N86-24733

A

(NASft-CS-178820) ~SPACE PiaiFCEH EIPEMDABIES

BESUPPL* COMCEPI DEFINITION STDDY. VOLOHE 1?

EXECUTE SMHAiX Beport, Mar. - Dec. 198^'

(Rockwell International Corp.) 36 p ' nr,^i = HC &03/HF AQ1 Dnclas

^

 /Sl-a0J CSC I _2_2B_G3/18 _ i» 29 3 7

SPACE PLATFORM EXPENDABLES RESUPPLY

CONCEPT DEFINITION STUDY

STS 85-0174

VOLUME I

EXECUTIVE SUMMARY

CONTRACT NAS8-35618

FOR PERIOD MARCH 1984 - DECEMBER 1984

206 5e
SPACE PLATFORM EXPENDABLES RESUPPLY

CONCEPT DEFINITION STUDY

STS 85-0174

VOLUME I

EXECUTIVE SUMMARY

CONTRACT NAS8-35618

FOR PERIOD MARCH 1984 - DECEMBER 1984

206 5e
This final report of the Space Platform Expendables Resupply Concept

Definition study was prepared by Ihe"Advanced Engineering organization of the

Space Transportation Systems Division of Rockwell International Corporation

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

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

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

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

into four documents:

Volume I - Executive Summary

Volume II - Study Results

Volume III - WBS and Dictionary

Volume IV - Cost Estimate

Personnel who made major contributions to the study effort are:

V. A. Weldon - Study Manager

R. M. Hayes - Effectiveness Analysis & Programmatics

R. J. Budica - Systems Engineering

S. N. Pace - Requirements

Y. Brayer - Requirements

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

D. R. Slocum - Business Management

W. A. McClure - 6EO Servicing Effectiveness Analysis

M. Pittman - Mission and Fluid Parametric Requirements

P. Forrester - Configuration Design

L. Livingston - Thermal Analysis

V. Mora - Propellant Feed Analysis

J. Nakamoto - Propellant Component Analysis •

J. J. Gernand - Related IR&D

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

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

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

2065e
Executive Summary Contents

Page

1.0 Introduction 1

2.0 Summary of Results 3

2.1 Major Conclusions 3

2.2 Selected Options 4

2.3 Recommended Approach 6

3.0 Study Results 8

3.1 System Requirements Analysis & Trades (Task 1) 8

3.2 Concept Definition (Task 2) 22

3.3 Programmatic and Developmental Planning (Task 3) 26

4.0 Recommendations for Technology Development 29

2065e . ii
~ Illustrations

Figure Page

1 Deleted

2 Study Flow 2

3 Selected Mission Scenarios 5

4 Resupply Module General Configuration (Without Servicer) 6

5 Resupply Module General Configuration (With Servicer) 7

6 Bi-Propellant/Helium Transfer Flight Demonstration Concept 7

7 Task 1 Activities 8

8 Low Earth Orbit Alternative Fluids Resupply Scenarios 12

9 Effectiveness Analysis Groundrules and Assumptions 13

10 Satellite Systems Availability Through Servicing 14

11 Net Economics Benefits Summary for NASA LEO ETR Spacecraft 15

12 Economic Benefits of Full Servicing Relative to Resupply Only 16

13 Lifetime Extension Options (10-15 yrs) Relative Annual Annuities 16

14 Expendables Resupply in GEO Potential Operating Modes 18

15 Storable Propellant Refueling from Orbiter OMS Pod Tanks and/or

Cargo Bay Tank 19

16 LEO Remote Refueling of Storable Stage Candidate Scenario 20

17 Top-Off of Storable Bi-Propellant Integral Propulsion Satellite 21

18 Potential LEO Top Off of DoD High Energy Orbit Integral

Storable Propulsion Satellites 21

19 Mid-Term Review DRM Recommendations 22

20 Bi-Propellant/Helium Transfer Flight Demonstration Concept 25

21 Resupply Module Flight Demonstration Program Supplier Schematic 25

22 Status of Technology Required for ERM Program 26

23 Development Test Schedule for Concept Feasibility 27

24 Technology Development Flight Demonstration, and Resupply 28

25 Cost Estimate for Technology Development and Flight

Demonstration Programs 29

2065e iii
Tables

Table Page

1 Fluid Resupply Mission Model 9

2 Fluid Parametric Requirements Summary 10

3 Candidate Spacecraft/Platform Selections 11

2065e iv
Acronym Dictionary

DRM Design Reference Mission

ERM Expendables Resupply Module

ETR Eastern Test Range

GEO Geosynchronous Orbit

GSE Ground Support Equipment

IP Integral Propoulsion

JSC Johnson Space Center

LEO Low Earth Orbit

MMH Monomethylhydrazine

MSFC Marshall Space Flight Center

NTO Nitrogentetroxide

OMS Orbital Maneuvering System

OMV Orbital Maneuvering Vehicle

OTV Orbit Transfer Vehicle

ROM Rough Order of Magnitude

SDI Strategic Defense Initiative

SPER Space Platform Expendables Resupply

USAF SD United States Air Force Space Division

206 5e
1.0 INTRODUCTION

S/013S"

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

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

operational efficiency and cost/effectiveness. The emerging Orbital

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

will introduce the capability for fluid resupply enabling satellite lifetime

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

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

1.1 Background

Numerous studies have been performed in recent years concerning the transfer

of fluids to satellites within the Orbiter payload bay. These studies have

mainly focused on the transfer of hydrazine (^H^). On a recent Shuttle

flight, transfer of hydrazine using man-in-the-loop was demonstrated within

the Orbiter payload bay. Also the Air Force is considering incorporation of

Interface requirements allowing on-orbit fluid transfer in future satellites

(with notable applications to SDI). It is apparent that the era of fluid

transfer on-orbit is emerging and that remote expendables resupply will becon

a recognized requirement for future satellites and/or propulsion modules.

1.2 Objectives and Scope

The overall objectives of the study are summarized as follows:

1. Develop Performance & Operational Requirements Task 1

associated with critical liquids, gases, and lubricants > * 60Z of study

resupply

2. Develop Design Requirements for the resupply module

(ERM) & equipment needed in conjunction with OKV to

perform remote fluid resupply functions

3. Develop resupply module Concept Designs (including

tankage, transfer systems & supporting mechanisms) &

associated spacecraft adaptation (user interface Task 2

emphasis) f - 30Z of study

A. Define a Flight Demonstration Program for automated

remote fluid resupply

5. Develop Program Planning Data, including cost, -\ Task 3

schedule & preliminary supporting development / " 10Z of study

program

2065e/ - 1 -
The study scope was initially limited to the definition of remote fluid

resupply concepts for satellites/spa'ce platforms; keeping in mind a- concurrent

servicing system for repair, maintenance, and module change-out. This scope

includes fluid transfer to propulsion modules for LEO and higher energy orbit

satellites/platforms (expendable or reusable), but not to the space station.

This assessment of resupply benefits, with and without other types of

servicing was intended to confirm a comprehensive justification for future

remote resupply operations.

1.3 Guidelines and Assumptions

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

study are summarized below:

o Utilize available data from related study/development activities,

including propellant transfer experiments, upper stage design,

operations studies, and space station/platform studies.

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

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

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

Shuttle, space station, or separate depot facilities,

o The study scope was initially limited to investigation of fluid

resupply; keeping in mind a concurrent servicing system for repair,

maintenance, and module change out.

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

spacecraft servicing, the impact of remote spacecraft servicing on

nominal fluid resupply modes will be considered,

o The proposed system must meet all manned safety criteria during initial

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

station after a resupply mission.

1.4 Study Flow and Methodology

The study tasks and their relationship to each others are shown on figure 2.

T««k1

1.1 Develop

Requirements

lor Resupply

.

