A

fNASA-CK-178819) SPACE PLA1FCEH EXPENDABLES ~ ~N86-247~32

RESOPPLY CONCEPT DEFINITION SIODY Final

Report, Jan, - Oct. 1S85 {Bocksell

international Corp.) 116 p He &06/8F A01 Dnclas

v CSC1 22B G3/18 42931

STS85-017A

ADDENDUM

FINAL REPORT

SPACE PLATFORM EXPENDABLES RESUPPLY

CONCEPT DEFINITION STUDY

FOR PERIOD JANUARY 1985 - OCTOBER 1985

DECEMBER 1985

CONTRACT NAS8-35618

Rockwell International

Space Transportation

Systems Division
Addenda To:

Space Platform Expendables Resupply

Concept Definition Study

Final Report

STS85-0174

Contract MAS 8-35618

For Period January 1985 - October 1985

347 4e -1 -
CONTENTS

Section Title Page

1.0 Executive Sunmary 9

1.1 Introduction 9

1.2 Background 9

1.3 Supplemental Study Scope and Objectives 12

1.4 Sunmary of Results 12

2.0 Study Results 28

2.1 System Requirements and Scenarios (Task 1 ) 28

2.2 Concept Definition (Task 2) 42

2.3 Programmatic and Development Planning (Task 3) 96

3474e -2-
ILLUSTRATIONS

Fi gure Page

1 RM/OMOV Configured to Meet User Needs 10

(Developed in Original Study)

2 Bi-Propel!ant/Helium Transfer Flight Demonstration 11

Concept (Developed in Original Study)

3 Requirements Depend on Evolving Space Operations 13

4 Fluid Requirements Driven By Space Operations 15

5 Three Types of Surveys Conducted 16

6 User Need Determines SPER Module Sizing 17

7 Technical Studies Determine Demonstration Data 20

Requirements

8 Mid-Deck Ullage Transfer Experiment 21

9 Payload Bay Flight Demonstration and Ground Test 22

Article

10 Comparison of Advanced Development/Flight 24

Demonstration Options

11 Advanced Development and Flight Demonstration 27

Program (Master Schedule)

12 Requirements Depend on Evolving Space Operations 29

13 Projected Annual Fluid Transfer Engagements 34

14 Fluid Requirements Driven By Space Operations 35

15 Three Types of Surveys Conducted 36

16 Resupply Survey Questionaire Results to Date 40

(October 1985)

17 User Need Determines SPER Module Sizing 41

18 Fluid Transfer Schematic, Pressurant Purge Design 47

18a Fluid Transfer Schematic, Pressurant Purge Design, 47

Without Self Refueling Plumging

19 Fluid Transfer Schematic, Dribble Volume Purge 48

Design

20 Fluid Transfer Schematic, Vacuum Pump Purge Design 48

3474e -3-
ILLUSTRATIONS (Continued)

Figure Page

21 Required Gear Pump Power at Various Delta P for MMH 49

22 Required Gear Pump Power at Various Delta P for NTO 50

23 Required Gear Pump Power at Various Delta P for NTO 51

and MMH

24 Required Gear Pump Power at Various Delta P for 52

Hydrazine

25 Power to Pump at Various Delta P for MMH 54

26 Power to Pump 55

27 Power to Pump at Various Delta P for NTO and MMH 56

28 Power to Pump at Various Delta P for Hydrazine 57

29 Final Receiver Pressure Vs. No. Supply Tanks 60

30 System Weight Comparison Cascade Only Resupply 63

Method

31 System Weight Comparison Cascade With Compressor 64

Resupply Method

32 Temperature and Pressure Location 68

33 Representative Interaction of Remote Resupply 69

Development

34 Standardized Refueling Coupling Schematic 71

35 Fairchild Manual Coupling 72

36 Mid-Deck Ullage Transfer Experiment Schematic

37 Flight Demonstration Test Vehicle 82

Schematic - MMH HALF

38 Flight Demonstration Test Vehicle 83

Schematic - NTO HALF

39 Expendable Resupply Module Test Program 86

40 & 41 Component Test Schematic 89

42 Flight Demonstration Test Vehicle (Preliminary 91

Design Concept)

3474e -4-
ILLUSTRATIONS (Continued)

Fi gure Page

43 Comparison of Advanced Development/Flight TOO

Demonstration Options

44 Advanced Development and Flight Demonstration 103

Program (Master Schedule)

45 Mid-Deck Ullage Transfer Experiment Schedule 104

and Manpower Estimate

46 Mid-Deck Ullage Transfer Experiment 106

47 Flight Demonstration Test Schedule and 112

Manpower Estimate

48 In-Bay Flight Transfer Experiment 114

3474e -5-
TABLES

Table Page

1 Mission Model Integrates All Fluid Transfer Users 30

2 Resupply Requirements Summary 32

3 Update of SPER Module Mission Model Data Base - NASA 33

4 Commercial Firms/Consultants Surveyed 38

5 United States Government and Foreign Governments 39

Surveyed

6 Space Platform Expendables Resupply Module Sizing 43

Options and Drivers

7 Disconnect/Line Purge Options, Major Advantages 46

and Disadvantages

8 System Volume Comparison 59

9 System Weight Comparison 59

10 Battery System and Power Requirements for Cascade 62

With Compressor Method

11 Pressure Instrumentation 66

12 Temperature Instrumentation 67

13 Recommended Tests 75

14 Matrix of Test Objectives 76

15 MUTE System Weight 76

16 Recommended Tests 79

17 Matrix of Test Objectives 80

18 Flight Demonstration Test Vehicle System Weight 81

19 Compressor B/B Test Matrix 88

20 Programmatics Overview 97

21 In the Early Flight Demonstration Option Has 99

Superior Advantages and only Minor Disadvantages

22 Mid-Deck Ullage Transfer Experiment Significant 102

Characteristics

3474e -6-
TABLES (Continued) •

Tab!e Page

23 Mid-Deck Ullage Transfer Experiment Cost Range 107

Estimate

24 Flight Demonstration Test Significant 109

Characteristics

25 Flight Demonstration Test Cost Range Estimate 110

&

111

3474e -7-
ACRONYM DICTIONARY

AXAF Advanced X-Ray Astrophysics Facility

B/B Brassboard

BOL Beginning of Life

CREP Cosmic Ray Experiment Program

DDT&E Design, Develop, Test and Engineering

DMSP Defense Meteorlogical Satellite Program

DRM Design Reference Mission

DOD Department of Defense

DOMSAT Domestic Satellite

EOS Earth Observing System

ERM Expendables Resupply Module

ETR Eastern Test Range

EUVE Extreme Ultraviolet Explorer

EVA Extra Vehicular Activity

FDTA Flight Demonstration Test Article

GEO Geosynchronous Orbit

GP-B Gravity Probe-B

GPM Gallons Per Minute

GRO Gamma Ray Observatory

GSE Ground Support Equipment

GSTDN Ground Space Flight Tracking and Data Network

He Helium

JPL Jet Propulsion Lab

JSC Johnson Space Center

LaRC Langley Research Center

LDR Large Deployable Reflector

LEO Low Earth Qrbit

MIL Man-in-Loop

MMH Monomethylhydrazine

MPS Materials Processing In Space

MSFC Marshall Space Flight Center

MUTE Mid-deck Ullage Transfer Experiment

NASA National Aeronautics and Space Administration

NTO Nitrogentetroxide (N204.)

N2 Nitrogen

N2H4 Hydrazine

OMV Orbital Maneuvering Vehicle

OTV Orbit Transfer Vehicle

PMD Propellant Management Device

PVT Pressure, Volume, and Temperature

RCS Reaction Control Subsystem

RF Radio Frequency

RM Resupply Module

RMS Remote Manipulator Subsystem

ROM Rough Order of Magnitude

S/C Spacecraft

SDI Strategic Defense Initiative

SIRTF Shuttle Infrared Telescope Facility

SPAS Shuttle Pallet Satellite

SPER Space Platform Expendables Resupply

STS Space Transportation System

TFU Theoretical First Unit

TDRSS Tracking and Data Relay Satellite System

USAF SD United States Air Force Space Division

WTR Western Test Range

3474e -8-
1.0 EXECUTIVE SUMMARY

1.1 INTRODUCTION

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

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

operational efficiency and cost/effectiveness. The emerging Orbital

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

will introduce the capability for fluid resupply enabling satellite lifetime

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

summarizes a supplemental study to the original Phase A study and is presented

as addenda to that study.

1.2 Background

As background an overview of key results from the original study are

presented. The original Phase A study was a $340K effort which ran from March

1984 through January 1985. Long term objectives for this activity consisted

of the definition of both an operational Expendables Besupply Module (ERM) and

the definition of a flight demonstration for remote resupply. To this end the

specific study objectives for the Phase A study consisted of the definition of

user needs and requirements, the conceptual design definition of a

representative resupply module, a conceptual approach to the flight

demonstration of remote resupply, and the definition of a program plan to

satisfy the long term objectives.

One key result of this original study activity was the design development of a

resupply'module concept. This concept and its key features are presented in

Figure 1. The key features of this configuration include the capability of

the SRM to be utilized as an ERM or an OMV extended mission kit. Six

stretched QMS tanks contain 4-5,500 Ibs. of usable propellant for resupply or

OMV propulsive use. The dry weight of the ERM for bi-propellant resupply only

is 7,960 Ibs. For helium and bi-propellant resupply it is 9,14-0 Ibs. The dry

weight of the ERM to be utilized as an OMV extended mission kit would be 7,200

Ibs. This includes the elimination of helium compressors and associated

batteries and simplified propeilant management devices to meet OMV propellant

transfer requirements.

Also as a result of the original study, a concept for a Bi-propellant/Helium

transfer flight demonstration was developed. This concept is presented in

Figure 2. The principle test objective for the flight demonstration is to

validate storable bi-propellant fluid transfer in a zero-gravity environment.

Since the ERM engagement of a spacecraft is accomplished by the OMV with a

remote manipulating arm end effector, the remote manipulator systems (RMS) on

the Orbiter can approximate the OMV engagement capability. The systems on

both the receiver and supplier test articles can be attached to either a MPESS

or a SPAS structure (structure only without any subsystems). The Orbiter RMS

will lift the receiver test article out of the cargo bay and engage the

supplier test article. Repeated fuel transfer tests will then be performed.

Since the supplier test article has the greater need for power and electrical

links to the Orbiter, it was recommended it remain seated in the cargo bay.

This minimizes the interface complexity of power and electrical links running

through the orbiter. The supplier test article would be autonomous with its

own power source.

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Some additional key conclusions from the original study are as follows:

o Remote resupply provides substantial benefits

o In GEO, mainly with concurrent servicing

o But . . . far term horizon

o In LEO, in conjunction with storable propellant depot

o Supports wide scope of potential operations

o Single "core" resupply module design is adaptable to both requirement

categories

o Total program cost including adequate technology development & flight

demonstration is low/affordable

o Remote resupply can be operational by early 1990's LEO operations

o LEO operations projected first need

o Evolves to support space station growth

1.3 Supplemental Study Scope and Objectives

The supplemental study was a $98K effort conducted from January 1985 through

October 1985. In the supplemental study the requirements task was expanded

beyond the original study to include consideration of the evolving space

transportation infrastructure. The scope of the design section, in the

supplemental study, was limited to the consideration of the transfer process

itself and to flight demonstration concepts. The scope of the supplemental

study programmatics task was limited to the generation of cost estimates of

the resupply options and to provide support in identifying the "Best"

demonstration program.

The objectives for this supplemental study were separated into three tasks:

requirements, design, and programmatics. The primary objective was to select

a preferred flight demonstration program option. This was accomplished by

defining user needs, assessing the state of the development of the technology,

defining demonstration concepts, and costing those concepts. User needs such

as fluid type, quantity, and schedule enable module sizing to be defined.

This information was then used as input to the design tasks where the

demonstration design requirements and preliminary designs are generated. The

programmatics task develops the cost estimates for the resupply concepts and

provides decision support for selecting the "Best" demonstration program.

1.4 S""""ary of Results

Systems Requirements and Scenarios (Task l)

The system requirements of the original task 1 were revised to include a

projection of elements from the evolving space infrastructure. Figure 3

depicts an evaluation of possible key elements of the space infrastructure.

These depicted elements are the ones of primary interest to the task of

expendables resupply. In accordance with the expanded scope of this contract,

a closer look has been taken at the resupply needs and interactions of

infrastructure elements.

