and yaw positioning at the 

grapple mechanism interface was selected as the extendible boom for the 

system. 

Its most important function is to provide safe clearance between the OMV and a 

spinning target. 

recovery support elements that are obstructed to an approach by OMV due to 

deployed target elements, such as solar arrays, antennas and experiment 

packages. 

orbit transfer thrust vector following capture and prior to orbit transfer. 

A 

This mechanism is capable of folding into a compact stowed position. 

The manipulator arm provides the capability to reach 

It also enables alignment of a target's center of mass with the OMV 

A third major component is the spin table. The spin table will house a direct 

current reversible torque motor and a tachometer unit to enable precise 

control of the spin-up of the recovery system and maintenance, within close 

tolerance, of the spin rate and phasing of the system during recovery 

operations. The spin table can be mounted efficiently to the interface flange 

on the end of the extendible boom. 

table is a boresight, wide angle view television camera. On this base, it can 

be mounted in a fixed configuration, or configured to spin at the same rate as 

the spin table. During initial independent research in satellite recovery 

simulations, it was unclear whether the operator was supported best by a fixed 

or spin rate matching boresight camera. This optional camera configuration 

was included to support development of a flexible system design architecture. 

. 

Also shown configured within the spin 

. 

In this representation of the recovery system, both the System B and System C 

recovery kits are shown. 

mechanism attached to the flanged grapple mechanism interface device. This 

system will be capable of recovering satellites with what was previously 

defined as the more complex tumble motion to be expected. This motion will be 

created by a satellite failure producing some level of angular momentum that 

exceeds satel 1 i te atti tude control capabi 1 i ty. The resul ting tumbl e motion 

will be general motion initially, with multi-axis spin; however, it is 

expected to evolve to single axis spin about the satellite's major principal 

axi s. 

The full-up System C has the MMC enveloper grapple 

10-4 
The Sysa B conf,gration is also represen-ed on Figure 10.2-1. This system 

includes the structural and el ectrical/C&DM interface with OMV, the extendible 

boom and a small gripper, connected to the system with the grapple mechanism 

interface flange. 

attached to the smaller gripper to provide localized viewing of the attachment 

to the hard point of a remote, controllable satellite. 

amenable this design is to subsystem interchange and its capacity for 

accomnodati ng a wide variety of recovery missions. 

0 

In addition, a close proximity television camera is 

This portrays how 

10.3 TSR System - Compact Design 

As expressed earlier, one of the key design drivers was the requirement to 

make all the TSR kit elements as compact as possible for efficient transport 

with OMV in the Orbiter cargo bay. In Figure 10.3-1, this presentation of the 

full-up recovery System C shows all of the subsystem mechanisms in folded or 

non- deployed configurations. All of the subsystem a1 ternatives shown here 

and evaluated in Section 9.0 were the most compact of the alternatives 

considered. The single exception is the case of the extendible boom, where 

the 4-bar mechanism is considered equally compact when compared with the 

foldable manipulator arm. The scale of this set of mechanisms is accurate as 

shown, however, these designs are conceptual and have not been optimized for 

compactness. 

0 

Extendible Boom-Arm 

(Folded Position) Spinlhpin fable 

Grapple Mechanism 

(Folded Position) 

Interface Mechanism 

Figure 10.3-1 TSk - A Compact Design 

10-5 
Another example of how this conceptual TSR system was designed for compactness 

is shown in Figure 10.3-2. One of the key features of this grapple mechanism 

was the designed-in capacity to fold each of the outer two grappler segments, 

in each of the two arms of the grapple mechanism, into the adjacent segment of 

the arm. This design feature provides a highly compact grapple mechanism. 

None of the other grapple mechanism candidates were considered as having even 

a reasonable degree of compactness and that was one of the reasons the study 

team concentrated on the conceptual design of the MMC enveloper grapple 

mechanism. 

76 in. 

t-70 in.-- c 

t 

I 

1 

I 

I 

----,,, 

22 in. , I 24 in. 

I20 in. 

------- J 

The site of the grapple mechanism envelope is illustrated in Figure 10.3-2. 

The length of one of the pair of grapple arms, which is sized to envelope a 

fifteen-foot diameter target is nearly 19 feet. The inner segments of the 

grapple elements are sufficiently larger than the adjacent outer elements to 

allow each of the two outer segments to fold inside the other. This design 

approach was used to provide maximum compactness for the folded grappler. 

