CR-17484 3

Februar y 1985

Advanced Orbit Transfer Vehicle

Propulsion System Study

Final Report

Prepare d by

Marti n M a r i'e 11 a

P.O. Box 179

Denver ,

Colorad o 8020 1

(IASA-CB-174843) ADVANCE D CEEI1 ISANSJfEB

VEHICLE PEOPUXSIG N ••••fcAD CAP S

LANGUAGE««««£c«n h }SYSTEM STUCI final Report

(Martin Marietta Aerospace) 17*4 p

HC A08/flf API C3CL 21H G3/2U

Prepare d Under

Contrac t NAS3-2385 8

fo r

NAS A Lewi s Researc h Cente r

21000 Brookpar k Rd.

Cleveland ,

Ohio 44135

Aerospac e

:

A7XI/77VSV AfXI/7/f T7XI
PREFACE

The purpose of this report is to document the results of the

Advanced Orbit Transfer Vehicle Propulsion System Study performed

for NASA Lewis Research Center under Contract NAS3-23858. Dr.

Larry Cooper of NASA Lewis Research Center and Mr. James Brown of

Pratt and Whitney Aircraft provided technical direction to the

study. The Martin Marietta Aerospace Program Manager, Mr. Larry

Redd, was directly supported by the following personnel:

J. A. Cathcart

T. W. Cooper

R. M. Corringrato

S. T. Cronau

S. C. Forgie

M. J. Harder

J. G. McAllister

T. J. Rudman

V. W. Stoneback

C. A. Beers and Eagle Engineering, Houston, TX

PRECEDING PAGE BLANK NOT FILMED

11
CONTENTS

PAGE

1.0 SUMMARY 1-1

2.0- INTRODUCTION 2-1

3.0 PHASE I - BASELINE OTV CONCEPT DEFINITION 3-1

3.1 Ground Rules and Assumptions Development 3-1

3.1.1 Mission Model Assessment ..... 3-1

3.1.2 Recommended Vehicle Characteristics 3-1

3.1.3 Propellant Requirements for Rev 6 Mission Model 3-4

3.2 Mission Description and Mission Requirements 3-7

3.2.1 Mission Description 3-7

3.2.2 Mission Requirements 3-10

3.3 System Requirements and Baseline Subsystem Selection. . . . 3-11

3.3.1 System Requirements 3-11

3.3.2 Baseline Subsystem Selection 3-11

3.4 Tradeoff Studies Performed 3-15

3.4.1 Major Configurational Characteristics of a Space-Based

Reusable OTV 3-16

3.4.2 Baseline Configuration Selection 3-18

3.4.3 Engine Design Criteria for Aerocapture 3-24

3.4.4 Main Propulsion System Tradeoff Studies 3-29

3.4.5 Baseline Vehicle Configuration 3-32

3.4.6 Space-Based OTV Propulsion Subsystems Design 3-36

3.5 Phase I Conclusions and Recommendations 3-43

4.0 PHASE II - SERVICING, RELIABILITY AND LIFE CYCLE COST ... 4-1

4.1 Maintainability 4-1

4.1.1 Baseline Vehicle Servicing Character 4-1

4.1.2 Failure Modes and Effects Analysis 4-2

4.1.3 Component Grouping for Servicing 4-13

4.1.4 Space-Based OTV Servicing Considerations 4-19

4.2 Servicing Operations 4-33

4.2.1 Serviceability and Basing of the Baseline Vehicle 4-33

4.2.2 General Servicing Considerations 4-34

4.2.3 Service Operations 4-37

4.3 Reliability Analysis and Effects of Nonindependent

Failures 4-72

4.4 Life Cycle Cost (LCC) Analysis 4-75

4.4.1 Life Cycle Cost Model Development 4-75

4.4.2 Life Cycle Cost Results 4-85

4.4.3 Life Cycle Cost Profiles 4-107

4.4.4 Level of Vehicle Modularization ..." 4-115

4.5 Phase II Conclusions and Recommendations 4-115

iii
List of Figures

3.1-1 OTV Parameters 3-2

3.1-2 Vehicle System Mass vs Specific Impulse 3-4

3.1-3 Costs for GEO Payloads 3-5

3.1-4 Spectrum of OTV Propellant Requirements 3-6

3.2-1 OTV Trajectory Plot (Delivery) 3-9

3.2-2 OTV Trajectory Plot (Aerobrake) 3-9

3.3-1 Aerobrake Design Parametrics 3-12

3.4-1 Vehicle Fleet Comparisons 3-17

3.4-2 Two Engine Baseline Weight Statement 3-20

3.4-3 Four Engine Baseline Weight Statement 3-20

3.4-4 Two Engine Feed System Schematic 3-21

3.4-5 Four Engine Feed System Schematic 3-22

3.4-6 Specific Impulse vs Mixture Ratio 3-23

3.4-7 Effect of Mixture Ratio on Propellant Requirements .... 3-24

3.4-8 Aerocapture Maneuver with Two Main Engines 3-26

3.4-9 Aerocapture Maneuver with Four Main Engines 3-26

3.4-10 Trajectory Simulations - Aerospike Control 3-28

3.4-11 Contingency Operation Propellant Margin 3-30

3.4-12 Initial Weight vs Number of Engines (P&W Engine Data) . . . 3-31

3.4-13 Initial Weight vs Number of Engines (Aerojet Engine Data) . 3-32

3.4-14 Two-Engine Baseline Layout 3-34

3.4-15 Four-Engine Baseline Layout 3-35

3.4-16 G02/GH2 Subsystem Schematic 3-39

3.4-17 Auxiliary Power Unit Schematic 3-40

4.1-1 OTV Propulsion System Functional Outline 4-3

4.1-2 Main Feed System Modules 4-14

4.1-3 Gaseous Subsystem Modules 4-15

4.1-4 Space Based OTV Servicing Facility 4-30

4.2-1 Baseline Two Engine OTV Layout 4-33

4.2-2 Conceptual Space Station OTV Servicing Facility 4-34

4.2-3 Work Organization Development for OTV Servicing 4-36

4.2-4 Representative Servicing Scenarios 4-38

4.2-5 Top-Level Turnaround Flow for Space-Based OTV 4-39

4.2-6 OTV Proximity Operations 4-40

4.2-7 Preparation Operations 4-43

4.2-8 Onboard Mobility Operations . 4-45

4.2-9 OTV Grappling Fixtures 4-46

4.2-10 The RMS Standard Grapple Fixture 4-47

4.2-11 Detanking Operations 4-51

4.2-12 Inspection and Checkout Operations 4-56

4.2-13 OTV Maintenance—Time/Process Relationships 4-61

4.2-14 The Shuttle RMS Teleoperator System 4-65

4.2-15 The Relationship of Servicing Technologies 4-65

4.2-16 Characteristics of Automated Tasks 4-66

4.2-17 Mechanical Module Interface Options 4-68

4.2-18 Engine-OTV Interface Concepts 4-69

4.2-19 System Design for Multimode Maintenance Operations .... 4-71

iv
List.of Figures (Continued)

Page

4.3-1 Main Propulsion System Reliability (0.994 Single-Engine

Reliability) 4-73

4.3-2 Main Propulsion System Reliability (0.996 Single-Engine

Reliability) 4-74

4.3-3 Main Propulsion System Reliability (0.999 Single-Engine

Reliability) 4-75

4.4-1 OTV Life-Cycle Cost Model Scenario 4-76

4.4-2 Reliability Impact on LCC (Bum Rel - 0.994, NIFR =0)

Pratt & Whitney 4-86

4.4-3 Reliability Impact on LCC (Burn Rel - 0.994, NIFR - 0.03)

Pratt & Whitney 4-87

4.4-4 Reliability Impact on LCC (Burn Rel - 0.994, NIFR - 0.05)

Pratt & Whitney 4-88

4.4-5 Reliability Impact on LCC (Burn Rel - 0.996, NIFR = 0)

Pratt & Whitney 4-89

4.4-6 Reliability Impact on LCC (Burn Rel - 0.996, NIFR ^ 0.03)

Pratt & Whitney 4-90

4.4-7 Reliability Impact on LCC (Burn Rel - 0.996, NIFR = 0.05)

Pratt & Whitney 4-91

4.4-8 Reliability Impact on LCC (Burn Rel - 0.999, NIFR =0)

Pratt & Whitney 4-92

4.4-9 Reliability Impact on LCC (Burn Rel - 0.999, NIFR = 0.03)

Pratt & Whitney 4-93

4.4-10 Reliability Impact on LCC (Burn Rel - 0.999, NIFR = 0.05)

Pratt & Whitney 4-94

4.4-11 Reliability Impact on LCC (Burn Rel - 0.994, NIFR - 0)

Aerojet 4-95

4.4-12 Reliability Impact on LCC (Burn Rel - 0.994, NIFR = 0.03)

Aerojet 4-96

4.4-13 Reliability Impact on LCC (Burn Rel - 0.994, NIFR = 0.05)

Aerojet 4-97

4.4-14 Reliability Impact on LCC (Burn Rel - 0.996, NIFR - 0)

Aerojet 4-98

4.4-15 Reliability Impact on LCC (Burn Rel = 0.996, NIFR = 0.03)

Aerojet 4-99

4.4-16 Reliability Impact on LCC (Burn Rel - 0.996, NIFR - 0.05)

Aerojet 4-100

4.4-17 Reliability Impact on LCC (Burn Rel = 0.999, NIFR =0)

Aerojet 4-101

4.4-18 Reliability Impact on LCC (Burn Rel = 0.999, NIFR - 0.03)

Aerojet 4-102

4.4-19 Reliability Impact on LCC (Burn Rel - 0.999, NIFR = 0.05)

Aerojet 4-103

4.4-20 Missions/Overhaul Sensitivity 4-108

4.4-21 Life-Cycle Cost Profile 4-110

4.4-22 DDT&E Cost Profile 4-111

4.4-23 O&S Cost Profile 4-112

4.4-24 Production Cost Profile 4-113

4.4-25 Advanced OTV Cost Drivers 4-114

4.4-26 LCC as a Function of No. of Modules 4-117
List of Tables

3.2-1 Mission Timeline for Reference Mission 3-8

3.2-2 Mission Requirements 3-10

3.3-1 System Requirements 3-11

3.4-1 Main Propulsion System Tradeoff Study Conclusions and

Engine Operational Requirements 3-33

3.4-2 APU Mass Estimates 3-41

3.4-3 Cold Side APU Parameters 3-42

3.4-4 Hot Side APU Performance 3-42

4.1-1 Failure Modes and Effects Analysis (FMEA) Results 4-4

4.1-2 FMEA Severity Results Recommendations 4-12

4.1-3 Degree of Onorbit Servicing 4-13

4.1-4 Vehicle Module Descriptions 4-17

4.1-5 Space-Basing Study Assumptions 4-21

4.1-6 OTV Mission Operations Flow 4-21

4.1-7 OTV Servicing Overview 4-22

4.1-8 Engine Servicing Overview 4-23

4.1-9 Example Module Remove/Replace Times 4-26

4.1-10 Space Station OTV Servicing Facilities 4-29

4.2-1 Levels of Decomposition 4-36

4.2-2 Possible Onboard Moves 4-45

4.2-3 Grappling Activity Task Flow 4-46

4.2-4 Positioning (and Mounting/Berthing) Task Flow 4-49

4.2-5 Umbilical Attachment Activity Task Flow 4-52

4.2-6 Umbilical Detachment Activity Task Flow 4-54

4.2-7 NDI Technique Advantages and Disadvantages in Space .... 4-57

4.2-8 Spacecraft Maintenance Mode, Approach and Concept 4-60

4.2-9 Characteristics of Scheduled and Unscheduled Maintenance

Operations 4-62

4.4-1 LCC Model Development Stages 4-76

4.4-2 Pratt & Whitney Optimal LCC Propulsion Configuration . . . 4-104

4.4-3 Aerojet Optimal LCC Propulsion Configuration 4-105

4.4-4 LCC Sensitivities Summary (Pratt & Whitney) 4-106

4.4-5 LCC Sensitivities Summary (Aerojet) 4-107

4.4-6 Level of Servicing and LCC 4-116

5.0 REFERENCES 5-1

vi
Abbreviations

ACC Aft Cargo Carrier

ACS "Attitude Control System

AI Artificial Intelligence

APU Auxiliary Power Unit

Btu British Thermal Units

eg Center of Gravity

CCTV Closed-Circuit Television

DDT&E Design, Development, Test and Evaluation

DOD Department of Defense

deg Degree

dia Diameter

EMU Extravehicular Maneuvering Unit

EPS Electric Power System

EVA Extravehicular Activity

elec Electrical

FMEA Failure Modes and Effects Analysis

fps Feet per Second

ft Feet

g Gravity

GEO Geosynchronous Equatorial Orbit

GG Gas Generator

GH2 Gaseous Hydrogen

G(>2 Gaseous Oxygen

GPS Global Positioning Satellite

GRD Ground

Guid/

Nav Guidance and Navigation

H2 Hydrogen

He Helium

hp Horse Power

h Hour

l&CO Inspection & Checkout

I/F Interface

IOC Initial Operating Capability

IUS Inertial Upper Stage

IVA Intravehicular Activity

Igp Specific Impulse

k Thousand

L/D Lift to Drag

LCC Life-Cycle Cost

LEO Low Earth Orbit

LH2 Liquid Hydrogen

L02 Liquid Oxygen

LOS Line of Sight

LRU Line-Replaceable Unit

Ibf Pounds force

Ibm Pounds Mass

MES Main Engine Start

MLI Multilayer Insulation

MMS Multimission Modular Spacecraft

MMU Manned Maneuvering Unit

vii
Abbreviations (continued)

MPS Main Propulsion System

MR Mixture Ratio

MSFC Marshall Space Flight Center

mech Mechanical

min Minutes

N2H2 Hydrazine

NASA National Aeronautics and Space Administration

NDI Nondestructive Inspection

NIFR Nonindependent Failure Rate

nmi Nautical Mile

02 Oxygen

O&S Operations and Support

OMV Orbital Maneuvering Vehicle

ORU Orbit-Replaceable Unit

OTV Orbit Transfer Vehicle

Ops Operations

P&W Pratt and Whitney Aircraft

PU Propellant Utilization

P/L Payload

PAM Payload Assist Module

PIM Pumped Idle Mode

psf Pounds per Square Foot

psi Pounds per Square Inch

QDs Quick Disconnects

R/R Remove and Replace

RCS Reaction Control System

RF Radio Frequency

RMS Remote Manipulator System

RN Retractable Nozzle(s)

Rel Reliability

SE&I Systems Engineering and Integration

SEE Special End-Effector

SN Standard (Fixed) Nozzle(s)

SS Space Station

STS Space Transportation System

s Seconds

TDM Technology Development Mission

TDRSS Tracking and Data Relay Satellite System

THFS Teleoperator Human Factors Study

THIM Tank Head Idle Mode

TVS Thermodynamic Vent System

F Degrees Fahrenheit

R Degrees Rankine

o

viii
1.0 SUMMARY

Future NASA and DOD missions will benefit from high-performance

reusable orbit transfer vehicles. With the advent of a space

station, advanced engine technology, and various new vehicle

concepts, reusable orbit transfer vehicles that provide significant

economic benefits and mission capability improvements will be

realized. Engine and vehicle design criteria have previously

lacked definition with regard to such issues as space basing and

servicing, manrating and reliability, performance, mission

flexibility, and life-cycle cost for a reusable vehicle.

This design study has resulted in the definition of a reusable

orbit transfer vehicle concept and subsequent recommendations for

the design criteria of an advanced L02/LH2 engine. These

design criteria include number of engines per vehicle, thrust

level(s) per engine, mixture ratio, nozzle design, etc. The major

characteristics of the vehicle preliminary design include a low

lift-to-drag aerocapture capability, main propulsion system failure

criteria of fail operational/fail safe, and either two main engines

with an attitude control system for backup or three main engines to

meet these failure criteria. A maintenance and servicing approach

has also been established for the advanced vehicle and engine

concepts. Design tradeoff study conclusions were based on the

consideration of reliability, performance, life-cycle cost, and

mission flexibility.

This program was accomplished in two phases. In Phase I tradeoff

studies and analyses were conducted to define a baseline

L02/LH2 propulsion system at the major component level based on

the requirements of mission scenarios for future NASA and DOD space

architectures. Major engine assembly characteristics (number,

nozzle design, thrust, mixture ratio, throttling range, etc) were

determined using engine contractor parametric performance data.

In Phase II subsystem definition, servicing operational procedure

description, reliability, and life-cycle cost and sensitivity

analyses were performed to assess the design optimization and space

basing impacts on the baseline propulsion system established in

Phase I. The sensitivity studies provided overall system design

criteria and the rationale for choice of engine number, component

and subsystem development recommendations, servicing/replacement

methods, and recommendations for manrating criteria.

1-1
2.0 INTRODUCTION

The design requirements for a space-based, reusable, manrateable

orbit transfer vehicle (OTV) are substantially different from the

requirements typical of a ground-based expendable upper stage. The

mission may still require a delivery capability to geosynchronous

equatorial orbit (GEO) of a large payload. However, retrieval of

payloads from GEO may be required in addition to returning the

vehicle itself so it may be reused. If astronauts are on board,

the vehicle must obviously be manrated for that application.

The geometry constraints of ground launch for an upper stage can be

relaxed for a space-based OTV. But, depending on the mode of

return from a high orbit, geometric concerns are still an issue if

the aeroheating effects during an aerobraking maneuver are

considered, for example. Because the engine design requirements

must be compatible with the operation of an advanced OTV and a

suitable approach to servicing and maintaining the vehicle is also

essential, we must understand the issues surrounding the design

requirements of a space-based reusable OTV and appropriately

resolve them.

Previous OTV concept studies have investigated reusable propulsion

systems associated with aeroassisted vehicle concepts with moderate

and low lift-to-drag (L/D) ratios. There has also been interest in

a vehicle that can be packaged into an aft cargo carrier (ACC) on

the space shuttle orbiter external tank for delivery to orbit. The

trend in overall vehicle evolution is to move from ground-based

expendable vehicles to space-based reusable and manned vehicles.

This study focused on the "ultimate" vehicle that would be

space-based, reusable, aeroassisted, and manrateable. No attempt

has been made to show an evolution from ground basing to space

basing (however, initial delivery to the space station was

considered). Before proceeding into the conceptual design study,

the following previous related studies were reviewed.

Vehicle; - OTV Concept Definition Studies, General Dynamics

Contract NAS8-33533 and Boeing Contract NAS8-33532.

Future OTV Technology Study, Boeing Contract

NAS1-16088.

In-house activity on ACC OTV, Martin Marietta.

Space Station and Servicing - OTV Servicing Study, General

Dynamics Contract NAS8-35039.

2-1
Definition of Technology Development Missions for Early Space

Station, Martin Marietta Contract

NAS8-35042.

Space Station Needs, Attributes,

and Architectural Options, Martin

Marietta Contract NASW-3686.

Aeroassist - System Technology Analysis of Aeroassisted OTVs for

Moderate and Low L/D, General Electric Contract

NAS8-35096, and Boeing Contract NAS8-35095.

Technology Identification for Aero-Configured OTVs,

Boeing Contract F33615-82-C-3014.

2-2
3.0 PHASE I - BASELINE OTV CONCEPT DEFINITION

This section describes the mission ground rules and requirements,

system requirements and baseline vehicle/engine identification.

The results of vehicle and engine performance and geometry tradeoff

studies are also shown.

3.1 GROUND RULES AND ASSUMPTIONS DEVELOPMENT

3.1.1 Mission Model Assessment

Space missions through the year 2000 include the delivery of many

large payloads, require manned presence, and will significantly

benefit from the use of a reusable advanced orbit transfer

vehicle. The MSFC nominal mission model PS-01 (Rev 6) for the

period from 1994 through 2000 has been used in this study. The

years 1994-2000 were used because 1994 is the earliest projected

initial operational capability (IOC) of a space-based OTV. Two of

the mission types involving large payloads in the 1994-2000 time

frame include:

1) Delivery of 5000 Ibm to geosynchronous orbit (GEO) and return

of 2000 Ibm (8 total missions);

2) Manned sortie with delivery of 13,000 Ibm to geosynchronous

orbit and return of 13,000 Ibm (7 total missions).

The remaining of the 47 LEO/GEO missions concern payload delivery

(up to 16,000 Ibm) to GEO only.

These large payload missions will require a propulsion system with

high total energy capabilities. Other requirements include that

the vehicle be manrateable. The space station may provide a

suitable location for basing, servicing, refueling, and launching a

reusable OTV. The following discussion presents arguments for

configuring such a vehicle.

3.1.2 Recommended Vehicle Characteristics

The dry weight of a propulsion system can be expressed by a

constant value of hardware weight (independent of total system

size) plus a weight of hardware that depends on total propellant

loaded. This is expressed as:

Mdry = A + B (

where :

3-1
Mdry

 = vehicle dry mass

A - constant hardware mass

B = .variable hardware factor

Mprop = loaded propellent mass

Figure 3.1-1 shows the characteristics of single-stage liquid

propulsion vehicles and their limitations. The x-axis is the

nondimensional parameter of delta V divided by specific impulse

(I8p). The y-axis is the ratio of vehicle and payload initial

mass (including propellant) divided by payload mass and vehicle

constant hardware mass (A).

B = 0.1 0.01

o.os

103r10'

10

x for Low L/D

- Storable -1.88 (341 s)

- Cryo-1.4 (460s)

10

,-2

AV - 20,585 fps (LEO-GEO-LEO)

Figure 3.1-1 0 T V Parameters

3-2
Several conclusions may be drawn from the figure. The first is

that for a large enough delta V, there is an asymptotic value for a

given propellant combination and engine technology (lsp) for

which no amount of propellant can deliver any size payload.

Therefore, to complete a mission that imposes a delta V above the

asymptotic value, either a higher energy propellant must be chosen

or the required delta V must be reduced (as with an aerocapture

maneuver of some sort). For a low-earth orbit (LEO) to

geosynchronous equatorial orbit (GEO) payload delivery with an

serocapture return, the required delta V is about 20,000 fps for a

vehicle with a lift to drag ratio (L/D) of zero. Also, a

reasonable estimate for B for large propulsion systems is about 0.1

for storable propellants (N204/MMH) and about 0.125 for

IX>2/LH2 (ref Magnetoplasmadynamic Thruster Definition Study

performed for AFRPL, Contract F04611-82-C-0049). So, for a

storable propellant combination, the delta V of 20,000 fps places

the value of x very near the asymptotic value for total vehicle

weight. In other words, small savings in delta V and/or increases

in I8p result in very great savings in total vehicle mass for a

given payload mass. For example, providing a vehicle with a L/D of

1.0 will result in a possible inclination change capability on LEO

to GEO return of 14 deg and a delta V savings of only about 800

fps. This, however, is valuable in terms of reducing storable

propellant requirements and total size of the resulting vehicle

even with a subsequent increase in dry weight because of the higher

L/D ratio. Therefore the higher L/D approach appears suitable for

a vehicle with a storable propellant combination.

A L02/LH2 propulsion system with an aerobrake and a L/D of zero

provides a delta V savings of about 7500 fps when returning from

GEO over an all-propulsive system and results in an x value that is

further away from the asymptotic value for total vehicle mass.

Therefore the reduction in total vehicle mass is not as dramatic

for reductions in delta V by going to L/D ratios greater than zero

as for the storable propellant system. This is a result of being

on a "flatter" part of the curve relating x and total system mass.

For instance, in providing a 800 fps reduction in delta V during

return from GEO (via L/D of 1.0), the resulting L02/LH2 total

propulsion system mass shows only a reduction of about 5%. This

also assumes no increase in dry weight as a result of providing the

vehicle with a higher L/D capability. Therefore a mid to high L/D

configuration is not recommended for a L02/LH2 vehicle because

an inclination change capability does not appear to benefit the

system to a great degree. The concept of a low L/D reusable

aerobrake, however, is consistent with materials technology

evolution for the time frame of interest and the resulting vehicle

configuration may be more amenable to component accessibility and

servicing.

3-3
Figure 3.1-2 shows the relationship between total vehicle system

mass (propulsion system dry mass plus propellant) and specific

impulse. One can see that in going from an all-propulsive system

to one with aerobraking (L/D = 0) the aerobraked system is less

sensitive to specific impulse (Isp) in the range applicable to

L02/LH2 propellents. Detailed vehicle studies assessing the

differences in vehicle dry mass for low vs high L/D ratios would be

required before determining the benefits of additional delta V

savings.

300

I 200

.s

I

100

Reusable OTV

LEO-to-GEO Mission

All-Propulsive

(MF = 0.875)

Aerobraked

(MF = 0.86)

7,500 Ibm

Deploy/Retrieve

7,500 Ibm

Deploy/Retrieve

12,000 Ibm

Deploy

440 450 460

Specific Impulse, s

470 480

Figure 3.1-2 " Vehicle System Mass vs Specific Impulse

3.1.3 Propellant Requirements for Rev 6 Mission Model

In addition to the large payloads in the Rev 6 mission model that

require a high-energy OTV, numerous missions require much less

energy for each payload. This results in a wide range of payload

weights and propellant requirements for all missions. A single

reusable space-based OTV design for performing all the Rev 6

missions requiring orbit transfer would substantially reduce launch

costs because of eliminating the launch of the upper stages for

these small payloads. However, a single-stage design is not

feasible for performing single payload delivery missions for such a

wide range of payload masses.

3-4
Figure 3.1-3 (ref Magnetoplasmadynamic Thruster Definition Study,

AFRPL Contract F04611-82-C-0049) shows the trend in cost for

putting various sizes of payloads into GEO. The two curves show

the differences in payload delivery cost depending on whether

propellant (LC>2/LH2) would be available from scavenging

external tank residual and surplus propellant and reducing launch

costs accordingly. The main conclusion from the figure is that the

larger the payload and subsequent delivery vehicle (OTV) the

greater the cost savings will be. Therefore, fewer, larger

payloads to GEO would be desirable over many small payloads and a

few large ones.

401—

Reusable Aerobraked

LO

30

W

o

8 20

10

With No Propellant

Delivery Cost

_L

1,000 10,000

Payload Delivery to GEO, kg

Figure .3.1-3 Costs for GEO Payloads

The small payloads in the Rev 6 mission model (PAM, IUS, and LEASAT

class) may lend themselves to grouping for delivery. Distinct

groups that are intended to fly in the same month could possibly be

grouped and delivered to GEO with the same OTV flight. This would

also narrow the range of total propellant requirements over which

to design an appropriate OTV capable of performing the Rev 6

missions. Figure 3.1-4 shows the propellant requirements for all

(except three small PAM-delivered payloads that could not be

grouped by month) of the Rev 6 LEO to GEO missions from 1994 to

2000. The darkened bars represent the large single payload

deliveries by an OTV. The hollow bars are the propellant

requirements for the "grouped" payload missions. The ground rules

for logical grouping of these payloads are as follows:

3-5
Packaged with I

Grouped Missions -\ \

\i \

I

in

Advanced LO^L»2 OTV

(Aerobraked and Reusable)

Mass Fraction = 0.88

V

483s

 •

Rev 6 LEO-to-GEO AOTV Missions •

(1994 to 2000) •

Heavy Single Payloads •

n

Small, Grouped Payloads I I

(PAM, IUS and LEASAT Class)

I „ I I

10 20 30 40 50 60

Dvnnall*k«* DM«»!vii>B«n.»*j» < A** !!.«&

Figure 3.1-4 Spectrum of OTV Propellant Requirements

1) All non-DOD;

2) LEO-to-GEO delivery only;

3) Fly during same month.

The groupings of the smaller payloads and subsequent packaged

delivery to GEO results in propellent requirements in the same

range as those associated with the single large payload delivery

missions. However, the wide spectrum of propellant requirements

for the single large payloads and for the packaged groups leads to

the need for multiple sizes for OTVs, staging, drop tanks, or else

a totally modular approach to the design of an OTV.

In comparing these propellant requirements to estimates of

propellant available from scavenging residual and surplus space

shuttle external tank propeHants, it appears that a large

percentage of these propellants could be available. For instance,

if the total propellant requirement for 47 LEO to GEO missions

between 1994 and 2000 were from 1.5 to 2.0 million Ibm, as much as

two thirds of this propellant could be available from scavenging

(Ref STS Propellant Scavenging Systems Study, NAS8-35614). The

result here is that for the high-energy OTV missions, the average

propellant delivery cost may be significantly reduced by taking

advantage of scavengeable propellant.

3-6
Based on review of previous study conclusions, the mission model,

and appropriate vehicle characteristics, the ground rules and

assumptions to be followed during the baseline vehicle development

of this study were:

1) Space-based, reusable L02/LH2 vehicle;

2) Space station-berthed and -serviced;

3) Single STS delivery (deliver empty);

4) Aerobraked return (low lift to drag ratio);

5) Manrateable;

6) Able to capture Rev 6 missions (LEO to CEO).

3.2 MISSION DESCRIPTION AND MISSION REQUIREMENTS

3.2.1 Mission Description

The reference mission (Table 3.2-1) is a single payload delivered

to GEO and vehicle return empty with an aerocapture maneuver at

LEO. The aerobrake maneuver will nominally be a single pass.

Variations of this mission will include up to three passes, manned

servicing mission to GEO, and unmanned servicing at GEO. Timelines

for the mission are shown in the table. Figure 3.2-1 shows the

trajectory of the delivery portion and Figure 3.2-2 shows the

aerobrake return. For the purpose of system sizing, 325 fps are

included as gravity losses and 50 fps contingency for aerospike

operation. Additionally 500 fps is included in the perigee raising

and phasing scenario to adjust the altitude and inclination of the

orbit following atmospheric exit. The velocity taken out by the

brake was derived from work done at Martin Marietta on an aft cargo

carrier OTV.

The timelines were derived assuming a single shift on the space

station is required to load cryogenic propellants and launch the

OTV. A Hohmann transfer is used, and the return flight occurs

after the second nodal crossing allowing for one day at GEO for

operations. The OTV navigation and guidance is updated three times

before reentry. The OTV is then safed after achieving a proper

phasing orbit before the ACS is used to rendezvous with the space

station/OMV. Communications with the space station is used during

launch and rendezvous, and communications with the ground through

the TDRSS below approximately 6000 nmi is used for actual control

of the OTV operations.