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In this analysis the shuttle is being used as the basis for an in-bay

propellant tank (under JSC) and later the arrival of the OMV with associated

kits, such as the resupply module, servicer kit and capture kit. The OMV and

associated kits can be deployed from the ground, and/or space-based at 28.5

degrees or polar orbits. These spacecraft will then be augmented by permanent

orbital facilities primarily the space station with its accommodations for

OMVs and OTVs. Associated with the station, but separate from it for safety

requirements, will be a Leo depot, possibly supplied by propellant scavenging

from the STS. Eventually, this depot should be duplicated in polar orbits

with its own OMV and associated kits.

Around the turn of the century, new developments such as the shuttle block

change and a fully reusable OTV will likely alter the economics of supplying

propellant in space. The Shuttle block change could affect the use of

propellant scavening, and the OTV will generate requirements for greater

amounts of fuels, raising the need for sophisticated storage facilities in

space. A secondary space transport and Launch Vehicle will likely arrive in

support of SDI. SDI elements are not sufficiently developed yet to predict

their impact on orbital fluid requirements and have, therefore, been left out

of this projection.

Figure 4 depicts, in graphic form, the fluids to be transfered on-orbit to the

elements of the space infrastructure. Parametric techniques were used to make

estimates of the types, amounts and yearly fluid usage rates required by the

resupply mission model. Small quantities of associated pressurants such as He

and N2 were also required, but their amounts were relatively minor in

comparison to the primary fluids. In some cases, depending on the resupply

technique used, no additional pressurants are required for the transfer

process. In the latter 1990's, bi-propellants and cryogens are the dominant

required fluids. Hydrazine, water and liquid helium constitute a relatively

stable base of demand.

In updating the resupply mission requirements extensive contacts were made

with the potential user community. In addition to in-house data base and

literature searches, three related surveys were conducted Figure 5- They were

telephone contacts with various NASA program contacts used in the development

of the resupply mission model. The updated data did not contain any major

surprises and represented a maturing of several programs which may hold a more

positive attitude toward fluid resupply.

In January of 1985, a series of presentations were made to satellite

manufacturers TRW, Hughes, and Ford. The presentations were intended to

direct feedback on the proposed demonstration program and user benefit

analysis. A survey questionaire, prepared under Rockwell IR&D, was sent to

over 90 members of the space community in the U.S. and aboard. The contacts

included domestic and foreign space agencies, aerospace and communications

companies, and space insurers. The survey attempted to find new candidate

users of fluid transfer services for either remote or Orbiter-based service.

Thirteen useful responses have been received to date, with most supporting the

concept of concurrent servicing with resupply. As a result of the

requirements re-evaluation and discussions with potential resupply users, six

scenarios were developed for the ERM. These scenarios are depicted in Figure

6. In Scenario 1, the expendables resupply module provides the OMV with

propellant

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for transportation to GEO. This mode was found to be more economical than

using an expendable OTV to transport the OMV to GEO. Propellant transfer (and

concurrent module exchange servicing, when required) is accomplished directly

from the OMV after separation from the resupply module in Scenario 2. This

uses a reusable cryogenic OTV for Geo transport.

In Scenarios 3 through 6, the EEM operates in LEO and is reusable.

Some ERM's will be parked at the fuel depot (Scenario 5) and

participate in fuel scavenging operations from the STS (Scenario 6).

Some EEM1s will perform supply and/or top-off operations:

The major conclusions from the update of the requirements and scenarios is

presented as follows:

o Fluid transfer should be considered as an integral part of space

infrastructure

o Major customer for fluid transfer are other infrastructure elements:

OMV, OTV, Space Station

o All potential users support concurrent servicing

o Selection of MMH/NTO as baseline fluid still justified by user needs

o Need for resupply of other fluids

o Module sizing determined by operational scenarios

o Decoupling of tankage from resupply module may be advantageous to

development effort

o Adequate MMH/NTO available from OMV initially

o Focus on key technical/operational risk elements

o Lower front-end cost for timely development

The spectrum of resupply missions possible has been summarized into a set of

operational scenarios with drivers of resupply module sizing. This results in

a confirmation of the selection of MMH/NTO as the baseline fluid while

down-sizing the earlier resupply module design (45,500 Ibs of bi-propellant).

An ESM of some 7 klb of bi-propellants was deemed sufficient, while the larger

version serves as a growth version of the design.

A decoupling of tankage from the resupply module was discovered as a potential

means for concentrating on the technical issues associated with fluid transfer

while preparing flexibility in an evolving market for fluid transfer.

3474e -18-
Concept Definition

The Concept Definition task of the supplemental study was divided into two

major areas. These areas were: (l) the refinement of the transfer subsystem

for the purpose of determining the engineering objectives of the fluid

demonstration tests and (2) to develop a fluid demonstration test concept(s)

which would satisfy the engineering test requirements.

Several technical areas were studied to refine system requirements for a

propellant and pressurant transfer. The technical areas examined are as

follows: umbilical purge, pressurant transfer, propellant pump selection,

NASA quick disconnect development, and pressure and temperature

instrumentation. The results of these studies are presented in Figure 7. Two

flight test concepts were developed to provide proof of concept for on-orbit

remote expendables resupply. The first is the Mid-Deck Ullage Transfer

Experiment (MUTE) which is presented in Figure 8. MUTE's fundamental

objective are as follows:

1) To demonstrate the concept of ullage transfer in a micro-gravity

environment.

2) To demonstrate the propellant management devices (PMD's) ability to

position and control the ullage bubble.

3) To demonstrate the quantity gaging system's accuracy in a

micro-gravity environment.

The demonstration of a micro-gravity quantity gaging system is not critical to

the success of this flight test. However, MUTE does provide an opportunity to

conduct an on-orbit test of the micro-gravity gaging system that is being

developed as a separate NASA program.

The second flight test concept is the payload bay Flight Demonstration Test

Article (FDTA) which is shown in Figure 9. The Flight Demonstration and

Ground Test Article will be used in the orbiter cargo bay to demonstrate a

micro-gravity remote fluid transfer. FDTA consists of two platforms capable

of being separated or docked using the remote manipulator subsystem (RMS).

The upper portion of the test article, the receiver platform, piggy-backs on

the supply platform through deployable payload latches. The RMS system is

utilized to separate the receiver platform, by use of a grapple fixture, to

simulate remote docking and mating of two free-flying spacecraft. A snare

type end effector, incorporated at the fluid transfer interface of the supply

vehicle, accomplished final mating of the fluid transfer interface panels on

each test platform.

Should it become necessary to reduce the cost of the FDTA several options are

available. One option is to conduct the tests with only one fluid either MMH

or NTO. This enables a reduction in the size thereby lowering system weight

and material/component procurement and fabrication costs.

Another approach to lowering flight demonstration cost is through the

reduction in test objectives. Requirements analyses have indicated that a

compressor may not be need during the early phases of on-orbit resupply

operations. Elimination of the compressor from the advanced development

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program and the flight test article could result in considerable program cost

reductions. However, as on-orbit resupply operations mature this capacity

will become required and therefore, must eventually be developed.

The concept definition task conclusions are summarized as follows:

o Expendables Resupply Module (ERM) Design

o Pressurant purge of disconnects preferred

o Cascaded pressurant resupply preferred over compressor approach

for small systems

o Gear pump most effective in ERM flow & pressure regions

o Disconnects & liquid quantity gauging being developed under

other programs

o Mid-deck Ullage Transfer Experiment (MUTE)

o Verifies propellant management device (PMD) & quantity gauging

o Simple, benign fluids, & max. use of existing equipment

o Flight Demonstration Test Article (FDTA)

o Verifies PMD, pumps, disconnects, pressurant transfer

thermal & system control

o Operational fluids & max. use of existing equipment

o Ground test and flight test

Programmatic and Development Planning

For purposes of comparability, the schedules for advanced development and

flight demonstration recommended from the original and supplemental studies

have been laid on a common calendar scale (Figure 10). In the initial study

we recommended a late demonstration option, because it resulted in a one-year

earlier ERM phase C/D start.

As a result, of our supplemental study, we now recommend an earlier flight

demonstration option. This is accomplished by using the FDTA as the ground

systems test and verification test bed. This resulted in an earlier ERM phase

C/D program at approximately the same cost.

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The ideal schedule leading to the flight demonstration and, hence, the ERM

program start is presented in Figure 11. The initial effort is to specify the

hardware requirements for the ERM program. With the ERM requirements defined,

the design of the flight experiments and the development of the critical

technology components can start. Ideally, MUTE can fly in 22 to 24 months

after the program starts and the flight demonstration test is scheduled to fly

8 to 12 months afterwards.

The timing in development of several related fluid transfer components could

affect this schedule. The micro-gravity gaging system (NASA, JSC) would be an

ideal component to include on the Flight Demonstration Experiment. If the ERM

and the micro-gravity gaging system development schedules could be made

compatible then, could be the Flight Demonstration Experiment expenses could

be shared. The function of the micro-gravity gaging system, if not available

for the Flight Demonstration Test, could be performed by flow meters.

The Quick Disconnect (NASA, JSC) is essential to demonstrate remote resupply

in the Flight Demonstration Experiment. It is currently scheduled to be

available in mid-FY 1988, and its schedule may also need to be accelerated

slightly or the advanced development program stretched slightly.

The programmatic and development planning conclusions are presented as follows:

o Early Flight Demonstration - Before ERM phase C/D start

o Two Experiments - Mid deck ullage experiment & flight demonstration

test (orbiter cargo bay experiment)

o An Integrated Program - Same team for development & test

- Same development & flight hardware

o 312.5M Budget - for Advanced Development & Flight Demonstration

Programs

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2.0 Study Results

2.1 System Requirements and Scenarios (Task l)

A review of the task 1 study results is presented in this section of the final

report. The topics discussed in this review are the subtasks for task 1 which

are summarized as follows:

1) Expand mission model to include space infrastructure elements

2) Define parametric fluid quantities as determined by mission model

3) Survey Potential on-orbit resupply users

4) Reevaluate ERM scenarios from original study

5) Identify ERM sizing options and their driving requirements

Figure 12 depicts an evaluation of key elements of the space infrastructure.

These elements depicted are the ones of primary interest to the task of

expendables resupply. In accordance with the expanded scope of this

supplemental study, a closer look has been taken at the resupply needs and

interactions of infrastructure elements.

In this analysis we have the Shuttle being used as the basis for an in-bay

propellant tank (under JSC) and later the arrival of the OMV with associated

kits, such as the resupply module, servicer kit and capture kit. The OMV and

associated kits can be deployed from the ground, and/or space-based at 28.5

degrees or polar orbits. These spacecraft will then be augmented by permanent

orbital facilities primarily the space station with its accommodations for

OMVs and OTVs. Associated with the station, but separate from it for safety

requirements will be a Leo depot, possibly supplied by propellant scavenging

from the STS. Eventually, this depot should be duplicated in polar orbits

with its own OMV and associated kits.

Around the turn of the century, new developments such as the shuttle shock

change and a fully reusable cryogenic OTV will likely alter the economics of

supplying propellant in space. Certainly, the Shuttle block change will

affect the use of propellant scavenging, and the OTV will generate

requirements for greater amounts of cryo fuels, raising the need for

sophisticated gtorage facilities in space. A secondary space transport

element, the Unmanned Launch Vehicle, will likely arrive in support of SDI

deployments and increased lunar surface activity. The logistics for these

initiatives are not sufficiently developed yet to predict their impact on

orbital fluid requirements and have, therefore, been left out of this

projection.

The expanded mission model (Table l) reflects a current assessment of the

future demand for operational fluid transfer missions. Utilizing resupply

modules, OMV's and OTV's, fluid transfer operations are carried out to

on-orbit spacecraft and facilities. This model integrates the results of

separate OMV, OTV, and Space Station mission models into an identification of

fluid users and a schedule of missions. Many of the missions are not purely

fluid resupply, but resupply missions conducted in conjunction with other

satellite servicing tasks. It is not clear what the division will be between

fluid transfer operations occuring at the Orbiter and those occuring in a

remote in situ mode. This revision of the orbital fluid supply mission model

incorporates the updates and schedule changes since its last publication in

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September of 1984. Changes have been relatively minor; involving shifts in

the communications market, new DOD launch schedules, and the addition of some

potential foreign missions. The greatest uncertainty involved in this

assessment is the impact of Space Station operations.

