MCR-85-1365

Contract NAS8-35625

DPD 650

DR-5

(**?»

Final

Technical

Report December 1985

Servicer System

Demonstration

Plan and Capability

Development

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MARTIN MARIETTA

DENVER AEROSPACE

P.O. Box 179

Denver, Colorado 80201
MCR-85-1365

Contract NAS8-35625

DPD 650

DR-5

Final

Technical

Report December 1985

SERVICER SYSTEM

DEMONSTRATION PLAN AND

CAPABILITY DEVELOPMENT

Prepared by:

M.A. Bulboaca

J. A. Cuseo

W. L. DeRocher, Jr.

R.W. Maples

P. C. Reynolds

R. A. Sterrett

Approved

W. L. DeRocher, Jr.

Program Manager

*

MARTIN MARIETTA

DENVER AEROSPACE

P.O.Box 179

Denver, Colorado 80201
FOREWORD

This document was prepared by Martin Marietta Corporation under Contract

NAS8-35625, Servicer System Demonstration Plan and Capability Development,

Data Procurement Document 650, Data Requirement DR-5, Final Technical Report.

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

the National Aeronautics and Space Administration under the technical

direction of the Orbital Maneuvering Vehicle Task Team, with Mr. James R.

Turner as the Contract Technical Manager.

11
CONTENTS

Foreword ii

Contents ill

1.0 SUMMARY 1- 1

1.1 Introduction 1- 3

1.2 Study Objectives 1- 6

1.3 Relationship to Other NASA Efforts 1-10

1.4 Study Approach 1-14

1.5 Significant Conclusions and Recommendations 1-16

1.5.1 On-Orbit Servicing Development 1-16

1.5.2 Multi-Mission Modular Spacecraft Servicing 1-17

1.5.3 Ground Demonstrations 1-18

1.5.4 Cargo-Bay Demonstration 1-22

1.5.5 Free-Flight Verification 1-25

1.5.6 Servicer/MMS 1-g Demonstration Plan 1-27

1.5.7 Servicer/MMS 1-g Demonstration Equipment 1-29

1.5.8 Servicer Control Software - Basic Modules 1-31

1.5.9 Servicer Control Software - MMS" Modules 1-33

1.5.10 Servicer Software Demonstrations 1-35

1.6 Suggested Additional Effort 1-38

1.6.1 Servicing Tasks 1-39

1.6.2 Servicing Mechanism 1-39

1.6.3 Demonstrations 1-40

2.0 INTRODUCTION AND BACKGROUND 2- 1

2.1 Study Objectives 2- 3

2.2 Background 2- 8

2.3 Study Approach 2-15

3.0 MMS SERVICING CONSIDERATIONS 3-1

3.1 MMS Servicing Requirements 3- 7

3.2 MST Adaptation for 1-g 3-11

3.2.1 Modified Module Servicing Tool Requirements 3-11

3.2.2 Modified MST Mechanical Concept Definition 3-14

3.2.3 Modified MST Electrical Concept Definition 3-16

3.3 Module Exchange Options 3-19

3.3.1 Servicer Configuration 3-19

3.3.2 Alternative Modules 3-32

3.4 Extended Remote Maintenance/Resupply Concepts 3-37

iii
CONTENTS (Coat.)

4.0

4.1

4.2

4.2.1

4.2.2

4.2.3

4.2.4

4.3

4.3.1

4.3.2

4.3.3

4.3.4

4.4

4.4.1

4.4.2

4.4.3

4.5

5.0

5.1

5.2

5.2.1

5.2.2

5.2.3

5.2.4

5.2.5

5.2.6

5.2.7

5.2.8

5.2.9

5.2.10

5.2.11

5.2.12

5.2.13

5.3

5.4

6.0

6.1

6.2

6.2.1

6.2.2

6.2.3

6.2.4

6.2.5

~

DEVELOPMENT PLAN SUMMARY

Cargo-Bay Demonstration Schedule and Cost Estimate ....

Free-Flight Verification

Free-Flight Verification Objective

Free-Flight Verification Schedule and Cost Estimates . . .

1-g DEMONSTRATION PLAN ...'."

Multi-Mission Modular Spacecraft Module Exchange

Automatic Target Recognition and Error Correction

Convert ETU Control System from Analog to Digital

Optional Support Tasks

CARGO-BAY DEMONSTRATION PLAN

Page

4- 1

4- 1

4- 2

4- 3

4- 5

4- 5

4- 7

4- 10

4- 11

4- 12

4- 13

4- 15

4- 19

4- 19

4- 20

4- 22

4- 25

5- 1

5- 4

5- 6

5- 7

5- 8

5- 9

5- 13

5- 14

5- 15

5- 17

5- 17

5- 18

5- 18

5- 19

5- 19

5- 21

5- 21

5- 23

6- 1

6- 4

6- 10

6- 14

6- 14

6- 18

6- 21

6- 22

iv
CONTENTS (Cont.)

6.3 Cargo-Bay Demonstration Configuration Alternatives .... 6-23

6.3.1 Alternative Configurations 6-23

6.3.2 Configuration Descriptions 6-26

6.3.3 Use of the RMS During 0-g Demonstrations 6-34

6.3.4 Configuration Selection Criteria 6-38

6.3.5 Configuration Selection 6-41

6.4 Cargo-Bay Demonstration Schedule 6-44

6.5 Cargo-Bay Demonstration Cost Estimate 6-48

7.0 FREE-FLIGHT VERIFICATION PLAN 7- 1

7.1 Free-Flight Verification Requirements 7-3

7.2 Free-Flight Verification Elements 7- 9

7.2.1 Space Transportation System Support 7-12

7.2.2 OMV With Servicer System . 7-13

7.2.3 Serviceable Spacecraft 7-14

7.2.4 Airborne Support Equipment 7-15

7.2.5 Ground System 7-15

7.3 Free-Flight Verification Schedule 7-17

7.4 Free-Flight Verification Cost Estimate 7-19

8.0 SERVICER/MMS 1-g DEMONSTRATION 8-1

8.1 Servicer/MMS 1-g Demonstration Plan 8-1

8.1.1 Subsystem Requirements 8- 2

8.1.2 Torque Loads on ETU Drives for MMS Servicing Demonstrations 8- 9

8.1.3 Concept Definition of the MMS Module Mockup 8-37

8.1.4 Concept Definition of the Modified MST 8-49

8.1.5 Spacecraft Mockup Configuration Selection 8-50

8.1.6 Stowage Rack Mockup Configuration Selection 8-55

8.1.7 Servicer System Configuration - 1-g Demonstrations .... 8-65

8.1.8 Schedule and Cost Considerations 8-67

8.2 Servicer/MMS 1-g Demonstration Design 8-71

8.2.1 MMS Module Mockup 8-74

8.2.2 Stowage Rack Module Support 8-76

8.2.3 MMS Spacecraft Mockup 8-80

8.2.4 Connector Positioner Mechanism 8-81

8.2.5 MST Storage Rack 8-84

8.2.6 Optical Targets 8-86

8.2.7 Electrical Design 8-88

8.2.8 Parts and Materials Procurement 8-88

8.2.9 ETU/MST/MMS Interface Coordination 8-88

8.2.10 Design Coordination 8-89

8.3 Servicer/MMS 1-g Demonstration Equipment 8-90

8.4 Conclusions 8-96
CONTENTS (Cont.)

9.0 SERVICER CONTROL SOFTWARE - BASIC MODULES 9-1

9.1 Software Requirements 9- 6

9.1.1 System Requirements 9-12

9.1.2 Coordinate System Definition 9-17

9.1.3 Supervisory Mode 9-21

9.1.4 Supervisory Trajectory Hierarchy 9-24

9.1.5 Supervisory Operations and Displays 9-37

9.1.6 Trajectory Display Construction 9-44

9.1.7 Manual-Augmented Control Mode 9-48

9.1.8 Interface Definition 9-51

9.2 Software Preparation 9-53

9.2.1 Software Development 9-53

9.2.2 Software User's Manual 9-55

10.0 SERVICER CONTROL SOFTWARE - MMS MODULES 10- 1

10.1 Software Requirements 10- 3

10.1.1 System Requirements 10- 5

10.1.2 Supervisory Trajectory Hierarchy 10- 11

10.1.3 Supervisory Operations and Displays 10- 18

10.1.4 Interface Definition 10-21

10.2 Software Preparation 10- 23

10.2.1 Software Development 10- 23

10.2.2 Software User's Manual 10-24

11.0 SERVICER 1-g DEMONSTRATIONS 11-1

11.1 Basic Module Exchange Demonstration 11- 3

11.1.1 Basic Module Software Installation 11- 3

11.1.2 Basic Module Control System Checkout 11- 5

11.1.3 Basic Module Demonstrations 11- 9

11.2 MMS Module Exchange Demonstration 11- 10

11.2.1 MMS Module Exchange Equipment Installation and Checkout . . 11- 11

11.2.2 MMS Module Software Installation 11- 18

11.2.3 MMS Module Control System Checkout 11- 18

11.2.4 MMS Module Demonstrations 11- 29

12.0 CONCLUSIONS AND RECOMMENDATIONS 12-1

12.1 On-Orbit Servicing Development 12- 1

12.2 Multi-Mission Modular Spacecraft Servicing 12- 2

12.3 Ground Demonstrations 12- 3

12.4 Cargo-Bay Demonstration 12- 4

12.5 Free-Flight Verification 12- 6

12.6 Servicer/MMS 1-g Demonstration Plan 12- 7

12.7 Servicer/MMS 1-g Demonstration Equipment 12- 9

12.8 Servicer Control Software - Basic Modules 12- 10

12.9 Servicer Control Software - MMS Modules 12- 12

12.10 Servicer Software Demonstrations 12- 13

A Cargo-Bay Demonstration Work Breakdown Structure ..... A- 1

vi
Figures

1-1 IOSS On-Orbit Servicer Configuration 1- 4

1-2 Engineering Test Unit 1- 5

1-3 Study Task Flow Chart 1-15

1-4 On-Orbit Servicing Development Schedule 1-17

1-5 Multi-Mission Modular Spacecraft Mechanical System .... 1-18

1-6 MMS Module Exchange Using Straight Docking Probe

Adapter and Tool Adapter 1-19

1-7 Engineering Test Unit Adaptation for MMS Servicing .... 1-20

1-8 Ground Demonstrations Schedule 1-21

1-9 Artists Concept of Cargo-Bay Demonstration 1-22

1-10 Servicer Cargo-Bay Demonstration Schedule 1-24

1-11 The Operational Servicer with the OMV 1-25

1-12 Free-Flight Verification Program Schedule 1-27

1-13 Servicer System Configuration - 1-g Demonstrations .... 1-28

1-14 Connector Positioner Mechanism 1-30

1-15 MMS Module Mockup 1-31

1-16 Demonstration System Block Diagram 1-32

1-17 Representative MMS Software Menus 1-35

1-18 Module Servicing Tool for Ground Demonstrations of

MMS Module Exchange 1-36

1-19 Ground Demonstration of MMS Module Exchange 1-37

2-1 On-Orbit Servicer Configuration 2- 2

2.2-1 Serviceable Spacecraft Designs from the 70's 2- 9

2.2-2 Servicer Mechanism Working Volume 2-11

2.2-3 Engineering Test Unit 2-14

2.3-1 Study Task Flow Chart 2-17

3-1 Solar Maximum Mission Version of MMS 3-2

3-2 Multi-Mission Modular Spacecraft Mechanical System .... 3-3

3-3 MMS Module Structure with Module Retention System 3-4

3-4 MMS Module Exchange Using Straight Docking Probe

Adapter and Tool Adapter 3- 6

3-5 Docking and Berthing Pin Combined Design 3-7

3.2-1 Modified MST Configuration for Ground Demonstrations ... 3-15

3.2-2 MST Related Electrical Design Concepts 3-18

3.3-1 MMS Docking Probe Adapter 3-23

3.3-2 Revised MMS Docking Probe Adapter 3-24

3.3-3 Baseline MMS Servicing Configuration 3-26

3.3-4 Alternative MMS Servicing Configuration 3-27

3.3-5 Module Characteristics Histogram 3-35

3.4-1 Candidate Regions for Extension of Remote

Maintenance/Resupply 3-38

3.4-2 Candidate Uses of On-Orbit Servicer at Space Station ... 3-40

vii
Figures (Cont.)

4.2-1 The Engineering Test Unit of the IOSS 4- 3

4.2-2 Ground Demonstrations Schedule 4- 9

4.3-1 Artists Concept-of the Cargo-Bay Demonstration 4-10

4.3-2 Servicer Cargo-Bay Demonstration Schedule 4-16

4.4-1 The Operational Servicer with the OMV 4-19

4.4-2 Free-Flight Verification Program Schedule 4-23

5-1 The Engineering Test Unit of the IOSS 5-3

5-2 Ground Demonstrations Schedule 5-22

6-1 Cargo-Bay Development Program Plan Approach 6- 2

6.2-1 Cargo-Bay Demonstration Equipment 6-12

6.2-2 Cargo-Bay Demonstration Hardware Breakdown 6-13

6.2-3 Representative Servicer Mechanism Configuration 6-15

6.2-4 Stowage Rack Configuration 6-17

6.2-5 Module Support Structure 6-18

6.2-6 0-g Spacecraft Mockup Arrangement 6-20

6.3-1 Cargo-Bay Equipment Arrangement 2 6-27

6.3-2 Cargo-Bay Equipment Arrangement 3 6-29

6.3-3 Cargo-Bay Equipment Arrangement 9 6-31

6.3-4 Cargo-Bay Equipment Arrangement 11 6-33

6.3-5 Cargo-Bay Equipment Arrangement 16 6-35

6.3-6 Cargo-Bay Equipment Arrangement 2 Variation 6-36

6.4-1 Servicer Cargo-Bay Demonstration Schedule 6-45

7.2-1 Free-Flight Verification Equipment 7-11

7.2-2 The Operational Servicer with the OMV 7-13

7.3-1 Free-Flight Verification Program Schedule 7-17

8-1 Servicer End Effector/Modified MST Interface 8-7

8-2 Existing ETU Configuration and its Reference

Coordinates System 8-10

8-3 ETU Arm Configuration 8-12

8-4 End Effector Position in Cylindrical Coordinates 8-13

8-5 Cylindrical Coordinates 8-14

8-6 End Effector Attitude Definition 8-14

8-7 Reference Position of ETU (U = W = Y = Z = 0) 8-16

8-8 ETU Wrist Joints in the Reference Position (W =• Y = Z = 0). 8-17

8-9 Fastener Orientation Angle Alpha 8-19

8-10 Relationship between Z, Theta, and Alpha Angles 8-19

8-11 Y-Drive Position Relative to Modified MST and MMS Module . 8- 21

8-12 Allowable Modified MST Weight 8-23

8-13 Module Orientation Relative to the Arm for FLIP

Starting Position . .• 8-25

8-14 MMS Module Mockup and Modified MST Geometry 8-27

viii
Figures (Cont.)

8-15 MMS Module Mockup in Mid-FLIP Position 8-28

8-16 Alternative Locations for FLIP for MMS Servicing 8-29

8-17 Torque Load vs. Speed of the "Y" Drive 8-32

8-18 ETU Counterbalance System 8-33

8-19 "U" Drive Actuator Load - Cases 1 and 2 8-35

8-20 Load vs. Speed of the "U" Drive 8-36

8-21 MMS Module Structure with Module Retention System 8-37

8-22 MMS Module Outside Dimensions, Including the Insulation . . 8-38

8-23 Existing MMS Module Retention Fastener 8-39

8-24 Modified Module Attachment for 1-g Demonstration 8-40

8-25 Maximum Axial Force in the Fastener 8-41

8-26 Accidental Load Case 8-43

8-27 MMS Module Mockup Truss Structure 8-44

8-28 MMS Module Mockup with Torque Box Structure 8-46

8-29 MMS Module Mockup with Central Torque Box Structure .... 8-47

8-30 MMS Module Partial Mockup 8-49

8-31 Configuration 1) of the Spacecraft Mockup 8-52

8-32 Configuration 2) with Added MMS Mockup 8-53

8-33 Configuration 3) with MMS Mockup and Payload Section ... 8-54

8-34 Configuration 4) Interchangeable Mockups 8-55

8-35 Stowage Rack Alternative Arrangements 8-57

8-36 Trajectory Sequence for MMS Module Exchange 8-60

8-37 Variant 1 - Trajectory Sequence for Minimum MMS

Servicing Time 8-63

8-38 Recommended Stowage Rack Arrangement 8-64

8-39 Servicer System Configuration - 1-g Demonstrations .... 8-66

8-40 Change Order 3 Activity Schedule 8-70

8-41 Servicer/MMS Equipment Drawing Tree 8-73

8-42 MMS Module Assembly 8-74

8-43 Stowage Rack Module Support 8-77

8-44 Floating Nut Unit 8-78

8-45 Electrical Connector Receptacle Mockup 8-79

8-46 MMS Spacecraft Mockup 8-81

8-47 Connector Positioner Mechanism 8-82

8-48 MST Storage Rack 8-85

8-49 Optical Target Assembly 8-87

8-50 MMS Module Mockup Foam Structure 8-92

8-51 MMS Module Mockup, Front View 8-92

8-52 MMS Module Mockup, Partial Back View 8-93

8-53 MMS Spacecraft Mockup Assembly 8-93

8-54 MMS Module Support for Stowage Rack 8-94

8-55 MMS Module Mockup on Stowage Rack Support 8-95

8-56 Connector Positioner Mechanism 8-96

ix
Figures (Cont.)

9-1 Software Development Approach 9- 3

9-2 Demonstration System Block Diagram 9- 5

9.1-1 Servicer Servo Drive Console Interfaces 9-13

9.1-2 Mockup Configuration Layout 9-18

9.1-3 Cylindrical Coordinates 9-19

9.1-4 Servicer Mechanism Gimbal Coordinates 9-21

9.1-5 Analytical Block Diagram - Supervisory Mode 9-25

9.1-6 Analytical Block Diagram - Manual-Augmented Mode 9-50

11-1 ETU Configuration for Demonstration of Basic and

MMS Module Exchange 11- 2

11-2 Operator's Console Arrangement 11- 4

11-3 Engineering Test Unit Configuration for Ground

Demonstrations of Basic or MMS Module Exchange 11- 13

11-4 Stowage Rack Supports for "Temporary" Locations of

Basic and MMS Modules 11-15

11-5 Stowage Rack "Good" Locations for Basic and MMS Modules . . 11- 16

11-6 Module Servicing Tool for Ground Demonstrations 11- 17

11-7 ETU End Effector and Its Connector Positioner

Mated with MST 11-19

11-8 ETU End Effector Standing Off from MST 11-28

11-9 MST Probe Entering MMS Module Guide 11-31

A-l Cargo-Bay Demonstration Work Breakdown Structure A- 2
Tables

Page

1-1 Study and Program Objectives 1- 7

2.1-1 Study and Program Objectives 2- 4

3.3-1 MMS Module Servicing Method Alternatives 3-20

3.3-2 Servicer Configuration Evaluation Criteria 3-21

3.3-3 On-Orbit Servicer Characteristics 3-33

3.3-4 Candidate Alternative Modules . . 3-36

4.2-1 Candidate 1-g Demonstration Issues 4- 5

4.2-2 Recommended Activities for the Ground Demonstrations ... 4-6

4.2-3 Optional Activities for the Ground Demonstrations 4-7

4.2-4 Estimated Costs of the Ground Demonstrations 4-9

4.3-1 Candidate Cargo-Bay Demonstration Activities 4-12

4.3-2 Servicer Cargo-Bay Demonstration Characteristics 4-13

4.3-3 Cost Breakdown for the Cargo-Bay Demonstration 4-18

4.4-1 Free-Flight Verification Characteristics 4-22

4.4-2 Free-Flight Verification Cost Breakdown 4-24

5-1 Recommended Activities for the Ground Demonstrations ... 5-19

5-2 Optional Activities for the Ground Demonstrations 5-21

5-3 Estimated Costs of the Ground Demonstrations 5-24

6.1-1 Candidate Cargo-Bay Demonstration Activities 6-6

6.2-1 Space Transportation System Equipment 6-14

6.3-1 Candidate 0-g Demonstration Arrangements 6-25

6-3-2 Equally Applicable Cargo-Bay Demonstration Requirements . . 6-39

6.3-3 Alternative Arrangement Comparison 6-40

6.3-4 Alternative Arrangement Evaluation 6-42

6.5-1 Cost Breakdown for the Cargo-Bay Demonstration 6-49

7.1-1 Free-Flight Verification Activities 7-4

7.1-2 Free-Flight Verification Characteristics 7-9

7.2-1 Free-Flight Verification Equipment 7-12

7.2-2 Space Transportation System Equipment 7-12

7.4-1 Free-Flight Verification Cost Breakdown 7-20

8-1 Weight Data for "Y" Drive Torque Load Calculation 8-31

8-2 Coordinates and Joint Angles 8-59

8-3 Times Necessary to Perform the Six Steps 8-62

8-4 Change Order 3 Activities 8-67

8-5 Change Order 3 Deliverable Mockups 8-68

8-6 Assembly Drawings for Servicer/MMS 1-g Demonstrations ... 8-72

8-7 Change Order 3 Equipment List 8-90

xi
Tables (Cont.)

9.1-1 Software Components 9- 6

9.1-2 Geometrical Approach 9- 7

9.1-3 Associated Functions 9-11

9.1-4 Simulation Initialization 9-38

9.1-5 Total Trajectory Initialization 9-41

9.1-6 Shut-Down Procedure 9-42

9.1-7 Example Radial Total Trajectory Display 9-45

9.1-8 Example Axial Total Trajectory Display 9-47

10.1-1 Total Trajectory Initialization 10- 20

10.1-2 Shut-Down Procedure 10- 21

11.1-1 Basic Module Demonstrations 11- 9

11.2-1 MMS Module Exchange Equipment 11- 11

xii
1.0 SUMMARY

An upgraded spacecraft servicing demonstration plan, based on a

preliminary plan from contract NAS8-35496, has been prepared that leads

to a fully verified operational on-orbit servicing system based on the

module exchange, refueling, and resupply technologies by late 1992.

The resulting system can be applied at the Space Station, in low earth

orbit with an Orbital Maneuvering Vehicle (OMV), or be carried with an

OMV to geosynchronous orbit by an Orbital Transfer Vehicle (OTV). The

three recommended overlapping phases are:

1) Ground demonstrations using the MSEC Robotics Laboratory;

2) Cargo-bay demonstrations in the Orbiter using the Remote

Manipulator System to dock a Multi-Mission Modular Spacecraft (MMS)

mock up, carried to and from orbit on the MMS Flight Support System,

to the servicer and spare module stowage rack. Two forms of module

exchange and fluid resupply are recommended for demonstration on a

single Orbiter flight;

3) Free-flight verification using the OMV as the carrier vehicle and a

rented spacecraft bus to carry the MMS serviceable spacecraft

mockup.

The plan emphasizes the exchange of Multi-Mission Modular Spacecraft

modules as the MMS is a significant ongoing program involving

space-repairable satellites.

Three servicer mechanism configurations are included in the plan:

1) The Engineering Test Unit currently in use at MSFC would be used

for early ground demonstrations, procedures development, and

training;

2) A proto-flight quality unit would be used for the demonstration

flight in the Orbiter cargo bay and subsequently for ground

demonstrations, procedures development, and training;

1-1
3) One fully operational unit that has been qualified and documented

would be used in the free-flight verification activity.

The plan balances costs and risks by overlapping study phases,

utilizing existing equipment for the ground demonstrations, maximizing

use of existing MMS equipment, taking advantage of the ongoing NASA-JSC

orbital refueling program, and rental of a spacecraft bus rather than

building a new unit for a one-time use in the free-flight

verifications. The preliminary funding estimate is $1.0M for the

ground demonstrations, $9.3M for the cargo-bay demonstrations, $35M for

the free-flight verifications, and a total of $45.3M in 1985 dollars.

The plan must be significant and long-term to encourage users and

spacecraft designers to include on-orbit servicing in the form of

module exchange in their plans.

The capability development portion of the study effort had two parts -

software development and provision of MMS module exchange demonstration

mockup equipment. Software was developed, documented and successfully

demonstrated for the separate exchange of basic (24 in. cube) modules

and MMS (48 in. square by 20.5 in. deep) modules. Exchange of each

type of module was demonstrated in three control modes.

A plan for the demonstration of the exchange of MMS modules using the

servicer mechanism Engineering Test Unit (ETU) was prepared and

executed. The plan included: (1) establishment of requirements, (2)

conceptual design, (3) selection of MMS spacecraft mockup

configuration, (4) selection of MMS module mockup configuration, (5)

evaluation of adequacy of ETU load capability, and (6) selection of a

stowage rack arrangement.

The MMS module exchange demonstration mockup equipment was designed,

fabricated, checked out, shipped, installed, and demonstrated in the

MSFC Robotics Laboratory.

1-2
1.1 INTRODUCTION

Many studies and demonstrations during the past decade have clearly

proven the overwhelming cost effectiveness benefits of an unmanned

on-orbit satellite servicing capability. The ability to change out

failed or worn-out satellite modules and to replenish fuels and other

expendable commodities offers satellite programs a greatly reduced

operating cost when compared with replacement of an entire satellite.

Development activities that will eventually lead to routine orbital

servicing operations were initiated in the early 1970's. Several

alternative servicing systems, including satellite modules and component

design approaches, were defined and evaluated during this period.

With the Space Transportation System now operational, the capability

exists to deliver and retrieve an operational servicer system. It was

thus appropriate to initiate in 1983 the planning that will lead

directly to the operational servicing capability.

Since the early 1970's various alternatives for satellite maintenance

have been identified, conceptualized, and evaluated—unmanned orbital

servicing systems, manned extravehicular activities, highly reliable

expendable designs, and retrieval for ground refurbishment and return to

orbit. The first Integrated Orbital Servicing System (IOSS) study

completed in September 1975 along with a parallel study, Integrated

Orbital Servicing and Payloads Study, conducted by COMSAT Laboratories

of the Communications Satellite Corporation, jointly concluded:

1) On-orbit servicing is the most cost-effective satellite maintenance

approach;

2) Development of a single on-orbit servicer maintenance system is

compatible with many spacecraft programs;

3) Spacecraft can be designed to be serviceable with acceptable design,

weight, volume, and cost effects;

1-3
4) The evolving Space Transportation. System (STS) is designed to

support on-orbit maintenance;

5) Users need guarantees that servicing will be available and

assurances that it will be cost effective.

As satellite designs continue to evolve and the Space Station era

approaches, it becomes apparent that there is room for virtually all

the alternatives of satellite maintenance at one point or other in the

future. However, to minimize servicer system development costs, the

IOSS follow-on study, completed in June 1978, recommended that a single

servicer system having the capability to accommodate both low earth and

geosynchronous orbit applications should be evolved. This requirement

has been satisfied effectively by the servicer mechanization (Fig. 1-1)

conceptualized during the IOSS studies. The single design is

compatible with maintenance of most, spacecraft of the Space

Transportation System era. Adapters may be used to accommodate support

structure differences across the applications.

MODULE

STOWAGE RACK

END EFFECTOR

AND WRIST DRIVES

SERVICER STOWAGE/

DEPLOYMENT MECHANI

OMV OR ORBITER

INTERFACE

ELBOW ROLL DRIVE

DOCKING

MECHANISM

SHOULDER DRIVES

TEMPORARY ORU

STOWAGE LOCATION

INTERFACE

MECHANISMS

Figure 1-1 IOSS On-Orbit Servicer Configuration

1-4
ORIGINAL PAGE IS

OF POOR QUAUTY

An effective interface between each orbital replaceable unit (ORU), or

module, and the spacecraft and the servicer was defined and

breadboarded. The interface mechanism provides a logical and cost

effective method of incorporating orbital replaceable units (ORU) for

module exchange in all spacecraft.

The value of demonstrations in furthering on-orbit servicing

development was recognized in the decision to build a 1-g version of

the Integrated Orbital Servicing System of Figure 1-1. The result is

the Engineering Test Unit (ETU) of the IOSS shown in the photograph of

Figure 1-2. This unit was built and delivered to MSFC in 1978. It has

been used for over 350 demonstrations during the intervening seven

years.

Figure 1-2 Engineering Test Unit

1-5
Considerable interest in spacecraft maintenance was expressed by both

the Department of Defense and the commercial sector, however, the

general tenor of their support was that a demonstration of orbital

maintenance must be conducted prior to any commitment on their part. A

flight demonstration of the all-up maintenance capability is also a

NASA requirement prior to wholesale commitment to the concept.

However, a reduced capability test that exercises the basic concept and

exchanger capability can and should be demonstrated prior to the time

that a full capability will exist. With this background material in

hand, and with renewed interest by the space flight community, it was

appropriate to perform the prior study (Contract NAS8-35496) that

defined a path leading to demonstration of the servicing capability.

The cargo-bay demonstration part of the development plan from the prior

study was felt to be too expensive so it has been extensively revised

in the current study. The 1-g part of the development plan was found

acceptable and the portions of it having to do with basic and MMS

module software development and with the preparation of mockup

equipment for the demonstration of MMS module exchange were performed

as part of the subject contract activity. This software and mockup

equipment activity led to a series of successful 1-g module exchange

demonstrations.

1.2 STUDY OBJECTIVES

The objectives of this Servicer System Demonstration Plan and

Capability Development study are to identify all major elements and

characteristics of an on-orbit servicing development program and to

integrate them into a coherent set of demonstrations, to upgrade the

Engineering Test Unit control system for basic and MMS module exchange

demonstrations, and to upgrade the MSFC 1-g servicing demonstration

facility mockups to permit the exchange of MMS modules. These

objectives, along with the program objectives, are summarized in Table

1-1. The on-orbit servicing development plan was to be a revision of

the plan prepared during the prior study with increased emphasis on low

cost and use of MMS equipment. The revisions primarily addressed the

cargo-bay demonstrations.

1-6
Table 1-1 Study and Program Objectives

Study Objectives

To identify and integrate the major characteristics of an on-orbit

servicing demonstration program plan.

To upgrade the engineering test unit control system for MMS and basic

module exchange demonstrations.

To upgrade the 1-g demonstration facility to permit exchange of MMS

modules.

Program Objectives

Fully verified and documented operational on-orbit servicing system

a) Based on module exchange and fluid resupply technologies,

b) Suitable for use with Space Station.

Major issue is balance between the number and complexity of

development activities and cost.

The goal of the development program is a fully verified operational

on-orbit servicing system based on the module exchange and fluid

resupply technologies that is also suitable for use with and at the

Space Station. The plan must be significant and long-term to encourage

users and spacecraft designers to include on-orbit servicing in the

form of module exchange in their plans.

The second study objective involves the development of two software

programs - one for the exchange of basic modules and one for the

exchange of MMS modules - and the demonstration of the exchange of both

types of modules. The demonstrations were to be performed for three

different control modes which are:

1) Supervisory with minimal operator assistance;

2) Supervisory with operator assistance at each action;

3) Manual-Augmented.

1-7
The third study objective involved the design, fabrication, and

installation of MMS demonstration equipment. The Martin Marietta

provided equipment included:

1) Two MMS module mockups;

2) One spacecraft mounted module receptacle;

3) Two stowage rack mounted module receptacles;

4) A connector positioner drive;

5) An MST storage rack;

6) A set of MMS module targets;

7) A set of related wiring.

A light weight form of the Module Servicing Tool, which was adapted to

the Engineering Test Unit and modified for remote location of its

control system was provided by Goddard Space Flight Center. The

mechanical and electrical design was performed by Fairchild Space

Company, the mechanical equipment was built by GSFC, and the electrical

equipment was built by Fairchild.

The combination of the second and third study objectives amounted to

the goal of adapting the on-orbit servicer ETU to exchange MMS modules

and conducting successful demonstrations in 1-g. The challenge of this

goal was accepted and accomplished.

The first of two key study issues was the need to balance the number

and complexity of development activities against available funds. The

proposed approach, recommended in the Spacecraft Servicing

Demonstration Plan (SSDP) study, is to lay out a program with most of

the desired features, that overlaps the 1-g, 0-g, and operational

servicer demonstrations, and attempts to get an early operational

1-8
capability. It minimizes costs by taking advantage of parallel

activities such as the JSC refueling program, and advocates renting a

spacecraft bus rather than buying a new one. The program was also

scoped large enough to become a recognized part of NASA's long-range

plans. The promise of a clear plan by NASA to develop and use module

exchange for many years will encourage the user, or spacecraft

designer, to incorporate module exchange in his plans.

The second key study issue was the need to maintain a close working

relationship between MSFC and Martin Marietta personnel during servicer

control software development. A number of interfaces were defined so

both organizations could work towards the same goal:

1) Computer and interface electronics operations;

2) Functions to and from the Servicer Servo Drive Console and the ETU;

3) Functions to and from the control station for Manual-Augmented

control mode implementation.

The close working relationship between MSFC and Martin Marietta

personnel was particularly effective during the installation of the MMS

module exchange equipment and rework of the spacecraft mockup. The

cooperation and assistance of MSFC personnel, especially the Contract

Technical Monitor, Mr. James Turner and Messers. Tom Bryan and Don

Scott of the Robotics Laboratory, in obtaining needed materials and

performing the installation and rework resulted in the effort being

completed early and with better results than had been originally

planned. Their efforts are greatly appreciated.

The servicer system development plan was prepared to provide

implementors and users with a single development approach that will

culminate in orbital servicing operations. The plan is necessary

because only by providing a planned development program will both

development and user support be focused on the servicing issue.

Current planning for the Orbital Maneuvering Vehicle is such that

1-9
servlcer development must be started soon if a servicing capability is

to exist shortly after the OMV reaches an operational status.

Verification of a servicing capability with the OMV will result in a

well-proven system being available for potential use with the Space

Station. Many prior and current studies have addressed individual

elements of servicing. Many tools and support hardware elements have

been defined that will aid a future servicing program. These efforts,

however, have not resulted in a general move on the part of the user

community to incorporate serviceability in the form of module exchange

into their spacecraft designs. It is only through the implementation

of a development program that produces a demonstrated on-orbit

servicing capability that the benefits of this program will be realized

in future spacecraft operations. The upgraded development program plan

described in this report was prepared to satisfy this need.

1.3 RELATIONSHIP TO OTHER NASA EFFORTS

Prior and ongoing NASA activities, as well as future plans, in the area

of satellite servicing are discussed in relation to the objectives and

approach of this servicer system demonstration plan and capability

development study.

Servicing development activities were initiated in the early 1970's and

continue through the present time. Studies and development work have

been performed by NASA, other government agencies, and contractors.

Early study results concluded that on-orbit servicing was a more cost

effective approach than ground refurbishment of satellites.

Recommendations included that spacecraft be designed for servicing and

that module exchange was the most cost-effective method of servicing.

During the Integrated Orbital Servicing System study an Engineering

Test Unit was designed and built and has been in use at MSFC since 1978

for ground demonstrations of remote satellite servicing and other

development activities. A wealth of experimental data was accumulated

during that servicer demonstration and development program and

constitutes the basis for further development of an on-orbit satellite

servicing capability. -

1-10
As the Space Transportation System is operational, satellites in low

earth orbit are accessible for on-orbit maintenance and repair. Many

NASA efforts are now directed towards definition of the requirements,

interfaces and programmatic aspects of the three main approaches to

satellite servicing: (1) manned, using extravehicular activities, (2)

remote servicing, using a simple specialized mechanism for module

exchange, refueling, and resupply and controlled in manual and

automated modes, and (3) remote servicing operations using telepresence

technology and artificial intelligence.

EVA satellite servicing participated in a successful demonstration

during the Solar Maximum Repair Mission when equipment modules were

exchanged on a Multi-Mission Modular Spacecraft utilizing the Orblter

Remote Manipulator System (RMS), the Manned Maneuvering Unit (MMU) and

a module servicing tool (MST). Many tools and auxiliary devices have

been developed for use by the Shuttle or Space Station EVA crews to

perform various servicing tasks. The accumulated EVA experience

emphasizes the need for simple, easy maintenance and repair tasks,

ample clearances to accommodate the rather bulky EVA suit, and

provision for handrails and foot restraint brackets. Due to EVA time

and space limitations and the high cost and risk involved, basellning

EVA for maintenance, repair and refueling/resupply of spacecraft needs

to be determined by the user on an individual basis. Because of man's

direct involvement in the operations, the safety aspects are

particularly important and difficult to resolve. However, EVA remains

the main back-up system for repair in contingency situations at the

Orbiter and Space Station, due to its superior flexibility and ability

to perform unscheduled and unplanned repair operations.

An Orbital Maneuvering Vehicle is being developed by NASA-MSFC, with

the participation of other NASA centers, to supplement the STS for

satellite delivery, retrieval and on-orbit servicing. It will utilize

the Orbiter for launch and will have applications in both low earth

orbits (LEO) and geostationary earth orbit (GEO), when transported to

GEO by an Orbital Transfer Vehicle (OTV) or other orbit transfer

stage. Early availability of the OMV as a reusable vehicle will

1-11
obviate the necessity of including integral propulsion in many new

space initiatives for satellite deployment or retrieval. The OMV will

have a man-in-the-loop control capability from a ground control station

(6CS). Rendezvous and docking capability and an OMV compatible

servicer kit can be developed in subsequent phases to add satellite

retrieval and on-orbit servicing capabilities. The incongruity between

desirable polar orbits and the STS capability can be eased by use of an

OMV.

Servicing functions and approaches are being investigated by NASA and

its contractors in connection with Space Station operations.

Maintenance and repair missions are being evaluated for the Space

Station. For the proximity operations an RMS may be used, with manual

control from a special servicing platform. For LEO satellite

deployment and retrieval, the OMV will be used. In situ satellite

servicing at LEO can be performed using an OMV and a servicer from the

Space Station. Similar operations at GEO can use an OTV from the Space

Station to deploy and retrieve the OMV and the servicer. The control

of the servicer can be from the Space Station or from the ground.

Operating the OMV/servicer or OTV/OMV/servicer from the Space Station

can provide better availability of servicing and can reduce launch

costs.

The promise of advanced automation, including capabilities such as

telepresence and artificial intelligence, is being examined by the MSFC

Space Station project as part of a "smart front end" for the OMV. The

requirements for the smart front end include those functions that the

10SS approach can do, and much more. However, it is also possible to

refine the 10SS design so that the basic mechanism can also be used on

a smart front end. The major differences in the two approaches are in

the sensors used and complexity of the control systems. It thus seems

that the IOSS concepts should be considered for inclusion in the

development process leading to a Space Station applicable smart front

end for the OMV.

1-12
The Astro Electronics part of the RCA Government Systems Division has

contracted with Martin Marietta to help RCA investigate the application

of IOSS concepts to Polar Platform configurations. The information

provided by Martin Marietta is to help RCA decide on a servicing system

approach for this element of the Space Station program.

Many studies during the past decade proved the cost benefits of

on-orbit fluid resupply. The areas of fluid management requiring new

technology have been identified. Cargo-bay experiments are now planned

by NASA-JSC to demonstrate fluid transfer in 0-g and to test new quick

disconnects and sensors. For these first experiments, EVA operations

are planned. Safety aspects are of prime concern. Standardization of

the fluid resupply interface is an important issue affecting the

economics and ultimately the success of the satellite fluid resupply

activities. An interface standardization project is being pursued by

NASA-JSC. The objective is to develop a standard propellant servicing

interface for all satellites. A committee will be formed consisting of

appropriate NASA elements, the DoD and those industrial firms active in

the design and fabrication of satellite propulsion stages. This

committee will define the fluid interconnects, mechanical attachment

hardware, isolation philosophy, data format requirements, and

instrumentation and control interfaces consistent with safety

requirements and minimization of crew time lines. The program

objectives are to develop and certify a standardized disconnect design

for on-orbit resupply of earth storable, gaseous and cryogenic fluids

and to provide earth storable fluid disconnect flight hardware for the

Gamma Ray Observatory by March 1986.

The prior study, Spacecraft Servicing Demonstration Plan, made use of

the experience accumulated during the IOSS demonstrations and expanded

its scope to encompass demonstrations of Multi-Mission Modular Spacecraft servicing, other module and component exchange, and refueling

demonstrations utilizing the present state of the art technology.

Timing of various planned activities was such that it could take

advantage of the results of the NASA-JSC refueling development effort

and match the milestones of the OMV development program schedule.

1-13
A simple, proven servicer mechanism, with a standardized end effector

interface and supplemented by specialized adapters and interface

mechanisms, like the IOSS, can be built today with the present

technology. It can provide a much needed satellite servicing

capability now and the ability to test and develop the elements of

future generation servicers.

1.4 STUDY APPROACH

Our approach to the proposed study was to use the four tasks identified

in the contract statement of work. These are:

1) Task 1 - Servicer Development Program Plan;

2) Task 2 - Servicer Control Software;

3) Task 3 - Servicer Demonstration;

4) Task 4 - Program Management.

Figure 1-3 shows an overall logic flow for the four study tasks. Other

than the program management task, the work divides naturally into two

parts - preparation of the Servicer Development Program Plan (Task 1)

and generation of the servicer control software as well as conducting

servicer demonstrations at MSFC (Tasks 2 and 3). The MMS 1-g servicing

demonstration definition effort of Change Order 1 and the MMS 1-g

demonstration equipment drawing, fabrication, checkout, and

installation effort of Change Order 3 were included in Task 1. The MMS

module software requirements, programming, and user's manual

preparation effort of Change Order 3 were included in Task 2, while the

software installation and MMS module exchange activity of Change Order

3 were included in Task 3. MMS module exchange demonstrations required

the availability of a GSFC MMS Module Servicing Tool designed for use

in 1-g with the MSFC Engineering Test Unit. A more detailed

description of the approach to each of the four tasks of this study is

presented in Section 2.3.

1-14
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1-15
1.5 SIGNIFICANT CONCLUSIONS AND RECOMMENDATIONS

The significant conclusions and recommendations from this Servicer

System Demonstration Plan and Capability Development activity are

presented below. Many secondary conclusions and recommendations are

given in Sections 3.0 through 11.0. The conclusions and

recommendations which span the study are given first.

1.5.1 On-Orbit Servicing Development

The following conclusions and recommendations apply to the overall

on-orbit servicing development:

1) The recommended plan leads to the free-flight verification of an

operational servicer suitable for use with the OMV and the Space

Station;

2) The plan has three phases

- Ground demonstrations,

- Cargo-bay demonstration,

- Free-flight verification;

3) The free-flight verification can be completed by late 1992 (see

Figure 1-4);

4) The total estimated cost is 45.3 million 1985 dollars;

5) The plan includes three servicer mechanism configurations:

- The Engineering Test Unit currently in use at MSFC would be used

for early ground demonstrations, procedures development, and

training for the cargo-bay demonstration,

- A proto-flight quality unit would be used for the demonstration

flight in the Orbiter cargo bay and for procedures development

and training related to the operational servicer,

1-16
DESCRIPTION

OMV Design &

Development

OMV Operations

OMV Supporting

Development

Ground

Demonstrations

Cargo-Bay

Demonstration

Free-Flight

Verification

1985

PHASE i

1986

ATP PO

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R

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1987

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1988 1989

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1990

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3RD F

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1993

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Figure 1-4 On-Orbit Servicing Development Schedule

- One fully operational unit that has been qualified and

documented for use in the free-flight verification activity and

in subsequent operations;

6) The plan is based on use of proven IOSS designs and test hardware;

7) Areas for application of the module exchange form of on-orbit

servicing to the Space Station were identified.

1.5.2 Multi-Mission Modular Spacecraft Servicing

The following conclusions and recommendations apply to the involvement

of MMS equipment in the demonstration plan and in subsequent operations:

1) Primary emphasis would be on demonstrating the exchange of MMS

modules (see Figure 1-5);

2) The MMS Module Servicing Tool should be adapted to work with the

servicer end effector for the exchange of MMS modules;

3) A set of requirements for the MST adaptation was prepared;

1-17
ACS MODULE

TRANSITION

ADAPTER

PM-IA AUXILIARY

TANK KIT (ATK)

POWER

MODULE

FLUID

INTERCONNECT

C&DH MODULE

PM-I

PROPULSION

MODULE

Figure 1-5 Multi-Mission Modular Spacecraft Mechanical System

4) Light weight MMS module mockups with dimensionally correct standard

MMS attachment fixtures and connector shells should be used for

ground demonstrations;

5) On-orbit servicing of MMS modules should be effected by use of

lateral docking with a straight docking probe adapter, tool adapter

and modified stowage rack (see Figure 1-6).

1.5.3 Ground Demonstrations

The following conclusions and recommendations were developed during the

ground demonstration analyses:

1) The servicer system Engineering Test Unit, shown in Figure 1-2,

should be used as the mechanism for early ground demonstrations;

2) Continue the ability to demonstrate separately the exchange of both

basic and MMS modules;

1-18
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1-19
ORIGINAL PAGE f5

OF POOR QUALITY

3) The control system software of the MSFC servicing demonstration

facility has been upgraded;

4) MMS module exchange under computer control has been demonstrated

(see Figure 1-7) ;

5) Control mode analysis and testing for exchange of both module types

should be continued;

6) Approaches for the cargo-bay demonstration and for free-flight

verification should be developed;

7) Fluid resupply hardware should be developed and the process

demonstrated;

Figure 1-7 Engineering Test Unit Adaptation for MMS Servicing

1-20
8) The exchange of batteries or other individual components should be

demonstrated along with thermal blanket/access cover removal and

replacement;

9) An automatic target recognition and error correction system should

be developed and demonstrated;

10) The MSFC servicing demonstration facility should be made available

for support of flight operations in terms of simulations,

procedures development, training, and problem solving. The

facility should also be made available as a laboratory development

tool;

11) The first five ground demonstration activities can be accomplished

by late 1986 (Figure 1-8) if they are funded in time.

DESCRIPTION

OMV Design &

Development

Multi-Mission Modular

Spacecraft Modifications

(Spacecraft Mockup and

Stowage Rack)

Module Exchange

Control Mode Analysis/

Testing

Battery Exchange

Demonstration

Resupply Demonstration

Automatic Target

Recognition

Support for Cargo-Bay

Demonstration

Cargo-Bay Demonstration

1985

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1986

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FLIGHT „

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Figure 1-8 Ground Demonstrations Schedule

1-21
1.5.4 Cargo-Bay Demonstration

The following conclusions and recommendations were developed during the

cargo-bay demonstration analyses:

1) A proto-flight quality servicer mechanism should be built for use

in the single cargo-bay demonstration flight;

2) The MMS Flight Support System should be used to support the MMS

spacecraft representation during the cargo-bay demonstration;

3) The Orbiter Remote Manipulator System end effector should be used

for a docking system;

4) A specific arrangement of servicing demonstration elements in the

Orbiter cargo bay was selected and recommended for use (see Figure

1-9);

Figure 1-9 Artists Concept of the Cargo-Bay Demonstration

1-22
5) The characteristics of the recommended servicer cargo-bay

demonstration are:

- MMS mockup dock and undock by RMS,

- Supply of power, attitude control, thermal control and

communications by Orbiter,

- Servicer control station in Orbiter,

- Docking rigidization by servicer docking probe,

- Electrical connection between servicer and spacecraft via the

docking mechanism,

- Use of MMS triangular module support structure,

- Module exchange demonstration,

- Fluid resupply demonstration,

- Servicing equipment performance demonstration,

- Unassisted Supervisory control mode,

- Man-machine interaction evaluations,

- Compliance with Orbiter system safety requirements,

- Servicer spare module stowage rack mounted in trunnions in

Orbiter cargo bay,

- Use of representative servicing operational equipment,

- Operator training;

6) The hardware for the fluid resupply demonstrations should be

obtained from the ongoing Johnson Space Center refueling

demonstration flight program;

7) The first cargo-bay demonstration flight can be completed by late

1988 (Figure 1-10);

8) The recommended activities for the test flight are:

- The replacement of a Multi-Mission Modular Spacecraft type

module using an MMS Module Servicing Tool, incorporating an

electrical connector, and mounted so that the module moves

axially,

1-23
DESCRIPTION

OMV Design &

Development

Ground Demonstrations

System Engineering

Design & Development

Safety

Software Dev & C/O

Procurement

Fabrication

Assembly & C/O

Acceptance Testing

System Integration &

Demonstration

Verification Phase B

1985

"i

PHASE B

1986

ATP_

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1987

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REVIEW PH

rrRATTONS

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Figure 1-10 Servlcer Cargo-Bay Demonstration Schedule

- The replacement of a battery module on a light weight side

interface mechanism using an electrical connector and with a

near-radial module motion direction,

- The transfer of a fluid using a multiple line fluid resupply

module including a fluid interface unit and a hose and cable

management device mounted in a far-axial direction;

9) The cargo-bay demonstration servicer mechanism, after its flight

use, should be used to replace the ETU for ground demonstrations,

procedures development, and operator training.

1-24
1.5.5 Free-Flight Verification

The following conclusions and recommendations were developed during the

free-flight verification analyses:

1) A fully operational servicer system (Figure 1-11) that has been

qualified and documented should be built for use in the free-flight

verification activity;

2) One servicer system should be built;

A

Figure 1-11 The Operational Servicer with the OMV

1-25
3) The unassisted Supervisory control mode should be used;

4) A spacecraft bus, such as the SPAS-01, should be rented rather than

a new spacecraft being built for this one-time application;

5) The characteristics of the recommended servicer free-flight

verification are:

- One verification flight,

- Serviceable satellite mockup supported by a rented spacecraft

bus,

- Supply of power, attitude control, communications, and thermal

protection and control of the servicer from the OMV,

- Use of OMV for rendezvous and docking of servicer to the

serviceable spacecraft mockup,

- Use of serviceable spacecraft mockup and modules from cargo-bay

demonstration,

- Two way communication links to ground through TDRSS,

- Servicer control station at OMV ground control station,

- Docking rigidization by servicer docking probe,

- Deployment of stowed servicer mechanism and docking probe,

- MMS module exchange demonstration,

- Fluid resupply demonstration,

- Servicing equipment performance verification,

- Control mode verification,

- Operator training;

6) The recommended flight verification activities are:

- Exchange of MMS module,

- Exchange of other representative modules,

- Fluid transfer;

7) The free-flight verification of an operational servicer can be

completed by late 1992 (Figure 1-12).

1-26
DESCRIPTION

OMV Design &

Development

OMV Operations

Cargo-Bay

Demonstration

Definiion Study

Design & Development

Software Dev & C/O

Tooling & GSE

System Test Equipment

Control Station

Flight Unit & ASE

System Qualification

System Integration &

Verification

1988

f

\

1989

CIR

r-7

V

PHASE C/D

DEMONSTRATION

7 I FLIGHT

I

A

PHASE B *

1990

FIRST FU

V|

1991

GHT

2ND FLIGHT

I V

n* TOR

7 r?

i

DEVELOPMENT

1

1

1

1

COR C

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1992

3RD FLIGHT

f— 7

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FaGHT ,_,

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Figure 1-12 Free-Flight Verification Program Schedule

1.5.6 Servicer/MMS 1-g Demonstration Plan

The following conclusions and recommendations were developed during the

preparation of the servicer/MMS 1-g demonstration plan:

1) The servicer/MMS 1-g demonstration subsystem requirements were

identified for the MMS module mockup, spacecraft mockup, stowage

rack mockup, electrical connector positioner mechanism, and optical

targets;

2) A preliminary system concept design was performed and the relative

positions of the main components were established as shown in

Figure 1-13;

1-27
SPACECRAFT

MOCKUP

SUPPORT

STRUCTURE

rMODIFIED MST

STORED

24 IN. CUBE MODULE

MOCKUPS

STOWAGE

RACK

MOCKUP

Figure 1-13 Servlcer System Configuration - 1-g Demonstrations

3) Several characteristics of the servicer/MMS demonstration equipment

were selected:

- MMS bolt tightening torque of 10 + 1 ft-lbs and loosening torque

of 20 + 1 ft-lbs,

- Maximum torque of 50 ft-lbs for the wrist pitch (Y) drive of ETU,

- Maximum weight of 12.5 Ibs for MMS module mockup,

- Maximum distance of 7.25 in. between the end effector interface

and module latch interface,

- Maximum weight of 15 Ibs for the modified MST;

4) A light weight configuration and a structural concept were selected

for the MMS module mockup;

1-28
5) A simple, straightforward configuration was selected for the

spacecraft mockup, that emphasizes the MMS module while providing

realistic MMS servicing trajectories and preserving the existing

basic module exchange capability;

6) The arangement of the MMS module mockups, basic module mockups and

MST storage rack on the ETU stowage rack was selected based on:

- Minimum modification of the existing stowage rack,

- Minimum MMS servicing demonstration time,

- No system reconfiguration between MMS module and basic module

exchange demonstrations.

1.5.7 Servicer/MMS 1-g Demonstration Equipment

The following conclusions and recommendations were developed as part of

the servicer/MMS demonstration equipment design and fabrication

activities:

1) The design effort included:

- Drawing preparation,

- Coordination of MST integration,

- Design coordination,

- Materials and components procurement;

2) The connector positioner mechanism (see Figure 1-14) features:

- A compact, eccentric type mechanism,

- Accurate linear ball slide,

- 5/8 in. mating stroke,

- 20 Ib connector mating/demating force,

- Adjustable position for end of stroke,

- Simple interface with ETU end effector;

1-29
3) The optical target features:

- Common design for all MMS fastener locations and for the MST

storage rack,

- Compliant attachment to its support,

- Minimal resetting in case of accidental displacement;

Figure 1-14 Connector Positioner Mechanism

4) The weight of the fabricated and assembled MMS module mockup

(Figure 1-15) is 10.0 Ibs, compared to the 12.5 Ibs maximum design

limit;

5) The fabricated and assembled connector positioner mechanism:

- Was tested on a special bracket, prior to shipment to MSFC,

Smoothly mated and demated with the electrical connector,

- The mating and demating times were within the design goals.

1-30
PAGE ft

OF POOR QUALITY

Figure 1-15 MMS Module Mockup

1.5.8 Servicer Control Software - Basic Modules

The following conclusions and recommendations were identified during

development of the servicer control software for the demonstration of

basic module exchange:

1) Three control modes were implemented;

2) Software requirements were explicitly defined and documented;

3) All required interfaces between the computer and the electrical

equipment were defined and documented (see Figure 1-16);

1-31
6

0

(0

• HQuoS<u 4JCOXCOcoCOcOE0) QvO60 •H

Figure 1-16

1-32
4) The characterlcti.es of the Supervisory control mode trajectory

hierarchy for basic modules are:

- Four total trajectories,

- Twenty trajectories,

- Nine steps,

- Eight actions,

- Each hierarchy level is composed of elements below it in the

hierarchy,

- Four types of coordinate transformations,

- Closed loop operation of ETU joints,

- Control of end effector and interface mechanism drives,

- Operator assisted and unassisted modes;

5) Software program is menu driven;

6) Procedures and trajectory sequences for the Manual-Augmented

control mode were documented;

7) Simulated hardware characteristics are included in software so

program can be run independent of servicer hardware;

8) A test program for verifying the computer to servicer hardware

interfaces was provided;

9) A separate Software User's Manual was prepared for the basic module

software.

1.5.9 Servicer Control Software - MMS Modules

The following conclusions and recommendations were identified during

development of the servicer control software for the demonstration of

MMS module exchange:

1) The MMS module software follows the basic patterns and philosophy

of the basic module software;

1-33
2) Three control modes were Implemented;

3) Software requirements were explicitly defined and documented;

4) All required interfaces between the computer and the electrical

equipment were defined and documented;

5) The characteristics of the Supervisory control mode trajectory

hierarchy for MMS modules are:

- One total trajectory,

- Nine trajectories,

- Thirteen steps,

- Ten actions,

- Each hierarchy level is composed of elements below it in the

hierarchy,

- Four types of coordinate transformations,

- Closed loop operation of ETU joints,

- Control of end effector, connector positioner drive, and MST

latch and bolt drives,

- Operator assisted and unassisted modes;

6) Software program is menu driven (Figure 1-17);

7) Procedures and trajectory sequences for the Manual-Augmented

control mode were documented;

8) Simulated hardware characteristics are Included in software so

program can be run independent of servicer hardware;

9) A test program for verifying the computer to servicer hardware

interfaces was provided;

10) A separate Software User's Manual was prepared for the MMS module

software.

1-34
IOSS MAIN MENU - MMS

1. Run Setup Menu;

2. Mode Selection Menu;

3. Module Data Collection Menu;

4. Hardware Calibration Menu;

5. Exit to MCR.

Enter Item Number:

MODE SELECTION MENU

1. Unassisted Supervisory Mode;

2. Assisted Supervisory Mode;

3. Manual-Augmented Mode;

4. Manual-Direct Mode;

5. Return to IOSS Main Menu -

MMS.

Enter Item Number:

Figure 1-17 Representative MMS Software Menus

1.5.10 Servicer Software Demonstrations

The following conclusions and recommendations were identified during

the conduct of the basic and MMS module exchange demonstrations using

the two servicer software programs:

1) All of the demonstration equipment operated satisfactorily and was

comprised of:

- ETU and associated electronics by MSFC,

- PDP-11/34 computer with D/A and A/D's by MSFC,

- MMS modules, spacecraft mockup, and stowage rack modifications

by Martin Marietta,

- Connector positioner and wiring changes by Martin Marietta,

- 1-g Module Servicing Tool by GSFC (see Figure 1-18),

- MST electronics by Fairchild Space Co;

1-35
Figure 1-18 Module Servicing Tool for Ground Demonstrations

of MMS Module Exchange

2) Specific module location data could be readily collected for use in

the software program and in the Manual-Augmented trajectory

sequences using the procedures that were developed;

3) Separate demonstrations of basic and MMS module exchange were

successfully made in all three control modes (Figure 1-19);

4) Conduct of demonstrations in the Supervisory control mode in the

operator assisted or unassisted modes was easy to learn. Operation

in the Manual-Augmented control mode takes a little longer to

learn, as was expected;

5) Motion of the ETU during module exchanges in either Supervisory

mode was very smooth and precision was well within the basic module

equipment capture volumes and just within the tighter MMS equipment

capture volumes;

6) Integration of the MST was accomplished by operating philosophy

revisions, software modifications, and hardware adjustments;

1-36
ORIGINAL PAGE IS

OF POOR QUALfTY

Figure 1-19 Ground Demonstration of MMS Module Exchange

1-37
7) Non-orthogonality of the MMS module with respect to the docking

post (axial cylindrical coordinate) when the module top bolt is

tight and the bottom bolt is loose were accommodated by the

addition of a pseudo-combined motion capability where all six

cylindrical coordinates are changed together in a step-wise fashion

to approximate the desired path;

8) System operating techniques were identified for overcoming

anomalies so that the system should not be thought of as a

pre-programmed entity that cannot continue past the first anomaly.

Rather it is a system with three levels of control that can be used

interchangeably to get the job done in spite of a variety of

anomalies.

1.6 SUGGESTED ADDITIONAL EFFORT

A review of the study efforts and conclusions identified a number of

areas that merit consideration for additional effort. In addition to

the items listed below, it is assumed that the TDRSS program and the

OMV program including a docking system, payload rigidization system,

and ground control station will continue.

On-orbit servicing of spacecraft has become a part of the Space Station

development activity over the last year, as it should. These

activities are on-going at several NASA centers and at their Space

Station contractors. The work has been emphasizing robotics and

automation because of a Congressional directive to assign 10% of the

Space Station budget to advanced automation activities so that

operating costs can be reduced. The module exchange activities, which

are the subject of this report, and the associated equipment and

software should be brought to the attention of the Space Station

project for consideration. The IOSS concepts could form the first

phase of an on-orbit servicing capability and then evolve into a smart

front end with telepresence and artificial intelligence, as the needs

are understood and the technology is developed. It is recommended that

the IOSS concepts of module exchange for the on-orbit repair of

1-38
spacecraft be fully considered for its potential application to Space

Station.

1.6.1 Servicing Tasks

The following additional efforts are related to servicing tasks and in

particular to the Multi-Mission Modular Spacecraft, fluid resupply

demonstrations, and representative satellite modules:

1) The Module Servicing Tool and the servicer mechanism end effector

should be adapted to work together for the exchange of MMS modules

in 0-g;

2) The fluid resupply interface should be standardized;

3) The fluid resupply demonstration equipment should be based on the

NASA-JSC standardization effort;

4) Thermal cover removal/replace mechanisms and sensors for fastener

and attach interface status need to be developed;

5) A small, light interface mechanism or a tool adapter to remove

conventional captive fasteners should be developed.

1.6.2 Servicing Mechanism

The following additional efforts are related to the servicing mechanism:

1) The interface between the servicer end effector and the interface

mechanism, tools, and adapters should be standardized;

2) Special adapters should be developed as required for other types of

modules or servicing tasks;

3) An activity for continuing repair and maintenance of the 1-g

servicing demonstration equipment, including documentation and

configuration control, should be established.

1-39
1.6.3 Demonstrations

The following additional efforts are related to the ground and

cargo-bay demonstrations or to the free-flight verification:

1) Servicing control modes for the 1-g servicer should be analyzed and

investigated. Nine candidate subjects are discussed in Section

5.2.3 of this report;

2) Refueling and resupply hardware should be developed and the process

demonstrated;

3) An automatic target recognition and error correction system should

be developed and demonstrated;

4) Status of the MSFC Robotics Laboratory computer facilities should

be addressed to identify and implement an approach to obtaining a

higher level of reliability. This review should also consider use

of a 14 or 16 bit analog to digital converter;

5) Definition of the cargo-bay demonstration equipment should be

continued in the areas of servicer mechanism definition and

identification of the microprocessor and associated peripherals;

6) Additional development areas include:

- Special refueling disconnects for cryogenics or high pressures,

and self aligning conical electrical connectors,

- Development of in-line fluid couplings for replacement of tanks

and other propulsion system components,

Demonstration of other servicing tasks specific to Space

Station operations;

7) Demonstration of the mating of the servicer stowage rack to the OMV

should be a part of the Space Station technology development

missions.

1-40
2.0 INTRODUCTION AND BACKGROUND

The study activities that this report documents are part of a

resurgence of interest in on-orbit servicing that is based on the many

studies performed in the past. Those studies were able to clearly show

that orbital maintenance functions can be supported by the Space

Transportation System (STS) to effect large reductions in the cost of

spacecraft programs. This was found to be true both in geosynchronous

and low earth orbits. These economic benefits were augmented by

significant operational benefits, the totality of which implied that

the development of an on-orbit servicing capability should be

undertaken by the NASA.

Orbital servicing has a number of applications. The servicer and the

Orbital Manuevering Vehicle (OMV) can be carried to geosynchronous

earth orbit (GEO) on an Orbital Transfer Vehicle (OTV). Communications

satellites are typical geosynchronous spacecraft that can realize cost

benefits from servicing. In low earth orbit the OMV can be used as the

carrier vehicle for the servicer system. Where contamination or

thruster impingement effects are a concern, the cold gas propulsion

system of the OMV could be used. For spacecraft in different orbits

(altitude or inclination) the larger propulsive capability versions of

the OMV, or the OTV with OMV, are appropriate. The servicer system can

also be deployed in the Orbiter cargo bay and the failed spacecraft

docked to it using the Remote Manipulator System (RMS). A major

opportunity for the use of orbital maintenance technologies is the

emerging Space Station. The Space Station can be used as a base for

the OMV and OTV, which can transport a remotely controlled servicer

system to the failed spacecraft for repair in situ. Alternatively,

failed spacecraft can be returned by the OMV and OTV to the Space

Station for repair. Spacecraft repair at the Space Station can be by a

variety of techniques including remotely controlled module exchange.

One of the early servicing studies, the Integrated Orbital Servicing

System (IOSS) series, recommended that, to minimize servicer system

development costs, a single servicer system having the capability to

2-1
accommodate both low and high earth orbit applications should be

evolved. This requirement has been satisfied effectively by the

servicer mechanization (Figure 2-1) conceptualized during the IOSS

studies. The single design is compatible with maintenance of most

spacecraft of the STS era. Adapters are used to accommodate support

structure differences across the applications. An effective interface

between the spacecraft and the servicer was defined and breadboarded.

The interface mechanism provides a logical and cost effective method of

incorporating orbital replaceable units (ORU) for module exchange in

all spacecraft and can be applied to the Space Station itself.

MODULE

STOWAGE RACK

SERVICER STOWAGE/

DEPLOYMENT MEGHANI

OMV OR ORBITER

INTERFACE

END ETFECTOR

AND WRIST DRIVES

ELBOW ROLL DRIVE

DOCKING

MECHANISM

SHOULDER DRIVES

INTERFACE

MECHANISMS

Figure 2-1 On-Orbit Servicer Configuration

Considerable interest in spacecraft maintenance was expressed by both

the Department of Defense and the commercial sector; however, the

general tenor of their support was that a demonstration of orbital

2-2
maintenance must be conducted prior to any commitment on their part. A

flight demonstration of the all-up maintenance capability is also a

NASA requirement prior to wholesale commitment to the concept.

However, a reduced capability test that exercises the basic concept and

exchanger capability can and should be demonstrated prior to the time

when the full capability will exist. With this background material in

hand, and with renewed interest by the space flight community, it was

appropriate to perform the prior study (NAS8-35496), which defines a

path culminating in the demonstration of an on-orbit servicing

capability. The objective of that study was to provide a single

unified development program for both servicing implementors and users

to guide their future development and operational plans for this

important technology. The objectives of the current study are to

refine the servicer development plan and to begin the 1-g testing in

terms of demonstrating basic and MMS module exchange in three control

modes.

2.1 STUDY OBJECTIVES

The objectives of this Servicer System Demonstration Plan and

Capability Development study are to identify all major elements and

characteristics of an on-orbit servicing development program and to

integrate them into a coherent set of demonstrations, to upgrade the

Engineering Test Unit control system for basic and MMS module exchange

demonstrations, and to upgrade the MSEC 1-g servicing demonstration

facility mockups to permit the exchange of MMS modules. These

objectives, along with the program objectives, are summarized in Table

2.1-1. The on-orbit servicing development plan was to be a revision of

the plan prepared during the prior study with increased emphasis on low

cost and use of MMS equipment. The revisions primarily addressed the

cargo-bay demonstrations.

The goal of the development program is a fully verified operational

on-orblt servicing system based on the module exchange and fluid

resupply technologies that is also suitable for use with and at the

Space Station. A ground demonstration plan is envisioned that will

provide confidence in the development and operation of the on-orbit

2-3
system.. The servicing ground demonstrations cover a range of satellite

module sizes and Include the ability to service propellant systems.

They also include a servicing mechanism configuration that is

representative of an eventual flight unit. Qnphasis was placed on the

exchange of MMS modules.

Table 2.1-1 Study and Program Objectives

Study Objectiv

To identify and integrate the major characteristics of an on-orbit

servicing demonstration program plan.

To upgrade the engineering test unit control system for MMS and basic

module exchange demonstrations.

To upgrade the 1-g demonstration facility to permit exchange of MMS

modules.

Program Objectives

Fully verified and documented operational on-orbit servicing system

a) Based on module exchange and fluid resupply technologies,

b) Suitable for use with Space Station.

Major issue is balance between the number and complexity of

development activities and cost.

The Orbiter cargo-bay demonstrations utilize a protoflight version of

the servicer mechanism to reduce project costs. A single flight is

planned to demonstrate the exchange of a variety of modules, adequacy

of control from the Orbiter using the Supervisory control mode, and

accuracy of spacecraft to stowage rack alignment when the Remote

Manipulator System end effector is used as a docking mechanism. A

free-flyer (Orbital Maneuvering Vehicle) demonstration is planned as a

way of verifying the capabilities of an operational servicer. The plan

must be significant and long-term to encourage users and spacecraft

designers to include on-orbit servicing in the form of module exchange

in their plans.

The second study objective involves the development of two software

programs - one for the exchange of basic modules and one for the

2-4
exchange of MMS modules - and the demonstration of the exchange of both

types of modules. The demonstrations were to be performed for three

different control modes which are:

1) Supervisory with minimal operator assistance;

2) Supervisory with operator assistance at each action;

3) Manual-Augmented.

The basic module is a 24 in. cube and uses a side interface mechanism

to provide the structural interface between the module and spacecraft,

or stowage rack. Basic module exchanges were of four types:

1) A failed module in a spacecraft axial location being replaced;

2) A failed module in a spacecraft radial location being replaced;

3) A module being transferred from a spacecraft axial to a radial

location;

4) A module being transferred from a spacecraft radial to an axial

location.

The last two transfer types were used to simplify setting the

demonstration equipment up for either of the first two replacement

activities.

The MMS module has a 48 in. square plan form and is 20.5 in. deep. It

is fastened in place with two bolts. The bolts are tightened or

loosened with a Module Servicing Tool. MMS module exchanges involved

replacing a failed module in a spacecraft axial location with a good

module from the spare module stowage rack.

2-5
The third study objective involved the design, fabrication, and

installation of MMS demonstration equipment. The Martin Marietta

provided equipment included:

1) Two MMS module mockups;

2) One spacecraft mounted module receptacle;

3) Two stowage rack mounted module receptacles;

4) A connector positioner drive;

5) An MST storage rack;

6) A set of MMS module targets;

7) A set of related wiring.

A light weight form of the Module Servicing Tool, which was adapted to

the Engineering Test Unit and modified for remote location of its

control system was provided by Goddard Space Flight Center. The

mechanical and electrical design was performed by Fairchild Space

Company, the mechanical equipment was built by GSFC, and the electrical

equipment was built by Fairchild.

The first of two key study issues was the need to balance the number

and complexity of development activities against available funds. The

proposed approach, recommended in the Spacecraft Servicing

Demonstration Plan (SSDP) study, is to lay out a program with most of

the desired features, that overlaps the 1-g, 0-g, and operational

servicer demonstrations, and attempts to get an early operational

capability. It minimizes costs by taking advantage of parallel

activities such as the JSC refueling program, and advocates renting a

spacecraft bus rather than buying a new one. The program was also

scoped large enough to become a recognized part of NASA's long-range

plans. The promise of a clear plan by NASA to develop and use module

2-6
exchange for many years will encourage the user, or spacecraft

designer, to incorporate module exchange in his plans.

In evolving the SSDP recommended approach, a range of alternatives was

considered. At the high end of the spectrum was a servicer development

program to demonstrate several forms of module exchange, several cover

door opening or removal approaches, three or four approaches to

refueling (fluid resupply), and several approaches to cryogenic

resupply in each of three areas — l~g> 0-g, and free-flight. The

three phases were put in series so full advantage of prior work could

be incorporated in subsequent activities; this resulted in a long and

expensive program. Additionally, on-orbit servicing opportunities

would be lost with a concurrent loss of potential savings.

At the other end of the spectrum was a minimum cost program where

minimum cost equated to fewer functions being demonstrated, fewer

demonstrations, and a higher level of risk acceptance. However, a

significant failure could be enough to delay development of module

exchange by 10 years, or possibly ending it forever. It was decided to

go with a low cost program and mitigate risks by overdesign and use of

existing equipment where possible.

The second key study issue was the need to maintain a close working

relationship between MSFC and Martin Marietta personnel during servicer

control software development. A number of interfaces were defined so

both organizations could work towards the same goal:

1) Computer and interface electronics operations;

2) Functions to and from the Servicer Servo Drive Console and the ETU;

3) Functions to and from the control station for Manual-Augmented

control mode implementation.

2-7
An initial defintion of these interfaces was discussed at the

orientation meeting. The interface defintion was then refined and

documented as software requirements were specified and the software

code was prepared. When the software was delivered to MSFC, these

interfaces were verified before control system demonstrations were

initiated. Effective interface definitions led to their being quickly

verified.