3-7
ORIGINAL PAGE 13

OF POOR QUALITY

Table 3.2-1 Mission Timeline for Reference Mission

Mission Operation

Launch

Transfer to Loading Area

Chill Down and Load LHj/LO.,

Terminate and Leak Check

Disconnect Lines

Perform Final Prelaunch Operations

OTV Payload Move from Space

Station

Activate Main Engine

Navigation Update Position for Burn

GEO Transfer MES-1 (1st mam engine

start)

Transfer, Coast and Thermal Rolls

Navigation Update

Midcourse

Position for Burn

GEO Injection MES-2

Spacecraft Deployment Checkout

Operations

OTV/Payload Separation

GEO Operations. Thermal Rolls

^

Navigation Update

Position for Burn

OTV Deorbit Burn (MES-3)

Guidance/Navigation Update

Midcourse No. 1

Guidance/Navigation Update

Midcourse No. 2

Navigation Update (MES-4)

Atmosphere Entry

Atmosphere Exit

Apogee Boost No. 1 (MES-5)

Navigation Update

Apogee Boost No. 2 (MES-6)

Safe OTV

ACS Moves OTV to Station

Space Station Grapples and Berths OTV

OTV

Time,

h mm

from MES-1

7 30

7:15

• 3:15

3-00

• 2.15

• 1 15

• 30

• -20

3.00

3 15

5:00

5.15

6:00

6-15

27:45

28.50

29:05

29.20

32-00

32-15

34-00

34 15

34.50

35:00

35.05

35:40

36:55

37:10

41-00

41-30

42:00

System

Driver

Launch in one shift after

loading cryogens.

Loading accuracy impact on

margins.

Valve redundancy.

No venting near space

station.

10-mi separation.

Temperature of insulation

and components.

TDRSS below 6000 nmi;

NASA 85 ground station

above 10,000 nmi.

Control of 50% full OTV.

Temperature of MLI,

components.

Aerobrake perigee 45 nmi.

Atmospheric variations.

Base heating aerospike.

Perigee 4 nmi, aerobrake

and manrating.

Rendezvous with space

station orbit.

Safe operation. ACS

near space station.

Safe ACS system.

System

Impact

Zero-g gage on space station.

Seals.

Space station and TDRSS

links.

ACS sizing.

Go into tank head idle

mode (THIM).

ACS sizing.

ACS control.

ACS sizing.

Attitude adjustment for

GPS link acquisition

ACS sizing.

ACS.

Collision avoidance— ACS.

ACS sizing nodal crossing

every 12 h.

Ground monitor NASA 85

ground station.

ACS sizing.

ACS.

ACS.

Start engine THIM.

Structure, TPS, engine

throttling, ACS control.

Redundance, space

station rendezvous.

80-fps allowance for inclination adjustment.

Backup fuel cells.

ACS sizing.

ACS system design.

Remarks

Purge lines between valves.

Communication links,

verify systems "go "

Phasing orbit for targeting

during proximity operations.

Partial chilldown and loading

Ground Monitor

Burn ttme = 17 mm, 7820 fps.

0.1 deg/s to 0.5 deg/s>

Ground control monitor OTV

Delta V = 30 to 40 fps.

Burn time = 8 mm; delta V =

5840 fps.

Communication links, test.

Will use 2nd (24-h) nodal

crossing.

Delta V = 6050 fps; burn time

= 2 mm.

Delta V = 30 to 40 fps.

Delta V = 30 to 40 fps.

Delta V - 7800 fps. Viking

shell, low L/D.

Delta V = 50 fps for aerospike

Delta V = 250 fps, phasing

orbit.

Delta V = 250 fps, circularize

and adjust inclination.

4-h allowance.

3-8
UPDATE

GEO

UPLE6:

THERMAL ROLLS

&

S/C COM 01 POUTS

M I.

AvL=325 FPS

(LOSSES)

Figure 3.2-1 OTV Trajectory Plot (Delivery)

SPACE STATION 250 N.MI.

APOGEE

BOOST 1 & 2

Figure 3.2-2 OTV Trajectory Plot (Aerobrake)

(CONTINGENCY &

AEROSPIKE)

3-9
3.2.2 Mission Requirements

Table 3.2-2 shows the mission requirements. These were derived

from the reference mission scenario. A maximum time of two weeks

at GEO was selected for the manned mission to be compatible with

one month between missions and two-week servicing at the space

station.

Table 3.2-2 Mission Requirements

Mission

Operation Requirement Remarks

Delta V (Main Engine)

Payload Range

No. of Burns

Throttling

Maximum Time at GEO

Safety—Manned Flight, Space Station

Design Life

Time Between Overhauls

Aerobrake

No. of Aero Passes

Acceleration

Onorbit Thermal Environment

Guidance/Navigation

IOC

Mission Time

Initial (T/W)

13,985 fps up, 6,600 fps down.

16k Ibm deliver to GEO. 13k Ibm

deliver/return (GEO).

6 nominal.

During aerospike or large space

system.

2 weeks.

Fail operational/fail safe Shuttle

safety.

30 missions.

10 hours.

Yes (low L/D), with aerospike.

Up to 3 (1 nominal).

3.2 g (maximum).

Minimize boiloff.

GPS and Ground (and TDRSS)

tracking communications.

1994 to 1997.

Compatible with 2-week servicing

period at space station.

0.15 (minimum), 0.22 (nominal).

250-nmi space station; aerobrake - 7800 fps.

Rev 6 September 1982 13k manned sortie.

Single perigee and aerospike return.

TBD.

Manned servicing—one month turnaround

(2-week servicing).

ICD-2-19001 and NHB 1700.7A.

TBD Phase 11.

Engine life paces overhaul.

Use Viking configuration 70-deg half-cone

<CD = 1.6).

Minimize transfer time, reduce heating to

brake and stage.

Shuttle ascent.

Thermal roll, 0.1 deg/s to 0.5 deg/s.

No control from space station. GPS state

vector.

Space station IOC (advanced engine IOC).

One month between launch.

Minimize delta V for single perigee burn.

Near the space station, orbiter safety outlined in NHB 1700.7A was

used. Manned missions were assumed to be consistent with fail

operational/fail safe. At least in the early stages of this study

program, fail operational/fail safe was adopted as a "starting

point" in developing an appropriate vehicle design philosophy.

This allows the astronauts the option to return after one failure,

and the ability to return to the space station, STS, OMV retrieval,

etc after sustaining two failures. A 30-mission life is compatible

with two OTVs and approximately 60 missions. A 3.2 maximum g-level

was selected to be compatible with a shuttle launch. However, the

OTV will probably be delivered to the space station empty.

Therefore, the loads during orbit transfer (main engine burn and/or

the aerobraking loads) will probably be the structural design

drivers.

3-10
3.3 SYSTEM REQUIREMENTS AND BASELINE SUBSYSTEM SELECTION

3.3.1 System Requirements

Table 3.3-1 shows the OTV requirements for the various vehicle

systems. These were derived from the mission requirements, study

ground rules, and the statement of work.

Table 3.3-1 System Requirements

Aerobrake Requirements

- Withstand 2500 F and 15 psf

- Compatible with Space Station Servicing

- Reusable

— Compatible with Manned Mission Return

Engine Requirements

— Provide Impulse for GEO Transfer and Return

— Provide Aerospike Capability (TBD)

— Operate At a Mixture Ratio Between 5 and 7

- Reusable

— Provide Main Propellant Tank Pressurization

- Use LHj/LOj Propellents

- Throttlable (TBD)

ACS Requirements

— Provide Positioning and Alignment Capability of OTV

During All Mission Phases

— Provide Translation Near Space Station

— Perform Midcourse and Final Rendezvous Maneuver

— Safe Operation Near Space Station

- 10% Margin

— Reusable

— Space-Based

Main Propulsion Requirements

— Provide Subcooled LH_/LO_ to the Main Engine

System (Engine NPSHTlequirements)

- Provide GH^GOj to Fuel Cell System (or TBD)

— Provide a Fill/Drain System Compatible with Space

Station

— No-Vent Operation Near Space Station Except When

Provision for Collecting Vapor Is Available

— Provide for Safe System Shutdown 20 mi from

Station (or TBD)

- Manrated and STS Safety (Fail Operational/Fail Safe)

- Propellant Utilization (PU) System

Avionics

- Provide Following Capabilities on Orbit

— Guidance, Navigation and Control Compatible with

GPS

— Telemetry, Tracking and Communications Compatible

with TDRSS, Ground Station and Space Station

— Data Management

- Safety Status and Control

— Subsystem Status Indication

— System Checkout

— Generate, Control and Distribute Electric Power to All

OTV Systems

— Minimize Single-Point Failures of Safety- and/or MissionCritical Functions

— Provide Safety-Critical Data to Orbiter

Structure

— Ultimate Factor of Safety = 1.40 or Greater

- Pressure Vessels Ml L-STD-1522 (Factor of Safety 2.0 on

Ultimate)

Thermal Protection

— Provide Protection for Aerobrake

— Provide Protection on Base Heating of OTV

— Provide Thermal Protection of Cryogenic Tanks on Orbit

— Provide Protection of Components, Lines, Avionics, ACS,

and Engine on Orbit

3.3.2 Baseline Subsystem Selection

This study was intended to consider the issues of space basing and

engine operational design criteria for a reusable OTV. Most of the

vehicle subsystems were identified and fixed in order to proceed

with the key issues and tradeoff studies of importance. Previous

study results were used in subsystem selection. The fixed

subsystems and the rationale for their selection follows.

3-11
Aerobrake - The baseline aerobrake configuration has the following

characteristics:

Shape

C

D

Aerobrake Surface Emissivity, e

Maximum Temperature (Nextel

Synthetic Fabric)

Maximum Loading

Maximum Stage Temperature

(e = 0.5 )

Viking type, 70-deg half-cone angla.

1.6 without aerospike (highest drag loads),

0.8

2500°F.

15 psf (approx 3.0-g loading).

750°F.

Past work at Martin Marietta (Ref 1 and 2) was used to derive

reasonable diameter and stage lengths considering the above

characteristics and convective heating. The effects of an

aerospike were not considered. From Reference 1, the sensitivity

of temperature to reentry weight and diameter were considered.

Reference 3 indicates that W/CDA should be constant for a given

stagnation point heating. However, given the state of aerobrake

analysis for an OTV, further detail was not attempted. With

information from References 1 through 3, Figure 3.3-1 was prepared

to relate ballistic coefficient and stage/aerobrake geometry.

Additional thermal protection could be required or multiple passes

could be used to reduce heating and brake diameter. The data in

Reference 3 indicate that the stagnation point heat flux can be

reduced about 30% with two passes, and up to 50% for three passes.

70deg

1.5

1.0

X

0.5

Assumptions:

— Viking Brake Configuration

- CD= 1.6 (Constant)

- Stage € =0.5 0

— Max Temperature Stage 750 l=

— Max Temperature Brake 2500 F

10

W/CDA = B psf

Figure 3.3-1 Aerobrake Design Parametric*

W = Reentry Weight

3-12
ACS - A G02/GH2 ACS was selected because of ease of servicing

and refueling. Commonality with space station attitude control

propellant would be a desirable feature for the OTV. Contamination

control would make a G02/GH2 ACS attractive for OTV instead of

a monopropellant or bipropellant system, especially if the OTV is

to fly itself into berth at the space station. Also, as a possible

backup for the main propulsion system, the high performance as well

as the commonality of propellants make G02/GH2 a logical

choice. Sixteen thrusters were chosen for the full three-axis

control capability and limited redundancy.

Avionics - A system compatible with the GPS, TDRSS, and space

station was baselined. The OTV should be capable of rendezvousing

with space station, performing targeting for the aerobraking and

phasing maneuvers, and performing midcourse corrections. The

electronics on board the OTV will also need to handle the

diagnostics information to be recorded or transmitted to the ground

for postflight or inflight checkout of the subsystems requiring

health monitoring.

Electric Power - The major power source for the baseline OTV was

fuel cells with feed from the G02/GH2 accumulator system.

Small batteries used as a backup of the fuel cell power system and

for use during times when the main propellant tanks are empty were

baselined.

Gaging and Propellant Utilization - A point sensor system of gaging

during flight operation to monitor the level of propellant in the

tanks was baselined. During refill of propellant in low-g at the

space station, a flowmeter or other suitable gaging device is

recommended to monitor the level of propellants.

A propellant utilization system to minimize the propellant outage

was baselined for the OTV design. For a multiple-tank

configuration, the outage between tanks must be minimized and the

mixture ratio at the engine must be controlled for proper outflow

of oxidizer and fuel. Centaur data, space tug studies, and later

OTV studies provided a rationale for the performance advantages a

propellant utilization system offers a cryogenic vehicle.

Pressurization - An autogenous pressurization system for a

L02/LH2 OTV would provide a simple, lightweight means of

pressurizing the main propellant tanks with little or no

contamination. Therefore, autogenous pressurization was chosen for

the study baseline vehicle.

Thermal - Based on the results of a low-thrust chemical

orbit-to-orbit propulsion system propellant management study

(Contract NAS3-21954); conceptual design and analysis of orbital

cryogenic liquid storage and supply systems, (NAS3-22264); and an

IR&D D-24S project concerning advanced propulsion concepts (a study

performed at Martin Marietta) a multilayer insulation (MLI)

3-13
thickness of 1/2 in was chosen for the baseline vehicle main

propellant tanks. The previous studies show that a thicker

insulation is desirable for vehicles that experience boiloff before

launch and during ascent from the ground. However, because the

baseline vehicle of this study was intended for space operations

alone, less than 1 in. of MLI is optimum from a total vehicle

performance standpoint.

Venting - A thermodynamic vent system (TVS) with an active

component (small pumping unit) may be the lowest weight method of

venting for a large cryogenic vehicle as opposed to a passive TVS

that which would require a relatively heavy and sizable heat

exchanger. However, the active TVS presents a servicing concern

for access to the pump unit for inspection and maintenance. The

alternative to the two TVS methods mentioned is to simply settle

the propellants with the ACS and vent vapor from the tanks.

For this study baseline vehicle, a weight representative of a TVS

with pumping units in each tank was defined. Servicing of the unit

is addressed later in this report, noting possible modifications

and concerns (Section 4.1).

Tankage - A ground rule for this study was that the vehicle would

be delivered to orbit in the STS. Although it would be desirable

to deliver the space-based OTV to orbit in its fully assembled

configuration, assembly on orbit is consistent with the required

recurring operations. Other geometry constraints for a space based

OTV are associated with the problem of aeroheating during the

aerocapture maneuver. In other words, a short stage is desirable

to minimize requirements on aerobrake size and thermal protection

for the forward end of the stage. Also the space station hangar

will certainly not be an unlimited envelope of volume in which to

perform vehicle inspection and servicing. Therefore, in addition

to the classic shuttle constraints, other constraints drive the OTV

toward reasonable compactness.

In addition to the overall geometric considerations for the OTV

(tankage and packaging), the concerns include an inspection and

replacement capability for many of the components on board the

vehicle. Modular arrangement of the tankage would be functional

from a contingency replacement standpoint in case of tank damage

for one reason or another (meteroid penetration, collision, MLI

replacement, etc). A routine interchange of tankage may present

problems with regard to the complex interface requirements and

handling. Nonetheless, an "open pack" arrangement of all

subsystems, including tankage, must be inherent in the design of a

space-based OTV.

An alternative to total vehicle modularity (where the vehicle size

can be scaled up or down to "fit" the mission/payload) is to have a

mixed fleet with at least two different sizes of vehicles.

Although these vehicles' tanks would not be intended for routine

changeout for mission/configuration flexibility, they would lend

themselves to system modularity for inspection, subsystem access,

3-14
and contingency replacement. The total number of eventual

missions, along with vehicle reusability limitations, may warrant

at least a two-vehicle fleet. For instance, a two-stage concept

capable of performing the large payload LEO-to-GEO missions and

also capable of returning each stage to LEO via aerobraking is a

possible solution. The two stages may not provide significant

performance advantages over a single large vehicle for a large

payload (if at all), but would provide a method of performing the

wide spectrum of missions with minimum offloaded propellant. In

other words, this additional configurational flexibility of two

differently sized compact vehicles, along with the ability to stage

them, would provide a much more efficient use of propellant for the

range of missions by flying with a higher mass fraction than with

the earlier vehicle/fleet (single large stage or modular vehicle)

concepts mentioned. Also, from a servicing standpoint, the two

compact stageable vehicle fleet would simplify the hangar and

vehicle handling equipment design because:

1) The two stageable vehicles are the same size (or more nearly)

than for the two-vehicle fleet concept consisting of a large

vehicle and a small one;

2) The two stageable vehicles are reduced in size from the large

single-stage vehicle. Their smaller size is also important

from an aeroheating standpoint during the aerocapture maneuver;

3) Two small vehicles may ease turnaround servicing time

requirements depending on mission frequency.

3.4 TRADEOFF STUDIES PERFOBMED

A number of tradeoff studies were performed to appropriately define

a vehicle and engine system that are to be space-based and

reusable. The following list itemizes the pertinent issues

identified:

1) Mixture ratio vs stage performance;

2) Number of engines;

3) Thrust level(s);

4) Throttling range(s);

5) Gimbaling vs differential throttling;

6) Fixed vs extendable/retractable nozzles;

7) Modular stage(s) vs staged vehicle vs mixed fleet (and

corresponding tankage and structure).

3-15
The difficulty involving most of these tradeoff studies is that

they are interrelated. For instance, the choice of engine number

was one of the most important tradeoffs made and directly affected

the selection of engine thrust, vehicle servicing, whether

gimbaling may be required, packaging concerns with the nozzle(s),

and strong consideration of engine reliability requirements.

Understanding the synergism of all these factors was important in

arriving at a justified conclusion as to the appropriate number of

engines for a space-based OTV.

The judgment criteria considered during the performance of these

tradeoff studies included:

1) Overall stage performance;

2) Individual subsystem impact;

3) Impact on servicing timeline, equipment, tools, and manned

involvement;

4) Probability for mission completion and for safe crew return

(reliability and acceptable failure criteria);

5) Mission/configuration flexibility;

6) Life-cycle cost, including all of the above criteria.

3.4.1 Major Configurational Characteristics of a Space-Based, Reusable OTV

Developing an OTV design that will be cost effective in capturing

all the future space missions involves more than simply designing a

single propulsion stage and providing it with reusability and

serviceability. Because of the wide range of mission payload

masses and mission requirements (such as deploy only or

deploy/retrieve), a single-stage design would not result in an

efficient use of propellent compared with other concepts.

Figure 3.4-1 shows the propellent requirement for completing the

Rev 6 LEO to GEO missions (47) between 1994-2000 for various

vehicle fleet concepts. The single-stage approach to capturing the

missions is the most expensive in terms of propellant usage. The

modular stage was sized so its half-tankage version (two L0£

tanks and two LH2^ would hold about 33,000 Ibm propellant and

deliver 11,000 Ibm to GEO. The full-tankage modular version holds

twice this amount of propellant and is capable of performing the

13,000 Ibm GEO deployment and retrieval mission.

3-16
2.21—

2.1

I

to

° 2.0

I

OC

CM

(O

K

1.9

1.8

1.7

Rev 6.1994 to 2000

47 LEO to GEO Missions

Reusable, Aerobraked

(Low Lift/Drag)

Single Stage

Modular Stage

Sfageable Fleet

Mixed Fleet

460 480

Figure 3.4-1 Vehicle Fleet Comparisons

The stageable fleet was sized so the first stage would deliver

16,000 Ibm to GEO. It will also perform the perigee burn for

delivery of the 13,000-Ibm payload. The first stage would then

return to LEO via an aerocapture maneuver. The second stage would

perform the GEO circularization burn and then the decircularization

burn and aerocapture maneuver for retrieval of the 13,000-lbm

payload. The two stages would be used separately for the smaller

mission payloads according to their capabilities. The propellent

capacities of the first and second stages are about 40,000 Ibm and

32,000 Ibm respectively. The smaller stage (second stage for the

13,000 Ibm deploy/retrieve mission) was sized to deliver 11,000 Ibm

to GEO because this size of payload and its frequency of flight

appeared to be an effective breakpoint for propellant savings.

However, both stages being sized to deliver 16,000 Ibm to GEO would

still provide an efficient means of flying all missions.

3-17
The mixed fleet simply consists of a large stage sized for the

13,000 Ibm deploy/retrieve mission and a small stage sized for the

smaller missions. The smaller stage was sized for a 14,000 Ibm

delivery to CEO. Unlike the smaller stage of the stageable fleet

that was sized for an 11000 Ibm delivery to GEO, the smaller stage

of the mixed fleet was sized for a larger payload because of the

difference in overall stage sizes of the mixed fleet. The

breakpoint in propelIant capacity for sizing the smaller stage of

the mixed fleet is also a result of the large number of payloads of

14,000 Ibm.

The main conclusion associated with Figure 3.4-1 is that a single

large stage design sized for the largest payload/mission is an

inefficient method of performing the missions from a propellent

usage standpoint (compared to the possible alternatives). The

modular stage, of course, has associated "scar" weight through its

modularity that the mixed fleet (two or more stages) does not

have. The stageable fleet, of course, would have a scar weight

associated with the staging interfaces.

An example of stage performance and character vs service life and

servicing requirements is selection of tankage. For instance, one

choice would be to provide the vehicle with complete meteoroid

protection for the vehicle's design life so the probability of a

penetration would be less than that acceptable. This design might

also alleviate any need to reproof-check (reverify) the tank(s)

between missions (unless an unforeseen incident such as a collision

occurred). Also, the removal of MLI for tank inspection would not

be required. The effect of dry weight on this particular design

would result in noticeably greater propellent usage than for a

higher performance (less protective dry weight) design. However,

the higher performing vehicle would demand a much higher level of

servicing activity. Therefore, the nature of a space-based

reusable vehicle is constrained by its impacts on the servicing

requirements.

3.4-2 Baseline Configuration Selection

The OTV characteristics ground-ruled early in this study included:

1) Deliverable in STS cargo bay;

2) L02/LH2 propellents;

3) Aerobraked (low lift-to-drag ratio);

4) Space-based and reusable;

5) Manrateable;

6) Able to capture Rev 6 missions.

3-18
A desirable characteristic to add to the list is that of high

performance to minimize the use of propellents (and therefore their

launch cost). However, the amount of onorbit servicing the

resulting vehicle would require must be considered. Without prior

knowledge of the results of a servicing/performance tradeoff study,

a vehicle configuration was chosen that represents a reasonable

compromise of performance and service intensity as well as

complying with the study ground rules.

The vehicle concepts depicted in Figure 3.4-2 and 3.4-3 are the

study baseline vehicle concepts. Schematics of the recommended

feed systems are shown in Figures 3.4-4 and 3.4-5. The vehicles

were sized to perform the largest Rev 6 delivery mission of 16,000

Ibm to GEO. The only missions requiring more total impulse (a

significant amount more) are the 13,000-lbm manned deploy/retrieve

missions to GEO. The rationale for sizing to the 16,000-lbm

delivery to GEO mainly lies in the advantages of performing the

profitable missions that comprise most of the mission model.

The low frequency of 13,000-lbm manned missions does not provide

the rationale for a dedicated large single stage. The baseline

configuration, however, is capable of providing the perigee boost

to the 13,000-lbm manned payload and an associated propulsion

system (cryogenic or storable). The manned payload with its own

propulsion system for GEO insertion and decircularization would

seem reasonable from several standpoints. First, if it were a

storable system, the boiloff problem might be eliminated for the

long onorbit time. Also, if the system were pressure fed, the

manrating complexities would be reduced. Because of the low number

of projected manned missions, the performance penalties associated

with the storable propulsion system concepts previously presented

would be minor relative to the overall mission model. This

scenario would leave the OTV as simply a delivery workhorse, for

the most part, and may even have implications of reducing the

operational requirements with regard to manrating.

Let us assume, however, that the 13,000-lbm manned payload will not

have its own dedicated propulsion system for GEO insertion,

decircularization, and possible reentry with no need to return to

the space station. The baseline vehicles shown in Figures 3.4-4

and 3.4-5 will serve well as the larger of two stages for a

stageable/mixed fleet. A case exists for a stageable and/or mixed

fleet from the standpoints of propellant savings, mission frequency

and turnaround time, rescue capability, etc. In any case, for the

purposes of this study, the baseline configurations provided

reasonable starting points for performing main propulsion system

tradeoff studies and determining the implications of space basing

and reusability on an OTV.

3-19
ORIGINAL PAGE S3

OF POOR QUALITY

VEHICLE DEFINITION CONFIGURATION SKETCH

Propulsion System: Space Based. Reusable ,

Propellant:__LOr/LH2 , MR: 6.0

Main Engine: Thrust 150QQ Ib . ISP; 480

Payloa<J Delivery to CEO 16000 _lbs, ASE Ib

ITEM

AVIONICS

STRUCTURES

THERMAL

PROPULSION

FU Tank .(2)

OX Tank (2)

Propellane Feed

Pressurlzaclon

Engine Assy (2)

40.

ACPS & ACCUMULATORS

AEROBRAKE

CONTINGENCY (15%)

DRY WEIGHT

PROPELLANT, NON-USABLE

ntlRNOUT WEIGHT

UO11,0 FT

PROI'KI.LAX I t.OAI)i:i)

STAGE INITIAL WT.

PROPELLANT, USABLE

MASS FRACTION

WEIGHT, LBS

972

362

418

217

310

63

738

350

1090

770

5900

640

6540

125

43318

49218

42333

.860

REMARKS

INTERNAL CORE

0.5" MLI. TVS

2219 AL, .025"

1.5 NON-OPTIMUM

INCLUDES PNEUMATICS

AUTOGENOUS

7500 Ibf EACH

INCLUDES GAS GEN

ft DIA

MARGIN AND RESIDUALS

42 HR MISSION

220 ACS G02/GH2

Main

USABLE/INITIAL

RETRACTABLE

NOZZLES

Figure 3.4-2 Space-Based Reusable Vehicle Definition, 2 Engines

VEHICLE DEFINITION CONFIGURATION SKETCH

Propulsion System: Space Based, Reusable .

Propellant; LO^/LH, . MR: 6.0

Main Engine: Thrust 15000 jb . ISP: 480

Paylond Delivery to CEO 16000 Ibs, ASE Ib

ITEM

AVIONICS AND POWER

STRUCTURES

THERMAL CONTROL AND VENT

PKOPUI.SJO.N'

FU TJII'K (2)

OX Tank (2)

PropelInnC Feed & P.U

I'ressuiizatIon

Engine Assy (4)

46 Ibm - ACTUATORS

ACPS & ACCUMULATORS

AEROBRAKE

CONTINGENCY (15%)

DRY WEIGH I

PROPELLANT, NON-USABLE

RURNOUT WEIGHT

I'KOIT.I.I.AMI LOAD) U

STAGE INITIAL UI.

1'ROPELI.ANl , USAHLi:

MASS hRACI ION

WEIGHT, LBS

972

610

362

410

213

434

63

528

300

855

712

5459

640

6099

42087

47546

41102

.864

REMARKS

INTERNAL CORE

0.5" MLI. TVS

2219 AL, .025"

1.5 NON-OPTIMUM

INCLUDES PNEUMATICS

AUTOGENOUS

3750 Ibf EACH

38 FT DIA

MARGIN AND RESIDUALS

42 HR MISSION

220 ACS G02/GH2

MAIN

irSABLE/INTTrAI.

FIXED NOZZLES -

Figure 3.4-3 Space-Based Reusable Vehicle Definition, 4 Engines

3-20
o

Tank Isolation

and PU Valves

Refuel 00

Notes: A) High pressure feed for GH2/G02 accumilator

B) Feed for contingency gas g«n«rator/turbo-pui<p/hMt exchanger unit for accunuletor charging

Legend

rtain Engine

with HE valves

Pneunatic

Valve

Quick Disconnect

Pressure

Regulator

El Latch Valve

EH F/D Valve

O Pressure Transducer

0 Filter

Figure 3.4-4 Main Feed System—2 Main Engine Schematic

Modular spherical tanks were chosen over parallel cylinders because

of increased performance and decreased surface area and subsequent

propellant boiloff. Also, a shorter length is highly desirable for

the vehicle to minimize aerodynamic base heating. The vehicle,

however, would not be deliverable in its flight configuration in

the STS orbiter bay. It would be assembled after initial delivery

to LEO if delivered in the cargo bay. The configuration of modular

tankage is based on contingency replacement of tankage only. For

3-21
example, if a tank suffered a meteoroid penetration that was

detected in a routine leak test, the tank would be replaced. The

vehicle is not intended for routine configuration modularity for

mission flexibility because at this time fluid disconnect

technology will not allow such routine freedom. Therefore only

contingency replacement of tankage is recommended.

Tank Isolation

and PU Valves

HoUt: *) High pressure feed for CH2/C02 eocumilator

B) FMd for contingency gas generator/turbo-punp/heet exchanger unit for accunulator charging

(lain Engine

with RE vtlves

2

-IPPneuiMtlc

Valve

Quick Disconnect

Pressure

Regulator

E3 Latch valve

EH F/D Valve

0- Pressure Transducer

Q Filter

Figure 3.4-5 Main Feed System—4 Main Engine Schematic

3-22
Vehicle configurations designed to require no tank modularity for

contingency replacement are a possibility. In other words, the

exterior structural shell and meteoroid shielding provide the OTV

tankage with the protection against probable penetration during the

vehicle's design lifetime. Based on current estimates, the weight

penalty, however, would significantly increase propellant usage

over the baseline configuration and overshadow the costs of

periodic tank replacement.

Based on the results of vehicle sizing over a range of mixture

ratios, the minimum vehicle weight consistently resulted in a

mixture ratio of 6.0. Engine performance vs mixture ratio

information from Pratt and Whitney (Fig 3.4-6) was used for the

vehicle sizing. Figure 3.4-7 shows the results for various

missions.

490 I—

480

I'

a

« 470

c

I

o

I

> 460

450

i Note: No Tank Pressunzation Flow I

5.0 5.5 6.0

Inlet Mixture Ratio

6.5 7.0

Figure 3.4-6

Estimated Effect of Inlet Mixture Ratio on Vacuum Specific Impulse

at Full Thrust

3-23
60 1

55 •

50 •

35

EFFECT OF MIXTURE RATIO

ON PROPELLANT

13K

DELIVER/RETURN

16K

RETURN'EITTY

5K DELIVER

2K RETURN

«

—r

5.5 6.0 6.5 7.0

MIXTURE RATIO (Mo/Hp) USABLE

Figure 3.4-7 Effect of Mixture Ratio on Propellant

3.4.3 Engine Design Criteria for Aerocapture

One of the main drivers of engine design criteria for the study

baseline vehicle may be the operational requirements during the

aerocapture maneuver. Whether aerospike operation is a feasible

means of controlling vehicle trajectory during aerocapture remains

an open issue.

