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AfNASA-CK-178819) SPACE PLA1FCEH EXPENDABLES ~ ~N86-247~32RESOPPLY CONCEPT DEFINITION SIODY FinalReport, Jan, - Oct. 1S85 {Bocksellinternational Corp.) 116 p He &06/8F A01 Dnclas
v CSC1 22B G3/18 42931
STS85-017AADDENDUM
FINAL REPORT
SPACE PLATFORM EXPENDABLES RESUPPLYCONCEPT DEFINITION STUDY
FOR PERIOD JANUARY 1985 - OCTOBER 1985
DECEMBER 1985
CONTRACT NAS8-35618
Rockwell InternationalSpace TransportationSystems Division
https://ntrs.nasa.gov/search.jsp?R=19860015261 2020-07-02T11:42:25+00:00Z
Addenda To:Space Platform Expendables Resupply
Concept Definition StudyFinal Report
STS85-0174Contract MAS 8-35618
For Period January 1985 - October 1985
347 4e -1 -
CONTENTS
Section Title Page
1.0 Executive Sunmary 9
1.1 Introduction 91.2 Background 91.3 Supplemental Study Scope and Objectives 121.4 Sunmary of Results 12
2.0 Study Results 28
2.1 System Requirements and Scenarios (Task 1 ) 282.2 Concept Definition (Task 2) 422.3 Programmatic and Development Planning (Task 3) 96
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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 11Concept (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 20Requirements
8 Mid-Deck Ullage Transfer Experiment 21
9 Payload Bay Flight Demonstration and Ground Test 22Article
10 Comparison of Advanced Development/Flight 24Demonstration Options
11 Advanced Development and Flight Demonstration 27Program (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, 47Without Self Refueling Plumging
19 Fluid Transfer Schematic, Dribble Volume Purge 48Design
20 Fluid Transfer Schematic, Vacuum Pump Purge Design 48
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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 51and MMH
24 Required Gear Pump Power at Various Delta P for 52Hydrazine
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 63Method
31 System Weight Comparison Cascade With Compressor 64Resupply Method
32 Temperature and Pressure Location 68
33 Representative Interaction of Remote Resupply 69Development
34 Standardized Refueling Coupling Schematic 71
35 Fairchild Manual Coupling 72
36 Mid-Deck Ullage Transfer Experiment Schematic
37 Flight Demonstration Test Vehicle 82Schematic - MMH HALF
38 Flight Demonstration Test Vehicle 83Schematic - NTO HALF
39 Expendable Resupply Module Test Program 86
40 & 41 Component Test Schematic 89
42 Flight Demonstration Test Vehicle (Preliminary 91Design Concept)
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ILLUSTRATIONS (Continued)
Fi gure Page
43 Comparison of Advanced Development/Flight TOODemonstration Options
44 Advanced Development and Flight Demonstration 103Program (Master Schedule)
45 Mid-Deck Ullage Transfer Experiment Schedule 104and Manpower Estimate
46 Mid-Deck Ullage Transfer Experiment 106
47 Flight Demonstration Test Schedule and 112Manpower Estimate
48 In-Bay Flight Transfer Experiment 114
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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 39Surveyed
6 Space Platform Expendables Resupply Module Sizing 43Options and Drivers
7 Disconnect/Line Purge Options, Major Advantages 46and Disadvantages
8 System Volume Comparison 59
9 System Weight Comparison 59
10 Battery System and Power Requirements for Cascade 62With 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 99Superior Advantages and only Minor Disadvantages
22 Mid-Deck Ullage Transfer Experiment Significant 102Characteristics
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TABLES (Continued) •
Tab!e Page
23 Mid-Deck Ullage Transfer Experiment Cost Range 107Estimate
24 Flight Demonstration Test Significant 109Characteristics
25 Flight Demonstration Test Cost Range Estimate 110&
111
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ACRONYM DICTIONARY
AXAF Advanced X-Ray Astrophysics FacilityB/B BrassboardBOL Beginning of LifeCREP Cosmic Ray Experiment ProgramDDT&E Design, Develop, Test and EngineeringDMSP Defense Meteorlogical Satellite ProgramDRM Design Reference MissionDOD Department of DefenseDOMSAT Domestic SatelliteEOS Earth Observing SystemERM Expendables Resupply ModuleETR Eastern Test RangeEUVE Extreme Ultraviolet ExplorerEVA Extra Vehicular ActivityFDTA Flight Demonstration Test ArticleGEO Geosynchronous OrbitGP-B Gravity Probe-BGPM Gallons Per MinuteGRO Gamma Ray ObservatoryGSE Ground Support EquipmentGSTDN Ground Space Flight Tracking and Data NetworkHe HeliumJPL Jet Propulsion LabJSC Johnson Space CenterLaRC Langley Research CenterLDR Large Deployable ReflectorLEO Low Earth QrbitMIL Man-in-LoopMMH MonomethylhydrazineMPS Materials Processing In SpaceMSFC Marshall Space Flight CenterMUTE Mid-deck Ullage Transfer ExperimentNASA National Aeronautics and Space AdministrationNTO Nitrogentetroxide (N204.)N2 NitrogenN2H4 HydrazineOMV Orbital Maneuvering VehicleOTV Orbit Transfer VehiclePMD Propellant Management DevicePVT Pressure, Volume, and TemperatureRCS Reaction Control SubsystemRF Radio FrequencyRM Resupply ModuleRMS Remote Manipulator SubsystemROM Rough Order of MagnitudeS/C SpacecraftSDI Strategic Defense InitiativeSIRTF Shuttle Infrared Telescope FacilitySPAS Shuttle Pallet SatelliteSPER Space Platform Expendables ResupplySTS Space Transportation SystemTFU Theoretical First UnitTDRSS Tracking and Data Relay Satellite SystemUSAF SD United States Air Force Space DivisionWTR Western Test Range3474e -8-
1.0 EXECUTIVE SUMMARY
1.1 INTRODUCTION
NASA has recognized that the capability for remote resupply of space platformexpendable fluids will help transition space utilization into a new era ofoperational efficiency and cost/effectiveness. The emerging OrbitalManeuvering System (OMV) in conjunction with an expendables resupply modulewill introduce the capability for fluid resupply enabling satellite lifetimeextension at locations beyond the range of the Orbiter. This reportsummarizes a supplemental study to the original Phase A study and is presentedas addenda to that study.
1.2 Background
As background an overview of key results from the original study arepresented. The original Phase A study was a $340K effort which ran from March1984 through January 1985. Long term objectives for this activity consistedof the definition of both an operational Expendables Besupply Module (ERM) andthe definition of a flight demonstration for remote resupply. To this end thespecific study objectives for the Phase A study consisted of the definition ofuser needs and requirements, the conceptual design definition of arepresentative resupply module, a conceptual approach to the flightdemonstration of remote resupply, and the definition of a program plan tosatisfy the long term objectives.
One key result of this original study activity was the design development of aresupply'module concept. This concept and its key features are presented inFigure 1. The key features of this configuration include the capability ofthe SRM to be utilized as an ERM or an OMV extended mission kit. Sixstretched QMS tanks contain 4-5,500 Ibs. of usable propellant for resupply orOMV propulsive use. The dry weight of the ERM for bi-propellant resupply onlyis 7,960 Ibs. For helium and bi-propellant resupply it is 9,14-0 Ibs. The dryweight of the ERM to be utilized as an OMV extended mission kit would be 7,200Ibs. This includes the elimination of helium compressors and associatedbatteries and simplified propeilant management devices to meet OMV propellanttransfer requirements.
Also as a result of the original study, a concept for a Bi-propellant/Heliumtransfer flight demonstration was developed. This concept is presented inFigure 2. The principle test objective for the flight demonstration is tovalidate storable bi-propellant fluid transfer in a zero-gravity environment.Since the ERM engagement of a spacecraft is accomplished by the OMV with aremote manipulating arm end effector, the remote manipulator systems (RMS) onthe Orbiter can approximate the OMV engagement capability. The systems onboth the receiver and supplier test articles can be attached to either a MPESSor a SPAS structure (structure only without any subsystems). The Orbiter RMSwill lift the receiver test article out of the cargo bay and engage thesupplier test article. Repeated fuel transfer tests will then be performed.
Since the supplier test article has the greater need for power and electricallinks to the Orbiter, it was recommended it remain seated in the cargo bay.This minimizes the interface complexity of power and electrical links runningthrough the orbiter. The supplier test article would be autonomous with itsown power source.
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ORIGINAL PAGE ISOF POOR QUALITY
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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 servicingo But . . . far term horizon
o In LEO, in conjunction with storable propellant depoto Supports wide scope of potential operations
o Single "core" resupply module design is adaptable to both requirementcategories
o Total program cost including adequate technology development & flightdemonstration 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 throughOctober 1985. In the supplemental study the requirements task was expandedbeyond the original study to include consideration of the evolving spacetransportation infrastructure. The scope of the design section, in thesupplemental study, was limited to the consideration of the transfer processitself and to flight demonstration concepts. The scope of the supplementalstudy programmatics task was limited to the generation of cost estimates ofthe 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 selecta preferred flight demonstration program option. This was accomplished bydefining user needs, assessing the state of the development of the technology,defining demonstration concepts, and costing those concepts. User needs suchas fluid type, quantity, and schedule enable module sizing to be defined.This information was then used as input to the design tasks where thedemonstration design requirements and preliminary designs are generated. Theprogrammatics task develops the cost estimates for the resupply concepts andprovides 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 aprojection of elements from the evolving space infrastructure. Figure 3depicts an evaluation of possible key elements of the space infrastructure.These depicted elements are the ones of primary interest to the task ofexpendables resupply. In accordance with the expanded scope of this contract,a closer look has been taken at the resupply needs and interactions ofinfrastructure elements.
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In this analysis the shuttle is being used as the basis for an in-baypropellant tank (under JSC) and later the arrival of the OMV with associatedkits, such as the resupply module, servicer kit and capture kit. The OMV andassociated kits can be deployed from the ground, and/or space-based at 28.5degrees or polar orbits. These spacecraft will then be augmented by permanentorbital facilities primarily the space station with its accommodations forOMVs and OTVs. Associated with the station, but separate from it for safetyrequirements, will be a Leo depot, possibly supplied by propellant scavengingfrom the STS. Eventually, this depot should be duplicated in polar orbitswith its own OMV and associated kits.