--- ------ -__ , 

-

_____------- 

226 in. 

- 

i 

- i

30 in. 

Figure 10.3-2 MMC Enveloper Grapple Mechanism 

10-6 
Another of the principal design drivers for the recovery system was minimizing 

risk to the OMV and TSR kit during recovery. 

cmpactness, as shown in Figure 10.3-3, the half-deployed extendible boom was 

incorporated into the system to reduce risk to OMV. 

shown in Figure: 10.3-3 as a four-degree-of-freedom manipulator arm, is 

extended to provide clearance between the OMV and a rotating satellite with 

spinning appendages. 

In addition to providing ' 

0 

The extendible boom, 

10.4 MMC Envel oper Grappl e Mechanism 

The conceptual design of the MMC enveloper, which was selected as the grapple 

mechanism element for one of the conceptual System C recovery configurations, 

was influenced by an increasing concern on the part of the study team 

regarding the potential impact of contact dynamics between the TSR kit and the 

target during the grapple phase. 

Dapin Table Extendible Boom-Am 

(Folded Position) 

\ I 

Grapple 

Interface Mechanism 

Figure 10.3-3 TSR - OMV Risk Reduction 

Mechanism 

10-7 
Even with perfect conditions during recovery, with no major target 

protuberances/appendages and gi ven a reasonably symmetric target for recovery , 

as the operator begins to grapple and rigidize, the grapple mechanism will 

begin a series of contacts with the target. 

relative position changes between the target and recovery system that are 

expected to be complex, and have not yet been modeled. When initiating this 

grapple mechanism closure operation with a two- or even three-point gripping 

device, such as the C-clamp, in which the target is not enveloped, It is 

possible that the target position will change in such a manner that a new 

approach and grapple positioning setup will be required following each 

contact. 

would be untenable. 

be accomplished more feasibly using an envelopment approach. 

spinning target with a spinning grapple mechanism would provide a higher 

probability of successful grapple and rigiditation. 

shown in a grapple configuration in Figure 10.4-1. 

These contacts will produce 

If this situation were to become prominent, this type of grapple 

For this reason, grapple of a spinning target appeared to 

Thus, the MMC ' enveloper was designed from the operational concept that envelopment of a 

I The MMC enveloper is 

- Thrn Pairs of Grapple Elmmts Indopmd.ntlv ConVolhd 

- Grapplr Mechanism Deployad to Opnmum Emdopnwnt CMlfigurrtion 

- Targrt EnveIopod-Elmnonts Clod Slowlv to Minimizm Contact Dynamics 

- Gr8pplr Machniun Rigidized for hpIn and Tmaf8r toss or 

- dc Torque Motors-Harmonic. Plv~prY 

- fhon Spun Up-Minimin Deplovmmt Oynrmss 

Figure 10.4-1 TSR Grapple Envelope 
Thi s grappl e mec hani sm is configured wi th three pai rs of independently 

controlled grapple segments. Each of the segment pairs will be controlled 

with direct current (DC) torque motors with backdrivable gearboxes. The 

extension and retraction operations of each pair of segments will be 

synchronized, within close to1 erances. This is essential to maintain a stab1 e 

system configuration when opening and closing each of the segment pairs, while 

the TSR system is spinning to match rates with a spinning target. 

@ 

During recovery operations. the envel oper will be unfolded gradual ly and 

deployed to an optimum envelopment configuration, prior to the spin-up of the 

system. This procedure will minimize the dynamics of the deployment of the 

TSR kit, a non-trivial matter. Deployment dynamics of the TSR system could 

impact OW atti tude control requirements and deployment dynamics is considered 

a subject for early research and study. The three segment arms will provide 

the operator with extended flexibility in aligning the grappler to envelop a 

target with complex shape. Following spin-up of the system and target spin 

rate and phase matching, the OMV ground controller and the TSR operator will 

Position the grapple mechanism to envelop the target, at the target's center 

of mass. 

the target with the OMV, and maneuvering the TSR kit to a position of 

envelopment of the target. The TSR ground controller will then operate each 

pai r of envel oper segments independently to grapple the sate1 1 i te, whi 1 e using 

care to minimize pre-engagement contact and the resulting irregular relative 

motion. As shown in Figure 10.4-1, the target will be enveloped, avoiding 

target contact, and the ground controller will commence closure for grapple 

and rigidiration of the .target. The grapple mechanism will be capable of 

softly grasping the target, accommodating the resulting forces and torques 

generated by target contact, and rigidizing the target grapple. The grapple 

points on the MMC enveloper will be rubber coated to absorb forces, minimize 

relative motion induced by contact, and reduce damage to the target. 

