L 

4 

Contract NAS 9-17852 

DRL-6 

DRD MA-183TJ 

F88-04 

SUPERFLUID HEL- TANKER (SFHT) STUDY 

Final Report 

1 October 1988 

(&AS-CEI- 1721 16) SUPLBPLDID GELIUB TAYKBE 889-185 18 

($€E¶) S'IODY (Iall Aerospace SyEtems Div-) 

157 p CSCL 228 

Unclar 

63/38 0 19 1256 
F88-04 

SUPERFLUID HELIUM TANKER (SFHT) STUDY 

Final Report 

F88-04 

Approved by: 

1 October 1988 

Contract NAS 9-17852 

DU-6 

DRD MA-183TJ 

Richard Hopkins 

Study Manager 

F8884112AB8-1 
. F88-04 

TABLE OF CONTENTS 

Section 

1 INTRODUCTION ............................................... 2 DESIGN DESCRIPTION ......................................... 2.1 DEWAR SUBSYSTEM ....................................... 2.1.1 Major Configuration Trades ..................... 2.1.2 Mechanical/Structural Subsystem ................ 2.1.3 Fluid Subsystem ................................ 2.1.4 Liquid Acquisition System ...................... 2.1.5 Thermal Subsystem .............................. 2.1.6 Dewar Instrumentation and Cabling .............. 2.2 FLIGHT AVIONICS SUBSYSm ............................. 2.2.1 System Architecture ............................. 2.2.2 Demands on Host Vehicle Resources ............... 2.2.3 Role of Artificial Intelligence ................. 2.2.4 System Control Station (SCS) .................... 2.2.5 Dewar Monitor and Control Avionics .............. 2.2.6 Operations ...................................... 2.2.7 Redundancy ...................................... 2.2.8 Adapting to Different Host Vehicles ............. 2.2.9 Deviations from the OSCRS Concept ............... 2.2.10 Development Issues .............................. 2.3 GROUND SUPPORT EQUIPMENT .............................. 2.3.1 GSE Design Concept ............................. 2.3.2 Ground Operations/Facilities .................... 3 HSSION OPERATIONS ......................................... 3.1 HOST/USER INTERFACES AND OPTIONS ....................... 3.1.1 Initial Capabilities ........................... 3.1.2 Automatic Resupply Capability .................. 3.2 RESUPPLY OPERATIONS ................................... 3.3 USER IMPACTS ........................................... 3.4 SAFETY ................................................. 3.4.1 Cryogen Tank Pressure Relief .................... 3.4.2 Vacuum Shell Pressure Relief .................... 3.4.3 Plumbing Pressure Relief ........................ 4 ANALYSIS .................................................... 4.1 STRUCTURAL ANALYSIS ................................... 4.1.1 Sizing of Shells ............................... 4.1.2 Siring of Cryogen Tank Support Straps .......... 4.2 THERMAL ANALYSIS ...................................... 4.2.1 Model Description .............................. 4.2.2 Baseline System Performance .................... 4.2.3 Sensitivity Studies ............................ 4.2.4 Configuration Tradeoffs ........................ 4.3 VENTING ANALYSIS ...................................... 4.3.1 Low-Flow (Standby) Vent ........................ 4.3.2 High-Flow Vent ................................. 4.3.3 Emergency Vent ................................. 4.4 SUPERFLUID HELIUM TRANSFER ANALYSIS ................... 4.4.1 Modeling Ground Rules .......................... 4.4.2 Baseline System Description .................... 4.4.3 Transfer Analysis Results ....................... 4.4.4 Conclusions from Transfer Analysis .............. 

* 

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F8804112AB3-ii ii 
F88-04 

TABLE OF CONTENTS (Continued) 

Section 

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8 

Figure 

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COUONALITY ISSUES ......................................... 5.1 LIQUID ACQUISITION DE VICE .............................. TECHNOLOGY DEYELOPMENT ..................................... PROGRAM PLAN ................................................ 7.1 HARDWARE TO BE DEVELOPED ............................... 7.2 PROGRAM SCHEDULE ....................................... 7.3 RISK ................................................ 7.4 WKE-OR-BUY PLAN ....................................... 7.5 LONG-LEAD PROCUREMENTS ................................. 7.6 FACILITIES ............................................. CONCLUSIONS AND RECOMMENDATIONS ............................ REFERENCES .................................................. 

Study logic flow ........................................... Dewar configuration for SS-based resupply ................... SFHT system block diagram .................................. Dewar configuration for SS-based resupply .................. Dewar configuration for shuttle-based resupply ............. Oblate spheroid dewar configuration ........................ Cutaway view of dewar ...................................... Access to internal burst discs ............................. Fluid management schematic ................................. Liquid acquisition requirements ............................ .Liquid acquisition design options .......................... Thermomechanical pump concept .............................. Two components of He I1 ..................................... Why LAD pumps .............................................. LAD design criteria ........................................ Insulation system design ................................... Concept for dewar/mechanical cooler interface .............. Locations of dewar instrumentation ......................... Flowmeter concept .......................................... SFHT avionics has two segments ............................. SCS to STS interface ....................................... Dewar/SCS electrical interface ............................. System control station ...................................... SFHT control panel .......................................... Dewar monitor and control .................................. Three string - majority voting of commands .................. SS-SFHT electrical interface ............................... OW-SFHT electronic interface ............................... Alternative fill process .................................... COBE GSE ................................................ Active cooling using external J-T valve and external 

heat exchanger .............................................. Active cooling using internal J-T valve and heat exchanger .. Active cooling using internal J-T valve, heat exchanger, 

and ejector ................................................ Internally produced refrigeration using pressurized 

external storage tanker ..................................... 

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... F8804112AB3-iii 111 
F88-04 

TABLE OF CONTENTS (Continued) 

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Temperature-entropy diagram of helium showing the cooling 

process ................................................ Preparation of SFHT for launch through the vehicle 

processing facility ......................................... Schedule to prepare SFHT for launch with vertical 

procesesing ................................................ Preparation of SFHT for launch through the horizontal 

processing facility ......................................... Schedule to prepare SFHT for launch with horizontal 

processing ................................................ Initial capability . EVA required to connect avionics and 

fluid lines ................................................ Automatic connection of avionics and fluid lines ........... SFHT resupply process flow chart ........................... Operations for shuttle-based resupply ...................... AXAF XRS fluid management schematic ......................... Strap fatigue test results at R=0.05 and room temperature .. Schematic of thermal math model ............................ Ground lock-up capability .................................. Orbital pumpdown time ...................................... Standby (low-f low) vent .................................... Orbital pumpdown (high-flow vent) .......................... Emergency vent ............................................. Transfer system plumbing schematic ......................... Transfer system diameter profile ........................... Thermomechanical pump configuration ........................ Full-f low (1 , OOO t/hr) temperature profile .................. Full-f low (1 . OOO t/hr) pressure profile ........ i ........... Full-flow (1. OOO &/hr) heat flux profile ................... Minimum-flow (70 t/hr) heat flu profile .................... Minimum-f low (70 &/hr) temperature profile .................. Minimum-flow (70 &/hr) pressure profile ..................... Lad performance in oxygen and hydrogen for outflow rates 

of 10 and 100 percent of tank volume per hour ............... SFHT flight hardware breakdown .............................. SFHT program schedule ....................................... SFHT dewar manufactuirng process and facility usage ......... 

ESTIMATED LAUNCH COST FOR A VARIETY OF DEWAR SHAPES ........ FLUID OPERATIONS OVERVIEW .................................. VALVE POSITIONS FOR VARIOUS OPERATIONS ..................... EQUATIONS GOVERNING SUPERFLUID ............................. EXPECTED LAD PERFORMANCE ................................... DEWAR SENSORS AND CABLING .................................. POWER BUDGET ............................................... COMPATDILITY WITH SPACE STATION/OMV ....................... POWER BUDGET ............................................... STRUCTURAL DESIGN CONDITIONS AND CRITERIA .................. BREAKDOWN OF HELIUM USAGE .................................. SENSITMTY ANALYSES ....................................... 

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F8804112AB3-iv iv 
F88-04 

TABLE OF CONTENTS (Concluded) 

- Table 

4-4 EFFECT OF MECHANICAL COOLER FAILURE ON DEWAR PERFORMANCE ... 4-10 

4-5 MASS PENALTIES ASSOCIATED WITH CONFIGURATION TRADEOFFS ..... 4-11 

4-6 RESULTS SUMMARY 4-23 

4-7 RESULTS SUMMARY ............................................ 4-25 

5-1 ASSESSMENT OF USE OF SFET ELEMENTS ON 

SUBCRITICAL/SUPERCRITICAL TANKERS .......................... 5-1 

6-1 DETERMINING SFHT TECHNOLOGY DEYELOPMENT REQUIREMENTS ....... 6-2 

............................................ 

F8884112AB3-v V 
F88-04 

FOREWORD 

This report documents results of the Superfluid Helium Tanker Study performed 

for NASA Lyndon E. Johnson Spaceflight Center by Ball Aerospace Systems Group 

in Boulder, Colorado. The NASA technical manager for this contract, NAS 

9-17852, is Mr. William C. Boyd of the JSC Propulsion and Power Division. As 

required by the Statement of Work, the study effort included: Task 1 - Collection of Requirements, Task 2 - Fluid Subsystem Conceptual Design, Task 3 - 

System Analysis and Design, Task 4 - Commonality Assessment and Technology 

Development Recommendations, and Task 5 - Development Program Plan. 

