^NASA-CB-178821) SPACE PLiTFCfiia EXPENDABLES N86-24734

EESUPPLY CCNCIPT DEFINITION STUDY, VOLUME 2

Technical Report, Mar. - Dec, 1S84 (Bocksell

Internatioaal Corp.) 295 p BC A13/MP A01 Unclas

. CSCL 22B G3/18 «2939

SPACE PLATFORM EXPENDABLES RESUPPLY

CONCEPT DEFINITION STUDY

STS 85-0174

VOLUME II

TECHNICAL REPORT

CONTRACT NAS8-35618

FOR PERIOD MARCH 1984 - DECEMBER 1984

2082e/
SPACE PLATFORM EXPENDABLES RESUPPLY

CONCEPT DEFINITION STUDY

STS 85-0174

VOLUME II

TECHNICAL REPORT

CONTRACT NAS8-35618

FOR PERIOD MARCH 1984 - DECEMBER 1984

2082e/
This final report of the Space Platform Expendables Resupply Concept

Definition study was prepared by the Advanced Engineering organization of the

Space Transportation Systems Division-of Rockwell International Corporation

for the Marshall Space Flight Center (MSFC) of NASA in accordance with

Contract NAS8-35618. The study was conducted under the direction of the MSFC

Contracting Officer Representative (COR), Mr. Wilbur Thompson, during the

period from March, 1984 through December, 1984. The final report is organized

into four documents:

Volume I - Executive Summary

Volume II - Study Results

Volume III - WBS and Dictionary

Volume IV - Cost Estimate

Personnel who made major contributions to the study effort are:

V. A. Weldon - Study Manager

R. M. Hayes - Effectiveness Analysis & Programmatics

R. J. Budica - Systems Engineering

S. N. Pace - Requirements

Y. Brayer - Requirements

F. 0. Chandler - Operational & Demonstration Resupply System Design

D. R. Slocum - Business Management

W. A. McClure - GEO Servicing Effectiveness Analysis

M. Plttman - Mission and Fluid Parametric Requirements

P. Forrester - Configuration Design

L. Livingston - Thermal Analysis

V. Mora - Propellant Feed Analysis

J. Nakamoto - Propellant Component Analysis

J. J. Gernand - Related IR&D

The study was also supported by a subcontract to Science Applications, Inc.

personnel, under the leadership of Dr. Brian O'Leary. Useful information was

also received from SPAR Incorporated, Seton-Wilson Corporation, and MBB-ERNO.

2065e
TABLE OF CONTENTS

1.0 Introduction 1

1.1 Background 1

1.2 Objectives and Scope 1

1.3 Guidelines & Assumptions 2

1.4 Study Flow and Methodology 2

2.0 Summary of Results 3

2.1 Major Conclusions 3

2.2 Selected Options 4

2.3 Recommended Approaches 6

3.0 Study Results 7

3.1 Requirements and Benefits Analysis 7

3.1.1 Mission Model & Fluid Parametric Requirements 8

3.1.2 Candidate Spacecraft & Resupply Scenarios 25

3.1.3 Effectiveness Analysis 32

3.1.4 Design Reference Mission (DRM)/System Requirements

Selection 81

3.2 Concept Design (Task 2) 95

3.2.1 Operational System Configuration Selection 95

3.2.2 Safety and Contamination Control Requirements and

Considerations 99

3.2.3 Pumped vs. Pressure Regulated Propellant Transfer 102

3.2.4 Propellant Transfer Process Selection 105

3.2.5 Conceptual Designs for Ullage Exchange

Bi-Propellant and Pressurant Transfer Process 110

3.2.6 Propellant Transfer Pumps 115

3.2.7 Pressurant Transfer 123

3.2.8 Battery Power Sources for Pumps and Compressors 132

3.2.9 Thermal Management of Pump and Compressor Systems 136

3.2.10 Pressure Drop Analysis 143

3.2.11 Propellant Management Considerations 148

3.2.12 Propellant Quantity Gauging & System Health Monitoring 158

3.2.13 Quick Disconnect Selection 162

3.2.14 Avionics & Power Subsystems 168

3.2.15 Docking Hardware Selection 174

3.2.16 ERM System Weights 182

3.2.17 Receiver Spacecraft Impact Assessment 182

3.2.18 Advanced Development Requirements 192

3.2.19 Demonstration Flight Program 194

3.3 Programmatic and Development Planning 197

3.3.1 Ground Rules & Assumptions 197

3.3.2 Advanced Development and Demonstration Flight 198

3.3.3 Work Breakdown Structure 202

3.3.4 Baseline Program 206

3.3.5 ERM Developmental Test Program 208

3.3.6 Qualification Test Program 209

3.3.7 Build and Flow Plan 211

3.3.8 Advanced Development and Flight Demonstration Cost 211

3.3.9 Baseline Program Cost 214

3.3.10 Cost Estimate for Satellite Mod Kit to Enable Resupply 219

Appendix 225

Expendables Resupply Module Requirements Definition Document

2082e ii
ILLUSTRATIONS

Figure Page

1 Deleted

2 Study Flow 2

3 Selected Mission Scenarios 5

4 Resupply Module General Configuration (Without Servicer) 6

5 Resupply Module General Configuration (With Servicer) 6

6 Bi-Propellant/Helium Transfer Flight Demonstration Concept 7

7 Task 1 Activities 8

8 Low Earth Orbit Alternative Fluids Resupply Scenarios 27

9 Resupply Module Basing Options 28

10 Summary of Basing Mode Conclusions 29

11 SIRTF Servicing Mission Timeline (Ground Based OMV) 28.5° Orbit 30

12 Deleted

13 GEO Platform Resupply Using Centaur G 31

14 DOD-1 and DMSP Multi-Servicing Mission Timeline 32

15 Effectiveness Analysis Groundrules and Assumptions . 35

16 Satellite System Availability Through Servicing 36

17 Expendable Resupply Relationship for Spacecraft Design Life 37

18 Design Life (Consumables Depleted) Estimating Gross Economic

Benefit and Resupply Without Servicing 38

19 Relative Residual Value of Generic Satellites 40

20 Potential Need of Surveillance Programs for On-Orbit Remote

Consumables Resupply 40

21 Net Economic Benefits Summary for NASA LEO ETR Spacecraft 42

22 Economic Benefits of Full Servicing Relative to Resupply Only

for Selected LEO Satellites • 43

23 GEOSAT ACS Annual Usage 46

24 Equivalent GEOSAT Transponders as a Function of Communications

Payload Weight 48

25 Transponder Capability Growth 50

26 Transportation Price as a Function of GEOSAT Weight 51

27 GEO Comsat Lifetime Extension 56

28 IRR of Lifetime Extension 57

29 Lifetime Extension Options (10-15 yrs) Relative Annual Annuities (A) 58

30 Lifetime Extension Options (10-15 yrs) Relative Annual Annuities (B) 59

31 GEOSAT Contingency Resupply Vs. Replacement 60

32 Internal Rate of Return GEOSAT Contingency Resupply Vs. Replacement 61

33 Transponder Replacement Through Resupply 63

34 GEOSAT Transponder Replacement 64

35 GEO Resupply & Servicing Std. Transponder Mix 68

36 GEO Resupply & Servicing Transponder/ACS Tradeoff 69

37 Insurance Has a Significant Impact on GEO Resupply/Servicing 70

38 Effect of GEO Servicing Cost/Lb on Satcom Values 72

39 Effect of Satcom Obsolescence Rates on GEO Resupply/Servicing 73

40 Impact of Varying Levels of Revenue Restoration from Servicing 75

41 Impact of Charges in Transponder Prices 76

42 Lifetime Extension Options (10-15 yrs) 78

43 Equivalent GEO P/L Capability With OMV/RM Reusable Perigee Stage 79

44 Representative Resupply Module/OMV Operations in Conjunction With

LEO Depot and Reusable Perigee Stage 80

2082e iii
ILLUSTRATIONS (Cont'd.)

Figure

45 LEO Remote Refueling of Storable Stage Candidate Scenario 81

46 Low Earth Orbit Alternative Fluid Resupply Scenarios 82

47 Relationship Between Considerations for Minimizing Resupply Cost

to the Potential User 83

48 Expendables Resupply Module ROM Acquisition Cost 85

49 Relative Cost Effectiveness Estimates for Selected LEO Resupply

Scenarios 87

50 Risk Considerations for Scenario Evaluation 88

51 AHP Ranking of Risk Evaluation Criteria for Selected LEO Resupply

Scenarios 90

52 Risk/Economic Benefit Array for Candidate LEO Resupply Scenarios 91

53 Efficient Set of Candidate LEO, Low Inclination Design Reference

Missions 92

54 Estimated Risk Versus Net Economic Benefit for Selected DRMs Without

Concurrent Servicing 94

55 Mid-Term Review DRM Recommendations 94

56 Summary of Resupply Module Scenarios/Fluid Transfer Requirements 96

57 RM/OMV Configuration 97

58 Resupply Structural Concept . 98

59 Example Disconnect/Line Purge and Waste Storage Subsystem 103

60 Ullage Recompresslon Propellant Resupply 107

61 Ullage Recompression Thermal and Pressure Profiles 108

62 Ullage Exchange Propellant Resupply 108

63 Waste Storage Subsystem Tank Volume Requirements 110

64 Expendables Resupply System Schematic With Helium Transfer 111

65 Expendables Resupply System Schematic Without Helium Transfer 113

66 Pump Power for MMH Transfer 115

67 Pump Power for NTO Transfer 116

68 Pump Power for MMH and NTO Transfer 117

69 Power to Pump for MMH 118

70 Power to Pump for NTO 119

71 Power to Pump for MMH & NTO 120

72 Power to Pump for Hydrazine 121

73 Number of Supply Tanks Vs. Tank System Weight 124

74 Flow Rates of a 1.5 HP Diaphragm Compressor at Various Inlet Pressures 125

75 Flow Rates of 3.0 HP Diaphragm Compressor at Various Inlet Pressures 126

76 Time Required for Mass Transfer; Compressor Energy Required for Mass Transfer

Using a 2-Stage Compressor 127

77 Typical Assembly of a Diaphragm Compressor 130

78 Monogroove Heat Pipe 137

79 Axially Grooved Heat Pipe 138

80 Heat Transport Loop Using Mass Storage 140

81 Heat Transport Loop Using Radiator 140

82 Pumped Fluid Radiator 141

83 Heat Pipe Radiator 141

84 Heat Transport Loop Using a Two-Phased System 142

85 Simplified Resupply Module Receiver Vehicle Fluid Line Schematic 144

86 System Pressure Loss Vs. Flowrate 147

87 Analytic and Fluid Line Schematic Used in Fluid Transient Analysis 149

88 Peak Surge Pressure Vs. Initial Propellant Flowrate 150

89 Viking Orbiter Baseline Propellant Management Device 154

90 Shuttle QMS Tank Screen Configuration 158

2082e iv
ILLUSTRATIONS (Cont'd.)

Figure Page

91 Partial Pressure Balanced Coupling 163

92. 1/2 Inch Pressure Balanced Coupling 164

93 Conventional Q/D Coupling 165

94 Tube Sleeve & Poppet Quick Disconnect Coupling 166

95 Poppet Sleeve and Poppet Quick-Disconnect Coupling 167

96 Electrical Power Distribution 169

97 Avionics/Data Processing System 170

98 ERM/Carrier Vehicle Avionics Interface 171

99 Resupply Module Power Duty Cycle 172

100 Standard Square Type End Effector 176

101 Grapple Fixture/Target Assembly 177

102 Fluid Interface Panel Latch 178

103 Receiver Vehicle Fluid Interface Panel 179

104 Resupply Vehicle Fluid Interface Panel 180

105 Side View - Fluid Interface Panel 181

106 Impact on Hydrazine Slowdown Propulsion Systems for Ullage Recompression

On-Orbit Resupply 185

107 Impact on Bipropellant Slowdown Propulsion Systems for Ullage Recompression

On-Orbit Resupply 187

108 Impact on Pressure Regulated Bipropellant Propulsion Systems for Ullage

Exchange Resupply 188

109 Impact on Bipropellant Pump-fed Propulsion Systems for Ullage Exchange

Resupply 191

110 Development Test Schedule for Concept Feasibility 194

111 Bipropellant/Helium Transfer Flight Demonstration Concept 195

112 Remote Resupply Flight Demonstration Configuration 196

113 Resupply Module Flight Demonstration Program Supplier Schematic 197

114 Status of Technology Required for ERM Program 198

115 Early Vs. Late Demonstration Option Schedules/Costs 199

116 Late Demonstration Flight Schedule 200

117 We Recommend A Late Flight Demo & Early ERM Program Start 201

118 Work Breakdown Structure for ERM Program 203

119 ERM Subsystems Work Breakdown Structure 204

120 ERM System Weights 205

121 Potential Resupply Scenarios and Estimated Number of Expendable

Resupply Modules by Year 2000 207

122 ERM Developmental Test Program 208

123 ERM Qualification Test Program 209

124 First ERM Flight Hardware 210

125 Second ERM Flight Hardware 211

126 ERM Manufacturing Schedule 212

127 Advanced Development Program Supports the Flight Demonstration Program 213

128 Cost Estimate for Technology Development and Flight Demonstration

Programs 214

129 Parametric Cost Estimate for ERM Program 215

130 Parametric Cost Estimate - Baseline Program 216

131 Detailed Engineering Cost Estimate for ERM Program 217

132 Detailed Engineering Cost Estimate Range 218

133 Confidence/Probability of Reaching Baseline ERM Program Cost Target 219

134 Cost Breakdown for the ERM Baseline Program 220

135 Time Phased Cost Estimate for Baseline ERM Program 221

136 ERM Production Costs 222

137 Cost Estimate for Satellite Mod Kit to Enable Resupply 223

138 Expected Value Cost Estimate For Satellite Mod Kit to Enable Resupply 224

2082e v
. Tables

Table

1 Fluid Resupply Mission Model 9

2 Military Satellite Propulsion Requirements 10

3 DoD Resupply Requirements 10

4 Fluid Parametric Requirements Summary 25

5 Candidate Spacecraft/Platform Selections 26

6 Economic Comparison of Resupply Only Versus Replacement for

Candidate Spacecraft 38

7 ROM On-Orbit Value of NASA Spacecraft/Platform Programs 39

8 Fluids Resupply (Without Maintenance) Benefits for Polar Orbits 41

9 LEO Resupply Cost Drivers 84

10 ROM Acquisition and Total Costs for Expendables Resupply Module 85

11 Relative Cost Effectiveness of Selected LEO Resupply Scenarios 86

12 Results of Risk Evaluation for Selected LEO Resupply Scenarios

Using AHP 88

13 AHP Evaluation of Combined Risk and Complexity Results 89

14 Efficient Set of Remote Fluids Resupply Candidate DRMs 92

15 Deleted

16 Propellant Transfer Pump System Weights 103

17 6000 Psi Helium Tank Combinations 104

18 Helium System Weight Using 6000 Psi Tank Combinations 104

19 Combination Pump & Pressurant System 105

20 On-Orbit Propellant Resupply Options 106

21 Conceptual Resupply System Design, Component Legend 114

22 Power Required by Pump System for Various Propellants 122

23 Parameters of Motos and Pumps 123

24 Metal Diaphragm Compressor Requirements to Fill Various Pressure

Bottles 128

25 Two-stage Piston Compressor Requirements to Fill Various Pressure

Bottles From a 500 Psia Source 128

26 Parameters of the Compressor/Motor 128

27 Compressor Options 129

28 Flow Rates of a Diaphragm Compressor at Various Inlet and Outlet

Pressures 131

29 Battery Parameters 133

30 Resupply Module Energy Weight Requirements to Transfer 45,500 Ibs of

Storable Bipropellants and Fill 12 OMPS Bottles 135

31 Deployable Photovoltaic Panels 136

32 Heat Pipe Parameters 136

33 Weight Comparison of Each Heat Transport Loop 142

34 Total System Pressure Loss: Component Breakdown (Psid) at 10 GPM

Flowrate 146

35 Resupply Modules and OMV Extended Mission Kits Tank Requirements

" ~ Summary 151

36 Summary of Propellant Acquisition Concepts 151

37 Resupply Module design Reference Mission Scenarios Vs. PMD 160

38 Resupply Module Electrical Power Requirements 173

39 Resupply Module Power Users Source Designation 175

40 ERM System/Component Weights 183

41 Hydrazine Slowdown Satellite Weight Impact (Ullage Recompression) 184

42 Weight Impact on Bipropellant Slowdown Propulsion Systems for Ullage

Recompression Resupply 186

2082e vi
Tables (Cont'd.)

Table Page

43 Weight Impact on Pressure Regulated Propulsion Systems for Ullage

Exchange Resupply (Bipropellant) 189

44 Weight Impact on Pump-Fed Storable Bipropellant Propulsion System for

Ullage Exchange Resupply 190

45 Status of Technology for the Resupply Module Progam 193

2082e vii
Acronym Dictionary

Ah Ampere Hour

AXAF Advanced X-Kay Astrophysics Facility

BBQ Barbecue (Rotation of Spacecraft for Thermal Management

BOL Beginning of Life

CREP Cosmic Ray Experiment Program

DDT&E Design, Develop, Test and Engineering

DMSP Defense Meteorlogical Satellite Program

DRM Design Reference Mission

V Delta Velocity

DOD Department of Defense

DOMSAT Domestic Satellite

DPP Deployable Photovoltaic Panel

DPS Data Processing Subsystem

EOS Earth Observing System

EPHP Externally Pumped Heat Pipe

ERM Expendables Resupply Module

ETR Eastern Test Range

EUVE Extreme Ultraviolet Explorer

EVA Extra Vehicular Activity

GEO Geosynchronous Orbit

GN&C Guidance Navigation and Control

GP-B Gravity Probe-B

GPM Gallons Per Minute

GRO Gamma Ray Observatory

GSE Ground Support Equipment

GSTDN Ground Space Flight Tracking and Data Network

He Helium

IP Integral Propulsion

IRR Internal Rate of Return

Isp Specific Impulse

JSC Johnson Space Center

L Life of Satellite In Years

LaRC Langley Research Center

LDR Large Deployable Reflector

LEO Low Earth Orbit

Li Lithium

MIL Man-in-Loop

MMD Mean Mission Duration

MMH Monomethylhydrazine

MPS Materials Processing In Space

MSFC Marshall Space Flight Center

NASA National Aeronautics and Space Administration

NTO Nitrogentetroxide (^04)

N2 Nitrogen

N

2^4 Hydrazine

OMPS Orbiter Main Propulsion System

OMS Orbital Maneuvering System (Shuttle Orbiter)

OMV Orbital Maneuvering Vehicle

ORCS Orbiter Reaction Control Subsystem

OTV Orbit Transfer Vehicle

PVT Pressure, Volume, and Temperature

RCS Reaction Control Subsystem

RF Radio Frequency

2082e viii
Acronym Dictionary (Cont'd.)

ROM Rough Order of Mangitude

S Sulfur

SAI Science Applications Inc.

S/C Spacecraft

SDI Strategic Defense Initiative

SIRTF Shuttle Infrared Telescope Facility

SPER Space Platform Expendables Resupply

STS Space Transportation System

TFU Theoretical First Unit

TDRSS Tracking and Data Relay Satellite System

Ti Titanium

USAF SO United States Air Force Space Division

WTR Western Test Range

2082e ix
1.0 INTRODUCTION

(5,/A ^ NASA has recognized that the capability for remote resupply of space platform

expendable fluids will help transition space utilization into a new era of

operational efficiency and cost/effectiveness. The emerging Orbital

Maneuvering System (OMV) in conjunction with an expendables resupply module

will introduce the capability for fluid resupply enabling satellite lifetime

extension at locations beyond the range of the Orbiter. This report

summarizes a Phase A study of a remote resupply module for the OMV.

-tr-1—Background

It.

Numerous studies have been performed in recent years concerning the

transfer of fluids to satellites within the Orbiter payload bay. These

studies have mainly focused on the transfer of hydrazine (^H^). On a

recent Shuttle flight, transfer of hydrazine using man-in-the-loop was

demonstrated within the Orbiter payload bay. Also the Air Force is

considering incorporation of interface requirements allowing on-orbit fluid

transfer in future satellites (with notable applications to SDI). It is

apparent that the era of fluid transfer on-orbit is emerging and that remote

expendables resupply will become a recognized requirement for future

satellites and/or propulsion modules.

1.2 Objectives and Scope

The overall objectives of the study are summarized as follows:

1. Develop Performance & Operational Requirements

associated with critical liquids, gases, and lubricants

resupply

2. Develop Design Requirements for the resupply module

(RM) & equipment needed in conjunction with OMV to

perform remote fluid resupply functions

3. Develop resupply module Concept Designs (including

tankage, transfer systems & supporting mechanisms) &

associated spacecraft adaptation (user interface

emphasis)

4. Define a Flight Demonstration Program for automated

remote fluid resupply

5. Develop Program Planning Data, including cost,

schedule & preliminary supporting development

program

Task 1

- 60$ of study

Task 2

of study

Task 3

of study

The study scope was initially limited to the definition of remote fluid resupply

concepts for satellites/space platforms. This scope includes fluid transfer to

propulsion modules for LEO and higher energy orbit satellites/platforms

(expendable or reusable), but not to the space station. This assessment of

resupply benefits, with and without other types of servicing was intended to

confirm a comprehensive justification for future remote resupply operations.

2082e/4-15-85 - 1 -
1.0 INTRODUCTION

NASA has recognized^that the capability for remote resupply of space platform

expendable fluids will help transition space utilization into a new era of

operational efficiency and cost/effectiveness. The emerging Orbital

Maneuvering System (OMV) in conjunction with an expendables resupply module

will introduce the capability for fluid resupply enabling satellite lifetime

extension at locations beyond the range of the Orbiter. This report

summarizes a Phase A study of a remote resupply module for the OMV.

1.1 Background

Numerous studies have been performed in recent years concerning the

transfer of fluids to satellites within the Orbiter payload bay. These

studies have mainly focused on the transfer of hydrazine (^H^). On a

recent Shuttle flight, transfer of hydrazine using man-in-the-loop was

demonstrated within the Orbiter payload bay. Also the Air Force is

considering incorporation of interface requirements allowing on-orbit fluid

transfer in future satellites (with notable applications to SDI). It is

apparent that the era of fluid transfer on-orbit is emerging and that remote

expendables resupply will become a recognized requirement for future

satellites and/or propulsion modules.

1.2 Objectives and Scope

The overall objectives of the study are summarized as follows:

1. Develop Performance & Operational Requirements

associated with critical liquids, gases, and lubricants

resupply

2. Develop Design Requirements for the resupply module

(RM) & equipment needed in conjunction with OMV to

perform remote fluid resupply functions

3. Develop resupply module Concept Designs (including

tankage, transfer systems & supporting mechanisms) &

associated spacecraft adaptation (user interface

emphasis)

4. Define a Flight Demonstration Program for automated

remote fluid resupply

Task 1

= 60% of study

Task 2

= 30$ of study

5« Develop Program Planning Data, including cost, I Task 3

schedule & preliminary supporting development f - 1.0% of study

program J

The study scope was initially limited to the definition of remote fluid resupply

concepts for satellites/space platforms. This scope includes fluid transfer to

propulsion modules for LEO and higher energy orbit satellites/platforms

(expendable or reusable), but not to the space station. This assessment of

resupply benefits, with and without other types of servicing was intended to

confirm a comprehensive justification for future remote resupply operations.

2082e/4-15-85
1.3 Guidelines and Assumptions

The guidelines and assumptions established at the start of and during the

study are summarized below:

o Utilize available data from related study/development activities,

including propellant transfer experiments, upper stage design,

operations studies, and space station/platform studies.

o Incorporate results of In-Bay Fluid Transfer Experiment from STS 41-G.

o Insure coordination with space station/OTV fluid resupply studies.

o Propellant supply facilities will be assumed to exist as part of the

Shuttle, space station, or separate depot facilities.

o The study scope was initially limited to investigation of fluid

resupply only, but the requirements were later expanded to include

satellite servicing concurrent with refueling.

o Assuming that fluid resupply may be required prior to full-capability

spacecraft servicing, the impact of remote spacecraft servicing on

nominal fluid resupply modes will be considered.

o The proposed system must meet all manned safety criteria during

initial transport to low earth orbit and if returned to the orbiter or

manned station after a resupply mission.

1.4 Study Flow and Methodology

The study tasks and their relationship to each others are shown on

Figure 2.

Task 1

System Requirements

Analyses and -Trades

Requirements' "

for Resupply — i

L

Rockwel

° ^l ~\

Input Q

Data Base ^

<

1

1.2 Develop

Alternative

Resupply

Scenarios

1.3 Perfor * m

Effectiveness

Analysts

< >

• Benefits

• Mission Operations Requirements

• System/Subsystem Requirements

Task 2 i ' * Interface Requirements

Concept Definition

2.1 Develop

Resupply

Module

Design

t

 ~ , 2.2 Generate

_t> Demonstration -

Concept

• Design Concept

> i '

Approach

• Demonstration

Task 3 i ' Concept

Programmatic and

Planning . panning

• Schedule

2082e/4-15-85

Figure 2. Study Flow

- 2 -
Task 1, System Requirements, Analysis and Trades, was the focus of the

mid-term review. In subtask 1.1, resupply requirements were generated in-the

- form ~of a~ mission model. Next, fluid parametric requirements were generated

for this mission model to determine fluid types, quantities and usage rates

for resupply. This data was then examined to allow selection of candidate

spacecraft and platforms for resupply. In Subtask 1.2, alternative scenarios

were developed and screened for effectiveness. Mission operational and

system/subsystem requirements were next generated for these selected

scenarios, to provide data in support of the effectiveness analysis (and for

the Task 2 design analysis after selection of the Design Reference Missions).

The study effort after the mid-term review concentrated on the

development of resupply module concept designs to meet the approved DRM's

resulting from this review (Task 2). Continued effectiveness analysis

(subtask 1.3) resulted in the definition of specific requirements for these

DRM's in LEO and 6EO as well as expanded benefits assessments. In addition

the definition of a flight demonstration program and a supporting technology

program was accomplished.

Finally Task 3 involved programmatic planning and cost/schedule data

development for the technology and flight demonstrations as well as the

initial operational segments of the total program.

2.0 Summay of Results

2.1 Major Conclusions

The remote resupply alone of LEO satellites is found to be of potential

economic benefit but resupply combined with servicing is much more

advantageous. The latter case can be most effectively accomplished at the

Orbiter. Hence, remote expendables resupply alone was not recommended for

most of these satellites. Certain DoD satellites in high inclination LEO

orbits will require relatively frequent expendables resupply to avoid costly

replacement. Since it will likely be desirable to accomplish these resupply

operations without dependence on the Orbiter, the need for a LEO propellant

storage depot with a space-based OMV/RM is emphasized.

The major commercial industry drivers for GEO resupply relate to

communications satellite revenues. They are a function of the total number of

operating transponders and the number of transponders per satellite,

transportation costs, satellite production costs, insurance costs, and

technological obsolescence.

Lifetime extension of communications satellites through Attitude Control

Subsystem (ACS) resupply is of marginal benefit without concurrent servicing

of the satellite electronics. This is due to the relatively rapid rate (12/6

annual) of technological obsolescence. Planned lifetime extension where the

satellite's reliability is increased (and also the initial satellite cost)

turns out to be the least attractive scenario because of the increased

up-front costs. The contingency resupply of satellites which have lost their

fuel from malfunctions is always cost effective, if the problem can be

isolated and resolved early in the satellite's life when interruption of the

revenue stream can have the greatest impact.

Though adding additional transponders at the expense of ACS fuel weight

increases the toal revenue from a satellite, it is not cost effective

primarily because of the risk involved in assuring future operation. Less

2082e/4-15-85 - 3 -
onboard fuel will require more frequent (Optimally 2 in a 10 year period)

resupply missions, thus increasing the risk to the satellite revenue stream if

the resupply fails. Insurance premiums of less than 15% are required to make

this and other resupply cases viable.

Servicing large satellites is generally more cost effective than smaller

satellites because large satellites house more transponders which may be

serviced on only one mission. Multiple manifesting of missions achieves the

same benefit when the transponders are spread over several smaller satellites.

If a Comsat undergoes a low rate of technological obsolesence there is

less incentive to resupply and upgrade it. If this rate can be held to 6% or

lower it is unlikely that resupply/servicing will be beneficial unless a

revenue restoration level greater than 100$ can be realized. At a 12%

obsolesence rate a revenue restoration of 1Q% is required to assure a

non-negative impact to the satellite revenue stream.

The resupply system can be configured for reasonably low impact on the

receiver spacecraft. The impact is primarily required interface provisions.

The ullage exchange propellant transfer process has the inherent versatility

to adapt to different receiver propulsion systems; regulated, pressure fed,

blowdown, or pump fed without severe weight impacts to the receiver spacecraft,

The DDT&E costs for the resupply module program could be reduced by

exploiting the synergism with the NASA-JSC ^Ify in-bay tanker program.

Although the JSC program is being configured for different functional

requirements and missions, it will potentially result in certain common

technology development or modifiable hardware. Such hardware could include

interface components, Mission Peculiar Equipment Support Structure (MPESS) (as

tankage support for in-bay tanker and resupply module flight demonstration) ,

and propellant pumps.

2.2 Selected Options

Ullage exchange was selected as the best option for the N204/MMH

fluid transfer process. This approach is applicable to all potential receiver

propulsion subsystem and acquisition types through appropriate modifications.

It minimizes pressurant resupply requirements, involves no adiabatic

compression (explosion hazard), requires no waste or hazardous effluent

scavenging, and provides constant pressure resupply.

Bi-propellant ^O^MMH fluids were chosen for the resupply module

since they are compatible with some GEO propulsion systems, with the OMV

propulsion subsystem, and they are scavengable from orbiter QMS tanks.

Figure 3 presents ten scenarios of bi-propellant use (some without

helium) which were selected as design reference missions. Scenarios 1 and 2

are respectively contingency and planned resupply in GEO. Scenarios 6 and 7

are similar operations to 1 and 2 but in LEO. Scenario three is top-off of a

satellite integral propulsion subsystem. Five is refueling of a perigee stage

from the OMV/RM; eight is similar to five but also includes an apogee burn,

Resupply Module staging, and OMV aerobraking maneuver on return. In scenario

nine only the OMV is being resupplied. Four is a coupling of resupply modules

in LEO which function as a depot. Finally, in ten propellant is being

transferred from the orbiter to the depot through the OMV.

2082e/ - 4 -
A propellent depot becomes most attractive if one uses the basic resupply

module as the "core" element of_the_depot. Several of these modules could be

used to achieve the desired storable propellant capacity. This could reduce

development costs for the depot.

Propellant top-off capability in LEO via remote resupply would allow

launch in the STS of multiple satellites with one or more in an off-loaded

condition. Near term use of propellant top-off capability may also be

applicable to currently projected larger DOD satellites which if launched

fully loaded, would exceed STS lift capacity.

An additional application of the EM might be to fuel a high performance

reusable perigee stage, as is being studied at McDonnel Douglas. This stage

would be capable of launching a 12K Ib payload into GEO after fuelling by a

resupply module.

OF POOR QUALITY

POTENTIAL

MISSION

SCENARIO

CONTINGENCY MULTIPLE S/C

RESUPPLY FROM OMV

PLANNED RESUPPLY/MODULE

CHANGE-OUT FROM OMV FOR

TRANSPONDER MAXIMIZATION

PROPELLANT TRANSFER

FROM RM TO OMV FOR MORE

ECONOMICAL TRANSFER TO GEO

TOP-OFF

COMPACT INTEGRAL

PROPUL S/C

DEPOT OF

RM'S

REFUEL REUSABLE

PERIGEE STAGE

OPERATE OMV/RM

FROM LEO DEPOT

FOR PROPELLANT

TRANSFER OPERATIONS

RM/OMV USE FOR

PROPELLANT TRANSFER

FROM STS TO DEPOT

REFUEL

SPACE-BASED

OMV

RESUPPLY FROM

OMV

EXAMPLE

RM

PROPELLANT

LBS.

38.000

(1,000*)

AS. 500

(2,000*)

20,000

45,000

45,500

6,000

RM

HELIUM

LBS.

30 .

90

12

2082e/

Figure J. Selected Mission Scenarios

- 5 -
2.3 Recommended Approach

The selected resupply module configuration is shown in Figures 4 and 5

and is presented in two configurations; without and with the satellite

servlcer. This concept was selected from among several competing approaches

primarily on the basis of overall structural efficiency and for it's use of

existing hardware providng low development cost. The resupply module is

supported at its forward end by an existing IUS upper stage forward cradle.

This cradle includes load equalization capability which reduces the structural

redundancy between the resupply module and the Orbiter. It also allows a

TniniTniiyn of weight impact on the resupply module for attachment to the Orbiter

payload bay longerons and keel at its forward end. The resupply module uses

six stretched QMS tanks with a modified ullage positioning Propellant

Management Device (PMD) for ullage bubble position control.

STS Orbiter MFS Hi-Press

Heliun Tank* (12 max.)

'Prop/He Tank forward •^^

Forward Umbilical

(ooveable)

End Effector

Figure 4. Resupply Module General Configuration (Without Servicer)

Explosive Bole

iHeliua Resupply Tanks

(on RH-ExCended

Mission)

Forward _

ASE Cradle

End Effector and

Umbilical (OMV)

Satellite

Servicer

X 1135.87 (RN/OMV Ht 32,000 Ibs.)

o

Umbilical Panel

(Fixed)

i

Grapple Fixture

Figure 5- Resupply Module General Configuration (With Servicer)

2082e/4-15-85 - 6 -
ORIGINAL

OF POOR QUALITY

The selected flight demonstration approach uses hardware developed

through a low-cost Technology Validation .Program and-many off-the-shel-f

components" in conjunction with a low-cost structural framework. Two modules,

both of nearly identical configuration, would be docked together in the

payload bay using the RMS. Subsequently, storable bipropellants plus helium

transfer would be demonstrated without man-in-the-loop. One module would

represent the receiver spacecraft and the other the resupply module. Most of

the components, except the main tankage (because of its small size) would

later be used for operational resupply units to allow significant program cost

savings. The selected flight demonstration concept is depicted in Figure 6.

SPAS OR WESS STRUCTURE FOR

RECEIVER AND SUPPLIER

SIMULATORS

RMS USED TO DOCK

RECEIVER TO SUPPLIER IN

PAYLOAD BAY

• MOST OF DEMO SUBSYSTEMS LATER

USED FOR OPERATIONAL FLIGHT

UNITS FOR COST SAVINGS.

Figure 6. Bi-Propellant/Helium Transfer Flight Demonstration Concept

3.0 STUDY RESULTS

This section presents the analyses associated with the Space Platform

Expendables Resupply Concept Definition Study.

3«1 System Requirements Analysis and Trades (Task l)

The overall activities of the system requirements analysis and trades of the

study are shown in Figure 7> followed by a description of the main study

accomplishments for Task 1.

2082e/4-15-85 - 7 -
TASX1

MID-TERM REVIEW

V

SYSTEM REQUIREMENTS

ANALYSES « TRADES

1.1 DEVELOP

REQUIREMENTS FOR

RESUPPUT

1.1.1 GENERATE RESUPPIY

MISSION MOOR

1.1.2 DERIVE RESUPPUT

PARAMETRIC

REQUIREMENTS

1.1.3 SELECT CANDIDATE

SPACECRAFT 4

PLATFORMS

1.2 DEVELOP ALTERNATIVE RESUPPLY

SCENARIOS

1.2.1 6ENERATE ALTERNATE RESUPPLY

SCENARIOS

1.2.2 DERIVE MISSION OPER

REQUIREMENTS

1.2.3 SPECIFY SYSTEM/SUBSYS

REQUIREMENTS

1.2.4 IDENTIFY INTERFACE

REQUIREMENTS

±

1.3 PERFORM EFFECTIVENESS ANALYSIS

1.3.1 FORMULATE METHODOLOGY &

CRITERIA

1.3.2 GENERATE EFFECTIVENESS DATA

1.3.3 SELECT BASELINE ORMs

• MISSION OreUTHMt MQUfllEMEMTS

> I8IEFTTS • SYSTEM/ IUUYSTEM) U9UMCHEMTS

• INTERFACE KOU«ttiS(TS

Figure 7- Task 1 Activities

3.1.1 Mission Model and Fluid Parametric Requirements

A detailed mission model of fluids resupply engagement was developed in

Activity 1.1.1. The model was developed from earlier projections of OMV

servicing tasks and estimates of future STS launch rates and is shown in Table

1. All spacecraft that could take advantage of fluid resupply were included

in this engagement model.

Those missions which would actually be recommended for resupply are discussed

later in this report.

Estimates of the number of resupply engagements for each spacecraft/platform

were based on discussions with individual program offices. Such discussions

covered estimates of spacecraft technological obsolescence and value of

extended time on-orbit. Classified data on the make up of the DoD segment is

not available under this contract. The potential line of Table 1 includes

engagements which were considered speculative and beyond the nominal

categorization of the other fluid transfer engagements. They included

additional DoD users and engagements in GEO.

This mission model assumed the resupply capability would be incorporated in

follow-ons to existing large communications satellite programs. Estimates of

the satellites propellant requirements were made since the commercial users

were considered only a potential category. The propellant requirements are

for on-orbit station keeping and pointing. But these vary with satellite size

and pointing accuracy specifications.

2082e/ - 8 -
MISSION LOCATION 1989 1990 1991 1992 1993-1994 1995 1996 1997 1998 1999 2000 2001 COMMENT

GRO

AXAF

EUVE

CREP

PROTEUS I't

LOR

SIRTF

6RAV PROBE-B

EOS

MPS "I

GEO PLATFORM

CONTINGENCY

CIVILIAN TOTAL

NMI

227

320

324

216?

216

486

486

520

380

320

19323

TBO

OEG

28.5 0

28.5

28.5 0

28.5

28.5

28.5

98

90

99.8

28.5

0

TBD

1

D

1

D

D D 1

1 3 5

1"

2 5 6

1

1

1

D

0

6

9

1

1

D

6

1

9

1

1

D

1

1

5

1

10

1

2

3

6

1

1

1

1

1

5

1

1 1 1

1

t

• 2

1

2 4 3

1

1

1

3

3 AT ORBITER

3

1 AT ORBITER

1

10, PLATFORMS

2

4

1

4, PLATFORMS

30, PLATFORMS

EXPT SERVICING

20% OF OMV

CONTINGENCIES

NEED FLUIDS

64

DOD 3 47

NOMINAL TOTAL

CUMULATIVE

POTENTIAL LINE

POTENTIAL TOTAL

CUMULATIVE

4

4

4

4

8

12

8

12

8

20

8

20

13

33

13

33

14

47

14

47

17

64

3

20

67

7 12

71 83

10 7

17 19

84 103

4 8

87 95

8 12

12 20

115 135

10 6

105 111

8 12

18 18

153 171

111

60

171

- FIRST MISSION MAN-TENDED AT ORBITER

- DONE AT ORBITER MAN-TENDED

D -DEPLOYMENT

• 111 FLUID RESUPPLY ENGAGEMENTS FOR DOO, NASA AND COMMERCIAL USERS

FROM 1989-2001 (NOT INCLUDING SPACE BASED OTV, OMV AND STATION)

Table 1. Fluid Eesupply Mission Model

Based on the SAI subcontract study, DoD resupply requirements are at present

predicted to be primarily propellents. Evasion and on-orbit maneuvering

constitute a large percentage of the propellant requirements. The SAI study

also indicated that the DoD favors the bi-propellant combinations of ^04.

and A-50 for future system designs, although MMH may eventually be

substituted. A storable bi-propellant combination was used in the parametric

requirements for those satellites not currently baselined for hydrazine

propulsion.

Because of the confidential nature of DoD programs, parametric requirements

are for this portion of the model limited in scope and detail. General

propellant requirements are presented in the form of total impulse

requirements for the various missions in Table 2.

On-orbit propellant consumption for a few of the nominal and potential

missions was calculated by using Isp's of 230 and 290 seconds for hydrazine

and N20^/A-50 bi-propellant systems, respectively. These Isp's were

assumed for existing engine capability when utilzed for both delta V and

attitude control. Requirements for the other missions were also based

primarily on the SAI results. The DoD resupply requirements are summarized in

Table J>.

2082e/4-15-85 - 9 -
ALTITUDE

TOTAL IMPULSE REQUIRED

SECONDS)

INSERTION OM-ORBIT

COMMENTS

LOW EARTH ORBIT

HIGH EARTH ORBIT

GEOSYNCHRONOUS

2,0

10,0 - 20,0

20,0

~ 2,0

0,3 - 1,0

0,3 - 1,0

INSERTION AND ON-ORBIT

REQ'TS ARE SIMILAR

10,0 X 1$ SEC REQUIRED

FOR 1000 NM INSERTION

10,000 LB TO GEO,

MANEUVERING TO SERVICE

6 SATELLITES

*Includes Maneuver and Station-Keeping Capability

Table 2. Military Satellite Propulsion Requirements

CANDIDATE SATELLITE

NOMINAL

DMSP (1)

DOD - 1 (2)

DOD - 2 (2) (3)

DOD - 3 (2)

POTENTIAL^

DOD - 6 (2)

DOD - 7

DOD - 8

DOD - 9

RESUPPLY QUANTITY (LB)

70 (N2H4), 5 (NITROGEN)

7000 (N204/A-50)

8700 (N2H4)/ 1200(HELIUM)

7000 (N204/A-50)

7000 (N204/A-50).

250 (N2H4)

350 (N2H4)

1000 (N204/A-50)

RESUPPLY INTERVAL (YR)

2

1

3

3

3

5

7

5

(1) Propellant quantity shown is for remote resupply. Prior to 1994

retrieval to the orbiter is necessary for resupply.

(2) Propellant quantity calculated based on military propulsion

requirements (Table 2)

(3) Cryogenic helium quantity based on SIRTF (80$ of SIRTP capacity,

extrapolated for 3 year life)

2082e/4-15-85

Table 3« DoD Resupply Requirements

- 10 -
Resupply parametric requirements were developed for non-DoD missions. The

candidate"missions' selected were those likely to require or benefit from

consumables resupply during the fiscal years 1989-2001.

The NASA space systems technology model provided the starting point for

developing the parametric requirements. The appropriate program and/or

project managers were then contacted for more specific information. In many

instances, recently published documents were available on the individul

projects. One advantage in establishing personal contacts was the information

gained with respect to realistic mission launch dates and their attitude

toward utilizing resupply capabilities.

Results from the user contact efforts are summarized on the following pages

for the missions listed below:

Advanced X-Ray Astrophysics Facility (AXAF) - Objectives of the 28.5°

free-flyer are to determine the positions of X-ray sources, their physical

properties as composition and structure, and the process involved in X-ray

photon production.

Cosmic Ray Experiment Program (CREP) - Utilizing various NASA capabilities

(such as balloons, Spacelab, LDEF, special spacecraft), the CHEF will improve

upon previous terrestrial and interplanetary measurements by an order of

magnitude.

Earth Observing System (EOS) - The objective of the sun-synchronous polar

platform is integrated scientific observation of the Earth and its environment

including the biological, chemical, and physical characteristics of the land,

oceans, and atmosphere.

Extreme Ultraviolet Explorer (EUVE) - The EUVE explorer will make the first

all-sky map to discover, obtain accurate positions, and determine the spectral

energy distribution for all detectable EUV sources.

Geostationary Platform (Experimental) - A permanent platform for co-locating

compatible geostationary payloads such as communication and maritime systems.

Gravity Probe-B (GP-B) - High precision measurement of the relativistic

spin-orbit and spin-spin coupling will provide a fundamental step in

discriminating between viable competing theories of gravitation.

Gamma Ray Observatory (GRO) - Scientific objectives include: l) Study of

discrete objectives; 2) Search for evidence of nucleo-synthesis; 3) Survey

of the Galaxy in gamma rays; 4) Search for cosmological effects and possible

primordial "black hole emissions; and 5) observation of gamma ray bursts.

Large Deployable Reflector (LDR) - A dedicated astronomical observatory

operating in the spectral region between 30 micrometers and 1 milimeter.

Materials Processing in Space (MPS) - Free -flying MPS factory modules

designed to satisfy microgravity processing requirements of the commerical

sector.

Proteus - Long duration LEO system derived from Leasecraft concepts and

designed to support consecutive short duration Explorer - class payloads.

2082e/
Shuttle Infrared Telescope Facility (SIHTF) - The objective is to conduct

definitive high sensitivity infrared photo-metric and spectroscopic of a wide

range of astrophysical phenomena.

Data obtained for these programs is summarized on the following pages.

2082e/4-15-85 - 12 -
CONSUMAJBLES_REQUIREMENTS

PROGRAM; AXAF

SIZE; DIA (ft) 14 LENGTH (ft) 43 MASS: 19,000 LB

S/C REPLACEMENT COST; $200 M LIFETIME: > 10 YR

CONSUMABLES; TYPE GAS MIXTURE USAGE 85 LB OVER 3 YR

Xe, 10£ METHANE)

S/C STORAGE CONDITIONS: TITANIUM RESERVOIRS AT VARIOUS PRESSURES

RESUPPLY TRANSFER CONDITIONS: COULD BE ACCOMPLISHED BY A PRESSURE OR PUMP FED

UMBILICAL SYSTEM, CURRENTLY 952 paia IS THE

HIGHEST INSTRUMENT PRESSURE REQUIREMENT.

RESUPPLY TIMES: APPROXIMATELY 3 HR AFTER 3 YR OF OPERATION

ORBITAL PARAMETERS; h (nmi) - 320 (205 AFTER 3 YR) i(deg) - 28-5

, JL (deg) - UNDETERMINED CO (deg) - UNDETERMINED

f

j

S/C PROPULSIVE CAPABILITY: CURRENTLY NEEDS AN OMV, MAY USE STS,

INTEGRAL PROPULSION ¥ILL BE INVESTIGATED

COMMENTS; PROJECT IN PHASE A, PHASE B WILL BE FALL OF 1984, ATP EXPECTED

IN 1987 WITH IOC IN 1991; DESIGN OPTION SUPPORTED BY SAO IS TO

PLACE CONSUMMABLE REQ'T ON INDIVIDUAL DETECTORS (LEADS TO

INSTRUMENT C/0)

ATTITUDE TOWARD RESUPPLY; CURRENTLY BEING DESIGNED FOR MAINTENANCE AND

RESUPPLY, MAY INCLUDE EXPERIMENT CHANGE

CONTACT; CARROLL DAILY (MSFC) 205-453-2788

BILL McKINNEN (SAO) FTS-830-7358

ED MCLAUGHLIN (SAO) FTs-830-7362

NOTE: RESUPPLY OF AXAF FROZEN COOLANTS IS CONSIDERED TOO HAZARDOUS,

REQUIRES INSTRUMENT CHANGEOUT

2082e/ - 13 -
CONSUMABLES REQUIREMENTS

PROGRAM; CREP

SIZE; DIA (ft) UNDETERMINED LENGTH (ft) UNDETERMINED MASS: UNDETERMINED

S/C REPLACEMENT COST; UNDETERMINED LIFETIME; 4 YH

CONSUMABLES; TYPE N^ (S/C BUS PROPELLANT) USAGE a»550LB OVER 2 YR (N^T

He (SUPERCONDUCTING MAGNETS) «*160LB OVER 2 YR (LHe)

Xe-METHANE (GAS COUNTERS) *&?0 Ib over 2 YR

S/C STORAGE CONDITIONS: UNDETERMINED

RESUPPLY TRANSFER CONDITIONS: PROBABLY PRESSURE OR PUMP FED UMBILICAL

FOR HYDRAZINE AND GAS MIXTURE, REQUIRES ADVANCED

TECHNOLOGY FOR SUPERFLUID HELIUM

RESUPPLY TIMES: ROUGHLY 12 HR TO RESUPPLY AFTER 2 YR OF OPERATION

ORBITAL PARAMETERS; h (nmi) - 216 i(deg) 2.85

•*• (deg) - UNDETERMINED t*> (deg) - UNDETERMINED

S/C PROPULSIVE CAPABILITY: WILL PROBABLY REQUIRE A BUS/PLATFORM SUCH AS

PROTEUS FOR THE FINAL PHASES OF THE PROGRAM

COMMENTS; PROJECT IN PRELIMINARY DESIGN PHASE, MULTI-PHASE PROGRAM

USING VARIOUS NASA CAPABILITIES

ATTITUDE TOWARD RESUPPLY; WILL PROBABLY DESIGN FINAL PHASE OF PROGRAM TO

UTILIZE RESUPPLY CAPABILITIES

CONTACT; Dr. Jon Ormes (NASA HQ) 202-453-1462

(GSFC) 301-344-4793

2082e/4-15-85 - 14 -
CONSUMABLES REQUIREMENTS

PROGRAM: EOS

SIZE; DIA (ft) UNDETERMINED LENGTH (ft) UNDETERMINED MASS: UNDETERMINED

S/C REPLACEMENT COST; $50M LIFETIME: >10YR

CONSUMABLES; TYPE N2H4 (PLATFORM PROPULSION) USAGE 2000LB OVER 3YR (FUNCTION

OF SYSTEM DESIGN AND

NUMBER OF PAYLOADS)

S/C STORAGE CONDITIONS: PRESSURE TANKS OR DIAPHRAGM TANKS

RESUPPLY TRANSFER CONDITIONS: MAY USE PRESSURE FED UMBILICALS

RESUPPLY TIMES: APPROXIMATELY 4 HI TO RESUPPLY AFTER 3 YR OPERATION

ORBITAL PARAMETERS; h (nmi) - 380 i(deg) - 99-8

^l-(deg) - UNDETEMINED 6J (deg) - UNDETERMINED

.S/C PROPULSIVE CAPABILITY: INTEGRAL PROPULSION PROBABLE, OMV IS AN OPTION

COMMENTS; PROJECT IN CONCEPTUAL PHASE, CURRENTLY BASELINED WITH SPACE

STATION POLAR PLATFORM, DEMONSTRATION SERVICING LIKELY EARLY

IN OPERATIONAL PHASE

ATTITUDE TOWARD RESUPPLY; WILL BE DESIGNED FOR RESUPPLY OF EOS PROPELLANTS

CONTACT; CHUCK MacKENZIE (GSFC) 301-344-6008

2082e/4-15-85 - 15 -
CONSUMABLES REQUIREMENTS

PROGRAM: EUVE

SIZE; DIA (ft) UNDETERMINED LENGTH (ft) UNDETERMNED MASS: 882LB

S/C REPLACEMENT COST; $24M LIFETIME: 2YR

CONSUMABLES;; TYPE POSSIBLY N2H4 USAGE 400LB OVER 1YR

S/C STORAGE CONDITIONS: PROBABLY PRESSURE TANKS OR DIAPHRAGM TANKS

RESUPPLY TIMES: APPROXIMATELY ONE HOUR AFTER ONE YEAR OF OPERATION

ORBITAL PARAMETERS; h (nmi) - 324 i(deg) - 28.5

_n.(deg) - UNDETERMINED CO (deg) - UNDETERMINED

S/C PROPULSIVE CAPABILITY: MAY USE INTEGRAL PROPULSION FOR STS RENDEZVOUS

COMMENTS; PHASE A RFP DUE IN JULY 1984, OPERATES AT "NIGHT" ONLY, MIRROR

CONTAMINATION CRITICAL DURING RESUPPLY OPERATION

ATTITUDE TOWARD RESUPPLY; NASA JPL WOULD LIKE INCREASED LIFETIME BY RESUPPLY/

SERVICING, CURRENT PROGRAM BASELINED FOR 1YR OF OPERATION

CONTACT; R. A. PIOZAJ (NASA JPL) 818-354-7447

2082e/4-15-85 - 16 -
CONSUMABLES REQUIREMENTS

PROGRAM; GEO PLATFORM (EXPERIMENTAL)

SIZE; DIA (ft) 110 LENGTH (ft) 164 MASS: 12.000LB

(ON-ORBIT) (ON-ORBIT)

S/C REPLACEMENT COST; $175M LIFETIME: 7YR

CONSUMABLES;; TYPE N^ USAGE 1764 LB OVER 7 YR

S/C STORAGE CONDITIONS: MAY USE PRESSURE TANK OR DIAPHRAGM TANK

RESUPPLY TRANSFER CONDITONS: COULD BE ACCOMPLISHED BY PRESSURE OR PUMP FED

UMBILICAL SYSTEM

RESUPPLY TIMES: APPROXIMATELY 4 HR AFTER 7 YR OF OPERATION

ORBITAL PARAMETERS; h (nmi) - 19323 8 100° WEST i(deg) - 0

-A. (deg) - N/A CO (deg) - N/A

S/C PROPULSIVE CAPABILITY: OTV NECESSARY, ACS USING A SINGLE SLOWDOWN RCS

COMMENTS; PRELIMINARY ARCHITECTURE STUDIES COMPLETE, PAYLOAD STUDIES ON-GOING,

SERVICING DEMONSTRATION PROBABLY BEFORE 7 YR OF OPERATION

ATTITUDE TOWARD RESUPPLY; RESUPPLY/SERVICING SUPPORT (i.e. BY SPACE STATION)

WILL BE MANDATORY

CONTACT; GEORGE KNOUSE (NASA HQ) 202-453-1515

2082e/4-15-85 - 17 -
CONSUMABLES REQUIREMENTS

PROGRAM: GP-B

SIZE; DIA (ft) 15 LENGTH (ft) 8 MASS: 4000 LB

(ON-ORBIT) (ON-ORBIT)

S/C REPLACEMENT COST: $?2M LIFETIME: 2YR MAXIMUM

CONSUMABLES;; TYPE SUPERFLUID He USAGE 397LB OVER 1 YE

S/C STORAGE CONDITIONS: DEWAE CONFIGURATION @ 1.8K

RESUPPLY TRANSFER CONDITONS: WOULD REQUIRE SIRTF-TYPE ADVANCED

TRANSFER TECHNOLOGY

RESUPPLY TIMES: APPROXIMATELY 6 HE AFTER 1 YR OF OPERATION

ORBITAL PARAMETERS; h (nmi) - 520 i(deg) - 90

-A- (deg) - UNDETERMINED Co (deg) - UNDETERMINED

S/C PROPULSIVE CAPABILITY: ACS ONLY (He BOILOFF), USED TO COUNTER ATMOSPHERIC

DRAG AND SOLAR WIND EFFECTS

COMMENTS; ONLY 1 YR PROGRAM IF ADVANCED TECHNOLOGY GYROS "PROVE-OUT",

STATE-OF-ART GYROS WOULD RESULT IN 2-3 YS PROGRAM, FIRST

FLIGHT ON STS WILL BE TECHNOLOGY DEMONSTRATION

ATTITUDE TOWARD RESUPPLY; CURRENT DESIGN HAS BEEN DOWN-SIZED, IF STATE-OF-ART

GYROS ARE USED GP-B WILL PROBABLY REQUIRE RESUPPLY

CONTACT: JOYCE NEIGHBORS (MSFC) 205-453-5584

2082e/4-15-85 - 18 -
CONSUMABLES REQUIREMENTS

PROGRAM; GRO

SIZE; DIA (ft) 15 LENGTH (ft) 17 MASS: 33,000 LB

(ON-ORBIT) (ON-ORBIT)

S/C REPLACEMENT COST; $180M LIFETIME: 27 MO.