Based on the expanded mission model Figure 13 shows a projection of all

on-orbit fluid transfer engagements for the 1989-2001 time frame. These

engagements include the needs of the DoD, large platforms, space station,

space-based OTV's and OMV's, as well as other NASA and commercial spacecraft.

Engagements allocated to OMV servicing missions means the use of OMV as a

carrier vehicle for the resupply module. Materials processing in space (MPS)

engagements are based on estimates of factory modules aboard free-flying

leasecraft. The number of engagements decline in the latter time frames as

MPS operations transition to space station proximity. Further fluid transfers

are then accomplished through routine station resupply flights. SDI needs are

currently unknown.

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

yearly fluid usage rates required by the resupply mission model (Table 2).

Small quantities of associated pressurants such as He and N£ were also

required, but their amounts were relatively minor in comparison to the primary

fluids. In some cases, depending on the resupply technique used, no

additional pressurants are required for the transfer process. Figure 14

depicts, in graphic for the same information.

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

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

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

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

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

In updating the resupply mission, extensive contacts were made with the

potential user community. In addition to in-house data base and literature

searches, three related surveys were conducted. They were telephone contacts

with various NASA program contacts used in the development of the resupply

mission model. The updated data did not contain any major surprises and

represented a maturing of several programs which may hold a more positive

attitude fluid resupply. In January of 1985, a series of presentations were

made to satellite manufactures TRW, Hughes, and Ford. The presentations were

intended to direct feedback on the proposed demonstration program and

financial analysis. A survey question, prepared under Rockwell IR&D, was send

to over 90 members of the space community in the U.S. and aboard. The

contacts included domestic and foreign space agencies, aerospace and

communications companies, and space insurers. The survey attempted to find

new candidate users of fluid transfer services for either remote or

Orbiter-based service. Thirteen useful responses have to be received to date,

with most supporting the concept of concurrent servicing with resupply. A

summary of the results of these surveys is presented in Figure 15.

Table 3 presents a more detailed chart of NASA missions with potential

requirements for the use of ERM. The data for this survey resulted from

telephone contacts with approximately 20 programs or project offices. Some

additional information was obtained from the NASA Space Station mission data.

Many of the missions are expected to be resupplied by Space Station, using an

OMV. A total of twelve missions were identified on being candidate users.

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A aeries of on-site presentations were made to a series of satellite

suppliers. At the request of NASA HQ, ERM contract results were presented,

with emphasis on Geo satellite economic analysis, technical inputs to

resupplied satellites, and acceptable resupply demonstrations. A series of

technical and economic comments were received and documented, some comments

were common to more than one supplier. Summary conclusions from these

contacts are presented as follows:

o Need sophisticated resupply/services module and logistics support

system

o Some propellant transfer advanced development necessary

o Early performance of flight test segments appears to be a sufficient

d emons tr a ti on

o Near-term resupply of government spacecraft should convince insurers

and satellite users of concept viability

o Directly involve satellite suppliers for further technical/economic

studies

Over ninety survey forms, related to potential users of fluid transfer

services, were sent out in May. The contacts included domestic and foreign

space agencies, aerospace and communications companies, and space insurers.

Of the twenty-one responses received to date, twelve were considered to have

useful data. Figure 16 summarizes those contacts which submitted useful

responses and what the general nature of these comments were. No major

surprises were received, although some speculative missions were identified.

Strong interest in Orbiter-based resupply was voiced for the Canadian Radarsat

program and remote resupply for the TDRS program, depending on potential

costs. Tables 4 and 5 list recipients of the fluid transfer survey. The

check marks denote who provided responses and the check mark in parenthesis

denotes a response still in work. No check mark denotes no further contact of

this date.

Eased on the resupply mission model and the potential users contacts six ERM

scenarios were developed (Figure 17). In Scenario 1, the resupply module

provides the OMV with propellant for transportation to GEO. This mode was

found to be more economical than using an expendable OTV to transport the OMV

to GEO. Propellant transfer (and concurrent module exchange servicing, when

required) is accomplished directly from the OMV after separation from the

resupply module in Scenario 2. This uses a reusable cryogenic OTV for Geo

transport.

In Scenarios 3 through 6, the resupply module operates in LEO and is reusable.

Some ERM's will be parked at the fuel depot (Scenario 5) and

participate in fuel scavenging operations from the STS (Scenario 6).

Some ERM's will perform supply and/or top-off operations.

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From examining the resupply traffic model, fluid requirements and operational

scenarios, a variety of sizes are possible for the resupply module (Table 6).

• Our perception of the sizing requirement for he Orbiter in-bay tanker has been

included for comparison purposes. The range of sizes for all these options

are roughly similar (i.e. 2-5 klbs for hydrazine, 2-7 klbs for bi-propellant)

as the same general market is being serviced. Operational distinctions, such

as the need for concurrent servicing, Orbiter flexibility, and resupply

locations, will determine the type of remote or in-situ resupply.

The major conclusions from Task 1 are summarized as follows:

o Fluid transfer an integral part of space infrastructure

o Major customer for fluid transfer are other infrastructure elements:

OMV, OTV, Space Station

o Other potential users support concurrent servicing

o Selection of MMH/NTO as baseline fluid still justified by user needs

o Need for resupply of other fluids

o Module sizing determined by operational scenarios

o Decoupling of tankage from resupply module maybe advantageous to

development effort

o Adequate MMH/NTO available from OMV initially

o Focus on key technical/operational risk elements

o Lower front-end cost for timely development

The updating of the mission model provided no surprise but the expansion of

contract scope to include space infrastructure assets emphasized major area of

fluid transfer requirements. End user satellites typically required

concurrent servicing with resupply, while assets such as the OTV, OMV and

Space Station required resupply along (while acquiring their payloads on

separate mission). The spectrum of resupply missions possible was summarized

into a set of operational scenarios with drivers of resupply module sizing.

This resulted in a confirmation of the selection of MMH/NTO as the baseline

fluid while down-sizing the earlier resupply module design (45,500 Ibs of

bi-propellant). An ERM of some 7 klb of bi-propellants was deemed sufficient,

while the larger version serves as a growth version of this design. The

decoupling of tankage from the resupply module was also discovered as a

potential means for concentrating on the technical issues associated with

fluid transfer while preparing flexibility in an ending market for fluid

transfer.

2.2 Concept Definition (Task 2)

A review of the Task 2 Study results is presented in this section of the final

report. A summary of the eight subtasks is listed below.

l) The conceptual propellant and pressurant resupply system defined in

support of the Space Platforms Expendables Eesupply NASA Contract

Study (NASA8-35618), has been reevaluated and revised to incorporate

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new subsystem design options and operational requirements. In

addition to fluid resupply subsystem design changes, which include 3

separate disconnect/line purging designs, the bi-propellant resupply

quantity of the Expendables Resupply Module (E.R.M. ) has been reduced

from the original 45,500 Ibs. to 7,000 Ibs.

2) The on-orbit transfer of propellants with pumps will make possible

propellant scavenging and extend satellite life. Three different

types of pumps were considered and compared: centrifugal, gear and

peristaltic. This subtask reports the power requirements versus flow

rate at different P's for the propellants NTO, MMH, and ^Ify..

Power requirements for the centrifugal pump were determined from

reference (3)- Power requirements for the gear pump were determined

from data supplied by the Sundstrand Corporation.

3) An analysis of two pressurant transfer methods (cascade and

compressor) was performed and the optimum method and supply

configuration for resuppling a 7000 Ib bi-propellant's pressurant

system was determined.

4) Subtask 4 requires that the instrumentation requirements for

microprocessor control and system status monitoring during

resupply/quiescent periods be defined. Reference 10 was used as a

baseline. The types of instrumentation considered are pressure

transducers, thermocouples, valve position indication feedback

switches, flowmeters, and contact sensors.

5) Commonalities with on-going programs were assessed and presented.

6) The objectives and tests to verify these objectives are presented for

the Middeck Ullage Transfer Experiment (MUTE).

7) The test objectives and tests to verify these objectives are

presented for the Flight Demonstration Test Vehicle (FDTV).

8) The purpose of the ERM ground test programs is to develop the Flight

Demonstration Test Vehicle to be used to conduct the fluid transfer

during the ERM flight demonstration.

A re-evaluation of the ERM module's design reference mission scenario (Task l)

(Reference l), has redefined the expendables resupply capability of the

vehicle. The ERM., developed under the SPER. Concept Definition Study, is now

foreseen as a potential growth version (i.e., OMV extended missions kit, space

based propellant depot) of a more modestly sized module design. A ERM of, at

best, 7,000 Ibs. of bi-propellant will adequately cover resupply needs through

the year 2010. The 7,000 Ibs. bi-propellant capacity will require a tank

volume approximately equal to 6 Shuttle Orbiter RCS propellant tanks.

In addition to the changes in the expendables resupply capacity, minor changes

have been incorporated into the conceptual ERM fluid transfer system design.

The system design changes include l) isolating pressurant tanks, 2) regulating

compressor inlet pressure, 3) additional pump and compressor plumbing for ERM

propellant and pressurant tank refueling, and 4) three separate

disconnect/line purge subsystem designs.

3474e -44-
High pressure solenoid valves have been positioned at each pressurant tank

outlet for a more efficient management of resupply pressurant. Isolating

pressurant tanks will ensure a sufficient supply of high pressure helium or

nitrogen gas, for multiple engagement resupply missions.

In-house data, concerning the operational characteristics of high pressure gas

compressors, reveals that regulating the compressor inlet pressure will be

necessary, to obtain the desired design compression ratio and outlet pressure.

A rearrangement of the conceptual fluid system schematic's pump and compressor

subsystem plumbing allows the use of these components in the refueling of the

EEM's propellant and pressurant tanks. This modification is illustrated in

Figures 18, 19 and 20.

Three separate disconnect/line purge subsystems are illustrated in the -

enclosed figures. Figures 18 and 18a illustrate a pressurant purge design,

the same design as illustrated in the SPER Conceptual Design Study. Figure 19

illustrates a vacuum pump purge design, and Figure 20 a dribble volume purging

system. The fluid transfer schematics illustrated in Figures 19 and 20 also

use a quick disconnect design similar to the NASA/Fairchild, triple inhibit,

standardized refueling coupling (NASA 9-17333 umbilical). A detailed

description of the operational characteristics of each disconnect/line purge

design is documented in reference 2. Table 7 describes the major advantages

and disadvantages of each purge subsystem design.

Evaluation of Propellant Resupply Pumps

Propellant transfer can be accomplished by one of the following three

methods: ullage recompression, ullage vent and ullage exchange. In the

ullage recompression method, the flow rate is determined by the amount of heat

that the propellant tank can dissipate. The flow rates for the other two

methods are usually limited by the type of propellant acquisition devise

(vane, screen, diaphragm) in the receiver tank. Due to the different flow

rates required, the power requirements in this report were determined for

several P values at flow rates up to 10 GPM.

Although a peristaltic pump can provide enough P for both an ullage vent and

ullage exchange transfer it cannot accomplish an ullage recompression

transfer. This is due to the fact that a head rise of at least 250 psi is

required to complete an ullage recompression propellant transfer.

Consequently, the peristaltic pump was not longer considered as a candidate

for propellant transfer since it is capable of a maximum head rise of 100 psi.

Figures 21, 22, 23 and 24 present gear pump power requirements versus flow

rate at various head rise values for the propellants MMH, NTO, a system of MMH

and NTO, and hydrazine. Data supplied by Sundstrand Corporation was

extrapolated to complete the family of head curves seen in Figures 21, 22, 23

and 24. These curves are for a 10 GPM optimum flow rate gear pump and

represent power required to pump and do not include motor efficiency.

According to a representative of Motronics, motor efficiency could vary

anywhere from 70 to 90$ depending on how the motor is cooled, what materials

are used to make the motor, lamination, air flow (if any), heat dissipation,

etc. Due to all of the above factors, the motor efficiency will not be taken

into account in the power consumption calculated for the 10 GPM optimum gear

pump.

3474e -45-
Line Pressurant Purge

Vacuum Pump Purge

Dribble Volume Purge

Purge receiver and ERM

QDs, plus some line

leading to QDs.

Quick Purge Time

No Dribble Volume

Low System Weight

No Dribble Volume

Purges Receiver and

ERM QDs

Low System Weight

Simple Design

Quick Purge Time

High System Weight

More QDs to Mate

- Long Purge Times

- Problems with

Compressive Heating

Limited Life Pumps

- Receiver and ERM QDs

are not purged.