F8804112ABS-vi vi 
F88-04 

ACRONYMS AND ABBREVIATIONS 

AFD 

AP 

AUM 

AXAF 

BUD 

CCTV 

CF 

CITE 

COBE 

CET 

CT 

CWEA 

EGSE 

Ec 

ELV 

EPROM 

EVA 

GPC 

GSE 

IRAS 

J-T 

JSC 

LAD 

LCD 

LM 

LTV 

MLI 

MMS/FSS 

MMPS/&PF 

MRMS 

ON 

OMV 

OPF 

OSCRS 

PGEM 

PHSF 

PPF 

RCS 

RMS 

RSS 

RV 

SAEF 2 

scs 

SEE 

Sf He 

SFHT 

SHOOT 

SINDA 

SIP 

SIRTF 

SMRM 

SRV 

ss 

F8804112AB6-vii 

AFT Flight Deck 

Attached Payload 

Automatic Umbilical Mechanism 

Advanced X-Ray Astrophysics Facility 

Ball Aerospace Systems Division 

Closed Circuit Television 

Centrifugal 

Cargo Integration Test Equipment 

Cosmic Background Explorer 

Cathode Ray Tube 

Cryogen Tank 

Caution and Warning Electronics Assembly 

Electrical Ground Support Equipment 

Electro-luminescent 

Expendable Launch Vehicle 

Erasable Programmable Read Only Memory 

Extravehicular Activity 

General Purpose Computer 

Ground Support Equipment 

Infrared Astronomical Satellite 

Joule-Thomson 

Johnson Space Center 

Liquid Acquisition Device 

Liquid Crystal Display 

Lab Module 

Low-thrust Vent 

Multilayer Insulation 

Multimission Modular Spacecraft/Flight Support System 

Microgravity and Materials Processing Facility/Critical Point 

Phenomena Facility 

Mobile Remote Manipulator System 

Operations and Checkout Building 

Orbital Maneuvering Vehicle 

Orbiter Processing Facility 

Orbital Spacecraft Consumable Resupply System 

Payload Ground Handling Mechanism 

Payload Hazardous Servicing Facility 

Payload Processing Facility 

Reaction Control System 

Remote Manipulator System 

Rotating Service Structure 

Relief Valve 

Spacecraft Assembly and Encapsulation Facility, No. 2 

System Control Station 

Standard End Effector 

Superfluid Helium 

Superfluid Helium Tanker 

Superfluid Helium On-Orbit Transfer 

Systems Improved Numerical Differencing Analyzer 

Standard Interface Panels 

Space Infrared Telescope Facility 

Solar Maximum Repair Mission 

Short Range Vehicle 

Space Station 

vii 
F88-04 

SSF 

SSP 

STE 

STICCR 

STS 

TDRSS 

TM 

TPDM 

vcs 

VPF 

WED 

XRS 

Satellite Servicing Facility 

Standard Switch Panel 

Special Test Equipment 

SIRTF Telescope Instrument Changeout and Cryogen Replenishment 

Space Transportation System 

Tracking and Data Relay Satellite System 

Thermomechanical 

Three Point Docking Mechanism 

Vapor-Cooled Shield 

Vertical Processing Facility 

Vertical Payload Handling Device 

X-Ray Spectrometer 

viii 
ORIGINAL PAGE IS 

OF POOR QUALITY Section 1 

F88-04 

SUMMARY 

On-orbit fluid resupply is an economical approach to several aspects of the 

United States space program in the near future. The first is providing propellants, primarily hydrazines in the near future and eventually also liquid 

hydrogen, to extend satellite lifetimes by replenishing their station-keeping 

fuel and to refuel orbital maneuvering and transport vehicles needed to support manned presence in space. The second is replenishment of consumables at 

the space station (SS), such as nitrogen needed for the life support system. 

The current baseline is to carry nitrogen to the SS in the form of supercritical (cryogenic) nitrogen. This can be done with current technology. Launch 

mass and cost can be reduced by carrying liquid (two-phase), rather than 

supercritical, cryogen into space. For that approach technology development 

in the areas of liquid acquisition and liquid/vapor phase separation is needed. The third major application for fluid resupply is to replenish instrument systems using stored cryogens for cooling. Lifetimes of these systems 

are limited by the amount of cryogen that can be carried into orbit within 

mass and envelope constraints. This third application, specifically helium 

resupply, is addressed by this study. 

Development of helium resupply technology is most urgent compared to resupply 

of other fluids because the potential applications are more immediate. The 

nearest and highest priority applications are the Particle Astrophysics Magnet Facility (Astromag), under development at NASA Goddard Space Flight Center, and the Space Infrared Telescope Facility (SIRTF), under development at 

NASA Ames Research Center. Resupply of these systems is scheduled to occur 

about the year 2000. SIRTF is the major driver in designing the Superfluid 

Helium Tanker (SFBT) because it requires the greatest quantity of helium. To 

maximize .the value of this study, effort was focused on those issues and 

design features that impact cost (development and operations), risk 

(technical and programmatic), and interface parameters. This includes items 

such as dewar size, transfer system, and fluid operations. Minimal effort 

will be spent on the detailed definition of items that require 

straightforward technology and don’t significantly impact the design of any 

other parts of the system. 

This report documents accomplishments and recommendations of the two-phase 

SFHT study. During the first phase of the study, the emphasis was on defining a comprehensive set of user requirements, establishing SFHT interface 

parameters and design requirements, and selecting a fluid subsystem design 

concept. During the second phase, an overall system design concept has been 

constructed based on appropriate analyses and more detailed definition of 

requirements. Achieving commonality with the orbital spacecraft consumables 

resupply system (OSCRS) design has been a consideration, especially related 

to the avionics. Modifications needed to extend the baseline for use with 

cryogens other than SfHe have been determined, and technology development 

needs related to the recommended design have been assessed. 

Figure 1-1 shows the logic flow of the study. Some cost estimates have been 

made in support of configuration and operation tradeoffs; these are shown. 

During the final study phase, analysis of the on-orbit transfer system has 

been conducted in greater detail than originally planned to verify concept 

feasibility, identify design drivers, and establish performance parameters. 

This was advisable since the major design uncertainties and technology development issues are related to this aspect of the system. 

F8804112AA5-1 1-1 
F88-04 

IDENTIFY MAJOR SUBSYSTEMS 

* USERS 

LAUNCH VEHICLE 

.SITE OF RESUPPLY 

OMV 

GSE 

USER REQUIREMENTS 

SUBSYSTEM 

INTERFACES 

AND 

DESIGN CONSTRAINTS 

Task 1 

1 *wo PRELIMINARY DESIGN 

IC 

SFHT DESIGN 

ALTERNATIVES 

DESIGN AND 

REQUIREMENT 

ITERATIONS 

c 

DESIGN TRADEOFFS 

* COST 

* COMMONALITY 

I * RISK 

FINAL 

REVIEW 

DESIGN TRADEOFFS 

FLUID 

Task 2 

I 

RECOMMENDED 

IMPACTS TO PERFORMANCE 

I T8Sk 3 I 

I 

RECOMMENDATIONS 

PLAN DESIGN AND DEVELOPMENT AND 

DEVELOPMENT RE TECHNOLOGY 

PERFORMANCE COMMONALITY 

I 

Taak 5 I Task 4 r---------------------- 

USING RESULTS FROM STICCR STUDIES 

** USING RESULTS FROM SHOOT PROGRAM 

Figure 1-1 Study logic flow . 

F8804112AA5-2 1-2 
F88-04 

Major system design requirements are summarized below: 

0 Provide 8,900 L (based on resupply of warm SIRTF) of SfHe after 9 

months of orbital standby (at the SS). 

0 Launch on STS with Titan IV as an alternative. 

0 Perform resupply in the shuttle bay with the SS as an alternative. 

0 Resupply a variety of users, with primary consideration given to 

Astromag, SIRTF, AXAF XRS, MMPF/ CPPF, and Lambda-point 

Experiment. 

0 Perform host and user mating functions by EVA in the initial concept, but consider extension to automatic mating. 

0 Be compatible with all associated interface and design constraints 

imposed by other involved systems. 

0 Emphasize low-risk, low-cost approaches. 

The baseline design concept is shown in Figure 1-2 for resupply at the SS. 

For shuttle-based resupply, the configuration is essentially identical except 

for deletion of the mechanical coolers and associated electronics. Design 

features are as follows: 

0 A dewar of 10,000-E capacity containing SfHe at launch (this 

capacity is adequate for resupply of a warm SIRTF or for resupply 

of several smaller users). 

0 A spherical dewar configuration to minimize mass, launch cost, and 

A TM pump and a porous medium distributed within the tank that 

development cost. 

0 

uses the thermomechanical effect driven by heat from the pump to 

accomplish liquid acquisition. 

0 A state-of-the-art insulation system using carefully-constructed, 

double-aluminized mylar MLI and alumina/epoxy tension strap 

-supports. 

0 A plumbing system that provides operational flexibility and meets 

all safety-related requirements. 

0 Four Stirling cycle mechanical coolers to provide a mass-optimized 

system (and cost-optimized program) for a mission including 

9-month standby. 

0 An interfacing approach (mechanical, electrical, fluid, and avionics) that is compatible with resupply either in the shuttle bay or 

at the SS, using EVA-assisted electrical and fluid hookups. 