(>27MO DESIRABLE)

CONSUMABLES; TYPE N2H4 USAGE 3830 LB OVER 27 MO

S/C STORAGE CONDITIONS: PRESSURE TANKS AT 60°F, INITIAL PRESSURE OF

350 psia DECREASING TO 100 psia AT END-OF-CYCLE

RESUPPLY TRANSFER CONDITIONS: PRESSURE FED UMBILICAL SYSTEM, 200°F AND

>400 psia

RESUPPLY TIMES: APPROXIMATELY 10 HR AFTER 27 MO OF OPERATION

ORBITAL PARAMETERS; h (nmi) - 227 ( 216 AFTER @ YR) i(deg) - 28.5

(deg) - UNDETERMINED OJ (deg) - UNDETERMINED

S/C PROPULSIVE CAPABILITY: INTEGRAL PROPULSION MAY BE USED TO LOWER

COMMENTS; APPROXIMATELY 50 LB He FOR SINGLE SLOWDOWN SYSTEM; ALTERNATE

RESUPPLY SCENARIO - BLEED PRESSURANT, REFILL NgH., REFILL

PRESSURANT; EARLIER DESIGN USED ARGON

ATTITUDE TOWARD RESUPPLY; RESUPPLY HIGHLY DESIRABLE TO EXTEND OPERATIONAL

LIFE

CONTACT; RON McCULLAR (OSSA) 202-453-1428

2082e/ - 19 -
CONSUMABLES REQUIREMENTS

PROGRAM; LDR

SIZE; DIA (ft) 66 LENGTH (ft) 100 MASS: 88,000 LB

S/C REPLACEMENT COST; $250M LIFETIME: 10 YR.

CONSUMABLES; TYPE SUPERFLUID He USAGE 1320 LB OVER 3 YR

S/C STORAGE CONDITIONS: DE¥AR CONFIGURATION (CLOSED SYSTEM) AT 4 K

RESUPPLY TRANSFER CONDITONS: WOULD REQUIRE SIRTF-TYPE ADVANCED TRANSFER

TECHNOLOGY

RESUPPLY TIMES: APPROXIMATELY 10 HR AFTER 3 YR OF OPERATION

ORBITAL PARAMETERS; h (nmi) - 486 i(deg) - 28.5

-0. (deg) - UNDETERMINED CO Ueg) - UNDETERMINED

S/C PROPULSIVE CAPABILITY: WILL BE INVESTIGATED,

COMMENTS; WOULD LIKE SUN-SYNC ORBIT, CRYO USAGE DERIVED FROM ESTIMATED

THERMAL REQUIREMENTS (HEAT FLOW RATE) ASSUMING A CLOSED SYSTEM,

•^HC

 = >65 Cal/cc» Q = 1 ¥ATT

 @ 4 K.

ATTITUDE TOWARD RESUPPLY; DESIGN TO INCLUDE RESUPPLY AND CHANGEOUT

CONTACT; BRUCE PITTMAN (NASA ARC) FTS 448-5692

2082e/4-15-85 - 20 -
CONSUMABLES REQUIREMENTS

PROGRAM; MPS

SIZE; DIA (ft) UNDETERMINED LENGTH (ft) UNDETERMINED MASS: 25,000 LB

S/C REPLACEMENT COST; $60M LIFETIME: 5 YR

CONSUMABLES; TYPE HgO USAGE 5000 LB OVER 6 MO

S/C STORAGE CONDITIONS; ELECTROPHORESIS SYSTEM

RESUPPLY TRANSFER CONDITONS; PROBABLY CHANGEOUT OF ELECTROPHORESIS TANKS

RESUPPLY TIMES: APPROXIMATELY 5 HR AFTER 6 MO OF OPERATION

ORBITAL PARAMETERS; h (ami) - 486 i(deg) - 28-5

-A-(deg) - UNDETERMINED £O (deg) - UNDETERMINED

S/C PROPULSIVE CAPABILITY: IF NECESSARY, MMS TYPE PROPULSIVE MODULE COULD

BE USED FOR ORBIT CHANGES ONLY (PROBABLY NOT

NECESSARY)

COMMENTS; PRIMARY EFFORTS ARE TO LOCATE FACTORY MODULES ON SPACE STATION.

TIMING MAY BE SUCH THAT FREE-FLYER FACTORY MODULES WILL BE IN

OPERATION PRIOR TO SPACE STATION.

ATTITUDE TOWARD RESUPPLY; IF DESIGNED AS FREE-FLYER, RESUPPLY/SERVICING

WILL BE MANDATORY

CONTACT; DAVE RICHMAN (McDONNELL DOUGLAS) 3U-232-6967

2082e/4-15-85 - 21 -
CONSUMABLES REQUIREMENTS

PROGRAM; PROTEUS

SIZE; DIA (ft) VAEIABLE LENGTH (ft) VARIABLE MASS: 6000 LB AVG.

(W/PAYLOAD)

S/C REPLACEMENT COST; $25M LIFETIME; 10 YR

(2 YR MAX PER USER)

-

CONSUMABLES ; TYPE N.H. USAGE 1100 LB OVER 2 YR

 2 4

(ORBIT TRANSFER & STATION KEEPING)

S/C STORAGE CONDITIONS; PRESSURE TANKS

HESUPPLY TRANSFER CONDITONS; MAY USE PRESSURE OR PUMP FED UMBILICAL SYSTEM

RESUPPLY TIMES; APPROXIMATELY 2.5 HR AFTER 2 YR MISSIONS

ORBITAL PARAMETERS; h (nmi) - 216 i(deg) - 28.5

(deg) - VARIABLE <$J (deg) - VARIABLE

S/C PROPULSIVE CAPABILITY: PROBABLY INTEGRAL, MAY USE OMV

COMMENTS ; PRELIMINARY DESIGN PHASE, MAY BE DERIVATIVE OF MMS, TWO S/C

BASELINED

ATTITUDE TOWARD RESUPPLY; DESIGN WILL INCLUDE RESUPPLY AND SERVICING

CONTACT; W. HIBBARD (GSFC) 301-344-7510

2082e/ - 22 -
CONSUMABLES REQUIREMENTS

PROGRAM: SIRTF

SIZE; DIA (ft)^#12 LENGTH (ft) 27

S/C REPLACEMENT COST; $300M

MASS: 16,300 LB

¥/0 PROPULSION

LIFETIME: 10 YR

CONSUMABLES; TYPE SUPERFLUID He (SOLID ^ WOULD

BE USED IP OPEATING AT LOW INCL.)

USAGE 992 LB He OVER 2 YR

235 He OVER 2 YR

S/C STORAGE CONDITIONS: DEWAR CONFIGURATION, He 0 1.8K, H 0 7-10K,

He COLS WITHIN H COULD KEEP IT SOLID

•

RESUPPLY TRANSFER CONDITONS: UMBILICALS POSSIBLE, NEW TECHNOLOGY REQUIRED

FOR TRANSFER, EXPECT 80-90# OF He CAN BE

REPLENISHED AND 80% OF H2

RESUPPLY TIMES: APPROXIMATELY 7 HR AFTER 2 YR OF OPERATION

10 HR FOR He AND H AT LOW INCLINATION)

ORBITAL PARAMETERS; h (nmi) - 378-486 i(deg) - 93

(SECOND CHOICE-378 NMI 0 28.5°)

./>- (deg) - VARIABLE cJ (deg) - VARIABLE

S/C PROPULSIVE CAPABILITY: NONE

COMMENTS; TECHNOLOGY DEVELOPMENT FOR CONSUMABLES TRANSFER IS INCLUDED IN

RFP, POLAR ORBIT FIRST PREFERENCE, ONLY 600 LB SUPERFLUID He

REQUIRED FOR LOW INCLINATION

ATTITUDE TOWARD RESUPPLY; RESUPPLY WOULD BE RISKY, MAXIMUM OPERATIONAL

LIFETIME (2 YR) DESIRED BEFORE RESUPPLY

CONTACT; WALT BROOKS (NASA ARC) FTS 448-6530

JIM MURPHY (NASA ARC) FTS 448-6643

2082e/4-15-85 - 23 -
Due to the conceptual or preliminary state of many of the candidate missions,

particularly those in the late 1990's, many assumptions had to be made to

develop representative requirements. Those assumptions were:

o Hydrazine propulsion using a single blowdown system was assumed for

those missions requiring propulsion, but which had no spacecraft

design baselined. Although no currently baselined candidate

spacecraft designs utilize bi-propellants or require resupply of

pressurants, future spacecraft designs may include these as an option

for trades analyses.

o Once rendezvous with candidate spacecraft has been accomplished, one

hour is required for each 500 Ib of hydrazine to be resupplied

(includes actual transfer and associated operations).

o The spacecraft replacement cost was estimated to be 40$ of the total

program cost when the actual replacement costs could not be

determined.

o Mission operational lifetime was extended beyond the program baseline

if it would be technically or scientifically advantageous and

consumables resupply was the primary constraint.

o A hybrid cryogenic system was assumed for the large deployable

reflector to determine consumable quantities (eg. a combination

closed dewar and passive radiator system).

o Candidate missions scheduled prior to an operational OMV were assumed

to incorporate integral propulsion or use a spacecraft bus to

rendezvous with the orbiter for resupply.

o Materials processing in space conducted on free-flying factory

modules would be primarily involved in electrophoretic operations.

Results from the user contact efforts are summarized on the following pages

for the missions listed below:

Summary of Fluid Transfer Parametric Requirements Study Results

Parametric techniques were used to make estimates of the types, amounts and

yearly fluid usage rates required by the resupply mission model. Small

quantities of associated pressurants such as He and N2 were also required,

but their amounts were relatively minor in comparison to the primary fluids.

In some cases, depending on the receiver system design, no additional

pressurants are required for the transfer process.

While water is a significantly required fluid, it is involved in a modular

changeout in MPS factories and is thus not ideally suitable for transfer by

the resupply module. As expected, hydrazine and storable bi-propellant are

the dominant required fluids. In addition, another significant fluid of

interest was found to be liquid (primarily superfluid state) helium. The

original study scope excluded the consideration of OMV and OTV in the

development of the fluid transfer parametric requirements. This was changed

at the mid-term briefing but the requirements phase had already been

completed. However, OMV and OTV requirements will be added during the

follow-on. The results of the fluid transfer parametric requirements analysis

are shown in Table 4.

2082e/4-15-85 - 24 -
USER AND

CONSUMABLE TYPE

NASA

HYDRAZINE

LIQUID HELIUM

WATER

Xt-METHANE MIXTURE

COMMERCIAL

(POTENTIAL)

BI-PROP (N904/MMH)

HYDRAZINE

DOD*

BI-PROP (N204/A-SO)

HYDRAZINE

LIOUID HELIUM

DOD* (POTENTIAL)

BI-PROP (N204/A-50)

HYDRAZINE

CONTINGENCY"

1989

—

—

——

^

—

_

—

—_

—

1990

400

—

5,000

—

—

14,00_0

—

—_

—

1991

3.83_0

15,000

—

—

14,000

70

—

—_

1,500

1992

1,10_0

25,000

—

^

7,000

70

—

—_

—

1993

5,430

160

30,000

30

—

21,000

70

—

—_

—

RESUPP

1994

1,10_0

30,000

85

TM

14,000

17.470

2,400

—_

1.500

LY AMOI

1995

6,694

1,389

25,000

—

818

28,000

17.470

2,400

8,00_0

—

JNT (LB

1996

5.100

.

15,000

2,482

500

_

70

—

15,000

950

—

1997

1,100

992

5,000

85

2,307

1,000

35,000

8,770

1,200

2.000

250

1,500

1998

1,100

1,320

—

'

3,318

500

7.000

70

1.000

600

—

1999

5,100

992

—

4,694

500

7,000

17,470

2,400

15,000

950

2000

1,100

^

85

2.414

500

28.000

17.470

2,400

2,000

600

1,500

2001

1,100

2,312

—

4,694

500

14,000

70

—

14,000

950

—

•SMALL QUANTITIES OF PflESSURANTS ALSO REQUIRED FOR SOME SATELLITES - IMPORTANT FOR DESIGN REOT

•ESTIMATE OF PROPELLANT RESUPPLY REQUIRED IN CONTINGENCY SITUATIONS

• HYDRAZINE USED FOR THE MOST NUMBER OF UMBILICAL ENGAGEMENTS

| • GREATEST MASS OF FLUID FOR UMBILICAL TRANSFER IS BI-PROP"]

• WATER & SPECIAL GASES ARE TRANSFERRED VIA MODULE CHANGEOUT

Table 4- Fluid Parametric Requirements Summary

3.1.2 Candidate Spacecraft and Resupply Scenarios

Activity 1.1.3 was concerned with screening all the missions from the mission

model and selecting a smaller number for use in the development of alternate

resupply scenarios and mission operations requirements. The selection of a

representative set of missions occurred in two parts. The first was the

selection of missions which included the technical performance requirements of

all possible resupply candidates (e.g., LEO, polar, GEO, etc.). The second

step was a screening for missions that are likely to benefit from resupply.

The four general mission characteristics which were used are shown in Table 5«

Nine candidate spacecraft/platforms were selected. They occupied the three

general resupply locations of low-inclination LEO, high-inclination LEO and

GEO. Various single and multi-servicing scenario combinations were then

examined for these candidates. All of the candidates either require resupply

or benefit highly from its use and thus are promising missions to examine

closer for developing alternative resupply scenarios and requirements.

Using as input data the candidate spacecraft and space platforms selected for

Activity 1.1.3, a comprehensive set of resupply scenarios was generatd

(Activity 1.2.1). These scenarios included non-remote resupply options. In

addition, certain scenario trade issues, such as multiple versus single

spacecraft resupply advantage and disadvantages, wet versus dry launch

considerations, etc., were addressed as part of this activity. These data

were used in the concurrent effectiveness analysis to select the most

promising scenarios for further definition. Mission' operational requirements

were derived for these scenarios (in Activity 1.2.2) to develop

system/subsystem/interface requirements (Activities 1.2.3 and 1.2.4) for both

the effectiveness analysis and the Design Reference Missions (DRM) selected

2082e/4-15-85 - 25 -
GENERIC SPACECRAFT CLASS

• SCIENTIFIC EXPLORERS

• ASTRONOMIC OBSERVATORIES

• EARTH/WEATHER OBSERVATION

• HEWN/SURVEILLANCE

• NAVI6ATION

• COMMUNICATION-DELTA

• ENVIRONMENTAL OBSERVATION

• EARLY WARNIN8 000

• COMMUNICATION-IUS/TOS

• COMPLAT

• SMALL PLATFORMS

EUVE

CREP

6RO

AXAF

LANOSAT

OMSP

GPS

H.S.376

GOES

TORS

INTEL VI

PLANNED

PROTEUS

LEASE

CRAFT

MISSION OBJECTIVE

DURATION

SHORT 8 MO-24 MO

CONTINUOUS

CONTINUOUS

CONTINUOUS

CONTINUOUS

CONTINUOUS

CONTINUOUS

CONTINUOUS

CONTINUOUS

CONTINUOUS

MEDIUM ' 120 MO

UNIJ ON-ORBIT VALUE

JM84 SPACECRAFT

VERY LOW $50-100

VERT HIGH SSOO-1200

LOW-HIGH S80-300

HIGH-VERT HIGH

VERT LOW $50-100

LOW $75-125

7

7

MODERATE-HIGH $250-300

VERT HIGH

MODERATE $100-200

DURATION CONSTRAINT

(HISTORIC)

EXPERIMENT, FLUIDS

EXPERIMENT, INSTRUMENT

SENSORS, POWER

UNKNOWN

TECHNOLOGY OBSOLETE

7

7

TECHNOLOGY OBSOLETE

DESIGN REQUIREMENTS

CHANGEOUT/INVENTORY

COST OF ACCESS

(STS)

VERY LOW-MODERATE

LOW

MODERATE

MODERATE

HIGH

VERY HIGH

VERY HIGH

VERY HIGH

VERY HIGH

VERY HIGH

VERY LOW-MODERATE

LOW INCLINATION LEO

•AXAF

• MRS

•GRO

. • PROTEUS

HIGH INCLINATION LEO

• SIRTF

• DMSP

• 000-1

GEO

GEO PLATFORM

COMSATS

VARIOUS SINGLE & MULTI-SERVICING SCENARIO

COMBINATIONS EXAMINED FOR THESE CANDIDATES

Table 5. Candidate Spacecraft/Platform Selections

from this analysis. The system/subsystem requirements data defined as part of

1.2 also enabled the development cost/risk benefits data within the

effectiveness analysis to support justification of the DRM's.

Ten general resupply scenarios, shown in Figure 8, were developed for

operations in LEO. These scenarios included not only remote resupply, but use

of Orbiter-tended operations in order to insure a comprehensive examination of

alternatives. A spacecraft in LEO can be resupplied using a'resupply module

on a ground-based OMV, a space-based OMV or an OMV based at a space station.

These constitute the remote resupply options. The spacecraft can also be

resupplied using in-bay equipment as are being developed at JSC. Getting the

spacecraft to the Orbiter bay can be accomplished using the spacecraft's own

integral propulsion systems (IP), the OMV, or the Orbiter propulsion

capability. In each of the remote resupply options, the resupply module can

perform multiple or single engagements.

Several trade studies were identified in the original proposal related to

evaluating the advantages/disadvantages of the alternative resupply

scenarios. These studies focused on key discriminators among the various

resupply scenarios and were intended to develop data for the screening of

these alternatives in the effectiveness analysis. There were six small

studies conducted within Activity 1.2.1. The first examined the utility of

nodal regression effects in multiple engagement mission planning. The second

examined alternative space-basing options relative to ground-basing the

resupply module. This task also used much of the material developed in the

first study on multiple engagements. The third study looked at the pros and

cons of dry and wet launches of the resupply module itself. This had to

2082e/4-15-85 - 26 -
ORIGINAL PAGE fS

OF POOR QUALITY

SCOPED TO REMOTE vs

ORBITER-BASED

SPACECRAR

RETURNS

FOR EVA

EARTH

RESUPPLY ORBITER TO

SPACECRAFT

FOR EVA

d)

^f

3

D

I)

<-v

D

*-v

D

5)

OMV,n>

OMV,n =

OMV, n>'

OMV, n =

OMVMn>

OMV,,n =

i_ . . _

OMV GROUND

BA

1

1

SED

r (2)

k

REM

RES

OMV

SPACE

1 BASED

(41

1

! .:

r

' ^

__ ~ 4

k

OTE

JPPLY

(8)

i

SPACE

RESUPPLY

r

IS) l

w

»

IN-BAY

RESUPPLY

0 IN-BAY IP

T) IN-BAY OMS

RTEOMS

Figure 8. Low Earth Orbit Alternative Fluids Resupply Scenarios

include options for pumping from the OMV tankage, use in contingency

operations where the required fluid may not be known well in advance, and the

resulting on-orbit complexities involved if the resupply module arrives in

space dry. Another study was oriented toward GEO operations and examined the

weight sensitivities involved with expending versus reusing elements of a GEO

resupply system (OTV, OMV, EM). Cost factors were also identified for use by

the effectiveness analysis.

The final two studies supported development of system/subsystem/interface

requirements. Issues as to the appropriate levels of autonomy and redundancy

capabilities for the resupply module were addressed early in the study. We

evaluated both the OMV's capabilities and the potential role of

man-in-the-loop (M-I-L). Finally, we searched for critical issues in fluid

transfer technologies relative to the identified and required fluids.

Superfluid helium was found to have a variety of issues associated with its

transfer and thus would require future study.

Multiple Engagement Planning

Prior work on use of the Orbital Maneuvering Vehicle (OMV) for satellite

servicing raised the possibility of using nodal regression effects in multiple

satellite engagements. Nodal regression is due to the non-sphericity of the

earth and can result in changes of up to 9 degrees per day at low altitude and

low inclination orbits. Such effects decrease by a cosine term as orbital

inclinations increase and effectively cease for polar orbits.

2082e/4-15-85 - 27 -
An optimization algorithm was developed for planning multiple resupply

engagements. For real world situations of 2-3 engagements, at most on a

single resupply mission, closed form solutions are possible. These solutions

specify what sequence of orbital engagements is optimum, depending on whether

it is desired to minimize time or fuel usage. For a greater number of

engagements, the particular number varies depending on the mission complexity,

closed form optimal solutions are not possible and resort must be made to

various search techniques.

Resupply Module Basing Options

Figure 9 depicts the range of basing options examined in Activity 1.2.1.

Starting with filled resupply modules, we have two general ground-based

options and three space-based options. These options serve the four general

user groups denoted at the top of the page. Earlier work on space versus

ground-basing studies with respect to the OMV identified many of the current

issues related to the costs, risks and operational requirements of

space-basing.

USERS

SPACE

BASED

LOCATED IN

POLAR ORBITS

(LEO)

LOCATED IN

HIGH-INCLINATION

ORBITS (LEO)

L

SPACE-BASED AT

STATION IN THIS

INCLINATION

STANDARD ORBIT

SPACE-BASED AT

SPACE STATION

28.5° STANDARD

ORBIT

SPACE-BASED AT

STATION IN THIS

POLAR ORBIT

GROUND-BASED

SUPPLY OF WET RESUPPLY

MODULE (FILLED)

SUPPLY SOURCE

Figure 9« Resupply Module Basing Options

The specific conclusions drawn by the space vs. ground-basing study are shown

in Figure 10. Essentially, neither system is completely preferred, but rather

there needs to be a mix of both. The use of space-basing, particularly with a

space-based OMV, has significant economic advantages in the saving of

earth-to-space transport costs. Contingency missions, however, are likely to

require the sending up of a ground-based resupply module in that the specific

2082e/4-15-85 - 28 -
ja_eeds_ of jthjs mission may not be available on-orbit within an acceptable time.

The economics of space-basing argue for its implementation as operational

space experience is built up. This increase in experience will allow later

decisions on the exact number of vehicles that should be space-based and what

their optimal locations should be. Aside from the choice of a 28.5 degree

inclination, other parameters such as additional inclinations, altitudes and

nodal crossing times of space-based assets are still in question.

• EM BASING NEEDS TO BE CONSIDERED IN VIEW OF THE OPTIMALITY OF THE WHOLE

SYSTEM EFFECTIVENESS, NOT IN SEPARATE PIECES

• SPACE- & GROUND-BASING MODES ARE COMPETING & COOPERATIVE CONCEPTS

• SPACE-BASING IS PREFERRED FOR MEETING DETERMINISTIC RESUPPLY DEMANDS;

I.E., MISSIONS WHOSE RESUPPLY SCHEDULE IS KNOWN WELL IN ADVANCE

• GROUND-BASING IS PREFERRED FOR MEETING PROBABILISTIC DEMANDS, I.E.,

MISSIONS WHOSE SPECIFIC RESUPPLY REQUIREMENTS ARE NOT KNOWN WELL IN

ADVANCE, AS WITH CONTINGENCIES; ALSO* DUE TO LACK OF EFFECTIVE NODAL

REGRESSION EFFECTS, SPACE-BASING IN POLAR OR NEAR-POLAR ORBITS IS NOT

PREFERRED

• NODAL COLOCATION OF SPACECRAFT ("BEADS ON A STRING") SHOULD BE ENCOURAGED

FOR RESUPPLY EASE

• LONG-TERM SPACE EXPERIENCE WILL ACCURATELY DEFINE THE DETERMINISTIC

DEMAND FOR RESUPPLY & THUS THE RELATIVE EMPHASIS NEEDED ON SPACE-BASING

THE RM

Figure 10. Summary of Basing Mode Conclusions

Mission Timeline Analysis

Activity 1.2.2 focused on the development of mission operational

requirements. Specific spacecraft were used to develop alternative scenarios

to a greater level of detail. Starting with the functional descriptions

developed in Activity 1.2.1, mission timelines, delta V requirements, and

propellant weights were developed. The purpose of this was to take a closer

look at what the various alternative scenarios actually involved and to create

a base of data and experience for use in the effectiveness analysis. All of

the remote resupply scenarios developed in Activity 1.2.1 for LEO and GEO

operations were detailed. The candidate spacecraft/platforms selected in

Activity 1.1.3 were used to derive specific timelines and propellant

requirements. The use of the Orbiter in resupply operations was considered to

be outside of contract scope for this task. Resupply with the Orbiter was

examined in the effectiveness analysis as an option the remote resupply

scenarios had to compete with.

Figure 11 shows an example of a resupply mission which uses a ground-based OMV

in a single engagement with the Shuttle Infrared Telescope Facility.

2082e/4-15-85 - 29 -
486

6 7

1

1.

2.

3.

4.

5.

6.

7.

8.

^

2

 /

/1/1*'1 1 ' 1

0 2 y 24 26 28

PERFORMANCE EVENT

LAUNCH ORBITER, 28.5°

DEPLOY OMV/RM

BOOST FROM ORBITER ALTITUDE

CIRCULARIZE (COPLANAR)

TERMINAL RENDEZVOUS

PROXIMITY OPERATIONS

RESUPPLY SIRTF

DISENGAGE SIRTF

1 1 1 1

V

11 12

 A

I I A I I I I A I

30 32 34 36 38 40

 v 46 48 50 52 V 78

MISSION ELAPSED TIME (HR)

V Wp PERFORMANCE EVENT V

FT/SEC LB

561 379

548 350

• DEFINED 20 SCENARIOS TO GREATER LEVELS OF DETAIL

TIMELINES, AV'S, PROPELLANT REQUIREMENTS, ETC

9. RETROBURN W/.493"

PLANE CHANGE 586

10. CIRCULARIZE AT 150 NMI

W/437" PLANE CHANGE

11. ORBITER/OMV RENDEZVOUS 10

12. RETRIEVE OMV/RM

13. DEOR8IT

14. LAND

INCLUDING TOTAL OMV PROPELLANT

13

V

80

Wp

298

5

14231

1742 L

Figure 11.

SIRTF Servicing Mission Timeline (Ground Based OMV) 28.5° Orbit

2082e/4-15-85
Figure 15 shows a resupply mission to a large GEO platform using a Centaur G.

-The-OMV-used-is -ground-based-and-stored-at-GEO-for later-servicing

opportunities. The EM may also be eventually retrieved by a reusable OTV.

GEO (19.323) 0

134 138 142 146 ISO 154

MISSION ELAPSED TIME (HOURS)

PERFORMANCE EVENT

AV Wp

ft/sec Ib

LAUNCH ORBITER

DEPLOY CENTAUR/OMV/RM

BOOST FROM OBBJTER ALTITUDE

CIRCULARIZE AT 19,300 NMI

DRIFT TO PLATFORM LONGIT

i DEOHBIT

8,134

5,852

20

20,370

8.578

21

PERFORMANCE EVENT

158 162

tt/sec

WP

Ib

LAND ORBITER

RENDEZVOUS AT GEO ALT 5 5

WMWWWn OPCTATKJHS

RESUPPLY PLATFORM (1.764 Ib)

SEPARATE TO DORMANT ORBIT 22 23

STORE OMV AT Ah = 100 NMI

FOR LATER SERVICING

OPPORTUNITY

TOTAL OMV PROPELLANT 2,350 Ib

TOTAL RM WT (W/FLUIDS) 4.278 Ib

INITIAL MISSION WT (W/0 ASE) 46,796 Ib

Figure 13 GEO Platform Resupply Using Centaur G.

2082e/4-15-85
Figure 14 shows a multiple satellite engagement in polar orbits with a ground

based OMV. The DoD spacecraft being resupplied have not only different nodal

crossing times, but slightly different inclinations as well.

Ground Based — Initial Nodal Separation = 6.86 deg

-M/-1-

40 42 66 '76

 v 56 58 60 62 64

MISSION ELAPSED TIME. (HOURS)

78 82 98 100

PERFORMANCE EVENT Wp

LAUNCH ORBITER

DEPLOY OMV/RM

BOOST FROM ORBITER ALTITUDE 88 101

CIRCULARIZE (CO-PLANAR) 87 99

RENDEZVOUS & SERVICE DOD-1 10 11

BOOST TO DMSP W/2.380 1134 1104

PLANE CHANGE

CIRCULARIZE W/4.620 1992 1660

PLANE CHANGE

PERFORMANCE EVENT AV Wp

RENDEZVOUS & SERVICE OMSP 10 8

RETROBURNW/4.51" PLANE CHANGE 1965 1244

CIRCULARIZE AT 150 NM W/2.490 1221 658

PLANE CHANGE

OR8ITER/OMV RENDEZVOUS

RETRIEVE OMV/RM

OEORBIT

LAND

TOTAL OMV PROPELLANT 5.085 Ib

TOTAL RM WT (W/FLUIDS) 2.250 Ib

INITIAL MISSION WT (W/0 ASE) 11.306 Ib

Figure 14. DoD-1 and DMSP Multi-Servicing Mission Timeline

This chart and the proceeding ones are meant as examples of the range of

operational requirements that might be required of a resupply module and its

carrier. The data developed in producing these specific scenarios was used in

creating the functional allocation matrices for aystem/subsystem/interface

requirements definition. A major purpose of the development of alternative

scenarios was to establish the technical feasibility of the resupply options

and provide data for the effectiveness analysis evaluation discussed next.

3.1.3 Effectiveness Analysis

The Statement of ¥ork for the Spacecraft/Platform Expendables Resupply

contract requires that "....one or more spacecraft or platforms will be

jointly selected by the contractor and NASA for typical Design Reference

Missions (DEH)."

The dual objectives of Subtask 1.3 were:

2082e/4-15-85
1.- To select the-"best". DRMs for .subsequent use in Resupply Jfodule (Task 2)

design efforts, and,

2. To develop rationale and supporting data to justify authorization for

remote resupply demonstration.

The two primary outputs of the Effectiveness Analysis were:

1. Recommended DRMs for NASA/MSFC approval/concurrence, and

2. Justification for Remote Fluids Resupply demonstration.

These outputs have been developed by an iterative, "rough order of magnitude"

(ROM) parametric benefit/cost/risk assessment of resupply requirements for

each of 12 spacecraft programs developed in Subtask 1.1; in combination with

the mission profiles for each of 24 alternate fluids resupply scenarios

developed in Subtask 1.2.

Completion of the Effectiveness Analysis subtask required three sequential

activities:

1.3.1 Formulate Effectiveness Methodology

1.3«2 Generate Effectiveness Data

1.3*3 Recommend Design Reference Mission(s)

The methodology selected, and subsequently approved by NASA/MSFC, for use in

the Expendables Resupply study was developed by Rockwell International. The

methodology follows the (A-109 required) "justification for new start" process

by defining user NEED as a function of benefits provided, and estimating

concept AFFORDABILITY as a function of technical performance, cost, risk and

other factors.

The NEED/AFFORDABILITY method has been used on several similar NASA concept

evaluation study contracts (Alternate Thermal Protection System, NASA/LaRC;

Shuttle Derived Launch Vehicle, NASA/JSC; and Teleoperator Maneuvering Systems

Benefits, NASA/MSFC) and for internal Rockwell International Space

Transportation Systems Division project evaluations in the past four years.

A formal review of the methodology, supplemented by computer demonstrations

and NASA/MSFC participation in the prioritization of decision criteria for the

Expendables Resupply study, was completed on 11 May 84*

Each Candidate Design Reference Mission (DRM) subjected to the Effectiveness

Analysis consists of two major elements:

1. A candidate spacecraft/platform program in which the user's "NEED" for

Expendables Resupply has been estimated as a function of the economic

and/or operational benefits accruing to the program if fluids resupply

were available, and

2. an alternate fluids resupply scenario, for which parametric cost, risk

and complexity estimates are combined to determine that concept's

AFFORDABILITY" to the user program.

2082e/4-15-85 - 33 -
As an example of the NEED/AFFORDABILITY evaluation method applied to the

design reference mission selection, ROM parametric estimates of gross economic

and operational benefits have been made for each candidate user

spacecraft/platform defined in the potential resupply engagement model (from

subtask 1.1). An initial list of 30-40 potential receiver spacecraft has been

iteratively reduced by filtering out those programs in which significant

economic/operational benefits could NOT be identified.

ROM cost, risk and complexity estimates were made for each of approximately 20

alternate fluids resupply scenarios for Low Earth Orbit (LEO) missions, and

another four resupply scenarios for Geosynchronous Equatorial Orbit (GEO)

missions.

Parametric estimates were refined with each iteration to produce a "PRIME

CANDIDATE" list of approximately 12 receiver spacecraft programs for which

substantial economic/operational benefits could be estimated and to which

several of the 10 most cost/risk-effective resupply scenarios would be

applicable.

Groundrules and Assumptions

The USAF Space Division's "Unmanned Spacecraft Cost Model" was used to develop

many of the parametric cost estimates in the Expendables Resupply study, and

ranges have been applied to the Cost Estimating Relationships (CERs) to

indicate the degree of uncertainty associated with the estimates. Where

alternative sources were felt to provide better CERs, they have been used;

every reasonable effort has been made to reconcile CER estimates against

available programmatic data on a spacecraft-by-spacecraft basis.

The "Analytical Hierarchy Process" (AHP) technique (after T.L. Saaty) has been

used extensively throughout the study to convert subjective (engineering

judgment or managerial preference) data into quantitative "proxy" form. The

technique, with potential applications to the study, was discussed in

considerable detail during the May, 1984 methodology review with NASA/MSFC.

Prime applications of AHP have been, (l) to establish decision goals,

measurement criteria and the relative priority of each, (2) to estimate the

risk and complexity preference parameters for each alternate fluids resupply

scenario, and (3) to express the "contribution" of each alternative Candidate

Design Reference Mission to each "best DRM" decision goal.

The groundrules and assumptions for the effectiveness analysis are summarized

in Figure 15.

2082e/4-15-85 - 34 -
o ALL ECONOMIC & COST ESTIMATES IN CONSTANT 1984 $ MILLIONS

o PARAMETRIC ESTIMATES ARE ROUGH ORDER OF MAGNITUDE (ROM)

o USAF SD UNMANNED SPACECRAFT COST MODEL CERs USED

OMV ACQUISITION (MSFC BASELINE CONFIGURATION)

EXPENDABLES RESUPPLY MODULE (ERM) ACQUISITION

CANDIDATE SPACECRAFT/PLATFORM ACQUISITION

o ROCKWELL ORBITER DATA BASE CERs & INDEPENDENT ESTIMATES

LAUNCH (STS) & INSERTION STAGES

FLIGHT & MISSION OPERATIONS/SUPPORT

o RISK & COMPLEXITY ANALYTIC HIERARCHY PROCESS

BEST DESIGN REFERENCE MISSION DECISION RULE: SELECT CANDIDATE WITH

HIGHEST OVERALL CONTRIBUTION TO CONFLICTING GOALS

Figure 15. Effectiveness Analysis Groundrules and Assumptions

LEO Benefits Summary - NASA LEO ETR Spacecraft

As shown JLn Figure 16, expendables resupply capability is a necessary

condition for enabling satellite system availability through on-orbit

maintenance (which offers substantial cost efficiencies compared to the

traditional "proliferation" and/or "endurance" approaches).

Expendables Resupply as a stand-alone capability (without concurrent repair,

technology upgrade or routine maintenance), may not be sufficient except in

those cases where the only constraint on the spacecraft's function is

depletion of its on-board consumables supply. Situations where this is the

only constraint appear rare.

Rockwell's prior (1982-83) work with MSFC on the Teleoperator Maneuvering

System (now, Orbital Maneuvering Vehicle) Benefits Assessment identified five

general criteria which were useful in identifying the best candidates for

generic on-orbit servicing:

o Continuous Mission Objectives

o High Unit-Value Spacecraft

o Unit Spacecraft Life Limited

o Low Cost Access

o Non-Competitive Program Sponsors

All five are applicable to Expendables Resupply (a subset of generic on-orbit

servicing capability) and were used in the initial screening of the candidate

spacecraft list.

The economic value of concurrent maintenance and fluids resupply is determined

by the relationship between, (1) the spacecraft's remaining on-board fluids,

and (2) the spacecraft unit's functional availability retention as described

by a Weibull survival curve, for the nominally design spacecraft (Figure 17).

2113e/ - 35 -
d COSTS INCURRED

• ADDED PRODUCTION COSTS

• EXTRA TRANSPORT COSTS

SATELLITE SYSTEM

AVAILABILITY

PERCENT FUNCTIONAL

CAPABILITY

J_

ACHIEVEMENT OF

MISSION GOALS

ENDURANCE.

SATELLITE

MAINTAINABILITY

t)

ACOSTS INCURRED

• REDUNDANCY

• QUALITY

SERVICING CAPABILITY

SATELLITES DESIGNED

FOR SERVICING

1

i

RESUPPLY

1

REPAIR

REFURBISH

1

TECHNOLOGY

UPGRADE

1

MAINTENANCE

Figure 16. Satellite System Availability through Servicing

Only a small fraction of the spacecraft's initial functional capability can be

expected to survive beyond its nominal design life (at mean mission duration,

the probability of spacecraft function is approximately a/e = 37$). This

effectively limits the expected residual economic value of resupply WITHOUT

concurrent maintenance to the integrated area under the Weibull curve beyond

t

 = mmd>

To receive greater benefit from resupply without concurrent maintenance, the

spacecraft could be initially designed with a longer design life mmd, but at

exponentially higher unit acquisition cost for internal redundancy and

component quality.

Alternatively, the spacecraft might be designed for full serviceability with

the nominal design life mmd for only a modest increase in acquisition cost,

but with substantial increases in both spacecraft structural mass and volume

to accommodate on-orbit servicing by either EVA or remotely with the OM7.

2113e/ - 36 -
P(MS) %

P(MS) = e

NOMINAL DESIGN

• Wl = f (MMO')

P (SENSORS)*

P (SPACECRAFT).

HMD'

TIME

> Wf OPTIMIZED FOR MMD'

50

RESUPPLIABLE DESIGN

* A ACQUISITION COST-DRIVER

A MASS VOUJME SMAU

A TRANSPORT COST MINOR

MMD'

MAINTENANCE/REPAIR

TIME

SERVICEABLE DESIGN

A ACQUISITION COST MINOR

A MASS. VOLUME LARGE

A TRANSPORT COST-DRIVER

'MMD'^ TIME

Figure 17. Expendable Resupply Relationship for Spacecraft Design Life

Gross economic benefit of Expendables Resupply without maintenance were

estimated as the integrated area under Weibull survival curve past mmd for

each of the 12 prime candidate spacecraft/platforms. (Figure 18).

2113e/ - 37 -
ON-ORBIT VALUE (COST)

OBSERVATORY SPACECRAFT JMB4

UUNCH

VALUE SURVIVAL

P (MISSION SUCCESS)

' • SPACECRAFT

• MISSION EQUIPMENT

NASA OESIG RESUPPLY REMOVES CONSTRAINT ON LIFE

MISSION

PLATFORM THEORETICAL FIRST UNIT

PLATFORM DESIGN DEVELOP & TEST

E! I

GROSS ECONOMIC VALUE REMAINING

AFTER DEPLETION OF CONSUMABLES

,1 (INITIAL VALUE. 0)

DESIGN LIFE (CONSUMABLES DEPLETED)

Figure 18. Design Life (CONSUMABLES DEPLETED)

Estimating Gross Economic Benefit of Resupply Without Maintenance

Results of the analysis of reestimated gross economic benefit of fluids

resupply only (not including maintenance/refrain) for nine candidate

spacecraft are presented in Table 6.

AXAF

EOS

EUVE

GEOPLAT

GrPr B

GRO

LDR

PROTEUS

SIRTF

RESUPPLY ONLY

(No Maintenance)

ORIGINAL 1st Unit

On Orbit & Area

Value $M (t) mad $M

(Beta)

2.6 1,145 0.096 = 110

3.0 728 0.066 = 48

1.6 195 0.173 = 34

1.2 686 0.205 = 141

2.0 297 0.142 = 42

2.4 881 0.111 = 98

1.8 1,369 * 0.158 = 216

1.8 254 * 0.158 = 40

2.8 693 * 0.081 = 56

REPLACE SPACECRAFT

(Identical Technology)

REPLACEMENT ( 2nd ) Uni t

(TFU+Payload) *CRC+L&I = Cost

Limit $ Benefit of Service

(142 + 223) * .6 +110 = 329

( 77 + 173) * -6 +145 = 295

( 41 + ? ) * .8 + 10 = 43

( 92 + ? ) * .8 +185 = 259

( 43 + 93) * .6 + 11 = 95

( 94 + 272) * .6 +110 = 330

(143 + 314) * .6 +110 = 384

( 51 + ? ) * .8 + 13 = 54

(85+115) * .8 +145 = 305

Table 6. Economic Comparison of Resupply only versus

Replacement for Candidate Spacecraft

2113e/ - 38 -
Initial on-orbit values for each spacecraft were estimated from available

platform mass data and special mission equipment characteristic descriptions.

An imputed "value-added" for NASA design and test efforts and launch/insertion

costs were included. (Table 7) •

PROGRAM

AXAP

EOS

EUVE

GEOPLAT

GrPr B

GRO

LDR

PROTEUS

SIRTF

SP TEL

* Platform

PLATFORM

dry mass

(Ibs)

20,010

9,660

4,000

12,000

4,170

11,, 835

23,485

5,390

10,625

22,555

Contract

(estimates in $ 1984 million)

++Contractor++ NASA Payload

NR TFT DDT&E DDT&E

$M $M $M $M Ibs $M

135

75

44

89

44

84

139

50

79

152

Value

142

77

41

92

43

94

143

51

85

155

CER ($M84)

227

152

85

181

87

178

282

101

164

307

- 36

535

258

99

320

106

321

663

140

269

601

.7+. 012

3010

11000?

not

not

12100

26115

not

7440

1800

* dry

223

173

incl

incl

93

272

314

incl

115

?

mass

OnOrbit

L&I VALUE

$M $M

110

145

10

185

11

110

110

13

145

110

Ibs

1,145

728

195

686

297

881

1,369

254

693

1,310

Table 7. ROM ON-ORBIT VALUE OF NASA SPACECRAFT/PLATFORM PROGRAMS

Design life mmd were estimated from design provisions for consumables.

Residual value (beyond nominal design life) depends not only on estimated mmd

and Beta parameters, but on initial spacecraft installed (on-orbit) value.

From the wide variety of generic candidate spacecraft programs, the highest

unit residual values—and the best bets for servicing and resupply—are found

in the astronomic observatory, reconniassance/surveillance and earth

observation satellite classes. (Figure 19)-

Significant operational benefits can be expected to accrue to those programs

where Expendables Resupply removes (or relaxes) an existing operational

constraint, such as for surveillance programs (Figure 20).

2113e/ - 39 -
PRIME CANDIDATES FOB SERVICE 4 RESUPPLY

• LONG-TERM MISSION OBJECTIVE

• HIGH RESIDUAL VALUE

• USEFUL LIFE CONSTRAINED

WEAROUT, FAILURE

CONSUMABLES DEPLETION

TECHNOLOGIC OBSOLESCENCE

LOW COST ACCESS FOR SERVICING

"PERMANENT"

ASTRONOMIC

OBSERVATORY

..NAVIGATION OBSERVATORY

01Z34S678 9

Figure 19. Relative Residual Value of Generic Satellites

NOTE: SPECIFIC ECONOMIC & OPERATIONAL BENEFIT ESTIMATES CAN BE MADE - BUT MAY NOT BE USED

RELATIVE COST RATIO

2.0

TIME

1.0

CO

e

o

SURVEILLANCE B

\ PLANNED MANEUVERING CONSTRAINED

^ SURVEILL

\

\

\

\

4.0

3.0

2.0

1.0

• CONTIN

- • HIGH Ul

• DURATK

JOUS OBJECTIVE

m VALUE

)N CONSTRAINED

LAUNCH

i

LAUNCH «

tJ

INITIAL SATEl

I |

Mil

REFUEL #1

LITE

1

}

LAUNCH #4

} MINI

-Sl|$.

s

iTm

JiiLL TREFUEL #m2

{LAUNCH #2

i i i i

L\ SURVEILLANCE

A i

^

\X V PLANNEDVWP REMAINING

ACTU

WP

WITH PLANNED MANEUVERING

I

\

 Nl

TIME

SIGNIFICANT OPERATION BENEFIT ]

K

REFUEL

• REMOVES CONSTRAINT

ON MANEUVERING

• INCREASES SYSTEM

AVAILABILITY

TIME

•CONTINGENCY (UNPLANNED) MANEUVERING

TIME

Figure 20. Potential Need of Surveillance Programs For

On - Orbit Remote Consumables Resupply

2113e/ - 40 -
Prime candidate spacecraft in low earth, high inclinastion orbits include

selected DoD programs and, for superfluid helium transfer technology

demonstration rather than for economic benefit motivations such as the SIRTF

program as shown in Table 8.

PROGRAM

(WF)

EOS

2,000 LB; N2H4

DOD

? LB ?

GrPrB

400 LB; N2H4

SIRTF

1000LB; LHe

GROSS

ECONOMIC

BENEFIT

48

42

56

MAJOR

OPERATIONAL

BENEFIT

MANEUVERING

SURVIVABILITY

TECHNOLOGY

DEMONSTRATION

APPROXIMATE COST

RESUPPLY SCENARIO

1

38

28

32

2

52

44

46

7

23

12

12

8

23

12

12

MAX NET

ECONOMIC

BENEFIT

S84M

<28

<30

<44

NOTES: (Estimates in $ 1984 Million)

1. SCENARIO 2 NEVER COST EFFECTIVE AT WTR

2. SCENARIOS 3, 4, 5, & 6 NOT FEASIBLE AT WTR

3. SCENARIOS 9, & 10 COST EFFECTIVE FOR GrPrB

4. DIFFERENT NODAL CROSSING TIMES MAY PRECLUDE MULTIPLE

RESUPPLY UNLESS SPACECRAFT MANEUVERS

Table 8. Fluids Resupply (Without Maintenance) Benefits for Polar Orbits

Neither of the space-based OMV/RM alternate fluid resupply scenarios are

thought to be technically feasible in high (polar) inclination orbits, due to

the absence of high relative nodal regression rates which are required to

achieve ascending node congruity between a space-based resupply carrier

vehicle and the receiver spacecraft. Use of the ground-based OMV/RM mode for

single resupply engagements is not cost-effective at WTR (Vandenburg AFB) due

to the high transportation charge (approximately 250^ of ETR) on the inert

mass of the OMV/RM cargo unit and the technical feasibility of dual resupply

engagements in high inclinations using a ground-based OMV is an unresolved

issue*

Results of the parametric economic benefits analyses (Figure 21) indicate that

NASA LEO ETR spacecraft programs could receive a net economic benefit by

utilizing an Expendables Resupply capability, even without concurrent "full

servicing". However, the marginal benefit to each of these NASA ETR

representative programs which would be received by taking advantage of

comprehensive repair, instrument changeout and fluids resupply is so

substantial there will be a strong motivation to opt for Orbiter-based,

man-in-loop repair until teleoperation maintenance provides routine remote

servicing capability.

2113e/ - 41 -
RESUPPLY (W/0 CONCURRENT MAINTENANCE & REPAIR) (+) NET BENEFIT

V OBSERVATORIES & EXPLORERS

• HIGH UNIT-VALUE SATELLITES

• LOW COST ACCESS FOR RESUPPLY

V COST-EFFECTIVE ONCE PER PROGRAM

SPACECRAFT

AXAF

EUVE

GRO

LOR

PROTEUS

GROSS ECON

BENEFIT

$ffl84

RESUPPLY ONLY

73-110

23-34

65-98

144-216

27-40

COST TO

HESUPPLY

Sffl84

6-20

• 7-21

15-29

9-43

9-23

NET ECON

BENEFIT

SM84

RESUPPLY

53-104

2-27

36-83

101-207

4-31

196-452

ONE

BESUPPLY

RESUPPLY WITH CONCURRENT MAINTENANCE & REPAIR^ NET BENEFIT

y' ORDER OF MAGNITUDE

INCREASE IN BENEFITS

y' "FULL SERVICE" WILL

BE PREFERRED OPTION

SPACECRAFT

AXAF

EUVE

GRO

LOR

PROTEUS

GROSS

ECON

BENEFIT

SM84

219-329

29-43

220-330

256-384

36-54

COST TO

RESUPPLY &

REPAIR

SfflB4

23-80

8-14

32-87

31-119

10-17

NET BENEFIT

PER REPAIR

MISSION

Sffl84

139-306

15-35

133-298

137-353

19-44

NO

ENGAGEMENTS

PER

PROGRAM

4-6

1-2

2-4

2-3

3-4

TOTAL PROGRAM

NET ECONOMIC

BENEFIT

Jffl84

558-1836

15-70

267-1194

275-1060

56-177

1171-4337

MULTIPLE

ENGAGEMENTS

Figure 21. Net Economic Benefits Summary for

NASA LEO ETR Spacecraft

Even though the value of full servicing—repair, resupply, technology

changeout and routine maintenance—is limited to the spacecraft's replacement

coat (second unit + replacement launch and insertion), the net economic

benefits received from full servicing in major NASA ETR programs far exceed

the benefits of Expendables Resupply alone. Since stand-alone resupply can be

cost-effective only once during most programs, while on-orbit maintenance and

repair can be cost-effective many times. Figure 22 presents the results of

Figure 21 in the form of columns to more clearly illustrate the relative

benefits of the two approaches (Resupply alone versus Resupply with

Maintenance/Repair Servicing.)

GEO Satellite Resupply/Servicing Economics

Commercial operations in Earth orbit are primarly concentrated in

Geosynchronous Earth Orbit (GEO), located approximately 19,500 nautical miles

above the Earth's Equator. In this location, there are currently about 70

commercial communications satellites which provide television, telephone,

telepgraph, data transmission, and other communications services world-wide.

They originate from a vareity of countries and organizations, and represent an

investment of several billions of dollars. Over the next decade the number of

satellites in GEO will increase, until by the year 1990 it is estimated that

there will be over 150 commercial satellites representing a direct investment

of over 10 billion (thousand million) dollars in GEO.

2113e/ - 42 -
400

300

200

100

NET ECONOMIC BENEFIT TO

NASA LEO PROGRAMS

PER ENGAGEMENT

FULL SERVICING

• RESUPPLY

• REPAIR/MAINTENANCE

TECHNOLOGY UPGRADE'

TOTAL BENEFIT 5 PROGRAMS

FULL SERVICING

• RESUPPLY

REPAIR/MAINTENANCE

TECHNOLOGY UPGRADE

RESUPPLY

ONLY

4.0

3.0

2.0

1.0

Figure 22. Economic Benefits of Full Servicing Relative to

Eesupply Only For Selected LEO Satellites.

A typical GEO satellite is designed to reliably provide service over a 7 to 10

year period, at which time it is retired from service by boosting its orbit

several hundred miles into a "junk orbit" to free its valuable GEO location

over a specific longitude, for a new user. Due to the high orbit, no effort

are made to return the satellite for reburishment, nor are any methods

currently available to repair a malfunctioning satellite.

The large amount of investment located in this orbit offers a unique

opportunity, in that, within this one orbit is located the vast majority of

current commercial investment in space. Interest has been expressed in the

possibility of supplying services to GEO satellites; in particular to resupply

them with attitude control system fuel (ACS) needed to maintain the

orientation of the satellite and aim its antennas, and to replace or upgrade

the electronics on board.

A commercial communications satellite venture makes its revenues from

providing a service to customers; that is, it allows the usage of the GEO

satellite's transponders for a fee. These service conracts may be for a very

few hours, for several years, or the customer may buy an exclusive right to

use the transponder for the expected life of the satellite.

To provide this service, the commercial satellite operator must first:

2113e/ - 43 -
o establish financial backing for his venture,

o apply for and receive a permit to operate from the appropriate regulatory

agency,

o book launch space on a launch vehicle,

o arrange financing for the launch of their satellite,

o arrange launch and operations insurance,

o establish an organization for monitoring and operating the satellite over

its expected life.

Almost all of these milestones are performed in the private sector, and

several private sector service or manufacturing operations are avalable to a

new commercial satellite operator to accomplish these functions in each area.

Only the regulatory function is being exclusively retained as a governmental

responsibility. A GEO satellite is run as a business venture, and future

operations will be governed by standard business decision analyses. The

desirability of GEO servicing and resupply to a commercial user in the future

will be determined by these decision methods.

GEO SATELLITE COSTS AND REVENUES

In order to assess the desirability of GEOsat servicing or resupply to a

commercial satellite operator, it is first necessary to determine the expected

costs and revenues they are likely to encounter. Parametric relationships

were determined between fundamental characteristics of the GEOsats wherever

possible in order to assess the sensitivity of these factors upon the user.

Satellite Production Costs

The initial cost of buying a GEO satellie is a function of several factors,

the chief being the physical size of the satellite, whether or not it must be

designed uniquely for a specific commercial customer, and the required

lifetime of the GEO satellite. In this analysis, the basic satellite cost was

taken to be:

TFU cost = 2.02xm°-55 (i)

where m = GEOsat BOL weight in pounds and TFU cost is in millions of 1984 $•

This relationship is based upon a cost estmating relationship derived by D. E.

Koelle in AIAA-84-0704, "GEO Space Platform Economics", and is in close

agreement to ROM estimating relationships used internally. Application to

Intelsat VI and DBSC satellite systems fitted quite well.

The actual cost to the user is further dependent upon the number of satellites

built, and the development costs which went into constructing the first

satellite of this type. A Design, Development, Test and Engineering (DDT&E)

cost equal to 130$ of the TPU cost was assumed to be incurred prior to the

first satellite constructed, and then the DDT&E cost was spread evenly between

all the satellites.

The experience of producing the GEO satellites will allow a reduction in the

per satellite cost, dependent upon the number of satellites produced. Based

upon quoted prices for Ford Aerospace's "Supersat" series, and the Hughes

Satellite System's "Intelsat VI" series, an 80% learning curve was taken to be

typical. Since orders for Hughes Satellite Systems "HS-376" series of

satellites has reached 30, with Intelsat expected to buy from 5-14 of the

Hughes "Intelsat VI" series satellites, and Ford Astronautics expected to

build at least 5 of their "Supersat" series, a lot size of 10 satellites was

taken to be typical, and the average cost for a 10 satellite batch was assumed.

2113e/ - 44 -
Including DDT&E costs and learning curve effects into the satellite's

procurement cost yields:

PC = (2.02 m 0-55) H(ln(.8)/ln(2)) [1+1.3/1] (2)

where N is the number of satellites over which the costs are averaged.

This relationship was primarily derived for current generation communications

satellites, and does not include any sensitivity of the satellite cost to

increased reliability for a longer satellite life. As the satellite's life on

orbit increases, incorporating more reliable components and systems will

increase the overall satellite costs. The relative variation in TFU cost as

life increases was taken to be:

TFU' = TFU x [0.15 + 0.85 e(°.32L . 0.64)] (3)

where L = the mean mission duration in years, and the 0.15 and 0.85 factors

represent a 15% and 85% split between the non-electronic and electronic

subsystems. This relationship was based upon a board and component level

analysis of spacecraft subsystems during the TMS Benefit Assessment contract.

It is documented in Voume II, of that contract report.