Some propellant will

go into the

environment 0.2 cc

TABLE 7. DISCONNECT/LINE PURGE OPTIONS, MAJOR ADVANTAGES AND DISADVANTAGES

3474e -46-
ORIGINAL PAGE IS

OF. POOR QUALITY

FUD PANEL

He Trms

Ull Line

Purge

Ull Trani

MTO Line

Purge

•TO Trans

NTO Line

Purge

NTO Trans

Servicing, Checkout, f.i

Fill/Dram Lines have bee

omitted for clarity

GEHSh-C

Waste Storage Tank

FIGURE 18. FLUID TRANSFER SCHEMATIC, PRESSURANT PURGE DESIGN

Servicing. Checkojt,

Fill/Drair Lmes *i*

OPitted for c'arit>

Waste Storage Tank

FIGURE 18a. FLUID TRANSFER SCHEMATIC, PRESSURANT PURGE

DESIGN, WITHOUT SELF REFUELING PLUMBING

3474e -47-
-- FWO -PANEL

He Trans > ^

UH Trans

NTO Trins

AFT PANEL

M* Trans

Ull Trans

KTO Trans

Waste

Tarn

Storage ( )

nk \ /

ServlC'rj, rh»:lo.

F<ll/Dre'" L">"! n

b*«r wutea for c

FIGURE 19. FLUID TRANSFER SCHEMATIC, DRIBBLE VOLUME PURGE DESIGN

waste Storage^Tank

FIGURE 20. FLUID TRANSFER SCHEMATIC, VACUUM PUMP PURGE DESIGN

3474e -48-
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The power requirement is accurate at the optimum flow rate of 10 GPM but only

an approximation at the other flow rates. The justification for the above

approximation is that gear pumps are usually only run at their optimum flow

rates, and even though they can be run at off-design speeds, no data was found

for this case. A major advantage of the gear pump is its ability to handle a

wide range of head values. This 10 GPM optimum flow rate pump weighs

approximately five pounds (not including motor weight).

Figures 25, 26, 27 and 28 present pump power requirements of a centrifugal

pump versus flow rate at various head rise values for the propellents MMH,

NTO, a system of MMH and NTO, and hydrazine. Figures 25, 26, 27 and 28 were

duplicated from reference 3*

As can be seen in Figure 23, the power to pump NTO and MMH with a gear pump at

a flow rate of 9 GPM and a P of 250 psi is about 2900 watts. The power to

pump NTO and MMH with a centrifugal pump under the same conditions is about

2700 watts (Figure 27). The difference in power required by a centrifugal and

by a gear pump is not significant; therefore, further study into the power

consumption of the motors that will be coupled to these pumps is required.

The major advantage of the gear pump is its ability to provide a wider range

of P's than a centrifugal pump can provide. A gear pump can provide P's

up to 1500 psid while the centrifugal pump is not capable of much more than

delta P's in the 350 psid range at its optimum operating speed. Therefore,

centrifugal pumps are inefficient when run at off-design speeds. Variations

in operating speed are not critical for the gear pump.

Micropump Corporation believes they can make a magnetically coupled gear pump

to pump at 3 GPM with a P of 300 psia. To run the motor and pump

combination requires 1730 watts to pump MMH and 2880 watts to pump NTO. This

pump (with motor) weighs about twenty pounds, is 7-1/2 inches in diameter and

is about twenty inches long. To achieve pressure rises greater than 300 psia,

multiple stage pumps will be required.

An analysis of two pressurant transfer methods was performed and the optimum

method and supply configuration for resuppling a 70001b bi-propellant's

pressurant system was determined.

The two transfer methods analyzed were a cascade only method and a cascade

with compressor method. The cascade only method involves multiple supply

tanks at a higher pressure being opened sequentially to a receiver tank until

pressure equalization occurs; however, this transfer requires considerably

larger quantities of helium than is required for the resupply. Therefore,

larger and heavier tanks are required for this method. In contrast, the

cascade with compressor method requires only enough helium for the resupply

since it performs a cascade transfer first followed by the use of a compressor

to complete pressurization. However, this method requires a battery system

and two compressors for redundancy which increase system weight.

Supply system volumes and weights were calculated for three supply tank

pressures of 6000, 5500, and 5000 psia and 1 to 6 supply tanks. The optimum

transfer method and supply tank configuration was determined to be as

follows: A cascade only transfer with four 4200 cubic inch supply tanks at

6000 psia for a total system weight of 277 Ibs.

3474e -53-
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3474e -57-
The Concept Definition Study, Ref. 7, identified two possible pressurant

transfer approaches: a cascade or blowdown only method and a cascade with

compressor method. The cascade only method involves multiple supply tanks at

a higher pressure being opened one at a time to a receiver tank until pressure

equalization occurs, thus pressurizing the receiver from 500 psia and 70 F to

a level of 3600 psia and 70 F. The advantages of this method are its simple

operating procedure, sequential valve openings, and its minimal equipment

requirements, tanks and isolation valves only. However, due to this pressure

equalization technique, a larger volume of helium is required to be carried

along than is required for transfer; thus adding more system weight in the

form of larger tanks and excess helium. The cascade with compressor method,

although, requires only the amount of helium necessary for the resupply since

it performs a cascade transfer initially then utilizes a compressor to finish

receiver pressurization. However, this method also has its disadvantages; it

requires a battery and two compressors, for redundancy, which add considerably

to the system weight and complexity.

The supply system was sized for a total receiver volume of 13000 cubic inches

since it was determined that this volume of helium at 3600 psia would expel a

worst-case amount of 7000 Ibs of bi-propellant and still have the required 500

psia residual. A previous task analysis, Ref. 4 had determined that during

the helium transfer the supply tank would cool to approximately 40 F and the

receiver tank will heat up to 160 F. Since this occurs in relatively short

periods of time when compared to the overall transfer time, it was assumed

that the transfer was made isothermally at 40F and 160 F. This provides an

important worst-case assumption in that by assuming the receiver tank is at

160 F instead of 70 F, less helium was calculated to have been transferred

than actually was transferred due to the pressure difference between the two

temperatures, therefore, a larger supply system size and weight was

calculated to perform the transfer. Figure 29 depicts this pressure

difference for the cascade with compressor transfer method.

The actual sizing procedure for the cascade only method used the computer

program utilized in the analysis performed for Ref. 6 and, in particular,

involved guessing a total supply system volume and calculating the final

receiver tank pressure from the ideal gas equation with compressibility then

comparing that to the desired level of 3600 psia. System volumes for the

three supply tank pressures of 6000, 5500, and 5000 psia and 1 to 6 supply

tanks were calculated. Table 8 presents these values.

The sizing of the supply system for the cascade with compressor method was

easily accomplished once the three supply pressures were known. Since this

transfer method utilizes a compressor to achieve the desired receiver

pressure, only the amount of helium required to resupply the receiver is

needed in the supply system. Therefore, by using the ideal gas law and

knowing the receiver requirements and supply system pressure, the system

volume was calculated. Table 9 also presents these values.

The weight of a supply tank was calculated from an equation obtained during a

phone conversation with Mr. Newhouse, Ref. 8.

(Pb*V)/675,000=Wt

3474e -58-
I I

I I -

TOTAL SYSTEM VOLUME (CUBIC INCHES)

NUMBER | | 6001

OF | | SUPPLY

1 1 CASCADE ONLY

1

1 | | 26800

II

2 | | 19400

3 | | 18000

1 1 II

4 | | 16800

I I

5 1 1 16500

I I

6 | | 16200

) PSIA

PRESSURE

CASCADE W/ COHP

7800

7800

7800

7800

7800

7800

550C

SUPPLY

CASCADE ONLY

37500

24800

23400

21200

20500

19800

) PSIA

PRESSURE

CASCADE W/ COMP

8500

8500

8500

8500

8500

8500

500C

SUPPLY

CASCADE ONLY

61000

35000

31500

28000

27500

27000

) PSIA

PRESSURE

CASCADE U/ COHP

9400

9400

9400

9400

9400

9400

TABLE 8. SYSTEM VOLUME COMPARISON

II

IISYSTEM WEIGHT UBS)

NUMBER 1 1 6001

OF | | SUPPLY

|| CASCADE ONLY

1 || 417

1 1 1 1

2 || 308

I I

3 || 291

I I

4 || -277

I I

5 || 277

I I

6 || 277

} PSIA

PRESSURE

CASCADE U/ COMP

380

• 376

378

380

384

388

5501

SUPPLY

CASCADE ONLY

535

360

342

• 318

313

311

3 PSIA

PRESSURE

CASCADE W/ COMP

384

• 380

382

385

388

392

500!

SUPPLY

CASCADE ONLY

790

460

419

• 379

377

375

) PSIA

PRESSURE

CASCADE U/ COMP

389

* 385

337

389

393

397

• • ASTERISK DENOTES OPTIMUM CONFIGURATION

TABLE 9. SYSTEM WEIGHT COMPARISON

3474e. -59-
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3474e -60-
where:

Pb = Tank Burst Pressure (psia) = Max Operating Press*!.5 (S.F.)

V = Tank Volume (cubic inches)

Wt = Tank Weight (ibs)

The total system weight for the cascade only method includes the weight of the

tanks, helium, and 2 isolation valves per tank. The total weight for the

cascade with compressor method includes tanks, helium, 2 compressors

(redundancy), a battery system, and 2 isolation valves per tank. The

compressors were assumed to weigh 100 Ibs each based upon estimations by

various companies for a 2 scfm flowrate and 2/3 hp compressor. The tank

solenoid valves were assumed to weigh 2.3 Ibs each based upon valves currently

being used in the aerospace industry. The battery system weight varied

according to the amount of helium required to be transferred. This weight

included the weight of the Li/TiS battery, the coolant system, the power

conditioning system, and the system's structure. To calculate the battery

system weight the following equation was used:

(WtHe * Pcomp)/Pconv = Battery system wt

where:

WtHe = Mass of Helium required to be transferred (ibm He)

Pcomp

 a Power required by compressor per pound helium (Wh/lbm He)

Pconv = Efficiency of battery system (Wh/lbm battery)

The power required by the compressor was determined to be 4-02.6 Watt-hours per

pound of helium transferred. The efficiency of the battery system was

determined to be 40 Wh per pound of battery once the coolant system, power

conditioning system, and system structure was included. Table 10 presents

this system weight, the mass of helium required to be transferred, and the

amount of power required by the compressor.

Figures 30, 31 and Table 9 present the total system weight vs. the number of

supply tanks for the two transfer methods. A line has been drawn through the

optimum tank configuration in the figures and an asterisk denotes them in the

table. Beferring to Table 8 and Table 9, these optimum configurations are:

four 4200 cubic inch supply tanks (16800 cubic inches total) at 6000 psia with

a system weight of 277 Ibs for the cascade only method and two 3900 cubic inch

supply tanks (7800 cubic inches total) at 6000 psia with a system weight of

376 Ibs for the cascade with compressor method.

The overall optimum transfer method and tank configuration can now be

identified to be a cascade only transfer with four 4200 cubic tanks at 6000

psia since it weighs 99 Ibs less than the other method.

Instrumentation Requirements for Microprocessor Control

Pressure Transducers; Pressure Transducers were identified to monitor helium

tank inlet pressure, propellant tank ullage pressure, compressors/pumps inlet

and outlet pressures, quick disconnect inlet pressure, and waste tank

pressure. The minimum pressures to be monitored within the receiver vehicle

during resupply are helium and propellant tank pressures. The pressure

3474e -61-
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transducer locations are shown in Figure 32. A brief description of each

instrument and how it will be used is shown in Table 11. The pressure

transducers on the quick disconnect that will be used for leak detection

(pressure decay) are not discussed in this letter as they are recommended to

be an integral part of the disconnect. The sampling rate for adequate

microprocessor control is recommended to be 100 samples per second during

dynamic resupply operations and then reduced to one sample per second during

quiescent periods.

Thermocouples; Thermocouples were identified to support PVT gaging, helium

loading, pump/compressor operation, and vent operations. The minimum

temperatures to be monitored within the receiver vehicles during resupply are

helium and propellant tank locations. The thermocouple locations are shown in

Figure 32 and a brief description of each instrument and how it will be used

is shown in Table 12. The thermocouples on the quick disconnects are not

discussed in this letter as they are recommended to be an integral part of the

disconnect. The sampling rate for adequate microprocessor control is

recommended to be one sample per second during resupply and quiescent periods.

Valve Position Indication Feedback Switches - Each valve will be equipped with

redundant position indication feedback switches.