Around the turn of the century, new developments such as the shuttle blockchange and a fully reusable OTV will likely alter the economics of supplyingpropellant in space. The Shuttle block change could affect the use ofpropellant scavening, and the OTV will generate requirements for greateramounts of fuels, raising the need for sophisticated storage facilities inspace. A secondary space transport and Launch Vehicle will likely arrive insupport of SDI. SDI elements are not sufficiently developed yet to predicttheir impact on orbital fluid requirements and have, therefore, been left outof this projection.
Figure 4 depicts, in graphic form, the fluids to be transfered on-orbit to theelements of the space infrastructure. Parametric techniques were used to makeestimates of the types, amounts and yearly fluid usage rates required by theresupply mission model. Small quantities of associated pressurants such as Heand N2 were also required, but their amounts were relatively minor incomparison to the primary fluids. In some cases, depending on the resupplytechnique used, no additional pressurants are required for the transferprocess. In the latter 1990's, bi-propellants and cryogens are the dominantrequired fluids. Hydrazine, water and liquid helium constitute a relativelystable base of demand.
In updating the resupply mission requirements extensive contacts were madewith the potential user community. In addition to in-house data base andliterature searches, three related surveys were conducted Figure 5- They weretelephone contacts with various NASA program contacts used in the developmentof the resupply mission model. The updated data did not contain any majorsurprises and represented a maturing of several programs which may hold a morepositive attitude toward fluid resupply.
In January of 1985, a series of presentations were made to satellitemanufacturers TRW, Hughes, and Ford. The presentations were intended todirect feedback on the proposed demonstration program and user benefitanalysis. A survey questionaire, prepared under Rockwell IR&D, was sent toover 90 members of the space community in the U.S. and aboard. The contactsincluded domestic and foreign space agencies, aerospace and communicationscompanies, and space insurers. The survey attempted to find new candidateusers of fluid transfer services for either remote or Orbiter-based service.Thirteen useful responses have been received to date, with most supporting theconcept of concurrent servicing with resupply. As a result of therequirements re-evaluation and discussions with potential resupply users, sixscenarios were developed for the ERM. These scenarios are depicted in Figure6. In Scenario 1, the expendables resupply module provides the OMV withpropellant
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for transportation to GEO. This mode was found to be more economical thanusing an expendable OTV to transport the OMV to GEO. Propellant transfer (andconcurrent module exchange servicing, when required) is accomplished directlyfrom the OMV after separation from the resupply module in Scenario 2. Thisuses 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) andparticipate 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 ispresented as follows:
o Fluid transfer should be considered as an integral part of spaceinfrastructure
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 fluidso Module sizing determined by operational scenarios
o Decoupling of tankage from resupply module may be advantageous todevelopment effort
o Adequate MMH/NTO available from OMV initiallyo Focus on key technical/operational risk elementso Lower front-end cost for timely development
The spectrum of resupply missions possible has been summarized into a set ofoperational scenarios with drivers of resupply module sizing. This results ina confirmation of the selection of MMH/NTO as the baseline fluid whiledown-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 largerversion serves as a growth version of the design.
A decoupling of tankage from the resupply module was discovered as a potentialmeans for concentrating on the technical issues associated with fluid transferwhile preparing flexibility in an evolving market for fluid transfer.
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Concept Definition
The Concept Definition task of the supplemental study was divided into twomajor areas. These areas were: (l) the refinement of the transfer subsystemfor the purpose of determining the engineering objectives of the fluiddemonstration 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 apropellant and pressurant transfer. The technical areas examined are asfollows: umbilical purge, pressurant transfer, propellant pump selection,NASA quick disconnect development, and pressure and temperatureinstrumentation. The results of these studies are presented in Figure 7. Twoflight test concepts were developed to provide proof of concept for on-orbitremote expendables resupply. The first is the Mid-Deck Ullage TransferExperiment (MUTE) which is presented in Figure 8. MUTE's fundamentalobjective are as follows:
1) To demonstrate the concept of ullage transfer in a micro-gravityenvironment.
2) To demonstrate the propellant management devices (PMD's) ability toposition and control the ullage bubble.
3) To demonstrate the quantity gaging system's accuracy in amicro-gravity environment.
The demonstration of a micro-gravity quantity gaging system is not critical tothe success of this flight test. However, MUTE does provide an opportunity toconduct an on-orbit test of the micro-gravity gaging system that is beingdeveloped as a separate NASA program.
The second flight test concept is the payload bay Flight Demonstration TestArticle (FDTA) which is shown in Figure 9. The Flight Demonstration andGround Test Article will be used in the orbiter cargo bay to demonstrate amicro-gravity remote fluid transfer. FDTA consists of two platforms capableof being separated or docked using the remote manipulator subsystem (RMS).The upper portion of the test article, the receiver platform, piggy-backs onthe supply platform through deployable payload latches. The RMS system isutilized to separate the receiver platform, by use of a grapple fixture, tosimulate remote docking and mating of two free-flying spacecraft. A snaretype end effector, incorporated at the fluid transfer interface of the supplyvehicle, accomplished final mating of the fluid transfer interface panels oneach test platform.
Should it become necessary to reduce the cost of the FDTA several options areavailable. One option is to conduct the tests with only one fluid either MMHor NTO. This enables a reduction in the size thereby lowering system weightand material/component procurement and fabrication costs.
Another approach to lowering flight demonstration cost is through thereduction in test objectives. Requirements analyses have indicated that acompressor may not be need during the early phases of on-orbit resupplyoperations. Elimination of the compressor from the advanced development
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program and the flight test article could result in considerable program costreductions. However, as on-orbit resupply operations mature this capacitywill 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 approachfor small systems
o Gear pump most effective in ERM flow & pressure regions
o Disconnects & liquid quantity gauging being developed underother 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 transferthermal & 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 andflight demonstration recommended from the original and supplemental studieshave been laid on a common calendar scale (Figure 10). In the initial studywe recommended a late demonstration option, because it resulted in a one-yearearlier ERM phase C/D start.
As a result, of our supplemental study, we now recommend an earlier flightdemonstration option. This is accomplished by using the FDTA as the groundsystems test and verification test bed. This resulted in an earlier ERM phaseC/D program at approximately the same cost.
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The ideal schedule leading to the flight demonstration and, hence, the ERMprogram start is presented in Figure 11. The initial effort is to specify thehardware requirements for the ERM program. With the ERM requirements defined,the design of the flight experiments and the development of the criticaltechnology components can start. Ideally, MUTE can fly in 22 to 24 monthsafter the program starts and the flight demonstration test is scheduled to fly8 to 12 months afterwards.
The timing in development of several related fluid transfer components couldaffect this schedule. The micro-gravity gaging system (NASA, JSC) would be anideal component to include on the Flight Demonstration Experiment. If the ERMand the micro-gravity gaging system development schedules could be madecompatible then, could be the Flight Demonstration Experiment expenses couldbe shared. The function of the micro-gravity gaging system, if not availablefor the Flight Demonstration Test, could be performed by flow meters.
The Quick Disconnect (NASA, JSC) is essential to demonstrate remote resupplyin the Flight Demonstration Experiment. It is currently scheduled to beavailable in mid-FY 1988, and its schedule may also need to be acceleratedslightly 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 demonstrationtest (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 DemonstrationPrograms
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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 finalreport. The topics discussed in this review are the subtasks for task 1 whichare summarized as follows:
1) Expand mission model to include space infrastructure elements2) Define parametric fluid quantities as determined by mission model3) Survey Potential on-orbit resupply users4) Reevaluate ERM scenarios from original study5) 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 ofexpendables resupply. In accordance with the expanded scope of thissupplemental study, a closer look has been taken at the resupply needs andinteractions of infrastructure elements.
In this analysis we have the Shuttle being used as the basis for an in-baypropellant tank (under JSC) and later the arrival of the OMV with associatedkits, such as the resupply module, servicer kit and capture kit. The OMV andassociated kits can be deployed from the ground, and/or space-based at 28.5degrees or polar orbits. These spacecraft will then be augmented by permanentorbital facilities primarily the space station with its accommodations forOMVs and OTVs. Associated with the station, but separate from it for safetyrequirements will be a Leo depot, possibly supplied by propellant scavengingfrom the STS. Eventually, this depot should be duplicated in polar orbitswith its own OMV and associated kits.
Around the turn of the century, new developments such as the shuttle shockchange and a fully reusable cryogenic OTV will likely alter the economics ofsupplying propellant in space. Certainly, the Shuttle block change willaffect the use of propellant scavenging, and the OTV will generaterequirements for greater amounts of cryo fuels, raising the need forsophisticated gtorage facilities in space. A secondary space transportelement, the Unmanned Launch Vehicle, will likely arrive in support of SDIdeployments and increased lunar surface activity. The logistics for theseinitiatives are not sufficiently developed yet to predict their impact onorbital fluid requirements and have, therefore, been left out of thisprojection.
The expanded mission model (Table l) reflects a current assessment of thefuture demand for operational fluid transfer missions. Utilizing resupplymodules, OMV's and OTV's, fluid transfer operations are carried out toon-orbit spacecraft and facilities. This model integrates the results ofseparate OMV, OTV, and Space Station mission models into an identification offluid users and a schedule of missions. Many of the missions are not purelyfluid resupply, but resupply missions conducted in conjunction with othersatellite servicing tasks. It is not clear what the division will be betweenfluid transfer operations occuring at the Orbiter and those occuring in aremote in situ mode. This revision of the orbital fluid supply mission modelincorporates the updates and schedule changes since its last publication in
3474e -28-
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September of 1984. Changes have been relatively minor; involving shifts inthe communications market, new DOD launch schedules, and the addition of somepotential foreign missions. The greatest uncertainty involved in thisassessment is the impact of Space Station operations.
Based on the expanded mission model Figure 13 shows a projection of allon-orbit fluid transfer engagements for the 1989-2001 time frame. Theseengagements 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 acarrier vehicle for the resupply module. Materials processing in space (MPS)engagements are based on estimates of factory modules aboard free-flyingleasecraft. The number of engagements decline in the latter time frames asMPS operations transition to space station proximity. Further fluid transfersare then accomplished through routine station resupply flights. SDI needs arecurrently unknown.