ACS Usage and Number of Transponders

Annual ACS usage was derived from the following data, obtained from several

sources:

Satellite

Galaxy (HS376)

HS299

ACTS

MSAT

TDRS

SBS

STC

BSGP Platform

(GD Study)

ACS Fuel (Ibs) Lifetime (yrs) S/C Wt (ibs) Annual Usage

300

139

344

494

1,300

250

305

1,764

9

10

8

10

10

7

7

7

1,145

725

1,968

2,909

3,700

1,050

1,105

7,092

33-3

5-2

43-0

49-4

130.5

35.7

44.0

252.0

From this data the trend line was determined to be

ACS = 15.6 + .036m (4)

where m - GEOsat BOL weight in Ibs and ACS = annual usage of ACS fuel in Ibs

per year. See Figure 23«

However, the fraction of the GEO satellite that is occupied by the

communications payload varies with the mass of the satellite. Using the

estimating relation developed by Koelle, the following equation is best found

to estimate the fraction of communications payload to total GEOsat BOL weight

x = -.012 + .032 In m (5)

where m is again the BOL weight in Ibs. Thus the estimated weight of the

communications payload is F, transponder plus antenna weight.

P = -.012m + .032m (in m) (6)

2113e/ - 45 -
200

UJ

o

co

Ysi -s~

100

X

X

X

X

D

X

X TREND LINE (R2 - 0.94)

X ACS = -15.6 + 0.036 M

'1

1000 2000 3000 4000 5000 6000 7000

S/C BOL WEIGHT (Ibs)

Figure 23- GEOSAT ACS ANNUAL USAGE

The actual number of transponders in the payload is again variable. However,

to fully and precisely account for the number of transponders a very high

level of detail is required including the frequency, number, type, bandwidth,

and encoding scheme used on each channel. In the time available for this

effort, the detailed data was not available and the "ROM" estimators of 4/6

GHZ (C bands) transponders - 1 equivalent, 12/14 GHZ (Ku Band) - 2.5

equivalent, and 30/20 GHZ (KA Band) - 5 equivalent transponders were

used. Thus we find

2113e/ - 46 -
Sat

ABCS

Amersat

Fordsat

Galaxy

Gstar

HS399

SatcomK

Spacenet

Telstar

ECS

TelecomAnik C

Anik B

Anik D

No. Transponders

16KU

18C + 6Ku

24C + 24K

24C

16Ku

12C

16K

u

18C + 6K

u

24C

12Ku

5C + 6K

16Ku

12C + 6K

u

24C

Equivalent

Transponders

40

33

84

24

40

12

40

33

24

30

20

40

27

24

BOL wt.

(Ibs)

2,400

1,380

3,300

1,146

1,380

725

1,463

1,260

1,435

1,345

1,150

1,250

1,014

1,350

Comm. P/L

¥T (Ibs)

569

303

816

245

303

144

324

273

317

294

246

270

212

295

which may be best approximated by the relationship •

H « 4.197 + .089?

(7)

where N = number of equivalent transponders, and P = communicatons payload

mass in Ibs. (see Figure 24)

Satellite Revenues and Technological Obsolescence

Revenues from each GEOsat are dependent upon the number of transponders on the

GEOsat, and the price charged per hour or usage. Based upon quoted rates for

Satcom IV and Telstar 3 services, a rate of $200 per hour over a period of

time is typical. For short term leases of transponder facilities, this rate

will be much higher, but the proportion of time that the transponder is

actually in use may be much lower, depending on market conditions.

2113e/ - 47 -
80-

2 60

S40

X

x

X

X

X

X

X

X

X

X

X

X

0.089 P

X

x

20-

TOO 200 300 400 500 600 700 800 900

COMMUNICATIONS PAYLOAD WEIGHT (Ibs)

Figure 24. Equivalent GEOSAT Transponders as a Function

of Communications Payload Weight

The actual level of reimbursement for usage of each transponder is determined

by the supply and demand for transponder services in the maketplace, and the

individual capabilities of each transponder. To get a workable answer without

digressing into a treatise, the supply and demand levels for transponder

services are assumed to be equal, maintaining the price per transponder hour

at a constant level. Excess capabilitity is assumed to be filled by a growing

demand for GEO satellite services. Thus, without changes in the capability of

a transponder (a constant technology level), the revenues generate, and the

competitive position of a satellite will remain constant in the market.

If an "improved" transponder, possessing a higher throughout capability than

current technology transponders, is placed in GEO and its services are assumed

to be offered at the market rate of $200 per hour, it will attract customers

from existing systems. This will force a cut in the market price of the

existing transponders to match capabilities to market price.

2113e/ - 48 -
Historical changes in capability per transponder and future projections are

shown in Figure 25- Overall, this represents about a 12% annual growth in

transponder capability. Will this rate decline? Probably so; the most recent

growth in capability has been at about 6% annually, but recent developments

indicate there is still quite a bit of improvement achievable in GEOsats in

the future. This analysis assumes the historical rate of improvement in

transponder capabilities will continue.

If the asking price for transponders is assumed to remain constant, this rate

of improvement may be taken to be a technological obsolescence factor. In

other words, suppose a satellite operator launches a GEOsat in June, 1984 and

charges $200 per hour for use of its transponders. If another GEOsat operator

else puts June 1985 technology in their satellite and launches it, while

charging the marke price of $200 per hour, the June 1985 GEOsat operator will

be supplying (based on past trends) about 12$ more capability for the same

price. If the first operator doesn't take action, all their customers wil

switch to the new service to achieve more capability at the same price. To

keep them on board the first operator will have to discount their asking price

by 12$ to match offered capabilities to market price. If someone else

launches a June 1986 satellite, it will have an additonal 12% of capability,

and the June 1984 satellite operator will have to again discount their GEOsat

services price by another 12$ to match.

As time goes on, a satellite in GEO will continue to generate revenues, but

the rate at which the revenues are generated must be discounted to account for

more efficient competitors. Obviously, year by year changes are almost

impossible to predice while the overall effect averages out with longer times.

Mathematically, this is represented as

*

L

Total Revenues = £„., ft (/fr,) o h (8)

where R0 = initial revenues

= (# transponders) ($200/hr) (8750 hours/year)

rip = annual obsolescence rate

L = life of satellite in years

since this process may be generalized to be a continuous process

Total Revenues = C T> f>~re*i.

-U

 K°e d*

where rc = continuous discount rate

rc •» In (l + rj)

if rj = 12$ then rc = 11.33$

Solving,

Total Revenues = Rp_ (l - e~rc

L) (9)

RC

Note that to derive total revenues, one must only know the BOL weight of the

GEOsat. Knowing this, one may find

- the orommunications payload weight (eqn 6)

- the number of transponders (eqn 7)

- the total revenues over the satellite life (eqn 9)

2113e/ - 49 -
ce

UJ

o

a.

Ui

7000

CO

C3

UJ

a.

V)

5000

S£ 3000

S 1000

7200

21.7% ANNUAL

GROWTH

6.5% ANNUAL

GROWTH JX /

/

w

S 11.65% ANNUAL

X GROWTH LINE

1200

8.5% ANNUAL

GROWTH

I I

1970 1975 1980

Figure 25- Transponder Capability Growth

1985 1990

Total Costa and Revenues

Total costs to operate the satellite over its lifetime are comprised of TFU

(production cost), operations, insurance and transportation. An evaluation of

TFU cost and its sensitivity to satellite lifetime is given in eqns 1 and 2.

Operations cost may be taken to $2 million per year, (from typical current

systems) so total operations cost is 2L. Insurace may be assumed to be

(TFU •*• Transp.) (10)

for just replacement value. The insurace rate, rj_, is around 10% (although

current rates are higher due to recent upper stage failures).

An estimator for transportation price was obtained from the data in Figure 26

which summarizes the results of a previous analysis which compared the total

transportation costs to the user (excluding insurance) for various space

launch systems. This incorporates launch costs, payload integration and

processing costs and the procurement of an upper stage system if required.

2ll3e/ - 50 -
iza

100

UJ

u

ce^

|i 75

«/sz

zo

§^ 50

1

25

0

1

^•^-293.19 + 44.59 Ln m

x

X

X

X

• ,/•

/ *

1

1 1 1 I 1 1 1 1 I 1 1 1 1 I

3 1 2 3 4 5 6 7 8 9 10 11 12 13 14

Figure 26.

GEO WEIGHT DELIVERED (THOUSANDS OF Ibs)

Transportation Price As A Function Of

GEOSAT Weight

2113e/ - 51 -
Transportation price was derived from the following data:

Systems GEO Wt. Transportation

(Ibs) Price (1984 $)

Ariane - 3 (single) 2,800 60-3 \

(dual) 1,350 35-5

Ariane - 4 (single) 5,130 85-2

(dual) 2-340 47-3

Atlas/Centaur 2,500 61-5

Atlas-G/Centaur 2,600 50.5

Atlas II/Centaur 6,700 88.7

Delta 3920 1,400 41.4

Titan 34D/Transtage 3,950 106.5

STS/Centuar F 13,600 129-4

STS/Centuar G 10,600 129.4

STS/IUS-1 (HI) 5,000 90.1

STS/IUS-1 (Hughes) 4,410 72.0

STS/PAM-A 2,200 45-9

STS/PAM-D 1,350 23-7

STS/PAM-D2 1,800 31.6

This data may be approximated by

Transp = -293-19 + 44.69 In m (11)

Thus, the toal cost of launching and operating a GEOsat is found to be:

Cost = TFU + Ops + Insur + Transp.

- 2.02m 0.55 + 2L + .l(2.02m°'55 - 293-19 + 44-69 In m)

-293-19 + 44-69 In m

- 2.22m °-55 + 2L + 49.16 In m - 322.51 (12)

where m = GEO sat BOL weight in pounds, and L = lifetime in years.

If we want to service a satellite at GEO, we need to also consider how the

costs and revenues increase. Without technological updating of the GEOsat,

the total revenues may be found using eqn 9, and just substiuting the new

satellite lifetime L for the life of the satellite. The production costs

change substantially, however. As before,

Cost = TFU1 + Ops' + Insur' + Transp1 (13)

TFU cost increases since there may be a longer life required for the

satellite. This may be derived from Eqn (3)- The increased TFU cost, TFU1,

may be expressed as

TFU' = TFU [0.15 + 0.85e (0-32L1 - 0.64)1

[0.15 + 0.85e (0-32(10)_- 0.64);

= TFU (.0897) (0.15 13]

TFU [.0135 + .0763 e (0-32L1 - 0.64)] (14)

Operations cost is much simpler

Ops' - 2L1

Transportation cost for a resupply may be taken to be

Transp1 = Transp +• Resupply (15)

Where Transp is the transportaton cost from equation (ll) and Resupply cost is

Resupply = (annual fuel usage)(Years resupplied)(Cost/lbs)

2113e/ - 52 -
Since resupply may be made to seveal GEOsats on one mission, we cannot use equ

(11) here, and shold use the minimum relasitic cost per pound to deliver fuel

in GEO. This derivation is~ shown below:

Weights (Ibs) Cost (1984 $M

Payload to GEO 13,800 84M

(full STS flight required)

Centaur G' 40M

OMV

(assumed 20 flights

into 25 flight life) (3,500) 10M

OMV fuel (1,000)

ERM (1,000) 5M

Tankset (1,000) 5M

Total GEO fuel 7,300 Ibs

Total cost $154M

The minimum cost of resupplied fuel in GEO is then about $21,100/lb, and eqn

(15) becomes (in millions)

Transp' - -293.19 + 44.69 In m + A(L' - L) (.021) (16)

where A is the annual fuel usage from eqn (3)

Transp' - 293.19 + 44.69 In m + (6.45 + 0.21m) (L' - L) (.021) (17)

The GEO fuel weight of 7,300 Ibs is the total mass installed in a GEO

satellite. The weight allocation may be used for some combination of fuel and

replacement parts, rather then fuel alone. This needs to be considered when

performing concurrent servicing with resupply as discussed later.

Insurance is a bit more complicated. The insurance will have to include a

premimum on replacement cost, the resupply cost, and possibly expected future

revenues. Ignoring future revenue insurance, the insurance cost would be

Insur' - rx (TFU' + Transp') (18)

- .1 (TFU' + Transp')

where Transp' and TFU' are defined from eqns (17) and (14).

Combining all these equations, the resupplied satellited system's costs are

Cost = (l.l)TFU [.0135 + .0763 e(-32L' ~ -

64>]

-I- 2L' - 293.19 + 44.69 In m

+ (6.45 + .021m) (Lf - L) (.021) (1.1)

- 2.22m-55 [.0135 + .0763e(-32L' ~ -

64>

+ 2 L' - 293.19 + 44.69 In m

+ (.149 + .0005 m)(L' - L) (19)

This is a somewhat complicated equation, but it only needs GEOsat BOL mass m,

extended lifetime of the satellite L', and normal satellie life » L (usually,

L = 10 years), to determine GEOsat mission cost.

The total revenues achieved from this operation are found to be:

Revenues = ^ 1.75(4.197 + .089[-.Q12 + .032m +.032m(ln m)] (20)

2113e/ - 53 -
The total return from the satellite is

Total Return » Revenues - Cost (21)

and this may be calculated from the above equations. However, this

formulation does not include any recognition of the time value of money. Net

Present Value recognizes this effect, and may be found from

L

NPV=< Cn/fl-f-K.') -X0 (22)

where C± « is the annual cash flow from the project (usually the revenues

less expenses), IQ is the initial investment, and k is the appropriate

opportunity cost of capital (here taken to be 20%).

To compare projects on a percentage basis, the IRR or "Internal Rate of

Return" "oh each project may be found from:

0-t t£\*\« "To <23>

Projects with a higher IRR than their alternates should be chosen.

IMPACT OF GEO RESUPPLY AND SERVICING

When the GEOsat is merely resupplied, there will be no technological updating

of the satellite's electronics, and the market position of the GEOsat will

continue to decline after the satellite resupply. Revenues will be earned,

but the stream will continue to decline with time. Extra cost may be incurred

if Increased reliability is added to the satellite before launch to ensure

that it will continue to operate past its initial lifetime, through the period

for which it is resupplied.

If GEO servicing is performed upon the GEOsat, then it may be assumed that the

technological obsolescence in the satellite's communications payload wil be

eliminated at that time, and the revenue stream earned by the GEOsat will

return to the market position. There will be an increased cost of performing

the the mission, as extra weight will be carried into GEO and the required

electronics boxes will cost more than that of the fuel alone. However, the

increased costs of including extra reliability to ensure a longer satellite

lifetime will be avoided, as the GEO servicing capability should also allow

replacement of any failed subsystems, returning the system to the same

reliability as when it was launched.

Three basic methods of operating a GEO satellite system with GEO servicing

and/or resupply were considered:

Lifetime Extension - was considered to be the extension of an exisiting GEO

satellite's life via GEO resupply. This in turn may be broken down into

several variations including:

o Planned Resupply: the GEOsat has been built with the extra redundancy

and reliability needed for the longer service life.

o Opportunity Resupply: a GEOsat not built with the extra redundancy and

reliability for extended service life is found to be operating

satisfactorily; and is resupplied in GEO to extend its life.

o Resupply and Concurrent Servicing: besides resupply of ACS to extend

the GEOsat's lifetime, the electronics in the communications payload

are updated to remove the technological obsolescence which has occured

since launch or the last servicing.

2113e/ - 54 -
Graphically, these modes are shown in Figure 27. The Internal Rates of Return

(IRR) for these modes are shown below and graphically in Figure 28.

Geosat Planned Opportunity Resupply

(Ibs) Baseline Resupply Resupply Service

1000 16.7%2.5%17.3%20.5 %

2000 24.8 8.8 25.3 27.8

5000 52.0 25.3 52.1 53.1

7000 67.5 40.0 67.5 68.1

10000 87.9 45.1 87.9 88.1

15000 118.1 61.1 118.0 118.1

20000 114.6 75.1 144.5 144.6

Since the favored project has the highest IRR, planned resupply is seen to be

not preferred over the alternatives. The highest IRR is obtained from

resupply and service.

IRR (Internal Rate of Return) calculations do not always give an accurate

assessment for projects of different durations. IRR implicitly discounts cash

flows occurring in the future and tends to underestimate a longer term project

compared to a shorter tern. To resolve this, a comparison of the "Equivalent

Annual Annuity" for each project was made as shown on Figure 29.

This was done by first calculating the Net Present Value (NPV) of he project

using

NPV =^ Ct - C0 + G! + C2 + + CL

4=0 ~^^t ~O+r7 (1+r)^ (1-t-r)1-

where Ct - the total cash flow (revenues less costs) in each year, and the

cash flows are discounted at r = Weighted Average Cost of Capital for the

project (This may be thought of as the equivalent interest rate earned by this

money in alternative investments plus the required profit margin in the

company). A company should take only projects with a positive NPV (i.e.,

those that earn more than investing in riskless securities and meet minimum

profitability levels to cover their riskiness).

However, NPV as such does not easily compare projects of unequal duration.

Let us assume profitable projects will be repeated at the end of their life

(it's reasonable to assume that a profitable commercial comsat will be

replaced at the end of its lifetime).

2113e/ - 55 -
•INCLUDES EXTRA

REDUNDANCY FOR L

Figure 27. GEO Comsat Lifetime Extension

A NPV of say, "x" (NPV » X), may be considered to be equal to a constant

annuity of A. That is, having an NPV = X is the same as earning A dollars per

year each year the project is run. The equivalent annual anuity is found from

L i. r,

A . =/A£ i •- A -

This allows the comparison of projects of different duration, Projects with a

higher constant annual cash flow are preferred.

2113e/ - 56 -
80

70-

LU

ee

u.

50

y ^*40*

h-

<30

cc

LU

520

10

0

RESUPPLY & SERVICING

OPPORTUNITY RESUPPLY

'PLANNED RESUPPLY

1 2 5 7 1 0

GEOSAT WEIGHT (THOUSANDS OF Ibs)

Figure 28. IRB of Lifetime Extension

Figure 30 is an extension of the material of Figure 29 modified to indicate an

example of the impact of various penalties for including servicing/resupply

capability into the satellite design. Three basic cases were examined in this

analysis and the data was linearly extrapolated over the range of satellites

examined.

The data used was determined by two data points in each model:

15

MODEL

Low

Medium

High

Geosat Vt.

(LBS)

3,000

3,000

3,000

Vt. Impact

(*)

5

10

20

Geosat Vt.

LBS

5,000

5,000

5,000

Vt. Impact

(*)

2.5

5

10

2113e/ - 57 -
400

300

200

100

RESUPPLY &

SERVICING

OPPORTUNITY BASELINE

RESUPPLY

PLANNED RESUPPLY

1 2 7 ^ -" 10 15

. - " GEOSAT WEIGHT (THOUSANDS OF Ibs)

-100 L- Figure 29, Lifetime Extension Options (10-15 Yrs)

Relative Annual Annuities (A)

Using the constant annual annuity method of comparison, the following data was

generated, comparing the annuity of these three options to the two cases of

the baseline 10 year non-serviced, non-resupplied GEOsat case and the 15 year

nonserviced, but resupplied GEOaat (these satellites are designed with extra

reliability to last 15 years).

2113e/ - 58 -
LOW ~

CASE

1.5

46.7

330.6

567.8

952.4

MEDIUM

CASE

-4.0

41.6

328.4

567.8

952.4

HIGH

CASE

-14.2

32.3

323.9

567.8

952.4

BASELINE

(10 YEARS)

-10.7

12.6

223.1

407.3

713-5

RESUPPLY

(15 YEARS)

-158

-279

-334

-271

-62

WEIGHT

1,000

2,000

5,000

7,000

10,000

What this analysis indicates is that the penalties of designing the satellite

for GEO servicing is not a major consideration to the user in comparison to

400

'300

v»

O—O 15 YR RESUPPLY & SERVICING •

LOW

MED

HIGH

Jbs.

3000

3000

3000

%

5

10

20

Ibs

5000

5000

5000

1

MED !• WEIGHT

HIGH I PENALTY

OPPORTUNITY

RESUPPLY

2.5

5.0

10.0

VERY SLIGHT

DIFFERENCE

200

PLANNED

RESUPPLY &

SERVICING BASELINE

V)

o

o

100

PLANNED

RESUPPLY

10 15

GEOSAT WEIGHT (THOUSANDS OF Ibs)

-100 LFigure 30. Lifetime Extension Options (10-15 Yrs)

Relative Annual Annuities (B)

2113e/ - 59 -
the penalty of designing the ezta redundancy into the satellite to increase

its life by 50$. Under the assumptions used here, the most penalties are

imposed upon the smaller GEO satellites in order to make them serviceable.

Contingency Resupply - was considered to be the resupply of ACS fuel to a

GEOsat whose fuel has been deplete through a failure, in comparison to the

replacement of this satellite from the ground. Assuming a rapid, unplanned

depletion of fuel at the end of the first year, a GEO-based resupply system

should be be able to rendezvous and resupply the depleted satellite within 90

days with enough fuel to complete its 10 year life. If the satellite was to

be replaced, it was estimated it would take about IS months to checkout,

integrate, and launch a ground spare which would then have a 10 year

operational life. This is shown in Figure 31.

.FAIL AT I

RESUPPLY AT! » 1.25

.FAIL AT t « 1

REPLACE AT t = 2.5

V*

o

C9

GEOSAT CONTINGENCY

RESUPPLY

GEOSAT CONTINGENCY

REPLACEMENT

Figure 31. GSOSAT Contingency

Resupply Versus Replacement

2113e/ - 60 -
Comparing these two projects, the IRR is found to be:

GEOsat

Weight

1,000

2,000

5,000

7,000

10,000

15,000

20,000

Contingency

Reaupply

17.6

31.0

37.6

45.6

56.0

64.2

Contingency

Replace

8.6$

13-4

24.

30.

36.

45.

.7

.2

• 9

.7

52.9

This is plotted graphically in Figure 32.

70 r

60

£ 50

£

oc 40

u.

o

UJ

^

I 30

LU

i 20

10

0

SATELLITE RESUPPLY

(FAILS t - 1, RESUPPLY t = 1.25)

SATELLITE REPLACEMENT

(FAILS t -1, REPLACE t= 2.5)

5 10 15

GEOSAT WEIGHT (THOUSANDS OF Ibs)

Figure 32. Internal Hate of Return GEOSAT Contingency

Resupply Versus Replacement

20

2113e/ - 61 -
Transponder Replacement

Transponder replacement refers to the potential of replacing some portion of

the on-board ACS fuel with extra transponders. In this analysis, 6Q% of the

ACS fuel (equivalent to 6 years life) was replaced with transponders.

Resupply of the ACS took place at t » 3 and t » 6 years. The cost of the

satellite was increased to reflect the increased number of transponders. If

that 25 transponders are added to a 40 transponder satellite (and realizing

85$ of a GEOsat typically arises from electronics costs), the GEOsat cost

would increase to

Sat coat - (orig. sat cost) (1 * .85(25/40))

Insurance coverage was increased to reflect this riskier operation mode.

Besides reflecting TFU and transportation price, it must reflect the

possibility that the revenues from years 4-10 may not be gained if the

resupply operation fails. An insurance rate of 1056 (reflecting a 90$ overall

probability of resupply success) was used. Each time a resupply operation was

performed, this insurance factor was applied to recognize the risk to future

revenues. Thus for the first resupply, insurance will equal:

Insur (t - 3) - .

where R± is the expected revenue in year i.

In comparison to the baseline (no service, no resupply) case, the following

IRE'a were obtained:

GEOsat Transponder

Vt (ibs) Baseline Replacement

1,000 16.7? 15.9?

2,000 24.8 23.6

5,000 52.0 45.8

7,000 67.5 58.2

10,000 118.1 97.4

20,000 144.6 117.8

These are plotted graphically in Figure 34 (shown schematically in Figure 33).

RESULTS

Lifetime Extension

Figure 29 indicates the results of the analysis for the various lifetime

extension options listed above. The comparison is done using the constant

annuity method of comparison, due to the 15 year satellite life with lifetime

extension, in comparison to the 10 year baseline satellite life. This

indicates that:

t.

o planned resupply does not offer increassed benefits to the user over

the baseline case. This is due to the required extra cost of the

GEOsat required to ensure it will survive the increased operational

life.

o opportunity resupply offers a net benefit over the baseline case, but

must be tempered by the realization that no assurance of the GEOsat

surviving the increased lifetime can be given in this case.

2113e/ - 62 -
INCREASED REVENUES FROM

ADDITIONAL TRANSPONDERS

SAT

SAT

TRANS

INSUR

RESUP

INSUR

RESUP

INSUR

10-

Pigure 33« Transponder Replacement

Through Resupply

a combination of GEO resupply and servicing offers the greatest

benefit to the user over that of the baseline case. GEO resupply

continues the revenue stream to the user by adding ACS propellant to

increase satellite operational lifetime. GEO servicing capability

removes the accrued technological obsolescence of the payload, thus

increasing the total revenues, and avoids the requirement for an

increased satellite cost for the extra lifetime reliability.

2113e/
160

140 h

~ 120

*

1 100

UJ

oe •

£ 80

ui

5 60

«<

| 40

20

• 60% ACS REPLACED

• RESUPPLY AT t - 3, t * 6 YRS

• INCREASED INSURANCE FOR

RISKIER OPERATIONS

BASELINE

(NO RESUPPLY)V

TRANSPONDER

REPLACEMENT

5 10 15

GEOSAT WEIGHT (THOUSANDS OF Ibs)

Figure 34. GEOSAT Transponder Replacement

Contingency Resupply

Figure 32 indicates the IRR comparison of the contingency GEOsat resupply case

in comparison to the baseline replacement option. This implies:

o GEO resupply without servicing offers a benefit to GEOsat operators

over satellite replacement. This result may only be applicable in a

few cases where ACS fuel has been depleted while the status of the

communications payload remains unchanged.

Transponder Replacement

Figure 34 indicates the IRR comparison of the caae where additional

transponders are added to the GEOsat communications payload as a replacement

for unneeded ACS fuel assuming a GEO resupply capability exists. This result

indicates that:

20

2113e/ - 64 -
o transponder addition to GEOsats in lieu of ACS fuel does not offer a

net benefit over that of the baseline case. The increased revenues

from the increased transponders ~is" not sufficient to overcome "the

increased risks of this operation, which subordinates future profits

to the success of a resupply operation.

It appears possible that the demonstration of a very highly reliable GEO

resupply system could reverse this conclusion in some cases by reducing the

risk to future profits from the GEO resupply.

Caveats

The assumption used that transponder usage price remain constant with time was

chosen to simplify the analysis and is probably not reflective of actual

future price performance. The actual price of a transponder is highly

dependent upon a variety of factors - most important being supply and demand.

Rather than digressing into a projection of GEO transponder supply versus

demand, the results of this assumption shold be approximately true in most

cases; particularly in the ranking of alternatives. What is important is the

discounting of obsolete transponder value over new transponder value which

models real world behavior. Discussions in-house and with communications

satellite manufactures raised several questions over the assumption used in

defining baseline numbers (e.g., cost/lb to GEO, insurance rates, transponder

growth rates, etc). The major focus of questions, as expected, was the factor

which produced the 12 percent per year decline in saellite revenue

generation. There was general agreement that the revenue does decline with

time, but there were differing opinions as to why the decline occurred and

what its rate might actually be. Major decline factors suggested were:

o Technological obsolescence, as originally suggested.

o Over-supply of transponders leading to price cuts.

o Learning curve effects which allow the deployment of more

cost-effective satellites which in turn allow price reductions, i.e.,

the economic obsolescence of older systems.

o Failure of satellite componenets resulting in a reduction of revenue

earned.

o Changes in the competitive position of the communications ground

segments.

.Revenue decline can occur as a result of both lower prices and loss of revenue

generating components. These declines are also not all caused by the space

segment of the communications systems. Several observers pointed out the

importance of segmenting revenue declines due to the space segment from

changes occurring in the ground segment. What we really need to know is how

much of the revenue decline is due to the space segment and how much of that

consists of items we can actually do something about as part of satellite

servicing.

The cost of making a satellite capable of accepting resupply or servicing in

GEO has not been included in the satellite costs used here. Mostly, these

costs are expected to add to the non-communications and non-electronics costs

of the satellite where their impact is less .than that of the electronics

reliability costs or the changes in the revenue streams. New, larger

satellite designs incorporating resupply and servicing options are believed

to have a proportionally lower cost than retrofitting these options onto

exisitng smaller satellite designs optimized for current operations.

2113e/ - 65 -
CONCLUSION

The major benefit to GEO satellite operators seems to be the incorporation of

both GEO servicing and resupply options in their satellite designs. This

offers them:

o a longer life, avoiding recurring satellite procurement and

transportation costs,

o a means to return their offerring to the current market position by

replacing obsolete electronics for current technology electronics.

o the possibility of avoiding contingency failures by allowing GEO

repair in place of replacement of the entire satellite.

Replacing ACS tankage and assuming GEO resupply in order to add additional

transponders does not appear to offer a benefit over current operations, since

the increased revenues are accompanied by an increase in risk of operations.

A very highly reliable GEO resupply system could possibly reverse this

conclusion in some cases. This subject and further sensitivity analysis are

discussed in the following sections.

Evaluation of Optimum GEO Servicing/Resupply Invervals

While the previous section concluded that servicing in conjunction with

resupply was a desirable mode of operation, a key factor was left undefined

That is, what is the optimum servicing interval as defined by economic

requirements?

Methodology

Based on the spreadsheet analysis programs developed earlier,

resupply/servicing intervals were varied to examine their affect on the

economic value of the satcoms. Annual annuities were calculated for

commercial satcoms in the range of 1 to 10 klb. BOL as a function of servicing

intervals of 2 to 10 years. (The final case being equivalent to no servicing).

Two general cases were examined. The first assumed GEO resupply/servicing of

a satcom with a standard transponder mix as appropriate to its BOL mass. The

second case assumed a tradeoff of ACS fuel for additional transponders. That

is, only enough ACS fuel was alloted to last to the next resupply time. The

weight savings were then assumed to be used for additional revenue-generating

transponders. In both cases, the result of performing a servicing mission was

to restore the satcom"s revenue to its year 1 value (and thereafter experience

the 12% annual decline till the next service). The cost of performing the

mission to GEO was assumed to consist of three general elements:

o Transportation

o Use of the OMV/RM

o Cost of resupplied fuel and "black boxes"

The cost of resupplied fuel was assumed negligible and the cost of the

replaced "black boxes" (technological upgrading of the communications payload

and replacement of life-critical components) was included as a delta cost of

production. Transportation and OMV/RM usage fees were combined as part of a

net cost of getting to GEO; approximately $21,000/lb. in the cases examined.

2113e/ - 66 -
Crucial assumptions are related to the reliability of GEO operations and the

role of insurance. The reliability of satcom components was assumed to be the

same as those used in typical 10 year life satellites. Some decrease in

reliability may be allowable if routine visits are possible, but the cost

impact of this factor is uncertain and thus a more conservative assumption was

made. If a GEO mission fails, that failure may keep the satcom from earning

revenue until the next flight can be performed or, if severe enough, the

failure can result in the total loss of the satcom. While the cost of the

satcom itself and its transportation are insured against at launch, the

performance of GEO resupply/servicing will likely require insurance of the

satcom's projected revenue stream. Since the revenue stream, except for near

the end of the satellite's operational life, is much greater than the original

cost of the satcom and its transport to GEO, assumptions on the appropriate

insurance rate contribute directly to the cost of the mission. We assumed a

10% risk premium in our calculation and this was felt to be a conservative

assumption. Actual operational experience may eventually allow the decrease

of this rate to 5% or even 2%.

Results

Figure 35 is a plot of the expected annual annuities as a function of the

servicing interval for a range of satcom BOL masses. As expected, with

greater weight and more transponders, the economic value of the satellite

increases. These satcoms have a standard transponder mix and contain generous

safety margins of ACS fuel. Peak annual annuities for all satcom masses were

found at 6 years. Servicing every 5 was so close in value as to be easily

within the uncertainty range of the CERs we have used. As expected, servicing

more infrequently did not carry a large penalty. Compared to baseline 10 year

life cases for each satcom weight, servicing had positive advantages for all

intervals of 3 years and up.

Figure 36 shows a similar plot for the case of maximizing onboard transponders

through the off-loading of ACS fuel. This dramatically increases

revenue-generating capability, but at the risk of no safety margin for the ACS

fuel. The solid lines relate to the annual annuities scale to the left while

the dashed lines (for the larger satcoms) relate to the scale at the right.

This combination was necessary to avoid compressing the curves of the smaller

satcoms. It is interesting to note that an optimum servicing interval exists

in these cases as well, but that it occurs earlier at 4 years. The peak is

also more sharply defined, as would be expected from satcoms with greater

numbers of transponders to upgrade. Compared to baseline cases, this

transponder substitution is superior at all servicing intervals greater than

two years. A caution should be raised here, however, Adding a greater supply

of transponder to a specific satellite does not mean they can be sold for the

same price as the baseline situation. The number of satellite transponders to

be emplaced is less a technial capabilities question and more a function of

market forces. No attempt has been made in this analysis to account for the

price elasticity of transponder demand (which itself is a variable with

time). It is unlikely that all additional transponders would have the same

revenue-generating capability.

2113e/ - 67 -
$H)

ANNUAL

ANNUITIES

250 r

200

150

100

50

• SATCOM VALUES MAXIMIZED WITH 6 YEAR SERVICE INTERVAL

UK IB BOL

TK LB aa.

SK LB BOL

3K Lfl SOL

a u BX

IK Lfl SOL

456 7

SERVICING INTERVAL (YEARS)

10

Figure 35- GEO Resupply & Servicing

Std. Transponder Mix

Conclusion

Single optimum servicing intervals exist for the case of resupply and

servicing of commercial GEO communications satellites. Consistent acros the

existing range of satcoms (1-10 klbs), satcoms with standard transponder mixes

should be serviced every 5-6 years. In cases where transponders have been

added to use the weight allowance of off-loaded ACS fuel, that optimum

interval is every 4 years. The optimum points denote a peak in the equivalent

annual annuity value of the satcom as a function of varying intervals. The

performance of resupply/servicing enhanced the economic value of the satcom

project in all cases.

2113e/ - 68 -
(84 $M)«

MAXIMIZING

 TRANSPONDER LOAD SHORTENS SERVICING INTERVAL TO 4 YEARS

ANNUAL

ANNUITIES

100 r

50

L8 BO.

ANNUAL

ANNUITIES

400

350

300

250

200

150

100

50

3456 7

SERVICING INTERVAL (YEARS)

10

Figure 36. GEO Resupply and Servicing

Transponder/ACS Tradeoff

Sensitivity Analysis of Optimum GEO Servicing/Resupply Intervals

After the determination that such optimum servicing intervals existed, the

next step was to examine the sensitivity of these results to changes in key

baseline assumptions. Three key assumptions were examined:

o insurance rates for resupply/servicing operations

o the cost per Ib. of performing operations in GEO

o the obsolesence rate for satcoms, reflecting the yearly decline in

their revenue-generating capability.

It should be cautioned that these analyses and those in the preceeding

references were performed on a brefore-tax basis. No attempt was made to

simulate the effects of differing depreciation rates or changes in tax law.

2113e/ - 69 -
Effect of Insurance Sates on Optimum Geo Service Intervals

A significant portj.cn of the cost of performing resupply/sarvicing missions to

GEO satcoms is the requisite insurance. In the event of damage to the

satcoms, ita future revenue stream will be reduced or eliminated and insurance

will likely be needed against this possibility. Usiag the previously

developed spreadsheet analysis programs, the baseline insurance rate of 10$

was varied across the range of 2% to 15$. The resulting changes in the

satcom's annual annuity value and associated optimum servicing interval were

noted and plotted.

Figure 37 shows the optimum servicing intervals as a function of the insurance

rates for the resupply/aervicing operations. Insurance for the launch and

deployment phase, in contrast, was maintained at the baseline 1Q% in all

cases. As the insurance premium drops, performance of a CEO resupply/service

mission decreases and it becomes advantageous to perform such missions more

often. A key result here is that if insurance rates cannot be brought to 15%

or less, GEO resupply/servicing does not carry a strong economic benefit.

Figure 37 also shows how differing insurance rates affect the satcom's annual

annuity value as a function of the satellite's BOL Mass. Lower insurance

rates increase the value of the satcom's annual annuities-with the greatest

benefits occurring to the smaller satcoms. Conversely, the negative impact of

higher rates is greatest for the smaller satcoms.

• LOWER INSURANCE RATES

SHORTEN OPTIMUM

SERVICING INTERVAL

40

•£ LU •«<.

cap 30

20

uu

CJ

0

-10

• POSITIVE IMPACT OF LOW INSURANCE RATES

IS GREATEST FOR SMALLER SATCOMS

is 10

:u

LUO

O.U.

BASELINE

I I I I

6 8 10

OPTIMUM SERVICING

INTERVAL(YEARS)

10% BASELINE RATE

-15%

4567 8

SATCOM BOL MASS (klb)

Figure 37. Insurance Has A Significant Impact On

Geo Resupply/Servicing

2113e/ - 70 -
Effect of the Cost Per Lb. for Seo Operations

The costs of transportation to GEO and the use of the OMV/EM were combined

into a net cost of putting the resupplied fuel and "black boxes" on the

satcom. A variety of scenarios exist for getting fuel and components to a GEO

satcom; rather than attempting to use specific scenarios, parametric cost

estimates were used as target figures. For a financial analysis, how the

resupply and servicing is accomplished is of less importance than the

associated risk (see the case above) and the net cost of getting the actual

ACS fuel and new components on the satcom. In the baseline case, the net

charge was placed at $21,100/lb.

The cost of performing GEO resupply/servicing was varied across the range of

$10,000/lb. to $100,000/lb. The associated change in the optimum servicing

interval is shown in Figure 38. A decrease in cost/lb. makes it optimal to

service more often and increases in cost/lb. cause a lengthening in the

optimal interval. While these conclusions seem rather obvious, it is

important to note how insensitive the optimal servicing point is over a

relatively broad range of cost figures. From about $20k/lb. to $60k/lb., the

optimum interval stays at about 6 years. Only at the extreme figure of

$100k/lb. does the optimal servicing interval stretch to a point where

resupply/servicing is no longer desirable. On the other hand, dramatic

"decreases in cost/lb. do not result in significantly more frequent GEO

operations.

The effect of variences in the cost/lb. assumption on the satcom's annual

annuity value is shown in Figure 38. As expected, cutting the cost/lb.

increases the satcom's value and raising the cost/lb. decreases the satcom's

annual annuity. The effects are a relative insensitive function of satcom BOL

mass with smaller satellites being more affected than larger ones. The

benefit of cutting the baseline cost/lb. figure in half is, for example, less

than 5%. The disbenefit of raising it an order of magnitude is mostly less

than 15$« The original baseline figure of $21.1k/lb. to Geo was felt to be a

conservative and defendable estimate. The key result from this is that the

cost/lb. for GEO resupply/servicing has only a minor effect on the

desirability and timing of these operations.

Effect of Satcom Obsolesence Rate on GEO Resupply/Servicing

A crucial assumption of the spreadsheet analyses we have been performing is

that the constant dollar revenues from the satcoms decreases at a constant

average annual rate. In the baseline case, this has been assumed to be 12$

per year and is based on the historical growth rate of transponder demand.

The major benefit of planned resupply and servicing is the performance of

technological upgrading. The assumption being that such technical upgrading

is not only feasible but that it results in restoring the satcom's

revenue-generating capability to its year 1 value.

The 12$ figure has been questioned as being overly ambitious. The contention

is that such rates of decline are not occuring with current systems and that

the decline rates that do occur are only partially related to the space

segment-with ground services and market forces being the predominant

influence. To test the sensitivity of this assumption, obsolesence rates of

6% to 12$ were used across the range of 1-10 klbs. satellite BOL mass.

2113e/ - 71 -
+5

Ul

CO Q

Z -10

o

g -20

UJ

u

OB

UJ

a.

.30

^

i

 10

IU

a 8

ig 6

eo^"

10K 5/lb

21.1K S/lb

SDK 5/lb

100K 5/lb

ANNUAL ANNUITIES ARE RELATIVELY INSENSITIVE

TO THE COST/lb FOR GEO RESUPPLY/SERVICING

I I I

6 7 8 9

SATCOM BOL MASS (klb)

10

^BASELINE

THE OPTIMUM SERVICING INTERVAL IS INSENSITIVE

FOR A BROAO RANGE OF COST/lb ASSUMPTIONS

I I _L I

10 20 30 40

Figure 38.

50 60. 70 80

K S/lb IN GEO

90 100

Effect of GEO Servicing Cost/Lb.

On Satcom Values

The effect of different obsolesence rates on the optimum GEO

resupply/servicing interval are shown in Figure 39- If the satcom suffers

less obsolesence, there is less incentive to resupply and upgrade it. If a

satcom's revenue-generation decline can be held to a 6% rate or lower, it is

unlikely that resupply/servicing will present significant advantages. At

about a °i% rate, the optimum servicing interval has moved out about a year.

This move however is still well within the uncertainty range of our CER's and

not highly significant.

2113e/ - 72 -
ta

12

10

8

6

4

HIGHER OBSOLESCENCE RATES

ENCOURAGE EARLIER SERVICING

60 -

COCO

BASELINE

I I I

50

40

•30

i

i

20

10

0

6 8 10

OPTIMUM SERVICING INTERVAL (YEARS)

• LOWER OBSOLESCENCE RATES MOST

BENEFIT SMALLER SATCOMS

I i j i

9%

112% BASELINE RATE

8 10

SATCOM BOL MASS (klbs)

Figure 39. Effect of SATCOM Obsolesence Rates on

GEO Resupply/Servicing

Figure 39 also shows the effect of two different obsolesence rates on a

satcom's annual annuity value. Obviously the lower the rate of revenue

decline, the higher the equivalent annual annuity. The significant point

here, however, is the dramatic percentage increase in the annual annuity for

satcoms below about 3klbs. This can be expected from the relatively lower

mass devoted to communications payload in the smaller satellites and the

higher leverage of the obsolesence rates. The key result here is not only

1 that the obsolesence rate significantly affects the economic value of the

satcom, .but that it is surrounded by a large degree of uncertainty. It is not

obvious that technological upgrading will be able to fully restore year 1

revenue-generating capability or that changes on the ground might lessen the

need for such actions.

The financial model we have been using has been shown to produce stable

results across a wide spectrum of potential changes in key baseline figures.

No changes to earlier conclusions need to be made. Several key new

conclusions from the sensitivity analysis can be drawn:

o Insurance rates have a high impact on the economic viability of

resupply/servicing operations for satcoms. Rates of \5% or lower are

required with a consequent impact to the overall reliability

requirements for GEO operations. Based on historical experience, an

overall reliability level of 95% or better is required from launch

through GEO operations.

2113e/ - 73 -
o The cost/lb. figure for GEO resupply/servicing has a relatively minor

influence on the timing of optimum servicing intervals and the

satcom's annual annuity value. Resupply/servicing scenarios only

need to keep their cost/lb. figures below about $50k/lb. to be

economically viable. This includes not only the cost of getting to

GEO, but associated usage fees for required elements (e.g. OMV, OTV,

servicer kit, etc).

o If satcom revenue decline rates average less than 6% over their

lifetime, resupply/servicing is not likely to be attractive. Decline

rates of about 9% and above make technological upgrading potentially

attractive. The key questions here are the technical feasibility of

revenue restoration with upgrading and what combinations of space and

ground segment modification are required.

Sensitivity Analysis of Satellite Industry Concerns

Based on discussions held with THW, Hughes Aircraft and Ford Aerospace,

several additional economic sensitivity analyses were suggested. These

analyses were to further examine the effect of varying assumptions used in the

spreadsheet modeling of communications satellite resupply and servicing.

Three general situations were examined:

1. The first question addressed the assumption that satellite servicing

would restore the satcom's revenue-generating capability to its Year

1 value. What happens if full restoration is not possible by

satellite servicing, either due to some technical impossibility, or

problems with the ground segment? What happens if the revenue rate

is increased? The latter possibly would use coordinated changes in

the ground and space segments.

To examine this case, the level of revenue restoration was varied in

the spreadsheets. Instead of 100$, values of 30$, 50$, 10%, 90$, and

even 110$ and 120$ were used. Each time satellite servicing

occurred, revenue levels were restored to some percentage of their

Year 1 value.

The results are displayed in Figure 40. For lower restoration rates,

the optimum servicing interval lengthened. At restorations above

100$, the servicing interval did not decrease below the baseline case

(about 5 years) while the annual annuity value of the satcom

increased. As expected, annual annuity values decreased for less

than 100$ restoration. These results held across the examined range

of satellite BOL masses (l-lOK Ib). The relation of restoration rate

to the obsolescence rate (the yearly decline in satcom

revenue-generating capability) was that satcoms with higher

obsolescence rates could tolerate less restoration and still find

servicing viable. If your earning power is dropping fast enough, you

can be less picky about how much your revenues need to be enhanced.

2113e/ - 74 -
co

LU _

OB O

u

I

120

100

80

60

40

', REVENUE RESTORATION vs OPTIMUM SERVICE INTERVAL & ANNUITY IMPACTS

• RESTORATION ABOVE 100% DOESN'T CHANGE THE OPTIMUM SERVICING INTERVAL

• HIGH OBSOLESCENCE RATES REQUIRE LOWER RESTORATION LEVELS FOR ECONOMIC

BENEFITS'

LINE A. 12% REVENUE DECLINE RATE

LINE B. 9% REVENUE DECLINE RATE

LINE C. 6% REVENUE DECLINE RATE

SKIb

NO OPTIMUM SERVICE INTERVAL

8 10

OPTIMUM SERVICING INTERVALS (YRS)

BQL MASSES MOKIb

12'/. REVENUE DECLINE RATE

BASELINE RESUPPLY &

SERVICING POINT

.24 -

SERVICING DESIRABLE

(OPTIMUM INTERVALS EXIST)

20 40 60 80 100

% RESTORATION

120

Figure 40. Impact of Varying Levels of Revenue

Restoration from Servicing

Conclusions to be drawn from this case are:

o Greater than 100$ restoration is beneficial, but doesn't change the

servicing interval.

o Revenue restoration rates required for economic viability are not so

much a function of the satellite mass as of the obsolescence rate. A

baseline revenue decline of 12% per year requires a minimum of 70$

revenue restoration from servicing. A 6% decline requires a greater

than 100$ restoration level. The less obsolescence suffered, the

greater the required benefit from servicing.

2118e/ - 75 -
2. The second question addressed the assumption of $200/hr as the price

for a transponder. Industry representatives pointed out that

transponder prices vary greatly depending on how they are leased,

either long-term or on the spot market. If changed, how would this

assumption affect the selection of servicing intervals?

The spreadsheets were modified to use transponder prices of $100 to

$600/hr, reflecting current possible pricing extremes. The results

were that the optimum servicing interval varied very little.

Servicing intervals were every five years for up to $600/hr and every

six years for as lov as $100/hr. This is essentially the same

result, within the uncertainty of the data, as the baseline cases.

Also, as expected, satellie annual annuities rose with increases in

the transponder price; the greatest percent benefits being for the

smaller satcoms.

While the tansponder price greatly affects the annual annuity value

of satcoms, see Figure 4-1, it does not significantly affect the

timing of servicing operations. The earning power of the satcoms

relative to the current market, as reflected in the obsolescence

factor, is the more important issue to watch.

UJ

co

10

8

6

4

2

+ 200

S ul -i-100

^ e/j

UJ «£

a co

UJ

OL

100 200 300 400 500 600

S/Hfl

•100

$300/hr

$150/hr

S100/hr

i i i

2 4 6 3 1 0

SATELLITE MASS (K!b BOL)

Figure 41. Impact of Changes in Transponder Prices

2118e/ - 76 -
. ._3«- The third_ question. addressed the_ assumption of the cos_t_impac_t of _ __

producing longer life satcoms for the case of planned resupply. In

the planned resupply case, 10-year satcoms were modified to have a

15-year life at the same level of reliability. At the 10-year point,

a resupply operation occurs, providing ACS fuel for the remaining

five years of life. No other servicing occurs at this time. This

case contrasts with the opportunity resupply case in which a 10-year

satcom was refueled for an additional five years, but no costs were

incurred upfront to insure it would last the full 15 years.

Essentially, since it was working at 10 years, it was assumed to

continue to work if given sufficient ACS fuel.

Since the planned resupply case was not found to be economically

viable, the question was raised as to the relative importance of the

extra upfront costs incurred for the longer lifetime. The planned

resupply case made some conservative assumptions about the cost

penalty of extra life reliability, thus making it and opportunity

resupply (no cost penalty) two extreme cases. To examine the

significance of the penalty, it was reduced in percentages from the

planned resupply case (100$) to that of opportunity resupply (0?) and

the impacts to the annual annuity levels were noted.

The results of this case are shown in Figure 42. The baseline case

of a standard comsat, 10-year life with no resupply or servicing, is

superior in almost all cases to that of planned resupply. This holds

until the cost penalty for extra life reliability is reduced to 20%

of its original value. Thus, very optimistic assumptions on the cost

penalty for extra satcom life are necessary to make planned resupply

alone econmically attractive. This conclusion further emphasizes the

importance of performing servicing in conjunction with resupply,

particularly in the case of GEO operations.

The three cases examined here do not change any of our original conclusions on

the viability of GEO resupply and servicing. They do, however, provide

further data and insight into the particular financial model we have been

using. The results emphasize the importance of minimizing the cost impact of

making satcoms serviceable and of the need to restore/enhance their

revenue-generating capabilities for GEO operations to be desirable.

Potentially Resupply Uses in GEO Operations

The economical remote resupply of satellites and platforms in GEO depends on

the capability for technology upgrading and maintenance in GEO. An

economically desirable way of accomplishing this uses the resupply module as a

propellant tanker for the OMV. The primary alternative to use of the OMV

would be a space-based reusable OTV, if available.

In addition to the remote resupply of assets in GEO, the resupply module can

support some GEO operations while employed in LEO. A preliminary screening of

possible concepts included:

o Use the OMV/RM combination as a reusable perigee kick stage prior to

the development of a fully reusable OTV. See Figure 43 •

o Develop a LEO propellant depot to support the space-based operation

of the OMV/RM. Such operations would include propellant transfer to

elements heading for GEO. Supplies of depot propellant may be

economically available by scavenging orbiter OMS/RCS fluids. The

OMV/RM could also be used for such scavenging. See Figure 44.

2118e/ - 77 -
Figure 42. Lifetime Extension Options (10-15 Yrs)

2118e/

(w$) saumNNV ivnNNv INVISNOO

- 78 -
Figure_43. Equivalent GEO P/L Capability With

OMV/HM Reusable Perigee Stage

ew

X

§e i sn s e

.-

_i . in

si dT

(S«H)3WlldlUlAW01V101

2118e/

(Bit 030 N! IM Vd

- 79 -
REUSABLE STORABLE PERIGEE STAGE

RETURNS TO GENERAL VICINITY OF LEO

DEPOT AFTER PL DEPLOY TO GTO &

STATION KEEPS

ORBITER BRINGS PL TO OTV & USES RMS

TO AFFECT DOCKING. MAN-TENDED

CHECKOUT AVAILABLE AS REQUIRED

PU OTV

OHBITER RELEASES PL/OTV & MOVES

AWAY. OMV/RM TRANSFERS

PROPELLANT & PRESSURANT FROM DEPOT

TO OTV (AWAY FROM STS FOR RELIEF

FROM REDUNDANCY REQUIREMENTS FOR

MANNED SAFETY)

OTV/PU DEPLOYS TO GTO

HANGER DEPOT

OMV/RM

RENDEZVOUS

WITH ORBITER

FOR PROPELLANT

SCAVENGING

r OMV/RM

TRANSFERS

PROPELLANT

TO DEPOT &

FOR LONG

DURATION

BULK CONDITIONING

& ORBITER

RETURNS TO

GROUND

SEE (V

Figure 44. Representative Resupply Module/OMV Operations in

Conjunction with LEO Depot and Reusable Perigee Stage

o Use the OMV/RM to top-off satellites that have reached LEO, but are

still to be placed on station at higher altitudes (e.g., 12 hr

orbits, GEO) such top-off operations could ease weight limits on an

individual shuttle flight (i.e., the orbiter capacity is used to lift

a greater amount of dry mass). This concept is related to earlier

proposals to dry launch satellites and stages and consequently to

reoptimize their design. Such reoptimization would presumably result

in more efficient launch installations.

These concepts show that the resupply module can be used for more than the

support of specific spacecraft and platforms. In its support of GEO

operations, the RM becomes an important part of the developing space

infrastructure. See Figure 45« LEO remote resupply can serve to increase the

efficiency of space operations by introducing users to space-based operational

modes prior to the IOC dates of space station and the reusable OTV. The

OMV/RM combination is not only an effective one for GEO remote resupply and

servicing, but it can support the economic deployment of those assets to GEO

itself. Further examination of these possibilities in cooperation with

potential users is desirable.

2118e/ - 80 -
ORIGINAL PAGE IS

OF POOR QUALITY

SPACECRAFT MATED

TO PROPULSION MODULE

IN LEO USING ORSITER

TENDING STORABLE PHOPELLANT DEPOT

IN LEO FILLED FROM ORBITEH

QMS POO VUk OMV/RM t.4M LB

AVAILABLE IN 1990s AT S11T1B

OMV.RM FOR

TRANSFER TO UPPER

STAGE IN LEO

^••— f «^^«&> K'? *?§V

; OBBITER BRINGS SPACECRAFT TO

''LEO WITHOUT PROPULSION MODULE.

QMS PAOPELLAKT SCAVENGED FROM

ORBITER i TRANSFERRED TO

DEPOT IT OMV/RM

Figure 45 • LEO Remote Refueling of Storable Stage Candidate Scenario

3.1.4 Design Reference Mission (DfllQ/System Requirements Selections

Approximately 20 LEO fluids resupply scenarios and 4 GEO fluids resupply

scenarios were initially developed, and mission profiles defined for each in

Subtask 1.2. For ROM cost, risk and complexity estimating purposes, the

scenario list was "scoped" to remote scenarios versus "Orbiter-based"

scenarios.

The six remote resupply scenarios, two in-bay scenarios and two

return-to-earth" scenarios indicated in Figure 46 were used to compile initial

parametric affordability estimates. These were subsequently combined with the

needs (gross economic or operational benefits) assessments for each of the

twelve screened spacecraft programs to develop the semi-final list of

candidate design reference misisons. This section address how cost, risk,

complexity and resupply benefit factors were derived and used in selecting

design reference missions (DRMs) for the remote resupply module.

2118e/ - 81 -
• SCOPED TO REMOTE «

ORBITER-BASED

SPACECRAFT

RETURNS

FOR EVA

ORBITER TO

SPACECRAFT

FOR EVA

D OMV,

7

I) OMV,

2) OMV,

»-v

«) OMV,

I) OMV.

X> OMV.

n>1 — ,

OMV GROUND

- BAJ

n = 1 j

^

n>1

ON

>ED

' (2)

M^^^

k

REN

RES1

V

SPACE

n=1 BASED

Ul

n>1

i

\ -F

F

_^«w ^

^^H

__ i t

DTE

JPPLY

(<)

SPACE

RESUPPUT

r

^^^

IB|

7) IN-BAY IP

I) IN-BAY QMS

IN-BAY

RESUPP

..