0 A "smart" avionics subsystem that uses a three-string, majorityvoting approach for control of the SFHT. 

F8804112AA5-3 1-3 

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F88 -04 

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A SFHT design concept has been recommended that meets all mission requirements using low-cost and low-risk approaches. The all-aluminum, 10,000 & 

spherical dewar will perform a warm SIRTF resupply at either the shuttle or 

the space station (after a 9-month standby). The insulation system design is 

optimized to provide minimum launch mass based on the 9-month standby 

requirement; mechanical coolers are used to cool the dewar outer VCS during 

the extended SS standby. 

An oblate spheroid configuration was investigated to see if it would reduce 

shuttle launch cost; it has essentially the same launch cost and thermal 

performance as the spherical dewar, which is 1,600 lb lighter. It requires 

about 19 in. less shuttle bay length, providing better manifesting flexibility. Although the spherical shape has been selected as the baseline, there 

is not a strong preference. 

The study has emphasized the areas of technology risk, which center around 

the fluid subsystem design and operations. Technology development is needed 

for the follow-on automatic system, but only a cursory evaluation of those 

items was appropriate within the scope of this study. When the SFHT hardware 

is developed, state-of-the-art, space-qualified technology will be used for 

the avionics. Therefore, little study effort was spent on this topic. Regarding avionics, the study was more concerned with assuring that SFHT power 

and data requirements are compatible with shuttle and SS resources. Considerable software development will be required for the avionics "smart" 

computer system. 

BASG has a proven capability for designing and fabricating spaceborne dewars 

with high thermal performance. Although this is not a technology risk area, 

considerable study effort was spent on providing valid mass and thermal performance estimates for the dewar. This was necessary to make legitimate 

configuration tradeoffs and plan operations, both ground and flight. Also, 

minimizing the parasitic heat input to the cryogen tank will minimize the 

SfHe bath temperature and maximize the transfer efficiency. 

The primary technology development area for the SFHT is the transfer system, 

consisting of the liquid acquisition device, thermomechanical pump, SFHT 

plumbing, flexible transfer line, and user plumbing. Demonstration of this 

system is the main function of the SHOOT experiment. However, there are 

important scaling factors and detailed design impacts that can only be determined through a rigorous analytical understanding of the system. Regardless 

of ground component testing and SHOOT flight testing, a valid analytical 

model of the entire transfer system is required to establish the SFHT design 

with any confidence. 

Our approach has been to take a very careful end-to-end look at a point 

transfer system design based on the fundamental physics involved, therefore 

verifying the feasibility of performing on-orbit superfluid helium resupply. 

We conclude that it can be done within the imposed design constraints, but 

the system design and operation must be very carefully planned and verified. 

The next step is to construct a comprehensive analytic tool for understanding 

the system better, particularly the transient aspects of the resupply 

process. 

That analytic capability does not yet exist. 

F8804112AA5-5 1-5 
F88-04 

Section 2 

DESIGN DESCRIPTION 

The SFHT consists of two major subsystems: 

seen in Figure 2-1. 

and transport the SFHT and to prepare the dewar for launch. 

the dewar and the avionics, as 

There is also a large complement of GSE needed to handle 

The dewar is a 10,000 C. spherical configuration with the cryogen tank and 

vacuum shell made of aluminum (Figure 2-2). Mounted on the vacuum shell for 

SS-based resupply are four Stirling cycle coolers, which provide 9 W of 

cooling at the outer VCS. Also mounted on the vacuum shell are the 

electronics associated with the coolers, the dewar-based portion of the 

avionics, and the redundant transfer lines, shown in the stored position for 

launch. The dewar mounts into the shuttle bay with standard sill and keel 

fittings, which are adaptable for launch on a Titan IV. Not shown in the 

figure is the avionics remote control station, consisting of computers, 

displays, touch panels, and manual switches. 

Figure 2-3 shows the configuration for shuttle-based resupply with the mechanical coolers and extra grapple deleted. 

This section describes the major configuration trades conducted and the recommended design concepts for the dewar, avionics, and GSE. 

2.1 DEWAR SUBSYSTEM 

The dewar is divided into structural/mechanical, thermal, fluid, and instruBecause of the importance of the pump and LAD, 

these are discussed in a separate section even though they are part of the 

fluid subsystem. 

. mentation/cabling subsystems. 

2.1.1 Major Configuration Trades 

Analyses used to predict system performance characteristics and to aid the 

configuration trades are discussed in Section 4. Most of the trades relating 

to the fluid subsystem were performed during the first phase of the study and 

documented in the Interim Progress Report; the most notable trade results are 

selection of the thermomechanical pump, the helium is superfluid at launch, 

and the system must land with a full helium load. 

Mechanical Cooler 

Fluid loss during the 9-month orbital standby and pressure rise rate during 

ground lock-up are minimized by cooling the dewar outer VCS with mechanical 

coolers. The British Aerospace Stirling cycle machine is chosen for this 

purpose. A unit producing 1 W of cooling at 93 K has been qualified for the 

ISMS instrument to be flown on the Upper Atmosphere Research Satellite. 

This same cooler is baselined as part of the AXAF X-ray Spectrometer, under 

development at NASA/GSFC. The cooling power achievable by simple extension 

of the existing technology is 3-4 W . Our baseline uses a cooler producing 

3 W at 93 K; the system includes four of these operating simultaneously. As 

shown in Figure 3-7 of the Interim Program Report, the mass savings achieved 

by using a mechanical cooler continues to increase up to a cooling power of 

F8804112AA8-1 2- 1 
F88-04 

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F8804112AA8-2 

Figure 2-1 SFHT system block diagram 

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at least 12 W. 

four. 

pursuing the optimization further. 

The optimum number of coolers may indeed be greater than 

They are costly items, and four were chosen for the baseline without 

As discussed in 4.2.4, the estimated SFHT launch mass savings resulting from 

use of the coolers is about 600 lb. At a cost of $5,00O/lb per launch, this 

is a savings of $3 M per launch. The estimated development program cost 

impact (through delivery of the first flight system) resulting from the mechanical cooler approach baselined is therefore approximately compensated in 

the first three launches. Since the system is designed for 50 launches, the 

cost advantage of using the coolers is clear. 

Dewar Shape 

STS launch cost is based on a% algorithm accounting for the system mass and 

the length of bay space used . Cost is determined either by mass or bay 

space, whichever is the critical parameter according to the algorithm. Approximate mass and launch cost estimates were generated for a variety of 

dewar shapes, as shown in Table 2-1. These estimates indicate that the special shape is the lightest (as expected), but the d2 oblate spheroid shape 

may have minimum launch cost. More detailed mass estimates and corresponding 

launch cost estimates were then made for these two shapes. The oblate spheroid configuration is shown in Figure 2-4. Results indicate essentially 

equivalent launch costs and a spherical shape mass savings of 1,600 lb. 

ESTIMATED LAUNCH 

CONFIGURATION SPHERE 

Fluid Mass, lb 3,200 

Dry Mass, lb 3.820 

Total Mass, lb 7,020 

Bay Length, in. 117 

Launch Cost, $M 31.9 

Table 2-1 

COST FOR A VARIETY OF DEWAR SHAPES 

EL/2 : 1 EL/ 1.4 : 1 TOR/2r TOR/1 .7r 

3,200 3,200 3,200 3,200 

7,290 5,030 9,920 7,810 

10; 490 8,230 13,120 11,010 

88 98 80 86 

39.3 31.9 47.9 41 .O 

The spherical shape is advantageous with regard to mass, center-of-gravity 

location relative to the shuttle launch axis, and probably fabrication cost. 

The only advantage of the more compact oblate spheroid shape is potentially 

superior manifesting flexibility. Thermal performance is only slightly better, especially if mechanical coolers are used, for the spherical shape. 

Our choice is the spherical shape. 

Mechanical Interfacing 

Two docking and berthing mechanical interfaces were evaluated for the SFHT. 

Both have been successfully used in the space program. The standard end 

effector, as used on the orbiter RMS, and the three point docking mechanism, 

as used on the flight support system, were selected to be used on the SFHT. 

The SEE on the RMS has been used to remove large (as much as 32,000 lb) 

payload from the orbiter payload bay. The grapple is the mating interface to 

F8804112AA8-5 2-5 
F88-04 

I 

Figure 2-4 Oblate spheroid dewar configuration 

F8804112AA8-6 2-6 
F88-04 

the SEE. 

FSS and its TPDM in 1984. 

The Solar Maximim Mission (SMM) program was the first user of the 

The SEE does not have the capability of providing the forces required to 

automatically connect and disconnect a number of electrical and fluid connections. Yet it is simple to operate, making it a good choice for earlier SFHT 

missions where fluid and electrical connections will be made by EVA. One 

special purpose electrical connector can be provided on the SEE remote connection, but this provides an insufficient number of pins for the SFHT. 

The TPDM in its present configuration incorporates two 56 pin connectors, 

which is sufficient for the user to SFHT interface. There is sufficient 

space in the TPDM to add the automatic fluid connections. The TPDM is recommended for future missions after development of an automatic fluid coupling. 

The SFHT can easily be designed to facilitate the use of either the SEE or 

TPDM. Both would be easily removed and replaced on the SFHT to accommodate 

mating with user facilities. Failure tolerance of both approaches should be 

examined more closely to ensure that SFHT requirements are met. 