Alternative

Resupply

Scenarios

1.3 Pnform

Effectiveness

Analysis

Rockwell

Input

Dan Base

• Benefits

• OHM'S

• Minion Opinions Requirements

• Svstero/Subsystem Rtqwnmnts

T«»k 2 ' ' *

Cmcaft Drfmitim

2.1 Develop

Resupply

Moduli

Design

2.2 Generate

Demonstration

Concept

• Design Concept

• System Concept

Task 31

• Verification

Approach

• Demonstration

Concept

PnwrwRMtk and

PlMininf

Figure 2. Study Flow

• Planning

• Cost

• Schedule

2065e/ - 2 -
Task 1, System Requirements, Analysis and Trades, was the focus of the

mid-term review. In subtask 1-.-1, -resupply requirements were generated in

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

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

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

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

were developed and screened for effectiveness. Mission operational and

system/subsystem requirements were next generated for these sleeted scenarios,

to provide data in support of the effectiveness analysis (and for the Task 2

design analysis after selection of the Design Reference Missions).

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

resupply module concept designs to meet the approved Design Reference Missions

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

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

DRM's in LEG and GEO as well as expanded benefits assessments. In addition

the definition of a flight demonstration program and a supporting technology

program was accomplished.

Finally Task 3 involved programmatic planning and cost/schedule data

development for the technology and flight demonstrations as well as the

initial operational segments of the total program.

2.0 SUMMARY OP RESULTS

2.1 Major Conclusions

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

economic benefit but resupply combined with servicing is much more

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

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

these satellites. Certain DoD satellites in high inclinatioin LEO orbits will

require relatively frequent expendables resupply to avoid costly replacement.

Since, it will likely be desirable to accomplish these resupply operations

without dependence on the orbiter the need for a LEO propellant storage depot

with a space-based OMV/ERM is emphasized.

The major commercial industry drivers for GEO resupply relate to satellite

revenues. They are a function of the total number of operating transponders

and the number of transponders per satellite, transportation costs, satellite

production costs, insurance costs, and technological obsolescence.

Lifetime extension of satellites through attitude control subsystem (ACS)

resupply is of marginal benefit without concurrent servicing of the satellite

electronics. This is due to the relatively rapid rate (assumed 12% annual) of

technological obsolescence. Planned lifetime extension where the satellites

reliability is increased (and also the initial satellite cost) turns out to be

the least attractive scenario because of the increase up front development and

production costs. Additionally contingency resupply of satellites which have

depleted their fuel from malfunctions is always cost effective, if the problem

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

of the revenue stream can have the greatest impact.

Though adding additional transponders at the expense of ACS fuel increases the

total revenue from a satellite, it is not cost effective primarily because of

the risk involved in assuring future operation. Less onboard fuel will

require more frequent (Optimally 2 in a 10 year period) resupply missions

2065e/ - 3 -
increasing the risk to the satellite revenue stream. Insurance premiums of

less than 15% are required to make this resupply case viable.

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

receiver spacecraft. The impact is primarily required interface provisions.

The ullage exchange propellant transfer process has the inherent versatility

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

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

Servicing large satellites is generally more cost effective than smaller

satellites because large satellites house more transponders which may be

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

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

If a Comsat undergoes a low rate of revenue loss due to technological

obsolesence there is less incentive to resupply and upgrade it. If this rate

can be held to 6% or lower it is unlikely that resupply/servicing will be

beneficial unless a revenue restoration level greater than 100$ can be

realized. At a 12% obsolesence rate a revenue restoration of 70% is required

to assure a positive impact to the satellite revenue stream.

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

by exploiting the synergysm with the NASA-JSC Orbiter in-bay earth storable

propellant tanker program. Although the JSC program is being configured for

different functional requirements and missions, it will potentially result in

certain common technology development or modifiable hardware. Such hardware

could include interface components, Mission Peculiar Equipment Support

Structure (MPESS) (as tankage support for in-bay tanker and resupply module

flight demonstration), and propellant pumps..

2.2 Selected Options

Ullage exchange was selected as the best option for the ^O^MMH fluid

transfer process. This approach is applicable to all potential receiver

propulsion subsystem and acquisition types through appropriate modifications.

It minimizes pressurant resupply requirements, involves no adiabatic

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

scavenging, and provides constant pressure resupply.

Bi-propellant N204/MMH fluids were chosen for the resupply module since

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

subsystem, and they are scavengable from orbiter QMS tanks.

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

which were selected as the design reference missions. Scenarios 1 and 2 are

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

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

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

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

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

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

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

transferred from the orbiter to the depot through the OMV.

A propellant depot becomes most attractive if one uses the basic resupply

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

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

development costs for the depot.

2065e/ - 4 -
Propellant top-off capability in LEO via remote resupply would allow launch in

the STS of up to four satellites, with one or more in an off-loaded condition,

this could significantly exceed the current payload capability of the

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

applicable to currently projected larger DOD satellites which if launched

fully loaded, would exceed STS lift capacity.

An additional application of the ERM would be to fuel a high performance

reusable perigee stage being studied at MeDonneU Douglas. This stage would be

capable of launching a 12K lb payload into GEO after fuelling by a resupply

module•

CONTINGENCY MULTIPLE S/C

RESUPPLY FROM OMV •

POTENTIAL

MISSION

SCENARIO

PLANNED RESUPPLY/NODULE

CHANGE-OUT FROM OMV -

PROPELLANT TRANSFER

FROM RM TO OMV FOR.MORE

ECONOMICAL TRANSFER TO GEO

DEPOT OF

"MOO RM'S

REFUEL REUSABLE

PERIGEE STAGE

\ERVICER

TOP-OFF

COMPACT INTEGRAL

PROPUL S/C

PRECURSOR I

REUSABLE STAG9

OPERATE OMV/RM

FROM LEO DEPOT

FOR PROPELLANT

TRANSFER OPERATIONS

RM/OMV USE FOR

PROPELLANT TRANSFER

FROM STS TO DEPOT

s/c :

J

REFUEL

SPACE-BASED

OMV

Figure 3. Selected Mission Scenarios

2065e/ - 5 -
2.3 Recommended Approach

ORIGINAL PAGE 83

OF POOR QUALmf

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

presented in two configurations; without and with the satellite servicer.

This concept was selected from among several competing approaches primarily on

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

hardware providng low development cost. The resupply module is supported at

its forward end by an existing IUS upper stage forward cradle. This cradle

includes load equalization capability which reduces the structural redundancy

between the resupply module and the orbiter. It also allows a minimum of

weight impact on the resupply module for attachment to the orbiter payload bay

longerons and keel at its forward end. The resupply module uses six stretched

QMS tanks with a modified ullage positioning propellant management device

(PMD) for ullage bubble position control.

r

Transfer Pumps

STS Orbiter MPS Hi-Press

Helium Tanks (12 max.)

'Prop/He Tank forward —

Forward Umbilical •

(moveable)

End Effector

Oxid Transfer

Module

Cnbilieal Panel

(Fixed)

Grapple Fixture

Figure 4. Resupply Module General Configuration (Without Servicer)

2065e/ - 6 -
Explosive Bolt

iHelium Resupply Tanks

(on RM-Extended .

Mission)

End Effector and

Umbilical (OMV)

Satellite

Senricer

Forward

ASE Cradle

OMV

Aft Keel Pin

X 1135.87 (RM/OMV Mt 32,000 Ibs.)

o

Figure 5« Resupply Module General Configuration (With Servicer)

The selected flight demonstraton approach uses hardware developed through a

low-cost Technology Validation Program and many off-the-shelf components in

conjunction with a low-cost structural framework. Two modules, both of nearly

identical configuration, would be docked together in the payload bay using the

RMS. Subsequently, storable bipropellants plus helium transfer would be

demonstrated without man-in-the-loop. One module would represent the receiver

spacecraft and the other the resupply module. Most of the components, except

the main tankage (because of its small size) would later be used for

operational resupply units to allow significant program cost savings. The

selected flight demonstration concept is depicted in Figure 6.

SPAS OR MPESS STRUCTURE FOR

RECEIVER AND SUPPLIER

SIMULATORS

RMS USED TO DOCK

RECEIVER TO SUPPLIER IN

PAYLOAD BAY

Figure 6.

2065e/

MOST OF DEMO SUBSYSTEMS LATER

USED FOR OPERATIONAL FLIGHT

UNITS FOR COST SAVINGS

Bi-Propellant/Helium Transfer Flight Demonstration Concept

- 7 -
3.0 STUDY RESULTS

This section presents the analyses associated with the Space Platform

Expendables Resupply Concept Definition Study.