Parametric techniques were used to make estimates of the types, amounts andyearly fluid usage rates required by the resupply mission model (Table 2).Small quantities of associated pressurants such as He and N£ were alsorequired, but their amounts were relatively minor in comparison to the primaryfluids. In some cases, depending on the resupply technique used, noadditional pressurants are required for the transfer process. Figure 14depicts, in graphic for the same information.
While water is a significantly required fluid, it is involved in a modularchangeout in MPS factories and is thus not ideally suitable for transfer bythe resupply module. As expected, hydrazine and bi-propellant are thedominant required fluids. In addition, another significant fluid of interestwas found to be liquid (primarily superfluid state) helium.
In updating the resupply mission, extensive contacts were made with thepotential user community. In addition to in-house data base and literaturesearches, three related surveys were conducted. They were telephone contactswith various NASA program contacts used in the development of the resupplymission model. The updated data did not contain any major surprises andrepresented a maturing of several programs which may hold a more positiveattitude fluid resupply. In January of 1985, a series of presentations weremade to satellite manufactures TRW, Hughes, and Ford. The presentations wereintended to direct feedback on the proposed demonstration program andfinancial analysis. A survey question, prepared under Rockwell IR&D, was sendto over 90 members of the space community in the U.S. and aboard. Thecontacts included domestic and foreign space agencies, aerospace andcommunications companies, and space insurers. The survey attempted to findnew candidate users of fluid transfer services for either remote orOrbiter-based service. Thirteen useful responses have to be received to date,with most supporting the concept of concurrent servicing with resupply. Asummary of the results of these surveys is presented in Figure 15.
Table 3 presents a more detailed chart of NASA missions with potentialrequirements for the use of ERM. The data for this survey resulted fromtelephone contacts with approximately 20 programs or project offices. Someadditional information was obtained from the NASA Space Station mission data.Many of the missions are expected to be resupplied by Space Station, using anOMV. 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 satellitesuppliers. At the request of NASA HQ, ERM contract results were presented,with emphasis on Geo satellite economic analysis, technical inputs toresupplied satellites, and acceptable resupply demonstrations. A series oftechnical and economic comments were received and documented, some commentswere common to more than one supplier. Summary conclusions from thesecontacts are presented as follows:
o Need sophisticated resupply/services module and logistics supportsystem
o Some propellant transfer advanced development necessary
o Early performance of flight test segments appears to be a sufficientd emons tr a ti on
o Near-term resupply of government spacecraft should convince insurersand satellite users of concept viability
o Directly involve satellite suppliers for further technical/economicstudies
Over ninety survey forms, related to potential users of fluid transferservices, were sent out in May. The contacts included domestic and foreignspace agencies, aerospace and communications companies, and space insurers.Of the twenty-one responses received to date, twelve were considered to haveuseful data. Figure 16 summarizes those contacts which submitted usefulresponses and what the general nature of these comments were. No majorsurprises were received, although some speculative missions were identified.Strong interest in Orbiter-based resupply was voiced for the Canadian Radarsatprogram and remote resupply for the TDRS program, depending on potentialcosts. Tables 4 and 5 list recipients of the fluid transfer survey. Thecheck marks denote who provided responses and the check mark in parenthesisdenotes a response still in work. No check mark denotes no further contact ofthis date.
Eased on the resupply mission model and the potential users contacts six ERMscenarios were developed (Figure 17). In Scenario 1, the resupply moduleprovides the OMV with propellant for transportation to GEO. This mode wasfound to be more economical than using an expendable OTV to transport the OMVto GEO. Propellant transfer (and concurrent module exchange servicing, whenrequired) is accomplished directly from the OMV after separation from theresupply module in Scenario 2. This uses a reusable cryogenic OTV for Geotransport.
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) andparticipate 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 operationalscenarios, 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 beenincluded for comparison purposes. The range of sizes for all these optionsare 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, suchas the need for concurrent servicing, Orbiter flexibility, and resupplylocations, 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 todevelopment effort
o Adequate MMH/NTO available from OMV initiallyo Focus on key technical/operational risk elementso Lower front-end cost for timely development
The updating of the mission model provided no surprise but the expansion ofcontract scope to include space infrastructure assets emphasized major area offluid transfer requirements. End user satellites typically requiredconcurrent servicing with resupply, while assets such as the OTV, OMV andSpace Station required resupply along (while acquiring their payloads onseparate mission). The spectrum of resupply missions possible was summarizedinto a set of operational scenarios with drivers of resupply module sizing.This resulted in a confirmation of the selection of MMH/NTO as the baselinefluid while down-sizing the earlier resupply module design (45,500 Ibs ofbi-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. Thedecoupling of tankage from the resupply module was also discovered as apotential means for concentrating on the technical issues associated withfluid transfer while preparing flexibility in an ending market for fluidtransfer.
2.2 Concept Definition (Task 2)
A review of the Task 2 Study results is presented in this section of the finalreport. A summary of the eight subtasks is listed below.
l) The conceptual propellant and pressurant resupply system defined insupport of the Space Platforms Expendables Eesupply NASA ContractStudy (NASA8-35618), has been reevaluated and revised to incorporate
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new subsystem design options and operational requirements. Inaddition to fluid resupply subsystem design changes, which include 3separate disconnect/line purging designs, the bi-propellant resupplyquantity of the Expendables Resupply Module (E.R.M. ) has been reducedfrom the original 45,500 Ibs. to 7,000 Ibs.
2) The on-orbit transfer of propellants with pumps will make possiblepropellant scavenging and extend satellite life. Three differenttypes of pumps were considered and compared: centrifugal, gear andperistaltic. This subtask reports the power requirements versus flowrate at different P's for the propellants NTO, MMH, and Ify..Power requirements for the centrifugal pump were determined fromreference (3)- Power requirements for the gear pump were determinedfrom data supplied by the Sundstrand Corporation.
3) An analysis of two pressurant transfer methods (cascade andcompressor) was performed and the optimum method and supplyconfiguration for resuppling a 7000 Ib bi-propellant's pressurantsystem was determined.
4) Subtask 4 requires that the instrumentation requirements formicroprocessor control and system status monitoring duringresupply/quiescent periods be defined. Reference 10 was used as abaseline. The types of instrumentation considered are pressuretransducers, thermocouples, valve position indication feedbackswitches, 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 forthe Middeck Ullage Transfer Experiment (MUTE).
7) The test objectives and tests to verify these objectives arepresented for the Flight Demonstration Test Vehicle (FDTV).
8) The purpose of the ERM ground test programs is to develop the FlightDemonstration Test Vehicle to be used to conduct the fluid transferduring 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 thevehicle. The ERM., developed under the SPER. Concept Definition Study, is nowforeseen as a potential growth version (i.e., OMV extended missions kit, spacebased propellant depot) of a more modestly sized module design. A ERM of, atbest, 7,000 Ibs. of bi-propellant will adequately cover resupply needs throughthe year 2010. The 7,000 Ibs. bi-propellant capacity will require a tankvolume approximately equal to 6 Shuttle Orbiter RCS propellant tanks.
In addition to the changes in the expendables resupply capacity, minor changeshave been incorporated into the conceptual ERM fluid transfer system design.The system design changes include l) isolating pressurant tanks, 2) regulatingcompressor inlet pressure, 3) additional pump and compressor plumbing for ERMpropellant and pressurant tank refueling, and 4) three separatedisconnect/line purge subsystem designs.
3474e -44-
High pressure solenoid valves have been positioned at each pressurant tankoutlet for a more efficient management of resupply pressurant. Isolatingpressurant tanks will ensure a sufficient supply of high pressure helium ornitrogen gas, for multiple engagement resupply missions.
In-house data, concerning the operational characteristics of high pressure gascompressors, reveals that regulating the compressor inlet pressure will benecessary, to obtain the desired design compression ratio and outlet pressure.
A rearrangement of the conceptual fluid system schematic's pump and compressorsubsystem plumbing allows the use of these components in the refueling of theEEM's propellant and pressurant tanks. This modification is illustrated inFigures 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 19illustrates a vacuum pump purge design, and Figure 20 a dribble volume purgingsystem. The fluid transfer schematics illustrated in Figures 19 and 20 alsouse a quick disconnect design similar to the NASA/Fairchild, triple inhibit,standardized refueling coupling (NASA 9-17333 umbilical). A detaileddescription of the operational characteristics of each disconnect/line purgedesign is documented in reference 2. Table 7 describes the major advantagesand disadvantages of each purge subsystem design.
Evaluation of Propellant Resupply Pumps
Propellant transfer can be accomplished by one of the following threemethods: ullage recompression, ullage vent and ullage exchange. In theullage recompression method, the flow rate is determined by the amount of heatthat the propellant tank can dissipate. The flow rates for the other twomethods are usually limited by the type of propellant acquisition devise(vane, screen, diaphragm) in the receiver tank. Due to the different flowrates required, the power requirements in this report were determined forseveral P values at flow rates up to 10 GPM.
Although a peristaltic pump can provide enough P for both an ullage vent andullage exchange transfer it cannot accomplish an ullage recompressiontransfer. This is due to the fact that a head rise of at least 250 psi isrequired to complete an ullage recompression propellant transfer.Consequently, the peristaltic pump was not longer considered as a candidatefor 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 flowrate at various head rise values for the propellants MMH, NTO, a system of MMHand NTO, and hydrazine. Data supplied by Sundstrand Corporation wasextrapolated to complete the family of head curves seen in Figures 21, 22, 23and 24. These curves are for a 10 GPM optimum flow rate gear pump andrepresent power required to pump and do not include motor efficiency.According to a representative of Motronics, motor efficiency could varyanywhere from 70 to 90$ depending on how the motor is cooled, what materialsare 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 takeninto account in the power consumption calculated for the 10 GPM optimum gearpump.
3474e -45-
Line Pressurant Purge
Vacuum Pump Purge
Dribble Volume Purge
Purge receiver and ERMQDs, plus some lineleading to QDs.