The design philosophy for the aerobrake and engine system will

include several considerations. Minimizing the aeroheating

environment of the vehicle is a major concern in keeping the

thermal protection system weight for the vehicle to a minimum.

Therefore, the aerobrake must be properly designed to withstand the

predicted heat load (2500°F) and radiate it. In addition, short

vehicle length must be emphasized to minimize the base heating in

the wake of the flow around the aerobrake. The aerobrake must also

be configured to be compatible with the predicted airloads (up to

3-24
15 psf). Also, the aerobrake/vehicle configuration and operation

during the aerocapture maneuver should lend themselves to

aerodynamic stability to minimize the control authority

requirements.

The overall vehicle, aerobrake, and main propulsion system design

and subsequent operation must provide adequate control for

successful return from high earth orbit during normal main engine

system operation as well as during an engine-out condition.

Because the aerospike operation (engine operation during the

aerocapture maneuver) may be very effective for drag modulation

with blunt vehicle configurations as well as help minimize vehicle

structural and heat loads, this particular concept of operation and

brake design was carried through the main propulsion system

tradeoff studies. Based on studies performed to date, a 70-deg

Viking-shaped cone aerobrake appears to be a reasonable compromise

between drag and stability requirements.

Other aeroassist concept candidates include mechanically actuated

surface(s) for either lift capability or surface area variation for

drag modulation. A viable concept is one of vehicle eg offset and

roll for an angle of attack capability. This concept would most

probably provide a significant degree of control with a minimum

amount of probable failures.

Because of the wide variations expected in atmospheric

characteristics, it is questionable that a single degree of vehicle

control will provide total assurance for successful aerocapture on

a routine basis. For example, angle of attack (or rather, some

degree of lift capability) in addition to drag modulation may be

required for an adequate control capability. Possible failure

modes of the vehicle must also be considered, particularly for the

aerocapture maneuver because of the critical nature of steering

through a "thin" corridor and avoiding the obvious hazards of lack

of control and augering into the lower atmosphere.

Aerospike operation, particularly with two or more engines, may

provide this additional degree of control capability. Martin

Marietta Idea Report 83YD59, dated August 11, 1983, suggests that

differential throttling of multiple engines may provide lateral

motion (via induced angle of attack) by tailoring the bow shock and

subsequent streamlines around the vehicle. Whether this degree of

vehicle control is necessary or even feasible remains an open

issue. Figures 3.4-8 and 3.4-9 depict this method of control for

two- and four-engine vehicle concepts respectively.

3-25
Operations

- ACS for Roll

- One Engine on (THIM or PIM)

and One Engine Off; or One -

Engine in THIM and One

Engine in PIM

Attributes

- Main Engine Failure-Tolerant

— No Mechanical Actuators

- Drag Modulation and Angle

of Attack

Figure 3.4-8 Aerocapture Maneuver with Two Main Engines

Operations

- Differentially Throttle

All Engines

- ACS for Roll with

Engine Out

Attributes

— Main Engine Failure-Tolerant

— No Mechanical Actuators

— Drag Modulation and Angle

of Attack

Figure 3.4-9 Aerocapture Maneuver with Four Main Engines

3-26
Other open issues with regard to aerobrake design include:

1) Lack of data relative to variation of aerodynamic drag

coefficient and flow stability during aerospike operation;

2) Effects of aerospike on reducing aeroheating;

3) Base heating from aerobrake radiation as well as convective

heating;

4) Necessity of continuous throttling vs thrust steps for drag

modulation during aerospike operation.

The following discussion presents the recommendations for

aerobrake/main propulsion system design criteria based on the

analyses performed.

As discussed earlier, the aerospike operation has the potential of

providing a vehicle control capability in the upper atmosphere

(drag modulation), aerothermal shielding for the vehicle, and (in

conjunction with drag modulation) loads alleviation on the

aerobrake structure. Figure 3.4-10 shows the acceleration levels

estimated for various combinations of low-thrust TRIM (tank head

idle mode) burns and a high-thrust PIM (pumped idle mode) level.

Notice that two low-thrust burns before and following pumped idle

thrust help minimize the loads that affect the aerobrake

structure. The thrust range for aerospike operation is estimated

(based on present analyses) to be between 100 and 800 Ibf. For

thrust greater than 800 Ibf, the rocket effect begins to take over

and the drag reduction effect is negated.

Two or more main engines have advantages as well as disadvantages

with regard to the aerocapture maneuver. Among the advantages are:

1) Shutting down selected engines (or differential throttling

them) to add to control capability of vehicle via streamline

(bow shock) shaping and possible use of ACS for roll-controlled

steering;

2) Achieving several thrust "steps" at low levels for drag

modulation by various combinations of engine(s) operation(s)

such as off, TRIM, or PIM.

The disadvantages of having more than one engine include:

1) Size of opening in aerobrake;

2) Thrust vector offset with engine out;

3) Vehicle control complexity.

3-27
HP = 41.0nmi

High Thrust = 1500 Ib

4.0

3.0-

1.0

41.0nmi

FHigh Thrust = 1500 Ib

Low Thrust

HP = 41.0nmi

40 High Thrust = 500 Ib

£ 2.0

Z 1.0

HP = 41.0nmi

High Thrust = 500 Ib

r Low>

Thrust^

350 355 0 5

Relative Orbital Location, deg

10 350 355 0 5 10

Relative Orbital Location, deg

Figure 3.4-10 Trajectory Simulations—Aerospike Control

Contingency operation of performing the aerocapture maneuver by

using ACS aft-pointing thrusters to create an aerospike effect may

be possible. To target for a multiple-pass (2 or 3 passes) return

trajectory without main engine(s) operating, a certain amount of

vehicle control is still required. Present estimates suggest that

a maximum of 200 to 300 Ibf may be adequate for performing a 2- to

3-pass return. Therefore the use of ACS thrusters for the return

via multiple passes is a very feasible alternative.

Depending upon the results of vehicle/engine optimization,

considering engine length, diameter, and performance, nozzle

retraction may be required before entering the atmosphere. With

the emphasis on short vehicle length to minimize base heating in

the wake of the aerobrake, the long high-performance (high area

ratio) nozzles of the main engines will most likely protrude

through the aerobrake during the burns before entering the

atmosphere (up through GEO decircularization). The phasing and

rendezvous burns following the aerocapture maneuver are small and

the performance penalty for not reextending the nozzles may not be

significant.

3-28
3.4.4 Main Propulsion System Tradeoff Studies

The determination of major engine characteristics should depend on

several criteria—propulsion system reliability, performance,

serviceability, cost, etc. During Phase I of this study, geometry

and performance were given highest priority, keeping in mind the

implications of cost, manrating, reliability, and serviceability.

Cost and serviceability were considered in Phase II (Section 4.0)

when a life-cycle cost model of the OTV service life was developed

and tradeoff studies were performed. The main propulsion system

trade study ground rules were:

1) Total of 15,000 Ibf thrust;

2) 1, 2, 3, 4, 5, and 3 inline engines;

3) Fail operational/fail safe included in weight comparisons

(exception: single engine that is fail-safe only);

4) Long (retractable) and short (fixed) nozzles for each engine

system;

5) Two gimbal angles, 10, 20 deg (except for the 5 and 3 inline

engine cases which do not require large gimbal angles);

6) Long engine length chosen where length penalty balances

performance gain;

7) Adjacent nozzles separated by 1/2 ft.

The performance sensitivities for the baseline stage are as follows:

2.8 Ibm prope11ant and 146 Ibm propellant

Ibm dry weight second Isp

These were used in comparing propellant requirements and total

vehicle weights for various numbers of engines, varying gimbal

angles, nozzle design, and engine arrangements. Aerobrake size

(and consequently weight) must increase as it is moved aft to

package the engine(s) because of the base heating that would

otherwise occur at the forward end of the vehicle. Therefore the

engine envelope is a strong driver in overall vehicle configuration

and performance. (Aerobrake and structure weight sensitivity is 57

Ibm/ft further aft.)

Engine lengths for the retractable nozzle cases were chosen so

performance gains were balanced by the penalty of longer nozzle

length (in retracted configuration). Two axis gimbaling is

required for up to four-engine configurations for a two-engine-out

capability. The exception is for three engines inline with an ACS

for backup for decircularization at GEO and return. In this case,

the two outboard engines require one-axis gimbaling and the middle

3-29
one requires two axes. Five engines provide a two-engine-out

capability with only one engine requiring two-axis gimbaling. The

ground rule for adjacent packaging of engines required a spacing of

0.5 ft between nozzles.

To provide equivalence in failure criteria (fail operational/fail

safe or two-engine-out capability) for the two-engine system vs the

multiple engine (3 to 5) cases, the performance penalty of extra

loaded propellant for ACS return of the stage only from GEO was

calculated. Because the ACS system (G02/GH2) will operate at a

mixture ratio of approximately 5 to 1 and a lower ISp than the

main engines, extra propellant is required for contingency as a

backup mode as opposed to extra engine(s). Figure 3.4-11 shows

this weight penalty as a function of main propulsion system

performance. For instance, loss of the main propulsion system

(MPS) with a performance of 490 s results in greater contingency

propellant requirements than for MPS performance of 480 s because

of the greater difference in MPS and ACS performance.

500

400

fi 300

J

1

I

i

 2°°

.E

J

100

Contingency Operation Propellant Margin

- ACS or THIM Return from GEO

- I =450 s, MR = 4.0, AV = 6550 fps

sp

— Aerocapture: 125 Ibf Average Thrust for 10 minutes

(200 to 300 Ibf Maximum Thrust Required for 2 to 3-Pass Trajectory)

Total

Total for I = 455 s

4.69 MR

470 480

Main Engine I , s

_J

490

Figure 3.4-11 Contingency Operation Propellant Margin

3-30
The parametric vehicle data points were computed using the

sensitivities mentioned and engine data from Pratt and Whitney

Aircraft and Aerojet Tech Systems. These two sets of engine data

appeared to "bound" the variation of engine performance with thrust

level. In other words, an attempt was made to include the range of

proposed engine concepts (large to small) and their subsequent

inherent characteristics and estimated performance. Figure 3.4-12

shows the total vehicle weights resulting from the Pratt and

Whitney data and the results for the Aerojet data are shown in

Figure 3.4-13.

54

I

53

52

I

CO

 "

"£

J 50

49

..J3

cr

Failure

Criteria

Waived

Legend:

n 12-ft Retractable

(I, 3 Inline, 5)

A 12-ft Fixed

(1,3 Inline. 5)

• 10-deg Retractable

A 20-deg Retractable

O 10-deg Fixed

• 20-deg Fixed

2 3 4

Number of Engines

Figure 3.4-12 Initial Weight »s Number of Engines

3-31
53

52

51

ai

I 49

FAILURE

CRITERIA

WAIVED

Legend:

[] 12-ft Retractable

(I. 3 Inline, 5)

A 12-ft Fixed

(1. 3 Inline. 5)

B 10-deg Reinctnble

A 20-cley Rotiactable

O 10 (leg Fixed

• 20-deg Fixed

1234 5

3.4-13 Initial Weight vs Number of Engines (Aerojet Engine Data)

Example notations for the charts are:

12-ft retractable 1, 3 inline, or 5 engines that are 12 ft

aft of the tankage/accumulator subsystem

interface, with retractable nozzles;

10-deg Fixed - Engines allowed to gimbal up to 10 deg for

an engine-out capability with fixed

(standard) nozzles.

Table 3.4-1 shows the resulting conclusions and engine operational

requirements recommended from all Phase I activity.

3.4.5 Baseline Vehicle Configuration

Two baseline configurations were developed (Fig 3.4-14 and 3.4-15)

for two engines and four engines respectively. The two-engine

configuration is representative of an optimized vehicle

incorporating Pratt and Whitney engines with retractable nozzles

(as per the main propulsion system tradeoff study). Because the

tradeoff study showed short fixed nozzles to be optimum for the

Aerojet engine data and less penalty for lower thrust, the

four-engine configuration is representative of an optimized vehicle

incorporating four Aerojet engines.

3-32
Table 3.4-1

Phase I Main Propulsion System Tradeoff Study Conclusions and Engine

Operational Requirements

Conclusions

Results are sensitive to engine performance and geometry data.

Further investigate 1, 2, 3, or 4 engines in Phase II

— Vehicle subsystem impacts

— I triplications of reusability/servicing

— Mission flexibility

— Life-cycle cost.

Nozzle design

— Fixed at 60 in. for diameter reasons (Aerojet)

- Retractable to 60 in. for length reasons (Pratt & Whitney).,

- Benefits of retractable nozzles are more sensitive to available gimbal angle than fixed.

- Differential throttling shows little or no benefit.

— 5 or 3 inline engines not competitive with other configurations.

Requirements

— Duty cycle

— 6 burns/mission (21-minute burn time)

— Throttling (or step thrust levels)

— Flow stability and drag modulation during aerocapture maneuver (aerospike).

- Reduce thrust after perigee burn to extend engine life (TBD).

— Low-thrust missions (LSS).

- Gimbal

— Up to 20 deg provides significant benefits in total vehicle mass and propellent consumption.

— Specific impulse

— Greater than 460 s.

— Lifetime and reliability implications.

- Mixture ratio

— 6.0 oxidizer/fuel provides lowest weight stage and propellent usage.

The baseline main propellant tanks are supported at two

locations—forward and aft of each tank (Fig. 3.4-14 and 3.4-15).

The forward supports react forces in the X and Y directions while

the aft supports react against in the X, Y, and Z directions. The

structural attachments are designed to be easily detached and all

fluid and electrical lines are equipped with quick-disconnects

(QDs) to facilitate onorbit construction, servicing and/or

replacement of the tanks.

The baseline configuration employs a welded tubular aluminum

primary structure. Tubular members were selected on the basis of

their high strength-to-weight and stiffness-to-weight ratios.

Additional stiffness gains and weight savings may be realized if an

all-composite structure is used. While an investigation of this

possibility should eventually be instituted, it was outside the

scope of this study.

The desire to minimize vehicle length and diameter placed strict

limitations on the size and envelope of the primary structure

located between the main propellant tanks. To avoid an undesirably

heavy structure, the primary launch and aerobrake loads on the

tanks in the X, Y, and Z directions are reacted through the tank

aft structural attachments. The structure forward of the tank aft

ends then need only react tank loads in the X and Y directions in

addition to payload and equipment module loads in the X, Y, and Z

directions. This concept greatly reduces the size and weight of

this forward structure.

3-33
ORIGINAL PAGE !S

QUALfTY

I

a

3-34
ORIGINAL PAGE®

OF POOR

3-35
The equipment module is attached to the primary support structure

forward of the main propellant tanks. The GH2 accumulators are

internal to the equipment module and the G02 accumulators are

supported off of the module's aft end. The forward bulkhead of the

module serves as the payload interface. Detachable units

containing the OTV ACS, avionics, pressurant regulation and

control, and power systems surround the equipment module. Each of

these detachable units is attached at four structural QD points in

addition to any fluid or electrical QDs that may be required.

These units are serviced by removal and replacement with a new or

refurbished module.

In addition to the ACS units attached to the equipment module, four

aft-pointing ACS thrusters are attached to the aerobrake support

structure. These aft-pointing thrusters penetrate through the

aerobrake, permitting them to provide a limited aerospike in the

event of a main propulsion system failure. This thruster location

eliminates any thruster efficiency losses that would result from

thruster plume impingement on the aerobrake if these thrusters were

located forward of the aerobrake (near the payload interface plane).

An OMV is attached to the OTV aerobrake and the aerobrake is

attached to its support structure with the same type mechanisms and

at the same relative locations. This concept enables the OMV to be

attached to the OTV without the aerobrake attached. Berthing of

the OTV with the refueling station also employs these same

mechanisms, with an RMS accomplishing the actual berthing

operation. Propellant handling is accomplished through fluid lines

with QDs that could penetrate through the aerobrake if it were left

attached.

3.4.6 Space-Based OTV Propulsion Subsystems Design

A space-based OTV will require a unique design to accommodate the

needs of onorbit basing. This is necessitated by both space basing

and the desire for reusability. These two requirements are beyond

those imposed on previous upper stages. The space station will

provide the logistics and manned presence. From a servicing

standpoint the space station design will also influence the OTV

design. Such items as hangar design, resupply accommodations, and

the astronauts' servicing capabilities will all influence the

ultimate design. Because logistic considerations such as

consumables resupply and spares inventory will represent a large

portion of the routine operating expenses, the OTV design should

account for these requirements. It is also desirable to reduce

consumables both in quantity as well as in number. Fewer

resuppliable quantities on the OTV will simplify the servicing

operations. The subsystems design in the following description

will attempt to address these concerns in addition to such

"traditional" ones as minimum mass, cost, and maximum performance.

3-36
Current LH2/L02 upper stage designs have been highly optimized

for the present expendable mode of operation. High performance/low

weight were the major design drivers. In addition to the main

propellant, previous upper stages have also carried several fluids

to be used for various functions, e.g., ^H/^ for attitude

control, and He for purging, pressurization and valve actuation.

This represented the best compromise toward the goals of highest

attainable performance and reliability. Reusability will shift the

optimum breakpoints from one-time usage design life toward periodic

replacement for certain high-maintenance or life-limited items such

as the main engines. These items would become prohibitively

expensive to make last the life of an OTV. Space basing adds yet

another constraint in that the number of feasible maintenance tasks

is severely limited by the available manpower and support equipment.

It is anticipated that the upcoming space station will develop

technology for GH2/G02 thrusters and associated components in

the thrust range of interest for an OTV to achieve a GH2/G02

attitude control system (ACS). Once these gases are stored on

board, they could be made available to serve other functions as

well. The electric power system (EPS) fuel cells could obtain

their supply from the GH2/G02 storage bottles (accumulators),

eliminating separate storage and servicing. (Assuming fuel cells

compatible with propellant-grade cryogenics.) GH2 could be used

in place of He as the actuator pressurant. The GH2/G02 system

would be recharged by a simple check-valve arrangement tied to the

autogenous pressurant lines between the engine(s) and the tanks.

The ACS could also be used as a backup for a failed main engine if

a contingency means for charging the GH2/G02 accumulators were

provided. On the basis of these arguments, the use of GH2/G02

on a space-based OTV is proposed. The following subsystem

descriptions will further illustrate the advantages of a fully

integrated cryogenic OTV design with a preliminary point design.

The main propulsion system design is shown schematically in Figures

3.4-4 and 3.4-5 for two and four main engines, respectively. The

two-engine and four-engine cases have been selected as the

baseline. Only the number of engine isolation valves (propellant

and pressurant) and propellant manifolding are affected in

comparing the two schematics. A single valve is installed at each

tank outlet to provide tank isolation and for use by the propellant

utilization (PU) subsystem to control tank-to-tank outage. A

pneumatic valve with sufficient metering authority for the PU

system will be required. The downstream manifold communicates with

the main engine(s), an overboard dump, refueling quick-disconnects,

and the feed for a contingency auxiliary power unit (APU) used to

charge the GH2/G02 accumulators. Each main engine is provided

with isolation valves as a backup for the main engine valves.

These valves are also needed to facilitate removal of the main

engines. (There may be a blanket pressure in the lines.) A dump

system is shown on the liquid side in case the main propellents

need to be evacuated (anticipated to be a contingency event).

Residuals left at the end of a mission would be recovered. A

3-37
single QD and valve are shown for the refueling lines. Parallel

QDs may be needed depending on the unit's reliability, which cannot

be resolved at this point. Also not shown are filters for the main

propellants. Presumably, each main engine inlet will have a filter

that is serviced each time the main engines are changed out. The

propellants should also be highly filtered by the time they are

placed on board the OTV. Autogeneous pressurization will be

provided by the main engine(s) at a temperature and pressure

dictated by the autogenous pressurization and accumulator systems.

The gases will be reduced in temperature and pressure by redundant

regulators although a simple valve and orifice may be sufficient.

A thermodynamic vent system (TVS) will most probably be needed but

is not shown because the exact requirements for the TVS are not

known at this time and it should have a design similar to those

used on existing stage designs. Recovery of the TVS vent gases is

not anticipated because the relatively small amount of mass is not

expected to be enough to warrant recovery. A pump would be needed

to boost their pressure from near zero to the approximately 2000

psi needed for the accumulators. The mass and power requirements

of these pumps are believed to outweigh any benefit gained by

recovering the gas mass. However, future design efforts should

reexamine this conclusion as the systems become available and more

realistic analyses can be made. The TVS will present some

servicing complications because most of the components will be

mounted inside the propellant tanks. A high degree of reliability

and/or access to the tanks will be needed if servicing the TVS

components is necessary. Alternative means of controlling tank

pressures and temperatures should be investigated.

The gaseous subsystem is shown in Figure 3.4-16. Essentially, four

accumulators are fed through check valves from the main tank

autogeneous pressurization lines. From these accumulators, gases

are regulated and distributed to four "user" subsystems—the ACS

thrusters, the EPS fuel cells, the APU, and GH2 for pneumatic

actuators. Four accumulators were chosen so the system could be

divided in half, providing full redundancy. It may become

desirable to use only half the system for particular missions. The

accumulators will be sized by the worst-case mismatch between

supply and demand. Isolation valves are provided for either of the

two gas sources (autogeneous or APU) so recharging can be done by

either. Each tank is provided with its own filter/check-valve unit

in full isolation from the other tank. A redundant filter/

check-valve unit is provided should either of the primary units

fail. Similarly, separate regulators are used downstream from the

tanks so the system can be run with only one tank (per gas). These

regulator units are manifolded on their downstream sides so they

may feed the redundant isolation valve pairs controlling

distribution to the four end uses. A common regulated pressure has

been specified because this is the simplest and lightest

arrangement. The total system is expected to be designed at one

time so all the individual units could be optimized together. It

is not anticipated that "off-the-shelf" components will be

3-38
available to drive the design. The EPS fuel cells will probably

further reduce the pressure from that supplied by the system

shown. A nominal regulated pressure of 500 psi is envisioned.

This pressure will likely be determined by the needs of the

pneumatic actuators and/or the ACS thrusters. The APU unit

depicted between the two tank pairs and will be discussed

separately in the following paragraphs.

Fron CH2

Press Line

0 L*J

Ha—na-Ha' j EH

RCS Pnuenatic EPS

Systen

EPS PCS

Legend

OPressurant tank

Pressure Regulator

Latch valve

P Liquid Punp

I

 T I Turbine

EH F/D valve

O Pressure Transducer

Q Filter

.^*>

U Relief value

2 Quick Disconnect

3.4-16 GH2/GO2 Subsystem Schematic

3-39
Nominal operation of the accumulator system will be as follows. At

the beginning of a mission, at least one of the tanks for each gas

will have a pressure sufficient to start the APU and allow

activation of the engine valves. Normally, the APU would not be

used to charge the accumulators. The firing of the main engines

will accomplish this. The APU would be used as a backup for the

main engines to sufficiently charge the accumulators to enable

returning the OTV. Each main engine firing following the first

will keep the accumulators properly charged. A long stay at GEO

may require that the APU be started to recharge the accumulators.

On returning to the space station, the accumulators could be

offloaded to a safe blanket pressure. Alternatively, they could be

run down to a safe pressure just before docking with the space

station. In either case, they are not expected to require complete

evacuation, except when the system needs servicing. Routine OTV

turnaround would not require that they be empty, only that they be

at a safe pressure. Evacuation after each mission would require a

lot of time and energy in handling. In the space environment a

small leak will not constitute a safety hazard in itself. Rather,

because a small leak will need to be repaired to prevent the loss

of propellants, the blanket pressure will provide the gas needed to

detect the leak. Once detected, however, the leak repair may

require evacuation of the system. For this reason, the space

station may require a means of charging the accumulators back to

the necessary premission conditions. Potentially the APU could

also provide this if the gas generator could bootstrap itself. A

schematic for the APU is shown in Figure 3.4-17.

Hon-piop

Vent

Figure 3A-17 Auxiliary Power Unit Schematic

Gaseous propellants from the accumulators are combined in the gas

generator from where they are expanded through a turbine and twin

heat exchangers before being released overboard. Liquid

propellants from the main engine feedlines are pumped to a high

pressure and vaporized in the cold side of the twin heat exchangers

3-40
before being introduced into the accumulators. The size of the APU

will depend on its intended use. If it will serve as a backup for

the main engines, its reliability will need to be high. Redundant

APUs may be needed in this case. Steady-state operation vs duty

cycle operation will also affect its ultimate size. The expected

worst case is where the APU is used with the ACS as a backup for a

failed main engine(s). The primary mass use would be for the ACS

propellant but additional propellant will be needed to provide tank

pressurization and to power the APU. However, because of the

difficulty in building a very small turbopump unit, a better

compromise between unit mass and efficiency may well be a larger,

more efficient unit running at a low duty cycle. The following

first-level analysis explores this tradeoff.

The APU is to be used primarily as a backup system for a failed

main engine. As such, a premium will be placed on its

performance. As currently envisioned, the APU exhausts its

products nonpropulsively and thereby "wastes" them. The design

must therefore make the most out of what it uses. This means

minimum mass use for the APU. However, dry weight must also be

kept to a minimum. A balance must be struck between the APU dry

mass and the mass consumption. A target weight of 50 Ibm was

used. This is felt to be a reasonable penalty for the OTV to pay

for the flexibility gained by the APU. Approximate masses are

shown in Table 3.4-2, based on a steady-state APU design.

Table 3.4-2 APU Mass Estimates

Component

Gas Generator and Valves

Turbopump Unit

02 Heat Exchanger (HX)

Hj Heat Exchanger

Miscellaneous

Mass, Ibm

8

22

3

12

5

Individual performance will be covered in detail in the following

paragraphs. The design presented below is somewhat conservative

because it is based on a first-order analysis. An EPS use rate of

0.85 Ibm/h for 48 h plus 78 Ibm of ACS propellant at a 5:1 MR was

used to size the accumulators. The mixture ratio through the cold

side (liquid pumps and heat exchangers) of the APU was sized for

steady-state ACS usage. The hot side (GG, turbine and HXs) was

sized on a per Ibm/s of cold side flow basis so some scaling could

be performed.

The APU cold side consists of the liquid pumps, the heat exchangers

and the associated valving. Flow rates, heat fluxes, temperatures,

pressures and power requirements are shown in Table 3.4-3. For the

steady state design point shown, the pumping requirements are quite

modest. Liquids at or below the saturated temperatures were used

for the inlet conditions to the pumps. Vaporization was assumed to

3-41
take place entirely in the heat exchangers. Accumulator inlet

conditions of 200°R and 735 psi were chosen to minimize the heat

flux jjeeded. Efficiencies of 90% were used for the pumps. The

JANNAF tables were used to determine the enthalpy needed by the

fluids. From this information, the power needed per Ibm/s of cold

side flow could be determined. As shown in Figure 3.4-17, a single

turbine is used to drive the two pumps to facilitate the analysis.

This establishes a cold side mixture ratio. Separate turbines and

pumps would allow the mixture ratio to the accumulators to be

varied at a cost in simplicity. Future analyses may justify the

additional complexity this would entail. Also shown is a

gear-driven L0£ pump. Again this was done for ease of analysis.

A detailed turbopump design was beyond the scope of this study.

Table 3.4-3 Cold Side Parameters

Fuel

°2

H

2

Mass Flow, Ibm/s

0.833

0.167

Turbop

Ti,°R

165

38

[imp

Pi, psi

20

20

Po, psi

750

750

Ah, Btu/lb

1.59

5.04

Heat Exchanger

Ti,°R

167

42

To,°R

200

200

Pi, psi

750

750

Po, psi

740

740

q, Btu/lb

11.24

98.54

The APU hot side consists of the gas generator, the turbine, the

heat exchangers and the associated plumbing. Table 3.4-4 shows the

hot side parameters. Gas generator operation at 2500°R and 150 psi

was used. The operating pressure was chosen to give a good

pressure drop for the injector. The temperature was selected as

the highest possible without generating excessive thermal material

concerns. A higher operating temperature would be desireable to

reduce the mass flow needed to condition the cold side fluids.

These operating conditions yield a gas generator fuel-rich mixture

ratio of 1.35:1. The turbine needs 14.9 hp to drive the pumps,

assuming a turbine efficiency of 70%. The heat fluxes shown in

Table 3.4-3 size the mass flows to the hot sides of the heat

exchangers. The fluxes shown in Table 3.4-4 include 5% extra for

losses to the environment from the gas generator and heat

exchangers. A heat exchanger exit temperature of 500°R was chosen

to avoid freezing the water in the gas generator exhaust products.

The resulting heat exchanger effectivenesses are 80% for Q£ and

76% or H£ based on these temperatures and flows. This implies

rather large heat exchangers with both boiling and convective heat

transfer. The heat exchanger design was carried no further than

needed to estimate weight.

Table 3.4-4 Hot Side APU Performance.

Fuel

°2

H

2

Gas Generator

Ti,°R

200

200

Pi, psi

735

735

m, Ib/s

0.021

0.016

Turbopump

Ti/To, °R

2500

2333

2500

2333

Pi/Po, psi

150

113

150

113

Ah, Btu/lb

10.5

Heat Exchanger

Ti/To, °R

2333

500

2333

500

Pi/Po, psi

113

103

113

103

q, Btu/lb

11.80

103.2

m, Ib/s

0.004

0.034

3-42
As mentioned earlier, the design is based on steady-state operation

per Ibm/s of cold side flow. For four 100-lbf thrusters firing to

return the OTV from GEO orbit, a mass flow to the thrusters of

0.847 Ibm/s is needed at an Isp of 472 s. The resulting APU mass

usage of 0.0315 Ibm/s leads to an effective Isp of 455 s at a MR

of 4.7. There are several potential means of improving these

results. All involve reducing the hot side mass flow. Lowering

the accumulator inlet temperature and/or raising the gas generator

operating temperature would both be effective. Improving the

efficiencies of the pumps and the turbine would also help, of

course. However, these have already been chosen close to the

physical limits. Only percent improvements could be gained.

Running the gas generator oxidizer-rich (MR of 89:1) improves the

mixture ratio to 6.4 but reduces the ISp to 376 s. This is a

result of the higher molecular mass of oxygen, which may also

present material problems. A means of potentially improving the

efficiencies and simultaneously increasing the gas generator

operating temperature would be to operate the APU at a low enough

duty cycle to improve the efficiencies while not incurring

transient losses. The accumulators would need to be sized to

accommodate this as well as most of the rest of the gaseous

subsystem. As a backup mode to a failed engine, using the APU and

ACS thrusters offers improved performance (455 s and 4.7 MR vs 450

s and 4.0 MR) over tank head idle mode. Another route to

substantial improvement is an alternative means of heating the

gases after they have been pressurized. The turbine and pumps use

only 8% of the power used to condition the gases. A solar-powered

evaporator that supplies 100 Btu/s to the fluids could be used

although it would be quite large and massive. The APU will justify

itself only as a backup for totally failed main engines (2-engine

case) or where the operational advantages previously mentioned

become needed.