3.1 System Requirements Analysis and Trades (Task 1)

The overall activities of the system requirements analysis and trade studies

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

accomplishments.

TASK1

MID-TERM REVIEW

V

SYSTEM REQUIREMENTS

ANALYSES & TRADES

1.1 DEVELOP

REQUIREMENTS FOR

RESUPPLY

1.1.1 GENERATE RESUPPIY

MISSION MODEL

1.1.2 DERIVE RESUPPLY

PARAMETRIC

REQUIREMENTS

1.1.3 SELECT CANDIDATE

SPACECRAFT 4

PLATFORMS

1.2 DEVELOP ALTERNATIVE RESUPPLY

SCENARIOS

1.2.1 GENERATE ALTERNATE RESUPPLY

SCENARIOS

1.2.2 DERIVE MISSION OPER

REQUIREMENTS

1.2.3 SPECIFY SYSTEM/SUBSYS

REQUIREMENTS

1.2.4 IDENTIFY INTERFACE

REQUIREMENTS

t

1.3 PERFORM EFFECTIVENESS ANALYSIS

1.3.1 FORMULATE METHODOLOGY &

CRITERIA

1.3.2 GENERATE EFFECTIVENESS DATA

1.3.3 SELECT BASELINE DRMl

RENEFTTI

• MISSION OPERATIONS REQUIREMENTS

• SVSTEM/SUISYSTEM REQUIREMENTS

• INTERFACE REQUIREMENTS

Figure 7. Task 1 Activities

Mission Model

A detailed mission model of fluids resupply was developed in Activity 1.1.1.

The model, shown in Table 1, was developed from earlier projections of OMV

servicing tasks and estimates of future STS launch rates. All spacecraft that

could take advantage of fluid resupply were included. The actual selection of

Design Reference Missions was accomplished later in the study.

2065e/ - 8 -
MISSION LOCATION 1989 1990 1991 1992 1993 1994 1995 1996 1997 1998 1999 2000 2001 COMMENT

GRO

AXAF

EUVE

CREP

PROTEUS ni

LDR

SIRTF

GRAV PROBE-B

EOS

MPS <*>

GEO PLATFORM

CONTINGENCY

CIVILIAN TOTAL

NMI

227

320

324

216?

216

486

486

520

380

320

19323

TBD

DEG

28.5 D

28.5

28.5 D

28.5

28.5

28.5

98

90

99.8

28.5

0

TBD

1 1

D

1

D 1

D D 1 1

D

D

135 6

1

m

256 9

1

1

D

6

1

9

1

1

D

1

1

5

1

10

1

2

3.

6

1 1

111 1

1

1 1

2

1

1 1

524 3

1

1

1

3

3 AT ORBITER

3

1 AT ORBITER

1

10, PLATFORMS

2

4

1

4, PLATFORMS

30, PLATFORMS

EXPT SERVICING

20% OF OMV

CONTINGENCIES

NEED FLUIDS

64

DOD 3 47

NOMINAL TOTAL

CUMULATIVE

POTENTIAL LINE

POTENTIAL TOTAL

CUMULATIVE

4

4

4

4

8

12

8

12

8

20

8

20

13

33

13

33

14

47

14

47

17

64

3

20

67

7 12

71 83

10 7

17 19

84 103

4 8

87 95

8 12

12 20

115 135

10 6

105 111

8 12

18 18

153 171

111

60

171

(1) - FIRST MISSION MAN-TENDED AT ORBITER

(2) - DONE AT ORBITER MAN-TENDED

D -DEPLOYMENT

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

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

Table 1. Fluid Resupply Mission Model

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

were based on discussions with individual program offices. Such discussions

covered estimates of spacecraft technological obsolescence and the value of

extended time on-orbit. Classified data on the make up of DoD missions is not

available under this contract.

Due to the conceptual nature of many of the candidate missions, particularly

those in the late 1990's, a number of assumptions were made in developing

representative requirements. Those assumptions are:

o Hydrazine propulsion using a single blowdown system was assumed for

those missions requiring propulsion, but which had no spacecraft

design baseline. Though none of the current candidate satellites

utilize bi-propellants or require resupply of pressurants, future

spacecraft designs could include these as an option for trade analyses.

o Once rendezvous with candidate spacecraft has been accomplished, one

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

(includes actual transfer and associated operations).

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

program cost when the actual replacement costs could not be determined.

2065e/ - 9 -
o Mission operational lifetime was extended beyond the program baseline

if- it would be technically or scientifically advantageous and

consumables resupply was the primary constraint.

o A hybrid cryogenic system was assumed for the Large Deployable

Reflector (LDR) to determine consumable quantities (e.g., a

combination closed dewar and passive radiator system).

o Candidate missions scheduled prior to an operational OMV were assumed

to incorporate integral propulsion or use a spacecraft bus to

rendezvous with the orbiter for resupply.

Fluid Transfer Parameteric Requirements

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

annual fluid usage rates required by the resupply mission model. Quantities

of associated pressurants such as He and N2 were required in relatively

minor amounts. In some cases, depending on the receiver propulsion subsystem

used, no additional pressurants are required in the transfer process.

As expected, hydrazine and storable bi-propellents are the primary fluids to

be carried. In addition, another significant fluid of interest was found to

be liquid (primarily superfluid state) helium. Water, as used in MPS

factories, is more suited to module replacement due to the need to retrieve

products. The results of the fluid transfer parametric requirements analysis

are shown in Table 2.

USER AND

CONSUMABLE TYPE

NASA

HYDRAZINE

LIQUID HELIUM

WATER

Xe-METHANE MIXTURE

COMMERCIAL

(POTENTIAL)

BI-PROP (N204/MMH)

HYDRAZINE

DOD*

BI-PROP (N204/A-50)

HYDRAZINE

LIQUID HELIUM

DOD* (POTENTIAL)

BI-PROP (N204/A-50)

HYDRAZINE

CONTINGENCY"

1989

_

_

_

—

——

_

_

—

—_

—

1990

•40_0

5,000

—

——

14,00_0

—

—_

—

1991

3,83_0

15.000

—

——

14,000

70

—

—_

1,500

1992

1,100_

25,000

—

——

7,000

70

—

—

——

1993

5,430

160

30,000

30

——

21,000

70

—

—_

—

RESUPP

1994

1,100_

30,000

85

——

14,000

17,470

2,400

_—

1,500

LYAMOI

1995

6.694

1,389

25,000

—

818

—

28,000

17,470

2,400

8,000

^

—

JNT(LB

1996

5,100

15,000

—

2,482

500

_

70

15,000

950

—

1997

1,100

992

5,000

85

2,307

1,000

35.000

8.770

1,200

2,000

250

1,500

1998

1,100

1.320

—

3.318

500

7,000

70

1,000

600

—

1999

5,100

992

—

4,694

500

7,000

17,470

2,400

15,000

950

—

2000

1,100

_

85

2,414

500

28,000

17,470

2,400

2,000

600

1,500

2001

1,100

2,312

—

4,694

500

14,000

70

—

14,000

950

—

•SMALL QUANTITIES OF PRESSURANTS ALSO REQUIRED FOR SOME SATELLITES - IMPORTANT FOR DESIGN REQT

•ESTIMATE OF PROPELLANT RESUPPLT REQUIRED IN CONTINGENCY SITUATIONS

• HYDRAZINE USED FOR THE MOST NUMBER OF UMBILICAL ENGAGEMENTS

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

• WATER & SPECIAL GASES ARE TRANSFERRED VIA MODULE CHANGEOUT

Table 2. Fluid Parametric Requirements Summary

2065e/ - 10 -
Candidate Spacecraft and Resupply Scenarios and Trade Studies

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

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

resupply scenarios and mission operations requirements. The selection of a

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

selection of missions which included the technical performance requirements of

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

second step was a screening for missions most likely to benefit from

resupply. The four general mission characteristics used are shown in Table 3.