Fill Procedure 

The recommended fill procedure is the same as that used for IRAS and COBE. 

This is a proven approach that has minimal impact on the SFHT design. The 

alternatives considered and the rationale for this choice are discussed in 

2.3.1. 

2.1.2 Mechanical/Structural Subsystem 

The spherical dewar is mounted in the shuttle bay with standard sill and keel 

trunnion fittings, as shown in Figure 2-2. There is an additional ground 

handling trunnion that is removed before flight. Host and user interfaces 

are discussed further in Section 3.1. 

The cryogen tank consists of two spun and machined 5083 aluminum hemispheres 

welded together to make a 10,000 sphere, as shown in Figure 2-5. A 

thickened area is provided around the equator for the weld and for simple 

bolted attachment of internal plumbing components. Thickened rings are 

provided on each hemisphere at an elevation of about 30 deg for attachment of 

the twelve straps which support the cryogen tank from the vacuum shell. 

Aluminum-to-stainless steel transition joints are used where the various 

plumbing lines penetrate the tank. 

The vacuum shell is made of spun and machined 5083 aluminum spherical domes 

welded to a girth ring. The cryogen tank support straps are attached inside 

the girth ring. Pins and clevises are used at each end of the straps. The 

straps are made of alumina/epoxy to provide maximum thermal isolation. 

The two vapor-cooled shields are 0.04-in. thick and made of 1100-0 aluminum 

to provide high lateral thermal conductivity. They are supported by attachment to the support straps, and in turn they cool the straps. 

The four trunnions are properly positioned by mounting them on plates bolted 

to the girth ring. Mechanical interfacing with the Titan IV can be simply 

F8804112AA8-7 2-7 
F8804112AA8-8 

F88-04 

Figure 2-5 Cutaway view of dewar 

2-8 

~ 
F88-04 

accomplished by providing the appropriate interface structure bolted to the 

girth ring. 

All of the external hardware, i.e., the standard end effector, the transfer 

lines, and the dewar-based avionics, are mounted on the girth ring. If resupply is to be performed at the SS, there is additional equipment mounted to 

the girth ring: a grapple for attaching the SFHT to the SS, the four Stirling cycle coolers, and the five electronic boxes for the coolers. Mounting 

of this equipment is shown in Figure 2-2. 

The internal plumbing manifold, consisting of sixteen valves, two burst 

discs, and three electrical feedthroughs, is attached to the equator of the 

cryogen tank over a span of about 90 deg. The most critical items for access 

are the burst discs, although access to the entire manifold is certainly 

feasible. Providing access to all components of the assembly would result in 

a small system mass penalty. As shown in Figure 2-6, access is gained by 

removing a section of the girth ring, folding the MLI back like petals of a 

flower, and removing bolted sections from the vapor-cooled shields. Sealing 

of the girth ring removable plate can be by grind-down weld or by '0' ring. 

To ensure a reliable seal and because the need for access will be infrequent 

or never, the grind-down weld is the chosen approach. This may seem cumbersome, but it is a well-established technique and is probably a small effort 

compared to repairing the cause of the loss of guard vacuum and the potential 

resultant damage to the insulation system. Access to the two warm burst 

discs located in the girth ring is provided through this same cover. 

2.1.3 Fluid Subsystem 

This section summarizes the required fluid operations and describes the overall fluid subsystem' design. Further details regarding the liquid acquisition 

device and thermomechanical pump are provided in the next section. Detailed 

configuration trades leading to the recommended fluid subsystem design were 

performed during the first phase of the study and documented in the Interim 

Progress Report. 

The fl 

Table 

shown 

.uid operations occurring during each phse of the mission are shown in 

2-2. The fluid management system needed to perform these operations is 

in Figure 2-7. This design is based on the following considerations: 

0 Two-fault tolerance for safety and one-fault tolerance for mission 

success 

0 Maximum use of existing and low-risk technology 

0 Minimum number of components and valve actuations 

0 Minimal degradation of dewar thermal performance 

0 Access to critical components 

0 All hermetic seals between helium and guard vacuum 

The GSE connects to the bayonet couplings at the fill and high-flow vent 

lines. Valve positions for the various operations are given in Table 2-3. 

F8804112AA8-9 2-9 
7 

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c( 

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W 

v> 

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w 

c3 W 

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a W Y c1s 

CY I 

r--- 

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F88-04 

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F8804112AA8-10 2-10 
F88-04 

Table 2-2 

FLUID OPERATIONS OVERVIEW 

TEST AND LAUNCH PREF 

(BEFORE PIL I MAINTENANCE I INTEGRATION 

PRE-LAUNCH 

(STS OR ELV) 

POST-LAUNCH 

(STS OR ELV) 

ORBITAL 

TRANSPORT 

(OMV) 

DESCENT 

SHUTTLE RESUPPLY 

OPERATION 

NORMAL flLL 

BATH PUMPDOWN 

(ON GROUND) 

INSULATION 

COOLDOWN 

LOW-PRESSURE 

TOPOFF~FILL~ 

STANDBY VENTING I Ttut:E I TO GSE 

PUMP 

TO 

AMBIENT 

TO 

AMBIENT 

TO 

AMBIENT 

TO AMBIENT 

OR 

OVERBOARD 

I0 

TO 

AMBIENT 

TO 

AMBIENT 

TO 

AMBIENT 

TO AMBIENT 

OR 

CONT. VENT 

TO 

AMBIENT 

TO AMBIENT 

OR 

OVERBOARD 

TO 

AMBIENT 

TO AMBIENT 

OR 

OVERBOARD 

AMBIENT AMBIENT 

AMBIENT AMBIENT 

(NHo BATH) 

TO 

AMBIENT 

BATH PUMPDOWN 

(COOLDOWNIFILL) 

1 MAYBE NOT NEEDED IF GSE PROVIDES SUPERFLUID TRANSFER 

2 FLUID TRANSFERRED NEAR OR BELOW THE UMBDA POINT 

F8804112AA8-11 2-11 
ORfGlNAL PAGE IS 

OF ~OOR QUALITY 

-- WAC 1 WAC I 

STAN 0 BY 

LTV P 

BC - Commercial Bayonet Coupling 

BD -BurstDlac 

EC - Electrlcal Connector 

EF - Eiectrkal Feedthru 

EVAC - EVA Coupling 

HFPP - High-Flow Porous Plug IVCS - Inner Vapor-cooled Shield 

LAD - Liquid A ulstkm Devh 

OVCS - Outer Vaporcooled Shield 8 - Motor-operated Value RD - Relief DevicdPumpsut Port 

TM - Thermomechanlal 

PV - Purge Valve (Manual) 

LFPP - Low-Flow "b orous Plug 

VACUUM snu 

t 

EMERGENCY HIGH-FLOW 

VENT VENT 

lnl! c 

I I 

I 

I 

j 

CRYOGEN TANK 

Wl 01M 07 

Figure 2-7 Fluid management schematic 

F8804112AA8-12 2-12 
Table 2-3 

VALVE POSITIONS FOR VARIOUS OPERATIONS 

F88-04 

OPERATION 

COOLDOWN AND 

NORMAL FILL 

PU MPDOW N (GROUND) 

INSULATION COOLDOWN 

LOW-p TOPOFF/FILL 

STANDBY VENTING AND 

ORBITAL PUMPDOWN 

SAFING 

EMERGENCY VENTING 

USER RESUPPLY 

AM 9203MD135.010 

F8804112AA8-13 2-13 
F88-04 

Motor-operated valves will be either the 3/4-inch design or 1/2-inch design 

developed for the SHOOT program. In most valve locations the greater flow 

conductance of the 3/4-inch valve is needed to achieve desired flow rates and 

vent system conductance. Although the system mass would be slightly less if 

1/2-inch valves were used where possible. the development cost would probably 

be greater. For the current baseline we assume ail valves are 3j4-inch: 

Relief valves can be the Circleseal design used on the COBE dewar, and burst 

discs are assumed to be the Ametek-Striza design used on COBE .and SHOOT. 

Suitable electrical feedthroughs are made by Ceramaseal, similar to those 

used in the COBE dewar. A variety of vacuum tight electrical connectors for 

room temperature use are available, such as Deutsch and Cannon. The manual 

purge valve is a simple and readily available item. The combination relief 

device/pumpout port is a piston and "0" ring arrangement made by Cryolab and 

used on the COBE dewar. 

Currently available bayonet couplings, such as used for the COBE dewar, have 

a heat leak of about 1 W, compared to an estimated total on-orbit transfer 

system heat leak on the order of 5 W. The EVA coupling is assumed to be the 

coupling currently under development by JSC for t$e SHOOT program. The heat 

leak of this coupling is estimated by Moog at 1 W ; it is unlikely it can be 

made lower unless requirements, especially the thermal isolator design 

pressure, are changed. As discussed in 4.4.4, reduction of the coupling heat 

leaks would be beneficial to the transfer process, but further analysis and 

testing is needed to adequately determine the importance of this benefit. 

Flowrates through the standby vent and safing vent will be low (<8 mg/s) and 

vapor will exit at room temperature, so that no protective measures are required. Given the relatively large payload thrust allowables of the orbiter 

and the SS, design of the low thrust vents at these locations will be simple. 

A A low-thrust vent will be installed on the high-flow vent 

after the dewar is filled for launch. The flowrate out this vent during 

orbital transfer will be about 1.0 g/s maximum at a temperature less than 

10 K. This vent must be located in a protected area to prevent injury to 

personnel or damage to nearby equipment. The nTT" design is still adequate 

for this vent. 

is adequate. 