2.2 BACKGROUND

One of the justifications for the Space Transportation System was its

potential for supporting the repair or recovery of failed spacecraft.

This approach was extended to the concept of making less expensive

spacecraft, accepting the higher predicted failure rates, and using the

Shuttle to permit repair of those spacecraft that did fail. This

spawned a large number of government, academic, and industry studies on

how spacecraft might be configured for on-orbit servicing. Figure

2.2-1 illustrates the variety of concepts that were documented. The

whole gamut from recovery and ground refurbishment, through repair at

the Orbiter, through remote operations in low earth orbit, to repair in

geosynchronous orbit were addressed. All of the concepts discussed now

were addressed then except for Space Station related operations. The

long cylindrical spacecraft represents the Space Tug whose missions are

now to be handled by the Orbital Maneuvering Vehicle (OMV) and the

Orbital Transfer Vehicle (OTV). A good summary of the early work is

given in Proceedings of the Second Conference on Payload Interfaces,

MDC G4818, McDonnell Douglas Astronautics Company, Huntington Beach,

California, September 6-7, 1973.

Benefits from orbital servicing of spacecraft were identified in these

early studies and potential savings continue to be well recognized.

The three general approaches to orbital servicing are:

1) Man on extravehicular activity (EVA);

2-8
ORIGINAL PAGE IS

OF POOR QUALTtt

Figure 2.2-1 Serviceable Spacecraft Designs from the 70's

2) Operations remote from the man using telepresence technology;

3) Module exchange and refueling/resupply using a simple remotely

controlled mechanism.

Remote module exchange and refueling/resupply for on-orbit servicing

were examined and had the broadest application in the near-term of the

three servicing approaches. They are not as time- or space-limited as

EVA is. The technology is here and available while telepresence

technology is still in the research stage. Therefore, it is

appropriate that work on the module exchange and refueling/resupply

form of on-orbit servicing be continued. Martin Marietta has been

active in this technology since 1974 and is committed to actively

promoting it.

2-9
The extensive resource base was used in the 1974 through 1978

Integrated Orbital Servicing System (IOSS) study conducted by Martin

Marietta for Marshall Space Flight Center. The IOSS study initially

used the 1973 NASA mission model as a basis for establishing cost

benefits. The model included 47 NASA satellite programs for which

maintenance was applicable. Applicability of maintenance was based

on: spacecraft fleet size on orbit, program lifetime, and need for

equipment replacement.

If a satellite program was short, or the spacecraft value was low, then

maintenance was not attempted. Cost comparisons were made between:

1) Expendable spacecraft;

2) Return to the ground for refurbishment;

3) Return to the Orbiter for refurbishment;

4) Module exchange in the operational orbit (in situ servicing).

Generally, module exchange in the operational orbit was most cost

effective. If spacecraft are cheap, then it is cost effective to

expend them. The costs of returning a spacecraft to the ground and

relaunching were high enough to rule out ground refurbishment. Orbit

phasing effects and the launch costs related to propellant usage in

bringing spacecraft, especially geosynchronous spacecraft, back to the

Orbiter ruled out maintenance at the Orbiter. However, the orbits of

some spacecraft make this an acceptable approach. There were

significant cost savings from repair by module exchange in the

spacecraft's operational orbit. These savings are larger than the

costs of servicer system development. The same results were obtained

using much smaller mission models. These study results are applicable

to current-day situations. Some specific satelLite programs have

changed since these study results were generated; however, the

conclusions on cost effectiveness are as applicable to today's

satellite programs as they were to the program projected in 1973.

2-10
A wide variety of servicer mechanism configurations were identified in

the literature. They ranged from simple one degree-of-freedom (DOF)

devices, through a three DOF rectangular travel system, to two-arm

concepts, each with 7 DOF. The IOSS selected approach started with the

Shuttle launch cost rules that favored flat disk-shaped spacecraft such

as the Orbital Maneuvering Vehicle. From this, the servicer working

volume and observations shown in Figure 2.2-2 were developed.

Module Attachment

Location

Module

Stowage Rack

Spacecraft

Observations:

Separation Distance

• The module attachment locations form a surface of revolution about the

spacecraft centeriine.

• The first servicer degree of freedom should be roll about the base of the

docking probe.

• The need for minimum arm length and separation distance implies the

servicer mechanism must "reach around" the spacecraft and module

surfaces.

Figure 2.2-2 Servicer Mechanism Working Volume

The shaded area on Figure 2.2-2 represents the regions where the

servicer mechanism end effector must reach. The direction of module

removal is generally perpendicular to the shaded surface. The

applicability of a roll rotation for the first degree-of-freedom is

quite apparent. As the separation distance between the spacecraft and

stowage rack is reduced, the space available for servicer mechanism

elements near the base is reduced and the "reach-around" problem

becomes more difficult. The minimum separation distance was taken as

60 in., which allows for a 40 in. module, a 10 in. end effector, and a

5 in. clearance on each end. The "reach-around" problem leads to use

of a redundant degree-of-freedom.

2-11
Figure 2.2-2 implies that two layers, or tiers, of modules could be

incorporated at a single docking location. It was later decided to

simplify the servicer design to permit module exchange only from the

first tier and to wait until a specific need is identified before the

servicer configuration is grown to handle the second tier.

An extensive review and analysis of servicer mechanism configurations

and 28 serviceable spacecraft configurations was performed to arrive at

the selected servicer configuration shown in Figure 2-1. From the

review and analysis, extensive sets of requirements were prepared and

refined. All servicer configurations involving one or two arm segments

and many three arm segment configurations were considered.

The Figure 2-1 design has only two major components: 1) a servicer

mechanism and 2) a stowage rack for module transport. A docking

mechanism is also shown for referen.ce. The servicer mechanism and the

stowage rack were designed separately with interfaces for individual

removal and replacement. Stowage racks can be configured and loaded

for particular flights prior to attachment to the carrier vehicle. It

may be desirable to have available several stowage racks for this

purpose. The stowage rack shown mounts directly to an upper stage such

as the Orbital Maneuvering Vehicle. A flight support structure has

been designed to adapt the stowage rack shown to the Orbiter.

The entire design of the servicer system has been predicated on the

simple nature of the module exchange task as compared to the broader

variety of tasks that a general purpose manipulator is called upon to

perform. The simple activities of remove, flip, relocate, and insert

modules, when combined with the facts that all aspects of the module

trajectories are known far in advance of use and that the work volume

is a simple solid of revolution, were used in many ways to result in a

basically simple design in terms of mechanism configuration, control

system design, and operations approach. This simplicity was

accentuated by performing the mechanism and control system designs

concurrently in an integrated manner so that each of the needed

2-12
c •*"•«

functions was allocated to the system that could most effectively

accomplish it.

Three modes of control were included. The Supervisory mode of control

was proposed as the normal mode of operation. All servicer arm motions

and trajectories are determined before flight and stored on board. A

Manual-Direct mode is provided as a totally unsophisticated means of

backup control. It sends commands directly to the joints themselves.

The Manual-Augmented mode has man doing most of the arm control as in

the Manual-Direct mode only using hand controllers instead of panel

switches.

The value of demonstrations in furthering on-orbit servicing

development was recognized in the decision to build a 1-g version of

the Integrated Orbital Servicing System of Figure 2-1. The result is

the Engineering Test Unit (ETU) shown in the photograph of Figure

2.2-3. This unit was built and delivered to MSFC in 1978. It has been

used for over 350 demonstrations during the intervening seven years.

The ETU has shorter segment lengths than the IOSS as it was designed

initially for axial module exchange only. The later addition of a

sixth degree-of-freedom extended the ETU's capability to radial module

removal, albeit at a radius less than that of the Orbiter cargo-bay.

To date, satellite systems in general have not been designed and built

with the capability of changeout of subsystem or component modules.

The only satellite family that is currently in use and has a module

exchange capability is the Goddard Space Flight Center's Multi-mission

Modular Spacecraft (MMS). This satellite family is in operation in

several programs and is projected for continued use throughout* the

remainder of this century. The Marshall Space Flight Center's Space

Telescope has been designed for on-orbit repair by an astronaut on EVA

and is expected to fly soon. The U.S. Air Force has also shown

interest in the design of serviceable spacecraft, although the

particulars are not known to the authors.

2-13
ORIGINAL PAGE 18

OF POOR QUALITY

Figure 2.2-3 Engineering Test Unit

Several demonstrations and investigations of on-orbit refueling

capability are currently being planned. These efforts will include

definition and demonstration of connect/disconnect devices in support

of the transfer of fluids. Electrical umbilicals and connectors have

been developed in conjunction with the MMS subsystem modules as well as

on other programs.

The emerging Space Station program with its use as a base for many

spacecraft and with its associated polar platforms is a unique

opportunity to develop and implement on-orbit servicing in the form of

module exchange. The simple forms of a servicer and control system can

be used first and they can then be modified to the more advanced forms

of telepresence and artificial intelligence as these technologies

become available.

2-14
2.3 STUDY APPROACH

The work completed under contract NAS8-35496 indicated that a viable

plan for module exchange demonstration in the Orbiter cargo bay could

be prepared and integrated with a free-flyer based verification of an

operational servicer system. One of the first steps of the servicer

demonstration plan is upgrading the control system for the ETU to have

all three control modes - Supervisory, Manual-Augmented, and

Manual-Direct - operable. The incorporation of control equations

specifically generated for the unique geometrical configuration of the

ETU promised to result in smooth demonstrations with the ETU.

Our approach to the proposed study was to use the four tasks identified

in the contract statement of work. These are:

1) Task 1 - Servicer Development Program Plan;

2) Task 2 - Servicer Control Software;

3) Task 3 - Servicer Demonstration;

4) Task 4 - Program Management.

Other than the program management task, the work divides naturally into

two parts - preparation of the Servicer Development Program Plan (Task

1) and generation of the servicer control software as well as

conducting servicer demonstrations at MSFC (Tasks 2 and 3). The MMS

1-g servicing demonstration definition effort of Change Order 1 and the

MMS 1-g demonstration equipment drawing, fabrication, checkout, and

installation effort of Change Order 3 were included in Task 1. The MMS

module software requirements, programming, and user's manual

preparation effort of Change Order 3 were included in Task 2, while the

software installation and MMS module exchange activity of Change Order

3 were included in Task 3. MMS module exchange demonstrations required

the availability of a GSFC MMS Module Servicing Tool designed for use

in 1-g with the MSFC Engineering Test Unit.

2-15
Figure 2.3-1 shows an overall logic flow for the four study tasks.

Task 1 was to expand the Servicer Development Program Plan to include

detail planning and cost estimating for ground, in-bay, and free-flight

servicer demonstrations using a servicer system compatible with the

Orbital Maneuvering Vehicle (OMV). Special emphasis was to be devoted

to MMS servicing demonstrations. The Servicer Development Program Plan

produced under this task was to be used as input for the reports to be

produced under Task 4.

Preparation of the Servicer Development Program Plan was a natural

outgrowth of our work on Contract NAS8-35496, Spacecraft Servicing

Demonstration Plan. Both versions of the plan involve:

1) Use of the existing ETU at MSFC to demonstrate remotely controlled

exchange of a variety of modules and fluid resupply in 1-g;

2) Demonstrations of module exchange and fluid resupply in the Orbiter

cargo bay in 0-g;

3) A free-flight demonstration of module exchange and fluid resupply

using the OMV to bring the servicer to a spacecraft bus that

supports the modules to be exchanged.

The work emphasized exchange of Multi-Mission Modular Spacecraft

modules and the adaptation of the MMS Module Servicing Tool to the ETU.

Task 1 also included the servicer system Multi-mission Modular

Spacecraft 1-g demonstration design and plan. The overall

configuration and specific design requirements necessary to effect the

MMS 1-g module exchange demonstration were identified. Also in Task 1

was the fabrication, delivery, and installation of a set of mockup

equipment to aid in the demonstration of MMS module exchange in 1-g.

This equipment was integrated with the light weight Module Servicing

Tool provided by GSFC.

2-16
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2-17
The first six subtasks of Task 2 were completed before the first three

subtasks of Task 3 could be started. Similarly, the last three

subtasks of Task 2 were completed before the last three subtasks of

Task 3 could be started. The main inputs to Task 3 (the basic and MMS

control software programs and the control Software User's Manuals) were

produced under Task 2 at Martin Marietta Aerospace. As indicated in

Figure 2.3-1, the initial set of software for the exchange of modules

using the side interface mechanism is designated as "basic" software,

while that used for MMS module exchange is designated as "MMS" module

software.

Our approach to generating the software for servicer control system

upgrading was to start with the software used during the ETU design

acceptance review conducted at Martin Marietta. The requirements for

that software were modified extensively to handle the more complete

trajectories required, to enable us.e, or avoidance, of operator steps

when in the Supervisory mode, and to incorporate the end effector and

interface mechanism latching functions. The software was modified to

operate on the PDP-11/34 computer system at MSFC. Additional

modifications were made for operation with the MSFC electronic

interfacing equipment. All software modifications were checked out on

a PDP/11-34 computer at Denver Aerospace. These checks included use of

integrators to simulate joint drive motion.

In this way, it was possible to check out the software in a closed-loop

manner. Equations and instructions for both the Supervisory and

Manual-Augmented servicer control modes were provided. Software User's

Manuals were prepared under Subtasks 2.5 and 2.9 for basic and MMS

module exchange demonstrations. The necessary interfaces between the

new software and the MSFC equipment were initially established at the

Orientation meeting for the basic module software and at a Design

Coordination meeting for the MMS module software.

The basic software was checked out and demonstrated at MSFC on the

existing Engineering Test Unit (ETU), under Task 3. The MMS software

was checked out and demonstrated at MSFC on the existing ETU using the

2-18
MMS module mockups prepared under Task 1 and the 1-g version of the

Module Servicing Tool provided by GSFC.

Our initial approach to the servicer demonstrations was to have the

study manager and a software engineer travel to MSFC for a one week

period for each demonstration. However, several additional trips were

found to be necessary. During those visits the following occured:

1) New software was installed on the MSFC PDP-11/34 and certain

functional checks were conducted;

2) The software was interfaced with the Servicer Servo Drive Console,

the MSFC hand controller, and the ETU;

3) Servicer system operation using the new control system software was

demonstrated as defined in Task 3.

MSFC personnel were directly involved in these activities and obtained

hands-on training in using the new software. The cooperation and help

of the MSFC Robotics Laboratory personnel in performing the

demonstrations is appreciated very much. Without their efforts, we

would not have been able to successfully conduct the demonstrations.

Task 4 received timely inputs from Tasks 1, 2, and 3 as needed for

monthly progress reports, the final report, and the midterm and final

presentations. Updating of the Study Plan was also done under Task 4.

2-19
3.0 MMS SERVICING CONSIDERATIONS

The objective of this phase of the study was to review and update the

Multi-Mission Modular Spacecraft work of the prior study with the view

of placing increased emphasis on MMS module exchange and identifying an

approach to adapting the MMS module servicing tool to the ETU end

effector for ground, cargo-bay, and free-flight demonstrations. The

MMS, and MMS modules, have been, or will be, used in a number of

satellites that are currently flying, such as the Solar Maximum Mission

(Figure 3-1), that are being designed and built, such as the Landsat-D

and some defense systems, and they are being considered for other

spacecraft programs. The MMS evolved from a desire to standardize

subsystems and thus obtain low costs. The concept also involved an

ability to remove and replace modules while in orbit. Although the

initial concepts included a remotely controlled mechanism to exchange

modules, recent emphasis has been on use of man on EVA to exchange MMS

modules. This was demonstrated in 1984 during the Solar Max Repair

Mission.

Thus, one of the best ways to advance the satellite servicing

technology, using module exchange techniques, is to demonstrate an MMS

module exchange. However, the pivoting arm form of servicer mechanism

used in the Integrated Orbital Servicing Study was not explicitly

designed to interface with the MMS module attachment system. The

capability of the IOSS servicer to exchange MMS modules was developed

in the prior study. That work demonstrated the usefulness and

adaptability of the pivoting arm configuration.

The basic MMS spacecraft (Figure 3-2) consists of three standard

spacecraft subsystem modules and a mechanical structure that supports

the spacecraft subsystem modules. The structure also provides the

support for the instrument (payload) module, which is not part of the

MMS. The standard spacecraft subsystem modules are a communications

and data handling (C&DH) module, an attitude control subsystem (ACS),

and a modular power subsystem (MPS). The instrument module, which

includes the payload instruments and other mission unique equipment

3-1
Figure 3-1 Solar Maximum Mission Version of MMS

(such as solar arrays, high-gain antennas, etc.), attaches to a

transition adapter frame on the forward end of the MMS. A propulsion

module (PM) or a high gain antenna may be added to the aft end of the

MMS as a mission option. A signal conditioning and control unit

(SC&CU) and the electrical interconnecting harness complete the basic

MMS.

Associated with the MMS is a flight support system. The FSS is mounted

in the Orbiter and is used to support the MMS during launch and

return. It can also be used to elevate the MMS for direct deployment

or for deployment, or retrieval, by the Orbiter remote manipulator

system (RMS). The FSS was used to hold the MMS during EVA exchange of

modules during the Solar Max Repair Mission.

3-2
ACS MODULE

TRANSITION

ADAPTER

GRAPPLE POINT

MODULE

RETENTION

HARDWARE

THERMAL

LOUVERS

PMIA AUXILIARY

TANK KIT (ATK)

C& DH MODULE

PM-I

PROPULSION

MODULE

Figure 3-2 Multi-Mission Modular Spacecraft Mechanical System

Each MMS module is fastened with two bolts (module retention system).

Despite the radial arrangement of the modules with respect to the MMS

centerline, they cannot be removed in a direction perpendicular to the

servicer docking axis because each end of the central triangular

structure is blocked: one end by the transition adapter and experiments; the other end by a propulsion system or a high-gain antenna.

Each of the three large replaceable MMS modules is similar in external

configuration as shown in Figure 3-3. The maximum weight of an MMS

module is 500 Ibs and the module structure (frame, cover, module

retention system (MRS) and thermal hardware) weighs approximately 95

Ibs. For ground demonstrations, the module retention system and the

electrical connectors may be used in a MMS module mockup weighing

approximately 15 Ibs. The upper module retention system consists of a

beam, two restraint sockets and a preload bolt. The lower MRS consists

of a beam, two optional snubbers, and a preload bolt with reaction

pads. The overall result is a non-redundant attachment if the two

optional snubbers are not used.

3-3
EXTENDER

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RESTRAINT

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MODULE

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UPPER

PRELOAD

BOLT

CONNECTOR &

BRACKET C)

MODULE COVER

(EQUIPMENT

BASEPLATE)

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BRACKETS (2)

(OPTIONAL!

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PRELOAD

BOLT

MRS LOWER

HOUSING

CORNER

BRACKETS i2)

47.07 in.

(1.196m)

CORNcr"

DOUBLERSfg)

Figure 3-3 MMS Module Structure with Module Retention System

A module servicing tool (MST) was designed and built as a battery

powered EVA hand tool. It was designed to loosen and tighten the MMS

module retention hardware to predetermined torques of up to 160 ft-lb.

It provides a means for locking onto the modules in a manner that

avoids reaction torques on the crew member. Power is supplied by a

battery housed in the tool assembly. This tool has three separate

motors to perform two functions (1) latch the MST to the module to

react bolt-driving torques and to provide a handle for maneuvering the

module (2 motors) and (2) drive the preload bolts in and out. The MST

is quite bulky and heavy because of the self-contained batteries.

However, it can be provided with a servicer standard interface and can

be used as an adapter for exchanging the MMS modules using the servicer.

3-4
Various alternative servicing methods for a Multi-Mission Modular

Spacecraft were analyzed in the prior study. The recommended method

for remote, on-orbit servicing of the MMS was to use the standard

servicer configuration fitted with a straight docking probe adapter, a

modified Module Servicing Tool (MST) and a modified stowage rack

(Figure 3-4). The servicer docks with the MMS laterally, on its

existing grapple fixture or on a grapple fixture/berthing pin

combination that replaces the existing berthing pin between the power

module and the C&DH module, as shown in Figure 3-5. The docking probe

adapter is designed to be compatible with the servicer docking probe

interface at one end and with the MMS docking aid interface at the

other. A joint similar in design to the other servicer joints is

included in the docking probe adapter to allow tilting of the servicer

with respect to the MMS after docking to bring the servicer mechanism

into a plane parallel to the face of the module to be exchanged. The

joint is powered through an electrical connection across the servicer

docking interface. This feature allows the simple, axial mode of

operation of the servicer without modifying its basic configuration.

Either one of the two modules adjacent to the grapple fixtures can be

serviced in one docking. Two grapple devices, the standard one and the

grapple fixture/berthing pin combination, are required if it is desired

to be able to service any of the three modules. No modifications of

the MMS modules or module retention system (MRS) are required.

Instead, a modified MST compatible with the existing MRS and with the

servicer standard end effector interface is recommended.

3-5
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3-6
ACS MODULE

SUPPORT STRUCTURE

C4DI1 MODULE

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SC&CU

POWER MODULE -

SUPPORT

SC&CU STRUCTURE

GRAPPLE AND

BERTHING PIN

COMBINATION

N O W

GRAPPLE AND

BERTHING PIN

COMBINATION

LOWER FASTENER

CENTERLINE

Figure 3-5 Docking and Berthing Pin Combined Design

The remainder of this section includes:

1) A discussion of MMS servicing requirements with emphasis on the

operational case;

2) A discussion of an approach to modifying the module servicing tool

for use in a 1-g servicing demonstration;

3) An analysis of MMS module exchange options;

4) A discussion of extended remote maintenance and resupply concepts.

3.1 MMS SERVICING REQUIREMENTS

During the prior study, a set of requirements for operationally

servicing spacecraft with MMS modules was derived. These requirements

were reviewed and complemented in this study. The resulting

requirements to be placed on the servicing system are:

3-7
1) Minimum modification of the present configuration of the MMS

modules and/or module support structure;

2) Minimum modification of the standard configuration of the

servicer. An adapter may be used in connection with the standard

end effector interface in order to service MMS modules;

3) The method of removal/attachment of the MMS module shall be

compatible with the demating/mating of the existing electrical

connector(s) situated on the back of the module;

4) Adequate clearance shall be provided at all times between module

and satellite structure or other components;

5) The servicer shall clear the propulsion module or high gain antenna

at the lower end of the MMS support structure. A clearance

envelope of 86 in. by 103 in. diameter is required for satellites

using the Mark II propulsion system. The servicer should also

clear the payload envelope, including solar panels, antennas and

other appendages;

6) The number of times the servicer docks with the MMS in order to

perform all the servicing tasks shall be kept to a minimum;

7) The accuracy in positioning the servicer for module engagement

shall be within the capture envelope of the module retention

system. The adapter design shall be such as to minimize the errors

and the softness of the coupling at the interface;

8) Demonstration of other servicing tasks such as battery exchange,

other types of module exchange and expendable resupply, in addition

to MMS module changeout, should be performable on the same mission

without need for system reconfiguration.

3-8
The basic philosophy behind these requirements is that there should be

minimum changes required for any MMS components. The MMS is a fully

developed, operational system. Therefore, MMS design changes to

accommodate servicer existing interfaces or other servicer requirements

affect existing hardware and tooling and their implementation is

expensive. This cost element was considered in defining a servicer

system capable of exchanging MMS modules. However, some changes are

necessary and it may not be possible to service those MMS spacecraft

that have not been configured for remotely controlled servicing. The

requirements placed on the MMS spacecraft for remotely controlled

on-orbit servicing are:

1) Attitude stability that is within the capture envelope of the

Orbital Maneuvering Vehicle docking system;

2) RMS grapple fixture(s) adjacent, to the module(s) to be exchanged;

3) Electrical power control and an RMS grapple fixture electrical

connector that allows the servicer to control electrical power to

the module to be replaced and to put the MMS attitude control

system into a safe condition;

4) A target that allows the servicer to estimate post-docking

alignment errors;

5) Attachment targets near each module retention device if a

Manual-Augmented control mode is to be used;

6) A method of transferring module "ready to latch", "latch", and

"unlatched" signals to the servicer control system.

Each MMS is fitted with electrical connectors that are mateable with

corresponding connectors in the flight support system. The MMS

electrical power distribution and attitude control system can be

controlled through these connectors. While these berthing connectors

are not easily reachable by the servicer, it might be practical to

parallel the necessary leads to connectors at the grapple fixtures.

3-9
The provision of electrical connections between the MMS and the

servicer for control, power and status is the most significant

requirement on MMS spacecraft. The provision of the various targets

should not be difficult.

In the prior study, it was decided to use the MMS module servicing tool

as a basis for the interface between the servicer and the MMS modules.

As the MST is self-contained and designed to be used by an astronaut on

EVA, it will require some modification for use in an operational

servicer configuration. The MST adaptation requirements are:

1) The modified MST shall be compatible with the servicer end effector

interface (including the electrical disconnect). The translation

mechanism for mating/demating the electrical disconnect should be

located on the servicer end effector;

2) The number of electrical connections between the end effector and

the modified MST should be minimized;

3) The hand controls, electronics and power supply of the modified MST

should be in a remote location;

4) The modified MST interfaces with the end effector and the module

should be capable of transmitting all the moments and loads with

adequate margin for stiffness and safety;

5) The modified MST should stand 400 complete cycles of demonstration

operations without failure;

6) The modified MST should clear the TV camera, the lights, and all

other servicer components at all times during MMS servicing;

7) The modified MST should allow full view of the optical targets and

obstruct as little as possible the TV camera field of view;

3-10
8) The modified MST should have a ready-to-latch sensor at the

Interface with the module;

9) The distance between the end effector interface and the MMS module

latch interface shall be 17.25 in. for 0-g demonstrations;

10) The modified MST unit for cargo-bay servicing operations shall be

provided with an EVA override for manual unlatching.

The above requirements are complemented by additional requirements in

the sections on the 1-g, 0-g, and free-flight demonstration and

verification plans and in the next section involving adaptation of the

module servicing tool.

3.2 MST ADAPTATION FOR 1-g

A brief analysis was conducted to evaluate whether it was better to

modify the servicer engineering test unit (ETU) or the module servicing

tool for 1-g operation. It quickly became obvious that the load

carrying capability of the ETU would have to be increased by more than

an order of magnitude unless the MST weight was severely reduced.

Concurrently, it was observed that certain parts of the MST could be

readily removed, certain functions must be remotely located, and that

MST performance (bolt driving torque) could be greatly reduced for the

1-g demonstrations. These observations led to a decision to minimize

changes to the ETU and to make most of the changes to the MST. The

logic and the associated requirements were discussed with Goddard Space

Flight Center personnel and were accepted.

3.2.1 Modified Module Servicing Tool Requirements

The use of the MMS Module Servicing Tool as it might be adapted or

modified for use with the on-orbit servicer system, especially for 1-g

demonstrations was evaluated. The basic premise was that there would

be minimum changes to the MMS functional attachments and to the

Engineering Test Unit. However, an electrical connection interface

3-11
between the ETU end effector and the modified MST, with the connector

mated and demated by a connector positioner on the end effector, would

be provided. For the 1-g demonstrations, an electronic power supply

(driven by 110 vac, 60 Hz power) would be used in place of the battery

system. The following requirements apply to the modified MST as an

on-orbit servicer end effector adapter:

1) The 1-g and 0-g versions of the modified MST may be different but

should have the same ETU and MMS interfaces and approximately the

same overall dimensions;

2) The modified MST shall be compatible with the ETU end effector

interface (including the electrical disconnect). The positioner

mechanism for mating/demating the electrical disconnect should be

located on the ETU end effector;

3) The number of electrical connections between the end effector and

the modified MST should be minimized;

4) The hand controls, electronics and power supply of the modified MST

should be in a remote location;

5) The modified MST for the ground demonstration should have reduced

torque capability and longer operating time in order to save weight;

6) The modified MST interfaces with the end effector and the module

should be capable of transmitting all the moments and loads with

adequate margin for stiffness and safety;

7) The modified MST should stand 400 complete cycles of ground

demonstrations without failure;

8) The modified MST should clear the TV camera, the lights, and all

other servicer components at all times during MMS servicing;

3-12
9) The modified MST should allow full view of the optical targets and

obstruct as little as possible the TV camera field of view;

10) The modified MST should be as compact as practical to allow

maneuvering in volume restricted zones;

11) The modified MST should have a ready-to-latch sensor at the

interface with the module;

12) The distance between the end effector interface and the MMS module

latch interface, for the ground demonstration, shall be 7.25 in.;

13) The weight of the modified MST for the ground demonstrations shall

be less than 15 Ibs;

14) The modified MST unit for cargo^bay servicing demonstrations shall

be provided with an EVA override for manual unlatching;

15) For the 1-g demonstration, MST control should be selectable from

either an MST control panel, mounted in the MSFC control station,

or from a computer;

16) The MST latches would be modified to delete the inward motion

function, and to hold the module tightly during module transfer

operations.

The weight and size requirements for the 1-g version of the modified

MST were derived from an analysis of the ETU drive torque capabilities

(see Section 8.1.2). The critical ETU actuator is the wrist pitch (Y)

drive. The torque required from this drive must not exceed 50 ft-lb to

avoid overheating and to provide an acceptable speed. For a 50 ft-lb

load torque, the drive will take 15 sec to flip a module upside down as

compared to 5 sec with no load. .

The above requirements were discussed with GSFC and Fairchild Space

Company personnel and were used as a basis for their modified MST

design and fabrication activity.

3-13
3.2.2 Modified MST Mechanical Concept Definition

The weight of the existing MST, as configured for EVA use, is

approximately 67 Ibs. The modified MST to be used for ground

demonstrations needs to be redesigned for a drastic weight reduction

and reduction of the distance between the end effector interface and

the module latch interface (called the "B" distance).

For the 0-g MMS servicing demonstration, existing MST hardware can be

used if the battery, battery case, EVA handles and the controls are

removed and a standard ETU end effector interface, including an

electrical disconnect, is added opposite to the tool/latch end. A

distance "B" of 17.25 in. between the two interface reference planes

can be accommodated without modifications to the EVA overrides or tool

drive. Minor modifications to the latch mechanism will be required to

provide a firm grip on the module during exchange operations.

For the 1-g demonstrations, however, this simple adaptation is not

appropriate. In the analyses of Section 8.0, it was determined that

for a module weight of 12.5 Ibs and B = 17.25 in. the maximum allowable

weight of the modified MST would be only 3.5 Ibs, which is not

feasible. Therefore, the shape and the design of the modified MST for

the ground demonstrations should be different from the one used in 0-g.

Several ways of achieving the required light weight and compactness

were identified:

1) Reduction of fastening torque from 170 ft-lb to 20 ft-lb;

2) Longer latch and wrench cycle time to reduce the motors power and

size;

3) Offset motor/gearbox;

4) Use of composite and other light weight materials;

3-14
5) Elimination of the bolt drive extension simple removal feature.

A maximum module mockup weight of 12.5 Ibs and a "B" distance of 7.25

in. were selected for the ground demonstrations. From the Section 8.0

analyses, the maximum allowable weight of the modified MST is 15 Ibs.

An allowance of approximately 0.5 Ib was made for the electrical

connector positioner mechanism to be added to the end effector. Other

applicable requirements are given in Section 3.2.1.

In Figure 3.2-1 the general configuration of the modified MST for the

ground demonstrations, the critical dimensions required for providing

adequate clearances and the approximate e.g. position are given.

MODULE LATCH

INTERFACE

OPTICAL TARGET

AREA

3.0 MIN

END EFFECTOR

INTERFACE

TV CAMERA

"Z" DRIVE

MMS MODULE

MOCKUP

Y" DRIVE

MODIFIED MST (15 LBS)

Figure 3.2-1 Modified MST Configuration for Ground Demonstrations

Adequate clearance between the modified MST and the servicer arm and

the MMS module mockup must be provided. The tool should not obstruct

the field of view of the TV camera, inside of the 10° cone shown,

except for the very tip of the tool. The servicer operator should have

a good view of the module latch interface on the TV screen prior to

tool insertion and of the optical target during module attachment.

3-15
A sketch of one approach to the design of a modified MST for 1-g use

that appeared to satisfy the above requirements was prepared. Parts

were generally sized to provide the desired torque and speed. Separate

motors were used for the bolt and latch drives with one motor driving

both latches through a cogged belt. Copies of the sketch were provided

to GSFC and Fairchild personnel to encourage them that the 15 Ib weight

allowance could be met.

3.2.3 Modified MST Electrical Concept Definition

A conceptual arrangement of the modified MST wiring and its interface

with the Engineering Test Unit was prepared to better define the

interfaces between the four parties involved - Marshall Space Flight

Center, Goddard Space Flight Center, Fairchild Space Company, and

Martin Marietta Corporation. This arrangement was reviewed with the

four parties and accepted as a basis for hardware design and

fabrication. The MSFC responsibility included:

1) Computer interface equipment;

2) MSFC breakout box changes;

3) Location of MST control panel.

The GSFC/Fairchild responsibility included:

1) MST wiring;

2) MST electronics box;

3) MST power supply;

4) MST control panel;

5) Wiring between GSFC provided components (except MST) and between

MST electronics box and existing ETU junction box.

3-16
The Martin Marietta responsibilities included:

1) ETU connector positioner wiring;

2) Cabling changes along ETU mechanism;

3) Changes in existing ETU junction box;

4) Changes in existing Servicer Servo Drive Console wiring;

5) Changes to existing Servicer Control Panel to add connector

positioner functions;

6) Wiring and cabling to MST tool storage rack;

7) Changes in existing ETU cabling;

8) Wiring of a new MMS junction box.

These areas of responsibility and the cabling involved are shown in

Figure 3.2-2 along with the MMS module receptacle wiring suggested for

the MMS module exchange demonstrations. The MST power supply is to be

plugged into an existing 60 Hz receptacle in the Servicer Servo Drive

Console (SSDC) that is switched off when the ETU main power is turned

off. In this way, it will be possible to shut down the entire MMS 1-g

demonstration equipment (except for computer) by operating one switch.

The arrangement shown in the figure requires that cable connector P2 be

moved from the existing ETU junction box to the new MMS junction box

when MMS module exchange demonstrations are to be made. An earlier

alternative required the exchange of six basic module location

connectors with six MMS module location connectors. However, the

earlier design would have taken too much time between module exchange

demonstrations so the version shown in the figure was used.

3-17
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3.3 MODULE EXCHANGE OPTIONS

The results of the prior study with respect to adapting the servicer to

the MMS were reviewed to determine if any simplifications or increases

in capability could be identified. The primary areas of interest were

the servicer configuration including the docking probe and alternative

modules that could be involved. The basic approach is to use the

Integrated Orbital Servicing System (IOSS) concept as the reference and

then to modify the reference as necessary for MMS servicing. The

acronym IOSS is used to denote both the system and the study. The

context of use will generally indicate which meaning was intended. In

this way, the axial and near-radial capabilities of the IOSS for

exchanging modules from 15 ft-diameter spacecraft would be maintained

along with the capability to extend to radial exchange of a second tier

of modules.

3.3.1 Servicer Configuration

The prior study examined seven different alternative methods for

adapting the IOSS servicer to the servicing of MMS spacecraft. The

variations involved docking direction, servicer arm segment lengths,

type of adapter tool between servicer end effector and MMS module,

interface mechanism type, docking probe adapter type, and stowage rack

modifications. The selected configuration involved a docking probe

adapter with a hinge, Engineering Test Unit arm lengths and a modified

form of the MMS module servicing tool. The selected configuration is

shown in Figure 3-4.

3.3.1.1 Prior Study Results - The seven alternatives and three

variations listed in Table 3.3-1 were addressed in the prior study to

select an adaptation of the IOSS servicer to service MMS spacecraft.

All of the alternatives used the 45 in. lower and upper arm lengths of

the servicer Engineering Test Unit rather than the 79 in. arm lengths

of the IOSS. The ETU lengths were used because the prior study started

with the 1-g equipment and evaluated that equipment for the cargo-bay

demonstrations and the operational situation. The results of the prior

3-19
study servicer configuration trade study are not affected by a change

to the longer arm lengths of the IOSS. Some of the details may change,

but the results should be the same.

Table 3.3-1 MMS Module Servicing Method Alternatives

Axial Docking Methods

1) Modified servicer end effector and specialized adapter tool

2) Use of existing side interface mechanism

3) Use of alternative interface mechanisms

a) Single power takeoff

b) Dual power takeoff

c) Latches directly actuated with electric motors

4) Use of one latch mechanism in back of modified MMS module

5) Use of one active latch at bottom of modified MMS module and a

passive hook-up point at the top

Lateral Docking Methods

6) Use of an offset docking probe adapter and tool adapter

7) Use of straight docking probe adapter, tool adapter and modified

stowage rack.

The prior study recommended the use of alternative 7) of Table 3.3-1.

The variety of MMS configurations to be addressed was a driving

consideration. In particular, the solar arrays of the Solar Maximum

Mission and the Mark II propulsion module were difficult to work

around. Those servicer configurations involving docking at one end of

the MMS and using the MMS berthing pins for docking were cumbersome and

were judged to have a low probability of successful docking. These

criteria ruled out the first five candidates, including the variants,

of Table 3.3-1. A list of the criteria used for servicer configuration

selection is given in Table 3.3-2. There are few criteria associated

with changes to the MMS as it was strongly desired to avoid any

significant changes to the MMS. Some changes will be necessary as

noted in Section 3.1. The large number of criteria listed under

3-20
Impacts on Servicer Design result from a willingness to look at

different servicer alternatives. Alternative 7) of Table 3.3-1 was

selected over Alternative 6) because it provided better clearance

between the servicer and the MMS, has better accuracy and is stiffer,

has an easier-to-use docking system, is mechanically simpler, and is

lighter; even though the Alternative 7) stowage rack must be modified.

Table 3.3-2 Servicer Configuration Evaluation Criteria

General

Ability to service MM spacecraft fitted with Mark II propulsion

system

Loss of capability to service other spacecraft types

Use of three berthing pins for docking

Use of proven module latch mechanism

Ease of operation of both MMS module attachment bolts

Number of MMS module attachmen-t bolt engagements

Adaptability to 1-g demonstrations

Applicability to MMS type satellites currently in orbit

Impacts on MMS Design

- MMS structure configuration changes

- MMS module configuration changes

- Module servicing tool changes

- Weight increases

Impacts on Servicer Design

- ETU mechanism configuration changes

Inability to use standard interfaces

- Docking system used

- Docking system stiffness and accuracy

Increases in operational complexity

Increases in mechanical complexity

- Decreases in servicer dexterity

- Weight increases

Need for docking probe adapter

- Need for stowage rack modifications

- Arm segment length increases

- Servicer arm stiffness and accuracy

3-21
Both alternatives require that electrical connections be made between

the servicer and the module servicing tool and between the servicer and

the MMS. One, and preferably two, electrical grapple fixtures must be

mounted on the MMS structure for docking. The module servicing tool

must be modified to move its control functions to a remote location and

to interface mechanically and electrically with the servicer.

3.3.1.2 MMS Docking Probe Adapter - The MMS docking probe adapter is shown in

Figure 3.3-1 installed in the IOSS docking probe. Both docking probes

are adaptations of the Orbiter RMS standard end effector. The hinge is

used so that the two docking probes can be colinear for docking with an

MMS and can be placed at the proper orientation for MMS module

exchange. When the MMS docking probe is oriented to the other side of

the servicer centerline, then a second MMS module can be exchanged.

The hinge also allows the MMS docking probe to be folded down for

storage during launch in the Orbiter cargo bay.

The concept of a docking probe adapter was selected, rather than a

bolt-on connection, so that the adapter could be installed, or removed,

during a servicing mission. This permits the IOSS to service both an

MMS and a serviceable spacecraft with standard side interface

mechanisms on the same mission (mixed spacecraft missions). The

servicer arm has the reach and controllability to remove and install

the MMS docking probe adapter. The IOSS is also capable of being used

from the Space Station. The MMS docking probe adapter concept of

Figure 3.3-1 would simplify its installation by astronauts on EVA for

MMS servicing missions.

The work on the cargo bay demonstrations of Section 6.0 indicated that

the use of two docking probes would increase the flexibility and

inaccuracy of the connection between the spacecraft and servicer.

Unlike the steady state misalignments, which can be measured and

accounted for in the servicer computer, the flexibility effects are not

so easy to handle. The RMS standard end effector has a connector with

a limited number of pins. The IOSS docking probe must handle all of

the signals between the servicer and the spacecraft, plus the signals

3-22
MMS

DOCKING

PROBE

ADAPTER

MMS DOCKING INTERFACE

MMS DOCKING PROBE

HINGE

IOSS DOCKING INTERFACE

IOSS DOCKING PROBE

SERVICER MECHANISM

STOWAGE RACK INTERFACE

Figure 3.3-1 MMS Docking Probe Adapter

to and from the hinge joint. A rough estimate of the number of wires

needed indicates that the standard RMS connector is marginal.

These considerations lead to a recommendation that the IOSS docking

probe be replaced by a separation collar as shown in Figure 3.3-2. The

separation collar would be designed for easy operation by an astronaut

on EVA. The pin limitation of the RMS connector would no longer be of

concern. The separation collar would permit installation of either the

MMS type docking probe adapter or the IOSS docking probe. The ability

to conduct mixed spacecraft missions would be lost. However, a much

3-23
MMS DOCKING INTERFACE

REVISED

MMS

DOCKING

PROBE

ADAPTER

HINGE

SEPARATION COLLAR

SERVICER MECHANISM

t \Vin

> STOWAGE RACK INTERFAC

Figure 3.3-2 Revised MMS Docking Probe Adapter

stiffer structure would result. It is recommended that the revised MMS

docking probe adapter be used until a firm requirement to conduct mixed

spacecraft missions is identified.

3.3.1.3 Docking Probe Hinge - Because the MMS docking probe hinge, or

orientation joint, represents an additional cost, its deletion was

reviewed. An approach to servicing the Solar Maximum Mission

Spacecraft, without using the docking probe hinge, was identified and

evaluated. The conclusion of this analysis was to retain the docking

probe hinge as the advantages of its removal do not clearly outweigh

the corresponding disadvantages.

3-24
The baseline servicer configuration for servicing the MMS in 0-g is

shown in Figure 3.3-3. The servicer is fitted with a straight docking

probe adapter, a modified Module Servicing Tool (MST) and a modified

stowage rack. The servicer docks with the MMS laterally, on its

existing electrical grapple fixture or on a grapple fixture/berthing

pin combination that replaces an existing berthing pin. An orientation

joint similar in design to the other servicer joints is included in the

docking probe adapter to allow tilting of the servicer with respect to

the MMS after docking to bring the servicer mechanism into a plane

parallel to the face of the module to be exchanged. The joint is

powered through an electrical connection across the servicer docking

interface. This feature permits the simple, axial mode of operation of

the servicer without modifying its basic configuration. Either one of

the two modules adjacent to a grapple fixture can be serviced in one

docking. Adequate clearance is provided between the servicer and the

solar panels or other spacecraft elements.

For MMS servicing missions the standard docking probe would be detached

from the servicer at the servicer mechanism interface and replaced by a

MMS type docking probe that is longer and includes the orientation

joint. The changeover to the MMS docking probe can be done at the same

time as the stowage rack is set up for the MMS servicing mission. No

modifications of the MMS modules or module retention system (MRS) are

required. Instead, a modified MST compatible with the existing MRS and

with the servicer standard end effector interface is recommended.

An alternative MMS servicing configuration for 0-g is shown in Figure

3.3-4. The servicer docks with a Solar Maximum Mission version of the

MMS laterally on its existing electrical grapple fixture or on an added

grapple fixture/ berthing pin combination that replaces an existing

berthing pin. Unlike the baseline configuration, the docking probe

does not have an orientation joint. A simple, straight docking probe,

with an RMS end effector is used. Both modules adjacent to a grapple

fixture can be serviced in one docking. Several modifications to the

servicer mechanism and its control software are necessary in order to

increase the axial reach and allow operating the two fasteners of each

3-25
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3-26

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3-27
module with the end effector at a 60° angle relative to the docking

probe. As in the baseline configuration a modified MST is used to

adapt the standard end effector of the servicer arm to the standard MMS

module retention system.

The stowage rack is modified by eliminating one modular beam in order

to clear the spacecraft solar panels. Both segments of the docking

probe, between the stowage rack and the arm and between the arm and the

RMS end effector are longer in order to increase the arm reach in the

axial direction and to provide clearance between the stowage rack and

spacecraft. These modifications can be implemented at the same time as

the stowage rack is set up for the MMS servicing mission.

The tradeoff study on elimination of the orientation joint, or hinge,

from the MMS docking probe involved layout preparation and analysis of

interferences of the servicer with -the solar arrays and other

spacecraft elements. To minimize interference between the Solar

Maximum Mission solar array and the OMV, the docking post must be

reworked to put the shoulder pitch joint halfway along the lengthened

docking post. This increases the number of changes involved in going

from a basic to an MMS servicing mission. A modified servicer

mechanism, with increased shoulder pitch joint travel, can reach the

MMS module latches. While the 45 in. arm segment length of the ETU

could be used, the long sides of the parallelogram would have to be

spaced farther apart because of the larger shoulder pitch joint angular

travel required. Figure 3.3-4 shows an arm segment length between that

of the ETU and that of the IOSS.

While the above mechanical changes permit the servicer mechanism end

effector to reach all points along the desired paths, certain control

system complexities arise when combined motions are involved. Module

exchange as shown in the figure means that the MMS module must move

straight along a line at a 60° angle with respect to the docking post.

3-28
This is a form of combined (axial and radial) motion. The servicer

control and trajectory generation system has been developed so that it

is only necessary to change one cylindrical coordinate at a time when a

module is being moved. For combined motion, it is necessary to change

more than one cylindrical coordinate at the same time and to

synchronize the variations in the three cylindrical coordinates. While

this is generally possible, an extra level of complexity is added to

the module exchange trajectory generation system. Additionally, in

order to orient the end effector properly while moving the MST along

the centerline of an MMS module fastener, all three wrist joints of the

arm must also be actuated at the same time at varying rates. Control

software modifications are needed to perform these more complex

maneuvers.

The Manual-Direct (joint-by-joint) control mode is more difficult,

almost impossible, to apply because, of the need for joint

synchronization. Without this simple back-up control mode, the

reliability of the servicer system is lower, as compared to the

baseline. One possible solution may be to redesign the wrist yaw and

shoulder pitch joints for backdrivability.

The required arm modifications needed for reaching the module were

defined and some of the equations governing the arm motion were

established. Important software modifications would be required as

well as design changes affecting the shoulder segment parallelogram

arrangement. The time allocated to equation derivation did not permit

us to obtain an explicit solution for each of the joint angles. The

solution form was implicit, which means that an iterative solution

approach would be required. The solution is complicated by the fact

that multiple solutions are possible at certain steps because of the

range of allowable angles involved. All potential solutions must be

followed and checked until the correct solution is identified. Once a

correct solution is identified for a particular trajectory action, then

subsequent solutions can use the first solution to simplify the

solution "tree" because the functions are continuous and the servicer

does not move far between solutions. It may also be that the iteration

3-29
converges quickly. However, with additional effort, it is likely that

a better approach to solving the equations can be found. It is clear

that the resulting equation solution technique will be more complex

than that used for the simple radial and axial exchange used. The

radial and axial motion equations, each involve only four degrees of

freedom as opposed to the six involved in combined motions.

The advantages of removing the orientation joint from the docking post

do not clearly outweigh the disadvantages that have been identified.

These servicer system modifications address only the servicing of one

form of spacecraft utilizing MMS modules. Servicing the MMS module in

an axial mode can be applicable to other spacecraft, as well as MMS.

Additional analyses are required to further define the hardware and

software modifications and the reliability aspects of the orientation

joint elimination.

For the ground and cargo-bay MMS servicing demonstrations, the baseline

configuration of the servicer shown in Figure 3.3-3 is recommended.

3.3.1.4 Servicer to MMS Electrical Connections - Certain functions on MMS and

other spacecraft being serviced on orbit need to be safed or controlled

during the servicing operations. While some of these functions might

be controlled from the ground if the spacecraft communications system

was working, a more direct approach is to provide an umbilical

connection between the servicer and the spacecraft. Representative

functions that will require control include:

1) The spacecraft attitude control system should have its actuators

(thrusters) turned off so they do not fight the OMV attitude

control system and waste propellant;

2) Orientable appendages, e.g., solar panels or communication

antennas, should be fixed in position if they could physically

interfere with the servicing operation;

3-30
3) Flow of fluids in lines to and from quick disconnects and

associated purging functions will need to be controlled;

4) Pressures, temperatures, and fluid flow rates in and out of tanks

will need to be controlled;

5) Electrical power will need to be provided to continue certain

functions or to keep modules warm while spacecraft power is off

during spacecraft component replacement;

6) Electrical power to modules being replaced should be controlled to

prevent arcing between pins as module electrical connectors are

disengaged;

7) Preliminary checkout of spacecraft after modules are replaced

including a check that connector continuity exists;

8) Provision to the servicer of ready-to-latch, latched, and unlatched

signals for each module location. These signals are used in the

servicer control system.

The concept of an electrical umbilical connection between the servicer

and the spacecraft being serviced is not new. The MMS flight support

system (FSS) has two umbilical connectors that can be remotely

actuated. These umbilical connections were used during the Solar

Maximum Repair Mission and are used during normal launch operations for

the above types of functions as well as for caution and warning

functions. MMS functions can be controlled and monitored on the

Orbiter or on the ground by using the Orbiter communications system.

The FSS umbilical connector actuators can be controlled from the

Orbiter along with other FSS functions.

The recommended approach for on-orbit servicer operations is to use the

electrical connector on the RMS end effector used for servicer docking

to provide the necessary connections. Because of the limited number of

pins in this connector, it will be necessary to use signal multiplexing

3-31
to obtain enough control and sensing functions. Spacecraft control

would be from the ground, or Space Station, using the OMV

communications link with the signals passing through the servicer. For

MMS vehicles the RMS grapple fixture (servicer docking system)

connections should be integrated with the FSS umbilical connection

signals so that the spacecraft could be controlled through either

connection.

For those cases where the pins in the RMS end effector connector are

insufficient in number or in current-carrying capacity, the servicer

mechanism can be used to mate an auxiliary umbilical connection between

the servicer and the spacecraft. This auxiliary electrical umbilical

connection could use some of the concepts of the fluid resupply

umbilical connection approach.

3.3.2 Alternative Modules

The emphasis given to MMS modules and spacecraft should not be

Interpreted to imply that less emphasis has been placed on the full

range of capabilities for which the flight version of the IOSS was

designed. The Engineering Test Unit is a small (45 in. arm segment

lengths) version of the flight design (79 in. arm segment lengths).

The flight servicer was designed to exchange up to 40 in. cube and 700

Ib modules in axial and radial directions where the radial attachment

point was near the docking end of the spacecraft. The flight servicer

capability can be extended to removal of two tiers of modules in a

radial direction when the need is established. Table 3.3-3 lists a few

of the on-orbit servicer characteristics along with the accommodations

required from a carrier vehicle.

3-32
Table 3.3-3 On-Orbit Servicer Characteristics

Compatible with operations at/with Orbiter, OMV, OMV/OTV, Space

Station

Multiple spacecraft servicing per mission

Axial module replacement

Radial module replacement — attach locations in a common plane

Maximum operating radius - 7.5 ft radial, 11.2 ft axial

Module mass - 10 to 700 Ibs

Module size - 17 in. cube to 40 in. cube

Provides failed module temporary stowage

Degrees of freedom - 6

Stowed length - 27 in.

Tip force > 20 Ibs

Latch actuator located in end effector

Time to replace one module - 10 minutes

Compatible with automated, supervisory and remotely manned control

Accommodations required from carrier vehicle

- Transport capability

- Rendezvous and docking

- Electrical power

Two way communications

- Attitude control

- Assistance in thermal control

- Data processing

The on-orbit servicer design requirements were developed from an

extensive review and analysis of servicer mechanism configurations and

28 serviceable spacecraft configurations from the literature available

in 1975. Ten preliminary servicer configurations were analyzed, mocked

up with a one-tenth scale model of a spacecraft and stowage rack, and

evaluated. The initial recommendation was judged to be too capable and

not simple enough. A major point in the selection was that the

eventual form of servicer that will become accepted and used could not

be identified, so the selection had to be made on the basis of a best

estimate of the probable situation. It was decided to go with a

relatively simple configuration that has natural and easy growth

options.

3-33
The selected design has only two major components: (1) a servicer

mechanism, and (2) a stowage rack for module transport. The servicer

mechanism and the stowage rack were designed separately with interfaces

for individual removal and replacement. This allows for simple removal

for maintenance and also for quick ground reconfiguration. Stowage

racks can be configured and loaded for particular flights prior to

attachment to the carrier vehicle. It may be desirable to have

available several stowage racks for this purpose. The stowage rack

mounts directly to an upper stage such as the Orbital Maneuvering

Vehicle. A flight support structure has been designed to adapt the

stowage rack to the Orbiter.

Figure 3.3-5 is a histogram of data on 683 modules from 30 different

serviceable spacecraft. One graph shows the average weight vs largest

module dimension and the second graph shows the 90 percentile weight vs

largest module dimension. This data was used to select representative

module size and weight sets as indicated by the design value curve.

The design value sets are:

1) 40 in. cube - up to 400 Ib;

2) 26 in. cube - up to 200 Ib;

3) 17 in. cube - up to 75 Ib.

These sets were used to establish reference characteristics for

different interface mechanism sizes. The interface mechanism provides

the structural attachment between a module and the stowage rack. It

also provides the alignment and mating/demating forces for the

electrical connectors. The above data suggested the development of an

interface mechanism as a two-part kit in perhaps three sizes. These

standard interface mechanisms could be made available to spacecraft

designers. Each designer could then make his choice within his own set

of design and economic constraints. The recommended interface

mechanism standard sizes thus became - 17 in., 26 in., and 40 in.

3-34
100

70

40

ca

ca

CO

00

UJ

20

10

DESIGN VALUE

0.0055 LBS/IN.

10 20 40

MODULE WEIGHT (IBS)

70 100 200 400 700

Figure 3.3-5 Module Characteristics Histogram

These correspond to modules no larger than a cube of the indicated

dimension. As the interfaces between the interface mechanism and the

module and the spacecraft both seem to lie within the spacecraft

designer's usual responsibilities, it would be possible to leave these

design aspects up to the spacecraft designer. However, the interface

with the servicer mechanism end effector and its mechanical drive

system would have to be standardized across all interface mechanisms.

Similarly, the method for attaching the interface mechanism baseplate

receptacle alternatives into the stowage rack would also have to be

standardized. In this way, a single—or few—stowage rack designs

could be used for all missions.

While the above discussion appears to limit the size and weight of

orbital replaceable units to 40 in. cubes and 400 Ibs, this is not

really the case. Larger and heavier modules can be handled if they are

advantageously located with respect to the servicer's operating

volume. The spacecraft designer should look at the options of reducing

module size and weight or of locating the module and its attachment

points for easier handling by the servicer. In some cases, it may be

3-35
necessary to go to a seven degree of freedom growth version of the

servicer. Using these approaches, it should be possible to exchange

instruments comparable in size to those on the Space Telescope.

A number of other alternative modules were identified in the Spacecraft

Servicing Demonstration Plan study. They are listed in Table 3.3-4 and

sketches of some can be found in the referenced study. The battery

module is used to represent a small heavy module. Batteries will need

to be replaced because of their limited and somewhat unpredictable

lifetimes. The fluid tank module with an in-line coupling is an

alternative to the fluid umbilical interface unit. Replacing an empty

tank with a full one may be appropriate for smaller tanks or where it

is difficult to obtain required high pressures on orbit. The in-line

coupling term is used to indicate that the coupling must seal for long

periods and that a quick disconnect may not provide adequate sealing.

The in-line coupling would be tightened using a mechanical drive. Two

types of umbilical interface units are listed - fluid and electrical.

They are similar and they both require connections (cables or hoses)

back to the stowage rack that must be managed.

Table 3.3-4 Candidate Alternative Modules

Battery module

Fluid tank module with in-line coupling

Electrical umbilical interface unit

Fluid umbilical interface unit

Access door

Hinged access cover drive

Combinations of electrical connection and refueling/resupply umbilicals

have also been proposed. While small electrical connectors may be

mated using a simple interface mechanism, large electrical connectors

and the fluid disconnects will likely require a translation device to

provide the high mating and demating forces required. Dust covers with

their removal mechanisms may be required on both the spacecraft and

servicer sides of the fluid and electrical umbilical interface units.

3-36
The access door is listed as a module type to show that access covers

or doors can be treated as a module where the interface mechanism is a

special configuration to properly secure the door. This form of access

door would be completely removed and temporarily stored to one side.

The underlying module would then be removed and replaced. The hinged

access cover drive is another approach to using covers over modules to

provide thermal protection. In this case, the cover is hinged to the

satellite and latched down. The servicer end effector attaches to a

fitting on the satellite near the door. The interface mechanism drive,

or end effector power takeoff, is used to power a mechanism that frees

the access cover latches and drives the cover to an open position. The

end effector jaws are then opened and the servicer can be used to

remove the uncovered module in the normal way. After the module has

been replaced, the access door can be driven closed and latched by

using the servicer end effector and interface mechanism drive.

3.4 EXTENDED REMOTE MAINTENANCE/RESUPPLY CONCEPTS

The extension of remote maintenance/resupply can occur in three

directions or regions - applications, orbits, and functions. These are

illustrated schematically in Figure 3.4-1. The Integrated Orbital

Servicing System was initially considered as a front end for the Space

Tug and is now considered as a kit for the Orbital Maneuvering

Vehicle. As such, its primary application has been to free-flying

spacecraft. However, with the advent of the Space Station, the Space

Station itself and its attendant free-flying platforms become potential

areas of application. The connotation of platforms being larger than

spacecraft implies that there will be more docking locations and/or

larger modules associated with platforms. As the IOSS was designed to

work with objects taken to orbit by the Orbiter, and platforms are

taken to orbit by the Orbiter, then it is likely that the IOSS

capabilties will also fit the elements of free-flying platforms. The

slow, deliberate motions of module exchange should minimize Induced

platform accelerations. The potential applications at the Space

Station are extensive and are discussed below.

3-37
APPLICATIONS

/ INTEGRATED ORBITAL

I SERVICING SYSTEM

ORBITS

FUNCTIONS

- SPACECRAFT

- PLATFORMS

- SPACE STATION

- NEAR SPACE STATION WITH OMV

- GEOSYNCHRONOUS WITH OMV/OTV

- OTHER LEO WITH STS/OMV

- REPAIR/MAINTENANCE

- UPGRADING

- RESUPPLY

- PRODUCT RETURN

Figure 3.4-1 Candidate Regions for Extension of Remote

Maintenance/Resupply

The orbit applicability of the Integrated Orbital Servicing System is

dependent on the carrier vehicles available. It is straightforward to

use the Orbiter, OMV and the OMV with OTV as carrier vehicles for the

IOSS. This should provide the orbit coverage listed in the figure

along with the polar orbits that are reachable with the Orbiter/OMV

combination. The polar orbits are particularly significant because

there are not current plans to have a Space Station in polar orbit,

just free-flying platforms. These orbits are the same as have been

extensively considered in the past. It still appears that propellant

costs are too high to consider the extension of on-orbit servicing to

planetary probes. Perhaps if an extensive Lunar presence occurs in the

future, then it may be appropriate to extend servicing to Lunar orbits.

The servicing and maintenance functions that can be performed with

module exchange and umbilical connection are shown in the figure. The

first three functions have been addressed for some time and include

such things as inspection, fault isolation, and clean up. The supply

3-38
of materials for making products is similar to resupply, and can

include solids, powders, or fluids. The new element is the return of

completed product to earth. The IOSS can transfer properly packaged

product to the Space Station or to the Orbiter for return to earth from

free-flying platforms or even between platforms should that become a

requirement. The return of a failed module has much in common with the

return of a processed product except that the product may require

additional environmental control or monitoring functions. This type of

thing should be within the capabilities of an extended servicer system.

The IOSS functions of module exchange and umbilical connection for

electrical signal or fluid transfer are widely applicable to the Space

Station as shown in Figure 3.4-2. The sketch on the left hand side of

the figure is a Martin Marietta concept for servicing of objects that

are brought to the Space Station. Examples of servicing functions that

can be performed by the IOSS are listed on the right.

The IOSS could be involved in the assembly process by bringing modules

to prepared locations on the deployed trusswork. The prepared

locations also make it easy to replace any subsystems that subsequently

fail or otherwise become obsolete. The servicer could also be used as

a small manipulator during the assembly process. The repair of

spacecraft at the Space Station could be similar to the remote

maintenance of spacecraft. However, the servicer would need to be

supported by a mobility unit to move it around in the servicing

facility, particularly to transport replacement modules and special

adapters and tools.

The portable manipulator function is appropriate with the servicer in

the servicing facility and mounted on the mobile remote manipulator

system. The mobile RMS could be fitted with power and communications

links so the servicer could be operated while supported on the mobile

RMS. The concept of an operating servicer on the mobile RMS could also

be applied to the installation and servicing of experiments at the

Space Station. This feature could be valuable for those experiments

3-39
STAGING

AREA

SERVICING FACILITY

MOBILE

RMS

SERVICER STORAGE

FUEL DEPOT

OMV BERTHING RING

• ASSEMBLY PROCESS

• REPAIR OF SPACECRAFT

• PORTABLE MANIPULATOR

• EXPERIMENT INSTALLATION WITH MRMS

• EXPERIMENT SERVICING WITH MRMS

• FLUID' UMBILICAL CONNECTIONS

• REPAIR OF SPACE STATION

• REPAIR OF OMV

• REPAIR OF OTV

Figure 3.4-2 Candidate Uses of On-orbit Servicer at Space Station

located on the Space Station, but far from the habitable areas. As is

discussed elsewhere in this report, the servicer can be used to

resupply fluid or make umbilical connections. These features could be

used at the Space Station fuel depot. However, the use of dedicated

actuators for the fluid umbilical connection could be more

appropriate. Much depends on the variety of umbilical locations and

types that are likely to be used. This is an obvious area for

standardization.

Another possibility, is to incorporate the IOSS concepts into the warehouses that store replacement modules much as trucks and fork-lifts are

used in terrestrial warehouses. These and similar concepts could be

used to reduce EVA workloads, especially those that are repetitive or

hazardous.

3-40
The repair of the OMV and OTV are also candidates for application of

the servicer as they are just spacecraft and can have many of their

components configured for module exchange. Some repairs will be better

done by EVA because of their type or component locations on the

spacecraft. It may also be useful to configure parts of the Space

Station itself in modular form, especially the functions outside the

habitable areas. The servicer, perhaps on the mobile RMS, could then

be used to replace or resupply external Space Station components.

The configuration of spacecraft to be on-orbit serviceable also

simplifies pre-launch operations. The systems used for fault location

and diagnosis while on orbit can also be applied during pre-launch

checkout. Once a fault has been traced to a specific module, then that

module can be replaced simply and easily. Thus, spacecraft configured

for on-orbit servicing will be much easier to operate with during

pre-launch preparations.

3-41
4.0 DEVELOPMENT PLAN SUMMARY

The objective of this section of the study final report is to summarize

the work reported in MCR-85-1313, Servicer Development Program Plan,

Martin Marietta Aerospace, Denver, CO, July 1985. It also serves as an

introduction to the material in Sections 5.0, 6.0, and 7.0 and shows

how they integrate together.

4.1 DEVELOPMENT PROGRAM PLAN INTRODUCTION

The concept of remote servicing involves the refurbishment of a

spacecraft in its normal orbit without the direct use of EVA

personnel. A typical mission would include rendezvous and docking with

the disabled spacecraft, performing the refurbishing operations and

returning the vehicle to its normal operation. Several spacecraft can

be serviced during a single mission". The basis of our approach is to

use module exchange. It should be thought of in a most general way -

the replacement of one object with another. The objects need not be

the same, nor need they perform the same function. They just need to

have a similar interface. Fluids inside tanks can even be resupplied

in the form of module exchange. This approach can fulfill many

on-orbit servicing needs. It can be applied to the Shuttle, Space

Station, LEO satellites, free-flying platforms, and geosynchronous

satellites.

The objectives of the spacecraft servicing demonstration plan are to

identify all major elements and characteristics of an on-orbit

servicing development program and to integrate them into a coherent

plan. The extent to which these objectives can be met is defined by

the available funding and how well the development program is planned

and executed. The selection of the elements to be included must be

thoroughly thought through to maximize the benefits of the

demonstrations and meet the cost goals.

4-1
Program cost is a critical aspect of this development plan. Innovative

approaches have been used to reduce the cost while maintaining high

technical standards. The ultimate goal of the development program is a

fully verified operational, on-orbit servicing system based on the

module exchange, refueling, and resupply technologies.

The availability of an on-orbit servicing capability can be

convincingly demonstrated to the user community only through flight

tests. The acceptance of on-orbit servicing methods by the spacecraft

designer is also linked to the financial and programmatic commitment of

NASA for timely development of the operational capability.

The development program plan is an improvement on the plan proposed as

part of the final report of the previous contract (NAS8-35496,

Spacecraft Servicing Demonstration Plan, MCR84-1866, Martin Marietta

Corporation, Denver, CO July 1984).- Although the ground demonstrations

and free-flight verification have been modified, the majority of the

effort has been on the cargo-bay demonstration. The major criticism of

the previous plan was the high cost of this demonstration. We have

attempted to lower the cost by better defining the tasks and

concentrating on the primary objectives.

The work on the ground demonstrations and the free-flight verification

consisted of a critical review of the previous plans. The plan was

updated to reflect recent program changes. The schedules and cost

estimates were revised based on these minor changes, which are

described in the sections associated with these two activities.

4.2 1-g DEMONSTRATIONS

The ground demonstrations are designed to develop and demonstrate the

basic techniques of remotely controlled servicing of spacecraft. The

Engineering Test Unit (ETU) of the IOSS was selected as the servicer

mechanism for ground demonstrations based on the results of the

tradeoff study done as part of the prior contract. The actual ETU is

shown In Figure 4.2-1. Details of this selection process as well as

the selection of the related hardware is documented in Section 3 of the

4-2
ORIGINAL PAGE FS

OF POOR QUAirrV

"Spacecraft Servicing Demonstration Plan - Final Report," July 1984

(MCR-84-1866).

Figure 4.2-1 The Engineering Test Unit of the IOSS

Work on the ground demonstrations is a continuation of the ongoing work

with the ETU. Some of the elements included are part of the current

activities and have been included in the plan for completeness. The

cost estimate reflects the work to be done and does not include the

currently funded tasks.

The main role of the servicing ground demonstrations is to support

further flight demonstrations.

4.2.1 1-g Demonstration Objectives

The ground demonstrations are the next step in gaining acceptance of

the concepts of remote servicing of spacecraft. They are a

4-3
continuation of the on-going activities that have developed the

techniques of module exchange. The principal objectives of the

servicer ground demonstrations, using a modified ETU, are:

1) To demonstrate the adaptability and flexibility of the module

exchange concept - This can best be done by demonstrating an

exchange of the MMS module mockup, because it is the only on-orbit

serviceable spacecraft modular concept that is operational and it

was designed for a different servicing interface. Additional

demonstrations should be conducted to show that the IOSS is a

flexible servicing system, and does not impose significant

constraints on spacecraft design;

2) To evaluate approaches for the cargo-bay demonstration and freeflight verification - The ground demonstration will be used to

select the control strategies to be used in the demonstration and

verification tasks. Operating procedures can be varied and checked

for overall effectiveness in the use of available resources. These

types of issues are best answered with relatively inexpensive

ground testing rather than expensive flight experiments;

3) To demonstrate the use of the ground servicer as a laboratory tool

- The development of new servicing concepts, new hardware, and

software before further flight testing and operational

implementation can be readily investigated on the ground. A good

example is the development of a satellite remote propellant

resupply capability.

If problems arise during the flight tests or operational servicing,

the ground demonstration unit could be used for finding and

checking out solutions;

4) To demonstrate the use of the ground servicer as a training

facility - Training of the operators for the flight demonstrations

as well as for actual servicing operations can be done using the

ground servicer system.

4-4
4.2.2 1-g Demonstration Issues

Table 4.2-1 lists the issues that were considered in selecting the

approach for the ground demonstrations. These tasks include a broad

range of activities that could be done with the ETU. These issues, or

elements, are described in Section 5.0. The table should be considered

as a shopping list of activities that could be done if there was

sufficient funding, as opposed to what must be done in order to meet

the requirements.

Table 4.2-1 Candidate 1-g Demonstration Issues

Multi-Mission Modular Spacecraft Module Exchange

Fluid Resupply Demonstrations

Control Mode Analysis and Testing

AXAF Module Exchange Demonstrations

Battery Module Exchange Demonstration

Automatic Target Recognition and Error Correction

Engineering Test Unit Electronics Update

Conversion of ETU Control System from Analog to Digital

Tank Exchange

Interface Mechanism Development

Adapter Tools for Specialized Tasks

4.2.3 1-g Demonstration Recommended Approach

The activities listed in Table 4.2-1 were reviewed and screened. Those

selected for the baseline approach are necessary to support the

cargo-bay demonstration and free-flight verification. The activities

selected are shown in Table 4.2-2.

4-5
Table 4.2-2 Recommended Activities for the Ground Demonstrations

Activity Description

MMS Module Exchange*

Fluid Resupply Demonstrations

Control Mode Analysis and Testing

Battery Module Exchange Demonstration

Automatic Target Recognition and Error Correction

* Being worked as part of the current contract

The three primary servicer functions are included in the recommended

activities. These include the exchange of modules in the axial and

radial directions and the demonstration of fluid resupply operations.

The exchange of the MMS modules was selected over the exchange of the

AXAF modules as it was felt to be a more widely accepted application.

There are MMS satellites currently being flown, such as the Solar

Maximum Mission and Landsat. By showing that the servicer can adapt to

an existing design, its flexibility will be demonstrated.

The battery module was selected to demonstrate radial exchange as it is

a practical application of servicing. The normal servicing mission

will include the exchange of failed modules (as demonstrated with the

exchange of the MMS module) and the routine replacement of components

that have a limited life. A battery module is a primary example of the

second case. The exchange of a battery module also demonstrates the

demating and mating of an electrical connector that carries a

substantial amount of current. This type of connection can cause

arcing problems if not properly deenergized.

The fluid resupply demonstration was selected as it is one of the

primary servicing functions. Its demonstration is timely, as a number

of fluid disconnect interface units are being developed and they can be

adapted to operate with the servicer.

4-6
The activities of the control mode analysis and the automatic target

recognition system will advance the design of the servicer to a point

that it is possible to have an operational unit. The control mode

analyses will improve this system and lessen the risk of failure. By

better understanding the controls of the servicer, future versions can

be improved. The automatic target recognition and error correction

system will permit automatic operation of the servicer. It will also

lead to a reduction in cost by permitting a less accurate docking

mechanism to be used and compensated for with the error correction

function.

The analyses and demonstrations identified in Table 4.2-3 will enhance

the results, but were felt to be nonessential to meeting the primary

objectives of a cost conscious program. Many of these tasks may be

funded as part of other programs. It is felt that the ETU servicer

will be upgraded to a digital system if the cargo-bay demonstrations

are carried out. The tasks such as the tank exchange and special tools

and interfaces will be included as part of the programs requiring their

use. It is beyond the scope of this program to accommodate every case

of remote servicing.

Table 4.2-3 Optional Activities for the Ground Demonstrations

Activity Description

AXAF Module Exchange Demonstrations

Engineering Test Unit Electronics Update

Convert ETU Control System from Analog to Digital

Tank Exchange

Interface Mechanism Development

Adapter Tools for Specialized Tasks

4.2.4 1-g Demonstration Schedule and Cost Estimates

The schedules and cost estimates were updated from those developed

during the prior contract. Data is presented for the activities in the

4-7
baseline approach. These activities are designed to support the

demonstration of remotely controlled module exchange and fluid resupply

in the Orbiter cargo-bay.