GENERIC SPACECRAFT CLASS

• SCIENTIFIC EXPLORERS

• ASTRONOMIC OBSERVATORIES

• EARTH/WEATHER OBSERVATION

• RECON/SURVBLLANCE

• NAVIBATION

• COMMUNICATION-OEUA

• ENVIRONMENTAL OBSERVATION

• EARLY WARNINB ODD

• COMMUNICATION-IUS/TOS

• COMPLAT

• SMALL PLATFORMS

EUVE

CREP

GRO

AXAF

LANDSAT

OMSP

GPS

H.S. 376

GOES

TORS

INTEL VI

PLANNED

PROTEUS

LEASE

CRAFT

MISSION OBJECTIVE

DURATION

SHORT 6 MO-24 MO

CONTINUOUS

CONTINUOUS

CONTINUOUS

CONTINUOUS

CONTINUOUS

CONTINUOUS

CONTINUOUS

CONTINUOUS

CONTINUOUS

MEDIUM 120 MO

UNJI ON-ORBIT VALUE

SMM SPACECRAFT

VERY LOW $50-100

VERY HIGH $500-1200

LOW-HIGH $80-300

HIGH-VERY HIGH

VERY LOW $50-100

LOW $75-125

7

7

MODERATE-HIGH $250-300

VERY HIGH

MODERATE $100-200

DURATION CONSTRAINT

(HISTORIC)

EXPERIMENT, FLUIDS

EXPERIMENT, INSTRUMENT

SENSORS. POWER

UNKNOWN

TECHNOLOGY OBSOLETE

7

7

TECHNOLOGY OBSOLETE

DESIGN REQUIREMENTS

CHAN6EOUT/INVENTORY

COST OF ACCESS

(STS|

VERY LOW-MODERATE

LOW

MODERATE

MODERATE

HIGH

VERY HIGH

VERY HIGH

VERY HIGH

VERY HIGH

VERY HIGH

VERY LOW-MODERATE

LOW INCLINATION LEO

•AXAF

•MRS

•GRO

• PROTEUS

HIGH INCLINATION LEO

•SIRTF

• DMSP

•DOD-1

GEO

• GEO PLATFORM

• COMSATS

VARIOUS SINGLE & MULTI-SERVICING SCENARIO

COMBINATIONS EXAMINED FOR THESE CANDIDATES

Table 3. Candidate Spacecraft/Platform Selections

Nine candidate spacecraft/platforms were selected for scenario development.

They occupied the three general resupply locatons of low-inclination LEO,

high-inclination LEO, and GEO. Various single and multi-servicing scenario

combinations were then examined for these candidates. All of the candidates

either require resupply or benefit highly from its use.

2065e/ - 11 -
ORIGINAL PAGE IS

OF POOR QUALFTY

Using as input data the candidate spacecraft and space platforms, a

comprehensive set ,of resupply scenarios was generated. Among the scenarios

were some non-remote resupply options. In addition, certain scenario trade

issues were addressed such as multiple versus single spacecraft configurations

and wet versus dry launch. This data was used in the concurrent effectiveness

anaysis to select the most promising scenarios for further definition.

Mission operatonal requirements were derived for these scenario (and used) to

develop system/subsystem/interface requirements for both the effectiveness

analysis and the Design Reference Missions (DEM). The system/subsystem

requirements data were also used to develop cost/risk benefits for the

effectiveness analysis to support justification of the DRM's.

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

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

orbiter-tended operations. A spacecraft in LEO can be resupplied using a

resupply module with an OMV. Each of the remote supply options may perform

multiple or single engagements. The spacecraft can also be resupplied using

in-bay equipment being developed at JSC. Delivering the spacecraft to the

orbiter bay can be accomplished using the spacecraft's own integral propulsion

system (IP) or the orbiter1s OMV capability.

• SCOPED TO REMOTE vs

ORBITER-BASED

SPACECRAFT

RETURNS

FOR EVA

ORBITER TO

SPACECRAFT

FOR EVA

OMV, n>1

_, i

OMV GROUND

OMV, n = 1

OMV, n>1

OMV, n = 1

OMV,,n>1

OMVan = 1

BA

j

i

ON

SED

L21

L

REN

RES

V

SPACE

BASED

(4)

2 r

1

 _J

. t

k

OTE

UPPLY

(6)

t

SPACE

RESUPPLY

(8)

2) IN-BAY IP

T) IN-BAY OMS

IN-BAY

RESUPPLY

1'

Figure 8. Low Earth Orbit Alternative Fluids Resupply Scenarios

2065e/ - 12 -
Six trade studies were identified in the original proposal for evaluating

resupply scenarios. The first examined the utility of nodal regression

effects in multiple engagement mission planning. The second examined

alternative space-basing options for the resupply module relative to

ground-basing. This task also used much of the material developed in the

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

cons of dry and wet launches of the resupply module itself. The fourth study

was oriented toward GEO operations and examined the weight sensitivities

involved with expending versus reusing elements of a GEO resupply system (OTV,

OMV, ERM). Cost factors were also identified for use by the effectiveness

analysis. The final two studies supported development of

system/subsystem/interface requirements. Issues as to the appropriate levels

of autonomy and redundancy for the resupply module were addressed. Both the

OMV's capabilities and the potential role of man-in-the-loop were evaluated.

Finally, critical issues in fluid transfer technologies concerning the

required fluids were identified. Superfluid Helium was found to have a

variety of issues associated with its transfer and thus would require closer

study.

Effectiveness Analysis Summary

The groundrules and asumptions for the effectiveness analysis are

summarized in Figure 9.

• ALL ECONOMIC ft COST ESTIMATES IN CONSTANT 1984 S MILLIONS

• PARAMETRIC ESTIMATES ARE ROUGH ORDER OF MAGNITUDE (ROM)

• USAF SD UNMANNED SPACECRAFT COST MODEL CERs USED

• OMV ACQUISITION (MSFC BASELINE CONFIGURATION)

• EXPENDABLES RESUPPLY MODULE (ERM) ACQUISITION

• CANDIDATE SPACECRAFT/PLATFORM ACQUISITION

• ROCKWELL ORBITER DATA BASE CERs & INDEPENDENT ESTIMATES

• LAUNCH (STS) ft INSERTION STAGES

• FLIGHT ft MISSION OPERATIONS/SUPPORT

• RISK ft COMPLEXITY ANALYTIC HIERARCHY PROCESS

BEST DESIGN REFERENCE MISSION DECISION RULE: SELECT CANDIDATE

WITH HIGHEST OVERALL CONTRIBUTION TO CONFLICTING GOALS

Figure 9. Effectiveness Analysis Groundrules and Assumptions

2065e/ - 13 -
LEO Benefits

As shown in Figure 10, expendables resupply capability is a necessary

condition for enabling satellite system availability through on-orbit

maintenance (which offers substantial cost efficiencies compared to the

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

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

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

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

depletion of its on-board consumables supply.

System Availability Mission Accomplishment

SATELLITE SYSTEM

AVAILABILITY

PERCENT FUNCTIONAL

CAPABILITY

A COSTS INCURRED

• ADDED PRODUCTION COSTS

• EXTRA TRANSPORT COSTS

ACHIEVEMENT OF

MISSION GOALS

SATELLITE

MAINTAINABILITY

t)

ACOSTS INCURRED

• REDUNDANCY

• QUALITY

• SERVICING CAPABILITY

• SATELLITES DESIGNED

FOR SERVICING

1

1

RESUPPLY

1

REPAIR

REFURBISH

1

TECHNOLOGY

UPGRADE

1

MAINTENANCE

Figure 10. Satellite System Availability through Servicing

2065e/
Results of the parametric economic benefits analyses shown in Figure 11

indicate that NASA LEO ETR Spacecraft programs could receive a net economic

benefit~by~utilizing an Expendables Resupply capability, even without

concurrent "full servicing". However, the marginal benefit to these programs

received by taking advantage of comprehensive repair, instrument changeout,

and fluids resupply is so substantial (nearly an order of magnitude

improvement in net benefits) that there should be a strong motivation to opt

for orbiter-based, man-in-loop repair until teleoperation maintenance is

possible.