Figure 46. Low Earth Orbit Alternative Fluids Resupply Scenarios

Cost Factors

Major "Sough Order of Magnitude" (ROM) cost estimates were generated for each

of 10 LEO and 4 GEO alternate fluids resupply scenarios, with minor

adjustments made to reflect program differences at the candidate design

reference mission level. Average total coats per resupply engagement

(initially assuming 111 engagements) were netted against spacecraft gross

economic benefits to produce net economic benefits for each candidate design

reference mission. Figure 47 shows the relationship between the various

factors contributing to the average total cost of a resupply mission.

Space Transportation System (STS) level transportation costs were estimated

using average cost per flight (not "user fee") relationships for the 1990 -

1998 time frame, with specific estimates for each launch site:

o ETH — S110M cost per flight

o WTH — ?145M cost per flight

^2255/pound *

^6050/pound *

It was assumed that LEO spacecraft designs would be Shuttle-optimized in the

post-1990 period, therefore, "length" and "weight" transportation cost factors

would be identical.

2118e/ - 82 -
Objective — Minimize Resupply Cost to End User

111 ENGAGEMENTS

AVERAGE TOTAL COST

RESUPPUf

L»

RM USER FEE

RESUPPU

MODULE

COSTS

STSLEVEt

COSTS

TECHNOLOGY

ODTSE

PRODUCTION

OPERATIONS

& SUPPORT

J_

USER

ACCOMMODATION

COSTS

INTEGRATION

FLIGHT OPS

USER FEES

OMV

EVA

TRANSPORT

DESIGN

HMD

PRODUCTION

MASS

VOLUME

OPERATIONS

ETR

WTR

A TRANSPORT COST

A INSURANCE

GROUND STATION

Figure 47- Relationship Between Considerations for Minimizing

Resupply Cost to the Potential User

User "accommodation" costs (to benefit from resupply without concurrent

maintenance) are known to be significant," but were not specifically

estimated. It is quite unlikely that a program sponsor would deliberately

increase the spacecraft design life (at very large marginal cost) to obtain

the .benefits of resupply alone, without accommodating for full-scale on-orbit

servicing.

It was found that transportation and insertion costs dominate all other cost

elements for the LEO alternate fluids resupply scenarios by a substantial

margin. For the cost-minimization goal, those scenarios which rely on

transporting the OMV or spacecraft itself to/from low earth orbit are very

heavily penalized—particularly at WTR. Table 9 shows the LEO resupply

operations cost drivers identified by this analysis.

2118e/ - 83 -
Expendables Resupply Life-Cycle Cost Drivers

(| million 1984)

Carrier (OMV) User Pee

$1.0-$1.5 / Engagement

Hesupply User fee

$0.5-?1.0 / Engagement

STS-Level Transport

o OMV inert 3200#-4000#

o EEM inert 800#-1600#

o Integration Charge

o Launch & Plight Ops

o STS-Level Cost

OMV Acquisition / # OMV Engagements

1300 (4 Units) 200-300

EBM Acquisition / # EBM Engagements

$55-$110 110

ETR

$7.2

1.8

1.0

1.2

11.2

$2250/lb. WTR

$9.0

3.6

1.0

1.4

15.0

$19.4

4-8

1.0

1.2

26.4

$6050/lb.

$24.2

9-7

1.0

1.4

36.3

* STS Cost Per Plight

* Capacity * 0.75

* STS Cost/Pound

$110M

48,750 Ibs.

$2250

$145M

24,000 Ibs.

$6050

Table 9. LEO Resupply Cost Drivers

Rough Order of Magnitude (ROM) acquisition costs for the Resupply Module (RM)

were estimated using parametric Cost-Estimating-Relationships (CERs) from the

USAJ? Space Division's "Unmanned Spacecraft Cost Model" for a range of RM inert

weights between 800 pounds and 1600 pounds. Pending design definition and

trades in Task 2 of the Expendables Resupply study, no attempt was made to

estimate software development at this time. Table 10 presents the results of

this analysis.

Figure 48 presents the estimated total acquisition cost as a function of

number of units acquired which was projected using the preliminary work

breakdown structure also depicted in this figure. Further definition of RM

development and acquisition costs were performed as part of the Programmatics

effort, as found in Section 3.3«

An important "select the best DRM" affordability decision goal is to minimize

the total cost of resupply to the user (receiver spacecraft) program. Except

in those very few cases in which the satellite itself has an extremely low

mass and/or low volume, STS transportation costs preclude either of the

"return-to-earth" fluids resupply scenarios, unless concurrent full-servicing

is also accomplished.

At ETR (due East launch), all of the remaining eight LEO on-orbit fluids

resupply scenarios are sufficiently cost-effective to generate positive net

economic benefits for each candidate spacecraft in the final list. Table 11

presents representative cost data for these scenarios.

2118e/ - 84 -
Expendables

Structure, Thermal

Tele, Track, Comm

Elec Power & Dist

Program - Level

min

(Ibs)

700

50

50

Resupply Module

($ million,

max

(Ibs)

1300

150

150

ROM Acquisition

1984)

Cost

mi nitmiTn

12.8

3.7

0.7

8.0

2.

1.

1.

2.

2

8

0

3

maximum

17.

9.

2.

13.

6

0

2

4

3

4

2

4

.2

.8

.1

.7

DDT&E (incl TFU) 800 1600 32.5 57.1

ROM Total Acquisition Cost

($ million, 1984)

Quantity:

Total Acquisition $:

Average Acquisition $:

Total Acquisition $:

Average Acquisition $:

1

32.5

57.1

2 '

39-1

19.5

70.5

35.3

4

51.6

12.9

95-9

24.0

8

75.4

9-4

144.3

18.0

Table 10. ROM Acquisition and Total Costs for

Expendables Resupply Module

.1

.1

.3

.3.1

.3.2

.3.3

.3.4

.3.7

.3.8

.3.9

.3.10

.4

.4.1

.4.2

.4.3

.4.4

.4.7

.4.8

.4.9

.4.10

.4.11

.S

TOTAL ACQUISITION (SM 1984)

125 r

WBS

PROJECT MANAGEMENT

SYS ENG & INTEG

DESIGN & DEVELOP (NR)

STRUCTURES & MECHANISMS

ELEC PWR & DIST

THERMAL CONTROL

COMM & DATA M6MT

SOFTWARE

6RO SUP EQUIP

AIR SUP EQUIP

GROUND CONTROL

PRODUCTION (TFU)

STRUCTURES & MECHANISMS

ELEC RNR & DIST

THERMAL CONTROL

COMM & DATA MGMT

SOFTWARE

GRO SUP EQUIP

AIR SUP EQUIP

GROUND CONTROL

VEHICLE ASSEMBLY

SYS VERIFICATION & TEST '

ALLOCATED

ALLOCATED

650 LB - 1250 LB

100 -

50 LB. 200W - 150 LB, 500W

50 LB - 100 LB

50 LB - 150 LB

NOT INCLUDED

NOT INCLUDED

ALLOCATED

ACOUISIDON COST/ENGAGEMENT

- S1H 1984

ACQUISITION COST/ENGAGEMENT

3 4 5

» UNITS ACQUIRED

Figure 48. Expendables Resupply Module ROM Acquisition Cost

2118e/ - 85 -
Secnario (See Pig.

OMV User Pee

ERM User Fee

OMV (Wi) Transport

EHM (Wi) Transport

OMV (Wp) Transport

STS Integration

STS Flight Ops

OMV Refuel Fee

In-Bay EVA Fee

EBM ASE Transport

Cost to ETR User

NOTES: 1. OMV

Cost

47)

1

4

1

2

1

11

¥p 1250

2 . Space-based

3. ERM ASE 200

(Per

1

.2

.8

.0

.3

.8

.5

.0

7?

Resupply J to User —

($ million, 1984

2

1.2

.8

8.1

2.7

2.8

1.0

1.3

.5

17.9

1

1

2

1

"6

Ib/engagement

OMV

Ib.

($400M, 4

3

.5

.0

.4

.1

.0

.5

.8

.0

7s

•

9

4

1.5

1.0

.4

.1

2.8

1.0

1.4

.5

~9~T2~

17751b/2

units )/300

ETR

1

1

2

1

~6

5

.5

.0

.4

.1

.0

.8

.0

~

1

1

2

1

1

~8

6 7

.5

.0

.4

.1 2.7

.8

1.0

.4 .4

.0 1.0

.4

T2 5T~

8

2.7

1.0

.5

.4

5-6

engagements

engagements

Table 11. Relative Cost Effectiveness of Selected LEO Resupply Scenarios

At ¥TR (sun-synchronous), only the "in-bay" fluids resupply scenarios are

cost-effective across the entire polar candidate spacecraft list, and the

"in-bay" scenarios may require the spacecraft to perform its own orbit

transfer and/or planar/nodal maneuvering.

Figure 49 presents the relative cost effectiveness of all ten LEO resupply

scenarios shown in Figure 47.

Risk and Complexity Factors

Avoidance of (and/or manageability of) risk is an important affordability

goal, and each of the above 10 LEO fluids resupply scenarios was subjected to

a comprehensive risk + complexity analysis using the Analytical Hierarchy

Process technique accepted at the methodology review at NASA/MSFC early in the

study.

Major risk + complexity issues were identified and then prioritized through

several "modified Delphi" working sessions during the latter phases of Task

1. Three issues, which remain unresolved, dominated the combined risk +

complexity analysis:

l) CREW SAFETY, particularly the potential exposure of astronauts and

their life-support systems to toxins as a result of human involvement

in propellant transfers, either in the Orbiter payload bay or at the

Space Station.

2118e/ - 86 -
COST/HESUPPLY

$ffl~84

COST/RESUPPLY

12345678 9 1 0

SCENARIO

WTR 98°

1.0 < 10

1 2345678 9 1 0

SCENARIO

Figure 49. Relative Cost Effectiveness Estimates for Selected

LEO Resupply Scenarios

2) CONFLAGRATION POTENTIAL, particularly the consequences of inadvertant

combinations of bi-propellant reactants in the vicinity of either the

Orbiter vehicle or the Space Station.

5) OPERATIONAL COMPLEXITY, particularly the number of orbital inclinations

and ascending node crossing times involved in the multiple servicing

fluids resupply scenarios.

Final resupply scenario risk evaluation results obtained using the Analytical

Hierarchy Process are shown on Table 12.

2118e/ - 87 -
g ground baaed; 3 = apace baaed;

Crew

n = no. of engagements; as - apace station

Safety

0.61

Explosion

Potential

0.26

Rendez

/Dock

0.07

Contamination

0.06 RISK

OMV g n=2

OMV g n=l

OMV 3 n=2

OMV a n-1

OMV SS n=2

OMV SS n=l

IN-BAY IP

IN-BAY OMS

R-T-E IP

R-T-E OMS

.07

.08

.07

.06

.04

.04

.01

.01

.10

.13

.02

.03

.02

.03

.01

.02

.01

.01

.04

.07

.00

.01

.01

.01

.01

.01

.00

.00

.00

.00

.00

.01

.00

.00

.00

.01

.01

.01

.01

.01

0.10

0.12

0.10

0.10

0.07

0.07

0.04

0.04

0.16

0.21

moderate

moderate

moderate

moderate

mod-high

mod-high

high

high

low

very low

Table 12. Results of Risk Evaluation for Selected

LEO Resupply Scenarios Using AHP

Considerations for risk include the factors shown in Figure 50.

RISK

MATERIALS

PROCESSES

TESTS

METHODS

PROTOTYPES

TECHNOLOGY

RISK

cnmcALirr

TECHNOLOGY-'PROJECT

DDT4E

NETWORK

BST

NEEDS

•CONCEPT

DEVELOPMENT

•FSED

. DETAIL DESIGN

. DUAL TEST

•FLIGHT TEST

RISK

IMPACTS

ALTERNATE

TECHNOLOGY

- TECHNOLOGY

BASE

- APPLICATION

- SCALE-UP

- LOAD TIME

RISK

REDUCTION

- STATUS

- NETWORK

- R&T NEEDS

- RESOURCES

MISSION

PERFORMANCE

•CONCURRENCY

•PARALLEL

•EXPERIENCE

•PAST PERFORM.

PROJECT

' COST

PROJECT

SCHEDULE

-SAFETY

-OPERATION

-SYSTEM

•PAYLOAO

I- ODT&E £2

[- PROD £2

LO&S A2

j-OOTAE <

UlOC

L-FOC

MAJOR RISK ISSUES

• CREW SAFETY

• EXPLOSION POTENTIAL

• PAYLOAD CONTAMINATION

• RENDEZVOUS/DOCKING

COMPLEXITY FACTORS

•HARDWARE

•SORWARE

•CONTROL

•OPERATIONS

Figure 50. Risk Considerations for Scenario Evaluation

2118e/ - 88 -
The combined risk plus complexity preference rankings for the 10 LEO alternate

fluid resupply scenarios were developed by AHP specification of risk criteria

importance and subsequent AHP evaluation of the contribution of each scenario

to each risk goal. Mission complexity was incorporated as a candidate design

reference mission level parameter, given the specifics of receiver spacecraft

and fluids resupply scenario.

Combined risk plus complexity evaluation results using the Analytic Hierarchy

Process as shown in Table 15«

OMV g n=2

OMV g n-1

OMV s n=2

OMV s n-1

OMV SS n=2

OMV SS n-1

in-bay IP

in-bay QMS

R-T-E IP

S-T-E QMS

Crew

Safety

0.06

0.07

0.07

0.08

0.01

0.01

0.01

0.01

Explosion On-Orbit

Potential Operations

0.07

0.07

0.07

0.07

0.01

0.01

0.01

0.02

eliminated— not

eliminated—not

0.01

0.04

0.01

0.02

0.01

0.02

0.09

0.13

cost-effective

cost-effective

RISK+CMPLX

OVERALL

0.14

0.19

0.15

0.17

0.03

0.05

0.11

0.16

RANK

7

3

6

4

10

9

8

5

2

1

Table 13. AHP Evaluation of Combined Risk and Complexity Results

Significant risk avoidance preferences were indicated for the two

"re'turn-to-earth" resupply modes, which were clearly not cost effective for

the majority of candidate spacecraft, and for the ground-based OMV single

engagement scenario.

The risk evaluation criteria used for the pair-wise comparisons (which

characterize the Analytic Hierarchy Process) are shown in Figure 51 along with

their respective rankings as determined from the indicated results.

Considering the combination of risk and return enabled elimination of the

poorest set of candidate design reference missions and retention of the

most-efficient (highest return for any given level of risk, or lowest risk for

any given level of return) set of candidate DRMs.

Combined estimates of return (net economic benefit) and risk for five NASA

astronomic observatory and explorer programs in low inclination orbits were

arrayed as shown in Figure 52 and "dominated" candidate DRMs were eliminated.

For example, if the Gamma Ray Observatory program had been selected for a

resupply design reference mission, either fluids resupply scenario #2

(ground-based OMV, single engagement) or scenario #4 (space-based OMV, single

engagement) or scenario #8 (in-bay, using Orbiter maneuvering) would have been

preferred on combined return/risk estimates.

2118e/ - 89 -
RANK ORDER RANK ORDER

NO. SCENARIO

1

2

3

4

5

6

7

.8

9

10

OMVg

OMVg

OMVs

OMVs

OMVss

OMVss

IN BAY

IN BAY

RTE

RTE

RISK/RANK

n>1

n = 1

n>1

n = 1

n>1

n « 1

IP

OMS

IP

OMS

4

3

5

6

7

a

10

9

2

1

5

3

6

4

8

7

10

9

2

1

6

5

4

3

2

1

8

10

7

9

8

5

9

6

10

7

2

1

4

3

MOD/6

MOD/3

MOD/4

MOO/4

MOD-HI/7

MOO-HI/7

HIGH/10

HIGH/9

LOW/2

LOW/1

COMPLEXITY

6

6

6

6

9

9

3

3

1

1

9

6

9

6

9

6

4

2

2

1

7

7

7

7

7

7

3

3

3

3

8

5

9

6

9

6

4

3

3

1

MOD-HI/8

MOD/5

HIGH/ 10

MOO/6

HIGH/10

MOO/6

LOW-MOD/4

LOW/3

LOW/ 2

VERY LOW/1

RISK

COMPLEXITY

MOD-HI/7

LOW/3

MOO-HI/6

LOW-MOD/4

VERY HIGH/10

HIGH/9

HIGH/8

MOD/5

LOW/2

VERY-LOW/1

•6 RESUPPLY SCENARIOS EXHIBIT "MANAGEABLE" RISK

• MULTIPLE ENGAGEMENTS/ MISSION DIFFICULT

Figure 51. AHP Ranking of Risk Evaluation Criteria for Selected

LEO Resupply Scenarios

Overall, the Advanced X-Ray Astronomic Facility program exhibited the lowest

risk range for on-orbit resupply, since a Xenon/Methane transfer diminished

the Crew Safety and Explosion Potential concerns. Conversely, the Large

Deployable Reflector program exhibited the highest risk range, reflecting the

unknowns involved with superfluid helium transfer technology.

An efficient set (return versus risk avoidance) of NASA low inclination orbit

candidate design reference missions was developed by eliminating both the low

payoff spacecraft and the high risk scenarios from further consideration as

shown in Figure 53-

2118e/ - 90 -
Efficient Return/Risk f) Complexity Eliminates

Worst Selections

250

SM NET ECONOMIC BENEFIT

100

50

8 5

SM & *C

$M NET ECONOMIC

BENEFIT

*r i 4* A 8 HD

9

R 136t*7

 * 200.

04 e-6

"100 « 1 DOMINATED

AXAF DETAIL

02 *

-90 , i i 80

L-M | M

L

— Q

M-H

0 0 £©

 AXAF

 X. °

• . GRO f •

• N2H4 § —^70

;

•8 J7 ,

-4* * 6

GRO DETAIL

2

} L-M M M-H

l i 1 j

DOMINATED

l

RISK n COMPLEXITY

LOW LOW-MED MODERATE MOD-HIGH HIGH EXTREME

DRM-1 BACKUP

Figure 52. Risk/Economic Benefit Array for

Candidate LEO Resupplj Scenarios

The same return/risk methodology was applied to NASA and DoD candidates in

polar inclinations and to NASA, commercial and DoD candidates in higher energy

(12 hour and GEO) orbits to define a refined set of candidate design reference

missions. For example, it was determined that the contingency resupply of

consumables to GEO satellites in the event of their premature depletion (as

opposed to the typical recovery plan of building and launching a replacement

unit) was benefical.

2118e/ - 91 -
N

200

150

100

50

ET ECONOMIC BENEFIT

SM 84/RESUPPLY

Low Earth Orbit, Low Inclination

"EFFICIENT" CANDIDATE DRMs,

NASA

RESUPPLY SCENARIO

INCREASING RETURN

_

DIMINISHING

RISK

MSFC AXAF

PROTEUS^

N2H4

A A 4 A Q a PUVE Y a

SPACECRAFT

AXAF

EUVE

GRD

LDR

PROTEUS

1 2

A/

A/

A/

3 4

v'

y'

V

^

*

5 6 7 8

v'

v^

v'

v'

•

9

v'

•

10

<J

•

RISK (7 COMPLEXITY

LOW MODERATE HIGH

Figure 53- Efficient Set of Candidate LEO, Low Inclination

Design Reference Missions

The refined set of candidate DRMs is shown in Table 14.

AXAF #2

AXAF #4

EUVE #3

GRO #2

GRO #4

LDR #2

LDR #4

PROTEUS #4

DoD #2

EOS #2

IUS ComSat

GEO ComPlat

Net Econ

Benefit

92

101

24

71

80

195

204

28

?

10

60

70

Operation

Benefit

special

purpose

contingency

contingency

contingency

contingency

contingency

contingency

operational

contingency

contingency

design rqmt

Scenario

Risk

low

low-mid

moderate

low-mod

moderate

high

high

moderate

low-mod

low-mod

moderate

low-mod

Service

Orbiter

YES

YES

9

YES

YES

9

9

9

9

9

NO

NO

Potential

Sponsor

MSFC

MSFC

JPL

GSFC

GSFC

ARC

ARC

GSFC

DoD

GSFC

Commercial

MSFC

Table 14. Efficient Set of Remote Fluids Resupply Candidate DRMs

2118e/ - 92 -
This semi-filial list of the most (return/risk) efficient candidate design

reference missions for expendables resupply without concurrent servicing was

comprised of nine individual spacecraft/platform programs and three alternate

fluids resupply scenarios.

Three programs (AXAF, GBO and LDR) were eliminated from selection as the

"best" DEM, since the extremely high marginal value of concurrent servicing to

each of these programs is very likely to strongly motivate the program

sponsors to choose an "in-bay" maintenance plus technology upgrade plus

resupply scenario.

The AXAF spacecraft is typical of many large and complex LEO systems which

will require resupply for an adequate design life. However, other types of

servicing will also be accomplished concurrent with the resupply operations.

(In fact, for AXAF, the resupply and module changeout will likely be

accomplished without fluid transfer across an umbilical interface since its

design uses experiment modules which are removeable/replaceable and contain

specific fluids). Such spacecraft will likely be returned to the Orbiter for

fluid transfer/modular exchange servicing since, as previously discussed, the

resulting economic beneifts are much greater. Therefore, they are not

recommended for remote resupply using the OMV with an add-on resupply module

kit.

The final list of spacecraft candidate DEMs without concurrent maintenance all

involve propellant transfer (either hydrazine or storable bi-propellant) using

the remote (in-situ) resupply capability of the Orbital Maneuvering Vehicle.

The Extreme Ultra Violet Explorer (EUVE), Earth Observation System (EOS) and

PROTEUS programs all exhibit limited economic payoff (relative to other

candidates) and, given identical risk levels associated with on-orbit

propellant transfer, were eliminated in favor of the two selected "best"

design reference missions— and Geosynchronous Equatorial Orbit Spacecraft

programs in which consumables resupply is likely to be a specific design

requirement to allow increased revenue from operations in GEO or reduced

transportation costs to LEO through propulsion module propellant top-off or

resupply of a reusable version of the satellite/propulsion module. These LEO

operations could both be accomplished using an OMV/EM based at a LEO depot.

Figure 54 indicates these "resupply only" DEM selections, along with their

common fluids, NTO/MMH bi-prqpellant, on a plot of estimated risk versus net

economic benefit.

Ample evidence currently exists (e.g., Insat, Intelsat VI, and NASA/MSFC's

baseline OMV) to indicate an evolution from hydrazine to storable

bi-propellant forms of energy storage on-orbit. The study mid-term review

recommended the DRMs are shown in Figure 55«

2118e/ - 93 -
Recommended DRMs for Remote Fluids

Resupply Without Maintenance

NET ECONOMIC BENEFIT

(RESUPPLY W/0 MAINTENANCE)

JM84

200 r

100

«

LOR

000

OPERATIONAL

BENEFIT

OMV61

BI-PROPELLANT

i

OMV S11

AXAF

OMV 61

OMV 61

OMVS1

GRO-

® SEOPLATJ

® COMSAT

PROTEUS 9

OMV 51 ® EUVEOMVS1

EOS OMV 61

LOW MODERATE HIGH RISK

Figure 54. Estimated Risk versus Net Economic Benefit for Selected

DRMs without Concurrent Servicing

Recommended Design Reference Missions

• REF FLUIDS FOR RESUPPLY ARE BI-PROPELLANTS

• NTO & MMH

• RATIONALE:

• LARGEST PROJECTED FUTURE QUANTITY

• NATURAL EVOLUTION FROM N2H4 IN-BAY DEMONSTRATIONS

• BASIC PROPELLANT FOR NUMEROUS FUTURE SPACECRAFT WHICH CAN

BENEFIT FROM RESUPPLY

LARGE COMSATS (ALREADY USED ON INTELSAT VI)

FUTURE LARGE GEO PLATFORM

REFUELING OMV & HIGH PERFORMANCE STORABLE STAGE

POTENTIAL RESUPPLY DEPOT REQUIREMENT

GEO

LEO

• REFERENCE QUANTITY TBD

• CANDIDATE MISSIONS

</ LARGE GEO PLATFORM

V FUTURE LARGE COMSATS

VRESUPPLY IN LEO

• DOD (MANEUVERING)

• ORBITAL VEHICLES

Figure 55 • Mid-Term Review DRM Recommendations

2118e/ - 94 -
Based on these recommendations and comments at the mid-term review, additional

work was performed to summarize the range of fluid requirements. The initial

focus of the effectiveness analysis was on the end users of resupply services,

such as particular satellites. In the course of this study, it became

apparent that the resupply module was most effective when considered an

integral part of the entire space infrastructure. The variety of potential

uses for a resupply module, other than simple resupply, presorted in the

proceeding pages are summarized in Figure 56.

The resupply scenarios developed in the effectiveness analysis for LEO and GEO

operations were used to create mission event timelines and to better

understand the operational issues in remote resupply. They were the basis of

the eventual final set of Design Reference Missions shown in Figure 56. This

set of DRMs included a broader performance range of potential mission

scenarios (with a consequent sizing impact for the resupply module). The

broad operational range and complexity that might be required of a resupply

module necessitated a large number of Design Reference Missions, without

specifying particular satellite programs. We identified near and far-term

resupply candidates with priority requirements, bounded the range of

performance requirements, and proceeded with the concept Definition Task.

3.2 Concept Design (Task 2)

The Concept Definition task was divided into two subtasks: 2.1 - Develop

Resupply Module Design, and 2.2 - Generate Demonstration Concept. These tasks

incorporated inputs from Task 1 including: Design Reference Missions and

Scenarios, Missions Operations Requirements, System/Subsystem interface

requirements, and systems concepts. The objective of Subtask 2.1 - Concept

Design - was to develop a resupply module concept design including provisions

for: propellant tankage, pressurant storage bottles, propellant and

pressurant transfer systems and supporting mechanisms which are consistant

with the results of Task 1. In conjunction with this design definition those

impacts to the user spacecraft, which are required to support on-orbit remote

resupply, were identified. The objective of subtask 2.2 was to develop a

flight demonstration concept which would provide sufficient validation of the

resupply module concept defined in subtask 2.1. The design and test concepts

defined in subtask 2 are intended to serve as a basis for the programmatics

section of the study.

3.2.1 Operational System Configuration Selection

The ten basic mission scenarios/DRM's indicated in Figure 56 were evaluated

for sizing requirements to the Expendables Resupply Module propellant

requirements from 6,000 to 45,000 Ibs. and helium requirements from zero to 90

Ibs. Scenario two is the driving requirement for GEO operations. In this

mission the full resupply module capacity is used to supply enough propellant

to transfer an OMV plus its servicer kit to GEO. Resupply in this case would

be directly from the OMV but the OMV would not be recommend from GEO.

Scenario five is the driving requirement for LEO remote resupply operations.

This scenario is the refueling of a high performance reusable storable OTV.

Lesser levels of resupply module capability are provided by off-loading the

module rather than removing tankage. The objective is to adapt one basic

"core" resupply module design to many different uses.

2118e/ - 95 -
^

CONTINGENCY MULTIPLE S/C

RESUPPLY FROM OMV

PLANNED RESUPPLY/MODULE

CHANGE-OUT FROM OMV .

ERVICER

PROPELLANT TRANSFER

FROM RM TO OMV FOR.MORE

ECONOMICAL TRANSFER TO

DEPOT OF

RM'S

TOP-OFF

COMPACT INTEGRAL

PROPUL S.'C

WIM 3%%%

REFUEL REUSABLE

PERIGEE STAGE

OPERATE OMV/RM

FROM LEO DEPOT

FOR PROPELLANT

TRANSFER OPERATIONS

1QJ

 RM/OMV USE FOR

^^ PROPELLANT TRANSFER

FROM STS TO DEPOT

RESUPPLY FROM

OMV

POTENTIAL

MISSION

SCENARIO

-*SfT\ ^>(J

=OQ

0»Q

EXAM

RM

PROPELLANT

LBS.

38,000

(1,000*)

45,500

(2,000*)

L 5 fiAft

45,500

6,000

PLE

RM

HELIUM

LBS.

----

4*

70

90

12

Figure 56. Summary of Resupply Module Scenarios/Fluid Transfer

Requirements

The ERM in the various scenarios served as an OMV extended mission kit, a LEO

refueling SRM, a LEO propellant and helium storage depot, a precursor

reusdable upper stage, and a LEO propellant depot for STS scavenged storable

propellants.

A versatile, low-cost resupply module concept resulted from the trade analyses

and design definition phases of the study. Some of the various trades and

design characteristics included propellant transfer alternatives, weight

trades for a pump-fed versus pressure regulated transfer systems, propellant

system schematics, transient and stady state pressure drop characteristics,

fluid interface panel design, thermal control provisions, electrical power

requirements and distribution, avionics and data processing interfaces, major

subsystem weight breakdowns, user spacecraft impact assessment, general

configuration layouts and key features. To provide the required propellant

capacity for the envelope of propellant quantity requirements associated with

the selected resupply module scenarios, the resupply module configuration

depicted on Figure 51 was selected. This concept was selected from among

several competing approaches mainly on the basis of overall structural

efficiency in combination with utilization of existing hardware for low

development cost. The resupply module is supported at its forward end by an

existing IUS upper stage forward cradle. This cradle includes load

equalization capability which reduces the structural redundancy between the

2118e/ - 96 -
resupply module and the orbiter by one degree. This allows a minimum of

structural weight on the forward end of the resupply module for attachment to

the orbiter payload bay. The resupply module uses modified Orbiter QMS

tanks. The modifications include a lengthened cyclinderical section and a

zero "g" PHD.

RM/OMV INTERFACE

OMV

QMS

PROPELLANT

TANKS (6)

(MODIFIED)

'FORWARD UMBILICAL

(REMOVEABLE)

KEY CHARACTERISTTCS

• DRY WT FOR RESUPPLY MODULE ONLY

(AS SHOWN) = 7960 LB

- BI-PROPELLANT TRANSFER

ONLY IN LEO (FROM DEPOT)

- 45,000 LB USABLE PROP

- ADD 1171 LB FOR HE RESUPPLY

CAPABILITY

• (6) STRETCHED OMS TANKS WITH NEW

PMD FOR 0-6 PROP MGMT AND ULLAGE

POSITIONING

• IUS FWD CRADLE USED IN CONJUNCTION

WITH INTEGRAL AFT LONGERON ATTACH

PTS FOR MINIMUM FLIGHT WEIGHT

• OMV SUPPORTED FROM RM IN PAYLOAD BAY

• DRY WT AS PROPELLANT KIT FOR OMV

ONLY (INCLUDES SIMPLIFIED PMD'S AND

ELIM OF COMPRESSORS/BATTERIES FOR)

- 7200 LB

Figure 57. EM/OMV Configuration

As shown on Figure 58, the basic structural components of this Resupply Module

are very simple, contributing to meeting the objective of low cost. Yet, the

configuration is very efficient structurally. All fore and aft loads and part

of the vertical loads are reacted at two payload bay longeron attachment

points in the main structural bulkhead. As a representative attachment to the

OMV, six bolts are provided. For the case where the Resupply Module is

detached from the OMV in GEO, these bolts may be explosive. The main bulkhead

also had a keel attachment for the Orbiter Payload Bay.

The tanks are supported by the main bulkhead for fore and aft, as well as

lateral and vertical loads. They are stabilized at their forward ends by

machined fittings which attach to the cylindrical center body of the Resupply

Module. This body provides the main bending stiffness of the Resupply

Module. NASTRAN finite element analysis showed that this concept meets the

Orbiter fundamental response stiffness requrement of 7 Hz even when fully

2134e/ - 97 -
loaded with propellant. Moat elements of this Resupply Module structure are

machined aluminum plate, including the center body which assembled from

quarter sections that are rolled to proper shape after machining.

PROP. TANK FWD.

SUPPORT FITTING

EXPLOSIVE BOLT

PLCS. (BASELINE)-

PROP./He TANK

FORWARD SUPPORT

FITTING

STS ORBITER

MPS HI-PRESS."

HELIUM TANKS

12 REQD.

%/

•s./ NJ r^*^-3

^er** Vtr-i

0

{* t—&\y -/C?NS ^ ^«<r

^- -^.vl ^

SAFTKEELPIN

"RM" MAIN STRUCTURE

r\

A

Figure 58. Resupply Module Structural Concept

A design layout (Figures 4 & 5) was produced of the selected adaptable

Resupply Module configuration. These layouts were used to develop weight

estimates for the resupply module structure. Two adaptations are depicted.

In both cases the OMV is supported from the resupply module in the orbiter

payload bay (OMV support fittings could be removed to reduce weight and

cost). The combination OMV/ERM would be deployed from the payload bay using

the orbiter RMS. The adaptation depicted with the truss interstage allows

sufficient volume between the OMV and ERM for the OMV's servicer kit (for

missions in which concurrent module change-out type servicing is required).

2134e/ - 98 -
3-2.2 Safety and Contamination Control Requirements and Considerations

The general requirements presented in Chapter 2 of "Safety Policy &

Requirements for Payloads Using the Space Transportation System, NHB 1700.7A,"

were used as a guideline in defining the Resupply Module subsystem's

redundancy and safety requirements. Specifically, subsystem/functions that

could result in a catastrophic hazard (i.e., potential for personal injury,

loss of orbiter, space station, or other free flying vehicles) such as a

failure in the propellant transfer or hazardous waste scavenging subsystems,

should be controlled by a minimum of three independent mechanical inhibits.

Also, since the pressurant system is considered a critical hazard (i.e.,

potential for damage to the orbiter, space station, or other free flying

vehicles) and, as such, requires to be controlled by two independent inhibits.

The safety requirements presented in Chapter 2 of "Safety Policy and

Requirements for Payloads Using the Space Transportation Systms," NHB 1200.7A,

December 9» 1980 are used as the safety and redundancy design criteria for all

orbiter payloads. Contact GSFC NAS5-26152 (Multi-Mission Modular Spacecraft

In-Orbit Refueling Study) results, which included the Orbital Resupply

System's (ORS) interpretations of the NHB 1700.7 safety requirements, was used

as a guideline in explicating the Resupply Module (RM) System's redundancy and

safety tolerances, defined by HHB 1700.7- Some of the more critical RM

related items are listed as follows:

a) For hazardous functions, the Resupply Module must tolerate a minimum

of credible failures and/or operator errors determined by the hazard

level. This criterion applies when the loss of a function or the

inadvertent occurrence of a function results in a hazardous event.

b) No single failure or operator error shall result in damage to STS

equipment, space station, satellite, or resupply module; or in the

use of contingency or emergency procedures.

c) No combination of two failures, operator errors, and RF (radio

frequency) signals shall result in the potential for personal injury,

loss of the orbiter, space station, satellites, resupply module,

ground facilities, or STS equipment.

d) No single mechanical or electrical failure will endanger personnel,

orbiter, space station, satellite, resupply module, or shuttle

payload. As a possible design goal for the resupply module, no

single mechanical or electrical failure will preclude the transfer of

propellant.

e) A function that could result in a critical hazard must be controlled

by two independent inhibits. For orbiter and possibly space station

deployed payloads (e.g. Resupply Module), monitoring and safing of

the two inhibits will not be required when both inhibit power and

— control circuits for the function are connected to a bus that is not

energized until the payload reaches a safe distance from the orbiter

or space station.

2134e/ - 99 -
1. Presaurant Considerations - The pressurant system ia

considered a critical hazard and, as such, requires two

independent inhibits (e.g. isolation valves, burst disks). A

design goal of three independent inhibits was considered for

the pressurant fill and ullage vent systems of the Orbital

Besupply System. However, system cost and complexity may

force the use of only two shutoffs for the pressurant fill and

vent system for the resupply module.

2. Hazard Detection & Safing - The need for hazard detection and

safing by the flight crew to control time - critical hazards

shall be minimized and will be implemented only when an

alternate means of reduction or control of hazardous

conditions is not available.

f) A function that could result in a catastrophic hazard must be

controlled by a minimum of three independent inhibits, whenever the

hazard potential exists. Paragraph 202-2 of NHB 1700.7 states that

one of the three inhibits must preclude operation by HP commands.

When remotely controlled vehicles (e.g. OMV, Satellites, RM) are

resupplied by the Orbiter and/or Space Station, paragraph 202-2

imposes the requirement for an electrical data/command interface to

exist between the transfer and receiver vehicles; hence, establishing

manual control(through the Orbiter and/or Space Station) over the

propellant transfer process and eliminating the use of RF commands.

However, when remotely controlled vehicles are resupplied by other

remotely controlled vehicles (e.g.. RM resupplying a satellite or OMV,

OMV resupplying a satellite) the entire propellant transfer process

would have to be controlled through RF commands. Under this

situation, paragraph 202-2 may not be considered applicable and may

be modified accordingly.

1. Propellant Transfer - The Expendables Resupply Module will

rquire three independent mechanical inhibits (e.g. isolation

valves, manifold valves, QD's) for its transfer system. Any

residual propellant in the receiver tank that requires

disposal prior to, during, or subsequent to transfer operation

may be captured and stored in disposal tanks aboard the

resupply module. However, whenever applicable, decompostion

reactors could be used to disassociate the residual propellant

into innocuous gases, which in turn would be vented

overboard. Propellant contaminated ullage gases in the

receiver tank that require disposal prior to transfer

operations should be handled using the same techniques

discussed for residual propellant disposal.

2. Deployment and/or Separation - The premature deployment or

separation of the ERM to a condition where it cannot withstand

STS-induced loads (including landing) or will prevent safe

entry of the Orbiter is a catastrophic hazard. The inhibit

mechanisms must be designed to preclude inadvertent mechanical

operation in the induced environments.

2134e/ - 100 -
3. Deployment/Extension Preventing Payload Bay Door Closure - If

during planned EM operations an element of the EEM or any ERM

support equipment violates the payload bay door envelope, the

hazard of preventing door closure must be controlled by

independent primary and backup methods, and this combination

must be two failure tolerant.

4. Retrieval of Expendables Resupply Module) - EEM should have

the capability to return hazardous systems to a safe

condition, that is, to meet the requirements under sections

2.1 and 2.2 of NHB 1700.7A. EEM shall provide verification to

the Orbiter, Space Station, or the ground that safing has been

accomplished prior to its retrieval, and while still at a safe

distance from the Orbiter or Space Station.

5« EF Energy Badiation - EEM produced radiated fields from EEM

transmitter antenna systems in excess of levels defined for

payload bay door open operatioins in JSC 07700, Volume XIV,

Attachment I is a catastrophic hazard. Three independent

inhibits to radiation in excess of these levels must be

provided.

Structural safety requirements listed in the GSFC Study, Multimission

Modular Spacecraft In-Orbit Eefueling NAS5-26152, as applicable to EEM,

are given below.

a. The structural design will provide ultimate factors of safety no

less than 1.4- for all mission phases except emergency landing.

When failure of structure can result in a catastrophic event

design will be based on fracture control procedures.

b. The structural design will comply with the ultimate design load

factors for emergency landing loads are specified in JSC07700,

Volume XIV, Attachment I.

c. The structural design will comply with the stress corrosion

requirements of MSFC-SPEC-522.

d. All pressurant and propellant tanks will have a minimum

burst-to-operating pressure ratio of 4:1 in the prescence of

ground personnel. The other considerations described in NHB

1700.7 for pressure vessels also apply.

e. All lines and fittings will have an ultimate safety factor of no

less than four. Components such as valves, regulators, etc., will

have an ultimate safety factor of no less than 2.5-

A burst disk in series with a relief valve should be considered for

thermal relief mechanisms. Each element should be designed to operate

when the internal pressure level exceeds 1.1 times the maximum operating

pressure. The relief valve should be located such that it can be checked

without violtating the remainder of the system. The propellant relief

assembly should be considered for connection to a disposal container.

To insure the integrity of the Shuttle Orbiter1s Payload Bay environment and

the local environment of the Space Station and other free-flying spacecraft, a

disconnect/line purging system should be incorporated into the propellant

transfer systems of space vehicles requiring on-orbit refueling.

2134e/ - 101 -
Any propellant or contaminated ullage gas leaks, spills, and/or venting would

create a tenuous atmosphere around the space vehicles involved. The

contamination to the local environments of these spacecraft could obstruct the

field-of-view of the optical telescopes and highly sensitive cryogenically

cooled sensors on these vehicles.

In addtion to obstructing the field-of-view, if vented residual

propellents and ullage gases were to come in contract with experiments on

board the Orbiter, Space Station, or satellites, the caustic and toxic

characteristics of monopropellants and hypergolic propellants could destroy

the optical and/or the cryogenically cooled surfaces of these experiments.

Another area of concern, in conjunction with propellant spills and venting,

are Extra Vehicular Activities (EVA's). EVA's may necessitate maneuvers in

and/or near the vicinity of vented propellant clouds or spills. Astronauts

could conceivably return to the Orbiter or Space Station in propellant

contaminated space suits, creating a potentially hazardous condition within

the interior environment of these vehicles.

Venting of residual propellant and ullage prior to propellant transfer

operations within or in close proximity to the Orbiter payload bay is not

acceptable. According to orbiter safety policies, hazardous materials shall

not be released or ejected in or near the Orbiter.

Because of the numerous concerns associated with propellant/ullage leaks,

spills, and venting, it may be in the best Interest of both the supply and

receiver vehicles to implement a hazardous waste storage/purge subsystem.

Regulated pressure gas (which would be supplied by a resupply module, or any

other resupply system), would flow through the propellant and ullage transfer

purge circuits and clear all refueling disconnects and adjoining lines of

residual propellants, prior to disengaging the resupply and receiver vehicles.

A simplified disconnect/line purge circuit is illustrated in Figure 59. The

safing of these propellant and ullage transfer circuits would help minimize

the possibility of a leak or spill emanating from the propellant/ullage

transfer interface panel.

3.2.3 Pump Versus Pressure Regulated Propellant Transfer

Two different transfer systems were studied to transfer 45,500 Ibs of

bipropellant from the Expendables Resupply Module. The first system is a

pressure regulated system with propellant tanks at 250 psia. This system will

be referred to as the pressure regulated system. The pressure in the

pressurant tanks in this pressure regulated bipropellant transfer system was

allowed to drop from the maximum operating pressure to 500 psia.

Two alternatives for the first system studied looked at pressure regulated

system weights using existing pressurant tanks with a maximum operating

pressure of 4,500 psia. A new 6,000 psia maximum operating pressure tank

system was also evaluated.

2134e/ - 102 -
Receive* VCHICU nesumr VEHICIC

— — — UlUgc rurge Circuit

——•— PropcllMt Purge Circuit

Waste

Stortgi

I**

Figure 59« Example Disconnect/Line Purge and Waste Storage Subsystem

The second bipropellant transfer system used a bipropellant pump along with a

pressurant system as shown in Table 15- This system will be referred to as

the combination pump and pressurant system. In this system the propellant

tanks are pressure regulated to 50 psia. The pressurant supply pressure is

allowed to drop from the maximum operating pressure of the pressurant tank to

300 psia. In this system the propellant is transferred with a pump and a

sufficient amount of helium must be carried to replace the volume vacated by

the transferred bipropellant and also pressurize the propellant tank to 50

psia. This lower propellant tank pressure requires fewer pressurant storage

tanks and allows a pressurant system weight savings. The combination pump and

pressurant system weight has two major elements.

One element comprises the pump system and the other the pressurant tank system

weight. The pump system weight depends heavily on the power source. The

battery is the heaviest component of a pump system as can be seen in Table 16

Parameter

Pump and Motors

Battery

Electronics

Heat Pipe

Total

Table 16. Propellant Transfer Pump System Weights

Types of Batteries

Ag/Zn

30 Ibs

540 Ibs

40 Ibs

15 Ibs

625 Ibs

Li/Ti S2 Li/Tj_ S2 (Developed)

30 Ibs 30 Ibs

378 Ibs 330 Ibs

40 Ibs 40 Ibs

15 Ibs 15 Ibs

463 Ibs 415 Ibs

2134e/ - 103 -
These pump system weights are sized to transfer 45,5000 rbs of bipropellant in

approximately 7 hours.

Two alternatives for the combination pump and pressurant systems weight using

existing pressurant tanks and new 6,000 psia pressurant tanks were also be

evaluated. Alternative number one is to transfer 45,500 Ibs of bipropellant

with a pressure regulated systems using exisiting pressure tanks. To

accomplish this, 10 Orbiter Main Propulsion System (OMPS) helium tanks with a

total pressure regulated system weight of 925 Ibs are required, (see

Table 15) Alternative number two is to transfer 45,500 Ibs of bipropellant

with a pressure regulated system using 6,000 psia pressurant tanks. Four

combinations of 6,000 psia tanks were considered and are presented in Table 17.

Total pressure regulated

No. of Tanks Volume of each tank, in^ system weigt, Ibm

2 25,170 609

3 16,648 608

4 12,508 602

5 10,034 608

Table 17. 6000 Psi Helium Tank Combinations

Alternative number three is to transfer 45,500 Ibs of bipropellant using the

combination pump and pressurant system with existing pressurant tanks. Two

OMPS helium tanks with a weight (tanks + helium) of 185 Ibs are required. The

total weight of the combination pump and pressurant system would be the weight

of the tanks plus the weight of the pump system. The weight of the developed

Li/TiS2 battery powered pump will be used since it should be attainable in

the near future. Therefore, the total weight (tanks plus helium plus pump

system) to do the above transfer wil be 600 Ibs for the system with 2 OMPS

tanks (See able 15),

Alternative number four is to transfer 45,500 Ibs of bipropellant using the

combination pump and pressurant system with new 6,000 psia pressurant tanks.

The weights and volumes of the different 6,000 psi tank combinations can be

seen below in Table 18.

No. of Tanks Volume of each tank, in^ Helium system weight, Ibm

2 5,189 125

3 3,435 124

4 2,591 125

5 2,081 130

Table 18. Helium System Weight Using 6000 Psi Tank Combinations

The above pressurant supply went from an operating pressure of 6,000 pisa to a

minimum pressure of 300 psia -co displace 45,500 Ibs of bipropellant. The toal

combination pump and pressurant system weight (weight of tanks plus helium

plus pump system) for the four different combination of tanks can be seen in

Table 19.

2134e/ - 104 -
No. of Tanks- Total, Combination Pump and Pressurant System Wt. (Lbs)

2 540

^3 __... 539

4 540

5 545

Table 19. Combination Pump and Pressurant System Weights

Using the combination pump and pressurant system to transfer 45,500 Ibs of

space storable bipropellants yields 542 savings in weight over a pressure

regulated transfer system which uses existing MPS helium bottles. The new

6,000 psi composite helium bottles are used in the transfer process however

the weight savings are not significant (Table 15). It is concluded that for

systems utilizing existing helium bottles, a combination pump and low pressure

regulated propellant tank is most weight efficient.

3.2.4 Propellant Transfer Process Selection

Numerous propellant transfer scenarios and timelines can be projected for the

on-orbit resupply of propellants to satellites, the OMV, propellant tank

farms, etc. The various scenarios depend upon the type of tank being

resupplied (i.e. screen, vane, or diaphragm), the type of tank being used to

supply propellant, and the type of propellant being resupplied (NTO, MMH or

^H^). All of the scenarios however can be placed in one of three general

categories: ullage recompression, ullage vent, and ullage exchange. A brief

discussion of all three methods of propellant transfer are presented in the

following sections. A summary of their key advantages and disadvantages are

tabulated in Table 20. The rationale for selecting the ullage exchange

scenario as the principle propellant transfer process for the resupply module

is also presented.

The ullage recompression method of propellant transfer will be used to

resupply tank systems which operate in blowdown. This is assumed feasible for

either screen, vane, or diaphragm tanks with either mono or bi-propellants.

The resupply would Ideally be performed from a system pressure regulated at a

pressure equal to or greater than the beginning of life blowdown pressure in

the receiver systems (typically about 350 psia). If this is not possible a

liquid pump will be required in the resupply system to obtain the higher

pressures in the receiver system. The major advantages to propellant transfer

using this method are (1) its mechanical and operational simplicity and (2)

resupply of the pressurant gas is not required, though it may be necessary to

periodically make up the pressurant gas which is lost through propellant

saturation. Figure 60 depicts a ullage recompression type of resupply.

A thermal analysis of the adiabatlc recompression of propellant tank ullage

gas reveals that the heat generated by this process can exceed the maximum

operating temperature limits of existing propellant tanks, the

auto-decomposition temperature of hydrazine (310°F), and the

auto-decomposition temperature monomethyl-hydrazine (400°F). A

representative adiabatic,

2134e/ - 105 -
Table 20. On-Orbit Propellant Resupply Options

ORIGINAL PAGE IS

OF POOR QUALfTY

1 1

> Transfer Options 1

I 1

1 1

1 1

1 1

1 Ullage Recoipression 1

1

1

1

1

1

1

1

1

1

1

1 Ullage Ventinq

1 1

1 1

1 1

1

1

1

1

1

1

1

1

1

1

1

1

1

1

Ullage Exchange 1

1

1

1

1

1

1

1

1

1

1

1

1

1

1

1

1

Hain Advantages

t Mechanical Siiplicity; No provisions required

for venting and repressurizing.

t Operational Siiplicity, Only Propel 1 ant is

transfer ed.

t Loner heat dissipation requireaents than

than ullage recoipression resupplies.

t Ion pressure propeilant resupply.

t Constant pressure resupply.

t Negligable coipressve heating effects.

t Receiver tank ullage is transfered over to

the supply tank, and reused. A large waste

scavenging subsystei nould not be required.

t Host versatile of all transfer options

-usable in other resupply scenarios, depending

on receiver vehicle requireients.

1

1

1

1

1

1

1

1

1

1

1

1

1

I

1

1

1

1

1

1

1

I

1

1

1

1

1

I

1

I

1

1

1

1

1

1

1

1

1

1

1

1

1

1

1

1

1

1

1

Rain Disadvantages

t Recotpressian of the receiver tanks

ullage bubble requires a large heat

dissipation.

t The pressure of the resuoply propeilant lust be higher than the

beginning-of-life receiver tank

pressure.

t Mechanically isre cesplex than ullage

recoipression resupplies.

t Operationally sore coin! ex than ullage

recoipressisn resupplies.

t Pressurant resupply required.

t Overboard venting of propel last

saturated ullage gas could be a contaiination concern. Iiplesentation

of a waste scavenging suisystsa far

storage of ullage gas :s net feasanla.

1 Can utilize only those PAD' 3 capable

of positioning the prop, tank's ullage

bubble.

t Propellant tetering required.

t Mechanically lore coaplex than ullage

recoipression resupplies.

t Operationally acre cocplex than ullage

recsapression resupplies.

t Pressurant resupply required.

t Propellant transfer pusps necessary

for this transfer option.

(puips are a possible option far

ullage venting fc recoapresion)

t Can utilize only those PAD'S capable

of positioning the prop, tank's ullage

bubble.

t Propellant letenng required.

1

1

1

1

1

ii

i

1

1

1

1

1

1

1

i

1

1

1

1

1

I

1

1

i

1

1

1

1

1

;*

i

i

i

i

i

i

i

i

i

i

i

i

i

i

i

i

i

i

i

i

i

i

i

i

i

i

i

2134e/ - 106 -
RECEIVER VEHICLE I RESUPPLY MODULE

RECEIVER

TAMC ULL«fiE

PROPELLANT

TRANSFER

SYSTEM

PROPELUANT

OPTIONAL

PROPELLANT

PUXP

Figure 60. Ullagae Eecompression Propellant Resupply

thermal loading profile for ullage recompression resupply is illustrated in

Figure 61. Even though the extreme temperatures calculated by the adiabatic

analysis are not an exact representation of real-life compressive heating, the

trends illustrated by the thermal profiles show that excessive amounts of heat

could be generated by a ullage recompression type of propellant resupply.

To transfer propellant using the ullage recompression method only one

mechanical connection is required, that being the propellant fill line.

Additional propellant disconnects would provide for system redundancy,

purging, and safing. The propellant quantity transferred could be measured

with flow meters or quantity gaging on the supplier and receiver.

The ullage exchange scenario is used for resupplying pressure regulated

propulsion systems. A simple pictorial representation of this resupply

scenario is depicted in Figure 62. Ullage gas is displaced out of the

receiver vehicle's propellant tanks as the incoming resupply propellant fills

the tank. The ullage gas is then transferred to the resupply module's

propellant tanks, filling the volume left by the transferred propellant.

This resupply scenario is a constant pressure process, requiring the use of

propellant pumps for the transfer of resupply propellant. However, current

bipropellant seal technology may find it difficult to produce a seal which can

endure the cycle life and endurance specifications required by the resupply

module's propellant pumps. An advantage to having a constant pressure,

propellant transfer process is the absence of extreme compressive heating

effects. The receiver tank ullage gas is displaced out of the tank rather

than recompressed; thus eliminating the need to dissipate any large quantities

of heat (as would be generated if the ullage were recompressed).

2134e/ - 107 -
1500 f 1000 f

0 .1 .2 .3 .4 .5 .6 .7

TIM ( hour* )

.8 .9 1 1.1 1.2

Figure 61. Ullage Recompression Thermal and Pressure

Profiles (Resupply Volume = 144,500 cu.in.)

RECEIVER VEHICLE RESUPPLY MODULE

PBESSKUNT

TAJ*

FCCEIVER

TAMC ULLAGE

PRESSURANT

TRANSFER

SUBSYSTEM

PROFELLANT

TRANSFER

SUBSYSTEM

Figure 62. Ullage Exchange Propellaat Resupply

2134e/ - 108 -
Flowmeters would be used to calculate the quantity of propellant transferred

to the receiver vehicle. A pressure, volume, and temperature (PVT) gaging

system will not work on a resupply system of this type, due to the

interconnections of the ullages.

Resupply modules designed for ullage exchange resupply scenarios are flexible

in that they can also resupply propulsion systems operating in a blowdown node

using a ullage recompression type process. However, unlike ullage

recompression propellant transfers, the ullage exchange process must utilize

only those propellant acquisition devices (PAD) capable of positioning the

ullage bubble within the receiver propellant tank. If the PAD is unable to do

so, resupply propellant could unknowingly be transferred (through the ullage

exchange lines) back into the resupply module.

In addition to the transfer of propellant, a pressure regulated propulsion

system requires the resupply of high-pressure helium gas. The development of

a high-pressure gas compressor would be most beneficial in improving the

volumetric efficiency of the resupply module's pressurant supply tanks. If a

contamination-free gas compressor is not technically and/or economically

feasible, the pressurant transfer procedure would be reduced to operating as

high-pressure blow-down process.

Ullage Venting Resupply transfer can be used in resupply pressure regulated

propulsion systems. With this method, the receiver tank is vented to near the

propellant vapor pressure before transferring propellant. The required amount

of propellant is then resupplied. The tank is then pressurized collapsing any

propellant vapor bubbles trapped in the propellant acquisition device. Just

as with the ullage exchange resupply process, ullage venting resupply also

requires the transfer of high-pressure helium gas, which is used by the

pressure regulated propulsion system.