The schedule for the ground demonstrations is shown in Figure 4.2-2.

The OMV design and development schedule is shown as a reference. The

MMS modification and module exchange activities shown on the schedule

will be completed as part of the current contract. The control mode

activities will cover a period of approximately one year. The schedule

permits the results to be used in establishing the approach to the

cargo-bay demonstration flight. The battery exchange demonstration,

the refueling demonstration, and the automatic target recognition and

error correction activities will run concurrent with the control modes

studies. These elements will involve the design and fabrication of the

necessary hardware, integration into the ETU and a short period of

checkout and demonstration. There should be no conflict in schedules

due to the availability of the ETU as each element should only require

a few weeks of actual testing.

The last ground demonstration element is supporting the cargo-bay

demonstration flight. This includes operations simulations, flight

training and problem solving.

Cost estimates were made for each of the elements of the recommended

approach for the ground demonstrations. These are engineering

estimates based on our previous experience in preparing for

demonstrations with the ETU. The estimates are based on the schedule

of Figure 4.2-2. The estimates can be revised to reflect changes in

the schedule.

The cost estimates for the ground demonstrations are presented in Table

4.2-4. As previously noted, the costs for the MMS modifications and

module exchange demonstrations are included as part of the current

funding. The cost for supporting the cargo-bay demonstrations is

included as part of that task.

4-8
DESCRIPTION

OMV Design &

Development

Multi-Mission Modular

Spacecraft Modifications

(Spacecraft Mockup and

Stowage Rack)

Module Exchange

Control Mode Analysis/

Testing

Battery Exchange

Demonstration

Resupply Demonstration

Automatic Target

Recognition

Support for Cargo-Bay

Demonstration

Cargo-Bay Demonstration

1985

i

PHASE B

I I

BASIC MMS

TV7 r"7

 VI

1986

ATP

 POR«TT

fv v^

I

DEMONSTRAT

n

V

DEMO

r— f I V

1987

cm

v^

PHASE C/D

EON

ISTRATION

DEMONSTRATION

I ^

1

DEVELOPMENT

1988

TRAINING 4

OPERATIONS

DEMONSTRATION

FLIGHT r_,

V

Figure 4.2-2 Ground Demonstrations Schedule

Table 4.2-4 Estimated Costs of the Ground Demonstrations

Description

Control Mode Analysis/Testing

Battery Module Exchange

Fluid Resupply Demonstration

Automatic Target Recognition

Total Estimated Cost

Estimated Cost*

250

250

300

150

950

* In Thousands of 1985 Dollars

4-9
4.3 CARGO-BAY DEMONSTRATION

An experimental version of the spacecraft servicer will be demonstrated

in the zero-g environment of the Orbiter cargo-bay. An artists concept

of the demonstration is shown in Figure 4.3-1. A concerted effort was

made to reduce the cost of this demonstration. For example our

approach in the prior contract was to mount the servicer and stowage

rack on the flight support system of the MMS and to use a generic

spacecraft. It was suggested by Goddard Space Flight Center that we

use the MMS as our spacecraft and to mount it on the flight support

system. This approach helps reduce the cost by making use of existing

equipment. A mockup spacecraft consisting of a MMS triangular support

structure, a MMS module, a battery module and a fluid storage tank will

be serviced. The Orbiter RMS will be used to place the MMS mockup on

the servicer and return it when the experiment is completed. The

experiment will be controlled from the Orbiter's -aft flight deck by an

astronaut using a minicomputer.

Figure 4.3-1 Artists Concept of the Cargo-Bay Demonstration

4-10
The revisions to this task represent the majority of the work done on

this version of the development plan. The work has involved reducing

the cost of the program while still including the major servicing

functions.

4.3.1 Cargo-Bay Demonstration Objectives

The cargo-bay demonstrations are to demonstrate the servicing

techniques developed on the ground in the controlled 0-g environment of

the Orbiter. This demonstration is a key step in the process of

convincing the user that remote servicing is a viable technology. The

principal objectives of the servicer cargo-bay demonstration are:

1) To enhance users acceptance of on-orbit servicing by demonstrating

the servicer in the zero-g environment of the Orbiter cargo-bay.

The users are likely to look favorably on the concept if they are

shown there are no anomalies in the zero-g environment. It is

important to demonstrate the key servicing operations;

2) To incorporate representative servicing equipment of the type that

could be used for the free-flight verification. The degree to

which the equipment is operational will have an effect on the user

and his acceptance of the servicing concept;

3) To limit risk of failure by making extensive use of procedure

verification, analysis and 1-g simulations. A failure, even if it

were not directly related to the servicer, would hurt its

acceptance by the users;

4) To demonstrate the servicers adaptability to change by including

designs that show the servicers flexibility to accommodate changes

due to technology advances or variations in user requirements;

5) To show the compatibility of the OMV and servicer schedules. The

development of the servicer should correspond with the development

4-11
schedule of the OMV. A primary function of the OMV will be to repair

and maintain spacecraft using the servicer. It is important that the

development schedule of the servicer is such as to complement the OMV

development schedule.

4.3.2 Cargo-Bay Demonstration Issues

The major issues of remote servicing were considered when the approach

for the cargo-bay demonstration was established. The ideal situation

would be to address all of the issues and demonstrate the full

capabilities of the servicer. A representative list of what might be

included is presented in Table 4.3-1. The characteristics of the

different categories are summarized in Section 6.0.

Table 4.3-1 Candidate Cargo-Bay Demonstration Activities

Module Type

- Battery module

Multi-mission Modular Spacecraft type module

- Propellant tank module

Electrical connection interface unit

Propellant resupply module with interface unit

- Access door

- Electrical connector

- Fluid in-line coupling

- Wave guide connector

- Fiber optics connector

- Thermal connector

Interface Mechanism Type

Lightweight side interface mechanism

Alternative interface mechanism concepts

- Hinged access cover drive

Special Tools

- MMS module servicing tool

Other interchangeable adapter tools

Refueling/resupply interface unit

- Hose or cable management device

- Propellant in-line coupling drive

Direction of Module Motion

Near axial

- Far axial

- Near radial

- Compound motions

4-12
4.3.3 Cargo-Bay Demonstration Recommended Approach

The approach used for the cargo-bay demonstration attempts to match the

technical requirements with the cost constraints. The prior study,

NAS8-35496, recommended two separate Orbiter flights and a cargo-bay

arrangement with on-orbit servicing equipment attached to the MMS

flight support system. The first flight was to demonstrate module

exchange in a variety of ways while the second flight was to

demonstrate the resupply of fluids. A disk shaped spacecraft mockup

was recommended to support the modules to be exchanged. The resulting

0-g flight demonstrations were judged to be too expensive.

The emphasis has been changed for this study to a minimum cost 0-g

demonstration that incorporates a few important servicing functions.

However, a range of additional servicing activities were identified

that can be incorporated into the plan. The demonstration of MMS

servicing also received increased emphasis in this study. The general

characteristics of the recommended approach are listed in Table 4.3-2.

Table 4.3-2 Servicer Cargo-Bay Demonstration Characteristics

One demonstration flight

MMS triangular structure for spacecraft mockup

Axial MMS module exchange

Radial battery module exchange

Propellant transfer to PM-1 propulsion module

MMS to and from orbit on flight support system

MMS dock and undock by RMS

Supply of power, attitude control, thermal control, and

communications by Orbiter

Servicer control station in Orbiter

Unassisted Supervisory control mode

Docking rigidization by servicer docking probe

Electrical connection between servicer and spacecraft via the

docking mechanism

Servicing equipment performance demonstration

Man-machine interactions included

Compliance with Orbiter system safety requirements

Use of representative servicing operational equipment

Operator training

Servicer docking probe normal to Orbiter wing plane

4-13
A single flight experiment will be used to demonstrate the axial

exchange of a MMS module, the radial exchange of a battery module, and

fluid transfer. The servicer and stowage rack will be stored and

operated in the cargo-bay. The spacecraft mockup will consist of a MMS

triangular support structure, a MMS module, a battery module and a PM-1

propulsion module. The spacecraft will be stored in the cargo-bay on a

Flight Support System (FSS) provided by GSFC along with the MMS

hardware. The Orbiter RMS will be used to remove the spacecraft from

the FSS and to dock the spacecraft to the servicer.

The arm of the servicer will unlatch and remove the MMS module from the

spacecraft and store it on the stowage rack. A replacement module will

be taken from the stowage rack and installed on the spacecraft mockup.

The MMS module from the spacecraft is then moved from the temporary

location on the stowage rack to the location previously occupied by the

replacement module. Once secured, the process is repeated for the

battery module. The difference in the process is that the MMS modules

are removed axially and the battery modules are removed radially from

the spacecraft mockup.

The servicer will then get the fluid resupply interface from the

stowage rack and physically connect it to the spacecraft. A hose

management system will be provided to prevent the fluid and electrical

lines from becoming entangled in the servicer or spacecraft structure.

The fluid connections are mated to an interface near the PM-1

propulsion module using the servicer mechanism. After the mating is

completed a low hazard fluid( e.g., alcohol) will be transferred to the

propulsion module. The fluid resupply interface will then be demated

and returned to the stowage rack. The spacecraft mockup is removed

from the servicer and returned to the FSS using the Orbiter RMS.

The entire process is controlled from the aft flight deck by an

astronaut who uses a small computer terminal to interface with a

mini-computer, which is located on the stowage rack. The operator

4-14
monitors the experiment with video images from the Orbiter cameras in

the cargo-bay and cameras mounted on the RMS and the servicer

mechanism. Data and the video signals are sent to the ground station

using the standard Space Transportation System communications links.

The ground support technicians will be available for trouble shooting

but can not control the servicer directly.

The servicer mechanism will be similar to the Engineering Test Unit

(ETU) used in the ground demonstrations with design modifications for

0-g operation. It will be based on the preliminary design for the

on-orbit servicer developed during the IOSS study effort. The servicer

control system will be upgraded to include digital circuits.

The demonstration has been considered as a flight experiment rather

than as an operational activity. This approach reduces the cost by

limiting the amount of the formal review process and testing. The

hardware used can be a commercial grade that is modified to meet the

requirements. The safety requirements of the Shuttle become a driving

force in the design. Many of the requirements are satisfied by over

designing the part and performing extensive analyses. The testing is

limited to those items that need qualification for safety or are

critical development items. This approach helps to reduce the cost,

but increases the chance for an unsuccessful flight. We have tried to

balance the risk of failure with the cost constraints.

4.3.4 Cargo-Bay Demonstration Schedule and Cost Estimate

New schedules and cost estimates were developed for the cargo-bay

demonstration based on the approach described in the previous section.

The schedule was made compact in order to reduce costs.

The servicer cargo-bay demonstration schedule was developed from an OMV

development schedule. The key point from the OMV schedule was an OMV

authority to proceed (ATP) for Phases C and D in late April of 1986.

This date is shown on Figure 4.3-2 along with other OMV milestones.

4-15
The January 1989 date selected as the beginning of Phase B for the

free-flight verification or operational servicer development

corresponds with the end of the OMV supporting development of a

servicer kit. This approach integrated well with the use of

representative time spans for the various demonstrations and

verification activities. It was decided that the results of the

servicer cargo-bay demonstration would be most useful if the cargo-bay

demonstration was completed at the start of the operational servicer

development in January 1989. The cargo-bay flight is completed before

the operational servicer preliminary design review (PDR) so that the

cargo-bay demonstration results can be factored into the operational

servicer design and development.

DESCRIPTION

OMV Design &

Development

Ground Demonstrations

System Engineering

Design & Development

Safety

Software Dev & C/O

Procurement

Fabrication

Assembly & C/O

Acceptance Testing

System Integration &

Demonstration

Verification Phase B

1985

i

PHASE B

1986

ATP TM

[V V

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Figure 4.3-2 Servicer Cargo-Bay Demonstration Schedule

A preliminary cost estimate was prepared for the proposed cargo-bay

demonstration. A primary function of this effort was to reduce the

cost of the cargo-bay demonstration.

4-16
The cargo-bay demonstration was treated as a flight experiment rather

than a flight qualified unit. This changes the technical approach used

in the program, which in turn affects the procedures to estimate the

costs. The approach used was to estimate the program cost of a flight

qualification program using the normal cost estimating relationships

(CER). To these results a complexity factor was applied to account for

the difference between experimental and flight qualified units. The

data from the equipment lists and the work breakdown structure were

used to establish the inputs to the CERs.

The complexity factor was arrived at by reviewing recent programs of a

similar nature. These include the Westar/Palapa rescue mission, the

Solar Max repair mission and a Shuttle demonstration of storable fluid

management techniques. From this data, a complexity factor of 0.4 was

derived.

The total estimated cost of the servicer cargo-bay demonstrations is

approximately $9.3 million. A breakdown of this estimate and a summary

of the basis of cost is presented in Table 4.3-3.

The estimated cost of 9.3 million dollars is more than a 50 % reduction

over the previous estimate. These savings result from reducing the

number of Shuttle flights, utilizing as much existing equipment as

possible, and treating the demonstration as a flight experiment.

Although these measures reduce the cost, there is a price associated

with them. The risk of an unsuccessful mission from the technical

standpoint is increased (the plan does not increase the risk to the

Shuttle or the crew). The depth to which the servicing techniques will

be demonstrated has been reduced, but all the major, activities have

been included. The proposed plan has been designed to demonstrate the

major servicing functions in the most economical manner possible.

4-17
Table 4.3-3

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4-18
4.4 FREE-FLIGHT VERIFICATION

ORIGINAL PAGE IS

POOR QUALfTY

The free-flight verification is intended to demonstrate the operational

capability of the servicer. Figure 4.4-1 shows an artists concept of

the servicer and the OMV just prior to docking with a spacecraft. The

servicer will be flight qualified and ready for operations. It will be

compatible with both the OMV and the Space Station.

e

Figure 4.4-1 The Operational Servicer with the OMV

4.4.1 Free-Flight Verification Objective

The objective of the free-flight verification tests is to verify that

an operational servicer capability exists and is available for use by

the user community. These verification tests should also increase

confidence that the servicer can be used at the Orbiter, at, or near,

the Space Station, and in geosynchronous orbit. The servicer used for

4-19
the verification flight will become the first operational servicer. It

will be necessary to fully qualify the unit and document its

development, fabrication and test.

The servicer is intended to be used as a kit for the OMV. The

verification will demonstrate the two vehicles compatibility and how

they form an effective system for performing servicing operations.

Both will be controlled from the OMV ground station. The servicer

extends the capability of the OMV by enabling it to perform module

exchange and fluid resupply operations.

The servicer will be a valuable tool for the Space Station. A servicer

unit could be located in the servicing bay to perform repairs to

spacecraft and equipment. The operators would be located on the

ground, freeing the astronauts for other tasks. The OMV/servicer could

be used as a backup to the station's RMS, moving equipment around the

station and using the servicer mechanism to install it. It could

replace EVA personnel in many circumstances, especially where hazardous

materials are involved. For example, the servicer could be used to

refuel the OMV and OTV. The design of the operational servicer will

consider the needs of the Space Station.

4.4.2 Free-Flight Verification Recommended Approach

The approach to be used for the free-flight verification is in the

formulative stages. As a result, the distinction between the issues to

be considered and the baseline plan are not as clearly defined as for

the ground or cargo-bay demonstrations. We have attempted to identify

the major elements needed to meet the objectives of the free-flight

verification. We feel the verification flight will include the same

basic servicing functions demonstrated in the cargo-bay, adding the

features of operations with the OMV and control from a ground station.

The servicer system must go through the full design and development

process including the preparation of production tooling so that further

units may be produced. The cost estimate is based on the production of

4-20
a single servicer system. A single verification flight is

recommended. It is possible to use less fully qualified components for

the serviceable spacecraft and the modules to be exchanged as the

serviceable spacecraft is to be used only once. It may be possible to

rent a spacecraft bus and use it to support the serviceable

spacecraft. The serviceable spacecraft and module mockups are the

units used for the cargo-bay demonstrations.

The proposed plan calls for the servicer and OMV to be mated on the

ground and carried to space in the Orbiter. Table 4.4-1 presents the

major characteristics of the verification flight. The servicer's

stowage rack will carry the modules, tools, and stored fluids necessary

to service the spacecraft mockup.

Once in space, the OMV with the servicer and the spacecraft mockup will

be separately deployed from the Orbiter cargo-bay. After a reasonable

amount of separation has been achieved, the OMV will rendezvous and

dock with the spacecraft. The servicer will exchange a MMS module

(axial) and a battery module (radial). It will demonstrate fluid

resupply operations by mating a fluid interface device to the

spacecraft and transferring fluid to the spacecraft. The vehicles will

be returned to the Orbiter and transported to the ground.

4-21
Table 4.4-1 Free-Flight Verification Characteristics

One verification flight

MMS triangular structure for spacecraft mockup

Spacecraft mockup attached to rented spacecraft bus

Axial MMS module exchange

Radial battery module exchange

Fluid transfer to PM-1 type propulsion module

Servicer's supply of power, attitude control, thermal

control, and communications by OMV

Spacecraft mockup*s supply of power, attitude control,

thermal control, and communications by spacecraft bus

Servicer control station on ground

Unassisted Supervisory control mode

Docking rigidization by servicer docking probe

Electrical connection between servicer and spacecraft via

the docking mechanism

Servicing equipment performance demonstration

OMV-servicer interactions included

Man-machine interactions included

Compliance with Orbiter system safety requirements

Use of representative servicing operational equipment

Operator training (ground personnel)

4.4.3 Free-Flight Verification Schedule and Cost Estimates

A set of schedules and cost estimates have been defined for the

proposed verification program. These numbers are based on the approach

defined in the previous section. This approach is basically the same

as proposed in the earlier study, except only one unit will be built.

The cost for production tooling has been kept in the schedule and cost

estimates.

The development of the operational servicer is to be coordinated with

the development of the OMV. Being an "OMV kit", it is desirable to

have the servicer flight qualified at the start of OMV operations. A

schedule showing the development of the servicer and the major OMV

milestones is shown on Figure 4.4-2. The key OMV milestones are the

first flight occurring in early 1990 and the start of normal operations

in mid 1992.

4-22
DESCRIPTION

OMV Design &

Development

OMV Operations

Cargo-Bay

Demonstration

Definition Study

Design & Development

Software Dev & C/O

Tooling & GSE

System Test Equipment

Control Station

Fight Unit & ASE

System Qualification

System Integration &

Verification

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Figure 4.4-2 Free-Flight Verification Program Schedule

The phase B study for the servicer has been shortened to 9 months due

to the extensive amount of definition work that has been done as part

of earlier contracts. A three year period has been allocated for the C

and D phases, which is a representative timespan for operational

equipment of this complexity. The development schedules of the

servicer and the serviceable spacecraft have been combined. It is felt

that the definition period of the spacecraft may be somewhat longer

than that of the servicer as little work has been done in this area.

The design and development of the servicer will take comparatively

longer as the spacecraft will be used only once and will probably

consist of a combination of existing equipment. The development of a

servicer ground control station has been included. Due to the "kit"

relationship with the OMV, the servicer control station will likely be

4-23
collocated with the OMV ground control station, and will utilize much

of their support equipment.

A preliminary cost estimate was prepared for the free-flight servicer

system verification. The costing was based on estimated weights of the

equipment for the free-flight servicer verification. The cost estimate

was developed using cost estimating relationships (CER) contained in

the Martin Marietta Aerospace Cost Data Base and in several NASA

pricing models.

A breakdown of the cost estimate is shown in Table 4.4-2. The total

estimated cost of the engineering effort and building the servicer

system for free-flight verification is approximately $35 million.

The estimated cost of 35 million dollars is the same as was previously

estimated. The cost of the servicer unit was reduced by using the

proto-flight approach. This approach permits a single unit to be used

for qualification testing arid operations. The tests are performed on

the flight unit in such a manner as to verify the capability, but not

damage the unit. The savings were offset by increased costs due to

inflation and the need to incorporate several items that were

previously included in the cargo-bay demonstrations.

Table 4.4-2 Free-Flight Verification Cost Breakdown

Flight Equipment

Servicer Mechanism

Airborne Support Equip.

Servicer I/F Equip.

Stowage Rack

Stowage Rack I/F

Spacecraft Bus Rental

Mockup Equipment

Docking Probe

Subtotal

Program Total

* Costs are in millions of

8.0

8.2

3.0

4.0

0.5

2.2

0.8

4.5

31.2

35.0*

1985 dollars

Ground Equipment

Servicer Checkout

Stowage Rack Checkout

Spacecraft Checkout

Ground Control Station

Personnel

Subtotal

0.2

0.1

0.1

3.3

0.1

3.8

4-24
4.5 USER COMMUNITY COORDINATION

The user community coordination activity has been primarily with NASA

personnel as many United States non-DoD spacecraft programs look to

NASA for direction. Emphasis has been on presenting the remotely

controlled module exchange capabilities to personnel within the Space

Station program as that program is at a stage where these on-orbit

servicing concepts could be accepted and implemented.

In September 1984 a presentation was made to Goddard Space Flight

Center personnel involved in the Multi-mission Modular Spacecraft

program. The presentation described the early studies, servicer

configuration, economic advantages, and remote servicer demonstration

plan. This meeting led to increased emphasis on the inclusion of MMS

module exchange for near-term activities.

The March 1985 Mid-Term Review presentation was expanded to include a

section on servicing background. Presentations were made to Space

Station program personnel at MSFC and at NASA Headquarters. GSFC

personnel also attended the NASA Headquarters meeting. The

presentation material included: fundamentals/perspective on spacecraft

servicing, servicing alternatives, evaluation of alternative

maintenance modes, guidelines for application of spacecraft servicing,

example serviceable spacecraft configurations, implications on

serviceable spacecraft design, on-orbit servicer configuration, control

mode description, on-orbit servicer Engineering Test Unit, and

candidate concepts for use of on-orbit servicing with and at the Space

Station. The presentations were well received.

A Servicer Development Program Plan report was prepared and distributed

to a wide variety of potential users. The draft plan was distributed

in April 1985 and the final version was distributed in July 1985. This

plan proposes a sequence of ground demonstrations, a 0-g demonstration

in the Orbiter cargo-bay, and a free-flight verification with the

servicer system mounted on the Orbital Maneuvering Vehicle. This plan

4-25
leads to an operational on-orbit servicing capability that can be used

on the OMV or at the Space Station.

A presentation was made to the MSFC Propulsion and Vehicle

Accommodations Splinter Group during the Space Station Orientation

meetings at the request of the Martin Marietta Space Station team in

April of 1985. The presentation included: OMV kits with application

to Space Station, description of Integrated Orbital Servicing System,

on-orbit servicer characteristics, on-orbit servicer control modes,

on-orbit servicing fundamentals, on-orbit servicer development plan

summary, description of the Engineering Test Unit of the on-orbit

servicer, Orbiter cargo-bay demonstration of on-orbit servicing, fluid

resupply with on-orbit servicer, automatic umbilical isometric,

automatic umbilical characteristics, on-orbit servicer system at Space

Station, remote satellite servicing mission activities, suggested basic

approaches to accommodation of on-orbit servicing kit at Space Station,

general accommodation requirements for on-orbit servicer, additional

accommodations requirements for OMV kit, for mobile RMS, and for

servicing facility, summary of current related contract research and

development, summary of current related independent research and

development activities, and a listing of open issues. The presentation

was well received and additional material and assistance has been

provided to the Martin Marietta Space Station team.

A number of discussions were held with a representative of RCA

Government Systems Division of Princeton, NJ during the first half of

1985. A number of documents were also sent to RCA. Subjects discussed

included: on-orbit servicing in the form of module exchange,

spacecraft design for servicing, application of servicing to Space

Station, application of servicing to polar platforms, and use of OMV in

polar orbits to service platforms. The discussions are continuing.

A presentation was made to Fairchild Space Company personnel on June 5,

1985. This presentation included a summary of prior work on on-orbit

servicing, objectives and progress on the then-current study, and

4-26
advantages of involving the MMS and the Module Servicing Tool in the

1-g and Orbiter cargo-bay demonstrations.

A presentation was made to the MSFC Science and Engineering Technology

Manager for Teleoperation and Automation on June 6, 1985. This

presentation included: a general on-orbit servicing background

discussion, objectives and progress on the then-current study, and a

summary of the work being done and planned for the Engineering Test

Unit in his laboratory. The presentation was well received.

A paper entitled "Spacecraft Design for Servicing" was presented at the

Satellite Services Workshop II. The paper was well received. The

workshop was held at the Johnson Space Center in Houston, Texas on

November 6, 7 and 8, 1985.

4-27
5.0 1-g DEMONSTRATION PLAN

The objective of this study activity was to update and expand the

ground demonstrations section of the servicer development program plan

proposed under the previous contract. The 1-g demonstration plan

emphasizes the MMS servicing demonstrations along with the development

and the demonstration of the basic techniques of remote on-orbit

servicing. The specific MMS servicing demonstration requirements, the

available options, and the approach selected were discussed in Section

3.0. In this section, the general requirements of the 1-g

demonstrations, as well as elements of the ground demonstration plan

developed under the previous contract are updated and expanded. The

schedule of the ground demonstrations and a cost estimate are also

presented.

The ground demonstrations are conducted first as they are less

expensive than the 0-g demonstrations, the equipment is more accessible

and is easier to reconfigure, a wider range of tests can be conducted,

and the data is easier to collect.

The objectives of the ground demonstrations are to obtain a better

understanding of on-orbit servicing so that the cargo-bay

demonstrations may be better focused and to increase user confidence in

the technology and in the program. These objectives can be expanded as:

1) To demonstrate the adaptability and flexibility of the module

exchange concept - This can best be done by demonstrating an

exchange of the MMS module mockup, because it is the only on-orbit

serviceable spacecraft modular concept that is operational and it

was designed for a different servicing interface. Additional

demonstrations should be conducted to show that the IOSS is a

flexible servicing system, and does not impose significant

constraints on spacecraft design. Exchange of equipment at the

individual component level, such as battery replacement, including

the opening or removal of an access door/thermal protection cover

can further demonstrate the versatility of this servicer system;

5-1
2) To evaluate approaches for the cargo-bay demonstration and

free-flight verification - The ground demonstration will be used to

select the control strategies to be used in the demonstration/

verification tasks. Through the use of a parametric analysis of

the identified control strategies the optimal control system can be

selected. Operating procedures can be varied and checked for

overall effectiveness in the use of available resources. This

includes such things as ground support, astronaut time, and Shuttle

interfaces. These types of issues are best answered with

relatively inexpensive ground testing rather than expensive flight

experiments;

3) To demonstrate the use of the ground servicer as a laboratory tool

- The development of new servicing concepts, new hardware, and

software can be readily investigated on the ground before flight

testing and operational implementation. A good example is the

development of a satellite remote fluid resupply capability. The

ground servicer can also be used as an integration and checkout

facility. Development of an automatic target recognition and error

correction system, of new controls or of new tools and adapters can

benefit from the use of the ground servicing demonstration system

as a laboratory tool. New sensors, sophisticated end effectors,

and other elements of the next generation of servicing systems can

be developed using the ground and the flight demonstration units.

If problems arise during the flight tests or operational servicing,

the ground demonstration unit could be used for finding and/or

checking out solutions;

4) To demonstrate the use of the ground servicer as a training

facility - Training of the operators for the flight demonstrations

as well as for actual servicing operations can be done using the

ground servicer system. For this reason, it is important that

hardware and software commonality with the flight units be designed

5-2
ORIGINAL PAGE IS

OF POOR QUALITY

into the ground demonstration servicer. This will also make possible

more convincing, high fidelity, ground servicing demonstrations.

The main role of the servicing ground demonstrations is to support

further flight demonstrations. The availability of an on-orbit

servicing capability can be convincingly demonstrated to the user

community only through flight tests. The acceptance of on-orbit

servicing methods by the spacecraft designer is also linked to the

financial and programmatic commitment of NASA for timely development of

the operational capability.

The Engineering Test Unit (ETU) of the IOSS was selected as the

servicer mechanism for ground demonstrations based on the results of a

tradeoff study done as part of the prior contract. The actual ETU is

shown in Figure 5-1. Details of this selection process as well as the

selection of the related hardware are documented in Section 3 of the

"Spacecraft Servicing Demonstration Plan - Final Report", July 1984

(MCR-84-1866).

Figure 5-1 The Engineering Test Unit of the IOSS

5-3
The activities selected for the ground demonstration plan represent a

continuation of the ongoing demonstration tasks, using the ETU. Some

of the elements included are part of the current activities and have

been included in the plan for completeness. The cost estimate reflects

the work to be done and does not include the currently funded tasks.

5.1 1-g DEMONSTRATION REQUIREMENTS

The system requirements applicable to the servicing ground

demonstration equipment defined in the prior study were analyzed and

expanded. The top level requirements used for the definition and the

selection of the elements of the 1-g demonstration plan are presented

below. Additional subsystem requirements used in the design of the

servicer/MMS 1-g demonstration equipment are presented in Section 8.0.

The specific requirements developed for the Module Servicing Tool (MST)

adaptation for 1-g demonstrations and other MMS servicing requirements,

with emphasis on the operational case are listed in Section 3.1.

The following requirements, or guidelines, apply to the ground

demonstration servicing system:

1) The existing Engineering Test Unit of the IOSS will be used for all

the ground servicing demonstration activities;

2) Minimum modifications will be made to the existing ETU

configuration. An adapter shall be used in connection with the

standard end effector interface in order to service MMS modules or

for fluid resupply;

3) The ground demonstration servicer system shall be capable of 1-g

demonstration of MMS module exchange in addition to the present

capability of exchanging the 24 in. cube module with side interface

mechanism in axial and radial directions, without any system

reconfiguration. The design of the servicer system shall provide a

capability for later expansion to include ground demonstrations of

fluid resupply and demonstration of other servicing tasks;

5-4
4) The trajectories used during ground demonstrations of MMS module

exchange and the relative position of the servicing system elements

shall provide a good representation of the on-orbit servicing of an

MMS, using lateral docking and axial module exchange;

5) The increased end effector load due to MMS module mockup, modified

MST, and other servicer modifications, such as for fluid resupply,

shall not exceed the servicer design load capability;

6) The positioning accuracy of the servicer arm with modified MST and

with or without module mockup shall be within the capture envelope

of the MMS module retention system;

7) Adequate clearance shall be provided between all servicer system

elements;

8) The ground demonstration servicer system shall be capable of 400

complete cycles of MMS module exchange demonstrations without

refurbishment;

9) Optical targets shall be provided for all locations where the

servicer end effector engages module attach interfaces, fluid

resupply interfaces, or adapters, at their storage locations;

10) The adapters used for performing special servicing tasks shall

obstruct as little as possible of the field of view of the existing

TV camera and lights;

11) The 1-g demonstrations of fluid resupply shall be designed so that

this operation can be performed as part of the same overall

demonstration with other servicing activities, such as module

exchange or inspection;

12) Initial ground demonstrations of fluid resupply may use water and

air pressurant in order to minimize risk and cost;

5-5
13) The fluid resupply servicing interface should conform with the

industry established fluid resupply standard interface;

14) The mate/demate subsystem of the fluid resupply interface shall

include an auto-indexing feature to assure the correct mating of

the disconnect valves;

15) The hose/cable management system for 1-g fluid resupply

demonstration shall be counterbalanced and shall assure servicing

at all required locations;

16) The following real time control functions shall be provided as a

minimum for the fluid resupply 1-g demonstrations:

a) Control of disconnect mate, demate, leak test, and purge

functions,

b) Control of liquid and gas pressures,

c) Valve position indication;

17) The servicer control modes, Supervisory, Manual-Augmented and

Manual-Direct, and the associated control software should be common

to all ground servicing demonstrations;

18) Separate specific software programs for each demonstration/activity

are permissible.

5.2 1-g DEMONSTRATION ELEMENTS

Several ground demonstration tasks were identified as possible

candidates for a representative 1-g demonstration plan and they are

described in the following paragraphs. The ground demonstration plan

elements were selected from among these candidates considering the

development work needed to support the next phases of cargo-bay

servicing demonstration and free-flight verification. An effort was

5-6
made to minimize the cost by using existing equipment and control

software whenever possible.

5.2.1 Multi-Mission Modular Spacecraft Module Exchange

This task involves the fabrication of two lightweight MMS modules,

modification of the existing MMS module servicing tool, adapting the

existing spacecraft and stowage rack, and providing the related

software.

A lightweight mockup of the MMS module will be designed and two units

built. The design goal is a maximum weight of 13 Ib for the module,

including the electrical connector(s) and the module retention

hardware. The module mockup will be a full size representation of the

outside shape and dimensions of the MMS module, will have the same

attachment interface and provide adequate structural support for the

two attaching fasteners and for the two latch receiving brackets. A

lightweight version of the module retention and connector mounting

hardware will be designed and fabricated for the mockup. The fastener

operating torque will be reduced to the minimum value that will secure

the module in place.

A modified, lightweight Module Servicing Tool (MST) will be designed

and built. It will be compatible with the ETU end effector interface,

including an electrical disconnect and it will have the same interface

with the MMS module retention system. The MST latches will hold the

module tightly during the module transfer operations, being modified to

delete the inward motion function of the existing EVA tool. The

electronics, power supply and controls will be in a remote location and

the MST will be designed for a lower operating torque to reduce its

weight to less than 15 Ibs, as the ETU has a limited lifting

capability. The modified MST will be provided with a "ready to latch"

sensor.

A lightweight and compact connector positioner will be attached to the

ETU end effector and will be used for mating and demating the

electrical connection to the MST.

5-7
Modification of the present spacecraft mockup Is necessary In order to

incorporate support structure, compatible attachment interfaces,

connectors and sensors for one MMS module. Modification of the stowage

rack mockup is necessary for receiving the MMS module in two

locations. One location is for temporary stowage of the failed module

from the spacecraft mockup and the other for the replacement module

that is to be transferred and fastened to the spacecraft mockup.

The MMS module exchange activity has been performed as part of the

existing contract. It was supported by the MST redesign and

fabrication work by the GSFC, which was assisted by Fairchild Space

Company. The demonstration plan, design, and equipment are discussed

in Section 8.0. The servicer control software for MMS module exchange

is described in Section 10.0 and the MMS module exchange demonstrations

are described in Section 11.2.

5.2.2 Fluid Resupply Demonstrations

For initial ground demonstrations, a nonhazardous fluid will be

transferred from the stowage rack to the spacecraft mockup. This will

demonstrate fluid transfer in an economical and safe manner. The added

expense of transferring a hazardous fluid provides no additional

Information to the program. Those types of activities are better

suited for programs specifically addressing propellant transfer rather

than this program, which deals with remote servicing.

A servicer fluid resupply module mockup comprised of a fluid tank, air

tank, piping, flow control, hose/cable management system, fluid

transfer Interface unit, instrumentation, controls, and support

structure should be assembled. The fluid transfer interface unit

should carry disconnects for liquids, gases and electric cables. It

should also have a translation mechanism, attachment alignment

mechanism, and a dust cover removal mechanism. Such a device will be

obtained either from within Martin Marietta or from NASA. The

development of such an interface is not within the scope of this

program. Goddard Space Flight Center has recommended that the

5-8
refueling interface being developed by Johnson Space Center be used for

the demonstration. This option will be evaluated, as using existing

hardware will reduce the program cost. Simplified functional mockups

should be built for disconnect valves with leak test and purge

capabilities.

Modifications of the spacecraft mockup are necessary in order to

accommodate the fluid and pressurized air tanks, piping, valves,

instrumentation and controls, and the fluid transfer interface. The

same approaches used on the fluid resupply module mockup will be

applied to the spacecraft mockup.

5.2.3 Control Mode Analysis and Testing

A series of tests and analyses will be conducted to evaluate the

effectiveness of the different control modes. Variations in the

standard approaches will be tried to shorten the time requirements and

improve the accuracy. The result of this activity will be

recommendations on the control modes to be used in the flight

experiments.

5.2.3.1 Reconfigurable Control Console - The ETU controls and displays have

been set up around the Servicer Servo Drive Console as this arrangement

was convenient during checkout of the basic module and MMS module

exchange software. However, the various input devices and displays are

not well placed from a human factors viewpoint. The result is that the

system is not as easy to work with as it might be. Also the effects of

rearranging the controls and displays is not easy to evaluate.

It is recommended that the pertinent equipment be moved to one of the

available reconfigurable control consoles at MSFC. The pertinent

equipment includes:

1) Servicer Control Panel;

2) TV Monitor;

5-9
3) Computer Monitor and Keyboard;

4) Hand Controllers for Manual-Augmented Control Mode;

5) MST Control Panel.

Provisions should be made for an ability to use one six

degree-of-freedom hand controller or two three degree-of-freedom hand

controllers. The equipment should be designed to accommodate

additional servicing demonstrations, such as fluid resupply, at a later

time.

5.2.3.2 Module Trajectory Modifications - The trajectories used to exchange

the basic and MMS modules were developed primarily on the basis of

collision avoidance. Few attempts were made to refine these

trajectories to reduce the demonstration time. It is recommended that

the trajectories be reexamined to determine if the demonstration time

can be significantly reduced. Items for consideration include: 1)

module location; 2) clearances required; 3) velocity limits. The

Supervisory control mode software limits the system velocities at both

the cylindrical coordinate and the joint drive levels. It may be

possible to raise these velocity limits and still operate along true

cylindrical coordinates.

5.2.3.3 Supervisory Control Mode Assistance Level - The current Supervisory

control mode incorporates two levels of operator assistance. The

"unassisted" level only requires operator actions at safety-related

points once the basic module exchange trajectory has been Initiated.

There are three inputs for basic module exchange and six inputs for MMS

module exchange. The "operator assisted" level requires operator

responses for each and every action listed on the computer display. An

analysis should be conducted to identify if there is a better form for

the "operator assisted" mode and to define that form.

5.2.3.4 Manual-Augmented Control Mode Variations - One form of the

Manual-Augmented control mode and one improved form of the related

trajectory sequences have been implemented. No data has been collected

5-10
to determine if the forms being used are among the better forms. It is

recommended that a series of analyses and test activities with

different operators be conducted to identify forms of the

Manual-Augmented control mode and related trajectory sequences that

will speed up the module exchange process, produce more accurate

trajectories while in the guides, and ease the learning process. Items

for consideration include:

1) Specific trajectory sequences;

2) Form of displays used;

3) Use of rotational hand controller.

The form of the targets and TV monitor reticle used has not been

examined in any detail. They were assembled quickly from convenient

materials. It is now difficult to watch the error meters and the

target/reticle positions at the same time because of their relative

locations. Consideration should be given to overlaying the error meter

data on the TV screen so the operator can respond more easily to the

display information.

At the study Orientation Meeting it was decided to use a

single-degree-of-freedom hand controller for the wrist roll function.

When it became impractical to incorporate the rotational hand

controller into the setup, it was decided to use the rocker switches on

the Servicer Control Panel to command end effector rotations. The

result is clearly non-optimum. It is recommended that the subject of

rotational control be reviewed, better approaches tested, and a

preferred system selected.

5.2.3.5 RMS Control Modes - The Orbiter Remote Manipulator System incorporates

a set of control modes that have evolved over the years and that are

familiar to the astronauts. It is recommended that information be

obtained on these control modes, the rationale for their use, and what

5-11
characteristics are preferred by the astronauts. This information

would then be used to determine if the servicer control modes should be

made similar to the RMS control modes. An easy correspondence between

the control modes would make the learning process easier because of

similar use patterns and operational philosophies.

5.2.3.6 Error Correction via Manual-Augmented Control Mode - As discussed in

Section 5.2.6, the errors associated with the docking system can cause

the end effector to fail to attach to the interface mechanism or the

interface mechanism to fail to be captured by its guides. One way to

overcome these errors is to use the Manual-Augmented mode to realign

the target and reticle and thus generate bias signals that can be used

in the Supervisory control system. There are a number of problems to

be addressed, but solutions should be possible. Areas of concern are:

1) Changes to Supervisory mode software to permit returning to a

specific point in a trajectory sequence after switching to the

Manual-Augmented mode;

2) Philosophy of and method for computing the error signals and

incorporating them into the module location data. For example,

should the corrections be applied to all module locations or just

to the one where the errors were measured;

3) Method for measuring the errors;

4) Target and reticle changes to permit identifying all six error

components.

Use of the Manual-Augmented control mode for error measurement can be

considered as an alternative to, or as a complement to, the automatic

target recognition and error correction system of Section 5.2.6.

5.2.3.7 Software for Combined Motions - While the servicer mechanism is

electromechanically capable of performing combinations of radial and

axial motions along straight lines of arbitary direction, the control

5-12
system has not been designed for combined motions. The premise for

control system development has been to keep it simple. Combined

motions add significantly to control system complexity. However, the

analyses of MMS module exchange indicated that it would be possible to

delete the orientation joint if combined motions could be effected. It

is suggested that the effects of incorporating combined motions in the

control software be assessed and if the complexities are not too severe

then a combined motion capability should be developed.

5.2.3.8 Error Correction via Automatic System - The automatic target

recognition and error correction system activity of Section 5.2.6 would

primarily address the error measurement and transformation of errors to

a reference coordinate system. This activity would develop a

philosophy and methodology for how the generated errors would be used

by the system and how they would be incorporated in the software.

5.2.3.9 Inadvertent Module Release Warning System - As a result of an

Orientation Meeting Action Item, an approach for warning the operator

of an inadvertent end effector jaw opening was generated. The transfer

of MMS modules introduces another modality where modules could be

inadvertently dropped; namely inadvertent unlatching by the 1-g Module

Servicing Tool. It is suggested that this problem be readdressed and

if felt to be significant then a warning system should be implemented.

5.2.4 AXAF Module Exchange Demonstrations

A demonstration of focal plane instrument module exchange, such as for

the AXAF, requires building two large volume, lightweight module

mockups. The module retention system could be a lightweight version of

the base mounting interface mechanism. The design should include

electrical and fluid disconnects.

Modification of the spacecraft mockup is required to accommodate radial

removal of the AXAF module mockup including a hinged thermal cover with

an unlatching/opening mechanism, actuated by the power takeoff

(interface mechanism drive) of the servicer.

5-13
Modification of the stowage rack to accommodate the AXAF module in two

locations is needed to provide structural support and latch

interfaces. A hinged thermal cover, similar to the one on the

spacecraft mockup should be fitted on one of the two stowage rack

locations. The second location is used for temporary storage.

The above discussion assumes that the focal plane instruments are

selected for AXAF remote servicing demonstration. This is a reasonable

assumption as the focal plane instruments are likely to be the most

difficult to service. However, AXAF project personnel should assist in

determining which specific AXAF equipment should be selected for a

remote servicing demonstration. They should also assist in defining

other parameters of the demonstration such as need for covers,

allowable time without electrical power, and need to be compatible with

EVA module exchange. The specific equipment selected will generally be

required in the quantities and at the locations identified above.

5.2.5 Battery Module Exchange Demonstration

Battery module exchange can be used to demonstrate radial exchange of

equipment at the component level, in order to prove the operational

flexibility of the ETU servicer. A representative battery module

mockup should be designed and two units should be built. The mockup

should have an electrical disconnect and a lightweight latch mechanism

capable of mating/demating the disconnect. The latch mechanism will be

developed from the existing side interface mechanism concept. As an

alternative, the module battery mockup could be attached to the base

structure using captive fasteners. An adapter tool should then be

built to actuate the fasteners, mate or demate the disconnect, and

support the battery during transfer operations. MMS type fasteners and

the MMS adapter tool could be used instead of standard captive

fasteners.

The stowage rack and the spacecraft should be modified to receive the

battery modules: one for the replacement module and one for temporary

attachment to the stowage rack.

5-14
Data resulting from this demonstration will be used for the design of

the cargo-bay demonstration of battery exchange in the radial direction.

5.2.6 Automatic Target Recognition and Error Correction

A prior study* of the expected error of mechanical arms, conducted by

Martin Marietta Aerospace with internal funding, shows that errors

approaching + 0.80 in. can occur for a system like the IOSS. This

number should not be compared with the ETU repeatability of 1/8 in.,

which is only one small component of the overall error. Among the

dominant sources of error considered, is the vehicle docking

misalignment. Without special provisions, docking misalignment can be

on the order of degrees. Docking misalignment not exceeding 0.3 deg in

any of the three axes was considered in the above-mentioned study.

However, if the standard RMS end effector is used as a docking probe,

post rigidization accuracy of + 0.4 deg is expected in the roll

direction and + 0.15 deg in the pitch and yaw directions**. Roll is

the most critical error component and is unfortunately the most

difficult to reduce for small central-type docking systems. Large roll

errors about the docking axis can be easily measured and easily

compensated for when performing an axial module removal. However, the

roll error is hard to measure and compensate for during a radial module

removal.

The IOSS end effector capture envelope is + 0.75 in. and the guide

capture capability of the side mounting interface mechanism is + 0.50

in. These capture capabilities are marginal, when using the RMS end

effector as a docking probe. The use of adapters for the ETU end

* Orbital Inflight Maintenance (Project 27D) Vol. 2 - Accuracy Capability

of Mechnical Arms, Martin Marietta Aerospace, Report No. D76-48727-002,

December 1976.

** R. G. Daniell, et al., "The Design and Development of an End Effector

for the Shuttle Remote Manipulator System" 16th Aerospace Mechanisms

Symposium (J. F. Kennedy Space Center, Florida) May 13-14, 1982, NASA

Conference Publication 2221.

5-15
effector and/or docking probe enhances the system operational

flexiblity, but at the same time, may appreciably decrease its accuracy

below the minimum acceptable level. For radial module removal up to

63% larger errors are expected.

In manual control modes the operator can make the required corrections

before engaging the module or the end effector, by using the video

camera capability.

In the Supervisory mode, however, an equivalent capability needs to be

developed, in the form of an automatic target recognition and error

correction system. The system can use the existing video equipment,

and the on-board computing capability, to scan and interpret the TV

image prior to engagement, detect the error, issue the required

commands for correction to the servicer control system, verify the

results, and then command the final engagement.

All six components of the relative location error must be identified -

three in linear displacement and three in angular displacement. As

most sensors have an image plane, they can only measure three error

components. If the targets have a known size, then stadiametric

ranging can be used to obtain a fourth error component. If the target

is also given some depth (as for the RMS target), then all six error

components can be measured. The sensor system must know, and be able

to recognize, the geometry of the target. Another approach, applicable

to the servicer mechanism with its degrees-of-freedom, is to use two

planar targets of known geometry at right angles to each other.

Another consideration is the gross error situation such as docking at

the incorrect docking port or having a 120 deg roll docking error.

These kinds of errors can be recognized by properly coding each of the

specific targets.

An autonomous video rendezvous and docking system development was

initiated by Martin Marietta Aerospace under a contract from MSFC***.

*** Development of an Autonomous Video Rendezvous and Docking System,

Martin Marietta Aerospace MCR-83-584, Phase 2, June 1983, MSFC Contract

NAS8-34679

5-16
It requires a modified optical target with three reflective spots,

special software and a special computer interface box to handle the

data processing. The system has been operated in the Space Operations

Simulation Laboratory of Martin Marietta Aerospace and the technology

is readily applicable to the servicer. However, the Martin Marietta

autonomous video rendezvous and docking system has a larger capability

than is needed for servicer error measurement so alternative, simpler

devices should also be considered. The range of servicer system errors

is expected to be small and this fact should be considered. The

expanding use of robotics has led to the development of small optical

devices that may be applicable to the error measurement problem. Also

charge coupled device (CCD) cameras with sensing arrays that are fixed

in size and are at the camera's focal plane are better than vidicons

for accurate measurement.

5.2.7 Engineering Test Unit Electronics Update

Improvements in the reliability of the Engineering Test Unit can be

obtained by updating some of its controls electronics, such as

replacement of relays with solid state switches, replacement of wire

wrapped boards with printed circuit boards, and by improving some other

circuit elements. The ETU has had few failures compared to the

expected level for equipment of its complexity. Most of the recorded

problems are linked to failures of electronic components. The changes

proposed have a potential for improving the ETU if it is to be used

extensively in the future.

5.2.8 Convert ETU Control System from Analog to Digital

The modifications list includes digital sensors (like optical encoders

inside the joints), digital Inputs and displays, microprocessor

computations, and a new control panel. These modifications will

improve the accuracy and the stability of the controls system. Process

5-17
controllers are available off-the-shelf for use as microprocessors. It

is likely that the control electronics for the operational servicer

system will use digital technology. This activity could also assist in

the design of the servicer system for the cargo-bay demonstration.

5.2.9 Tank Exchange

As the capability of the servicer expands, the replacement of

individual components will be possible. It will be desirable to change

components that contain fluids such as tanks, regulators, valves, and

thrusters. In order to change out these components, it will be

necessary to develop an on-orbit serviceable in-line coupling that will

not leak under pressure, vibration, and mechanical load for the entire

life of the spacecraft, when connected or when disconnected. If fluid

were allowed to escape, there could be a significant contamination

problem as well as fluid pressure loss. The device will have to be

designed to be operated by the servicer. The actual process involved

with the changeout of a fluid component will probably require special

adapter tools.

5.2.10 Interface Mechanism Development

The development of interface mechanisms for other spacecraft modules is

appropriate. These interface mechanisms would be for special

applications where existing interfaces don't meet the requirements.

The requirements might include unusual geometry situations, special

alignment, or unique connector problems. These interface devices

should include a standard end effector interface, so that the servicer

does not need to be modified to exchange the unusual module. This does

not limit changes in the design of the actual interface mechanism.

t

As part of this activity, it may be appropriate to form a committee

that would establish standards for the interface between the end

effector and the interface mechanisms. These standards would include

physical, functional, and electrical considerations. Committee members

should be drawn from all interested groups.

5-18
5.2.11 Adapter Tools for Specialized Tasks

This development may include adapter tools for removing special

fasteners, or a PFMA type end effector for gripping and deploying an

antenna or performing other contingency tasks. These specialized tools

will include a standard interface to the servicer end effector allowing

for a simple exchange of the tool in the servicing environment. The

tool would be carried on the stowage rack and mounted to the servicing

arm when required. These tools will greatly add to the capability of

the servicer and its ability to adapt to change. The MMS Module

Servicing Tool adaptation is an excellent example of this kind of

specialized task.

5.2.12 Recommended Approach

The proposed activities were reviewed and screened. Those selected for

the baseline approach are necessary to support the cargo-bay

demonstration and free-flight verification. The activities selected

are shown in Table 5-1.

Table 5-1 Recommended Activities for the Ground Demonstrations

Activity Description

MMS Module Exchange

Fluid Resupply Demonstrations

Control Mode Analysis and Testing

Battery Module Exchange Demonstration

Automatic Target Recognition & Error Correction

Reference

Section

5.2.1*

5.2.2

5.2.3

5.2.5

5.2.6

* Worked as part of the current contract.

The three primary servicer functions are included in the recommended

activities. Among them are the exchange of modules in the axial and

radial directions and the demonstration of fluid resupply operations.

5-19
The exchange of the MMS modules was selected over the exchange of the

AXAF modules as it was felt to be a more widely accepted application.

There are MMS satellites currently being flown, such as the Solar

Maximum Mission and Landsat. By showing that the servicer can adapt to

an existing design, its operational flexibility will be demonstrated.

The battery module was selected to demonstrate radial exchange as it is

a practical application of servicing. The normal servicing mission

will include the exchange of failed modules (as demonstrated with the

exchange of the MMS module) and the routine replacement of components

that have a limited life. A battery module is a primary example of the

second case. The exchange of a battery module also demonstrates the

demating and mating of an electrical connector that carries a

substantial amount of current. This type of connection can cause

arcing problems if not properly deenergized.

The fluid resupply demonstration was selected as it is one of the

primary servicing functions. Its demonstration is timely, as a number

of fluid disconnect interface units are being developed, and they can

be adapted to operate with the servicer.

The activities of the control mode analysis and the automatic target

recognition system will advance the design of the servicer to a point

that it is possible to have an operational unit. The control mode

analyses will improve this system and lessen the risk of failure. By

better understanding the controls of the servicer, future versions can

be improved. No attempt has been made to prioritize the nine suggested

control system activities. Further discussion with NASA personnel is

recommended. The automatic target recognition and error correction

system will permit automatic operation of the servicer. It will also

lead to a reduction in cost by permitting a less accurate docking

mechanism to be used and compensated for with the error correction

function.

5-20
5.2.13 Optional Support Tasks

The analyses and demonstrations identified in Table 5-2 will enhance

the development of on-orbit servicing, but were felt to be nonessential

to meeting the primary objectives of a cost conscious program. Many of

these tasks may be funded as part of other programs. It is felt that

the ETU servicer will be upgraded to a digital system if the cargo-bay

demonstrations are carried out. The tasks such as the tank exchange

and special tools and interfaces will be included as part of the

programs requiring their use. It is beyond the scope of this program

to accommodate every aspect of remote servicing development.

Table 5-2 Optional Activities for the Ground Demonstrations

Activity Description

AXAF Module Exchange Demonstrations

Engineering Test Unit Electronics Update

Convert ETU Control System from Analog to

Digital

Tank Exchange

Interface Mechanism Development

Adapter Tools for Specialized Tasks

Reference

Section

5.2.4

5.2.7

5.2.8

5.2.9

5.2.10

5.2.11

5.3 1-g DEMONSTRATION SCHEDULE

The schedules were updated from those developed for the prior

contract. Data is presented for the activities in the baseline

approach. These activities are designed to support the demonstration

of remotely controlled module exchange and fluid resupply in the

Orbiter cargo bay.

The schedule for the ground demonstrations is shown in Figure 5-2. The

OMV design and development schedule is shown as a reference. The MMS

modification and module exchange activities shown on the schedule were

5-21
completed as part of the current contract. The control mode activities

will cover a period of approximately one year. The schedule permits

the results to be used in establishing the approach to the cargo-bay

demonstration flight. The battery exchange demonstration, the

refueling demonstration, and the automatic target recognition and error

correction activities will run concurrent with the control mode

studies. These elements will involve the design and fabrication of the

necessary hardware, integration into the ETU, and a short period of

checkout and demonstration. There should be no conflict in schedules

due to the availability of the ETU as each element should only require

a few weeks of actual testing.

DESCRIPTION

OMV Design &

Development

Multi-Mission Modular

Spacecraft Modifications

(Spacecraft Mockup and

Stowage Rack)

Module Exchange

Control Mode Analysis/

Testing

Battery Exchange

Demonstration

Resupply Demonstration

Automatic Target

Recognition

Support for Cargo-Bay

Demonstration

Cargo-Bay Demonstration

1985

i

PHASE B

I I

BASIC HHS

TT T~f V VI

1986

ATP POR

fv V

I i

I

OEMONSTRAT

. . T~7

V

DEMO

*— F I V

1987

CDR^,

V

PHASE CIO

ION

<ST RATION

DEMONSTRATION

I ^

DEVELOPMENT

1988

TRAINING t

OPERATIONS

DEMONSTRATION

FUGHT ,_,

V

Figure 5-2 Ground Demonstrations Schedule

The schedule of Figure 5-2 was prepared early in the contract. At the

time of the Final Report printing, it appears that the Figure 5-2

schedule is not compatible with funding constraints.

5-22
The last ground demonstration element is supporting the cargo-bay

demonstration flight. This includes operations simulations, flight

training, and problem solving. Other activities that could be

supported by the ground servicer demonstrations during the breaks in

the flight support activities are:

1) Development of special refueling and electrical disconnects such as

cryogenic or high pressure disconnects, self aligning conical

electrical disconnects, etc.;

2) Development of in-line fluid couplings for replacement of tanks and

other propulsion system components;

3) Demonstration of other servicing tasks specific to the Space

Station operations such as resupply of other fluids, space

maintenance and assembly tasks, and return of products from a space

manufacturing facility.

5.4 1-g DEMONSTRATION COST ESTIMATE

Cost estimates were made for each of the elements of the recommended

approach for the ground demonstrations. These are engineering

estimates based on our previous experience in preparing for

demonstrations with the ETU. The estimates are based on the schedule

presented in Section 5.3. The estimates can be revised to reflect

changes in the schedule. The following assumptions were made in

estimating the costs:

1) All costs are in 1985 dollars;

2) Costs include the design, development, and fabrication of the

experiment hardware;

3) The labor to perform the actual testing of the control modes has

been assumed to be provided by NASA personnel. The analysis of the

results is assumed to be part of the contractors work.

5-23
The cost estimates for the ground demonstrations are presented in Table

5-3. The estimated cost shown for control mode analysis and testing is

not adequate to perform all of the work identified under Section 5.2.3,

rather it is an estimate of what might validly be attempted during the

next contract phase. The other Section 5.2.3 activities could be

performed in later years. As previously noted, the costs for the MMS

modifications and module exchange demonstrations are included as part

of the current funding. The cost for supporting the cargo-bay

demonstrations is included as part of that task.

Table 5-3 Estimated Costs of the Ground Demonstrations

Description

Control Mode Analysis and Testing

Battery Module Exchange

Fluid Resupply Demonstration

Automatic Target Recognition

Total Estimated Cost

Estimated Cost*

250

250

300

150

950

* In Thousands of 1985 Dollars

5-24
6.0 CARGO-BAY DEMONSTRATION PLAN

The objective of this phase of the work activity was to identify and

define the major elements of an on-orbit servicing demonstration in the

Orbiter cargo bay. The objective of the cargo-bay demonstration is to

help convince satellite designers that on-orbit servicing in the form

of remote module exchange and fluid supply can be done on orbit and

that the major elements of the system can be designed, built, and

operated.

The cargo-bay demonstrations are conducted after the ground

demonstrations so they can benefit from the results of the ground

demonstrations. A smaller number of demonstrations will be required

for the cargo bay and a set of equipment that satisfies the

requirements of the experiments to be conducted in the Orbiter was

selected. The derived objectives of the cargo-bay demonstrations are

to confirm the ground tests, show that there are no anomalies, to

demonstrate that module exchange and on-orbit refueling can be done, to

incorporate servicing equipment that is representative of expected

operational equipment, to limit risk of failure, to demonstrate the

servicers adaptability to change, to show compatibility of the OMV and

servicer schedule, and thereby to increase confidence in the technology

and the program. The cargo-bay demonstrations are considered to be a

significant step on the path to obtaining user acceptance of on-orbit

servicing.

A flow chart of the approach used for revising the cargo-bay

demonstration plan is shown in Figure 6-1. This process reviews the

technical, safety and cost aspects of the project, selects an

arrangement of equipment in the Orbiter cargo-bay, and combines the

results into a coherent plan. The first step in revising the plan for

the cargo-bay demonstrations from the prior work was to review the

existing plan and determine the key cost factors. The objectives of

the plan and the associated requirements were reviewed and updated to

reflect the current views on the development of the servicer. The

cargo-bay plan will meet the objectives of demonstrating the basic

6-1
activities of module exchange and fluid resupply. Additional

activities are limited to those items necessary to support these

primary servicing tasks. Optional elements that can enhance the

program are described, but are not included in the basic program.

Several trade studies were performed to select the activities and

configuration for the cargo-bay demonstration.

Figure 6-1 Cargo-Bay Development Program Plan Approach

A review of the appropriate Shuttle safety requirements was made to

determine their effect on the program and whether or not there were

alternatives that could reduce the cost. For example, it was

recommended that the use of pyrotechnic devices and propellants should

be avoided, as they required a significant amount of safety work due to

their hazardous nature.

A review of similar type programs was made to determine how they were

able to reduce their program costs. These included the Weststar/Palapa

rescue mission, the Solar Max repair mission and a demonstration of

storable fluid management techniques. Cost and technical information

was collected for these programs. The basic plan was formulated with

the information gathered as part of these reviews. The basic approach

was used to generate a work breakdown structure and equipment lists

(see Appendix A). These were made to a level of detail that permitted

the major analysis and test activities to be identified.

6-2
A review of the plan was made by one of Martin Marietta's Shuttle

integration safety engineers, to insure that all major activities

necessary to meet the safety requirements were included. We recognize

the importance of safety and have included the necessary tasks in our

plan.

These documents and information were used to produce the plan for the

cargo-bay demonstrations. Schedules and cost estimates were developed

to meet the requirements of the plan. These were reviewed to assure

that they were realistic and fulfilled the objectives of the program.

The final step was to document the work as part of the development

program plan and this study final report.

The prior study addressed the source and eventual utilization of the

cargo-bay demonstration hardware, in particular the servicer mechanism

itself. The early suggested approach had been to upgrade the on-orbit

servicer Engineering Test Unit (ETU). The recommended approach for

this study is to design and fabricate a new set of equipment that will

be better than the ETU but will not be as expensive as an operational

unit. The ETU was built in 1977 using early 1970's technology. The

cargo-bay servicer should be more representative of what will be used

for the operational system that will probably use late 1980's

technology. The differences are in sensors, motor types, digital vs.

analog computations, and local area networks for communications. The

ETU would also have been in use for ten years by the time it was

flown. A ground unit is needed for support during the cargo-bay

demonstration flights. The cargo-bay demonstration equipment can be

used as the ground unit during use of the operational servicer. While

the servicer mechanism is a significant cost item, it is not a dominant

one. A 50% reduction in the servicer mechanism cost corresponds to a

14% reduction in total cargo-bay demonstration costs. The

recommendation is to build a new servicer mechanism and associated

electronics that will provide a high confidence in a successful

cargo-bay demonstration yet will be significantly cheaper than an

operational unit. The differences in cost will be due to lower program

requirements, fewer reviews, and less extensive testing.

6-3
6.1 CARGO-BAY DEMONSTRATION REQUIREMENTS

The prior study, NAS8-35496, recommended two separate Orbiter flights

and a cargo-bay arrangement with the on-orbit servicing equipment

attached to the Multi-Mission Modular Spacecraft (MMS) flight support

system (FSS). The first flight was to demonstrate module exchange in a

variety of ways while the second flight was to demonstrate propellant

resupply. A disk shaped spacecraft mockup was recommended to support

the modules to be exchanged. The resulting 0-g demonstration was

judged to be too expensive.

The emphasis has been changed for this study to a minimum cost 0-g

demonstration that incorporates a few important servicing functions.

However, a range of additional servicing functions are identified that

can be incorporated into the plan. The demonstration of MMS servicing

also received increased emphasis in this study. As a result, the

following guidelines, or requirements, for development of the 0-g

flight demonstrations were identified:

1) There will be only one demonstration flight;

2) The MMS triangular structure will be used as the basis for the

spacecraft mockup;

3) The MMS module exchange direction will be axial with respect to the

servicer;

4) A battery module with side Interface mechanism will be used to

demonstrate radial direction exchange;

5) Access to the PM-1 propulsion module will be provided for

demonstrating propellant resupply;

6) The short (45 in.) servicer arms will be used instead of the long

(79 in.) servicer arms;

6-4
7) It is desirable to carry the MMS equipment to orbit on its flight

support system;

8) The MMS docking probe adapter will be attached directly to the

servicer without the servicer's primary docking equipment.

The single demonstration flight is recommended as a way of reducing

costs. Use of the MMS triangular structure as a basis for the

spacecraft mockup increases the involvement of MMS equipment in the

demonstration. It also should reduce cost because a set of flight

qualified hardware is available at GSFC. The MMS module exchange

direction from the spacecraft was selected to be axial as that is the

direction recommended for normal operations (see Section 3.0 and Figure

3-4). The servicer docking probe is fitted with a hinged adapter so

the servicer base (docking post) can be swung to either side permitting

the MMS modules to be exchanged in -an axial direction. This simplifies

the module exchange process and permits exchanging modules on either

side of the grapple fixture (used for docking). If all three modules

are to be exchanged, then a second docking must be performed. It was

recommended that each future MMS spacecraft be fitted with two grapple

fixtures so that any of the three modules could be exchanged.

With regard to requirements 3), 4), and 5) a representative list of

candidate activities that might be included is presented in Table

6.1-1. The following paragraphs summarize the different categories.

The module types include the several typical modules and various types

of connectors. The two basic types of module exchange are represented

by a MMS module (axial exchange) and a battery module (radial

exchange). The propellant tank module is considered a special case as

fluids are more likely to be resupplied than their container replaced.

A remote servicer will be capable of opening access doors and

performing simple servicing functions. These might involve the

replacement of batteries or the removal of samples for return to earth.

6-5
Table 6.1-1 Candidate Cargo-Bay Demonstration Activities

Module Type

- Battery module

- Multi-Mission Modular Spacecraft type module

- Propellant tank module

- Electrical connection interface unit

- Propellant resupply module with interface unit

- Access door

- Electrical connector

- Fluid in-line coupling

- Wave guide connector

- Fiber optics connector

- Thermal connector

Interface Mechanism Type

- Lightweight side interface mechanism

- Alternative interface mechanism concepts

- Hinged access cover drive

Special Tools

- MMS module servicing tool

- Other interchangeable adapter tools

- Refueling/resupply interface unit

- Hose or cable management device

- Propellant in-line coupling drive

Direction of Module Motion

- Near axial

- Far axial

- Near radial

- Compound motions

The two most used types of interface connectors are electrical and

fluid. Most module interfaces include one or the other of these two.

The electrical interfaces can include electrical power,

instrumentation, data, and control signals. The size and number of

connectors varies with the application. Fluid connectors can handle

propellant, coolants, pressurants, and fluid supply. The transferring

of cryogenic fluids is considered a special subset of fluid

6-6
connectors. There are a. number of advanced types of connectors that

are not widely used. These include connectors for wave guides, fiber

optics, and thermal energy transfer.

There are a number of different types of interface mechanisms that have

been developed for use with modules. Most of them have been developed

for specific applications. The servicer is not affected by the actual

type of mechanism used, but it does need to have a standard interface

or an adapter tool to operate the mechanism. A mechanism was developed

as part of an earlier IOSS contract that attaches to the side of the

module. The servicer arm interface consists of an attachment point and

a drive unit. There is a need to standardize these interfaces,

allowing for the necessary functions. The three primary functions are

the physical attachment, a method of operating the interface mechanism,

and an electrical connector to communicate with and control the

module. It is possible to replace -some of these functions with the

interfaces of the docking probe of the servicer, but a standard

interface unit should have all three.

Special tools are required to perform some servicing activities. This

is especially true when the original spacecraft element was not

designed for remote servicing. Their use permits the servicer to be

extremely adaptive to a variety of needs. There are a number of tools

that have been developed for special functions. For example, an

interface unit has been developed by NASA - JSC which permits the

resupply of fluid systems. In addition to the actual interface

mechanisms, a system is required to manage the hoses. These types of

special tools greatly enhance the potential of the remote servicer.

There are a number of ways by which a module can be removed. The

first, and probably the easiest is to remove the module in the axial

direction. The trajectory of the module being removed is parallel to

the axis of the docking probe of the servicer. The second basic

trajectory is to remove the modules in the radial direction. Again

this is with respect to the docking probe axis. These two cases

require the servicer to move in a single direction during the removal

6-7
process. By being along one of the principal axes of the servicer, the

control is reasonably simple. The third case is where the module is

removed along a direction that is not parallel to one of the principal

axes of the servicer. This requires the servicer to move in a compound

motion. This type of operation requires much more complicated control

software. Although the IOSS can perform any of the three trajectories,

given the proper software, it is recommended that the first two

approaches be used whenever possible.

A minimum set of activities from Table 6.1-1 is:

1) Exchange of an MMS module in an axial direction;

2) Exchange of a battery module with a side interface mechanism;

3) Resupply of propellant through -an umbilical connection.

These three activities are the most important ones and cover all of the

basic servicing functions. The other candidate activities of Table

6.1-1 can be considered for addition to the basic plan at a later time.

The short (45 in.) servicer arm lengths are recommended for the

cargo-bay demonstration rather than the longer (79 in.) arm lengths

recommended for the operational servicer. All three recommended

activities can be accomplished with the short arms and the resulting

configuration should be less expensive. Shorter arms imply lower joint

torque and thus smaller joints and lower electrical power

requirements. The shorter arm lengths are also appropriate if it is

decided to use the cargo-bay demonstration servicer as the 1-g unit

after completing its flight objectives.

With regard to Requirement 7, the decision to use the MMS triangular

structure as a basis for the spacecraft mockup leads to the

desirability of using the MMS flight support system to support the MMS

in the Orbiter cargo-bay during launch and return. The FSS was

6-8
designed to support the MMS and its use avoids the need for designing

and flight qualifying an alternative support structure. A fully

qualified FSS is available for use.

Figure 3-4 shows two docking mechanisms used between the servicer and

the MMS, one on either side of an orientation joint. The docking

mechanism closest to the servicer is all that is needed for most

spacecraft servicing missions. However, the triangular configuration

of the MMS led to the need for the orientation joint so that the MMS

modules could be removed axially with respect to the servicer. The

orientation joint permits the docking post to be straight during

docking so that this operation is easier. The joint also permits

reorientating the stowage rack so that it can be aligned for axial

module removal of the modules on either side of the grapple fixture.

The concept is that the orientation joint, and the second docking

mechanism, would be an adapter that- would only be used on those flights

involving an MMS. However, as the 0-g demonstration flight will only

involve an MMS configuration there is no need to be able to dock

without the orientation joint. Thus, it is recommended that the

previously proposed docking mechanism closest to the stowage rack be

deleted in order to reduce cost and to increase stiffness of the

docking post (Requirement 8).

An electrical connection is necessary to demonstrate that the servicer

can take control of a disabled spacecraft and properly shut off and

turn on spacecraft equipment to conduct orderly module exchanges. The

recommended docking system is the RMS end effector and the grapple

fixture with electrical connections. These equipments have a standard

connector that can be used to provide the electrical connections

necessary for control of the spacecraft subsystems.

A number of requirements regarding involvement of support systems for

the on-orbit servicer were identified in the prior study. These are:

6-9
1) The Orbiter Remote Manipulator System (RMS) should be involved in

the demonstration because the RMS will be part of many servicer

operations scenarios;

2) Electrical power, attitude control, data transfer, and thermal

control assistance should all be provided by the Orbiter as the

Orbiter is the carrier vehicle, which normally provides these

functions;

3) Servicer control station should be provided in the Orbiter to

simplify the communications paths for control;

4) The adequacy of ridigizing the docking connection as well as its

alignment accuracy should be checked as part of the cargo-bay

demonstrations.

Additionally, the cargo-bay demonstration objectives of Section 4.3.1

can be thought of as programmatic requirements and they were considered

in developing the definition of the cargo-bay demonstration.

There will also be other sets of requirements placed on the equipment

by other interfacing systems. One example is the set of safety

requirements placed on any equipment that is carried or operated in the

Orbiter. The Johnson Space Center has established a process for the

safety review of all such equipment. A series of four safety reviews

are conducted on each project. Areas of concern include: structural

strength; potential for debris in cargo-bay; rupture of sealed

containers; flammability and toxic outgassing of materials used in the

crew cabin; and effects of STS environments especially vibroacoustics.

These requirements must be met to the satisfaction of JSC safety

personnel.

6.2 CARGO-BAY DEMONSTRATION ELEMENTS

The elements involved in the cargo-bay demonstration can be separated

into the following groups:

6-10
1) Space Transportation System Support;

2) Servicer System;

3) Spacecraft Mockup;

4) Airborne Support Equipment;

5) Ground System.

Each of these groups are discussed in turn. Certain of the cargo-bay

demonstration elements are shown in Figure 6.2-1 so they may be more

readily visualized in the following discussion. The flight support

system is part of the airborne support equipment and will be provided

by GSFC. The Multi-Mission Modular Spacecraft is the spacecraft mockup

where the basic triangular structure, the MMS module, and the PM-1

propulsion module will be provided by GSFC. The ground monitoring

station can be located at the JSC Mission Operations Control Center or

at the MSFC Huntsville Operations Support Center.