The most critical vent is the emergency vent, which may have a flowrate as 

high as 250 g/s at a temperature less than 10 K. This vent must be located 

in a protected area, and the low-thrust vent design may be more complex than 

the others. This flowrate is small enough that it can be vented directly 

into the shuttle bay without causing overpressirization; it therefore need 

not be routed into the overboard cryogenic vent. Analysis of the emergency 

vent system is presented in 4.3.3. As discussed in the Interim Progress 

Report, vent collection at the SS is neither necessary nor desirable. A 

collection system compatible with SFHT operations would be extremely large 

and costly. Also, helium is inert and the gas cloud formed during transfer 

operations will dissipate very quickly; local venting of large amounts of 

helium does not violate SS cleanliness requirements. While on the ground, 

emergency venting is to ambient via the low-thrust vent. 

Since a single failure could prevent mating of the EVA coupling as currently 

designed, redundant fill lines and couplings are provided. During on-orbit 

transfer, the interface with the user is at the EVA coupling. It is assumed 

the mating half of the coupling is welded into the user vacuum shell. For 

SFHT ground operations, the flexible transfer lines are removed, and the GSE I 

F8804112AA8-14 2-14 
F88-04 

connects to bayonet BC1 or BC2. 

vacuum pump for system vacuum purging and bath pumping. 

The dewar orientation required for fill and topoff operations is dictated by 

the location of the high-flow vent within the cryogen tank. Three alternatives are possible: 

The high-flow vent is connected to the GSE 

1. The vent can be located so that 100 percent fill is achievable in 

the given orientation and 50 percent is achievable in the 

orthogonal orientation. 

2. It can be located so that 90 percent fill is achievable in two 

orthogonal orientations. 

3. Four valves (for redundancy) and another plumbing penetration into 

the cryogen tank can be added to the design to allow 100 percent 

fill in two orthogonal orientation. 

Because of the very long lock-up capability of our design (4 months with 

mechanical coolers running and 25 days without coolers), fill operations can 

be performed in a facility where there will be no constraints on dewar orientation. Our approach is to use alternative 1. above; this will minimize cost 

and operational risk factors. 

2.1.4 Liquid Acquisition System 

Acquisition of superfluid helium (SfHe) in zero gravity for transfer at a 

high flowrate is a unique requirement with no established technical approach 

based on flight experience. The unique physical properties and behavior of 

SfHe make experience gained with other fluids not very useful. We have 

therefore taken a top-down system design approach to arriving at the conceptual design of this critical SFHT subsystem. 

. 

The baseline we have chosen looks very much like a conventional screen-chan- . 

ne1 liquid acquisition device (LAD), but its design is significantly different. The properties of the SfHe, combined with the heat flux from the thermomechanical (TM) pump, cause the LAD to assist with pumping, and can assure 

that cavitation in the liquid will not occur. This LAD uses either a finemesh metallic screen or a porous material (either porous "plugs" or membranes) with much higher flow impedances than would be acceptable for other 

fluids. 

2.1.4.1 Subsystem Requirements 

Figure 2-8 breaks down the fundamental performance required of the LAD into 

its component parts. The fundamental performance requirement is that liquid 

be delivered to the pump at 1,000 L/hr, or more accurately at a rate such 

that the overall transfer operation can be completed within the timeline 

based on a nominal rate of 1,000 L/hr. This means that the flow rate can be 

allowed to fall off near the end of the transfer, or even that occasional 

interruptions can be tolerated if necessary. Unlike many propulsion 

applications that involve a high-speed turbine pump, the SFHT can even allow 

gas or a gas/liquid mix to appear at the inlet of the TM pump occasionally, 

because this type of pump recovers gracefully without damage. 

F8804112AAE-15 2-15 
F88-04 

b 

ENVIRONMENTAL 

CONDITIONS 

SUPPLY SFHE 

TO PUMP 

b 

TRANSFER 

CONDITIONS 

. 

FLOW ALLOWABLE TEMPERATURE 

RATE RESIDUAL RANGE 

L, - 

NN 902YMD135.041 

DIRECTION) SELF-FILLING 

DURING OPERATION SYSTEM 

Figure 2-8 Liquid acquisition requirements 

F8804112AA8-16 2-16 

1000 LIHR 

(AVERAGE 

DURING TRANSFER 

-40% AT 1000 1.4 2.1K 

LIHR 
F88-04 

The LAD must drain most of the tank contents to allow replenishing the user 

system with minimum SFHT launch mass and/or maximum flexibility in servicing 

other users later. This requirement should be stated in terms of the residual allowed when transferring at the full 1,OOO t/hr, plus a smaller allowed 

residual when "topping off" the user at the lowest practical transfer rate. 

Based on the SHOOT baseline, we are assuming 10% allowable residual at the 

full transfer rate, and 3% at the end of the reduced-rate topoff. 

The SFHT will operate over a range of temperatures during the user cooldown 

and fill operations, and the LAD must function over this range. The lowest 

supply tank temperature anticipated is 1.4 K, and the downstream side of the 

TM pump (and therefore the supply tank) must be kept below the lambda point 

(2.18 K) to allow the pump to function. The overall SFHT system should be 

optimized to keep the operating temperature as low as possible to minimize 

the mass loss associated with the TM pump operation. Therefore, the LAD 

should be designed to operate from 1.4 K to 2.1 K, with any tradeoffs made to 

optimize it for the low end of this range. 

During the transfer operation, our analysis of measurements made on the 

orbiter show that we should anticipate random gccelerations of up to g, 

with impulses up to 0.1 g during docking, etc. When operating on the SS and 

the OW, we anticipate the random accelegations to be reduced. The appropriate design requirement is therefore 10- g. Occasional impulses above this 

level can be tolerated without design impact because both the TM pump and the 

selected design approach for the LAD will recover gracefully from any momentary loss of liquid flow. 

A conventional screen-channel LAD would face the additional requirement that 

it retain liquid during the launch acceleration because such systems do not 

recover automatically if vapor is ingested into the LAD. This is frequently 

a driving requirement for the LAD design. Fortunately, the unique behavior 

of SfHe causes the selected LAD to refill itself automatically when used. 

2.1.4.2 Design Options 

Figure 2-9 shows all of the design approaches considered for the LAD, and 

classifies them in a way which simplifies the selection of the baseline. 

Positive displacement approaches (bellows, diaphragms, pistons, etc.) can be 

rejected immediately as impractical at cryogenic temperatures or as having 

serious weight penalties. The fundamental choice therefore becomes one between LAD approaches which supply liquid to the pump at a pressure either 

lower or higher than that of the main bath. 

Classical LAD schemes which depend on the surface tension of the liquid to 

retain it in the galleries all cause a finite liquid pressure drop. Openchannel systems (e.g., vane structures) cause very little pressure drop, but 

analysis has shown them to be capable of only withstanding accelerations in 

the micro-g range, and therefore unusable by SFHT. Fundamentally this is 

because the surface tension of SfHe is more than an order of magnitude lower 

than that of other cryogens. 

Analysis of a screen-channel LAD operating by itself in SfHe7, even with 

corrections for the properties of SfHe, predicts a pressure at the pump inlet 

7 

F8804112AA8-17 2-17 
F88-04 

PROWOE LlQUlO 

TO PUMP 

II I i i_l FOUNTAM a ACCELERATION 

Figure 2-9 Liquid acquisition design options 

F8804112AA8-18 2-18 
F88-04 

below the saturation pressure of the liquid. Viscous pressure drop for the 

normal component of the SfHe (see below) traveling downstream through the 

impedance of the screens is significant. When the TM pump is considered 

jointly with the LAD, however, we find that the normal component travels 

upstream away from the pump, and the viscous drag causes the impedance of the 

screen to raise the pressure of the fluid inside the LAD slightly. This will 

be discussed below. 

In conventional fluid systems that use screen-channel LADS, the fluid is 

forced through the LAD by a pressurizing gas. As long as the fluid is kept 

above saturation pressure everywhere inside the LAD channels, it cannot form 

bubbles (cavitate). It is possible to pressurize cryogens other than SfHe 

because either: 

0 The pressurant gas is helium, which does not condense at the temperature of the cryogen, or 

0 It is pressurized with its own vapor, but the low thermal conductivity of the liquid limits the rate at which the vapor condenses 

Pressurizing a mixture of SfHe and gaseous helium above the saturation line 

is impossible because the effectively infinite thermal conductivity of large 

bodies of SfHe causes the gas to condense almost instantly, returning the 

system to the saturation line. Any pressure drop in the LAD therefore brings 

the liquid in the channels below its vapor pressure, and raises the possibility of cavitation. 

In general, liquids will withstand pressures below their vapor pressure for a. 

short time without forming bubbles because of the time required for random 

fluctuations of thermal energy .to allow creation of a bubble.larger than a 

certain critical radius. Testing has shown that SfHe can withstand a pressure head of about 8 cm below saturation, but that this sustainable pressure 

falls off at high flow rates. Other work has shown that this sustainable 

head can fall to near zero under certain high flowrate conditions. The 

extent to which, or whether, the LAD or SFHT can be allowed to count on operating correctly below the SfHe saturation line is uncertain at this point. 