3.5 PHASE I CONCLUSIONS AND RECOMMENDATIONS

Based on the configuration and performance tradeoff studies of

Phase I, the following major conclusions resulted in addition to

the main propulsion system tradeoff study conclusions shown in

Table 3.4-1:

1) Investigate the range of vehicle/engine systems further (Phase

ID,

a) Implication of reusability/servicing,

b) Life-cycle cost;

2) Spherical tankage with contingency replacement;

3) Retractable nozzles benefit only with large (15- to 20-deg)

gimbal angles;

3-43
4) GH2/G02 accumulator system supplying ACS, EPS, pneumatics

and contingency APU;

5) Choice of engine number and nozzle design strongly depends on

engine performance characteristics and manrating failure

criteria;

6) Stageable fleet concept should be given further consideration.

These conclusions remain consistent with the philosophy of the

baseline vehicle development of Phase I.

3-44
4.0 PHASE II - SERVICING, RELIABILITY AND LCC

This section describes the space based OTV servicing scenario,

engine configuration reliabilities and life-cycle costs (LCC) for

the respective engine configurations.

4.1 MAINTAINABILITY

4.1.1 Baseline Vehicle Servicing Character

The baseline vehicle configuration(s) developed during Phase I were

used in defining an appropriate approach to servicing a space-based

reusable orbit transfer vehicle. The objective was to develop a

reference servicing operational scenario for the life of the OTV

that appears reasonable considering life-cycle cost (LCC),

logistics, space station facilities, technology availability, etc.

Relevant tradeoff studies were then conducted to verify that the

reference scenario is the lowest LCC approach.

A reasonable approach to servicing a space-based reusable OTV

onorbit involves more than simply modularizing the vehicle so

everything imaginable may be easily removed and replaced.

Modularization seems to be a key word in the rhetoric surrounding

the issue of space servicing; however, the understanding of the

word leaves much to be desired. In examining the baseline vehicle

configurations, one finds items that should be included in a

simplified servicing approach of removal and replacement at the end

of a reasonable service life or following a failure of some kind.

These items include:

1) Avionics;

2) Batteries or fuel cells;

3) ACS thrusters or cluster modules;

4) Main engine(s).

These subsystems/components lend themselves to "modularization"

either because of their inherent geometric peripheral packaging,

their character of being at the "edge of things" schematically, or

their simple interfaces. Also, and maybe coincidentally, these

appear to be items that have shorter expected lifetimes than such

items as structure, tankage, isolation valves, etc. So it would

seem that providing for easy servicing of these short-life items

would be adequate. In further investigation, however, this

approach of designing the OTV for long life with the exception of

the easy service items makes little sense. For instance, if a

component other than those intended for easy serviceability were to

4-1
fail, the option of returning the "space-serviceable" vehicle to

earth for unscheduled servicing does not appear attractive from a

cost standpoint. Typical components/subsystems that are not

easy-access items for simple remove and replace operations include:

1) Main tankage that may suffer meteoroid penetration and require

replacement;

2) Thermodynamic vent system that may include moving parts and

require replacement (internal to main propelIant tanks);

3) Multilayer insulation that may degrade as a function of time

on-orbit or may degrade as function of meteoroid penetrations

(that may leave microscopic holes on the exterior but literally

shred the underlying layers);

4) Valves, filters, regulators, vents, etc that are part of the

main propulsion feed and pressurization systems; and also the

accumulator system that feeds the ACS, EPS, and pneumatics

(that ideally would be welded into the system).

These typical items are not necessarily buried within the

propulsion system in all cases. However, their integration in the

propulsion system provides difficulty in arranging them for

individual removal/replacement or servicing in the space

environment. The ability to detect degradation or failures in each

of these items separately is also a significant design problem.

Some compromises must be made as to the design of the vehicle for

modularization (and proper degree thereof), component redundancy,

fault detection and isolation, component design lifetime, and

service intensity requirements.

The first step with the baseline vehicle configuration was to

examine the design's components and subsystems for estimated

lifetimes, failure modes, failure effects, and failure frequency

estimates. Grouping the components into reasonable modules for

contingency removal and replacement in addition to defining the

module interfaces, geometry, and servicing character was essential

for further development of the appropriate servicing scenario.

Therefore a failure modes and effects analysis was performed for

the baseline design as a starting point in addressing the

maintainability aspects of the OTV.

4.1.2 Failure Modes and Effects Analysis (FMEA)

A FMEA was performed to analyze the proposed OTV propulsion system

concept. Although the data available were not sufficient to

perform a complete FMEA, this method of analysis was used to

further refine the design by considering the effects of incipient

failure modes and to provide input to recommend replacement

intervals and component grouping arrangements for

remove-and-replace operations.

4-2
The FMEA was structured according to the functional block outline

illustrated in Figure-4.1-1. Wherever possible and considered

appropriate, the OTV propulsion system was broken down to the

subassembly level.

1.0 Propulsion System ' .

1.1 Aerobrake

1.2 ACS and Accumulator

1.2.1 ACS

1.2.2 APU

1.2.3 Accumulator

1,3 Main Propulsion

1.3.1 Main Engine Assemblies

1.3.2 Propellant Feed

1.3.3 Pressunzation

1.3.4 Tank Assemblies

1.4 Avionics and Power

1.4.1 Guidance and Navigation

1.4.2 Communications

1.4.3 Electric Power

Figure 4.1-1 OTV Propulsion System Functional Outline

Design features and operating characteristics for the OTV

propulsion system were evaluated to identify the various failure

modes. The failure modes were then evaluated as to their potential

severity and likelihood of occurrence. These classifications are

tabulated.

Failure Severity

Severity

Classification

4

3

2

1

Failure Likelihood

Failure

Likelihood

Classification

1

2

3

Severity Category

Category I— Catastrophic. Failures that may cause death or system loss.

Category II— Critical. Failures that may cause severe injury or major system damage that will

result in mission loss.

Category III— Marginal. Failures that may cause minor injury or minor system damage that

will result in delay or mission degradation.

Category IV— Minor. Failures not serious enough to cause injury or system damage, but that

could result in unscheduled maintenance or repair.

Description

Low probability of occurrence some time during life of system.

Moderate probability of occurrence some during life of system.

High probability of occurrence some time during life of system..

It is assumed the OTV will require that the system be designed to

fail operational in the event of a single failure mode and fail

safe in the event of two failure modes. This assumption was used

as a criterion for evaluating the compensating provisions for each

failure mode and assigning the appropriate severity

classifications. Table 4.1-1 lists the results of the FMEA

exercise. Table 4.1-2 lists items identified as catastrophic and

critical failures and recommends the design considerations to

downgrade the severity classifications.

4-3
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4-11
Table 4.1-2 Severity Results and Recommendations Based on FMEA

ORIGINAL PAGE IS

OF POOR QUALrTY

Control Coeponent

Nuobter Daicrlptlon

Bovarity

Failure Claiilflnodea Failure Elltct an Syatoe citlan

Racoaandattona

To Oonnqrtdl Bavarlty

1.2.1.1 laolatlan Eitornal Loak Loit of Proptllint

Lttch Vilvt

Cataatrophlc Bick up llolltlon vilvaa.

1.2.3.1 Litch Valvoe Eitarnal Leak Lou of prepillint. Cataatrophlc lick up vtlvu.

1.3.2.3 Englni

lioUtlon

Vilvt

Eitornal Luk Loti fuol. Cataetrophlc Racoaaond latch v«lvt upatroaa

to dagrado affect to loiing

function of only am angina.

1.3.2.a *pu Cryo Fall Cloaatf No contingency aoda Cataatrephic

Faad Valva tccuaulitor oparatton.

Entarnal Leak Lou of propallant and Cataatrophlc

laaa of altalon.

Daalgn ihould raflact

radundant vilvaa in parillal.

Contidar lelectlng/dovelaplng

aara reliable coaponanta or

raconflgurtng ayataa.

1.3.3.4 EVA Taat Part Eitarnal Laak Loia of praaaurliatlon. Cataatraphlc Canildar talacting/davaloping

aora raliabla coaponanti or

reconfiguring ayttaa.

1.3.4.1 LO^ Storage Natarald

Tank Panatratlon

1.3.4.2 LH,

Tank

Minor loit of propallant,

potentially aara aarloua

if not detected.

HI I Deterioration Unknom it thla tlee.

1.3.4.2 LH, Storage Utterold

Penetration

Minor leaa of propallant.

potentially aora aarlaua

if not detected.

ML1 Deterioration Unknonn at thla tlee.

Critical

Critical

Tanka ihould ba verified by

tilting and fracture aachanice

imlyale. Alao, aatabliah a

aathod of Ink datactian.

Naad to datarelna affecta of

deterioration over a prolonged

period of ttoo.

Ttnka ahoutd ba vartflad by

tilting and fracture eechanlca

inilytla. nlao, aatabllih a

aathod of laak datactlon.

Naad to datarelna i««tcti of

deterI or aTlon over a prolonged

period of tlee.

1.3.4.9 Fluid Puep Puap Failure

k lapallar

Loaa of propel 1 ant by

vantlng or burning.

Critical Eniure eathod of aarvlcablllty

or coeponanta raliablllty

aicaada tank life.

The FMEA was used as an input source to estimate the number of

missions before recommended component replacement and/or overhaul.

These estimations were based on the failure modes, likelihood of

failure, and best engineering judgment of realistic, obtainable

goals for scheduled replacement to ensure safe operation and

availability of the vehicle. Evaluation of the estimated

replacement intervals disclosed that the missions to replacement

fell near 30 mission intervals for several subsystems (actually

from about 30 to 50 missions). Examples of estimated lifetimes are

(1) main engines - 30-mission life, (2) ACS thrusters - 40 to 55

missions, and (3) main tankage - 100 to 200 missions.

4-12
4.1.3 Component Grouping for Servicing

The preliminary OTV baseline design prepared during the Phase I

effort has been further refined to identify orbit replaceable units

(ORUs) that enable the OTV to be serviced at the space station. A

study was performed to determine what varying degrees of

serviceability would imply for the modularity of the entire

vehicle. Table 4.1-3 shows the results. Fluid system components

were divided into logical groups that should minimize the logistic

costs of servicing while maximizing the capability to handle

unforeseen failures. All major electronic components were left as

separate ORUs, there being no great advantage to a larger

grouping. Because no single component has an expected lifetime of

less than 30 missions, the grouping of similar life components into

modules was not needed as earlier anticipated. Figures 4.1-2 and

4.1-3 show the grouping of fluid system components and Table 4.1-4

depicts the module interfaces, geometry, accessibility and

operations anticipated.

The seven modules shown in Figure 4.1-2 for the main propulsion

system include all fluid components except the main engine, some

transducers and the pressurization system vent valves. Vent valves

will likely be permanently installed in the plumbing and located

remotely for functional reasons. If their reliability proves low,

they will be serviced as single components.

Five modules are envisioned for the GH2/G02 subsystem as shown

in Figure 4.1-3. The check valves and regulators are expected to

have the shortest life. Thirty missions are shown for these

components and they are therefore included as part of large

modules. If they do prove troublesome they would be isolated into

smaller modules so they could be replaced individually after they

have failed. Designing for more and smaller modules is undesirable

because of the increased number of fluid connectors necessitated

(more potential leaks). The APU is the most difficult item to be

serviced because of its many and different fluid connections.

Despite its rotating machinery, its life is expected to be at least

30 missions because of its contingency nature (backup for failed

main engines or possibly during excessively long onorbit hold

periods).

Table 4.1-3 Degree of Onorbit Servicing

Level

1

2

3

4

5

Type

Resupply

Subsystem-Level Module R/R,

Mechanical & Electrical

Subsystem-Level Module R/R,

Mechanical, Electrical, Fluid

Major Component Remove/Replace, Mechanical, Electrical,

Fluid

Piece Parts

Cumulative

No. of Items

Serviced

2 to 5 Fluids

5 to 15

20 to 50

60 to 100

?

Typical Items

L02, LH2, He, N2H4/ GN2

RCS, EPS, Avionics, Instrumentation, Etc

Main Engine, RCS Thrusters, Fuel Cell(s),

Tanks, PCA, Etc

Inlet Valves, Tank Valves, TVS, Turbopump

Assembly, MLI, Etc

Seals, Bearings, PC Boards, Etc

4-13
QRiQJMAl PAGE 63

OF POOR QUALITY

Space Baaed OTV - Main Ited system - Z Main Engine Schematic

Regulation and

Control- Typ of 7

Pressurization

Isolation - Typ of 4

Pressur.

Isolation Valves

Typ of 2

Quick Disconnect

Notes: A) High pressure feed for CH2/C02 accunulator

6) FMd for contingency gas generator/turbo-puRp/heat exchanger unit for accunulator charging

Legend

(lain Engine

with HE valves

Pneunatic

Valve

Quick Disconnect

nodule Contents

S3 Latch Valve

0 Pressure Transducer

0 Filter

0 Oriface

Figure 4.1-2 Main Feed System Modules

4-14
Legend

oPressurant tank

Pressure Regulator

Latch Valve

Liquid PIMP

Turbine

O Pressure Transducer

M Filter

U Relief Valve

3 Quick Disconnect

Figure 4.1-3 Gaseous Subsystem Modules

The modules depicted in Figures 4.1-2 and 4.1-3 are further defined

in Table 4.1-4. The first column corresponds to the fluid modules

identified in addition to the avionics, fuel cells and thruster

clusters. The second column defines the number and type of

interfaces needed. Rough geometry, access and operations type are

shown in the third column. Modules needing frequent servicing such

as engines, ACS thrusters and main tanks are given an access rating

of 1 tno obstructions) to simplify the remove and replace (R/R)

operation as much as possible. The other modules were judged

4-15
Table 4.1-4

Table 4.1-4

Vehicle Module Descriptions

Type and

Number Of

Modules

Type &

Number of

Interfaces

Geometry, Access,*

Packaging, Operations,

etc

Main

engines

(2 to 4)

Accumulator

check valve

assembly (2)

1 Mech

1 Elec

5 fluid

[2 very high

pressure for

G02 and GH2

autogenous pressures,

1 for GH2 pneumatics,

(high pressure),

2 for L02 and LI^l

60 x 60 in DIA

Access - 1

Ops - extensive

fluid R/&R

Tanks

L02 (2)

LH2 (2)

ACS U6)

4 valves per

thruster

APU (1)

(gas

generator)

Mech, Elec

1 liquid

low pressure,

1 low pressure

gas

Mech, Elec,

2 high -

pressure gas

2 very highpressure gas,

2 liquid

(low pressure) ,

Elec & Mech

LH2 - 11 ft DIA

L02 - 8 ft DIA

Access - 1

Ops - fluid R/&R

10 x 5 in DIA

Access - 1

Ops - fluid R/&R

1 ft x 1 ft x 2 ft

Access - 2

Ops - extensive

fluid R/&R

3 highpressure gas,

Elec & Mech

12 x 12 x 6 in

Access - 3

Ops - fluid R/&R

Accumulator

pressure

regulator

assembly (2)

Main pressure

control

assembly (2)

3 highpressure

Elec & Mech

1 high-pressure

gas,

1 low-pressure

gas,

Mech, Elec

12 x 12 x 10 in

Access - 3

Ops - fluid R/&R

Accumulator

latch valve

set (7)

2 very highpressure gas,

Elec, Mech

6 x 6 x 4 in

Access - 3

Ops - fluid R&R

8 x 6 x 12 in

Access - 3

Ops - fluid R/&R

Access:

1 No obstructions

2 Remove access Panel

3 Remove other Items

4-16
Table 4.1-4 Vehicle Module Descriptions (Concluded)

Type &

Number Of

Modules

Type &

Number of

Interfaces

Geometry, Access,*

Packaging, Operations,

etc.

Main tank

pressure

isolation

assembly (4)

2 lowpressure gas,

Elec, Mech

Access:

1 No obstructions

2 Remove access Panel

3 Remove other Items

6 x 6 x 8 in

Access - 2

Ops - fluid R/&R

Vent and

fill QD

assembly (2)

Tank isolation

and PU flow

control (4)

Engine

Isolation

valve (4)

Autogenous

pressurization

isolation

valves (2)

Fuel cells u;

Avionics

(4)

1 low-pressure

liquid,

1 high-pressure

pneumatic

gas,

Elec, Mech

2 liquid

(low pressure),

1 high-pressure

pneumatic gas,

Elec, Mech

2 lowpressure

liquid,

1 highpressure

pneumatic gas,

Elec, Mech

2 very highpressure gas ,

Mech & Elec

2 highpressure gas ,

1 low-pressure

liquid (water),

Elec, Mech

Elec, Mech

3 x 1 ft Dia

Access - 1

Ops - fluid R/&R

6 x 8 in Dia

Access - 1

Ops - fluid R&R

6 x 8 in Dia

Access - 2 (3)

Ops - fluid R/&R

4 x 4 x 12 in

Access - 3

Ops - fluid R/&R

1 ft x 2 ft x 2 ft

Access - 1

Ops - extensive

fluid R/&R

1 x 1 ft x 6 in

Access - 1

Ops - fluid R/&R

4-17
according to their anticipated physical location. The operation

type was determined by considering both the access and interface

complexity of the module. Any fluid R/R implies inerting of the

plumbing on both sides of the module, the removal and replacement

of the module, and finally checkout of the system to verify

successful operation. An "extensive fluid R/R" is used for modules

affecting more than one fluid circuit such as the APU or main

engines.

Determination of the module component groupings basically played

the benefits of few fluid connections against the opposing

logistics concern to capture the most components with the lowest

launch mass, i.e., replace the entire fluid system as opposed to

replacing each component. Further considerations were physical

proximity, functional similarity, failure frequency, physical size,

and such operational factors as requiring venting of the entire

fluid system (the OTV is stored with a residual blanket pressure in

the tanks and fluid lines to aid in leak checking and to minimize

propellant losses). Examples of the above are tank isolation

valves, which are are modules because of their physical location

and their replacement as individual units requires the evacuation

of only one tank (Fig 4.1-2), the vent and fill Q/D modules, which

have only two fluid interfaces, contain four components and are

located together on the aerobrake (Fig 4.1-3), and the APU, which

does not include the control latch valves because this would

require the evacuation of parts of both the GH2/G02 and main

propulsion systems. In addition the latch valve pairs are grouped

as modules. The low-pressure latch valve pairs are shown mainly

because of their number and the inability to group them into a

larger set (Fig 4.1-3). Their number allows one spare to back up

several potential failures. On the other hand, the high-pressure

latch valve pairs are identical for the GH2 and G02 sides and

therefore lend themselves to grouping into a larger module, which

minimizes the number of fluid connections required. This is

important because of the this system's high pressure.

4-18
It must be emphasized that the fluid system modules shown are

preliminary at this point and should be further refined as a more

detailed cost analysis is undertaken to understand more fully the

logistic influences. This will aid in determining a cost effective

means of providing the necessary spares. A more extensive

operational analysis will also refine the necessary human factor

considerations, space station facility and service requirements,

and estimate the time required to complete the module servicing

operations. The component failure frequency rates can be further

considered as more accurate component life data are developed,

presumably when the actual components are in their development

phase.

The results of the FMEA and component grouping include:

1) Recommendations for component and subsystem redundancy;

2) Recommendations for fault detection instrumentation and methods

of fault isolation;

3)' Recommendations for reasonable groupings of components into

remove-and-replace modules;

4) Design considerations for both mission success and vehicle

serviceability;

5) General maintenance philosophy.

4.1.4 Space-Based OTV Servicing Considerations

Space-based servicing of an OTV has been outlined in sufficient

detail to arrive at OTV and support system servicing requirements.

The space station facilities needed and their functional

requirements have been identified. The Impact of logistics and

space-serviceable design on the OTV design has been detailed.

Using the space station (SS) as a launch and refueling platform

will allow the decoupling of the space transportation system (STS)

earth to low earth orbit-(LEO) and the OTV LEO to geosynchronous

equatorial orbit (GEO) legs of payload delivery to GEO. The

shuttle will no longer be forced to launch in a window dictated by

the payload delivery, but rather on a periodic basis that would

allow optimization of ground resources for routine flow. The

burden of meeting the launch window then falls upon the SS/OTV

system. This implies the need for a highly dependable OTV and OTV

support system if the launch windows are to be reliably met.

The OTV support system will in part consist of SS facilities

capable of performing routine maintenance and certain contingency

repair procedures. It will also need an efficient logistics

function to provide needed spares and consumables in a cost

effective, timely manner. Implied by this is a highly developed

health monitoring system for the OTV and its subsystems. This

system must be capable of diagnosing items in need of attention

4-19
early enough so the necessary preventive action can be scheduled

and lengthy downtimes avoided. All this is made very challenging

by the fact that the SS will be able to provide only very limited

manned support because of the restricted number of men available,

the difficulty of working in the space environment, and the demands

of other SS activities.

Because none of the hardware actually exists, it was necessary to

make a few assumptions and establish sensible ground rules to allow

the study to proceed. These are shown in Table 4.1-5 and are

briefly discussed. Because the objective of this study was to

identify engine impacts with regard to servicing, detailed design

of the SS support facilities, etc was not attempted. For instance,

all refueling operations are assumed to be performed on the SS

instead of at a remote propellant farm. Operationally, the only

impact is on the timeline. The operations to be performed remain

similar. The major assumptions show up in Table 4.1-5 while many

of the smaller assumptions will be noted in the text as

appropriate. The study ground rules were use of the space station

as the OTV base, STS shuttle as the launch vehicle, manrating of

the OTV, L02/LH2 propellants, and the use of an aerobrake with

a low lift-to-drag 'ratio. From a servicing standpoint, L02/LH2

propellants, manrating, and the aerobrake present the greatest

drivers. While the aerobrake itself may not need much servicing, a

fixed aerobrake restricts OTV maneuvering about the SS, drives

hangar design, and complicates engine servicing. Manrating implies

a high degree of reliability/redundancy which in turn affects the

integrity of servicing operations. The L02/LH2 propellants

have a major impact on the propellant storage and transfer systems

and to a lesser extent on the engine servicing requirements.

Principally the latter was concerned only with engine changeout

implications and the required health monitoring system and its

requirements.

As previously mentioned, all space-based OTV servicing was assumed

to be at the SS and a means to maneuver the OTV about the SS was to

be provided. Specifically, the hangar and refueling depot were

assumed attached and controlled from a permanent OTV control

station at the SS. The OTV control station will control all

OTV-related operations—data handling, refueling, line-of-sight

(LOS) proximity operations, maintenance scheduling and procedures

[except extravehicular activity (EVA)], and SS inventory control.

The OTV was assumed to be under ground control for the LEO-GEO-LEO

.phase of the delivery missions. Both the baseline Rev 6 mission

model and the SS mission model (Ref 5, Vol 3) indicate an OTV

launch frequency of one every two weeks to one month. Therefore, a

two-week turnaround was used as the ground rule.

Given the above assumptions and ground rules, the general OTV

servicing flow was estimated as shown in Table 4.1-6. From this

list of operations, those pertinent to engine and OTV servicing

were further broken out to allow an operational and functional

analysis that revealed the SS facilities needs and the engine

servicing impacts. These were used as a baseline against which

alternative servicing concepts were explored/evaluated. Also,

contingency operations such as unscheduled maintenance are

discussed relative to the impact on the baseline functional flow.

4-20
Table 4.1-5 Space-Basing Study Assumptions

Space Station to provide up to A men/day , 8 hrs/day

Hangar, refueling area, and storage facilities attached to Space Station

OTV ground controled except when within Space Station line-of-sight

2 EVA/11VA men per EVA operation

Routine tasks automated as much as possible

Space Station to provide storage for necessary spares

OTV missions to average two week intervals

The OTV to be moved about SS by SS RMS(s)

Times for "routine" operation, not development period

OTV RCS to provide OTV control for "prox ops"

Payload mated to OTV prior to OTV refueling

Table 4.1-6 OTV Mission Flow

Operations Timeline

Premission Operations

— Payload and Payload ASE Delivery to Space Station by STS

— Uplink Mission-Specific Software to Space Station Computer

Mission Preparation Operations

— Move Payload to Hangar—Perform Payload Checkout

— Move OTV to Hangar

— Verify OTV Readiness-Update OTV Mission Computer

- Mate OTV and Payload-Verify OTV/Payload Interfaces

- Move OTV/Payload to Refueling Area

— Secure OTV/Payload and Connect Umbilicals

— Perform Propulsion System Check

— Chill Down and Fill Main Tanks with Required Propellants

— Resupply RCS and Pressurant (ff Necessary)

— Disconnect Umbilicals-Release OTV/Payload/Deploy from Space Station

— "Small" RCS Burn to Separate OTV and Space Station

— Pass Mission Control from Space Station to Ground

Perform Mission(Return to Space Station Line-of-Sight

Postmission Processing

— Pass Mission Control from Ground to Space Station

— Safe Main Propulsion at TBD Miles from Space Station

— RCS Burn to Space Station Rendezvous

— Space Station Capture of OTV/Payload, Berth at Refueling Area

— Connect Umbilicals and Offload Propellant Residuals

— Downlink OTV Mission Data to Space Station and Ground Computers

— OTV Exterior Visual Inspection

— Perform Postmission Propulsion System Checks

— Disconnect Umbilicals and Move OTV to Hangar

- Demate Payload (If Attached)

— Prepare Payload for Ground Return (If Necessary)

— Space Station and Ground Computers Return OTV Status and Service

Requirements

- Perform OTV Service As Required

— Prepare OTV for Storage and Move to Storage

- 7 Days

- 7 Days

- 7 Days

- 2 Days

-24 h

-20 h

-19 h

-18 h

-17 h

-16.5 h

- 8h

- 7h

- 6h

- 1 h

- 0.5 h

0.0 h

0 to 3 Days

3 to 8 Days

80.0 h

80.5 h

81.5 h

81.5 h

82.0 h

83.0 h

83.0 h

86.0 h

87.0 h

96.0 h

100.0h

120.0 h

122.0 h

168.0 h

4-21
A "top-down" approach was first used to divide up the nominal

two-week turnaround so the maximum available time to do tasks could

be delineated. Next, specific individual tasks were considered

"bottom-up" in that actual times and equipment needed to perform

comparable tasks on the ground were determined. In this fashion,

areas of further research were identified. For the purposes of

this study, the shorter of the two times was used to assemble the

timelines shown. Included with the operational analysis are

columns indicating facility needs, intravehicular activity (IVA),

EVA and delta time.

Tables 4.1-7 and 4.1-8 indicate tasks, facilities, and time data

for the baseline OTV turnaround for vehicle servicing and engine

servicing respectively. Complete mission turnaround is shown to

take approximately 10 days. This is driven primarily by the

LEO-GEO-LEO time and the OTV postmission processing.

Table 4.1-7 OTV Servicing Overview

Operation

Main Propellant— Resupply

Avionics— Scheduled Maintenance

— Module Test

— Module Replacement

- ACS Update

Avionics— Health Monitoring

Maintenance

— Module Replacement

— Module Repair

Avionics— Mission-Peculiar

— Module Replacement

— Reconfiguration

Tanks— Scheduled Maintenance

— External Inspection

-" = PU and TVS System

Tanks— Unscheduled Maintenance

— Tank Removal

— Insulation Repair

— Transducer Replacement

- PU and TVS System

Tanks— Mission-Peculiar

— Tank Reconfiguration

RCS-Scheduled Maintenance

— Leak Check

— Transducer Check

RCS— Resupply

RCS-Health Maintenance

— Transducer Replacement

— Thruster Replacement

Structure and Aerobrake— Health

Monitoring Maintenance

— Aerobrake Refurbishment

— Structure Repair

Facilities

Resupply Area, RMS

Cryo Tanks, Umbilicals

SS Hangar

SS Computer

EMU, HPA, Lighting

SS Computer

SS Hangar

EMU, HPA, Lighting

SS Workshop

SS Hangar

EMU, HPA, Lighting

EMU, HPA, Lighting

Resupply Area

CCTV Monitor

SS Computer

SS Hangar

Resupply Area, RMS Console

EMU, HPA, Lighting

EMU, HPA, Lighting

SS Hangar, EMU, HPA

Resupply Area

Resupply Area, RMS Console

Resupply Area, Umbilicals

SS Computer, Pressure

SS Computer

Resupply Area, Umbilicals

SS Hangar

EMU, HPA, Lighting

Hangar, EMU, HPA, Lighting

—

Resupply Area, EMU, MMU

SS Hangar, EMU, HPA

Tools

Resupply Software,

Control System

-

Test— Access

LRU ASE, Removal

SS & ACS Software

-

LRU ASE, Removal

Electronics

—

LRU ASE, Removal

LRU ASE, Removal

_

RMS + CCTV

Test + Access

..

RMS, CCTV, Tank ASE

Insulation Repair Kit

LRU ASE, Removal

LRU ASE, Removal

—

RMS, CCTV, Tank ASE

-

RCS Software

RCS Software

RCS Software

—

LRU ASE, Removal

LRU ASE, Removal

—

Aerobrake Repair Kit

LRU ASE, Removal

Delta

Time

7.0

6.0

3.0

3.0*

0.5

-

1.0»

1.0»

—

1.0+

2.0+

2.0

1.7

0.3

—

2.0

3.0

1.0

2.0

—

2.0

1.0

0.5

0.5

2.0

~

2.0

2.0

—

3.0

2.0

Space Station

(SS) Manhours

IVA

7.0

6.0

3.0

3.0

0.5

-

1.0

1.0

-

1.0+

2.0+

2.0

1.7

0.3

—

2.0

3.0

1.0

2.0

—

2.0

1.0

0.5

0.5

2.0

-

2.0

2.0

—

3.0

2.0

EVA

-

—

6.0

—

-

2.0

2.0

-

2.0+

4.0+

—

-

-

—

4.0t

6.0

2.0

4.0

—

4.0t

-

~

—

—

—

4.0

4.0

—

6.0

4.0

Total

7.0

12.0

3.0

9.0

0.5

-

3.0

3.0

—

3.0+

6.0+

2.0

1.7

0.3

-

6.0

9.0

3.0

6.0

—

6.0

1.0

0.5

0.5

2.0

-

6.0

6.0

--

9.0

6.0

Note:

* Mission average expected for all avionic modules, 1 hour for contingency.

t Tank replacement only for modular OTV; some EVA assistance anticipated.