It should be noted, that the overboard venting of propellant saturated ullage

gas could cause the degradation of the local environment and/or damage the

sensors and experiments of nearby spacecraft. It has been suggested that the

implementation of a waste scavenging subsystem on the resupply module, could

safely contain and store the vented ullage gas. However, an analysis on the

feasibility of a wste scavenging subsystem reveals that the waste storage tank

volume required for storage of the resupply receiver vehicle's propellant tank

ullage gas could be substantial. Figure 65 presents the required waste

storage tank volume as a function of the resupply receiver tank's ullage

volume. For an empty shuttle RCS tank to be vented down to 8$ of its

operational pressure, slightly over two QMS tanks would be required to receive

the vented ullage gas. Clearly the waste storage tank volume requirements

make the implementation of a waste scavenging subsystem of this approach

impracticable. Contamination sensitive spacecraft would have to make its own

provisions in protecting itself during propellant resupply operations, or

disallow ullage venting propellant resupplies.

An analysis of the various on-orbit propellant resupply scenarios suggest that

the resupply module's propellant transfer subsystems be designed to carry out

"ullage exchange" type of propellant transfers.

2134e/ - 109 -
OF

30 '

20 •

15 •

5 '

111 !

^^^

s

s

i i n i

_ . ..

>

V

n ~ -

^ !

S

•A

1 1 1 1

J/\

S

{ ' " t

X (

I " '«

tA

a

^

a

1

1 1 — 1—

i

i

>1

Xl '

r ' J :

Assiaattons

rocess assuned

a ot Isothermal.

eceiver tank,

ln* > O.OB

operating

xamle- For 250

sia mitait operting, final press

s 20 ps la

1 1 1 1 I 1

A

1 1 1

0 SO tOO 150 200 250 300

WASTE STORAGE TNK VOL HEQO (ft1)

Figure 63- Waste Storage Subsystem Tank Volume Requirements

Ullage venting resupply scenarios were the least acceptable of the three

transfer methods analyzed. The overboard venting of propellant saturated

ullage gas could be a contamination concern. Implementation of a waste

scavenging subsystem for storage of ullage gas was found to be impracticable.

Since the mechanical and operational characteristics of ullage venting

scenarios are similar to ullage exchange scenarios, it would be most

beneficial for contamination sensitive spacecraft to convert from ullage

venting to ullage exchange propellant resupplies.

Even though ullage exchange scenarios are mechanically and operationally more

complex than ullage recompression scenarios, a resupply module designed for

the ullage exchange process can also resupply vehicles' using the ullage

recompression process. The versatility of the ullage exchange process is it

biggest asset.

3.2.5 Conceptual Designs For Ullage Exchange Bi-Propellant and Pressurant

Transfer Proces "~~~~~

Preliminary conceptual design system schematics for the on-orbit resupply of

bi-propellants the using ullage exchange process and their accompanying

pressurants were defined and are presented in Figures 64 and 65. The

differences between the two conceptual resupply schematics, are in the design

of the pressurant subsystems. The high-pressure pressurization/pressurant

resupply subsystem defined as Option "A" (Figure 64-), stores and provides

gaseous helium or nitrogen to the propellant transfer, waste storage/purge

subsystem, and the pressurant transfer susystem. Pressurant for the

propellant transfer and waste storage/purge subsystems is supplied via

2134e/ - 110 -
Figure 64. Option "A: Expendable Resupply System Schematic,

with the Helium Transfer

OF POOR

S3

5

 £ s * s » i

fi ° t S.

«» • M*M

IMil

II IP

Wlpllglli

2l34e/

-SV—T-V®

e 5 5: & s: £ S S* fc •.»

ii!

Ill
parallel and series redundant regulators, while pressurant transfer is

controlled by high-pressure gas compressors. The presurant subsystem defined

as Option "B" (Figure 65). was not designed for on-orbit resupply of

pressurant. The subsystem provides helium or nitrogen for propellant transfer

and disconnect/line purging only.

Contingent on the type of propellant management device used in the resupply

receiver vehicle's propellant tanks (capillary screen, surface tension vane,

diaphragm), all three types of propellant resupply schemes: ullage transfer,

ullage recompression, and ullage venting, could be resupplied by the Option

"A" transfer systems. Without the addition of a pressurant resupply module,

the Option "B" transfer system will be limited to ullage recompression

resupplies, or to enhance an OMV's mission capabilities. The propellant

transfer sysltem's propellant pumps, or a pressure regulated procedure could

be used in the resupply of mono-propellant or space storable bi-propellant

propulsion systems. Due to the hypergolic nature of these bi-propellants two

independent propellant and waste storage subsystems (one for the transfer of

fuel and the other for oxidizer) would be required. Two independent

pressurant systems, one for the fuel system and the other for the oxidizer

system, may be necessary since fuel and oxidizer vapors cannot be restrained

from migrating into common lines.

The safety and redundancy requirements defined by "Safety Policy and

Requirements for Payloads Using the Space Transportation System," NHB 1700.7A

require two independent inhibits to control/restrain all presurant transfer

paths leading to overboard outlets. In the case of propellant transfer and

wase management functions, safety guidelines require three independent inhibis

controlling propellant paths leading to overboard outlets.

The majority of the components in the preliminary resupply design schematic

are existing qualified components or utilize existing component technology to

manufacture new hardware. Table 21 of the enclosure presents the component

legend for the resupply system schematics. However, there are three major

components in the conceptual resupply system design that require advanced

development. One of these new components is a high-pressure gas compressor

used in pressurant transfer missions (e.g. ullage transfer, ullage venting).

The main obstacles in the development of such a component are:-(l) gas

contamination by lubricants during the compression process, and (2) reducing

the compressor weight.

A propellant transfer pump is another component requiring advanced

development. ¥ith current propellant seal development it may be difficult to

produce a seal which can endure the cycle life and endurance specifications

required by the propellant transfer pumps.

If the development of a contamination-free gas compressor or propellant

transfer pump are not technically and/or economically feasible, the resupply

system design would be limited in its transfer capabilities; the pressurant

system would be reduced to operating as high-pressure blow-down system, and

the propellant transfer system would not be able to accommodate the ullage

transfer resupply scenario.

2134e/ - 112 -
Figure 65. Option "B" Expendable Resupply System Schematic,

Without Helium Transfer

9 9

21?4e/
Table 21. Conceptual Resupply System Design, Component Legend

SYMBOL COHPONET

Cheek Valve

Disconnect/ Fluid Transfer

Disconnect/Cap/ Fluid Servicing £ Check-Out

Dual Burst Disk

Dual Pressure Regulator

Filter

Flcwneter

Gas Compressor

Non-Propulsive Vent

Propellant Pump

Quad-Check Valve

Relief Valve

Tanks: i. Propellant* Stretched QMS

ii. Pressurant* Shuttle MPS

iii. Waste Storage

Vapor Bubble Detector

Valve

2l34e/ - 114 -
ORSGS2W PAGE fS

OP POOR QUALITY

The third new component, fluid transfer disconnect. This disconnect is

expected to be a remotely operated spin-off of the manual valve presently

being developed by NASA JSC.

3-2.6 Propellant Transfer Pumps

propellant trasnsfer may be utilized in any of the three general category

scenarios of propellant resupply (ullage recompression, vent, and exchange).

The general requirements of the pump are determined by the type of receiver

subsystem (vane, screen, and diaphragm). Threfore, the power requirements in

this report were determined for several head values at flow rates up to 10 GPM.

Figure 66, 67, and 68 show the pump power requirements versus the flow rate

for MMH, NTO and the MMH-NTO propellant transfer with a head rise of 250 PSI.

Tvo commercial pumps (Garrett and Marquardt) are compared. The Marquardt pump

can be used up to a flow rate of 21 GPM. Marquardt is the preferred pump for

flow rates greater than 10 GPM. The Garrett pump has a lower power

requirement at flow rates between 1 and 10 GPM, therefore, it is the preferred

3500

3000

2500

2000

1500

1000

500

0

POWER TO PUMP (WATTS)

CHANGE IN PRESSURE 250 PSI

MARQUARDT

PUMP

GARRETT

PUMP

10 15

FLOW RATE (GPM)

20 25

Figure 66. Pump Power For MMH Transfer

- 115 -
pump in the lower flow rate region. Similarity relations were used on the

Garrett pump to redesign a pump with optimum flow rate at 1 GPM. The result

is shown in Figure 67 (labeled similarity pump). The results of the

similarity calculations indicate that the design of a 1 GPM optimum pump would

not result in sufficient energy reductions compared to the Garrett pump.

5000

4500

4000

3500

3000

2500

2000

1500

1000

500

0

PUMP POWER (WATTS)

CHANGE IN PRESSURE 250 PSl

MARQUARDT

PUMP

GARRETT

PUMP

SIMIUkRITY

PUMP

10 15

FLOW RATE (GPM)

20 2

B

w

Figure 67. Pump Power For NTO Transfer

2134e/ - 116 -
8000

7000

6000

5000

4000

3000

2000

1000

0

PUMP POWER (WATTS)

CHANGE IN PRESSURE 250 PSI

MARQUAROT

PUMP

GARRETT

PUMP

10 15

FLOW RATE (6PM)

20 25

Figure 68. Pump Power For MMH and NTO Transfer

Figures 69, 70, 71, and 72 present the pump power requirements versus the flow

rate at various head rise values for the propellants MMH, NTO, a system of MMH

& NTO, and hydrazine. The Garrett pump data was eztapolated to complete the

family of head curves seen in Figures 69, 70, and 72. The data has an

increasing variance with decreasing head and flowrates. The method of data

extrapolation incorporated the available performance pump curves with

similarity relations at lower pump speeds. The power requirements to pump

hydrazine at different heads was determined by using the Garrett pump data and

correcting for the higher density and viscosity of the hydrazine as compared

to MMH.

2134e/ - 117 -
Table 22 lists the power required by the pump system (including motor,

efficiency - 0.90) for MMH, NTO, MMH & NTO, and hydrazine. The values in

Table 22 differ from Figures 69, 70, 71, and 72 as a result of the inclusion

of the motor efficiency. The parameters of the pump's motor are listed in

Table 23. In particular it lists the input requirements of the motor, and

total system weight of the pump and motor system.

300 PSI 2SOPS1 200 PSI 150 PS! 100 PSI

2000

1800

1600

1400

1200

1000

800

600

400

200

0

POWER TO PUMP (WATTS)

4 6

ROW RATE (GPM)

8 10 12

Figure 69. Power to Pump For MMH

2134e/ - 118 -
300 PSI 250 PS1 200 PSI 150PSI 100 PSI

2500

2000

1500

1000

500

POWER TO PUMP (WATTS)

4 6 8

FLOW RATE (GPM)

10 12

Figure 70. Power to Pump For NTO

2134e/ - 119 -
300 PS1 250 PSI 200 PSI 150 PSI 100 PSI

4000

3500

3000

2500

2000

1500

1000

500

POWER TO PUMP (WATTS)

0 1 2345678 9 1 0 1 1

FLOW RATE (GPM)

Figure 71. Power to Pump For NTO & MMH

2134e/ - 120 -
300 PSI 250 PS1 200 PSI 150 PSI 100 PSI

2000

1800

1600

1400

1200

1000

800

600

400

200

0

POWER TO PUMP (WATTS)

4 6

FLOW RATE (CPM)

8 10 12

Figure 72. Power to Pump For Hydrazine

- 121 -
Head Rise Flow Rate Power (watts)

(PSI) (GPM) MMH NTO MMH 4 NTO HYDRA7TNE

300 10

8

6

4

2

1620

1360

1090

800

445

2330

2030

1710

1270

800

3950

3390

2800

2070

1245

1904

1590

1280

940-

522

250 10 1340 1780 3120 1580

8 1110 1640 2750 1300

6 940 1370 2310 1100

4 655 1070 1725 770

2 367 589 956 430

200 8 890 1290 2180 1040

6 690 1070 1760 809

4 528 800 1330 620

2 289 480 769 339

150 7 556 833 1390 652

6 500 756 1260 587

4 378 589 967 444

2 233 367 600 274

100 7 411 556 967 483

6 352 533 885 414

4 267 413 680 313

2 178 278 456 209

* Pump system consists of one motor for both pumps.

Table 22. Power Required by Pump System for

Various Propellants

2134e/ . 122 -
Table 23. Parameters of Motors and Pumps

Parameters ~ Garrett Marquardt

pump type Centrifugal Centrifugal

operating speed (EPM) 23,664 80,000

optimum flowrate (6FM) 10 21

suction specific speed MMH (11,089) MMH (12,500)

(EPM) NTO (16,013) NTO (8,500)

motor type AC Induction AC Induction

Input Voltage 212 VAC 225 AC

frequency 420 Hz 1400 Hz

current 25 amp 35 amp

operating time >15 hr 1 hr (Grease Pack)

20,000 hr (mist)

efficiency of motor 0.90 0.90

weight of pump 30 30

and motor (Ibs)

3.2.7 Pressurant Transfer

Two helium receiver volumes were investigated. The first receiver volume to

be investigated was an QMS helium tank design case. The second receiver

volume investigated was a volume equivalent to two RCS helium tanks. The

volume of two RCS tanks is in the size range of required helium to expel 2,000

pounds of storable bipropellants.

Different existing helium tanks were studied for resupplying one QMS helium

tank from 500 psi and 70°F. end of life (EOL) to approximately 3600 psi and

70°F. beginning of life (BOL). Figure 73 shows the number and the type of

tanks to resupply one QMS helium tank. When multiple supply tanks were used,

one supply tank at.a time was opened to the receiver tank. This process was

found to increase the efficiency. In Figure 73 the number to the right of the

ID symbol represents the final pressure in the receiver and in the last supply

tank. The supply tanks' initial pressure was taken as the maximum operating

pressure of the particular tank being used. Figure 73 also shows the wide

variation in number and weight of the resupply systems using existing tanks.

For this particular resupply requirement of resupplying one QMS He tank to

3600 psi and 70°F., a system of 6 MPS helium bottles proved to be optimum.

The space transfer of pressurants can be accomplished by two basic approaches

when using a compressor:

1) Using a variable flow rate/inlet pressure, or

2) Using a set flow rate/inlet pressure.

A metal diaphragm compressor will accept a wide range of inlet pressures where

as a 2-stage piston compressor will accept a specific inlet pressure or a

specific flow rate. Flow rates, energy requrements and fill times for these

compressors are presented in Figures 74, 75, and 76. Specific cases of gas

transfer are presented in Tables 24 and 25. A summary of their major

advantages and disadvantages are tabulated in Table 27. The rationale for the

tentative selection of the metal diaphragm compressor in general pressurant

2134e/ - 123 -
Figure ?3« Number of Supply Tanks vs. Tank System Weight ORIGINAL PAGE IS

OF POOR QUALITY

—-^-.-^.:--—;• •:- ~F U. S ":—--.r—.i-^t-^'=-P O =3 •

7^srJ=:t. 't====3; LU ~ o

-A--^Q ^ U>

£233--:*:—*

:""=^-J.i-i-5lZa—;.3 oo o o

----^^,—

*iH9I3M W31SAS 5WV1 -.-.-

2134e/ - 124 -
500PS1A 1000PS1A 1500 PSIA 2000 PSIA

45

40

35

30

25

20

15

10

FLOW RATE (SCFM)

1000 1500 2000 2500 3000 3500 4000

OUTLET PRESSURE (PSIA)

4500 5000

Figure 74. Flow Rates of a 1.5 HP Diaphragm Compressor

at Various Inlet Pressures

2134e/ - 125 -
500PS1A 1000 PSIA 1500 PSIA 2000 PSIA

70

60

50

40

30

20

10

FLOW RATE (SCFM)

1000 1500 2000 2500 3000 3500

OUTLET PRESSURE (PSIA)

4000 4500 5000

?igure 75- Flow Rates of a 3-0 HP Diaphragm Compressor

at Various Inlet Pressures

2134e/ - 126 -
3 I

w

i!

\

S

(ft

12

\

\

\

\

\

\

\

\

\

\

\

\

\

\

\

\

\

\

\

\

\

\

\

\

\

\

S

CM

O

o

2®

a

o u

00

 uS*

O

HO

si

C

s

f

Figure 76. Time Required for Mass Transfer;

Compressor Energy Required for Mass Transfer

Using A 2-Stage Piston Compressor

2134e/ - 127 -
Table 24. Metal Diaphragm Compressor Requirements to Pill

Various Pressure Bottles

Type of Pressure

Bottle

RCS

MPS

QMS

Power of

Compressor

1.5 HP

3.0 HP

1.5 HP

3.0 HP

1.5 HP

3.0 HP

Pill

Time

0.50 hr

0.26 hr

1.4 hr

0.92 hr

5.5 hr

3.6 hr

Total Energy

to Compressor

0.56 kwh

0.58 kwh

1.55 kwh

2.06 kwh

6.09 kwh

8.05 kwh

Required Heat

Rejection Rate

3300 BTU/hr

6320 BTU/hr

3000 BTU/hr

6410 BTU/hr

3000 BTU/hr

6400 BTU/hr

Table 25• Two-Stage Piston Compressor Requirements to Fill

Various Pressure Bottles from a 500 Psia Source

Type of Pressure

Bottle

Power of Fill

Compressor Time

Total Energy

to Compressor

Required Heat

Rejection Rate

RCS

MPS

QMS

2.0 HP

2.0 HP

2.0 HP

0.70 hr

2.52 hr

8.4 hr

1.04 kwh

3.76 kwh

12.5 kwh

.4551 BTU/hr

4552 BTU/hr

4557 BTU/hr

Parameter

Table 26. Parameters of the Compressor/Motor

2-Stage Piston Metal Diaphragm Compressor

Motor

Flow Rate

System Weight

System Volume

2.0 HP

6 SCFM

175 Ibs

1.1 ft3

1.5 HP

20-*-3.0 SCFM

150 Ibs

3.75 ft3

3.0 HP

33-»-4.7 SCFM

200 Ibs

5.6 ft3

2134e/ - 128 -
Table 27. Compressor Options

. Metal Diaphragm Compressor

Advantages

• very low gas contamination

- variable Inlet pressure/flow rates

- less energy storage required

per Ib. of pressurant transferred

- shorter transfer times

Disadvantages

higher compressor/motor weight

greater compressor/motor volume

- lower operating period

possible diaphragm failure

2-Stage Piston Compresor

Advantages

- smaller system volume

- lower compressor/motor weight

- longer operating time

Disadvantages

fixed flow rate/inlet pressure

higher energy storage

required per Ibm of pressurant

longer fill times required

possible gas contamination

transfer is also presented. It is further recommended that prototypes of both

compressors be built and tested. In particular note Table 26 which presents

three compressors with 1.5, 2.0, and 3.0 HP motors and their respective

weights of 150, 175, and 200 Ibs.

A metal diaphragm compressor allows a variable inlet pressure and thus a

variable flow rate. Figure 77 shows a typical metal diaphragm compressor.

Operation starts with the diaphragm deflectd to the bottom of the cavity by

the gas inlet pressure with the hydraulic piston at bottom dead-center

position. As the crankshaft rotates and the hydraulic piston moves toward top

dead-center, the hydraulic pressure increases. When the hydraulic pressure

reaches the pressure level of the inlet gas in the cavity, the diaphragm moves

up compressing the inlet gas.

When the pressure of the inlet gas in the cavity reaches the pressure level

downstream of the discharge check valve, the valve opens and the inlet gas is

discharged into the tank. Since the hydraulic system has slightly more

displacement capacity than the gas system, all gas is displaced. Excess

hydraulic fluid is then released through the hydraulic relief valve (pressure

level is set at just above outlet pressure).

2134e/ - 129 -
UPPCft (GAS) H6AO

DIAPHRAGM GROUP

DIAPHRAGM

CONSTRUCTION

LOWER

(HYOPAUUO

H6AO

AUTOMATIC

HVORAUUC

INJECTION

PUMP

Figure 77. Typical Assembly of a Diaphragm Compressor

Figure 74, Figure 75 and Table 28 present expected flow rates at varying inlet

and outlet pressures for a 1.5 HP and a 3«0 HP motor operating at 800 HPM. In

the process of transferring pressurant from the transfer tank to the receiver

tank the first step will be free transfer, until the two tanks reach a

equilibrium pressure. When the equilibrium pressure is obtained the

compressor is activated and pressurant is transferred until the required

receiver pressure is obtained. Using the variable flow rates (at a variable

inlet pressure) one can determine the time and electrical energy requirements

to fill any tank(s).

Table 24 presents results based on the assumption that the transfer and

receiver tanks are the same type tank. For example, one OMPS transfer tank

transfers pressurant to one OMPS receiver tank. The energy required to run a

3-0 HP motor for 0.92 hr is 2.1 kwh with a required heat rejection of 6410

BTU/hr.

One of the disadvantages of a diaphragm compressor is its higher mass and

volume required for respectable flow rates. Fluitron (a metal diaphragm

compressor company) is testing compressors at higher operating speeds. This

will result in higher flow rates with a small increase in weight. Another way

to double capacity but only increasse the weight by about 22$ is to use a

duplex compressor. A duplex consists of two compressor heads driven by a

single motor, the two compressor heads operate in parallel. Expected weights

2134e/ - 130 -
Table 28. Flow Rates of a Diaphragm Compressor

~at "Various Inlet and Outlet Pressures

Inlet Outlet Pressures

(PSIA)

Flow Rate (SCFM)

1.5 HP 3.0 HP Motor

500— * 1000

1500

2000

2500

3000

3500

4000

4500

5000

8.5

7.0

6.4

5.3

4.5

4.0

3.7

3.2

3.0

13

11

10

7.5

7.0

6.3

5.7

5.0

4.7

1000

1500

2000

1500

2000

2500

3000

3500

4000

4500

5000

2000

2500

3000

3500

4000

4500

5000

2000

2500

3000

3500

4000

4500

5000

15

14

12

11

10.6

10.0

9.4

8.5

25

21

18

16

15.2

15.0

14.7

42

31

26

24

22.4

21

20

24

22

19

18

17

16.2

15.0

13.6

41

34

29

26

24.4

24.1

23.7

69

51

43

39

37

34.5

33

and volumes are presentd in Table 26 based on present technology and assuming

a one-third earth usage weight using light weight.high strength materials and

the removal of structually excess material.

A 2-stage piston compressor is confined to a narrow inlet pressure range and

thus a fixed flow rate. Normally the inlet pressure varies during pressurant

transfer, thus a pressure regulator will be required at the inlet of the

piston compressor. The only advantage of a fixed flow rate is that the system

is quickly in steady state so that energy, time, and heat transfer values are

linear with respect to the weight of gas transfered (Figure 76). Figure 76

was created on the assumed values of a 2.0 HP motor with a flow rate of 6

SCFM. Note in Figure 76 that the mass of nitrogen transferred is greater than

the mass of helium tranferred for the same stored energy. This can be

understood if one realizes that the work done is pressure-volume work and not

2134e/ - 151 -
mass transfer at high pressures. Table 25 presents estimated time, enegy,

heat rejection rates to fill one ORCS, OMPS, or QMS pressure bottles from a

500 psia source using a 2.0 HP motor at a 6 SCFM flow rate.

The major advantages of the 2-stage compressor are: l) a smaller system

volume, 2) a slightly lower weight, and 3) a longer operating time between

servicing. The smaller volume exists due to the ability to build the

compressor into a cylindrical package 10 inches in diameter and 2 feet long.

The diaphragm compressor is large at the diaphragm (1.2 feet in diameter) plus

it requires a larger gear drive. The mtal diaphragm compressor has a lower

expected operating life due to the metal diaphragm (1000 hours as compared to

2000 hours for the 2-stage piston). The diaphragm life can be increased by

using a layered diaphragm of metal sheets or combind metal and elastomer

layers (to 2000 hours).

The major diadvantages of the 2-stage piston compressor are: l) longer fill

times, 2) a higher energy requirement, 3) possible gas contamination, and 4) a

fixed inlet pressure. The fixed inlet pressure will require a pressure

regulator (adds about 5 pounds) and also leads to longer fill times and higher

energy requirements. The longer fill times exist due to the fixed scfm that

enters the piston chamber at a given pressure, where the diaphragm compressor

accepts higher scfm at higher pressures. The 2-piston compressor also loses

the higher pressure gas (a loss of stored energy) due to the fixed

inlet pressure. Gas contamination will be greater problems with the 2-stage

compressor because of the direct contact of the piston with the gas. The

least case contamination will be teflon dust and a worst case will be

hydrocarbon contamination from the grease packs.

Table 27 lists the advantages and disadvantages of the two candidate

compressors. The advantages of the diaphragm compressor include; very low gas

contamination, less energy required to transfer a given mass of gas, shorter

tank transfer times, and a variable inlet pressure, the disadvantages of the

diaphragm compressor include; a greater weight, a larger volume, and a lower

operating life. The metal diaphragm compressor is recommended for use in

cases of transfer between tanks of similar size. The advantages of the

2-stage piston compressor include; a small volume, a lesser weight and a

longer operating life. The disdvantages of the 2-stage piston compressor

include; possible gas contamination, longer fill times and greater energy

requirements because of the fixed inlet pressure. The 2-stage piston

compressor is recommended for usage when the transfer tank is large in

comparison to the receiver tank.

3.2.8 Battery Power Sources for Pumps and Compressors

There exist two types of batteries, the primary battery and the secondary

battery, that can be used as energy storage devices. The primary battery is

to be used in cases where the system is used only once such as an expendable

stage. A primary battery cannot be recharged, the secondary battery can be

recharged and should be applied in cases of multiple use such as LEO where the

resupply module can be recovered.

2134e/ - 132 -
01

-4 O)

O <*•(

X 1-4

•H b.

01 O

f. O

« e'-

03

u

01

4J

01

CO

p

CO

0-

u

01

CO

CQ

ON

CM

s- 4J i-H

en -^

e

o •"

u c

01

X P

td P

O £

CO <

O. -M/

CO

U

0)

6

CO

z

(U

a

o

o

CM

O

o

o

vC

CD

CO

O

-1

^*s

en

P

P

CO

01

O

CO

to

0

jj

&•€

vD

i-H

4J

0)

*

o

vO

P CO

3 0)

O O

O CM O\ CM ON

CM

•

OS

«j\ m \o o\ oo

in r~» i-t o 1-1

iH i-t CM CM CM

CM

O

o O

CM

r*. \o 10 o >A

iH CM ^ CM

O

CM

O

o o o o o

o o o o o o

O -3- O O O O

in iH CM 00 CM 'H

o

o

-aO

o o 1-1 in

P <-»

O « 4J

i-l i-l iH

6

i-l O

01 >

C to m

o •*•*•'

•H Q

—B C MM

JB CA

0)

u

(0

a

O

•o >

01

4-> m

CO "">

3

CO

0)

T5

•H

X

o

Ol

w o 1-1

•^ ao to

o <; 3

e ^ c

•H C CO

SS3 N Z

E

3 10 e c

•H • 0)

C i-l 60 iCO >-*• O s

•U P

•H "Q

B CM ^

3 CO iH

iH -H CU CM

A H ^ X *J -^ O -^

iH 1-1 -rt t4

J J Z Z

o

en

a

a

CO

O

CM

to

CO

01

CO

4J

3

c

60

CO

CU

p

CU

CO

to

cu

60

o

o.

c

0)

p

to

CO

01

•a

O)

CO

I—I

3

c

01

a.

o

•H

U

a,

P

CO

•O

O

O

01

tn

p

CO

•o

o

o

01

co cu

.0 T3

O T-4

e ^H

O *H

•H -H

^J ^J

U CO

tl M

2145e/ - 133 -
Lithium/Thionyl Chloride has the highest energy density of the developed

primary batteries, as can be seen in Table 29 The battery has a 90$ depth of

discharge and a 90$ efficiency in chemical to electrical conversion. The

battery has one disadvantage; it tends to overheat when physically abused.

Physical abuse is clasifed as puncturing, opening, and mutilating. Battery

overheating can also occur by exceeding current drain capabilities or fusing.

However, this disadvantage is of concern only when batteries energy density

exceeds a 5000 Ah capacity.

Zinc/Silver Oxide is another primary battery but at lower energy density than

the Lithium/Thionyl Chloride battery. An advantage Zn/AgO has over Li/SOCl2

is that it can dischage a massive amount of energy without the concern of an

exothermic dischage. Shelf life is a major disadvantage of the battery where

a 16$ energy loss per year occurs as a wet cell. A solution to this problem

is to introduce the electrolyte just before application, either manually or

remotely.

Lithium/Titanium Bisulfide has the highest energy density of the

near-developed secondary batteries. Secondary batteries have a depth of

discharge of 80$ and a chemical to electrical conversion efficiency of about

85$. The battery can be recharged about 200 times before replacement. The

battery is recommended for use when the resupply module can be recovered and

then reused or in conjunction with a photovoltaic system. One disadvantage is

that the battery has not been flight tested since it is still in the

development stage. Simulated flight tests are in progress.

The Nickel-Hydrogen battery has an outstanding cycle life (about 6000

charges/discharges before replacement) for a secondary battery. This allows

the battery to mate well with photovoltaic panels over long time periods. One

drawback of the Ni/Ho system is low energy density (20Wh/lb) as compared to

the Li-Ti 82 (100 Wh/lb). The Ni/H2 system can last 20 years as long as

is cylced lesa than 6000 times. The Ni/H2 battery plus photovoltaic panels

is best used for a continuous power source such as a tank heating system.

Table 30 presents energy weight requirements for the resupply module to

transfer 45,500 Ibs of fuel, fill 12 MPS bottles, run the heaters, etc. The

deployable photovoltaic panel (DPP), used in all scenarios, is the same as

presented in Table 31 using the Si 100 mm cell and a 100 mm cover. The

batteries come from Table 29 and specific amp-hour batteries are used

depending on the application. The system weight includes batteries, DPP,

voltage regulators, charge/discharge unit, and power hardware.

In LEO (scenario 1 Table 30), it was assumed that the battery would be

discharged 60 minutes and recharged 30 minutes and function for one year. The

Ni-H2 battery is preferred in this case over the Li/Ti 82 because of the

number of recharging cycles required in LEO (one every 90 minutes). One 50 Ah

Ni-N2 battery can be used for 5000 to 6000 recharging cycles and still

supply the needed power. Whereas four 400 Ah Li/Ti 82 batteries are

required because they can only be recharged about 200 times this requires more

batteries with a higher energy density to last 1 year of usage.

In GEO (scenario 2 Table 30), it was assumed that the battery would be

cyclically charged for for 12 hours and discharged for 12 hours and be

functional for 1 year. The Li/Ti 82 battery is the preferred battery in

this case. Five 100 Ah Ni-H2 batteries are required to supply power during

-J

2145e/ - 134 -
the eclipse period compared to two 400 an-Li/Ti 83 batteries (with power to

spare). The cycle life of the battery is now of secondary consideration.

What is of concern is to have a large energy density so that the battery

system weight is a minimum during the 12 hour discharging. Using two 400 Ah

Li/Ti 82 gives an expected life of 373 days and weighs 460 Ibs, but using

the five 100 Ah Ni-H2 batteries gives an expected life of about 14 years and

weighs 936 Ibs. Since the scenario only exists for one year the lower weight

Li/Ti 82 battery system is preferred.

Scenario 3 (Table 30) uses only primary batteries with an eight day mission.

Since the Li/SOCl2 cannot be recharged the continuous energy requirements

sets the 8 day mission life. One 8000 Ah battery, weighing 1037 Ibs satisfies

the above requirement. This demonstrates the validity of using a rechargeable

system for continuous energy requirements.

Table 30 Resupply Module Energy Weight Requirements To Transfer

45,500 Ibs of Bipropellant and Pill 12 OMPS Bottles

Scenario 1 - LEO - 1 Year Life

a) Continuous Power - Use Ni-I?2 Batteries, and a 2.2K¥e

Deployable Photovoltaic Panel (DPP)

Peak Power - Use Li/Ti 82 Batteries, and a 1.2 Ktfe DPP

System Weight 1208 Ibs

b) Continuous Power - Use Li/Ti S2 Batteries, and a 1.6 KWe DPP

Peak Power - Use Li/Ti 82 Batteries, and a 1.1 KWe DPP

System Weight 1510 Lbs

Scenario 2 - GEO - 1 Year Life

a) Continuous Power - Use Ni-H2 Batteries, and a 2.2 KWe DPP

Peak Power - Use Li/SOCl2 Batteries

System Weight 1540 Ibs

b) Continuous Power - Use Li/Ti 82 Batteries, and a 2.2 KWe DPP

Peak Power - Use Li/SOCl2 Batteries

System Weight 1064 Ibs

Scenario 3 - LEO or GEO - 8 Day Mission

Use One 8000 Ah Li/SOCl2 Battery

System Weight 1037 Ibs

2145e/ - 135 -
Table 31 Deployable Photovoltaic Panels

Name

Deployable Rigid Solar Arrays

(DRSA)

Deployable Hybrid Solar Arrays

(DHSA)

Solar Electric Propulsion System

(SEPS)

Power Extension Package

(PEP)

Power

(Kite)

1.0

3-7

12.5

26.8

Weight

(lb)

76.4

176

416

1022

BOL-W/lb

13

21

30

32

Deployed A:

(«2)

110

460

1335

3020

BOL - Beginning of Life

3.2.9 Thermal Management of Pump and Compressor Systems

Heat from the bipropellant pumps can be rejected into the large heat

reservoir available in the pumped MMH and NTO. The MMH is used to cool the

pump. When 45,500 Ibs of bipropellant is transferred at a flow rate of 6 GPM

in seven hours, the pumped MMH temperature will increase about 3*F and the

NTO temperature will increase about 2*F. As a result of the small

temperature change of the space storable bipropellant, no further heat

rejection is required.

Heat from the battery and the power conditioner can be removed by several

heat pipes. Chemical to electrical conversion in the battery has an

efficiency of at least 85 percent. By pumping 45F500 Ibs of bipropellant and

transferring helium from 12 MPS bottles, about 850 W of heat must be removed

from the batteries and 530 W from the power conditioner (with a 81%

efficiency). Table 32 presents two types of heat pipes that are candidates

for use and the total weight of the requird heat tubes to remove 1380 watts.

Table 32. HEAT PIPE PARAMETERS

PARAMETERS MONOGROOVB HEAT PIPE AXIALLY GROOVED HEAT PIPES

Length (M) 0.75 0.55 1

Weight (Ibs) 5 3 6

Heat Transferred 215 200 150

T (evap - cond) (Fe) 20 7-2 10.8

Tube material Al Al Al

Number of grooves 1 20 27

Transport material ammonia ammonia ammonia

Required tubes to

remove 1380W 7 7 10

Total Weight (Ibs) 35 21 60

2145e/ - 136 -
Ammonia is selected as the coolant,in the aluminum heat tubes. At room

temperature ammonia has a good surface tension (20 dyne / cm), high thermal

conductivity as a gas (0.0144 BTU/hr ft °F), and a high heat of vaporiztion

(509 BTU/lbm). The aluminum heat tube is selected because it is compatible

with ammonia, it is light weight, and it has a high thermal conductivity

(118 BTU/hr ft "F).

Figures 78 and 79 present the monogroove heat pipe and the axially grooved

beat pipe. The basic monogroove design contains two large axial channels, one

for vapor and one for liquid. The small slot separating the channels creates

a high capillary pressure difference which, coupled with a minimized flow

resistance of the two separate phase channels, results in a high axial heat

flux. The evaporation and condensation film coefficients are provided by

circumferential grooves in the walls of the vapor channel. It is important

that a continuous liquid flow path exists between the primary axial groove and

the circumferential wall grooves in the evaporator sectin (otherwise dry out

will occur).

Attached Flange

Vapor Channel

Circumferential Grooves

Liquid Channel with

Wire Mesh

Axial Slot, as narrow as possible

to maximize capillary

Hi

Evaporator

Section

i f f

V

Q Out

Adiabatic

Section

Condenser

Section

Figure 78. Monogroove Heat Pipe

The monogroove operating principle is characterized by two differential

pressure relationships which must be satisfied simultaneously. The monogroove

slot must develop enough capillary rise to overcome the vapor and liquid

channel viscous losses.

^P (monogroove slot)>^P (vapor channel) +4P (liquid channel)

2145e/ - 137 -
Evaporator

Section

Adiabatic

Section

Condenser

Section

Axially Grooved Liquid Channel

Vapor Channel

Figure 79- Axially Grooved Heat Pipe

The primary relationship requires the wall wick capillary pressure rise to

offset the cumulative viscous pressure losses in the vapor channel, liquid

channel, and circumferential wall grooves.

(wall cap) (vapor channel) (liquid channel) +4P (wall wick)

The internal configuration of an axiaily grooved heat pipe consists of a

series of flow channels fabricated as a part of the tube wall and parallel to

the longitudinal axis of the heat pipe. The large open-flow channel of such a

design offers a low resistance to liquid flow. The large open-flow channels,

however, are sensitive to liquid/vapor shear interaction.

Unlike the monogroove heat pipe, the axial groove heat pipe operating

principle is characterized by one diffrential pressure relationship. The

capillary pressure difference must equal the cumulative viscous pressure

losses in the vapor channel, liquid channel, and the additional pressure drop

in the liquid phase due to the countercurrent vapor flow.

/IP (cap) =4P (vapor) + £? (liquid) + ^P (shear)

A potential advantage of the monogroove heat pipe is the anticipated 2 K¥ heat

flux. At this time, very few zero "g" heat tests have been performed on heat

pipes. Zero "g" has been simulated, for 20 sec., by flying in a parabola.

The test proved that the pipe would prime and the liquid would be drawn into

the proper channel. A quickly designed test was performed on the eighth

shuttle flight (when a payload was cancelled at the last minute). As a result

of a shortage of design time (6 weeks) liquid crystal tapes that change colors

at different temperature were used instead of installed thermocouples. The

data was comparable to predictions and a heat flux of 70 watts was

demonstrated.

Axially grooved heat-pipe technology has been developed extensively in the

last 15 years for zero "g" work. An aluminum axiaily grooved heat pipe flow

aboard the Orbiting Astronomical Observatory was still functioning after more

than four years in orbit. A total of 55 aluminum axiaily grooved heat pipes

2145e/ - 138 -
were used on the Application Technology Satellite for almost three years of

continuous flight operation. Since the technology of the axial grooved heat

pipe is further developed than the technology on the monogroove heat pipe, the

axial groove heat pipe is the suggested choice for the EHM.

There are three approaches that may be utilized to cool the compressor:

Approach A The excess heat may be absorbed in either a mass storage device

or a phase change material.

Approach B The excess heat may be absorbed by a coolant loop and then

rejected by either a pumped fluid radiator or a heat pipe

radiator.

Approach C The excess heat may be absorbed by a two-phase system and then

rejected by either a pumped fluid radiator or a heat pipe

radiator.

A short description of each of the three approaches follows.

APPROACH A; Figure 80 presents a heat transport loop using MMH as the mass

storage device. An extended QMS tank holds 853 gal. of MMH. A

2 HP compressor running 30 hours will incrase the temperature of

the extended QMS tank by 25"F. An advantage of using the MMH

tank is that no radiator is required, which also indicates a

disadvantage - you need a full MMH tank available when the

helium is compressed. This approach can only be considered a

viable approach when the helium can be transferred before the

MMH.

APPROACH B; Figure 81 presents a heat transport loop using a pumped fluid as

the coolant and either a pumped fluid radiator or a heat pipe

radiator as the heat rejection device. An advantage of a pumped

fluid system is the excellent thermal control that one has in

the heat management. Flow rates and thus heat dissipation

control is done by in-line valves. A disadvantage of a pumped

fluid system is that it contains active components (such as

pumps and valves) which will require periodic replacement,

maintenance, and redundancy components.

Figures 82 and 83 present a pumped fluid radiator and a heat pipe radiator,

respectively (both with an area of 53 F2). The fluid radiator has the same

advantages as a pumped fluid system; it is more reliable and better thermally

controlled than a heat pipe system. A disadvantage of the pumped fluid system

is if a leak occurs (due to micrometeoroid damage, etc.) complete system loss

will occur; but, a leak in a heat pipe radiator will result in the loss of

only the heat pipe that is damaged. This leaves the majority of the system

intact and functional.

Freon-21 (CHC^F) was selected as the coolant fluid. Its normal boiling and

freezing points are respectively 48*F and -211°F. The heat capacity of the

liquid at 77*F is 0.256 BTU/lbm°R. Assuming a 2HP compressor is running for

30 ;hours and the maximum temperature of the compressor and compressed He gas

is 110'F would require a flow rate of 1/2 GPM for the Freon-21 cooling loop.

Another advantage of the Freon-21 is its excellent compatibility with other

materials.

2145e/ - 139 -
Pumps-"

Compressor

Heat

Exchanger

Cold IT H He OuOutt I 4 Hot He In

Temperature Control Valve

Figure 80. Heat Transport Loop Using Mass Storage

Radiator

Compressor

Heat

Exchanger

t Cold T He Out j Hot He In

Pumps

Figure 81. Heat Transport Loop Using Radiator

2145e/ - 140 -
Manifold

Freon 21 Out

Flow Tubes

1 t

lol I

>

r

T lol

Aluminum Extrusion Honeycomb

Figure 82. Pumped Fluid Radiator

Freon

21 In

1 , , ' 1 1 ' 1

 ' ! ' ' ! i , , i i i i

i i • i • ; ' i i ' , | ' ,,1,11 1

NH3 Heat Tubes

Aluminum Extrusion Honeycomb

Freon

21 Out

Figure 83- Heat Pipe Radiator

APPROACH C; Figure 84 presents a two-phase transport cooling loop. The

primary function of an externally pumped heat pipe (EPHP) is to

transport large amounts of heat across a small temperature

gradient, thereby consuming a small amount of pumping power.

The fluid circulation rate can be reduced by at least an order

of magnitude because the heat of vaporization is much greater

than the amount of heat the fluid can absorb just warming up.

Cooling a 2-HP compressor operating for 30 hours will require a

flow rate of 1/20 of a GPM and a 5-10 watt pump. A limitation

or specific system requirement is needed by a capillary

structure (porous wicking or machined grooves) near the exit of

the radiator (condenser). Externally pumped heat pipes appear

quite promising for the transport of large heat loads over long

distances; but some analytical development, but no new

technology, will be required before implementaion of the system.

2145e/ - 141 -
Heat

Exchanger

Evaporator

Structure

Liquid

Figure 84. Heat Transport Loop Using a Two-Phase System

Ammonia is preferred over freon-21 in th EPHPs because of better surface

tension, thermal conductivity, and a higher heat of vaporization. Since

ammonia is to be used, the tube walls should be aluminum for compatibility.

Table 33 presents a weight comparison of each heat transport loop mentioned

above. Approach A (heat storage in MMH) has the lowst system weight of the

three aproaches. Approach A is also the only approach that does not have a 53

ft^ radiator. Approach B (a liquid coolant loop) haa listed both applicable

radiators. As shown, there is very little weight difference between the two

radiators, and thus very little weight difference between the complete

systems. The approach C system (a vapor/liqud coolant loop) is of lower

weight than the Approach B system because of the lower flow rates and power

requirements.

PARAMETER

Radiator (ibs)

Heat Exchanger (ibs)

Temperature Control

Valve (Ibs)

Two pumps (ibs)

Tubing (ibs)

Extra battery weight

(Ibs)

Total (Ibs)

APPROACH A

10

5

10

30

15

70

APPROACH B

Pumped Fluid, Heat Pipe

66

10

5

10

30

15

136

64

10

5

10

30

15

134

APPROACH C

Heat Pipe Radiator

64

12

5

4

30

2

117

Table 33« Weight Comparison of Heat Transport Loop

The key results of this analysis are as follows:

2145e/ - 142 -
1. No heat dissipation devices are required for the space storable

bipropellant pump. All heat will be absorbed in the transferred MMH

and NTO. The MMH will cool the pump and increase in temperature of

(~ about 3*F and the NTO temperature will increase about 2*F.

2. Heat from the battery and the power conditioner can be removed by

several ammonia heat pipes. In the case examined, the peak heat of

1380 watts can be removed by using 7 axially grooved heat pipes each

transferring 200 watts to the structure.

3- Heat removal from the compressor system will occur at two separate

areas. Heat must be removed from the compressor and the compressed

gas. At this time, the best approach is to use a freon-21 coolant

loop with an ammonia heat pipe radiator. The complete heat rejection

system has a weigh of about 135 Ibs. The externally pump heat pipe

represents a potential 13^ weight reduction over the above system but

is not at present fully developed.

3.2.10 Pressure Drop Analysis

An analysis of th'e pressure drop occurring during a resupply module refueling

of an OMV was accomplished. The resupply line schematic defined in 3*2.5 was

used to describe the resupply design. The fill/drain lines for the OMV were

taken from "OMV Alternative Systems Design Concepts" Proposal (February 14,

1985). Two different line sizes, 0.5 inch and .75 inch, were used for the

resupply lines. Pressure drops were calculated for a resupply system using

solenoid valves exclusively and then for a system using ball valves

exclusively. Losses were calculated only for the propellant transfer lines

and ullage return line. Purge and vent line losses were not examined in this

study.

To begin with, the line lengths and number of bends were estimated for the

resupply module. The resupply module was assumed to have 30 feet of line

running from the supplier tank to the propellant quick disconnect (40 feet

including lines to the propellant pumps) and a total number of line bends

equivalent to twenty 90 degree bends. Using vendor supplied data the pressure

loss across a solenoid valve (Parker P/N 5770021) for NTO and MMH flowing at

10 gpm was found. For NTO, a drop of 36 psid was found; for MMH a drop of 22

psid was found to occur across a solenoid valve. Ball valves have virtually

no flow losses and hence were assumed have 0 psi pressure drop. Data on a

Symetrics 0.5 inch quick disconnect give a pressure loss of 9«8 psid for NTO

and 6.2 psid for MMH based upon a 10 gpm flowrate.

Tanks were assumed to have a 5 psid drop from the ullage gas to the tank

outlet and flexline was assumed to have a five fold increase in pressure loss

over an equivalent length of smooth pipe. Figure 85 presents a simplified

resupply module schematic of all propellant lines and of the ullage return

line used in the ullage transfer resupply configuration.

With 1/2 inch line and solenoid valves the pressure drop from the resupply

tank through the propellant transfer disconnect was 252/163 psid for ox/fuel

(this includes a 5 psid drop encountered in the tank). The pressure drop

through the OMV's propellant transfer disconnect to the receiver tank's ullage

was 196/126 psid for ox/fuel respectively. The pressure drop in the ullage

2145e/ - 143 -
Ullage Exchange

Lines

direction

Propellant

Transfer Lines

/

L.X

,. ,,.. , .

HXh"XJ

Propellant

Pump

RESUPPLY MODULE OMV

Figure 85. Simplified Resupply Module/Receiver

Vehicle Fluid Line Schematic

return line was approximately 2 paid for 100$ helium in the line. The total

pressure loss was 4-50/291 psid for a ox/fuel system with 1/2" lines, solenoid

valves, and 100$ ullage gas in the ullage return lines. All of the pressure

losses were at a flowrate of 10 gpm.

¥ith 3/4 inch line and solenoid valves the pressure drop from the resupply

tank through the propellant transfer disconnect was 175/110 psid for ox/fuel

(this includes a 5 psid drop encountered in the tank). The pressure drop

through the OMV's propellant transfer disconnect to the receiver tank's ullage

was 136/86 psid for oz/fuel respectively. The pressure drop in the ullage

return line was approximately 1 psid for 100$ helium in the line. The total

pressure loss was 312/197 psid for a ox/fuel system with 3/4" lines, solenoid

valves, and 100$ ullage gas in the ullage return lines. All of the pressure

losses were at a flowrate of 10 gpm.

With 1/2 inch line and ball valves the pressure drop from the resupply tank

through the propellant transfer disconnect was 108/75 psid for ox/fuel (this

includes a 5 psid drop encountered in the tank). The pressure drop through

the OMV's propellant transfer disconnect to the receiver tank's ullage was

2145e/ - 144 -
88/60 psid for ox/fuel respectively. The pressure drop in the ullage return

line was approximately 2 psid for 100$ helium in the line. The total pressure

loss was 198/137 psid for a ox/fuel system with 1/2" lines, ball valves, and

100$ ullage gas in the ullage return lines. All of the pressure losses were

at a flowrate of 10 gpm.

With 3/4 inch line and ball valves the pressure drop from the resupply tank

through the propellant transfer disconnect was 31/22 psid for ox/fuel (this

include a 5 psid drop encountered in the tank). The pressure drop through the

OMV's propellant transfer disconnect to the receiver tank's ullage was 28/20

psid for ox/fuel respectively. The pressure drop in the ullage return line

was approximately 1 psid for 100$ helium in the line. The total pressure loss

was 60/43 psid for a ox/fuel system with 3/4" lines, ball valves, and 100$

ullage gas in the ullage return lines. All of the pressure losses were at a

flowrate of 10 gpm.

A breakdown of the total pressure losses by component is listed in Table 34.

As can be seen in this table, the solenoid valves dominate the pressure

losses. Line losses and bend losses are of the same order followed by small

losses in the flexlines, tanks, and quick disconnects. The total system

pressure losses are listed at the bottom of the table.

The data for the 3/4" dia. line system pressure loss is plotted in Figure 86.

This figure suggests the magnitude of pressure losses for propellant type and

valve type. The maximum achievable pressure rise that can be achieved with

near-term propellant pumps is in the range of 300 psid. This maximum

achievable pressure rise limits the achievable flowrates to between 11 and 17

GPM for a system with solenoid valves and between 26 and 30 GPM for a system

with ball valves.

At flowrates in the range of 1 to 7 gpm, adequate performance could be

obtained with a system using solenoid valves. A major advantage in using

solenoid valves rather than AC-motor operated ball valves, is in the

components weight savings. A solenoid valve weighs 2.3 Ibs, while a ball

valve weighs 4«1 Ibs. ¥ith the high number of valves required for the

operation of a remotely operated resupply module, the use of solenoid valves

rather than ball valves could mean a weight savings of as much as 250 Ibs.

An analysis of-the pressure transients experienced by the propellant transfer

system during shutdown was conducted. To carry out the analysis a simplified

propellant transfer system was simulated using an in-house fluid transient

math model. The resupply module and receiver vehicle combination was

simulated by connecting a constant pressure tank to a constant chamber

pressure engine. The engine was used to simulate the receiver tank since the

program that was used could not simulate propellant transfer between tanks.

The isolation valve located upstream of the injector (i.e., receiver tank

isolation valve) was closed at a specified rate while the engine chamber

pressure and supply tank pressures were held constant. As the isolation valve

closed a fluid transient was initiated, and the time history of the fluid

pressure immediately upstream of the isolation valve was observed. Figure 87

compares the analytical model and the simplified fluid line schematic.

2145e/ - 145 -
CO

D.

as

as

o w

C/3 O

CO J

O U.

M Oi

06 CJ

W O

CO rH

Cl3

at; H

ft. <

H

CO

CO

o

H

sr

01

to

H

a

rH S

CO 01

> JJ.

CO

rH en o

co H

pa z

0)

c

•H

01

> e

<u TJ jj

•H CO

O >>

c co

41

rH 0

O H

CO Z

.

g

01

rH S

CO 01

CO

rH >,

rH CO O

CO H

M Z

01

1-1

,J

in gj

o <u z

CO

> S

cu

•e w

•H CO

O rX

c co

01

rH 0

0 H

CO Z

±J

c

01

o

e.

E

cS

sr CM o

O r* rH rH

rH rH

sr m m

• • •

O ON rH 00

rH rH

sr CM o

O r» rH rH

rH rH

sr m in

• • •

O ON rH OO

rH rH

ON rH vO

O ON r» m

rH sr in

rH in OO

• • •

O m ON rH

rH SO ON

ON rH SO

O ON l^ m

rH sr m

rH in OO

• • •

O m ON rH

rH SO ON

CU

c

J

TJ

cu

cu

CO

CU U

CO C

CO CO CU

0 rH C

J rH -H

CU rH CO

X 0. X T3

c o cj c

!0 *J rH 4)

H a. t»- xi

sr

CN O rH

rH

sO

* C7\ O rH

rH

CN ST rH

rH in

rH

SO

•

ON CN rH

rH in

CN

CM O CN

rH

sO

ON O CN

rH

sr

CM ST CN

rH in

rH

SO

CN CN CN

rH m

CN

01 CU

c e

^4 *^4

CO iJ rH

4J

U 01 C

cu ao -H

c c

C m 6

0 f S

CJ O -H

CO X rH

•H W 01

•o *

CO <U

J«! 0) SO 3^S

•-! > 10 O

— •" H t-H O

£_ -5 —i -H

o

vO

r-.

ON

oo

ON

O

m

sr

oo

so

o

CO

CO

«—4

a

w

o

2145e/

- 146 -
M

Q.

a

o

L.

Of

3

CO

M

a>

t.

Q.

in

x

ID

+J

O

1000

ORIGINAL PAGE IS

OF POOR QUALITY

10

s ( » i i 10

FLowrate (gpm)

i ; i « 100

Figure 86. System Pressure Loss Vs. Flowrate

Figure 88 shows the maximum surge pressure as a function of initial flowrate.

With a valve closing time of 8 msec, the surge pressure slope is 83 psid/GPM.

While these values are preliminary, they do suggest the magnitude of the

transients that can be expected in the resupply system. During a pumped

propellant transfer, the shutdown transients can be minimized by tapering the

2145e/ - 147 -
pumping rate down to a low value 'and'^hen shutting the system down. However,

if the propellant transfer operations-'had to be terminated immediately an

isolation valve would be closed. The sudden closure of a valve would initiate

a pressure transient with a magnitude similar to the values predicted above,

and the fluid system would have to be designed to withstand fluid transients

of this magnitude.

Prom this analysis, it can be concluded that solenoid valves can be used in

the resupply module for fluid management. Using a pump with a 300 psid

pressure rise in a resupply module that used solenoid valves would yield

maximum flowrates in the range of 10 to 15 gpm. Most fluid transfer

operations will have maximum flowrates limited by propellant management

devices and not by resupply system pressure losses. Fluid transients during

shutdown were examined, and it was observed that for worst case conditions

(high flowrate, fast acting valve) a peak surge pressure rise of up to 83

psid/GPM above steady state conditions was obtainable.

3.2.11 Propellant Management Considerations

The selection of propellant tanks is an important step in the design of any

low-g propellant handling system. In many cases the tanks constrain the

vehicle's performance capabilities, including constraining the number of

attitude control thrusters allowed to fire, the duration of thruster burns,

minimum time between burns and allowable maneuvers along some axes. The tanks

will also define the vehicle's capability to either resupply or to be

resupplied under low-g or zero-g acceleration fields. The selection tanks

also helps determine the design of the rest of the system, including the need

for hazardous waste scavenging system, liquid or gas pumps, ullage transfer

systems, etc.

Propellant tank and propellant acquisition/management device requirements for

tanks operating as either an OMV extended kit or propellant resupply module

were determined and are summarized in Table 35. State-of-the-art proipellant

tank concepts were then investigated to determine their potential for meeting

the identified requirements. Finally, current tanks representative of each of

the concepts were researched to determine estimate of the capabilities of each

concept.