Quick Purge Time
No Dribble Volume
Low System Weight
No Dribble Volume
Purges Receiver andERM QDs
Low System Weight
Simple Design
Quick Purge Time
High System Weight
More QDs to Mate
- Long Purge Times
- Problems withCompressive Heating
Limited Life Pumps
- Receiver and ERM QDsare not purged.
Some propellant willgo into theenvironment 0.2 cc
TABLE 7. DISCONNECT/LINE PURGE OPTIONS, MAJOR ADVANTAGES AND DISADVANTAGES
3474e -46-
ORIGINAL PAGE ISOF. POOR QUALITY
FUD PANEL
He Trms
Ull LinePurge
Ull Trani
MTO LinePurge
•TO Trans
NTO LinePurge
NTO Trans
Servicing, Checkout, f.iFill/Dram Lines have beeomitted 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 PURGEDESIGN, WITHOUT SELF REFUELING PLUMBING
3474e -47-
-- FWO -PANEL
He Trans > ^
UH Trans
NTO Trins
AFT PANEL
M* Trans
Ull Trans
KTO Trans
WasteTarn
Storage ( )nk \ /
ServlC'rj, rh»:lo.F<ll/Dre'" L">"! nb*«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 onlyan approximation at the other flow rates. The justification for the aboveapproximation is that gear pumps are usually only run at their optimum flowrates, and even though they can be run at off-design speeds, no data was foundfor this case. A major advantage of the gear pump is its ability to handle awide range of head values. This 10 GPM optimum flow rate pump weighsapproximately five pounds (not including motor weight).
Figures 25, 26, 27 and 28 present pump power requirements of a centrifugalpump 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 wereduplicated from reference 3*
As can be seen in Figure 23, the power to pump NTO and MMH with a gear pump ata flow rate of 9 GPM and a P of 250 psi is about 2900 watts. The power topump NTO and MMH with a centrifugal pump under the same conditions is about2700 watts (Figure 27). The difference in power required by a centrifugal andby a gear pump is not significant; therefore, further study into the powerconsumption 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 rangeof P's than a centrifugal pump can provide. A gear pump can provide P'sup to 1500 psid while the centrifugal pump is not capable of much more thandelta P's in the 350 psid range at its optimum operating speed. Therefore,centrifugal pumps are inefficient when run at off-design speeds. Variationsin operating speed are not critical for the gear pump.
Micropump Corporation believes they can make a magnetically coupled gear pumpto pump at 3 GPM with a P of 300 psia. To run the motor and pumpcombination requires 1730 watts to pump MMH and 2880 watts to pump NTO. Thispump (with motor) weighs about twenty pounds, is 7-1/2 inches in diameter andis 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 optimummethod and supply configuration for resuppling a 70001b bi-propellant'spressurant system was determined.
The two transfer methods analyzed were a cascade only method and a cascadewith compressor method. The cascade only method involves multiple supplytanks at a higher pressure being opened sequentially to a receiver tank untilpressure equalization occurs; however, this transfer requires considerablylarger quantities of helium than is required for the resupply. Therefore,larger and heavier tanks are required for this method. In contrast, thecascade with compressor method requires only enough helium for the resupplysince it performs a cascade transfer first followed by the use of a compressorto complete pressurization. However, this method requires a battery systemand two compressors for redundancy which increase system weight.
Supply system volumes and weights were calculated for three supply tankpressures of 6000, 5500, and 5000 psia and 1 to 6 supply tanks. The optimumtransfer method and supply tank configuration was determined to be asfollows: A cascade only transfer with four 4200 cubic inch supply tanks at6000 psia for a total system weight of 277 Ibs.
3474e -53-
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The Concept Definition Study, Ref. 7, identified two possible pressuranttransfer approaches: a cascade or blowdown only method and a cascade withcompressor method. The cascade only method involves multiple supply tanks ata higher pressure being opened one at a time to a receiver tank until pressureequalization occurs, thus pressurizing the receiver from 500 psia and 70 F toa level of 3600 psia and 70 F. The advantages of this method are its simpleoperating procedure, sequential valve openings, and its minimal equipmentrequirements, tanks and isolation valves only. However, due to this pressureequalization technique, a larger volume of helium is required to be carriedalong than is required for transfer; thus adding more system weight in theform of larger tanks and excess helium. The cascade with compressor method,although, requires only the amount of helium necessary for the resupply sinceit performs a cascade transfer initially then utilizes a compressor to finishreceiver pressurization. However, this method also has its disadvantages; itrequires a battery and two compressors, for redundancy, which add considerablyto the system weight and complexity.
The supply system was sized for a total receiver volume of 13000 cubic inchessince it was determined that this volume of helium at 3600 psia would expel aworst-case amount of 7000 Ibs of bi-propellant and still have the required 500psia residual. A previous task analysis, Ref. 4 had determined that duringthe helium transfer the supply tank would cool to approximately 40 F and thereceiver tank will heat up to 160 F. Since this occurs in relatively shortperiods of time when compared to the overall transfer time, it was assumedthat the transfer was made isothermally at 40F and 160 F. This provides animportant worst-case assumption in that by assuming the receiver tank is at160 F instead of 70 F, less helium was calculated to have been transferredthan actually was transferred due to the pressure difference between the twotemperatures, therefore, a larger supply system size and weight wascalculated to perform the transfer. Figure 29 depicts this pressuredifference for the cascade with compressor transfer method.
The actual sizing procedure for the cascade only method used the computerprogram utilized in the analysis performed for Ref. 6 and, in particular,involved guessing a total supply system volume and calculating the finalreceiver tank pressure from the ideal gas equation with compressibility thencomparing that to the desired level of 3600 psia. System volumes for thethree supply tank pressures of 6000, 5500, and 5000 psia and 1 to 6 supplytanks were calculated. Table 8 presents these values.
The sizing of the supply system for the cascade with compressor method waseasily accomplished once the three supply pressures were known. Since thistransfer method utilizes a compressor to achieve the desired receiverpressure, only the amount of helium required to resupply the receiver isneeded in the supply system. Therefore, by using the ideal gas law andknowing the receiver requirements and supply system pressure, the systemvolume was calculated. Table 9 also presents these values.
The weight of a supply tank was calculated from an equation obtained during aphone conversation with Mr. Newhouse, Ref. 8.
(Pb*V)/675,000=Wt
3474e -58-
I II I -
TOTAL SYSTEM VOLUME (CUBIC INCHES)
NUMBER | | 6001OF | | SUPPLY
1 1 CASCADE ONLY
11 | | 26800
II2 | | 19400
3 | | 180001 1II
4 | | 16800
I I5 1 1 16500
I I6 | | 16200
) PSIAPRESSURE
CASCADE W/ COHP
7800
7800
7800
7800
7800
7800
550CSUPPLY
CASCADE ONLY
37500
24800
23400
21200
20500
19800
) PSIAPRESSURE
CASCADE W/ COMP
8500
8500
8500
8500
8500
8500
500CSUPPLY
CASCADE ONLY
61000
35000
31500
28000
27500
27000
) PSIAPRESSURE
CASCADE U/ COHP
9400
9400
9400
9400
9400
9400
TABLE 8. SYSTEM VOLUME COMPARISON
IIII-
SYSTEM WEIGHT UBS)
NUMBER 1 1 6001OF | | SUPPLY
|| CASCADE ONLY
1 || 4171 11 1
2 || 308
I I3 || 291
II4 || -277
II5 || 277
I I6 || 277
} PSIA
PRESSURE
CASCADE U/ COMP
380
• 376
378
380
384
388
5501SUPPLY
CASCADE ONLY
535
360
342
• 318
313
311
3 PSIAPRESSURE
CASCADE W/ COMP
384
• 380
382
385
388
392
500!SUPPLY
CASCADE ONLY
790
460
419
• 379
377
375
) PSIAPRESSURE
CASCADE U/ COMP
389
* 385
337
389
393
397
• • ASTERISK DENOTES OPTIMUM CONFIGURATION
TABLE 9. SYSTEM WEIGHT COMPARISON
3474e. -59-
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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 thetanks, helium, and 2 isolation valves per tank. The total weight for thecascade with compressor method includes tanks, helium, 2 compressors(redundancy), a battery system, and 2 isolation valves per tank. Thecompressors were assumed to weigh 100 Ibs each based upon estimations byvarious companies for a 2 scfm flowrate and 2/3 hp compressor. The tanksolenoid valves were assumed to weigh 2.3 Ibs each based upon valves currentlybeing used in the aerospace industry. The battery system weight variedaccording to the amount of helium required to be transferred. This weightincluded the weight of the Li/TiS battery, the coolant system, the powerconditioning system, and the system's structure. To calculate the batterysystem 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 perpound of helium transferred. The efficiency of the battery system wasdetermined to be 40 Wh per pound of battery once the coolant system, powerconditioning system, and system structure was included. Table 10 presentsthis system weight, the mass of helium required to be transferred, and theamount of power required by the compressor.
Figures 30, 31 and Table 9 present the total system weight vs. the number ofsupply tanks for the two transfer methods. A line has been drawn through theoptimum tank configuration in the figures and an asterisk denotes them in thetable. Beferring to Table 8 and Table 9, these optimum configurations are:four 4200 cubic inch supply tanks (16800 cubic inches total) at 6000 psia witha system weight of 277 Ibs for the cascade only method and two 3900 cubic inchsupply tanks (7800 cubic inches total) at 6000 psia with a system weight of376 Ibs for the cascade with compressor method.
The overall optimum transfer method and tank configuration can now beidentified to be a cascade only transfer with four 4200 cubic tanks at 6000psia since it weighs 99 Ibs less than the other method.
Instrumentation Requirements for Microprocessor Control
Pressure Transducers; Pressure Transducers were identified to monitor heliumtank inlet pressure, propellant tank ullage pressure, compressors/pumps inletand outlet pressures, quick disconnect inlet pressure, and waste tankpressure. The minimum pressures to be monitored within the receiver vehicleduring resupply are helium and propellant tank pressures. The pressure
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transducer locations are shown in Figure 32. A brief description of eachinstrument and how it will be used is shown in Table 11. The pressuretransducers on the quick disconnect that will be used for leak detection(pressure decay) are not discussed in this letter as they are recommended tobe an integral part of the disconnect. The sampling rate for adequatemicroprocessor control is recommended to be 100 samples per second duringdynamic resupply operations and then reduced to one sample per second duringquiescent periods.