RESUPPLY (W/0 CONCURRENT MAINTENANCE & REPAIR) 0 NET BENEFIT

V

7 OBSERVATORIES & EXPLORERS

• HIGH UNIT-VALUE SATELLITES

• LOW COST ACCESS FOR RESUPPLY

V COST-EFFECTIVE ONCE PER PROGRAM

SPACECRAFT

AXAF

EUVE

GRO

LOR

PROTEUS

GROSS ECON

BENEFIT

SRI84

RESUPPLY ONLY

73-110

23-34

65-98

144-216

27-40

COST TO

RESUPPLY

Sffl84

6-20

- 7-21

15-29

9-43

9-23

NET ECON

BENEFIT

SM84

RESUPPLY

53-104

2-27

36-83

101-207

4-31

196-452

ONE

RESUPPLY

RESUPPLY WITH CONCURRENT MAINTENANCE & REPAIR® NET BENEFIT

ORDER OF MAGNITUDE

INCREASE IN BENEFITS

"FULL SERVICE" WILL

BE PREFERRED OPTION

SPACECRAFT

AXAF

EUVE

GRO

LOR

PROTEUS

GROSS

ECON

BENEFIT

SH84

219-329

29-43

220-330

256-384

36-54

COST TO

RESUPPLT &

REPAIR

Sffl84

23-80

8-14

32-87

31-119

10-17

NET BENEFIT

PER REPAIR

MISSION

$H84

139-306

15-35

133-298

137-353

19-44

NO

ENGAGEMENTS

PER

PROGRAM

4-6

1-2

2-4

2-3

3-4

•

TOTAL PROGRAM

NET ECONOMIC

BENEFIT

SR84

558-1836

15-70

267-1194

275-1060

56-177

1171-4337

MULTIPLE

ENGAGEMENTS

Figure 11. Net Economics Benefits Summary for

NASA LEO ETR Spacecraft

The value of servicing with resupply is limited to the spacecraft's

replacement cost (second unit + replacement launch and insertion). Still the

net economic benefits received from full servicing in major NASA ETR programs

far exceed the benefits of expendables resupply alone.

Since stand-alone resupply can be cost-effective only once during most

programs, while on-orbit maintenance and repair can be cost-effective many

times, the total benefits available from full servicing should drive most

program offices to orbiter-based servicing until the advent of teleoperator

in-situ servicing.

Figure 12 graphically presents the results of Figure 11 in the form of columns

to more clearly illustrate the relative benefits of the two approaches.

2065e/ - 15 -
3

Z

5

o

u

400

300

200

100

NET ECONOMIC BENEFIT TO

NASA LEO PROGRAMS

PER ENGAGEMENT

FULL SERVICING

• RESUPPLY

• REPAIR/MAINTENANCE

• TECHNOLOGY UPGRADE

TOTAL BENEFIT 5 PROGRAMS

FULL SERVICING

•RESUPPLY

• REPAIR/MAINTENANCE

• TECHNOLOGY UPGRADE

Figure 12. Economic Benefits of Pull Servicing Relative to

Resupply Only

GEO Benefits

400 . o—o

S 300 .

200 .

100

• 100

GEOSAT VC1GHT (THOUSWCS Of LBS)

Figure 13- Lifetime Extension Options (10-15 Yrs) Relative Annual Annunities

2065e/ - 16 -
Figure 13 depicts the result of an economic analysis performed to compare the

relative annual annuities of a spacecraft which is resupplied and

technologically upgraded (life incrase of 50%) versus a conventional ten-year

life system (baseline). Rockwell studies indicate that the total weight

penalty to enable resupply/concurrent module change-out servicing can be less

than 10%. Even for this level of penalty, it is seen that as the mass of the

spacecraft increases the economic advantage of resupply/servicing relative to

the baseline approach becomes correspondingly greater.

The financial model used in the study was shown to produce stable results

across a wide range of potential changes in key baseline numbers. Several

implications for the development of GEO Resupply/Servicing can be drawn:

o Insurance rates have a high impact on the economic viability of

resupply/servicing operations for satcom's. Rates of 15% or lower are

required to make GEO operations beneficial. Based on historical

experience, an overall confidence level of 95% or better is necessary

for launch through competion of GEO operations.

o The cost/lb. figure for GEO resupply/servicing has a relatively minor

influence on optimum servicing intervals and on the satcom's annual

annuity value. Resupply/Servicing scenarios only need to keep their

cost/lb. figures below about $50K/lb. to be economically viable. This

includes not only the costs of getting to GEO, but associated usage

fees (e.g., OMV, OTV, Servicer Kit, etc.).

o If satcom revenue decline rates average less than 6% over their

lifetime, resupply/servicing is unlikely to be attractive.

Obsolescence rates of about 9% and above make technological upgrading

attractive. The key questions here are the technical feasibility of

revenue restoration with upgrading and what combinations of space and

ground segment modifications are required.

Further conclusions drawn for GEO satcom Remote Resupply include:

1) GEO Resupply without servicing offers a major benefit for contingency

operations, including multiple resupply in GEO of several satellites

from one STS launch. Lack of concurrent servicing, however, results

in little benefit to resupplying GEO satellites. This is primarily

due to the rapid (12% annual) rate of technological obsolescence.

2) The cost effectiveness of GEO operations increases by servicing larger

satcoms. This analysis is, therefore, most applicable to large

concentrations of capital as would occur with satellite clusters or a

GEO platform. Lifetime extension is cost-effective with satcoms that

are still functional at the end of their lives and only requiring

additional ACS fuel. This, however, is quite expensive and is not

seen as desirable without technological upgrading during the

satellite's life.

2065e/ - 17 -
3) Transponder replacement incurs increased satellite costs and insurance

risks which are not completely covered by the. increased

revenue-generation capability.

4) The major unknown remaining from these studies are the actual costs of

incorporating a spacecraft resupply subsystem into a communications

satellite. Parametric estimates have shown that the weight/cost

penalties associated with resupply and servicing are likely to be much

less than the costs of providing for longer life without resupply and

servicing.

The more satellites that can be resupplled from one flight, the greater the

economic benefits. If an operational mode can be implemented before the

advent of reusable OTV operations the most economical means of accomplishing

the transfer to 6EO would be to use the resupply module as a propellant tanker

for the OMV (rather than expending a relatively expensive upper stage such as

the Centaur). The propellant transfer would be made directly from the OMV

after detaching from the resupply module (Figure

WITHOUT CONCURRENT SERVICING

• GEO RESUPPLY OF

BI-PROPELLANT FOR

LIFETIME EXTENSION OF

MULTIPLE SATELLITES

RESUPPLY DIRECTLY

FROM OMV

EXCEPT FOR VERY FAR TERM POTENTIAL

MISSION SHOWN ABOVE RESUPPLY IN GEO

REQUIRES CONCURRENT SERVICING FOR

JUSTIFICATION ,

WITH CONCURRENT SERVICING

• GEO RESUPPLY OF

BI-PROPELLANT & HELIUM

PRESSURANT TO PLATFORM FOR

MAJOR LIFETIME EXTENSION

RESUPPLY &

MODULE EXCHANGE

TRANSPONDER

MAXIMIZATION THROUGH

ON-ORBIT PROPELLANT

OFFLOAD

Figure 14. Expendables Resupply in GEO Potential

Operating Modes

In this mode the servicer is attached to the front end of the OMV. Thus a

truss structure separates the resupply module from the OMV to allow sufficient

volume for the servicer. For this case the resupply (probably also including

Helium) is directly from the OMV after separation from the resupply module.

2065e/ - 18 -
Communication satellite transponder maximization can be accomplished by ACS

propellant. off-loading. _ If the same strategy described above is employed in

conunction with technology up-dating at about the mid-point of the satellite

( design life, significant economic benefits can result. This mode has the

x potential to occur sooner than those described above, perhaps well before the

availability of a space-based reusable OTV.

This study was conducted in coordination with related studies at Rockwell.

One of these, NAS9-16994, "STS Propellant Scavenging System Study", performed

for NASA-JSC in FY 84 and FY 85 has identified the potential for about 1.4

million pounds of storable bi-propellant (NTO/MMH) which could be delivered to

28° inclination LEO by the Orbiter throughout the 1990's at very low cost

(117$/lb). Most Orblters will not be fully loaded with payload, and therefore

QMS pod and payload bay tankage would be available for proepllant storage and

scavenging. Additional work performed under IR&D has identified relatively

simple low-cost provisions by which the Orbiter can transfer propellants to

the OMV/ERM (Figure 15). This propellant could then be transferred to a depot

for later support of remote resupply operations in LEO.