Three propellant acquisition concepts were investigated: elastomeric

diaphragms, surface tension vane devices, and surface tension screen

acquisition devices. A summary of each type's advantages and disadvantages is

presented in Table 36. It was determined that the concept best suited to the

needs identified in this study, including operating capabilities and the

capability of being resupplied on-orbit, was elastomeric diaphragms. However,

no suitable diaphragm has been flight qualified for multiple cycles when used

with NTO. It was, therefore, recommended that the developments in this area

be closely monitored. Of the remaining two concepts, vane device tanks have

the very desirable capability of controlling the position of the ullage bubble

in zero-g. This capability allows the venting and transfer of the ullage to

the resupply tank during propellant transfers, eliminating the need for

hazardous waste scavenging systems in the extended mission kit or resupply

module. Vane devices may not be suitable for operation in acceleration fields

such as will be required of OMV and OMV extended mission kit tanks.

Satisfactory performance may be achieved by placing a propellant retention

sponge type device over the tank outlet, although this would require

2145e/ - 148 -
Analytical Model Schematic

Fluid Line Schematic

constant pressure

supply tank

constant pressure

supply tank

line 1

line 2

nr-ifirm ^

line 3

line 4

Iso. valve ]

f~

7.

i

i

7.

7.

i

t

7.

5 ft

r

i

5 ft

t

k

5 ft

r

k

5 ft

• 1

A

Note;

the orifice 1s sized

to simulate line losses,

bend losses, valve losses,

and QD losses.

Iso valves

30 ft, 15x90 deg

bends

chamber pressure 1s

constant while the

the 1so valve closes

(simulates receiver

tank)

\/

constant pressure

receiver tank

constant chamber

pressure engine

QO

1so valves

Figure 87. Analytic and Fluid Line Schematic Used

in Fluid Transient Analysis

( 0.5 inch line diameter )

2145e/ - 149 -
1400

1200

1000

42 800

Wl

a.

•

t

S 600

2

o.

<u

o

400

200

1 Reference*

|Flu1d: NTO -;

_ instantaneous r •

valve closing

_j g msec valve -

._... f _ ... _ closing . j. ..., . __ _,

40 msec valve

j closing ( ... f

steady state

system pressure

246 8

Initial Flowrate, 6PM

10

Figure 88. Peak Surge Pressure Vs. Initial Propellant Flowrate

constraints be placed on OMV maneuvers, such asduration of maneuvers producing

adverse accelerations, and time between maneuvers to allow the sponge to

refill. Surface tension screen acquisition device tanks perform very well

under the environments required of an OMV or extended mission kit tank. The

problem if there is no known screen device tank which has the capability of

positioning the ullage bubble. It is felt however that a screen tank can be

modified to obtain this capability.

2145e/ - 150 -
Table 35. RESUPPLY MODULES AND OMV EXTENDED

MISSION KITS TANK REQUIREMENTS

SUMMARY

o Acceleration Level

o +X

o Omni directional

o Steady Coast Resupply

o Expulsion Rate

o +X

o Omnidirectional

o Steady Coast Resupply

o Propellant Temperature

o Operating Pressure

o Volume

o Vibration

o Shock

o Expulsion Cycles

o Pressure Cycles

Extended

Mission

Kit

(Bi-Propellants)

0.16 g

0.0048 g

1 x 10-?g

1.39 Ib/sec NTO.0/85 Ib/sec MMH

0.048 Ib/sec NOT,0.028 Ib/sec MMH

1.0 Ib/sec NOT,0.6 Ib/sec MMH

40 - 100° P.

257 psia - TBD

200,448 in?*

100 Missions Shuttle

Launch & Ascent

Shuttle Crash

Landing Loads

100

200

Resupply

Modules

N/A

N/A

1 x 10~7g

N/A

N/A

0.6 Ib/sec

40 - 100° F.

TBD

N/A

100 Missions

Shuttle Launch

& Ascent

Shuttle Crash

Landing Loads

100

200

TABLE 36. Summary of Propellant Acquisition Concepts

Advantages

Disadvantage

Examples of Users

DIAPHRAGM

o Positive Ullage Positioning o

o Can be resupplled with or

without venting.

o Operates over wide range of

acceleration levels.

o No potential for gas

ingestlon.

o High Expulsion Efficiency

o Diaphragm Life Limits

o Not applicable to NTO.

o Shuttle APU

o TDRSS

VANE DEVICE

Positive Ullage

Positioning

o Can be resupplied

with or without

venting.

SCREEN DEVICE

o Operates over wide

range of acceleration

levels.

o Few Operational

Constraints

o Does not require

propellant to be in

specific location.

o Operates only under o No ullage positioning.

extremely low

accelerations. o Must vent prior to

refilling,

o Must allow time for

propellant to resettle

between maneuvers.

o Viking Orbiter

o INSAT

o ARABSAT

o Shuttle RCS

o Shuttle OMS

o INTELSAT

2145e/ - 151 -
Before an examination of propellant tanks and propellant acquisition device

concepts can be made, it is first necessary to estimate the propellant tanks

operating environment and performance requirements. One of the most important

parameters is the acceleration field under which the tank must supply

propellants. The Rockwell OMV design consists of 6 main engines, 110 Ibf

thrust each, all firing in the +X direction (+X translation), four 5 Ibf

thrust RCS engines in both the plus and minus X directions, and two 5 Ibf

thrust RCS engines in each of the +Y and +Z directions. The approximate dry

weight of the OMV is 4200 Ibs., including an estimate for residual

propellants. Since sizes and weights of resupply modules and OMV extended

mission kits have not been finalized to date, these weights will not be

included in this analysis. With one of these kits attached to the OMV, it can

be expected that the accelerations will be 25 to 50% less. The maximum

expected linear acceleration then is along the x-axis and is approximately:

max

•( 6 • 110 Ibf •*• 4 * 5 Ibf ) 32.174 Ibm ft ,5m2t* = 0.16 g's

4200 Ibm Ibf s2 S2

Any tank selected would be installed along this axis so that this acceleration

accentuates the greatest capability of the tank. The largest adverse

acceleration would then be in the -X direction due to the 4 RCS engines and

would be approximately:

Aadverse = (4*5 Ibf) * 32.174 ltm ft = 0.15 ft = 0.0048 g's

max ———— 4200 Ib—m —^ Ibf s—2 — s2—

Any translation in any other direction would impose an acceleration of

0.075ft/s2 (0.0024 g's) corresponding to 2 RCS-engines firing. The maximum

rotational rate about any axis is estimated to be 0.052 rad/sec (3°/sec).

The radius of rotation (i.e. OMV - extended mission kit e.g. to the center of

the tanks) is estimated to be 5 feet for x-axis rotations (i.e. roll) and 10

feet for pitch and yaw. This yields centripetal accelerations (rw2) of

0.013, 0.027, and 0.027 ft/sec2 for rotations about the roll, putch and yaw

axes. Since it is expected that pitch and yaw rotations will not be sustained

to settle propellant, especially in conjunction with the worst case

translational accelerations, and since the translational acceleration

requirements are much more severe, pitch and yaw maneuvers need not now be

considered. Likewise, assuming roll maneuvers will not be performed in

conjunction with translations (except possible along the x axis), and since

the translational requirements are again much more severe (0.15 or 0.075

ft/s2 vs. 0.013 ft/s2) roll maneuvers will also not be considered at this

time.

Another important design requirement is the tank's expulsion rate. This

requirement can be determined by either the propellant flow rate required by

the thrusters of the time available to resupply either the OMV or satellites.

The propellant flow rate to the OMV's six main thrusters (Igp=312 sec)

simultaneously is:

(6 * 110) Ib

= 2.12

312 sec

2145e/ - 152 -
at a mixture ratio of NTO to MMH of 1.65. It is also desired to fire up to

six 5 Ibf RCS engines (Isp = 260 sec) in conjunction with the main engines

resulting in an additional flow of:

(6 * 5) lb

260 sec

or a total flow of 2.24 Ibs/sec (1.39 Ibs/sec NTO, 0.85 Ibs/sec MMH). Again

it is assumed that the tanks will be installed such that the settling

acceleration caused by the main engines requiring this flow rate causes the

propellant to settle in the most advantageous location in the tank (i.e.

probably over the tank outlet). The worst case flow rate required under an

adverse acceleration would be that required to support either 2 or 4 RCS

engines, or 0.038 (0.024 Ibs/sec NTO, 0.014 Ibs/sec MMH) or 0.076

(0.048 Ibs/sec NTO, 0.028 Ibs/sec MMH) respectively.

Propellant temperature is another important parameter in the performance of

any surface tension driven propellant acquisition device. For this study it

will be assumed that the temperature will be controlled to between 40 and

100°F.

The OMV extended mission kit will be required to supply propellants to the

OMV. The operating pressure of the OMV is planned to be 257 psia, the same as

the Space Shuttle QMS regulated pressure. It may be desired, however, to have

different operating pressures for extended mission kits and resupply modules

depending on the type of propellant transfer which is selected. This will be

determined later and for now will not constrain the selection of any tank.

Since the OMV, SPER modules, and extended mission kits will be launched and

returned to earth in the Space Shuttle, They should each be capable of

sustaining a maximum of 100 missions worth of launch, ascent and random

vibration. Also, any propellant tank which flies on the Orbiter is required

to be certified to be able to withstand crash shock loads. Various tanks,

including the Shuttle OMS and RCS screen acquisition device propellant tanks

and the Shuttle APU diaphragm tanks, have been certified to these

requirements. It will therefore be assumed that these requirements will not

preclude the use of any tank, although it should be noted that they may

require that off-the-shelf tanks be redesigned and requalified. Additionally,

it was assumed that each tank will be required to be capable of the equivalent

of a maximum of 200 expulsion cycles and 200 pressure cycles.

Three low-g propellant tank concepts exist which have the potential of meeting

the above requirements. These include elastomeric diaphragms, surface tension

screen acquisition devices, and surface tension vane devices. Each of these

concepts will be discussed in the following sections with an attempt made to

identify some of the advantages of each type. Elastomeric expulsion devices

have received wide use throughout the industry in mono-propellant (hydrazine)

systems. To date, although much effort is being expended, no diaphragm has

been qualified for repeatable service with NTO. Examples of operating systems

currently employing elastomeric diaphragms include the tracking and data relay

satellites (TDRS) and the Space Shuttle Auxiliary Power Units (APU). The use

of these diaphragms in the Shuttle APU tanks has demonstrated the ability of

the tanks to meet the requirements of Table 35 for mono-propellant resupply

module, specifically the vibration, shock, pressure cycles, expulsion cycles

and expulsion rate requirements.

2145e/ - 153 -
The major advantage of employing elaatometric diaphragm tanks in a resupply

module is the ease with which these tanks can be reaupplied in low "g"

environment. A system using these tanks can operate either in a blowdown mode

or in a pressure regulated system. In both cases, either recompressing the

ullage or venting the ullage during a refill is comparatively easy. Other

advantages of elastomeric diaphragm tanks include: typically exhibiting

expulsion capabilities of 99$ and greater; the independence of expulsion on

acceleration levels; the independence of performance on on-orbit propellant

sloshing; and the improbability of pressurant gas ingestion to the thrusters.

The major disadvantage to elastomeric diaphragm tanks is the inability to

qualify a diaphragm compatible with NTO. There is a good possibility,

however, that a NTO diaphragm could be developed in the near future. Work in

this area should be closely monitored. If a diaphragm is developed which

meets all of the requirements of Table 35» it would be a leading candidate for

any tank system developed in the future.

A second propellant tank concept capable of managing propellant in a low "g"

environment utilizes surface tension vane devices. Vane devices, manufactured

of sheet metal, rely on the very small surface tension forces of the

propellant to orient the propellant over the outlet. A typical vane device is

shown in Figure 89* (This device was developed for the Viking 75 Orbiter.)

Figure 89- Viking Orbiter Baseline Propellant Management Device

2286e/ - 154 -
The design of the vanes, including their height, number and—shape (i.e.,

distance between the tank wall and the vane) determine the forces acting on an

off-center ullage bubble and the stable location of the ullage. Because the

surface tension forces are so low, these tanks are designed to operate under

an extremely low acceleration field, corresponding to a bond number (ratio of

acceleration forces to surface tension forces) of approximately 1, or in an

acceleration field which causes the bulk propellant to settle over the

outlet. For a 36-inch spherical or cylindrical tak, a bond number of 1

corresponds to an acceleration of approximately 1 x 10~5 g's. Because of

this, the use of these tanks is normally confined to providing propellant

under zero "g" to start an engine which will then provide thrust to settle the

propellant over the outlet. The capability of a vane device can be enhanced

by designing a propellant retention honeycomb type sponge over the tank

outlet. If this sponge is designed large enough to supply propellant during

any adverse acceleration and then allowed to refill before any subsequent

thruster firings, satisfactory tank performance could be obtained.

In many cases this is impractical as the time necessary to resettle the

propellant becomes unacceptable. If the propellant becomes reoriented such

that it loses contact with the vane device, surface tension forces will be

unable to act to resettle the propellant over the tank outlet. For this

situation a communication channel is designed into the tank, extending from

the top of the tank to the bottom. For propellants with approximately a zero

contact angle (angle formed at the intersection of the propellant interface

and a solid surface, measured within the propellant) such as NTO and MMH,

their surface tension will force propellant into the channel and then wick

along the channel to the tank's bottom. In near zero "g" this can take

several hours. This can be acceptable in a system such as a satellite when

thruster burns are required infrequently, but would be unacceptable to a

maneuvering vehicle such as an OMV or Space Shuttle. For propellants with a

high contact angle such as hydrazine, the performance of a communication

channel is very much degraded. The alternative would be to perform a settling

burn of approximately 10~3 g's, which could resettle the propellant in about

30 seconds.

The vane device used in the V075 tanks was certified to an expulsion rate

capability of 0.64 Ib/sec NTO, 0-42 Ib/sec MMH, and 1.4 Ib/sec N2H4. It

is expected that these rates could be greater under a settling acceleration.

The only operating advantage of a vane device tank over a surface tension

screen acquisition device tank is its ability to control the location of the

ullage bubble. This is necessary if it is desired to vent the tank such as

could be required in an emergency situation or during a resupply of the tank

on-orbit. Low "g" reloading of tanks of this nature has been experimentally

shown possible as long as the inertial forces of the incoming propellant do

not exceed its surface tension forces. The dimensionless Weber number is used

_to define this critical range. The Weber number is defined as the ratio of

the inertia forces to the surface tension forces by:

2286e/ - 155 -
Where: Q - Volumetric Flow Rate

R » Tank Inlet Radius

V » Incoming Liquid Velocity

Q » Kinematic Surface Tension

For the V075 vane tank, it has been shown experimentally through drop tower

testing that the propellant-gas interface remains stable (i.e., reloading is

possible) for Weber number less than 7. This traslates into volumetric flow

rates of 2.81 in3/sec NTO, 4.13 in3/sec MMH, or 5.37 in3/sec N21?4.

For the V075 tank (43,000 in3 volume) this means a resupply of 4.25 hrs.,

2.89 hrs. and 2.22 hrs. for NTO, MMH, and N2H2, respectively. If this

design was to be revised to Incorporate baffles over the tank inlet this time

could be reduced by at least a factor of 2.

The major disadvantage of vane device propellant tanks is, as was discussed

previously, the inability of the propellant management device to maintain

propellant over the tank outlet under adverse accelerations greater than

approximately 10~5 g's. This capability can be enhanced by incorporating a

propellant retention sponge over the tank outlet although this will

necessitate placing maneuvering constraints on the duration of adverse

accelerations and time between maneuvers to allow propellant to resettle if a

tank of this type is used on an OMV or an OMV extended mission kit. However,

a vane device tank would work very well in a resupply module as long as the

propellant resupply occurs in near zero gravity.

Another problem exists while outflowing vane device tanks in low "g"

environments. That is the tanks tend to outflow the propellant nearest the

tank outlet resulting in premature ingestion of gas while a significant amount

of propellant remain in the tanks. For propellants with low contact angles

(NTO and MMH) which have a low "g" configuration of approximately a column

over the outlet in the PMD, outflow in low "g" causes a "necking" of this

column at the outlet, allowing gas to enter the outlet. Drop tower tests have

shown that outflow in this condition will result in two-phase flow with a

series of bubbles with a diameter the size of the outlet line. For

propellants with a high contact angle (hydrazine) which have a low "g"

configuration of approximately a flat interface with propellant settled over

the outlet, outflow will cause "suction dip", or a pull through of gas through

the propellant. Either case can result in approximately a 10% residual

trapped in the tank. For the Viking Orbiter missions, however, with a

relatively high-g (O.lg) settling propellant over the outlet, the expulsion

efficiency of the tanks was as near to 100% as the on-board instrumentation

could measure. Other problems relating to premature gas ingestion have been

experienced using vane device tanks. One problem was caused by sloshing of

propellants upon engine start-up which causes propellants to leave the outlet

area. For the Viking Orbiter missions this gas ingestion was estimted to be

20 in3 per engine start.

The ability of vane devices to withstand the vibration and shock requirements

is now know. During vibration testing of the V075 tanks "several baffle

elements experienced localized yielding as a result of three corner folds

induced by the slosh motion." Assuming the Viking Orbiter One Mission

requirements were not as severe as 100 missions of Shuttle launch vibration, a

serious problem could exist. Further, the V075 tests were conducted with

water. It was discovered during the Shuttle RCS Tank Certification Program

that NTO propellant loads during vibration can be an order of magnitude

greater than those in water. Also, it may be impossible to support

2286e/ - 156 -
the vanes with stiffeners or struts without impacting the tanks' ability to

move and position the ullage bubble, therefore, much attention should be

given to this area before a vane device tank is chosen for the SRM, OMV, or

OMV extended mission kit.

Although there appears at this time to be several serious disadvantages to

using vane device tanks for the applications of interest to this study, the

advantage of being capable of positioning the ullage may be of such benefit

that these type tanks should receive thorough study. Since there is an

infinite number of possible vane designs, and since each vane design is chosen

for a specific set of requirements, it is possible that an acceptable and

effective vane design can be found for these applications.

Surface tension screen acquisition device propellant tanks again rely on the

very low surface tension forces of the propellant to separate the pressurant

gas from the propellant. In this case, barriers and flow channels made of

very fine screens (pores measured in microns) allow passage of propellants but

because of the surface tension forces, restrict the flow of gas. Propellant

then flows beneath the screen barriers or through the channels to the tank

outlet. Because of the very small pore size, a screen device tank is capable

of operating in much higher acceleration fields than are the vane tanks.

There are an unlimited number of possible screen device designs, varying in

complexity from a simple start basket over the outlet to a total low "g"

propellant acquisition system. The Shuttle Orbiter QMS tank is designed to

maintain propellant over the tank outlet during adverse accelerations such as

would be caused by RCS Maneuvers, and then to provide propellant for QMS

engine start, which then provides thrust to settle the propellant. To do

this, the Shuttle Orbiter QMS tank propellant management device consists of a

screen bulkhead isolating the outlet end of the tank which contains a screen

gallery system. The screen bulkhead is designed to provide a barrier to hold

propellants, sufficient for approximately 18 QMS engine starts, at the AFT end

of the tank. The QMS tank propellant management system configuration is shown

in Figure 90.

The greatest advantage of screen device tanks is their capability to operate

in a wide range of acceleration levels, from essentially zero-g to greater

than 0.07 g's. They can be designed to operate with propellant settled in a

specified location due to engine thrust or under omni-directional

accelerations which may settle propellant in any location in the tank. This

makes screen tanks free from many operational constraints. There is also

minimum time requirement between burns to allow propellant to settle or

reorient. The Shuttle Orbiter OMS tank has shown some capability during past

Shuttle flights for extremely low outflow rates (0.056 Ib/sec NTO, 0.035

Ib/sec MMH) in zero "g". With a settling acceperation it outflows 12 Ib/sec

NTO and 7.3 Ib/sec MMH to the OMS engine.

The disadvantage of screen device tanks is their inability to position the

ullage bubble in the tank. Without this ability, it is impossible to vent the

ullage from the tanks while propellant is being resupplied to them. This

makes it necessary to vent the tanks, both ullage and remaining propellant, to

near vacuum before reloading, and then vacuum loading. This venting of ullage

gases and propellant could necessitate the incorporation of a hazardous waste

scavenging system in any resupply module or extended mission kit. It may be

possible to eliminate this problem by designing a stand pipe in the tank which

would locate the ullage in a desirable position under near zero gravity.

Using a stand pipe, in conjunction with a liquid/vapor separating device,

2286s/ - 157 -
AFT WGtW PROBE GAILERT ASSY

MAIN PEED DISCHARGE

COHUNICATTON SCXCEN

•PUCSSUtt INLET

Figure 90. Shuttle QMS Tank Screen Configuration

could allow the tank to be vented during propellant loading. Although there

is no known tank of this type available now, it may be a rather simple

modification to existing tanks. This should be thoroughly investigated. If a

screen tank can be made with this capability it would have many advantages

over a vane device tank.

A matrix of resupply mission scenarios and propellant acquisitions devices are

arranged in Tabl« 37. The worst case resupply scenario (requiring ullage

positioning and operational capability under varying acceleration levels),

dictates that a propellant acquisition device integrating a vane arrangement

and a capillary screen system would be necessary.

3«2.12 Propellant Quantity Gauging and System Health Monitoring

Tank quantity gauging and flowmeters are the two principal means of

determining the quantity of propellant transferred during refueling. Although

tank gaging does not directly measure the transferred amount (but infers it

from tank quantity), it has the advantage of directly verifying total tank

load, which is important for mission confidence. A flowmeter, by itself,

cannot do this unless tank load is accurately known prior to transfer. In the

case of spacecraft returning from its mission, this is not known unless a tank

gauging system is provided, or the spacecraft tanks are off-loaded to zero

quantity before transfer.

2286e/ - 158 -
Positive displacement integrating fluid meters similar to commercial water

meters or gas meters are not considered acceptable due to their

close-tolerance moving parts and a general lack of development for

withstanding launch vibrational environments. Turbine meters have been

developed for rocket engine application and would be acceptable in terms of

cost, accuracy and reliability. Consideration should also be given to the

solid state type meters such as vortex-shedding flowmeters and swirl meters.

The former consists simply of a stationary rod or bar placed across the

diameter of the flow stream such that trailing edge (Von Karman) vortices are

generated in the downstreams flow which can be detected by presssure sensors

in the wall of the duct. The quantity of flow per vortex is repeatable over a

15:1 range of flowrate. Similar performance is provided by the swirlmeter

(Fischer & Porter) design in which an axial vortex is generated by stationary

radial vanes and caused to precess downstream by a sudden enlargement of the

flow duct. These vortex precessions are sensed by a pressure pickup and

electronically counted as an indication of total flow.

Spaceborn system accuracies realistically achievable by the above flowmeter

types are +0.1$ positive displacement meters, + 0.50$ for small turbine meters

and +.75$ for the shedding-vortex and swirlmeter types. These values are

stated as a percentage of volumetric flow rate over a 10:1 flow range; with

the benefit of calibration vs. flowrate, temperature and pressure.

The principal zero "g" tank gaging methods are listed as follows:

o PVT (Pressure, Volume, Temperature) - Ullage volume is inferred from

the perfect gas law, PV=MRT, where the mass of gas (M) and the gas

constant (R) are known, and the gas temperature (T) and pressure (P)

are directly measured. Propellent quantity is obtained by

subtracting the inferred ullage volume from the total tank volume.

o RF (radio frequency) - An antenna mounted in the tank is used to

determine the number of resonant RF frequencies inside the tank

enclosure as an index of propellant mass (Bendix).

o Acoustic Resonance - An acoustic transponder is used to determine the

fundamental resonant frequency of the ullage bubble as an indicator

of ullage volume and, hence, propellant quantity. For accurate

results, this ullage bubble should be positioned at the transponder

by means of a diaphragm, capillary device, or a slight settling

thrust.

o Nuclear Gaging - This technique uses a number of radioactive sources

(typ-infilly gamma ray) mounted on the outside of the tank, and infers

the mass of propellant in the tank by the amount of radiation

absorbed.

o Capacitance Probe Gaging - The Shuttle QMS tanks use full length

capacitance electrodes to sense the level of propellant. This is not

a true zero "g" system since appreciable settling thrust, is required

to prevent propellant from filling the narrow space between the

electrodes by capillary action, and giving a false "full" signal.

Furnishing this amount of settling force, by either engine thrust or

vehicle spinning, may not be acceptable.

2286e/ - 159 -
Table 37. Resupply Module Design Reference Missions Scenarios Vs. PMD

Prop Acq Devices

Screen Vanes

1. ERM serves as an OMV extended mission kit to get req'd req'd

the OMV to GEO with sufficient propellant to do

5 or 6 resupplies of GEO satellites (propellant

resupply only).

- The OMV will provide the resupply

2. ERM serves as an OMV extended mission kit to req'd req'd

place the OMV + Servicer + Helium Servicer into

GEO to supply and service 1 GEO platform to

extend its life by 5 years. This includes

helium servicing.

- The OMV will provide the resupply

3. The ERM with the OMV as the propulsive stage req'd

provides top-off to an off-loaded, launched

integral propulsion satellite requiring 30K

Ibs of propellant. The ERM + OMV take the

propellant from a propellant depot at the Space

Station and resupply the integrated propulsion

system satellite.

4. The ERM serves as the propellant depot attached req'd

to the Space Station. It accepts propellant

from the orbiter by orbiter bi-propellant

scavenging. It refuels other ERM's or OMV's or

both.

5. The ERM with the OMV as the propulsive stage req'd

refuels a GEO kick stage.

6. The ERM + OMV load another ERM + OMV (which will req'd

do scenario 1) which have been launched offloaded in LEO.

7. The ERM + OMV load or top-off another ERM + OMV req'd

(which will do Scenario 2) which have been

launched off-loaded in LEO.

8. The ERM + OMV serve as a refuelable/reusable req'd

periges kick stage. The ERM + OMV are loaded

in LEO and provide the kick for GEO, then use

aerobrake to return to LEO to be refueled to do

another GEO kick.

2286e/ - 160 -
Of these techniques,-PVT is the only zero "g" method wich has reached an

acceptable state of hardware development and flight experience. (Apollo, GPS,

Viking and Orbiter, etc.)- In addition, the RF and nuclear systems are quite

complex, with the latter requiring special care in the handling of radioactive

materials.

The tank gaging achievable by PVT methods can be as good as 0.5% of full load,

but more typically is around 3-5-5$, depending chiefly on dynamic temperature

sensing errors and the accuracy of tank calibration. The principal source of

error is the practical difficulty of sensing the true bulk temperature of the

ullage gas under transient thermal conditions, which result from changing

ambient conditions, or from isentropic compression or expansion of the ullage.

Another source of error in the PVT Method is solubility of the pressurizing

gas in propellents, and its gradual loss through the engines. A nominal

correction can be made for this effect in the PVT calculations.

The PVT system developed in the early 1970's and currently in use gaging the

Shuttle Orbiter RCS propellant quantities has a worst case analytical gaging

accuracy of 5«4$. This system includes two pressure transducers on both he

pressurant tank and the propellant tanks with repeatability accuracy of 1.4$

(+0.5$ tansducer, +0.6$ noise and impedence, +1.00$ signal conditioner and

+0.5$ multiplexer demultiplexer encoding accuracy) and one temperature probe

in the pressurant tank and one temperature sensor on the propellant tank, each

with 1.5$ repeatability accuracy (+0.75$ transducer). Shuttle Orbiter flight

experience has shown this systems actual gaging accuracy to be approximately

J.5$« This is a relatively inexpensive system using off-the-shelf

transducers. Better accuracy can be gained although the price will begin to

increase rapidly for the more accurate transducer. Approximately a 1$

increase in accuracy can be gained without much increase in cost if the

temperature sensor on the propellant tank skin is replaced with a temperature

probe into the tank. This is not possible with diaphragm tanks and probably

not desired with tanks with ullage positioning capability. Better accuracy

can also be gained by adding additional redundant pressure and temperature

transducers. It is believed, however, that a system similar to the Shuttle

Orbiters RCS will be acceptable for resupply module applications. It will be

necessary, however, to add redundant temperature transducers on the pressurant

and propellant taks.

To measure propellant transferred quantities, a redundant integrating type

flowmeter system should be incorporated into the propellant transfer lines. A

turbine type of flowmeter system will probably be the least expensive in terms

of overall cost and can provide good reliability, accuracy, flow rangeability,

and the ability to indicate reverse flow in the event of off-loading the

receiver tanks. The requirement to off-load could occur as a result of

inadvertent overfilling of the receiver tanks or as a prelude to refueling the

receiver tanks. The guard against flowmeter failure or disagreement between

flowmeters, and since PVT may not be available as backup, it is desirable to

include three flowmeters in series. The three can be averaged or a 2 out of 3

voting^network can be utilized to determine actual propellant quantities

transferred. The overall loading accuracy of this system is estimated to be

approximately 1$. It will be necessary to place controlling valves and

orifices downstream of the flowmeters to prevent propellant outgassing near

the flowmeters which will significantly affect their accuracy.

2286e/ - 161 -
System Health Monitoring system health status should he computer monitored

during all phases of the ERM systems orbital operation. Pressures and

temperatures of components, pressurant and propellant are verified to be

within operational limits. Corrective action for environmental control is

also available through system heaters. All temperature and pressure sensors

incorporated in the resupply system must be operable during the general health

monitoring mode.

During resupply operations, critical temperature and pressure measurements

will be programmed by the computer for fluids and components of the resupply

system to assure the resupply operation selected will not cause system

opration outside established tolerance limits.

Fault isolation requires pressure and temperature sensors, placed at

appropriate resupply system locations, to be operative. These may be the same

sensors (or redundant to) used to monitor system health. These sensors have

been positioned on (in) pressurant tanks, propellant tanks, upstream and

downstream of pressurant and propellant isolation valves and on (in) all

transfer quick disconnects. These sensors provide the capability to detect

leakage, fault isolation and safe system operation through fault isolation.

Using state-of-the-art microprocessors to integrate these pressure and

temperature parameters, the system status is processed on-board. The status

presents system health, fault isolation, and safe system configuration after

fault isolation.

A computer provides control to the transfer of pressurants and propellants

during resupply, assuring system limits are not violated. Computer control of

the systems health and safety configuration is presently achievable as is the

capability of the on-board computer to work in harmony with the computer

controlling the vehicle being resupplied.

In summary, the PVT method was the only state-of-the-art zero "g" tank

quantity gaging system identified. Also, an integrating flowmeter was the

only method available to measure propellant transferred if a quantity gaging

system was inoperative (i.e., during an ullage transfer type of resupply). A

computer controlled Health Monitoring system should also be incorporated into

the ERM to assess the system performance at all times.

3.2.13 Quick Disconnect Selection

The difference between the pressure balanced versus non pressure balanced

fluid disconnects lies in the interface panel separation loads. Quick

disconnects coupling designs utilizing pressure balanced fluid disconnects

yield low interface panel separation loads. Therefore, latches may not be

reqired and the end effector-grapple fixture mechanism may provide all of the

connecting force requirements. For fluid disconnects that are non or

partially pressure balanced the flow rates and pressures encountered by the

disconnects, during fluid transfer operations, may produce sufficient forces

to require latching mechanisms to secure the interface panels. ^

Representative propellant disconnect couplines are shown in Figures 91, 92,

93> 94, and 95. Figure 91 is a partial pressure pressure balanced

configuration. This coupling has been designed with the entrance and exit

fluid force to be almost counter-balanced. The design shown is a

NASA-sponsored Fairchild design which features low connection forces and very

low dribble volume upon disconnection. Figure 92 represents a design of a

2286e/ - 162 -
currently existing pressure balanced design by Purolator and has essentially

zero pressure connection forces. The dribble volume may be greater than the

NASA design but is expected to be low. In both designs figures 91 and 92,

venting the front face of the male poppet to vacuum or local ambient is

reuired. Figure 93 presents & conventional quick disconnect doupling with

high connection forces and potentially excessive dribble volume. Figures 94

and 95 present existing quick disconnect couplings with high connection forces

and small dribble volumes.

NASA-JSC is currently in the procurement cycle to develop a quick disconnect

for manual and remote actuation operation. The coupling design concept, as

specified by the NASA RFP, should exhibit versatility of use by a variety of

different spacecraft and fluids. In addition, the coupling design should be

mechanically simple, highly reliable, light weight, easy to engage and

disengage, and readily adaptable to automatic operation. The quick disconnect

which will be selected for the resupply module will be a derivation of this

NASA design and development contract.

-— 2.2S — »- .75 ••— -

SPHERICAL ALIGNMEN T BEARIN6-

Figure 91. Partial Pressure Balanced Coupling

2286e/ - 163 -
• WEI6HTAPPROX. 8LBS.

• OPERATING TEMP 4P*F-'OCT (toilT w KBWW El£)

• LOAD 20LPS. TO COMPRESS

• CN&AGEMENT 1/4 INCH

• LIMIT ALIGNMENT OF FACES S

FEMALE HALF

1.5 DM

Figure 92. 1/2 Inch Pressure Balanced Coupling

(Balanced Spring Loclcup and Release)

2286e/ - 164 -
ORIGINAL PAGE 6S

OF POOR QUALrTY

AAAX.

•J.572 L//O:;FLLC!r.\ST

±.053

Figure 93- Conventional Q/D Coupling

2286e/ - 165 -
VMVl SUlVt

romi

SIMS

nuo

ICCfcfl

Figure 94. Tube Sleeve and Poppet Quick Disconnect Coupling

2286e/ - 166 -
ORIGINAL PAGE IS

OF POOR QUALITY

POPPET SLEEVE AND POPPET QUICK-DISCONNECT COUPLINGS

AOCJitl HUG

,«CIIVMING tINC

IOCHNG Mil

Figure 95- Poppet Sleeve and Poppet Quick-Disconnect Coupling

2286e/ - 167 -
3.2.14 Avionics and Power Subsystems Definition

The ERM is baselined to operate with a smart host vehicle. Where the OMV acts

as the host vehicle an extended capability OMV is baselined. This OMV is

expected provide the following functions which drive ERM system requirements:

o GN&C

o Communication and tracking

o Vehicle docking

o Primary power, supply (quiescent mode)

o Primary data processing

These interfaces are derived from the ERM mission scenarios. The prime

mission scenario driving these interface requirements is the ERM operating as

an OMV extended mission kit enabling the OMV to be placed in GEO where, the

OMV alone would provide the resupply and/or servicing operations. The second

major scenario is the ERM acting as the Space Station propellant depot

accepting Orbiter scavenged propellant.

GUIDANCE, NAVIGATIONS CONTROL

The GN&C activities are primarily an OMV function. This is due to the remote

servicing/resupply operation requiring the OMV as the carrier vehicle.

DOCKING

The docking sybsystem is also provided by the OMV although some ERM interfaces

will be required. The present configuration mounts the ERM to the forward

structure of the OMV as shown in Figure 4. This requires the docking

mechanism, an RMS end effector, to be mounted on the ERM to facilitate the

umbilical connections. The video and lighting which are part of the OMV

docking subsystem would be occluded in the present OMV configuration.

Therefore some mounting hardware and signal links are required on the ERM to

use the OMV docking subsystem.

COMMUNICATIONS & TRACKING

Communications and Tracking subsystem is provided by the OMV and provides

compatibility with the TDRSS, Space Station, GSTDN and the STS.

POWER

The power subsystem design is driven by the electrical load profile. Two

operational phases of the ERM mission include; a quiescent mode of status

monitoring and propellant storage thermal conditioning, and the hi-operating

mode during fluid transfer activities. During the 7 day ERM mission, 90% of

the time spent is in the quiescent power mode.

During quiescent mode the ERM power source will be the OMV power system. The

ERM thermal control system both active and passive will be designed to

function within the OMV power subsystem capability.

During fluid transfer, the high power requirements must be provided by a

dedicated ERM power supply. This power supply should be rechargeable

2198e/ - 168 -
batteries to provide multi-mission capability and on-orbit recharging. The

-power distribution sybsystem shall provideoswitching between both internal and

external power generation. It will also provide for energy storage management

to recharge the EEM batteries between missions. The batteries are sized to

provide all the power requirements of the transfer mode which include the

umbilical mating, .system checkout, resupply operations, system purge,

post-activity checkout, and demating activities. The electrical power

distribution system as shown in Figure 96.

ELECTRICAL

POWER - OMV

(QUIESCENT)

OMV

EXPENDABLES

VALVE

CONTROL UN IT

EVCU

HEATER

CONTROL

UNIT

HCU

SATELLITE

COUPLING

SYSTEM

SCS

DISTANCE

MEASURING

SYSTEM

0

M

V

DMS

POWER DISTRIBUTION ASSY

ELECTRICAL

POWER-BATTERS

(HI-POWER)

1 BATT

POWER

REGULATION

SYSTEM

1 PRS

g o * M i

DATA

PROCESSING

SYSTEM

JDPS

ILLUMINATION

CONTROL

| ICU

0

M

V

0

M

\l

TV

CAMERA

TV-CAM

TV

REMOTE

CONTROL UNIT

TV-RCl

Figure 96. Electrical Power Distribution

Data Processing Subsystem

The data management function is performed by the Data Processing System (DPS)

controlling all events and sequencing all operations on-board the ERM. The

ERM DPS provides for all software.necessary to perform dedicated ERM functions

and is connected to the OMV's DPS via a 2-way data bus. This allows for

maximum utilization of OMV software and processors for communications (uplink

& downline) and rendezvous/docking operations.

The DPS architecture is modular in nature with a distributed, standardized

processor arrangement integrated by a power and databus oriented structure

shown in Figure 97. The DPS operates in 3-modes: as the executive DPS

controlling all dedicated ERM functions; as a subsidiary DPS providing the OMV

DPS with all critical information and as a relay for all man-in the loop

control data uplinked via the TDRSS and DOMSAT systems.

The DPS provides for the following functions:

2198e/ - 169 -
o System checkout o

o Sequencing events o

o Automated malfunction o

procedures

o purging & venting operations o

o power distribution and o

control

System monitoring

Fault detection & annunciation

Expendables resupply control -

Environmental control

Satellite receiver system checkout

EXPENDABLES TANA

& LINES

PURGE

I TANKS & \

I LINES /"

HEATER

CONTROL

UNIT

TV

CAMERA

& TV

REMOTE

CONTROL

ERM DATA PROCESSING SYSTEM

TV SIGNAL

DISCRETES irrr

POWER

DISTRIBUTION

ASSEMBLY

Figure 97. Avionics/Data Processing System

ERM/CARRIES VEHICLE AVIONICS INTERFACE

The ERM Avionics will be designed such that when it is employed operationally,

subsequent to the demonstration experiment, minimum changes will be required.

An exception will be the additional software required with the DPS to

interface with the Orbiter for the demonstration phase.

The design of the Avionics system maximizes available carrier vehicle

capability while maintaining key ERM functional autonomy. The extended OMV

Avionics will provide the communications and tracking system and the

rendezvous and docking system.

To mimimize ERM cost and weight the OMV power system will provide necessary

system maintenance power during quiescent mode operations. The electrical

interface between the ERM Avionics and the Orbiter (demonstration program)

will be identical to that between the ERM Avionics and the carrier vehicle

(operational program). The DPS on-board the ERM will provide software for all

dedicated ERM functions and take advantage of the OMV's DPS where desirable

2198e/ - 170 -
via a two-way data bus. The OMV's DPS will provide for master data processing

for the EEM acting as an extended kit to the OMV.

Figure 98 shows the ERM/Carrier vehicle avionics for the operational and

demonstration Expendables Rsupply Module.

SATELLITE

& MONITOR

DATE INTERFACE

CMOS

FDA

COUPLING VERIF

PRESS

TEMP

QUANTITY

RATE

VALVE POSITION

DISCRETE CMDS

TV

ERM

(

3

— '

i

i

1

1

ELEC

E

/•

1

POWER

DATA INTERFACE

» /•

te

i

'

I

to

-

1

1

DOWNLINK/

DOWNLIST &

UPLINK

ORBITER ONLY

DISPLAYS & CONTROIS

CMDS

FDA

COUPLING VERIF

PRESS

TEMP

QUANTITY

RATE

VALVE POSITION

MONITOR

DISCRETE CMDS

TV MONITOR

CARRIER

VEHICLE

-V- GCS

Figure 98. ERM/Carrier Vehicle Avionics Interface

Result of the Expendables Resupply Module power requirements analysis

indicates that five (5) Li/TiS2 batteries having a total capability of 43«4

kwh and weighing approximately 675 Ibs. would provide power adequate for the

helium compressor and the bipropellant pump for the mission duty cycle shown

in Figure 99-

The mission selected represents a full STS duration OMV out-and-back

excursion. The hosekeeping power level was determined to be 815 watts, based

on the identified subsystem load requirements, as shown in Table 38. The

power required for the heaters is high and would drive the battery

requirements for a long 7-day mission even though proper thermal control with

insulating blankets and coatings was taken into account.

2198e/ - 171 -
Figure 99- Resupply Module Power Duty Cycle

t*rP

&

£

CO

LD

oo

tn

CSJ

(MX)

2198e/ - 172 -
Table 38. Resupply Module Electrical Power Requirements

EQUIPMENT POWER (WATTS)

QUIESCENT PEAK

DATA MANAGEMENT SYSTEM

THERMAL CONTROLS - HEATERS

FOR TANKS, PUMPS, VALVES & OTHER COMPONENTS

MECHANICAL ACTUATORS

LATCHES, END EFFECTOR

FLUID TRANSFER SYSTEM

PUMPS

COMPRESSORS

VALVES (5)

VENT/PURGE VALVES (4)

X-DUCERS/SIGNAL CONDITIONING

MARGIN

TOTAL POWER

50

670

75

2200

300

20

J5.

815

2300

1107

500

400

-15fl

7072

2198e/ - 173 -
The heaters were sized for a side to the sun GEO mission with a 50% BBQ. With

this orientation constraint, the power required was reduced to 670 watts

required to hold proper thermal temperature limits versus a 1000 watt

requirement with the resupply module/OMS end to the sun orientation. The

heaters were sized for the worst case to the sun orientation at a 50% duty

cycle or for a total of 2000 watts. With six tanks, this reduced to 350 watts

required per tank.

The heater requirement of 670 watts for a 7-day mission would drive the

battery sizing requirement; therefore, a power source user designation for

each different mission scenario was constructed and shown in Table 39.

3.2.15 Docking Hardware Selection

Several fluid transfer umbilical plate concept designs have been generated.

Each concept utilizes the standard RMS and effector (Figure 100) and grapple

fixture (Figure 101) to perform the docking and final closure mating. In

designs utilizing pressure balanced fluid disconnects to yield low interface

separation loads, latches may not be required and the end effetor-grapple

fixture mechanism may satisfy all of the connecting force requirements. It is

established that the closing force realized from the end effector-grapple

fixture is between 700 to 900 pounds. Fluid disconnects that are not or only

partially pressure balanced would require panel latches (Figure 102) to meet

the connecting force requirements.

Figure 103 represents a typical bipropellant fluid interface panel on a

receiver vehicle. This panel would contain fluid disconnects for propellant

transfer, electrical disconnects and a grapple fixture. The mating panel, to

be used on the resupply vehicle is shown on Figure 104. It contains mating

fluid and electrical disconnects and a standard end effector.

The sequence to be used for fluid interconnect would be the following:

1. Contact would be made with the end effector and grapple fixture.

2. The end effector would snare the grapple fixture and pull the

satellite or payload to be refueled to the resupply vehicle.

3. The end effector would bring the two vehicles together and would

engage the interface umbilical panel alignment pins.

4. The end effector would bottom out the grapple fixture plate to the end

effector face properly aligned by the alignment guides or pins.

5. The latches (if used) would actuate and clamp both halves of the

umbilical panel togeher.

6. The fluid disconnect would be actuated to engage to their final open

positon.

Umbilical Interconnect and Latching Device

The orientation of the latches and an estimate for the actuation travel

required for the fluid disconnects is shown in Figure 105. The engagement

travel distance was taken from a representative fluid disconnect.

2198e/ - 174 -
Table_39» Resupply Module Power Users Source Designation

219''

LU

Qi

O

CO

LU

CD

o_

CD

1 —

u_..

CO

LU

Q_

CO

«c

Qi

Qi

<c

CD

CO

s:

CD

LU

CD

LU

X

LU

CO

cr:

LU

f—

<c

CQ

CO

1 —

Q_

CD

Q_

O_

CO

1 —

nz

Q_

CD

Q_

Q_

co

i —

zc

Q_

CD

Q_

Q_

co s:

LU LU

CD <_>

CO

\

\

.

s. s. s. v \ \ ^

V

v V V S. \ \ s.

II

1—

C3

LU

CD

LU

LU

«=C

CD

CJD

CO

LU

CD

CO

CO

LU

H—4

OH

LU

«=c

CQ
ORIGINAL PAGE

OF POOR

EVA HANDHOLD

SNARE CABLES

SHOWN CLOSED

AROUND PHANTOM

OUTLINE OF

GRAPPLE FIXTURE

'- -/• SHAFT

ELECTRICAL CONNECTOR

Figure 100. Standard Snare Type End Effector

2198e/ - 176 -
Finish Matte Black

FED-STI>-595a 37038

3/8 In. Dia. Rod 4 In. Long

Matte Black with Matte

White Tip

— Matte White Lines

/ 0.2 In. Wide

1

 FED-STD-595a 37875

y-rl 11 j A | .020

r

Adjacent Figure for'

Target Details

NOTES: 1. Align 3/8 in. dia.

rod with RMS CCTV

camera £ ref. SPARRMS-ICD.017

2. Target plate cutout

and paint mask

(line width and

location) accuracy

is + .04 in.

Finish Dark

Gray Matte I

FED-STD-595a 36173

Grapple Fixture Target

None

Figure 101. Grapple Fixture/Target Assembly

2198e/ - 177 -
SPRING LOADED

SEPARATION

DEVICE

63

FLUID TRANSFER SUPPLY PANEL

(MALE COUPLING SIDE)

'FLUID TRANSFER RECEIVER PANEL

(FEMALE COUPLING SIDE)

Figure 102. Fluid Interface Panel Latch

2198e/ - 178 -
LATCH - ALIGNMENT RECEIVER

END EFFECTOR ALIGNMENT SURFACE

FLUID COUPLINGS.

ELECTRICAL DISCONNECT

GRAPPLE FIXTURE

Figure 103- Receiver Vehicle

Fluid Interface Panel

2198e/ - 179 -
END EFFECTOR ALIGNMENT SURFACE

END EFFECTOR

FLUID COUPLINGS

ALIGNMENT LATCH

Figure 104. Hesupply Vehicle

Fluid Interface Panel

2198e/ - 180 -
Figure 105. Side View - Fluid Interface Panel

2198e/ - 181 -
3.2.16 ERM System Weights

Systems and component weights for the ERM are presented in this section. The

dry weight of the resupply module was determined for two cases, a high and a

low. (See Table 40).

These high and low weight estimates represent the bonds of a 90% confidence

range. This indicates that there is less than a 10% probability of the ERM

dry mass exceeding 9168 Ibs or being less than 8259 Ibs.

3.2.17 Receiver Spacecraft Impact Assessment

Four different types of satellite propulsion systems were analyzed as part of

the spacecraft impact assessment: hydrazine blowdown, bi-propellant blowdown,

pressure regulated (bi-propellant), and pump-fed (bi-propellant).

The modifications required for the on-orbit refueling of spacecaft propulsion

systems were minimal. A list of these modifications are as follows:

1. All satellites were assumed to be launched and deployed by the Space

Transportation Systems. Therefore, all satellite propulsion systems were

modified to meet STS payload safety and contamination requirements

including: a) propellant lines leading to an overboard outlet should be

controlled by a minimum of three independent mechanical inhibits, and b)

pressurant lines leading to an overboard outlet should be controlled by a

minimum of two independent mechanical inhibits.

2. Multiple tank systems/subsystems (hydrazine systems, MMH subsystems, and

NTO subsystems), require tank Isolation valves for propellant

loading/ullage transfer control.

3. To insure the integrity of the payload's environment and the local

environment of the space station, a disconnect/line purge circuit should

be incorporated into the propellant transfer systems of all space vehicles

requiring on-orbit refueling operations. High pressure gas (which would

be supplied by a resupply module, or any other resupply system), would

flow through the purge circuit and clear all refueling disconnects and

adjoining lines of residual propellant, prior to disengaging the receiver

and resupply vehicles.

The component and weight impact assessment of these modifications, on the four

previously mentioned satellite propulsion systems are summarized in Tables 41

through 44. Schematics of these modified propulsion systems are illustrated

in figures 106 through 109.

2198e/ - 182 -
i ORIGINAL PAGE IS

Table 40. ERM System/Component Weights OF POOR QUALITY

i

IA

v>

a

IA

a

IA

IA

2198e/

I

S 3

8

e «•

§13

s

a a

i i

35 £*•

a o

•• ••

• •

o o

Ul Ul

S IA

3 t

5 I

Ob >•

a »

,- I

z &.

a a

i i i

8 8

S

HI

k- M "*

a *» 5 * o

55225

0 PLA

T

EATERS

2

5 g g i>j t2

•O .» IM O O f» «r • * m

i ui r

- K S ~

3

Z W

l/l UJ «

Z "• Z V» u. v»

•• X O _J « -J

k> OC O. Ul W Ul

IM

IA

a

O

!>

e

af

S £

e oc

S

O O k.

^ Z w

• Ul U •• <

O Ml M •

2££8|

& « ui a 5

01 v>

tn IM

IM

irv ^*

§ S

•« *. »M

» S: 5

O !»!»- O p. «

1

 1 S2

 s

k- V>

IA a

M C z VI Z M

m
IS

BE POOR QUALITY

Table 41'. Hydrazine Slowdown Satellite Weight Impacts

(Ullage Recompression)

1.

Itpict oft Sitillitt Prop. Syitnt

Hydraiint Blotriewi

Pr opal lion SyitH

I. UlUgi Rtcoip. fftdificttioni

i. Inttrfact Pmtl

1. Pmtl

2. ffiuonntctt

3. EltttricaJ Conntctiont

4. Srapplf Fiitart

ii. V«lvt PmtU

1. Pmtl i

2. Vtlvti

iii. Othtr Cotpontnti

1. Vilvn (Tank Iitlatien)

SyitH Htiaht Incrttw

b. Ullagt RKOIP. Modification*,

OiKonntct/Unt Purjt Option

i. Inttrftct Pmtl

1. Pmtl

2. Diteonntets

3. ElKtncil Conntetions

4. Srapplt Fiitart

ii. Vilvt Pmtli

1. Pmtl i

2. Vilvti

iii. Othtr Cotpontnti

1. Vilvti (Tmfc Iiolition)

Syittt Htight Incrttw

Vt./Unit

llbi.)

18. S

4.0

1.0

22.0

25.0

2.3

2.3

25.0

4.0

1.0

22.0

25.0

2.3

2.3

Sturct o* In* or ut ion

Enginttrino, Htiittt

LEU dtiign » 4

Enginttring titiatt

Shuttlt I/F control doc.

Eng. Ht. buttf on Shuttlt

Shuttlt coaponmt

Shuttlt Cotpontnt

Enginttrina ntiutt

LEH dnign § 4

Enginttring ntititt

Shuttlt I/F control doc.

Eng. tft. b»td on Shuttlt

Shuttlt cotpontnt

Shuttlt cotpontnt

« of Units

Rtquirtd

1 • (1/vthiclt)

1

2 (p«r,d»tJ>

1 (i/vthiclt)

1

2

4 (l/tmkl

1 (1/VtfllClt)

2

2 (pw,dati)

1 (1/vthiclt)

1

3

4 (I/tank)

1

Total Ht.

(lb».)

18.8

4.0

2.0

22.0

25.0

4.«

9.2

85. i

25.0

8.0

2.0

22.0

25.0

6.9

».2

78.1

2198e/ - 184 -

c -
Figure 106. Impact on Hydrazine Slowdown Propulsion Systems

For "Ullage Recompression" On-Orbit Resupply

(0

— t» --•

-

 1 r

 -
Table 42. Weight Impact On Bipropellant Slowdown Propulsion

Systems for Ullage Recompression Resupply

2.

laoact M Sttilliti Prop. SyitiM

W-froptllint 81 orient

Propulsion SyitH

I. Ullaqi Ricgtp. HodHicationi

i. Intwfaei Pint!

1. Pintl

2. Diuonnicti

3. ElictricaJ ConnKtiont

4. Srapplt Fixturi

ii. Valvi Pintlt

1. Ptntli

2. Valvis

iii. Othtr Coapontnts

I. Vilvn (Tink tiolition)

SyitM Vtiqitt Incrtaii

b. Ullaqt Rtcotp. Notification*,

OifconnKt/Lint Purqt Option

i. Intirfici Pinil

1. Pintl

2. Oitconntcti

3. ElKtrieal CannKtions

4. Sripplt Fiiturt

ii. Vtlvt Pmtli

J. Pintli

2. Vtlvn

iii. Othir Coipenmti

1. Vilvti (Ttnk Iiolitioit)

Syfttt K»ioht Incrinc

It. /Unit

Obi.)

29.0

4.0

1.0

22.0

23.0

2.3

2.3

37.3

4.0

1.0

22.0

2S.O

2.3

2.3

Sourei ti taforution

Enqinnrinq Htiuti

LEH tniyi » 4

Enqinrtring ntiuti

Shuttli I/F control doe.

Eftq. Mt. butd on Shuttlt

Shuttlt eoaponmt

Shuttlt Coiponmt

'

Enginttrino, nttiati

LCT dniqn » 4

Enqinttrinq ntiutt

Shuttli I/F control dec.

Eng. tit. boitd on Shuttli

Shuttli coopontnt

Shuttli coipenmt

1 of Uniti

tequirid

_ —

1 (1/vihicli)

2 (loi/Hu>

2 (pw.diti)

1 U/vihieli)

1

4 (2oi/2fu)

0

1 (1/vihiclt)

4 I2ox/2fu)

2 (p*r,ditJ)

1 (1/vihicli)

1

6 (3oi/3fu>

0

Total «.

(Ibt.)

25.0

9.0

2.0

22.0

29.0

1.2

0.0

»1.2

37.5

16.0

2.0

22.0

25.0

13.8

0.0

114.3

2198e/ - 136 -
Figure 107. Impact on Bi-Propellant Slowdown Propulsion

For Ullage Recompreasion On-Orbit Resupply

-sCO

CO

K

CO

UJ

CO

CO

-s1

Gu en O < ui

g S°=

u, fct

K ca

2198e/

- 187 -
Figure 108. Impact on Pressure Regulated Bipropellant Propulsion

Systems for Ullage Exchange Resupply

Non-Return

Valve

Press. Regulator

"

Relief

Valve

To

Thrusters

To

Thrusters

-sn- -03-1

To

Thrusters

He

Trans.

"

LlageCNTO>

Trans.

S-G

MMH

Trans.