Thermocouples; Thermocouples were identified to support PVT gaging, heliumloading, pump/compressor operation, and vent operations. The minimumtemperatures to be monitored within the receiver vehicles during resupply arehelium and propellant tank locations. The thermocouple locations are shown inFigure 32 and a brief description of each instrument and how it will be usedis shown in Table 12. The thermocouples on the quick disconnects are notdiscussed in this letter as they are recommended to be an integral part of thedisconnect. The sampling rate for adequate microprocessor control isrecommended to be one sample per second during resupply and quiescent periods.
Valve Position Indication Feedback Switches - Each valve will be equipped withredundant position indication feedback switches.
Flowmeters - Two flowmeter feedback channels are required to monitor flow inboth the forward and reverse flow directions.
Contact Sensors - Contact sensors are necessary to verify proper mating of thequick 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 programsand a few flight demonstrations. The two programs of interest are: theorbital maneuvering vehicle (OMV) and the orbital spacecraft consumablesresupply (OSCES). As is seen in Figure 33, OSCRS, OMV, and the ERM areexpected to be operational by 1990, 1991-2, and 1993 +, respectively.Therefore, a reduction in development testing of the ERM can be expected byusing 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, connectionreliability with leakage monitoring. Major test components -propellant pump, coupling, and tank acquisition device with ullagecontrol.
E) Pressurant transfer - basic function, multiple transfers, connectionreliability 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, OSCRSwill demonstrate the use of the MMS berthing system.
2) OSCRS will perform a propellant transfer in which a manual couplingdevice will be used. This will allow testing/demonstration of thecoupling system.
3) Heat transfer determinations during ullage recompression in thereceiver diaphragm tank will help to establish optimum flow rates.
A manual coupling is being developed by Fairchild. The coupling is to bedeveloped as a standardized coupling for Ify, MMH, NTO, GN£, GHe, andCgCl^F^ (trichlorotrifluroethane). OSCRS' usage of the coupling will beto transfer hydrazine to the GRO in 1990, but by using appropriate seals alllisted fluids can be transferred. The Fairchild coupling will be available bymid 1986 to transfer Ify and it could be available by the end of 1987 totransfer other fluids. The coupling is manually engaged and five inhibitsmust 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 andS1-S3 for the spacecraft side. The inhibits cannot be opened until the twohalves of the coupling are engaged or the coupling disengaged before theinhibits are manually closed. The manual engagement and disengagement forceis about 36 pounds. A rough leak detection system will be performed bypressure transducers at the indicated ports shown in Figure 34. A weightanalysis indicates that the tanker and spacecraft coupling will weight 13.0and 10.8 pounds, respectively. Fairchild also has funding to startpreliminary designs for the automated coupling in early 1987. Fairchildindicated that the present manual design will not be automated as is.
Another possible area of compatibility that OSCRS may have with the ERM is thedevelopment of a propellant pump. At this time the use of a propellant pumpis only considered a option for hydrazine transfer.
The OMV is a Shuttle-based/launched, remotely controlled, free-flying vehiclecapable of performing a wide range of on-orbit services. The initial OMV(launched about mid 1991) will perform payload placement and/or retrieval inLEO when delivered to standard STS orbit by the orbiter. As the space programadvances and mission needs materialize the reference design OMV can beupgraded by modifications to perform more sophisticated missions, to bespace-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 asfollows.
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. Alldevelopment required for rendezvous and remote docking will be performed onthe OMV before the ERM.
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FIGURE 35. FAIRCHILD MANUAL COUPLIHG
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The ERM will be able to use only limited developmental work from OSCRS and theOMV. OSCES will provide information on the propellant coupling, propellanttransfer, leak detection, and docking with the MMS berthing system. OMV willprovide information on the rendezvous and remote docking and release withother spacecraft.
On-Orbit Middeck Ullage Transfer Experiment
MUTE ia an Orbiter middeck experiment designed to investigate micro-gravityfluid transfer dynamics using a water solution, capillary/screen type tanks,micro-gravity gaging system, and a small reversible fluid pump. Thefundamental objectives of MUTE are as follows:
1) To demonstrate the concept of ullage transfer in a micro-gravityenvironment.
2) To demonstrate the propellant management device's (PMD's) ability toposition and control the ullage bubble by observing the ullagetransfer process. The PMD will include a ullage/propellant separatorand will be placed in each tank.
3) To demonstrate a micro-gravity quantity gaging system's accuracy in amicr o-gravi ty envir onment.
These objectives are to be met by using off-the-shelf equipment and bymodeling the structure and tanks after the storable fluid managementdemonstration (SFMD).
MUTE consists of a supply tank, a receiver tank, a PMD in both the supply andreceiver tank, a reversible pump, a micro-gravity gaging system, a flow meter,and associated valves, filters, lines, pressure gauges, and temperaturesensors. MUTE has no power or command interfaces with the Orbiter; and isbasically self-contained, with the exception of an overboard vent connected tothe Orbiter waste management system. The entire experiment will be designedto fit in place of several lockers in the orbiter middeck. Support equipment,from STS, will include cameras and associated lights, power to be supplied byOrbiter outlets, (for the cameras and associated lights) a micro-gravityaccelerometer measurement system (MGAMS), and a data recorder.
Control of MUTE on-orbit will be accomplished by the crew through a sequenceof manual valve and switch operations. The small fluid pump is the motiveforce 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 datainputs will be correlated.
A preliminary schematic to perform the recommended tests is presented inFigure 36. The recommended tests are presented in Table 13, with expectedtest objectives in Table 14. The following is a few general operatingcomments:
l) Prior to tests, the crew will connect the MUTE's vent line to theOrbiter waste water overboard vent system.
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Test No. Transfer Fluid from Supply Tankto Receiver Tank
Transfer Fluid from Receiver Tankto Supply Tank
1. Transfer to dry tank at low flowrate with ullage exchange. Stopfilling at each 10% of masstransfer.
2. Transfer to wet tank at low flowrate with ullage recompressionuntil tank is 50$ full.
3. Transfer to wet tank at moderateflow rate with ullage exchange.
4. Transfer to wet tank at highflow rate with ullage exchange.
5. Transfer to wet tank at low,moderate, and high flow rateswith ullage exchange.
6. Transfer 75-802 of fluid towet tank at moderate flow ratewith ullage exchange duringvarious Orbiter accelerations.Stop flow and accelerate untilscreen breakdown. Wait forrewetting of screen to occurbefore transferring remainingfluid.
Transfer fluid back to supplytank at a low flow rate withullage exchange.
Transfer fluid back to supplytank at a low flow rate withullage recompression.
Transfer fluid back to supplytank at a moderate flow ratewith ullage exchange.
Transfer fluid back to supplytank at a high flow rate withullage exchange.
Transfer back to supply tankat a moderate flow rate withullage exchange.
Transfer fluid back to supplytank at a moderate flow ratewith ullage exchange.
TABLE 13. RECOMMENDED TESTS
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Objective Test No.
1) To evaluate PMD's Ullage Controlpropel1 ant/ullage separation X X X X Xwetting pattern of PMD X X X X Xwetting time of PMD X X X X Xorbiter acceleration effects X
2) To evaluate Zero-Gravity Gaging Systemaccuracy determinations
steady state flow rate X X X Xvariable flow rate X Xorbiter acceleration X
response timesteady state flow rate X X X Xvariable flow rate X Xorbiter accelerations X
comparison to PV gaging system X
TABLE 14. MATRIX OF TEST OBJECTIVES
Component Weight (Ibs)
each system
SFMD 250 Ibsremoval of valves, gas cylinder, ."id cylinder -50 Ibs
pressure transducers 2 12temperature sensor 1 7pump 5batteries 25flow meters 3 6zero-gravity gaging system 40+10% 33
Total 363 Ibs
TABLE 15. MUTE SYSTEM WEIGHT
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•*;2) After eactetest the test fluid will be returned to the supply tank'/,_ before going on the next test.
-s3) 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 afterSFMD, SFMD's system weight was used as a baseline by removing weights ofcomponents not recftfired and then adding in new component weights plus an extra10%. 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-gravityremote'Tluid transfer. There will be two brass board vehicles capable ofbeing separated or'docked using the remote manipulator subsystem (RMS). Theupper portion of th% test vehicle, the receiver vehicle, piggybacks on thesupply vehicle throtigh deployable payload latches. The RMS system is utilizedto separate the receiver vehicle, by use of a grapple fixture attached to thereceiver vehicle, to simulate remote docking and mating of the two vehicles.A snare type end effector, incorporated at the fluid transfer interface of thesupply vehicle, accomplishes final mating of the fluid transfer interfacepanels 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 transferr- interface during a remote operation.
2) To demonstrate proper QD mating, leak integrity verification,operation^"and purging with safing, before, during, and after a fluidtransfer.
3) To evaluate^ the PMD's ullage control and separation capability duringa propel!ant transfer.
4)*> To evaluate propel!ant transfer by the methods of ullage transfer andullage recornpression. Different pumped flow rates, and the abilityof the tanker to resupply itself will also be examined during theullage transfer tests.
5) To evaluate-' pressurant transfer by two methods: 1 ) a pressuranttank cascade and 2) by compressor usage.
6) To evaluatep-the micro-gravity quantity gaging system for bothaccuracy 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 componentsand as many off-the-shelf components as possible.
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FDTV contains four forward RCS propellent tanks, consisting of two receiverand two supply tanks, for the MMH/NTO propellent transfer. All four tankswill be aligned with the +X-axis of the Orbiter's coordinate system duringlaunch and contain a 6Q% nominal load (55% of total volume). Only minormodifications to the pressurant inlet area of the RCS propellant tank will bemade so that ullage/propel!ant separation will occur. The propellant transferwill be accomplished by the use of three small propellant pumps. As is seenin Figure 37, the MMH side will provide pumping to the receiver tank and willalso 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 receivervehicle and one on the supply vehicle with simpler valving. There are twodifferent pressurant transfer systems for comparison and demonstrationpurposes. The MMH side will perform a pressurant cascade from four 4800 psiatanks to fill a RCS helium tank to 3600 psia, while the NTO side .will use acompressor to perform the pressurant transfer after a blowdown pressureequalization transfer.