This will be accomplished by translatfng the cantilevered TSR kit to a 

Following accomplishment of a firm grapple, the TSR kit operator will reverse 

the current of the TSR spin/despin motor (generator) and despin the 

satellite. The OMV will provide adequate reaction control forces to dissipate 

the target's excess angular momentum and to stabilize the target. The TSR 

10-9 
system will then repositfon the extendible boom, the OMV operator will 

initiate a short series of translation maneuvers to test the new "system" 

center of mass offset, and fire main thrusters for return to the operating 

on-orbit base, Shuttle or Space Station. 

10.5 TSR Capture Envelope Flexibility 

Many of the grapple mechanisms evaluated in the concept definition^ task were 

not capable of grappling a wide range of target satellite sires. 

assessment of the TSR mission model revealed a substantial range of target 

capture envelopes, with satellites ranging from four to fifteen feet in 

diameter. As shown in Figure 10.5-1, the MMC enveloper grappler mechanism is 

capable of capturing a broad spectrum of target satellite sizes, and would be 

limited solely by the ultimate decision on length of grapple elements. 

The 

The conceptual System C, non-control 1 ab1 e, tumbl i ng satel 1 i te recovery system 

with the MMC enveloper is shown in Figure 10.5-2, capturing the recently 

disabled and abandoned NOAA-8 satel 1 i te, and demonstrating a recovery that 

could be accomplished today if the OMV and TSR kit were available. 

lost recently, in January 1986 and its replacement, NOAA-10, was destroyed in 

a launch failure in May 1986. 

NOAA-8 was 

Figure 10.5-1 MMC Enveloper Target Capture Envelope Flexibility 

10-10 T+g&-- 
L 

L aJ 

a. 

0 

x 

c 

c, 

h 

aJ 

V 

E 

0 

u 

10-1 1 3 c-7 
10.6 OMV/TSR Kit Interfaces 

The design of the TSR system takes into account a number of interfaces with 

the OMV to ensure that the recovery kit can perform a variety of missions 

while connected to the OMV. 

interchangeable mechanisms, a data processor (microprocessor) to col1 eCt 

sensor data and format it for transmission via OMV communication links, a TV 

camera and lighting. All system support functions such as electrical power, 

telemetry and control comnunications, structure/mechanisms, attitude control 

and stabilization are interface requirements that must be met by the carrier 

vehicle, the OMV. 

accomnodations for the TSR kit is shown on Table 10.6-1. 

The basic TSR kit consists of a set of 

An outline of the general interface requirements/ 

Tab1 e 10.6-1 OMV/TSR Kit Interfaces/Accomnodations 

o Mechanical 

- Standard docking latches 

- Payload accommodation umbilical 

o Electrical Power 

- 28 vdc only 

- 250 watts operational 

- 50 watts standby 

o Telemetry 

- ' Sensor control data, system status - 64 kbps 

o Comnands 

- 1 kbps 

o Attitude Stabilization and Control 

deployment, envelopment, contact dynamics, despin, orbit 

transfer 

- Adequate for maintaining control during: -- 

The mechanical/physical interface for the conceptual TSR kit will be the 

standard OMV docking latches. 

pins on the TSR kit in mating operations in the Orbiter cargo bay during 

payl oad processing. 

electrical/ communication interconnection between the OMV and TSR kit and that 

will be the payl oad accomnodations umbilical. 

These latches will grapple standard FSS latch 

The other physical interface/accomodation is the 

10-1.2 
The electrical power requirement will be 28 volts direct current (DC). The 

torque motor for all mechanisms is baselined at 28 volts DC power. The 

maximum operational power level is anticipated to be required during despin. 

Power required during the despin operation, using a despin period of 500 

seconds, is approximately 175 watts. With an allowance of 75 watts for other 

electrical functions underway at that time, the maximum operational power 

requirement is approximately 250 watts. The standby power requirement of 50 

watts is required for heating motors, gearboxes and the microprocessor. 