IMUfl

1 TANK SIZE

SCAVENON6

FUBHTS

165

TAWABE

S1.I

COSTS (S M)

ELECTRICAl

14.4

INSnUHEMTATKM

3.4

mn

•on

40

OPS

liiWkm

41.3

OMV

nupELUurr

12.6

TOTAL

163.5

DELIVERED

PROPELLANT

(1MO LI)

1,409

PHOPELUUfT

COSTS

(»/LI)

1"

• IDENTIRED UNDER CONTRACT

• DEVELOPED UNDER IR&D

• BASELINE MEETS ALL REQUIREMENTS:

• MMH, NTO, & HELIUM TRANSFERRED

FROM QMS SYSTEM TO OMV

• NO VENTING OR CONTAMINATION

OF ENVIRONMENT

ASE RESUPPLY UMBILICAL

OXIDIZER 00 PANa

END EFFECTOR

ASE UMBILICAL

STOWED POSITION

424.00

OMV

ASE

MOUNTED ON

BULKHEAD 1307

END VIEW OF

ASE UMBILICAL

VIEW LOOKING AR AT X01214.50

& AT BULKHEAD X01307.00

ASE UMBILICAL

EXTENDED POSITION

COUPLING ENGAGEMENT

MECHANISM

Figure 15. Storable Propellant Refueling from Orbiter OMS Pod Tanks

and/or Cargo Bay Tank

2065e/ - 19 -
ORIGINAL PAGE S3

OF POOR QUAIJIY

The refueling scenario depicted in Figure 16 depicts a spacecraft being

refueled by an OMV/ERM operating from a LEO depot. The depot would be

supplied via storable propellant scavenging from many orbiter flights.

Implementation of this concept could given economic benefits similar to those

achievable by a Space Station-based OTV.

REPRESENTATIVE INTEGRATED

LEO SPACE OPERATIONS FOR

ECONOMIC BENEFITS & PRECURSOR

CAPABILITY FOR SPACE STATION

SPACECRAFT MATED

TO PROPULSION MODULE

IN LEO USING ORBITER

TENDING STORABLE PROPELLANT DEPOT

IN LEO FILLED FROM ORBITER

OMS POD VIA OMV/RM 1.4M LB

AVAILABLE IN 1990s AT S117/LB

OMV/RM FOR PROPELLANT

TRANSFER TO UPPER

STAGE IN LEO

ORBITER BRINGS SPACECRAFT TO

LEO WITHOUT PROPULSION MODULE.

OMS PROPELLANT SCAVENGED FROM

ORBITER & TRANSFERRED TO

DEPOT BY OMV/RM

REUSABLE STORABLE

PROPULSION MODULE

COURTESY MCDONNELL DOUGLAS CORP.

Figure 16. LEO Remote Refueling of Storable Stage Candidate Scenario

Figure 17 depicts a single satellite launch case for remote resupply in LEO

allowing economic benefits to spacecraft subsequently transferred to 6EO.

This approach is based on the reconfiguration of a single relatively heavy

integral propulsion satcom to take advantage of the full payload bay

diameter. Currently, except for LEASAT (Syncom IV - about 3,000 Ib. BOL mass

in GEO) communication satellites are limited to about 12 feet in diameter so

as to be Ariane compatible. This approach does not cause an STS

transportation cost penalty because installed weight, rather than length,

would determine the cost if a shorter installation was provided. The use of

propellant top-off in LEO through remote resupply offers the possibility of

achieving significant cost-savings through the use of a shorter installation

in the payload bay.

2065e/ - 20 -
BI-PROPELLANT

LEO DEPOT

/FILLED FROM

'

INCLUDING

APOGEE

ELEMENT

1

I

TRANSFER

COMBINATION

T06EO

DEPLOY

COMPACT SATELLITE

DESIGN TAKES FULL

ADVANTAGE OF PL

BAT OIA. PROPELLANT

OFF-LOAD FOR LENGTH

DRIVEN TRANSPORTATION

COST TO LEO

REPRESENTATIVE BENEFIT:

COULD REDUCE COST OF

DEPLOYING 8K LB BOL

MASS S/C TO GEO BY

JttM (

Figure 17- Top-Off of Storable Bi-Propellant Integral Propulsion Satellite

in LEO

Our Satellite System Division provided support to the remote resupply study

through presenting a satellite supplier's viewpoint. They assisted in

estimating spacecraft impacts for fluid resupply/module change-out servicing

and helped to identify potential applications for remote resupply* One

suggestion was that it might be possible to reduce STS payload weight by

off-loading fuel before launch and topping-off in LEO. For example, one

specific DOD satellite under study (including maneuver propellant) is too

heavy for the payload bay if four satellites are deployed to LEO via one STS

Launch. It is highly desirable that the satellite weight not be reduced and

that the capability to launch four satellites to LEO via one STS flight be

provided. An option to allow this, would be to off-load propellant from one

or all of the satellites and top-off the satellites using the OMV/resupply

module in conjunction with a LEO storable propellant depot (Figure 18).

• LAUNCH OF FOUR SATELLITES TO 150 NMI x 28.5° LEO IN ONE STS

DESIRED

(4) 10 FT DIA x 10 FT

LONG SATELLITES

A

• TOTAL INSTALLED WEIGHT IN PL BAY APPROXIMATELY 77,000 LB

SATELLITES FULLY FUELED

• OFF-LOADING 3000 LB PROPELLANT PER SATELLITE OR 12,000 LB

FROM ONE SATELLITE ENABLES LAUNCH TO LEO BY STS OF ALL 4

SATELLITES

Figure 18. Potential LEO Top Off of DoD High Energy Orbit

Integral Storable Propulsion Satellites

2065e/ - .21 -
The recommended mid-term review DRMs are shown in Figure 3« The common

denominator between the recommended "best" design reference missions is

bi-propellant -fluid resupply. Ample evidence exists (e.g., Insat, Intelsat

VI, and NASA/MSFC's baseline OMV) to indicate an evolution from hydrazine to

bi-propellant forms of propellant storage on-orbit. The specific impulse

advantage of bi-propellant, coupled with its availability through scavenging

from the orbiter's QMS tanks suggests that NTO/MMH bi-propellant may become

the "standard" of storable fluids (Figure 19).

• REF FLUIDS FOR RESUPPLY ARE BI-PROPELLANTS

• NTO & MMH

• RATIONALE:

• LARGEST PROJECTED FUTURE QUANTITY

• NATURAL EVOLUTION FROM N2H4 IN-BAY DEMONSTRATIONS

• BASIC PROPELLANT FOR NUMEROUS FUTURE SPACECRAFT WHICH CAN

BENEFIT FROM RESUPPLY

G£O | LARGE COMSATS (ALREADY USED ON INTELSAT VI)

I FUTURE LARGE GEO PLATFORM

LE£)

 ( REFUELING OMV & HIGH PERFORMANCE STORABLE STAGE

I POTENTIAL RESUPPLY DEPOT REQUIREMENT

• REFERENCE QUANTITY TBO

• CANDIDATE MISSIONS

V LARGE GEO PLATFORM

V FUTURE LARGE COMSATS

VRESUPPLY IN LEO

• DOO (MANEUVERING)

• ORBITAL VEHICLES

Figure 19- Mid-Term Review 1>RM Recommendations

3.2 Concept Definition

Selected Operational System Configuration

The ten basic mission scenarios/DRM's indicated in Figure 5 were evaluated for

sizing requirements to the Resupply Module propellant requirements from 6,000

to 45t000 Ibs. of storable bipropellants and Helium requirements from zero to

90 Ibs. Scenario two is the driving requirement for GEO operations (avoids

expending an OTV before the available timeframe of a reusable OTV). The full

resupply module capacity is used to supply enough propellant to transfer an

OMV plus its servicer kit to GEO. Resupply in this case would be directly

from the OMV. Scenario five is the driving requirement for LEO remote

resupply operations - refueling a high performance reusable storable OTV.