Thrusters

A, SATELLITE PROP, SYSTEM

(NO MODIFICATIONS FOR

ON-ORBIT RESUPPLY)

B, MODIFICATIONS FOR ULLAGE EXCHANGE

PROPELLANT RESUPPLIES

i—B-G

To To

Thrusters Thrusters

c, MODIFICATIONS FOR ULLAGE EXCHANGE

PROPELLANT RESUPPLIES

(DISCONNECT/LINE PURGE OPTION)

2198e/ - 188 -
Table 43. Weight Impact on Pressure Regulated Propulsion

Systems for Ullage Exchange Resupply

PAGE IS

OF POOR QUALITY

3. J

lepact en Sit tl lit* Prep. Syitm

Bi -Propel 1 ant, Prtsturi Regulated

Prepuliion Syitei

a. Ullage Exchange Modifications

i. Interface Panel

1. Panel

2. Ditconnecti

3. Electrical Cennectieni

4. Grapple Fixture

ii. Valve Panels

1. Panels

2. Valvei

ill. Other Ceapenenti

1. Valves (Tank Isolation)

Systti Height Increase

b. Ullage Exchange Hodificatiens,

Disconnect /line Purge Option

i, Interface Panel

i. Panel

2. Disconnects

3. Electrical Connections

4. Srapple Fixture

ii. Valve Panels

, ". 1. Panels

2. Valves

iii. Other Coeponents

1. Valves (Tank Isolation)

Systee Height Increase

Ht./Unit

libs.)

44.0

4.0

1.0

22.0

2S.O

2.3

2.3

—-—«-••

69.0

4.0

1.0

22.0

2S.O

2.3

2.3

Source of Information

Engineering estiaate

LEH design » 4

Engineering estiaate

Shuttle I/F control doc.

Eng. est. based on Shuttle

Shuttle component

Shuttle Component

Engineering estieate

LEH design f 4

Engineering estiiate

Shuttle I/F control doc.

Eng. est. based on Shuttle

Shuttle component

Shuttle coiponent

1 of Units

Required

1 (I/vehicle)

5 (2ox/2fu/lHe)

2 (par, data)

1 (I/vehicle)

2

9 (4ox,lHe),

(4fu)

0

— — •-•j'»""— •

1 (I/vehicle)

9 (4ox/4fu/lHe)

2 (pw.data)

1 (I/vehicle)

2

13 <6w,lHt),

(ifu)

0

Total Ht.

(Ibs.)

44,0

20.0

2.0

22.0

50.0

20.7

0.0

158.7

69.0

36.0

2,0

22.0

50.0

29.9

0.0

208.9

2198e/ - 189 -
Table 44. Weight Impact on Pump-Fed Storable Bipropellant

Propulsion Systems for Ullage Exchange Resupply

4.

Iiptct on Sattllitt Prop. Syitm

M-ProBtllmt, Putp-ftd

Pronuliioft Syttii

I. ID 1 191 Eichanqt Ntdifi cation*

i. Intirfact Pinil

1. Pwtl

2. OiKonnKtt

3. ElKtricil Conntctiont

4. Srapplt Fixturt

ii. ViJvt Pantli

1. Plfltlt

2. VilvH

iii. Othtr Coiponmtt

1. Vilvii (Tink Iiolition)

Syitu Htio,ht Inertiu

b. Ullaqt Exehmqt HodHieationt,

Diuonnict/Lini Purjt Option

i. Inttrfaet PintI

1. Pinil

2. Diuonntcti

3. Elietrical Connictioni

4. Sripplf Fixturt

ii. Vilvt Pintli

]. Pinili

2. VI!VH

iii. Other Cotponmtf

1. Vilvii (Tink Iiolition)

SyttH K»i?ht Incrtait

«. /Unit

(Ibi.)

SO.O

4.0

1.0

22.0

29.0

2.3

2.3

75.0

4.0

1.0

22.0

29.0

2.3

2.3

Sourei of Inforution

Engintiring ntititt

LEH dniqn « 4

Enginierino, Htiuti

Sknttli I/F control doe.

Enq. nt. bued OR Shuttlt

Shuttli eoipoRint

Shuttlt Cotpontnt

Eno, inuring titintt

LEU dnign » 4

Engintfrlnj iitiiitt

Shuttlt I/F control doc.

Eng. tit. baitd on Snuttlt

Shuttlt cotpontnt

Shuttlt cotpontnt

1 o^ Units

Riquirtd

1 (1/vtniclt)

« (2n/2ftt/2Ht)

2 (pw.diti)

1 U/Yiflicli)

2 (loi/lfu>

10 (4«i/lHt),

(4*u/!Ht)

4 (1/tink)

1 (I/vtfticlt)

10 (4oi/4fu/2Ht)

2 (pv.diti)

1 (1/vtfticlt)

2 lloi/Hu)

14 (6oi/lHt),

(ifu/lHt)

4 (1/tink)

Total ft.

(Ibi.)

50.0

24.0

2.0

22.0

50.0

23.0

».2

180.2

75.0

40.0

2.0

22.0

50.0

J2.2

«.2

230.4

2198e/ - 190 -
PAGE 53

OF POOR QUALITY

Figure 109- Impact on Bipropellant Pump-fed Propulsion

Systems for Ullage-Exchange Eesupply

A, SATELLITE PROPULSION SYSTEM

( NO MODIFICATIONS FOR

ON-ORBIT RESUPPLY )

KLLOHS

KCUNUTORS

C, MODIFICATIONS FOR ULLAGE EXCHANGE

PROPELLANT RESUPPLIES

( DISCONNECT/LINE PURGE OPTION )

B, MODIFICATIONS FOR ULLAGE EXCHANGE

PROPELLANT RESUPPLIES

2198e/ - 191 -
3.2.18 Advanced Development Requirements

The essential functions which are required to demonstrate the operational

interfaces of the resupply module include the final closure and remote fluid

disconnect engagement and actuation, proper ullage positioning of the receiver

and supplier's propellant tank PMD for the ullage exchange process, total

system operation and checkout utilizing microprocessor control and safe system

termination including disconnect disengagement, fluid transfer plate

unlatching and safe vehicle separation. Each function requires elements to be

included in the flight demonstration program as follows:

o Docking end-effector/grapple fixture snare and final closure to remote

fluid disconnect engagement.and actuation-demonstrates ability to

remotely provide close tolerance alignment for latching and remote

fluid disconnect engagement and actuation

o Ullate positioning of the receiver and supply tank's propellan

management device is required to exhibit proper ullage control during

propellant transfer by ullage transfer process.

o Microprocessor control of system status and health monitoring, fluid

transfer including the pumping system, valving, etc., and finally line

purging and system safing.

o system safe termination including remote disconnect disengagement,

fluid transfer plate unlatching, and finally safe vehicle separation.

Several advanced developments are required to provide the base needed for the

types of components to be used for remote resupply. These risk areas include

seal develoment for fluid quick-disconnects, propellant pumps, and gaseous

compressors, as well as fluid acquisition/ullage positioning in order to

provide a contamination free transfer process.

o The components utilizing these advanced developments would include the

following:

DEVELOPMENT RISK COMPONENTS

Propellant Pumps Pump shaft seals and lubrication Moderate

Gas Compressor Compressive element seals, staging Moderate

Requirements and heat rejection and

lubrications

Quick Disconnects Sliding redundant seals Low

Propellant Management Desin of low-g ullage positioning Low, low

device

The development risk of each technology element is inversely proportional to is

technology maturity level (as defined by standard NASA technology development

descriptions). The gas compressors, quick disconnects, and fluid pumps have been

identified for a special technology development effort that must preceed the ERM

program. Contamination control is not expected to be a special problem for the

fluid resupply operations anticipated due to the selection of the ullage exchange

process for the bi-propellant transfer. Table 49 summarizes the technology status

for the resupply module program.

2198e/ - 192 -
Table 45- Status of Technology for the

Resupply Module Program

ORIGINAL PAGE IS

OF POOR DUALITY

Technology Maturity Levels

Ho. Description

8 Operations

7.2 Engineering Model Teited in Space

7.1 Engineering Model Qualified

6.2 Prototype Tested in Teat Bed Unsunned

A.I Prototype Tested at Contractors

S.4 Preprotocype Tested at NASA

5.3 Preprototype Tested at Contractors

5.2 Major Function Tested at NASA

S.I Major Coatponents Tested at Con tractor a

4.J Critical Hardware tested

4.2 Critical Function Tested Over Tine

4.1 Critical Function Demonstrated

3.2 Conceptual Design Tested Exper mentally

3.1 Conceptual Design Tested Analytically

2 Conceptual Design Foraulated

1 Baaxc Principlea UDservefl and Deported

Development Kisk Assessaent

Technology Requirements

Fluid Contatcnent/Tranafer Design

Fluid Acquisitio

n

2

i

M

:

I

M

•0

Wtf

aPM

.

2

•

Q.

1

9

*a

•*t

a

^»

mm

3

I

•

u

L

>

m•

U

mm

4

*a>

U

1

1

•

mmt

O

M

U

•eo)

Sr

mm

3

x

« • *rf *•

I Flou/Quanl Maeaureoier

1

1

•e

1 Leakage Monitor ii

mmmm

• "Ofl

2

'io

n Control Contamlnal mm

*•

3

re Control

a

««

•

|

w

2-

1

• —•

• w •

5 ? S ±

0 .0 0 £

3 B 0. 0

use *

iun — a. u 5 u —

Very Low, 3-

1

Denote

Operations

•

•

wb —0 o• S k •>

U b

m e t

I 3 . ?,

5 S . -2 « ^« *J e u

•o jt o o *

c u s -> S

• o « » o

x Q a: < a.

Low, 4* Moderate

1

Three major components can be identified as requiring special developmenal testing

to provide concept feasibility. These are the propellant pump, the helium

compressor, and fluid disconnect.

The major test/demonstration items required to verify satisfactory component

operation and performance are as follows:

o Evaluate component design concept feasibility with a workhorse-type test

unit.

o Verify operation and performance of a flight-weight prototype test unit to

flight application requirements.

The scope of the Test Program should encompass the following:-

o Breadboard-type tests using referee fluids and a development (workhorse)

component test unit shall provide the basic data base for component

functional design evaluation.

o Development tests using referee fluids, propellants and prototype

flight-weight test components shall provide the basis for verification of

the component design concept feasibility and fluid transfer usage. In

addition, component life shall also be demonstrated.

2198e/ - 193 -
The Development Test Schedule for the components requiring special developmental

testing to provide concept feasibility has four major sections. The first two

include the generation of specifications and requirements and the design phase for

the developmental hardware and test setup. The third section includes the

breadboard testing of workhorse units for pumps and compressors and runs from 2nd

quarter of FY87 through 2nd quarter of FY88. The fourth section -

Development/Verification Testing, provides the verification of operation and

performance of flight-weight prototype test units for the pumps and compressors and

runs from 3rd quarter FY87 through 4th quarter FY88. Figure 110 presents the

overall schedule for development test.

At the conclusion of the development/verification testing flight-weight design

components will have been shown to be feasible for operation when subjected to

flight environments, ground servicing and fluid transfer usage.

j FY86 FY87 FY88 FY89

PHASE

TFST1N6 PHASF

n (R/tt)

1. COMPRESSOR

2. PUMP

ft > T

r —i

-- j

» f

1, COMPRESSOR

2, PUMP

3, DISCONNECT

-+-

» f T

T

A

0 - DESIGN COMPLETE

F - FABRICATION COMPLETE

T - TEST COMPLETE

A- FINAL REPORT

[""> B/B LOW LEVEL

"" TEST SUPPORT

MANUALLY

ACTUATED

PR IVFRFO

MATED DISCONNECT

QUAL

-A

BI-PROP AQUAL

Figure 110. Development Test Schedule for Concept Feasibility

3.2.19 Demonstration Flight Program

The flight demonstration concept will include sufficient fidelity to test the

essential functions of remote on-orbit resupply and demonstrate to the user

community the viability and attractiveness of such a concept. In order to show

proof-of-concept, specific elements will be required on each side of the resupply

interface.

2198e/ - 194 -
Simplicity and economy are emphasized in the selected flight demonstration program

(Figure "111). The principle test objective is to validate bi-propellant fluid

transfer in a zero-gravity environment. Since the ERM engagement of a spacecraft

is accomplished by the OMV with a remote manipulating arm, the remote manipulator

system (RMS) on the Orbiter can approximate the OMV engagement capability. The

systems on both the receiver and supplier test articles can be attached to either a

MPESS (NASA owned) or a SPAS structure (structure only without any subsystems).

The Orbiter RMS will lift the receiver test article out of the cargo bay and engage

the supplier test article. Repeated fuel transfer tests will then be performed.

Since the supplier test article has the greater need for power and electrical links

to the Orbiter, we recommend it remain seated in the cargo bay. This avoids the

need for complicated power and electrical links running through the RMS. The

supplier test article would be autonomous with its own power source.

The system on both the receiver and supplier test articles (Figure 112) can be

attached to either a MPESS (NASA-owned) or a SPAS structure (shown). Grapple

fixtures are provided on the receiver simulated vehicle for use in docking with the

supplier vehicle and by the RMS to transport the receiver vehicle from a berthed

position to the docked position with the supplier vehicle. Fluid transfer system

hardware will incorporate to the greatest extend possible final design pumps,

compressors, ullage control tankage and disconnects.

Figure 113 presents the resupply module flight demonstration supplies schematic.

SPAS OR MPESS STRUCTURE FOR

RECEIVER AND SUPPLIER

SIMULATORS

RMS USED TO DOCK

RECEIVER TO SUPPLIER IN

PAYLOAD BAY

MOST OF DEMO SUBSYSTEMS LATER

USED FOR OPERATIONAL FLIGHT

UNITS FOR COST SAVINGS

2198e/

Figure 111.

Bi-Propellant/Helium Transfer Flight Demonstration Concept

- 195 -
RECEIVER VEHICLE SUPPLY VEHICLE

FUburfK*

Cametar Atuafarf

Greppki future

Interim Conntctor. -m./^, End Effector

Asumntyv

Intirtaci ConMctor

Asujnbly

Initrfact Support

Slruciuri

SHE) \

Ononal

SPAS-01

Structure \

Imafacf Comenar. End Effector

Inioftct Connector

Interface Support

Structure

Figure 112. Remote Resupply Plight Demonstration Configuration

2198e/ - 196 -
PRESSURANT

CHECKOUT

d) PANEL

INTERFACE

PANEL

PROPELLANT

I SERVICING &

PURGE/DRAIN

PANEL

Figure 113.

Resupply Module Flight Demonstration Program Supplier Schematic

3-3 Programmatic and Developmental Planning

The programmatic and developmental planning task included the definition of

cost and schedule requirements for the resupply module operational program,

technology validation and flight demonstration programs.

3«3«1 Groundrules -and Assumptions

Task 3«1> Technology Development, Flight Demo, and ERM Program

The Baseline program consists of two test articles (Developmental and

Qualification) and two flight units (one from salvaging systems from the test

articles. The earliest possible start for the phase C/D of the ERM program is

the fourth quarter FY 1987, and the earliest possible first operational flight

is the last quarter of FY 1991.

Task 3.2, Estimated Program Cost

All costs are in constant 1984 dollars. STS launch costs are not included for

either the flight demonstration cost or the first ERM operation flight. Costs

for GSE and payload support, sustained mission operations/training,

maintenance and refurbishment, and additional production units are not

included in the baseline ERM program.

2198e/ - 197 -
3.3.2 Advanced Development and Demonstration Flight Plans

ORIGINAL 53

OF POOR QUALFTY

Two options were considered for the flight demonstration program: An early

flight demonstration by the end of the FY 1988 and a late flight demonstration

that is part of the ERM phase C/D program that flies a year later at the end

of FY 1989.

The technology development program would precede both the flight demonstration

program and the ERM phase C/D. All the technology requirements and their

maturity status are shown in Figure 114. Standard NASA definitions were used

to assess each technology's maturity level. By definition, those technologies

with low maturities are the ones that need to be emphasized in the advanced

development program. The critical developments are, (1) the gas compressor,

(2) fluid pumps, (3) quick disconnects, contamination control, and leakage

monitoring.

Technology Maturity Levels

No. Description

8 Operations

7.2 Engineering Model Tested in Space

7.1 Engineering Model Qualified

b.J Prototype Developed to quality

6.2 Prototype Teated in Teat Bed Unmanned

6.1 Prototype Teated at Contractora

5.4 Preprototype Teated at NASA

5.3 Preprototype Tested at Contractors

5.2 Major Function Tested at NASA

S.I Major Components Teated at Contractors

4.J critical Hardware Tested

4.2 Critical Function Tested Over Time

4.1 Critical Function Demonstrated

3.2 Conceptual Design Tested Experimentally

3.1 Conceptual Design Tested Analytically

2 Conceptual Design Formulated

1 Basic Principles Observed and Reported

Development tisk Assessment

Technology Requirements

Fluid Containment/Transfer

§

f4

W

f*

m*•«

3

0*

tj

<

T3

a

u.

2

t

9

•

1

1

*

)

>«4

9

m

2

^i

3

••4

3

Ml

3

m

>

If

^1

>

)

•

u

ami

4

2

t

9

)

U

^4

1

t

a

.••

3

cr

••

3

>«

t m

•* u

' C

: «

• a

2. • 7 u

~. 3

* •

O •

(•• X

1

00

e

••4

U

B O

X U

• •**

jt e

• o

« X .1

•^^

,

2

a

4^

e

o

j

:

3

*^

J

•

•**

B

0

^

s

u

3

o

j

3

J

II

3

^

D

M

II

a.

a

H

1

Design

• «••

- « c o

3 C « ->

rfU f — O C U *>

3 B Q. 0

2 2 g i M a. u u

1

Remote

Operation*

•

•

» ~* u

a. o • 3 b <M W 14

• C •

§

O M

0 • C,

> «a u <•?

» c • -<

V ^4 4J fi to

9 J< O O •

c o B -• 5 II O II > O K O 0! < 0.

1

1- Hone, 2- Vary Low. 3- Low, 4- Moderate

Figure 114. Status of Technology

Required For ERM Program

2196e/ - 198 -
The schedule for the early flight demonstration option Is shown In Figure

115. Breadboard testing precedes the verification testing. The verification

testing could culminate with the flight demonstration.

\ 1986

EARLT FLIGHT

DEMO OPTION

1 1987

12H

* 1 1988 I

FLIGHT

| TECHNOLOGY DEVELOPMENT *

1989

DEMO

c

1 1990

REDESIGN &

1

4H

1991 1

VALIDATION |

of critical technology

components

PHASE C/D START

ERM PROGRAM DESIGN

Ground Test

£££££ 7* Validation ,„ PLIGHT DEMO

TECHNOLOGY DEVELOPMENT ^ I REDESIGN & VALIDATION

PHASE C/D START

ERM PROGRAM DESIGN

Figure 115

Early Vs. Late Demo Option Schedules/Costs

The schedule for the late flight demonstration option is shown in Figure 116.

The technology development phase culminates with the successful verification

of key predefined ground tests with breadboard test apparatus, probably by the

end of FY 1987.

This schedule is shown in Figure 116. Given a successful verification at the

breadboard level, the ERM phase C/D could start. After the specification and

requirements for the ERM program are defined, the flight demonstration program

would begin. Changes to previous specification and requirements defined in

the advanced development program will probably result in a

redesign/modification to the critical components in parallel with the design

of the flight demonstration.

Each of these two options for the flight demonstration program have their

advantages and disadvantages. The cost of the two options are nearly the same

as shown in Figure 113. The early estimated cost for the flight demonstration

program is $12M with $4M required to modify and validate any specification or

requirement changes once the ERM phase C/D program starts. The estimated cost

for the late flight demonstration is $7M for the technology development and

then $9M for the flight demonstration. Both options—technology development

and flight demonstration combined—cost nearly $16M.

The other considerations in closing between the two options are, (1) risk, and

whether an early ERM phase C/D program start is more desirable than an early

customer acceptance/commitment to the ERM program. The advantages and

disadvantages of the two options are displayed in Figure 117. Many issues are

still unanswered before making a selection between these two options or the

creation of other options:

2196e/ - 199 -
1986 1987 1988 1989 1990 1991

[TECHHOIQGT DEVELOPMENT PROGRAM [

SPECIFICATION/REQUIREMENTS

BREADBOARD DESIGN

BREADBOARD FAB I TEST

CRITICAL TECHNOLOGY COMPONENT

| FLIGHT OEHOHSTMTlCm PROGRAM |

SPECIFICATION/REQUIREMENTS

FLIGHT DEMO DESIGN

FLIGHT DEMO FA» i TEST

CRITICAL TECHNOLOGY COMPONENT

I CONCEPT VERIFICATION

DESIGN

| FAB

(DESIGN) FAB |

| TEST

TEST

FROM ERH PROGRAM

I DESIGN I TEST t C/0

ERM PROGRAM

SPEC I F I CAT I ON/REQU I REMENTS

DETAIL DESIGN

DEVELOPMENTAL TEST ARTICLE

QUALIFICATION TEST ARTICLE

FIRST FLIGHT HARDWARE

MISSION SUPPORT I ANALYSIS

DESIGN

~®

FA, | \ <

FAB | TEST

DESIGN ]

f

FAB

/

0

TEST i REFURB SYS l^"'"""* '»'"*

TEST | X ^-n FLIGHT HARDWARE

X

 F*» ^ *<J)

TEST I C/0^ OPERATIONS {

LEGEND

1 SPECIFICATIONS FOR BREADBOARD DESIGN

2 BIPROPELLANT TRANSFER VALIDATED WITH GROUND TESTING LEADS TO

ERH PROGRAM PHASE C/0 START

3 ERM SPECIFICATIONS DRIVES FLIGHT DEMONSTRATION PROGRAM

t FLIGHT DEMONSTRATION HARDWARE AVAILABLE FOR REUSE ON FIRST

ERH FLIGHT UNIT

S FIRST ERH FLIGHT

Figure 116. Late Demonstration Flight Schedule

a) Is the flight demonstration needed to gain customer

acceptance/commitment?

b) Is the flight demonstration needed to reduce the technical risk of

the program? or both

If the user community could be persuaded to support and commit to the ECM

program without the flight demonstration, then we recommend the late flight

demonstration option and the schedule shown in Figure 116. A technology

development program would be initiated prior to the start of an operational

expendables resupply program. The Technology Development Program would use

ground testing to develop results sufficient to initiate the operational

program.

Preferably, the specifications and requirements for the Technology Development

Program would be oriented to the ERM Program so that the critical technology

components (pump, compressor, and quick disconnects) developed would be usable

without extensive redesign. More likely, however, considerable learning will

take place during the Technology Development Program and the Phase B of the

ERM Program. Thus, the critical technology components will probably require

some redesign and, consequently, recertified in the Flight Demonstration

2196e/ - 200 -
Figure 117. We Recommend a Late Flight Demo & Early ERM Program Start

EARLY FLIGHT DEMO

(END OF FY 1988)

LATE FLIGHT DEMO

(END OF FY 1989)

GO

LU

CRITICAL TECHNOLOGIES TESTED

IN SPACE - LOWER RISK

EARLY CUSTOMER ACCEPTANCE/

COMMITMENT TO ERM PROGRAM,

EARLY ERM PROGRAM START

(3RD QUARTER FY 1987)

ERM SYSTEMS TESTED ON

FLIGHT DEMO - LOWER RISK

GO

GO

ONE YEAR DELAY IN ERM PROGRAM

START

DOES NOT TEST SYSTEM DESIGNED

FOR THE ERM PROGRAM - HIGHER

RISK

ERM PHASE C/D START BEFORE

CRITICAL TECHNOLOGIES ARE

TESTED IN SPACE - HIGHER RISK

LATER CUSTOMER ACCEPTANCE/

COMMITMENT TO ERM PROGRAM

2196e/ - 201 -
Program. Of critical importance, is the fact that all the systems and

hardware in the Flight Demonstration Program reflect the specifications and

requirements of the ERM Program. Then, systems tested in zero-gravity would

be those of the operational expendables resupply program. When flight

demonstration is completed at the end of FY 1989, the systems and hardware can

be salvaged for reuse on the ERM first flight unit, which is scheduled for

launch two years later.

Successful completion of predefined key ground tests, that raise the

development of the critical technologies to an acceptable level, lead to the

start of the ERM Program. Six months later, after the ERM specifications and

requirements are defined, the Flight Demonstration Program would begin,

designed to use ERM systems and hardware.

3.3.3 Work Breakdown Structure

The Space Platform Expendable Resupply WBS, as defined herein, was selected

mainly to support the program planning, scheduling, and costing performed in

the programmatlcs Task 3. The top level WBS is shown in Figure 118. The

selected work breakdown structure for the ERM program consists of the

following elements:

1.1 Program management and system engineering is the core of the

program and provides the overall program perspective necessary for

continuity and decision-making.

1.2 Engineering design

1.3 Developmental, qualification, flight demonstration, and functional

testing.

1.4 Developmental, qualification, flight demonstration, and functional

testing.

1.5 Mission planning, flight support, and check-out necessary to

support the first launch and post-flight analysis.

1.6 Payload support for the resupply module in the orbiter payload bay

is GFE (this hardware is available from other programs).

1.7 & 1.8 Operational support beyond the baseline program has not been costed.

The WBS for the Expendables Resupply Module Systems are Orbiter/STS

designations, as shown in Figure 119.

The weights for end ERM systems range from a low to a high estimate,

reflecting an 80 percent confidence that the actual design weight will be

within this range. The total weight of the ERM will range between a low of

8260 and a high of 9170 pounds. The system weights by WBS are given in

Figure 120.

2196e/ - 202 -
Figure 118. Work Breakdown Structure For ERM Program

-'

•0

•••

in

mmm

ro

mmm

ru

•^

-

MISSIO

N

OPERATION

S

I

GSE &

PAYLOA

D

SUPPOR

T

LAUNC

H

OPERATION

S

HIMB^HHi^^H

1 0

p«- 1 g < W»

UJ W Ul

t- M *-

10 Us- « aa

v* as

^

PRODUCTIO

N

MANUFACTUR- IN

G

DESIG

N &

JEVELOPMEN

T

H> dt

^9 O

X Z

Ul

3 §

^9 tfl

oo

MAINTENANCE 1REFURBISH- 1MEN

T 1

a

2

1

"

o

CO*

1FIRS

T FL

T

HARDWARE

TES

T 8 C/

0

DEVELOP- MENTAL

TESTIN

G

i — —

DEVELOP- I

MENTAL IES

- ARTICLE !

1 PRELI

M

DESIG

N

PROGRA

M

MG

T

ST

1

5 F1TT1N(

* LATCHE

S

v> o a.

a: •» a.

Zs!2

5 «

«*- z

M »4

_» z*-

< O «fl

3 M Ul

0 *- HUl

u> _J

< W) »-

3 u< ae

O H- <

riS< M

*~ w.

UJ Ul

a a SYS ENGR'( ft NTEGRATIO

K

H5 o »-

•* g to

^1 ^y ^y

U. o H"

uj 52 $ j ui uu o

H- - l/> (K

^^

Ul

2196e/ - 203 -
Figure 119. ESM Subsystems Work Breakdown Structure

^

i

M

•

^•1

—

—

•*•••

^MH

i^M

ENVI

R CONTR1

SYSTEMS

VIONICS

^

1 -

on

*" 9

O fid

ea

u > Gd en

Gd

U >•

M M

P" as

Ss

^1 O

Bd a*

en

a P M en

* >• CIIANICA

L

SYSTEMS

aTRUCTURA

L

YSTEMS

en en

in

*T

«*

•— •

A_

«.

2

eo

j

COL

D PLATES

& HEATERS

VIDE

O

1.3.4.

3

fREGULATIO

N

& DIST.

as

ad

cdH

H

M

OJ

M

1

of

0*

I

•a

en

FRAM

E

SUPPORT!

1

I

THERMA

L

PROTECTIO

N

OVOTU

U

ZINSTRUM

E

TATIO

N

1.3.4.

4

ISEQUENC

E

CONTRI.RS

2

at

au

a

<

W

LINES &

COMPONEN

T

1

13

Soa uB

1

C3 fid

Z *

£2 M CO

9b

1

•

•=r

9

MB

1

ad

! BLANKETS i

»n

4

6

4

1UMBILIC

A .

rDOCKIN

G 1

|

COMPONEt |

COVERS UPROCESS I

N

r

^

•

•<

— i•

M

t

... i

fm

Cl)INTERFA

C

5

| ACTUAT

O

M

.

z

fl£ h^ QD

£5 Hoj . xsnc

n u of j . 2z«^

n a o«-» -IS

3 z a-« « aHr

e H c»j •- s •

- «3 CO —

1

 !, I

^ ^^^^ ?^i

a atf jt! fe ^ ad M

a, 5J

 u

1 1 1

V

H

M

fid U

u a

CO

*x

,.-IJ g

> ' M

0

O

2196e/ - 204 -
Figure 120. ERM System Weights

«B

•M

*

LU

V)

a

HS

W

52

IA

«A

-J

|

1 COMPONEN

T

M)

i

0]

••«

LU

Q

«/s

O

**

In

e

o

z

ht

M

1

»B-^-___«

i

» o s

s s « OD

1

3

LU

Z

3

1 UJ

1 <0 ELEC. POWER

• BATTERIE

S

(4AKU/Hr)

to

2 eo

N. OO

•0 <4T

M ;*

in «o

ee

' 2

a

1 LU

«-

M

M

STRUCTURE

• FRAME I

SUPT

S

04

•••

ee

o

Q.

UJ

UJ

•o

.

Z

•-

M

U,

•

^^

i

s

•

.

ri

•a

a

g

rvi

IM

00

5

=

<M

frOe

a

e «*

•O «M

^4 ^0

sIM ^s

^% ^J

a. a.

a a

LU LU

O LU

1 1

s

a <-g>

o o

LU LU

« IA

,0

"°

sin

s

~

O

(V

• COMPONEN

T

PANEL

S

• INTERFACE

i

•

i

IA

O

in

O

in

=

s

(M

LU

a.

•

<° 9 • <o S i CNTL

S 1 DISPLA

Y

INSTRUMENTATIO

N

a a :

OO' I

i t

z :

LU •

O a: o i

LU *" "• « I

£S5^ J

> -• ^- a £

•*•

m m

rg fu

0 0 IA in

IA

1 a

 i

a

LU

ee

' i

•

n

&

• BLANKET

S

t COVER

S

(V

«

!•

A

X

^

A

a

X

j

.

E

«

o

n

u

j

c

SEPARATIO

N 1

0

z

u

i

OB

Z

UJ

IT

i

w•

«*••

<M

C

1

1

1MECHANISMS'

S

^

i

1

•

1SYSTE

M 1

i

„

<eOR8/E

T SE

P

a

o

(M

O

&.

s

IEC/LS

S

a

o

•

• COL

D PLATE

t HEATER

S

«*

*•

in

IA

oc

a

o

L.

a

• RM

S EN

D

EFFECTOR

S

DOCKIN

G «o

o

u

Ul

ee

a

3

LU

*"

•

<

i

3

in

IM

1

1

1WEIGMT-L

B 1

(X

oo

Kl

rv

1

u

s

1

IA

a

2

-

«M in «O f- rg

(M

00 -O O rw o *- oo oo ^- -o

»- fO ^ •-

M

IA LU V)

X •- * Z IA S Z

« ^ <

»- IA •-

IA IA • «A 3

S

c a. LU

0 X Z

a i IA 3 <A 3

O "* *" LU >~ UJ

C ^ IA Z IA Z

•o o oo rg 2

v» IAO .^SaSov S

- 205 -
Potential Demand for ERM Units

In Scenarios 1 and 2, the resupply module provides the OMV with propellant for

transportation to GEO. This mode was found to be more economical than using

an expendable OTV to transport the OMV to GEO. The propellant transfer (and

concurrent module exchange servicing, when required), is accomplished directly

from the OMV after separation from the resupply module (RM). These missions

are expected until the late 1990's. When they do occur, there will be a

maximum of three, placing three QMS into orbits where there are clusters of

satellites (over North America, Europe, and Japan). In Scenarios 3 through

10, the resupply module operates in LEO and is reusable.

- Some RM's will be parked at the propellant depot (Scenario 4) and

participate in propellant scavenging operations from the STS (Scenario

10).

- Some RM's will perform resupply and/or top-off operations:

o S/C with integral propulsion systems going to GEO (Scenario 3)

o Reusable perigee vehicles (Scenario 5)

o Reusable OMV/RM perigee vehicles (Scenario 8)

o Space-based OMV (Scenario 9)

o Top-off of RM for Scenarios 1 and 2 (Scenarios 6 and 7)

At a minimum, two ERM's are required to provide a threshold level of service.

One ERM will be located at the depot and participate in propellant scavenging

from the orbiter. One ERM will perform the propellant transfer and top-off

operations. These ERM's are interchangeable and probably rotate functions.

When one ERM is filled with propellant from the scavenging operations, it will

switch roles with a near empty ERM to perform resupply and top-off missions.

The potential resupply scenarios for the ERM program and the number of ERM

units required initially in the early 1990's, and later in the late 1990's, is

shown in Figure 121.

3.3.4 Baseline Program

The ERM baseline program includes two test articles, the developmental and

qualification test articles, so that testing can be performed in parallel.

This gives the program schedule slack and reduces schedule risk. The ERM

baseline program also includes two flight units. One of these flight units is

assembled from the salvageable system on the two test articles. The ERM phase

C/D program could start as early as the third quarter of FY 1987 and have its

first operation flight by the end of FY 1991. After two periods of docking,

scavenging and fluid transfer demonstrations, expendables resupply operations

could commence in FY 1992.

2196e/ - 206 -
Figure 121. Potential Resupply Scenarios And Estimated Number of

Expendables Resupply Modules by Year 2000

OF POOR QUALITY

PROPELLANT TRANSFER

FROM RM TO OMV FOR.MORE

ECONOMICAL TRANSFER TO GEO

4]LEO DEPOT OF

MOO RM'S

CONTINGENCY MULTIPLE S/C

RESUPPLY FROM OHV •

PLANNED RESUPPLY/MODULE

CHANGE-OUT FROM OMV FOR

TRANSPONDER HAX1MUATIOH

OPERATE OMV/RM

FROM LEO DEPOT

FOR PROPELLANT

TRANSFER OPERATIONS

\

O RM/ONV USE FOR

 v

.PROPELLANT TRANSFER

FROM STS TO DEPOT

REFUEL REUSABLE

PERIGEE STAGE

\ERVICER

TOP-OFF

COMPACT INTEGRAL

PROPUL S/C

ECURSOR I

USABLE S1AGB

1 I

s/c ::

J

REFUEL

SPACE-BASED

OMV

POTENTIAL

MISSION

SCENARIO

-v>\^

—-v^V^ *V?vJ

__^^ ^^^\

e>©

-^Q

ERN

TYPE

EXPEND

REUSE

TOTALS

ESTIMATED NO.

Of RM'S REW.

HIN

EARLY

• 1990 's

0

2

2

MAX

EARLY

1990>

3

4

7

2196e/ - 207 -
3.3.5 ERM Developmental Test Program

The WBS systems required to perform developmental testing as shown in

Figure 122 are the mechanical, fluid, electrical and avionics. The body

structure will resemble the final RM configuration because it will be

fabricated from the machine proofings preceding the machining of parts for the

flight qualified hardware.

• DEVEI QPMFNTAL TEST ARTICLE

- BODY STRUCTURE (1.3.1.1) FROM LOW COST MATERIAL RUN THROUGH MANUFACTURING

MACHINES TO VERIFY NUMERICAL CONTROL SETTING

- HECAHN1CAI n.3.7.8) - EXCEPT PYROS

- FLUID SYSTEMS (1.3.2.?)

- ELECTRICAL POWER DISTRIBUTION/CONTROL (1.3.4.6)

- AVIONICS (1.3.M.3)

• SCHEDULE

1987 1988 1989 1990 1991 I

PHASE C/D AWARD

PRELIMINARY DESIGN C I

DETAIL'DESIGN [ \

DEVELOPMENT TEST j FAB | TESTIM \

• DEVELOPMENTAL TEST OBJECTIVES

- FLUID TRANSFER TESTS

- DOCKING AND RELEASE

- PRESSURANT TRANSFER - BASIC FUNCTION, MULTIPLE REPEATS. CONNECTION RELIABILITY

LEAKAGE MONITORING

- RESUPPLY MODULE/ORBITER COMPATIBILITY TESTS

- INSTALLATION/INTERFERENCE

Figure 122. ESM Developmental Test Program

Pyro, environmental control, thermal protection and batteries are not needed

for the developmental test program. Parallel testing permits added

flexibility to the developmental test program. This test article has a

year-and-one-half slack before its systems are needed for assembly into the

second flight unit. Therefore, developmental testing could continue for

longer than the anticipated one-year period if necessary without disrupting

the program.

2196e/ - 208 -
3.3.6 Qualification Test Program

The systems required on the qualification test article are shown in Figure

123. This test program is likely to leave most sensitive systems unusable for

flight hardware. Thus, inexpensive mass simulators will be substituted

whenever possible.

• QUALIFCATION TEST ARTICLE

- BODY STRUCTURE (1.3.1.1) - FIRST PRODUCTION UNIT

- COLD PLATES/HEATERS (1.3.5) & MASS SIMULATORS

- THERMAL PROTECTION BLANKET (1.3.1.2)

- FLUID SYSTEMS (1.3.2.3)

- ELECTRICAL POWER DISTRIBUTION & CONTROL (1.3,4.6) & MASS SIMULATORS

- MASS SIMULATORS FOR MECHANICAL SYSTEMS. AVIONICS. BATTERIES

• SCHEDULE

1987 198B 1989

 199° 1991

PHASE C/D AWARD ---

PRELIMINARY DESIGN

DETAIL DESIGN - -

DEVELOPMENTAL TEST I FAB. I TES.

T

QUALIFICATION TEST » MB » TE5T i°FF LIHITS

QUALIFICATION TEST OBJECTIVES

- ACOUSTIC VIBRATION TESTS

- THERMAL VACUUM TESTS

- RANDOM VIBRATION TESTS

- FLUID TRANSFER OVER SPECIFIC OFF-LIMIT RANGES

Figure 123. ERM Qualification Test Program

The parallel testing approach to this program allows sufficient time for

off-limits testing to validate the ERM for operations outside its design range.

FIRST ERM FLIGHT HARDWARE

The first flight unit can be produced six months early as shown in Figure 124

and available if necessary to supplement the test validation program before it

is needed for flight check-out. The first operational flight will occur in

the last quarter of FY 1991.

2196e/ - 209 -
Figure 124. First ERM Flight Hardware

sp

se

^>*

r^i

EM

*

!

§*gj^

*

^=j

JgS

pjj

1^"

<C/3

JE*

^

*2

CO

z

^^

1—

1

O

5

5

^

1

OS

LU

cs»

o

LU

,-^ir

LU

H3:

CD

^^

HQ_

LU

x:

LU

* «P

LU

co

CO

en

CO LU

_J =3

«a: LU

O

1 CO

•

•••

o»

o>

••

a

o

£

•^•f

09

O

r-

^^

OO

aa

o

*"

^.

^»

OO

^

^"""

e

_ o

•^

a

u

01

e

o

a

«

o

b.

*~~P*

g^^_

«_

— g.

1 ^ ^

••••

ca co

C£ LU

<C <=J 2=

< >C£-

 C —O

ea <c LU

"^ 2: ca

3C. —1

LU — —

CO _J «C

•a: LU t— =: ce LU

o_ o. ca

en

LU

03

Ik

h—

CO

LU

t—

H5

|

LU

LU

far

at

ZZ

UJ

_S

C/3

HX

Ik

Ik

Q

HV)

IM

—

s

MM

1—

CO

LU

hO

<

LU

O^^3

•^•i

^

UJ

*~

00

0

^^

U

M

CD

^HH

LU

cn

3

cj

t~r*

^

Co

LU

^^»

CO

••••

LU

<+-

. CO

1 1

iUJ

ca

o

CO

LU

^^

_J

^ZZ

O

LMM

. .

2J

LU

\

CO

y*

LU

HCO

^~CO

o

"9*

^c

=3

O

• 1

LU

IS. <_>

—_J K =—:

^^ ^D

^^B ^J

CO LU

cx:

^•^

u- •

CO

o

H^

^1 1^ i

c_

O CO

LU

CO =>

~"

 <*~l

h- LU

t^ ^ t_ j

2= CD

"*^ g><

LU Q.

CO 2=

^** <**^

LU —

CO <— >

>- 2:

CO =3

LU

<c s:

Ij co

<C LU

1 1

2196e/ - 210 -
SECOND ERM FLIGHT HARDWARE

The second ERM flight unit is assembled from the systems salvaged from the

development and qualification test articles as shown in Figure 125. This will

avoid $12 to 18 million in additional costs. Thus, the additional cost to

assemble the second flight unit is only $4 to 6 million.

($12-18M)

SUBSYSTEMS SALVAGED FROM

DEV. & QUAL TEST ARTICLES

+($4-6M)

OTHER SYSTEMS AND

ASSEMBLY & C/0

- SECOND ERM FLIGHT

HARDWARE

0 DEVELOPMENTAL TEST ARTICLE;

-BOD Y STRUCTURE (1.3.1.1) - PROOFING/EXPENDABLE — ®

MECHANICAL SYSTEMS (1.3.7.8) - EXCEPT PIBO

- FLUID SYSTEM (1.3.2.3)

- ELECTRICAL PO¥ER DISTRIBUTION/CONTROL (1.3.4.6)

- AVIONICS (1.3.4.2)

0 QUALIFICATION TEST ARTICLE;

- BODY STRUCTURE (1.3.1.1) - FIRST PRODUCTION UNIT

COLD PLATE/HEATERS (1.3-5)—®

- THERMAL PROTECTION BLANKET (1.3.1.2)

- FLUID SYSTEMS (1.3.2.3)—®

- ELECTRICAL PO¥ER DISTRIBUTION 4 CONTROL (l.3.4.6)-PAHTIAL0 OTHER SUBSYSTEMS:

BATTERIES (1.3.3)

MECHANICAL - PYROS (1.3.7.8)

COLD PLATES/HEATERS (1.3-5)

INTEGRATE

REUSABLE

SUBSYSTEMS

INTO A

SECOND ERM

FLIGHT

HARDWARE

FIGURE 125. Second ERM Flight Hardware

3.3.7 Build and Flow

The manufacturing plan for fabricating the ERM structure is shown in

Figure 126. The planning, calculating the numerical control settings for the

machines, and proofing (running low cost material through the machines to

verify the adequacy of the numerical control settings) are only performed

once. The first production unit is the qualification test article which will

be later assembled into the second flight unit. The second production unit

will take 42 weeks to complete and will be used as the first flight unit.

3.3.8 Advanced Development and Flight Demonstration Cost

Ideally, the same personnel involved in the Advanced Development Program will

roll over to support the Flight Demonstration Program. They will also

indirectly support the ERM Program by helping to shape its specifications and

requirements.

2196e/ - 211 -
Figure 126. Expendables Resupply Module Manufacturing Schedule

.<*•

/.$

w

»

1

1 o

s

5

a

S *

1

 ~* &

- i< »

3 •*> Z "• v* "~

35

 ^ "g^o

d

d j

y?

J5*i

a*«

tfJ^tf

II

*3

ii

fay,

I

Q

•^5

JC

\4

£Sutr. Smuts

sji * 5

IS;

J ui --

3

-2

« s

J

4
There is time delay between the Advanced Development Program and the Flight

Demonstration Program while the ERM Specifications and Requirements are

defined as shown in Figure 127. If the user/customer community can be

persuaded to support/commit to the ERM program without an early flight

demonstration program, then we recommend the late flight demonstration

option. This schedule is preferred over a continuous Advanced Development

Program that culminates with a flight demonstration for two reasons:

1. If the ERM Program is deferred until a flight demonstration, the

program cannot start until the end of FY 1988 or 12 to 15 months later.

2. The systems and concepts verified on the flight demonstration may be

different than those eventually specified for the ERM Program.

TECHNOLOGY DEVELOPMENT PROGRAM

SPECIFICATION/REQUIREMENTS

DESIGN

ANALYSIS & TEST

BREADBOARD FAB ft TEST

CRITICAL TECHNOLOGY COMPONENT,

DIRECT

HEAOCOUNT

20

10

.FLIGHT DEMOMSTHATIQN PROGRAM ,

'BREADBOARD

J

FLIGHT DEMO

s_

VALIDATION OF FLUID TRANSFER CONCEPT.

FLIGHT DEMO FAB j .TEST,

B/B FAB VB/B+ TEST / * / *

DEMONSTRATION

DESIGN/FAB | TESTING FAB [TESTING'S,

PHASE C/D START OF ERM PROGRAM

^ 1986 I 1987 j «988 \ 1969 1990

30p CONCEPT VALIDATION • FLIGHT DEMONSTRATION^

TECHNOLOGY DEVELOPMENT

PROGRAM

l

1966 1987 1988

FISCAL TEAR

J989 J990

Figure 127. Advanced Development Program Supports

the Flight Demonstration Program

A rough best estimate for the total cost of the Advanced Development Program

and Flight Demonstration Program is $7M and $9M, respectively. Details of the

$16M, or cost estimate are presented in Figure 128 from manpower requirements,

labor rates, and the percentage of the program subcontracted. The labor rate

includes direct and fringe benefits, overhead, and G&A. None of the cost

estimates include subcontractor fees. The labor rate is uncertain because of

the level of Indirect support allocated to overhead.

2196e/ - 213 -
TECHNOL06T DEVELOPMENT MO FU16HT OCMmSTMTIOH

DIRECT HOURS

LABOR RATE

I SUBCONTRACTED

SPAS AND TANKS

LOW BEST ESTIMATE HIGH

133.400 HR -101 148.200 HR »20I 177.800 Hft

J50/HR J55/HR J60/HR

301 401 SOX

9.5M

1.9ft

13.6M

2.5M

21.3*

3.4H

n.4fl 16.1H 24.7N

TECHNOLOGT DEVELOPMENT

DIRECT HOURS

LABOR RATE

X CONTRACTED

TANKS

TOTAL

LOW

51.900 HR -10X

J50/HR

30X

3.7M

1.2*

4.9M

BEST ESTIMATE

57.700 HR »20X

J55/HR

40X

S.3M

1.6M

1.9*

PROGRAM

HIGH

69.200 HR

S60/HR

SOX

8.3*

2.4*

10.7*

DIRECT HOURS

LABOR RATE

I CONTRACTED

SPAS

TOTAL

LOW

81.500 HR

SSO/NR

30X

5.8*

0.7!)

«.S»

aiSHT DEMONSTRATION

BEST ESTIMATE

PROGRAM

HIGH

10X 90.500 HR »20X 108.600HR

J55/HR {60/W

401 SOf

8.3M

0.9*

«.2«

13.0M

l.OM

14.0*

Figure 128. Cost Estimate for Technology Development

and Flight Demonstration Programs

If the specifications and requirements for the critical technology components

do not change in the ERM Program, then these components need not be

redesigned. The cost of the Flight Demonstration Program will be $4M less if

it can use the components from the Technology Development Program and "borrow"

systems/parts from the first ERM flight unit. Thus, if the ERM specifications

and requirements do not change, the total cost for the technology development

and flight demonstration is |l2M.

3.3.9 Baseline Program Cost

The ERM baseline program cost was estimated by two separate and independent

methods: Parametric and detailed engineering. The purpose was to confirm our

cost estimate for using two methods.

Parametric Estimate for ERM Program

The parametric cost estimating relationships for each subsystem (except

software) depends on weight, a factor for degree of design modification (a new

design has a factor of 1.0), and design complexity relative to a baseline

program for which costs are known (a system as complex as the Orbiter for

instance would rate a factor of 1.0).

System DDT&E

System Production

Cost Estimate

Cost Estimate

A * Nd * Dc * (weight)3

C * PC * (weight)0

2196e/ - 214 -
The high parametric cost for each system reflects the worst combination of

weights and complexity factors and the low parametric cost reflects the best.

These costs are presented in Figure 129.

ORBITER

was

1.3.1.1

1.3.1.2

1.3.7.8

1.3.2.3

1.3.3

1.3.4.6

1.3.4.3-5

1.3.5

COMPONENT

BOOT STRUCTURE

THERM PROTECT BUT

SEPARATE/MECHANISM

FLUID SYSTEMS

ADV. BATTERIES

EL8C PUR DIST/CONTR

AVIONICS

COLD PLATE/HEATERS

1.3.1.1

1.3.7.8

1.3.2.3

TRUSS STRUCT

SEPARATION BOLTS

HEUIUM TANKS

SOURCE OF COST

ESTIMATING

RELATIONSHIP

ORBITER ( FACTORS

SPACE STATION

ORBITER t FACTORS

ORBITER t FACTORS

NtHj QUOTE

SPACE DIV CER

•83 PROCURE

SPACE STATION

ORBITER I FACTORS

ORBITER t FACTORS

•83 PROCURE

HEIGHT

RANGE (Lfl

2534-2789

250-275

201-241

3721-3842

800-960

360-504

291-407

100-120

SUBTOTAL

148-163

36-43

160

SUBTOTAL

ftOTCE COST

(i OF 84JJ

7.7-24.2

.1-.3

.3-2.8

6.8-21 .2

j.a-4.5

.5-1.7

1.2-1.5

.8-1.6

21 .2-57.9

1.9-6.1

.1-.1

.3-.3

2.3-6.5

SOFTWARE I.1-.2

TOTAL |23. 6-64.6

HEU DESIGN

FACTOR

1.0

.2-.3

.2-.3

.4-. 6

1.0

1.0

-

1.0

1.0

DESIGN

COMPLEXITY

.1-.3

.1-.3

.1-.3

.25- .5

PURCHASE

.1-.3

PURCHASE

.1-.2

. -.3

PURCHASE

PURCHASE

THEORETICAL

FIRST AMT

<H OF 84S1

5.0-7.7

.2-.3

1.3-2.2 '

4.1-5.9

1.9-3.2

1.1-2.0

1.2-1.5

.8-1 .6

15.6-24.4

1.4-2.2

.1-.1

.3-.3

1 .8-2.6

17.4-27 iT

PRODUCTION

COMPLEXITY

.4-.6

.S-.7

.5-.7

.S-.7

.S-.7

.S-.7

PURCHASE

.1-.2

.4-.6

PURCHASE

PURCHASE

I

Figure 129. Parametric Cost Estimate for ERM Program

To account for special features in the ERM baseline program as given in

Figure 130, several cost elements must be added to the low estimate: $15.6M

for the first flight hardware and J5M to assemble the systems on the

developmental and qualification test articles into a second flight unit.

Since the range of factor values in the cost estimating relationships are

defined over a range with an 80Z confidence internal, the high $89M and low

$44M parametric cost estimates also approximate an 80% confidence internal

range.

2196e/ - 215 -
• MILLIONS OF 84 DOLLARS

ESTIMATE

DDTlE

FIRST RIGHT HARDWARE

REFURBISH DEV I DUAL TEST ARTICLES INTO SECOND

FLIGHT HARDWARE

TOTAL COST RANGE

PROBABILITY THAT THE ACTUAL COST HILL LIE HITH1N

THIS RANGE IS 80Z

PROBABILITY THAT THE ACTUAL COST HILL BE LESS

THAN THE LOW ESTIMATE

PROBABILITY THAT THE ACTUAL COST HILL EXCEED THE

HIGH ESTIMATE

LOH

$23,5

$15.6

$ 5

W4 •

HIGH

$64.6

$21.1

INCLUDED

$89.0

101

10Z

Figure 130. Parametric Cost Estimate - Basline Program

Detailed Engineering Cost Estimate for the ERM Baseline Program

This cost estimating method is based on a program funding curve that has a

definitive start, a slope rising to a peak and falling thereafter, and

concluding with a definitive finish as illustrated in Figure 131. The ERM

program funding profile was estimated from direct personnel expected to

support this program. With a direct dollar labor rate and an estimate for the

portion of the program that is subcontracted, the total cost was estimated at

$64M (Figure 132).

The total direct hours, the direct labor rate, and the percentage of work

subcontracted is uncertain and their estimates vary over a range of likely

values. The set of worst and best conditions determines the high-and-low-cost

estimate, respectively. The combination of high estimates for labor hours,

labor rate, and portion of the effort subcontracted results in a high cost

estimate of $100M for the ESM program. The combination of low estimates for

labor hours, labor rate, and portion of the effort subcontracted results in a

low-cost estimate of $45M for the EBM program.

Expected Cost - Best Estimate

The cost estimate range for both the parametric and detailed engineering

methods were calcuated so that there was only a 10% chance that the actual

cost would be outside the ange, either above the high or below the low. Thus,

there is an 80% confidence/probability that the actual cost will be between

2196e/ - 216 -
FISCAL YEAR

SCHEDULE ITEMS

PHASF r/n AWARD

PRELIMINARY DESIGN

DETAIL DESIGN

lE\aj(PENT TESTING

QUALIFICATION TESTING

FIRST FUfflT ARTICLE

IKSTfli, 00, 1NTEG &

TESTING

FIRST OPERATIONAL FLT

FIT (PS & SUPPORT

1987

< i

/

19

ol.M'M"

\

^v^

1 FAB 1ES

38

CDR

t

SUSTAINING

r HARDWARE "\^

I FAB

1989

ENGINEERU

TESHNG^V^

1

1990

6

TESTING"*^

FAB

1991

3

^KfUPB

\c^ip\

\

r

1992

•

> FOF

' 1

DIRECT tEADCOJNT

Figure 131.

Detailed Engineering Cost Estimate for ERM Program

the low and high cost estimate. The probability distribution between the high

and low cost estimate is spread by a Beta Distribution. A 65/35 Beta

Distribution was selected to reflect the fact that the cost distribution is

skewed upward. The expected cost is the estimate at the 502

confidence/probability point by definition. Both the parametric ($62M

expected cost) and the detailed engineering ($66M expected cost) methods led

to similar cost projections as shown in Figure 133. The midpoint of these two

expected cost projects - $64M - is our best estimate for the cost of the

baseline ERM program. Thus, there is a 502 chance that the program could be

less than $64M, but there is also a 502 chance that the baseline ERM program

could be greater than $6AM. Furthermore, there is only a 102 chance that the

baseline ERM program cold be completed for $45M, and there is only a 102

chance that the baseline ERM program could cost more than J95M.

2196e/ - 217 -
• MILLIONS OF 81 DOLLARS

TOTAL DIRECT HOURS

Z EFFORT SUBCONTRACTED (MATERIAL &

PURCHASED PARTS)

DIRECT LABOR RATE (FRINGE BENEFIT &

OVERHEAD & G&A)

TOTAL ERM PROGRAM COST

RANGE COST ESTIMATE

LOW ESTIMATE

-10X

311.760 •*-

65Z

J50/HR

TM.7mfcn _ „-

'_" - *LH

.35

BEST ESTIMATE

- 346,400 -

701

J55/HR

y&.utrfez _ r,,-

.30

HIGH ESTIMATE

+20X-

—415.680

75Z

mm

415.fi80xGO,irrc =TMTfllM S1O*I

.£>

Figure 132.

Detailed Engineering Cost Estimate Range

The $64M best estimate for the ERM baseline program was allocated to its main

elements so that other program variations can be costed. Some other program

variations are offered in Figure 134. For instance, if the first flight

hardware is removed from the baseline program, the expected cost would drop by

nearly $18M, down to $46M. The first operational flight could still occur

that last quarter of FY 1991, if the off-limits testing of the qualification

test article was eliminated. The flight article would be assembled from the

system salvaged from the developmental qualification, and flight demonstration

test articles. This flight article would be available for functional testing

and check-out. On the other hand, the production of flight units 3 and 4

could be included in the baseline program for an additional $35M and a total

program cost of $99M.