0 

Telemetry data requirements include sensor ana system status data sent to the 

ground and is estimated at 64 kilobits per second (kbps). 

tions will require a telemetry capability of 256 kbps. 

Video communicaThe attitude control and stabilization requirement is varied and extensive in 

scope. The OMV ACS system and control laws must be adequate to'maintain 

control of the OMV, TSR kit and disabled satellite during all recovery 

operations including deployment of mechanisms. The completion of recommended 

studies and experiments related to deployment and spin-up of the recovery kit 

elements, and experiments designed to examine contact dynamics forces and 

torques during grapple operations, will support the determination of ACS 

requirements for kit operations. 

J) 

10.7 ConceDtual Recovery Systems - SummarY 

The recommended recovery system architecture and conceptual system designs are 

presented in this section in a format that will illustrate the efficacy of the 

MMC modular, interchangeable element approach. 

Shown in Figure 10.7-1 are the system configurations for System 8, for both of 

the recovery scenarios described in DRM 3 and DRM 4. For Case 1, the scenario 

is a controllable, stable target, with a recovery support element (RCE), an 

RMS grapple fixture that is obstructed from a direct OMV approach by a 

deployed solar panel. For this recovery candidate, the conceptual TSR system 

consists of the structural/mechanism interface element, a mu1 tiple 

degree-of-freedom manipulator arm (to gain access to the grapple fixture), a 

grapple mechanism interface flange and an RMS end effector. This System 6 

configuration has the requi site capabi 1 i ty to recovery di sabl ed sate1 1 i tes i n 

this category of recovery scenario. 

10-1 3 
System B-Target Controllable, Recovery Support Element (RSE) Obstructed 

RMS End Effector 

System B- Target Controllable, No RSE Available 

Camera 

Small Gripper 

Figure 10.7-1 Conceptual TSR System B 
For the second System B scenario, the case of a controllable satel 1 i te with no 

grapple fixture or flight support system latch pins, the recovery system is 

also shown in Figure 10.7-1. The mechanical interface provides for ready 

attachment to the OW, and in this configuration a small gripper, for 

attachment to target hard points, is attached to the interface flange and the 

extendible boom. Again, the basic system is tailored to the specific recovery 

scenario. In the System 8 cases, cost benefits will be derived from the lower 

weight required for these recovery missions, when compared with the full-up 

system shown next. 

@ 

The recovery system configurations for both System C recovery scenarios are 

illustrated in Figure 10.7-2. In the first case, with a controllable, Spin 

stabilized target such as INTELSAT-6, the mechanical interfaces (both the 

structural and umbilical), the extendible boom and the spin table are 

included. The "stinger" type grapple mechanism, attached to the grapple 

mechanism interface flange, will be used to secure a solid grip on the 

INTELSAT kick motor, when the OMV and TSR operators have matched the spin rate 

and positioned the grappler for attachment. 

is configured or tailored for this mission through assembly and checkout 

operations at the OW kit storage area. 

Again, the basic modular system 

0 

Finally, the most difficult recovery scenario is the full-up System C 

scenario, which the previously described analysis indicates is most likely a 

tumblinghpinning satellite. This is the actual "complex" motion case in 

which the satel 1 i te is non-control1 ab1 e, due to excessive angul ar momentum 

being introduced into the system through some torque inducing failure, and as 

a result is tumbling or spinning about a major axis. 

requires the largest complement of modular system elements, including an 

envel oper-type grappler. 

with the spinning satellite and the grappler is positioned in phase with the 

geometric mass of the target, the OMV and TSR operators maneuver the system to 

me1 op the target and manipulate each of the independently operated grappler 

element pairs to effect a smooth, firm grapple. 

This recovery scenario 

Once this TSR kit configuration has matched rates 

Thus, the conceptual modular design does contain a1 1 of the "fundamental I' 

accommodations to enable recovery of the full range of the identified and e defined System B and System C mission requirements. 

10-1 5 
System C-Target Controllable, Spin-Stabilized 

Stinger 

System C-Target Noncontrollable, Tumbling/Spinning 

Figure 10.7-2 Conceptual TSR - System C 

Large Enveloper 

Grapple Mechanism 

10-1 6 
As a final summary note, the conceptual modular TSR kit just described was 

compared to the initial recovery concepts, as shown in Figure 10.7-3. 

evaluation, which does include critical design factors, such as the 

requirement for compactness, demonstrates the substantial increase in 

effectiveness of the alternate approach selected by MSFC and MMC in completing 

the concept definition phase, Task 2. 

concepts was evaluated at 80% of the maximum possible score, the selected 

concept was evaluated at 96% of that maximum possible score. 