The servicer control station will be mounted on the Orbiter aft flight

deck so that the operator can look out one of the windows and view the

module exchange operations. A special stowage rack configuration has

been selected for the cargo-bay demonstration to minimize the launch

and return length in the cargo-bay. The servicer mechanism is shown

mounted on a representative docking post. An RMS end-effector is used

as a docking mechanism and is mounted at the end of the docking post

along with the orientation joint used for MMS module exchange. Also

shown on the stowage rack is the fluid resupply module with its fluid

tanks and hose management device.

The hardware that must be specially built or obtained for the cargo-bay

demonstration is shown in Figure 6.2-2 organized into the groups used

in the following discussion.

6-11
•

I

H3a- U§I

CSISc

Figvire 6.2-1

6-12
Total System

Flight System Ground System

Servleer Spac

r ~i

ecraft Airborne Ground Ground

System Mockup Support Support Monitoring

— Servicer Unit

Structure

Mechanisms

Electronics

Docking Probe

— Stowage Rack

Structure

Module Fixtures

Fluid Supply

Fluid Resupply I/F

MMS Servicing Tool

— MMS Structure — Flight Support

— MMS Modules 1— STS Control Station

Modules Computer

Fixtures Video Display

— Battery Modules

Modules

Fixtures

— Propulsion Module

Fluid Resupply I/F

— Grapple Fixtures

Modules Included

as part of S/C

— Video

Figure 6.2-2 Cargo-Bay Demonstration Hardware Breakdown

The equipment organization of Figure 6.2-2 was also used for the Work

Breakdown Structure (WBS), which is presented in Appendix A. The WBS

was prepared to aid in organizing the project so that major cost items

and activities such as environmental tests would not be overlooked.

During the prior study, demonstration sequences were prepared for each

of the two cargo-bay demonstration flights. The first flight involved

an estimated 1010 min and the second involved an estimated 725 min.

Much of that time involved using the manual control mode to perform

servicing operations. With the exchange of fewer modules and use of

only the unassisted Supervisory control mode, the flight operations

should take approximately nine hours. Thus, the servicer system

demonstration should be effected in one day by the astronaut crew.

6-13
6.2.1 Space Transportation System Support

Certain of the equipment required for the servicer cargo-bay

demonstration Is auxiliary equipment available for use on or with the

Orbiter as part of the Space Transportation System. This equipment is

listed in Table 6.2-1. Its provision, control, and use should present

no difficulties.

Table 6.2-1 Space Transportation System Equipment

Space Shuttle including Orbiter

Remote Manipulator System

Attitude control, electrical power, data processing, and thermal

control assistance

Communication links

Cargo-bay cameras

RMS cameras

Control center consoles

As control of the module exchange process will be totally at the

Orbiter, the need for continuous two-way communications with the ground

is lessened. Ground personnel will monitor the servicing operations,

however, continuous monitoring is probably not required. Ground

personnel can only tell astronauts what is happening and advise them of

suggested actions. They can not control the servicer mechanism

directly.

It may be useful to have a second astronaut, perhaps the RMS operator,

controlling the RMS and cargo-bay cameras. The second astronaut could

also operate a hand-held motion picture camera to supplement data

collection by the video cameras.

6.2.2 Servicer System

The servicer system includes the servicer unit, stowage rack, and video

system. The servicer unit includes its structure, the mechanism,

6-14
ORIGINAL PAGE IS

OF POOR

electronics, and docking probe. The servicer system structure is the

tube connecting the mechanism to the stowage rack.

A representation of the servicer system mechanism is shown in Figure

6.2-3, which is a photograph of the Engineering Test Unit with the

counterbalances blacked out. The 0-g servicer mechanism will

externally look much like the ETU and will have the same length arm

segments (45 in.). The joint ordering will be the same although the

design of some joints, e.g., shoulder pitch, may well be different.

The position feedback elements will also be different.

Figure 6.2-3 Representative Servicer Mechanism Configuration

The 0-g electronics will be packaged for mounting on the stowage rack.

The electronics will include most of the Servicer Servo Drive Console

functions such as: motor drive amplifiers; mode control logic; signal

conditioning; precision power supplies; and interfaces with the

computer and the Orbiter data processing and communications systems.

6-15
The docking probe will include an RMS end-effector with electrical

connector and the orientation joint, or hinge, for aligning the stowage

rack with the appropriate faces of the MMS triangular structure. The

orientation joint will have additional travel so that the docking probe

can be folded back on itself. This will allow the stowage rack to be

mounted higher in the cargo bay for better viewing of the module

exchange process while providing clearance for closing the cargo-bay

doors.

The stowage rack configuration is shown in Figure 6.2-4. It is a

simple truss configuration made especially for the cargo-bay

demonstration. The stowage rack mounts in the Orbiter sill trunnions

and uses one keel fitting to provide lateral location. In addition to

the "good" battery and MMS modules, temporary locations for each of

these modules is provided. There is also a storage location for a 0-g

Module Servicing Tool modified for use with the servicer system. The

fluid resupply system is shown as a two tank module on the forward end

of the stowage rack fastened to the main structural cross member. A

hose management system and umbilical connection mechanism are included

in the fluid resupply module. The Intent Is to overdesign the stowage

rack to avoid the need for environmental test except perhaps for the

non-operational vibroacoustics test.

The video system will consist of a charge coupled device (CCD) video

camera mounted on the servicer mechanism end-effector and a camera

control box located on the stowage rack. A CCD camera is used because

it will not be damaged if it inadvertently looks at the sun and because

the sensor has a very short length that satisfies the packaging

requirements of the servicer end-effector. The resolution of the video

camera need only be on the order of a 200 line TV system.

6-16
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60

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60

Figure 6.2-4

6-17
6.2.3 Spacecraft Mockup

The spacecraft mockup is based on the use of a MMS module support

structure as shown in Figure 6.2-5. A development version of this

triangular structure is available at GSFC and it could be used for the

cargo-bay demonstration. Its use not only saves design and fabrication

costs, but also should save some testing cost as the unit was designed

and built to meet the STS environments. While the figure shows one

grapple fixture, a second grapple fixture is needed for the cargo-bay

demonstration.

Figure 6.2-5 Module Support Structure

The spacecraft mockup will also require a MMS module, a battery module,

a fluid tank, fluid connections and control, wiring, and an electrical

umbilical at one of the grapple fixtures. The MMS module will also be

provided by GSFC. A complete MMS module is not required. Only the

physical and electrical interfaces and the module structure are

6-18
required. Several mockup modules are available at GSFC that should be

adequate. The battery module with its side interface mechanism for

radial module removal and electrical connections will be new. The

fluid tank can be a PM-1, a PM-1A, or any similar tank that is

available and can be easily qualified for flight. A set of fluid

piping equipment including flow control, flow measurement, quick

disconnect, tank pressurization, etc. will also be required. It is

expected that this equipment can be obtained on loan from GSFC, or from

the on-going JSC fluid resupply program. Some work will be required to

integrate the fluid handling equipment into the MMS module support

structure. The wiring will be for control, status monitoring and

thermal control. It will be designed and fabricated for this specific

use.

One method for arranging the three candidate activities on the MMS

triangular structure is shown in Figure 6.2-6. The MMS module would be

attached on side A in the normal way. Side B would have the battery

module mounted to demonstrate radial module exchange. The PM-1

propulsion module umbilical connection is shown near the end of the

structure where the PM-1 is normally fitted. A special fitting would

need to be provided to demonstrate propellant resupply.

Figure 6.2-6 also shows the 20 in. minimum operating radius and the 82

in. maximum operating radius that correspond to the 45 in. arm segment

lengths of the servicer Engineering Test Unit. A servicer mechanism

with these arm lengths is recommended as it will be less expensive that

one with the 79 in. arm segment lengths of the flight version of the

Integrated Orbital Servicer System mechanism. One effect of the short

arm lengths is that the battery module will have to be one to three

inches smaller than the 24 in. cube used for the 1-g demonstrations.

However, this is not felt to be a significant factor.

6-19
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Figure 6.2-6

6-20
A preliminary layout of the allowable battery module size based on ETU

mechanism arm lengths was made. In addition to requiring a smaller

module size, it was found that the module attach point would have to be

approximately five inches back from the MMS module centerline. There

is also an interaction of the relative locations of the MMS module

attachments, the battery module attachment, and the second grapple

fixture location. It is desired to have the battery module motion

along a radial line from the docking post. The primary MMS grapple

fixture is in a plane half way between the MMS module attach points.

Use of the secondary grapple fixture location, nearer the lower

trunnion fittings, as shown in Figure 6.2-6, provides a larger space

for the location of additional servicing functions.

6.2.4 Airborne Support Equipment

The airborne support equipment consists of the MMS Flight Support

System and a control station on the Orbiter Aft Flight Deck. The FSS

is shown on Figure 6.2-1 and consists of cradles A, A' and B. All

three cradles are necessary to properly support the spacecraft mockup

and to deploy it for easy capture and transfer. Operations with the

FSS have been developed and all of the necessary ancillary equipment is

available. It is also fully redundant in terms of operational

functions. Equipment for connecting into Orbiter systems and a control

and display panel exist.

The prior study recommended that the servicer control station be

located on the ground because that case is representative of the

operational case and so that the effects of the two way communication

links could be brought into the demonstration. This was not judged to

be a low cost approach because of the ground station costs and the

Orbiter integration costs. The minimum cost case will be for servicer

control from the Orbiter if the crew training and integration costs can

be constrained. The increased involvement of MMS equipment, especially

the flight support system, and the RMS involvement imply that a certain

amount of crew training will be necessary in any case.

6-21
A minimum cost approach would Involve use of only a few control modes

during the 0-g demonstrations. The backup, or Manual-Direct, control

mode is necessary to overcome computer failures. It is recommended

that various forms of the Supervisory and Manual-Augmented control

modes be investigated in the 1-g demonstrations and that one control

mode, in addition to Manual-Direct, be selected for the 0-g

demonstrations. It will be assumed here that the selected mode will be

a form of the unassisted Supervisory control mode. If another control

mode is selected, then appropriate changes in the 0-g demonstrations

will need to be made.

The use of a ground control station, inclusion of the two way

communication link effects, and involvement of other control modes can

be considered as alternative servicing functions that can be brought

into the 0-g demonstrations if time and funding permit.

The resulting control station at the Orbiter Aft Flight Deck will

Include a minicomputer, a keyboard and display unit, and a television

monitor for display of the servicer end effector camera picture.

Standard Orbiter interconnecting cables and interfaces to the Orbiter

communication links will be used. Where possible, equipment will be

borrowed, or rented. It is expected that 10 to 12 thousand lines of

code will be required in the airborne computer. Depending on the

computer selected, 256K bytes of Internal memory and 5 M bytes of hard

disk memory will be required.

6.2.5 Ground System

The ground system consists of ground support and ground monitoring.

The ground support consists of the usual development, test, and

handling equipment for the flight articles as well as use of the 1-g

servicer system demonstration facility at MSFC. Development, test, and

handling equipment for GSFC supplied flight equipment will also be

provided by GSFC. The remainder of this equipment will be new. It is

expected that the MSFC 1-g servicer demonstration facility will be set

up to represent the flight situation and that it will be used for

procedures development, flight crew training and anomaly investigation.

6-22
The ground monitoring equipment will consist of consoles and display

equipment located either at JSC or at MSFC. Existing equipment should

be adequate with some software development to be able to generate

appropriate displays. Data will consist of joint angles, trajectory

coordinates, TV views, and discrete events. The ground monitoring

personnel will follow the planned scenarios and check that all planned

events occur in the proper order and that no unplanned events occur.

Ground personnel will not be able to control the servicer demonstration

directly. Rather they will be able to inform the flight operator of

any pertinent observations.

6.3 CARGO-BAY DEMONSTRATION CONFIGURATION ALTERNATIVES

The revised basis for the cargo-bay demonstrations indicated that the

configuration, or arrangement of the flight equipment should be

reevaluated. This section of the study final report identifies a range

of alternative arrangements, does a preliminary screening, describes

the resulting configurations, identifies selection criteria, and

recommends a specific configuration.

6.3.1 Alternative Configurations

The major elements to be arranged in the Orbiter's cargo bay for 0-g

demonstration of on-orbit servicing are:

1) The servicer mechanism and stowage rack;

2) The spacecraft mockup (based on the MMS triangular structure);

3) The MMS flight support system.

The arrangement for the servicing demonstration operations was

addressed first and the arrangement for launch and return was evolved

from the operational arrangement. The range of operational

arrangements considered first were based on the following

considerations:

6-23
1) The MMS triangular structure could be horizontal, vertical, or

transverse with respect to the Orbiter. Variations from these

cardinal directions were to be considered later if there were any

specific advantages identified. The one variation identified had

potential interference problems with the Orbiter cargo-bay doors

and was dropped from further consideration. The MMS centerline

directions result in three possibilities;

2) The servicer docking post, at the stowage rack, must be

perpendicular to the MMS centerline. This is necessary so that the

MMS is oriented for axial MMS module removal. The servicer docking

post orientation increases the number of possibilities by a factor

of two;

3) The servicer can be attached to the Orbiter, FSS, or MMS.

Attachment to the MMS would be via the docking probe. The servicer

attachment variations increase the number of possibilities by a

factor of three;

4) The MMS will be attached to the servicer. However, the MMS could

also be attached to the FSS if the servicer is not attached to the

Orbiter or the FSS. No new arrangement possibilities are

introduced by this consideration.

The number of possible arrangements is 18. Each of the candidates were

defined and evaluated as shown in Table 6.3-1. The results of the

evaluation are:

1) Five candidates are valid arrangements.

These are: 2, 3, 9, 11, and 16;

2) Four candidates were variants of valid arrangements. Each variant

was a 90 degree rotation of the MMS centerline or of the docking

post direction. The variants are: 1 of 16, 8 of 11, 15 of 9, and

18 of 3;

6-24
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6-25
3) Four candidates were inconsistent in that an element orientation

was transverse and it was to be attached to the FSS. However, the

FSS cannot accept a transverse orientation. The inconsistent

arrangements are: 4, 10, 14, and 17;

4) Three arrangements would not fit because the docking post had a

transverse orientation and the stowage rack was to be down in the

Orbiter cargo bay. These arrangements are: 5, 6, and 12;

5) Two arrangements had the servicer docking post axis in a horizontal

orientation with the stowage rack attached to the FSS. These

arrangements used only a minimum of the FSS capabilities. These

arrangements are: 7 and 13.

The five valid candidate arrangements were selected for further

evaluation.

6.3.2 Configuration Descriptions

Simple layouts were made for each of the five valid configurations and

for a variant of arrangement 2. Figure 6.3-1 shows arrangement 2 in

the operational and stowed configurations. The MMS is attached to the

FSS and the centerline of the MMS is horizontal in the cargo bay. The

servicer is cantilevered off the MMS and is not otherwise supported.

There is marginal clearance around the mockups for servicer mechanism

operation. When it is desired to go from demonstrations of MMS module

exchange to demonstrations of battery module exchange, it is necessary

to rotate the MMS using the FSS rotational system. As the MMS is

rotated, the orientation joint in the servicer docking post is driven

in the opposite direction so the stowage rack remains in a horizontal

plane with respect to the Orbiter. This process is called a dual

rotation to differentiate it from those arrangements that only involve

operation of the servicer docking post orientation drive when switching

from MMS to battery module demonstrations.

6-26
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6-27
The stowed configuration shows the use of the FSS cradles A, B, and A

prime to properly support the MMS during launch and return and to

reorient the MMS during operations. The orientation joint on the

servicer is used to fold the docking mechanism out of the way and

thereby reduce the overall length of the stowed configuration. The

docking post is not folded back. That servicing function is left as an

alternative for later consideration. It would be good to be able to

test the servicer mechanism deployment mechanism and it is advantageous

to decrease the stowed length of equipment in the Orbiter cargo bay.

However, the recommended approach is to delete servicer functions and

reduce cost.

There are a number of difficulties with arrangement 2. It will be

necessary to have a long umbilical from the servicer to the Orbiter for

power, commands, and data transfer. It is not as easy to arrange for

jettisoning this configuration as the docking mechanism, or the

servicer mechanism, may be connecting the stowage rack to the MMS. Two

sets of separation devices must be provided or else the MMS, and

possibly the FSS, must be jettisoned with the servicer.

The most flexible part of the docking post and mechanism is the RMS end

effector to grapple fixture attachment and it is next to the MMS. This

means that the large mass of the stowage rack is cantilevered far out

from the soft point in the structure. The result is an arrangement

with a low stiffness and a low natural frequency.

Arrangement 3 is shown in Figure 6.3-2. The MMS is docked to the

servicer that is fastened into the Orbiter cargo bay. The docking post

is vertical and the MMS centerline is parallel to the Orbiter

centerline. The MMS centerline could also be transverse to the Orbiter

centerline, but viewing from the aft flight deck windows would not be

as good. The use of intermediate angular positions to improve the

viewing can be investigated during detail design. The operational

clearance is excellent and the full range of arm motions can be

exercised.

6-28
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6-29
The stowed configuration for the MMS is similar to that for arrangement

2. The stowed arrangement for the stowage rack is the same as for the

operational configuration except that the docking mechanism is folded

back to allow the cargo-bay doors to close. A special form of the

stowage rack is shown that is specially configured for the 0-g

demonstration equipment and module arrangements. The arrangement 3

configuration of the stowage rack can also be adapted to some of the

other arrangements.

The operational configuration of arrangement 3 does not include having

the MMS fastened to the FSS. In this sense, some of the capabilities

of the FSS are not well used. The stowage rack configuration does not

lend itself to natural containment of loose modules during reentry and

landing. Thus it will be necessary to carefully address how to assure

that all modules and the servicer mechanism are well secured or

jettisoned before deorbit is initiated.

Arrangement 9 is shown in Figure 6.3-3. The MMS is docked to the

servicer that is fastened into the Orbiter cargo bay with the docking

post horizontal. The MMS centerline is shown vertical to the Orbiter

for better viewing from the aft flight deck, although the MMS

centerline could also be transverse to the Orbiter centerline. The

operational clearance is acceptable but there are some combinations of

servicer mechanism position and MSS module orientation that must be

avoided.

The stowed configuration of the MMS involves only the FSS cradle A

prime^ The other cradles were deleted to reduce length in the cargo

bay. The servicer system stowage is similar to arrangement 2 except

that, it is stowed farther away from the FSS. This was necessary to

provide room for the RMS end effector when docking the MMS to the

servicer. The stowed length is almost 60 in. longer than the other

arrangements. There are several ways to reduce the stowed length if it

should become critical. The most obvious way is to have the stowage

rack near the FSS during launch and then move it back, using the RMS,

when setting up for the servicing demonstrations on orbit.

6-30
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6-31
The view of a module exchange from the Orbiter aft flight deck will be

obscured either by the stowage rack or by the FSS. The MMS is not

attached to the FSS during operations so the full capability of the FSS

is not used. There is a concern that the RMS reach might not be

adequate to dock the MMS to the servicer. A second grapple fixture is

required.

Arrangement 11 is shown in Figure 6.3-4. The MMS is attached to the

FSS with the MMS centerline vertical with respect to the Orbiter. The

servicer is docked to the MMS with the docking post axis transverse to

the Orbiter. The docking post could have been parallel to the Orbiter

centerline except that the view from the aft flight deck would be

degraded. While not obvious from the figure, a portion of the stowage

rack has been cut off so that an adequate clearance exists over the

cargo-bay sill.

It is necessary to have an electrical umbilical between the cargo bay

and the servicer to conduct commands, data, and power. It is difficult

to jettison the servicer without undocking the MMS should the servicer

fail while it is docked and attached to a module. The stowed length is

near average for all arrangements and is such that the stowage rack

will readily contain any improperly stowed parts during reentry and

landing. As the servicer and stowage rack are cantilevered from the

MMS, the docking system stiffness is low and could be a problem when

the Orbiter thrusters fire. This is especially dif-ficult if the arm is

trying to place a module in the stowage rack when the thrusters fire.

The location of the docking post, only 38 in. above the Orbiter sill

clearance line, restricts the range of operation of the servicer arm

and will require careful planning to assure that proper clearance is

maintained. This means that it is unlikely that representative

trajectories can be used. As with arrangement 2, the servicer

orientation hinge must be rotated in opposition to the FSS rotational

drive when switching from modules on one side of the MMS to modules on

the other side.

6-32
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6-33
Arrangement 16 is shown in Figure 6.3-5. This was the preferred

arrangement from the prior study before the MMS was used as the

spacecraft structure. The stowage rack is mounted on the FSS and the

FSS is used to elevate the stowage rack out of the cargo bay and to

rotate the stowage rack and MMS so that it can be seen better from the

Orbiter aft flight deck. The MMS could also be positioned transverse

with respect to the Orbiter. However, there is little or no difference

in the two arrangements. This configuration has the shortest stowed

length as the MMS can be fitted in next to the servicer docking probe.

A separate, new structure is needed to support the MMS in the cargo bay.

Wiring from the Orbiter to the servicer can be routed directly through

the FSS cable management system and umbilical connections. This

arrangement may cost a little more than the other arrangements because

of the need to provide a separate launch and return support system for

the MMS. The servicer stowage rack, will require additional structure

and adapters to properly interface with the FSS. The MMS is not used

with the FSS although these two were designed to be used together. The

MMS also needs a second grapple fixture.

A variation of arrangement 2 is shown in Figure 6.3-6. In this case

the MMS grapple fixture is vertical in the Orbiter cargo bay and the

docking post is tipped 60 degrees away from the Orbiter vertical. The

variation avoids the dual rotation requirement of arrangement 2.

However, there is an interference between the stowage rack and the

Orbiter sill clearance volume. This means that the stowage rack must

be cut off on both sides. There are also restrictions on the servicer

arm motion and on the trajectories that can be accomodated. Because of

these restrictions, it was decided to drop the variant from further

consideration and to retain arrangement 2 for further evaluation.

6.3.3 Use of the RMS During 0-g Demonstrations

It is recommended that the Orbiter remote manipulator system (RMS) be

used for docking the MMS to the servicer or the servicer to the MMS.

6-34
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6-35
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6-36
The RMS will be available for the 0-g demonstrations and will also

probably be used for assembling the servicer to the OMV for some

missions. It could also be used for docking failed spacecraft to the

servicer in the Orbiter cargo bay as an alternative to EVA servicing.

When the RMS is being used for docking or berthing a spacecraft, the

RMS wrist camera view of the berthing location will be obscured by the

spacecraft being handled. The RMS has a capability for moving its end

effector along a straight line path from point A to point B under

computer control. The A and B points can be input before launch or

during flight. More than one segment can be input and used. However,

there are many error contributors including thermal deflections of the

cargo-bay sill at the RMS attachment point. The result is that the

location of the berthing point on the spacecraft relative to the

berthing latch locations are not well known. In particular, the RMS

errors can exceed the capture volume of the recommended docking

mechanism. The recommended docking, mechanism is the same as for the

Orbital Maneuvering Vehicle (OMV) and is the RMS end effector and

grapple device.

One advantage of the recommended docking process for the cargo-bay

demonstration is that one side of the docking interface is fixed and

not moving. An approach for assisting the docking is to use the

servicer camera. The orientation drive can be positioned so that the

docking mechanism is colinear with the docking post. The servicer arm

would be positioned so that the servicer camera views a similarly

positioned target. When the target is aligned in the servicer camera

field of view, then the docking mechanism and the grapple fixture would

also be lined up. This approach is better if the servicer camera

picture could be put on the RMS wrist camera display. Another aid

would be to transform the RMS hand controller signals into the servicer

camera coordinates. However, this last transformation might involve

too much interfacing between the RMS and servicer systems. It should

be possible to orient the servicer TV camera so that its axes are

parallel to the docked position grapple fixture axes. This might be a

simple way to reduce the RMS to servicer system TV display cross

coupling effects.

6-37
The auxiliary cargo-bay cameras can also be positioned so their view

helps in the docking operation. These cameras can additionally be used

for viewing and recording the module exchange operations. After the

RMS has completed the docking operation, it should be disconnected from

the spacecraft to simplify operations when the MMS is rotated to those

positions where module exchange can take place. The RMS wrist and

elbow TV cameras could also be used to monitor and record the module

exchange operations.

6.3.4 Configuration Selection Criteria

The arrangement, or configuration, selection criteria were identified

by reviewing the criteria used in the prior study and considering the

advantages and disadvantages of the candidate arrangements. It was

possible to sort the criteria into two categories—those that apply

equally to all five arrangements and those that can be used in a

comparative evaluation of the five candidate arrangements. Those

requirements that apply equally to all five candidates are listed in

Table 6.3-2. These requirements were discussed in some detail in the

prior study and in "Integrated Orbital Servicing Study Follow-On

Implementation and Test Program Plan", MCR-76-258, Martin Marietta .

Corporation, April 1978. The Table 6.3-2 requirements should be

considered in subsequent levels of definition of the Orbiter cargo bay

servicer demonstrations.

Those criteria and requirements that can be used to help select a 0-g

servicer demonstration arrangement are listed in Table 6.3-3. Certain

of these requirements were carried over from the prior study while

others are more directly related to this particular evaluation.

Emphasis has been given to the use of MMS program equipment in the

manner for which it was designed. The comparative data was obtained by

comparing the layouts that went into the preparation of Figures 6.3-1

through 6.3-6. Generally the items near the top of Table 6.3-3 are

considered to be more important. The data entries of Table 6.3-3 have

been addressed as part of the Section 6.3.3 discussion for each

arrangement.

6-38
Table 6.3-2 Equally Applicable Cargo-Bay Demonstration Requirements

1. Enhance User Acceptance of On-Orbit Servicing

2. Incorporate Representative Servicing Operational Equipment

3. Include Verification of Procedures, Analysis Techniques, and

1-g Simulations

4. Adaptability to Changes in Knowledge Level

5. Compatible with OMV Development Schedule

6. Costs That Are Phased to User Acceptability

7. MMS Involvement

8. Provision of Electrical Power, Attitude Control, Thermal

Control, and Communications

9. Spacecraft Deployment and Retrieval by RMS

10. Use of RMS for Docking

11. Docking Misalignment Effects

12. Spacecraft to Servicer Alignment

13. Servicer Mechanism Performance

14. System Force and Torque Levels

15. Repeatability Accuracy (Electro/Mechanical)

16. End Effector Capture

17. Methods of Accommodating Attach Errors (compliance)

18. Interface Mechanism Performance

19. Interface Mechanism Capability for Capture, Latch, Unlatch

and Release

20. Connector Performance including Mate and Demate -

Electrical, Waveguide, Thermal, and Fluids

21. Control System Mode Validation

22. Man Machine Interaction

23. Servicer Control Station Location

24. Communications Link

25. Lighting

26. Malfunction Mode/Backup System

27. Mission/Man/STS System Safety

28. Pre and Post Module Exchange Condition Analysis

29. Supplementary Visual Aids

30. Supplementary TV Cameras

31. Deployment of Servicer Mechanism and Docking Probe

32. Similarity to 1-g Arrangement

33. Sequence of Specific On-Orbit Activities

34. Module Stowage Space Availability

35. Routing of Wiring to MMS

6-39
Table 6.3-3 Alternative Arrangement Comparison

Consideration

1. Similarity to Operational Use

2. Docking System

Stiffness

3. Ease of Jettisoning

Equipment

4. View of Operations

from Aft Flight

Deck

5. MMS Launch/Return

Support System

6. MMS on FSS During

Operations

7. Stowed Length in

Cargo Bay ( in. )

8. RMS Reach Concerns

9. Routing of Orbiter

to Servicer Wiring

10. Launch/Return

Weight

11. Equipment List

Deltas

12. Relative Cost

13. Operational Clearance

14. Representative

Trajectory Accomodation

15. Range of Servicer

Arm Travel

16. Containment of

Loose Parts During

Return

17. Transition Between

MMS Sides

18. Need for Second

Grapple Fixture

19. Adaptability to

Deployment of

Stowage Rack

Arrangement

2

Good

Poor

Hard

Acceptable

FSS

Yes

231

None

Umbilical

Nominal

Nominal

Nominal

Poor

Acceptable

Full

OK

Dual

Rotation

No

No

3

Excellent

Good

OK

Excellent

FSS

No

226

None

Direct

Nominal

Nominal

Nominal

Excellent

Excellent

Full

Poor

Single

Rotation

Yes

Yes

9

Good

Good

Easy

Poor

FSS

No

290

Yes

Direct

No A or B

Cradles

No A or B

Cradles

Nominal

Acceptable

Acceptable

Full

OK

Single

Rotation

Yes

No

11

Good

Poor

Hard

Excellent

FSS

Yes

231

None

Umbilical

Nominal

Nominal

Nominal

Acceptable

Good

Reduced

OK

Dual

Rotation

No

No

16

Excellent

Good

Easy

Excellent

Separate

No

172

None

Via FSS

No A or B

Cradles

No A or B

Cradles

Add MMS

Support

Add Some

Excellent

Excellent

Full

OK

Single

Rotation

Yes

No

6-40
6.3.5 Configuration Selection

The data of Table 6.3-3 are given a quantitative assessment in Table

6.3-4. A first order comparative approach was used in that a nominal

case was assigned a zero, a better than nominal case was given a plus,

and a less than nominal case was given a minus. Totals for each

arrangement for number of pluses, nominals, and minuses are given at

the bottom of Table 6.3-4 along with the net pluses (total pluses less

total minuses). Arrangement 3 is seen to score a little better on this

basis than arrangement 16, which is the preferred arrangement from the

prior study. The data of Table 6.3-4 may be weighted in a wide variety

of ways. If the table is divided into three parts (6 items, 7 items, 6

items), then arrangement 3 is best in the top and middle parts and

comparable to arrangement 16 in the lowest part. This indicates that

arrangement 3 will be best for a wide range of weighting approaches

that give higher weights to the items near the top of Table 6.3-4.

Arrangement 3 is given a negative for item 6 of Table 6.3-4. However,

it is given a positive for Item 2. No way has been identified to get

high stiffness in the docking system while operating with the MMS

attached to the FSS. The next negative is for a low capability to

contain loose parts during reentry and landing. These loose parts

could be due to a failure of a module to be latched down or a failure

of the servicer mechanism to restow itself. If the problem is felt to

be severe during the detail design process, then corrective approaches

will be necessary. The last negative for arrangement 3 is item 18 that

is the need for a second grapple fixture. Again this negative is

concomitant with a plus for docking system stiffness.

Arrangement 3 (see Figure 6.3-2) is recommended. For this arrangement

the MMS is docked to the servicer that has its stowage rack fastened

into the Orbiter cargo bay. The docking post is vertical and the MMS

centerline is parallel to the Orbiter centerline for good viewing of

the module exchange operations from the Orbiter aft flight deck. This

arrangement is further described in Section 6.2 of this report.

6-41
Table 6.3-4 Alternative Arrangement Evaluation

Consideration

1. Similarity to Operational Use

2. Docking System

Stiffness

3. Ease of Jettisoning

Equipment

4. View of Operations

from Aft Flight

Deck

5. MMS Launch /Re turn

Support System

6. MMS on FSS During

Operations

7. Stowed Length in

Cargo Bay (in. )

8. RMS Reach Concerns

9. Routing of Wiring

to Servicer

10. Launch/Return

Weight

11. Equipment List

Deltas

12. Relative Cost

13. Operational Clearance

Arrangement

2

0

-

-

0

+

+

0

0

-

0

0

0

-

3

+

+

0

+

•

+

-

0

0

+

0

0

0

+

9

0

+

+

-

+

-

-

-

+

+

+

0

0

11

0

-

-

+

+

+

0

0

-

0

0

0

0

16

+

+

+

+

-

-

+

0

0

+

-

-

+

6-42
Table 6.3-4 Alternative Arrangement Evaluation (Cont)

Consideration

14. Representative

Trajectory Accommodation

15. Range of Servicer

Arm Travel

16. Containment of

Loose Parts During

Return

17. Transition Between

MMS Sides

18. Need for Second

Grapple Fixture

19. Adaptability to

Deployment of

Stowage Rack

Total Plus

Total Nominal

Total Minus

Net Plus

Arrangement

2

0

0

0

-

0

0

2

12

5

(3)

3

+

0

-

+

-

+

9

7

3

6

9

0

0

0

+

-

0

7

7

5

2

11

—

-

0

-

0

0

3

10

6

(3)

16

+

0

0

+

-

0

9

5

5

4

6-43
A failure analysis of arrangement 3 should be conducted during a

subsequent design activity to identify:

1) Critical failure modes;

2) Redundancy requirements;

3) Methods for overcoming failures;

4) Margins of safety;

5) Backup provisions.

This analysis should result in a more reliable system and will confirm

the reliability of arrangement 3 or indicate any necessary changes.

Similarly a risk assessment analysis is required during a subsequent

design activity to identify:

1) Risk elements or areas;

2) Probability of occurence;

3) Program schedule impacts;

4) Program cost impacts;

5) Risk management actions.

The risk assessment analysis will also help to confirm the viability of

arrangement 3 and will indicate any necessary changes.

6.4 CARGO-BAY DEMONSTRATION SCHEDULE

The servicer cargo-bay demonstration schedule (Figure 6.4-1) was based

on an OMV development schedule. The first two lines show the OMV

design and development schedule and the ground demonstrations schedule

for reference. The servicer cargo-bay demonstration schedule was

developed to correlate with this OMV schedule. The key point from the

OMV schedule was an OMV authority to proceed (ATP) for Phases C and D

in late April of 1986. This date is shown on Figure 6.4-1 along with

other OMV milestones. The start of the cargo-bay demonstration

corresponds with the ATP of Phase C/D of OMV. This start time permits

the completion of the demonstration by the start of the free-flight

6-44
DESCRIPTION

OMV Design &

Development

Ground Demonstrations

System Engineering

Design & Development

Safety

Software Dev & C/O

Procurement

Fabrication

Assembly & C/O

Acceptance Testing

System Integration &

Demonstration

Verification Phase B

1985

i

PHASE B

1986

ATP

[V VDEMONSTRATIONS

I_

1987

COR

1988

SUPPORT

1989

CIR

REQUIREMENTS PROCEDURES TRAINING & OPERATIONS

REVIEW PHASES 0 ll 2 i—v—-v—v

DEVELOPMENT CHECKOUT

DEMONSTRATION

FLIGHT

i^°

Figure 6.4-1 Servicer Cargo-Bay Demonstration Schedule

verification program. The January 1989 date, selected as the beginning

of Phase B for the free-flight verification or operational servicer

development, corresponds with the end of the OMV supporting development

of a servicer kit. This approach integrated well with the use of

representative time spans for the various demonstrations and

verification activities. It was decided that the results of the

servicer cargo-bay demonstration would be most useful if the cargo-bay

demonstration was completed at the start of the operational servicer

development in January 1989. The cargo-bay flight is completed before

the operational servicer preliminary design review (PDR) so that the

cargo-bay demonstration results can be factored into the operational

servicer design and development.

6-45
The cargo-bay demonstration flight program Is recommended to start in

1986 and has been treated more as a development experiment than as a

full fledged program. The normal phases of a program have been

combined into a single, 3 year program. Most of the early work has

been completed as part of the ground demonstration activity. Three

additional support tasks are required from the 1-g demonstration

equipment. The first is for design support and parallels the servicer

design. The second task is for procedures development and the third is

for astronaut training as system operators. Additional astronaut

training will be involved in terms of RMS and FSS operations, data

collection, and assistance in overcoming anomalies.

The safety review phases (0-3) required for Shuttle payloads are

shown on the schedule. The importance of crew and Orbiter safety have

been included in the plan. To insure no major safety related task was

overlooked, the plan was reviewed by one of Martin Marietta's Shuttle

integration safety engineers. The formal design review process has

been reduced, relying on the safety reviews and program annual reviews

to provide the needed customer interaction. This approach was selected

in order to better meet the funding allocations and fulfill the

technical objectives.

The schedule items are representative of this type of project. The

fabrication and assembly of the flight unit and airborne support

equipment (ASE) are estimated to take 13 months. Software development

and documentation for the ASE is included in the software development

block. The monitoring station software, and checkout has been

scheduled to start after the preliminary design is completed and to be

finished by the start of system qualification tests.

A single schedule has been shown for the total of the cargo-bay

demonstration equipment. Separate schedules will need to be prepared

for each of the major sets of equipment, which are:

1) Servicer System;

6-46
2) Spacecraft Mockup;

3) Airborne Support;

4) Ground Support;

5) Ground Monitoring.

Second level schedules were prepared for the demonstration servicer and

the spacecraft mockup during the prior contract. From that work

certain schedule implications can be drawn. The deletion of the second

cargo-bay demonstration flight would save 10 months. Also the time

from program initiation to first flight was reduced by a further nine

months to maintain correspondence with the OMV schedule. The result

was deletion of both Phase B and the separate procurement for Phase

C/D. Some of the Phase B work has been accomplished on the current

contract but the resulting time span is tight.

The schedule makes the assumption that the equipment to be provided by

GSFC can be provided to meet the implied schedule requirements, e.g.,

FSS available throughout 1988 for systems integration and

demonstration. It may be that GSFC has other plans for the FSS in this

time span. Also no attempt has been made to identify a specific

Shuttle flight where the current manifest would permit adding the

on-orbit servicer demonstration activity. These elements should

properly await the identification of an approved funding stream.

However, the decision to go with one Shuttle flight should ease this

type of schedule concern.

The schedule of Figure 6.4-1 assumed funding availability in Fiscal

Year 1986. This funding is not available, thus the Figure 6.4-1

schedule will slip until funding becomes available. If funding is not

obtained relatively soon, then the servicer may not be ready when it is

needed, both for the OMV and for the Space Station.

6-47
6.5 CARGO-BAY DEMONSTRATION COST ESTIMATE

A preliminary cost estimate was prepared for the recommended cargo-bay

demonstration. A primary function of this effort was to reduce the

cost of the cargo-bay demonstration.

The cargo-bay demonstration was treated as a flight experiment rather

than as a flight qualified unit. This changes the technical approach

used in the program, which in turn affects the procedures to estimate

the costs. The approach used was to estimate the program cost of a

flight qualification program using the normal cost estimating

relationships (CER). To these results a complexity factor was applied

to account for the difference between an experimental and flight

qualified unit. The data from the equipment lists of Section 6.2 and

the work breakdown structure of Appendix A were used to establish the

inputs to the CERs.

The complexity factor was arrived at by reviewing recent programs of a

similar nature. These include the Westar/Palapa rescue mission, the

Solar Max repair and a Shuttle demonstration of storable fluid

management techniques. From this data, a complexity factor of 0.4 was

derived.

The total estimated cost of the servicer cargo bay demonstrations is

approximately $9.3 million. A breakdown of this estimate and a summary

of the basis of cost is presented in Table 6.5-1. In estimating the

costs of the cargo-bay servicer demonstrations, the following

assumptions were made:

1) All costs are in 1985 dollars;

2) Costs include the design, development and fabrication of the

experiment hardware;

3) The costs for the Shuttle flight, the standard RMS, the RMS grapple

fixture and the crew time are not included in the cost estimate;

6-48
Table 6.5-1

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6-49
4) The costs associated with crew training are not included;

5) The following items will be provided by Goddard Space Flight Center

and are not included in the estimate:

a) Flight Support System (FSS);

b) FSS checkout, handling and transportation equipment;

c) Multi-Mission Modular Spacecraft (MMS) support structure;

d) MMS checkout, handling and transportation equipment;

e) 2 MMS modules and 3 interfaces;

f) Modified Module Servicing Tool;

NOTE: The cost of integrating the item 5 equipment into the

system is included in the estimate.

6) The ground monitoring station, which consists of a display console

and data handling equipment, is assumed to be an existing NASA

facility and is not included in the estimate;

7) The installation and removal of the experiment hardware into/from

the Orbiter cargo bay will be performed by NASA personnel. The

cost of the NASA portion is not included in the estimate. Martin

Marietta will have a limited number of personnel for support, and

their cost is included in the estimate.

A major objective of the demonstration plan activities, namely to

reduce the cost of the cargo-bay demonstration, has been achieved. The

estimated cost of 9.3 million dollars is more than a 50 % reduction

over the previous estimate. These savings result from reducing the

number of Shuttle flights, utilizing as much existing equipment as

possible, treating the demonstration as a flight experiment, limiting

the amount of qualification and development testing, and limiting the

formal design and review process. This new approach treats the

cargo-bay demonstration as a flight experiment rather than as a flight

qualified spacecraft. A number of changes result, some of which are

not readily apparent. The program structure is less formal, there are

6-50
fewer design reviews, less traceability and lowered customer/contractor

interaction. The standards on the hardware are also reduced.

Commercially available hardware is used, and modified to meet the

requirements if necessary. For example, a commercially available valve

might be used for the fluid resupply system, to meet the sealing

requirements the valve seat may be reworked or replaced, but the

overall cost is much less than for a flight qualified unit. Efforts

will be made to use previously qualified hardware if available. This

will include small components and large systems. The MMS structure and

empty modules will be loaned from Goddard, as well as the MMS flight

support system.

Although these measures reduce the cost, there is a price associated

with them. The risk of an unsuccessful mission from the technical

standpoint is increased (the plan does not increase the risk to the

Shuttle or the crew). This approach has successfully been used by

Martin Marietta in previous flight experiments. The depth to which the

servicing techniques will be demonstrated has been reduced, but all the

major activities have been included. The proposed plan has been

designed to demonstrate the major servicing functions in the most

economical manner possible.

6-51
7.0 FREE-FLIGHT VERIFICATION PLAN

While the objective of the Orbiter cargo-bay experiment was to

encourage potential users of on-orbit servicing in the form of module

exchange and show that the major elements of the system can be

designed, built and operated, the objective of a free-flight

verification is to verify that the equipment is operational and ready

for use. The free-flight verification tests are considered to be the

final proof that establishes an orbital servicing capability. Thus the

design, development, and test process must be suitable for operational

equipment. Similarly all the appropriate documentation must be

prepared so that the capability can be used by others. The prior study

recommended that at least two production units and adequate spares be

procured so there would be a higher availability of servicing equipment

for operational flights. However, the response to that recommendation

was that using programs should pay for all equipment past the first

set. Thus, this study includes costs for only the first production

servicer system.

The emphasis on redefinition of the cargo-bay demonstrations resulted

in little new work on the free-flight verification plan. This approach

is satisfactory as it is still three years until Phase B is scheduled

to start and further 1-g investigations and the evolving 0-g

demonstration will help to better define the free-flight verification

plan. The free-flight verification plan from the prior study was

reviewed and modified to reflect the identified changes. This led to

the revised schedule and cost estimates presented here.

Just as the cargo-bay demonstration will benefit from the ground

demonstrations so will the free-flight verification benefit from both

the ground and cargo-bay demonstrations. The module exchange and fluid

resupply parts of the verification are expected to be similar to the

cargo-bay demonstration activities and to use the same spacecraft

mockup. The control modes demonstrated are also expected to be similar

except for the incorporation of lessons learned from the cargo-bay

7-1
demonstration. It is recommended that the cargo-bay demonstration

servicer system hardware be modified and adapted to become the 1-g test

equipment. Not only will this provide a newer set of equipment but it

should be easier to make the cargo-bay equipment functionally similar

to the free-flight equipment. Functional similarity is important for

procedures development, operator training, and anomalies investigation.

The prior free-flight verification plan emphasized the operational use

of the servicer system as a kit on the Orbital Maneuvering Vehicle.

The emphasis within NASA on the Space Station led to consideration of

how the servicer system could be used on the Space Station. A number

of potential alternatives were identified and are introduced in Section

3.4. As the Space Station becomes better defined it may be appropriate

to include Space Station requirements in the free-flight verification

plan.

The approach to be used for the free-flight verification remains in the

formulative stages. As a result, the distinction between the issues to

be considered and the baseline plan are not as clearly defined as for

the ground or cargo-bay demonstrations. We have attempted to identify

the major elements needed to meet the objectives of the free-flight

verification. We feel the verification flight will include the same

basic servicing functions demonstrated in the cargo bay, adding the

features of operations with the OMV and control from a ground station.

The basic approach to the free-flight verification tests is based on

the desire to have a fully operational on-orbit servicer system at the

end of the program. This means that the servicer must go through the

full design and development process including obtaining production

tooling, which will be available for future units.

For the flight, the servicer can be mated with the OMV on the ground

and transported to space in the Orbiter or the servicer and OMV can be

mated in orbit near the Orbiter. A spacecraft mockup similar to the

one used on the cargo-bay demonstration is mounted on a rented

7-2
spacecraft bus and carried into space on the same Orbiter. The

spacecraft and the servicer with OMV are deployed separately from the

Orbiter. Both are controlled from ground stations. The OMV with

servicer would rendezvous with the spacecraft and dock with it, using

the grapple fixture of the MMS mockup. The servicer would exchange MMS

and battery modules and perform fluid resupply operations. This

activity would be controlled from the servicer ground station, which

would likely be collocated with the OMV ground station. The

servicer/OMV returns the spacecraft to the Orbiter. The RMS is used to

return both vehicles to the Orbiter bay where they are stowed for the

return to earth.

7.1 FREE-FLIGHT VERIFICATION REQUIREMENTS

The basic objectives of the free-flight verification process are to:

1) Establish a servicer system operational capability;

2) Increase confidence of potential users in the servicer concept;

3) Demonstrate compatibility of the OMV and servicer system;

4) Demonstrate rendezvous and docking techniques;

5) Demonstrate free-flight module exchange and fluid resupply.

These verification tests should also increase confidence that the

servicer can be used at the Orbiter, at or near the Space Station, in

other low earth orbits, and in geosynchronous orbit.

The free-flight verification tests are considered to be the final proof

that establishes an orbital servicing capability. Thus the design,

\

development, and test process must be suitable for operational

equipment. Similarly, all the appropriate documentation must be

prepared so that the capability can be used by others. It is

7-3
recommended that one prototype unit be built with production tooling so

additional units can be built to meet the operational requirements. A

preliminary list of the verification flight activities is presented in

Table 7.1-1. As the program matures a better definition of these

activities will be developed.

Table 7.1-1 Free-Flight Verification Activities

Demonstration of MMS Module Exchange

Demonstration of Other Module Exchange Activities

Demonstration of Refueling

Demonstration of Rendezvous and Docking

Communication Links

Control Station Location

Deployment of Servicer Docking Probe

Servicer Mechanism Performance

Interface Mechanism Performance

Connector Performance Including Mate And Demate—Electrical and Fluids

Methods of Accommodating Attach Errors (Compliance)

End Effector Capture

Interface Mechanism Capability for Capture, Latch, Unlatch, and Release

Repeatability Accuracy (Electro/Mechanical)

Spacecraft to Servicer Alignment

Control System Modes Validation

Man Machine Interaction

Malfunction Mode/Backup Systems

Mission/STS System Safety

Pre- and Post-Module Exchange Condition Analysis

The list of Table 7.1-1 was prepared from a similar list in the prior

study and from extended lists prepared during the IOSS activity. The

demonstration of rendezvous and docking is a necessary activity for

remote servicing of spacecraft. The OMV should have demonstrated

rendezvous and docking, as part of its spacecraft retrieval function,

several times before the servicer free-flight verification. However,

in this case it will be necessary to either use the servicer TV camera

or to arrange things so the OMV TV camera can properly see its target.

One potential solution is to mount the stowage rack on the OMV so that

the OMV TV camera, and other ranging equipment, looks through the

region reserved for temporary module storage as this region will be

open at the time of rendezvous and docking and is fairly large.

7-4
The OMV communication links will similarly have been checked out during

early OMV flights. It will only be necessary to extend the links

slightly on both ends with hard wires. It appears desireable to

collocate the servicer control station with the OMV control station as

they both use the same communication links.

It is desireable to include a servicer system docking probe deployment

device in the design. The docking probe extends 60 in. in front of the

servicer stowage rack in the operational configuration. An IOSS design

of a deployment mechanism reduced this extension to 28 in., which

implies a reduction in launch costs. Additionally, it is easier to

support the servicer mechanism so it can absorb Orbiter launch and

reentry loads when it is folded back as opposed to being deployed.

It is recommended that the cargo-bay demonstrations be limited to

operation in one form of Supervisory control and the Manual-Direct

control mode so that astronaut training and operating time could be

limited. For the free-flight verification, where control is from the

ground and the astronauts need not be involved, there is no strong need

to limit control mode selection. Thus, it is recommended that all four

control modes should be exercised and perhaps multiple module exchanges

in each control mode should be performed.

An example flight plan was prepared in the prior study. A slightly

modified form of that plan is presented here for completeness. The

example flight plan is considered to be representative and alternative

plans, with different initial assumptions, can also be prepared and

evaluated. One of the precursors to a free-flight verification is the

need to demonstrate mating of the servicer stowage rack and the OMV

while in orbit. This demonstration has been suggested as part of the

Space Station technology development missions (TDM). It is also

assumed that the OMV has progressed through its development program to

where an OMV is available and can be launched with adequate propellant

for the free-flight verification mission onboard.

7-5
The serviceable spacecraft is assumed to be a special spacecraft for

the verification. It might be that there is a failed spacecraft

requiring servicing of the kinds to be demonstrated when it is needed,

but it is very unlikely. So the plan is to obtain a special

serviceable spacecraft. In addition to the full size operable modules

to be exchanged, the spacecraft would require an attitude control

system, two-way communications to the ground through the TDRSS, a

docking receptacle, a translational thrust capability to put it on a

drift orbit with respect to the Orbiter, and the usual structure, power

supplies, and thermal control. This plan shows the serviceable

spacecraft being returned to earth (to avoid more space debris), but it

may be possible to use it for some other mission after servicing has

been verified. An. alternative to a special design and build of a

serviceable spacecraft is to use the Shuttle Pallet Satellite (SPAS-01)

built by Messerschmitt-Boelkow-Blohm (MBB). The SPAS-01 would have to

be reconfigured to this special use including the addition of the MMS

triangular support structure, modules, and propellant tankage and

handling system from the cargo-bay demonstrations. Additionally its

communications system must be upgraded to work with the TDRSS.

However, the SPAS-01 is an interesting alternative that should be

considered.

The flight plan starts out with the servicer, with replacement modules,

the OMV, and the serviceable spacecraft being launched in a single

Orbiter. The servicer would be fastened to the OMV before launch and

would be returned to earth with the OMV. At the appropriate time in

the mission, the servicing activities would be started. All

activities, other than those involving the Orbiter Remote Manipulator

System (RMS), are controlled from the ground. Because of the close

relationship between the servicer and the OMV, ground control of the

servicer could be from the OMV ground control station (GCS).

At the appropriate time, the spacecraft attitude control system must be

shut off so that it does not fight the OMV attitude control system.

Two types of servicing are recommended - module exchange and fluid

transfer. The servicing functions are the same as are planned for the

7-6
cargo-bay demonstration. Alternatives to the OMV for boosting the

spacecraft back towards the Orbiter would be to use the spacecraft

attitude control system thrusters to initiate the transfer or for the

OMV and spacecraft to return to near the Orbiter in the docked

configuration.

The OMV is put into a quiescent mode when it is near the Orbiter. The

Orbiter will then do whatever maneuvering is necessary for the RMS to

be able to reach out and retrieve the OMV and servicer. When both the

servicer and spacecraft have been stowed in the Orbiter, the Orbiter

crew can continue with their other mission tasks or Initiate reentry

and landing.

The mission duration time was estimated at 21 hrs. This time period

can be modified extensively depending on the desired separation between

the Orbiter and spacecraft and on whether time or propellant is used to

achieve and remove the separation distance. Of the total mission time

of 21 hrs, the Orbiter crew need only be involved for eight hrs. The

ground operations crew will need to be involved for the total mission

time. The module exchange and propellant transfer time allocations

were four hours. This time would be increased if it is desired to

verify all control modes repeatedly as is recommended. However, the

module exchange verification time would still be a small part of the

total time because the orbit transfer time is expected to be long.

As with the cargo-bay demonstration, a number of support systems will

be involved. They include .the Space Transportation System to get the

verification equipment to and from orbit. The OMV will be expected to

provide: 1) communications; 2) attitude control; 3) rendezvous and

docking; 4) electrical power; 5) assistance in thermal control; and 6)

servicer control station location. The RMS will be used to deploy and

retrieve the OMV, servicer system, and serviceable spacecraft. The

TDRSS will be used for communications by both the OMV and the

serviceable spacecraft.

7-7
It is anticipated that the OMV to servicer mechanical interface will be

made fairly rigid to withstand the OMV large engine thrusting and to

provide an adequate structural frequency. The alignment and

flexibility effects of the servicer to spacecraft docking system will

affect the module exchanges and must be properly accounted for in the

design.

Other requirements appropriate for an operational system will also be

applied. Example subjects include:

1) Safety;

2) Operability;

3) Reliability;

4) Maintainability.

Table 7.1-2 presents the major characteristics of the verification

flight. The servicer"s spare module stowage rack will carry the

replacement modules, tools, and stored fluids necessary to service the

spacecraft mockup.

Several alternatives have been proposed that demonstrate extended

capabilities of the servicer. For example, the servicer and OMV could

be carried to space separately and then mated in free space. This

would probably be more representative of an actual mission or

operations from the Space Station. These types of options will be

evaluated as more requirements are defined and plan funding becomes

more certain.

7-8
Table 7.1-2 Free-Flight Verification Characteristics

One verification flight

MMS triangular structure for spacecraft mockup

Spacecraft mockup attached to rented spacecraft bus <

Axial MMS module exchange

Radial battery module exchange

Fluid transfer to PM-1 type propulsion module

Servicer's supply of power, attitude control, thermal control, and

communications by OMV

Spacecraft mockup's supply of power, attitude control, thermal

control, and communications by spacecraft bus

Servicer control station on ground

Unassisted Supervisory control mode

Docking rigidization by servicer docking probe

Electrical connection between servicer and spacecraft via the

docking mechanism

Servicing equipment performance demonstration

OMV-servicer interactions included

Man-machine interactions included

Compliance with Orbiter system safety requirements

Use of representative servicing, operational equipment

Operator training (ground personnel)

7.2 FREE-FLIGHT VERIFICATION ELEMENTS

The elements involved in the verification flight can be separated into

the following groups:

1) Space Transportation System Support;

i

2) OMV with Servicer System;

3) Serviceable Spacecraft;

4) Airborne Support Equipment;

5) Ground System.

7-9
The two major pieces are the servicing system/OMV and the serviceable

spacecraft. These are shown in Figure 7.2-1. Each of them requires a

ground-based control station, interface equipment with the Orbiter, and

ground checkout equipment. It is recommended that the servicer system

be built to operational equipment standards and that planning include

the delivery of production tooling. The prototype approach will be

used for the servicer. This is to say, that the qualification testing

will be performed on the actual flight unit, rather than on a

duplicate. The servicer support equipment should also be designed to

operational standards for repeated use. The serviceable spacecraft and

its support equipment could be designed on a one-time use basis and

might even incorporate equipment from other programs into its design or

the SPAS-01 spacecraft built by MBB might be used. Items such as the

MMS modules and structure, the fluid resupply equipment, and the

Orbiter interfaces from the cargo-bay demonstrations will be used. The

cost estimates are based on rental.of a spacecraft bus and an allowance

for refurbishment of the rented spacecraft.

Certain "existing" equipment will also be needed. The Orbiter and

other parts of the STS, such as the TDRSS, will be used, as will

certain OMV related equipment including, the OMV, its ground control

station, and its docking and rigidization equipment. It is also

required that the OMV be on orbit and have sufficient propellant

onboard to perform the free-flight verification. It is assumed that a

servicer 1-g trainer is available from the servicer cargo-bay

demonstration program.

7-10
°*K*NAL PAGE fS

POOR QUALOY

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7-11
The free-flight verification equipment that must be procured

specifically for the free-flight verification project is listed in

Table 7.2-1. This list was used in preparation of the cost estimates,

Table 7.2-1 Free-Flight Verification Equipment

Integrated Orbital Servicing System

Servicer Control System in OMV Ground Control Station

Replacement Modules

Propellant Resupply Equipment

Servicer to Orbiter Interface Equipment

Servicer Ground Checkout Equipment

Serviceable Spacecraft

Spacecraft Control System in OMV Ground Control Station

Spacecraft to Orbiter Interface Equipment

Spacecraft Ground Checkout Equipment

7.2.1 Space Transportation System Support

Certain of the equipment required for the servicer free-flight

verification is auxiliary equipment available for use on or with the

Orbiter as part of the Space Transportation System. This equipment is

listed in Table 7.2-2. Its provision, control, and use should present

no difficulties.

Table 7.2-2 Space Transportation System Equipment

Space Shuttle including Orbiter

Remote Manipulator System

Cargo-bay cameras

RMS cameras

Control equipment at the Orbiter will be for control of the Orbiter

during proximity operations, control of the RMS during equipment

deployment and retrieval, operation of Orbiter trunnions, and safety

related caution and warning displays. Module exchange operations will

be conducted away from the Orbiter and thus it need not be involved in

these operations.

7-12
7.2.2 OMV With Servicer System

ORIGINAL PAGE FS

OF POOR QUALfTY

An artists concept of the servicer system mounted on the OMV just prior

to docking with a spacecraft is shown in Figure 7.2-2. The

verification flight would involve a different module complement in the

spare module stowage rack. A representative module complement, to

which a fluid transfer module must be added, is shown in Figure 6.3-2.

Similarly, the spacecraft for the verification flight would look more

like the sketch in the upper right hand corner of Figure 7.2-1.

Figure 7.2-2 The Operational Servicer with the OMV

The OMV is assumed to be far enough along in its development that all

of its flight tests have been completed and it is fully operational.

This means that it has been used to deploy and retrieve spacecraft.

All of its interfaces with the Orbiter and its elements such as the RMS

will have been fully verified as will the ability of the TDRSS to

7-13
provide satisfactory communications. Demonstration of spacecraft

retrieval by the OMV implies that operation of its rendezvous and

docking system has also been verified. Similarly, the OMV payload

rigidization system will also have been demonstrated. It will be

necessary to conduct a precursor test on the servicer to OMV

interfaces. This test can be conducted on the ground.

The on-orbit servicer will be more like that identified during the IOSS

program (see Figure 2-1), than the version shown for the cargo-bay

demonstration. In particular the longer (79 in.) arm segment length

should be used. The servicer docking post deployment system should

also be included because of its value in reducing stowed length and

thus launch costs for the operational system. The stowage rack

configuration should generally be like that identified during the IOSS

program except that it should contain the same module complement as for

the cargo-bay demonstrations.

7.2.3 Serviceable Spacecraft

A representation of a candidate serviceable spacecraft is shown in the

upper right hand corner of Figure 7.2-1. The spacecraft bus shown is

the SPAS-01, although other candidate spacecraft should be considered.

As this flight is a one-time use, it appears reasonable to rent a

spacecraft rather than to design and build a spacecraft for this one

use. The requirements on the spacecraft as identified in Section 7.1

are not unusual or stringent and thus it may not be difficult to

identify candidates. One possibility is to refurbish and use the

spacecraft, or parts of it, that was retrieved by the OMV during the

OMV flight test program. The spacecraft will be needed for 18 to 24

months to allow for rework, checkout, flight preparation, and

post-flight examinations and rework.

The spacecraft bus will be modified by adding the spacecraft mockup

from the cargo-bay tests. An MMS module support structure (see Figure

6.2-5) forms the basis of the serviceable part of the spacecraft. The

7-14
module complement and arrangement are shown in Figure 6.2-6 and

discussed in Section 6.2.3. Few changes in the module equipment used

for the cargo-bay test are expected to be required for the verification

flight. While only one electrical grapple fitting is required on the

serviceable spacecraft, it may be advantageous to leave the two grapple

fixtures on the spacecraft mockup from the cargo-bay demonstrations.

7.2.4 Airborne Support Equipment

Airborne support equipment for the free-flight verification is not

extensive. The primary items are devices to support the OMV and

serviceable spacecraft in the Orbiter along with electrical interfaces

and caution and warning equipment. It is expected that the servicer

spare module stowage rack will be fitted with trunnion adapters and a

keel fitting so that the servicer system can be mechanically installed

in standard Orbiter cargo-bay fittings. The electrical interface and

caution and warning provisions will need to be developed.

The OMV will be made compatible with the Orbiter as part of its

development process. Similarly, one criteria for selecting a

spacecraft bus is that it have a set of developed interfaces with the

Orbiter. Thus, development of airborne support equipment for the

free-flight verification should not be a major task.

7.2.5 Ground System

The ground system consists of ground support, ground monitoring, and

ground control. These functions will be available for the OMV and

their existence should be a criteria in selection of a serviceable

spacecraft bus. For the servicer system, the ground support consists

of the usual development, test, and handling equipment for the flight

articles as well as use of the 1-g servicer system demonstration

facility at MSFC. Development, test, and handling equipment for GSFC

supplied flight equipment can be provided by GSFC. The remainder of

7-15
this equipment will be new. It is expected that the MSFC 1-g servicer

demonstration facility will be set up to represent the flight situation

and that it will be used for procedures development, flight crew

training, and anomaly investigation.

The servicing system ground monitoring equipment will consist of

consoles and display equipment located near the control station. New

equipment will be required along with software development to be able

to generate appropriate displays. Data will consist of joint angles,

trajectory coordinates, TV views, and discrete events. The ground

monitoring personnel will follow the planned scenarios and check that

all planned events occur in the proper order and that no unplanned

events occur. Ground monitoring personnel will not be able to control

the servicer demonstration directly. Rather they will be able to

Inform the ground operator of any pertinent observations.

The servicer control station should be located on the ground and near

the OMV control station as all servicer communications pass through the

OMV links and some interfacing between OMV and servicer control

operators is likely to be required. The control station will include a

computer, a keyboard and display unit, and a television monitor for

display of the servicer end effector camera picture. Interconnecting

cables and interfaces to the OMV communication links will be used. It

is expected that 10 to 12,000 lines of code will be required in the

computer. Depending on the computer selected, 256 K bytes of internal

memory and 5 M bytes of hard disk memory will be required. These

computer requirements may be increased when redundancy, failure, fault

isolation, and trajectory generation requirements are considered.

All of the ground system equipment should be designed, fabricated,

tested, and controlled as operational equipment that will be used for

many years and upgraded as needs evolve. It is important that the

existing ETU be replaced by a more representative unit, possibly the

cargo-bay demonstration unit, for use as the operational 1-g training

system.

7-16
7.3 FREE-FLIGHT VERIFICATION SCHEDULE

The development of the operational servicer is to be coordinated with

the development of the OMV. Being an "OMV kit", it is desirable to

have the servicer flight qualified at the start of OMV operations. A

schedule showing the development of the servicer and the major OMV

milestones is shown on Figure 7.3-1. The key OMV milestones are the

first flight occurring in early 1990 and the start of normal operations

in mid 1992.

DESCRIPTION

OMV Design &

Development

OMV Operations

Cargo-Bay

Demonstration

Definition Study

Design & Development

Software Dev & C/O

Tooling & GSE

System Test Equipment

Control Station

Flight Uni & ASE

System Qualification

System Integration &

Verification

1988

\

1989

CIR

r-7

V

PHASE C/0

DEMONSTRATION

7 I FLIfiHT

I

PHASE B *i

1990

FIRST FU

V|

1991

EOT

2ND FLIGHT

i— »

rp POR

V r7

f

DEVELOPMENT

I

I

I

I

COR C

i—r r

V \

CHECKOUT

1

1

L

1992

3RD FLIGHT

T—p

V

:R FRR

? — ^

J

3

"3

i

VERIFICATION

FLIGHT ,_

! v

Figure 7.3-1 Free-Flight Verification Program Schedule

The first two lines of the figure show the OMV design and development

schedule and the OMV operations schedule for reference. A line

indicating the cargo-bay demonstration schedule is also shown. The

7-17
servlcer development Is coordinated with the OMV schedule to provide a

qualified "kit" for when the OMV becomes fully operational.

The phase B study for the servicer has been shortened to 9 months due

to the extensive amount of definition work that has been done as part

of earlier contracts. A three year period has been allocated for the C

and D phases, which is a representative timespan for operational

equipment of this complexity. The development schedules of the

servicer and the serviceable spacecraft have been combined. It is felt

that the definition period of the spacecraft may be somewhat longer

than that of the servicer as little work has been done in this area.

The design and development of the servicer will take longer as the

spacecraft will be used only once and will probably consist of a

combination of existing equipment. The development of a servicer

ground control station has been included. Due to the "kit"

relationship with the OMV, the station will likely be collocated with

the OMV ground control station, and will utilize much of its support

equipment.

The servicer system will be transported to space in the cargo bay of

the Orbiter, making it necessary to meet the safety standards of the

Shuttle. All four safety review phases have been included in the

plan. Additional reviews may be necessary after the servicer becomes

operational, depending on the type of operations to be performed.

The schedule includes time for the fabrication of production tooling,

although only a single unit is to be built as part of this effort.

This will permit the future production of servicer units in a timely

and economical manner.

The verification flight date of late 1992 corresponds to the

operational time frame of the OMV. The schedule for the servicer could

be accelerated if there was a need to do so, allowing for an earlier

flight date. The servicer schedule not only complements the

development of the OMV, but ties nicely to the schedule of the Space

Station.

7-18
7.4 FREE-FLIGHT VERIFICATION COST ESTIMATE

A preliminary cost estimate was prepared for the free-flight servicer

system verification. The costing was based on estimated weights of the

equipment for the free-flight servicer verification. The cost estimate

was developed using cost estimating relationships (CER) contained in

the Martin Marietta Aerospace Cost Data Base and in several NASA

pricing models.

In estimating the costs of the servicer free-flight verification, the

following assumptions were made:

1) All costs are in 1985 dollars;

2) Costs include the design, development, fabrication, testing, and

checkout of the verification hardware;

3) The costs for the Shuttle flight, the OMV, the standard RMS, the

RMS grapple fixture and the crew time are not included in the cost

estimate;

4) The costs associated with crew training are not included;

5) The following items will be provided by Goddard Space Flight Center

and are not included in the estimate:

a) Multi-Mission Modular Spacecraft (MMS) support structure;

b) MMS checkout, handling and transportation equipment;

c) Two MMS modules and three interfaces;

d) Modified Module Servicing Tool;

NOTE: The cost of integrating the Item 5) equipment into the

system is included in the estimate.

7-19
6) The design and build of one servicer for the free-flight

verification will use the traceability, configuration control and

qualification requirements of fully operational equipment. The

proto-flight approach will be used for the servicer;

7) Leasing of the SPAS-01 satellite, or its equivalent, and its

modification for the module exchange and fluid resupply functions

as the serviceable satellite was assumed to be possible for minimum

cost;

8) Replaceable modules and the fluid resupply module will be reused

and reworked as necessary from the cargo-bay servicer demonstration

tests;

9) The Installation and removal of the experiment hardware into/from

the Orbiter cargo bay will be performed by NASA personnel. The

cost of the NASA portion is not included in the estimate.

A breakdown of the cost estimate is shown in Table 7.4-1. The total

estimated cost of the engineering effort and building the servicer

system for free-flight verification will be approximately $35 million.

Table 7.4-1 Free-Flight Verification Cost Breakdown

Flight Equipment

Servicer Mechanism

Airborne Support Equip

Servicer I/F Equip

Stowage Rack

Stowage Rack I/F

Spacecraft Bus Rental

Mockup Equipment

Docking Probe

Subtotal

Program Total

8.0

8.2

3.0

4.0

0.5

2.2

0.8

4.5

31.2

35.0

Ground Equipment

Servicer Checkout

Stowage Rack Checkout

Spacecraft Checkout

Ground Control Station

Personnel

Subtotal

0.2

0.1

0.1

3.3

0.1

3.8

7-20
The estimated cost of 35 million dollars is the same as was previously

estimated. The cost of the servicer unit was reduced by using the

proto-flight approach. This approach permits a single unit to be used

for qualification testing and operations. The tests are performed on

the flight unit in such a manner as to verify the capability, but not

damage the unit. The savings were offset by increased costs due to

inflation and the need to incorporate several items that were

previously included in the cargo-bay demonstrations.

7-21
8.0 SERVICER/MMS 1-g DEMONSTRATION

During this phase of the study, the elements of the Servicer/MMS 1-g

Demonstration Plan presented in the Sections 4.0 and 5.0 were expanded

to provide the basis for the design and fabrication of the MMS

demonstration hardware.

The servicer configuration for 1-g demonstration was reevaluated to

include MMS module changeout. The requirements of the MMS servicing

ground demonstrations were analyzed and a preliminary system concept

design was performed to establish the relative position of the main

components. The configuration of the MMS module mockup and of the

modified MST were selected and the e.g. locations were calculated.

Unbalanced moment loads on the ETU drive motors for MMS ground

servicing demonstrations were analyzed and were found to be within the

existing mechanism capability. The servicing times corresponding to

various arrangements of the modules on the stowage rack were analyzed.

The MMS module locations were selected to minimize the servicing time.

Schedule specifics and cost considerations for the basic contract and

Change Orders 1 and 3 presented at the Midterm Review were included in

this Section, along with a description of the Servicer/MMS 1-g

Demonstration Equipment Drawings and of the equipment fabrication and

tests, as part of the Change Order 3 activity.

8.1 SERVICER/MMS 1-g DEMONSTRATION PLAN

This activity was a continuation of the study completed under the prior

contract, NAS8-35496, and expanded and updated the Servicer Development

Program Plan to include high fidelity ground demonstrations of a

servicer system. As part of Change Order 1 of the subject contract,

the servicer system/MMS 1-g demonstration was further defined as

described in this section. More detailed requirements, at the

subsystem level, were identified. An engineering analysis was

performed to verify that the loads on the drives of the ETU, during the

demonstrations of MMS module exchange, are within the existing drive

8-1
capability. Trade studies were conducted to select the overall

configuration of the servlcer system and a preliminary concept design

of specific demonstration equipment was performed. A detailed schedule

for the servicer/MMS 1-g demonstration equipment design and fabrication

was prepared along with a cost estimate.

8.1.1 Subsystem Requirements

The servicer/MMS 1-g demonstration system requirements presented in

Sections 3.1 and 5.1 were reviewed and subsystem requirements were

identified for the MMS module mockup, spacecraft mockup, stowage rack

mockup, electrical connector positioner mechanism, and optical

targets. The specific requirements developed for the Module Servicing

Tool (MST) adaptation for 1-g demonstrations are included in Section

3.1.

The selected configuration of the MMS 1-g demonstration servicer will

be the alternative that satisfies more of these requirements and to a

higher degree.

The following requirements apply to the MMS module mockup:

1) The MMS module mockup should have the same size, shape and physical

appearance as the actual module;

2) The module retention system shall have the same interface with the

support structure as the actual hardware except that lower torque

levels and different materials are allowable for the ground

demonstration in order to reduce weight;

3) The weight of the MMS module mockup shall be kept to a minimum;

4) The interface of the module mockup with the modified MST shall be

the same as for the actual hardware except that different materials

may be used for the ground demonstration in order to reduce weight;

8-2
5) The MMS module mockup shall be designed to withstand, with proper

margins of safety, all the loads generated during all phases of

servicing demonstrations. For the module mockups used in flight

servicing demonstrations, additional verification for the launch,

deployment and return loads is required;

6) The module mockup shall have provisions for operating "ready to

latch", "latched" and "unlatched" sensors for both attachment

fasteners;

7) The mockup of the module electrical connector shall have the same

interface with the spacecraft support structure mockup as the

actual hardware, including the misalignment capability, except

that, for the ground demonstration, different materials and a

smaller, rack and panel type connector may be used in order to

reduce the mate/demate forces and the weight. The number and size

of the wires across this interface shall be minimized;

8) The module mockup should be easily repairable in case of accidental

damage, using common tools and readily available materials;

9) The module mockup should withstand 400 complete cycles of ground

demonstrations without refurbishment.