Lesser levels of resupply module capability are provided by off-loading the

module rather than removing tankage. The objective is to adapt one basic

"core" resupply module design to many different uses.

2065e/ - 22 -
The ERM in the various scenarios served as an OMV extended mission kit, a LEO

refueling ERM, a LEO propellant and Helium storage depot, a precursor reusableupper stage, and"a means to transfer orbiter scavenged storable propellant to

a LEO depot.

A versatile, low-cost resupply module concept resulted from the trade analyses

and design definition phases of the study. Some of the various trades and

design characteristics included propellant transfer alternatives, weight

trades for a pump-fed versus pressure regulated transfer system, propellant

system schematics, transient and steady state pressure drop characteristics,

fluid interface panel design, thermal control provisions, electrical power

requirements and distribution, avionics and data processing interfaces, major

subsystem weight breakdowns, user spacecraft impact assessment, general

configuration layouts and key features. To provide the required propellant

capacity, the resupply module configuration depicted in Figure 3 was selected.

As discussed earlier efficiency in combination with utilization of existing

hardware for low development cost. The resupply module is supported at its

forward end by an existing IUS upper stage forward cradle. This cradle

includes load equalization capability which reduces the structural redundancy

between the resupply module and the orbiter by one degree and allows a minimum

of scar weight on the resupply module for attachment to the orbiter payload

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

stretched QMS tanks. Also the QMS-type PMD's are replaced by a zero-g concept

based on existing technology.

As shown in Figures 4 and 5, the basic structural components of the Resupply

Module are very simple. Yet, the configuration is very effficient

'structurally. All fore and aft loads and part of the vertical loads are

supported at two payload bay longeron attachment points in the main structural

bulkhead. As a representative attachment to the OMV, six bolts are provided.

For the case where the Resupply Module is detached from the OMV in GEO, these

bolts may be explosive. The main bulkhead also has a keel attachment for the

orbiter payload bay.

The tanks are supported by the main bulkhead fore and aft, as well as lateral

and vertical loads. They are stabilized at their forward ends by machined

fittings which attach to the cylindrical center body of the Resupply Module.

This body provides the main bending stiffness of the ERM. NASTRAK finite

element analysis showed that this concept meets the Orbiter fundamental

response stiffness requirement of 7 Hz even when fully loaded with

propellant. Most elements of the structure are machined aluminum plate,

including the center body which is rolled to its proper shape after machining

in quarter sections which are subsequently fastened together mechanically.

A design layout (Figures 4 and 5) was produced of the selected Resupply Module

configuration. This layout was used in conjunction with a finite element

analysis to develop a realistic weight estimate for the resupply module

structure. Two adaptations are depicted, with and without servicing. In both

cases the OMV is supported from the resupply module in the orbiter payload bay

(OMV support fittings could be removed to reduce weight and cost). The

combination OMV/ERM would be deployed from the payload bay using the orbiter

ERMS. The adaptation depicted with the truss interstage (Figure 5) allows

sufficient volume between the OMV and ERM for the OMV's servicer kit (for

missions in which concurrent module change-out type servicing is required).

2065e/ - 23 -
Flight Demonstration Program

The fright demonstration concept will include sufficient fidelity to test the

essential functions of remote on-orbit resupply and demonstrate to the user

community the viability and attractiveness of such a concept. In order to

show proof-of-concept specific elements will be required to be tested on each

side of the resupply interface.

Simplicity and economy are emphasized in the selected flight demonstration

program. The principle test objective is to validate bi-propellant fluid

transfer in a zero-gravity environment. Since the ERM engagement with a

spacecraft is accomplished by the OMV with a remote manipulating arm, the

remote manipulator system (EMS) on the Orbiter can approximate the OMV

engagement capability. The systems on both the receiver and supplier test

articles can be attached to either a MPESS (NASA owned) or a SPAS structure

(structure only without any subsystems). As demonstrated in Figure 20, the

Orbiter RMS will lift the receiver test article out of the cargo bay and

engage the supplier test article. Repeated fuel transfer tests will then be

performed.

Since the supplier test article has the greater need for f>ower and electrical

links to the Orbiter, we recommend it remain seated in the cargo bay. This

avoids the need for complicated power and electrical links running through the

RMS.

The system on both the receiver and supplier test articles can be attached to

either a MPESS (NASA-owned) or a SPAS structure (shown). Grapple fixtures are

provided on the receiver simulated vehicle for use in docking with the

supplier vehicle and by the ERMS to transport the receiver vehicle from a

berthed position to the docked position with the supplier vehicle. Tankage

employing ullage control PMD's will be utilized. Fluid transfer system

hardware will incorporate, to the greatest extent possible, final design

pumps, compressors and disconnects.

The resupply module flight demonstration program supplier vehicle will utilize

final design type hardware in the fluid transfer subsystems including the

pumps, compressors and disconnects. The propellent tankage will be

essentially off-the-shelf with PMD's that provide ullage positioning in zero-g

to allow the ullage exchange process during resupply. The system plumbing and

structural configuration of the supplier and receiver test vehicles will be to

sufficient fidelity and detail to provide operational performance for the

final resupply module flight article design. Figure 21 presents the resupply

module flight demonstration supplier module transfer subsystem schematic.

2065e/ - 24 -
IS

OF POOR QUALRY SPAS OR MPESS STRUCTURE FOR

RECEIVER AND SUPPLIER

SIMULATORS

UuMin SHS

in Demrannai

Mount

SUPPLIER

SIMULATOR

SPiS-01

taSTS-7

• first ff it Ryu rfon

Ortmif Piylout Biy

. PrownDockBjWrtnRMS

RECEIVER

SIMULATOR

RMS USED TO DOCK

RECEIVER TO SUPPLIER IN

PAYLOAD BAY

• MOST OF DEMO SUBSYSTEMS LATER

USED FOR OPERATIONAL FLIGHT

UNITS FOR COST SAVINGS

Figure 20. Bi-Propellant/Helium Transfer Flight Demonstration Concept

INTERFACE

PANEL

PRESSURANT

CHECKOUT

d) PANEL

PROPELLANT

_ „ SERVICING &

S - d) PURGE/DRAIN

Figure 21. Resupply Module Flight Demonstration Program Supplier Schematic

2065e/ - 25 -
3.3 Programmatic and Developmental Planning

The programmatic and' developmental planning task included the definition of

cost and schedule requirements for the resupply module operational program,

technology development and flight demonstration programs.

Technology Development, Flight Demo, and ERM Program

The technology development program would precede "both the flight demonstration

program and the ERM phase C/D. All the technology requirements and their

maturity status are shown in Figure 22. Standard NASA definitions were used

to assess each technology's maturity level. By definition, those technologies

with low maturities are the ones that need to be emphasized in the technology

development program. The critical technologies are, (l) the gas compressor,

(2) fluid pumps, (3) quick disconnects, contamination control, and leakage

monitoring.

Two options were considered for the flight demonstration program: An early

flight demonstration by the end of the FY 1988 and a late flight demonstration

that is part of the ERM phase C/D program that flies a year later at the end

of FY 1989• The schedule for the early flight demonstration option is shown

in Figure 23. Breadboard testing precedes the verification testing. The

verification testing could culminate with the flight demonstration.

Technology Maturity Levels

No. Description

8 Operation*

7.2 Engineering Model Tested in Space

7.1 Engineering Model Qualified

6.J Prototype Developed to quality

6.2 Prototype Tested in Test Bed Unmanned

6.1 Prototype Teated at Contractors

5.4 Preprototype Tested at NASA

5.3 Preprototype Tested at Contractora

5.2 Major Function Tested at NASA

5.1 Major Components Tested at Contractora

4.2 Critical Function Teated Over Time

4.1 Critical Function Demonstrated

3.2 Conceptual Design Tested Experimentally

1.1 Conceptual Design Tested Analytically

2 Conceptual Design Formulated

1 Basic Principles Observed and he ported

Technology Requirements

Fluid Containment/Transfer Design

*

Fluid Acquisition

••

•

IM

•

0

•o

f4

a

r*

••

—

•CL

3

3

•B

f4

3

n*

Im

mm

Gas Compressor mi

k Disconnects

•**

§•

••

/Quantity

urenenta

o «

= £

00

e

•*

u

0 O

9Z

• o

2*

••••

amlnation Control §

•i

erature Control

9)

• ~>

t) U •

w oo e u

ua E — v e u —

b 3 B «

u ^ o **

t/) 0. O M

Remote

Operations

•

«

to ** u

. o e

O U <M

u u

• C II

s o w

> O " C,

> ee u «•? M e « •*

v •* w c w 1 Jt O O «

» o « > o

e o at < o.