Control of FDTY on-orbitwill be accomplished by the crew through a sequenceof remote operations. The RMS will be the motive force during mating anddemating steps, requiring intensive crew attention; but the resupply processshould only require crew monitoring with a minimum interaction between thecrew and the FDTV.
Table 16 presents the recommended tests to be performed by FDTV to satisfy thetest objectives presented in Table 17. Each of the four tests are broken downinto their basic steps required for the completion of each test. The firsttest will examine the docking, mating, leak integrity verification, purging,safing, and the demating process. The second test will perform the pressuranttransfer by the cascade method and compressor method, and perform an ullagerecompression propellant transfer. The third test will perform an ullagetransfer at a low flow rate. The last test will demonstrate the ullagetransfer at a higher flow rate. Propel!ant transfer will occur back and forthbetween the receiver and supply tanks during each test.
FDTV's system weight is presented in Table 18. The expected dry weight is3309 Ibs. To this was added a 15% potential growth factor plus the requiredpropel!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 FlightDemonstration (ERM-FD). The FDTV shall be capable of conducting remote matingand demating operations, performing propel!ant transfer by means of "ullagetransfer" and "ullage recompression" methods, together with the transfer ofpressurant by means of "cascade" and compressor methods. In addition the testvehicle shall contain a propellant management device (PMD) capable ofseparating gas and liquid, (for ullage transfer tests) and a zero-gravityfluid quantity gaging system (gaging system) to determine the amount ofpropellant in the tanks. To date, the ERM study has defined the need todevelop three critical components for this transfer demonstration.Reiterating, these components are the propel!ant transfer pump, the pressurantcompressor and the mating fluid couplings. It should be noted that for thisprogram it shall be assumed that the NASA will develop the couplings for thisapplication (including remote configuration) and that this development effortshall produce qualified units applicable for the ERM-FD test vehicle. The
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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 verification6) 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 andreceiver propel 1 ant lines and tanks; then prepare for the ullagetransfer method.
2) Using the supply pump transfer MMH by ullage transfer from thereceiver tank back to the supply tank, from 90 to 30% of thefilled volume.
3) Using the receiver pump transfer NTO by ullage transfer from thereceiver tank back to the supply tank, from 90 to 30% of thefilled volume.
4) QD safing and purging.5) Demating of QD6) 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 filledvolume.
5) Using the supply pump resupply the NTO receiver tanks by ullagetransfer at a higher flow rate from 30 to 90% of the filledvolume.
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
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1)2)
3)
4)
5)
6)
7)
8)
Objectives
To Demonstrate Docking and Alignment of Interface
To Demonstrate QD Integrity Verification, which include:QD MatingLeak Integrity VerificationPurging and-Safing
To Evaluate the PMD's Ullage Control with:Propel! ant/Ullage SeparationPropellant Supply Capability
To Evaluate Propel! ant Transfer by:Ullage Transfer with:
Pump UsageDifferent FlowratesTanker Resupply
Ullage Recompression with:Thermal ControlPump Usage
To Evaluate Pressurant Transfer by the:Cascade MethodCompressor Method
To Evaluate Micro Gravity Gaging System for:Accuracy DeterminationsFlow rates Effects
To Demonstrate Thermal & System Control
To Demonstrate Demating and Separation of Vehicles
Test Numbers1 2 3 4
X X
X XX XX XX
X
X
X
XXX
XXX
XXX
X X
X
X
XXX
XXXXX
XXX
X
X
X
XXXX
XXX
XXXX
XXX
X
X
TABLE 17. MATRIX OF TEST OBJECTIVES
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Subsystems Weight
1) Fluid Transfer Systempropel1 ant tank 300pressurant tank 287waste tank 8pump 60compressor 200disconnects 240valves, regulators, etc. 325pressure end temperature sensors 90
2) Interface panelsend effector 65grapple fixture 44face plates 38latch 20electrical connects 8
3} Structural Systemframe and support 665fittings and latches 624thermal blankets and covers 125
4) *Avionics/PowerWire harness 40regulation & distribution 100thermal control 30on-board data management 40
S u b t o t a l 3 3 0 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
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gaging system, which will be developed under a separate contract, shall becustomized to the test vehicle tankage and verified during the ERM-GTP. Thedevelopment of the liquid/gas separator unit of the PMD shall be tested and/oranalytically assessed as a separate unit, with no plans to assess itsoperation 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 andfabrication of breadboard test units.
2) Component testing of pump and compressor breadboard units. Testingand/or evaluation of the liquid/gas separator together with thegaging system configuration requirements will be performed duringthis time period.
3) FDTV design and fabrication. During this time frame the test andcheckout of the coupling mating assemblies to assure alignment anddetermine engagement and separation loads will be performed.Procedures for coupling operation will be assessed.
4) FDTV development and verification test program. This period isdivided into two test series with allowance for potential modifica-tions between the two test series. During this period, low levelsupport using breadboard development test apparatus may be advanta-gous 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 ofprototype (pump and compressor) units to assess component features unique tospace application and for proof of concept of required advancedstate-of-the-art component features and (2) Brassboard testing of the FDTVtransfer system using flight components and hardware.
Breadboard Tests
Prior to these tests, a final assessment of component design from variouspotential suppliers shall be performed and the most promising of thesecomponents will be selected for breadboard (B/B) development. The breadboardtest results will be used to enable the selection of a unique component designfor each fluid transfer application. If, based upon the B/B tests there aremore than one viable component design for a unique application both designswill be further evaluated during the FDTV brassboard test program. Thebreadboard test results will also enable the finalization of component designspecification requirements.
The B/B test effort will be concentrated chiefly to the development of thepropel!ant pump, gas compressor (includes assessment of gas and compressorcooling requirements), and the liquid/gas separator device.
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Propel1 ant Pump Tests
The breadboard test setup shown in Figure 40 shall be the test-bed for theevaluation of potential pump configurations to be used in the FDTV. Flightconfiguration flowmeters, if available, shall also be incorporated in order toenhance test results. Three types of pumps shall be evaluated (Gear,multi-stage centrifugal, and piston), however, only the two types with thehighest potential will be selected for breadboard development. All breadboardtesting 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 transfertechnique (4 flow tests each at 5, 10, and 15 gpm), and 12 flow tests usingthe ullage recompression technique (4 flow tests each at 1, 3, and 5 gpm).
The primary test objectives are to assess the pump performance and, ifnecessary, recommend design changes based upon the test results. Detailedtest objectives are listed below:
1) Evaluate pump performance using ullage transfer method and variousflowrates and tank pressure conditions.
2) Evaluate pump performance using ullage recompression method andvarious flowrates and tank pressure conditions.
3) Define any pump deficiencies and provide recommended changes to theflight-type pump designs.
4) Obtain preliminary data on ullage recompression receiver tank ullagetemperature effects (secondary objective).
5) Pump power requirements assessment.
6) Determine the pump magnetic coupling drive performance at variouspump speeds and pressure differentials (optional only if magneticcoupling is used).
Gas Compressor Tests
The compressor breadboard test setup is shown in Figure 41 . A series ofhelium flow tests at various flowrates, temperatures and pressures will beperformed to assess compressor performance and design parameters. Theobjectives 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 atvarious flow conditions. (Optional only if magnetic coupling isused).
3489e -87-
TEST CONDITIONS TEST NO.
(Pressure « PSIA, 1,2 3 4 5 6 7 8 9,10Flowrates = SCFM)
1. COMPRESSOR
INLET PRESSUREA. 500; B. 700 AMAXIMUM OUTLET PRESSUREA. 3000; B. 5000 A, B
COMPRESSOR FLOWRATESA. 5; B. 10 A, B
GAS INITIAL INLET TEMPERATUREA. 70°F, B. 1000F; A2. AMBIENT CONDITIONS
PRESSUREA. 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
A A A A A A A , 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 recommendedchanges to the flight compressor design.
Liquid/Gas Separator Tests
The liquid/gas separator testing will be performed at the component level andwill be restricted to the verification that this unit will maintain itsstructural and mechanical integrity when subjected to launch environments.
Brassboard Ground Test Program
This test program will provide the final phase of the fluid transfer systemdevelopment and will also provide the verification tool to establish the FDTVfluid transfer capability for the flight demonstration.
Test Vehicle Description
The Test Vehicle Description (T/V) shall be designed and fabricated to beinstalled in the space shuttle payload bay and shall be in two separatesections; the tanker side, which will be fixed to the orbiter, and thereceiver side which will simulate a spacecraft being resupplied. The twosections shall be launched in the demated configuration. Once in orbit themating of the two sections is performed by the RMS and the grapple and endeffector (Figure 42). the T/V shall therefore be designed to performsatisfactorily after being exposed to launch and landing environments. TheFDTV fluid schematics are presented in Figures 37 and 38 for the fuel (MMH)and oxidizer (^Htty) systems respectively, the helium transfer systemsfor use with the compressor and cascade methods are also included. Apreliminary FDTY design concept showing the major component layout ispresented Figure 42. The test Vehicle (T/V) configuration shall be the sameas for the flight demonstration except that the Grapple Fixture and EndEffector need not be included for the conduct of the brassboard tests. TheTest Vehicle will include seven major hardware and component items critical tothe fluid transfer operation. These are; propel!ant tanks (4), heliumpressurant tanks (7), fluid couplings (6), gas compressor (1), zero-gravityquantity gaging system (4) including four flowmeters as backup, propellentpump (3), and miscellaneous si nail tanks (4) to purge and safe the couplings.
A heat exchanger will be required to cool the compressor oil, compressorhousing (requires circulating pump), and the compressed gas. The heatexchanger shall be designed to tie into one of the space shuttle payload baycooling loops using Freon-113 as the fluid media.
This circuit of the compressor heat exchanger will also require a smallcirculating pump. The ground tests will include a flowmeter in this circuitin 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 fluidtransfer operations.
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Test Program Description
The test program will consist of a series of tests directed at developing aflight test vehicle which will be utilized in a fluid transfer demonstrationaboard the space shuttle payload bay. The results from these tests willdevelop and assess the fluid transfer system capabilities as well as providethe tool to define the flight demonstration procedures and regimes.