This 

While the best of the initial recovery 

10-1 7 
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10-1 8 
11 .O SUPPORTING DEVELOPMENT PLAN 

'@ 11.1 Objectives and Summary 

The purpose of the supporting development plan is to outline the research and 

technol ogy devel opment, i ncl udi ng ground-based testing and simulation and 

on-orbit demonstration activities, and the flight hardware development needed 

to establish the technical readiness of an OMV tumbling satellite recovery 

front-end kit. 

The elements of the tumbl ing sate1 1 i te recovery development program i ncl ude 

ground based research and study efforts, Orbiter cargo bay or proximity 

operations experiments and the flight hardware development activities required 

to provide a validated, operational recovery system for future users. These 

activities must be integrated into a valid program for recovery kit planners, 

and must be coordinated with concurrent OW and Space Station development 

activities. The technology development issues identified in the Supporting 

Research and Technology (SRBT) report are addressed in either ground or 

flight-orfented experiments. io 

1 1.2 Ground Demonstrati on Acti vi ties 

The recomnended ground-based demonstration approach is out1 ined in Tab1 e 

11.2-1. The first recommended initiative is to design, develop and use a set 

of recovery kit ground demonstration units, incl uding software simul ation and 

hardware systems, capable of addressing the SRBT issues described in Volume I1 

of this final study report. 

hardware demonstration unit( s) will be used (1) to evaluate a1 ternative 

recovery concepts; (2) to examine system deployment characteristics; (3) to 

assess contact dynamics and resultant target .and recovery system reactions to 

recovery operations; and (4) to expand tests directed at determining operator 

capabilities (or limitations) in conducting recovery operations. 

As shown in Table 11.2-1, these programs and 

The plan would include the examination and possible use of existing MSFC/MMC 

1 aboratory configurations to conduct these types of ground-based technol ogy 

efforts. 4 

11 -1 
Table 11.2-1 Ground Demonstration Acti vi ties 

Design, Develop, & Exploit Recovery Kit Ground Demonstration Unit@) 

- Evaluate Concepts Feasibility 

- Recovery System Deployment Characteristics 

- Contact Dynamics in Recovery Operations 

- Recovery System Operations/Operator Assessment 

Utilize Existing MSFClMartin Marietta Simulation Capabilities to Address 

Identified Technology Issues 

- Contact Dynamics Concerns 

- Force 81 Moment Measurements, Resulting Position/Motion States 

- Computer Simulations Using Varying Configurations, Evaluate Human Factors 

Limitations 

Demonstrate Use of Recovery Demonstration Unit as Laboratory Tool 

- Evaluate Alternative Concepts 

- Evaluate Subsystem Mechanisms - Grapple Devices 

- Eventual Use as Astronaut Trainer for Flight Experiment 

- Identify Logical Flight Experiment Candidates 

A tni rd element of a ground demonstration program will be to use the new 

ground demonstration units or modifications or extensi ons of exi Sting 

facilities as development tools, as the program proceeds through the 

development process. 

revi sed concepts or subsystem mechani sms, including the extensi ve family of 

large and small potential grapple mechanisms. This equipment would be used 

eventually as training devices for astronauts conducting Orbiter cargo bay or 

proximity operations flight experiments. 

Developers will need to continue to evaluate new or 

As a final note, the conduct of ground-based experiments Elated to remote 

satellite recovery will support clarification of what tecnnology issues can be 

more efficiently addressed on the ground and which of these must be addressed 

oy flight experiments. 

The recommended schedule for development of ground demonstration software and 

hardware elements to.support the evolution of design and development of a 

tumbling satellite recovery system is provided in Figure 11.2-2. 

recommended that a series of requirements trades and analyses be conducted 

I 

It is I 

11-2 
. 