Flowmeters - Two flowmeter feedback channels are required to monitor flow in

both the forward and reverse flow directions.

Contact Sensors - Contact sensors are necessary to verify proper mating of the

quick disconnect panel and can be of the simple spring loaded type.

Commonalities with On-Going Programs and the EBM

The expendable resupply module (EBM) is compatible with two on-going programs

and a few flight demonstrations. The two programs of interest are: the

orbital maneuvering vehicle (OMV) and the orbital spacecraft consumables

resupply (OSCES). As is seen in Figure 33, OSCRS, OMV, and the ERM are

expected to be operational by 1990, 1991-2, and 1993 +, respectively.

Therefore, a reduction in development testing of the ERM can be expected by

using the developmental results of OSCRS and OMV.

The following development tests are required for the ERM.

A) Rendezvous with active or passive targets.

B) Docking with subsequent release.

C) Mating and alignment for the fluid interface with latching.

D) Propellant transfer - basic function, multiple transfers, connection

reliability with leakage monitoring. Major test components -

propellant pump, coupling, and tank acquisition device with ullage

control.

E) Pressurant transfer - basic function, multiple transfers, connection

reliability with leakage monitoring. Major test components -

compressor and coupling.

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A few compatible areas between the OSCRS and the ERM, are as follows:

1) Some form of docking mechanism will be required for the ERM, OSCRS

will demonstrate the use of the MMS berthing system.

2) OSCRS will perform a propellant transfer in which a manual coupling

device will be used. This will allow testing/demonstration of the

coupling system.

3) Heat transfer determinations during ullage recompression in the

receiver diaphragm tank will help to establish optimum flow rates.

A manual coupling is being developed by Fairchild. The coupling is to be

developed as a standardized coupling for ^Ify, MMH, NTO, GN£, GHe, and

CgCl^F^ (trichlorotrifluroethane). OSCRS' usage of the coupling will be

to transfer hydrazine to the GRO in 1990, but by using appropriate seals all

listed fluids can be transferred. The Fairchild coupling will be available by

mid 1986 to transfer ^Ify and it could be available by the end of 1987 to

transfer other fluids. The coupling is manually engaged and five inhibits

must be manually actuated in sequential order to allow a propellant transfer.

In Figures 34 & 35 the inhibits are labeled T1-T3 for the tanker side and

S1-S3 for the spacecraft side. The inhibits cannot be opened until the two

halves of the coupling are engaged or the coupling disengaged before the

inhibits are manually closed. The manual engagement and disengagement force

is about 36 pounds. A rough leak detection system will be performed by

pressure transducers at the indicated ports shown in Figure 34. A weight

analysis indicates that the tanker and spacecraft coupling will weight 13.0

and 10.8 pounds, respectively. Fairchild also has funding to start

preliminary designs for the automated coupling in early 1987. Fairchild

indicated that the present manual design will not be automated as is.

Another possible area of compatibility that OSCRS may have with the ERM is the

development of a propellant pump. At this time the use of a propellant pump

is only considered a option for hydrazine transfer.

The OMV is a Shuttle-based/launched, remotely controlled, free-flying vehicle

capable of performing a wide range of on-orbit services. The initial OMV

(launched about mid 1991) will perform payload placement and/or retrieval in

LEO when delivered to standard STS orbit by the orbiter. As the space program

advances and mission needs materialize the reference design OMV can be

upgraded by modifications to perform more sophisticated missions, to be

space-based at either LEO or GEO, and to accommodate DOD missions.

There are two areas of compatibility between the ERM and the OMV, they are as

follows.

1) Rendezvous with a cooperative or non-cooperative target.

2) Remote docking with subsequent release.

The first OMV docking will probably occur with the Shuttle in mid 1992. All

development required for rendezvous and remote docking will be performed on

the OMV before the ERM.

3489e -70-
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ORIGINAL PAGE fS

Of POOR QUALITY

FIGURE 35. FAIRCHILD MANUAL COUPLIHG

3489e -72-
The ERM will be able to use only limited developmental work from OSCRS and the

OMV. OSCES will provide information on the propellant coupling, propellant

transfer, leak detection, and docking with the MMS berthing system. OMV will

provide information on the rendezvous and remote docking and release with

other spacecraft.

On-Orbit Middeck Ullage Transfer Experiment

MUTE ia an Orbiter middeck experiment designed to investigate micro-gravity

fluid transfer dynamics using a water solution, capillary/screen type tanks,

micro-gravity gaging system, and a small reversible fluid pump. The

fundamental objectives of MUTE are as follows:

1) To demonstrate the concept of ullage transfer in a micro-gravity

environment.

2) To demonstrate the propellant management device's (PMD's) ability to

position and control the ullage bubble by observing the ullage

transfer process. The PMD will include a ullage/propellant separator

and will be placed in each tank.

3) To demonstrate a micro-gravity quantity gaging system's accuracy in a

micr o-gravi ty envir onment.

These objectives are to be met by using off-the-shelf equipment and by

modeling the structure and tanks after the storable fluid management

demonstration (SFMD).

MUTE consists of a supply tank, a receiver tank, a PMD in both the supply and

receiver tank, a reversible pump, a micro-gravity gaging system, a flow meter,

and associated valves, filters, lines, pressure gauges, and temperature

sensors. MUTE has no power or command interfaces with the Orbiter; and is

basically self-contained, with the exception of an overboard vent connected to

the Orbiter waste management system. The entire experiment will be designed

to fit in place of several lockers in the orbiter middeck. Support equipment,

from STS, will include cameras and associated lights, power to be supplied by

Orbiter outlets, (for the cameras and associated lights) a micro-gravity

accelerometer measurement system (MGAMS), and a data recorder.

Control of MUTE on-orbit will be accomplished by the crew through a sequence

of manual valve and switch operations. The small fluid pump is the motive

force for a series of test fluid transfers into and out of the receiver tank.

Data sources will consist of the video cassette recorder, 35 mm still photos,

MGAMS acceleration data, data recorder, and astronaut comments/logs. All data

inputs will be correlated.

A preliminary schematic to perform the recommended tests is presented in

Figure 36. The recommended tests are presented in Table 13, with expected

test objectives in Table 14. The following is a few general operating

comments:

l) Prior to tests, the crew will connect the MUTE's vent line to the

Orbiter waste water overboard vent system.

3489e.' -73-
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Test No. Transfer Fluid from Supply Tank

to Receiver Tank

Transfer Fluid from Receiver Tank

to Supply Tank

1. Transfer to dry tank at low flow

rate with ullage exchange. Stop

filling at each 10% of mass

transfer.

2. Transfer to wet tank at low flow

rate with ullage recompression

until tank is 50$ full.

3. Transfer to wet tank at moderate

flow rate with ullage exchange.

4. Transfer to wet tank at high

flow rate with ullage exchange.

5. Transfer to wet tank at low,

moderate, and high flow rates

with ullage exchange.

6. Transfer 75-802 of fluid to

wet tank at moderate flow rate

with ullage exchange during

various Orbiter accelerations.

Stop flow and accelerate until

screen breakdown. Wait for

rewetting of screen to occur

before transferring remaining

fluid.

Transfer fluid back to supply

tank at a low flow rate with

ullage exchange.

Transfer fluid back to supply

tank at a low flow rate with

ullage recompression.

Transfer fluid back to supply

tank at a moderate flow rate

with ullage exchange.

Transfer fluid back to supply

tank at a high flow rate with

ullage exchange.

Transfer back to supply tank

at a moderate flow rate with

ullage exchange.

Transfer fluid back to supply

tank at a moderate flow rate

with ullage exchange.

TABLE 13. RECOMMENDED TESTS

3489e -75-
Objective Test No.

1) To evaluate PMD's Ullage Control

propel1 ant/ullage separation X X X X X

wetting pattern of PMD X X X X X

wetting time of PMD X X X X X

orbiter acceleration effects X

2) To evaluate Zero-Gravity Gaging System

accuracy determinations

steady state flow rate X X X X

variable flow rate X X

orbiter acceleration X

response time

steady state flow rate X X X X

variable flow rate X X

orbiter accelerations X

comparison to PV gaging system X

TABLE 14. MATRIX OF TEST OBJECTIVES

Component Weight (Ibs)

each system

SFMD 250 Ibs

removal of valves, gas cylinder, ."id cylinder -50 Ibs

pressure transducers 2 12

temperature sensor 1 7

pump 5

batteries 25

flow meters 3 6

zero-gravity gaging system 40

+10% 33

Total 363 Ibs

TABLE 15. MUTE SYSTEM WEIGHT

3489e -76-
•*;

2) After eactetest the test fluid will be returned to the supply tank

'/,_ before going on the next test.

-s

3) All test operations are controlled manually by the crew using the

proper valves, switches, and instrumentation.

MUTE's system weighi!:is presented in Table 15. Since MUTE is modeled after

SFMD, SFMD's system weight was used as a baseline by removing weights of

components not recftfired and then adding in new component weights plus an extra

10%. MBTE's systefti Weight is expected to be about 360 Ibs.

Orbiter^argo Bay Fluid Transfer Flight Demonstration

FDTV wi'll be used i^ the Orbiter cargo bay to demonstrate a micro-gravity

remote'Tluid transfer. There will be two brass board vehicles capable of

being separated or'docked using the remote manipulator subsystem (RMS). The

upper portion of th% test vehicle, the receiver vehicle, piggybacks on the

supply vehicle throtigh deployable payload latches. The RMS system is utilized

to separate the receiver vehicle, by use of a grapple fixture attached to the

receiver vehicle, to simulate remote docking and mating of the two vehicles.

A snare type end effector, incorporated at the fluid transfer interface of the

supply vehicle, accomplishes final mating of the fluid transfer interface

panels on each brass board test vehicle.

The primary test objectives to be performed on the FDTV are as follows:

1) To demonstrate proper docking and alignment of the fluid transfer

r

- interface during a remote operation.

2) To demonstrate proper QD mating, leak integrity verification,

operation^"and purging with safing, before, during, and after a fluid

transfer.

3) To evaluate^ the PMD's ullage control and separation capability during

a propel!ant transfer.

4)*> To evaluate propel!ant transfer by the methods of ullage transfer and

ullage recornpression. Different pumped flow rates, and the ability

of the tanker to resupply itself will also be examined during the

ullage transfer tests.

5) To evaluate-' pressurant transfer by two methods: 1 ) a pressurant

tank cascade and 2) by compressor usage.

6) To evaluatep-the micro-gravity quantity gaging system for both

accuracy determinations and variable flow rate effects.

7) To demonstrate thermal and remote system control.

These test objective^ are to be met by using the brass board test components

and as many off-the-shelf components as possible.

3489e -77-
FDTV contains four forward RCS propellent tanks, consisting of two receiver

and two supply tanks, for the MMH/NTO propellent transfer. All four tanks

will be aligned with the +X-axis of the Orbiter's coordinate system during

launch and contain a 6Q% nominal load (55% of total volume). Only minor

modifications to the pressurant inlet area of the RCS propellant tank will be

made so that ullage/propel!ant separation will occur. The propellant transfer

will be accomplished by the use of three small propellant pumps. As is seen

in Figure 37, the MMH side will provide pumping to the receiver tank and will

also resupply the tanker side by using only one pump. The NTO side,

Figure 38, will transfer propellant by use of two pumps, one on the receiver

vehicle and one on the supply vehicle with simpler valving. There are two

different pressurant transfer systems for comparison and demonstration

purposes. The MMH side will perform a pressurant cascade from four 4800 psia

tanks to fill a RCS helium tank to 3600 psia, while the NTO side .will use a

compressor to perform the pressurant transfer after a blowdown pressure

equalization transfer.

Control of FDTY on-orbitwill be accomplished by the crew through a sequence

of remote operations. The RMS will be the motive force during mating and

demating steps, requiring intensive crew attention; but the resupply process

should only require crew monitoring with a minimum interaction between the

crew and the FDTV.

Table 16 presents the recommended tests to be performed by FDTV to satisfy the

test objectives presented in Table 17. Each of the four tests are broken down

into their basic steps required for the completion of each test. The first

test will examine the docking, mating, leak integrity verification, purging,

safing, and the demating process. The second test will perform the pressurant

transfer by the cascade method and compressor method, and perform an ullage

recompression propellant transfer. The third test will perform an ullage

transfer at a low flow rate. The last test will demonstrate the ullage

transfer at a higher flow rate. Propel!ant transfer will occur back and forth

between the receiver and supply tanks during each test.