The following requirements apply to the spacecraft mockup for the

ground MMS servicing demonstration:

1) The spacecraft mockup shall be provided with one MMS module

location in addition to the existing axial and radial module

locations;

2) The position of the MMS module mockup on the spacecraft mockup,

shall represent the expected relative position of the MMS module

with respect to the servicer when using lateral docking and axial

MMS module exchange;

8-3
3) The interface of the support structure with the module mockup shall

be the same as the actual hardware except that different materials

and a smaller electrical connector may be used;

4) The support structure interface shall include proper sensors to

indicate the "ready to latch", "latched" and "unlatched" conditions

for each of the two fasteners;

5) Provisions for later expansion of the spacecraft mockup to

incorporate fluid resupply demonstration hardware shall be made;

6) Adequate stiffness of the spacecraft mockup attachment to the

docking probe mockup shall be provided in order to maintain the

accuracy of positioning of the servicer arm and modified MST (with

and without module) within the capture envelope of the module

retention system or of the module latch interface;

7) Optical targets shall be provided on the spacecraft mockup, one for

each fastener (and fluid resupply interface, if applicable). The

targets shall be located so that they are clearly visible on the TV

camera monitor, during module/tool attachment/detachment;

8) The spacecraft mockup shall withstand 400 complete cycles of MMS

module exchange demonstrations without refurbishment.

The following requirements apply to the stowage rack mockup for the

ground demonstration of MMS servicing:

1) The stowage rack mockup shall accommodate two functional attachment

locations for MMS modules (a "temporary" and a "good" module

location), in addition to attachment locations for two 24 in. cube

modules and a modified MST storage rack;

2) The interface with the MMS module mockups shall be the same as for

the operational hardware except that lower torques, different

materials and a smaller electrical connector may be used. An

8-4
electrical connector mockup, incorporating a small electrical

connector, to verify proper mating, shall be provided at the "good"

module location only;

3) The support structure interface shall include sensors to indicate

"ready to latch", "latched" and "unlatched" conditions for each of

the two fasteners, at each module location;

4) Modifications to the current configuration of the ETU stowage rack

mockup shall be minimized;

5) A storage location for the modified MST shall be provided on the

stowage rack mockup having the same interface as the MMS module.

The modified MST shall be secured in this location using its

latches. A sensor for the presence of the modified MST in the

storage location shall be provided;

6) The location of the modified MST on the stowage rack mockup shall

be such that no interference with the arm operation occurs during

demonstrations and good visibility is assured for the operator and

other persons watching the MMS module exchange demonstrations;

7) The location of the MMS module mockups on the stowage rack shall

provide good visibility for the persons watching the servicing

demonstrations and a minimum total MMS servicing time;

8) Optical targets shall be provided on the stowage rack, one for each

MMS fastener and for the modified MST storage location, positioned

so that they are clearly visible on the TV camera monitor during

attachment/detachment;

9) The stowage rack mockup shall withstand 400 complete cycles of MMS

module exchange demonstrations without refurbishment.

The following requirements apply to the electrical connector positioner

mechanism to be installed on the existing ETU end effector:

8-5
1) The connector positioner shall provide proper alignment and

translation stroke for mating a subminiature, double density, type

"D" connector with 19 pins, for up to 22 AWG leads;

2) The stroke of the mechanism shall be 5/8 in.;

3) The insertion force capability of the mechanism shall be more than

20 Ibs;

4) The power supply for the mechanism actuator shall be 24 + 2 Vdc;

5) The total time of translation for the mating or demating stroke

shall be less than 6 sec;

6) The net weight of the mechanism (without cabling) shall be less

than 0.5 Ibs;

7) The connector positioner mechanism shall be as compact as

practical. Adequate clearance with respect to all servicer system

elements shall be provided. In the retracted position, the

mechanism shall allow at least +5° of angular misalignment of the

end effector with respect to its mating interface, prior to jaw

closing;

8) The mechanism or its wiring shall not obstruct the field of view of

the TV camera attached to the servicer end effector;

9) The connector positioner mechanism shall be attached to the

existing ETU end effector, opposite to the existing power takeoff

(interface mechanism drive). Its position relative to the ETU end

effector shall be in accordance with the Servicer End

Effector/Modified MST Interface Drawing shown in Figure 8-1. The

existing "ready-to-latch" sensor of the ETU end effector shall be

relocated to the position shown in the figure;

8-6
Figure 8-1

8-7
10) The connector positioner mechanism shall be provided with sensors

to indicate the completion of the stroke in each direction;

11) Mechanical stops shall be provided for the end of stroke in each

direction, capable of stalling the electrical actuator and

preventing mechanical overload of the mating connectors;

12) The "connected" end of stroke of the mechanism shall be adjustable

within 1/8 in.;

13) The connector positioner mechanical interface with the existing

ETU end effector shall be kept as simple as possible for ease of

installation;

14) The connector positioner shall be a self contained unit, capable

of being fully adjusted and tested prior to its installation on

the ETU;

15) The adjustment points for the end of stroke and sensors shall be

accessible before and after the installation of the mechanism on

the ETU. Common tools should be used for making the adjustment;

16) The connector positioner mechanism should withstand 400 complete

cycles of MMS module exchange demonstrations without refurbishment.

The following requirements apply to the optical targets:

1) An optical target shall be provided at each MMS fastener location

on the stowage rack and spacecraft mockups and at the MST storage

location, for verification of end effector alignment, using the

existing TV camera and lights;

2) The optical targets for the MMS servicing demonstrations shall be

of similar design and have patterns similar to the existing

targets for the basic module exchange. The horizontal line shall

be at least 7 in. long to take full advantage of the existing TV

monitor image size, for improved alignment capability;

8-8
, « » « « *.

3) The relative position of the optical target with respect to the end

effector interface and TV camera lens shall be the same for both

basic and MMS module exchange demonstrations;

4) The targets shall have a white background and black lines. The

line width shall be 0.1 in.;

5) The optical targets shall be mounted directly on the module

supports, both on the stowage rack and the spacecraft mockups, to

minimize the need for resetting in case of accidental module

support movement;

6) Adequate clearance between targets and the MMS modules shall be

provided at all times during module exchange. Compliant attachment

of the target plate to its bracket is recommended to minimize

damage to the MMS module mockup in case of accidental interference;

7) The optical target shall be adjustable in its plane within +3/8 in.

in all directions. Positional adjustment in the direction

perpendicular to the target plane is not required. The target

plate shall be perpendicular to the optical axis of the TV camera

within +2°;

8) The design of the optical target for MMS module exchange should be

such that its alignment can be done with the existing template used

for the basic module optical targets.

8.1.2 Torque Loads on ETU Drives for MMS Servicing Demonstrations

An analysis of the unbalanced torques acting on the ETU drive motors

during the ground demonstrations of MMS module exchange was performed.

It was found that all the loads are within the existing ETU

configuration capability.

The existing ETU coordinate system definition, the coordinate

transformation equations and the arm geometry were reviewed. The same

8-9
ORIGINAL PAGE IS

0£ POOR QUALITY

coordinate variables, equations and arm geometry will be used in the

conceptual design of the servicer configuration for ground

demonstrations of MMS module exchange.

The existing configuration of the ETU and the three axes (X, Y and Z)

of its reference system are shown in Figure 8-2. The origin is on the

docking probe centerline, midway between the upper and lower joints of

the parallelogram of the shoulder segment of the arm, which is 33.25

in. above the top of the stowage rack mockup beams. X is along the

Figure 8-2 Existing ETU Configuration and its Reference Coordinates

System

8-10
docking axis, positive towards the spacecraft mockup. Y is along the

middle rib (beam) of the stowage rack mockup and Z completes a right

handed system.

The ETU arm configuration is shown in Figure 8-3. The six joints and

their alphabetical designation, as well as the principal dimensions of

the arm are shown on the figure.

There are six degrees of freedom of the servicer mechanism. These six

can be combined into two groups of three each. One relates to the

position of the end effector, thinking of the end effector as a single

point. The second group relates to the attitude of the end effector,

now thinking of the end effector as a movable hand capable of rotating

about three different independent axes. These movements are

independent of the position of the end effector, which may or may not

change at all while the attitude of.the end effector is changed. The

definitions that follow treat end effector "position" and "attitude"

separately. Also treated separately are the cylindrical coordinates

from the specific joint coordinates of the ETU.

The variables in the cylindrical coordinate system, defining the end

effector position are x, r, and Theta, where:

1) x is along X, +x for +X, same origin as for the XYZ reference

system;

2) r is the radial distance, always positive;

3) Theta, the central angle, is defined in Figure 8-4. The +Y axis is

the zero reference. Theta is + towards +Z.

8-11
s

o

3

00

c

o

a

3

U

fl

00

01

1-

M

Figure 8-3

8-12
Figure 8-4 End Effector Position in Cylindrical Coordinates

Three variables are used to define the attitude of the end effector.

They are the angles Psi, Phi and Omega, which are shown in Figure 8-5.

While Figure. 8-5 shows the axes of rotation of Psl and Omega being

colinear and Indicating a sigularity, the singularity is not of

significance. Phi is +90 or -90 degrees for axial operation and the

singularity is avoided because the Psi and Phi axis are at 90 degrees

to each other. For radial motion Phi is 180 degrees, however, the

singularity can be ignored because of the way in which the Psi and

Omega cylindrical coordinates are used.

An auxiliary reference system, X1, Y' and Z' is considered attached to

the end effector as shown in Figure 8-5. The three rotations, Psi, Phi

and Omega have the desired meaning only when they are taken in sequence

- Psi, then Phi and then Omega. They modify the attitude of the end

effector auxiliary reference system with respect to its previous

intermediate step position as shown in Figure 8-6.

8-13
z

X (out of

paper)

Figure 8-5 Cylindrical Coordinates

X'

Y' STEP 1

FIRST ROTATION

OF * ANGLE

x

1

STEP 2

SECOND ROTATION

OF 4> ANGLE

Y' STEP 3 -X1

THIRD ROTATION

OF (2 ANGLE

Figure 8-6 End Effector Attitude Definition

8-14
The ETU joint coordinates expressed in terms of angles of rotation of

the six drives of the servicer mechanism are also used to define the

arm position.

The position and the attitude of the end effector is physically

changed by actuating the six drives T, U, V, W, Y and Z. The final

result is a function of the reference position, the value of the six

angles of joint rotation, also designated T, U, V, W, Y and Z

respectively, and the dimensions of the arm segments 1. t i

^3, lc and lg. A conventional reference position of the arm

was selected as shown in Figure 8-7 for the purpose of establishing

the origin or zero values for the six angles of rotation of the

joints. All joint angles are zero in the reference position, except

for the elbow "V" and the shoulder roll "T" drive angles, which are

measured as shown. The arm in the reference position is horizontal,

that is parallel to the YZ plane of the servicer reference system.

In the reference position the end effector faces outward as shown in

Figures 8-7 and 8-8. The latter shows also the positive directions

of wrist joint rotation for the W, Y and Z drives.

8-15
ORIGINAL PAGE fS

OF POOR QUALITY

'Z" DRIVE END EFFECTOR

"Y" DRIVE

"W" DRIVE

Figure 8-7 Reference Position of ETU (U=W=Y=Z = 0)

8-16
CENTER OF END EFFECTOR

"Z" DRIVE

"W" DRIVE

FOREARM

Figure 8-8 ETU Wrist Joints in the Reference Position (W=Y=Z=0 )

The length of ETU arm segments are:

1, =

45.0 in.

45.0 in.

12.875 in.

6.625 in.

3.375 in.

For MMS module exchange demonstrations the ETU operates in the axial

mode only, therefore, W = Psi = 0 at all times. Except during the FLIP

or IFLIP steps the end effector centerline is vertical. When it is

facing the stowage rack mockup Y = Phi = +90°. When facing the

spacecraft mockup Y = Phi = -90°.

Once the position of the MMS module mockup on the stowage rack or

spacecraft mockup is selected, the cylindrical coordinates x, r and

Theta of its fastener locations are defined. Also known is the angle

Alpha, between the Y axis and the orientation vector of the fastener

(see Figure 8-9). The equations used to calculate the corresponding

8-17
joint angles T, U, V and Z and the end effector attitude angle Omega

are given below. More detailed explanation is given in the Servicer

Simulation Software Requirements (MMS Modules) document, MCR85-1331,

Martin Marietta Corporation, June 1985.

U - sin^lp-J ; a ->y 122 + i32 „ 46>g l in- . g - tan'1 f —

, fr2 - a2 - (l! cos U + 15)

2

 - 16

2

~|

L 2ab J

wh

T - 9 - a + <$ -V

w - * - o

Y • <fr • ±90°

ere:

b • W (lj cos U + 15)

2

 +

6 - V - sin'1 ^ sin (V -f

li cos U + 15

»).».„-,

16

2

-]

I

Z = $ + Q -a and ft = Z - 6 when facing the stowage rack

Z = 180° -(6+9-0, ) an d n=Z+ 6 when facing the spacecraft

(see Figure 8-10)

8-18
MMS MODULE

MOCKUP

FASTENER

ORIENTATION

VECTOR

TV CAMERA ( OPTICAL TARGET )

POSITION

PARALLEL TO Y AXIS

VIEW LOOKING DOWN

ON STOWAGE RACK

Figure 8-9 Fastener Orientation Angle Alpha

END EFFECTOR FACING STOWAGE RACK

-a

z = 5 + 9 - a

END EFFECTOR FACING SPACECRAFT

Z = 180 - ( S + 9 -

NOTE: BOTH DIAGRAMS SHOW VIEWS OF ETU ARM, LOOKING DOWN ON STOWAGE RACK

Figure 8-10 Relationship between Z, Theta and Alpha Angles

8-19
The arm configuration and the alphabetical designation of its drives

was shown in Figure 8-3. Because the axes of the T and V drives remain

vertical at all times, the unbalanced gravity moments do not load their

drive motors (except for a negligible increase in bearing friction).

The large diameter bearings have sufficient load carrying capability

margin in all load cases for MMS servicing demonstrations.

The Z drive is also vertical at all times during MMS module exchange

demonstrations except during the FLIP or IFLIP steps, when the motor is

not actuated and the brake is applied.

The W drive motor is only actuated to correct for minor arm bending and

fabrication misalignments, as the ETU remains in the axial mode at all

times. W drive motions are less than 1 deg. The worm gear speed

reducer does not back drive so there in no need for a brake. The load

on the W drive motor was checked and found to be within the existing

mechanism capability at all times during the demonstration of the MMS

module exchange.

The remaining drive motors that are subject to torque loads due to

unbalanced gravity moments, during the ground demonstrations of MMS

module exchange, are the wrist pitch Y drive and shoulder pitch

(elevating) U drive motors.

8.1.2.1 Engineering Test Unit Load Limitations - A critical actuator of the

ETU is the wrist pitch (Y) drive. The torque required from this drive

must not exceed 50 ft-lb to avoid overheating and provide adequate

speed. Under a 50 ft-lb torque load the drive will require 15 sec for

a 180° flip, as compared to 5 sec with no load.

A number of sketches were prepared to identify the baseline shown in

Figure 8-11. The MMS module mockup is shown extending horizontally

from the "Y" drive centerline. The unbalanced moment on the servicer

wrist pitch "Y" drive is maximum in this position, occurring during the

module exchange demonstration. The torque value depends on the weight

of the module mockup (WM) and of the modified MST (WT), on the

8-20
END EFFECTOR ; I/F

MODIFIED

"Z" DRIVE-y

1.0"

y—MMS MODULE

7

 MOCKUP

^MODULE

LATCH I/F

Figure 8-11 Y-Drlve Position Relative to Modified MST and MMS Module

8-21
positions of the centers of gravity and on the distance "B" between the

end effector interface and the module latch interface. The e.g. of the

modified MST was assumed to be at a 0.6B distance from the end effector

interface. The position of the module mockup e.g. was calculated based

on a preliminary design. While the end effector and "Z" drive combined

e.g. is approximately on the "Y" drive centerline, the weight of the TV

camera and of the lights contribute to the total unbalanced torque on

the "Y" drive.

Using the weights and distances shown in Figure 8-11 and a 50 ft-lb

Y-drive load capability, calculations were made for the maximum

distance between the end effector and module latch interface (distance

"B" on the figure) as a function of the module mockup and modified MST

weights. The results of the calculations are plotted on the graph of

Figure 8-12.

If the MMS module mockup weighs 20 Ibs and the modified MST weighs only

7.5 Ibs, the maximum distance between the end effector interface and

module attach interface (distance "B" on Figure 8-11) is 3.5 in. This

distance can be 7.25 in. if the module mockup weight is 12.5 Ibs and

the modified MST weighs 15 Ibs. With B - 5 in. and module mockup

weight of 12.5 Ibs, the MST could weigh as much as 20 Ibs. After

discussions with GSFC and Fairchild Space Company, the selected

requirements became:

1) "B" distance 7.25 in.;

2) MST weight 15.0 Ibs maximum;

3) Module weight 12.5 Ibs maximum.

8-22
LLI 10

3 CD

CM

CO

00 =

3

vo o

z

o

Qi

o

o

LU

CVJ LJ_

•a•H

H

CO

2

<u

,—1

n

i-(

*

cs

IT)

(S81) 1H9I3M 1SW Q3IdIQOW HlflVHOIlV

01

M

•H

fc

Figure 8-12

8-23
8.1.2.2 Selection of the FLIP Position - The FLIP is a step in an ETU

trajectory that changes the attitude of the end effector from vertical,

facing the stowage rack to vertical, facing the spacecraft, by rotating

the Y drive from +90° to -90° while all other drives are at rest

with the brakes applied. This provides a simple motion of the module,

which is easy to anticipate, avoiding personnel injury and/or damage to

the module mockup due to Interference. The IFLIF is the inverse step,

from the end effector facing the spacecraft to facing the stowage rack.

Both the FLIP location, in terms of z, r and Theta values and the

module orientation with respect to the arm, in terms of the 1 value

need to be selected. The following requirements apply:

1) The FLIP location and the module orientation must be such that

adequate clearance is provided between all the servicer system

elements;

2) The FLIP location and the module orientation should be such that

the unbalanced gravity moment applied to the "Y" and "U" drive

motors is minimized;

3) Changes in the existing ETU configuration and control software in

order to demonstrate MMS module exchange shall be minimized;

4) The selected FLIP location and module orientation should be such

that the total module exchange time is minimized;

5) The selected FLIP location and module orientation should allow good

viewing by visitors, during demonstrations;

6) The selected FLIP location should allow later expansion of the

spacecraft mockup to add a propulsion module mockup for refueling

demonstrations.

Four module orientations relative to the arm were considered for FLIP.

They are the A, B, C and D orientations shown in Figure 8-13 in the

8-24
starting position for FLIP, when the module is facing the stowage

rack. During the FLIP the module rotates 180 degrees around the "Y"

drive axis, from Y = +90° to Y = -90°. The module must be rotated so

that it passes through Y = 0 position, for the B, C and D orientations

because, otherwise, the forearm is in the way. Passing through Y =

180° is possible for orientation A, however, this way of performing the

FLIP is not recommended because it requires extending the arm to a

larger radius in order to avoid interference with the spacecraft.

Also, the "Y" joint position sensor and limit switches does not allow

the joint to reach the 180° position.

The advantages and disadvantages of each of the four orientations were

analyzed, based on the above requirements in order to select the module

orientation with respect to the arm during the FLIP step. A layout of

the modified ETU for the MMS ground servicing demonstration was

prepared in order to compare the flip alternatives and check for

interferences.

DRIVE

Figure 8-13 Module Orientation Relative to the Arm for FLIP

Starting Position

8-25
The assumed geometry of the MMS module mockup and the modified MST for

the ground demonstrations and the e.g. positions are shown in Figure

8-14. The flip clearance radius of 62.1 in. for the B and D

orientations and the 11.0 in. dimension of the clearance envelope for

the end effector wire harness are shown. The clearance radius "R" for

the A and C orientations was shown on Figure 8-11. For B - 7.25 in.;

R - 41.5 in.

The B and D orientations require extending the arm to a larger radius

for FLIP compared to A and C, in order to avoid interference with the

module mockups in the stowage rack and with the spacecraft mockup.

Extending the arm and the module to a larger radius increases the

unbalanced torque on the "U" drive motor. For the B and D

orientations, in order to provide adequate clearance, the end effector

should be at a radius of at least 82 in., while the A and C

orientations require only a 75 in. radius. For the D orientation,

interference will occur unless the FLIP and IFLIP steps are performed

at different locations. This increases the complexity of the control

software unnecessarily.

The maximum unbalanced moment on the "Y" drive motor occurs at

mid-FLIP, when the combined e.g. of the module mockup and the modified

MST is in the same horizontal plane as the "Y" drive centerline (see

Figure 8-15).

For the B and D orientations the maximum unbalanced torque on the "Y"

drive motor is larger than in the case of the A and C orientations

because the horizontal distance between the combined e.g. and the drive

centerline is larger (22.08 in. for B and D vs. 18.69 in. for A and C).

Although the unbalanced torque on the "Y" drive is the same for both

the A and C orientations, A is better than C because it provides an

unobstructed view of the end effector. Also, the maximum unbalanced

moment on the "U" drive is smaller for the A orientation as compared to

C because the combined e.g. of the module mockup and the modified MST

is closer to the "U" drive centerline, as can be seen in Figure 8-15.

8-26
0.90

48.00

49.800

22.9

REF

24.00

MODULE LATCH I/F

37.79 R

MODIFIED MST

END EFFECTOR I/F

r-"Z" DRIVE

11.77

20.51

MMS MODULE MOCKUP

COMBINED C.G.

W=27.5

Ibs

•-10.4

Figure 8-14 MMS Module Mockup and Modified MST Geometry

8-27
Figure 8-15 MMS Module Mockup in Mid-FLIP Position

In conclusion, the A orientation provides minimum unbalanced torque

loads on both "U" and "Y" drives, results in a better view of the end

effector during demonstrations and there is adequate clearance with

less radial extension of the ETU arm. This orientation of the module

with respect to the arm was selected for the FLIP step of the MMS

servicing ground demonstrations.

The value of the Z angle during the FLIP, using the A orientation, is

-90 degrees.

The selection of the x, r and Theta coordinates for the FLIP was made

considering the existing ETU FLIP location and three alternative

locations.

8-28
The existing FLIP location of the ETU is at x = +20 in., r = 82 in. and

Theta + 45° (location 0 on Figure 8-16). A layout was made for the

servicer system configuration for ground demonstrations of MMS module

exchange and it showed that this location for FLIP is not feasible, due

to interference with the spacecraft mockup.

TO CENTER

OF MODULE

(TOP VIEW)

Figure 8-16 Alternative Locations for FLIP for MMS Servicing

8-29
Three alternative locations for FLIP were considered (see Figure

8-16):

1) Module parallel to MMS spacecraft module location,

x - 4-11.00 in.

r - 63.00 in.

Theta - 8.87°;

2) Module in front of spacecraft radial module location,

x - +13.00 in.

r - 75.00 in.

Theta - -70.00°;

3) Module parallel to Z axis, above middle rib of stowage rack,

x - +14.00 in.

r - 82.00 in.

Theta - -16.67°.

These alternative FLIP locations were analyzed and traded off against

the requirements. Location No. 1 provides too little clearance, only

1 to 2 inches between the module mockup and the spacecraft mockup or

the module in the stowage rack. Location No. 3 offers good

visibility of the FLIP but does not provide adequate room for later

addition of a propulsion module on the spacecraft mockup.

The recommended alternative is the FLIP Location No. 2. It provides

adequate clearance, allows an unobstructed view during demonstrations

and leaves ample room available for expansion of the spacecraft

mockup. During MMS module exchange system checkout, it was necessary

to increase the value of x to 19.00 in. to provide adequate clearance.

8-30
8.1.2.3 Torque Load on "Y" Drive - Maximum torque load occurs at mid-FLIP,

when the combined e.g. of the module mockup and modified MST Is in the

same horizontal plane with the "Y" drive centerline. The A orientation

of the module mockup was assumed as shown in Figure 8-15. A reference

system Xa, Ya, Za, attached to the end effector is also shown in the

figure. Other assumptions were the weight of the module mockup of 12.5

Ib, the weight of the modified MST of 15 Ib and a horizontal distance

of 18.69 in. from the combined e.g. to the centerline of the "Y" drive

(see Figure 8-14).

The weight and the e.g. position of the elements considered in the

calculations are given in Table 8-1 and were shown previously in Figure

8-11.

Table 8-1 Weight Data for "Y" Drive Torque Load Calculation

ELEMENT

End Effector and

"Z" Drive

Module and

Modified MST

TV Camera

Lights

WEIGHT

(Ib)

22.46

27.5

3.0

3.0

Xa

(in.)

0

-18.69

- 3.0

- 3.0

Za

(in.)

0

0

- 1.0

- 4.5

The calculated unbalanced torque applied to the "Y" drive output flange

is 44.35 ft-lb. The dual path internal gearing of the "Y" drive has a

ratio of 103.09:1 and an estimated efficiency of 90%. The

corresponding maximum motor torque load is 0.48 ft-lb. Assuming a

linear variation of the motor torque vs. motor speed (see Figure 8-17),

with a stall torque of the existing motor of 0.85 ft-lb and a no-load

speed of 67 rad/sec, the motor speed under the maximum torque load will

be 29.2 rad/sec. The corresponding speed for the output flange of the

"Y" drive will be 16 deg/sec, which is more than the minimum "Y" rate

of 12 deg/sec established in the Servicer Simulation Software

Requirements (MMS Modules) document.

8-31
T

Ft-Lbs

1.00

0.50 0.48 Ft-Lbs?

12.0 Deg/Sec -» N= 21.59 Rad/Sec (MOTOR)

•MAXIMUM TORQUE LOAD

(MMS SERVICING)

Y =16.0 Deg/Sec

29.2 Rad/Sec

(MOTOR)

10 20 30 40 50 60 70 80 N

Rad/Sec

(MOTOR)

Figure 8-17 Torque Load vs. Speed of the "Y" Drive

In conclusion, under the maximum unbalanced torque load, during the

ground demonstrations of MMS module exchange, the "Y" drive motor will

maintain enough speed to prevent overheating or stalling.

8.1.2.4 Torque Load on "U" Drive (Shoulder Elevation) - A top view of the

Engineering Test Unit servicing mechanism was shown in Figure 8-3. A

counterbalance system, comprised of the elbow and the shoulder

counterbalance, minimizes the load on the shoulder elevation drive in

1-g operation (see Figure 8-18).

Ideally, in a perfectly counterbalanced system, the "U" drive would

have no motor load except for inertia and friction. However, this

condition cannot be achieved mainly because during the arm operation

8-32
ELBOW COUNTERBALANCE

87 Lbs

SHOULDER COUNTERBALANCE

253 Lbs

SHOULDER ELEVATION

 24-°

"U" DRIVE ACTUATOR

Figure 8-18 ETU Counterbalance System

8-33
the module mockup and/or a tool adapter is not attached to the end

effector at all times. The elbow counterbalance weight of 87 Ibs was

set during the ETU design and testing activities to bring the e.g. of

the forearm as close as possible to the center of the elbow "V" drive

for all operational conditions. Twenty-two cases were considered then

to determine two extreme positions of the e.g. relative to the center

of the "V" drive. The counterbalance weight was set so that the center

of the "V drive is midway between these two extreme positions of the

forearm e.g., providing minimum variation of the unbalanced load on the

shoulder "U" drive. When a module is attached to the end effector, the

shoulder elevation drive is slightly unbalanced so the backlash is

always taken up. Without a module, the backlash is taken up in the

opposite direction in a similar way. This feature of the ETU

counterbalance system was used to increase the accuracy and made

possible the use of a less expensive linear actuator for the shoulder

elevation drive.

The twenty-two cases considered in the ETU design were reviewed to

determine the two extreme cases that give maximum unbalanced gravity

moments and motor torques on the "U" drive in the two directions during

the ground demonstrations of MMS module exchange. These two cases are

described below.

CASE 1 This case produces the maximum unbalanced load on the "U" drive

linear actuator in the "down" direction. The arm is in mid-FLIP

position with the modified MST and the MMS module mockup attached to

the end effector using orientation A, as shown in Figure 8-18. Maximum

loading of the "U" drive actuator is obtained by having the arm

horizontal and extended to the maximum radius r = 89.0 in. The loads

and dimensions applicable to Case 1 are shown in Figure 8-19. The

value of the load acting "down" on the "U" drive linear actuator is

FIT * 171.6 Ibs and was determined from the moment equilibrium around

the "U" drive centerline.

8-34
CASE 2 This case produces the maximum unbalanced load on the "U" drive

actuator in the "up" direction. The arm is in the mid-FLIP position as

shown in Figure 8-19 except that no module or modified MST is attached

to the end effector. In order to get the e.g. of the forearm as close

as possible to the "U" drive centerline, the end effector is set at a

maximum x (x = +20 in., U = 26.39°) and the end effector faces the

spacecraft and is rotated so that the e.g. of the TV camera and lights

is closest to the "U" drive (Y = -90°, Z = 160.54°). Other loads and

dimensions applicable to Case 2 are shown in Figure 8-19. The value of

the maximum load acting "up" on the "U" drive linear actuator,

determined from the moment equilibrium around the "U" drive centerline

s = -177.9 Ibs.

FOREARM C.G.

(CASE 2 - Wp=153.96 Lbs.)

DRIVE ACTUATOR

"U" DRIVE

CENTERLINE

FOREARM

=45.00 (CASE 1)

=40.31 (CASE 2)

d=3.98 (CASE 1)

rMAX=89.0 (CAS

r=84.73 (CASE 2)

r=rMAX=89.0 (CASE 1)

MODIFIED MST

SHOULDER

COUNTERBALANCE

253 Lbs.

HMS MODULE

MOCKUP

NOTE: CASE 1 CONFIGURATION SHOWN.

CASE 2 - NO MODULE OR MODIFIED MST

DIMENSIONS IN INCHES

Figure 8-19 "U" Drive Actuator Load - Cases 1 and 2

8-35
The maximum unbalanced force on the "U" drive actuator occurs in Case 2

and is -177.9 Ibs (up). It is practically equal in absolute value to

the maximum force down, of 171.6 Ib. This means that the existing ETU

counterbalance system is almost perfectly suited to the MMS module

exchange demonstrations and no modifications are necessary.

Assuming a 90% efficiency for the "U" drive, the equivalent load seen

by the motor is 197.7 Ibs. The existing "U" drive linear actuator is a

Duff-Norton Model MPD-6405 with a stall force of 500 Ibs at the tip and

a no-load speed of 68 in./min, corresponding to 7.22 deg/sec for the

"U" angle. Assuming linear variation of the speed with the load, as

shown in Figure 8-20, the "U" rate under maximum load will be 4.37

deg/sec, which is larger than the rate limit of 4 deg/sec, established

in the Servicer Simulation Software Requirements document.

4.0 "U" DRIVE RATE LIMIT

MAXIMUM LOAD

(MMS SERVICING)

Deg/Sec

Figure 8-20 Load vs. Speed of the "U" Drive

In conclusion, during the MMS module exchange demonstrations the "U"

drive motor will maintain enough speed, under the maximum load, to

prevent overheating and stalling and all the ETU drives will be loaded

within their capability.

8-36
8.1.3 Concept Definition of the MMS Module Mockup

The MMS module mockup for the ground servicing demonstrations should be

a realistic representation of the actual hardware (see Figures 8-21 and

8-22). It should have the same outside shape and dimensions.

EXTENDER:

MRS BEAM

TWO AXIS

RESTRAINT SOCKET

MODULE FRAME

UPPER FASTENER

CONNECTOR(S)-^

MODULE COVER

(EQUIPMENT

BASEPLATE)

THREE AXIS

RESTRAINT SOCKET

LOWER FASTENER

BACK VIEW

Figure 8-21 MMS Module Structure with Module Retention System

8-37
The module retention system (MRS) hardware, the latch interface, the

multi-layer insulation and the front louvers should be simulated as

closely as possible while their weight should be kept to an absolute

minimum.

THERMAL LOUVERS

48.00

Figure 8-22 MMS Module Outside Dimensions, Including the Insulation

Preliminary calculations have shown that a very lightweight MMS module

mockup weighing less than 15 Ibs is feasible. A modified, lightweight

fastener arrangement was used. Instead of the existing titanium bolt

(see Figure 8-23) an aluminum bolt having the same thread interface but

reduced shank diameter and hollow core was used (see Figure 8-24). The

disc spring arrangement was replaced by a spring washer. A sintered

bronze thrust washer was used under the bolt head to reduce the

tightening torque while providing enough friction to prevent self

loosening under load and vibration.

8-38
SPACECRAFT

FLOATING NUTDISC SPRING

ARRANGEMENT

TITANIUM FASTENER

Figure 8-23 Existing MMS Module Retention Fastener

8-39
co

•HWi01 GeoIoCS <QJ

i—l 3T3T3HJ

•il TJCNI000)oo •HCb

Figure 8-24

8-40
8.1.3.1 Fastener Tightening and Loosening Torques - In order to reduce the

weight of the module mockup and of the modified MST, the torque applied

to the fasteners when tightening and loosening them, and the

corresponding reaction torques applied to the module latch interface

should be as low as possible. However, enough clamping force should be

provided when tightening, to prevent self loosening under vibration.

The fastener tightening and loosening torques needed to be determined

before the module mockup structural concept could be established.

The critical case for the fastener tightening torque is shown in Figure

8-25. The servicer supports the weight of the module mockup before

starting to tighten the first fastener to the spacecraft mockup. The

weight of the overhanging module and the force resisting the mating of

the electrical connector produce slight deflections in the servicing

mechanism so that the first fastener to be tightened gets to be engaged

first with the spacecraft mockup interface. During fastening, a corner

of the module retention hardware contacts the spacecraft and the axial

load in the fastener acts to straighten the module.

A 10 Ib mating force was assumed for the small electrical connector

placed inside of the connector mockup.

10 Ibs

Figure 8-25 Maximum Axial Force in the Fastener

8-41
A maximum axial force In the fastener of 400 Ibs was calculated from

the moment equilibrium including an assumed 70 Ibs clamping force, when

the module is in a horizontal position, being attached to the

spacecraft mockup. A friction coefficient of 0.1 sliding and 0.3

static was assumed for the thread.

For the overhead position the required tightening torque is 6.8

ft.-lb. A tightening torque setting of 10 + 1 ft-lb for the modified

MST was chosen.

The critical case for loosening the fastener is when the module is

removed from the stowage rack. The weight of the module mockup is

supported by the stowage rack and is not helping to reduce the

loosening torque. With the torque setting for tightening established

from the spacecraft case, 650 Ibs clamping force is produced in the

fastener in this case. The corresponding loosening torque is 16.7

ft-lb. A loosening torque setting of 20 + 1 ft-lb was chosen for the

modified MST. This will be the maximum torque reacted by the module

mockup structure between the two fastener attachment locations and

either of the latch interfaces.

8.1.3.2 Module Accidental Load - In addition to the reacted torque from

fastener tightening and loosening and the loads from its own weight the

module should withstand reasonable accidental loads from handling and

improper operation without incurring permanent deflections or extensive

damage. The module structure should resist without damage a load of 25

Ibs applied on any corner and in any direction when an opposite face is

supported (see Figure 8-26). It corresponds to the dynamic load that

occurs when the module is dropped from a height equivalent to the

maximum elevation of the end effector.

8-42
25 1

Figure 8-26 Accidental Load Case

The structure should maintain its original dimensions and remain square

after the accidental load is removed. However, local damage to the

corners or outer surface finish is acceptable to allow absorption of

the energy of impact and protect the rest of the structure. The module

mockup should be easily repairable in case of accidental damage, using

common tools and readily available materials.

8.1.3.3 Module Mockup Structure - Various structural arrangements were

investigated:

1) Truss structure, using small diameter, thin wall aluminum tubing

(see Figure 8-27);

2) Box construction, using lightweight paper honeycomb/fiberglass

panels and formed balsa wood and fiberglass sandwich panels;

3) Torque tube structure connecting the module retention hardware, the

connector and the latch interface and use of foam board for

simulation of the outside surface (see Figure 8-28);

8-43
4) Use of foam board for both central torque box and outer surface,

and balsa wood/fiberglass for module latch interface and connector

mockup (see Figure 8-29).

The truss type module mockup (Variant No. 1) has a welded truss

structure, made of 3/8 in. diameter, thin wall tubing, connecting the

two module latch interfaces, the electrical connector and the two

fasteners. The outer tubes represent the faces and the edges of the

actual module. Fiberglass guide tubes between the latch interfaces and

fasteners are also used as structural members. A removable cover made

of medium weight metallized polyester film is attached to the structure

with Velcro type fasteners. The total weight of the module mockup is

14.2 Ibs compared to 12.5 Ibs for the foam board module (Variant No. 4).

THREE AXIS

RESTRAINT

SOCKET

UPPER FASTENEREXTENDERELECTRICAL CONNECTOR

MOCKUP

 TWO AXIS

RESTRAINT

SOCKET

MODULE

LATCH

INTERFACE

GUSSET

ALUMINUM

TUBING

(3/8"ODx

1/32"W)

LOWER

FASTENER

Figure 8-27 MMS Module Mockup Truss Structure

8-44
The cost of the truss structure module is higher because of the welding

fixtures required and of the difficulty of welding thin wall (1/32 in.)

aluminum tubing. Lower fidelity is obtained because the front louvers

are not simulated. However, it has the advantage of higher damage

resistance than the foam board module and allows the operator and the

observers to see through the module if the cover is removed.

The paper honeycomb/fiberglass module (Variant No. 2) is a box type

construction simulating the outside surface of the actual module and

having an internal torque tube structure. It is similar to Variant No.

3, shown in Figure 8-28, except for the material used for the box. The

honeycomb core is 3/8 in. thick and is made of phenolic impregnated

paper. The inner and outer skin are made of fiberglass 0.006 in.

thick. Large cutouts in the core and in the inner skin are provided in

order to reduce weight. Fiberglass skin and 1/8 in. thick balsa wood

core are used for the module latch interface support structure and for

the torque tube. The total weight is 17.4 Ibs and is almost 5 Ibs

heavier than the foam board type module (Variant No. 4). The paper

core is subject to damage from a high humidity atmosphere. A sample of

paper honeycomb core Impregnated and fully cured failed our 20 minutes

water submersion test. The cost is higher than Variant No. 4 because

of extra work required for wall panel fabrication. The structure is

stlffer but more subject to damage on impact than the foam board type.

It is also difficult to repair.

The torque tube structure, Variant No. 3, (see Figure 8-28) is made of

formed sandwich panels of fiberglass with balsa wood core 1/8 in. thick

and the outer surface of the module is simulated by a foam board

enclosure.

8-45
FOAM BOARD

ENCLOSURE

FIBERGLASS/BALSA WOOD UPPER STRUCTURE

ELECTRICAL

CONNECTOR

4 IN. DIA

TORQUE

TUBE

48.0"

FIBERGLASS/BALSA WOOD

INTERFACE BOX

Figure 8-28 MMS Module Mockup with Torque Box Structure

The material used for the box is 3/4 in. styrofoam board bonded with

urethane adhesive. Preliminary tests showed that this material

provides high resiliency, resistance to water, minimum weight and ease

of repair of localized accidental damage. Metallized polyester film is

bonded to the outside surface to simulate the multi-layer thermal

insulation. Because the foam board enclosure does not carry loads

except for its own weight and inertia, the module is heavier than the

Variant No. 4. It weighs 14.6 Ibs.

Minimum weight is obtained with an integral torque box construction

made of styrofoam board (Variant No. 4, see Figure 8-29), and this is

the recommended structure design for the MMS module mockup.

8-46
CENTRAL

TORQUE

BOX

SECTION CC VIEW B

LATCH INTERFACE BOX

r

r i

SECTION AA

Figure 8-29 MMS Module Mockup with Central Torque Box Structure

8-47
This arrangement gives a better weight to stiffness ratio, the central

box structure providing a direct connection between all attachment

points. The foam board is 3/4 in. thick styrofoam that can be cut and

bonded with ease and minimum fixturing. It deflects on impact without

shattering and accidentally damaged areas can be easily cut out and

rebuilt. A thin film of metallized polyester bonded to the outside

surface simulates the thermal insulation while rigidizing the

structure. Urethane foam is used to reinforce the edge bond lines.

The upper and lower latch Interface boxes are made of balsa

wood/fiberglass sandwich construction. Large cutouts are provided in

the foam board panels to keep down the weight. A weight estimate

showed that the module can be built for a total weight of 12.5 Ibs.

The e.g. is 10.4 in. aft of the module latch interface and 22.9 in.

below the upper fastener centerline as shown on Figure 8-14. The

module mockup can be made to simulate in more detail the outer shape

and size of the actual equipment than is possible in the case of the

truss structure.

A partial mockup, representing one quarter of the MMS module was

designed and built as a development version to validate the structural

concept, the assembly method and the combination of materials selected

(see Figure 8-30). It is within the weight allowance, has a nice

appearance and is rugged. Two foot drops on a corner and transport as

baggage on a commercial aircraft did not damage the mockup. The

approach was accepted by the MSFC personnel and was used for the

deliverable MMS module mockups.

8-48
ORIGINAL PAGE |S

POOR QUALfTY

Figure 8-30 MMS Module Partial Mockup

8.1.4 Concept Definition of the Modified MST

The existing Module Servicing Tool (MST) wei-ghs approximately 67 Ibs

and in order to be used for ground demonstrations needs to be

redesigned for a drastic weight reduction and reduction of the "B"

distance, between end effector interface and the module latch interface,

8-49
Several ways of achieving the required light weight and compactness

were identified. ..Ml'-riV*1

•.-»•• V

VT»,f •:•'.

A maximum module mockup weight of 12.5 Ibs. and a "B" distance of 7.25

in. were selected for the ground demonstrations. From the graph in

Figure 8-12, the maximum allowable weight of the modified MST is 15

Ibs. An allowance of approximately 0.5 Ib was made for the electrical

connector positioner mechanism to be added to the end effector. Other

applicable requirements are given in Section 2.1.

Details of the modified MST concept definition effort are reported in

Section 3.2.2. Other applicable MST requirements are given In Section

3.2.1.

In Figure 3.2-1 the general configuration of the modified MST for

ground demonstration, the critical dimensions required for providing

adequate clearances and the approximate e.g. position are given.

This figure also defines the mechanical Interface between the ETU, the

modified MST designed by Fairchild Space Company and built by GSFC and

the MMS module mockup provided by Martin Marietta Aerospace.

8.1.5 Spacecraft Mockup Configuration Selection

Four candidate configurations were identified and evaluated for the

spacecraft mockup of the servicer system for ground demonstrations of

MMS module exchange. These candidate configurations were:

1) Addition of a MMS module and simple support structure to the

existing ETU spacecraft mockup;

2) As In 1), except for an MMS mockup added around the support

structure to give a higher level of fidelity;

3) An MMS mockup replacing the existing ETU spacecraft mockup;

8-50
4) Use of two separate Interchangeable mockups.

This set of 1-g spacecraft mockup configuration alternatives was

reviewed with MSFC personnel, including the NASA Technical Manager. It

was agreed that these four alternatives spanned the range of interest

and provided an adequate variety of advantages and disadvantages.

The relative position of the MMS module mockup with respect to the

servicer arm location, in all these four alternative configurations,

corresponds to the actual module position when performing on-orbit

module exchange, using lateral docking of the servicer to the MMS and a

docking probe fitted with an orientation joint.

The advantages and disadvantages of these four candidate configurations

were analyzed, based on the requirements developed in Section 8.11, and

Configuration 1) was recommended for ground demonstrations of MMS

module exchange.

8.1.5.1 Configuration 1) - A MMS module mockup is bracketed off one side of

the existing spacecraft mockup as shown in Figure 8-31. The concept is

simple and straightforward. The existing boxes that are attached to

one side of the ETU spacecraft mockup are replaced by a support

structure, which was provided with the necessary module retention

hardware, sensors, connector and wiring to receive a MMS module

mockup. The existing axial and radial cubic module locations of the

ETU spacecraft mockup are retained together with the existing

capability of basic module exchange demonstrations. Ample room is

available for later expansion of the spacecraft mockup to include a

propulsion mockup for fluid resupply demonstrations. There is no

appearance of an MMS triangular structure configuration. However,

realistic trajectories of MMS module exchange can be demonstrated at a

relatively low cost. This spacecraft mockup configuration emphasizes

the use of MMS modules. The approach is consistent with the use of MMS

modules on spacecraft other than those using the MMS triangular

structure.

8-51
EXISTING ETU SPACECRAFT MOCKUP

SUPPORT STRUCTURE

FUNCTIONAL MMS MODULE MOCKUP

DOCKING PROBE MOCKUP

Figure 8-31 Configuration 1) of the Spacecraft Mockup

8.1.5.2 Configuration 2) - Configuration 1) is modified by adding an MMS

mockup around the support structure as shown in Figure 8-32. The

existing capability for basic module exchange demonstration is

retained. A good representation of the MMS helps understand the

positional relationship with the servicer and increases the fidelity of

the module exchange demonstrations by providing the same clearances

between the module and MMS support structure, grapple fixture and other

spacecraft elements. However, the cost is higher because of the

additional mockup, the orientation joint of the docking probe mockup is

not visible and it is somewhat difficult to explain the demonstration

setup because there are two spacecraft mockups looking like one mockup.

8-52
EXISTING ETU SPACECRAFT MOCKUP

MMS MOCKUP

n.^GRAPPLE FIXTURE MOCKUP

FUNCTIONAL MMS MODULE MOCKUP

DOCKING PROBE MOCKUP

Figure 8-32 Configuration 2) with Added MMS Mockup

8.1.5.3 Configuration 3) - The existing ETU spacecraft mockup is replaced

by one representing an MMS. The basic axial and radial module exchange

would be performed in the payload section of the MMS (see Figure

8-33). Because of the reach envelope limitations, the existing ETU arm

being shorter than the flight version of the servicer arm (45 in. vs.

79 in. length of shoulder or forearm segments), the basic radial module

insertion is performed in the wrong direction (out from, instead of in

towards the docking probe).

8-53
PAYLOAD SECTION

MMS MOCKUP

BASIC MODULE

RADIAL LOCATION

FUNCTIONAL MMS

MODULE MOCKUP

BASIC MODULE AXIAL LOCATION

DOCKING PROBE

_^ HOCKUP

Figure 8-33 Configuration 3) with MMS Mockup and Payload Section

This demonstration setup tends to decrease the apparent importance of

the basic radial and axial module exchange. The cost of the spacecraft

mockup is higher than in the case of Configurations 1) or 2) and

exceeds the available funds.

8.1.5.4 Configuration 4) - Two separate, interchangeable mockups are used:

the existing ETU spacecraft mockup and an MMS mockup as shown in Figure

8-34. The mockups would be interchanged using the new MSFC high

capacity manipulator system.

The concept of exchanging the two spacecraft mockups has significant

disadvantages. The large mockups are relatively fragile, the

realignment of the spacecraft mockup is difficult, additional time is

required for mockup interchange and the room to store the unused mockup

is not available.

8-54
EXISTING ETU

SPACECRAFT

MOCKUP

MMS MOCKUP

Figure 8-34 Configuration 4) Interchangeable Mockups

8.1.5.5 Configuration Selection - The advantages and disadvantages of the

four alternative configurations were analyzed and a trade study was

performed based on requirements. Configuration 1), shown in Figure

8-31, was selected and is recommended for the MMS module exchange

ground demonstrations.

8.1.6 Stowage Rack Mockup Configuration Selection

The existing ETU stowage rack can be used with minor modifications to

perform ground demonstrations of MMS module exchange and basic, 24 in.

cube, module exchange. An MST storage rack and supports to receive MMS

module mockups in two locations must be added to the existing stowage

rack mockup. The two existing 24 in. cube module supports also need

modification of their baseplates to permit their installation closer to

the end—of-beam triangular structure in order to make room for the

large MMS modules. The interface of the MMS module with the stowage

rack support should be identical to the one used on the spacecraft

mockup. The same module retention hardware, electrical connector and

sensors should be used. Only the "good" module location should be

8-55
provided with a connector Interface. The support structure should

allow small displacements parallel to the baseplate to avoid damage in

case of an accident. Ease of realignment of the supports and optical

targets is an important requirement. Wiring modifications are

necessary for interfacing between the Servicer Servo Drive Console and

the MMS supports and the modified MST. The MST storage rack should be

provided with a standard MMS module latch interface plate, guide tube,

sensor for MST presence and with an optical target. The MST latches

will be used for MST attachment to the storage rack.

The location of the modified MST storage rack and of the MMS and basic

modules were selected based on the requirements developed in Section

8.1.1.

8.1.6.1 Modified MST Storage Position Selection - The modified MST, when

stored in its storage rack, extends vertically about 12 in. above the

top of the stowage rack beams. Because of a potential interference

with the servicer arm operation and the need to provide good visibility

during the demonstrations, the MST storage location was selected

first. The location needs to be at a large radius, away from the

operating range of the servicer arm and close to the stowage rack beam

to minimize the support structure. The selected location is shown in

Figure 8-35 and the cylindrical coordinates are: r = 80.00 in. and

Theta = -87.38°. The latch interface plate of the MST storage rack is

at the same height as the module latch interfaces, corresponding to X =

-19.5 in. for the end effector coordinate.

8.1.6.2 Stowage Rack Alternative Arrangements - The MMS module mockups must

be located with one corner close to the docking probe (locations IV and

V in Figure 8-35) to be able to fit in, with adequate clearance for two

24 in. cube modules. Also the MMS module fasteners must not be in line

with either cube module to avoid interference. Because of these

requirements, only the two alternative arrangements, shown in Figure

8-35, are feasible (arrangements I-III-IV-V and II-III-IV-V).

8-56
The 24 in. cube module with side interface mechanism can be stowed in

two of the three locations marked I, II and III while the MMS module

mockup occupies locations IV and V.

MST STORAGE

Figure 8-35 Stowage Rack Alternative Arrangements

There are two possibilities for arranging the "good" and the

"temporary" MMS module locations, the "good" module can be either in

location IV or V, with the " temporary" position in the other location.

The top of the MMS module has a cross beam reinforcement for the

restraint sockets and is considerably more rigid than the bottom. In

order to improve the accuracy of positioning the module, the top

fastener location should be used for attachment to the modified MST and

servicer when handling the module. The arrows in Figure 8-35 indicate

the possible location of the top fasteners for each MMS module

location. There are four possible alternatives from this standpoint.

8-57
In all, there are 16 variants of the stowage rack arrangement that are

feasible and that were evaluated in a study of servicing times to

select the stowage rack mockup configuration.

8.1.6.3 Coordinates of Fastener Locations - The coordinates of the MMS module

fastener locations for all the positions used in the 16 variants of

module arrangement as well as of the modified MST storage rack and of

the FLIP locations were calculated and are shown in Table 8-2. Also

shown are the values of the joint angles and of the fastener

orientation angle alpha. The equations used to calculate these

coordinates and angles were presented in Section 8.1.2. For each MMS

fastener location of the stowage rack two different x coordinate values

were considered to allow proper clearance when the end effector has

only the modified MST attached (Tool case) or when both the module and

the modified MST are attached (Module case). The MMS module fastener

location in the top line of the table is identified by the module

location (IV or V) and by the fastener orientation with respect to the

module (N-E-S-W), the "north" direction being parallel to the "Z" axis

of the servicer. The values of the coordinates used in this phase of

the study were preliminary and vary slightly from the actual values

selected in a later phase. However, the results of this analysis are

not affected substantially by these variations and the conclusions

based on these calculations remain valid.

8-58
Table 8-2 Coordinates and Joint Angles

^"XLOCATION

COORD. ^X.

r (in.

Theta

X.

(in.)

Alpha

Delta

(deg)

T

(deg)

U

(deg)

V

(deg)

W

(deg)

Y

(deg)

Z

(deg)

Psi

(deg)

Phi

(deg)

Omega

(deg)

)

(deg)

Tool

Mod.

(deg)

Tool

Mod.

Tool

Mod.

Tool

Mod.

Tool

Mod.

Tool

Mod.

Tool

Mod.

Tool

Mod.

Tool

Mod.

Tool

Mod.

Tool

Mod.

IV N

63.35

+61.74

+5.50

+11.00

+90.00

+37.12

+35.84

+11.25

+10.99

+7.02

+14.15

+83.84

+82.75

0.00

0.00

+90.00

+90.00

+8.86

+7.58

0.00

0.00

+90.00

+90.00

-28.26

-28.26

IV

63

+28

+5

+11

-180

+37

+35

-22

-22

+7

+14

+83

+82

0

0

+90

+90

+245

+244

0

0

+90

+90

+208

+208

E

.35

.26

.50

.00

.00

.12

.84

.23

.49

.02

.15

.84

.75

.00

.00

.00

.00

.38

.10

.00

.00

.00

.00

.26

.26

IV

30

+11

+5

+11

-90

+64

+62

-56

-57

+7

+14

+128

+127

0

0

+90

+90

+165

+163

0

0

+90

+90

+101

+101

S

.59

.31

.50

.00

.00

.29

.16

.29

.94

.02

.15

.12

.57

.00

.00

.00

.00

.60

.47

.00

.00

.00

.00

.31

.31

IV W

30.59

+78.69

+5.50

+11.00

+0.00

+64.29

+62.16

+11.09

+9.44

+7.02

+14.15

+128.12

+127.57

0.00

0.00

+90.00

+90.00

+142.98

+140.85

0.00

0.00

+90.00

+90.00

+78.69

+78.69

V N

30.59

-11.31

+5.50

+11.00

+90.00

+64.29

+62.16

-78.91

-80.56

+7.02

+14.15

+128.12

+127.57

0.00

0.00

+90.00

+90.00

-37.02

-39.15

0.00

0.00

+90.00

+90.00

-101.31

-101.31

V S

63.35

-61.74

+5.50

+11.00

-90.00

+37.12

+35.84

-112.23

-112.49

+7.02

+14.15

+83.84

+82.75

0.00

0.00

+90.00

+90.00

+65.38

+64.10

0.00

0.00

+90.00

+90.00

+28.26

+28.26

FLIP

75.00

-69.95

+12.00

+12.00

+45.00

+24.96

+24.96

-111.54

-111.54

+15.47

+15.47

+62.68

+62.68

0.00

0.00

+90.00

+90.00

-89.99

-89.99

0.00

0.00

+90.00

+90.00

-114.95

-114.95

TOOL

80.00

-85.71

+5.50

n/a

+0.00

+21.44

n/a

-123.06

n/a

+7.02

n/a

+55.02

n/a

0.00

n/a

+90.00

n/a

-64.27

n/a

0.00

n/a

+90.00

n/a

-85.71

n/a

8-59
8.1.6.4 Representative Servicing Trajectory Sequence - The diagram shown in

Figure 8-36 represents schematically the 23 steps of the total

trajectory used to exchange an MMS module. The starting or ending

points of each step are the MMS module fastener locations on the

stowage rack and spacecraft mockups, the rest position of the arm, the

modified MST storage location or the FLIP location.

S/C11

LEGEND

• TOP FASTENER

• BOTTOM FASTENER

® TOOL ONLY. DOWN

® MODULE AND TOOL, DOWN

Figure 8-36 Trajectory Sequence for MMS Module Exchange

8-60
The arm moves from the rest position to the modified MST storage

location, picks up the tool and goes to the FLIP location. After the

FLIP step is completed, the arm goes to detach the bottom MMS fastener

of the spacecraft mockup. The next step is to go to and detach the top

fastener and then, with the MST latches still closed, to carry the

module through the FLIP step to the "temporary" MMS module location on

the stowage rack. Handling of the module is done using the top

fastener latch Interface. Module attachment is done in a reverse

sequence, the bottom fastener being tightened last. In a similar

manner, the "good" module is picked up from the stowage rack mockup and

is installed on the spacecraft, passing through the FLIP step.

Finally, the arm and the tool return to the stowage rack, move the

module from the "temporary" to the "good" location, then the tool is

stored and the arm returns to the rest position, completing the total

trajectory. The steps can be followed on the diagram using the arrows

and the sequential numbering.

Once the spacecraft mockup configuration, the FLIP location, the rest

position and the storage location of the modified MST are established,

the time necessary to complete the total trajectory depends on the

module arrangement on the stowage rack. When using a different variant

of module arrangement for the stowage rack, only 6 steps (shaded on the

diagram) out of 23 take a different length of time to be performed.

The total time to perform these 6 steps was calculated for each of the

16 variants of module arrangement in the stowage rack to determine

which arrangement corresponds to the minimum servicing time. A

computer was used in performing these calculations. The allowance of 3

sec for speeding up when starting and slowing down when ending each

step was neglected, being a constant for all variants. Four

coordinates (x, Theta, r and Z) were considered to be driven at the

same time at constant rates from their initial to final values, in each

step. The rates, established in the Servicer Simulation Software

Requirements document, were: x = 2 in./sec; 6=6 deg/sec; r = 2

in./sec and Z = 12 deg/sec. The time to perform each step was taken as

the largest of the time intervals for the individual coordinates. The

other three coordinates remain constant to the end of the step, upon

8-61
reaching their final values. For each variant, a total time necessary

to perform the six steps was calculated. The results are shown in

Table 8-3.

Table 8-3 Times Necessary to Perform the Six Steps

Variant 1

Variant 2

Variant 3

Variant 4

Variant 5

Variant 6

Variant 7

Variant 8

Variant 9

Variant 10

Variant 11

Variant 12

Variant 13

Variant 14

Variant 15

Variant 16

T = 109.9 sec

T = 119.2 sec

T = 114.6 sec

T = 114.5 sec

T = 132.6 sec

T = 120.2 sec

T • 115.7 sec .

T = 137.2 sec

T = 114.6 sec

T = 119.2 sec

T = 119.4 sec

T = 114.6 see

T - 137.4 sec

T = 120.3 sec

T - 115.7 sec

T = 132.6 sec

Variant 1 corresponds to the minimum servicing time. The module

arrangement and the trajectory sequence diagram is shown in Figure

8-37. The maximum amount of time saved over the variant taking the

longest time is 27.5 seconds. Most of the time saving comes from the

fact that all four MMS module fasteners are in line, reducing the aria

travel distances between them. Also step 22 returning the modified MST

to its storage location is the shortest possible.

8-62
S/C12

LEGEND

$ TOP FASTENER

• BOTTOM FASTENER

©TOOL ONLY, DOWN

©MODULE AND TOOL, DOWN

Figure 8-37 Variant 1 - Trajectory Sequence for Minimum

MMS Servicing Time

8-63
8.1.6.5 Selection of Stowage Rack Arrangement - The recommended stowage rack

arrangement is shown in Figure 8-38. It corresponds to the minimum

servicing time. The "temporary" MMS module location is "IV," both top

fasteners are in the 12 o'clock orientation (S/R 11 and S/R 13) and the

24 in. cube modules are in the "II" and "III" locations.

Figure 8-38 Recommended Stowage Rack Arrangement

The outline of the spacecraft mockup is shown in this top view using

phantom lines. Fastener locations, both for the stowage rack (S/R) and

for the spacecraft (S/C) are indicated. The modified MST storage

location is S/R 15. All the requirements, presented in Section 8.1.1,

for the stowage rack configuration are met.

8-64
The two existing 24 in. cube modules with side interface mechanisms and

their existing support structures are used in locations II and III.

The baseplates of these support structures needed to be modified to

prevent interference with the stowage rack beams when the modules are

located as shown in the figure. A corner of each baseplate was

removed to provide clearance with the stowage rack end-of-beam

triangular structure. Cutouts were provided in the bottom frame of the

baseplate that is used in location III to clear the radial structural

member adjacent to the bottom plate of the stowage rack.

The modified MST storage rack will include a module latch interface and

a guide tube similar to the ones used on the MMS module mockup and will

be attached to the stowage rack structure using a simple bracket. A

sensor for the presence of the modified MST in the storage rack and an

optical target will be provided. The modified MST will be secured in

this location using its latches.

8.1.7 Servicer System Configuration - 1-g Demonstrations

The servicer system configuration definition included trade studies to

select the spacecraft mockup configuration, the MMS module design and

the stowage rack module arrangement.

The servicer system configuration for 1-g demonstrations of MMS module

exchange is shown in Figure 8-39. It includes a spacecraft mockup with

a MMS module mockup bracketed off one side of the existing spacecraft

mockup, the servicer arm attached to the docking post and a stowage

rack mockup, modified to receive the MMS module in two locations, the

24 in. cube module with side interface mechanism in two locations and a

modified MST storage rack.

The arrangement of the modules in the stowage rack mockup was selected

based on the minimum servicing time by analyzing sixteen alternatives.

8-65
The selected servicer configuration permits going easily from

demonstrations of axial or radial module exchange using the 24 in. cube

modules to demonstrations of axial exchange of MMS module mockups or

vice versa. It will only be necessary to select the proper set of

software, make a change in one electrical connection, and select the

desired trajectory.

The spacecraft mockup configuration emphasizes the use of MMS module

exchange. This approach is consistent with the use of MMS modules on

spacecraft other than those that use the MMS triangular structure.

SPACECRAFT

MOCKUP

SUPPORT

STRUCTURE

24 IN. CUBE MODULE

MOCKUPS

STOWAGE

RACK

MOCKUP

Figure 8-39 Servicer System Configuration - 1-g Demonstrations

8-66
8.1.8. Schedule and Cost Considerations

The next step after the servicer system configuration for 1-g

demonstrations of MMS module exchange was selected was to develop the

schedule and estimate the cost of design, fabrication, and installation

of the new equipment and ETU modifications according to Change Order 3

to this contract. The Change Order 3 schedule and cost specifics were

presented to MSFC in March 1985 at the Midterm Review and are

summarized in the following paragraphs.

8.1.8.1 Change Order 3 Activities - Change Order 3 represented a major

increase in the contract activities as listed in Table 8-4. The first

activity involved drawing preparation and fabrication of most of the

mockup and electrical equipment needed to demonstrate the exchange of

Multi-Mission Modular Spacecraft modules using the Engineering Test

Unit.

Table 8-4 Change Order 3 Activities

Servicer/MMS 1-g Demonstration Equipment

- Drawing Preparation

- Fabrication and Assembly

- Installation at MSFC and Checkout

MMS Module Exchange Software

- Requirements Documentation

- Manual-Augmented Trajectory Sequence Documentation

- Software Preparation

- Software User's Manual

Servicer Demonstration

- Software Installation and Checkout

- MMS Module Exchange Demonstration

8-67
A 1-g version of the MMS Module Servicing Tool was provided by Goddard

Space Flight Center. The MST is used as an interface between the ETU

end effector and the MMS modules and provides the mechanical torque to

tighten and loosen the MMS attachment bolts. Requirements for design

of the 1-g version of the MST have been prepared. Upon completion of

fabrication at Martin Marietta, the Change Order 3 equipment was

checked out, shipped to MSFC, Installed, and checked out again.

The MMS module exchange software was developed along the same lines as

the basic module exchange software. The same form of trajectory

hierarchy was used, but with the necessary differences to allow for use

of the MST, two attachment bolts, and for the size of the modules. All

documentation and the software program itself for the MMS modules, were

separate from the basic module exchange software.

The software was delivered to MSFC and checked out on their PDP-11/34

computer with the Engineering Test Unit. Initial demonstrations of MMS

module exchange in all three control modes were made.

8.1.8.2 Change Order 3 Deliverable Mockups - The equipment listed in Table

8-5 was delivered and installed at MSFC as part of Change Order 3. Two

MMS module mockups were produced — one "failed" module for the

spacecraft mockup and one "good" module for the stowage rack.

Table 8-5 Change Order 3 Deliverable Mockups

MMS Module Mockups (2 sets)

Spacecraft Mockup with MMS Module Receptacle

MMS Module Receptacle with Support Structure (2 sets)

ETU Wiring Modifications

Electrical Connector Positioner

Module Servicing Tool Storage Rack

MMS Module Location Targets

8-68
A receptacle for an MMS module was produced and fastened to the

existing spacecraft mockup with brackets. There is no spacecraft

mockup that looks like the Multi-Mission Modular Spacecraft, rather the

visual emphasis is on MMS module exchange.

Two MMS module locations were required in the stowage rack — one for

temporary module storage and one for storage of the "good" module.

Each of these locations consists of a receptacle for the MMS module and

a support structure of the proper height.

The ETU wiring modifications included: 1) wiring to the MST; 2) wiring

to the electrical connector positioner; 3) revisions to the servicer

control panel; 4) junction box changes; and 5) wiring to indicators at

each fastener of each module location.

The electrical connector positioner was fastened on the ETU end

effector and is used to make an electrical connection between the ETU

and the MST after the MST has been grasped by the ETU end effector jaws.

8.1.8.3 Change Order 3 Schedule and Cost Specifics - Figure 8-40 shows the

major schedule milestones for the Change Order 3 activity as presented

to MSFC at the Midterm Review, in March 1985. The schedule was

generally followed as planned with one difference being that the MST

integration meeting was held at the beginning of June, one day before

the Design Coordination Meeting.

The Change Order 3 activity started on April 1, 1985 to provide time to

complete most of the work associated with the basic contract and Change

Order 1.

The equipment (Task 1) and software (Task 2) activities were conducted

in parallel and led to the software installation and initial MMS module

exchange demonstration in October of 1985. The equipment activity was

arranged as a waterfall of drawing, fabrication, and installation. The

installation subtask included assembly and the checkout at Martin

Marietta, shipping to MSFC, and installation and checkout at MSFC.

8-69
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8-70
The software activities consisting of requirements definition, software

development, and preparation of a Software User's Manual were scheduled

in the waterfall pattern. The second part of the software requirements

subtask was the preparation of the trajectory sequence documentation

necessary to conduct a module exchange in the Manual-Augmented control

mode. A time span was allowed in November for updating the Software

User's Manual. The MMS Software User's Manual was prepared to the

requirements of DR-6 of DPD650.

Two meetings were planned — one for MST integration at GSFC and one to

review the drawings at MSFC, in addition to the final review, which was

scheduled in mid-November of 1985.

The total cost of the Change Order 3 activities was estimated at $200 K

at the time of the proposal preparation. The cost of design,

fabrication, testing, and installation of the Modified MST by Fairchild

Space Company and GSFC was not Included in this amount, being funded

separately.

8.2 SERVICER/MMS 1-g DEMONSTRATION DESIGN

Change Order 3 activities included the design of the necessary

equipment for adding a MMS module exchange capability to the existing

Engineering Test Unit. With the preliminary concept design completed

under Change Order 1 of the contract, the effort consisted mainly of

design refinement and coordination and drawing preparation and checking.

The main assembly drawings produced are listed in Table 8-6 and the

drawing tree is presented in Figure 8-41. The drawings produced were

"Form 4" level for cost reduction, with the parts list and notes on the

same drawing and with parts call-outs on the drawing field, simplified

call-out of standard processes and some "not-to-scale" dimensions

permitted. Also the same person preparing the drawings was providing

all the engineering support functions, such as stress analysis,

materials and standard parts and processes selection, procurement and

drawing checking.

8-71
Table 8-6 Assembly Drawings for Servicer/MMS 1-g Demonstrations

DESCRIPTION

MMS Module Mockup Assembly

Spacecraft Mockup Assembly

Stowage Rack and Spacecraft Interface Assembly

Connector Positioner Mechanism Assembly

MST Storage Rack Assembly

Optical Targets Assembly

New Cable Assemblies

MMS Junction Box Assembly

DRAWING NO.

RES4100000

RES4200000

RES4300000

RES4400000

RES4500000

RES4600000

RES4XOOXXO

RES3159950

In addition to drawing preparation, this effort included preparation of

MST interface documents, attending MST Integration Meeting and Design

Review Meeting as well as incorporation of modifications to the

drawings and MST Interface Documents.

8-72
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8-73
8..2.1 MMS Module Mockup '. ' "

The mockup of the MMS module (see Figure 8-42) has the same outside

dimensions and shape as the actual module. The lightweight box

structure is reinforced by an internal torque box, directly connecting

the upper and lower fasteners. Not shown on the figure is a layer of

metallized polyester film, that covers the box and is bonded to it.

The plastic film is 0.002 in. thick and represents the thermal

multilayer insulation, while strengthening the box. The material used

for building the box and its internal structure is 3/4 in. thick

styrofoam board, with large cutouts for weight reduction and

accessibility during construction.

3 AXIS CONSTRAINT SOCKET

ELECTRICAL CONNECTOR

MOCKUP

r-TOP FASTENER

GUIDE TUBE

MST

INTERFACE

2 AXIS CONSTRAINT

SOCKET

BOTTOM FASTENER

Figure 8-42 MMS Module Assembly

8-74
Special attention was paid to the proper selection of the adhesives .

used for joining the styrofoam parts to themselves and to the outer

layer of film. The majority of adhesives readily available are not

suitable because they dissolve or etch the styrofoam or stain the

metallized film. A special styrofoam panel adhesive was used for the

box and urethane foam was used to reinforce the bond lines. Urethane

bond was used to install the metallized film, as well as the upper and

lower interface boxes and the electrical connector mockup, that are

made of balsa wood and covered with fiberglass.

The MST interface is the same as for the actual module except that

different materials were used. The front interface plate is made of

thin wall fiberglass with aluminum inserts for the latch attachment

area instead of solid aluminum. The guide tube is made of 0.015 in.

instead of 0.030 in. thick fiberglass.

The electrical connector mockup has the same shape, size, and relative

position as the plug side of the actual disconnect used in the back of

the MMS power module. A small, five pin connector half, housed within

the mockup, mates with the other half installed on the spacecraft or

stowage rack interface.

The module retention system has the same interface with the spacecraft

as the actual hardware. It is comprised of the upper and lower

fastener assembly and two constraint sockets. A description of the

materials used for the fastener assembly was given in Section 8.1.3 and

Figure 8-24. The hollow core aluminum bolt is permanently lubricated

with dry lube finish. A maximum loosening torque of 20+1 ft-lb was

considered for the design of the fastener, MST interface plate, and the

supporting structure.

The constraint sockets are thin aluminum shells riveted to the upper

beam of the module. The two-axis socket is mounted in slotted holes

allowing 3/8 in. horizontal float (parallel to the beam). The sockets

have approximately 0.020 in. preload when the central bolt is tightened.

8-75
The design total weight was 12.5 Ibs maximum. The lightweight and

resilient styrofoam material of the box and the box design assure the

ruggedness required to withstand 400 cycles of MMS module exchange

demonstrations and the accidental load due to a 2 ft drop on a flat

surface, without substantial damage. The mockup is easy to repair in

case of accidental damage using available styrofoam and urethane bond.

8.2.2 Stowage Rack Module Support

The two MMS module mockups built under Change Order 3 can

interchangeably be installed in two locations on the stowage rack

mockup, a "good" location and a "temporary" one, as well as in one

location on the spacecraft mockup. The design of the module interface

is the same for all three locations. Each of them is comprised of an

aluminum tubular frame supporting two floating nut units for the two

bolts of the module and two constraint balls to match the sockets of

the module (see Figure 8-43).

Near each floating nut unit, on the opposite side of the frame, there

is a sensor unit with three microswltches and a spring loaded rod. It

senses the "ready-to-latch", "latched", and "unlatched" condition of

the fastener by gaging the protrusion of the bolt through the floating

nut.

Four legs, braced to the module interface frame provide the support and

the proper elevation at both stowage rack MMS module locations. They

are provided with swivel leveling pads for ease of height adjustment

and leveling.