1 I I 1986 1987 1988 I 

Recovery System Description 

 systems Requirements 

,-hquirmenn-Trdm & Analyrrr 

-- Rwirements Mission Operations 

-Design Dwelopmrnt 

\-Design Trades & Analysis 

- Concept Solection & Approval 

- System Design 

I- Fabrication 

Acceptam testing & System InttJgmtion/Dmonrtratior: - 

Vorifiation & Continuing Research - 

Figuw 11.2-1 Ground Demonstration Schedule 

lo initially to support determination of the number and type of software and 

hardwam units that will cost effectively address those issues identified in 

Volume 11. This task would include evaluations of existing laboratory 

configurations in NASA and in industry. 

With ground demonstration objectives and requi rements identified, concept 

design and design trades and analyses are recommended to enable concept 

selection of the ground demonstration unit(s) in mid-1987 and approval by 

MSFC. With MSFC di rection to proceed, the system design, fabrication, 

checkout and demonstrati on of selected software and hardware elements for 

ground-based TSR technology support will be conducted and completed in FY 1988. 

Following production of these software and hardware units, they will be used 

throughout the TSR kit flight hardware defi ni tion and development periods to 

support resol uti on of evol vi ng tec hnol ogy i ssues. 
11.3 STS Cargo BaylProximity Operations 

I Tne definition of on-orbit flight experiments to support technology 

development of the TSR kit will evolve and be refined through expsrience with 

the ground demonstration units. 

be required to validate the recovery concept agreed upon for development, and 

to verify contact dynamics forces and torques and the impact of relative 

movement between the target and recovery system on the recovery operation. 

Candidate cargo bay/proximi ty operations experiments are outlined on Table 

It appears that an on-orbit experimnt will 

I 

I 11.3-1. 

Table 11.3-1 Cargo Bay/Proximi ty Operations Experiinents 

Define an STS Cargo BaylProximity Operations Equipment Set 

- Scaled Satellite Recovery System 

- Extendible Boom, Spin Table, Envelopment Grappler 

- Equipped with interface to STS & RMS 

- Scaled Composite Recovery Target 

Conduct Remote Recovery Experiments in Zero-G 

- Remote Recovery Operations 

-. Spin Axis Alignment, Spin Rate MatchingIPhasing 

- Operations, Operator Limitations 

- Recovery System Deployment Dynamics 

- Target-Recovery System Contact Dynamics 

Cargo Bay Experiments Should Be Phased to 

Support Flight Hardware Phase C/D CDR. 

I The on-orbit remote satellite recovery experiments should be conducted with 

I 

nign fidelity equipment to validate the system concept. 

the requirements and conceptual design of the scaled down experiment equipment 

should begin prior to the start of flight hardware Phase C/D for the TSR 

sy stem. 

Thus, definition of 

11-4 
-l c 

The experimental recovery equipment would be an extendible boom, a spin table 

and an envelopment grappler. The system would be designed to interface with 

the Orbiter RMS end-effector and equipped with an operating interface in the 

Orbiter. 

spacecraft bus, designed to be controllable for multiple tumble and spin modes 

and rates. 

ze ro-gravi ty remote recovery ope rati ons that Val i date ground-based 

experinents. Actual experience to be gained includes emote recovery 

operations, such as spin axis alignment, spin rate matching and phasing with 

target, recovery system deployment dynamics and activities conducted in 

reaction to target/ recovery system dynamics. 

@ The spacecraft target would be a modification of a current rented 

This experiment wuld enable operators to conduct the first 

The schedule for the cargo bay/proximity operations experiment is presented in 

Figure 11.3-1. The program phases are similar to those shown for the ground 

demonstration uni t(s) program. Again, the schedule is designed to provide a 

proof-of-concept experiment prior to flight hardware Critical Design Review. 

11.4 Flight Hardware Program 

The objective of .the actual OMV tumbling satellite recovery kit flight 

hardware program is to be prepared to conduct free flight operations in 1993 

with actual or simulated targets. This flight hardware will be developed on a 

schedule consistent with development plans for OMV and other OW front-end 

kits. 

conceptual, definition and development phases (A, 8, and C/D). 

It wi 11 be conducted usi ng the generally accepted NASA/MSFC approach of 

During the flight hardware program, the plan is to continue to examine and 

define requi rements for accommodating the Space Station and TSR kit for 

deployment to, and operations from, the Space Station. 

The Supporting Development Plan (SDP) schedule is provided in Figure 11.4-7. 

The schedule outlines an integrated TSR kit development program that includes 

ground-based and on-orbit STS flight experiments and a flight hardware program 

that provides for free flight operations in 1993. 

11-5 
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