FDTV's system weight is presented in Table 18. The expected dry weight is

3309 Ibs. To this was added a 15% potential growth factor plus the required

propel!ant and pressurant requirements for a total payload weight of 6735 Ibs.

The purpose of the ERM-GTP is to develop the Flight Demonstration test vehicle

(FDTV) to be used to conduct the fluid transfer during the ERM Flight

Demonstration (ERM-FD). The FDTV shall be capable of conducting remote mating

and demating operations, performing propel!ant transfer by means of "ullage

transfer" and "ullage recompression" methods, together with the transfer of

pressurant by means of "cascade" and compressor methods. In addition the test

vehicle shall contain a propellant management device (PMD) capable of

separating gas and liquid, (for ullage transfer tests) and a zero-gravity

fluid quantity gaging system (gaging system) to determine the amount of

propellant in the tanks. To date, the ERM study has defined the need to

develop three critical components for this transfer demonstration.

Reiterating, these components are the propel!ant transfer pump, the pressurant

compressor and the mating fluid couplings. It should be noted that for this

program it shall be assumed that the NASA will develop the couplings for this

application (including remote configuration) and that this development effort

shall produce qualified units applicable for the ERM-FD test vehicle. The

3489e -78-
Test No. Steps in Each Test

1 ) Unlatch receiver vehicle from the supply vehicle.

2) Using the RMS move receiver vehicle away from the supply vehicle

and back again for docking step.

3) Docking and alignment of interface.

4) QD mating.

5) QD leak integrity verification

6) QD purging and safing.

7) Demating of QD.

8) Using the RMS move receiver vehicle away from the supply vehicle

and back again for docking step.

1) Docking and alignment of interface.

2) QD mating.

3) QD leak integrity verification.

4) Transfer pressurant by cascade method on the MMH side.

5) Transfer pressurant by compressor method on the NTO side.

6) Using supply pump transfer MMH by the ullage recompression

method from 60 to 90% of the filled volume.

7) Using supply pump transfer NTO by the ullage recompression

method from 60 to 90% of the filled volume.

1) Allow pressurant to equilibrate the pressure on the supply and

receiver propel 1 ant lines and tanks; then prepare for the ullage

transfer method.

2) Using the supply pump transfer MMH by ullage transfer from the

receiver tank back to the supply tank, from 90 to 30% of the

filled volume.

3) Using the receiver pump transfer NTO by ullage transfer from the

receiver tank back to the supply tank, from 90 to 30% of the

filled volume.

4) QD safing and purging.

5) Demating of QD

6) Disengagement and separation of vehicles.

1) Docking and alignment of interface.

2) QD mating.

3) QD leak integrity verification.

4) Using the supply pump resupply the MMH receiver tanks by ullage

transfer at a higher flow rate from 30 to 90% of the filled

volume.

5) Using the supply pump resupply the NTO receiver tanks by ullage

transfer at a higher flow rate from 30 to 90% of the filled

volume.

6) QD safing and purging.

7) Demating of QD.

8) Disengagement of and separation of vehicles.

9) Latch receiver vehicle to supply vehicle.

TABLE 16. RECOMMENDED TESTS

3489e -79-
1)

2)

3)

4)

5)

6)

7)

8)

Objectives

To Demonstrate Docking and Alignment of Interface

To Demonstrate QD Integrity Verification, which include:

QD Mating

Leak Integrity Verification

Purging and-Safing

To Evaluate the PMD's Ullage Control with:

Propel! ant/Ullage Separation

Propellant Supply Capability

To Evaluate Propel! ant Transfer by:

Ullage Transfer with:

Pump Usage

Different Flowrates

Tanker Resupply

Ullage Recompression with:

Thermal Control

Pump Usage

To Evaluate Pressurant Transfer by the:

Cascade Method

Compressor Method

To Evaluate Micro Gravity Gaging System for:

Accuracy Determinations

Flow rates Effects

To Demonstrate Thermal & System Control

To Demonstrate Demating and Separation of Vehicles

Test Numbers

123 4

X X

X X

X X

X X

X

X

X

X

X

X

X

X

X

X

X

X

X

X X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

TABLE 17. MATRIX OF TEST OBJECTIVES

34896 -80-
Subsystems Weight

1) Fluid Transfer System

propel1 ant tank 300

pressurant tank 287

waste tank 8

pump 60

compressor 200

disconnects 240

valves, regulators, etc. 325

pressure end temperature sensors 90

2) Interface panels

end effector 65

grapple fixture 44

face plates 38

latch 20

electrical connects 8

3} Structural System

frame and support 665

fittings and latches 624

thermal blankets and covers 125

4) *Avionics/Power

Wire harness 40

regulation & distribution 100

thermal control 30

on-board data management 40

Subtotal330 9

+15% Growth 496

5) Propellent and Pressurant 2930

Total Payload Weight 673"5~lbs.

*Allocation - Analysis to be performed at a later date.

TABLE 18. FLIGHT DEMONSTRATION TEST VEHICLE SYSTEM WEIGHT

3489e -81 -
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3474e -83

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-i*H- check valve

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-£ji— compressor

-Q_QJ- dual burst disk

-j/1/1- dual pressure regulator

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—Q]— liquid propel!ant detector

manual valve

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{] servicing/test port

temperature sensor

valve

-84-
gaging system, which will be developed under a separate contract, shall be

customized to the test vehicle tankage and verified during the ERM-GTP. The

development of the liquid/gas separator unit of the PMD shall be tested and/or

analytically assessed as a separate unit, with no plans to assess its

operation during the system fluid transfer ground testing.

Ground Test Program (GTP)

The ERM-GTP will be divided into four major tasks:

1) Pump and compressor specifications, including supplier design and

fabrication of breadboard test units.

2) Component testing of pump and compressor breadboard units. Testing

and/or evaluation of the liquid/gas separator together with the

gaging system configuration requirements will be performed during

this time period.

3) FDTV design and fabrication. During this time frame the test and

checkout of the coupling mating assemblies to assure alignment and

determine engagement and separation loads will be performed.

Procedures for coupling operation will be assessed.

4) FDTV development and verification test program. This period is

divided into two test series with allowance for potential modifications between the two test series. During this period, low level

support using breadboard development test apparatus may be advantagous and cost effective to isolate unique component problem areas.

The ERM-GTP proposed timeline is presented in Figure 39.

Test Program

The Test Program is divided into two major areas (1} breadboard testing of

prototype (pump and compressor) units to assess component features unique to

space application and for proof of concept of required advanced

state-of-the-art component features and (2) Brassboard testing of the FDTV

transfer system using flight components and hardware.

Breadboard Tests

Prior to these tests, a final assessment of component design from various

potential suppliers shall be performed and the most promising of these

components will be selected for breadboard (B/B) development. The breadboard

test results will be used to enable the selection of a unique component design

for each fluid transfer application. If, based upon the B/B tests there are

more than one viable component design for a unique application both designs

will be further evaluated during the FDTV brassboard test program. The

breadboard test results will also enable the finalization of component design

specification requirements.

The B/B test effort will be concentrated chiefly to the development of the

propel!ant pump, gas compressor (includes assessment of gas and compressor

cooling requirements), and the liquid/gas separator device.

3489e -85-
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Propel1 ant Pump Tests

The breadboard test setup shown in Figure 40 shall be the test-bed for the

evaluation of potential pump configurations to be used in the FDTV. Flight

configuration flowmeters, if available, shall also be incorporated in order to

enhance test results. Three types of pumps shall be evaluated (Gear,

multi-stage centrifugal, and piston), however, only the two types with the

highest potential will be selected for breadboard development. All breadboard

testing will be performed utilizing water and/or freon-113 as the test fluid.

The test program will consist of 12 flow tests using the ullage transfer

technique (4 flow tests each at 5, 10, and 15 gpm), and 12 flow tests using

the ullage recompression technique (4 flow tests each at 1, 3, and 5 gpm).

The primary test objectives are to assess the pump performance and, if

necessary, recommend design changes based upon the test results. Detailed

test objectives are listed below:

1) Evaluate pump performance using ullage transfer method and various

flowrates and tank pressure conditions.

2) Evaluate pump performance using ullage recompression method and

various flowrates and tank pressure conditions.

3) Define any pump deficiencies and provide recommended changes to the

flight-type pump designs.

4) Obtain preliminary data on ullage recompression receiver tank ullage

temperature effects (secondary objective).

5) Pump power requirements assessment.

6) Determine the pump magnetic coupling drive performance at various

pump speeds and pressure differentials (optional only if magnetic

coupling is used).

Gas Compressor Tests

The compressor breadboard test setup is shown in Figure 41 . A series of

helium flow tests at various flowrates, temperatures and pressures will be

performed to assess compressor performance and design parameters. The

objectives for these tests are as follows:

1) Evaluate compressors performance at various temperatures, pressures,

and flowrates. See test matrix in Table 19.

2) Evaluate compressor cooling requirements.

3) Obtain preliminary data on requirements far receiver tank gas cooling.

4) Compressor power requirements assessment.

5) Evaluate the compressor magnetic coupling drive performance at

various flow conditions. (Optional only if magnetic coupling is

used).

3489e -87-
TEST CONDITIONS TEST NO.

(Pressure « PSIA, 1,2 3 4 5 6 7 8 9,10

Flowrates = SCFM)

1. COMPRESSOR

INLET PRESSURE

A. 500; B. 700 A

MAXIMUM OUTLET PRESSURE

A. 3000; B. 5000 A, B

COMPRESSOR FLOWRATES

A. 5; B. 10 A, B

GAS INITIAL INLET TEMPERATURE

A. 70°F, B. 1000F; A

2. AMBIENT CONDITIONS

PRESSURE

A. Ambient; B. Vacuum A, B

TEMPERATURE

A. 7QOF; B. 100°F A

3. HEAT EXCHANGER

FLUID TEMPERATURE

A. 400F; B. TBD A, B

LIQUID FLOWRATE

A. TBD A

A A A B B B B

A, B A, B A, B A, B A, B A, B A, B

A, B A, B A, B A, B A, B A, B A, B

B A B A B A B

AAAAAAA , B

B A B A B A B

A A, B A A, B A A, B A, B

A A A A A A A '

TABLE 19. COMPRESSOR B/B TEST MATRIX

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6. Define any compressor design deficiencies and provide recommended

changes to the flight compressor design.

Liquid/Gas Separator Tests

The liquid/gas separator testing will be performed at the component level and

will be restricted to the verification that this unit will maintain its

structural and mechanical integrity when subjected to launch environments.

Brassboard Ground Test Program

This test program will provide the final phase of the fluid transfer system

development and will also provide the verification tool to establish the FDTV

fluid transfer capability for the flight demonstration.

Test Vehicle Description

The Test Vehicle Description (T/V) shall be designed and fabricated to be

installed in the space shuttle payload bay and shall be in two separate

sections; the tanker side, which will be fixed to the orbiter, and the

receiver side which will simulate a spacecraft being resupplied. The two

sections shall be launched in the demated configuration. Once in orbit the

mating of the two sections is performed by the RMS and the grapple and end

effector (Figure 42). the T/V shall therefore be designed to perform

satisfactorily after being exposed to launch and landing environments. The

FDTV fluid schematics are presented in Figures 37 and 38 for the fuel (MMH)

and oxidizer (^Htty) systems respectively, the helium transfer systems

for use with the compressor and cascade methods are also included. A

preliminary FDTY design concept showing the major component layout is

presented Figure 42. The test Vehicle (T/V) configuration shall be the same

as for the flight demonstration except that the Grapple Fixture and End

Effector need not be included for the conduct of the brassboard tests. The

Test Vehicle will include seven major hardware and component items critical to

the fluid transfer operation. These are; propel!ant tanks (4), helium

pressurant tanks (7), fluid couplings (6), gas compressor (1), zero-gravity

quantity gaging system (4) including four flowmeters as backup, propellent

pump (3), and miscellaneous si nail tanks (4) to purge and safe the couplings.

A heat exchanger will be required to cool the compressor oil, compressor

housing (requires circulating pump), and the compressed gas. The heat

exchanger shall be designed to tie into one of the space shuttle payload bay

cooling loops using Freon-113 as the fluid media.

This circuit of the compressor heat exchanger will also require a small

circulating pump. The ground tests will include a flowmeter in this circuit

in order to assess flow requirements.

In addition the FDTV will include a microprocessor for data processing,

recording and display and an electrical panel consisting of switches,

component status indicators, displays, etc, to enable control of the fluid

transfer operations.