Final development of the gas compressor and propel!ant transfer pump will beaccomplished during this testing phase. In addition, operation, verificationand safing of the fluid couplings during mating and demating, together withthe verification of the gaging and flowmeter system will be demonstrated.
The test program will include ambient and vacuum testing (zero-gravity testswill not be performed). All thermal sensitive areas such as gas compressionduring helium gas transfer and during ullage recompression propellent transferwill be diagnosed in a vacuum atmosphere. Except for the zero-gravityenvironment, the proposed flight demonstration timeline an procedures will beperformed as verification and also as a tool for determining systempredictions 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 highflowrates.
1-B. Pressure conditions - Evaluate pump performance at variousinlet, outlet and delta pressures.
1-C. Temperature conditions - Evaluate pump performance at nominaland high temperature conditions (fluid and ambient).
2. Fluid coupling verification.
2-A. Mating/demating - Define mating and demating loads alignmentrequirements, and procedures. Electrical connector mating anddemating will also be evaluated and requirements established.
2-B. Verification of Interfaces - coupling interface verificationprocedures 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 andhigh fill rates.
3-B. Pressure Conditions - Determines compressor performance atvarious inlet, outlet, and delta pressure conditions.
3489e -92-
3-C. Temperature Conditions - Determine compressor performance atvarious temperature conditions.
3-D. Compressor/gas heat exchanger - Determine heat exchangerperformance (heat exchanger requirement established frombreadboard tests).
4. Zero-gravity quantity gaging system verification - This gaging systemshall be developed by NASA/JSC under a separate contract.
4.A Installation - The gaging system shall be installed andcalibrated to the FDTY propel1 ant tanks.
4.B Performance - The gaging system performance shall be verifiedduring flow tests using flowmeters and PVT methods.
4.C Zero gravity simulation - Gas ingestion and tank orientationtests 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 beanalytically verified for flight application (i.e. launchenvironment effects and design concept).
6. Helium Tanks.
6.A Heat exchanger - verify the performance of the helium tank heatexchanger. The need for the heat exchanger shall be determinedfrom the breadboard tests.
7. Helium Transfer
7.A Verify and demonstrate the optimum transfer configuration usingthe 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 theullage transfer and the ullage recompression transfer techniquesfor various flowrates, temperature and pressure conditions.
8.B Assess and establish the capability of the micro gravityquantity gaging system.
9. Flight experiment procedures.
9.A Establish and define the flight demonstration gas and propellenttransfer procedures and estimate the timeline required.
3489e -93-
9.B Develop a semi-automatic control of the transfer operationsutilizing a micro processor in conjunction with minimalastronaut involvement.
9.C Define orbiter interface and orbiter support requirements suchas compressor cooling (heat exchanger) operations.
10. Test Vehicle flight worthiness
10.A Flight environment - FDTY flight worthiness shall be verifiedanalytically with vibration/acoustic testing support performedonly if analysis deems it necessary.
10.B Instrumentation & Components - all instrumentation andcomponents shall be verified during testing, and calibrationdata together with component checkout data verified aftercompletion of the transfer tests. These data shall be part ofthe FDTV data package.
10.C System leak checks - The system shall be verified to be leakfree within the specified requirements prior to and aftercompletion 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 theconceptual propellent and pressurant resupply system defined insupport of the S.P.E.R. Concept Definition Study (MASA8-35618).Design changes include 1) isolating pressurant tanks, 2) regulatingcompressor inlet pressure, 3) additional pump and compressor plumbingfor E.R.M. self refueling, and compressor plumbing for E.R.M. selfrefueling, and 4) three separate disconnect/line purge designs.
2) An E.R.M. bi-propel!ant capacity of 7,000 Ibs., tank volumeapproximately equal to 6 RCS propel!ant tanks, will be adequate forresupply needs through the year 2010.
3) Power requirements are specified for MMH, NTO and Nofy atdifferent flow rates and head rises or both gear and centrifugalpumps.
4) Recommend selection of the gear pump over the centrifugal pump forits versatility over a wide range of P 's .
5) The tank volumes required to resupply a 13000 cubic inch receivertank to 3600 psia using two different pressurant transfer methods andthree different supply pressures were calculated and presented inTable 2.
6) The total transfer system weights were calculated from the determinedsystem volumes and are presented Table 3.
3489e -94-
7} The overall optimum transfer method and supply tank configuration wasdetermined to be a cascade only transfer with four 4200 cubic inchsupply tanks (16800 cubic inches total) at 6000 psia for a systemweight of 277 Ibs.
8} For a resupply only tanker, with a propellent load of 7000 Ibs, thecascade method of pressurant resupply is preferred, but for themaximum 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 ullagecontrol.
C) Pressurant Transfer - basic function, multiple transfers,connection reliability with leakage monitoring. Major testcomponents - compressor and automated coupling.
10) The Middeck Ullage Transfer Experiment (MUTE) is designed to test theconcept of ullage transfer and the associated PMD, and micro-gravitygaging system. These objectives were examined and a series of testsproposed to verify the ullage transfer process.
11) The Flight Demonstration Test Vehicle (FDTV) is designed to testseveral aspects of remote on-orbit fluid transfer. The fluids to betransferred will include NTO, MMH, and Helium. The proposeddemonstration will test docking, QD mating, leak verification,purging, safing, ullage control and separation during a propellanttransfer, cascade and compressor pressurant transfer, and thermal andremote system control.
12) The ERM ground test program will be divided into four major subtasks.•
A) Pump and compressor specifications, including supplier designand fabrication of breadboard test units.
B) Component testing of pump and compressor breadboard units.Testing and/or evaluation of the liquid/gas separator togetherwith the gaging system configuration requirements will beperformed.
C) FDTV design and fabrication. The test and checkout of thecoupling mating assemblies to assure alignment and determineengagement and separation loads will be performed. Proceduresfor coupling, operation will be assessed.
D) FDTV development and verification test program. Low levelsupport using breadboard development test apparatus may beadvantageous and cost effective to isolate unique componentproblem areas.
3489e -95-
2.3 Programmatics (Task 3)
Summary
Rockwell 's recommendations in the follow-on ERM Study for the advanceddevelopment and f l ight demonstration differ significantly from the initialstudy.
The significant difference between our original and current recommendationsare presented in Table 20. In this study, Rockwell defined and weighed fourcriteria for selecting the best advanced development and flight demonstrationprogram. These criteria and their weights were approved by the NASA Contract-ing Officer Representative at the mid-term review. Based on these criteria,Rockwell recommends the early fl ight demonstration program which culminates atthe end of the advanced development program. We reviewed 12 differentprograms and from these selected the one that maximized the NASA's goals, asdefined by the NASA approved criteria and weights we applied to the selectionprocess.
Rockwell 's recommended program includes two fl ight experiments as well as theadvanced development of the critical components to enable remote resupplyservicing. The first fl ight experiment (Mid-Deck Ullage Transfer Experiment -MUTE) is a low cost validation ($1.8 m i l l i o n ) of ullage transfer in amicro-gravity environment. Not only wil l this experiment remove theuncertainty from the propel 1 ant management devices ability to position andcontrol the ullage bubble, but it demonstrates to potential users NASA'sresolve to develop remote resupply servicing. Stimulating user support forremote resupply servicing is one of NASA's goals, and hence it was a criteriain selecting our recommended program.
The second f l ight experiment (Flight Demonstration Test - FDT) is thevalidation of the advanced developed for remote resupply servicing.Rockwell 's low cost approach is to "fly the brass board", the hardwaredeveloped and tested during the advanced development phase. The primaryfl ight test objective of FDT is to verify remote propel!ant transfer in amicro-gravity environment.
Groundrules and Assumptions
o Customer Goals:
* Least cost advanced development and f l ight demonstration program.
Budget target is $12M.Relative importance of goal is 50-60%.
* Reduce technical risk and stimulate user support for the ERMprogram.
This is to be accomplished by an advanced development andfl ight demonstration programs.
Relative importance of goal is 40-30%.
* Earliest ERM program start date.
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Relative importance of goal is 10$.
o All cost estimates are in constant 1985 dollars.
o STS launch costs are not directly included in the cost estimate, butlaunch costs were considered in selecting the top-and-bottom trussexperiment over the side-by-side truss experiment in the Orbitercargo bay.
o A nominal six month waiting period was assumed for Orbiter payloadintegration activity and this cost is not included except for directcontractor support.
o NASA costs, such as astronaut training and test facilities at KSC forpre-flight propellant transfer testing are not included.
Rationale For Recommendation
Rockwell has based its selection for the advanced development and flightdemonstration on the NASA's selection criteria and their relative weights.The most important criteria was cost (weight of 50 to 602). All options weredesigned to fit within the scope of NASA's target of $12 million. The mostcostly programs were eliminated in the selecting process. The survivingoptions were within $1 million of one another. So, cost was eliminated as acriteria in the subsequent selection process. The next two most importantcriteria were highly correlated, so they were combined into a single compositecriteria. Flight demonstration options of greater technical scope andchallenge tended to reduce the technical risk of the subsequent ERM program.And because of their greater visibility and realism, technical challengingdemonstrations were also more likely to increase user support for the NASA'sremote resupply program. The least important criteria was an early ERM phaseC/D start (weight of 10%). Thus, with cost eliminated, the dominate criteriawas technical risk and user support.
A comparison between our recommendations in the initial ERM study and thefollow-on study are shown in Table 21. Rockwell's current recommendation (theearly flight demonstration option) greatly contributes to most importantcriteria reducing technical risk and soliciting user support (weight of 75-80%once the cost criteria is removed). The only disadvantage is a later ERMphase C/D start. However, due to our efforts to find a low cost flightdemonstration, ("fly the brass board" discussed later), our program plan isshortened by six months. Thus, our low cost program cuts what would otherwisebe a 12 months delay in the ERM phase C/D start to only a 6 months delay (incomparison to the late flight demonstration option originally recommended).
The schedules of the early and late flight demonstration options from theinitial ERM study and the early flight demonstration option currentlyrecommended from the follow-on ERM study are displayed in Figure 43.