Development Risk Assessment 2

1- Hone, 2- Very Low, 3- Low, 4- Moderate

2 3 4 3 1 2 3 1 1 1

Figure 22. Status of Technology Required for ERM Program

2065e/ •/ L i - 26 -
I FY86 FY87 FY88 FY89

•SPFf/RFQIIIRFflFRTg

PHASE

TESTING PHASE

BREADBOARD (B/B)

1. COMPRESSOR

2. PUMP

nFV. /VERIFICATION

1. COMPRESSOR

2. PUMP

3. DISCONNECT

LEGEND:

D - DESIGN COMPLETE

F - FABRICATION COMPLETE

T - TEST COMPLETE

A- FINAL REPORT

~j- B/B LOW LEVEL

" TEST SUPPORT

ft * T

O F T

A

N2IU|

MANUALLY

ACTUATED

nFLIVERFD

QUAL

Figure 23. Development Schedule for Concept Feasibility

The schedule for the late flight demonstration option is shown in Figure 24.

The technology development phase culminates with the successful verification

of key predefined ground tests with breadboard test apparatus, probably by the

end of FY 1987. Given a successful verification at the breadboard level, the

ERM phase C/D could start. After the specification and requirements for the

ERM program are defined, the flight demonstration program would begin.

Changes to previous specification and requirements defined in the technology

development program will probably result in a redesign/modification to the

critical technology components in parallel with the design of the flight

demons tra t ion.

The Baseline Operational ERM program consists of two test articles

(Developmental and Qualification) and two flight units (one from salvaging

systems from the best articles. The earliest possible start for the phase C/D

of the ERM program is the fourth quarter FY 1987, and the earliest possible

operational flight is the last quarter of FY 1991.

Estimate Program Cost

Figure 25 presents cost estimates for the Technology Development and Flight

Demonstration Programs, with all costs in constant 1984 dollars. STS launch

costs are not included for either the flight demonstration cost or the first

ERM operation flight. Costs for GSE and payload support, sustained mission

operations/training, maintenance and refurbishment, and additional production

units are not included in the baseline ERM program.

2065e/ - 27 -
ORiGINAL PAGE 33

OF POOR QUALITY

1986 1987 1988 1989 1990 1991 199?

| TECHNOLOGY DEVELQPHEHT PROGRAM |

SPECIFICATION/REQUIREMENTS

BREADBOARD DESIGN

BREADBOARD FAB I TEST

CRITICAL TECHNOLOGY COMPONENT

| FLIGHT DEHOH5TRAT10H PROGRAM |

SPEC If ICAT ION/REQUIREMENTS

FLIGHT DEMO DESIGN

FLIGHT DENO FAB t TEST

CRITICAL TECHNOLOGY COMPONENT

ERH PROGRAM

SPEC I F I CAT I ON/REQU I REMENTS

DETAIL DESIGN

DEVELOPMENTAL TEST ARTICLE

QUALIFICATION TEST ARTICLE

FIRST FLIGHT HARDWARE

MISSION SUPPORT I ANALYSIS

12) CONCEPT VERIFICATION

DESIGN

| FAB

(DESIGN FAB|

| TEST

TEST

FROM ERM PROGRAM

TEST t. I/O

/

FAB | \ <

D

,PEC

ESIGN FAB | TEST

©

DESIGN^

®

/ FAB | TEST « KEFURB SYS y" "

/ FAB 1 TEST | X ^\

-ASSY OF SECOND

FLIGHT HARDWARE

TEST C OPERATIONS

LEGEND

1 SPECIFICATIONS FOR BREADBOARD DESIGN

2 BIPROPELLANT TRANSFER VALIDATED WITH GROUND TESTING LEADS TO

ERM PROGRAM PHASE C/D START

3 ERM SPECIFICATIONS DRIVES FLIGHT DEMONSTRATION PROGRAM

<• FLIGHT DEMONSTRATION HARDWARE AVAILABLE FOR REUSE ON FIRST

ERM FLIGHT UNIT

5 FIRST ERM FLIGHT

Figure 24

Advanced Development

Flight Demonstration, and Resupply Module Program

2065e/ - 28 -
TECHNOLOGY DEVELOPMENT AND

DIRECT HOURS

LABOR RATE

I SUBCONTRACTED

SPAS AND TANKS

LOW

133.400 HR -101

SSO/HR

30S

9.5M

1.9*

11.4R

FLIGHT DEMONSTRATION

BEST ESTIMATE

148.200 HR +20X

J55/HR

40X

13. 6M

Z.SH

16. 1M

HIGH

177,800 HR

J60/HR

SOX

21. 3M

3.4M"

24. 7M

TECHNOLOGY DEVELOPMENT

DIRECT HOURS

LABOR RATE

I CONTRACTED

TANKS

TOTAL

LOU

51.900HR -lOt

S&O/HR

301

3.7M

1.2M

4.9M

BEST ESTIMATE

57,700 HR +201

155/HR

40X

5.3M

1.6M

6.9N

PROGRAM

HIGH

69.200 HR

S60/HR

SOS

8.3M

2.4M

10. 7M

FLIGHT DEMONSTRATION PROGRAM

DIRECT HOURS

LABOR RATE

I CONTRACTED

SPAs

TOTAL

LOU

81.500 HR

SSO/HR

30X

5.BM

0.7M

6.5M

BEST ESTIMATE

10X 90.500 HR +201

155/HR

40X

8.3M

0.9H

9.2M

HIGH

108.600HR

I60/HR

SOX

13. OS

I.OM

14. OM

Figure 25

Cost Estimate for Technology Development and

Flight Demonstration Programs

4.0 RECOMMENDATIONS FOR TECHNOLOGY DEVELOPMENT

Three major components can be identified as requiring special developmental

testing to prove concept feasibility. These are the propellant pump, the

helium compressor, and fluid disconnect.

The major test/demonstration items required to verify satisfactory component

operation and performance are as follows:

o Evaluate component design concept feasibility with a workhorse-type

test unit,

o Verify operation and performance of a flight-weight prototype test

unit to flight application requirements.

The scope of the Test Program should encompass the following:

o Breadboard-type tests using reference fluids and a development

(workhorse) component test unit shall provide the basic data base for

component functional design evaluation.

o Development tests using reference fluids, propellents and prototype

flight-weight test components will provide the basis for verification

of the component design concept feasibility. These tests will assess

operation and performance capabilities when subjected to flight

environments, ground servicing and fluid transfer usage. In addition,

component life will also be demonstrated.

2065e/ - 29 -
5.0 Recommendations for Technology Development

The specifications and requirements for the Advanced Development Program must

be oriented to the Resupply Module Program so that the critical technology

components (pumps, compressors, and quick disconnects) developed are likely to

be usable without extensive redesign. Most likely, considerable learning will

take place during the Technology Validation Program and the Phase B study of

the program. Thus, the critical technology components will probably require

some redesign and consequently will be recertified in the Plight Validation

Program. Of critical importance, is that all systems and hardware in the

Plight Demonstration Program reflect the specifications and requirements of

the operational E1M Program.

Timely program development could lead to a flight demonstration by late 1989

given the availability of qualified storable bi-propellant disconnects

(derived from the GRO resupply tanker). The first operational remote resupply

mission (probably in LEO) would be supported by the first OMV docking

demonstration probably by late 1990 to early 1991* There is considerable

potential for commonality between the NASA-JSC earth storable propellant

tanker program (first refueling in-bay of GRO also planned for late 1990) and

the remote resupply flight demonstration program. This synergysm could reduce

the costs of both programs and any follow-on effort.

2065e/ - 30 -