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Test Program Description

The test program will consist of a series of tests directed at developing a

flight test vehicle which will be utilized in a fluid transfer demonstration

aboard the space shuttle payload bay. The results from these tests will

develop and assess the fluid transfer system capabilities as well as provide

the tool to define the flight demonstration procedures and regimes.

Final development of the gas compressor and propel!ant transfer pump will be

accomplished during this testing phase. In addition, operation, verification

and safing of the fluid couplings during mating and demating, together with

the verification of the gaging and flowmeter system will be demonstrated.

The test program will include ambient and vacuum testing (zero-gravity tests

will not be performed). All thermal sensitive areas such as gas compression

during helium gas transfer and during ullage recompression propellent transfer

will be diagnosed in a vacuum atmosphere. Except for the zero-gravity

environment, the proposed flight demonstration timeline an procedures will be

performed as verification and also as a tool for determining system

predictions and behavior.

Brassboard Ground Tests

These tests shall be designed to meet the following objectives:

1. Propellent pump verification.

1-A. nowrate - Evaluate pump performance at low average, and high

flowrates.

1-B. Pressure conditions - Evaluate pump performance at various

inlet, outlet and delta pressures.

1-C. Temperature conditions - Evaluate pump performance at nominal

and high temperature conditions (fluid and ambient).

2. Fluid coupling verification.

2-A. Mating/demating - Define mating and demating loads alignment

requirements, and procedures. Electrical connector mating and

demating will also be evaluated and requirements established.

2-B. Verification of Interfaces - coupling interface verification

procedures shall be established and demonstrated for mating,

fluid flow preparation, safing and demating operations.

3. Gas compressor verification.

3-A. Fi11 rates - Determine compressor performance at low, average and

high fill rates.

3-B. Pressure Conditions - Determines compressor performance at

various inlet, outlet, and delta pressure conditions.

3489e -92-
3-C. Temperature Conditions - Determine compressor performance at

various temperature conditions.

3-D. Compressor/gas heat exchanger - Determine heat exchanger

performance (heat exchanger requirement established from

breadboard tests).

4. Zero-gravity quantity gaging system verification - This gaging system

shall be developed by NASA/JSC under a separate contract.

4.A Installation - The gaging system shall be installed and

calibrated to the FDTY propel1 ant tanks.

4.B Performance - The gaging system performance shall be verified

during flow tests using flowmeters and PVT methods.

4.C Zero gravity simulation - Gas ingestion and tank orientation

tests will assess gaging sensitivity to varying fluid location.

5. Helium Tanks.

5.A Tank servicing procedures shall be developed and verified.

5.B The liquid/gas separator design and installation shall be

analytically verified for flight application (i.e. launch

environment effects and design concept).

6. Helium Tanks.

6.A Heat exchanger - verify the performance of the helium tank heat

exchanger. The need for the heat exchanger shall be determined

from the breadboard tests.

7. Helium Transfer

7.A Verify and demonstrate the optimum transfer configuration using

the cascade method.

7.B Determine transfer characteristics at various temperatures,

flowrates and pressures using a gas compressor.

8. Propel!ant Transfer

8.A Verify and demonstrate the optimum transfer procedures using the

ullage transfer and the ullage recompression transfer techniques

for various flowrates, temperature and pressure conditions.

8.B Assess and establish the capability of the micro gravity

quantity gaging system.

9. Flight experiment procedures.

9.A Establish and define the flight demonstration gas and propellent

transfer procedures and estimate the timeline required.

3489e -93-
9.B Develop a semi-automatic control of the transfer operations

utilizing a micro processor in conjunction with minimal

astronaut involvement.

9.C Define orbiter interface and orbiter support requirements such

as compressor cooling (heat exchanger) operations.

10. Test Vehicle flight worthiness

10.A Flight environment - FDTY flight worthiness shall be verified

analytically with vibration/acoustic testing support performed

only if analysis deems it necessary.

10.B Instrumentation & Components - all instrumentation and

components shall be verified during testing, and calibration

data together with component checkout data verified after

completion of the transfer tests. These data shall be part of

the FDTV data package.

10.C System leak checks - The system shall be verified to be leak

free within the specified requirements prior to and after

completion of the test program.

TASK 2 CONCLUSION:

The key results of this task analysis are as follows:

1) New subsystem design options have been incorporated into the

conceptual propellent and pressurant resupply system defined in

support of the S.P.E.R. Concept Definition Study (MASA8-35618).

Design changes include 1) isolating pressurant tanks, 2) regulating

compressor inlet pressure, 3) additional pump and compressor plumbing

for E.R.M. self refueling, and compressor plumbing for E.R.M. self

refueling, and 4) three separate disconnect/line purge designs.

2) An E.R.M. bi-propel!ant capacity of 7,000 Ibs., tank volume

approximately equal to 6 RCS propel!ant tanks, will be adequate for

resupply needs through the year 2010.

3) Power requirements are specified for MMH, NTO and Nofy at

different flow rates and head rises or both gear and centrifugal

pumps.

4) Recommend selection of the gear pump over the centrifugal pump for

its versatility over a wide range of P's.

5) The tank volumes required to resupply a 13000 cubic inch receiver

tank to 3600 psia using two different pressurant transfer methods and

three different supply pressures were calculated and presented in

Table 2.

6) The total transfer system weights were calculated from the determined

system volumes and are presented Table 3.

3489e -94-
7} The overall optimum transfer method and supply tank configuration was

determined to be a cascade only transfer with four 4200 cubic inch

supply tanks (16800 cubic inches total) at 6000 psia for a system

weight of 277 Ibs.

8} For a resupply only tanker, with a propellent load of 7000 Ibs, the

cascade method of pressurant resupply is preferred, but for the

maximum system versatility a compressor is required.

9) The ERM will still require the following development tests:

A) Mating and alignment for the fluid interface with latching.

B) Propellent Transfer - Major test components - propellant pump,

automated coupling, and tank acquisition device with ullage

control.

C) Pressurant Transfer - basic function, multiple transfers,

connection reliability with leakage monitoring. Major test

components - compressor and automated coupling.

10) The Middeck Ullage Transfer Experiment (MUTE) is designed to test the

concept of ullage transfer and the associated PMD, and micro-gravity

gaging system. These objectives were examined and a series of tests

proposed to verify the ullage transfer process.

11) The Flight Demonstration Test Vehicle (FDTV) is designed to test

several aspects of remote on-orbit fluid transfer. The fluids to be

transferred will include NTO, MMH, and Helium. The proposed

demonstration will test docking, QD mating, leak verification,

purging, safing, ullage control and separation during a propellant

transfer, cascade and compressor pressurant transfer, and thermal and

remote system control.

12) The ERM ground test program will be divided into four major subtasks.

•

A) Pump and compressor specifications, including supplier design

and fabrication of breadboard test units.

B) Component testing of pump and compressor breadboard units.

Testing and/or evaluation of the liquid/gas separator together

with the gaging system configuration requirements will be

performed.

C) FDTV design and fabrication. The test and checkout of the

coupling mating assemblies to assure alignment and determine

engagement and separation loads will be performed. Procedures

for coupling, operation will be assessed.

D) FDTV development and verification test program. Low level

support using breadboard development test apparatus may be

advantageous and cost effective to isolate unique component

problem areas.

3489e -95-
2.3 Programmatics (Task 3)

Summary

Rockwell's recommendations in the follow-on ERM Study for the advanced

development and flight demonstration differ significantly from the initial

study.

The significant difference between our original and current recommendations

are presented in Table 20. In this study, Rockwell defined and weighed four

criteria for selecting the best advanced development and flight demonstration

program. These criteria and their weights were approved by the NASA Contracting Officer Representative at the mid-term review. Based on these criteria,

Rockwell recommends the early flight demonstration program which culminates at

the end of the advanced development program. We reviewed 12 different

programs and from these selected the one that maximized the NASA's goals, as

defined by the NASA approved criteria and weights we applied to the selection

process.

Rockwell's recommended program includes two flight experiments as well as the

advanced development of the critical components to enable remote resupply

servicing. The first flight experiment (Mid-Deck Ullage Transfer Experiment -

MUTE) is a low cost validation ($1.8 million ) of ullage transfer in a

micro-gravity environment. Not only will this experiment remove the

uncertainty from the propel 1 ant management devices ability to position and

control the ullage bubble, but it demonstrates to potential users NASA's

resolve to develop remote resupply servicing. Stimulating user support for

remote resupply servicing is one of NASA's goals, and hence it was a criteria

in selecting our recommended program.

The second flight experiment (Flight Demonstration Test - FDT) is the

validation of the advanced developed for remote resupply servicing.

Rockwell's low cost approach is to "fly the brass board", the hardware

developed and tested during the advanced development phase. The primary

flight test objective of FDT is to verify remote propel!ant transfer in a

micro-gravity environment.

Groundrules and Assumptions

o Customer Goals:

* Least cost advanced development and flight demonstration program.

Budget target is $12M.

Relative importance of goal is 50-60%.

* Reduce technical risk and stimulate user support for the ERM

program.

This is to be accomplished by an advanced development and

flight demonstration programs.

Relative importance of goal is 40-30%.

* Earliest ERM program start date.

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Relative importance of goal is 10$.

o All cost estimates are in constant 1985 dollars.

o STS launch costs are not directly included in the cost estimate, but

launch costs were considered in selecting the top-and-bottom truss

experiment over the side-by-side truss experiment in the Orbiter

cargo bay.

o A nominal six month waiting period was assumed for Orbiter payload

integration activity and this cost is not included except for direct

contractor support.

o NASA costs, such as astronaut training and test facilities at KSC for

pre-flight propellant transfer testing are not included.

Rationale For Recommendation

Rockwell has based its selection for the advanced development and flight

demonstration on the NASA's selection criteria and their relative weights.

The most important criteria was cost (weight of 50 to 602). All options were

designed to fit within the scope of NASA's target of $12 million. The most

costly programs were eliminated in the selecting process. The surviving

options were within $1 million of one another. So, cost was eliminated as a

criteria in the subsequent selection process. The next two most important

criteria were highly correlated, so they were combined into a single composite

criteria. Flight demonstration options of greater technical scope and

challenge tended to reduce the technical risk of the subsequent ERM program.

And because of their greater visibility and realism, technical challenging

demonstrations were also more likely to increase user support for the NASA's

remote resupply program. The least important criteria was an early ERM phase

C/D start (weight of 10%). Thus, with cost eliminated, the dominate criteria

was technical risk and user support.

A comparison between our recommendations in the initial ERM study and the

follow-on study are shown in Table 21. Rockwell's current recommendation (the

early flight demonstration option) greatly contributes to most important

criteria reducing technical risk and soliciting user support (weight of 75-80%

once the cost criteria is removed). The only disadvantage is a later ERM

phase C/D start. However, due to our efforts to find a low cost flight

demonstration, ("fly the brass board" discussed later), our program plan is

shortened by six months. Thus, our low cost program cuts what would otherwise

be a 12 months delay in the ERM phase C/D start to only a 6 months delay (in

comparison to the late flight demonstration option originally recommended).

The schedules of the early and late flight demonstration options from the

initial ERM study and the early flight demonstration option currently

recommended from the follow-on ERM study are displayed in Figure 43.

For purposes of comparison, all program schedules have been laid on a common

calendar scale. Schedule differences for the flight demonstration and ERM

phase C/D start can be observed by comparing the milestones. The program with

the earliest flight demonstration option is the one we currently recommend.

3489e ^ -\ -98-
Early Flight

Demonstration Option

Late flight

Demonstration Option Criteria

Importance* of

Criteria

Technology Uncertainties

removed before ERM

program starts.

Accelerated by

18 months

6 months

Parallel effort exposes

ERM programs to risk

of schedule delay and

cost increases

IS month delay

6 month earlier

ERM program start

Technical

Risk

User

Support

ERM

Program

Start

75-802

25-20%

*Normalized from criteria approved by customer at mid-term

briefing (cost 60-50%, Support & Risk 30-40%, Program Start 10%)

TABLE 21 . THE EARLY FLIGHT DEMONSTRATION OPTION

HAS SUPERIOR ADVANTAGES AND ONLY MINOR DISADVANTAGES

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It is 18 months earlier than the program we proposed in the initial ERM

study. The ERM phase C/D start that we currently recommend is only six months

later than the program we recommended originally.

In summary, significantly more benefits have been gained (less technical risk

and 18 month earlier flight demonstration) than have been lost (six month

later ERM phase C/D start).