The floating nut unit (see Figure 8-44) is of the same construction as

the actual unit except for the use of aluminum instead of stainless

steel for the housing and cap.

8-76
The stainless steel nut is integral with the guiding cone. Its

floating capability is +0.100 in. radially and _+8.5° for angular

misalignment. Softer springs, as compared with the actual hardware,

were used to reduce the loads on the module mockup structure.

FLOATING NUT UNIT

CONNECTOR

RECEPTACLE

MOCKUP

TUBULAR

FRAME

•CONSTRAINT

BALL

co ;

CM

/- SWIVEL

TSr LEVELING

PAD

SENSOR UNIT BRACING

Figure 8-43 Stowage Rack Module Support

8-77
FLOATING NUT

GUIDING CONE

HOUSING

Figure 8-44 Floating Nut Unit

An electrical connector receptacle mockup is attached to the spacecraft

module interface and to the "good" location of the MMS module support

on the stowage rack. The receptacle (see Figure 8-45) is mounted in a

spring loaded housing that allows +0.20 in. mismatch and large angular

misalignment at the start of engagement with the connector mockup of

the module. The tapered walls of the receptacle guide the connector

mockup and provide alignment of the small electrical disconnects within

0.010 in. during their mating. Bevels on all four sides of the

receptacle flange and on the mounting plate help center the receptacle

under spring pressure after removing the module mockup. This compliant

mounting of the receptacle provides smooth engagement of the small

electrical disconnect that is housed within it and limits the loads on

8-78
the MMS module tnockup structure. An added safety feature is the

redundancy of contacts of the five-pin electrical disconnect. The pin

half of the disconnect that is mounted on the module has all the

contacts electrically connected with jumpers. On the receptacle side,

the socket half of the disconnect has three contacts connected to one

wire and two to the other. Thus the electrical circuit is closed upon

mating the disconnects even if one pin is accidentally broken.

The attachment brackets of the electrical connector receptacle mockup

are adjustable in all directions using slotted holes.

ELECTRICAL DISCONNECT

RECEPTACLE AMOUNTING

r L MI L.

BEVEL

HOUSING

SPRING

Figure 8-45 Electrical Connector Receptacle Mockup

8-79
8.2.3 MMS Spacecraft Mockup

In order to emphasize the MMS module exchange and not the triangular

configuration of the existing spacecraft, only a simple structure that

supports the module interface frame and allows its adjustment and

leveling was provided (see Figure 8-46). This is consistent with the

use of MMS modules on spacecraft other than those using the MMS

triangular structure. However, the relative position of the module

with respect to the ETU arm and stowage rack mockup represents the

actual module position when performing on-orbit module exchange, using

lateral docking of the servicer to the MMS and a docking probe fitted

with an orientation joint.

The MMS module interface frame is the same as the ones used for module

support on the stowage rack, except that instead of the four legs, four

adjustable rods are used for structural attachment. They permit

vertical adjustment and leveling of the module interface frame. Two

adjustable links, a subframe, two angle brackets, and two adjustable

clevises form the support structure that is attached to the existing

spacecraft mockup. Rework of the existing ETU spacecraft mockup to

receive the MMS module included deletion of the boxes on one side of

the mockup body and local internal structural reinforcement at the

attachment .points.

Later expansion of the MMS spacecraft mockup to incorporate fluid

resupply demonstration hardware is possible at each end of the

interface frame. Tanks and plumbing components can also be fitted

within the support structure of the MMS spacecraft mockup.

8-80
ANGLE BRACKET

EXISTING SPACECRAFT

MOCKUP

•ADJUSTABLE CLEVIS

•ADJUSTABLE LINK

SUBFRAME

REINFORCING

STRUCTURE ,

ADJUSTMENT

ROD

L

 2.0

NOM.

MMS MODULE

INTERFACE FRAME

Figure 8-46 MMS Spacecraft Mockup

8.2.4 Connector Positioner Mechanism

The connector positioner mechanism installed on the existing ETU end

effector provides proper alignment and translation stroke for mating a

subminiature, double density 19-pin connector, part number 2DE-19P, by

ITT Cannon. It connects up to 19 22 AWG leads across the interface

between the ETU end effector and the modified MST for the provision of

power, control, and monitoring circuits. The socket side of the

connector, part number 2DER-19S, is installed on the MST and has a

floating mount, permitting 0.030 in. of diametral float.

Figure 8-47 shows the main features of the connector positioner

mechanism. Several types of mechanisms were considered for the

connector positioner: a solenoid actuated mechanism with direct drive,

8-81
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8-82

ORIGINAL PAGE rs

OF POOR QUAUTY
a paralellogram linkage, a single lever arrangement with compliant

connector attachment, a ball screw and nut with direct drive and an

eccentric actuated mechanism, with linear ball slide guiding. These

candidate mechanisms were traded against the requirements given in

Section 8.1.1. The solenoid mechanism was too heavy, the parallelogram

linkage was too complex, not compact enough and difficult to set up and

test before installation on the ETU. The single lever and the ball

screw and nut mechanisms were too complex and not compact enough. The

eccentric mechanism best met all the requirements and was selected for

the connector postioner. The short stroke, high final force,

compactness, and light weight characteristics made the eccentric

mechanism well suited for this particular application.

The mechanism uses an electrical (28 Vdc) gearmotor, a 1.0 in. diameter

eccentric, a needle bearing mounted adjustable link and a linear ball

slide unit to provide 5/8 in. mating stroke and 20 Ibs mating/demating

force for the connector. It was designed as a compact, self-contained

unit that is bolted to one side of the end effector, opposite to the

existing power takeoff.. The simple interface with the ETU end effector

and the built-in adjustability simplified the integration at MSFC and

allowed adjusting and testing of the unit prior to delivery. The cost

and the weight of the mechanism was kept low by minimizing the number

of parts. A thin wall aluminum housing was used to position and

connect all the mechanism parts, to provide end-of-stroke solid stops

and to support cabling and adjustable limit switches. A standard

miniature ball slide unit was used to provide very accurate linear

guiding during the translation motion. The cumulative error of the end

effector and mechanism interface is smaller than the floating

capability of the disconnect half, mounted on the MST, providing a very

smooth engagement. The free end of the gearmotor shaft is supported by

a ball bearing that prevents overloading of the small gearmotor

bearings. Special seals, attached one each side of the eccentric help

prevent contamination of the needle bearing. The motor is stalled at

each end of the stroke by butting the link against a solid stop on the

housing. All the screws were installed with locking compound or lock

washers to prevent self loosening through vibration.

8-83
There are three places for adjustment of the connector positioner

mechanism. The link length is adjustable using a turnbuckle type screw

with two jam nuts. The smaller jam nut has a left hand thread. The

screw has a 5/32 in. hexagonal socket. Thus, the fully extended

position of the connector holder can be adjusted within 1/8 in. Two

subminiature microswltches are mounted stacked opposite to each other

inside the housing, using two screws in slotted holes. Their position

is adjustable within 0.088 in. using a Phillips screwdriver. The

adjustment can be made both before or after installation on the ETU. A

peelable shim was provided with the unit, to be installed between the

mechanism mounting flange and the ETU end effector for proper line-up

with the connector half on the MST. The installation required drilling

and tapping of three holes in the end effector body for the 6-32 size

screws. Two 1/8 in. diameter dowel pins, installed through the flange,

shim and end effector, take the shear loads and prevent shifting. The

existing "ready-to-latch" sensor of the ETU was relocated as shown on

Figure 8-47 and a 3/16 in. thick spacer was used for its installation

to prevent interference with the end effector.

Adequate clearances were provided with respect to all servicer system

elements. The compact mechanism configuration was critical in

providing the required clearance to the TV camera. The 5/8 in. stroke

of the mechanism is more than the approximately 1/4 in. stroke actually

required for mating/demating the connector and is needed for providing

+5° angular misalignment capability of the end effector with respect to

its mating interface, prior to jaw closing. The connector positioner

is compact enough and is installed such that it does not obstruct the

field of view of the TV camera.

8.2.5 MST Storage Sack

When not in use, the modified MST is stored in a special rack, attached

to the stowage rack mockup and it is secured using its latches. The

storage rack (see Figure 8-48) has the same interface with the MST as

8-84
the MMS module mockup. The same fiberglass interface plate with

aluminum inserts for latches and the same fiberglass guiding tube are

used, enclosed in a balsa wood and fiberglass box. The box is attached

to the stowage rack using a simple plywood bracket.

3.65

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79.92

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EXISTING ETU

STOWAGE RACK

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ADJUSTABLE

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SECTION A-A

20.15^

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PLYWOOD BRACKET

I ; ILL.

Figure 8-48 MST Storage Rack

A microswitch on an adjustable bracket senses the presence of the MST.

An optical target of the same design as the one used for MMS module

interface with the spacecraft or with the stowage rack is attached to a

wooden bracket on the MST rack. The relative position of the optical

target with respect to the latch interface plate is the same as for the

targets of the MMS modules.

8-85
8.2.6 Optical Targets

An optical target is installed at each MMS fastener location on the

stowage rack and spacecraft mockups and at the MST storage location,

for verification of the end effector alignment, using the existing TV

camera and lights (see Figure 8-49).

The target plate design is similar to the other targets of the ETU,

used for the basic module exchange and has a similar pattern. It

consists of a 2 1/4 in. square with its centerlines arranged such as to

indicate the top direction, towards the center of the MMS module

mockup. The lines are black on a flat white background and are 0.10

in. wide. The width of the line was selected to provide optimum

resolution when using the existing 244 x 188 pixels TV camera. The

horizontal centerline extends over the full 8 in. width of the target

plate to take full advantage of the existing TV monitor image size for

improved alignment capability. The relative position of the optical

target with respect to the end effector interface and TV camera

objective is approximately the same for both basic and MMS module

exchange demonstrations.

Approximately 1.5 in. clearance is provided between the optical target

and the MMS module mockup during the servicing demonstrations. Two

corners of the target plate are beveled to assure proper clearance

during module insertion or retrieval. In addition, the target plate

support has a compliant attachment to its mounting bracket. A spring

loaded double hinge arrangement was used to prevent damaging the MMS

module mockup in case of accidental interference and to return

automatically the target to its preset alignment. Because the mounting

bracket is attached directly to the MMS module interface frame, the

need for target resetting is minimized, in case of accidental module

support displacement.

8-86
.TOP OF

'TARGET

BEVELED CORNER

TARGET PLATE

HORIZONTAL

LINE

ADJUSTMENT

SLOTS

TARGET C.L.

A

DOUBLE

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ASSEMBLY

SPRING

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TARGET PLATE SUPPORT

END EFFECTOR I/F

"" 'T¥T~)" ir i *-

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4.50H- MOCKUP

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.65+—

 N- MODULE INTERFACE FRAME

Figure 8-49 Optical Target Assembly

8-87
The optical target is adjustable +3/8 in. in all directions in its

plane. The two mounting screws are installed in slotted holes in the

target plate and in its support. Target plate alignment can be done

using the existing template, made for the basic module optical

targets. A spacer block, 1 in. square by 2 in. approximate dimensions,

•

can be used for proper alignment of the MMS target, using the existing

template.

8.2.7 Electrical Design

The electrical design activity of Change Order 3 involved the revision

of certain drawings from MCR-78-535, SSDC Electronics Identification,

• March 15, 1978, the creation of new drawings, parts ordering, and

coordination with Fairchild Space Company. Eight drawings were revised

and eighteen new drawings were created. They incuded the new MST/MMS

connecting diagram, new cable drawings, Servicer Control Panel •

modification for connector positioner controls, and the MMS junction

box assembly drawing. The Lisa computer was used for the new drawings

because of the computer availability, ease of use, ease of

incorporating changes, and quality of output. The drawings generally

look like schematics but contain sufficient information for assembly

and wiring.

8.2.8 Parts and Materials Procurement

The design activity of Change Order 3 involved also the selection of

the materials and purchased components, contacting vendors and

preparation of the necessary procurement documents.

8.2.9 ETU/MST/MMS Interface Coordination

An important activity that was part of the Change Order 3 equipment

design involved coordination of the interfaces between the ETU and the

modified Module Servicing Tool and between the MST and the MMS module

8-88
mockup. Parties involved were the Marshall Space Flight Center

Contract Technical Manager, the MSFC Robotics Laboratory personnel

handling the modifications of the ETU control panel, the MMS office of

the Goddard Space Flight Center, in charge of building the modified

Module Servicing Tool and cabling, Fairchild Space Company performing

the design of the MST, as well as the design and manufacture of the MST

power supply and controls, and Martin Marietta Denver Aerospace

Division designing and building the rest of the equipment and the

related controls modifications for MMS servicing demonstrations with

the ETU.

Coordination of both the electrical and mechanical interfaces was

performed. It included requirements definition, connector selection,

pin assignment, cable length and MST control panel configuration. We

prepared and updated drawings of the mechanical interface between MST

and end effector and MMS module mockup. The interface coordination

involved exchange of drawings by mail, telephone conversations, as well

as an MST Interface Coordination Meeting held at Fairchild Space

Company, in Germantown, MD on June 5, 1985 and attended by

representatives of all parties involved.

8.2.10 Design Coordination

The design coordination between Martin Marietta and MSFC was assured

through regular telephone conversations between the Study Manager and

the MSFC Contract Technical Manager, a series of vists of Martin

Marietta personnel to Marshall Space Flight Center, document exchanges

through mail or telex and a Design Coordination Meeting held at MSFC on

June 6 and 7, 1985. During the meeting, we presented assembly and

detail drawings of the connector positioner mechanism, stowage rack

mockup, spacecraft mockup, MMS module mockup, and cabling. Comments

were received and go-ahead with parts manufacturing was approved.

Details of the MMS software/computer system and servicer electrical

interfaces between MSFC and Martin Marietta were discussed. The

clarification items identified at the Integration Meeting were expanded

and resolved at the Design Coordination Meeting.

8-89
Measurements of the existing spacecraft and stowage rack mockups of the

ETU were taken at MSFC to help verify the installation requirements for

the MST rack and optical targets as well as the necessary refurbishment

for MMS module exchange demonstrations.

8.3 SERVICER/MMS 1-g DEMONSTRATION EQUIPMENT"

The equipment fabricated, tested, and prepared for shipment as part of

the Change Order 3 (subtask 1-9) of this contract is listed in Table

8-7. The electrical equipment included cable assemblies, MMS junction

box and a servicer control panel modification kit for control of the

connector positioner.

Table 8-7 Change Order 3 Equipment List

- MMS Module Mockup

- MMS Spacecraft Mockup

- MMS Stowage Rack Support

- MMS Stowage Rack Support

- MMS Target Assembly

- MMS Target Assembly for MST Rack

- Connector Positioner

- MST Rack

- Electrical Equipment

Pt. No. RES4100000-009

Pt. No. RES4200000-009

Pt. No. RES4300000-009

Pt. No. RES4300000-029

Pt. No. RES4600000-009

Pt. No. RES4600000-019

Pt. No. RES4400000-009

Pt. No. RES4500000-009

2 units

1 unit

1 unit

1 unit

6 units

1 unit

1 unit

1 unit

12 pieces

Included in the shipment was a repair or maintenance kit comprised of

materials and spares that might have been necessary had any damage

occurred during transportation.

Fabrication of the Change Order 3 equipment started with the fiberglass

items (interface boxes, upper structure, and connector mockups) for the

MMS module mockup. Tooling design and fabrication, as well as aluminum

inserts and complete assemblies of balsa wood covered with fiberglass

were made in our Advanced Composite Materials Shop.

8-90
I'

Fabrication of all the assemblies of Table 8-7 , except for the

fiberglass items, wiring and electrical equipment was done in our

Prototype Development Shop.

The electrical equipment fabrication and testing, was done in our Space

Operations Simulation Laboratory.

The MMS module mockups fabrication started with cutting details of the

styrofoam box using a high speed, carbide tipped table saw. Important

weight savings were achieved by making precise, square cuts for the

styrofoam details, thus minimizing the glue line thickness and the

amount of glue required. Figure 8-50 shows the MMS module in the

process of assembly. Details of the foam structure, of the fiberglass

upper and lower interface boxes, and of the top fiberglass cross beam

are visible. During the next stage of assembly, a layer of metallized

polyester film was bonded to the outside surface of the box. A black

paint pattern was added on the front surface of the module to represent

the thermal louvers and to reduce glare. A contrasting black and white

pattern was painted on the MST interface plate. A front view of the

completed MMS module raockup is shown in Figure 8-51. A partial view of

the back side of the module , showing the upper bolt and the related

hardware, the two constraint sockets, riveted to the cross beam and the

electrical connector mockup, can be seen in Figure 8-52. The total

weight of the MMS module mockup is 10.0 Ibs, compared to the design

limit of 12.5 Ibs maximum.

The MMS spacecraft mockup assembly is shown in Figure 8-53. The module

interface frame is shown assembled with the two floating nut units, two

constraint balls, and an electrical connector receptacle mockup.

Attached to the back of the frame is the support structure for

installation on the ETU. It is comprised of a subframe, adjustable

links, and two angle brackets. Two optical targets, mounted on the

interface frame, one for each fastener location, are also shown in the

photograph.

8-91
ORIGINAL PAGE IS

OT POOR QUALITY

Figure 8-50 MMS Module Mockup Foam Structure

Figure 8-51 MMS Module Mockup, Front View

8-92
ORIGINAL PAGE IS

OF POOR QUALITY

Figure 8-52 MMS Module Mockup, Partial Back View

Figure 8-53 MMS Spacecraft Mockup Assembly

8-93
The MMS module support for the stowage rack mockup is shown in Figure

8-54. The same tubular interface frame that was used on the spacecraft

mockup is shown in this figure, fitted with four adjustable legs.

Similar arrangement of the floating nut units, constraint balls,

electrical receptacle mockup, and targets are shown. The sensor units,

behind the floating nut units, are also visible. The assembly shown is

for the "good" module location. The "temporary" one is similar, only

the electrical receptacle mockup was omitted. The MMS module assembled

on the stowage rack support is shown in Figure 8-55.

Figure 8-54 MMS Module Support for Stowage Rack

8-94
ORIGINAL PAGE IS

OF POOR QUALITY

Figure 8-55 MMS Module Mockup on Stowage Rack Support

Figure 8-56 shows the connector positioner mechanism in its fully

extended position. Details of the housing, eccentric actuated link,

connector, and linear ball slide are visible.

The connector positioner mechanism was tested using a special stand and

completely wired connector. It performed smoothly. Actuation time at

22 Vdc is approximately 3 sec. The microswltches were adjusted. No

operating problems were encountered, connector alignment was

satisfactory, and connector engagement was proper.

8-95
Figure 8-56 Connector Positioner Mechanism

8.4 CONCLUSIONS

The MST modifications for 1-g MMS module exchange demonstrations using

the ETU were analyzed. A set of subsystem requirements for the ground

demonstrations was established, including major design constraints such

as the allowable weight, maximum wrench torque, e.g. location, relative

position of the servicer and module latch interfaces and the required

sensors. Because of the need for a drastic weight reduction and

compact design for 1-g demontrations, using the flight MST hardware is

not feasible. However, for 0-g demonstrations the existing flight

qualified hardware can be used with relatively few modifications. The

interfaces of the MST with the end effector and the MMS module were

defined. The connector positioner mechanism was located on the ETU end

effector.

8-96
The servicer configuration selection for the 1-g demonstrations was

completed. The requirements were analyzed and defined. A simple,

straightforward configuration was selected for the spacecraft mockup,

that emphasizes the MSS module rather than the general MMS appearance

while providing realistic MMS servicing trajectories and preserving the

existing basic module exchange capability. The MMS module mockup

configuration and structural concept were selected. A development

version of a partial mockup was designed and built to validate the main

design characteristics. The loads on the ETU drives during the ground

demonstration of MMS module exchange were checked and found to be

within the existing design capability. The stowage rack configuration

and module arrangement were selected based on requirements. The

recommended configuration features minimal modification of the existing

stowage rack mockup and minimum MMS servicing demonstration time while

allowing the existing basic module exchange to be demonstrated without

system reconfiguration.

A detailed schedule for the servicer/MMS 1-g demonstration equipment

design and fabrication and a cost estimate were prepared.

The design effort of Change Order 3 included further refinement of the

preliminary design, drawing preparation and checking, stress analysis,

selection and procurement of components and materials, as well as MST

interface coordination and design coordination. The main design

features of the servicer/MMS 1-g demonstration equipment were

presented, followed by a description of the fabrication process and

testing. All the design requirements were met. The electrical

connector positioner mechanism performed smoothly during tests, within

the required accuracy, time of actuation, and current level. The

weight of the MMS module mockup is 10.0 Ibs, compared to the maximum

design limit of 12.5 Ibs. All the Change Order 3 equipment was

delivered to MSFC on schedule.

8-97
9.0 SERVICER CONTROL SOFTWARE -BASIC MODULES

The objective of the 1-g servicer software development is to modify the

existing ETU control software for smoother operation and additional

control modes. Smoother operation is partly dependent on the use of a

good set of geometrical relationships. The servicer arm configuration

was selected for the servicing operation and is different from that of

a general purpose manipulator. This means that equations specific to

the servicer, rather than a general set, must be used. Additionally,

there are a number of subtleties in the arrangement that must be

properly included. The basic module servicer control software

development was recorded in three documents: 1) MCR-85-1310, Servicer

Simulation Software Requirements (Basic Modules), Martin Marietta

Aerospace, Denver, CO, May 1985; 2) MCR-85-1311, Basic Servicer Control

Software User's Manual, Martin Marietta Aerospace, Denver, CO, July

1985; and 3) MCR-85-1312, Manual-Augmented Trajectory Sequences (Basic

Modules), Martin Marietta Aerospace, Denver, CO, August 1985.

The three control modes to be implemented in software are:

1) Supervisory without operator assistance;

2) Supervisory with operator assistance;

3) Manual-Augmented.

The fourth control mode — Manual-Direct — is implemented totally

within the servicer electronics and control panel hardware. This mode

is desired as a backup control, where each joint is driven

independently without the computer. Both Supervisory modes incorporate

a trajectory hierarchy in the software so that the software will

generate the correct trajectory after being told its general form

(axial or radial) and the pertinent module location coordinates. After

setting up the initial conditions, in the unassisted Supervisory mode,

9-1
the operator need only watch and provide inputs at a few safety related

points. Significantly more operator inputs are required during the

assisted Supervisory mode.

The Manual-Augmented mode is more of a piloted operation where the

operator inputs control signals via hand controllers and uses a TV

picture as the basic feedback information. The computer solves

equations so that hand controller inputs cause servicer end effector

motions that are parallel to the TV monitor coordinates.

The software program is used in conjunction with the satellite

servicing 1-g demonstration system, Engineering Test Unit (ETU)

servicing mechanism, control system servo drive electronics and

on-orbit vehicle mockups in order to accomplish satellite servicing

demonstrations at the Robotics Laboratory of the Marshall Space Flight

Center. The computer is a PDP-11/34 and an RCA APT 4801 terminal is

used. The software provides the sequencing functions and rate commands

for control of the ETU.

The software discussed in this section is for the demonstration of the

exchange of basic modules. Basic modules are approximately 24 in.

cubes and are fastened into the spacecraft or spare module stowage rack

with a side interface mechanism that requires only one drive action.

The side interface mechanism can be loosely thought of as being a

one-bolt attachment system. Section 10 of this report describes the

software for MMS module exchange where the attachment is via two bolts.

The approach used to develop software for the basic 1-g servicer

demonstration is shown in Figure 9-1. MSFC had been using a set of

software to demonstrate module exchange and this software was reviewed

to identify approaches and useful code. The Martin Marietta software

used for the Design Acceptance Review was also examined to identify

useful code. Some of the Martin Marietta code was in assembly language

and it was decided not to use that code.

9-2
REVIEW PRIOR

MMC SOFTWARE

PREPARE SOFTWARE

REQUIREMENTS DOCUMENT

REVIEW PRIOR

MSFC SOFTWARE

DEMONSTRATE MODULE

EXCHANGE

PREPARE SOFTWARE

USER'S MANUAL

Figure 9-1 Software Development Approach

The preparation of software requirements was found to be more extensive

than originally planned because the early work had only involved a few

simple trajectories. The new requirements involve 20 trajectories of

complexity similar to the prior work. The equations and control laws

from the prior work could be carried over directly. The concept of a

trajectory hierarchy was developed and used. This hierarchy is used to

establish trajectories and to present displays.

A PDP-11/34 computer was available at Martin Marietta for use in

developing the software. An earlier version of the computer operating

system and compiler were used for the Martin Marietta machine as

compared to the versions available on the MSFC machine. It was not

possible to use similar terminals so there were some differences in the

development displays.

A Software User's Manual was prepared and sent in draft form to MSFC

before the module exchange demonstrations. It was updated after the

demonstrations. The module exchange demonstrations included all three

control modes. The basic module Software User's Manual is discussed in

Section 9.2.2 and the basic module demonstrations are discussed in

Section 11.1.

9-3
The major elements and areas involved in the servicing demonstrations

are shown in Figure 9-2. To the right of the figure are the mockup

elements, the servicer mechanism (ETU), and the TV camera used in the

Manual-Augmented control mode. On the lower left hand side are the

elements in the MSFC control station. The control station at the MSFC

Robotics Laboratory was to be located across the hall from the mockup

area, but it was left adjacent to the ETU for this development work.

The Servicer Control Panel (SCP) is shown in the control station, but

it can also be installed in the Servicer Servo Drive Console (SSDC).

The SCP has knobs, switches, meters, and indicator lamps for

Manual-Direct control modes and it contains a Servo Power switch that

can be used to shut down the servicer mechanism in an emergency.

The PDP-11/34 computer, interface electronics and the Servicer Servo

Drive Console are located adjacent to the mockups. The SSDC is the

interface between the servicer mechanism and the rest of the system and

incorporates the servo electronics, switching logic, power supplies,

and a digital voltmeter to be used for trouble shooting.

The PDP-11/34 computer performs the functions indicated. It is

controlled from the RCA APT 4801,terminal and interfaces with the

hardware via the A/D and D/A converters. The shaded area shows those

signals that flow through an MSFC prepared breakout box. The breakout

box permits monitoring signals and contains relay logic for control of

the end effector jaw, interface mechanism, connector positioner, MST

latch, and MST bolt drives.

9-4
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9-5
9.1 SOFTWARE REQUIREMENTS

The software contains the components, or functions, listed in Table

9.1-1. Coordinate transformations are required both ways between the

cylindrical and joint coordinate systems and to transform the

cylindrical coordinate rates into joint angle rates as joint angle

rates are the inputs to the ETU control loops. The six outer control

loops are closed (error junction) in the cylindrical coordinate

system. The trajectory end points are generated in cylindrical

coordinates and compared with a transformation of the joint angles to

obtain error signals in cylindrical coordinates. The rate command

signals are limited so that the servos can follow the limited rates

with low errors.

Table 9.1-1 Software Components

Coordinate Transformations

Six Control Loops

Supervisory Trajectory Hierarchy

Display Generation

Manual-Augmented Transformations

Hardware Simulation

Operator Warnings

Module Location Data Collection

Associated Functions

The Supervisory mode trajectory hierarchy consists of four levels with

higher levels calling the lower levels much in the manner of calling

subroutines. Also called are rules that generate the appropriate

trajectory end points from the stored module location data. The

displays are generated in the same process and show the specific

trajectory data and where in the trajectory the servicer is at each

instant of time. A hardware simulation function (integration of joint

rates to obtain joint angles) is included in the software to assist in

software development. The joint angle integration occurs at a rate

determined by how fast the calculations can be performed so it is not

easily relatable to real time. Subroutines and procedures are included

9-6
to collect and store module location data in cylindrical coordinates.

The approach used avoids the need for the operator to explicitly make

readings and do calculations. Rather, the software does it for the

operator.

The various coordinate systems used in the software and the functions

expressed in each of the coordinate systems are listed in Table 9.1-2.

Cylindrical coordinates were chosen as they fit the system geometry

very well. The servicer working volume is a cylinder with its axis

along the docking post. Also, the initial pivoting arm form of

servicer used a translation drive along the docking post direction.

The module exchange directions (axial and radial) also parallel

specific cylindrical coordinates. Thus, module movement in the side

interface mechanism guides only involves changing one cylindrical

coordinate. This approach makes for a very simple trajectory

definition.

Table 9.1-2 Geometrical Approach

Cylindrical Coordinates

- Module location storage

Trajectories

- Computer displays

- Control Loops

- Modules move along one cylindrical coordinate at a time

Joint Coordinates

Hardware servo loops for each joint

Hardware commands are joint rates

Hardware feedbacks are joint angles and rates

TV Coordinates

Used for Manual-Augmented control mode

Hand controller motions parallel TV coordinates

9-7
The ETU hardware operates in terms of joint angles and joint rates.

The inner control loops are rate loops based on joint tachometers and

the loops are closed in the SSDC electronics. The Manual-Direct

control mode also is based on joint angles with the commanded values

being set in on angle set potentiometers and the electronics develops

the angle errors for display on the angle error meters. The TV

coordinates are used for the Manual-Augmented control mode. The hand

controller is mounted so that its motion directions parallel the TV

screen coordinates. The computer transforms the hand controller

signals into joint rate commands so that apparent TV camera motion

parallels the hand controller motion. When in the Hawk mode, hand

controller motion corresponds to radial and tangential motion.

The transformation and control loop elements expressed in software for

the Supervisory and Manual-Augmented control loops are:

1) Joint Angle to Cylindrical;

2) Cylindrical to Joint Angle;

3) Cylindrical to Joint Rate;

4) Display Term;

5) Error Computation;

6) Rate Generation;

7) Limiters;

8) Commanded Rate Limiters;

9) Display to Cylindrical Rate;

10) Error Meter Drive Signals.

9-8
The specific equations used for the above control loop elements were

derived during the IOSS study and verified during this study.

Analytical block diagrams showing the equations and their relationships

were developed for the Supervisory modes and for the Manual-Augmented

mode and are included in the software requirements document. To

simplify the equations and to obtain explicit solution forms the

geometry was divided into three separate cases. These are:

1) Axial with module towards the spacecraft;

2) Axial with module towards the stowage rack;

3) Radial.

For each of these three cases, only four degrees of freedom are

required, the other two joint angles are kept fixed. During

transitions between one case and another only one joint is allowed to

move while the other joints have zero commanded joint rates. In this

way the singularities that would result when going from one case to the

other are avoided. The rate generators in the Supervisory motion are

simple multiplications of the respective cylindrical coordinate error

signals by constants. The inverse of each constant is a first order

estimate of the control loop time constant. By selecting these time

constants to lie between 0.3 and 1 seconds,.the control loops filter

out high frequency components and the servicer mechanism moves

smoothly. The time constants are low enough to keep the position

errors within acceptable limits.

The purpose of the trajectory hierarchy for the Supervisory mode is to

provide a logical way of combining trajectory elements so that a

variety of trajectories can be formed from the elements in a systematic

way. The hierarchy is used for conducting module exchange, developing

displays, and for computer generation of trajectories. The hierarchy

uses some of the concepts of structured software development and thus

fits very well into a digital computer program. The four levels in the

hierarchy for the basic module exchanges are:

9-9
1) Total trajectories (4);

2) Trajectories (20);

3) Steps (9);

4) Actions (8).

Each level is constructed from elements in the levels below x^ith the

actions being the most basic level. A total trajectory is used for a

demonstration and an example is to start with the ETU at the rest

position, pick up a failed module from the spacecraft, put the failed

module in a temporary location in the stowage rack, move a good module

from the stowage rack to the spacecraft, move the failed module from

the temporary stowage rack location to the original good module

location in the stowage rack, and return to the rest position. A

trajectory is the path of motion and processes associated with moving

from one module location to another. An example is to move a module

from the stowage rack to the spacecraft. Steps are sequences of

actions that are used in different trajectories much as subroutines are

used in software programs. An example step is flipping a module upside

down so it may be readily inserted into an axial spacecraft location.

The actions are the basic elements and each is associated with a degree

of freedom. Three of the actions are the basic cylindrical coordinates

of the end effector, three are the attitudes of the end effector, and

the other two are the end effector jaw motion and the interface

mechanism drive motion. Every total trajectory is made up of a

sequence of actions. The other two levels were introduced to simplify

the construction process.

While the Supervisory control mode incorporates the trajectories into

the software, for the Manual-Augmented mode, the trajectories take the

form of printed lists. The same general format and trajectory

sequences are used for both modes. However, the Manual-Augmented

trajectory sequences have different forms of actions and explicitly

state the requirements for completing an action. Comments are also

provided to assist the operator in following the trajectory sequence.

9-10
A full description is provided in MCR-85-1312, Manual-Augmented

Trajectory Sequences (Basic Modules), Martin Marietta Aerospace,

Denver, CO, 80201, August 1985.

In addition to the functions described above, the associated functions

listed in Table 9.1-3 are provided to make the system easy to use. The

basic approach for selecting specifics is the use of numbered menus. A

selection is made by typing in the appropriate number and a carriage

return. Certain menus also provide instructions for the operator to do

things away from the keyboard, such as turn on servo power. Procedures

for trajectory initialization, start-up, and shut-down are provided in

the Software Requirements Manual.

The software also provides for interrogation of the A to D converters

for bringing hardware information into the program. Signals are sent

to the D to A converters to cause the hardware to move. Scale factors

and other geometrical data can be readily changed by accessing the

appropriate menu and changing the data. Recompiling is not necessary.

Module location data in the computer can be verified by calling up the

appropriate menu that will display the joint error signals. When the

end effector is moved to the appropriate location, then the displayed

joint error signals should go to zero. When the program is put into

hold, a menu appears that allows the operator to select a variety of

data groups for display. The operator can also elect to continue or to

abort, which will return the program to the mode selection menu.

Table 9.1-3 Associated Functions

Menus for Making Selections

Transferring Signals To and From Hardware

Trajectory Initialization

Start-up Procedure

Shut-down Procedure

Computer Hold

Demonstration Abort

Module Location Calibration

Data Display

Data Printout

9-11
9.1.1 System Requirements

A general block diagram of the Servicer Servo Drive Console and its

interface with the computer is shown in Figure 9.1-1. The software

requirements that pertain to the overall system operation and

capabilities are defined in the paragraphs below.

9.1.1.1 Mode Control - The mode control requirements are:

1) Computer modes shall be primarily controlled from the computer

terminal. The computer modes are operate and hold. It shall be

possible to start a run from the computer terminal as well as

temporarily "hold" the run, to continue the run after a "hold" and

to abort the run;

2) The run "hold" may be initiated by depressing the space bar on the

computer terminal or by operating the Computer Mode switch on the

Servicer Control Panel to the HOLD position;

3) The control modes are: 1) Supervisory with operator assistance; 2)

Supervisory without operator assistance; 3) Manual-Augmented; and

4) Manual-Direct. It shall be possible to select any one of these

four modes from a menu. When operating in any one mode, it shall

be possible to enter the "hold" computer mode to end operations in

one control mode and then revert to the Mode Selection Menu. When

the hardware is being used, the computer shall check that the

Servicer Control Panel MODE SELECT switch is in the proper

position. If the Servicer Control Panel MODE SELECT switch is not

in the correct position, then the computer shall remind the

operator, with an appropriate display, to put the switch in the

proper position;

4) The computer shall accept transferring from the Supervisory mode to

the Manual-Augmented mode and back;

9-12
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9-13
5) The end effector jaw and interface mechanism drive functions will

be computer controlled and displayed as an action in the

Supervisory mode, but will be manually controlled from the Servicer

Control Panel in the Manual-Augmented mode.

9.1.1.2 Sequencing - The sequencing requirements are:

1) The computer program shall be capable of providing a program "hold"

that suspends any action in progress until released. The hold will

be implemented to occur only at the end of a computation cycle;

2) Whenever the control mode has been changed from Supervisory,

Manual-Augmented, or Manual-Direct and then returned to the

Supervisory mode, the computer will check all joint positions for

agreement with the rest position. If they do not agree, the

computer will indicate this and prevent starting a trajectory

sequence until agreement is reached either by restarting the

program when hardware is not present or by repositioning the ETU

when hardware is present;

3) When hardware is not present, the joint angle integrators shall be

reset to the rest position whenever the IOSS computer program is

inititated.

9.1.1.3 Timing - A goal for the sampling or recycle frequency of the basic

program when hardware is present shall be 15 samples/sec. It is

pertinent to note that a single command in the coordinate reference

system used will result in multiple joints being commanded in a single

cycle of the computer program. When hardware is not present, a fast

integration alternative shall be used. The integration speed need not

be relateable to real time, but may be paced by the computer solution

speed.

9.1.1.4 General - The software shall be developed using structured

programming techniques. It shall be formatted so that it is not

9-14
necessary to recompile after changing constants, scale factors, or

trajectory end points. The software shall contain a file of all scale

factors.

Where the operator is requested to respond to a CO? or PR?, the

computer shall accept any of CO, PR, or Y and continue. Any other

response by the operator shall initiate a "hold".

9.1.1.5 Computer Operations - The software has been generated so that the

operator can select the specific items he is interested in. The

selection is performed using a set of hierarchical menus. The first

menu (IOSS Main Menu) appears when the IOSS program has been selected.

It takes the following form:

IOSS MAIN MENU

1. Run Setup Menu

2. Mode.Selection Menu

3. Module Data Collection Menu

4. Hardware Calibration Menu

5. Exit to MCR

The operator selects a menu by typing in a number and a carriage

return. The number 5 will end use of the IOSS program and return the

operator to the operating system for computer shutdown or selection of

ano ther pro gram.

The Run Setup Menu takes the following form:

1. Operator's Name - allows entry of up to 25 characters.

2. Run Title - allows entry of up to 50 characters to describe the run

to be made.

3. Run Number - allows entry of up to 10 characters.

4. Hardware Present? - allows the operator to enter a Y or N (yes or

no) to answer the question. The default value can be selected.

9-15
5. Save Constants to Disk? - allows the operator to select if he wants

to save changes in constants that are made, or just use the changed

values for one run.

6. Geometry Constants Menu - allows the operator to verify or change

the geometrical constants of the servicer mechanism as used in the

software. The configuration cannot be changed.

7. Limit Constants Menu - allows the operator to check or change 16

limit values used in the program.

8. A/D Scale Factors Menu - allows the operator to check or change 7

analog to digital scale factors.

9. D/A Scale Factors Menu - allows the operator to check or change 11

digital to analog scale factors.

10. Joint Position Threshold Menu - allows the operator to check or

change 12 joint error threshold values.

11. Simulated Hardware Characteristics Menu - allows the operator to

check or change 14 characteristics of the simulated hardware.

Examples are: time to latch, interface mechanism ready, and

specific initial values of the mechanism cylindrical coordinates.

The Mode Selection Menu contains the following items:

1. Unassisted Supervisory Mode

2. Assisted Supervisory Mode

3. Manual-Augmented Mode

4. Manual-Direct Mode

5. Return to IOSS Main Menu

The Module Data Collection Menu includes the following items:

1. Read a Spacecraft Location

2. Read a Stowage Rack Location

3. Read REST Position

4. Read Spacecraft Flip Location

5. Read Stowage Rack Iflip Location

6. Read Spacecraft Transition Location

7. Read Stowage Rack Inverse Transition Location

8. Return to IOSS Main Menu

9-16
For each of the specific locations, a designator is assigned (Numerical

and Set), mechanism joint angle data is collected and transformed to

cylindrical coordinates, and the data is filed for use in the

trajectory generation process. The cylindrical coordinate data is also

displayed for the operator.

The Hardware Calibration Menu includes the following items:

1. Check Spacecraft Location

2. Check Stowage Rack Location

3. Return to IOSS Main Menu

For each of the items 1 and 2, the operator enters the appropriate

designator (Numerical and Set). The computer displays the designator,

the corresponding cylindrical coordinates, and the differences between

desired and actual joint angles. '

9.1.2 Coordinate System Definition

The following paragraphs present a definition of the coordinate systems

used for the two modes - Supervisory and Manual-Augmented. A number of

terms, some variables, and some constants are used in describing the

requirements, specifically the equations. These terms are defined in

Section 11.0 of the Software Requirements Document. The physical

configuration of the ETU and accompanying mockups is shown in Figure

9.1-2 to aid in the discussion. The six joints and their alphabetical

designations are shown on the figure.

There are six degrees of freedom of the servicer mechanism. These six

can be combined into two groups of three each. One relates to the

position of the end effector, thinking of the end effector as a single

point. The second group relates to the attitude of the end effector,

now thinking of the end effector as a movable hand capable of rotating

about three different independent axes. These movements are

9-17
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9-18
independent of the position of the end effector, which may or may not

change at all while the attitude of the end effector is changed. The

definitions treat end effector "position" and "attitude" separately.

Also treated separately are the cylindrical coordinates from the

specific joint and arm coordinates of the ETU. This latter set is

referred to as ETU joint or gimbal coordinates

A right-handed cartesian reference coordinate system was defined with

respect to the stowage rack and spacecraft mockups. A cylindrical

coordinate system was defined with respect to the servicer mechanism

and its parameters are shown in Figure 9.1-3. The three positional

cylindrical coordinates are x (out of paper), r (radius), and Theta

(central angle). The three cylindrical coordinate attitude angles -

Psi, Phi, and Omega - are also shown on the figure. While the figure

appears to indicate that Psi and Omega are redundant, this is not

actually the case. For all operating siutations either Psi or Phi will

be 90 deg from the position shown and the redundancy is avoided.

X'

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paper)

Figure 9.1-3 Cylindrical Coordinates

The decision to emphasize axial and radial motion results in

significant simplification of the geometrical equations. For axial

motions, Psi is fixed at 0 deg and Phi is fixed at ^90 deg, depending

upon whether the servicer end effector is pointing towards the

9-19
spacecraft or towards the stowage rack. This means that relationships

only need be established for the remaining four degrees of freedom.

Similarly for radial motion, Psi is fixed at 90 deg and Omega is fixed

at -90 deg. Again geometric relationships for only four degrees of

freedom needed to be established. The equations for the full six

degrees of freedom have been derived, but they are much more complex.

This is especially true for the rate transformations. Also it is

necessary to use a more complex method, such as quaternions, for

keeping track of the solution quadrants for several of the angles.

Figure 9.1-4 shows the relationship for the "translation" gimbal, or

joint, angles of the servicer mechanism. These angles are:

1) Shoulder roll - T;

2) Shoulder pitch - U;

3) Elbow roll - V.

The servicer mechanism end effector joint angles are shown in Figure

9.1-2.

The servicer mechanism is designed so that the shoulder pitch joint

drives a parallelogram linkage that keeps the mechanism forearm (1«

of Figure 9.1-4) always parallel to the front face of the stowage

rack. This simplifies the geometrical transformations. However, the

geometrical transformations are complicated by the fact that the elbow

roll drive axis does not intersect the second parallelogram axis (Iand lg on Figure 9.1-4). Another parameter used.is the angle Delta,

which defines the relationship between the radius vector and the TV

camera reference axes. Delta is called the display term.

9-20
always in YZ plane

i cos U + St. 5

Figure 9.1-4 Servicer Mechanism Gimbal Coordinates

A constraint on system configuration is that the end effector wrist

, segment (1^ of Figure 9.1-4) must always be directed inboard towards

the docking axis so that the end effector can reach in close towards

the docking axis. The transfer between axial motion facing the stowage

rack and axial motion facing the spacecraft is called a module flip.

An analysis was conducted to identify the best location for the flip.

It is located in the +Y, -Z quadrant (see Figure 9.1-2). The transfer

between axial and radial motion is called a transition. An analysis

was conducted to identify the best sequence of actions and location for

the transition. The transition is conducted in front of the short end

of the spacecraft mockup (the +Y, +Z quadrant of Figure 9.1-2).

9.1.3 Supervisory Mode

The Supervisory mode of control is proposed as the normal mode of

operation. All servicer mechanism motions and trajectories are

determined beforehand. The trajectories are converted to sequences of

9-21
commands that are fed into the control laws. This results in the ETU

following the trajectory defined by the sequence of commands. A

combination of the trajectories being stored explicitly in the computer

and rules for constructing trajectories and specific step end point

data being stored in the computer were used in developing the software.

There are two forms of the Supervisory control mode — with operator

assistance and without operator assistance. A similar approach is used

for both forms in terms of trajectory definition and display format.

The unassisted mode is essentially automatic, while the operator must

input "GO" instructions before each action when in the assisted

Supervisory mode. The operator actions are explicitly requested by the

program through the terminal display. In the unassisted mode the

operator is only required to provide inputs during trajectory

initialization and system shutdown and to set some Servicer Control

Panel switches that involve safety.-

The unassisted mode does provide short pauses at certain points in each

trajectory to provide the operator time to make certain checks. An

audio tone sounds at the beginning of the pauses. The operator must

take an action to hold or abort the trajectory sequence.

In both the assisted and unassisted modes, the computer controls the

end effector and interface mechanism drives. The computer allows the

drive to operate until an end-of-travel signal is received. If a

pre-specified time expires before the end-of-travel signal is received,

the computer will hold the sequence and notify the operator. The

operator can initiate or remove a hold by operating the COMPUTER MODE

switch on the Servicer Control Panel or by computer terminal keyboard

operations.

The totality of the trajectory mode software is composed of two major

parts. One is the higher frequency computation of rate commands

necessary to drive an ETU joint or joints to a predefined position.

The other is a trajectory sequencing routine that establishes the

9-22
desired position the joints must be driven to. It does this in a

preestablished order of cylindrical coordinate motions to accomplish an

entire module exchange. The operation of the end effector and

interface mechanism drives are integrated into the sequence.

The trajectory sequence is based on the hierarchy described in Section

9.1.4. The hierarchy is a logical way of defining and combining

trajectory elements so that a variety of trajectories can be formed

from the elements in a systematic way. The hierarchy is used for

conducting module exchanges, developing displays, and for computer

generation of trajectories. The four levels in the basic module

exchange hierarchy are:

1) Total trajectories - 4 items;

2) Trajectories - 20 items;

3) Steps - 9 items;

4) Actions - 8 items.

Operations and displays for the two Supervisory control modes are

discussed in Section 9.1.5. The sequence of activities necessary to

perform a module exchange is described along with the specifics

necessary for generating the displays. Two representative displays and

the rules and guidelines for trajectory and display formatting are

presented in Section 9.1.6. The displays provide the operator with an

overview of the total trajectory and allow him to quickly identify

where he is, what is going on, what is next, and what activities

remain. By using the terminal's reverse video feature for element

highlighting, the operator is not burdened with excessive detail. The

only display differences between the assisted and unassisted modes are

that the assisted mode displays will include the terms CO? and PR? at

those places where operator assistance is required.

9-23
The generation of trajectory displays is addressed in Section 9.0 of

the Software Requirements Document. That section discusses how the

mockup equipment and the ETU are set up for collection of module

location data, the way in which trajectories are made specific, the

logic for generating trajectories and displays, and a method for

verifying that a module attach point location has not changed.

The software functions required to implement the Supervisory mode are

shown in block diagram form in Figure 9.1-5 with emphasis on

computation of the high frequency rate commands to drive the ETU

joints. This diagram does not represent the total software for

servicer control. A similar figure for the Manual-Augmented mode is

also necessary to show all modes. Some elements are, of course, common

between these two modes.

Computer control of the servicer ca-n generally be thought of as two

relatively independent software activities. One is the higher

frequency computation of rate commands necessary to drive a joint or

joints to a predefined position. The other is a trajectory sequencing

routine that establishes the desired position the joints must be driven

to. It does this in a preestablished order of cylindrical coordinate

motions to accomplish an entire module exchange. That routine is shown

as block 2 on the figure. This order and the associated displays on a

CRT terminal make up the bulk of the trajectory sequencing routine.

The basis for these desired sequences and displays are discussed in

more detail in Sections 7.0, 8.0 and 9.0 of the Software Requirements

Manual. The majority of Figure 9.1-5 is related to the joint drive

functions and shows the interrelationships of the sub-elements of joint

control.

9.1.4 Supervisory Trajectory Hierarchy

The purpose of the trajectory hierarchy for the Supervisory mode is to

provide a logical way of combining trajectory elements so that a

variety of trajectories can be formed from the elements in a systematic

way. The hierarchy is used for conducting module exchange, developing

9-24
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FOLOOUT

9-25

HOLDOUT FRAME
displays, and for computer generation of trajectories. The hierarchy

uses some of the concepts of structured software development and thus

fits very well into a digital computer program. The four levels in the

basic module exchange hierarchy are:

1) Total trajectories (4);

2) Trajectories (20);

3) Steps (9);

4) Actions (8).

Each level is constructed from elements in the level below with the

action being the most basic level. A total trajectory is used for a

demonstration and an example is to start with the ETU at the rest

position, pick up a failed module from the spacecraft, put the failed

module in a temporary location in the stowage rack, move a good module

from the stowage rack to the spacecraft, move the failed module from

the temporary stowage rack location to the original good module •

location in the stowage rack, and return to the rest position. A

trajectory is a portion of a total trajectory. An example is to move a

module from the stowage rack to the spacecraft. Steps are sequences of

actions that are used in different trajectories much as subroutines are

used in software programs. An example step is flipping a module upside

down so it may be readily inserted into an axial spacecraft location.

The actions are the basic elements and each is associated with a degree

of freedom. Three of the actions are the basic cylindrical coordinates

of the end effector, three are the attitudes of the end effector, and

the other two are the end effector jaw motion and the interface

mechanism drive motion. Every total trajectory is made up of a

sequence of actions. The other two levels were introduced to simplify

the construction process.

The same hierarchy is used for both the unassisted and operator

assisted forms of the Supervisory control mode. Trajectories are

9-26
expressed in cylindrical coordinates (see Section 9.1.2). In most

steps more than one action will be performed concurrently. The steps

have been defined so that at the end of most steps the end effector

jaws are open and there is at least three in. of separation distance

between the jaws and an interface mechanism or the interface mechanism

is free of its baseplate. This is so that if the trajectory is stopped

at the end of a step, a redundant load path will generally not exist.

If there is an unplanned motion while there is a load path, then

baseplate receptacles can be moved out of position and they would need

to be realigned before further demonstrations could be conducted.

The four levels of the hierarchy will be discussed from the bottom up

so that the method of conducting the higher levels is more meaningful.

9.1.4.1 Actions - Actions are the basic elements and each is associated with

a degree of freedom. Three of the actions are the basic cylindrical

coordinates of the end effector, three are the attitudes of the end

effector, and the other two are the end effector jaw motion and the

interface mechanism drive motion. Actions are the first level elements

and are used to form steps.

The following types of actions are used:

1) Drive x (distance along docking axis) to x command;

2) Drive r (distance from docking axis) to r command;

3) Drive Theta (central angle) to Theta command;

4) Drive Psi (end effector attitude angle about r vector) to Psi

command;

5) Drive Phi (end effector attitude angle about an axis perpendicular

to r and x, for axial configuration) to Phi command;

9-27
6) Drive Omega (end effector attitude angle about x axis, for axial

configuration) to Omega command;

7) Operate end effector jaw drive to open or close;

8) Operate interface mechanism drive to latch or unlatch.

Other operations that are used, but that are not actions are:

1) Set commands—is used to establish the values that functions should

have at the end of an action;

2) Set derivative to zero—is used to avoid driving one degree of

freedom (DOF) while another DOF goes through a range where the

first DOF equations change; e.g., in going from axial to radial;

3) Set derivative to normal—is used after the need for setting a

derivative to zero has passed;

4) Pause—is used to provide the operator with time to verify

something;

5) Operator instruction—is used to tell the operator to perform

certain functions such as setting a switch on the Servicer Control

Panel.

An action is complete when the error (difference between actual DOF and

DOF command) becomes less than a threshold. The angle thresholds are

+0.2 deg and the distance thresholds are +0.2 in. It is then

permissible to go to the next step.

Situations were identified where more than one action must be performed

at the same time, where it is permissible to combine actions, and where

constraints on combined actions should be applied. Each of these

situations have been defined and documented.

9-28
9.1.4.2 Steps - Steps are sequences of actions that are used in different

trajectories much as subroutines are used in software programs. An

example step is flipping a module upside down so it may be readily

inserted into an axial spacecraft location. Steps are constructed of

actions and are used to form trajectories.

The following types of steps are used:

1) PANDO - Position and orient end effector for next step or to rest

location;

2) ATTACH - Move in and attach end effector to module;

3) RELEAS - Release end effector from module and move back;

4) LATCH - Insert and latch module;

5) UNLAT - Unlatch and withdraw module;

6) FLIP - Flip module from stowage rack side to spacecraft

side;

7) IFLIP - Flip module from spacecraft side to stowage rack side;

8) TRANS - Transition module from axial motion with module facing

stowage rack to radial motion with module facing

spacecraft;

9) ITRANS - Transition module from radial motion with module facing

spacecraft to axial motion with module facing stowage rack,

The process of constructing each step type from the list of actions has

been defined. Also the initial and final conditions of each step were

defined so the need for intermediate steps could be identified when

steps were formed into trajectories.

9-29
An example of a step is the step ATTACH. Step ATTACH involves

attaching the end effector to the interface mechanism. The ATTACH step

starts with the end effector 3.0 in. away from the latch position in an

radial or axial direction with the other cylindrical coordinates equal

to the module location coordinates. For axial motions, the Omega

command is Z _+ Delta and for radial motion, the Phi command is Y +

Delta where Y and Z are module latched gimbal angles. The other

attitude angles are derived from the module latched coordinates. The

end effector jaws are open.

The following sequence of actions is used during the ATTACH steps:

1) Set commands

x = final value

r = final value

Theta = final value

Psi = final value

Phi = final value

Omega = final value;

2) Drive x, r, Theta, Psi, Phi, and Omega until errors are inside

threshold;

3) Verify end effector READY signal is present and jaws CLOSED signal

is not present;

4) Notify operator that end effector is ready, or is not ready, for

attach (display and audio);

5) Pause for 4 sec;

6) Operate end effector jaw drive until end effector jaw CLOSED signal

is present or for 15 sec (check that jaw CLOSED signal is present

for two successive computation cycles before accepting presence of

signal);

9-30
7) If CLOSED signal is not received after 15 sec, stop sequence and

notify operator.

At the completion of the ATTACH step, the end effector cylindrical

coordinates are equal to the module location coordinates, except that:

x = module location coordinte +0.5 in. for spacecraft axial

location;

x = module location coordinate -0.5 in. for stowage rack axial

location;

r = module location coordinate -0.5 in. for spacecraft radial

location.

For axial motions, Omega command will be Z + Delta and for radial

motion, Phi command will be Y + Delta where Y and Z are module latched

gimbal angles. The other attitude angles will be per the module

latched coordinates. The end effector jaws will be closed.

9.1.4.3 Trajectories - Trajectories are portions of total trajectories. An

example is to move a module from the stowage rack to the spacecraft.

Trajectories are constructed from steps and the actions of PANDO and

are used to construct total trajectories.

Each trajectory has a specific name. The names are constructed from

the following rules:

1) Each trajectory name will be made up of 5 characters;

2) The first character denotes whether a module is present or not.

E for no module, M for a module present;

3) The second and third characters define the starting point for the

trajectory:

9-31
RE for rest position

SX for spacecraft axial location

SR for spacecraft radial location

CK for stowage rack (always axial);

4) The fourth and fifth characters define the end point for the

trajectory using the same definitions as for the starting point.

Twenty trajectory types were identified to fully define the

possibilities. However, only fifteen are used for constructing the

total trajectories. The following types of trajectories are used:

1) ERESX - End effector only from rest position to spacecraft axial

offset location;

2) ERESR - End effector only from -rest position to spacecraft radial

offset position;

3) ERECK -"End effector only from rest position to stowage rack axial

offset location;

4) ESXRE - End effector only from spacecraft axial offset location to

rest position;

5) ESXSR - End effector only from spacecraft axial offset location to

spacecraft radial offset location;

6) ESXCK - End effector only from spacecraft axial offset location to

stowage rack axial offset location;

7) ESRRE - End effector only from spacecraft radial offset location to

rest position;

8) ESRSX - End effector only from spacecraft radial offset location to

spacecraft axial offset location;

9-32
9) ESRCK - End effector only from spacecraft radial offset location to

stowage rack axial offset location;

10) ECKRE - End effector only from stowage rack axial offset location

to rest position;

11) ECKSX - End effector only from stowage rack axial offset location

to spacecraft axial offset location;

12) ECKSR - End effector only from stowage rack axial offset location

to spacecraft radial offset location;

13) ECKCK - End effector only from one stowage rack axial offset

location to a second stowage rack axial offset location;

14) MSXSR - Move in from standoff location, attach end effector to

module at spacecraft axial location, unlatch and withdraw

module, move module to spacecraft radial location, insert

and latch module, release end effector, and move back to

standoff location;

15) MSXCK - Move in from standoff location, attach end effector to

module at spacecraft axial location, unlatch and withdraw

module, move module to stowage rack axial location, insert

and latch module, release end effector, and move back to

standoff location;

16) MSRSX - Move in from standoff location, attach end effector to

module at spacecraft radial location, unlatch and withdraw

module, move module to spacecraft axial location, insert

and latch module, release end effector, and move back to

standoff location;

9-33
17) MSRCK - Move in from standoff location, attach end effector to

module at spacecraft radial location, unlatch and withdraw

module, move module to stowage rack axial location, insert

and latch module, release end effector, and move back to

standoff location;

18) MCKSX - Move in from standoff location, attach end effector to

module at stowage rack axial location, unlatch and withdraw

module, move module to spacecraft axial location, insert

and latch module, release end effector, and move back to

standoff location;

19) MCKSR - Move in from standoff location, attach end effector to

module at stowage rack axial location, unlatch and withdraw

module, move module to spacecraft radial location, insert

and latch module, release end effector, and move back to

standoff location;

20) MCKCK - Move in from standoff location, attach end effector to

module at stowage rack axial location, unlatch and withdraw

module, move module to second stowage rack axial location,

insert and latch module, release end effector, and move

back to standoff location.

One of the simpler trajectories is used as an example. The MSXCK

trajectory involves moving a module from its spacecraft axial offset

location to the stowage rack axial offset location. The trajectory

initial conditions are the spacecraft axial offset location with the

end effector jaws open. The following sequence of steps was used:

1) ATTACH;

2) UNIAT;

9-34
3) PANDO to inverse flip initial conditions

- Set commands

x = -10 in.

r =8 2 in .

Theta = +45 deg

Psi = 0 deg

Phi = -90 deg

Omega = -78.3 deg

Drive x, r, Theta, Psi, Phi, and Omega until errors are inside

threshold;

4) IFLIP;

5) PANDO to latch initial conditions

- Set commands

x = latched module location + 30 in.

r = latched module location

Theta = latched module location

Psi = latched module location

Phi = latched module location

Omega = Z - Delta

Drive x, r, Theta, Psi, Phi, and Omega until errors are inside

threshold;

6) LATCH;

7) RELEAS.

The MSXCK trajectory final conditions are the stowage rack axial offset

location with the end effector jaws open.

9-35
9.1.4.4 Total Trajectories - Total trajectories are used for demonstrations.

An example is to start with the ETU at the rest position, pick up a

failed module from the spacecraft, put the failed module in a temporary

location in the stowage rack, move a good module from the stowage rack

to the spacecraft, move the failed module from the temporary stowage

rack location to the original good module location in the stowage rack,

and return to the rest position. Total trajectories are constructed

from trajectories. The following types of total trajectories are used:

1) Axial total trajectory;

2) Radial total trajectory;

3) Special total trajectory - axial to radial;

4) Special total trajectory - radial to axial.

The process of constructing total trajectories from trajectories was

defined. The process was particularly easy because of the "manner used

to define trajectories. It is not necessary to define specific initial

and final conditions as all total trajectories start and end at the

rest position.

The axial total trajectory involves the replacement of a failed module

that is initially in a spacecraft axial location. The sequence of

trajectories is:

ERESX - MSXCK - ECKCK - MCKSX - ESXCK - MCKCK - ECKRE.

The radial total trajectory involves the replacement of a failed module

that is initially in a spacecraft radial location. The sequence of

trajectories is:

ERESR- MSRCK - ECKCK - MCKSR - ESRCK - MCKCK - ECKRE.

9-36
Note that three trajectories are common between the axial and radial

total trajectories.

The special total trajectory - axial to radial - is used to move a

module from an axial spacecraft location to a radial spacecraft

location so that the radial rather than the axial total trajectory can

be demonstrated. The axial to radial special total trajectory is the

following sequence of trajectories:

ERESX - MSXSR - ESRRE.

The special total trajectory - radial to axial - is used to move a

module in the spacecraft mockup from the radial to the axial location

so that the axial rather than the radial total trajectory can be

demonstrated. The radial to axial special total trajectory is the

following sequence of trajectories:-

ERESR - MSRSX - ESXRE.

The following supplemental trajectories were not used in the total

trajectories listed above: ERECK, ESXSR, ESRSX, ECKSX, and ECKSR.

These trajectories were included so that a complete set are available

for use in potential future additions to the total trajectory list.

9.1.5 Supervisory Operations and Displays

The purpose of this section is to describe the sequence of activities

necessary to perform a module exchange so that the application of the

software requirements can be better understood and to describe the

specifics of the displays to be used during the unassisted and assisted

Supervisory modes. The sequence of activities is similar for both

modes except that the operator must make certain keyboard actions

before and after each action in the operator assisted mode.

9-37
The general order of performing a module exchange demonstration is;

1) Initialize the simulation;

2) Initialize the selected total trajectory;

3) Exchange the modules;

4) Shut the simulation down.

9.1.5.1 Simulation Initialization - The process of simulation initialization

is outlined in Table 9.1-4. It assumes that the system is completely

off to start with and this activity ends with the demonstration system

ready for initializing a total trajectory.

Table 9.1-4 Simulation Initialization

Step

No.

1.

2.

3.

4.

5.

6.

7.

8.

9.

10.

Action

Turn on computer and establish initial display

Request the IOSS program

Select Run Setup Menu if any run characteristics are to be

changed and enter specific data

Select Mode Selection Menu

Computer checks that Servicer Control Panel Mode Select

Switch is in the SUPERVISORY position

Start SSDC/ETU per Procedure II-l of MCR-78-531

(One-G Servicer System Operating Procedures)

Configure Servicer Control Panel (SCP) switches

for specific run to be made

Position ETU to rest position

Adjust TV camera

Initiate computer program start

The particulars for performing steps 1 through 4 were developed by the

programmer and are described in the Software User's Manual. Step 5.

involves the computer verifying that the Servicer Control Panel Mode

Select switch is in the SUPERVISORY position. If the proper signal is

9-38
not received, a display notifies the operator to reset the switch. The

program will not continue until the signal is received. Step 6.

follows the noted procedure and ends with power to the SSDC and camera

but with the servos off. For step 7. the Servicer Control Panel

switches should be set to a standby condition for the specific run to

be made (this is the rest position and Supervisory Mode). The SCP

switch positions are:

1) SHOULDER ROLL ANGLE SET pot - 468;

2) SHOULDER PITCH ANGLE SET pot - 500;

3) ELBOW ROLL ANGLE SET pot - 560;

4) WRIST YAW ANGLE SET pot - 497;

5) WRIST PITCH ANGLE SET pot - 730;

6) WRIST ROLL ANGLE SET pot - 169;

7) SHOULDER ROLL RATE LEVEL switch - MED;

8) SHOULDER PITCH RATE LEVEL switch - MED;

9) ELBOW ROLL RATE LEVEL switch - MED;

10) WRIST YAW RATE LEVEL switch - MED;

11) WRIST PITCH RATE LEVEL switch - MED;

12) WRIST ROLL RATE LEVEL switch - MED;

13) MODE SELECT switch - SUPERVISORY;

14) HAWK MODE switch - OFF;

9-39
15) COMPUTER MODE switch - HOLD;

16) MODULE LOCATION switch - S/C A;

17) INTERFACE MECH switch - OFF;

18) END EFFECTOR switch - OFF;

19) SERVO POWER switch - OFF (in).

The ETU is positioned to the rest position (step 8.) by pushing both

the SDP and SCP SERVO POWER switches to ON. Next the SCP DIRECT

CONTROL rocker switches are operated as necessary until the JOINT ERROR

lamps are OUT. The SCP SERVO POWER switch is pressed to OFF. The TV

focus should be adjusted in step 9. so that a black-on-white target

stays in focus from four ft to one ft distance from the TV camera

lens. In step 10. the computer program can be initiated using

procedures developed by the programmer and documented in the Software

User's Manual.

9.1.5.2 Total Trajectory Initialization - The process of total trajectory

initialization is outlined in Table 9.1-5. This activity starts with

the first menu displayed on the terminal screen and ends with the

trajectory being started in either the assisted or unassisted

Supervisory mode. This portion of the activity is totally computer

menu driven. The operator needs only to follow the prompts given and

to make selections appropriate to the demonstration he wants to

conduct. The computer also checks that certain initial conditions are

proper and notifies the operator if they are not suitable for the

selected operations.

9-40
Table 9.1-5 Total Trajectory Initialization

Step

No.

1.

2.

3.

4.

5.

6.

7.

8.

9.

10.

11.

Action

Select Unassisted or Assisted Supervisory Mode

The computer will display the Supervisory Mode Menu

Select the numerical and set designators for S/C A, S/C B,

S/R A, and S/R B

Use of the Module Placement Verification Sequence and its

associated menu will permit the operator to verify that the

module receptacle electrical connections are properly made

Turn servo power ON

Place SCP Computer Mode switch to OPERATE

Select total trajectory from four available

Computer checks that ETU is in rest position

Computer checks that modules are in proper position and that

the stowage rack temporary location is empty

Computer asks operator to verify the computer terminal type

in use

Computer starts trajectory sequence

There are two menu types of display to be used in the program. One has

the following format:

TITLE

1. Option 1

2. Option 2

3. Option 3

Enter Item Number:

To select a particular option, the number corresponding to it is typed

in and followed by a carriage return. The other main type of menu

display has the form.

9-41
TITLE

1. Text 1 Value 1

2. Text 2 Value 2

3. Text 3 Value 3

Enter Item Number:

This display format is used to display the current value of various

parameters and allows the user to change them if desired. To change a

particular item, the operator enters the appropriate line number on the

keyboard followed by a carriage return. A new line of text will then

appear telling the operator to enter the new data. The operator enters

the new numbers and presses the carriage return key. The data on the

appropriate line will then change. To continue on from this type of

display the line number associated with "Return to Prior Menu" plus a

carriage return is entered.

9.1.5.3 Shut-Down Procedure - This procedure can be entered at any time. It

should be used after an abort. The procedure is outlined in Table

9.1-6.

Table 9.1-6 Shut-Down Procedure

Step

No.

1.

2.

3.

4.

5.

Action

End module exchange by operating SCP Computer Mode switch to

HOLD or by operating space bar on computer terminal

Computer will cause a small auxiliary display to appear

Operator types ^in a 10 and carriage return to return to Mode

Selection Menu

Operator types in a 4 on Mode Selection Menu to put computer

in Manual-Direct mode

Select Manual-Direct mode on SCP

9-42
Table 9.1-6 Shut-Down Procedure (Cont.)

Step

No.

6.

7.

8.

9.

10.

11.

12.

13.

14.

15.

Action

Verify that ETU is in rest position

Configure Servo Drive Panel switches for shut down

Turn servo power off

Turn master power off

Operate space bar to end Manual-Direct mode on computer

Operate carriage return to return to Mode Selection Menu

Operate key 5 and carriage return to return to IOSS Main Menu

Operate key 5 and carriage return to exit from IOSS program

Log off by entering BYE and carriage return

Turn computer off

The small auxiliary display associated with the HOLD mode appears on

the right hand side of the main display. The format of this display is

as follows:

Display Parameters

1. Cyl. Cmd.

2. Jnt. Rate

3. Err. Mtr.

4. Jnt. Pos.

5. Cyl. Pos.

6. Cyl. Err.

7. Cyl. Rate

8. Jnt. Cmd.

9. Continue

10. Abort

This display can be used to call up the specific displays of interest

or it can be used to follow the shut-down procedure.

The hold function is provided as a method of stopping a module exchange

and then allowing it to continue later. A hold can be initiated by the

operator pressing the COMPUTER MODE switch on the Servicer Control

Panel to HOLD or by operating the space bar on the computer terminal.

9-43
It is possible to continue the program. This is effected by use of the

auxiliary menu described above by typing in a 9 and a carriage return.

The program will not continue until the Computer Mode switch on the

Servicer Control Panel is returned to OPERATE. The operator is

notified of a failure to return the Computer Mode switch to OPERATE.

An abort procedure should only be used in an emergency, when there is a

danger to personnel or equipment. Otherwise the hold and normal shut

down procedures should be used. An abort is initiated only by

operating either of the two SERVO POWER switches to OFF; one on the

Servicer Control Panel and one on the Servo Drive Panel. Operation of

the SERVO POWER switches to OFF causes a rapid discharge of the servo

compensation network capacitors that reduces the amplifier output to

zero very quickly and allows the brakes to engage. Use of any other

abort method may result in slow decay of the compensation network

capacitors and continuation of joint motion. This is particularly

hazardous if an unbalanced module is being supported by a backdrivable

wrist drive servo.

9.1.6 Trajectory Display Construction

A single display format is used. The details of the display depend on

total trajectory, step, etc. The approach is based on the use of the

hierarchy discussed in Section 9.1.4. The first line of the display is

the title of the total trajectory. The second line is a list of the

trajectories that make up the total trajectory where the current

trajectory is highlighted. The third line is a definition of the

current trajectory. The fourth line is a list of the steps in the

current trajectory with the current step highlighted. The fifth line

is a definition of the current step. Lines six through 24 are a list

of the actions and specific data involved in the current step with the

current action highlighted.

The displays for the assisted and unassisted modes are similar except

that the assisted mode will use the terms .CO? and PR? to prompt the

operator's assistance. The examples used below are for the assisted

mode. For the unassisted mode, the CO? and PR? terms are deleted.

9-44
For the computer terminal displays at MSFC the current operation

highlighting takes the form of reverse video. The main text is green

on black and the reverse video will be black on green. One blank space

before and one after is used to further emphasize the reverse video.

For this document, leading and trailing astericks are used to denote

highlighting. The display is limited to 80 characters in width and 24

lines.

9.1.6.1 Radial Total Trajectory Example - An example of a display for one

step of a radial total trajectory is shown in Table 9.1-7. The first

five lines are as discussed above, but the fourth line is abbreviated

as there is only one step in this trajectory. Note that the ESRCK

trajectory is highlighted in the second line and the PANDO step is

highlighted in the fourth line. Similarly the third action, set

commands and the specific commands, is highlighted. The actions are

numbered to make them stand out better.

Table 9.1-7 Example Radial Total Trajectory Display

RADIAL TOTAL TRAJECTORY

ERESR MSRGK ECKCK MCKSR *ESRCK* MCKCK ECKRE

MOVE END EFFECTOR FROM SPACECRAFT RADIAL (S/C B) TO STOWAGE RACK (S/R B)

*PANDO*

PANDO TO STOWAGE RACK OFFSET LOCATION

1) SET COMMANDS CO?

R = 70 IN.

PHI = 140.1 DEC.

2) DRIVE TO COMMANDS PR?

*3) SET COMMANDS CO?

X = -19 IN.

4)

5)

PSI = 0 DEG.*

DRIVE TO COMMANDS

SET COMMANDS

R

THETA =

PR?

CO?

PHI

OMEGA =

71 IN.

+86 DEG.

+90 DEG.

+86 DEG.

6) DRIVE TO COMMANDS PR?

9-45
When all of the actions on a display are complete, the computer steps

to the next display. For the last trajectory display, where the end

effector has been returned to the rest position, the computer returns

to the Supervisory Mode Menu. The operator then has a choice of

rerunning the same total trajectory, selecting a new total trajectory,

redesignating the modules to be moved, or returning to the Mode

Selection Menu or to higher level menus and shut down of the operations.