8 

2.1.4.3 Recommended Design Approach 

We recommend a LAD design based on the TM principle. Whether or not this 

system should be designed to maintain the SfHe pressure above Saturation 

under adverse accelerations is a tradeoff we discuss below. The overall LAD 

concept (Figures 2-5 and 2-10) looks very much like a conventional screenchannel LAD, but the flow impedances (i.e., the screen pore size and open 

area) are chosen specifically to function with the TM pump and SfHe. 

To understand how the SFHT LAD behaves, it is necessary to understand the 

physical principles governing the behavior of SfHe. The well-established 

two-fluid model describes SfHe as consisting of two components, a normal 

component that has viscosity and entropy, and a superfluid component that has 

neither. The overall SfHe density p is given by 

P = P, + P,, 

F8804112AA8-19 2-19 
F88-04 

GRLLERY 

(1 OF 4) 

CHRNNEL POROUS MRTERIFIL 

CRYOGEN TQNK 

GRIND-DOWN WEL 

FILL LINE 

* FLOW 

HERTER 

DETFlIL R NN 9203 .OS1 

D 

F8804112AA8-20 

Figure 2-10 Thermomechanical pump concept 

2-20 
F88-04 

where p, and p, are the densities of the two components. 

Just below the lambda point (the temperature near 2.18 K where the superfluid 

transition occurs), the SfHe consists almost entirely of the normal component, as shown in Figure 2-11. As the temperature is reduced to the range 

around 1.6 K where SFHT will operate, the liquid consists mostly of the superfluid component. 

To predict the behavior of the LAD requires analyzing the velocities Vn and 

VB of the normal and superfluid components separately. Table 2-4 shows the 

four equations that govern the velocities of the two components. Here i is 

the mass flow, A is the cro'ss-section area of the flow channel, x is the 

distance down the flow path, Q is the external heat input, and T, P, and S 

are the temperature, pressure, and entropy respectively. The meaning of the 

mass conservation equation is obvious, and the energy conservation equation 

is based on the fact that the energy carried by the liquid is given by its 

enthalpy plus the internal convection described below. The Gorter-Mellink 

equation is a well-accepted empirical relationship between the temperature 

and pressure gradients and the mutual friction between the two fluid 

components, using experimental values for the Gorter-Mellink constant, a. 

The pressure drop is tied only to the normal fluid component. 

The fifth equation describes the heat flux q, which is the energy transport 

within the fluid due to the difference between V This internal 

convection has no counterpart in normal fluids an% is the phenomenon that 

gives SfHe its anomalously high thermal conductivity. 

and VB. 

Figure 2-12. shows why it is that the LAD in the SFHT tends to pump the SfHe. 

The porous plug of an ideal TM pump would have pores so small that it would 

allow essentially none of the normal component to be forced back through by 

the high downstream pressure. In real materials, however, a finite Vn flows 

upstream through the LAD channel and screen. The flow impedance of the 

screen (or other porous material covering over the opening into the channel) 

therefore causes the pressure to be higher inside the LAD than in the main 

bath. In other words, the screen of the LAD tends to act as a TM pump, driven by the backflow of the normal component from the main TM pump. 

Figure 2-13 shows the design criteria for a conventional screen-channel LAD. 

The LAD must be able to draw fluid at the rate m from the screen located 

farthest away from.the pump inlet without either causing cavitation or sucking bubbles in through the screen closest to the pump. The pressure PLoc, 

at the pump inlet is 

- 'LOCAL - PBATH - PSCREEN - PcE~NEL - ~ACCELER*TION - PBE,N~uLLI 

where P EN and PCHM EL are the viscous pressure "lossn across the impedance of" Re screen ant the channel, PAcCELEB TI , is the "static headn that 

would be caused by an adverse acceleration of t\e vehicle, and PB,,,o,LI is 

due to the Bernoulli effect. Note that with normal fluids the viscous pressure drop term causes the pressure inside the screen to be lower than in the 

bath, but the fact that V, is flowing upstream causes these terms to produce 

a pressure increase. 

F8804112AA8-21 2-21 
F88-04 

1.0 

0.8 

a 0.6 

a 

v) 0.4 

U 

I- 

< 

E 

U 

B 

0.2 

0 

F8804112AA8-22 

COMPO NEW 

NORMAL 

0 0.5 1.0 1.5 2.0 

TEMPERATURE (K) 

A/N 5708 

Figure 2-11 Two components of He I1 

2-22 
F88-04 

Table 2-4 

EQUATIONS GOVERNING SUPERFLUID 

Mass Conservation: 

Energy Conservation: 

d - (mh + q) dx dx 

Gorter-Mellink Equation: 

Viscous Pressure Drop: 

dx .( D2 

0.316~V_~ 

2 D Re"4 

( 1 ami nar ) 

(turbulent) 

F8804112AAB-23 2-23 
F88-04 

L!alN! 

=P 

BATH SAT P 

P 

LOCAL i--f--i 

- "s BACKFLOW THROUGH 

PUMP 

- VISCOUS PRESSURE DROP OF 'LOCAL = BATH 

PASSING THROUGH PORES 

F8804112AA8-24 

Figure 2-12 Why LAD pumps 

2-24 

s + P""" 1 

AM 9203ND135.038 
F88-04 

ACCELERATION = 10 - G (111) 

=P 

BATH SAT 

P 

BUBBLE 

BREAKTHROUGH 

AVOID CAVITATION: 

CAV -P SAT 

>P 

(GOAL) SAT 

’P 

AVOID BUBBLE BREAKTHROUGH: 

0 

m, 

F8804112AA8-25 

Figure 2-13 LAD design criteria 

2-25 

AJN 9203lMD135.039 
F88-04 

The other principal design requirement for conventional screen-channel LADs, 

namely that launch accelerations will not cause bubble breakthrough, does not 

apply to this system using SfHe. As soon as the main TM pump is started and 

Vn starts to flow back from the pump toward the main bath, the pressure rise 

due to the TM pumping action will cause any bubbles within the LAD to collapse. This makes the LAD self-filling, and allows it to recover automatically if an anomalous acceleration does cause momentary cavitation. 

Table 2-5 compares the performance predicted for a SFHT LAD using conventional 325x2300 wire/inch Dutch twill stainless steel screen with one using a 

pogous membrane instead. Four channels were assumed, using a flow area of 80 

cm for both the channel and the screen (or porous membrane). The bubble 

points (the pressure difference required to cause bubble to break through the 

holes against the surface tension of the liquid) are high enough that bubble 

ingestion is clearly not a problem with either the screen or the membrane. 

The principal issue is possible cavitation. 

Table 2-5 

EXPECTED LAD PERFORMANCE 

PRESSURE CHANGE FROM BATH SCREEN MEMBRANE 

lEXPRESSED IN CM OF HELIUM HEAD1 J3-10 rrm) (0.2 rrm) 

Viscous Drop Across Screen +8. OxlO-' +2.1 

Viscous Drop Across Channel +6. Oxlo-' +6. Oxlo-' 

Acceleration (At lo-' g) -0.3 -0.3 

Bernoulli Pressure Drop -3.9x10-' -3. 9x10-a 

Total -0.3 +1.8 

Bubble Point of Screen/Yembrane 30 N/A 

With the screen, the pressure rise due to the TM pumping action is tiny compared to the pressure reduction from an adverse acceleration of lo-' g. With 

the finer pores of the membrane, however, it is easy to achieve a pressure 

rise that will keep the liquid securely above the saturation line under the 

worst-case design acceleration. 

The metallic screen used in this example is a well-developed technology used 

in conventional LADs. The porous membrane is a filter material used routinely for fairly high (e.g., 55 psi) pressure differentials, and which has been 

used in SfHe research. The standard screen technology presents essentially 

no hardware development risk, but makes the performance of the SFHT dependent 

on the partially-understood capability of SfHe to withstand pressures below 

saturation without cavitation. The porous membrane (or other porous material 

with pores smaller than about 1 micron) could be used to make SFHT immune to 

accelerations during transfer, but would require selection of the material 

and development of suitable mounting techniques. 

F8884112AA8-26 2-26 
F88-04 

2.1.4.4 Recommendations for Development 

We recommend that a LAD system like that of Figure 2-5 be baselined for SFHT, 

with a decision to be made between two possible approaches to the covering 

for the channel openings: 

Option 1: Metallic Screen 

The well-established manufacturing technology minimizes the hardware development effort, but the tiny TM pressure rise possible means that random accelerations during the transfer will cause the pressure in the channels to drop 

below the saturation pressure intermittently. Further understanding of the 

ability of SfHe to withstand these pressures may or may not show that cavitation is unlikely. If bubbles do form, they will collapse when the adverse 

acceleration is removed, and do so at a rate that can be calculated. When 

analyzing the acceptability of this approach, it is important to recognize 

that occasional cavitation and interruption of the mass flow may be acceptable because the system recovers gracefully. This is the approach currently 

baselined for SHOOT. 

Option 2: Fine-Pore Material 

Using either thin membranes or conventional porous plug materials allows the 

liquid to be kept above saturation under any accelerations likely to occur 

during transfer, but does require some hardware development effort. Previous 

experience with the plastic membranes at SfHe temperatures and separately in 

high-pressure filtering applications makes this effort seem quite feasible. 

The additional pressure rise within the LAD may have additional system benefits as well. 