Program Master Schedule

The master schedule for the program recommended by Rockwell, leading to the

earliest flight demonstrations is presented in figure 44. The initial effort

is to specify the hardware requirements for the ERM program. With the

requirements for the ERM program defined, the design of the flight experiments

and the development of the critical ERM components can start. Ideally, MUTE

can fly on schedule in 22 to 24 months after the program starts, and the

Flight Demonstration Test can fly in 8 to 12 months afterwards.

The timing in development of several related fluid transfer components could

affect this schedule. The micro-gravity gaging system (NASA, JSC) would be an

ideal component to include on the Flight Demonstration Test If the timing was

coincidental, the micro-gravity gaging system could be flight-tested on the

Flight Demonstration Test without the expense of an additional independent

flight demonstration of its own. This probably will not happen unless the

program is accelerated by six to nine months so that the flight hardware is

ready by early to mid FY 1988. The function of the micro-gravity gaging

system, if not available for the Flight Demonstration Test, could be performed

by fluid gaging.

The Quick Disconnect (NASA, Langley) is essential to demonstrate remote

resupply in the Flight Demonstration Experiment. It is currently scheduled to

be available in mid-FY 1988, and its schedule may also need to be accelerated

slightly.

Mid-Deck Ullage Transfer Experiment

The significant characteristics of the Mid-deck Ullage Transfer Experiment are

presented in Table 22. The experiment was selected because it met the NASA's

main criteria of low cost, reduction in technical risk, and potential to

estimate potential user interest and support in remote resupply servicing.

The time required to design, fabricate and assemble, and test MUTE is twelve

months. Ground test time is only scheduled for three to five months. Since

this experiment is designed mainly to test the effects of a micro-gravity

environment on the ullage transfer, only limited testing can be performed on

the ground. The schedule and manpower estimate is presented in Figure 45.

The schedule includes an effort to insure that the test plan is

human-engineered since the orbiter crew will be directly responsible for

running the experiment. Crew training is planned during the nominal six-month

period of payload integration activity at KSC.

3489e -101 -
Test Objective Verify ullage transfer with water

solution in a micro-gravity environment

Customer Benefit Reduce technical risk in Propellent

Transfer Experiment

Early flight demonstration to stimulate

potential user support low cost ($1.8M)

Test Schedule 21-24 months after start of advanced

development program.

TABLE 22. MID-DECK ULLAGE TRANSFER EXPERIMENT SIGNIFICANT CHARACTERISTICS

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In case of non-critical failures in the ullage transfer that might occur

during the flight experiment, there are twelve months available to diagnose

and correct them. The modifications would then be retested on the Flight

Demonstration Test. Thus, the back-up to MUTE is the Flight Demonstration

Test.

This schedule is based on a success-oriented program that assumes no major

difficulties are encountered.

The grass-roots (detailed engineering) cost estimate is uncertain because the

three key variables - direct labor hours, labor cost rates, and direct

material - are uncertain. To provide a cost range, likely values were

estimated for each variable as presented in Table 23. The best combination of

cost conditions results, by definition, in the low cost estimate. The worst

combination of cost conditions results, also by definition, in the high cost

estimate. The combination of most likely conditions results in the most

likely cost estimate.

These cost estimates only reflect the estimating error given the program plan

as reflected in the MUTE schedule and manpower estimate. The program plan,

and hence the cost estimate, reflects a success-oriented program with no major

difficulties encountered.

The low cost estimate reflects that there is only a 10% chance that the cost

of the experiment would be lower, and the high cost estimate reflects a 10%

chance that the cost would be higher. The probability curve between these two

extreme points was spread by a 35/65 Beta distribution function (ogive curve)

and is displayed in Figure 46. The expected cost (50% cummulative probability

that the actual cost could be less than or greater than this cost estimate) is

$1.87M. The best estimate cost from Table 23 was $1.71 million. The true

best estimate is somewhere in the vicinity of these two discrete estimate. If

the midpoint is taken, then the best estimate for the MUTE is 31.8M

(Figure 46).

In Rockwell's effort to meet NASA's primary goal of a low-cost advanced

development and flight demonstration program, MUTE was designed to low cost.

Some of the factors we have identified to control costs are listed as follows.

o Simple Experiment with Limited Test Objective

o Non-Rigorous Packaging Constraint - Design to Fit Mid-Deck Locker

Space

o Minimum Electronic and Microprocessor Control - Maximize Orbiter

Right Crew "Hands-On" Effort

o Human Engineering Effort to Simplify Mechanics of Experiment

o Maximum Use of GFE (Existing Lockers and Video-Recorder Equipment)

o Minimize Data Recording by Visual Inspection and Video-Recording

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Besides a simple, low cost experiment to verify that the ullage transfer works

as expected in a micro-gravity environment, MUTE also contributes to NASA's

objectives to reduce technical risk and to support the marketing effort to

gain user interest and support for remote resupply services.

Flight Demonstration Test

The significant characteristics of this experiment are presented in Table 24.

The Flight Demonstration Test represents the cummulation of the advanced

development. The cost of advanced development was previously estimated at

$7 million. Thus, the marginal cost to "fly the brassboard" as Rockwell

proposed, is an additional $3.7 million.

This flight demonstration meets the NASA's cost target. In addition, it

reduces the technical risk by verifying remote fluid transfer in a

micro-gravity environment. Furthermore, the second flight demonstration

builds on the momentum of the first flight experiment (MUTE) to gain potential

user interest and support for remote resupply services.

The schedules for the brassboard and the subcontracted effort for developing

the pump and compressor are integrated. The schedule and manpower estimate is

shown in Figure 47. The development effort by NASA for the remote quick

disconnects and micro-gravity gaging system are assumed available during the

brassboard assembly to support this schedule.

The component subcontractors are scheduled to deliver a prototype unit for the

pump and compressor sixteen months after program start. The flight

demonstration brassboard will test remote fluid transfer initially with a

prototype of the critical ERM components and water solution fluids. Four to

six months later, the flight test hardware will be delivered, integrated, and

tested on the brassboard.

During the nominal six-month period for payload integration at KSC, we plan to

perform fluid transfer testing with the actual hypergolic fluids planned for

the flight demonstration.

The Flight Demonstration Test is scheduled for 30 - 36 months after program

start, with a six-month post analysis of the results. Given a completely

successful flight test, the ERM phase C/D program could start as early as 30

months from the start of advanced development and flight demonstration

programs.

The grass-roots {detailed engineering) cost estimate is uncertain because the

four key variables - direct labor hours, labor rates, direct material , and the

subcontracted cost to develop the pump and compressor - are uncertain. To

provide a cost range, likely values were estimated for each variable and

presented in Table 25. The best and worst combinations of cost conditions

results in the low and high cost estimates, respectively. The most likely

combination of cost conditions results in the most likely cost estimate.

3489e -108-
Test Objective Verify remote propellent resupply with receiver and

supplier hardware.

Customer Benefit Reduce technical risk by verifying that all the

critical ERM components (quick disconnects, pumps,

compressor, and propellent management devices) work

as designed in a micro-gravity environment.

Early flight demonstration stimulates potential

user support.

Low cost (S10.7M) due to:

o Integration of advanced development and flight

demonstration program, and

o Less test hardware with our plan to "fly the

brassboard".

Test Schedule 30 - 36 months after start of advance development

program.

TABLE 24. FLIGHT DEMONSTRATION TEST SIGNIFICANT CHARACTERISTICS

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These cost estimates only reflect the estimating error, given the program plan

as reflected in the schedule and manpower estimates for the Flight

Demonstration Test. The program plan however, and hence the cost estimate,

reflects a success-oriented program with no significant difficulties

encountered.

The low cost estimate reflects that there is only a 10% chance that the cost

of the experiment could be lower and the high cost estimate reflects that

there is only a 10% chance that the cost could be higher. The probability

curve between these two extreme points was spread by a 35/65 Beta distribute

in function (ogive curve) and is displayed in Figure 48. The expected cost

(50% cummulative probability that the actual cost could be less than or

greater than this cost estimate) is $11.5 million. The best estimate cost

from Table 25 was $10.0 million. The true best estimate is somewhere in the

vicinity of these two discrete estimates. If the mid-point is taken, then the

best estimate for the Flight Demonstration Test is $10.7 million.

Given cost target of $12M for the Advanced Development and Flight

Demonstration Program, Rockwell has designed a program that not only to meet

this goal but has cut the cost of the Flight Demonstration Test sufficiently

to fund an additional flight test (the MUTE). The significant cost reduction

items are summarized as follows:

o Use Flight Demonstration Test Article as Systems Ground Test Article

o Avoid Duplication

o Reduces Early Demonstration Schedule by Six Months

o Borrow Orbiter RCS Propellent and Pressurant Tankage

o Less Structure than Side-by-Side Approach

o Lower Weight

o Use Lower Cost Valves and Regulators than Orbiter

o Potential Use of MPESS for Lower "Half" of Structure

The three most significant cost avoidance items account for most of the cost

reduction.

One, we propose to "fly the brassboard". Instead of two separate test

articles, one for advanced development and one for the flight demonstration,

will be developed we plan to develop the technology for remote fluid transfer,

using the structure designed for the Flight Demonstration Test.

Two, we propose to "borrow" the RCS propel!ant and Helium pressurant tanks

from the Orbiter logistical spares inventory instead of purchasing them

outright as a separate cost item. This could avoid costs of up to two million

dollars.

3489e -113-
ORIGINAL PAGE IS

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Third, we propose a top-and-bottom test structure rather than our original

side-by-side test structure. Not only will the weight and, hence, structure

cost be less, but the flight transportation expense will decline because the

space usage in the Orbiter cargo bay is cut in half. Presently, the width of

our Flight Demonstration Test structure is only slightly wider than the RCS

tank, or 39 inches.

The Flight Demonstration Test is a simple, low cost experiment to verify

remote propellent transfer between a simulated supplier and a receiver

spacecraft. This demonstration was selected because its contributes is to

NASA's objectives to reduce technical risk and gain user interest and support

for remote resupply services.

3489e -115-
REFERENCES:

1. Pace, S. N., "Development Scenario and Requirements for Expendables

Resupply Module," RI IL AE-MOA-85-178, June 3, 1985.

2. Saxton, H. A. Saxton, "Pump Transfer of Propellant, Task Analysis -

Fluid Transfer Study. A Transmittal of Results for Subtask 2.6, IRSD

Project 252, RI IL 287-201-85-111, August 22, 1985.

3. H. A. Saxton, "Pump Transfer of Propellant, Task Analysis - Fluid

Transfer Study. A Transmittal of REsults for IR4D Project 283,

November 21 , 1984.

4. Compressor Transfer of Pressurants, Task Analysis - Storable Fluid

System Conceptual Design, I.L. 287-201-84-143, November 20, 1984.

5. Development Scenario and Requirements for Expendable Resupply Module,

I.L. AE-MOA-85-178, June 3, 1985.

6. Transmittal of the Rationale for Pressurant Resupply Process, I.L.

287-201-84-114, October 16, 1984.

7. Space Platforms Expendables Resupply Concept Definition Study, STS

85-0174, April 1985.

8. Phone conversation on July 9, 1985 with Norm Newhouse,

Supervisor-Brunswick Corporation Product Design. (402) 464-3211.

9. "Establishment of Subtasks and Time Line for the Expendable Module

Follow-On Study." RI IL #287-201-85-063, F. 0. Chandler to J. J.

Gernand, April 25, 1985.

10. "System Status Sensing and Measurement Techniques for On-Orbit,

Propellant and Pressurant Transfer System - A Transmittal of Results

for subtask 2.5 IR&E Project 286; H. A. Saxton to J. J. Gernand, July

27, 1984.

11. "Orbital Maneuvering Vehicle - Preliminary Definition Study",

Marshall Space Flight Center, January 1985.

12. "Preliminary Design Review for Standardized Refueling Coupling",

Fairchild Control Systems Company, April 4, 1985.

13. Scoter, C. W., "Preliminary Flight Demonstration Options and Their

Relative Costs" IL AE-MOA-85-1'84, dated June 4, 1985.

14. Clement, J. L., "Cargo Systems Manual: SFMD", JSC-19484, July 1,

1984.

15. Scoter, C. W., "Preliminary Flight Demonstration Options and Their

Relative Costs" IL AE-MOA-85-184, dated June 4, 1985.

16. Matus, S. J., "Expendables Resupply Brass Board Test System" IL

385-121-85-065, dated June 11 , 1985.

3489e -116-