For purposes of comparison, all program schedules have been laid on a commoncalendar scale. Schedule differences for the flight demonstration and ERMphase C/D start can be observed by comparing the milestones. The program withthe earliest flight demonstration option is the one we currently recommend.
3489e ^ -\ -98-
Early FlightDemonstration Option
Late flightDemonstration Option Criteria
Importance* ofCriteria
Technology Uncertaintiesremoved before ERMprogram starts.
Accelerated by18 months
6 months
Parallel effort exposesERM programs to riskof schedule delay andcost increases
IS month delay
6 month earlierERM program start
TechnicalRisk
UserSupport
ERMProgramStart
75-802
25-20%
*Normalized from criteria approved by customer at mid-termbriefing (cost 60-50%, Support & Risk 30-40%, Program Start 10%)
TABLE 21 . THE EARLY FLIGHT DEMONSTRATION OPTIONHAS SUPERIOR ADVANTAGES AND ONLY MINOR DISADVANTAGES
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It is 18 months earlier than the program we proposed in the initial ERMstudy. The ERM phase C/D start that we currently recommend is only six monthslater than the program we recommended originally.
In summary, significantly more benefits have been gained (less technical riskand 18 month earlier flight demonstration) than have been lost (six monthlater ERM phase C/D start).
Program Master Schedule
The master schedule for the program recommended by Rockwell, leading to theearliest flight demonstrations is presented in figure 44. The initial effortis to specify the hardware requirements for the ERM program. With therequirements for the ERM program defined, the design of the flight experimentsand the development of the critical ERM components can start. Ideally, MUTEcan fly on schedule in 22 to 24 months after the program starts, and theFlight Demonstration Test can fly in 8 to 12 months afterwards.
The timing in development of several related fluid transfer components couldaffect this schedule. The micro-gravity gaging system (NASA, JSC) would be anideal component to include on the Flight Demonstration Test If the timing wascoincidental, the micro-gravity gaging system could be flight-tested on theFlight Demonstration Test without the expense of an additional independentflight demonstration of its own. This probably will not happen unless theprogram is accelerated by six to nine months so that the flight hardware isready by early to mid FY 1988. The function of the micro-gravity gagingsystem, if not available for the Flight Demonstration Test, could be performedby fluid gaging.
The Quick Disconnect (NASA, Langley) is essential to demonstrate remoteresupply in the Flight Demonstration Experiment. It is currently scheduled tobe available in mid-FY 1988, and its schedule may also need to be acceleratedslightly.
Mid-Deck Ullage Transfer Experiment
The significant characteristics of the Mid-deck Ullage Transfer Experiment arepresented in Table 22. The experiment was selected because it met the NASA'smain criteria of low cost, reduction in technical risk, and potential toestimate potential user interest and support in remote resupply servicing.
The time required to design, fabricate and assemble, and test MUTE is twelvemonths. Ground test time is only scheduled for three to five months. Sincethis experiment is designed mainly to test the effects of a micro-gravityenvironment on the ullage transfer, only limited testing can be performed onthe ground. The schedule and manpower estimate is presented in Figure 45.
The schedule includes an effort to insure that the test plan ishuman-engineered since the orbiter crew will be directly responsible forrunning the experiment. Crew training is planned during the nominal six-monthperiod of payload integration activity at KSC.
3489e -101 -
Test Objective Verify ullage transfer with watersolution in a micro-gravity environment
Customer Benefit Reduce technical risk in PropellentTransfer Experiment
Early flight demonstration to stimulatepotential user support low cost ($1.8M)
Test Schedule 21-24 months after start of advanceddevelopment 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 occurduring the flight experiment, there are twelve months available to diagnoseand correct them. The modifications would then be retested on the FlightDemonstration Test. Thus, the back-up to MUTE is the Flight DemonstrationTest.
This schedule is based on a success-oriented program that assumes no majordifficulties are encountered.
The grass-roots (detailed engineering) cost estimate is uncertain because thethree key variables - direct labor hours, labor cost rates, and directmaterial - are uncertain. To provide a cost range, likely values wereestimated for each variable as presented in Table 23. The best combination ofcost conditions results, by definition, in the low cost estimate. The worstcombination of cost conditions results, also by definition, in the high costestimate. The combination of most likely conditions results in the mostlikely cost estimate.
These cost estimates only reflect the estimating error given the program planas reflected in the MUTE schedule and manpower estimate. The program plan,and hence the cost estimate, reflects a success-oriented program with no majordifficulties encountered.
The low cost estimate reflects that there is only a 10% chance that the costof 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 twoextreme points was spread by a 35/65 Beta distribution function (ogive curve)and is displayed in Figure 46. The expected cost (50% cummulative probabilitythat 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 truebest estimate is somewhere in the vicinity of these two discrete estimate. Ifthe 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 advanceddevelopment 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 LockerSpace
o Minimum Electronic and Microprocessor Control - Maximize OrbiterRight 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 worksas expected in a micro-gravity environment, MUTE also contributes to NASA'sobjectives to reduce technical risk and to support the marketing effort togain 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 advanceddevelopment. The cost of advanced development was previously estimated at$7 million. Thus, the marginal cost to "fly the brassboard" as Rockwellproposed, is an additional $3.7 million.
This flight demonstration meets the NASA's cost target. In addition, itreduces the technical risk by verifying remote fluid transfer in amicro-gravity environment. Furthermore, the second flight demonstrationbuilds on the momentum of the first flight experiment (MUTE) to gain potentialuser interest and support for remote resupply services.
The schedules for the brassboard and the subcontracted effort for developingthe pump and compressor are integrated. The schedule and manpower estimate isshown in Figure 47. The development effort by NASA for the remote quickdisconnects and micro-gravity gaging system are assumed available during thebrassboard assembly to support this schedule.
The component subcontractors are scheduled to deliver a prototype unit for thepump and compressor sixteen months after program start. The flightdemonstration brassboard will test remote fluid transfer initially with aprototype of the critical ERM components and water solution fluids. Four tosix months later, the flight test hardware will be delivered, integrated, andtested on the brassboard.
During the nominal six-month period for payload integration at KSC, we plan toperform fluid transfer testing with the actual hypergolic fluids planned forthe flight demonstration.
The Flight Demonstration Test is scheduled for 30 - 36 months after programstart, with a six-month post analysis of the results. Given a completelysuccessful flight test, the ERM phase C/D program could start as early as 30months from the start of advanced development and flight demonstrationprograms.
The grass-roots {detailed engineering) cost estimate is uncertain because thefour key variables - direct labor hours, labor rates, direct material , and thesubcontracted cost to develop the pump and compressor - are uncertain. Toprovide a cost range, likely values were estimated for each variable andpresented in Table 25. The best and worst combinations of cost conditionsresults in the low and high cost estimates, respectively. The most likelycombination of cost conditions results in the most likely cost estimate.
3489e -108-
Test Objective Verify remote propellent resupply with receiver andsupplier hardware.
Customer Benefit Reduce technical risk by verifying that all thecritical ERM components (quick disconnects, pumps,compressor, and propellent management devices) workas designed in a micro-gravity environment.
Early flight demonstration stimulates potentialuser support.
Low cost (S10.7M) due to:
o Integration of advanced development and flightdemonstration program, and
o Less test hardware with our plan to "fly thebrassboard".
Test Schedule 30 - 36 months after start of advance developmentprogram.
TABLE 24. FLIGHT DEMONSTRATION TEST SIGNIFICANT CHARACTERISTICS
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These cost estimates only reflect the estimating error, given the program planas reflected in the schedule and manpower estimates for the FlightDemonstration Test. The program plan however, and hence the cost estimate,reflects a success-oriented program with no significant difficultiesencountered.
The low cost estimate reflects that there is only a 10% chance that the costof the experiment could be lower and the high cost estimate reflects thatthere is only a 10% chance that the cost could be higher. The probabilitycurve between these two extreme points was spread by a 35/65 Beta distributein function (ogive curve) and is displayed in Figure 48. The expected cost(50% cummulative probability that the actual cost could be less than orgreater than this cost estimate) is $11.5 million. The best estimate costfrom Table 25 was $10.0 million. The true best estimate is somewhere in thevicinity of these two discrete estimates. If the mid-point is taken, then thebest estimate for the Flight Demonstration Test is $10.7 million.
Given cost target of $12M for the Advanced Development and FlightDemonstration Program, Rockwell has designed a program that not only to meetthis goal but has cut the cost of the Flight Demonstration Test sufficientlyto fund an additional flight test (the MUTE). The significant cost reductionitems 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 costreduction.
One, we propose to "fly the brassboard". Instead of two separate testarticles, 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 tanksfrom the Orbiter logistical spares inventory instead of purchasing themoutright as a separate cost item. This could avoid costs of up to two milliondollars.
3489e -113-
ORIGINAL PAGE ISOf POOR QUALFTY
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3474e -114-
Third, we propose a top-and-bottom test structure rather than our originalside-by-side test structure. Not only will the weight and, hence, structurecost be less, but the flight transportation expense will decline because thespace usage in the Orbiter cargo bay is cut in half. Presently, the width ofour Flight Demonstration Test structure is only slightly wider than the RCStank, or 39 inches.
The Flight Demonstration Test is a simple, low cost experiment to verifyremote propellent transfer between a simulated supplier and a receiverspacecraft. This demonstration was selected because its contributes is toNASA's objectives to reduce technical risk and gain user interest and supportfor remote resupply services.
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REFERENCES:
1. Pace, S. N., "Development Scenario and Requirements for ExpendablesResupply 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, IRSDProject 252, RI IL 287-201-85-111, August 22, 1985.
3. H. A. Saxton, "Pump Transfer of Propellant, Task Analysis - FluidTransfer Study. A Transmittal of REsults for IR4D Project 283,November 21 , 1984.
4. Compressor Transfer of Pressurants, Task Analysis - Storable FluidSystem 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, STS85-0174, 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 ModuleFollow-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 Resultsfor subtask 2.5 IR&E Project 286; H. A. Saxton to J. J. Gernand, July27, 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 TheirRelative 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 TheirRelative Costs" IL AE-MOA-85-184, dated June 4, 1985.
16. Matus, S. J., "Expendables Resupply Brass Board Test System" IL385-121-85-065, dated June 11 , 1985.
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