The time-phased funding profile for the expected program cost of $64M was

calculated from the detailed engineering cost estimates. Peak funding occurs

in FY 1989 at $20M as shown in Figure 135.

ERM Production Cost

A 91% learning rate is possible for follow-on production units. The $35M cost

estimate for production units 3 and 4 assumes they are integrated with the

baseline program so that material part/systems procurement, manufacturing

set-up, and learning are not lost. Again, costs are estimated over a range

from high to low. These costs were estimated parametrically in Figure 136.

2196e/ - 218 -
100

90.

CONFIDENCE THAT 80.

BASELINE ERM

PROGRAM COST

HILL NOT EXCEED 70.

A GIVEN $ LEVEL

60.

SO.

40.

30.

20.

10.

PROBABILITY

••

PARAMETRIC EST

EXPECTED COST EST.

DETAILED ENGINEERING

JEA DISTRIBUTION 65/35

^ *•— ' THROUGH UIOi/LOV RANGE

. DETAILED ENCmERINC EST.

lo 20 30 40 60 70 80 90 100 110

BASELINE EM PROGRAM COST

(M M of M$)

r THE EXPECTED COST 101 1KB BIN BASELin PROGRAM

Figure 133. Confidence/Probability of Reaching Baseline

ERM Program Cost Target

3.3.10 Cost Estimate for Satellite Mod Kit to Enable Resupply

This cost estimate was computed both by parametric and by detailed engineering

techniques as shown In Figure 137. The parametric cost estimates were

calculated by adding the scar weight of the mod kit to a 2,000 pound satellite

and computing the Increase In cost from the cost estimating formulas.

There Is reasonably good agreement between the parametric and detailed

engineering cost estimate range for DDT&E. However, the production cost range

estimated by the parametric method is low because the additional parts in the

mod kit Include expensive quick disconnects and valves that are more expensive

than the average plumbing parts in the cost estimating formula. Thus, the

detailed engineering cost estimate more nearly represents the true cost of the

mod kit.

The expected DDT&E costs estimated by parametric and detailed engineering

methods, is between $1M to &1.25M as shown In Figure 138.

2196e/ - 219 -
ERM BASELINE PROGRAM

• DDT&E

• REFURBISH SYSTEMS ON DEV & QUAL TEST ARTICLES INTO A SECOND

FLIGHT HARDWARE

• FIRST FLIGHT HARDWARE, C/0 FUNCTIONS £ SHUTTLE INTEGRATION

• GSE & PAYLOAD SUPPORT

TOTAL EXPECTED COST

ERM EXTENDED PROGRAM

• PRODUCTION OF FLIGHT UNITS 3 & H

• MISSION OPERATIONS & TRAINING

t MAINTENANCE & REFURBISHMENT - SPARES

• PROGRAM MANAGEMENT & SUSTAINING ENGINEERING (LESS THAN IS)

EXPECTED COST

llfl

SH

18*

641

EXPECTED COST

35fi

NOT COSTED

NOT COSTED

NOMINAL

Figure 134. Cost Breakdown for the ERM Baseline Program

The production cost range, estimated parametrically, was assumed to reflect

the lower half of the parametrically confidence interval (below 50%). This

resulted in a new upper cost estimate of nearly $1,000,000 instead of $770,000

originally calculated. Still, the spread between the expected costs estimated

by parametric formulas and detailed engineering is wide, $770,000 to

$1,240,000, reflecting more uncertainty in this cost estimate. We believe the

true cost is towds the upper end, however.

Summary

If the user/customer can be influenced to support/commit to the ERM program

without an early flight demonstration, we recommend that the flight

demonstration be incorporated within the ERM program and fly a year later, if

at all.

2196e/ - 220 -
ANNUAL FUNDING

(in 8 Of 84 1)

20 '

18 '

16-

H12-

10-

8 -

6-

4.

2_

Ifi

1

 15M

20M

17M

IOM

1H

I

FISCAL YEAR "87 1988 1989 1990 1991 1992

Figure 135. Time Phased Cost Estimate for Baseline ERM Program

The expected cost of the flight demonstration program can be kept to $5M if:

1) performed in the orbiter payload bay, 2) the MPESS and SPAS structures are

used, 3) all major systems are salvaged for reuse on the first flight unit,

and 4) the critical technology components don't need to be redesigned.

The ERM baseline program consists of two test articles (development and

qualification) and two flight units.

The second flight unit is assembled from major systems salvaged from the

development and qualification test articles, thus cutting program costs by $12

to 18M.

The expected program cost is $64M with a peak funding of $20M in 1989 if the

phase C/D starts in the fourth quarter of FY 1987 and the first operational

flight occurs in the last quarter of FY 1991.

The average expected cost to produce additional ERM's is $18M per unit.

2196e/ - 221 -
ORIGINAL PAGE IS

QF. POOR QUALITY

UNIT PRODUCTION

(IN MILLIONS OF

30

25-

20.

15.

10.

9.

. 8.

5 7.

5.

^ .

•

• 11 . •

*- —

2=

1

ta

i .

:' L

.-

t •

"i

. t

H-,

r

:

:'

rtli.ii

*'\ '' • •il- :

Mil '

;•jH .....

i-piKj-i

"]$]<•

tJGJ^

4 1

•1

i.

i

{lisj;1

T

|

4-

•— >.

•

e

- — ;

'•' .'

':• '.—

•

';:'i

=i.'

I

- i

>• >.

.:•' —

^ »

•.-.r;.

•^r

:

^UT-:-

..."

•-•«

-

^^

-.

.-

"Jr.

'u~

1

 "•

^^

>•«•

•

;:H

!

ii'

:

t::

9M

TTT

,:.:

:~

:i:

„

M,

-„:

^

1

•i

M

1

= :

• .

H

LI

-

•.J

"•

' •.-:•"

1

1 :,

I6H CO.

M

1

3W

%

r~:

1

cc

b

r*

IS

—

5T ESTI

. . -IV

r ESTIM

.••-=•; ::

••'-:.-":

• :.-.:_.:.

.-:••;

EXPECTED COST OF

PRODUCTION UNITS 2&3

IS *35M

*

l

.f*!rs""*i

:_ il. '^3: r^t

Kli

l|::

t»-'

1:^

•P*

2'TOIRD PLIfflT POU1TH

|3^

h--''

HARDWARE

r

ir

ij

R.

TTM

',:

'•i?

r

IQ

•v

 =

UT

HARDUA8E

-=PT*1 1

= ;'-::

"- H:

-:

"i1

A

:i:

:L: —

.:

. .« :

• » . it •

i "^; !

. . ...

. ^-^ .._

•••»£?

***i* '

MATE

ATE

."T:;.

:•-•*-

•?=;••.."

in

f

li

:

:

JPJ.:

1!i;

.••,-::.!

— ••

I.S 2 2.5 3 4 5 i 7 8 9 10 1.5 2 tm pmncnoM ODAMTIIT

Figure 136. ERM Production Costs

2196e/ - 222 -
DDT8E

PRODUCTION

COSTS FDR

TO) InJ

PARAMETRIC

LOW

DESIGN Ctm£XITY=,3

3,3fi x .3

= jptfl-nm

HIGH

DESIGN COMlfXIT¥=.5

3,3fi x ,5

* $1 .f?fl.nm

COST ESTIMATING RELATIONSHIP

EXTRA PLUMBING COST ESTIMATE FROM

- USAF UNMANNED S/C

GRAPPLE TRUSS, DOCKING FIXTURE &

VALVE PANEL

- NASA-PRC

PRODUCTION

COMPLEXITY

.7

,77fi x .7

• $510.000

•

PRODUCTION

COMPLEXITY

1.0

.77fi x 1,0

= $770,000

DETAILED ENGINEERING

LOW

*50A*Yin.WTO ^

,75

= $690,000

QUICK DISCONNECT

5x75.000=$375.000

VALVES

9xW.OOO=$360.0CO

ELECT. CONNECTIONS

2x«.000=$80,000

GWPPLE FIXTUE i

P/V&S

$160,000

ASSY jim.nm

TOTAL $1,075,000

HIGH

$£50A* x IS.Gnn n»s

.60

= $1.560,000

QUICK DISCONNECT

5xlOO.aXH500,000

VALVES

9x50.COHt|50.000

ELEQ, CONfECTIGNS

2x50,000=$100.000

GWPLE FIXTURE S

PAfaS

$200,000

$15Q.fflQ

TOTAL $1.WO,000

Figure 137. Cost Estimate for Satellite Mod Kit to Enable Resupply

2196e/ - 223 -
CUM pioe

'.or

.9

.1.

.7,

.4

.5.

.4

.3

.2

.1

0

Distribution

65-3S

Detailed

Engineering

Expected Cost

Estimate

Parametric Est.

Expected Cost

$1,020*000-1,240,000

800 1000 1200 HOO 1600

00 TIE COST

(In 1000't of 84 S)

Beta

Distribution

SO-SO

400 600 800 1000 1200 1400

MODUCTIOM COST (TFU)

(in 1000's of 84 S)

Figure 138. Expected Value Cost Estimate for Satellite Mod

Kit to Enable Resupply

2196e/ - 224 -
Appendix

- 225 -

I785b
<p«rm Tr»n»oor*ttoi» ft :

- * MM IntOTiatfonai Corporation

12214 Ukowood Boulavmrt . . „ ,

y. c«iifomia 90241 International

EXPENDABLES RESUPPLY MODULE

REQUIREMENTS DEFINITION DOCUMENT

226
C/5 caUJUJ00ooUJ

S REQJ

<JI

f

Si*. UJ _J O LU U

J

>glo^

2fh£

«C O Uw

£ux2

O O O

STANDARD RECE

UMBILICAL I/FMONITOR RESUPOPS, STATUS &

MEASUREMENTS

CEIVER

<

o a (S

|

:S

O 0 R RECIP c! o

CKING

/TIME

TRAINED -LROL o

*S3

oS 8tuUDES DOCKING NICATIONS "53

C/?

TER P/L

ORT

REQUIRED

OPROVIDE LAUNCH ft

RETURN CAPABILITY1

OPROy/IDE POWER SEP

ft RMS HANDLING

OCONTROL PANEL

FLU ID TRANSFERft MONITOR

en5i

ku <

NG

C

oT

ft

OC

e§

iV)s hi-5 U Ul V> o e o 1

COMPATIBLE H/GS

FLUID LOADING ft

DRAI

ING

UID FILL ft SERVICING

REQUI 0 C

T

L

ON L

PANE

CATIO

CHECK

ION

ROL

MIS

CON

CO

SYSTEM

TOR

JNICATION LINK

TMV COm SYSTEM

</> UJ13

II

ii
o

Q

0

to i—T—i «

1

i

1

1

Oiu 1

MMIIMM

M ]

u I fcUikMMri 1 SKIIMI |

M

5 5

ll

fi1*

•

o

•

1 1

•

s

1

2

5

— 1 1

1 • 2 $ 3,1 I

in

"i

>nmVPl

228
c

o

1.2

0

tt

c

LU

ORIGINAL PAGE IS

OF POOR QUALrTY

IIMO

s»

i22?«

w**<

551

u,^

BTM«

:«*

isawj

*

K§

ill

I* MsW

§s

>a

41 ONIITf

R

AHACN

MECHANIIMI

I.I I

•Ml CN

O

CFFECIORI

4 OOCKINt/

UMIIIKAL

ACTUATOR!

1 MIIIIO

N irtciri

c

MICHAHISMf

It

lETAMTM

N

rVROI

JT

- S

i!

S 9

• *

=S

ii

aw

^ **

M

2=S

5~9E

•e«

mjj'«"s

M*

IW

2i

ts

»«

N

^

«

»!

^

«W

» i L_J

«•

J^^

laui

m

ri

t.t i

•EOUMTNWa

•IIIRMUIKM |

5

••

14

COMMUNICATION

INIEWACEI |

-I - - L

t.t

INERMAL

CONIROll

10114 f

IMUVfU

M

• 1

1 1

tf\

I

721
1.0 AVIONICS & POWER SUBSYSTEMS

1.1 Batteries

1.2 Regulation and Distribution

1.3 DPS

1.4 Communications Interfaces

1.5 Thermal Controls

1.6 Actuators and Pyros

1017./5
1.0 AVIONICS^ POWER SUBSYSTEMS DEFINITION

The avionics and power subsystems shall provide hardware and

software necessary to perform and control the following functions and

operations:

o Communication Processin

o EPDC

o Fluid Transfer Operations

o Data Processing

o Active Thermal Conditioning

o Umbilical Attach and Detach

Resupply mission functions which are performed by ERM system

interface elements are:

Communication - OMV subsystem

CN & C (inc. docking) - OMV subsystem

The avionics/power subsystem requirements are shown in tabular form

in Figure la.

1017e/6
1.1 Batterie*

s

The power generation system aboard the E8M shall provide the

electrical power requirement* for all operational phases of the resupply

mission. During quiescent phases of the resupply mission power may be

provided in the powered down mode.

Electrical power will provide for the following subsystems: data

processing, active thermal control, actuators, and fluid storage and

transfer.

When in the cargo bay power will be provided from the Orbiter

through the cargo interface.

Backup power to the ERM could be provided by the OMV carrier

vehicle thru an existing power interface necessary to power the OMV

docking camera and lights mounted on the ERM.

10l7e/7

232.
1.2 Power Regulation & Distribution

•

i

The Power Regulation and Distribution subsystem shall be capable of

regulating both Internal and external power generation. The autonomous

Internal power generation shall provide for an operational and quiescent

•ode. The quiescent power down node shall provide for system and status

monitoring and provide for active thermal control during non-transfer

operations. During the operational phase, power will be required to

operate valves and actuators, DPS and thermal control during the

transfer.

1017e/8

233
1.3 Data Processing Subsystem

The data processing subsystem shall provide hardware and software

capabilities for system performance monitoring on board and ground

checkout subsystem sequencing and master timing.

The software shall be capable of supporting in conjunction with GSE

software pre-mission system and interface checkout. The DPS shall be

able to request and accept digital data, process allorlthms and generate

necessary output commands for all ERM subsystems.

Data Management Subsystem, Section 2.0, will further elaborate on

the on-board data management and interface requirements.

10l7e/9
1.4 Communication Interfaces

V

The Communications Processor shall provide the interface with the

Carrier's Communication and Tracking Subsystem via the DPS. It shall

process ground command* and format telemetry data for communication

uplink and downlink. The required RF signals shall be compatible with

the Carrier's telemetry transmission and/or reception links including

direct link and TDRS/DOMSAT relay.

The communication processor must also provide interface with the

Orbiter's communication and tracking subsystem via a payload interface.

It shall be compatible with the Orbiter's telemetry transmission and/or

reception links.

Certain operational functions or sequence initiation may require

command control or monitoring by an operator. Communication coverage

regions and times must be understood prior to critical ERM operations.

1017e/10

235
1.5 Thermal Control

The them*! control subsystem is conjunction with the passive

thermal protection system shall maintain within acceptable limits

through all mission phases, the thermal environments, plume environment

and/or temperatures of all subsystems and components. This shall be

accomplished by utilizing techniques such as thermal insulation,

coatings, thermal isolators, heat sinks, heat conduction paths,

radiators, and heat sources. Prelaunch and post-abort landing ground

thermal conditioning of the vehicle shall be employed when possible to

assist the thermal control system in performing its functions.

Numerous subsystem alternatives, exist for thermal control of earth

storable propellents. The most common are thermostatically controlled

heaters, insulation, coating, isolators and heat sinks. Typical

temperature limits for all mission phases are 40 ? to 120 F.

Cryogenic thermal control is more difficult to resolve. Numerous

concepts have been developed for thermal control of cryogenic fluids.

Concepts that employ integrated structure and cryogenic insulation with

oultllayer radiation shield is a generally accepted approach. Super

insulation may be used in evacuated or purged systems. However, current

studies indicate that the uncompressed evacuated system has the lower

coefficient of thermal conductivity.

10l7eAl

236 //
1.6 Actuators & Pyros

The actuators and pyros shall control docking and unbilical

connects and disconnects. Electro/Mechanical actuators control the

fluid distribution and isolation system for transfer and safing

operations. The pyros can insure emergency separation of failed stack

docking and mechanical interfaces.

See Section 4.3 and 4.5 for further requirements on actuators and

pyros, respectively.

1017e/12

237
ORIGINAL

OF POOR QUAtrnr

CO

POWER GENERAT

ME

N DETACHE

D

ORBITE

R POWER

ER

REGULATE

S & DISTRIBUT

E

BATTER

Y POWER TO 'THE

DPS, VALVES, ACTIVATO

R

THER

M NTROL ELEME

N

DURI

N ESCEN

T R

OPER

L MODE

S

SUP

P OWER TO

RECE O

R T

OP

S

S

P

VE

S

C

O

QUI

TION

A

IE

S P

VE

R F

N

G

A

T

LI

IV

R

E

E

PA

C

T

S

UR

SP

PROTECT

S S

SUBSYSTEM

S

PLUME ENVI

5 Sg en- u.0-

oao < x o

Si§0a,,£5

-5P^=A_J'"i

o o

CE RFO

R

en d s

i

!££ , — en

CO

c/

>

c/

UJ O

CO

S*d£S*£S

00 OO O

uj

en

si

S^ i so<u • Z O 01 LU UJ

« PuenSts

UJO I I I I I

OO

£

>

2

00

UJ

_JOL a.

u. SUE

O OO

en

gen

o 2

j^ Uj — HU ^ O en 3 en

O Z a. i _; > O — uj Z u. en

OOP C O

ad

§ en

CO

UJ

en

a: =

a: s

UJ I—

2 tn

a.5

CM

en

S

cr

en

g

UJ

CO

03

CO

UJ

-J

VJ

\JL
2.0 ON BOARD. DATA MANAGEMENT

2.1 Sequencers/Control

2.2 Measurement/Monitoring Sensors

2.3 Interfaces

2.4 Manual Override

1017«/13
2.0 ON BOARD. DATA MANAGES NT SUBSYSTEM

The data management subsystem provides control of all events and

sequencing in the RM itself down to the component actuation level.

System and status monitoring fault detection and automated

malfunction procedures provide for assurance of subsystem health, while

maximizing mission success probability.

In addition, it controls the data uplinked and downlinked via the

OMV related to communication of engineering data, status data and the

automatic or man-inthe-loop control data relayed through the system met

by way of TDRSS and DOMSAT for extended coverage.

The Data Management Subsystem intefaces with the Receiver to

monitor and control critical fluid subsystems.

Figure 2a shows in a tabular format the onboard data management

subsystem requirements.

10l7e/14
2.1 Sequencers/Control

System and status monitoring will check the general health and

safety of the ERa during quiescent and operational phases. Leakage for

propellents and pressurants, and system pressures and teaperatures will

be statused as being within operational limits. During quiescent

periods constant confirmation and computer self monitoring should

determine system safety, heater operation, and fluids and components are

within acceptable limits.

The operational status of the fluid transfer system should be

verified upon disconnect engagement by the leakage indicator, pressures

at critical junctures, valve position indicators, temperatures of the

fluids and components. Upon disconnect engagement leak checks should be

performed. Electrical path checks should be verified with the

electrical umbilicals. During system operation, flow, pressure drops,

pressure and temperature measurements will be used to determine

satisfactory system performance and the transferred rate and quantity.

Durln transfer using the ullage recompressln method for hydrazine,

ullage temperatures must be continuously monitored to control the

adiabatic compressure heating of the ullage to within acceptable

limits. Sufficient instrumentation should be added to provide override

protection and process termination in the event of a malfunction.

The completion and safing monitoring of the transfer process will

require sufficient pressure and leak detector measurements to evaluate

the purging, depressurizlng and donating actions. Prior to system

safing sub-system and component valve positions, pressurs and

temperatures need to be verified.

The process of fault detection resulting in system shutdown should

be automatically controlled by an on-board computer. Computer

.controlled and monitored operations for transfer operations and fault

1017e/15
detection with autooatic shutdown will provide the aost rapid response

capability for remote operations. The operator monitoring the activity

can then access the situation and determine to restart the transfer

operation.

1017e/16

•24 Z
2.2 Measurement/Monitoring Sensors

The instrumentation subsystem shall sense, acquire, condition,

digitize, format and distribute data from the RM subsystems for

telemetry, checkout, performance monitoring and GSZ. Subsystem

measurements include temperatures, pressures, strains, acceleration,

currents, voltages, events, flow, position, quantity, rates and

electrical power.

Tank quantity gaging and flowmeterlng are the two principle means

of determining the quantity of propellant transferred during refueling.

Although tank gaging does not directly measure the transferred amount

(but infers it from tank quantity), it has the advantage of directly

verifying total tank load, which is important for mission confidence. A

flowmeter, by itself, cannot do this unless tank load is accurately

known prior to transfer. In the case of a spacecraft returning from its

mission, this is not known unless a tank gaging system is provided, or

the spacecraft tanks are off-loaded to zero quantity before transfer.

PVT is the only state-of-the-art zero-gaging method which has

reached an acceptable state of hardware development and flight

experience.

The various resupply fluids and transfer systems would warrant the

need for both a tank gaging system and a fluid transfer measurement.

Integrating flowmeters are the only method available to measure fluid

transferred if a quantity gaging system was inoperative (i.e., during an

ullage transfer fluid resupply operation). Since an RM should be able

to resupply fluid utilizing any of the fluid transfer methods, it is

recommended that both types of propellant measurement systems be

incorporated.

10l7e/17
2.3 Interfaces

The receiver satellite functions and subsystem status will need to

be monitored.

The satellite control center will configure the receiver satellite

for resupply operations. Confirmation of the receiver satellite systems

status (e.g. attitude control) will be required prior to initiating the

transfer operations.

Prior to fluid transfer a fluid quantity status of the receiver is

necessary to know resupply quantity requirements. During transfer

operations critical temperatures and pressures aboard the Receiver must

be monitored for accurate resupply quantity determination and to prevent

excessive adiabatic heating during an ullage recompression transfer.

See Section 3.OH) for further description of receiver satellite

requirements.

For transfer of OMV bi-propellants, a DPS interface will be

required to measure critical pressures and temperatures and status the

various subsystems.

1017e/18

i
2.4 Manual Override *

The operation of Che resupply system is envisioned Co include

operator controlled or monitored activity at an Operators' Control Panel

(OOP). The OOP would have function indicators, position indicators,

some critical pressure and temperature readouts for monitoring system

health, pressures, temperatures, flows and quantities at the various

mission phases. Emergency operation termination should be provided at

the OCP which would safe all fluid flow circuits.

1017e/19

-ZH6
S ERS SE

UB- ATI ALLOPE E

S

TO MEASSYSTEMS

S

Ift

oEMERGENCY OVE

CAPABILITY

CO

LU

£23

S^SS

£222

^il

w

U.5SJH

juj a: —-

<MO

Fjfil

fil o o a <l

x

COUJ

>en (/>«i

!Si | &S£

o o

a < s a.

a o o

a o o o

«»u*-:

is.

a a

S

yj CS

u. LU a:

u LU-—• a o o o

at w S

w

o o a

c/r

coUJ UJ

•Ss*3

M«>

<£:

c/j e §>LL L^ az u

UJ

CO

CO

CO

oCM

<NLUJLL
3.0 RESUPPLT SYSTJM

3.0-A Fluid Delivery & Resupply Operations

3.0-B Tanks &

3.0-C Contamination Control

3.OH) Receiver Spacecraft Design Considerations

3.1 Tanks & END

3.2 Lines

3.3 Valves & Vents

3.4 Interfaces Dnblllcals

1017«/20

2V7
3.0 8ESUPPLY FLUIDS SUBSYSTBi INTRODUCTION

Besupply fluid subsystem requirements are very auch dependent upon

Che nature of the resupply fluid, mission objectives, receiver

resupply requirements and the fluid acquisition system.

This section will discuss the operational flow of the fluid

transfer storable and cryogenic fluid transfer techniques, pressuranc

transfer techniques, propellant management devices (RID), contaminant

control, Receiver spacecraft Interface requirements.

Figures 3a, 3b and 3c summarize in tabular fora the resupply fluid

subsystem requirements for OMV bi-propellant, hydrazine and

super-fluid Helium transfer, respectively.

To effectively discuss the resupply fluids subsystem requirements,

a narrative format vas chose to present this section.

1017«/21

2V*
' 3.0-A Fluid Delivery and Resupply Operations

This section will cover the unique requirements for various

re sup ply fluids including, nonopropellants, bipropellants pressurants,

cryogenics and special fluids.

Resupply scenarios include ERM ground and space basing. Space

basing could require fluid or tank changeout at the Orbiter or at a

Space Station. Orbiter scavenging of OMV bi-propellants or MPS

cryogenics residuals would make the KM a candidate for those fluids.

OMV bi-propellants could be transferred through an RM (without tanks)

to satisfy receiver bi-propellant candidates.

10l7e/22
3. 0-A.l S tors bl«.Fluid Acquisition and Transfer

Methods to be considered for resupply of storable propellants

should include: ullage recoopression, ullage displacement, and ullage

vent/repressurlzation. These should be considered on the basis of

resupply fluid and ullage gas, mechanical and operational complexity,

time and safety considerations. Each refueling approach is discussed

in the next few paragraphs. Highlights of these methods are

summarized and schematics provided in figure 3d.

- Ullage Recompression

In this approach, the pressurant gas either remains in the

propellant tank or is returned to gas storage bottles during the

propellant resupply. Propellant is transferred against an increasing

ullage pressure as the incoming propellant compresses the ullage gas.

This creates adiabatic heating which must be regulated to permit heat

dissipation. This method is primarily applicable to resupplying

systems operating in blowdovn. The propellant will most likely be

resupplied from a regulated pressure propellant tank system.

Therefore, the quantity of propellant transferred can be determined

using a tank quantity gaging system. It will be necessary however to

know the initial load in the receiver tank so that the quantity to be

transferred and the final load is known. Therefore, receiver tanks

should also Incorporate a tank gaging system.

- Ullage Vent/Vacuum Fill

This type of propellant transfer will be necessary on regulated

pressure systems with tanks that do not have the capability to

position the ullage bubble. This is an undesirable method of transfer

1017e/23

1.60
because it will probably necessitate a hazardous waste scavenging

system be included in the resupply module/extended mission kit. With

this method, the receiver tank is vented to vapor pressure, probably

to a scavenging system, before transferring propellant. The required

amount of propellant is then resupplied, controlled either by flow

••

metering or a supply tank quantity gaging system, at a pressure that

need only be great enough to overcome the frictlonal losses in the

refueling lines. The tank is then pressurized collapsing any

propellant vapor bubbles trapped in the propellant acquisition

device. The critical issue is ga-liquid phase separation to preclude

liquid venting. The transfer control system is more elaborate because

of variable temperature/ variable pressure conditions. An alternative

would be to fill and vent at contant receiver ullage pressure.

- Ullage Transfer

The ullage transfer resupply method will beused for resupplying

pressure regulated propellant tanks that have the capability of

locating the ullage bubble. This is a constant pressure transfer

during which the ullage gas is vented off the receiver tank as

propellant is pumped in. The ullage is then transferred to the supply

tank to replace the transferred propellant. A system of this type

will require a propellant pump to overcome the transfer lines

frictional losses. A PVT propellant gaging system will not work on a

system of this type. It will therefore be necessary to use flowmeters

to calculate the quantity of propellant transferred. Again, a

receiver tank propellant gaging system will be required so that the

initial propellant load Is known. Ullage displacement offers many

advantages including the avoidance of pressurant gas loss, but is more

complex and therefore useful mainly for the resupply of large capacity

tanks, such as OTV's.
1017e/Z4

3.0-A. 2 Pressurant R«supply

The options for pressurant transfer include high pressure storage

supply bottles, noderate pressure source and high pressure pump, or

cryogenic storage and conditioned transfer. High pressure storage

supply offers the simplest method; however, this results in high

residuals after transfer. High pressure ganged bottle sequential

transfer has been evaluated to determine the lightest weight and least

residual configuration. Moderate pressure source and a high pressure

pump should also be evaluated as far as weight, volume and the design

complexity of a high pressure pump. Figure 3f shows schematics of the

above pressurant transfer techniques.

1017e/25

25Z
3.0-A. 3 Cryogenic Fluid Acquisition and Transfer

*

Acquisition of cryogenic liquids Bust depend on either a settling

acceleration or a CAD. Positive expulsion devices are not viable

alternatives. Also, supercritical feedout is not viable due to the

high energy requirements and the necessarily high energy state of the

•

fluid. For on-orbit storage and subsequent propellant transfer, the

primary areas of concern are (1) Thernodynamic control of the supply

tanks, (2) Liquid/vapor interface control, (3) Chllldovn and vent of

warm vapor andor inert gases from the receiver tank, and (A)

Propellant transfer and fill operations. In a low-gravity

environment, prolonged sub-critical storage of the cryogens in the

supply tanks is complicated by the lack of positive control of the

location of the liquid and vapor phases. Vapor venting must occur to

effectively control tank pressure as heat leaks into the tank. A

Thezmodynamic Vent System (TVS) for hydrogen will insure that only

vapor is vented.

10l7e/26

253
Heat is transferred from Che stored fluid into the throttled liquid

which is supplied to the TVS from « CAD. An internal heat exchanger

attached to the CAD would maintain the fluid la the CAD in a subcooled

state. Alternatively, an internal TVS with compact heat exchanger and

forced recirculation should be considered. An external heat exchanger

«

should be installed at the tank plumbing penetrations to intercept

concentrated heat leakage.

An example of a typical cryogenic transfer approach would be:

1. Propellant acquisition in the supply tank provided by simple

capillary sponge or vane type devices (Figure ) in the supply

tank rather than use of settling thrust.

2. LEL chilldown bolloff vapors from the receiver tank reabsorbed in

the bulk supply by reverse-flow cycling without a vapor return

line.

3. Fully autogeneoua pressutlzation of the supply tanks by means of a

capillary or pump-driven bootstrap loop eliminates the requirement

for helium pressurlzation.

4. Propellant-swirl techniques in the receiver tank can be used to

•

position the tank ullage space at the tank vent opening, facilitate

zero-g gaging of the tank by acoustic resonance techniques, and

sweep vapor bubbles from the wall to the vent port thereby reducing

or eliminating the need for a theroodynamic venting system and

coils.

5. Subcoollng of cryogenic propellants by helium bubbling on the

Shuttlelaunch pad or active refrigeration at a space base can

greatly reduce or eliminate boiloff losses in propellant transfer

and on long tankage hold periods (parking orbit, multiple sortie

missions, etc.).

10l7e/27
3.0-B Fluid & Pressurant Tank*

The fluid tanks selection for the resupply operations must consider

the nature of the resupply fluid, mission objectives, receiver satellite

resupply requirements and the fluid acquisition system on the receiver

satellite. Supply tank compatibility with receiver satellite tanks is

discussed in the fluid transfer operations. It is expected that an ERM

nodular design concept would provide for varying tank size and shaping.

Existing qualified tanks could be tailored to particular resupply

mission objectives in this manner reduce design/development costs for

the EB2I. A study of the various fluid acquisition system was conducted.

The results of the study indicate a preference for positive

expulsion devices (diaphragms) with those fluids that are compatible

with the operational elastomers (N

2

H

4

 wlch

 teflon, or Neoprene).

Although bladders and diaphragms are simple to operate and reliable,

they *re incompatible with reactive oxidizers (N?°4»

 for

 example)

and cryogenic liquids. Surface tension systems need to be used with the

storable oxldizers, and based on a philosophy of commonality of hardware

they very probably should be used with the related fuels as well (MMB,

UDMH or A-30, for example) on the same flights.

Cryogenic liquids present some special problems in the area of

liquid acquisition, principally related to thermal leakage and bubble

formation in the devices themselves. This can be avoided by appropriate

design minimizing thermal inputs into the PMD's, as well as the use of

thexaodynamic venting and cooling of the PMD themselves. This approach

is recommended for use with superfluid helium as well, together with the

rest of the special hardware compliments necessary for this fluid,

including vacuum jacket vapor cooled shields, porous plug evaporators

and thermoacoustlc dampers In the lines, as required.

Figure 3g shows the various acquisition devices studied and

highlights their main features.
3.0-C Contamination Control

Acceptance of fluid resupply by the satellite user community

critically depends on the alnlnlration of Impact by the operations on

those satellites, as well as economic considerations. Contamination of

sensitive instrumentation, chemically active surfacves or thermal covers

would compromise the operational effectiveness 'of the user satellite and

therefore reduce the attractiveness of resupply.

This study has concentrated on the conceptualization of resupply

system minimizing the impact on the user by starting with user-defined

requirements. A contamination control system approach has been

developed which minimizes contamination by 1) using zero-dribble-volume

quick disconnects, 2) design of a fluid interface which permits purge of

all surfaces contacting any fluid considered to be a contaminant, 3)

sleeving the Q-0's with purge jackets, 4) using catalytic vents with

N.H, to avoid liquid venting and the resulting contamination, and 5)

use of vacuum catch bottles for trapping purge products or 6) use of

chemical scrubbers to trap purge products.

The purge catch bottles could be vented to space between

engagements, at distances considered to be contamination-safe from any

sensitive vehicles, in preparation for the next engagement.

1017e/29
3.0-0 Receiving Spacecraft Design Considerations for Resupply

Minimizing the overall cost of resupply on-orbit requires that

careful consideration be given to design of the spacecraft and its

systeas to achieve maximum compatibility with resupply operations. It

is also desirable from a cost and operations standpoint that the

resupply interfaces of the various spacecraft and platforms be

standardized as much as possible.

The most obvious of the special provisions needed on such

spacecraft are the fluid disconnect fittings which interface with the

resupply module. One basic option in the use of such fittings is the

"hard docking" mode in which opposing disconnect halves are placed in

close proximity by the procedure • of mechanical docking between two

Interface rings. The disconnect halves are then engaged by special

drive mechanisms located on the resupply module.

One of the most important spacecraft resupply design factors is the

basic configuration of the propulsion system (RCS or ACS) to be

serviced. From the standpoint of operating cost and simplicity, the

preferred liquid propulsion system would be a conventional

monopropellant type using a single "blowdown" pressurization mode.

Besides the fact of requiring one rather than two propellant resupply

circuits, pressurant gas venting and resupply is not required since the

ullage gas is recompressed in the spacecraft propellant tanks by the act

of refilling (usually to a 2/3 or 3/4 full condition). Higher

propellant resupply pressures are required, however, to accomplish this

recospression.

If a storable propellant tank on the spacecraft uses regulated

pressurization for its ullage, it is necessary (prior to resupply) to

either vent this gas, or return it to the resupply module for

•

10l7e/30

25?
recompression or storage ia the propellant resupply tank ullage.

*

Venting usually requires a non-propulsive vent using opposed nozzles.

If contamination from storable propellant vapors is a problem, a

catalyst device may be required to decompose these into less harmful

products, and/or a mast may be needed to locate the vent nozzles at an

acceptable distance from the spacecraft.

Whenever gas is compressed in a tank during resupply, heat of

compression is generated which may have to be dissipated during

refueling and thus prolong the transfer operation. This can be

minimized by heat-sinking the affected tanks to the spacecraft

structure, or to a special radiator by means of heat pipes. Gas

orifices or swirl nozzles in the receiver tank can also be used to

promote heat transfer to the tank wall. In the case of a blowdown

propellant tank using capillary acquisition devices, liquid jets inside

the tank can maintain nearly isothermal ullage conditions during

refilling. However, if a diaphragm is used for liquid acquisition, such

forced heat transfer in the recompressed ullage is not feasible, and

several hours may be required for ullage cooling by gas conduction in

the zero-? environment.

In the case of cryofluid resupply, a special bypass vent circuit

may be needed in the spacecraft to dump chilldown vapors from the

Initially warm transfer plumbing directly overboard and avoid the

addition of thermal energy to the receiver tank.

Other potential spacecraft design requirements for resupply include

pressure and temperature instrumentation for propellant quantity gaging

by the PVT method. Safing of the spacecraft prior to docking requires

redundant isolation valving in the propellant lines and vent lines, and

appropriate control interlocks in systems such as RCS, ACS attitude

control, electrical and guidance.

1017e/31
&*»

*

OF POOR QUALTTIf

!

— • 4 «- • *

b « a>

o — g r

«» — 3 •

3 »

 m

• § b b C C

b w O • C — ??:-"

i.l.

• •i—*.

*•

 e

 ~

» • »

• •

« i£ •g C I J3

c •- b x b

*

h

 "»

Sr

-S5 S -^

.

a- »

u u

• *

^ 13cJ£'

S

 — e * i*

w£ I 3 &.

btud

ea

nes Prtsturlit

d f

1300 P

S 1)

flo

w rat

a

t HMII i 1

0 t

b

• NT

O $ 1

6 I

b

SyiU

a prccl

contact of r

g r" 8S5 — • • •. ~

o e a «• *• u S w Ob x c u* ®

b3««n«w> 2

s

-^ S

•5 •

o-g

S 3 —

w

 a

^" V*

i i s

UJ CO

UJ

on

CO

CO

— « c — c ^ • •

ae M a

§:.

w • a

b •

>

3 •

> a >

- c b i

A 03 O

-

CO

IU

I

•

tl

*

Y

b

V

UJ 5 is£

•

:

2 -

 M

 ^ • *

S

—• «

"3

 u

rO

UJ
i—r

•» a —

3 s

5 :t* — »> b» • e b b * • e -S

•• • a e * • • » o • b b

3

•• "c

 t •o 2 • S• 5

• 0 b b m U

i 2

«• ? "?

g.r 2

— w « e • «

a a «>«•"-*

• •

Prtssur

PS

Flo

w r

a

Transf

e

— »

O ' •• b

« • •

~ m 3 a

e co • *. w> • o

wt Ob >» e w* woe

b 3 » M • Wl ->.

SS^S:ST,1«2S >

i*cSll5§a

^ ^ a ««•*•• — *.

M »W> WlkCV ^

-

2 * SI

a • • •

—

c

3

5-

e e i i i s a

fcO

UJ

U X

03 — — «>b a

o .- e.

b

e

• ti« a

C » »•

S

C

e <•• — • t*

k k iC

11 a a. C7

a - — ^ b <

5^

* •

-

<f> m a

CO

UJ

UJ

— > S

b

«

UJ

</•)

-Q

rO

LsJ

->
ORIGINAL FA&H I*

Qf PQQR QUALITY

ri—r

Ul « 3 «•

8

S.ilpf

*!•

3 ~ 5

%~o a

• • •

f, ill

— a ••

g 8

w — _** 3*

IM

V

<B V ^ J

sl 5 t S.

Is 1 !|

" •* ° • 5 S

lfc «• ^ '

a —

»2 2 ^ •» o M • a.

o —

>» C IM *!

»» « v

i I

tm

a

C/)

b U

o •

a u §5

g

40

so !

(0 •• a. >-

s 5- s -

?1£?l^

e

§ s

• » o

3 ^ 0

a » a.

i • >

13

o

fO

u

Q/

VJ

U
(f

t aooooJ>a£u.o3So

^~ ^

5 »S»

3 j-S£

K ^ 'i

S

^5

-3«5«

3 3*-»••

*i

'J

f

3 5 3 S » • • " i|= C

 U

B5

s »:=

I

•• Mw siS

s

S

EL

5 x* ,*

IM IM 1 I i a s

iii

siiiif

• •

0 s

= § I• I5V

J

^ •",'
OF POOR
CO

CCs

ui

«

ceo

u

grS ns § ^

1

*

CCco

tllui

g

^l

Oo

2X

O

-£*! fn\t^ \

IM U

** •

IMtM

UJ

37
O

F POOR

0) O^O|2

ttoo

tn

(N

>M

eI Sags

;

EL

IUcrIL
C

OaO•OOO• ••M2OO

r4

•J)

•f.

Ji u amLL
4.0 MECHANICAL SUBSYSTEM

4.1 Orbiter Attach Mechanisms

4.2 RMS End Effectors/Gtapples

4.3 Docking Uabilical Actuators

4.4 Mission Specific Mechanisms

4.5 Separation Pyros

1017./32
4.0 MECHANICAL SUBSYSTEM INTRODUCTION

*

The ERM mission scenarios require various structural and mechanical

interfaces aaong the STS and resupply elements. The ERM design should

minimize interface hardware and maximize interface compatibility with

its elements. The required structural and/or mechanical interfaces

follow:

o ERM/GSE

o ERM/Orbiter Cargo Bay

o ERM/Orbiter RMS

o ERM/Space Station

o ERM/Carrier Vehicle (OMV)

o ERM/Recelver

A tabular summary of this section's requirements is provided in

Figure 4a.

1017e/33

A •».
4.1 Or biter Attach Mechanisms

The'ERM shall provide support trunnions so as to fit in the Orbiter

cargo bay as a deployable payload, conforming to ICD 2-19001, 'Shuttle

Orbiter/Cargo Standard Interfaces". The attached device shall allow the

EBM to be replaced in the cargo bay and secured for return to earth.

The saae support mechanism could be used to facilitate ground

handling and processing operations.

1017e/3A
4.2 RMS End Effactors/Grapples

An RMS grapple fixture will allow an Orbiter and/or Space Scation

RMS to handle the ERM within its cargo bay and facilitate mating and

deaating of the ERM from its carrier vehicle.

The ERM shall provide another grapple fixture coapatible with the

OMV carrier forward payload end effector sysCea to attach and detach

from the carrier vehicle.

1017e/35
4.3 Docking/Umbilical Actuators

*

The docking and uabiUcal actuators will be required to automate

the transfer of fluids from the RM to the Receiver. The primary design

approach will be the use of remotely controlled and automatic means for

docking, fluid and pressurant transfer and separation. The docking

system must allow for disconnect misalignment given the navigational

accuracy of the approach.

Design goals for remote engagement/disengagement of a self-sealing,

automatic open/close disconnect should include the following:

0 Eaploy "gang-on~/"Gang-off" design concept for mating/demoting all

disconnects simultaneously for docking/separation

0 Provide semi-balance/balanced design to minimize

engagement/disengagement loads

0 Provide swivel design to obtain maximum misalignment to simplify

docking maneuvers

0 Incorporate design features to preclude Jamming of the ganged

disconnects during docking maneuvers

0 Minimize trapped volumes, leakage rates, and retention/separation

forces

0 Minimize travel required for engagement/disengagement

0 Incorporate design features to preclude failure of disconnect to

disengage under designed retractional forces

A preliminary review of existing hardware indicates that a family

of disconnects is available for consideration in a Resupply System.

Modifications to existing hardware and demonstrations type testing are

considered required to meet the ERM requirements

Two previous disconnect study programs conducted, (MMS in-orbit

1017e/36
r«fueling study analysis and design for Goddard Spacecraft Center,

and Space Transportation System Disconnect Program for George C.

Marshall Space Flight Center) indicate that remote refueling la-orbit is

feasible, and can be defined and based on the use of currently qualified

space rated hardware. The only new design innovation on this program

for disconnects is docking remotely. However, this requirement is

considered within the state-of-the-art for disconnects.

The mating, checkout, operation, and donating of fluid disconnects

"is one of the more critical resupply operations.

1017e/37
4.4 Mission Specific Mechanisms

*

The design goal is to keep mission specific mechanism requirements

to a minimum. This would be accomplished through standardization of

receiver and carrier interface plates. This interface concept is shown

in Figure 4b. It is of flexible design to accommodate the full spectrum

of fluids which have been determined to be required by candidate

receivers. The concept simplicity would result in reliable operation

and reduced development cost.

The ERM requirements for retention mechanisms (discussed in

sections 4.1 - 4.3) are also driven by the design goal to maximize

interface compatibility. Where the EBM is loaded in space (e.g., STS

fluid scavenging) or used to transfer OMV bi-propellants from the

Carrier to the Receiver, unique requirements arise. The interface plate

to the Carrier (or space-based fluid source) would be required to

provide fluid and DPS umbilicals. Figure 3a in Section 3 summarizes the

fluid transfer system requirements for OMV bi-prop transfer.

1017./38
4.5 Separation Pyros

Provision oust be aide to effect separation of the attach

aechanisas should the system seize and fail to disengage. Pyrotechnic

devices Bay be necessary on both the ERM carrier interface (should the

carrier not provide for payload release) and the Receiver interface.

1017./39
ORIGINAL PAGE §3

OF POOR QUALITY

CO

Mr^M

^

Lb

U.

oe. MM—3

1

0

 1

Q£

y

U.

H—

eg

I

00

O

g

LU

LU

— M_

<

g

ec

LU

Q_

LU

^J

<

_-^r

~^~

*^J

^

J

^

 ~

oo

HLU

LU

WW

E

CD

w

•• «ov

-

.0

«* b X

b O «

0

 • "° z

c i oiui

b ~ It O W

*• £ u ** "

•O X"O b «• — o — o x

> — . > a 01

o a o a.o

b at b 3 3 a. -o a. at «>

at

^3

x <o

a x

£ A O

•M o tt

a 01 v

b *O

4f <8

— • U b

•— b •— • T3

« *• »l O

S

3 ff^L b « •

w o b a

•

X

*_l

Ji

"x-< » b « ^ b tl

a. « o a CA

x- o a.

b o « a. a

« — > 3

•• a*» -J _i

.a o < a. a.

b

O 1 II

•

X

H^-

^

oe ••

ui Z

^ «

O bt

S

UI

X

^wl

VI

^

|

e

b

b

3 b

o a

4* U

HI

•I £

Q

V) *<

(A •

41

^£ ^ ae a

• u

.C b _ b

« o •£ o

3 u x u

4» •* •

.a 41 .a at

« u « u

8-2 «2

• •

(A

(A •

§

— 21

« w a

b «A b

41 41

«•>•«•

XI « b b a <•

O (A W

• • •

i

W WJ

^j ^U

u. _J

ui a.

a «

Ul W

trt ^&

X Z

1

•

fc

f^

b % ^

> 41 C

•^ 3 «

41 O b

41

b 41 7

5 SH

"3 a ^-^

j< « —

o m a. B

a x a 3

c a

8 ^*

4> JB

Ol U X

« *• *•

Ol •— • 4)

«fi 4t

M • ^tf

§-5?-o

IS-&-?

5I!xu!

• • •

•

«*

9 ^

s

u. Ja

•" o

b b

« 41 41

> > b

1"«T w W 42*

U U X 4f 41 «^»

oe OE u.

• •

<

Ml

O

1

z

5

^J

0

"1

o

M

w

•^ **

• b

41 2 ll

41 ^ •o «

II

b

41

41 >

U •••

a 41

"3 ** ••

£ *•

4t *^ b

5"sr

S?3

J5§

•

0*»

3 «•

<» •

u

Ol f

ao^ ^-^

C £

s

 3

•

s

w

Ul

%£

§i

•• <

(A W

^

—2

X

W

e

01

O)

b

01

8

J

±2

«|

£ «

o

•• b

«**

X

u e

c ^

Ol w b <•

41 b

4> 41

U •—

41 01 01

«^ *• U

•«• • 41

Ul M b

•

Ik

"V

M W

b 41

41 e

> A

f u

« MIQ

U «•

41 M

OE <•

*

O

OE

(^

O

5

3

a.

Ul

(A

in

NT

.
o

a

i

o

o

o

o

•S

I

o

a

CO

ORIGINAL PAGE IS

OF POOR QUALITY

i

Us

I!1

1

i

i

= ^T _

li I"

-c-

^*

Uj

li.
5.0 STRUCTURES SUBSYSTEM

5.1 Frame

5.2 Subsystem Fittings/Brackets

5.3 Interface Attach Devices

5.4 Thermal Blankets & Covers

1017e/40
5.0 STRUCTURES SUBSYSTEM INTRODUCTION

The' structural configuration of the ERM aust accommodate a wide

variety of resupply aissions, fluids, tank shapes and sizing*. The ERM

•ygtea configurations which support the various mission scenarios

include:

o ERM/GSE

o ERM/Orbiter Cargo Bay

o ERM/Space Station Storage

o ERM/Carrier Vehicles

o ERM/Recelver and ERM/Carrier

o ERM/Carrier/Transfer Vehicles

Two of the basic structural requireaents shall be adaptability and

lightweight design. To achieve these requireaents, a number of concepts

have been developed and are shown in Figure Sa.

A tabular suawary of this section's requireaents is provided in

Figure 5b.

1017e/41
5.1 Franc

The* basic 8M framework shall distribute and react external and

internal loads resulting from all design on flight and ground loads and

their associated operational environments with any EHM configuration

defined in the Introduction.

The frane must be able to accommodate a vide variety of re supply

missions, fluids, tank shapes and sizings. The structure shall be

provided with adequate strength and stiffness in its design environment

to withstand limit loads without loss of operational capability and to

withstand ultimate loads at design temperature without failure. The ERM

shall satisfy the structural requirements as specified in Volume XIV,

Id) 2-19001, "Shuttle Orbiter/Cargo -Standard Interfaces" and In Section

5.0 of this document.

v.

!017e/42
5.2 Subsystem Fittings/Brackets

To 'achieve the basic structural requirements for adaptability and

lightweight design, the subsystem fittings and brackets oust provide for

mission design flexibility.

The subsystems requiring secondary structural attachments include

the avionics and power modules, resupply tanks and their associated

plumbing, and the interface hardware.

The RM may be required to provide attachment for the OMV docking

camera and lighting equipment which are otherwise excluded during

rendezvous approach with the current OMV configuration.

A modular design with a central core could provide for mission

specific hardware changeout, common, subsystea sharing, eary access and

quick reservicing or changeout, and a cost-effective design to meet

various mission objectives.

1017e/43
5.3 Interface Attach Device*

To accoaplish the reacte resupply objective* of the various mission

•cenarloa and to maximize the use of space transportation systeas,

numerous ERM systea configurations have been defined (see Section 5.1).

The RM shall provide a structural/aechanical interface for the following

configurations:

o BM/Shuttle Orblter Cargo Bay

o RH/Shuttle Orbiter RMS

o RM/Carrier Vehicle (OMV)

o KM/Receiver

o RM/Space Station

The structural interfaces will be required to support static and

dynaaie loads emanating froa launch flight and re-entry envlronaents.

Coupled dynaalc loads are introduced by the Orblter propulsion and

altitude control systea, RMS control maneuvers, the 'Carrier propulsion

and altitude control systea, the Receiver (if control systea is active),

gravity gradient torgues, aerodynamic drag forces, thruster exhaust

pluaes, and resupply fluid trnasfer and docking operations.

The structural/aechanical interface elements are discussed in

Section 4.0.

1017e/U
5.4 Thermal Blankets «ad Covers

Thermal blankets and covers shall provide the BM subsystems with

passive cbanal control to protect against the natural and induced

flight theraal environment.

The theraal environment includes Orbiter ascent, entry and abort

thermally induced loads, thruster exhaust plume heating loads, system

and subsystem generated heat loads and thermal radiation.

Theraal insulation, coatings and thermal isolators shall be used

where possible to reduce active thermal control requirements (see

Section 1.5).
•^

woaOaooouI5ccUJ

5

i3Ml .3 • •

»jg

ii

• •

^25

P="

is5

535

Ii

!

i

i

:e.

3S = w ^ • 5 g5

iiii

^_ ^t .^ ^»

i ^^

55 2 «i ^

||| |

iiiii

s2ii!

gM,5 iXn uu

in«A«Nin

6

inUJ C* 3—U.

SB'
ORIGINAL PAGE SS

OF POOR QUAJJTY

CO

§

%

UJ

£

OS

UJ

CO

UJ

u.

oc

UJ

^^

•"•

K-

^"255

UJ

^J

UJ

a:

g

CO

^ f^ *• M

1- "S-gJ •

k ** • If X

ik >^ oi <•» k

e o

g

£? ^ k *< 5 _.

w

 03 »U

X-»~ ~ ° C <• -

• •*•£. x i • k • J *» O» *• e k S •»

_i •"- •• 5 *• <•- • ^ • • —• 2 « • .c

at -*

 a

 — u u • o>

k£ • 5 k 52 Sc

II *< 3 n«CaC O

§*•» •**

^ 01

J3fC»u-ac«-~o )

k O — .a u O « ^ *< —

o o o o

W II

• u

T3 m* tt

o 5 -o

a "• "5

« u •

<• k **

•? * *

"S • • ft ^ *•

C k 3

OB « k

Ul k> O

0 O

IU

S

ae

Ik

*t

x o c •»

k i 5 •

• 5-»- tt

•O 3 • *•

e M • • 3 o— x •:J|

k •Q •

41 3 k

•Q *• • •!

» U 'B 3

§25 5

i-^§

o

5-

e toall:onfigurat1iIck changeo

u

nent

s

1§^i

*• -^ X O a M o u

<o «i — •

3 "•< "S

o o

••>

• •

% S •

-i c £- x

w S '

X* g

• ?_,

• 5 i ^ .A •

2 "a S

o

«n

uw

u

OE

C 81

tu *»

Si

Ul t3 *•*

Ul Ik

ai

^

5

r

 —

 M«

' 3

*•• C •

• Ok

*.^« ft

•~ it >

•* m — •*• k •

u ft u

• a ft Ik O k

o

*• -o

C ft •»

Sin M -k

010. — -- •

c c o -a — >

•>- k k <• *•

-• 3 « -a k u

• •Q-te C O ft

. <• a ^ k

o x c <-

i *• • • ft -•

5

 w U

.S

*** Qi^9 "^ *• C

O k 3 O ** «•

i-ri = s.

o e

4*

k

i

J(

U

O

k

>

•JJ

u

ae

o

s

s

ft w±

II U £

'i5

** > •*

2 k S •P> • a

«< ^ •

•Gfc S « • k

U> U **

O ft

•- uFor structur 'carrier ve

h

•^TM -^

^*

>

" U

 •J

kS

0. «

O

fc

o

u

ft

•

1k

»

k

o

1 C

M f"

C k

• X ft •fc k oe *•

o *• ui ^

.Ok f

S^S 0

S! 3-*

« at « o

S-S x

*<«.•*<

•"«!»«•»•

— •.«• e •^ *•*<•••

u J3 k w

u•.

 k o a 0 > —

o

— ui for structur 'arbiter 1 SI

Interface

s

^TM ^^

•o ^ "S

•^ U *f

§2.2 k *- •

0. « k

O

X

•°

^J

o

^•k

a

%

»

II

o o

X

w

<

* k

IU ft k k k» > ft ft

U. IU ft k —

ae u u k A •

ui M « m k H- > ae u o

z ui

:

a

 .

^^

•* »

II 0 k • 3 a

« 81"

Silo

*• a «<

• M C

• MM

• II •

*f 4tf _tf C

u • • o

« X • k

*< •» k f

O A II >

k 3 ^ C

a. «• *• «

o

*• *• «

3 k c k ft II

?-£ 5

§

O k

*

k

•

 0 > — 3

k C f O"

O 0 *• ft

<» u u k

•a « » o

•V S W k

§i^ c

if 52

O 0

„

• B

t • 3 *•

U X

3 «•

i 3

«1 Ul

Si

Ul IU

0 0

Ul

tu

I

81

« Ul

K ac

OE IU

IU >

X 0

^ w

•

K

in

M

I

0

•

CO

00

cc

oo

\n

U

fV