4-22
Table 4.1-8 Engine Servicing Overview

Operation

Engine— Turnaround Maintenance

— Analysis of Flight Data

— Lockup Pressure Decay

— Engine Valve Operation Check

— Nozzle Visual Inspection

— Nozzle Extension Check

— Gimbal Actuator Check

— Connect Umbilicals

— Turbopump Torque Check

— Ignition System Check

— Instrumentation Checkout

- Solenoid Checkout

— Disconnect Umbilicals

Engine— Periodic Maintenance

— Setup Operations

— Turbopump Boroscope

— Thrust Chamber Inspection

— Engine LRU Replacement

— Tool Stowage

Engine— OTV Engine Remove and

Replace

— Setup Tools

— Attach Engine Fixture

— Disconnect Engine

— Move Engine to Storage

— Pick Up Replacement

- Align and Attach

- Check/Verify QDs

— Store Tools

Engine— Unscheduled Maintenance

— Repair in Hangar

— Repair LRU in SS

— Repair on Ground

Facilities

SS Hangar

SS and Ground Computer

SS Computer, Refuel

SS Computer, Refuel

SS Computer, Refuel

SS Computer, Refuel

SS Computer, Refuel

SS Hangar

SS Computer

SS Computer

SS Computer

SS Computer

SS Hangar

SS Hangar

SS Hangar

Power, Lights

CCTV Monitor

Power, Lights, EMU, HPA

SS Hangar

SS Hangar, RMS, EMU, Foot

Restraint, Lighting

_

..

~

~

..

-

—

--

SS Hangar, RMS, EMU, HPA

SS Hangar, RMS, EMU, HPA

SS Hangar, RMS, EMU, HPA

Tools

..

Engine Software

Engine Software

Engine Software

RMS + CCTV

RMS + CCTV

RMS + CCTV

RMS

Engine Software

Engine Software

Enginr Software

Engine Software

RMS

-

Engine, LRU ASE

Boroscope

RMS, CCTV

Engine, LRU ASE

Engine, LRU ASE

Engine Fixture,

Engine Disconnect,

Protective Covers

-

-

-

—

-

-

-

--

Above

Above plus LRU ASE

Above plus Engine

ASE

Delta

Time

3.4 h

2 Days

0.5

0.5

0.6

0.2

0.2

0.3

0.3

0.3

0.5

0.3

0.2

4.0

0.5

1.0

1.0

2.0

0.5

5.0

0.5

0.5

0.5

0.2

0.1

0.7

2.0

0.5

2.0+

3.0+

2.0

Space Station

(SS) Manhours

IVA

5.4-

2.0

0.5

0.5

0.6

0.2

0.2

0.3

0.3

0.3

0.5

0.3

0.2

4.0

0.5

0.5

1.0

1.5

0.5

3.8

0.5

0.5

0.5

0.2

0.1

0.7

0.8

0.5

2.0+

4.0+

1.7

EVA

-

--

..

-

--

.-

..

-

-

--

--

..

-

6.0

1.0

1.0

•-

3.0

1.0

6.0

1.0

1.0

1.0

0.4

0.2

1.4

0.5

0.5

4.0+

4.0+

3.4

Total

5.4

2.0

0.5

0.5

0.6

0.2

0.2

0.3

0.3

0.3

0.5

0.3

0.2

10.0

1.5

1.5

1.0

4.5

1.5

9.8

1.5

1.5

1.5

0.6

0.3

2.1

1.3

1.0

6.0+

8.0+

5.1

The following discussion will cover the OTV mission flow. The

"generic" OTV mission was anticipated to begin early with the

mission planning activities and other operations by the payload

program. The SS begins its preparations two to three days before

the payload is delivered by the STS. The payload is delivered a

nominal one week early, principally at the convenience of the

shuttle, and is stored onboard the SS awaiting premission

processing. Facilities for handling the payload are presumed

available. Their exact manifestation is immaterial, but should

include a means to mechanically restrain the payload and provide

dormant power, data handling, and thermal protection.

A day before the mission the payload is moved into the servicing

hangar for final checkout operations. No EVA is anticipated, but

could be used if the payload had nonstandard interfaces or required

some minor contingency repair. For normal operations all

premission payload checkout operations will be handled remotely.

The four hours of checkout time are primarily to allow for payload

operations that may be more economically performed on the SS than

on the ground. For example, payloads could be launched without

fluids to relieve designing for launch loads.

4-23
Following successful payload checkout, the OTV will be moved to the

servicing hangar for mating with the payload. A final health check

will be made of the OTV and the mission parameters will be loaded

into the OTV main computer. The OTV-to-payload interface (I/F) is

assumed to be primarily mechanical with a minimal electrical I/F

provided. The electrical I/F would be standardized as well as the

mechanical I/F. If nonstandard I/Fs were used, the timeline would

need to be modified to allow for OTV I/F modification. No fluid

I/Fs are anticipated. Two payload I/Fs are implied—one for the

OTV and one for the STS. Once mated and the I/Fs verified, the OTV

and payload will be moved to the OTV refueling area.

Refueling is performed as the last major operation in the prelaunch

flow to avoid bringing a fully loaded OTV into the hangar and to

minimize boiloff. This implies a refueling area capable of

accommodating the OTV, aerobrake, and payload. The OTV is docked

and refueled on the aft end. A fixed aerobrake will complicate the

refueling area design. Presumably a door will be provided in the

aerobrake to allow the fluid umbilicals access to the OTV fluid

interfaces. The refueling operation itself is the subject of much

debate and is simplified here into a tank chilldown operation

followed by the bulk fluid transfer. Simultaneous fluid transfer

is assumed. Nonhypergolic fluids and "no-leak" quick-disconnects

(QDs) should allow this. Also, the attitude control system (ACS)

propellants and pressurants are resupplied in parallel with the

main propellants. Pressurant needs should be minimized as much as

possible because of the inordinate costs of resupplying pressurants.

Following resupply, a final OTV checkout can be performed (gimbal

actuators, pressure checks, etc). The OTV and payload are then

disconnected from the refueling area and deployed from the SS. The

timeline shown assumes that the SS remote manipulator system (RMS)

releases the OTV and payload combination with a small delta-V

relative to the SS. The OTV uses ACS burn to give additional

delta-V of about 3 fps, allowing swifter OTV and SS separation. At

a safe distance from the SS the OTV control is passed to the ground

and the delivery mission begins. An orbital maneuvering vehicle

(OMV) could be used to accomplish the same operation.

Space station control resumes following the return of the OTV to a

safe area within LOS of the SS where OTV safing is performed. This

may comprise venting the OTV propellant residuals. However, this

timeline assumes that the cost of propellants is sufficiently

important to warrant recovery. Safing would then primarily entail

deactivation of the main engines and the ACS if an OMV is used to

recover the OTV. The OTV is returned to the SS following safing

either by the OTV ACS or an OMV. The OTV is berthed at the

refueling area.

4-24
If safing were to entail venting of propellants, it may have a

major impact on the OTV. Nonpropulsive vents must be provided with

the appropriate valving and controls. Venting through the engines

would be possible but could impose undesirable characteristics on

the engine. Additionally, the resulting thrust would need to be

accounted for. An OMV would not be able to do this because the OMV

would likely be mated to the aft end of the OTV so its thrust can

act through the OTV/payload center of mass.

Fostmission processing is essentially the reverse of the premission

flow. The residual propellants are removed after docking at the

refueling area. Liquid propellants are returned to the SS cryogen

tanks and gaseous propellants are recovered for use by the SS. ACS

propellants would also be returned to storage to aid in the

accuracy of premission loading (mass measurement errors would

otherwise accumulate). It may be desirable to leave a blanket

pressure of propellant gases in the tanks for structural reasons.

During propellant offloading, the SS data handling system will

downlink mission data from the OTV and return the bulk of these

data to the ground where they will be processed. Additional data

will have already been sent to the ground during the mission. Some

data will also be retained by the SS computer to allow SS personnel

to begin postprocessing scheduling. Quick data analysis and

turnaround will be essential to efficient OTV servicing. The bulk

of the analysis software is assumed to reside on the ground because

it isn't cost effective to burden the SS computer or personnel with

this task. Two days are allowed for the ground to return a

preliminary postmission maintenance schedule to the SS. During

these two days, the OTV would be returned to the hangar if it still

has a payload attached. Otherwise, the OTV would be moved to its

storage area.

Because postmission OTV servicing highly depends on what

maintenance needs to be performed, the routine servicing flow will

be discussed along with a separate discussion of such major

contingency operations as engine removal or aerobrake repair.

Because crew time is expected to be a valuable commodity, routine

operations will be highly automated. The ground processing of

mission data will also perform an optimization of servicing tasks

and return a timetable detailing the exact operations to be

performed. An approximation is only possible now because both

routine (every mission) and contingency operations will be

interwoven to effect the optimization. This approximation appears

in Table 4.1-6 made up of the scheduled maintenance tasks from

Tables 4.1-7 and 4.1-8.

While the OTV is still berthed at the refueling area, a propulsion

system check will be performed. This check will be in support of

ground analysis of flight data to determine items in need of

maintenance and to execute tests designed to isolate any anomalies

detected in the flight data. The objective of this checkout is to

provide early detection of failures that can be remedied in the OTV

maintenance to follow. Also, tests that require pressurants will

4-25
need to be performed here. If the OTV were equipped with removable

tanks, the tank operations would be performed in this area.

All maintenance operations will be performed in the servicing

hangar after the schedule has been returned. The first operation

will be an overall OTV visual inspection. This could be done via

EVA, but will likely be done with a closed-circuit TV (CCTV) and

monitor. In this case, sufficient mobility must be given to the

CCTV to allow it to reach all areas of the OTV. Most likely, only

specific areas will routinely be inspected such as the engine

nozzles, aerobrake, and OTV exterior. CCTV movement could then

proceed in a preprogrammed manner and the crew would only override

to inspect questionable areas.

The servicing hangar is expected to provide for more extensive

checkout umbilicals than those provided at the refueling area so

specific tests of the avionics can be run. All umbilical actuation

will be automated to avoid EVA costs. EVA is anticipated only for

nonroutine module changeout operations, nonroutine inspection, and

other infrequently performed operations it wouldn't be cost

effective to automate. After checkout umbilicals are attached, the

avionics will be checked via checkout software and equipment

carried for this purpose. Any anomalies will be noted and factored

into the maintenance schedule relayed from the ground. Any EVA

operations would be performed following schedule finalization. EVA

module changeout would be performed on all items so identified in

the preceding checks. This assumes that the proper modules are

already on board the SS and that the modules were designed for EVA

replacement. Both of these assumptions will be discussed more

completely later. No modules that will require changeout after

every mission have yet been identified. If this were the case,

this would likely be accomplished robotically using only one IVA

crewman, once again to avoid EVA costs. Table 4.1-9 lists the

example EVA-replaceable modules. This table includes estimated

times and anticipated interfaces. Because ACS modules may involve

fluid disconnects, two operations are shown to illustrate the

differences. The fluid QDs lengthen the time because of the

additional effort required to assure the crew's safety

(installation of spill containment shrouds and checkout following

installation).

Table 4.1-9 EVA-Replaceable Modules

Subsystem

Avionics

Reaction Control System

Electric Power System

Structure & Aerobrake

Module

Main Computer

TT&C

C&DH

Guidance

RCS Module

Thrust er

Power Supply

Aerobrake

Contents

CPU, I/O Unit, Memory

Antenna(s)

RF Electronics

Gyros

Tanks, Valves, Thrusters

REA Valves, Thrust Chamber

Etc

Fuel Cells, Valves, Tanks,

Heat Exchanger & Pumps

Aerobrake Module

Interfaces

Mechanical, Electrical

Mechanical, Electrical

Mechanical, Electrical

Mechanical, Electrical

Mechanical, Electrical

Mechanical, Electrical & Fluid

Mechanical, Electrical

Mechanical

R/R Time,

hours

0.5

0.2

0.5

0.4

0.6

1.5

0.5

1.0

4-26
Two major contingency operations identified are engine removal

(which could also be routine) and aerobrake repair. Aerobrake

repair is included at this point as a possibility. It is too early

to say exactly what aerobrake repair implies or what type of

failure it may suffer. Holes could be repaired either by patching

or panel replacement. Aerobrake removal to ease servicing would be

desirable but isn't a contingency operation. This would be

included in overall processing flow near the end of premission

processing and the beginning of postmission processing.

Several levels of engine maintenance are identified as detailed in

Table 4.1-8. Two types of scheduled maintenance are

shown—operations performed after every flight and those performed

every 10 missions. These latter operations are more extensive and

performed in addition to the regularly scheduled maintenance. They

also include EVA operations (turbopump inspection and orbital

replaceable unit (ORU) replacement). The engine removal and

replacement operation is detailed as well as three possible

unscheduled engine repair operations. Unscheduled maintenance

could occur on the engine while it is attached to the OTV. This

would involve essentially replacement of an ORU that failed

prematurely. A removed engine could have a failed ORU repaired in

a SS workshop if future analysis showed this to be feasible and

cost-effective. Any major repair of the engine will entail removal

and return to earth for repair.

The tasks listed are indicative of the types of operations viewed

as feasible. When the engine is further defined, the tasks will

need to be reevaluated. The turnaround maintenance tasks are to be

fully automated so they may be performed with IVA. Because the

inspection tasks will need manned involvement, more manhours are

assigned to the tasks. These tasks are listed separately from the

OTV tasks previously discussed for ease of discussion. They would

be fully integrated with the OTV tasks as part of the ground

timeline optimization performed to arrive at the appropriate

maintenance schedule. If an engine removal were scheduled, the

inspection would be eliminated.

An experienced ground crew under ideal conditions (air-conditioned

test cell fully equipped with the necessary tools) can remove an

RL-10 in about five hours. The EVA crewmen are expected to replace

an engine in four hours in the SS hangar. This short time is a

goal to make sufficiently efficient OTV turnaround a possibility.

It will be necessary to simplify the OTV/engine interface to enable

both the engine removal itself and provide the necessary functional

integrity to the interface once the engine has been replaced. For

this reason, it is desirable to eliminate pressurant-activated

components because this eliminates a gaseous QD from the OTV/engine

interface. If the propellant tanks are left with a blanket

pressure, a set of valves will be needed on the vehicle side. The

main engine valves should remain with the engine so they can be

serviced after the engine has been removed (possibly on the SS,

likely on the ground). The simplest interface design has all QDs

aligned along a plane that separates the engine and the vehicle.

4-27
This design type would lend itself to remote engine removal, which

is a desirable feature. This approach would likely incur a weight

penalty relative to an approach that minimizes weight at the

expense of requiring EVA assistance. Cost modeling of OTV

servicing scenarios is expected to aid in recommending which

approach to use.

The functional and operational analysis just presented identified

five basic space station facilities that will be needed to support

a space-based OTV. The facilities are shown in Table 4.1-10.

While the facilities are treated as separate items dedicated

entirely to the OTV, in the actual space station they will be

general-purpose facilities designed to support the OTV, OMV and

other spacecraft designed for SS servicing. At this point, the

facilities are separated more for functional reasons than for

hardware reasons. The actual SS facilities will probably recombine

the functions into units logically arranged as part of the SS

design effort. Therefore, the following facility discussions

emphasize the needed functions divided functionally. Possible

overlaps are included in the individual discussions.

The servicing hangar will house all the necessary items used for

servicing the OTV and other spacecraft. It should be a

general-purpose facility with some dedicated items specifically for

servicing the OTV and the SS OMV because these two spacecraft will

comprise the majority of the servicing requirements. A means of

mechanically holding the various spacecraft will be needed. A

variety of umbilicals will also be needed, mostly electrical. It

may also be desirable to provide a pressurant umbilical.

Propellants and other hazardous fluids may be handled at another

facility. Power for lighting and power tools should be supplied as

well as a means of securing the astronaut, his tools, and any other

loose items necessary. One current hangar concept (Fig. 4.1-4)

involves a translation mechanism for the crewmen and a rotary

carriage for the spacecraft. This would allow the possibility of a

quasi-EMU (extravehiclular maneuvering unit) in which the EMU (or

spacesuit) shares the SS atmosphere through an umbilical carried

with the translation mechanism. In this hangar, total portability

would not be necessary because a combination of translation and

spacecraft rotation will allow access to all portions of the

spacecraft.

Because, as with the servicing hangar, many functions of the SS

computer system have already been mentioned, they will only be

summarized here. Only a small portion of the SS computers'

responsibilities will be represented by the OTV activities. The SS

computer will function primarily as a link between the OTV

computer, ground facilities, and the SS crewmen. OTV data stored

during the mission will be downlinked to the ground through the SS

computer with a portion being retained for the SS crewmen to act

upon (SS safety-related items, for instance). After ground

processing, an estimate of the OTV maintenance schedule will be

4-28
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4-30
returned to the SS. The SS computer will then factor in

maintenance tasks discovered during postmission processing of the

OTV and prepare a final maintenance schedule. The SS computer will

also handle loading of the OTV computer with mission-specific data

before the OTV mission. Part of the SS computer will also handle

control of the many automated servicing mechanisms. These will

include the SS RMS(s), refueling, and CCTV movement.

The above-mentioned functions may more logically be part of the OTV

control station. Certainly items that are entirely OTV-specific

will be functions of the OTV control station. The major item here

is OTV refueling and OTV LOS control. The SS computer will

probably just monitor safety-related items so it can respond

properly if an emergency should occur. The bulk of the OTV-related

software and systems will reside in the OTV control station

(functionally at best). The OTV control station will be the

primary man/machine link between the OTV and the SS crew. Several

OTV display and equipment controllers will be logically arranged to

enable efficient IVA control of the various phases of the OTV

mission. The OTV control station, as with the servicing hangar,

will probably share hardware with other spacecraft. That, however,

is a space station issue.

The OTV refueling area will work closely with the control station.

The primary function here is, obviously, refueling of the OTV.

However, several other propellant- and fluid-related functions will

also be accomplished here. Because the refueling area will

represent a significant portion of the SS mass, its location will

be critical to SS control. The disturbances caused by the

propellant transfer will also need to be accommodated.

The refueling area will house the cryogen tanks, an OTV mechanical

interface, and the necessary umbilicals to allow refueling of all

propellants and pressurants. An electrical umbilical is also

necessary to allow control of the OTV and the downlinking of OTV

data stored during the OTV mission. It is not envisioned that

other spacecraft will be able to employ this hardware for their

refueling, mainly because of the physical size of the OTV compared

to other spacecraft. Another refueling station will likely be

provided by the SS for these smaller spacecraft. (They are also

likely to require earth-storable propellants, not cryogens.)

Spacecraft wishing to use this facility will accommodate the OTV

and not vice versa. All the necessary control hardware will reside

here (valves, pumps, plumbing, etc) while the control software will

be housed at the OTV control station. One or more CCTVs will be

necessary if the refueling area is not visible from the control

station.

The space station will need to provide some sort of storage

facilities for both the OTV spares and the various payloads. These

facilities will at least provide mechanical holddown and minimal

power and data interfaces to sustain the vehicles in a dormant

mode. Desirable features would be thermal and meteoroid

protection. The servicing hangar could provide all of these at a

4-31
loss in utility. These are, of course, space station issues.

However, they are worth some discussion because there are several

possible modifications of the baseline timeline. For instance,

payload and OTV mating could be performed at the storage area if

the proper alignment capability existed. The payload checkout

could also be performed there, saving time and minimizing the

movement of masses about the SS thereby saving SS propellant.

As an aside, this brings up the subject of the multiple payload

interfaces necessary on the payload that it otherwise wouldn't

need. The STS interface now involves trunnion fittings and an

electrical umbilical. The OTV, on the other hand, would require

some sort of axially acting mechanical interface and an electrical

umbilical separate from that used for the shuttle. Presumably one

of these two interfaces could be used by the space station storage

facility. A tradeoff exists between requiring the payload to

supply these interfaces and scarring either the shuttle or OTV to

eliminate one of the interfaces. Because the payload is launched

only once while the STS and OTV make multiple trips, the mass

penalty may be best assigned to the payload. This is a subject for

further study.

The timelines discussed so far are for a routine mission where no

major failure has occurred that requires a delay to allow the STS

to bring up the needed spares or, worse yet, return of the OTV to

the ground for extensive servicing. Very few missions are likely

to be "routine" and may well require delays that affect the

baseline timeline. The learning curve is likely to extend through

much of the "routine" mission time frame of the early to late

1990s. A fully debugged OTV/SS system by 1994 is unrealistic and

an operational OTV by then is an ambitious goal. However, all the

mission analyses to date suggest large payoffs for the ability to

fly LEO-GEO missions on a two-week schedule. A case for an OTV

fleet is emerging.

The other response to downtime impacts is a sufficient spares

inventory at the SS to avoid the majority of the delays. Because

failures are by nature unpredictable, this implies storing many

spares that may never be needed. Unnecessary spares cost both in

launch mass for the spares and in the mass of the facilities needed

to house them. As a part of the evolving SS and OTV, a

comprehensive inventory management effort is recommended that will

simultaneously minimize the required mass at the space station and

the downtime incurred by the OTV. This would entail a

high-reliability OTV coupled with a component-by-component failure

analysis to pinpoint likely failures. In addition, grouping the

high failure items so they may be replaced as a unit(s) is

required. The OTV design must adhere to this modular philosophy to

some degree. One spare unit capable of remedying several failures

will be very valuable. A reusable space-based OTV cannot be

optimized alone; rather the OTV and its support system should be

optimized.

4-32
4.2 SERVICING OPERATIONS

OF POOR QUALSTY

4.2.1 Serviceability and Basing of the Baseline Vehicle

The conceptual design philosophy incorporated in the baseline

vehicle (Fig. 4.2-1) reflects the considerations for serviceability

based on the nature of the vehicle and its usage. This is

evidenced by the central core structure, removable tanks, removable

aerobrake, easily accessible avionics and propulsion subsystem

modules.

Optimum vehicle serviceability also implies consideration of the

space station servicing facilities and the operations required in

the servicing scenario. Figure 4.2-2 depicts a space station

servicing bay dedicated to accommodating the OTV. The orbiter-like

doors open for initial berthing of the OTV following a mission.

Both routine and contingency turnaround operations, including

aerobrake removal for ease of vehicle inspection and servicing,

will be conducted here. The doors may also be opened for radial

main propelIant tank removal/replacement. However, they can remain

closed for most remaining turnaround operations to provide the

needed meteoroid protection.

Figure 4.2-1 Baseline (Two-Engine) OTV Layout

4-33
Test and Checkout Equipment,

Data Processing and Communication

Control, and Servicing Area Control

/- Umbilical Connections

and Docking Interface

Mobile Work Station

with Data Display and

Controls for Carousel and

Manipulator Arm Operations

Tool and

Module Storage

Facility

Orbiter-Type

Servicing Bay

Doors

•Servicing Area

Strongback

Stowed

Aerobrake

Servicing Facility Dedicated Space

Crane Positions and Docks OTV

with Carousel. Also Maneuvers

Astronaut on MFR Type Device

and Aids in Tank and Aerobrake

Removal and Replacement

Figure 4.2-2 Conceptual Space Station OTV Servicing Facility

The space crane will be used for:

1) Docking the OTV with the rotisserie carousel;

2) Removal and replacement of tankage and aerobrake (tankage

replacement on a contingency basis);

3) Maneuvering an inspection and servicing platform for astronaut

use.

Additional intelligent pairs of manipulators for coordinated

anthropomorphic operation may be attached to the mobile work

platform(s) and translated to areas/subsystems of the vehicle that

lend themselves to such automated servicing.

4.2.2 General Servicing Considerations

Servicing will be key to successful space basing of an OTV. Though

traditionally defined as "fuel it and fix it," servicing involves

much more than this definition suggests, which the succeeding

discussion clearly demonstrates.

4-34
Technology development and utilization will, in turn, be the keys

to successful onorbit servicing. The existing technology is not

sufficient in some areas to allow the conduct of servicing

operations. New technologies will, however, emerge and, if

carefully applied, will support the evolution of satellite

servicing capabilities.

Goals and Guidelines for Future Servicing - The goals for OTV

servicing in the space station IOC time frame include:

1) Minimization of total system life-cycle cost;

2) Utilization of emerging technology (the state of art is a

function of time frame);

3) Development and utilization of systematic technology

projections that affect hardware evolution so evolutionary path

avoids deadends;

4) Evolution of the servicing mode toward increasingly automated

operations.

Other goals will be treated within the context of the detailed

discussions that follow. Our reference servicing scenario reflects

consideration of all of the above-mentioned goals.

Reference Servicing Scenario Development - Space-based OTV

servicing is expected to encompass a wide variety of preflight,

postflight, and eventually, inflight servicing operations. These

operations have been widely discussed in relevant literature from a

variety of viewpoints. Most of the operational descriptions have

been rather superficial. Some of the available analyses, however,

have attempted to go to a greater level of depth and identify

specific service needs and functions for both propulsion vehicles

and payloads.

The Teleoperator Human Factors Study (THFS) recently completed for

NASA-MSFC, took a different approach and made the assumptions

necessary to go to great detail. THFS identified several levels of

operational and servicing functions and, through iterative

analyses, compiled preliminary lists of low-level (task and

subtask) functions we seek to describe as generic, "primitive"

tasks. These primitive tasks are the basis for all higher level

tasks. The levels of operational and servicing functions

identified in THFS are sequence, activity, task, and subtask.

These terms are defined in Table 4.2-1.

4-35
Table 4.2-1 Levels of Decomposition

Term

Sequence

Activity

Task

Subtask

Definition

A group of goal-directed related activities comprising a major

portion of a mission.

A group of goal-directed related tasks that fulfill a limited purpose

within a sequence.

Basic unit of behavior; smallest logically definable set of

perceptions, decisions and responses to complete a task.

Each perception, decision, or response that is a task component.

Example

Deploy Spacecraft

Proximity Operations

Adjust Orientation

(Perception of Data, Decision,

Control Activation, Feedback

of Response Adequacy)

Initiate Adjustment

The functions at each level were identified by progressive

decomposition of scenarios for representative programs selected

from the NASA-OAST space systems technology model on the basis of

their widely varying natures.

The THFS study and succeeding work will ultimately lead to the

identification of a finite set of specific "primitive actions"

that, when assembled into complete service operations, can be

performed entirely by man, entirely by automation, or by some

combination of the two. How the operations will be performed is a

function of technology availability; the possibilities range from

EVA to IVA, from purely manual to hard automation, teleoperation

and flexible automation. Timeline knowledge and knowledge of

probable equipment capabilities and configurations, when combined

with the knowledge of the composition of the functions, will permit

a bottoms-up process of determining the work organization for each

servicing function and the overall servicing process.

Figure 4.2-3 presents an overview of this directly applicable

hierarchical work organization development approach in an OTV

servicing context.

Program

t

Project

.1

Servicing Functions Mission Work Organization Process

Activit

1

y

Task

1

Subtask

/

Personnel | Automation

1

» juu upuim * Work Cull

Layout

n

utifvi .« Equipment Task

Capabilities Organization

K

Transition \

V

rl

1 Subtask

or j[

Task Element

Figure 4.2-3 Work Organization Development for OTV Servicing

4-36
The lower level functions, those that go through the transition in

Figure 4.2-3, fall into three categories—manipulation, mobility,

and housekeeping. These primitive actions are, as stated

previously, the building blocks assembled into successively higher

level groups—subtasks, tasks, activities, and sequences.

Typical activities are:

1) Capture;

2) Berthing;

3) Leak check;

Typical sequences include:

1) Retrieve;

2) Detank;

3) Inspect/examine;

4) Check out;

5) Conduct preventive maintenance;

6) Prepare for deorbit;

4) Module R/R;

5) Umbilical connections.

7) Repair;

8) Resupply (fluids);

9) Reconfigure;

10) Store;

11) Mate payload;

12) Deploy.

The illustrative top-level scenarios combining these sequences for

specific operations are shown in Figure 4.2-4. These illustrative

scenarios have been integrated with other scenarios foreseen into a

comprehensive top-level turnaround flow for servicing a nominal

space-based OTV. Figure 4.2-5 depicts this flow.

4.2.3 Service Operations

The service operations depicted in Figure 4.2-5 are analyzed at

some length in the following paragraphs. These operations are

decomposed to the level necessary to identify relevant requirements

and issues. The requirements, issues and other associated

considerations identified in the decomposition process are then

organized and discussed separately for each of the operations.

Succeeding sections of this report will discuss the implications of

these requirements, issues, and considerations for OTV and space

station design. Although these are high-level discussions

incomplete at this early stage of OTV design, they represent a

first step toward organizing the information that will ultimately

result in the identification and justification of systemwide and

vehicle-specific design requirements.

4-37
Deployment

I OTV I

Preparation P

I 1

OF POOR QUALITY

Conduct

Proximity

Operations

Launch

' Perform |

"

Contingency Repair Operations

I 1

I Checkout f~

I 1

Inspection Analysis

Retrieval

' OTV I

I Preparation I

I I

Figure 4.2-4 Representative Servicing Scenarios

OTV Proximity Operations - This process, as nominally conceived,

encompasses maneuvering the OTV from its rendezvous and shutdown

position to the location where the RMS (or other grappling device

such as a "space crane" variation of the RMS) will assume control

of OTV movement. It could be preceded or interrupted by the

retraction of deployable components. The flow of the activities in

this sequence is depicted in Figure 4.2-6. Preceding and

succeeding sequences are also shown in this figure for perspective.

4-38
ORIGINAL PAGE !S

Of POOR QUALITY

Mission i

Operations |

L_ j

• Perform Mission

• Orbit Transfer

• Rendezvous

• Shutdown &

Sating of

Unneeded

Systems

1

Inspection

Operations

• Analysis of

Flight Data

• Inspection

• Checkout

1

Checkout __

Operations """

O Necessary

Verification

Only

Proximity

Operations

O Alternatives:

- ACS

- OMV

— Tether

Maintenance

Operations

• Module R&R

• General

Service

_ -oad _

** 2onsummables

Prepare for

Handling

Onboard

Operations

• Retract/Remove • Grappling

Appendages • Positioning

• Destack • Berthing

• Contingency

Preparations

Onboard

Operations

Prepare for

~ Launch ~

• Final Checkout

• Deploy

Appendages

- Storo OTV •—

Proximity

Operations

O Move to

Launch

Point

Detanking

Operations

• Operate Leak

Detectors

• Connect

Umbilicals

• Drain Tanks

• Vent/Purge

Tanks & Lines

• Umbilical

Disconnect

Checkout

Operations "

O Necessary

Verification

Only

_ Launch

Onboard

Operations

(Prepare 8? Prepare 8J

'Deliver 1 j Deliver {

lOMV J iPayload j

"\

Assembl

 7e

Stack

• Join OMV

to OTV

• Join Payload

(or Servicer

Kit) to OTV

j Mission

"i Operations

L. J

Legend:

• = Activities within Sequence

O = Comment/Discussion

Figure 4.2-5 Top-Level Turnaround Flow for Space-Based OTV

Options for transition of an OTV from free flight to space station

control include using its onboard attitude control system (ACS) to

maneuver and using an OMV to go out and retrieve the OTV. As can

be quickly deduced from Figure 4.2-6, use of the OTV's ACS would be

the simpler total system. This is the baseline recommendation.