Whichever approach is chosen, the LAD must be analyzed and developed jointly 

with the TM pump and transfer system. In the present example, the pressure 

and temperature rise within the LAD are so small that the pump and transfer 

system can be analyzed neglecting them. In an optimum system, however, it 

may turn out that there are advantages to sharing the pressure rise more 

equally between the LAD and the main TM pump. In the limit, the main TM pump 

plug could be eliminated entirely, relying on the LAD to provide the driving 

pressure and to contain the full pressure rise and temperature rise without 

structural failure or excessive heat conduction to the main bath. 

2.1.5 Thermal Subsystem 

The thermal subsystem consists of two parts: external (outside the vacuum 

shell) and internal (inside the vacuum shell). The external thermal system 

controls the vacuum shell temperature and maintains externally mounted equipment within required temperature limits. The internal system limits the heat 

input to the stored helium. 

Internal Thermal Control (Figure 2-14) has five major parts: 

1. Because of the low latent heat and high sensible heat of helium, 

it is essential for long-term liquid helium storage to use a 

vapor-cooling system that intercepts parasitic heat inputs. For 

the lifetime, size, and boundary conditions of the SFHT, two 

F8804112AA8-27 2-27 
F88-04 

ORIGINAL PAGE IS 

OF POOR QUALITY 

UECHRN ICRL COOLERS 

(8.4 Y) 

VFlCUUU SHELL (300 K) r 

\ TOUTER vcs 

CRYOGEN TFlNK (10.000 L 

7 FlLUUINR/EPOXY 

SUPPORT STRRPS 

I 

I t 

I 1 

I Y. I 

\\\\ 

F8804112AA8-28 

L49 LRYERS YLI 

Figure 2-14 Insulation system design 

2-28 

ITERS) 

A/N 9203 061 
F88-04 

vapor-cooled shields give a mass-optimized system. 

and conductive heat inputs are intercepted at each VCS. 

All radiative 

2. Radiation input is blocked by high-quality MLI blankets similar to 

those of the COBE dewar. The three blankets contain 20, 30, and 

40 layers (from cold to warm) of double-aluminized mylar with 

dacron net spacers. 

3. Fluid loss during the extended on-orbit standby is reduced by a 

factor of two by four Stirling cycle mechanical coolers which take 

heat from the outer VCS. The cost and mass trades related to 

these coolers are discussed in Section 2.1.1. Four British Aerospace Stirling cycle coolers with 3 W cooling power at 93 K each 

are baselined; efficiency is about 25 W/W. The cooler consists of 

a compressor and a displacer, where the cooling is produced. 

Figure 2-15 shows the interface between the four displacers and 

the dewar. It provides high thermal conductance by copper conduction, minimal loading on the fragile displacer cold tip by using 

copper braid, separation between the dewar guard vacuum and the 

displacers vacuum, and semipassive thermal switching to disconnect 

a given displacer from the cold finger in case a cooler fails. 

Cooling by the cold finger wi*ll produce the thermal contraction 

needed to make the switch close, but a heater is needed to force 

it open since multiple coolers are thermally manifolded. 

4. The cryogen tank is supported from the vacuum shell by a thermally-optimized, tension strap support system using aluminalepoxy 

straps. The thermal efficiency andlJ:fi risk of this approach make 

it the obvious choice for the SFHT. 

5. Gaseous heat conduction into the helium tank is essentially eliminated by a high-quality guard vacuum. The vacuum shell maintains air leakage at no more than lo-' sccs, and helium leakage is 

held below lo-* sccs by careful assembly of the tank and plumbing 

and in-process testing. 

The external thermal control system maintains heat dissipating components 

within required temperature ranges and shields sensitive items, such as the 

avionics, from solar illumination. In many long-life dewar applications, 

carefully designed combinations of high-emittance coatings on space-viewing 

surfaces and MLI blankets on sun-viewing surfaces are used to minimize the 

vacuum shell temperature and standby boil-off rate. For example, the COBE 

dewar vacuum shell is predicted to be at 150 K. In the case of the SFHT, 

periods of extended standby will be in a protected area, probably in the SS 

customer servicing facility. This enclosure will be maintained at a nominal 

temperature of 300 K.12 Consequently, our baseline design and performance 

estimates assume a vacuum shell temperature of 300 K. 

Heat producing components on the vacuum shell (i.e., electronics and mechanical coolers) will be cooled by attachment to the vacuum shell, by self radiation, and/or by attachment to the host heat rejection system. BF&hlJhe shuttle orbiter and the SS have adequate heat rejection capacity for the 

SFHT maximum total power level of about 400 W. The vacuum shell will have a 

high-emittancellow absorptance finish to provide radiative cooling, even when 

exposed to solar radiation. 

F8804112AA8-29 2-29 
F88-04 

I 22.7 I 

A/N 9203.036 

Figure 2-15 Concept for dewar/mechanical cooler interface 

. 

F8804112AA8-30 2-30 
F88-04 

2.1.6 Dewar Instrumentation and Cabling 

This section discusses instrumentation and cabling that are part of the dewar. The instrumentation is used to monitor the condition or health of the 

dewar and to aid control of the fluid operations. Instrumentation, primarily 

thermometers, used to monitor avionics components and other equipment external of the dewar is also required. 

Table 2-6 is an instrumentation list, showing wire counts, wire routing, and 

when the sensors are used (i.e.,. ground or flight). Figure 2-16 shows approximate locations of the various sensors, which are described further below. Related technology development requirements are discussed in Section 6. 

Thermometry 

Germanium resistance thermometers are used for the range of 1.5 - 10 K, and 

silicon diodes are used for the range of 10 - 350 K. The high sensitivity of 

the GRT is needed for the measurements around and in the helium tank. The 

large range of the SD makes it the choice for all of the warmer measurements. 

The thermometers are located in redundant pairs. GRTs are located in the 

helium bath, at two'locations within the liquid acquisition device, on the 

inner and outer surfaces of the TM pump, on the vent line just downstream of 

the low-flow porous plug, and on the outer surfaces of the two porous plug 

phase separators. SDs are located on the inner VCS,.on the outer VCS, on 

each of the four thermal switches connecting the mechanical coolers to the 

dewar, at three places on the vacuum shell, and on both of the EVA couplings. 

Pressure Transducers 

Two types of pressure transducers are used: high range (0 - 50 psi) and low 

range (0 - 2 psi). The high-range transducer primarily monitors tank pressure to check for anomalous over-pressure conditions; the Teledyne Taber unit 

being used by SHOOT is adequate. The low-range transducer is located just 

downstream of the TM pump and operates at the pump temperature (-2 K). It is 

part of the transfer flowrate measuring technique described below. The Validyne model APlO being used in the SHOOT venturi flowmeter is the choice here. 

F 1 owme t er s 

Two mass flowrate measurements are made: the stagdby boil-off rate (10 mg/s 

maximum) and the superfluid transfer rate (1.5~10 g/s maximum). The standby 

flowrate is determined by simply measuring the thermal capacitance (icp) of 

the flow, as shown in Figure 2-17, and knowing the specific heat of the gas. 

This approach works particularly well at the low temperatures and flowrates 

involved because a GRT can be used to make very sensitive measurement of 

temperature, there are no radiation effects, material thermal capacities are 

so low that their effect is negligible, and very little heat input (-50 mW) 

is needed to produce an accurately measurable temperature rise (-1 K). This 

flowrate is measured because it is one of the indicators of dewar "thermal 

healthn and to help determine if helium is leaking from the system via an 

unknown path. The latter is done by comparing the integrated flow measurements with the mass loss indicated by the mass gage described below. 

F8804112AA8-31 2-31 
F88-04 

ORIGINAL PjGE $5 

OF FObR QUALITY 

Table 2-6 

DEWAR SENSORS AND CABLING 

mrr 

LIQUID LEVEL, NORYAL (LL) 

SUPERFLUID YASS GAGE HEATER (LY) 

GERYANIUY RESISTANCE THERYOUETER (GT) 

SILICON DIODE THERYOYETER (SO) 

PRESSURE (P), 0.50 pa 

STRAIN GAGE (SO) 

ACCELEROUETER (A) 

VACUUU (V) 

VALVE POSITION 

VALVE POWER 

TU PUYP HEATER 

SfANDEY FLOWYETER HEATER (FY) 

SME TRANSFER FLOWYETER 

POSITION SENSOR, EVA COUPLING (PS) 

0.2 pJ 

OUANTITY 

2 

2 

12” 

20 

2 

2 

12 

2 

1 

26 

26 

2 

2 

2 

NO. Of WIRES 

8 (UO)-- 

4 (UO) 

10 11(28) 

48 (UO) 

40 (UO) 

10 

6 (Mol 

48 

12 (140) 

4 (HIGH VOLTAGE) 

130 (840) 

116 (128) 

6 (128) 

. 4 (128) 

NA * FLOWRATE BAS1 

4 

~ ~ ~ ~ 

VEM 

VACUUU 

VACUUU AND VENT 

VACUUU AND EXTERNAL 

EXTERNAL 

VACUUY 

VACUUY 

VACUUY 

EXTERNAL 

VACUUU 

VACUUY 

VENT 

VACUUY 

I OH G73, GTS AND P3 

EXTERNAL 

USE 

GSE 

BOTH 

BOTH 

BOTH 

BOTH 

BOTH 

GSE 

GSE 

GSE 

BOTH 

BOTH 

BOTH 

BOTH 

FLIGHT 

A/N 92031MD135.019 

F8804112AA8-32 2-32 
F88-04 

ORIG!NAL PAGE IS 

OFI POOR QUALITY 

ItrGT 

FM 

LL 

MG 

P 

PS 

SD 

SG 

GERMANIUM THERMOME 

FLOW METER 

LIQUID LEVEL 

PRESSURE 

POSITION SWITCH 

SILICON DIODE 

MASS GAGE, Liauio 

TER 

STRAIN GAGE,-STRAP I 

I 

or1 

air 

YO1 

uaa 

LL1 0 

LL2 

Figure 2-16 Locations of dewar instrumentation 

so5 - SO10 

V 

. 