Certain additional features are used in the operator assisted mode.

When the new step is displayed, the 1) Set commands CO?, and the

specific commands are highlighted. When the operator has typed in CO

and carriage return, then 2) Drive to commands PR? is highlighted.

When the operator has typed in PR and carriage return, then the

computer commands the ETU to drive to the commanded coordinates. When

the actual and commanded coordinates are within the threshold, then the

next step is highlighted. Only one. action is highlighted at a time.

The computer accepts CO, PR, or Y as valid responses to CO? or PR?.

Any other response puts the system in hold and notifies the operator by

a text line near the bottom of the screen.

9.1.6.2 Axial Total Trajectory Example - An example of a display of one step

of an axial total trajectory is shown in Table 9.1-8. The first five

lines are as discussed above. Note that the MSXCK trajectory is

highlighted in the second line and the LATCH step is highlighted in the

fourth line. Similarly the third action, Drive to Commands, is

highlighted. When the operator has typed in PR and carriage return,

the computer causes the ETU to drive to the commanded coordinates.

When the actual and commanded coordinates are within the threshold and

when the computer has checked to see that a READY signal is available

and the LATCHED signal is not present for the stowage rack B location,

then the next step is highlighted. If these check conditions are not

satisfied, then the operator will be notified and the sequence put in

the hold mode.

9-46
Table 9.1-8 Example Axial Total Trajectory Display

AXIAL TOTAL TRAJECTORY

ERESX *MSXCK* ECKCK MCKSX ESXCK MCKCK ECKRE

MOVE MODULE FROM SPACECRAFT AXIAL (S/C A) TO STOWAGE RACK (S/R B)

ATTACH UNLAT PANDO IFLIP PANDO *LATCH* RELEAS

LATCH MODULE IN STOWAGE RACK

1) OPERATOR SET SCP MODULE LOCATION SWITCH TO S/C B CO?

2) SET COMMANDS CO?

X = -20.3 IN.

R = 71.0 IN.

THETA = +4 DEC

PSI - 0 DEC

PHI = +90 DEC

OMEGA = -86 DEG

*3) DRIVE TO COMMANDS . . PR?*

4) INTERFACE MECHANISM IS READY CO?

5) PAUSE PR?

6) SET COMMANDS CO?

X = -22 IN.

7) LATCH INTERFACE MECHANISM AND DRIVE X PR?

8) INTERFACE MECHANISM IS LATCHED CO?

In action 1) the operator is requested to set the Servicer Control

Panel MODULE LOCATION switch to the S/R B position. The module is to

be placed in that location. The SCP switch is used in a set of

computer-independent logic associated with warning the operator of

inadvertent end effector jaw operation. This action is required for

both the assisted and unassisted modes. The MODULE LOCATION switch

should be repositioned before every latching step so that it is

properly positioned for the next step, which is RELEAS. If the switch

is not properly set, then the buzzer will sound unnecessarily.

The pause of step 5) is not needed for the operator assisted mode, but

is included to maintain similarity with the unassisted mode where it is

needed. Step 8) is included to tell the operator that the LATCH signal

9-47
is present. For the operator unassisted mode, each action is

highlighted for at least 1 sec before moving on so that the operator

has time to read the highlighted lines.

9.1.7 Manual-Augmented Control Mode

The Manual-Augmented control mode has the operator doing most of the

mechanism control, as in the Manual-Direct mode, only using hand

controllers instead of panel switches for some coordinates. In

addition, the computer is in the loop to facilitate the direction of

motion of the mechanism and to provide coordination of its motion with

respect to the displays provided. The most useful role for this mode

is to perform unscheduled motions to previously identified targets of

opportunity.

The Manual-Augumented control mode uses written procedures to define

the trajectory sequences so trajectory sequences are not a part of the

software requirements. The Servicer Servo Drive Console incorporates a

Hawk mode as part of the Manual-Augumented control mode. The Hawk mode

switches the inputs to the Wrist Pitch and Wrist Roll error meters in

the Servicer Control Panel to signals from the computer. Thus the

computer must develop these signals. They basically combine the

display term with the Wrist Pitch and Yaw error signals. In the axial

mode this provides an easy way of lining the display coordinates up

with the radial and tangential components of the cylindrical

coordinates. In the radial mode the Hawk function can be used to

position the Wrist Pitch joint and decouple the effect of Wrist Pitch

motions from central angle, Theta, motions. The display term, Delta,

is also used as part of the transformation between the display

coordinates and the cylindrical coordinates.

The cylindrical and gimbal coordinate systems defined in Section 9.1.2

for the Supervisory mode are applicable for the Manual-Augmented mode

as well. One additional coordinate system is required, however, and it

relates to the visual display the man uses as a reference for

commanding the mechanism. Both the hand controller motion and the TV

9-48
are coordinated to this one common reference, which is centered in the

TV monitor and is parallel to the monitor face coordinates. The

Manual-Augmented control mode is shown in block diagram form in Figure

9.1-6. While several of the blocks on Figure 9.1-6 are similar to

those on the corresponding Supervisory mode diagram of Figure 9.1-5,

most are unique to the Manual-Augmented control mode.

The Manual-Augmented start-up procedure is similar to that for the

Supervisory mode as given in Section 9.1.5 and Table 9.1-4. The

operator selects item 3 from the Mode Selection Menu by depressing key

3 and then the carriage return. The computer will check that the

Servicer Control Panel Mode Select Switch is in the Manual-Augmented

position and will notify the operator if it is not. A clock will be

started and the Manual-Augmented display will be presented. The

display states that the Manual-Augmented Mode was entered by

on at . The operator can then move the hand controllers and

cause the ETU mechanism to move.

The Manual-Augmented mode can be ended at any time by pressing the

computer terminal space bar or by operating the Servicer Control Panel

Computer Mode switch to HOLD. This will cause the Manual-Augmented

Control Mode Hold Display to appear. The Hold Display shows various

data groups for the operator's information. From the Hold Display the

operator can choose to abort the sequence, which will cause the

following display to appear:

Manual-Augmented Mode ending at

Elapsed time is minutes.

Press RETURN for previous menu.

Operation of the carriage return key will return the program to the

Mode Selection Menu. Operation of key 5 and carriage return will

return the program to the IOSS Main Menu. Operation of key 5 and

carriage return will return the program to the computer operating

system. The computer can then be shut down. Shut down of the ETU can

be effected by following steps 5, 6, 7, 8 and 9 of Table 9.1-6.

9-49
ORIGINAL PAGE IS

OF POOR QUALITY

Dl.pUy (TV) co CT1. Coordluc. Tru.fon (Axi.l) [ 3

•»•> « M «) 1*1 i-. co. (z (*J «)

Dl.pl.7 (IV) to Cyl. Coordliuc. Trm.term (t^Ul)

r • »t_ co. (T » 8) - Y .la (r > «)

TV

0 * 1,) .1. U e.. f

. I « 0' •

. T > 0* - F«elo| Jto««|« be*

. V > »!• -

. W < »5' -

. ill Irntth. lo InchM

TfttM ID

6. Ill <otlu lo «.( Bd

r.tM la

Olrcctlr to t

Joint Drlv*

Loop la S*rvlc«r

Control El«cttonic*

Control «nd Coajniter Kode

To Error H*tar«/Ll{ht Driver*

on S«rvlc«r Control F«a«l

9-50

Figure 9.1-6 Analytical Block Diagram - Manual-Augmented Mode
The Manual-Augmented mode can be put into "hold" at any time by

operating the Computer Mode switch of the Servicer Control Panel to

HOLD, or by operating the Mode Select Switch on the Servicer Control

Panel to the MANUAL-DIRECT position. The following display will result:

Manual Augmented HOLD Display

Joint Rate T = U = V =

Error Meter Y = Z = Delta =

Joint Position T = U = V =

W = Y = Z =

Cylindrical Rates X = R =

Theta =

Translation Control XTV = YTV =

ZTV =

Radial Position =

1) Continue

2) Return to Mode Selection Menu

Enter Item Number

The choices are indicated by the numbers. However, the program will

not continue unless the Computer Mode switch of the Servicer Control

Panel is in OPERATE. While the Manual-Augmented mode is in hold, all

six joint derivatives will be set to zero.

9.1.8 Interface Definition

A complete definition of all signals between the computer and the

Servicer Servo Drive Console of the Engineering Test Unit was made. A

separate computer program, called TESTH, was written to assist in

checking out these interfaces. TESTH provides the following functions:

1) An ability to select and send commands on each of the digital to

analog channels;

9-51
2) An ability to read each status signal from the analog to digital

converter. These signals are displayed in octal notation. Many of

the signals are also converted to engineering units or true/false

statements, as appropriate;

3) An ability to command the servicer mechanism to move along a single

cylindrical coordinate at a time at a selectable rate. This

function is very useful during system calibration and for

positioning the end effector during module location data collection.

The signals from the computer that pass through the digital to analog

converter are:

1) Joint rate commands (6);

2) Error meter commands (6);

3) End effector jaw drive commands;

4) Interface mechanism drive commands.

The signals to the computer that pass through the analog to digital

converter are:

1) Joint position feedback (6);

2) Hand controller commands (3);

3) Rate level selection;

4) Computer mode selection;

5) Control mode selection;

6) Module receptacle status (4);

7) End effector jaw status.

9-52
9.2 SOFTWARE PREPARATION

The basic servicer control software was initially to be based on an

early set of software developed on a PDP-11/45 computer. It developed

that a significant part of the early software was in assembly language

and thus could not be readily transferred to the PDP-11/34. Similarly

the MSFC software used a set of standard Jacobeans for coordinate

system transformation and these were not suitable for the unique

servicer arm geometry with its offsets. The A/D and D/A addressing was

transferred to the new software. The new software was developed to

satisfy the requirements of the Software Requirements Document.

Software for the basic modules is called the IOSS program.

9.2.1 Software Development

A PDP-11/34 computer was available at Martin Marietta and it was used

for the software development to reduce the potential for error when

transferring to the MSFC PDP-11/34 machine. The Martin Marietta

machine used Version 3.2 of the RSX-11M operating system with a Version

4.0 Fortran 77 compiler. The MSFC machine initially had the same

operating system, but it was later converted to Version 4.0 of the

RSX-11M operating system and a Version 5.0 Fortran 77 compiler so that

separate operating system and applications programs disks could be

used. Any attempt to put all of the applications programs and the

operating system on one disk would have resulted in excessive

development time because of the small storage volume available.

Additionally, the MMS programs would not have fit in with the basic

programs. It was necessary to revise the software slightly so it would

operate satisfactorily for both versions of operating system and

compiler.

The software was written entirely in Fortran. The memory limtations of

the PDP-11/34 required that an overlay system be developed and used.

It uses almost all of the 32K of available random access memory. The

program contains almost 6000 lines of executable code that reside in 93

routines. The routines fit into a hierarchical structure and they are

completely cross referenced.

9-53
The first part of the software hierarchy involves setting up the

computer and peripherals and reading certain data from the disk. The

second part involves a series of menus that gives the operator the

ability to change things such as the A/D scale factors. The third part

provides for selection of a control mode and execution of that mode.

Execution of the Supervisory control mode is the largest part of the

program. It includes total trajectores, trajectories, steps, actions,

Supervisory motion loop, and display generation. Display generation

involves a larger amount of software than was originally expected. The

Manual-Augmented control functions are also included in this part of

the software.

The next two sections of the software include provisions for collection

and storage of module location data and for verification of the module

location data. The last section of the software has to do with program

shutdown.

The software structure involves source files, object files, task files,

command files, overlay descriptor files, data files, and the various

commands needed to compile, link, and execute. The software is broken

up into different source files because it is easier to edit small files

and it is easier to move the different object file names within the

task builder file to create useable overlays. There may be more than

one routine inside a source file. But the file name will reflect the

purpose of all the different routines. An example is SCRUTL.FTN. This

file contains several routines that deal with SCReen UTiLities.

Compilation of software is done on a file by file basis. So for each

source file, there will be one object file. A compilation can take

many source files and create one object file, but we don't do this

because of the overlay scheme. A simple compilation command such as

"F77 SCRUTL=SCRUTL" will only create the object file SCRUTL.OBJ and any

error messages will be listed to the terminal. A compilation command

file F77.CMD allows the user to compile all the IOSS program source.

It is invoked by typing F77@F77. It is stopped by entering control C

and typing ABO F77.

9-54

C ~
Task building involves a task build command file (IOSS.TKB) and a task

build overlay descriptor file (IOSS.ODL). The task build command file

is rather simple. But the overlay descriptor file is complex.

Significant rearranging was done to create working overlays and even

now they are very large (almost to the 32K limit). Unfortunately, not

much expansion is available. The task builder is started, using the

command files, by issuing the MCR command TKB @IOSS. This approach

assumes that the machine has a floating point processor.

The software was prepared in Denver using the available PDP-11/34.

Provisions are included for simulating various hardware functions such

as joint rate integration and hardware switch positions. Extensive

checks of trajectories, displays, inhibit logic, equations, and

auxiliary functions were made before the software was taken to MSFC.

The software was put onto magnetic tape for transfer to MSFC. It was

then read off the tape and put on a. disk at MSFC. Various tests and

calibrations of the hardware to software interfaces were made before

any attempt was made to drive the hardware. See Section 11.1 for a

discussion of the hardware demonstrations.

9.2.2 Software User's Manual

The Software User's Manual provides a means of understanding the

structure of the IOSS software. It explains the installation

procedure, how to run the software, and describes the software. The

delivered software is over six thousand lines of executable code in 93

routines. The document has a hierarachical breakdown of these routines

and they are completely cross-referenced.

The first chapter covers installation and execution of the IOSS

software. The second chapter is a small one that describes the

relationship between the different types of files. The third chapter

presents the highest level of source description. The fourth chapter

presents an intermediate level of source description in block diagram

form. The fifth chapter presents the most detail in the form of

listings of the individual subroutines.

9-55
9.2.2.1 Installation - This section covers the installation of the IOSS

software. The software is shipped on 1/2 in. magnetic tape in a BRU

(Backup and Restore Utility) format. The directory the software

resides in is [100,1]. The software is restored first. Once the

software is on the disk, it is compiled to create the object files

(*.OBJ). After the compilation, the IOSS task is built. The task

builder uses two files. The first is the command file, IOSS.CMD. The

second file is IOSS.ODL and describes the overlay structure. Once the

task build is complete, the presence of the IOSS.TSK and IOSS.MAP files

is checked. The file IOSS.TSK is the executable load module and the

file IOSS.MAP is a map of the load module.

9.2.2.2 Running the IOSS Software - The IOSS software can be run from any

directory with one restriction. The software has to be executed where

the data files exist. These data files all have suffixes that end in

.DAT and are ASCII files that can be edited. It is not recommended

that they be edited directly, rather there is a procedure described in

the Software User's Manual that allows them to be updated in a simpler

manner.

9.2.2.3 Starting the IOSS Program from MCR - The IOSS software is run from

the directory containing the data files. If the data files are not

present, errors will result and the program will exit smoothly to a

higher level. The IOSS program is started from the UFD [100,1], by

issuing the following MCR command:

RUN IOSS

The IOSS program then asks the following question:

Terminal Type (VT52, RCA, Simple):

The user responds with the type of terminal he is working with. The

default is selected by simply pressing RETURN and is currently the RCA

terminal type. The simple terminal is used only for creating

hardcopies of the various menus and is not capable of copying the

9-56
Supervisory or Manual displays. Then the mode of the terminal is

changed to the slave mode. This change is normally invisible to the

user. But if the program aborts before a normal completion, the

computer will not respond to a terminal input. Next, the various

constants are read from a disk file (CONSTS.DAT). The most recent

version (the one with the highest version number) is read from disk.

If this file does not exist in the current UFD, the program will

abort. The name of the file is printed to the screen before it is read,

The final step in the initialization process is to read the module

locations from the disk. Again, only the most recent versions are read

and the files need to be in the current UFD. The data files are named

SCI.DAT through SC6.DAT, SRI.DAT through SR9.DAT, REST.DAT, FLIP.DAT,

IFLIP.DAT, TRANS.DAT, and ITRANS.DAT. Also the name of each data file

is printed to the terminal before it is read. Once the data files are

finished being read, the screen is cleared immediately and the IOSS

Main Menu is displayed.

9.2.2.4 IOSS Main Menu - The IOSS Main Menu is the highest menu in the

hierarchy of menus and it looks like the following:

IOSS MAIN MENU

1. Run Setup Menu

2. Mode Selection Menu

3. Module Data Collection Menu

4. Hardware Calibration Menu

5. Exit to MCR

Enter Item Number:

Selection of Item 5 returns the user to the operating system. This is

done when the user is finished using the IOSS program.

9.2.2.5 Run Setup Menu - The Run Setup Menu is used primarily to set up a new

run. It is also used to access the various constants. The menu looks

like the following:

9-57
RUN SETUP MENU

1. Operator's Name John Doe

2. Run Title A Test Run

3. Run Number 1

4. Hardware Present? T

5. Save Constants to Disk

6. Geometry Contants Menu

7. Limit Constants Menu

8. A/D Scale Factors Menu

9. D/A Scale Factors Menu

10. Joint Position Threshold Menu

11. Simulated Hardware Characteristics Menu

12. Return to IOSS Main Menu

Enter Item Number:

The data in this menu or in the listed next level menus can be changed

by following the prompts that are provided. When an item is changed,

the screen is rewritten to show the new data.

9.2.2.6 Mode Selection Menu - This menu allows the user to select one of the

four operating modes. The Mode Selection menu is:

MODE SELECTION MENU

1. Unassisted Supervisory Mode

2. Assisted Supervisory Mode

3. Manual-Augmented Mode

4. Manual-Direct Mode

5. Return to IOSS Main Menu

Enter Item Number:

Selection of Item 1 calls up the Supervisory mode menu. From this

menu, a total trajectory will run unassisted, as explained in the

Supervisory unassisted mode section. Selection of Item 2 calls up the

9-58
Assisted Supervisory mode menu. From this menu, a total trajectory

will run only with assistance, as explained in the Assisted Supervisory

mode section.

Selection of Item 3 starts the Manual-Augmented mode. Several items

are displayed (operator's name, time, and date, and then the

Manual-Augmented mode equations are executed in a 15 Hz loop. The

operator can then drive the ETU using the hand controllers. The

software also checks for a keypress or for the computer mode switch to

leave the Manual-Augmented position. Once either of these actions

happen, the computer will print a display of pertinent parameters and

then give the user the option to continue or leave the Manual-Augmented

mode.

Selection of Item 4 displays the operator's name, the time, and date.

Then the software starts cycling in. a 15 Hz loop waiting for a keypress

or for the computer mode switch to leave the Manual-Direct position.

Once either of these actions happen, the computer will print the ending

and elapsed time and then exit. - Selection of Item 5 returns the user

to the previous menu.

9.2.2.7 Supervisory Mode - Selection of Item 1 from the Mode Selection Menu

calls up the Supervisory mode menu for unassisted operation. The

Unassisted Supervisory mode menu is:

UNASSISTED SUPERVISORY MODE MENU

1. S/C A is S/C 1, Set 1

2. S/C B is S/C 3, Set 1

3. S/R A is S/R 1, Set 1

4. S/R B is S/R 3, Set 1

5. Module Placement Verification Sequence

Turn Servo Power On

Put SCP Computer Mode Switch to Operate

6. Perform Axial Total Trajectory

7. Perform Radial Total Trajectory

9-59
8. Perform Axial to Radial Total Trajectory

9. Perform Radial to Axial Total Trajectory

10. Return to Mode Selection Menu

Enter Item Number:

The text between Items 5 and 6 informs the operator to turn the servo

power on and set the computer mode switch to operate. If the computer

mode switch is not in operate, an immediate hold will be generated when

a total trajectory starts.

The Assisted Supervisory Mode menu looks and acts like the unassisted

menu, except the title has assisted in it. The differences are very

evident when a total trajectory is running.

Selection of Items 6, 7, 8, or 9 start a total trajectory. For a

complete description of total trajectories, consult Chapters 7, 8, and

9 of the Servicer Simulation Software Requirements document. Various

checks are performed at the very beginning of a total trajectory.

These checks include one to make sure the arm is at rest and the others

look for the various signals from module locations that insure their

correct position.

If a total trajectory is being run assisted, then the operator has to

tell the computer to perform the next step. The computer will pause at

the beginning of each action and wait for a PR, a CO, or a Y to

continue with the action. If the user responds with something other

than PR, CO, or Y, the computer will generate a hold command with the

effects described below. Hold commands are generated in both assisted

and unassisted modes and can be generated by the operator in the

following ways:

1) Doesn't answer with a CO, PR, or Y to an assisted mode prompt;

2) Presses a key while a movement action is executing;

3) Moves the computer mode switch to HOLD while a movement action is

executing.

9-60
Hold commands can be generated by the computer in the following ways:

1) Jaw open or close exceeds 12 seconds;

2) Latch or unlatch exceeds 28 seconds;

3) Ready signals from latch mechanism and jaw not correct.

Once a hold command is generated, all joint rates are zeroed, the latch

mechanism is stopped and jaw movement is stopped. Then a hold menu is

displayed on the right side of the screen (leaving the Supervisory

display untouched). This hold menu is:

Display Parameters

1. Cyl. Cmd.

2. Jnt. Rate

3. Err. Mtr.

4. Jnt. Pos.

(The Supervisory Display 5. Cyl. Pos.

would appear here.) 6. Cyl. Err.

7. Cyl. Rate

8. Jnt. Cmd.

9. Continue

10. Abort

Item:

Hold Due to (A hold message would appear here.)

Selection of one of Items 1 through 8 will result in the display of

different sets of parameters and their respective values. After

displaying the different parameters and the respective values, the

computer will prompt the user to press return to get back to the hold

menu. Selection of Item 9.erases the hold display menu and continues

the total trajectory from where it stopped. Selection of Item 10 will

abort the total trajectory and leave the arm at its current position.

9-61
The operator will be returned to the Supervisory mode menu (the

previous one) and from there he probably will go to a manual mode and

drive the arm to the rest position.

9.2.2.8 Software Hierarchy List - A list of all the major software modules

and how they relate with one another is given. The hierarchy list is

interpreted as follows. Each new column to the right represents a

lower level of subroutine calls. For example, IOSS is the highest

level program, the main program. IOSS.FTN first calls TTYPE, then

SLAVE and so on. IOSS calls MODLOC, which in turn calls RDLOC. Also,

a short description of each module follows the module name. Not all

the modules are listed here. Some of the lowest levels modules are

used many times and would clutter the list. The IOSS software

hierarchy list is:

IOSS

TTYPE

SLAVE

RDCONS

MODLOC

RDLOC

SETUP

WRCONS

GEOMK

LIMITK

ADSFK

DASFK

TOL

SIMHCH

MODSEL

SUPERV

MPSEQ

TTRAJS

. IOSS Main Program

Terminal Type Selection

Set Terminal .as a Slave

Read the Constants from Disk

Get the Module Locations

Read a Module Location from Disk

Run Setup Menu

Write Constants to the Disk

Geometry Constants Menu

Limit Constants Menu

A/D Scale Factors Menu

D/A Scale Factors Menu

Threshold Menu

Simulated Hardware Characteristics

Mode Selection Menu

Supervisory Control Menu

Module Placement Sequence

Total Trajectories

9-62
MANAUG Manual-Augmented Menu

READH Read the Hardware

TVCRAT TV to Cylindrical Rates

CJRATE Cylindrical to Joint Rates

MALIM Manual-Augmented Limits

MADONE Manual-Augmented Done Function

DRIVER Drive the Hardware

MAHOLD Manual-Augmented Hold

MANDIR Manual Direct Mode

HDCOL Module Data Collection Menu

READH Read the Hardware

WRLOC Write a Module Location to Disk

CALIBR Module Calibration Menu

CHRRDY Character from Keyboard Ready

READH Read the Hardware

JNTPOS . Calculate the Joint Position

ERROR Calculate the Joint Errors

DRIVEH Drive the Hardware

CALDSP Calibrate Screen Display

NOSLAV Set Terminal to Noslave Mode

The above hierarchy is four levels deep. However, TTRAJS has several

levels beneath it. They are:

TTRAJS 4 Total Trajectories

TRAJS 20 Trajectories

STEPS 9 Steps

SUPMOT Supervisory Motion Loop

SUPMOT also has several levels beneath it and they are listed below:

SUPMOT Supervisory Motion Loop

CYLJNT Cylindrical to Joint Positions

READH Read the Hardware

JNTPOS Calculate the Joint Positions

9-63
ERROR

CNTINU

SWHOLD

CHRRDY

KEYHOLD

CHKLAT

CHKJAW

ZRORAT

RATGEN

CJRATE

DRIVER

Calculate the Joint Errors

Supervisory Continue Function

Mode Switch Hold

Character from Keyboard Ready

Keyboard Hold

Check Latching Mechanism

Check Jaw Mechanism

Zero Rates

Rate Generation

Cylindrical to Joint Rates

Drive the Hardware

9-64
10.0 SERVICER CONTROL SOFTWARE - MMS MODULES

The objective of this activity is to obtain a set of servicer control

software for MMS modules that will allow the smooth demonstration of

the exchange of MMS modules using the same three control modes as for

basic module exchange. The MMS module servicer control software

development was recorded in three documents: 1) MCR-85-1331; Servicer

Simulation Software Requirements (MMS Modules), Martin Marietta

Aerospace, Denver, CO, June 1985; 2) MCR-85-1360, MMS Module Servicer

Control Software User's Manual, Martin Marietta Aerospace, Denver, CO,

December 1985; and 3) MCR-85-1361, Manual-Augmented Trajectory

Sequences (MMS Modules), Martin Marietta Aerospace, Denver, CO,

January 1986.

There are many similarities between the two sets of software. They are

implemented on the same computer, the same terminal is used, they work

with the 1-g servicer demonstration system equipment, the same general

trajectory hierarchy is used, and the same general approaches are used

for software interfaces.

There are also differences between the basic and MMS software

programs. The MMS software characteristics include:

1) One total trajectory is used. The MMS modules are only exchanged

axially in the recommended approach. This reduces the number of

total trajectories and trajectories required;

2) The MMS module is fastened with two bolts instead of one. This

increases the number of steps involved and the number of module

latch indicators;

3) A Module Servicing Tool is involved. This increases the complexity

of some steps, increases the number of trajectories, increases the

number of drive functions (actions) that must be commanded and also

increases the number of status indicators required;

10-1
4) The number of stored sets of module location data has been reduced

to one as the extra stored sets have not been used with the basic

modules;

5) The module flip location was moved as discussed in Section 8.0;

6) No radial motions are required although all of the equations were

carried over as that was easier and solution time is not a

problem. However, the transition and Inverse transition steps have

been deleted;

7) A more inclusive number of checks are made before the trajectory is

started to minimize errors; •

8) A Module Servicing Tool storage location is needed. This increases

the number of trajectories and status indicators.

The software discussed in this section is for the demonstration of the

exchange of MMS type modules. MMS modules are approximately 48 in.

square and 20.5 in. deep and are fastened into the spacecraft or spare

module with two bolts. These bolts are driven by a 1-g version of a

Module Servicing Tool (see Section 3.2) that has been adapted to work

with the servicer mechanism end effector. The arrangement of the MMS

modules on the spacecraft mockup and in the stowage rack is discussed

in Section 8.0. Section 9.0 describes the software used for basic

module exchange and contains supplemental information that is necessary

to understand the MMS software.

The approach used to develop the MMS software was to copy the basic

software to a new disk, delete the routines .that are not needed, revise

basic routines as necessary, and add routines specific to the MMS

modules. A software requirements document for the MMS module exchange

was prepared first using the delete, revise, and add process. The

software was then developed according to the stated requirements. This

10-2
approach was very effective in reducing development time and made the

resulting program easier to use because of its similarity with the

basic module program.

A Software User's Manual was prepared and sent in draft form to MSFC

before the module exchange demonstrations. The module exchange

demonstrations included all three control modes. The MMS module

Software User's Manual is discussed in Section 10.2.2 and the MMS

module demonstrations are discussed in Section 11.2.

The major elements and areas involved in the MMS servicing

demonstrations are similar to those shown in Figure 9-2 except that the

MST, a set of MST controls, a Connector Positioner, and the MST storage

rack have been added. These elements are described in Chapter 8.0.

10.1 SOFTWARE REQUIREMENTS

The software components, or functions, for MMS module exchange are

similar to those for basic module exchange as listed in Table 9.1-1.

The same geometrical approach is used for MMS and basic module exchange

as listed in Table 9.1-2. It was originally thought that MMS module

exchange could be effected by changing one cylindrical coordinate at a

time because of the way the MMS spacecraft is configured with respect

to the servicer by use of the docking adapter orientation joint (see

Setion 3.0), and that the MST could be inserted into, or withdrawn

from, the module by a purely axial motion. These statements are true

when both bolts are loose or tight. However, when the top bolt is

tight and the bottom bolt is loose, then the MMS module is tipped

enough that pure axial motion is not adequate. The solution to this

problem is addressed in Section 11.2.3. The same coordinate transfer

motions and control loop elements are used for the basic and MMS

software.

The Supervisory control mode trajectory hierarchy has the same purpose

and rationale for both MMS and basic module exchanges. However, the

four levels of the hierarchy are different in detail and are as follows:

10-3
1) Total trajectories (1);

2) Trajectories (9);

3) Steps (13);

4) Actions (10).

Definitions for each level are given in Section 9.1 along with how the

hierarchy is used in the Manual-Augmented control mode operations. The

MMS software involves only one total trajectory as no radial module

location is used. The number of trajectories has been reduced because

of the fewer total trajectories even though some new trajectories were

added. Four new steps have been added because of the two bolts per

module situation. Two actions have been added to operate the Connector

Positioner and the MST bolt drive. .

The associated functions listed in Table 9.1-3 for the basic modules

are also used for the MMS modules. The amount of checking performed

before certain actions, especially the start from rest, has been

increased to reduce the number of potential trajectory interruptions.

The same form of display construction is used for both MMS and basic

module exchanges to ease the operator work load when transferring

between module types.

The coordinate system defined for the basic modules in Section 9.1.2 is

also used for the MMS modules. However, the MMS module flip location

is off the short end of the spacecraft mockup as discussed in Section

8.0.

The Supervisory control mode is similar for both of the two forms of

module exchange except that for MMS modules the computer controls the

connector positioner, and the MST latch and bolt drives along with the

end effector. The interface mechanism drive is not used during MMS

module exchanges. Also the trajectory hierarchy is different between

the two modes as explained above. The equations and transformations

shown in Figure 9.1-5 are also used for both module forms.

10-4
Trajectory display construction is the same for both types of modules.

The specific displays are different because of the different

trajectories involved, but this does not affect the display

construction methodology. The MMS displays do require the use of two

lines to list all of the steps for those trajectories involving module

transport.

The same Manual-Augmented control mode software is used for both module

forms. The written trajectory sequences that the operator follows

during a demonstration are different. Also the operator will control

the Connector Positioner and the Module Servicing Tool directly from

their control panels.

10.1.1 System Requirements

A general block diagram of the Servicer Servo Drive Console and its

interface with the computer is shown in Figure 9.1-1. The software

requirements that pertain to the overall system operation and

capabilities are defined in the paragraphs below.

10.1.1.1 Mode Control - The mode control requirements are:

1) Computer modes shall be primarily controlled from the computer

terminal. The computer modes are operate and hold. It shall be

possible to start a run from the computer terminal as well as

temporarily "hold" the run, to continue the run after a "hold" and

to abort the run;

2) The run "hold" may be initiated by depressing the space bar on the

computer terminal or by operating the Computer Mode switch on the

Servicer Control Panel to the HOLD position;

3) The control modes are: (1) Supervisory with operator assistance,

(2) Supervisory without operator assistance, (3) Manual-Augmented,

and (4) Manual-Direct. It shall be possible to select any one of

these four modes from a menu. When operating in any one mode, it

10-5
shall be possible to enter the "hold" computer mode to end

operations in that control mode and then revert to the Mode

Selection Menu. When the hardware is being used, the computer

shall check that the Servicer Control Panel MODE SELECT switch is

in the proper position. If the Servicer Control Panel MODE SELECT

switch is not in the correct position, then the computer shall

remind the operator, with an appropriate display, to put the switch

in the proper position;

4) The computer shall accept transferring from the Supervisory mode to

the Manual-Augmented mode and back;

5) The end effector jaw drive function will be computer controlled and

displayed as an action in the Supervisory mode, but will be

manually controlled from the Servicer Control Panel in the

Manual-Augmented mode;

6) The Module Servicing Tool latch drive and bolt drive functions will

be computer controlled and displayed as an action in the

Supervisory mode, but will be manually controlled from the MST

Control Panel in the Manual-Augmented mode;

7) The ETU Connector Positioner drive function will be computer

controlled and displayed as an action in the Supervisory mode, but

will be manually controlled from the Servicer Control Panel in the

Manual-Augmented control mode.

10.1.1.2 Sequencing - The sequencing requirements are:

1) The computer program shall be capable of providing a program "hold"

that suspends any action in progress until released. The hold will

be implemented to occur only at the end of a computation cycle;

2) Whenever the control mode has been changed from Supervisory,

Manual-Augmented, or Manual-Direct and then returned to the

Supervisory mode, the computer will check all joint positions for

10-6
agreement with the rest position. If they do not agree, the

computer will indicate this and prevent starting a trajectory

sequence until agreement is reached either by restarting the

program when hardware is not present or by repositioning the ETU

when hardware is present;

3) When hardware is not present, the joint angle integrators shall be

reset to the rest position whenever the 10SS computer program is

inititated;

4) When the hardware is present, the digital to analog converter

outputs will be commanded to zero by the computer before the

Servicer Servo Drive Console is turned on.

10.1.1.3 Timing - A goal for the sampling or recycle frequency of the basic

program when hardware is present shall be 15 samples/sec. It is

pertinent to note that a single command in the coordinate reference

system used will result in multiple joints being commanded in a single

cycle of the computer program. When hardware is not present, a fast

integration alternative shall be used. The integration speed need not

be relateable to real time, but may be paced by the computer solution

speed.

10.1.1.4 General - The software shall be developed using structured

programming techniques. It shall be formatted so that it is not

necessary to recompile after changing constants, scale factors, or

trajectory end points. The software shall contain a file of all scale

factors.

Where the operator is requested to respond to a CO? or PR?, the

computer shall accept any of CO, PR, or Y and continue. Any other

response by the operator shall initiate a "hold".

10.1.1.5 Computer Operations - When the IOSS-MMS program is selected from the

POP 11/34 operating system by typing RUN IOSSMM and then a carriage

return, it will ask for the terminal type being used. The terminal

type used at MSFC is the RCA APT 4801 and the operator should respond

10-7
by typing in RCA followed by a carriage return.

In order to properly run a MMS module exchange demonstration, the MMS

module status indicators must be connected into the system. This is

accomplished by connecting P2 to J2A on the MMS junction box. As the

IOSSMM program is started, the following prequisite action appears:

Verify that P2 is connected to J2A on the MMS junction box.

Press RETURN when complete.

Because the MMS module exchange uses a different type of module, it has

its own set of module status indicators. For proper operation, the

prerequisite instructions shown must be followed.

The software has been generated so that the operator can select the

specific items he is interested in. The selection is performed using a

set of hierarchical menus. The first menu (IOSS Main Menu - MMS)

appears when the IOSSMM program has been selected. It takes the

following form:

IOSS MAIN MENU - MMS

1. Run Setup Menu

2. Mode Selection Menu

3. Module Data Collection Menu

4. Hardware Calibration Menu

5. Exit to MCR

The operator selects a menu by typing in a number and a carriage

return. The number 5 ends use of the IOSSMM program and returns the

operator to the operating system for computer shutdown or selection of

another program. The codes used for selection of each of the four IOSS

programs are different and easy to remember. The four IOSS programs

are:

1) IOSS - Basic Modules;

10-8
2) TESTH - Basic Modules;

3) IOSSMM - MMS Modules;

4) TESTHM - MMS Modules.

The Run Setup Menu takes the following form:

1. Operator's Name - allows entry of up to 25 characters.

2. Run Title - allows entry of up to 50 characters to describe the run

to be made.

3. Run Number - allows entry of up to 10 characters.

4. Hardware Present? - allows the operator to enter a T or F (true or

false) to answer the question. The default value can be selected.

5. Save Constants to Disk? - allows the operator to select if he wants

to save changes in constants that are made, or just use the changed

values for one run.

6. Geometry Constants Menu - allows the operator to verify or change

the geometrical constants of the servicer mechanism as used in the

software. The configuration cannot be changed.

7. Limit Constants Menu - allows the operator to check or change 16

limit values used in the program.

8. A/D Scale Factors Menu - allows the operator to check or change 20

analog to digital scale factors.

9. D/A Scale Factors Menu - allows the operator to check or change 15

digital to analog scale factors.

10. Joint Position Threshold Menu - allows the operator to check or

change 12 joint error threshold values.

11. Simulated Hardware Characteristics Menu - allows the operator to

check or change 30 characteristics of the simulated hardware.

Examples are: time to latch, interface mechanism ready, and

specific initial values of the mechanism cylindrical coordinates.

The data in this menu can be changed to verify the operability of

certain checks being made by the software.

10-9
The Mode Selection Menu contains the following items:

1. Unassisted Supervisory Mode

2. Assisted Supervisory Mode

3. Manual-Augmented Mode

4. Manual-Direct Mode

5. Return to IOSS Main Menu - MMS

The Module Data Collection Menu includes the following items:

1. Read a Spacecraft Top Bolt Location

2. Read a Spacecraft Bottom Bolt Location

3. Read a Stowage Rack Top Bolt Location

4. Read a Stowage Rack Bottom Bolt Location

5. Read a Spacecraft Top Bolt Offset Location

6. Read a Spacecraft Bottom Bolt Offset Location

7. Read a Stowage Rack Top Bolt Offset Location

8. Read a Stowage Rack Bottom Bolt Offset Location

9. Read a MST Storage Location

10. Read REST Position

11. Read Spacecraft Flip Location

12. Read Stowage Rack Iflip Location

13. Return to IOSS Main Menu - MMS

For each of the specific locations, angle data is collected and

transformed to cylindrical coordinates, and the data is filed for use

in the trajectory generation process. The cylindrical coordinate data

is also displayed for the operator. Items 1 through 4 of the Module

Data Collection Menu represent module bolt locations corresponding to

both module bolts tightened. Items 5 through 8 represent offset

locations used in the pseudo-combined motion technique described in

Section 11.2.2. These offset locations correspond to the module bottom

bolt loose and the module top bolt tight, with the socket of the MST

1/4 in. into the guiding cone and tipped to match the guiding cone

angle.

10-10
The Hardware Calibration Menu includes the following items:

1. Check Spacecraft Top Bolt Location

2. Check Spacecraft Bottom Bolt Location

3. Check Stowage Rack Top Bolt Location

4. Check Stowage Rack Bottom Bolt Location

5. Check MST Storage Location

6. Return to IOSS Main Menu - MMS

For each of the items 1 through 5, the computer displays the

corresponding cylindrical coordinates, and the differences between

desired and actual joint angles.

10.1.2 Supervisory Trajectory Hierarchy

The purpose of the trajectory hierarchy for the Supervisory mode is to

provide a logical way of combining trajectory elements so that a

variety of trajectories can be formed from the elements in a systematic

way. The hierarchy is used for conducting module exchange, developing

displays, and for computer generation of trajectories. The hierarchy

uses some of the concepts of structured software development and thus

fits very well into a digital computer program. The four levels in the

hierarchy are:

1) Total trajectories (1);

2) Trajectories (9);

3) Steps (13);

4) Actions (10).

Each level is constructed from elements in the level below with the

action being the most basic level. A total trajectory is used for a

demonstration and an example is to start with the ETU at the rest

position, pick up the MST, pick up a failed module from the spacecraft,

10-11
put the failed module in a temporary location in the stowage rack, move

a good module from the stowage rack to the spacecraft, move the failed

module from the temporary stowage rack location to the original good

module location in the stowage rack, return the MST to its storage

rack, and return to the rest position. A trajectory is a portion of a

total trajectory. An example is to move a module from the stowage rack

to the spacecraft. Steps are sequences of actions that are used in

different trajectories much as subroutines are used in software

programs. An example step is flipping a module upside down so it may

be readily inserted into an axial spacecraft location. The actions are

the basic elements and each is associated with a degree of freedom.

Three of the actions are the basic cylindrical coordinates of the end

effector, three are the attitudes of the end effector, and the other

three are the end effector jaw motion and the MST latch and bolt drive

motions. Every total trajectory is made up of a sequence of actions.

The other two levels were introduced to simplify the construction

process.

The same hierarchy is used for both the unassisted and operator

assisted forms of the Supervisory control mode. Trajectories are

expressed in cylindrical coordinates (see Section 9.1.2). In most

steps more than one action will be performed concurrently. The

trajectories have been defined so that at the end of most trajectories

the end effector jaws are open and there is at least three in. of

separation distance between the jaws and the MST attach fitting, the

MST is free of the module, or the module is free of its baseplate.

This is so that if the total trajectory is stopped at the end of a

trajectory, a redundant load path will generally not exist. If there

is an unplanned motion while there is a load path, then the baseplate

receptacles could be moved out of position and they would need to be

realigned before further demonstrations could be conducted.

The four levels of the hierarchy will be discussed from the bottom up

so that the method of constructing the higher levels is more meaningful.

10.1.2.1 Actions - Actions are the basic elements and each is associated with

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a degree of freedom. Three of the actions are the basic cylindrical

coordinates of the end effector, three are the attitudes of the end

effector, and the other four are the end effector jaw motion, the

connector positioner motion, and the MST latch and bolt drive motions.

Actions are the first level elements and are used to form steps.

The following types of actions are used:

1) Drive x (distance along docking axis) to x command;

2) Drive r (distance from docking axis) to r command;

3) Drive Theta (central angle) to Theta command;

4) Drive Psi (end effector attitude angle about r vector) to Psi

command;

5) Drive Phi (end effector attitude angle about an axis perpendicular

to r and x, for axial configuration) to Phi command;

6) Drive Omega (end effector attitude angle about x axis, for axial

configuration) to Omega command;

7) Operate end effector jaw drive to open or close;

8) Operate MST latch drive to latch or unlatch;

9) Operate MST bolt drive to tighten or loosen;

10) Operate connector positioner to connect or disconnect.

Other operations that are used, but that are not actions are:

1) Set commands—is used to establish the values that functions

should have at the end of an action;

10-13
2) Set derivative to zero—is used to avoid driving one degree of

freedom (DOF) while another DOF goes through a range where the

first DOF equations change; e.g., in going from axial facing the

spacecraft to axial facing the stowage rack;

3) Set derivative to normal—is used after the need for setting a

derivative to zero has passed;

4) Pause—is used to provide the operator with time to verify

something;

5) Operator instruction—is used to tell the operator to perform

certain functions such as setting a switch on the Servicer Control

Panel.

An action is complete when the error (difference between actual DOF and

DOF command) becomes less than a threshold. The angle thresholds are

+0.2 deg and the distance thresholds are +0.2 in. It is then

permissible to go to the next step.

Situations were defined where more than one action must be performed at

the same time, where it is permissible to combine actions, and where

constraints on combined actions should be applied. Each of these

situations has been defined and documented.

10.1.2.2 Steps - Steps are sequences of actions that are used in different

trajectories much as subroutines are used in software programs. An

example step is flipping a module upside down so it may be readily

inserted into an axial spacecraft location. Steps are constructed of

actions and are used to form trajectories.

The following types of steps are used:

1) PANDO - Position and orient end effector for next step or to rest

location;

2) ATTACH - Move in and attach end effector to MST;

10-14
3) RELEAS - Release end effector from MST and move back;

4) LATCH - Insert MST and latch MST to module;

5) UNLAT - Unlatch MST and withdraw it from module;

6) FLIP - Flip module or MST from stowage rack side to spacecraft

side;

7) IFLIP - Flip module or MST from spacecraft side to stowage rack

side;

8) FASTN - Use MST to tighten module bolt;

9) UNFAS - Use MST to loosen module bolt;

10) TOTBO - Transfer MST from module top bolt to module bottom bolt;

11) BOTTO - Transfer MST from module bottom bolt to module top bolt;

12) MOVIN - Move module in from standoff location to top bolt ready

for tightening position;

13) MOVBK - Move module back from top bolt loose (unfastened)

position to standoff location.

The process of constructing each step type from the list of actions has

been defined. Also the initial and final conditions of each step were

defined so the need for intermediate steps could be identified when

steps were formed into trajectories. An example of a step is given in

Section 9.1.4.2.

10.1.2.3 Trajectories - Trajectories are portions of total trajectories. An

example is to move a module from the stowage rack to the spacecraft.

Trajectories are constructed from steps and the actions of PANDO and

are used to construct total trajectories.

10-15
Each trajectory has a specific name. The names are constructed from

the following rules:

1) Each trajectory name will be made up of 5 characters;

2) The first character denotes whether a module or the MST is present

or not. E for no module or MST, M for a module and MST present,

and T for MST present;

3) The second and third characters define the starting point for the

trajectory:

RE'for rest position

SX for spacecraft axial location

TS for MST storage location

CK for stowage rack (always axial);

4) The fourth and fifth characters define the end point for the

trajectory using the same definitions as for the starting point.

Twenty trajectory types were identified in the basic module exchange

work to fully define the possibilities. However, only nine are used

for constructing the total MMS "trajectory and they are all different

from the basic module exchange trajectories. The following types of

trajectories (listed in alphabetical order) are used:

1) ERETS - Move end effector only from rest position to MST storage

rack axial offset location;

2) ETSRE - Move end effector only from MST storage rack axial offset

location to rest position;

3) MCKCK - Move MST in from standoff location, latch MST to module at

stowage rack axial location, unfasten both bolts and move

module back. Move module to second stowage rack axial

location, move module in, fasten both bolts, unlatch MST,

and -move MST back to standoff location;

10-16
4) MCKSX - Move MST in from standoff location, latch MST to module at

stowage rack axial location, unfasten both bolts, move

module back, and flip. Move module to spacecraft axial

location, move module in, fasten both bolts, unlatch MST,

and move MST back to standoff location;

5) MSXCK - Move MST in from standoff location, latch MST to module at

spacecraft axial location, unfasten both bolts, move module

back, inverse flip MST and module, move module to stowage

rack axial location, move module in, fasten both bolts,

unlatch MST, and move MST back to standoff location;

6) TCKCK - Move MST from one stowage rack axial offset location to a

second stowage rack axial offset location;

7) TCKTS - Move MST from stowage rack axial offset location to MST

storage rack axial offset location;

8) TSXCK - Move MST from spacecraft axial offset location to stowage

rack axial offset location;

9) TTSSX - Move MST from MST storage rack axial offset location to

spacecraft axial offset location.

An example trajectory is given in Section 9.1.4.3.

10.1.2.4 Total Trajectories - Total trajectories are used for

demonstrations. An example for a basic module is to start with the ETU

at the rest position, pick up a failed module from the spacecraft, put

the failed module in a temporary location in the stowage rack, move a

good module from the stowage rack to the spacecraft, move the failed

module from the temporary stowage rack location to the original good

module location in the stowage rack, and return to the rest position.

Total trajectories are constructed from trajectories. The following

type of total trajectory is used for MMS module exchange:

1) Axial total trajectory.

10-17
The process of constructing total trajectories from trajectories was

defined. The process is particularly easy because of the manner used

to define trajectories. It is not necessary to define specific initial

and final conditions as all total trajectories start and end at the

rest position.

The MMS axial total trajectory involves the replacement of a failed

module that is initially in a spacecraft axial location. The sequence

of trajectories is:

ERETS - TTSSX - MSXCK - TCKCK - MCKSX - TSXCK - MCKCK - TCKTS - ETSRE.

10.1.3 Supervisory Operations and Displays

The Supervisory operations and displays for MMS module exchange follow

the same pattern as for basic module exchange as described in Section

9.1.5. Detail differences are identified in this section.

The general order of performing a module exchange demonstration is:

1) Initialize the simulation;

2) Initialize the selected total trajectory;

3) Exchange the modules;

4) Shut the simulation down.

The MMS simulation initialization is similar to the basic module

exchange initialization of Section 9.1.5.1. However, the IOSSMM

program is called, the computer sets the D/A converter commands for

zero output, and the MST initial conditions are also checked in the MMS

procedure. The SCP panel switch positions are similar for both types

of demonstrations.

10-18
It Is necessary to check that the MST is ready for use. Checks are

made for the MST being installed in its storage rack and that the SCP

MST Storage Status PRESENT lamp is ON. For the MST Electronics Box, it

is verified that:

1) POWER switch is ON;

2) POWER LAMP is ON;

3) 28V LAMP is ON;

4) 10V LAMP is ON;

5) 0-10V LAMP is ON;

6) 0-32V LAMP is ON.

For the MST Control Panel, it is verified that:

1) POWER switch is ON;

2) POWER LAMP is ON;

3) All eight LED's come ON when LAMP TEST button is pressed;

4) FUNCTION SELECT switch is OFF;

5) FUNCTION EXECUTE switch is OFF;

6) READY TO LATCH LED is ON;

7) UNLATCHED LED is OFF;

8) LATCHED LED is ON;

9) TURNS INDEX LED may be ON or OFF.

10-19
The process of total trajectory initialization is somewhat different

for MMS module exchange and is outlined in Table 10.1-1. As with the

basic module software, this part of the activity is computer menu

driven. The operator needs only to follow the prompts given and to

make selections appropriate to the demonstrations he wants to conduct.

However, the MMS software makes a more extensive set of checks as

indicated in Step 8. of the table.

Table 10.1-1 Total Trajectory Initialization

Step

No. Action

1.

2.

3.

4.

5.

6.

7.

8.

9.

Select Unassisted or Assisted Supervisory Mode

The computer will display the Supervisory Mode Menu

Use of the Module Location Switch Correspondence and its associated

menu will permit the operator to remind himself of the

correspondence between the SCP Module Location switch positions and

the module bolt designations

Turn servo power ON

Place SCP Computer Mode switch to OPERATE

Tell program to perform axial total trajectory

Computer checks that ETU is in rest position

Computer checks that modules are in their proper positions, that the

stowage rack temporary location is empty, that the end effector jaws

are open, that the connector positioner is in the disconnected

position, that the MST latches are latched, and that the MST is in

its storage rack

Computer starts trajectory sequence

The shut-down procedure for an MMS module exchange demonstration is a

little simpler than that for the basic module exchange and is shown in

Table 10.1-2. The hold and abort processes are similar for both the

basic and MMS module exchange demonstrations. These processes are

described in Section 9.1.5.3.

10-20
Table 10.1-2 Shut-Down Procedure

Step

No.

1.

2.

3.

4.

5.

6.

7.

8.

9.

10.

11.

12.

Action

End module exchange by operating SCP Computer

Mode switch to HOLD or by operating space bar on

computer terminal

Computer will cause a small auxiliary display

to appear

Operator types in a 10 and carriage return to

return to Mode Selection Menu

Select Manual-Direct mode on SCP

Verify that ETU is in rest position and MST is

latched in its storage rack

Configure Servo Drive Panel switches for shut down

Turn servo power off

Turn master power off

Operate key 5 and carriage return to return to

IOSS Main Menu - MMS

Operate key 5 and carriage return to exit from

IOSSMM program

Log off by entering BYE and carriage return

Turn computer off

10.1.4 Interface Definition

A complete definition of all signals between the MMS software in the

computer and the Servicer Servo Drive Console of the Engineering Test

Unit was made. A separate computer program, called TESTHM, was written

to assist in checking out these interfaces. TESTHM provides the

following functions:

1) An ability to select and send commands on each of the digital to

analog channels;

2) An ability to read each status signal from the analog to digital

converter. These signals are displayed in octal notation. Most

of the signals are also converted to engineering units or

true/false statements, as appropriate;

10-21
3) An ability to command the servicer mechanism to move along one

cylindrical coordinate at a time at a selectable rate. This

function is very useful during system calibration and for

positioning the end effector during module location data collection.

The signals from the computer that pass through the digital to analog

converter are:

1) Joint rate commands (6);

2) Error meter commands (6);

3) End effector jaw drive commands;

4) Connector positioner drive commands;

5) MST latch motor drive commands;

6) MST bolt motor drive commands.

The signals to the computer that pass through the analog to digital

converter are:

1) Joint position feedback (6);

2) Hand controller commands (3);

3) Rate level selection (1);

4) Computer mode selection;

5) Control mode selection;

6) Module receptacle status (6);

7) End effector jaw status;

10-22
8) Module servicing tool status;

9) Module servicing tool presence;

10) Connector positioner status.

10.2 SOFTWARE PREPARATION

The MMS module servicer control software was based initially on the

basic module servicer control software and the servicer software

requirements document for MMS module exchange. The basic module

software (IOSS and TESTH) was first copied to a new disk. The

inappropriate routines were stripped out, new routines were added, and

other routines were modified. All of the new software was operated on

the POP 11/34 computer at Denver using the simulated hardware mode and

was carefully checked against the software requirements to the extent

possible. Software for MMS module exchange is identified as IOSSMM and

the test software as TESTHM.

10.2.1 Software Development

The MMS module software was developed on the same machine at Martin

Marietta as was used for basic module software development. The

software can be used with Versions 3.2 or 4.0 of the RSX-11M operating

system and Versions 4.0 or 5.0 of the Fortran F77 compiler. The

software is written entirely in Fortran and involves almost 6000 lines

- of code in 87 routines. The software is organized in the same way for

both programs (see Section 9.2.1).

The software structure involves source files, object files, task files,

command files, overlay descriptor files, data files, and the various

commands needed to compile, link, and execute. The software is broken

up into different source files because it is easier to edit small files

and it is easier to move the different object file names within the

task builder file to create useable overlays. There may be more than

one routine inside a source file. But the file name will reflect the

10-23
purpose of all the different routines. An example is SCRUTL.FTN. This

file contains several routines that deal with SCReen UTiLities.

Compilation of software is done on a file by file basis. So for each

source file, there will be one object file. A compilation can take

many source files and create one object file, but we don't do this

because of the overlay scheme. A simple compilation command such as

'F77 SCRUTL=SCRUTLI will only create the object file SCRUTL.OBJ and any

error messages will be listed to the terminal. A compilation command

file F77.CMD allows the user to compile all the IOSSMM program source.

It is invoked by typing F77 @F77. It is stopped by typing ABO F77.

Task building involves a task build command file (IOSSMM.TKB) and a

task build overlay descriptor file (IOSSMM.ODL). The task build

command file is rather simple. But the overlay descriptor file is

complex. Significant rearranging was done to create working overlays

and even now they are very large (almost to the 32K limit).

Unfortunately, not much expansion is available. The task builder is

started, using the command files, by issuing the MCR command TKB

@IOSSMM. This approach assumes that the machine has a floating point

processor.

The software was prepared in Denver using the available PDF 11/34.

Provisions are included for simulating various hardware functions such

as joint rate integration and hardware switch positions. Extensive

checks of trajectories, displays, inhibit logic, equations, and

auxiliary functions were made before the software was taken to MSFC.

The software was put onto magnetic tape for the transfer. It was then

read off the tape and put on a disk at MSFC. Various tests and

calibrations of the hardware to software interfaces were made before

any attempt was made to drive the hardware. See Section 11.2 for a

discussion of the hardware demonstrations.

10.2.2 Software User's Manual

The Software User's Manual provides a means of understanding the

structure of the IOSS-MMS (referred to therein as IOSSMM) software.

10-24
Included is a complete description of the software, which represents

over six thousand lines of executable code in 87 routines. The

document has a hierarchical breakdown of these routines that are

completely cross-referenced. The manual also explains the installation

procedure and how to run the software.

The first chapter covers installation and execution of the IOSSMM

software. The second chapter is a small one that describes the

relationship between the different types of files. The third chapter

presents the highest level of source description in the form of

software module hierarchy. The fourth chapter presents an intermediate

level of source description in block diagram form. The fifth chapter

presents the most detail in the form of listings of the individual

subroutines.

10.2.2.1 Installation - This section covers the installation of the IOSSMM

software. The software was shipped from Martin Marietta Aerospace,

Denver, CO on 1/2 inch magnetic tape in a BRU (Backup and Restore

Utility) format. The IOSSMM software resides in directory [101,1].

The software is restored first. Once the software is on the disk, it

is compiled to create the object files (*.OBJ). After the compilation,

the IOSSMM task is built. The task builder uses two files. The first

is the command file, IOSSMM.CMD. The second file is IOSSMM.ODL and

describes the overlay structure. Once the task build is complete, the

presence of the IOSSMM.TSK and IOSSMM.MAP files is checked. The file

IOSSMM.TSK is the executable load module and the file IOSSMM.MAP is a

map of the load module.

10.2.2.2 Running the IOSSMM Software - The IOSSMM software can be run from

any directory with one restriction. The software has to be executed

where the data files exist. These data files all have suffixes that

end in .DAT and are ASCII files that can be edited. It is not

recommended that they be edited directly, rather there is a procedure

described in the Software User's Manual that allows them to be updated

in a simpler manner.

10-25
10.2.2.3 Starting the IQSSMM Program from MCR - The IOSSMM software is run

from the directory containing the data files. If the data files are

not present, errors will result and the program will exit smoothly to a

higher level. The IOSSMM program is started from the UFD [101,1]

directory, by issuing the following MCR command:

RUN IOSSMM.

The IOSSMM program then asks the following question:

Terminal Type (VT52, RCA, Simple)?

The user responds with the type of terminal he is working with. The

default is selected by simply pressing RETURN and is currently the RCA

terminal type. The simple terminal is used only for creating

hardcopies of the various menus and. is not capable of copying the

Supervisory or Manual displays. Then the mode of the terminal is

changed to the slave mode. This change is normally invisible to the

user. But if the program aborts before a normal completion, the

computer will not respond to a terminal input. Next, the various

constants are read from a disk file (CONSTS.DAT). The most recent

version (the one with the highest version number) is read from disk.

If this file does not exist in the current UFD, the program will abort.

The next step in the initialization process is to read the module

locations from the disk. Again, only the most recent versions are read

and the files need to be in the current UFD. The data files are named

SCI.DAT, SC2.DAT, SC3.DAT, SC4.DAT, SRI.DAT, SR2.DAT, SR3.DAT, SR4.DAT,

SR5.DAT, SR6.DAT, SR7.DAT, SR8.DAT, MSI.DAT, REST.DAT, FLIP.DAT, and

IFLIP.DAT. The name of each data file is printed to the terminal

before it is read. Once the data files are finished being read, the

screen is cleared immediately and the following message appears:

Prerequisite to IOSS Main Menu - MMS;

Verify that P2 is connected to J2A on the MMS junction box;

Press RETURN when complete.

10-26
Once the operator has performed the verification and pressed the RETURN

key, the IOSS Main Menu - MMS is displayed.

10.2.2.4 IOSS Main Menu - MMS - The IOSS Main Menu - MMS is the highest menu

in the hierarchy of menus and it looks like the following:

IOSS MAIN MENU - MMS

1. Run Setup Menu;

2. Mode Selection Menu;

3. Module Data Collection Menu;

4. Hardware Calibration Menu;

5. Exit to MCR.

Enter Item Number:

Selection of Item 5 returns the user to the operating system. This

item is employed when the user is finished using the IOSSMM program.

10.2.2.5 Run Setup Menu - The Run Setup Menu is used primarily to set up a

new run. It is also used to access the various constants. The menu

looks like the following:

RUN SETUP MENU

1. Operator's Name John Doe

2. Run Title A Test Run

3. Run Number 1

4. Hardware Present? T

5. Save Constants to Disk

6. Geometry Constants Menu

7. Limit Constants Menu

8. A/D Scale Factors Menu

9. D/A Scale Factors Menu

10. Joint Position Threshold Menu

11. Simulated Hardware Characteristics Menu

12. Return to IOSS Main Menu - MMS

Enter Item Number:

10-27
The data in this menu or in the listed next level menus can be changed

by following the prompts that are provided. When an item is changed,

the screen is rewritten to show the new data. The MMS simulated

hardware characteristics menu is more extensive than the related menu

in the basic module software because of the larger number of statused

elements and to simplify simulated hardware operations.

10.2.2.6 Mode Selection Menu - This menu allows the user to select one of the

four operating modes. The Mode Selection menu is:

MODE SELECTION MENU

1. Unassisted Supervisory Mode;

2. Assisted Supervisory Mode;

3. Manual-Augmented Mode;

4. Manual-Direct Mode;

5. Return to IOSS Main Menu - MMS.

Enter Item Number:

Selection of Item 1 calls up the Unassisted Supervisory Mode menu.

From this menu, the total trajectory will run unassisted, as explained

in the Supervisory unassisted mode section. Selection of Item 2 calls

up the Assisted Supervisory Mode menu. From this menu, a total

trajectory will run only with assistance, as explained in the Assisted

Supervisory mode section.

Selection of Item 3 starts the Manual-Augmented mode. Several items

are displayed (operator's name, time, and date) and then the

Manual-Augmented mode equations are executed in a 15 Hz loop. The

operator can then drive the ETU using the hand controllers. The

software also checks for a keypress or the computer mode switch to

leave the Manual-Augmented position. Once either of these actions

happen, the computer will print a display of pertinent parameters and

then give the user the option to continue or leave the Manual-Augmented

mode.

10-28
Selection of Item 4 displays the operator's name, the time, and date.

Then the software starts cycling in a 15 Hz loop waiting for a

key-press or the computer mode switch to leave the Manual-Direct

position. Once either of these actions happen, the computer will print

the ending and elapsed time and then exit. Selection of Item 5 returns

the user to the previous menu.

10.2.2.7 Supervisory Mode - Selection of Item 1 from the mode selection menu

calls up the Supervisory mode menu for unassisted operation. The

Unassisted Supervisory mode menu is:

UNASSISTED SUPERVISORY MODE MENU

1. Module Location Switch Correspondence Menu

Turn Servo Power On

Put SCP Computer Mode Switch to Operate

2. Perform Axial Total Trajectory

3. Return to Mode Selection Menu

Enter Item Number:

The text between Items 1 and 2 informs the operator to turn the servo

power on and set the computer mode switch to operate. If the computer

mode switch is not in operate, an immediate hold will be generated when

the total trajectory starts.

The Assisted Supervisory Mode menu looks and acts like the unassisted

menu, except the title has "assisted" in it. The differences are very

evident when a total trajectory is running.

Selection of Item 2 starts the total trajectory. For a complete

description of the total trajectory, consult Chapters 7, 8, and 9 of

the Servicer Simulation Software Requirements (MMS Modules) document.

Various checks are performed at the very beginning of the total

trajectory. These checks include the following: making sure the arm

is at rest; checking that the mode switch is in operate; checking that

the MMS modules are in their proper positions; and checking that the

10-29
MST Presence, Jaw Open, and Connector Positioner disconnected signals

are present.

If the total trajectory is being run assisted, then the operator has to

tell the computer to perform the next step. The computer will pause at

the beginning of each action and wait for a PR, a CO, or a Y to

continue with the action. If the user responds with something other

than PR, CO, or Y, he will generate a hold command with the effects

described below. Hold commands are generated in both assisted and

unassisted modes and can be generated by the operator in the following

ways:

1) Answers with other than a CO, PR, or Y to an assisted mode prompt;

2) Presses a key while a movement action is executing;

3) Moves the computer mode switch to HOLD while a movement action is

executing.

Hold commands can be generated by the computer in the following ways:

1) Jaw open or close exceeds 15 seconds;

2) MST latch or unlatch exceeds 6 seconds;

3) CP connect or disconnect exceeds 4 seconds;

4) Bolt fasten or unfasten exceeds 28 seconds;

5) Ready signals from MST latch, jaw, CP, and MST bolt drive

mechanisms not correct.

Once a hold command is generated, all joint rates are zeroed, the latch

mechanism is stopped and jaw movement is stopped. Then a hold menu is

displayed on the right side of the screen (leaving the Supervisory

display untouched). This hold menu is:

10-30
Display Parameters

1. Cyl. Cmd.

2. Jnt. Rate

3. Err. Mtr.

4. Jnt. Pos.

(The Supervisory display 5. Cyl. Pos.

would appear here.) 6. Cyl. Err.

7. Cyl. Rate

8. Jnt. Cmd.

9. Continue

10. Abort

Item:

Hold Due to (A hold message would appear here.)

Selection of one of Items 1 through 8 will result in the display of

different sets of parameters and their respective values. After

displaying the different parameters and the respective values, the

computer will prompt the user to press return,to get back to the hold

menu. Selection of Item 9 erases the hold display menu and continues

the total trajectory from where it stopped. Selection of Item 10 will

abort the total trajectory and leave the arm at its current position.

The operator will be returned to the Supervisory mode menu (the

previous one) and from there he probably will go to a manual mode and

drive the arm to the rest position.

10.2.2.8 Software Module Hierarchy - The MMS software module hierarchy is

similar to the basic module software hierarchy except that MPSEQ is no

longer needed and the routines under TTRAJS are different. The MMS

software levels beneath TTRAJS are:

TTRAJS 1 Total Trajectory

TRAJS 9 Trajectories

STEPS 13 Steps

SUPMOT Supervisory Motion Loop

10-31
11.0 SERVICER 1-ij DEMONSTRATIONS

The objective of this activity was to demonstrate on the ground the

capability to exchange failed modules from a spacecraft with good

modules from a stowage rack. The exchange (at separate times) of both

basic, 24 in. cube, modules and MMS, 48 in. square by 20.5 in. deep,

modules were to be demonstrated. In addition, the ability to control

the system in the Supervisory mode with operator inputs between

actions, the Supervisory mode with minimal operator action, and the

Manual-Augmented mode was to be demonstrated. The demonstrations were

conducted in the Robotics Laboratory in Building 4619 at Marshall Space

Flight Center and primarily used MSFC equipment. The MMS module

exchange equipment was new, of which the MMS modules, the MMS

spacecraft and storage rack modifications, the connector positioner,

the optical targets, the MST storage rack and certain wiring were

prepared by Martin Marietta, the 1-g Module Servicing Tool was provided

by GSFC, and the MST electronics were prepared by Fairchild Space

Company. MSFC developed the electrical interconnections between the

A/D and D/A converters and the Servicer Servo Drive Console.

The servicer/MMS mockup equipment for Change Order 3, produced by

Martin Marietta, was shipped to MSFC and was received in good

condition. The existing spacecraft mockup at MSFC was reworked to

accept the MMS mockup and was repaired and painted to improve its

appearance. The MMS spacecraft mockup was installed and aligned. The

existing stowage rack mockup equipment was rearranged and modified so

that the MMS mockup equipment could be installed. Changes were made to

the Engineering Test Unit (ETU) and Servicer Servo Drive Console (SSDC)

electrical wiring to facilitate operation of the new MMS mockup

equipment, the connector positioner, and the 1-g Module Servicing Tool.

All installation, integration and alignment activities for the new

equipment were successfully accomplished. This effort included the

integration of the 1-g Module Servicing Tool, delivered by GSFC and

Fairchild Space Company, with the MMS mockup equipment, as well as the

11-1
ORIGINAL PAGE IS

Of! POOR QUALFTY

installation of the connector positioner mechanism on the ETU end

effector and its alignment with the MST. The new ETU configuration for

ground demonstration of basic and MMS module exchange is shown in

Figure 11-1.

Figure 11-1 ETU Configuration for Demonstration

of Basic and MMS Module Exchange

The cooperation and assistance of MSFC personnel, in particular the

Contract Technical Manager, Mr. James Turner and Messrs. Tom Bryan and

Don Scott of the Robotics Laboratory, in obtaining needed materials and

performing the above work was very helpful and is greatly appreciated.

11-2
The general demonstration block diagram is shown in Figure 9-2 and is

discussed in Section 9.0. The specifics for the MMS demonstration

equipment design and fabrication are discussed in Section 8.0. All of

the equipment was located in one room with the consoles being about 15

ft from the stowage rack mockup and arranged so that the operator had

to turn his head away from the displays to see the mockup equipment.

This arrangement was very convenient for the checkout activity. The

operator's console was arranged as shown in Figure 11-2. The control

and display equipment was arranged around the Servicer Servo Drive

Console. No attempt was made to develop an optimum control equipment

arrangement. The RCA APT 4801 computer terminal was placed on the desk

shelf of the SSDC. The computer display was placed on top of the SSDC

alongside the monitor for the servicer end effector TV camera. The

hand controller was mounted on a stand and placed next to the SSDC.

The MST control panel was mounted above the two CRTs and the MST

electronics were mounted in a remotely located electronics rack. The

resulting arrangement was satisfactory for the checkout and

demonstration activities.

11.1 BASIC MODULE EXCHANGE DEMONSTRATION

The basic module exchange demonstration activity was accomplished in

three phases: 1) the software was installed on the MSFC computer and

checked that it would go through exchanges in the simulated hardware

mode; 2) interfaces between the software and the SSDC were checked

using the TESTH program, the ETU scale factors were established, and

the module locations were measured and data stored in the computer; and

3) basic module exchanges were demonstrated. MSFC personnel were

involved in each phase to enable them to become familiar with use of

the software and to receive some hands-on training.

11.1.1 Basic Module Software Installation

The software was restored from a magnetic tape prepared at Martin

Marietta to a disk at MSFC on the PDP-11/34 system in the computer

room. The disk was then moved to the PDP-11/34 that is interfaced to

11-3
MST

CONTROL PANEL

COMPUTER

DISPLAY

TV

MONITOR

SERVICER CONTROL

PANEL

COMPUTER

TERMINAL

DESK SHELF

HAND

CONTROLLER

MST

ELECTRONICS

(LOCATED REMOTELY)

Figure 11-2 Operator's Console Arrangement

the ETU. Both the source and executable files were restored from the

tape. This operation involved close cooperation between contractor and

NASA personnel, and enabled MSFC personnel to become familiar with the

delivered software.

The first attempts to install the software were unsuccessful because of

a hidden privilege restriction. It was also decided that the program

11-4
was too large to be effectively checked out using one disk. This was

particularly difficult when the range of applications programs were

considered. Thus MSFC personnel arranged for a second disk and a

version of the RSX-11M operating system that could handle one disk for

the operating system and a separate disk for the applications

programs. It was necessary to modify the software slightly so that it

would run on the later versions of the operating system and F77

compiler.

11.1.2 Basic Module Control System Checkout

The basic module control system checkout involved checking and

modifying equipment alignment to vertical, A/D and D/A functions, and

ETU joint scale factors and zeros, setting module status indicators,

collecting module location data, and improving software logic and

Manual-Augmented control mode operations.

Hardware equipment alignment was checked using the direction of

gravity, as indicated by an accurate level, as a reference direction.

The verticality of the docking post, mechanism forearm perpendicular,

and all four interface mechanisms was checked and adjusted where

necessary. A 0.060 in. shim was installed between the end effector

base and the interface mechanism alignment cone to increase the

positional repeatability of this interface alignment.

The zeros and scale factors for each of the ETU joint feedback

indicators were checked. The SSDC voltmeter reading to angle data was

checked first and then the software to SSDC voltmeter calibration was

checked. Errors were found and corrected. Variations in the precision

+10 Vdc power supply were tracked and found to vary with room

temperature. The variations were within allowances.

The module locations were set and marked with pencil so the modules

could be easily returned to their desired locations.

The end effector TV camera lens was adjusted so that a good picture was

available over the desired range of travel. It was found necessary to

11-5
increase the width of the lines making up the Manual-Augmented targets

to 0.1 in., because of TV system degradation. New targets were made

for each of the four module locations used during the demonstrations.

These targets were aligned with respect to the interface mechanisms by

a template.