However, if an OMV is to be available for this purpose between

commitments to support other activities, it could probably be used

to pick up the OTV after shutdown and transfer it to its berth at

the detanking facility. Alternatively, the OMV or some other means

such as an EVA with the MMU could be used to ferry a cable (tether)

out to the OTV so it could be reeled in.

Use of OTV Attitude Control System - This option, as stated

earlier, involves the simplest system in terms of equipment,

interfaces, and procedures. It would require the OTV design to

provide the capability to operate ACS thrusters remotely after all

other systems have been shut down. Implementation of this scheme

could also require development, installation and support of a

remote control station in the space station if tradeoffs indicated

local control to be preferable to automated operations and/or

ground-based control. But such an integrated teleoperator/

4-39
supervisory control workstation would be justified by a number of

other applications such as OMV control from the orbiter or the

space station and OMV- or OTV-mounted servicer-kit operation—

although the latter application would considerably increase the

complexity and weight of the workstation. For this reason and

others, servicer-kit operations may best be left to ground-based

controllers where these operations can be satisfactorily conducted

if communication link delays can be overcome.

' Mission Operations

I r

I Perform

I Mission

I

J , J Return to Safe Main '

I Perform I Space Station Main Engine(s) __ Propulsion & _1

I Mission I Orbit & Shutdown other | I

I J Rendezvous Unneeded Systems | I

Proximity Operations

Use ACS to

Maneuver to

Enable RMS

Grapple

Safe

Remaining

OTV Systems

I

Fly!

i

•ether

TV

OMV Dock

with OTV

Dock Tether

to OTV

OMV/EVA

Release Tether

& Depart

OMV Transfer

OTV to Enahla

RMS Grapple

Position OTV

for Station

Handling

Preparations

I

•Prepare & ' u_±_ I Launch OMV |

Prepare & j J Dock &

1 Launch

OMV

• Deservice

I

• r ~»

I | Preparation'

—I for Station I

I Handling I

( or EVA | i OMV j ! OMV or EVAj

I II , ^uipment _,

Figure 4.2-6 OTV Proximity Operations

r~Docki_

I Deservice .

| OMV

\ 1

A number of relevant issues are involved in this option. The first

is the issue of control location—the determination of the space

station's role in the proximity operations process. Though many

argue that it is intuitive that local control would be preferable,

a formal tradeoff may introduce less apparent concerns that, in the

end, argue strongly for ground-based control in which the OTV would

be docked with a passive space station or for total automation of

proximity maneuvering.

4-40
The next issue with a direct impact on design of both the OTV and

the space station is control mode. Relevant questions regarding

this issue include: Will automated control be feasible? Desirable?

Supervisory control? Teleoperator control? What will be the

hardware, software, processing time, and other impacts of the mode

chosen? A discussion of these points is beyond the scope of this

study but they are being treated in a number of ongoing efforts.

Other points relating to this approach must also be considered.

The first is the contamination implication of OTV ACS and other

system propulsion products, venting and leakage—contributing to a

"brown cloud." The resultant system design requirements may affect

the development of OTV and space station hardware as well as

operational procedures. However, it should be noted that many

experts feel the oxygen/hydrogen system selected for the vehicle

can be used in the vicinity of the space station with no

significant contamination risk.

Another concern is plume impingement. Handling procedures and

space station surfaces will have to be designed to prevent damage

from ACS plumes if thrusters are used for close-in mobility. Some

minor damage to MLI surfaces was noted from MMU plume impingement

during the Solar Max rescue mission.

Yet another is illumination management. Obtaining proper levels of

lighting during the rendezvous and docking maneuvers may

necessitate some stationkeeping activity once proper illumination

levels are determined.

Lastly, the additional thruster use for these operations will

affect the OTV's ACS life and servicing requirements and the

significance of this effect must be determined. It could be that

formal tradeoffs will result in clear arguments for or against this

option based on such things as reliability impact, servicing

timeline optimization, and/or comparative ease of servicing one

vehicle or the other.

Use of OMV to Position OTV - Many issues and considerations of this

option are, as one should expect, identical to those involving use

of the OTV ACS for proximity maneuvering. Should control be

automated? Teleoperator controlled? From where: the ground? The

space station? It is assumed the OMV would be controlled from the

same workstation as the OTV. Will OMV outgassing present a

problem? Will plume impingement?

The impact of incremental use is also relevant, but in a broader

context than for the previous option. Not only are vehicle life

consumed and servicing requirements increased, but scheduling

conflicts may be created by the multiple demands on OMV time. Also

to be considered are the safety and control issues associated with

flying two vehicles simultaneously. While risks may be only

nominal, they will exist and will certainly be greater than those

associated with managing a single close-in free-flyer.

4-41
One possible advantage of using OMV in this application will result

from its greater maneuverability. This assumes it will be designed

for close-in operations and the OTV will be designed principally as

a bulk carrier. The factors that will determine whether this

increased maneuverability is really advantageous include whether it

is necessary for safety reasons or for positioning before RMS

capture of the OTV.

Design requirements that result from this proximity operations

concept include:

1) Interface (docking) provisions for joining OMV to OTV;

2) Possible simplification of the OTV ACS system because it will

not be employed in the vicinity of the space station.

Use of OMV (or MMU) to Attach Tether to OTV - Tether retrieval

would partially overcome the "brown cloud ' and plume impingement

concerns inherent in close-in thruster operation. However, it

would still require outward-bound propulsion and maneuvering as

well as the necessary support systems. Because it probably could

not be justified solely on the basis of the emission reductions

that would result, it is envisioned principally as a candidate for

backup mode operations. If, however, it is possible to somehow

eliminate the OMV/MMU role in tether retrieval, this option would

become clearly superior in terms of the criteria previously noted.

One advantage that could accrue from any tether approach employed

would be the elimination of RMS handling of an incoming vehicle.

If the tether system was based at the first service berth, it could

bring the vehicle directly to its mooring. Of course, such issues

as tether maneuvering (control) and deceleration must be

satisfactorily resolved before such a system becomes feasible.

An associated benefit is simplification of service bay design.

Elimination of the requirement to roll back a portion of the

service bay (cylinder) wall for RMS access may become feasible if

the vehicle to be serviced could be brought in through the open end

of the cylinder by a tether terminating inside.

Station design requirements resulting directly from this retrieval

approach include:

1) Addition of the tether system;

2) Accommodation of the tether within the service bay;

3) Integration of the tether system into on-station mobility

operations.

No impact on vehicle design is foreseen. The vehicle/tether

interface envisioned would employ the docking facilities and

operations discussed in the following paragraphs.

4-42
Preparation for Station Handling Operations - Though not really a

part of the retrieval process, these preparation sequences will be

briefly considered at this point because they interrupt the

retrieval process here—where they appear in the nominal process

flow. The first activity that will be considered is retraction of

deployable components. This discussion will be followed by

consideration of the destacking operations aspects and what could

be called contingency preparations. As illustrated in Figure

4.2-7, any or all of these activities could occur.

Proximity

Operations

.

fan

VIL

,

H/\

Retract/

Remove

Deployable

Components

Disassemble

Stack

Conduct

Preparations

1

1

1

1

1

1

r

1 i

1 I

I

i

i

i

•

l

J

'Onboard |

Mobility •

Operations '

Figure 4.2-7 Preparation Operations

These processes, depending on situation-specific variables, could

occur either before or after control over OTV motion is exerted by

the RMS.

Retraction or Removal of Deployable Components - This sequence

would be necessitated by conflicts between OTV appendages (such as

antennas, engine nozzles, and/or the aerobrake) and the clearance

at the service berths/bays. The number and complexity of the

activities involved would, of course, be a function of system

design. With proper forethought, any appendages that could not be

eliminated by design could be designed for automated retraction or

simplified removal operations. Furthermore, any problems in the

retraction process could (again assuming accommodating design) be

satisfied by a teleoperator or by an EVA sortie as noted in the

contingency preparations discussion.

Destacking Operations - This sequence would be necessitated if a

payload is too large to fit within the servicing bay or if safety

considerations or physical configuration require separation before

detanking. It could become the preferred method of operation,

rather than a contingency method, if vehicle/payload control issues

can be satisfactorily resolved.

4-43
If it is necessary to remove a payload or decouple the OMV/OTV pair

in free space rather than within the confines of the servicing bay,

these operations would best be treated as a separate sequence.

Whether the operations involve EVA or whether they can be

accomplished autonomously or with the intervention/participation of

a manipulator, their characteristics and associated issues will

differ in a uniform manner from operations inside the service bay.

For example, because the OTV (and payload) will probably not have

been detanked before these operations, the safety and operational

concerns associated with operating and working on fueled vehicles

will apply. Control during and after demating operations will be

another shared concern. Will the vehicle(s) be free-flying?

Tethered? Held by grappling arms?

Contingency Preparations - This sequence includes activities that

would be performed by a servicer (EVA or teleoperator) to correct

anomalies preventing routine station handling and processing.

Examples include:

1) Failure of deployables to automatically retract thus violating

envelope constraints;

2) Docking adapter malfunction;

3) Safety-related failures;

4) Attachment of grappling fixtures or other adapter(s) to the

payload returned by the OTV.

The issues, options, and alternatives associated with these and

other contingency operations would, of course, vary. But certainly

the design of hardware and procedures to accommodate multimode

operations (EVA, teleoperation, automation) will be a primary

concern.

Onboard Mobility Operations - Grappling the OTV is, in the truest

sense, a transitional element in the servicing process flow. It

both ends free-flight operations and begins onboard mobility

operations. This activity and the positioning activities are

presently envisioned as the components of the initial onboard

mobility sequence. Figure 4.2-8 depicts this sequence.

Grappling Operations - Grappling, with a minor variation (OTV

docked at the start instead of in flight), and positioning and

berthing will be repeated throughout the servicing process as the

OTV is moved from bay to bay and operation to operation around the

station. Table 4.2-2 lists the moves that could be involved; each

move represents a repetition of the sequence depicted in Figure

4.2-8.

4-44
Onboard Mobility Operations

•Proximity

I Operations

I I

Grappling Operations

II

T

Ii ^

Positioning Operations

—i1

ii

Tether Attached

• to OTV by OMV

or Alternative

Means

>rj

RMS Grapple

without Stack

— or\

Position OTV

& Stack for

Preparation for

Station Handling

1

1

1

i

(""Actuation & , • Preparation for ;

J Checkout of ' J Station Handling

j RMS J jj)perations

1 1

"v

y

Position £

Mount OTV

System

i

Position & 1 f" ~ ~ ~ T

Berth OTV IjDetanking 1

" at Fuel TjOperationsI

Depot 'i J

Sec

1

1

I

I

1

urn 1

RMS 1

! i i

Figure 4.2-8 Onboard Mobility Operations

Table 4.2-2 Possible Onboard Moves

From

Capture Point

Detanking Berth

Maintenance Bay

Storage Bay

Maintenance Bay

Stack Assembly Area

Tanking Berth

To

Detanking Berth

Maintenance Bay (for Postflight Maintenance)

Storage Bay

Maintenance Bay (for Preflight Maintenance)

Stack Assembly Area

Tanking Berth

Launch Point

All grappling activities associated with these moves, except

possibly those involving initial capture and those occurring before

movement to the launch point, could conceivably disappear if an

alternative onboard mobility system such as a rail or cable system

is developed. The task flow for this grappling activity would

remain unchanged whichever approach is chosen. This flow is shown

in Table 4.2-3, which also indicates the nature of the tasks

involved.

Grappling will drive a number of requirements in both OTV design

and station equipment design. OTV design will be affected in three

areas. The first, relatively minor, is the requirement to provide

grappling fixtures. Alternative approaches to satisfying this

requirement include the schemes illustrated in Figure 4.2-9. The

4-45
first involves using target fixtures located about the periphery of

the vehicle and the second approach employs a probe at the end of a

manipulator arm to attach to one of the vehicle's docking

fixtures. It is assumed the OTV will have a docking fixture at

each end to facilitate stacking operations, OMV docking, and

maintenance. The RMS standard grapple fixture described (Fig.

4.2-10) is an example of a concept that could be employed.

Table 4.2-3 Grappling Activity Task Flow

Task

1

2

3

4

5

6

7

8

Identify Target

Locate Target

Select Path to Target Vicinity

Move to Target Vicinity

Position (Rough) Manipulator(s) to Grasp OTV

Place (Final) Manipulator(s) to Grasp OTV

Grasp OTV

Verify Grasp

Nature of Task*

C

P

C/P

M

M

M

M

P

*C = Cognitive; P = Perceptual; M = Motor

'Stack Attachment Fitting Set

Docking/Grappling

Fixtures

Stack Attachment Fining Set'

Docking/Grappling Fixture

Figure 4.2-9 OTV Grappling Fixtures

4-46
Range and

Target

 Ro" Lines

(for Wrist TV)

Base Plate

Grapple

Shaft

Guide

Ramp

Electrical Connector

(Optional)

Figure 4.2-10

The RMS Standard Grappling Fixture

The standard grapple fixture consists of a rigid shaft, three

alignment cam arms, and a target fixture (Fig. 4.2-10). The rigid

shaft that is grappled by the special end-effector (SEE) provides

the structural support between the payload and the RMS. The

grapple target fixture is sighted by the RMS wrist camera and is

used to align the SEE with the grapple fixture before capture.

When the grapple fixture is within the capture envelope, the snares

of the SEE are closed about the rigid shaft and are withdrawn to

the end of the end-effector until a firm connection is made. The

grapple fixture cams are fitted into corresponding slots in the SEE

to rigidize the payload during manipulation.

Specifications:

Maximum weight: 22 Ib

Torsional moment about longitudinal axis of SEE: 450 Ib-ft

Bending moment to SEE: 1,200 Ib-ft

Shear force associated with bending moment: 50 Ib

Maximum payload weight: 32,000 Ib

The second OTV design requirement resulting from grappling

operations accommodation involves stabilization of the vehicle.

Methods of stabilization that minimize rotation must be developed

to support grabbing various peripherally located fixtures if this

capture approach is chosen/required. The necessary fixtures would

4-47
already exist because they will be necessary to satisfy other

operational requirements such as:

1) Stack stabilization (postdocking);

2) Remote (inflight) servicing;

3) Contingency grappling.

If the probe and docking fixture approach is chosen, the vehicle

could either be "normally" stabilized or spin-stabilized and would

simply require orientation to facilitate the approach of the

arm-mounted probe. Spin stabilization would be possible because

the docking fixtures are located at each terminus of the vehicle's

longitudinal rotational axis. This approach provides a

straightforward contingency capability that seems to contain few

risks and that would appear to be easy to perform either with a

remote manipulator or with total automation.

The third grappling-imposed requirement on the OTV would be to

accommodate the structural loads induced in the capture process.

The impact of this requirement can probably be minimized by proper

design and procedure development.

The station equipment affected by grappling all relates to the

manipulator arm. Specifically, its structure must be designed for

the loads involved and its controls must provide the response,

accuracy and flexibility necessary to (1) avoid damage to itself

and the OTV, and (2) satisfy established timelines. The controls

must meet these requirements in autonomous and teleoperator

operational modes and for normal and contingency situations.

The key issues involved in control of the grappling process include

speed and accuracy of response to control commands and design of

the feedback link for teleoperator situations where the human is in

the loop. The first two are a function of the evolving state of

the art in manipulator design and the situationally imposed

requirements (speed and accuracy required). The feedback issue is,

simply stated, determination of the video requirements—line of

sight, resolution, and 2-D or 3-D—necessary to accommodate

consistently satisfactory grappling performance. Future

investigations will have to resolve these issues.

Positioning and Mounting/Berthing Operations - As for grappling

operations, a fairly set task flow would be involved if an RMS-like

system is used for positioning (and mounting or berthing) the OTV

stack (Table 4.2-4).

4-48
Table 4.2-4 Positioning (and Mounting/Berthing Task Flow

Task

1

2

3

4

5

6

7

8

9

10

11

Clear Locale (Where Grappled)

Identify Destination

Locate Destination

Select Path to Destination

Move (Transfer) OTV to Destination's Vicinity

Position (Rough) OTV

Place (Final) OTV

Wait for (Assume Autonomous) Activation of Locks

Release Grasp

Verify Released

Clear Locale (Where Released)

Nature of Task*

C/P/M

C

P

C/P

M

M

P/M

_

M

P

C/P/M

*C = Cognitive, P = Perceptual, M = Motor.

Tasks 8 through 11 would be eliminated for the "position for

destack operations" sequence because it is assumed, for the purpose

of the present task flow, that destacking would occur with the OTV

"hanging" on the end of the manipulator arm. The remaining tasks

would be unchanged. Careful consideration of this sequence may

drive a requirement for more than one RMS. On the other hand, if

the tether attachment procedure is employed, the Table 4.2-4 steps

would be significantly changed—and hopefully simplified. Tether

attachment steps would represent the first tasks in rail or cable

mobility system operations. Under this scheme of operation some of

the tasks identified in Table 4.2-4 would no longer be necessary

and the nature of many of the tasks remaining would be changed.

Task 4 in the table, selecting a path, would no longer be

necessary; the cable or rail system would be the path from origin

to destination. Tasks 1 and 11 would become purely academic

because the set path would presumably not contain obstacles. The

transition from Task 6 to Task 7 could be simplified to nothing

more than a speed adjustment. And, finally, Tasks 8, 9 and 10

could be deleted because no manipulators would be involved.

The remaining tasks (2, 3, 5, 6 and 7) would be simplified enough

to permit automation using existing technologies and techniques. A

human role would, however, still be necessary until artificial

intelligence capabilities advance sufficiently to permit automated

task and path planning. This role would, at a minimum, involve

selecting and entering through a keyboard the appropriate code to

tell the automated system the destination of the vehicle.

The issue of automation, passed over lightly in the treatment of

grappling operations, bears further explanation at this time. The

three categories of tasks (cognitive, perceptual and motor)

identified in Tables 4.2-3 and 4.2-4 will be used to structure the

discussion that follows.

All the motor tasks readily lend themselves to full automation.

The only reservations that must be mentioned are the possible

impacts of control response speed and accuracy limitations imposed

by the state of the art in relevant technologies.

4-49
Perceptual tasks present an interesting mixture. Some, such as

"verify released," can be performed by electromechanical devices.

Others, such as "locate destination," require either some sort of

built-in and automatically updated position reference map or some

ability to interpret visual information. Both would require an

extensive computational capability to automate. Further, image

understanding is only an emerging discipline and will probably see

only limited use over the near term. An alternative technology

such as radar may, however, be able to perform many perceptual

tasks.

Cognitive tasks are the most complex of the three types. They

involve knowing and/or understanding, and often imply choices among

alternatives in varying, unstructured situations. Such descriptors

as "identify" and "select" are typical of cognitive tasks. There

is some overlap between the cognitive and perceptual

categories—"identify" typifies this group.

Some degree of intelligence—the ability to cope with a new

situation, to learn or understand—is required to perform the

cognitive tasks. An advanced understanding of the reasoning

process and extensive computational power will be required before

these tasks can be automated. Again, as with image understanding,

the necessary capabilities (often referred to as artificial or

machine intelligence) are beginning to emerge and may see limited

application over the near term.

In summary of this brief discussion, it can be said that for highly

structured tasks (e.g., "normal" docking), full automation is a

reasonable short-term goal. Indeed, it may be preferable to human

involvement. However, for less structured tasks (e.g.,

"contingency" docking or unplanned remove and replace operations)

where more complex perceptual and cognitive capabilities are

required, man-in-the-loop operation, either supervisory or

teleoperator, will be required over at least the short- and

intermediate-term goals.

Other considerations relevant to positioning and berthing

operations include positioning accuracy requirements and the impact

they will have on manipulator and berth design, vision requirements

(camera location) and accommodation of any operational

contingencies arising from equipment malfunction. The first two

subjects were adequately discussed under grappling operations. The

last implies accommodation of multimode operations by designing the

mobility system (whether a manipulator or a cable/rail system) to

readily permit a backup system such as another arm, another tether

(possibly with spread bar or sling), or an EVA astronaut to assume

control of the OTV and complete the desired transfer, positioning

or berthing process.

4-50
Detanking Operations - The detanking operations presumed necessary

for technical, economic, and probably safety reasons will be the

first servicing sequence performed aboard the station. Figure

4.2-11 shows the activities in the process. The issue of whether

multiple fluids would be transferred simultaneously or sequentially

is beyond the scope of this study; it should be addressed in future

fluid transfer studies.

j Detanking Operations

r- - -i ' 1 Onboard | | x""v

1 Mobility { i (*TM*f

1 Operations J J ^p

1 1 1

1

Connect

TM UmbilicaUs) -

1

i

Drain

Tanks

Vent&

' Tanks 8

Operate Leak

Detectors

Purge

i Lines

•M

Umbilical Is)

Disconnect

Onboard

Operations

I

Figure 4.2-11 Detanking Operations

Several fluids could potentially be involved in these detanking

operations. However, for this particular OTV design the number

will be minimized because the propellant and oxidizer will be used

for electric power reactants and for the ACS, the propellant will

serve as the pneumatic system fluid, tank pressurization will be

autogeneous, and a space station-provided means will be used for

purging fluid lines. This approach will permit operations

employing only two fluids—liquid hydrogen and liquid oxygen.

Umbilical Attachment - The first of these activities, umbilical

attachment, would preferably be a fully automated activity either

occurring at the time of docking or initiated on completion of

docking. (Though we believe the present space station baseline

envisions manual make/break, we foresee a possible change in this

position to take advantage of available techniques and

capabilities.) This choice will affect both space station and OTV

design.

Simultaneous umbilical connection could eliminate some mechanisms

from the berthing platform for positioning and inserting the

umbilicals and some tasks from the operation. However, it would

require precise positioning of the OTV and require the vehicle's

propulsion module to employ QD-type connectors and associated

plumbing specifically designed for onorbit refueling.

If umbilical attachment is to be initiated after completion of the

docking (berthing) activity, a number of additional cognitive,

perceptual, and motor tasks would be introduced. The aspects of

each type of task were previously discussed. But the specific set

of tasks involved cannot be identified until design iterations

determine such details as whether access panels will have to be

opened, whether connector covers will have to be removed and

stored, and so forth.

4-51
Table 4.2-5 presents a nominal task set. It ignores questions like

those above and focuses on the narrow issue of deploying the

umbilical. It is a "worst-case" task set and is described in a

manner intended to show adequate detail to illustrate that the

individual tasks (and hence the set) could be performed by hard

automation, by a sophisticated remotely controlled manipulator or

by an EVA individual. For hard automation, Tasks 1 through 3 would

be predetermined so they would be dropped from the flow.

Table 4.2-5 Umbilical Attachment Activity Task Flow

Task

1

2

3

4

5

6

7

8

9

Identify Target (Connection Point)

Locate Target (Connection Point)

Select Path to Destination

Free Restraints

Extend Umbilical (Move to Target Vicinity)

Position (Rough) Umbilical

Place (Final) Umbilical (Connect)

Lock (Turn?)

Verify Locked

Nature of Task*

C

P

C/P

M

M

M

M/P

M

P

*C = Cognitive, P = Perceptual, M = Manipulative.

The addition of this activity requires the station to provide and

accommodate the manipulator(s) capable of performing the included

tasks. Selection of the hard automation alternative to accomplish

the tasks would impose weight and maintenance penalties over the

flexible automation alternative because hard automation is, by

definition, specifically oriented toward a single purpose—it could

not be used for anything else. Selection of the flexible

automation alternative, in the form of a relocatable, multipurpose,

remotely controlled manipulator, would overcome these penalties but

impose penalties of its own in terms of operational complexity when

compared to simultaneously occurring docking and umbilical

attachment.

Whichever attachment method is chosen, some decisions will have to

be made about where the attachment(s) will be located. That is,

will plumbing within the OTV be used to group all connectors so a

single umbilical can simultaneously connect to all of them or will

multiple umbilicals be required to minimize OTV plumbing line runs

to satisfy weight, maintenance, safety, or technical requirements?

Leak Detection - Activated before or at the time of umbilical

attachment, leak detectors would be operated throughout the

detanking operations sequence. Presumably they would be tied into

caution and warning panels in the station and/or service bay. They

could also possibly activate automatic shutdown/recovery sequences

(contingency operations) if their reliability were high enough.

4-52
Tank Drainage - For the propellants this activity should involve no

more than (1) activating (opening) valves to allow the autogeneous

pressure to displace them, and (2) closing the valves when the

process is complete. Contingency situations, like inoperative

valves, would require the OTV either to provide a way to

mechanically actuate the valve or to provide a second access route

(valve) into each tank to permit depressurization/drainage of the

tank.

Inoperative (open) valves would also require some contingency

provision for closing the tank before umbilical disconnect to

prevent undesired venting. The same solutions, a second valve or a

local mechanical means of actuation, would apply. "Design for use"

considerations such as location (access), clear identification, and

force application requirements must be accommodated in designing

and locating these valves. They will have to accommodate all

possible contingency-mode operations as well as any changes caused

by evolutionary growth.

Thermal management of the tankage and lines throughout this process

will also be a concern of note. Because blockage of a line by

freezing during the process could prove to be a critical hazard, it

must be carefully considered in the design of lines and defueling

operations.

Venting and Purging - Clearing residuals from tankage and lines

after they are drained will involve two sets of considerations.

The first focuses on ensuring the residuals will not be hazardous

or contribute to the "brown cloud." Venting directly to space will

not be allowed to occur in the vicinity of the space station.

Rather, purged residuals will be processed for reuse through a

collection system. Whether this policy becomes flexible over time

will be a complex function of the seriousness of the contamination

problem in the station's environs and of the delivered "cost" of

the fluids as well as the weight, complexity, power consumption,

etc of a collection system.

The second set of considerations focuses on how the purging will be

accomplished. Two distinct phases or types of activity are

projected. The first is gas displacement (using autogeneous

pressure and, maybe, nitrogen) and the second is vacuum drying.

Combined, the two will ensure no residual fluid remains to

interfere with the refueling or remove and replace maintenance

operations that may become necessary. Both OTV and space station

systems will have to be designed to accommodate these processes.

The tasks associated with erecting and retracting the venting masts

intended to move the vented gases away from the facility will be

included in this venting and purging activity unless it proves

preferable and feasible to permanently erect the masts. The only

other tasks involved will be opening and closing valves

(essentially switching functions) that will be easy to command from

a remote site when in a normal operations mode.

4-53
Umbilical Disconnect - Once connecting lines are purged, the

umbilical(s) will be detached and stowed to complete the detanking

operations sequence. Table 4.2-6 shows the tasks that would be

involved in this activity.

Table 4.2-6

Umbilical Detachment Activity Task Flow

Task

Ref

1

2

3

4

5

6

Vent & Purge Operations

Unlock (Turn?)

Disconnect (Turn?) Umbilical

Retract Umbilical

Position (Rough) Umbilical

Place (Final) Umbilical

Secure Umbilical

Nature of Task*

C/P/M

M

M

M

M

P/M

M

*C = Cognitive, P = Perceptual, M = Manipulative.

Note that these tasks are all primarily manipulative in nature and

none are very complex. This tends to make them amenable to hard

automation. Any associated perceptual tasks could easily be

accomplished through the use of existing technology, e.g., sensors

and switches.

Of course, proper system design would have to allow for manual

(EVA) and/or teleoperator task conduct in contingency situations.

Issues and Considerations - A key issue was avoided in the

preceding discussion by assuming the OTV tanks will require

evacuation. However, a technology goal is to develop systems and

procedures that will eliminate this requirement. Two alternative

approaches come readily to mind:

1) Refilling partially full tanks without having to first empty

them;

2) Removing the "expended" tanks and replacing them with full

tanks.

The implications of such servicing alternatives in terms of OTV

design, space station design, operational hazards, timeline

reduction, etc, will require future investigation.

This discussion also avoids another key issue by not attempting to

define where the detanking facility should be located. It would be

convenient if detanking could be conducted at the same berth as

inspection and checkout (I&CO) and remove and replace (R/R)

maintenance functions. However, it is assumed that safety issues

and pollution considerations will drive this decision in spite of

timeline considerations. For example, safety requirements will

impose strict controls to prevent release of propellant and/or

oxidant near the orbiter; crew safety and potential contamination

of sensitive components are specific concerns. The treatment in

4-54
the following paragraphs applies whether the facility is remote

(tethered or a separate formation flyer) or a part of the main

space station structure (a separate bay or a hookup in the main

service bay).

A number of components, systems and technologies will be critical

to these detanking operation processes. Components that heavily

affect design, operation and maintenance include valves, filters,

couplings, and connectors. Key systems include drainage

interfaces, mass gaging, and component replacement systems. The

technologies identified as critical include mass gaging, commodity

transfer, slosh control and venting.

Interface management is the next area of concern. As has been

alluded to in previous discussions, any given interface must not

only accommodate automation, it must also be both EVA- and

manipulator-compatible.

Fluid management methods must be developed and definitized for the

scenario. Although chilldown procedures/techniques, particularly

for lines, are safety-critical, they have not been well defined.

Neither are the methods for measuring fluid levels in the tanks or

for handling residuals. Present technology mass and flow gaging

techniques are not yet very precise—_+10% is typical. It is

presumed that a two-step (gas displacement and vacuum cleaning)

residual purging process will be effective but the process and

hardware to effect it are unproven.

Finally, an assumption made early in the discussion of this

sequence was that EVA avoidance was a relatively high-priority

objective, at least for routine task performance. There are two

reasons. The first involves timeline impacts on baseline

operations. The total service timeline is extensive—prebreathing

before each EVA sortie and the manipulative limitations imposed by

the suit—are significant contributors to this total. Therefore,

minimization of this timeline through avoidance of EVA seems a

reasonable goal.

The second reason for assuming the desirability of EVA avoidance

involves the potential hazards inherent in management of fluids

and, specifically, management of cryogen transfer operations.

These operations can be performed by automation; there is no reason

to expose an EVA astronaut to the hazards involved under normal

operating conditions.

Inspection and Checkout Operations - Very simply stated, the

purpose of this sequence is determination of the "fitness" of the

OTV for continued operations. It involves two, possibly parallel,

activities—inspection and checkout—plus a third, analysis of

inflight data. It should occur before initiation of either

inspection or checkout. Figure 4.2-12 depicts the sequence.

4-55
Inspection

I

I Onboard ! I

{ Mobility I——

• Operations < |

I . 1 j

I

Analysis

of Flight

Data

t \

1

1

1

1

1

1

Checkout

! I Maintenance I

I i Operations I

I L !

Figure 4.2-12 Inspection and Checkout Operations

Inspection focuses on detection and interpretation ot anomalous

physical characteristics through sometimes augmented but typically

human sensory (principally visual) means.

Checkout, on the other hand, implies detecting and subsequently

isolating systems or components producing anomalous electronic

signals during simulated operating conditions if in a service bay,

and/or actual operating conditions if in flight. Data may

originate from specific sensors or from the capture of enroute

signals, depending on the system/component being checked.

These two activities could conceivably be conducted in parallel to

minimize timelines if technically compatible techniques are

employed. Undoubtedly they will often influence one another; data

from one will suggest the need for specific and possibly additional

attention to aspects of the other.

Inspection - The requirements for inspection will be a function of

the overall maintenance philosophy and its impact on vehicle

design. It is safe to assume that whatever the requirements, the

techniques employed will be nondestructive. Destructive test

techniques will presumably be reserved for unreparable

vehicle(s)/components and will, again presumably, be conducted by

"depot"-level ground-based activities.

Nondestructive inspection (NDI) techniques that could be considered

for use on an OTV at the space station include, but are not

necessarily limited to:

1) "Sniffer"-type leak detectors;

2) Unaided vision;

3) Image enhancement;

4) Multispectral imaging (e.g., thermal, electric field);

4-56
5) Dye penetrant;

6) X-ray radiographic;

7) Ultrasonic;

8) Eddy current;

9) Magnaflux;

10) Magnetic particle.

Each has advantages and disadvantages that affect its desirability

in space environment applications. For example, the magnaflux

technique requires inducement of a magnetic field and use of metal

flakes, both undesirable. Table 4.2-7 summarizes significant

strengths and weaknesses of these techniques in space-based

applications.

Table 4.2-7 NDI Technique Advantages and Disadvantages in Space

Technique

Sniffer(s)

Unaided Vision

Image

Enhancement

Multispectral

Imaging

Dye Penetrant

Radiographic

(X-Ray or

Gamma Ray)

Ultrasonic

Eddy Current

Magnaflux

(Magnetic

Particle)

Advantages

— Detect Small Concentrations

— Easy to Implement via Multiple Means (EVA

or Remote Camera)

- Permits "Seeing" Things That Normally Cannot

Be Seen

— Permits "Seeing" Things That Normally Cannot

Be Seen

— Enhances Visibility

— Does Not Require Elaborate Equipment

— Usable for Many Metals & Other Materials

— Film Provides Permanent Record

— Detects Surface & Subsurface Defects

— Simple Procedure

— Detects Surface & Subsurface Defects

— Can Provide a Permanent Record

— Simple Procedure

- Detects Surface & Subsurface Defects

— Simple Procedure

— Equipment is Portable

Disadvantages

— Requires "Sniffable" Medium

— Limited Capability

— Limited Consistency

— Equipment Required

— Equipment Required

- Limited Applicability

— Fluid Confinement

— Detects Only Surface Defects

— Weight of Equipment

— Dark Room Required

- Film Use

— Power Consumption

— Potential Safety Hazard

— Interface Medium (Coupling Agent)

Required

— Oscilloscope Required

— Interpretation Requires Highly Trained

Personnel

— Electric Current Required

— Does Not Give Physical Shape of

Discontinuity

— Employes Magnetism & Requires Loose

Particles

- Only Usable on Ferromagnetic Metals

— Electric Power Required

Inspection procedures, whatever they turn out to be, will probably

be conducted in the service bay. They will have a number of

implications for the OTV, for the facility and for support

equipment.

4-57
Until vehicle design becomes more firm, specific techniques and

requirements will be hard to identify but general requirements for

all techniques should be considered now. These include such things

as provisions for:

1) Access to all areas of the vehicle requiring inspection;

2) Lighting requirements;

3) Storage and handling of inspection aids (tools);

4) Minimization of inspection requirements and timelines.

Checkout - Three levels of checkout operations are envisioned. The

first, inflight operations, will focus on (1) identifying the

source of anomalous signals (unsatisfactory performance) so the

faulty component/system can, if possible, be brought back into

tolerance and, if not, be compensated for, and (2) recording the

anomaly for future analysis and correction. Compensation would be

accomplished by isolation and subsequent use of backup systems or

by other workarounds such as alternative approaches to the function.

The second, probably more extensive, level is postflight checkout.

These operations will focus on investigation (verification) of

inflight anomalies and the conduct of predictive investigations as

well as postmaintenance verification.

The third level of checkout operations occurs before operations

checkout. It could involve several separate procedures conducted

after unscheduled maintenance operations, after removal from

storage, after payload mating, after loading consumables and after

maneuvering to the launch point. Presumably these operations would

be comparatively brief and focus on ascertaining that specific

vehicle subsystems are ready for operation.

To be efficient, all these checkout processes must be effectively

automated. All three levels of checkout operations will require

similar interfaces and support, including:

1) Sensor/test equipment;

2) Communications link;

3) Data analysis and control hardware;

4) Data analysis and control software;

5) Data analysis and control executives.

The aspects of each will be discussed in the following paragraphs.

4-58
Automated checkout, particularly for inflight engine condition/

performance assessment, will demand substantial upgrading of

present diagnostic sensors. Ideally, these advanced diagnostic

sensors will be tied into local onboard or other no-delay

executives so they will be able to detect impending out-of-limit

situations and trigger immediate corrective action.

Diagnostic sensors will also have to be upgraded to support

effective automation of pre- and postflight checkout operations.

For these applications, attention will be focused on accurate and

timely diagnosis. Timeline minimization is seen as a key goal.

The use of these sensors and the issue of where they will

reside—built-in, plugged-in, or remotely located at the end of an

umbilical or RF communication link—will have a key impact on

system design. Overall system implications will be complex.

Turning from the sensors themselves to the handling of the data

they will produce, three additional sets of considerations become

relevant. These are the link from the sensors to the analytic

facility, the analytic facility including hardware, software, and

executives, and the process control conducted by the executives

and/or human supervisory controllers.

The communications link from the sensors to the analytic facility

can be very simple or very complex. A number of "whethers"

influence the issue. The first is whether the link will be direct

RF or through an umbilical into the space station. The next is

whether all diagnostic procedures will reside in programs at one

location or in several locations. Another involves whether time

delays induced by NASA's block-formatting can be worked around if

they are relevant.

Options for analytic facility basing include the ground, the space

station, and the OTV itself. Issues such as weight, information

processing requirements, equipment reliability and maintenance, and

communications (link) management will drive the basing decision.

In the end, multimode basing will probably be employed even though

it will increase the complexity of the communications system

because of the weight and maintenance advantages it will offer.

Process control, management of the checkout process, is seen as

presenting the most complex set of considerations. Relevant

technologies will still be evolutionary during the space station

era. Use of artificial (machine) intelligence techniques will

relieve the need for much of the present day's routine process

intervention (by humans) at simple decision points as well as

permit better use of other techniques such as parallel processing

to speed up the automatic checkout process. The human role will

evolve over time to that of supervisory control.

4-59
These machine intelligence technologies will also be directly

applicable to communications management. They will be able to make

routine and instantaneous decisions to (1) select the optimum

combination of OTV, space station, and ground-based information

processing resources, (2) select the optimum data link at a given

time and for a given situation, and (3) possibly impact the NASA

communication system block format-induced time delays. The end

result will be to speed up the checkout process while minimizing

OTV and space station information processing requirements.

Maintenance Operations - Maintenance has been defined as a specific

operational subset of "servicing" for the purpose of this

analysis. It is envisioned as involving assembly/disassembly and

upkeep action as taken to prolong system life, enhance system

operation, or prepare the system for a specific mission. As a

result, only two categories (types) of maintenance operation are

envisioned for this space-based reusable OTV. These categories are

module removal and replacement and general service operations.

Though there are only two categories of operations, a number of

specific (varying) characteristics within each category warrant

consideration. For example, module R/R could be scheduled or

unscheduled; it could be undertaken to restore operational

capabilities, to upgrade the vehicle, or to reconfigure it for a

specific purpose; and it could be undertaken by a number of

different methods or schemes. Potential onorbit maintenance

schemes include intravehicular activity (IVA) in a pressurized

compartment, IVA using remote manipulators, extravehicular activity

(EVA) using teleoperation, astronaut EVA using advanced

high-pressure "tactile" space suits, and IVA-based supervisory

control of advanced automation operating in either an IVA or EVA

environment. Table 4.2-8 summarizes the full range of spacecraft

maintenance method options. Onorbit maintenance approaches and

concepts are shown in some detail.

Table 4.2-8 Spacecraft Maintenance Mode, Approach and Concept

Mode

Approach

Concept

Expendable

Spacecraft

No

Maintenance

N/A

Ground—

Refurbishable

Spacecraft

Repair,

Refrubish, &

Reorbit

Total

Maintenance

Capability

Onorbit

Maintainable

Spacecraft

Space!

Mainte

Service

Bay

Level 1

Capability

EVA Automatic

Combination

itation

lance TM.^

Inside

Level 2

Capability

I

Manual

n

Visiting System

Maintenance

Orb

Service

Bay

Level 1

Capability.

/ JAutoEVA 1 mation

,|(TWS)

Combination

iter

Inside

Level

Capa

Mar

2

lility

lual

OMV/OTV

& Front-End

Kit

Level 1

Capability

Automation

Fetch

for

Orbiter

orS/S

I

Automation

.evel 1 Maint: Major Module Diagnostics and R&R; Fluids Replenishment; TBD Services (Kick It, Oil It, Lock/Unlock

and Reposition . . . )

.evel 2 Maint: Submodule Diagnostics and R&R; Submodule Repair (?); Limited Submodule Modification (?)

4-60
Because these schemes imply certain operational limitations and

impose specific requirements on space station OTV accommodations,

they will define maintenance and other service operation methods as

well as influence the OTV's configuration and the space station

facilities needed to accommodate the processes.

The numerous general service operations envisioned for the vehicle

involve many considerations. Some of these operations are wearout

prevention, cleaning, adjustment, recharging of energy storage

devices and refurbishment of thermal control coatings. Some of the

considerations involve access, modularization tradeoffs, definition

of advanced automation's role and design for multimode operations.

The many factors inherent in these operations and considerations

must be treated in any discussion of OTV maintenance. These will

be covered in subsequent paragraphs but first it seems appropriate

to briefly introduce Figure 4.2-13 that illustrates the

time/process relationships for OTV maintenance operations. This

depiction is not intended to be a linear representation of actual

time relationships; it does, however, identify and show the time

phasing of these operations.

CLOCK/CALENDAR TIME

INACTIVE TIME OPERATING TIME

PREVENTIVE

MAINTENANCE

TIME

STRESS

TIME

(SOiAflON

inuwstioetst

STRESS

TIME

[ DETECTION

HN$nCT10N«,

[CHECKOUT;

PREVENTIVE

MAINTENANCE

TIME

tSOLATKJN

(OIA6KQSB)

RECONFIGURATION

TIME

CORRECTIVE

(RESTORATIVE!

MAINTENANCE

TIME

ACTIVE

MAINTENANCE

TIME

ACTIVE

RESTORATION

' (REPAIR)

TIME

NON-ACTIVE

• MAINTENANCE

TIME

O PROGRAMMED

REMOVAL &

REPLACEMENT

OPERATIONS

O WEAROUT PREVENTION

• ALIGN

• ADJUST

• LUBRICATE

• OTHER FRICTION CONTROL

O CLEANING

0 CORRECTIVE

REMOVAL &

REPLACEMENT

OPERATIONS

O RETANKING

& OTHER

RELOADING

OPERATIONS

0 ADJUSTMENT

OPERATIONS

O PREPARATION FOR

MAINTENANCE

O WAIT FOR

• REPLACEMENT

MODULES

• INSTRUCTIONS

• TOOLS

• TEST EQUIPMENT

O ALL OTHER DELAYS

Figure 4.2-13 OTV Maintenance—Time/Process Relationships

The first aspect of OTV maintenance operations that will be

discussed is whether they are scheduled-or unscheduled. Scheduled

servicing includes planned functions conducted to permit continued

orbital operation of the OTV. Examples include component

replacement to avoid predicted breakdowns (design life limitations)

4-61
and to replace expendable items that have been consumed. The

necessary actions are known well in advance and the service crew

will have been familiarized and trained and will have conducted the

necessary verification simulations for these events before actual

operations. All crew aids/devices/tools and any other necessary

support equipment will also be in place and rehearsals will have

been conducted to support the scheduled servicing. Unscheduled

servicing involves the functions conducted on a contingency basis

to restore the OTV to an acceptable level of operational

capability. It includes repair or replacement of components that

fail randomly (unpredicted breakdowns) or because of "unscheduled

events" (accidents). It is also characterized by improvisation and

will, by definition, extend timelines. Table 4.2-9 summarizes

these points.

Table 4.2-9

Characteristics of Scheduled and Unscheduled Maintenance Operations

Scheduled Servicing

— Conducted to Permit Continued Orbital Operation

— Replaces/Refurbishes Design Life Limited Items and Consummable Items

— Programmed Operations

- Crew Well Trained and Verification Simulations Conducted

- Characterized by Use of Specifically Designed Tools and Support Equipment

Unscheduled Servicing

— Conducted to Restore the Vehicle to an Acceptable Level of Operational Capability

- Replaces/Refurbishes Items Which Have Failed Randomly or Due to Unscheduled Events

— Extend or Modify Timelines

- Characterized by Crew Improvization (Tools and Methods)

— Unprogrammed Operations

Given the tight turnaround schedule expected for the OTVs and the

characteristics of unscheduled servicing just listed, it would seem

a reasonable goal to seek to minimize unscheduled OTV servicing.

Maintenance Scenarios - The activity scenarios for maintenance

operations must be kept as simple and uncomplicated as possible to

accommodate present-day EVA processes and, at the same time,

promote evolution toward advanced automation. These scenarios will

be influenced by OTV characteristics (degree of modularization,

interfaces, retention hardware, etc) and by the evolution of

onorbit servicing equipment and techniques as we progress from the

present manual mode operational capability to increasingly

automated methods.

Detailed scenarios for these maintenance operations cannot be

developed at this time without a great deal of speculation about

the specific aspects of the modules, tools and processes involved.

However, experience gained through previous analyses of similar

situations (e.g., our remote orbital servicing satellite and

teleoperator human factors studies) indicates that, as for other

sequences and activities discussed earlier in this report, the

maintenance processes will be assembled from the known set of

4-62
"primitive actions" into higher level processes (subtasks, tasks,

activities, and sequences). This same experience also indicates

that module and tool design will significantly affect the ultimate

process design.

The Role of Automation and Artificial Intelligence - The purpose of

the present focused effort to apply advanced automation technology

to space station is not to replace man, but to augment or enhance

his capability in a technologically complex and sometimes hostile

environment. It is analogous to the development of diagnostic

tools (expert systems) for the medical profession. In this case,

there is no intent to replace the physician; rather the intent is

to increase his database and to enhance his diagnostic capability

using the experience of others that are "built into" the

user-friendly tool. Although we were able to design, build, and

operate flying machines long before the advent of the computer,

with the computer we have been able to extend our flight regime

well beyond the solar system.

A recent aerospace automation industry study group recommended four

areas for technology advancement. Three are relevant to OTV

servicing efforts at the space station:

1) Spacecraft services self-management (through AI);

2) Ground support automation;

3) Space station robotics for inspection, assembly, servicing and

repair, refurbishment and onorbit experiment interaction.

The first area involves the application of artificial intelligence

and expert systems to spacecraft services and self-management.

Current artificial intelligence research emphasizes the development

of self-modifying codes, that is, software that learns, modifies

itself, and becomes smarter and better. Useful results are being

obtained in this area and some expert systems that have been built

perform reasonably well. Some exploratory work is being conducted

at NASA Kennedy Space Center on the application of expert systems

to ground support operations (ground support automation). The real

benefit of expert systems that is directly relevant to onorbit

servicing is to reduce the number of human experts needed to do a

given job that requires human expert judgment. The human experts

cannot be eliminated, but they can be called on less frequently and

made more productive. This is the recurring theme for development

of automated design and analysis tools.

The following list summarizes potential expert systems applications

in space-based OTV servicing:

1) Assist crew in failure detection, diagnostics;

2) Direct test procedures;

4-63
3) Chart preferred servicing modes and task sequencing;

4) Assist crew in decision making (autonomous decision making

without ground support;

5) Direct safing procedures before crew access;

6) Assist crew in responding to emergencies;

7) Prevent crew-overload from massive data flow;

8) Enable autonomous remote servicing (a long-term prospect).

There is potentially a very large cost impact (savings) in this

area because the application is not understood (so no one knows how

to plan for it). It is important to understand the potential of

this automated technology before the OTV and space station program

costs are cast in concrete and budgeted.

Robotics/AI on the space station will be specifically useful for

OTV inspection, checkout, assembly, servicing and repair

processes. Here, robotics deals with both true robotics (machine

intelligence is applied and the mechanisms are autonomous) and

telepresence. In telepresence, communication techniques put man in

in a remote location so intimately in the loop that he loses

"sight" of his remote location and feels that he is physically

present at the worksite and is himself doing what his robot proxy

is actually doing. Figure 4.2-14 depicts the space shuttle's RMS

system. It is an existing teleoperator system that includes

necessary telepresence capabilities. Figure 4.2-15 summarizes the

relationship of these technologies to each other and to manual

operations and total autonomy and Figure 4.2-16 summarizes their

characteristics.

It is conceivable under certain scenarios that the use of presently

available (state-of-the-art) automated maintenance techniques could

eliminate the need for space station personnel to conduct EVA

operations in support of many OTV maintenance operations. Of

course some contingency situations will always remain that require

the versatility and capability to improvise provided by EVA

personnel.

For example, it should be possible to use existing teleoperation

capabilities to remove and replace (R/R) major assemblies

containing multiple submodules. Further, the mechanism could

transfer the major module assemblies to an interlock area where the

station crew could access the assemblies for submodule repair all

without any EVA. Extending this concept, the mechanism could

easily be equipped with a camera or other imaging system(s) to

enable the station crew to visually inspect the OTV from within the

space station.

4-64
Window

\ View /

Crew

Compartment

Bulkhead Maintenance

Bay

Elbow CCTV

on Pan & Tilt

Unit

Legend:

MCIU—Manipulator Controller Interface Unit

GPC—General-Purpose Computer

RHC—Rotational Hand Controller

THC-Translational Hand Controller

CRT-Cathode Ray Tube

Kybd-Keyboard

CCTV-Closed-Circuit TV

Retention

Devices

Figure 4.2-14 The Shuttle RMS Teleoperator System

Hands-On Remote Control

• Performance Growth

»-Greater Operator Sensitivity

»• Greater Operator Utility

Decrease in

Human Operators

Fewer Operators and

Greater transparency

Man Out of the Loop

Intelligent

Automation

Figure 4.2-15 The Relationship of Servicing Technologies

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Autonomy
Reduce Crew Involvement

Increases Precision in Parts Handling

Speeds Up Servicing Schedule

Artificial

Intelligence

(Supporting Man)

Teleoperations

and

Telepresence

(Man-in-the-Loop)

— Supports Trouble Shooting

— Saves Crew Time

— Increases Autonomy

— Guides Servicing Operations

— Minimizes EVA Tasks

— Enables Load Handling & Transfer

— Enables Remote (In-Situ) Servicing

- Adapts to Unforeseen Conditions

Figure 4.2-16 Characteristics of Automated Tasks

This mechanism would have to be flexible, probably programmable to

some degree, and retain the option to conduct remote control

operations. If the overall system engineering approach ensures the

OTV's design packaging complements the mechanism's capabilities,

programming needs can be minimized. For example, if the OTV were

designed to break down into five major assembly groups, the

maintenance mechanism would only require five program variants to

R&R the major assemblies. A potential maintenance scenario

employing these capabilities might be something like that described

in the following paragraph.

The space station's onboard computer detects a fault during the

preoperational checkout sequence and proceeds to isolate the

underlying failure to a specific submodule. A space station crew

member would subsequently note the error message and determine (if

the system had not already) that the submodule is in major assembly

group number two. The crew member could then install a

preprogrammed disk, tape or chip into the robot command computer or

simply authorize execution if the program were already on line.

The OTV's berthing platform would then rotate to its proper

orientation and the maintenance mechanism would move down its rail

and proceed to remove major assembly number two. It would then

carry the assembly into the interlock area. The interlock area

would then be pressurized so the space station crew could access

the assembly to R/R the failed submodule. The repaired major

assembly could then be returned to the vehicle and reinstalled by

the mechanism. The space station onboard checkout computer could

then verify the repair and continue the preoperational checkout

sequence.

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With respect to the development and application of automation and

machine intelligence technology, consider the following concerns:

1) Although we have built and tested expert systems and natural

language front-ends for various applications, we have only

recently become concerned about the real-time operation aspects

of expert systems. Much work is required over the near term if

we are to be able to successfully implement automated real-time

decision-making processes for a variety of space station-based

OTV maintenance applications such as we are doing in the

current automated management of power systems (expert systems

development) project;

2) With regard to robotics and teleoperators, once again, the

emphasis is not on replacing man in space but rather on

enabling a greater amount and variety of more dextrous, remote

and cost effective activities. To accomplish this, significant

laboratory work is required to evaluate such potential

techniques as task allocation schema, available control modes,

and human interaction optimization. It should be noted that in

many cases unique laboratory and space facilities will be

required to investigate these potential capabilities and to

define specific system requirements.

In summary, with specifically relevant and timely advanced

automation-related initiatives, we will be able to take a

significant step beyond present capabilities and extend man's reach

and utility through the use of integral system intelligence,

robotics, teleoperations, sensory perception, and unique processing

and packaging concepts (VHSIC). We must, and will, continue to

pursue optimized use of man with the goal of increased productivity

in space.

Maintenance Concept and Implications - The maintenance conceptenvisioned for this vehicle was discussed in some detail in Section

4.1. The aspects of the concept that warrant further discussion at

this time and in this context are treated in the following

paragraphs.

The concept is not firmly set, it is evolutionary and is expected

to develop over time as the knowledge (definition) of OTV and space

station operations and equipment improve. The philosophy

underlying the concept is to limit what is to be done at the

vehicle to minimize the complexity of service bay operations and

thereby reduce EVA requirements and maximize the near-term use of

automation. This philosophy leads to an interactive set of vehicle

design approach and maintenance concept considerations.

The first point of the maintenance concept is to perform only major

module removal/replacement in the service bay, not onsite component

repair or replacement. Such activity would be unnecessarily

difficult and time consuming when compared to available

alternatives. The next point involves "intermediate"-level

maintenance that is defined for this purpose as submodule R/R

4-67
operations. They are envisioned as possibly being conducted on an

as-necessary (contingency) basis inside the space station or in the

service bay if the parent module is too big to transit the

airlocks. The final point of the concept is to avoid returning the

entire OTV to the ground. If the vehicle is to be space-based and

if it can be highly modularized, it should only be necessary to

deliver and install one or more replacement modules to correct any

anomaly that occurs.

Several vehicle design approach considerations have been

identified. The first is component grouping (by life, type, and

location) into submodules and "major modules"—roughly comparable

to today's space-replaceable unit concept. Four types of

module-to-module and module-to-structure interfaces are

envisioned. These are fluid, electric power, mechanical, and

data. Mechanical module interface options (roughly illustrated in

Fig. 4.2-17) include multiple-fastener direct, single-fastener

direct (like MMS), and indirect through a transition plate. There

similarly are a number of alternative options for each of the other

module interface types, each with specific advantages and

disadvantages.

LDEF

Experiment

Tray

IOSS

Retrofit

Concept

Individual Fasteners

Module

Direct to

Spacecraft

MMS

Module

Single

Fastener

Mechanism

Module

Direct to

Spacecraft

Module

Indirect to

Spacecraft

Figure 4.2-17 Mechanical Module Interface Options

Although it will be highly desirable to standardize and minimize

the number of interfaces (particularly for the major modules), the

design process may possibly mandate that a variety of options be

employed for one or more of the interface types because of varied

operational and maintenance requirements and characteristics.

Figure 4.2-18 shows two concepts for the engine/OTV interface that

would both employ a variety of interface types.

4-68
Discrete Concept Planar Concept

OTV

Figure 4.2-18 Engine/OTV Interface Concepts

The implications of this standardization objective extend beyond

design of the OTV itself. We will need to carefully develop tasks,

processes and tools common to the widest possible variety of

space-based servicing operations. A few of the other space station

activities that would make use of these common facilities are

payload assembly and/or repair, OMV servicing, and expendable

propulsion vehicle servicing.

The implications of these points are:

1) To seek common onorbit servicing capabilities and requirements

for the widest possible range of applications;

2) To seek OTV, OMV, MMS, and other systems (vehicles and

payloads) commonality at the widest possible range of

levels—from major module to submodule, and, possibly, to the

component level.

The following basic maintainability principles also provide design

objectives for OTV system development:

1) Design to minimize the complexity of maintenance tasks such as

adjustment, inspections and calibrations;

2) Design for optimum interchangeability and use of standardized

(or commercial) items;

4-69
3) Design for rapid and positive recognition of equipment

malfunction or marginal performance;

4) Design for minimal resource requirements in accomplishing

maintenance tasks;

5) Design for rapid and positive identification of the replaceable

defective part, assembly or component;

6) Design to require minimum maintenance skills and training;

7) Design to require minimum numbers and types of tools and test

equipment (special and standard);

8) Design for optimum access to all equipment and components

requiring maintenance, inspection, removal or replacement;

9) Design for maximum safety for both equipment and personnel;

10) Design so the mean time to accomplish scheduled and unscheduled

maintenance is sufficiently low to assure the required

operational availability of the equipment;

11) Design to enhance and facilitate organizational- and

intermediate-level maintenance action;

12) Design to require a minimum number and type of repair parts and

assemblies;

13) Provide technical data for easy use (AI maintenance tutors,

etc) concurrently with the equipment;

14) Maximize the extent that performance can be verified,

malfunctions anticipated and located, and calibration performed.

In regard to the impact of advanced automation on the OTV

maintenance concept, a few additional points need to be made. The

first is that the OTV modules, support equipment and maintenance

procedures should be designed for multimode operations. Whether we

baseline EVA activity, automated module replacement, or some

intermediate capability, the need to accommodate evolution and

contingency operations will remain. As Figure 4.2-19 depicts, it

should ideally be possible to conduct these multimode operations

with only minimal mode-to-mode change—substitute a proxy for the

man but use the same work carriage and module service tool(s).

The next point is that module size and interfaces must be carefully

engineered to be compatible with the baseline (1990s) automation

capabilities. And the last point, a related one, is that small

components/assemblies must be carefully grouped to attain optimum

module and submodule size without inextricably confusing the

ability to develop reliability calculations and conduct maintenance.

4-70
MonuiF SERVICING (EVA)

nx

-x

/ ^—HANGAR

L^ LIGHTING

/ MODULE SERVICE TOOL

MODULE SERVICING (REMOTE)

-FOOT RESTRAINT

WORK PLATFORM

MULTIPLE POSITION

TRANSLATION CARRIAGE

MULTIPLE POSITION

TRANSLATION CARRIAGE

MODULE

SE«VICE

TOOL

GENERALPURPOSE

MANIPULATOR

SERVICER

figure 4.2-19 System Design for Multimode Maintenance Operations

OTV Storage - This sequence (if indeed it can qualify as a

sequence) is the transition between postflight and preflight

operations. The processes involved are expected to focus only on

thermal conditioning of the tanks and possibly conditioning

portions of the power system. It could, however, prove

advantageous to include provisions for data system access within

the storage facility to accommodate automated checkout and

reprogramming while the OTV is there. These provisions would

permit extended use of such other maintenance facility capabilities

as inspection and module changeout.

Transfer Stack Assembly - The transfer stack assembled would always

be a function of the mission to be performed. It would include a

payload for deployment missions, a servicer-kit and possibly an OMV

for service missions and some retrieval missions, and only the OTV

itself for many retrieval missions.

The sequence is envisioned as including two distinct activities.

The first would be employed only if a scenario using the OMV for

close-in mobility operations or remote recovery operations were

selected. This activity is mating of the OMV and the OTV. The

scenario for the activity is to use available manipulators in the

facility to position the two vehicles and hold them while the

4-71
docking interface was activated. Tasks in the activity would be

essentially similar to those in the destacking activity previously

discussed.

The second activity, joining the payload (or the servicing kit) to

the OTV, would routinely be performed. The single exception would

be for recovery missions that would not require front-end kit

(manipulator) preparation before orbit transfer. Again, this

activity contains essentially the same tasks as the destacking

activity. The noticeable difference would probably be in the

positioning accuracy required before docking.

Remaining Sequences - The balance of the sequence depicted in

Figure 4.2-5 are checkout operations, consumables loading,

preparation for launch, proximity operations, and launch (that

really begins the mission operations sequence). These will not be

treated further because they are either repeats of previously

discussed sequences (or portions thereof) or rearrangements of

activities treated in previously discussed sequences. They do not

introduce any new servicing considerations given the present depth

of analysis.

4.3 RELIABILITY ANALYSIS AND EFFECTS OF NONINDEPENDENT FAILURES

A reliability analysis was performed that considers the effects of

nonindependent failures (failures that render the entire main

propulsion system inoperative) on mission success probability. The

purpose of generating main propulsion system mission reliability

numbers (other than for the conclusions that can be drawn from them

alone) was to incorporate the delta reliability between various

engine configurations into the life-cycle cost comparisons of these

configurations to account for mission loss costs. This is

discussed in Section 4.4 of this report.

A range of engine single-burn reliability from 0.994 to 0.999 was

chosen to bound the problem. Also, a nonindependent failure rate

from 0 to 5% was selected. Then, for a six-burn mission, the main

propulsion system reliability for completing the mission was

calculated for various numbers of engines (from 1 to 4) and

engine-out capability. The results are shown in Figures 4.3-1

through 4.3-3. The obvious results show that multiple engines (3

or greater) provide the highest mission reliability for low (less

than about 2%) nonindependent failure rates. Correspondingly, two

engines provide the highest reliability for the higher

nonindependent failure rates. The crossover points of highest

mission reliability vary as a function of single burn reliability,

however. Agreement concerning reasonable estimates of engine

single-burn reliability and nonindependent failure rates will have

to be reached within the industry before firm conclusions can be

drawn. However, the trends shown here should aid in the definition

of engine and main propulsion system design decisions.