Operation of the 12 module location switches and the three end effector

switches was checked and switch operating points were reset. The

electronic interface logic that converted status indications to

voltages was reworked to provide better separation of the voltages for

repeatable software conversion of the voltages to status indications in

the computer.

All of the analog-to-digital and digital-to-analog functions were

checked for agreement with the software requirements document.

The logic and circuitry that converted end effector and interface

mechanism drive signals from the computer to polarized 28 Vdc power to

the motors was revised to give priority to the control panel switch

commands and to enable the computer generated commands only when the

control mode switch was set to Supervisory. The objective of this

revision was to decrease the likelihood of unexpected signals and to

provide a method of overriding unexpected signals.

Data was collected for each of four module locations and was verified

for the rest, flip initial and final, and transition initial and final

positions. The flip location was moved a little to provide better

clearance from the good module. This same type of data was collected

for use in the Manual-Augmented trajectory sequences.

The system was operated in the assisted and unassisted Supervisory

control mode to identify potential improvements. Some step end points

were adjusted to provide smoother transfers between steps. The

standoff distance at the end of each trajectory was increased to 3.0

in. from 2.0 in. to provide better clearance during the combined

motions of some end effector only trajectories.

11-6
A new logic was developed for driving x and r during the interface

mechanism latch and unlatch motions. The required x and r motions are

very non-linear with time because of the cam shape. The variation in

drive motor speed with load also increased the uncertainty. The

Interface mechanism linkages were disassembled, cleaned, greased, and

reassembled to obtain more consistent operation. The newly developed

logic worked very well.

The trajectories used when inserting a module in the radial location

were revised to provide a higher command trajectory. This additional

commanded height compensated for a twist in the servicer mechanism that

allowed the tip of the interface mechanism to sag below the guide

capture volume. The cardboard representation of one module was

replaced by a space frame so that both modules would have the same

appearance and weight. Removable fabric covers were made and were used

during motion picture filming so the modules were more visible.

The above checkout activities were conducted before the MMS equipment

was installed. Subsequent to the installation of the MMS equipment,

two other changes were made. The flip location was moved again so

there would be adequate clearance with the MMS temporary module

location in the stowage rack. Also, the good and temporary stowage

rack locations of the basic modules were moved as described in Section

8.0. Data were collected for all four basic module locations. This

demonstrated the usefulness of the new module location data collection

technique. The post-MMS installation locations of the basic modules in

the stowage rack resulted in slightly shorter module exchange times

(approximately 15 min) for the axial and radial total trajectories.

The system was operated in the Manual-Augmented control mode and a

number of improvements and corrections were identified. The trajectory

sequences were rewritten to incorporate the use of one hand controller

instead of the two that had been expected and to reflect the use of a

different ranging method.

It was determined that the size variation of the targets near the

standoff, attach, and unlatched positions was such that stadiametric

11-7
ranging could be used. The initial plan was to set values on the

shoulder pitch (or elbow roll for radial) angle set potentiometer and

then use the error meter to indicate the proper range. This method was

found to be cumbersome and distracted the operator from watching the TV

monitor. The stadiametric ranging method put the range data directly

on the TV monitor. A set of inner marks (at target width) was used to

define the 3.0 in. standoff distance and a set of outer marks (at

target width) was used to define the attached location. The unlatched

position is half way between the two sets of marks. This approach

worked quite well. The Manual-Augmented trajectory sequence

documentation was revised to incorporate the stadiametric ranging

approach.

The target motion directions, in end effector TV coordinates, were

checked against hand controller motion directions for a variety of end

effector attitudes. After switching polarity on two lines, all motion

direction correspondences were correct.

The initial approach for the display to cylindrical coordinate

transformation used in the Manual-Augmented control mode was to base

this transformation on wrist roll angle variation for both axial and

radial motions. However, it was found better to base the

transformation on wrist roll for axial motions and on wrist pitch for

radial motions. During radial motions wrist roll is kept constant and

wrist pitch is varied. Thus, the new approach provided a more useful

transformation.

The trajectory sequences were modified in some details to strengthen

the philosophy of using the Hawk mode for large motions where the

targets are not in the end effector camera field of view and using the

video system when the target is in the field of view. Initial hand

controller motion direction information was provided in the written

trajectory sequences at the beginning of each large motion so the

motion would start in the proper direction. These instructions were

found to be very useful as it is difficult for the operator to

visualize himself as being at the end effector camera when he can

easily turn and look at the entire mockup.

11-8
A draft version of the Manual-Augmented Trajectory Sequences (Basic

Modules) document was prepared for use during the preliminary

checkout. It was modified as the preliminary runs were made and the

changes discussed above were investigated. Also specific module

location data was collected and refined for use in a set of

representative trajectory sequence forms. These representative

trajectory sequences could then be used by MSFC personnel for

conducting demonstrations. The Manual-Augmented Trajectory Sequences

document was then updated and distributed.

11.1.3 Basic Module Demonstrations

Demonstrations of basic module exchange in all three control modes and

for radial and axial exchanges were successfully accomplished a number

of times. The specific combinations of exchanges and control modes are

shown in Table 11.1-1.

There was no need to perform the axial-to-radial or radial-to-axial

total trajectories in the Manual-Augmented mode as these trajectories

were only provided to simplify going between axial and radial

demonstrations.

Table 11.1-1 Basic Module Demonstrations

Control Mode

Supervisory Assisted

Supervisory Unassisted

Manual- Au gmen te d

Total Trajectory

Axial

X

X

X

Radial

X

X

X

Axial to

Radial

X

X

Radial to

Axial

X

X

11-9
The trajectories of Table 11.1-1 were repeated a number of times with

MSFC or Martin Marietta personnel at the controls. Each operator

quickly learned to perform the trajectories with the two Supervisory

control modes. These trajectories were very smooth. In many cases the

interface mechanisms would not touch the guides until they were almost

fully in place. An unassisted Supervisory mode axial or radial total

trajectory takes approximately 15 minutes. Each operator felt that the

form of display was very helpful in keeping track of where he was in

what is a moderately complex activity.

The primary basic module demonstrations were conducted before rework of

the mockups for integration of the MMS equipment. Subsequent to the

mockup rework, data were collected on the revised basic module

locations and all four trajectories were repeated in the Supervisory

mode without operator assistance.

After less than one hour of training, each operator could easily do

major motions in the Manual-Augmented mode and could insert and

withdraw modules in the guides. However, more practice will be

required for any operator to smoothly and efficiently perform module

exchanges using the Manual-Augmented mode. The need for this practice

had been expected.

A set of overall start-up and shut-down procedures were prepared for,

and provided to, MSFC personnel. These procedures primarily related to

the PDP-11/34 computer and its peripherals. The details of operating

the software, Engineering Test Unit, and Servicer Servo Drive Console

were provided in other documentation (see Introduction to Section 9.0).

11.2 MMS MODULE EXCHANGE DEMONSTRATION

The MMS module exchange demonstration activity started with the

installation of the Change Ordex 3 MMS mockup equipment, the

modification and refurbishment of the existing Engineering Test Unit,

and the integration of the GSFC and Fairchild Space Company supplied

MST equipment with the ETU. The software related activities were

11-10
performed in the same three phases used for the basic module exchange

demonstrations - equipment and software installation, checkout of

interfaces between software and hardware along with preliminary

operations, and module exchange demonstrations. MSFC personnel were

involved in each phase. The demonstration setup description for MMS

module exchange is similar to that used for basic module exchange and

is given in the introduction to this section (11.0) of the final report.

11.2.1 MMS Module Exchange Equipment Installation and Checkout

The servicer/MMS demonstration mechanical and electrical equipment for

MMS module exchange (per Change Order 3) that was shipped to Marshall

Space Flight Center by Martin Marietta is listed in Table 11.2-1.

Table 11.2-1 MMS Module Exchange Equipment

Item

- MMS Module Mockup

- MMS Spacecraft Mockup

- MMS Stowage Rack Support

with Receptacle

- MMS Stowage Rack Support

- MMS Target Assembly

- MMS Target Assembly

for MST Rack

- Connector Positioner

- MST Rack

- SCP Front Panel Mod.

- MMS J-Box

- Cable Assembly

- Cable Assembly

- Cable Assembly

- Cable Assembly

- Cable Assembly

- Cable Assembly

- Cable Assembly

- Cable Assembly

- Cable Assembly

- ETU J-Box Mod.

(J44, J45, J46)

Serial No.

01 and 02

01

01

02

—

—

—

—

—

—

—

—

—

—

—

—

—

—

—

— —

Part Number

RES4100000-009

RES4200000-009

RES4300000-009

RES4300000-029

RES4600000-009

RES4600000-019

RES4400000-009

RES4 500000-009

RES3159485

RES3159950

RES4700210

RES4700220

RES4700310

RES4700320

RES4700330

RES4700340

RES4700400

RES4700500

RES4700820

—

Quantity

2 units

1 unit

1 unit

1 unit

6 units

1 unit

1 unit

1 unit

1 unit

1 unit

1 unit

1 unit

1 unit

1 unit

1 unit

1 unit

1 unit

1 unit

1 unit

1 unit

11-11
The electrical equipment consisted of the cables, junction box, and

control panel kit required for the ETU and SSDC modification for MMS

servicing demonstrations. Included in the shipment was a repair or

maintenance kit comprised of materials and spares that might have been

necessary had any damage occurred during transportation or

installation. All the equipment arrived in good condition at MSFC.

The MMS demonstration mechanical and electrical equipment was installed

and checked out at MSFC. The equipment was first unpacked and

inspected in the Robotics Laboratory area of MSFC. Materials and tools

were prepared for ETU modification and repair and for MMS demonstration

equipment Installation. The majority of the necessary tools and

materials, other than those in our shipment, were provided by MSFC.

The stowage rack mockup was dismantled and removed to gain access to

the spacecraft mockup, servicer arm., and control cables. The existing

boxes on the front face of the spacecraft mockup as well as its front,

bottom, and end panels were then removed. The spacecraft mockup

structure was reinforced using 2x4 wooden beams and metal brackets.

The existing 1/8 in. thick panels of the spacecraft mockup, covering

the front, bottom, and end (radial module location), were replaced by

new panels made of 1/4 in. masonite. The new panels were painted white

and the corners were covered with white aluminum angle. The MMS

spacecraft mockup was then installed on the front panel and the module

interface frame was adjusted and leveled (see Figure 11-3). Three

non-functional module mockups were installed on the face of the ETU

spacecraft mockup that received the MMS module interface. White

stripes of vinyl tape were applied to the black docking post mockup so

that the docking post would stand out better.

The MMS equipment status and control cables were installed on the

spacecraft mockup and along the servicer arm. The MMS .junction box was

installed on top of the existing ETU junction box under a floor panel

and the module bolt status cables were connected. Connection

modifications on the existing ETU junction box and Servicer Servo Drive

11-12
ORIGINAL PAGE IS

OF POOR QUALITY

Figure 11-3 Engineering Test Unit Configuration for Ground

Demonstrations of Basic or MMS Module Exchange

11-13
Console were also made. The connector positioner control panel adapter

was installed on the existing Servicer Control Panel.

The stowage rack base was painted, reassembled and aligned. Installation of the stowage rack mockup was completed with the installation of

the stowage rack beams, MMS module supports, MST storage rack, basic

module supports and optical targets, as shown in Figure 11-4.

The two MMS module supports were positioned on the stowage rack and

leveled at the proper height. Their correct position was marked on the

stowage rack base. This marking took the form of disks cut from emery

cloth and glued face up on the floor of the stowage rack. The emery

cloth surface provided adequate friction so the nylon feet of the MMS

module supports would not slide too easily. The two basic module

receptacle units (for the side interface mechanism) were installed on

the stowage rack. Their wooden bases were modified to eliminate

interference with the stowage rack.

Figure 11-5 shows the arrangement of the MMS and the basic module

mockup "good" locations and also the installation of the MST storage

rack near the outer end of the left hand stowage rack beam.

The interface drawings produced by Martin Marietta as well as the

interface coordination activity described in Section 8.2.9, were

effective in assuring a smooth integration of the Module Servicing Tool

(Figure 11-6) and its control equipment with the Engineering Test Unit.

The Goddard Space Flight Center and Fairchild Space Company delivered

high quality equipment in time for its integration with the ETU. Their

representatives cooperated with the Marshall Space Flight Center and

Martin Marietta personnel for a successful integration. Only a few

minor installation problems were identified and they were quickly

corrected. A 0.19 in. shim was installed under the servicer attach

fitting of the MST to allow proper clearance for the locating cone of

the end effector. The slightly increased "B" distance, to 7.41 in., is

acceptable due to the MST weight falling below the allowable weight.

11-14
ORIGINAL PAGE IS

OF POOR QUALITY

Figure 11-4 Stowage Rack Supports for "Temporary" Locations of Basic

and MMS Modules

11-15
Figure 11-5 Stowage Rack "Good" Locations for Basic and

MMS Modules

The ground demonstrations MST has a total weight of 13.5 Ibs compared

with the design goal of 15.0 Ibs maximum. The ETU end effector

interface includes the attach fitting, an electrical disconnect mounted

to allow 0.030 in. of float, and a flat landing area for the

"ready-to-attach" sensor of the end effector. The interface of the MST

with the MMS module mockup includes two latches that mate with the

latch interface plate of the module, two "ready-to-latch" sensors and a

spring loaded 3/4 in. socket. The socket spring initially had an

approximate 2.7 Ib load at maximum compression and a load of 0.9 Ibs at

full extension.

11-16
ORIGINAL PAGE FS

OF POOR QUALITY

Figure 11-6 Module Servicing Tool for Ground Demonstrations

Correct mating of the tool with the MMS module, the ability to tighten

and loosen the bolt, and latch operations were verified, using the MST

power supply and control unit and a special test cable. These

operations were successfully repeated later using the revised ETU and

SSDC wiring. The tool interfaced correctly with the MMS module mockup

retention system and operated the bolts when the modules were attached

to the stowage rack supports. The MST bolt drive involves a worm gear

so the mechanism will not backdrive. Thus when a bolt is tightened the

torque causes the module structure to flex and a torque remains after

the motor voltage is removed. The residual torque and related friction

coefficients are high enough that it is difficult to remove the MST

from the module unless a pulse of loosening torque is applied to

relieve the torque loads. It was later found necessary to chamfer the

MMS module bolt heads, to deburr the square socket drive of the MST and

to adjust the MST socket spring force in order to obtain consistent

engagement of the MST socket on the MMS bolt heads.

11-17
;/ * • « » •

The correct mating of the MST with the ETU end effector was

successfully verified prior to the installation of the connector

positioner mechanism. Using the MST mated with the end effector as a

reference, the connector positioner was then aligned and fastened to

the ETU.

Upon completion of the installation of the electrical connector

positioner mechanism on the ETU end effector, connector mating tests

were successfully performed (see Figure 11-7). The connector

positioner adjustable link length was set to obtain the desired

connector pin engagement length and the raicroswitch settings were

verfied. Connector engagement and disengagement was smooth and proper

clearance between the MST and all end effector components was present

during end effector engagement.

11.2.2 MMS Module Software Installation

The MMS software was restored from a magnetic tape prepared at Martin

Marietta and transferred to the applications disk that contains the

basic module exchange software. Thus both sets of software are on the

same disk and reside in different directories. It is easy to transfer

between the basic and MMS sets of software. The applications disk was

installed in a disk drive connected to the PDP-11/34 that is interfaced

to the ETU. The operating system is on a separate disk and disk drive

connected to the same PDP-11/34 computer. Both the source and

executable files were restored from the tape. This operation involved

close cooperation between contractor and NASA personnel, and enabled

MSFC personnel to become familiar with the delivered software.

The MMS software installation went smoothly because the techniques had

been developed during the basic module exchange software installation.

11.2.3 MMS Module Control System Checkout

The MMS module control system checkout involved checking and modifying

A/D and D/A functions, adjusting module status indicators, collecting

11-18
ORIGINAL PAGE IS

OF POOR QUAUTY

Figure 11-7 ETU End Effector and Its Connector

Positioner Mated with the MST

11-19
module bolt location-data, improvement of software logic, refinement of

MST operations, addition of a pseudo-combined motion capability and the

related module offset location data collection and storage, addition of

a droop compensation system, and improvement of Manual-Augmented

control mode operations. The following discussion is arranged by

subject rather than chronologically.

MMS module control system checkout proved to be much more difficult

than expected because of the MST operations refinements, the need for

the pseudo-combined motion capability, the need for a droop

compensation device, and computer, D/A, and A/D operating problems.

While the PDP-11/34 computer and the D/A and A/D equipment generally

operated quite well, there were periods, starting in January, 1985,

when it would not operate properly. The conditions were erratic and

hard to repeat. The worst form of erratic operation was for the D/A to

send out full voltage signals, which caused the servicer mechanism to

move quickly in an unanticipated manner. Fortunately the operators

were attentive and equipment damage was minor. Erratic operation

occurred more often when room temperature exceeded 70°F.

In October two problems became repeatable enough that they could be

identified and corrected by MSFC personnel. One problem caused the

software to revert to the beginning of a trajectory. It was solved by

replacing the computer's math board (floating point accelerator).

Apparently the computer lost track of the numerical step it was on in

the trajectory and then went back to zero. The second problem was more

serious in that it caused the servicer mechanism to move quickly up and

to the right. Servo power had to be shut off before anything was

damaged. It was found that the digital to analog converters were

putting out full negative signals. This problem was more difficult to

solve, but the MSFC personnel were able to isolate the cause to two

integrated circuits that were replaced. Fortunately, the software

could be operated while this problem existed as long as the ETU servos

were off.

11-20
The erratic operation of the computer was also evidenced on our last

day of operation in December, 1985. The PDP-11/34 is not a new

computer. It is recommended that the PDP-11/34 computer and the

associated A/D and D/A converters be replaced with more reliable and

less temperature sensitive equipment. Care should be taken in

selection of new computer and peripheral hardware to minimize software

compatibility problems.

Another cause of the extended development time was the length of an MMS

module exchange (approximately 45 min). The software is such that it

can only be started from the beginning. It cannot be started part way

through a total trajectory because of the way the computer keeps track

of where it is in a trajectory. This is not a problem during normal

operation, but it results in a lot of unproductive time during

checkout. This is particularly true if the problem area is near the

end of the total trajectory. We were able to use the hold function to

help overcome this characteristic. However, it is recommended that any

new servicer software be designed so that operations can be started at

a number of different places within the overall total trajectory. It

is not necessary to be able to start anywhere in a total trajectory,

rather it is desirable to limit the unproductive time to less than 3 to

5 minutes per run.

The other overall constraint was the limited time and resources

available for problem solution. Generally a working solution was

obtained for each problem. However, resources did not permit

development of an understanding of each problem cause nor evolution of

better solutions. Once a workable solution was found, that solution

was used and the problem was set to one side. Little or no data was

collected to develop rationale for problem solutions. For example,

data on maximum and minimum forces for the selected MST socket return

spring were not collected. Rather springs were installed until one

worked and then we went on to the next problem.

11.2.3.1 A/D and D/A Checkout - Each of the analog-to-digital and

digital-to-analog functions were checked for agreement with the MMS

11-21
software requirements document. A number of modifications were made to

the breakout box to get consistency of operation and agreement with the

software requirements document. A diode was added in the breakout box

to correct for a sneak circuit that was causing the SSDC control mode

relays to stay energized after the normal source power had been

removed. A wire was added to the breakout box to provide a more

consistent ground return for the Supervisory mode control signal in the

breakout box.

MSFC personnel repaired one interface electronics board that is

involved in the signal path that drives the connector positioner to the

disconnect position. A method of interrogating the computer operating

system to ascertain the operability of the D/A and A/D converters was

tried. Unfortunately this caused an excessive wander in the ETU wrist

roll drive and a flashing of the computer terminal cursor. This

interrogation approach is no longer used.

11.2.3.2 Status Sensor Adjustments - The MMS module bolt status sensors were

adjusted to provide status signals at the proper points. As the

checkout operations continued, it was found necessary to reset these

microswitches more accurately. Flat washers were added to the top of

the indicator push rods to obtain more consistent contact with the

rounded ends of the MMS attachment bolts.

•

11.2.3.3 Module Bolt Location Data Collection - Data for each of the six MMS

module bolt locations and for the MST storage rack were collected and

stored in MMS software. Repeat checks of this data were quite good

with differences being less than the equivalent of 0.2 in. The TESTHM

program with the cylindrical position command mode was found to be very

useful for this data collection process. Complementary data was

collected for the six MMS module bolt offset locations after the need

for that data was established.

11.2.3.4 Software Logic Improvement - The large size and flexibility of the

MMS modules, when combined with the small MST insertion capture volume

and the MMS bolt and nut system capture volume resulted in the need for

11-22
a large number of software changes. One class of changes involved

obtaining module location data, and increments to that data, more

accurately than had been necessary in the past. Another class had to

do with overdriving the vertical position to compensate for MST and

module weight. A third class involved trimming trajectory mid points

for smoother operation and to provide greater clearances between MMS

modules and the mockups. Another class was MST operation oriented,

such as the need to pulse each bolt in the loosening direction, after

tightening, to release windup torques on the MST gears so the socket

would come off the bolt heads. Another class was the deletion of

redundant status checks to simplify status switch adjustments. Another

class involved ensuring that all D/A commands would be zeroed out at

program turn-on and if the function was not used in that program.

A number of MMS module exchanges were performed using the software in

the unassisted Supervisory control mode. Lists of discrepancies were

made for each module exchange activity, and then the discrepancies were

corrected before the next total trajectory was run. This process was

repeated a number of times until a complete trajectory had been run and

the discrepancy list was reduced to zero.

The digital signal level to the D/A discrete signal channels was raised

to obtain more consistent and positive discrete command signals.

Discrete signals are used to drive various elements such as the

connector positioner. The x direction motion rate was increased during

the unfasten step to better match the module motion due to bolt

loosening, and the amount of x travel was increased so the bolt thread

would be clear of the nut. When the bolt turns as it is just

unthreaded from the nut and the nut support springs are compressed, the

nut makes a loud banging sound that is distracting. The changes in x

travel and x rate minimized this noise.

The cylindrical and joint error tolerances used during the fasten and

unfasten steps were increased to avoid unnecessarily initiating the

hold mode. At times a MST latch signal was decoded by the computer

when the MST latch signal was not present. The software was revised to

11-23
require that the latch signal be decoded for two successive computation

cycles. This approach eliminated the erratic signal decoding.

11.2.3.5 MST Operation Refinements - It was found that the MST did not

consistently drive the MMS module bolts due to a failure of the MST

drive socket to drop onto the bolt heads. The problem was related to

module distortion with one loose bolt, module distortion due to MST

latch and unlatch action, and ETU positioning inaccuracies. The ETU

positions were trimmed up and the problem was eased, but not

eliminated. The MMS module bolt heads were chamfered strongly and

greased well. The MST socket spring was strengthened. These actions

further eased, but did not totally eliminate, the problem. It was then

decided to chamfer the square corners of the MST drive shaft near the

transition to the 1/4 in. diameter. Apparently there was an

interference with the end of the square hole broached into the socket

that caused the socket to stick. A number of different spring

configurations were tried until a sufficiently high force, but not too

high, was obtained. Too high a spring force resulted in an

inconsistent MST ready signal. The MST ready signal switches were also

adjusted. The result of these modifications is that consistent

engagement of the MST and the MMS module attachment bolts was obtained.

The MST latches were reworked to eliminate an interference during the

unlatch operation. The MST electronics contained an interlock that

stopped the unlatch action when a microswitch giving the unlatched

status operated. This microswitch could not be adjusted for consistent

operation, so the interlock was disabled after obtaining Fairchild

Space Company approval and directions. The software was then modified

to drive the MST latch mechanism for an additional two seconds to

assure that the latches were fully unlatched. It was also noted that

when the MST is latching to or unlatching from the module that the

module torque box is deflected significantly. The MST design is such

that both latches rotate the same way. The torque box deflection can

be eased by designing any new versions of the MST so that the two

latches rotate in opposite, instead of similar, directions.

11-24
It was found that it was difficult to consistently withdraw the MST

from the MMS module retention fitting. The module tended to hang up on

the MST. This situation was traced to the way the bolt tightening

torque and the non-backdriveable characteristic of the MST caused a

windup of the MST drive shaft and of the MMS module torque box, which

caused a high friction force between the MST and the module retention

fitting. The result was that it required excessive force to withdraw

the MST from the MMS. Attempts to ameliorate the problem by applying a

loosening torque pulse to the MST from the MST control panel were

successful, but this success could not be duplicated with loosening

torque pulses controlled from the computer. The MST maximum current

was reduced from 4.4 to 3.0 amperes. This eliminated the problem for

some of the fittings. The settings of the latch microswitches were

then adjusted so that the bolts would not be tightened as far in the

stowage rack fittings. The result was consistently easy extraction of

the MST from the MMS module retention fittings. It is recommended that

future MSTs be designed with a backdriveable bolt drive mechanism.

The MST interface fitting configuration is such that the ETU end

effector drive works harder when it is closing on the MST fitting than

on any of the other interface fittings. An attempt to relieve the

conical region did not reduce the excessive loads. It is desirable to

identify the source of these loads and to reduce the loads before the

ETU end effector jaw drive is damaged.

11.2.3.6 Pseudo-Combined Motion - The major software change was to adapt the

control system to the MMS module ends not being vertical when one bolt

was fastened and the other was unfastened. Each module end has a guide

for the MST probe as the MST socket is moved towards or away from the

MMS module bolt head. When the guides are not vertical, then the end

effector must move along a path where more than one cylindrical

coordinate changes. The basic and MMS software are not designed for

motion in a direction that requires coordination of two or more

cylindrical coordinates. The selected approach to overcome this

problem is to break the motion up into a series of actions so that the

relatively small lateral motions, for a given axial change, are

11-25
acceptable. This approach worked satisfactorily although it required

extensive software modifications to the latch and unlatch subroutines,

which affected 24 trajectory steps. An array of module bolt offset

location data was established for the data corresponding to the tipped

attitude. Data was collected for each tipped module/bolt location and

stored in the new array. Logic was written so the computer would

calculate end conditions for each of the steps along this

pseudo-combined (radial, tangential and axial) motion trajectory. The

resulting motion could be tailored for each bolt and was very

satisfactory.

Equations were developed to allow calculation of the cylindrical

coordinate attitude commands for the tipped condition. These attitude

commands were entered into the computer and the ETU was driven to the

commanded wrist attitude position. The MST probe angle was then

visually checked against the module attitude.

11.2.3.7 Droop Compensation - When an MMS module is fastened to the

spacecraft mockup by the top bolt and the bottom bolt is loose, the

bottom bolt end of the module will droop down due to the module weight

and the finite module stiffness. The bottom bolt also is held out of

its nut by the Acme thread diameter. The sag is made worse as the MST

is unlatched and withdrawn. System error variations can cause the MST

to hang up on the module and pull the module end down even farther,

which causes more binding. A droop at the bottom bolt causes a

horizontal motion of the latch interface at the top bolt. The motion

is like a rotation in a vertical plane about a point near the top bolt

head. The horizontal motion of the top bolt latch interface can result

in the MST probe not entering the corresponding hole. It was

recommended that this 1-g situation be analyzed, a set of requirements

be prepared, and a device made to control the module droop. The droop

effect is due to 1-g effects, and the support device should not be

needed in 0-g.

A droop compensation mechanism to support the bottom bolt end of the

module was designed and parts were fabricated by MSFC. The design was

11-26
• . .-. •• • *

based on certain parts that were available. The parts were assembled

and tested. The design approach is acceptable, but further effort is

needed to obtain reliable operation. Separate microswitches should be

installed near the bottom and top bolt status indicator cams. These

microswitches should then be used to directly control the AC relay.

This will eliminate the DC relays and their tendency to chatter and

provide better margins on microswitch operation. The test and enable

functions that were initially located near the mechanism should be

moved to near the control console. Also, consideration should be given

to replacing the AC solenoid with an AC motor and gear box of

appropriate size and configuration as the AC solenoid may not have

sufficient force available to properly lift the module for all droop

distances.

11.2.3.8 MMS Module Manual-Augmented Control Checkout - The six

degree-of-freedom hand controller was positioned adjacent to the

control console and the three translational degrees-of-freedom were

checked out using the TESTHM program. The Manual-Augmented control

mode software was checked to show that ETU motions were correct for

hand controller motions for a variety of end effector attitudes.

A set of reticles was taped to the TV monitor in a position

corresponding to the location of a basic module target. The seven MMS

targets were then aligned using the TV monitor reticle set. Figure

11-8 shows the ETU end effector approaching the MST that is secured in

a MMS module latch interface location. The relative position of the

Manual-Augmented target and the TV camera (just above the end effector

jaws) can be seen.

Several partial exchanges were made to show that the MST could be moved

and placed in the MMS module conical guides, that the MST could be

latched and unlatched, that bolts could be fastened and unfastened, and

that modules could be moved using the Manual-Augmented mode of control.

11.2.3.9 Miscellaneous - One of the two MMS module mockups was damaged during

the test activity. It was deduced that the styrofoam structure had

11-27
ORIGINAL PAGE IS

Of POOR QUALITY

Figure 11-8 ETU End Effector Standing off from MST

failed near the top bolt attachment. Wooden blocks were fabricated,

fitted, and glued in place. These glue blocks provided a better load

path for the loads from the top bolt attachment to the basic module

structure, which is a torque box. The repair eliminated the excessive

droop that had occurred after the damage and before the repair.

11-28
A relay on a Wrist Pitch circuit board was replaced to overcome erratic

operation. Two diodes on a Shoulder Pitch circuit board were replaced

so that consistent movement from the Shoulder Pitch up and down limits

could be obtained. The location of the temporary stowage rack top nut

resulted in the Wrist Roll angle being at a limit switch point with no

margin. The limit switch was moved to provide an adequate margin of

travel.

The number of minor failures in the demonstration facility equipment

indicates that an ongoing maintenance activity should be established.

11.2.4 MMS Module Demonstrations

Demonstrations of MMS module exchange in all three control modes were

successfully accomplished. A number of MMS module exchanges were

successfully demonstrated in the Supervisory control mode without

operator assistance by MSFC and Martin Marietta operators. An exchange

was also demonstrated using the operator assisted form of the

Supervisory control mode.

A trajectory was also successfully demonstrated in the Manual-Augmented

control mode. This control mode was easier to use than had been

expected. The striping on the top of each module could be used for

guidance during the bottom bolt to top bolt, and vice versa, steps. No

other control clues were needed. The socket on the end of the MST

probe can be seen well enough that it was easy to insert in and

withdraw from the MMS module attach fittings. The targets were

adequate for module alignment. Stadiametric ranging and the various

status indicators were as useful as they were for basic module

exchanges.

MSFC personnel videotaped most of two complete demonstrations.

Very little effort and time is required to switch back and forth

between demonstration of MMS module and basic module exchanges. All

that is required is to initialize the appropriate exchange on the

11-29
computer and to change one connector between the ETU and the MMS

junction boxes. These operations can easily be accomplished in less

than two minutes. The computer notifies the operator if the proper

electrical connection is not made.

As with the basic module exchanges, it was very easy to learn how to

conduct MMS module exchanges in the Supervisory modes. The

trajectories were very smooth and repeatable. When the MMS module had

both bolts tight, the MST would not touch the module guide until it was

almost to the ready to latch position (see Figure 11-9). An unassisted

Supervisory mode MMS axial total trajectory takes approximately 45

min. Each operator felt that the form of display was very helpful in

keeping track of where he was in what is a moderately complex

activity. Display of the total MMS trajectory, current trajectory,

current steps, and current actions, along with the associated

highlightings, was even more helpful than for basic module trajectories

because of the larger number (two lines vs one) of steps involved

during each module transport trajectory.

The extensive checkout activity led to the identification of different

ways to operate the system when anomalies occurred. For example, the

Acme thread on the MMS attachment bolts would sometimes hang up and not

engage properly. The computer would continue until the bolt fasten

timer timed out and then the computer would go to hold and notify the

operator of the excessive time. The operator would then check that the

indicator lamps on the Servicer Control Panel showed a ready but

unlatched condition for the bolt. He would recognize the problem, put

the SSDC and the MST control in the Manual-Direct mode, operate the MST

controls to first loosen the bolt slightly to align the threads, then

pulse the bolt in the tighten direction, monitor bolt drive current to

see that the bolt was not hung up, switch back to the Supervisory

control mode, and then tell the computer to continue the trajectory,

which it would do. While this process seems long, it takes less time

to do than to read about.

11-30
ORIGINAL PAGE IS

OF POOR QUALFTY

Figure 11-9 MST Probe Entering MMS Module Guide

11-31
A number of different ways were identified in which the operator could

use the capabilities of the equipment and of the computer program to

identify and overcome anomalies. The computer, and other displays,

would provide the required information and the operator had the

controls to effect a corrective action. System characteristics that

help in performing these workarounds are:

1) The computer will stay in the hold mode for an indefinite time;

2) The SSDC and MST can be switched to any of the three control modes

while the computer program remains in the Supervisory mode;

3) Servo power control is independent of computer operation;

4) Each ETU joint can be controlled independently;

5) The MST can be operated manually through the ETU wiring or through

a separate test cable.

The effect of learning how to do these workarounds is that the system

is not a pre-programmed entity that cannot continue past the first

anomaly. Rather it is a system with three levels of control that can

be used interchangeably to get the job done in spite of a variety of

anomalies.

11-32
12.0 CONCLUSIONS AND RECOMMENDATIONS

The significant conclusions and recommendations from this Servicer

System Demonstration Plan and Capability Development activity are

presented below. Many secondary conclusions and recommendations are

given in Sections 3.0 through 11.0. The conclusions and

recommendations which span the study are given first.

12.1 ON-ORBIT SERVICING DEVELOPMENT

The following conclusions and recommendations apply to the overall

on-orbit servicing development:

1) The recommended plan leads to the free-flight verification of an

operational servicer suitable for use with the OMV and the Space

Station;

2) The plan has three phases:

- Ground demonstrations,

- Cargo-bay demonstration,

- Free-flight verification;

3) The free-flight verification can be completed by late 1992;

4) The total estimated cost is 45.3 million 1985 dollars;

5) The plan includes three servicer mechanism configurations:

- The Engineering Test Unit currently in use at MSFC would be used

for early ground demonstrations, procedures development, and

training for the cargo-bay demonstration,

- A proto-flight quality unit would be used for the demonstration

flight in the Orbiter cargo bay and for procedures development

and training related to the operational servicer,

12-1
- One fully operational unit that has been qualified and

documented for use in the free-flight verification activity and

in subsequent operations;

6) The plan is based on use of proven IOSS designs and test hardware;

7) Areas for application of the module exchange form of on-orbit

servicing to the Space Station were identified;

8) A user's council should be formed to direct the implementation of

an on-orbit servicing capability.

12.2 MULTI-MISSION MODULAR SPACECRAFT SERVICING

The following conclusions and recommendations apply to the involvement

of MMS equipment in the demonstration plan and in subsequent operations:

1) Primary emphasis would be on demonstrating the exchange of MMS

modules;

2) The MMS Module Servicing Tool should be adapted to work with the

servicer end effector for the exchange of MMS modules;

3) A set of requirements for the MST adaptation was prepared;

4) A preliminary conceptual arrangement of the adapted MST for 1-g

demonstrations was prepared;

5) Lightweight MMS module mockups with dimensionally correct standard

MMS attachment fixtures and connector shells should be used for

ground demonstrations;

6) On-orbit servicing of MMS modules should be effected by use of

lateral docking with a straight docking probe adapter, tool adapter

and modified stowage rack;

12-2
7) The approach to the exchange of basic and MMS modules was reviewed

and a number of improvements, such as the deletion of a second

docking system and the need for a servicer to failed spacecraft

electrical interface were identified.

12.3 GROUND DEMONSTRATIONS

The following conclusions and recommendations were developed during the

ground demonstration analyses:

1) The servicer system Engineering Test Unit should be used as the

mechanism for early ground demonstrations;

2) Continue the ability to demonstrate separately the exchange of both

basic and MMS modules;

3) The control system software of the MSFC servicing demonstration

facility has been upgraded;

4) MMS module exchange under computer control has been demonstrated;

5) Control mode analysis and testing for exchange of both module types

should be continued;

6) Approaches for the cargo-bay demonstration and for free-flight

verification should be developed;

7) Fluid resupply hardware should be developed and the process

demonstrated;

8) The exchange of batteries or other individual components should be

demonstrated along with thermal blanket/access cover removal and

replacement;

9) An automatic target recognition and error correction system should

be developed and demonstrated;

12-3
10) The MSFC servicing demonstration facility should be made available

for support of flight operations in terms of simulations,

procedures development, training, and problem solving. The

facility should also be made available as a laboratory development

tool;

11) The exchange of other generic modules—AXAF and communications

satellite—should be coordinated with the respective project

offices and then demonstrated;

12) Additional development areas include:

- Demonstration of other servicing tasks specific to Space

Station operations,

- Continuation of interface mechanism development,

- Continuation of adapter tool development as needs are

identified,

- Special fluid resupply disconnects for cryogenics or high

pressures, and self aligning electrical connectors,

- Development of in-line fluid couplings for replacement of tanks

and other propulsion system components.

12.4 CARGO-BAY DEMONSTRATION

The following conclusions and recommendations were developed during the

cargo-bay demonstration analyses:

1) A proto-flight quality servicer mechanism should be built for use

in the single cargo-bay demonstration flight;

2) The MMS Flight Support System should be used to support the MMS

spacecraft representation during the cargo-bay demonstration;

3) The Orbiter Remote Manipulator System end effector should be used

for a docking system;

12-4
4) A specific arrangement of servicing demonstration elements in the

Orbiter cargo bay was selected and recommended for use;

5) The servicer control station should be on the Orbiter aft flight

deck;

6) The servicer should be exercised in the unassisted Supervisory

control mode;

7) The characteristics of the recommended servicer cargo-bay

demonstration are:

- MMS mockup dock and undock by RMS,

- Supply of power, attitude control, thermal control and

communications by Orbiter,

- Servicer control station in Orbiter,

- Docking rigidization by servicer docking probe,

- Electrical connection between servicer and spacecraft via the

.docking mechanism,

- Use of MMS triangular module support structure,

- Module exchange demonstration,

- Fluid resupply demonstration,

- Servicing equipment performance demonstration,

- Unassisted Supervisory control mode,

- Man-machine interaction evaluations,

- Compliance with Orbiter system safety requirements,

- Servicer spare module stowage rack mounted in trunnions in

Orbiter cargo bay,

- Use of representative servicing operational equipment,

- Operator training;

8) The hardware for the fluid resupply demonstrations should be

obtained from the ongoing Johnson Space Center refueling

demonstration flight program;

12-5
9) The recommended activities for the test flight are:

- The replacement of a Multi-Mission Modular Spacecraft type

module using an MMS Module Servicing Tool, incorporating an

electrical connector, and mounted so that the module moves

axially,

- The replacement of a battery module on a light weight side

interface mechanism using an electrical connector and with a

near-radial module motion direction,

- The transfer of a fluid using a multiple line fluid resupply

module including a fluid interface unit and a hose and cable

management device mounted in a far-axial direction;

10) The cargo-bay demonstration servicer mechanism, after its flight

use, should be used to replace the ETU for ground demonstrations,

procedures development, and operator training.

12.5 FREE-FLIGHT VERIFICATION

The following conclusions and recommendations were developed during the

free-flight verification analyses:

1) A fully operational servicer system that has been qualified and

documented should be built for use in the free-flight verification

activity;

2) The Orbital Maneuvering Vehicle should be the servicer carrier

vechicle;

3) One servicer system should be built;

4) The unassisted Supervisory control mode should be used;

5) The servicer control station should be located on the ground;

6) A spacecraft bus, such as the SPAS-01, should be rented rather

than a new spacecraft being built for this one-time application;

12-6
7) The characteristics of the recommended servicer free-flight

verification are:

- One verification flight,

- Serviceable satellite mockup supported by a rented spacecraft

bus,

- Supply of power, attitude control, communications, and thermal

protection and control of the servicer from the OMV,

- Use of OMV for rendezvous and docking of servicer to the

serviceable spacecraft mockup,

- Use of serviceable spacecraft mockup and modules from cargo-bay

demons tration,

- Two way communication links to ground through TDRSS,

- Servicer control station at OMV ground control station,

- Docking rigidization by servicer docking probe,

- Deployment of stowed servicer mechanism and docking probe,

- MMS module exchange demonstration,

- Fluid resupply demonstration,

- Servicing equipment performance verification,

- Control mode verification,

- Operator training;

8) Demonstration of the mating of the servicer stowage rack to the OMV

should be a part of the Space Station technology development

missions;

9) The recommended flight verification activities are:

- Exchange of MMS module,

- Exchange of other representative modules,

- Fluid transfer.

12.6 SERVICER/MMS 1-g DEMONSTRATION PLAN

The following conclusions and recommendations were developed during the

preparation of the servicer/MMS 1-g demonstration plan:

12-7
1) The servicer/MMS 1-g demonstration subsystem requirements were

identified for the MMS module mockup, spacecraft mockup, stowage

rack mockup, electrical connector positioner mechanism, and optical

targets;

2) A preliminary system concept design was performed and the relative

positions of the main components were established;

3) The location and orientation of the MMS module mockup during the

FLIP sequence were selected, based on requirements;

4) An analysis of the unbalanced torques acting on the ETU drive

motors during the ground demonstrations of MMS module exchange

showed that all the loads are within the existing ETU configuration

capability;

5) Several characteristics of the servicer/MMS demonstration equipment

were selected:

- MMS bolt tightening torque of 10 + 1 ft-lbs and loosening torque

of 20 + 1 ft-lbs,

- Maximum torque of 50 ft-lbs for the wrist pitch (Y) drive of ETU,

- Maximum weight of 12.5 Ibs for MMS module mockup,

- Maximum distance of 7.25 in. between the end effector interface

and module latch interface,

- Maximum weight of 15 Ibs for the modified MST;

6) A light weight configuration and a structural concept were selected

for the MMS module mockup;

7) A partial mockup, representing one quarter of the MMS module, was

designed and built as a development version to validate the concept;

8) A simple, straightforward configuration was selected for the

spacecraft mockup, that emphasizes the MMS module while providing

realistic MMS servicing trajectories and preserving the existing

basic module exchange capability;

12-8
9) The arrangement of the MMS module mockups, basic module mockups and

MST storage rack on the ETU stowage rack was selected based on:

- Minimum modification of the existing stowage rack,

- Minimum MMS servicing demonstration time,

- No system reconfiguration between MMS module and basic module

exchange demonstrations.

12.7 SERVICER/MMS 1-g DEMONSTRATION EQUIPMENT

The following conclusions and recommendations were developed as part of

the servicer/MMS demonstration equipment design and fabrication

activities:

1) The drawings for the production of the Change Order 3 equipment

were prepared at Form 4 level;

2) The design effort included:

- Drawing preparation,

- Coordination of MST integration,

- Design coordination,

- Materials and components procurement;

3) The MMS interface frame design is common for both spacecraft mockup

and stowage rack supports;

4) The MST latch interface box design is the same for both MMS module

fasteners and for the MST storage rack;

5) The connector positioner mechanism features:

- A compact, eccentric type mechanism,

- Accurate linear ball slide,

- 5/8 in. mating stroke,

- 20 Ib connector mating/demating force,

12-9
- Adjustable position for end of stroke,

- Simple interface with ETU end effector;

6) The optical targets feature:

- Common design for all MMS fastener locations and for the MST

storage rack,

- Compliant attachment to its support,

- Minimal resetting in case of accidental displacement;

7) The weight of the fabricated and assembled MMS module mockup is

10.0 Ibs, compared to the 12.5 Ibs maximum design limit;

8) The fabricated and assembled connector positioner mechanism:

- Was tested on a special bracket prior to shipment to MSFC,

- Smoothly mated and demated the electrical connector,

- The mating and demating times were within the design goals.

12.8 SERVICER CONTROL SOFTWARE - BASIC MODULES

The following conclusions and recommendations were identified during

development of the servicer control software for the demonstration of

basic module exchange:

1) Three control modes were implemented;

2) Geometrical equations specific to the ETU configuration were used;

3) Software can be used on PUP 11/34 computer with Version 3.2 or 4.0

of the RSX-11M operating system;

4) Software requirements were explicitly defined and documented;

5) All required interfaces between the computer and the electrical

equipment were defined and documented;

12-10
6) The characteristics of the Supervisory control mode trajectory

hierarchy for basic modules are:

- Four total trajectories,

- Twenty trajectories,

- Nine steps,

- Eight actions,

- Each hierarchy level is composed of elements below it in the

hierarchy,

- Four types of coordinate transformations,

- Closed loop operation of ETU joints,

- Control of end effector and interface mechanism drives,

- Operator assisted and unassisted modes;

7) Software program is menu driven;

8) Procedures and trajectory sequences for the Manual-Augmented

control mode were documented;

9) Manual-Augmented trajectory sequences use same approach as

Supervisory trajectories;

10) Coordinate transformations and error meter drive signals were

incorporated for Manual-Augmented control mode;

11) Procedures for all demonstration operations were provided;

12) Simulated hardware characteristics are included in software so

program can be run independent of servicer hardware;

13) A test program for verifying the computer to servicer hardware

interfaces was provided;

14) A separate Software User's Manual was prepared for the basic

module software.

12-11
12.9 SERVICES CONTROL SOFTWARE - MMS MODULES

The following conclusions and recommendations were identified during

development of the servicer control software for the demonstration of

MMS module exchange:

1) The MMS module software follows the basic patterns and philosophy

of the basic module software;

2) Three control modes were implemented;

3) Geometrical equations specific to the ETU configuration were used;

4) Software can be used on PDF 11/34 computer with Version 3.2 or 4.0

of the RSX-11M operating system;

5) Software requirements were explicitly defined and documented;

6) All required interfaces between the computer and the electrical

equipment were defined and documented;

7) The characterictics of the Supervisory control mode trajectory

hierarchy for MMS modules are:

- One total trajectory,

- Nine trajectories,

- Thirteen steps,

- Ten actions,

- Each hierarchy level is composed of elements below it in the

hierarchy,

- Four types of coordinate transformations,

- Closed loop operation of ETU joints,

- Control of end effector, connector positioner drive, and MST

latch and bolt drives,

- Operator assisted and unassisted modes;

12-12
8) Software program is menu driven;

9) Procedures and trajectory sequences for the Manual-Augmented

control mode were documented;

10) Manual-Augmented trajectory sequences use same approach as

Supervisory trajectories;

11) Coordinate transformations and error meter drive signals were

incorporated for Manual-Augmented control mode;

12) Procedures for all demonstration operations were provided;

13) Simulated hardware characteristics are included in software so

program can be run independent of servicer hardware;

14) A test program for verifying the computer to servicer hardware

interfaces was provided;

15) A separate Software User's Manual was prepared for the MMS module

software.

12.10 SERVICER SOFTWARE DEMONSTRATIONS

The following conclusions and recommendations were identified during

the conduct of the basic and MMS module exchange demonstrations using

the two servicer software programs:

1) All of the demonstration equipment operated satisfactorily and was

provided by:

- ETU and associated electronics by MSFC,

- PDP 11/34 computer with D/A's and A/D's by MSFC,

- MMS modules, spacecraft mockup, and stowage rack modifications

by Martin Marietta,

- Connector positioner and wiring changes by Martin Marietta,

12-13
- 1-g Module Servicing Tool by GSFC,

- MST electronics by Fairchild Space Co;

2) The overall appearance of the revised spacecraft mockup was

improved by a number of cosmetic changes;

3) The software could be readily installed and checked out on PDF

11/34 computer;

4) Scale factor and zero adjustments were made to obtain satisfactory

pure axial and pure radial motions of the ETU;

5) A number of changes were made in the software logic and in the

interface electronics to improve operations;

6) Specific module location data could be readily collected for use

in the software program and in the Manual-Augmented trajectory

sequences using the procedures that were developed;

7) Separate demonstrations of basic and MMS module exchange were

successfully made in all three control modes;

8) Conduct of demonstrations in the Supervisory control mode in the

operator assisted or unassisted modes was easy to learn.

Operation in the Manual-Augmented control mode takes a little

longer to learn, as was expected;

9) Motion of the ETU during module exchanges in either Supervisory

mode was very smooth and precision was well within the basic

module equipment capture volumes and just within the tighter MMS

equipment capture volumes;

10) Integration of the MST was accomplished by operating philosophy

revisions, software modifications, and hardware adjustments;

11) Non-orthogonality of the MMS module with respect to the docking

12-14
post (axial cylindrical coordinate) when the module top bolt is

tight and the bottom bolt is loose was accommodated by the

addition of a pseudo-combined motion capability where all six

cylindrical coordinates are changed together in a step-wise

fashion to approximate the desired path;

12) System operating techniques were identified for overcoming

anomalies so that the system should not be thought of as a

pre-programmed entity that cannot continue past the first

anomaly. Rather it is a system with three levels of control that

can be used interchangeably to get the job done in spite of a

variety of anomalies.

12-15
A. CARGO-BAY DEMONSTRATION WORK BREAKDOWN STRUCTURE

This appendix presents a work breakdown structure (WBS) and

associated dictionary for the cargo-bay demonstration activity.

The WBS was prepared to give an outline of the organization and

• * ' •"

activities involved and to aid in the identification of cost

sources. The selected WBS follows the pattern used in other

NASA activities of this same general size.

Figure A-l shows the WBS in graphical form. Level 1 is the

project itself, while level 2 shows the other organizations with

which the prime contractor must interface. Level 3 activities

are shown for the prime contract and include the seven major

activities arranged generally in a time sequence format with the

earliest activities to the left. Each of the prime contract

level 3 activities are expanded to level 4 and six of the level

4 activities are further expanded to level 5.

Each of the blocks on Figure A-l are defined in the WBS

dictionary that starts on page A-3. The paragraph number of the

dictionary corresponds to the number in the block of Figure A-l.

A-l
ORIGINAL PA6E tS

OF POOR QUAkitV

ORfGINAL r*

Level 1

Level 2

level 3

Level 4

Cargo B«y

Deaoosuition

Project

1.D

ProJQct

ilanagvnint

(HSfO

1.1

Syston

Engineering [

Inugritrion

(BSfC)

1.Z

Prim

Cantrict

1.3

Dlrborae

Support

Equipiunt

(CSfC)

1.4

«IV

(Inforngtion

only)

1.5

SIS

(Inforution

only)

t.E

Systen

1.3.3.

1 er Sptacrgft

(lockup

1.3.3.2

res,

ins C

.1

Structures,

nedienlsns C

Ihernal

1.3.3.2.1

Airborne

Support

Cquipiient

1.3.3.3

Flight

Support Sys

1.3.3.3.1

Ground Ground Softmrt

Doniuring Support

SUUoo(OlS) tijulphtnt

1.3.3.4 1.3.3.5 1.3.3.S

«irtornc

Support Equip

Dstea ( C/0

1.3.5.3

Structures,

nechinisns (

Tnernil

1.3.4.2.1

Figure A-l Cargo-Bay Demonstration Work Breakdown Structure

Fouxxrr A-2

FRAM£
SPACECRAFT SERVICER CARGO-BAY DEMONSTRATION

WBS DICTIONARY

1.0 SPACECRAFT SERVICER CARGO-BAY DEMONSTRATION PROJECT (Level 1)

This WBS element provides for all the necessary manpower,

materials, tooling and equipment, hardware, facilities, and

service required to design, develop, fabricate, assemble, test,

checkout, integrate and support the Spacecraft Servicer CargoBay Demonstration project. This includes all management,

engineering, and scientific activities of the entire project

both civil service and contractor.

1.1 PROJECT MANAGEMENT (Level 2)

This WBS element encompasses all project office (civil service)

activities and responsibilities including overall NASA mission

management of the entire Spacecraft Servicer Cargo-Bay

Demonstration project, including, all servicer hardware

development, integration, and operations.

1.2 SYSTEMS ENGINEERING AND INTEGRATION (Level 2)

This element includes all project office (civil service)

manpower required to ensure overall servicer performance,

including all servicer hardware and operational aspects of the

project.

A-3
1.3 SPACECRAFT SERVICER CARGO-BAY DEMONSTRATION PRIME CONTRACTOR

(LEVEL 2)

This element provides for the necessary manpower, materials,

equipment, hardware, facilities and services required to design,

develop, fabricate, assemble, test, checkout and provide

operational support to the Spacecraft Servicer Cargo-Bay

Demonstration project. Specifically, the development of the

servicer along with the necessary flight support equipment,

software, CMS, GSE, and overall system integration is included.

1.3.1 Project Management (Level 3)

This element encompasses the technical and business management

activities required to plan, execute, control and report

technical performance, schedules, and resources in a

cost-effective manner.

1.3.1.1 Planning and Control (Level 4) - This element provides

for the planning, authorizing and controlling of the contracted

effort. The planning and control activity shall provide timely

visibility into contract performance, cost, and schedule, and

shall be keyed to the contractor organization and to the Work

Breakdown Structure (WBS) and WBS Dictionary. Integrated

project schedules for the servicer are included herein. The

control function shall integrate cost, schedule, and performance

and shall relate progress and variance from the initial planning.

1.3.1.2 Data Management (Level 4) - This element pertains to

activities required to assure proper information control,

compatibility, availability, and relevancy; and to prepare and

deliver contractually required data to the Government. Included

are activities to identify, control and monitor the preparation,

reproduction, distribution and maintenance of internal and

deliverable documentation under this contract. (See list of

Contract Deliverable Items.)

A-4
1.3.1.3 Procurement Management (Level 4) - This element

includes responsibility for all subcontract activities,

procuring hardware as defined by procurement requirements,

assuring that selected subcontractors and suppliers perform and

deliver in accordance with approved requirements, providing the

project with performance surveillance, cost control, technical

direction, and status reporting on all subcontractors.

1.3.1.4 GFE Management (Level 4) - This element provides for

all tasks required to maintain a Government Furnished Equipment

(GFE) management for the Spacecraft Servicer Cargo-Bay

Demonstration project. This system shall be in accordance with

the contractor's existing standard operating procedures and in

compliance with the NASA procurement regulations. The system

includes identification of GFE items and required-delivery-dates

planning, control and technical monitoring. See equipment list

for GFE equipment.

1.3.2 Systems Engineering And Integration (Level 3)

This WBS element includes all labor, materials, and other

resources necessary to perform all the systems engineering

functions required to assure servicer performance. It includes

the effort necessary to define the requirements and

specifications needed to conduct the design and analysis of the

servicer, airborne and ground support equipment and the ground

monitoring station. Activities include analyses and trade

studies of system, system configuration and operations support

requirements. It also includes activities required to define

interfaces between all of the elements of the demonstration.

A-5
1.3.2.1 System Requirements and Analysis (Level 4) - This

element provides for the definition of technical requirements

necessary to conduct the design and development of the servicer,

airborne and ground support equipment, and the ground monitoring

station. It also includes analyses and trade studies required

to evaluate and define the detailed requirements of the

integration with its external interfaces including the STS.

This task will generate and maintain the Program Integration

Plan (PIP).

1.3.2.2 System Configuration (Level 4) - This WBb element

provides for the effort required to plan, implement, control,

and coordinate the servicer configuration in accordance with the

contractural requirements and specifications and provide the

related documentation required to support the reviews.

1.3.2.3 Safety Requirements and Analysis (Level 4) - This WBS

element provides for all effort required to perform analyses and

preparation of the documentation for the Safety Review Process.

A summary of the review process is provided in Table 1.3.2.3-1.

1.3.2.4 Operations Support Requirements (Level 4) - This

element provides for the definition of detailed launch site and

mission operations support requirements. It includes detailed

functional analyses to identify both pre-flight and post-flight

operations support as well as mission planning and post-mission

support analysis.

1.3.3 Design and Analysis (Level 3)

This WBS element includes all labor, materials, and other

resources required to perform the analyses, design, design

trades, and development testing activities for the cargo bay

demonstration.

A-6
Table 1.3.2.3-1 SUMMARY OF SAFE1Y REVIEW PROCESS

PHASE

0

I

II

III

TIMING

Payload/GSE

conceptual

design established

Payload/GSE

preliminary

design established

Payload/GSE

final design

established

Payload/GSE

fabrication

and testing

complete

PAYLOAD ORGANIZATION'S

SAFETY EFFORTS

1. Perform preliminary

safety analysis.

2. Prepare a ground operations concept (KSC).

1. Define and expand safety

analysis to reflect the

preliminary design:

a. Define hazards.

b. Define hazard causes.

c. Evaluate actions for

reducing or controlling

hazards.

d. Identify approach for

safety verification.

2. Prepare a mission (or ground

operations) scenario.

3. Determine compliance with

NHB 1700.7 and KHB 1700.7.

1. Refine and expand safety

analysis.

a. Evaluate interfaces

and mission (or

ground operations)

procedures, plans

and timelines.

b. Update hazard descriptions, causes, and

controls.

c. Finalize test plans,

analysis procedures, or

inspections for safety

verification.

2. Finalize description of

ground operations flow.

3. Determine compliance with

NHB 1700.7 and KHB 1700.7.

1. Complete safety analysis.

2. Prepare safety assessment

report.

3. Complete all safety verification tests, analyses,

and/or inspections.

4. Prepare safety compliance

data package.

PURPOSE OF REVIEW

1. Identify potential

hazards and applicable

safety requirements.

1. Assess the preliminary

design against NHB

1700.7 and KHB 1700.7.

2. Evaluate preliminary

hazard controls and

safety verification

methods.

1. Assess final design

against NHB 1700.7 and

and KHB 1700.7.

2. Concur on specific

hazard controls and

safety verification

methods.

1. Approval of safety

assessment report.

2. Review of safety compliance data package.

3. Identify open safety

items .

A-7
The hardware is divided into the servicer system, spacecraft

mockup, airborne support equipment, ground monitoring station

and ground support equipment. The preparation of design

drawings, parts lists, material analyses, wiring diagrams, and

the testing of critical components and subsystems is included in

this element.

1.3.3.1 Servicer System (Level 4) - This element includes all

labor, materials, and other resources required to perform the

analyses, design, design trades, development and qualification

testing activities for the servicer system. The servicer system

consists of the servicer unit, stowage rack, docking probe and

the video equipment. Although the modules reside at both the

stowage rack and the spacecraft, they are contained within the

spacecraft element. The fluid resupply equipment is divided

between the servicer system and the spacecraft mockup.

1.3.3.1.1 Structures, Mechanisms and Thermal Control (Level 5) -

This element provides for the design and development of the

servicer system structural and mechanical subsystem. The

specific tasks include:

a) Structural Analysis - The servicer system will be

overdesigned such that the safety requirements for the

operating load, emergency landing loads, and the cycle life

can be verified through the structural analysis. The

analysis must be of such detail as to fulfill this

requirement. The analyses will be performed for the case

of servicer arm locked in the stowed configuration as would

be the case in either the launch or landing. Analyses will

be performed for other cases identified in the safety

analyses.

A-8
b) Mechanisms - The mechanisms of the servicer system will be

verified through reviews of the design, inspection and

acceptance testing. Analyses will support the design

reviews and inspection portion of the verification

process. The modified module servicing tool (MST) will be

supplied by GSFC, who will perform any necessary

development or qualification testing.

c) Thermal Analysis - A thermal analysis will be performed to

determine the operating temperatures and size the heaters.

Detailed temperature profiles and time histories are not

required. Analyses will be performed for several steady

state cases including the expected extreme conditions, in

order to bound the problem. Models with a limited number

of nodes will be used and transient simulations will not be

performed.

d) Fluid Resupply System - Analyses and testing will be used

to verify the fluid system of the fluid resupply system.

Qualification testing will be used to verify containment

adequacy at the operating pressure. Analyses will be used

for the capacity, operating pressure, proof pressure, burst

pressure and cycle life.

1.3.3.1.2 Control and Electronics (Level 5) - The design

analyses of the servicers instrumentation, controls and

electronics shall support verification through reviews of the

design and acceptance testing. The interface between the

servicer system and the STS will be made through a NASA supplied

Standard Umbilical Release/Retract/Retention System (SURS),

which has been previously flight qualified.

1.3.3.1.3 Data Handling (Level 5) - This element provides for

the design and development of the servicer system data handling

subsystem excluding the design and development of the software.

A-9
1.3.3.1.4 Electrical Power Distribution (Level 5) - This

element provides for the design and development of the servicer

system electrical power and distribution subsystem, including

provision for the electrical power interface to the flight

support system. The STS interface will consist of a NASA

supplied SURS which is flight qualified.

1.3.3.1.5 Video (Level 5) - This element provides for the

design and development of the servicer system video and

alignment sensor subsystems. Efforts will be made to use a

previously flight qualified camera. The alignment sensor will

be developed as part of this program.

Analyses and development testing will be required for

verification.

1.3.3.2 Spacecraft Mockup (Level 4) - This element includes all

labor, materials, and other resources required to perform the

analyses, design, design trades, and development testing

activities for the Spacecraft Mockup. The spacecraft mockup

consists of a Multi-Mission Modular Spacecraft (MMS) structure,

MMS module, a fluid resupply module, battery module and grapple

fixtures. The MMS structure and modules are flight qualified

and will be supplied by GSFC. Fully developed fluid resupply

interfaces will be obtained from either within Martin Marietta

or NASA. The hardware to be built within this element consists

of the battery module, the fluid resupply module (storage tanks,

supply lines and supply system) and the necessary integration

hardware.

1.3.3.2.1 Structures, Mechanisms and Thermal Control (Level 5)-

This element provides for the design and development of the

spacecraft mockup structural and mechanical subsystem. The

specific tasks include:

A-10
a) Battery Module Structural Analysis - The structural

analysis of the battery module will be used to verify the

operating loads, emergency landing loads and cycle life.

The module will be tested for random vibration as part of

the servicer system.

b) Fluid Resupply Module Structural Analysis - The structural

analysis of the fluid resupply modules will be used to

verify the operating loads, emergency landing loads and

cycle life. The module will be tested for random vibration

as part of the servicer system.

c) Fluid Resupply Module Fluid System - Analyses and testing

will be used to verify the fluid system of the fluid

resupply module. Qualification testing will be used to

verify containment adequacy at the operating pressure.

Analyses will be used for the capacity, operating pressure,

proof pressure, burst pressure and cycle life.

1.3.3.2.2 Control and Electronics (Level 5) - The design

analyses of the spacecraft mockup instrumentation, controls and

electronics shall support verification through reviews of the

design and acceptance testing.

1.3.3.2.3. Data Handling (Level 5) - This element provides for

the design and development of the spacecraft mockup data

handling subsystem excluding the design and development of the

software.

1.3.3.2.4 Electrical Power Distribution (Level 5) - This

element provides for the design and development of the

spacecraft mockup electrical power and distribution subsystem

including provision for the electrical power requirements and

the electrical power interface to the servicer system.

A-ll
1.3.3.3 Airborne Support Equipment (Level 4) - This element

includes all labor, materials, and other resources required to

perform the analyses, design, design trades, and development

testing activities for the servicer airborne support equipment.

The airborne support equipment consists of the flight support

system, STS control station and the SIS cargo bay interface.

1.3.3.3.1 Flight Support System (Level 5) - This element

provides for the analysis of the existing Multi-Mission Modular

Spacecraft Flight Support System (FSS) and establishes the

interfaces between it and the servicer system, the spacecraft

mockup, and the STS cargo bay. The MMS flight support system is

flight qualified and will be supplied by GSFC.

1.3.3.3.2 STS Control Station (Level 5) - This element provides

for the definition of the STS" control and display station

interfaces. The control station will consist of a control

computer (equivalent to a militarized AT personal computer), a

safety shield to prevent damage should the CRT explode and a

control panel. The computer and control panel will require only

analysis, while development testing will verify the safety

shield.

1.3.3.3.3 STS Cargo Bay Interface (Level 5) - This element

provides the definition of the equipment needed to interface the

servicer and the flight support system into the STS cargo bay.

NASA supplied interfaces will be used.

1.3.3.4 Servicer Ground Monitoring Station (CMS) (Level 4) -

This element includes all labor, materials, and other resources

required to perform the analyses, and trades studies for the

servicer ground monitoring station.

A-12
1.3.3.5 Servicer Ground Support Equipment (Level 4) - This

element provides for the effort required to design, develop, and

test the ground support equipment required for the servicer

launch site integration activities. Included are transportation

and handling equipment, assembly equipment, mock-ups, and other

non-flight miscellaneous hardware.

1.3.3.6 Software (Level 4) - This element provides for the

design development, checkout and maintenance of the flight and

ground software required for the servicer. The flight software

includes all on-board software required for in-flight command

control and data handling for the servicer. Ground software

includes the software required to support the development and

ground checkout of the servicer and to support launch and

mission operations. Development testing will be used to verify

the software.

1.3.4 Hardware Procurement, Fabrication, Assembly & Checkout (Level 3)

This WBS element includes all labor, materials, and other

resources necessary to procure, fabricate, assemble and check

out subsystems, excluding software. Included is the design and

development of all shop drawings, the procurement of purchased

parts and materials, and the handling and staging of material

and purchased parts. This element includes all effort

associated with the manufacture and procurement of test hardware.

1.3.4.1 Servicer System (Level 4) - This element includes all

labor, materials, and other resources necessary to procure,

fabricate, assemble and check out the servicer system.

A-13
1.3.4.1.1 Structures, Mechanisms and Thermal Control (Level 5)-

This element provides for the effort necessary to procure,

fabricate, assemble, and check out the structural and mechanical

subsystems. Qualification tests will be performed for random

vibrations of the servicer system. Included in the system to be

tested are the servicer system, MMS module, fluid resupply

module, and battery module. The servicer arm will be in the

stowed position.

1.3.4.1.2 Control and Electronics (Level 5) - This element

provides for the effort necessary to procure, fabricate,

assemble, and check out the control interfaces and electronic

components.

1.3.4.1.3 Data Handling (Level 5) - This element provides for

the effort necessary to procure, fabricate, assemble, and check

out the data handling subsystem.

1.3.4.1.4 Electrical Power Distribution (Level 5) - This

element provides for the effort necessary to procure, fabricate,

assemble and check out the electrical power distribution

subsystem.

1.3.4.1.5 Video (Level 5) - This element provides for the

effort necessary to procure, fabricate, assemble, and check out

the video and alignment sensor subsystems.

1.3.4.2 Spacecraft Mockup (Level 4) - This element includes all

labor materials, and other resources necessary to procure,

fabricate, assemble and check out the spacecraft mockup

subsystems.

1.3.4.2.1 Structures, Mechanisms and Thermal Control (Level 5)-

This element provides for the effort necessary to procure,

fabricate, assemble, and check out the structural and mechanical

subsystem of the spacecraft mockup.

A-14
1.3.4.2.2 Control and Electronics (Level 5) - This element

provides for the effort necessary to procure, fabricate,

assemble and checkout the control interfaces and electronic

components of the spacecraft mockup.

1.3.4.2.3 Data Handling (Level 5) - This element provides for

the effort necessary to procure, fabricate, assemble, and check

out the data handling subsystem of the spacecraft mockup.

1.3.4.2.4 Electrical Power Distribution (Level 5) - This

element provides for the effort necessary to procure, fabricate,

assemble and check out the electrical power distribution

subsystem of the spacecraft mockup.

1.3.4.3 Airborne Support Equipment (Level 4) - This element

provides for the effort necessary to procure, fabricate,

assemble and check out the servicer airborne support equipment

subsystems.

1.3.4.3.1 Flight Support System (Level 5) - This element

provides for the effort necessary to interface with GSFC and

check out the flight support system and its interfaces to the

spacecraft mockup, servicer unit and STS cargo bay.

1.3.4.3.2 STS Control and Display Station (Level 5) - This

element provides for the effort necessary to procure, fabricate,

interface and checkout the servicer STS control and display

subsystem. Development testing of a CRT safety shield will be

performed as part of the element.

1.3.4.3.3 STS Cargo Bay Interface (Level 5) - This element

provides for the effort necessary to assemble, integrate and

check out the STS cargo bay interface subsystem provided by NASA.

1.3.4.4 Servicer Ground Monitoring Station (CMS) (Level 4) -

This element provides for the effort necessary to interface and

check out the servicer ground monitoring station subsystems.

A-15
1.3.4.5 Ground Support Equipment (Level 4) - This element

provides for the effort necessary to procure, fabricate,

assemble, and check out the servicer ground support equipment

subsystems.

1.3.5 System Assembly and Check out (Level 3)

This element includes all labor, materials and other resources

required to assemble the servicer and related equipment and then

conduct system level acceptance testing.

1.3.5.1 Servicer System Assembly and Check Out (Level 4) - This

system element provides for the resources to assemble the

servicer system and conduct check out tests.

1.3.5.2 Spacecraft Mockup Assembly and Checkout (Level 4) -

This element provides the resources to assemble the spacecraft

mockup and conduct checkout tests.

1.3.5.3 Airborne Support Equipment Assembly and Check Out

(Level 4) - This element provides for the resources to assemble

the airborne support equipment and to conduct check out tests.

1.3.5.4 Servicer Ground Monitoring Station Assembly and Check

Out (Level 4) - This element provides for the resources to

assemble and check out the servicer ground monitoring station.

1.3.5.5 Servicer Ground Support Equipment Assembly and Check

Out (Level 4) - This element provides for the resources to

assemble the servicer ground support equipment and to conduct

check out tests.

1.3.5.6 System Testing (Level 4) - This element provides for

the planning, documentation, materials and services required to

accomplish all system-level final acceptance testing of the

servicer system, spacecraft mockup and the flight support system.

A-16
1.3.6 Launch Operations Support (Level 3)

This element provides for the support required to plan, develop

and execute pre-flight and post-flight ground operations support

of the servicer. This activity shall include any training,

operations, integration activities and documentation support for

the servicer off-line activities as well as procedure

development and implementation support to the launch site for

on-line activities for the servicer checkout and integration.

1.3.6.1 Pre-Flight Operations Support (Level 4) - This WBS

element includes all labor, materials and other resources

required to support the pre-flight operations activity.

Included is all effort required to plan, organize, and execute

the pre-flight integration, checkout and handling of the

servicer and support equipment.

1.3.6.2 Post-Flight Operations Support (Level 4) - This element

includes all labor, materials, and other resources required to

support the post-flight operations activity. This task includes

all post-flight "safing" of the servicer hardware, removal from

the STS, inspection and checkout, preparation for shipment (if

applicable), and analysis and documentation of support

activities.

1.3.7 Mission Operations Support (Level 3)

This element includes all labor, materials, and other resources

required to support the mission operations activity. Included

is the support to mission planning, on-orbit mission operations,

ground control and flight crew training support and post-mission

analysis.

1-3.7.1 Mission Planning Support (Level 4) - This element

includes all labor, materials, and other resources required to

develop the sequencing of flight events necessary to complete

the prescribed mission. It also includes contingency planning

to respond to unprogrammed mission events.

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1.3.7.2 Mission Support (Level 4) - This element includes all

labor, materials, and other resources required to support ground

command, telemetry, tracking and communications during the

mission. Also included is the development of flight procedures

for activation, operation, and deactivation of the servicer as

well as procedures for trouble shooting of anomalies and

failures. In addition, all mission support contractor crew

training is included in this WBS element.

1.3.7.3 Post-Mission Analysis (Level 4) - This element includes

all labor, materials, and other resources required to analyze

and document the data generated during the mission relative to

the engineering instrumentation flown on the mission.

1.4 AIRBORNE SUPPORT EQUIPMENT (Level 2)

GSFC shall provide the Airborne Support System, MMS structure,

and MMS modules.

1.5 OMV (Level 2)

This element is to provide/obtain information only.

1.6 STS (Level 2)

This element is to provide/obtain information only.

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