4-72
.999990

.999975

.99995

CO

.975 -

.95 -

.90

No. Engines/Engine-Out Capability

2/1

•1/0

0.00 0.01 0.02 0.03 0.04 0.05

Non-lndepenaent Failure Rate

Figure 4.3-1

Main Propulsion System Reliability-Single-Burn Reliability = 0.994

4-73
.999990

CO

s

a

c

No. Engines/Engine-Out Capability

.975

.95 -

-1/0

.90 "- • ' i i i

0.00 0.01 0.02 0.03 0.04 0.05

Non-Independent Failure Rate

Figure 4.3-2

Main Propulsion System Reliability—Single-Burn Reliability - 0.996

4-74
.999990

.999975

.99995

No. Engines/Engine-Out Capability

i i i i *

0.00 0.01 0.02 0.03 0.04 0.05

Non-independent Failure Rate

Figure 4.3-3 Main Propulsion System Reliability-Single-Burn Reliability = 0.999

4.4 LIFE-CYCLE COST (LCC) ANALYSIS

The major efforts reported here include the development of a LCC

model for the development, production, and operation of the

baseline space-based reusable OTV. The cost trends and comparisons

that the LCC model was used to generate data for are also included.

4.4.1 LCC Model Development

Figure 4.4-1 depicts the methodology used for development of the

OTV LCC model. Table 4.4-1 defines the terms used for each of the

model development stages and this is followed by a description of

the modeling approach as applicable to the OTV.

4-75
OTV OPERATION AND

PROPELLANT USAGE

RETURNED

OTV MODULES

STS LAUNCH

GSE

OTV SERVICING (IVA/EVA)

SPARES

OPERATIONS CONTROL

MISSION PREPARATION

OVERHAULyCONTINGENCY R/R

PROPELLANT,

SPARES,ETC,

OTV PRODUCTION

MISSION

SUPPORT

Figure 4.4-1 OTV Life-Cycle Cost (LCC) Model Scenario

Table 4.4-1 LCC Model Development Stages

Stage Description

Problem Formulation

System Definition

Modeling

Model Formulation

Data Preparation

Model Translation

Validation

Experimentation

Interpretation

Implementation

Documentation

Identify the problem requiring solution.

Determine the boundaries, restrictions and measure of effectiveness to be used in

defining the system to be studied.

Determine the most suitable form <of modeling for solution of the problem.

Reduce the abstraction of the real system to a logic flow diagram.

Identify the data needed by the model and reduce the data to an appropriate form.

Describe the model in a language suitable for computer use.

Assure, to an acceptable level of confidence, that an inference drawn from the

model about the real system will be correct.

Execute the model to generate the desired data and then perform sensitivity

analyses.

Draw inferences from the data generated by the model.

Put the model and results to use.

Record the project activities and results and document the model and its use.

4-76
Problem Formulation - All the major cost elements associated with

the deployment, design, development, test and evaluation (DDT&E),

production, operation and maintenance of an orbital transfer

vehicle (OTV) were identified. Cost sensitivities were analyzed to

assist in obtaining an optimal design in terms of operational

capability and life-cycle cost.

System Definition - Figure 4.4-1 depicts the OTV development and

operating scenario. Total cost was developed for DDT&E and

production of two OTVs, ground preparation and launch via the

shuttle. The OTV was assumed to be docked at and operated from the

space station. All OTV missions were costed with propellant use

and EVA time. Scheduled and unscheduled repair actions, including

EVA, IVA, spares, ground resupply etc were costed against the OTV.

Modeling - The most appropriate approach for developing cost

predictions for conceptual design studies was through cost

estimating relationships (CERs) based on known design parameters

and factors from similar programs and reasonably substantiated

ground rules and assumptions. The CERs were developed for all

major cost elements associated with DDT&E, production and

life-cycle operation and maintenance of the OTV. Certain

parameters within the CERs may be fixed values, whereas others were

variables to enable sensitivity analyses to be performed. These

sensitivity analyses were beneficial in achieving the optimal

design at the least LCC.

Data Preparation - The following tabulates the baseline ground

rules and parameters used in development of the OTV LCC model.

All costs reported in millions of 1984 $.

One failure per two OTV missions (one OTV subsystem module R/R).

$30k per hour per man for EVA.

$10k per hour per man for IVA.

Seven missions per year for seven years.

30 missions between major overhauls (both modules and engines).

Four hours EVA R/R time for unscheduled failures.

Four hours IVA time for unscheduled failures.

Orbiter fees for OTV (and spares) transportation based on percent of total orbiter cargo weight capacity.

OTV ground processing factored from STS launch operations.

Flight operations at JSC $12k/hour.

STS transportation cost $1476/pound.

Average module weight 33 pounds.

13 module types (52 total modules).

50 manhours EVA turnaround time per mission.

112 manhours IVA turnaround time per mission.

One, two, three and four engine configurations.

Propulsion system total thrust 15,000 Ibf.

The following CERs have been developed for inclusion in the OTV

life cycle cost model. Each CER algorithm and source is presented,

followed by the validity range and a description of the associated

rationale.

4-77
Structure/Mechanisms (MSFC)

STRRDM = 0.821 x STRUG0'

437

STRUCK = 0.064 x STRUC0-

545

where: STRRDM = Structure/mechanism DDT&E cost

STRUCM = Structure/mechanism unit cost

STRUC = Structure/mechanism weight in Ibm.

Validity range: 60 to 2000/lbm

The structure/mechanism subsystem consists of a primary structure

of an aluminum tubular truss design. It includes trunnion fittings

and aluminum panels for the mounting of avionic components and

assemblies. Docking and interface mechanisms for payloads and the

orbiter are also included.

Thermal Control (MSFC)

THRMRDM = 0.821 x THERM0'

437

THRMUCM = 0.064 x THERM0'

545

where: THRMRDM = Thermal control subsystem DDT&E cost

THRMUCM = Thermal control subsystem unit cost

THERM = Thermal control subsystem weight in Ibm.

Validity range: 60 to 2000 Ibm

The thermal control subsystem maintains the temperature of the

spacecraft, mission equipment, and propellant within allowable

limits in certain orbital conditions. The thermal control

subsystem includes paint, insulation, radiators, heaters, louver

assemblies, temperature sensors, heat pipes, and a thermodynamic

vent system. Thermal control can be accomplished either passively

or actively. Passive control often means nothing more than a coat

of reflective paint. Active control ranges from mechanical

shutters to heat pipes running through the structure.

Propulsion (Space Division)

PROPRDM = 1.668 + (0.126 x PROP)

PROPUCM = 0.43 x PROP0-

494

where: PROPRDM = Propulsion subsystem DDT&E cost

PROPUCM = Propulsion subsystem unit cost

PROP = Propulsion subsystem weight in Ibm

Validity range: 200 to 4000 Ibm

The propulsion subsystem provides a reaction force for

orbit-keeping, orbit-changing, and attitude control. The

propulsion subsystem includes a reaction control system, fuel and

oxidizer tanks, pressurization and accumulator system, and flow

system hardware. Specifically excluded are costs for an advanced

engine and onorbit propellant costs.

4-78
Electric Power (MSFC)

EPSRDM = 0.597 x EPS0'

584

EPSUCM = 0.042 x EPS0'

784

where: EPSRDM = Electric power subsystem DDT&E cost

EPSUCM = Electric power subsystem unit cost

EPS = Electric power subsystem Weight in Ibm

Validity Range: 100 to 10,000 Ibm

The electric power subsystem stores, converts, regulates, and

distributes electrical energy to and between spacecraft

components. The electric power subsystem includes batteries, fuel

cells, regulators, converters, distribution units and wiring

harnesses.

Guidance, Navigation and Control (MSFC)

GNCRDM = 3.08 x GNC0'

516

GNCUCM = 0.5763 x GNC0'

494

where: GNCRDM = Guidance, navigation and control subsystem DDT&E

cost

GNCUCM = Guidance, navigation and control subsystem unit

cost

GNC = Guidance, navigation and control subsystem weight

in Ibm

Validity Range: 200 to 4000 Ibm

The guidance, navigation and control subsystem generates control

signals that interface with the propulsion subsystem to produce

thruster forces for translation and stabilization of the spacecraft

orbit. The guidance, navigation and control subsystem includes

inertial reference units, accelerometers, range-rate sensors, valve

control electronics, and attitude computers.

Data Management and Communications (MSFC)

DMCRDM = 0.615 x DMC0'

653

DMCUCM = 0.053 x DMC0'

917

where: DMCRDM = Data management and communications subsystem DDT&E

cost

DMCUCM = Data management and communications subsystem unit

cost

DMC = Data management and communications subsystem

weight in Ibm

Validity Range: 50 to 1000 Ibm

4-79
The data management and communications subsystem processes

information on spacecraft conditions and mission progress, stores

or transmits such data to the ground, and receives commands from

the ground and initiates their execution. The data management and

communications subsystem consists of processors, memory storage

units, coders and decoders, timing units, control consoles,

antennas, transmitters, and receivers. The software needed to

operate the data collecting and transmitting functions is also

included.

Aerobrake (MSFC)

ABRDM = 1.668 x (0.126 x ABWT)

ABUCM = 0.064 x ABWT0*

545

where: ABRDM = Aerobrake DDT&E cost

ABUCM = Aerobrake unit cost

ABWT = Aerobrake weight in Ibm

Validity Range: 60 to 2000 Ibm

The aerobrake CERs are the same as those used for

structure/mechanisms. As the design matures, an update should be

made to account for the aerobrake specifically.

Advanced Engine (Pratt & Whitney)

ENGUNITWT = 28.33 x THRUST0*

2825

ENGWT = ENGUNITWT x ENG

ENGDD = 306.25 + (0.0029 x THRUST)

ENGUNITCOST = 1.963 + (0.000103 x THRUST)

ENGUC = ENGUNITCOST x ENG

where: ENGUNITWT = The single engine weight in Ibm

THRUST = Single engine thrust in Ibf

ENGWT = Engine system weight in Ibm

ENG = Number of single engines on the OTV

ENGDD = Engine DDT&E cost

ENGUNITCOST = Unit cost of a single engine

ENGUC = Engine system unit cost

Validity Range: 250 to 500 Ibm

3000 to 15000 Ibf

The engine subsystem weight equations and CERs were developed from

data supplied by Pratt & Whitney. They represent the best-fit

regression curves of single data points.

4-80
Spacecraft DDT&E Cost

DDTE = STRRDM + THRMRDM + PROPRDM + EPSRDM + GNCRDM + DMCRDM +

ABRDM + ENGDD

where: DDTE = Spacecraft DDT&E cost

STRRDM = Structure/mechanism subsystem DDT&E cost

THRMRDM = Thermal control subsystem DDT&E cost

PROPRDM = Propulsion subsystem cost

EPSRDM = Electric power subsystem cost

GNCRDM = Guidance, navigation and control subsystem cost

DMCRDM = Data management and communications subsystem cost

ABRDM = Aerobrake DDT&E cost

ENGDD = Advanced engine DDT&E cost

The DDT&E equation gives design, development, test, and evaluation

cost of the total OTV. A detailed discussion of each subsystem is

given in the preceding pages.

Flight Hardware Cost

FH = STRUCM + THRMUCM + PROPUCM + EPSUCM + GNCUCM + DMCURM + ABUCM

+ ENGUC

where: FN = OTV flight hardware cost

STRUCM = Structure/mechanism subsystem DDT&E cost

THRMUCM = Thermal control subsystem unit cost

PROPUCM = Propulsion subsystem unit cost

EPSUCM = Electric power subsystem unit cost

GNCUCM = Guidance, navigation and control subsystem unit

cost

DMCUCM = Data management and communications subsystem

unit cost

ABUCM = Aerobrake unit cost

ENGUC = Advanced engine system unit cost

The flight hardware equation gives the unit cost of a single OTV.

Each subsystem is discussed in detail in the preceding pages.

Spacecraft Weight

SCWT = (STRUC + THERM + PROP + EPS + GNC + DMC + ABWT + ENGWT)xl.l5

where: SCWT = OTV Weight in Ibm

STRUC = Structure/mechanism subsystem weight in Ibm

THERM = Thermal control subsystem weight in Ibm

PROP = Propulsion subsystem weight in Ibm

EPS = Electric power subsystem weight in Ibm

GNC = Guidance, navigation and control subsystem

weight in Ibm

DMC = Data management and communications subsystem

weight in Ibm

ABWT = Aerobrake weight in Ibm

ENGWT = Advanced engine system weight in Ibm

4-81
The OTV weight equation gives the total weight for a single OTV.

Space Station Crew Cost

SMAINT = (SEVA x EVA + SIVA x IVA) x MISYR x NYO/106

USMAINT = (EVARR x EVA + IVARR x IVA) x FAIL x MISYR x NYO/106

where: SMAINT = Scheduled maintenance labor cost

SEVA = Scheduled EVA time in hours

EVA = IVA cost per hour

SIVA = Scheduled IVA time in hours

IVA = IVA cost per hour

MISYR = Number of missions flown by OTV per year

NYO = Number of years of OTV operation

USMAINT = Unscheduled maintenance labor cost

EVARR = Unscheduled EVA time in hours

IVARR = Unscheduled IVA time in hours

FAIL = Number of failures per OTV mission

The space station crew cost equations represent the cost of EVA and

IVA time while on orbit.

Program Support Elements (MSFC)

GSE = 0.10 x DDTE

SEID = 0.10 x (DDTE + GSE)

PMD = 0.05 x (DDTE + GSE + SEID)

IACP = 0.20 x FH

SEIP = 0.10 x (FH + IACP)

PMP = 0.05 x (FH + IACP + SEIP)

where: GSE = Ground support equipment cost

DDTE = OTV DDT&E cost

SEID = System engineering and integration cost

PMD = Program management cost for DDT&E phase

IACP = Installation, assembly, and checkout cost for

DDT&E phase for OTV subsystems

FH = OTV flight hardware cost

SEIP = System engineering and integration cost for

production phase

PMP = Program management cost and integration cost for

production phase

The program support equations were derived from NASA costing

information. They represent an average percentage factor of

certain elements.

4-82
Propellant Cost (Derived from Rev 6 Mission Propellant Requirements)

PROPWT = [BPROP + 148x (SCWT-5096) - 7660 x (ISP - 480)]

PROPLB = SHLC x 106 / 65000

PROPCOST = PROPLB x PROPWT / 106

where: PROPWT = Total propellent weight in Ibm

BPROP = Baseline propellant weight in Ibm

SCWT = Total OTV weight in Ibm

ISP = Advanced engine specific impulse in seconds

PROPLB = Propellant cost per Ibm

SHLC = Shuttle launch cost in millions of dollars

PROPCOST = Total propellant cost in millions of dollars

The propellant cost equation calculates the total propellant cost

based on the basic propellant usage plus the effects of changes in

OTV dry weight and engine specific impulse. The cost per pound is

calculated assuming a 65,000-lb cargo capacity for the shuttle.

Operation and Support Factors (Engine Replacement and Spares

Scenario)

ENGREPCOST = (ENGUNITCOST x ENGREP x MISYR x NYO) +

(M1SYR x NYO/MISOV x ENGUC)

SPARES = FAIL x MISYR x NYO

SPARESWT = SPARES x MODWT

OVWT = MODS x MODWT

MODULEWTS = SPARESWT + OVWT

SPARESCOST = (0.22 X MODULEWTS0'

44) + ENGREPCOST

LOGISTICS = (0.6 x FH) + (0.02 X SPARESCOST x NYO) +

(0.0005 X DDTE x NYO/3)

GC = 15 x NYO

where: ENGREPCOST = Total cost for engine replacements over the

life of the OTV

ENGUC = Advanced engines system weight in Ibm

ENGREP = Number of engines replaced per OTV mission

MISYR = Number of spacecraft missions per year

NYO = Number of years of OTV operation

MISOV = Number of OTV missions between scheduled

overhauls

SPARES = Number of replacement modules required over

the OTV lifetime

FAIL = Number of failures per OTV mission

SPARESWT = Contingency replacement module weight in Ibm

MODWT = Average replacement module weight in Ibm

OVWT = Replacement module weight, in Ibm, per OTV

overhaul

MODS = Number of replacement modules on the OTV

4-83
MODULEWTS = Total replacement module weight, in Ibm

SPARESCOST = Total cost of OTV spares

LOGISTICS = Total cost of training, technical data,

inventory control, and sustaining engineering

FH = OTV flight hardware cost

DDTE = OTV DDT&E cost

GC = Ground control cost of OTV operations

The operation and support equations represent the cost, except for

labor, to maintain and operate an OTV at the space station.

Mission Loss Cost (two loads propellant and turnaround operations)

MISLOSS = (PROPCOST + SMAINT + USMAINT + SPARESCOST + LOGISTICS) x

MPS x 2

where: MISLOSS = Cost to recover a lost mission

PROPCOST = Total propellant cost

SMAINT = Scheduled maintenance labor cost

USMAINT = Unscheduled maintenance labor cost

SPARESCOST = Total cost for spacecraft spares

LOGISTICS = Total cost for training, technical data,

inventory control, and sustaining engineering

MPS = 1 - Main propulsion system reliability

The mission loss equation represents the cost to complete failed

missions using a second OTV and to retrieve the disabled OTV on the

return trip. The number of mission losses is calculated by

subtracting the main propulsion system reliability, which is the

probability that the system will function properly, from 1.00

(which represents 100% success) and multiplying by the total number

of missions flown.

Life-Cycle Cost

LCC = DDTE + FH + SEID + SEIP + GSE + PMD + PMP + SPARESCOST +

SMAINT + USMAINT + LOGISTICS + GC + PROPCOST + MISLOSS + IACP

where: LCC = OTV life cycle cost

DDTE = OTV DDT&E cost

FH = OTV flight hardware cost

SEID = System engineering and integration cost for

the DDT&E phase

SEIP = System engineering and integration cost for

the production phase

GSE = Ground support equipment cost

PMD = Program management cost for the DDT&E phase

PMP = Program management cost for the production

phase

4-84
SPARESCOST - Total cost of OTV spares

SMAINT = Scheduled maintenance labor cost

USMAINT = Unscheduled maintenance labor cost

LOGISTICS = Total cost of training, technical data,

inventory control, and sustaining engineering

GC = Ground control cost for OTV operations

PROPCOST = Total propellant cost

MISLOSS = Cost to recover a lost mission

The life-cycle cost equation represents the entire cost of the OTV

from its inception throughout its useful life.

Model Translation - The CERs were translated into BASIC and entered

on an IBM personal computer.

Model Validation - Each CER was developed from data obtained from

numerous previous and current space programs. Historical data are

probably the most valid and accurate from which to develop CERs and

these data were used in the preparation of the OTV CERs. Much of

the data was contained in the Space Station Cost Model (PRC D-2115)

and the remainder were developed through analysis of OTV design

requirements.

Model outputs of dollar values for the cost groups of DDT&E,

production and operations/maintenance were compared with actuals or

other similar complexity programs and found to be of the same

nominal order of magnitude.

4.4.2 Life-Cycle Cost Results

Life-Cycle Costs for Engine Configurations - The LCC and propulsion

system performance reliabilities were developed for 1, 2, 3, and 4

engines with single-engine burn reliabilities of 0.994, 0.996, and

0.999, and nonindependent failure rates (NIFR) of 0, 0.03 and

0.05. Figures 4.4-2 through 4.4-19 reflect OTV LCC and propulsion

system reliability for the respective number of Pratt & Whitney and

Aerojet engines. The engines selected for LCC/performance analysis

were 1 engine with a 0 engine-out capability, 2 engines with a 1

engine-out capability, 3 engines with a 2 engine-out capability and

4 engines with a 3 engine-out capability.

4-85
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4-103
Based on an engine selection criteria of minimum LCC, Table 4.4-2

reflects the optimum number of engines for all combinations of

engine burn reliability and nonindependent failure rate for the

Pratt & Whitney engine data.

Table 4.4-2

Pratt & Whitney Optimal LCC Propulsion Configuration

Nonindependent Failure Factor

Burn Reliability

0.994

0.996

0.999

0.00

2 Engines

2 Engines

1 Engine

0.03

2 Engines

2 Engines

1 Engine

0.05

2 Engines

2 Engines

1 Engine

An examination of Table 4.4-2 shows two engines to be optimal under

the selection criteria for all but the 0.999 single-engine burn

situation. Further examination of Figures 4.4-2 through 4.4-10

demonstrates three or four engines to be only marginally superior

in reliability and significantly higher in LCC than two engines.

Nominal delta LCC values are $M70 for three engines and $M130 for

four engines over the two-engine configuration. A single engine

shows the least LCC only for a burn reliability of 0.999 (by a

delta of nominally $M10) but not for burn reliabilities of 0.994 or

0.996.

These data lead toward the conclusion that across the selected

spectrum of single-engine burn reliability (0.994-0.999) and

nonindependent failure factor (0.00-0.05), two engines demonstrate

near maximum reliability and minimum or near minimum LCC. In

addition, in making a two-engine configuration fail

operational/fail safe for the return of a manned capsule, an amount

of subsequent propellent cost must be added to the two-engine case

LCC for contingency operation of a G02/GH2 ACS system. For

seven 13,000-lbm manned payloads in the 1994-2000 time period (Ref

6 Mission Model), approximately $20M must be added to the

two-engine case to make it fail operational/fail safe. This

appears to be an affordable method of providing a high level of

manrating reliability and flexibility along with remaining the

lowest LCC configuration for the performance of the Rev 6 Mission

Model.

Table 4.4-3 reflects the optimum number of engines for all

combinations of single-engine burn reliability and nonindependent

failure rate for the Aerojet engine data. The results and

conclusions are the same as for the Pratt & Whitney data.

4-104
Table 4.4-3

Aerojet Optimal LCC Propulsion Configuration

Nonindependent Failure Factor

Burn Reliability

0.994

0.996

0.999

0.00

2 Engines

2 Engines

1 Engine

0.03

2 Engines

2 Engines

1 Engine

0.04

2 Engines

2 Engines

1 Engine

In comparing cost element outputs from the LCC model, most cost

elements (including DDT&E, unit production, GSE, program

management, spares, scheduled/unscheduled maintenance, overhaul

labor, and ground control) show only minor delta cost differences

between the various engine configurations. However, two cost

elements contribute toward major cost deltas between the engine

configurations—propellant cost and mission loss cost. PropelIant

cost is directly related to the propellant usage to achieve a

15,000-lbf thrust for the respective engine configurations.

Mission loss cost is related to the propulsion system reliability,

which in turn is derived from engine burn reliability, number of

engines and the nonindependent failure factor.

For Pratt & Whitney engine data, a greater number of engines

results in higher propellant usage and therefore higher LCC.

Conversely mission loss cost is highest for a single engine and

generally lowest for two engines (depending on burn reliability and

nonindependent failure factor - see previous discussion on optimal

propulsion system reliability considerations). The costs

associated with these two elements tend to be the determining

factors in the relative costs of the respective engine

configurations. The mission loss costs for two, three, and four

engines are relatively low and balance out, whereas one engine

generally has a very high mission loss cost. The resultant

summation of mission loss cost and propellant cost across the

selected ranges of single-engine burn reliability and

nonindependent failure factor shows two engines to be the optimal

or near optimal choice.

For Aerojet engine data, the propellant cost is less of a cost

driver because of the reduced propellant consumption differences

between the engine configurations. In view of this, the primary

cost driver for the Aerojet propulsion system becomes the mission

loss cost. The two-engine configuration demonstrates least mission

loss cost for all but the single-engine configuration with a burn

reliability of 0.999.

4-105
LCC Model Sensitivities - Tables 4.4-4 and 4.4-5 summarize the

sensitivities of cost parameter variations on LCC for Pratt &

Whitney and Aerojet. The two-engine burn reliability of 0.996 and

a nonindependent failure factor of 0.03 was selected as a baseline

for developing the cost parameter sensitivities. , The sensitivities

are generally the same for other engine configurations although the

LCC "actual values" differ. Each parameter has been varied over

ranges as wide as might be conceived at this time. Figure 4.4-20

shows the sensitivities of LCC to missions per overhaul (average

number of missions between remove and replace for the vehicles'

subsystems).

Table 4.4-4 LCC Sensitivities Summary (Pratt & Whitney)

(for 2 engines)

Parameter Sensitivity

(LCC $M per -)

Propellant Cost $182M per $100/lb

IVA Cost $59M per $10k/h

EVA Cost $28M per $10k/h

Failures/Mission $11M per failure/mission

Missions/Year $23M per mission/yr

(not including propellant)

Missions/Overhaul $3.8M to $0.33M per mission/overhaul

Specific Impulse $11M per second of Isp

4-106
TABLE 4.4-5 LCC Sensitivities Summary (Aerojet)

(for 2 engines)

PARAMETER SENSITIVITY

(LCC $M per )

Propellant Cost $176M per $100/lb

IVA Cost $59M per $10k/hr

EVA Cost $28M per $10k/hr

Failures/Mission $9.3M per failure/mission

Missions/Year $23M per mission/yr

(not including propellant)

Missions/Overhaul $4M to $0.33M per mission/overhaul

Specific Impulse $11M per second of Isp

The sensitivities for the Aerojet data are almost identical to

those shown for the Pratt & Whitney data.

Propellant cost variations provide the greatest impact on OTV LCC

even for only minor propellant cost deltas. This cost element

tends to outweigh all other cost elements considered in this

analysis in both actual magnitude and relative LCC sensitivity.

Secondary cost sensitivities are IVA, EVA, specific impulse and

missions per year. IVA and EVA variations contribute to LCC

sensitivity because of the significant amount of IVA and EVA time

involved in mission turnaround activity. Specific impulse changes

contribute significantly to LCC sensitivity because of the direct

relationship between specific impulse and propellant usage.

Missions per year changes drive LCC because of the direct

relationship with propellant usage, EVA, IVA and all other cost

elements. Variations in failures per mission and missions between

major engine overhauls are shown to provide minimal impact on LCC.

4.4.3 Life-Cycle Cost Profiles

The two-engine Pratt & Whitney propulsion system configuration with.,

a single-engine burn reliability of 0.996 and a nonindependent

failure factor of 0.03 was selected as a baseline for developing

LCC cost profiles. Figure 4.4-21 shows the baseline cost profile

with operations and support (O&S) costs being the major contributor

4-107
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with 74% of the total LCC. The majority of this O&S cost is

attributable to $M2685 propellant cost. Propellant represents 67%

of the total LCC and has been removed from consideration in

subsequent cost profiles to allow closer examination of the

secondary cost drivers.

The most significant LCC element is DDT&E (Fig. 4.4-22) of which

engines (34%), propulsion (18%), and aerobrake (14%) contribute 66%

to this element.

Operations and support costs (excluding propellant) is profiled in

Figure 4.4-23 where scheduled maintenance (mission turnaround) and

ground control comprise the majority of this cost element.

The third and least contributor to the LCC is production cost (Fig.

4.4-24), which accounts for only 2% of the total LCC.

Figure 4.4-25 shows the tall poles, or largest of the secondary

contributors to LCC, in descending order of magnitude (propellant

being excluded).
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4.4.4 Level of Vehicle Modularization

Section 4.1.3 describes the process used in modularizing the

baseline vehicle concept. The assessment criteria used in grouping

the vehicle components into modules include module size,

accessibility, interface complexity, module functional similarity,

failure frequency, physical proximity, operational concerns, etc.

This section provides a rationale from a LCC standpoint to

substantiate the level of serviceability (modularity) selected and

discussed earlier.

As demonstrated by the LCC model outputs and sensitivities, the

orbiter payload transportation costs are a significant cost

driver. Although space station storage costs have not been

computed, it is reasonable to expect that this type of storage (for

spares and equipment) will be at a premium. A review of Table

4.4-6 indicates a reduced cost for level 3 over the level 1 or

level 5 for these two major cost drivers (items A and 8). Analysis

of the LCC model outputs and sensitivities indicated that items C

through G in Table 4.4-6 were not major cost drivers. A cost

comparison for items C through F in Table 4.4-6 marginally favors

level 1 over level 3 and substantially over level 5. In summary,

level 5 probably has the least life-cycle cost. Levels 3 and 5

offer an advantage over the level 1 in that they can incur multiple

failures. Further analysis is required to determine which modules

require the most support because of their higher failure rate

and/or criticality to quantify these tradeoffs. Additional

maintainability and design engineering will enhance the location of

these modules for easier EVA R/R actions. Module design permits

technology upgrade more readily than either full assembly design

(via cost) or the subcomponent design (via interfaces).

Based on estimates of the major cost drivers, Figure 4.4-26

illustrates where the probable minimization of LCC will occur with

regard to vehicle and engine modularization.

4.5 MAJOR STUDY CONCLUSIONS

Phase II results served to verify the results and conclusions

arrived at during the Phase I study effort (see section 3.5). This

applies to development of the baseline vehicle(s), engine

operational requirements, and design recommendations.

4-115
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A summary of the study conclusions for Phase II exclusively are

tabulated.

o LCC Sensitive to:

MPS Reliability (Mission Loss)

Vehicle Dry Weight

'

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Propellent Cost (amount scavengable, etc)

o LCC Relatively Insensitive to:

- Number of Missions/Overhaul

Failures/Mission

EVA and IVA Cost

o Two or Three Engines Recommended Because of:

Highest Reliability

- Minimum LCC - Across Estimated Range of Dependant Failure

Rates and Engine Reliability

Fail Operational/Fail Safe Capability

(affordable manrating criteria)

o Onorbit Servicing is Cost Effective—Proper Degree of

Modularity Required:

Engine: 2 to 8 modules

Vehicle: 20 to 100 modules

4-118
5.0 REFERENCES

1. D. A. Schmitt: Parametric Aerothermal Curves for

"Viking-Shaped" Aerobrake. Martin Marietta Memo 0484/CE-1917,

March 1982.

2. D. A. Schmitt: Base Heating for "Viking-Shaped" Aerobrake.

Martin Marietta Memo 0484/CE-1917A, May 1982.

3. J. J. Rehder: Multiple-Pass Trajectories for an Aeroassisted

Orbital Transfer Vehicle. AIAA paper 84-0407.

4. Preliminary Test Plan for AOTV with Fluid Aerospike. Martin

Marietta Denver Aerospace, February 1984.

5. Space Station Needs, Attributes, and Architectural Options

Study. Final Report, Vol. Ill Mission Requirements. Martin

Marietta Aerospace, NASA Contract NASW-3686.

5-1