F8804112AA8-33 2-33 
ORIGltJAL PAGE IS 

OF POOR QUALITY 

T1 

p fil 

HEATER, 6 

I t 4 

T3 d-+ 

A I I GRT 

L-1ocm 

WHERE: A s FLOWRATE 

b I HEATER POWER 

h = ENTHALPY OF GAS 

F88-04 

3 

I I 

NN 9203/MD135.018 

Figure 2-17 Flowmeter concept 

F8804112AA8-34 2-34 
F88-04 

The on-orbit transfer rate can be determined to about 1 percent accuracy by 

measuring the temperatures upstream and downstream of the pump and pressure 

downstream of the pump. These values will be entered into the performance 

model of the pump, which will have been validated by ground characterization 

of the pump by test, and the flowrate calculated. 

Strain Gages 

Strain gages are used to measure tension in each of the 12 support straps, as 

was done in the IRAS and COBE dewars. These are used to pretension the 

straps during assembly, monitor stress levels during system testing, and 

ensure that proper pretension is maintained throughout the life of the system. It would also be desirable to measure the stress levels during one 

launch to verify the fatigue loading history assumed for sizing the straps. 

However, this capability is - not included in our avionics design since it is a 

one-time gathering of engineering information., 

Accelerometers 

Redundant triaxial accelerometers are located on the cryogen tank for system 

testing. The vacuum shell will be instrumented with laboratory accelerometers during testing. 

Position Sensors 

Each motor-operated valve has redundant position sensors to indicate "pen" 

or "closen. The EVA couplings also have position sensors for.electrica1 

verification of complete mating. 

Guard Vacuum 

A Vac-ion gage is used to monitor the guard vacuum level while on the ground, 

as done with the IRAS and COBE dewars. This is needed to verify the proper 

age when the dewar contains liquid helium. The gage is not used during 

flight. 

Mass Gaging 

Liquid level is measured on the ground with superconducting wire sensors, as 

was done in the IRAS and COBE dewars. The sensors are oriented orthogonally 

so that the liquid level can be measured with the dewar in two orientations. 

If this is not required, they will be parallel and redundant. 

Liquid mass will be measured on-orbit by applying a heat pulse and measuring 

the resultant temperature rise, as will be done in the SHOOT dewar. This 

approach can also be used on the ground when the bath.is superfluid. Redundant gages are used. Each gage consists of a simple wire-wound heater attached to the tank wall, a GRT, and the external electronics. The GRTs are 

interchangeable and therefore redundant. The design developed for SHOOT will 

be directly extended to the SFHT. The main difference is that the greater 

fluid mass requires a larger heat pulse. Based on the 0.02 K temperature 

rise recommended by the SHOOT program and a 10 minute pulse duration, the 

power needed is 100 W. Higher power and smaller pulse time could be used at 

F8804112AA8-35 2-35 
F88-04 

the expense of dewar thermal performance. 

will be based on the results of the SHOOT flight experiment. 

Detailed operational parameters 

2.2 FLIGHT AVIONIC SUBSYSTEM 

This section presents the conceptual design for the SFHT avionics. Since 

detail design was not in the scope of this study, specifics are minimized. 

Several areas are addressed which define the scope of avionic needs and requirements for the SFHT. 

The purpose of the SFHT avionics is to provide all control, monitor, and 

display functions necessary to transport cryogen and perform resupply operations in a flexible and efficient manner. As directed by JSC, our concept 

starts with the design from the Martin Marietta OSCRS study with modifications for improved operator interface, reduced power and parts count, increased reliability, and SFHT-specific requirements. We also expect the SFHT 

avionics to utilize both hardware and knowledge from the upcoming SHOOT 

experiment. 

Our baseline concept is primarily intended for resupply at the shuttle with 

additional considerations for SS and OW operation addressed. The SFHT will 

be used to resupply a variety of users, such as Astromag, SIRTF, AXAF, and 

XRS. We present the basic interface requirements between SFHT and the user 

to assist the start of planning for compatibility between them. 

Our avionics concept is a stand-alone system that uses the orbiter GPC for 

safety-related functions only, as discussed in 3.3. Growth provisions have 

been included to allow for future upgraded SFHT systems extending into the 

2010 time frame, as well as operation with other cryogenic systems. The SFHT 

avionics concept is designed to function with minimum, but critical, human 

oversight and control leading to growth into autonomous operation in the 

future. Closed-loop control is employed with an intelligent expert system 

combined with both sensor and actuator feedback to close the loop. 

2.2.1 System Architecture 

The SFHT avionics is broken into two functional blocks: the system control 

station (SCS) and the dewar monitor and control, in two distinct locations. 

Figure 2-18 illustrates the two functional blocks with their respective internal functions and 110 needs. 

The SCS avionics consists of those components required to make a man-machine 

interface (MMI) with the SFHT and to provide a communication link between the 

SFHT and its host. The SCS is located in an operator accessible area in the 

host, e.g., the AFD of the shuttle. The MMI elements consist of three electroluminescent color touch screen displays providing operator command input 

and data display feedback, plus a switch panel for hardwired manual inputs to 

the system. The SCS functions of receiving commands, processing and displaying data, and communication are orchestrated by three computers operating in 

a three string majority vote configuration. Electrical 1/0 to the SCS consists of power from the host, two-way communication with the host, a realtime clock (GMT) from the host, two-way communication with the dewar-mounted 

avionics, and hardwire controls for dewar operation and reliability. 

F8804112AA8-36 2-36 
SYSTEM 

CONTROL 

STATION 

AVlONlCS d DEWAR 

DEWAR SENSORS 

AND DEVICES 

DEWAR MONITOR 

AND CONTROL 

UPLINK SENSOR 

SENSOR 

COMMANDS COMMANDS 8 EXCITATION FLOW METERS 

MASS GAUGING DOWNLINK 

' /* COMPUTER * TEMPERATURE SENSORS 

MAJORITY VOTING 

4: . VALVE DANE K PRESSURE 

DATA COMPUTERS 

TM PUMP DRIVE POSITION DISPLAYS DATA 6 

c 

6 

CRITICAL 

WARNING 5 

/ 

REFRIGERATOR 

SENSOR ELECTRONICS VALVES 

TM PUMP 

TOUCH PANELS POWER CONDITIONING MANUAL SWITCHS 

MANUAL 

POWER 

AFDBASED 

AVIONICS 

DEWAR-BASED 

AVIONICS 

F8804112AA8-37 

Figure 2-18 SFHT avionics has two segments 

2-37 

&t4 9WMOl4203a 
F88-04 

Dewar monitor and control electronics provide the first line of control over 

the cryogen systems sensors and actuators. The design philosophy has been to 

put most of the SFHT controls near the cryogen system because this approach 

will best allow evolution to automatic operation, which is required of SFHT. 

Following this line of thinking, our concept places the dewar monitor and 

control avionics on the dewar structure near the cryogenic controls and sensors. This major functional block contains the following subfunctions: 

a Power conditioning of itself and those elements receiving power 

from it 

a Three computers plus a majority voting unit to complete the three 

string, majority vote, computing system 

a Valve drive circuitry for the cryogenic control valves 

a A proportional drive control for the TM pump 

a All sensor excitation and readout circuits for monitor of cryogenic and system status 

0 An active temperature control system to protect is own environment 

The dewar monitor and control 1/0 consists of the following: 

a Two way communication between its computers and the SCS computers 

a Power from the host 

a .Hardwire (manual) controls from the SCS switch panel 

.a Dewar sensors excitation and readout 

a Valve actuation for control of cryogen flow 

0 TM pump drive power 

a Switched host power to dewar mechanical coolers 

a Power to the user 

a Two way communications with the user 

a Hardwired emergency controls for pyrotechnic firings, etc. 

Instrumentation, discussed in 2.1.6, consists of sensors associated with 

measurement of the status of the system and of the actuators that can alter 

the status and/or condition of the system. The sensors include temperature, 

pressure, flow, and position while actuators include valves, latches, heaters, relays, and refrigerators. All electrical inputs and outputs for the 

instrumentation come from the dewar monitor and control electronics. Most of 

the instrumentation is located either inside the dewar or on the dewar 

plumbing. 

F8804112AA8-38 2-38 
F88-04 

2.2.2 Demands on Host Vehicle Resources 

The SFHT will be used with several different hosts, and even though we are 

primarily baselining for shuttle, a definition of demands on the host must be 

made to assure broad host compatibility. Three areas should be addressed to 

define the electrical demands. These are power requirements, data rates and 

interconnects. 

2.2.2.1 Power 

The SFHT power requirements are summarized in Table 2-7. Three levels of 

power requirements are shown. The first is a warning only, low power monitor 

mode in which a single computer string operates with no touchscreen input or 

display capabilities. Only anomalous condition alerts are displayed to the 

operator at the SCS. The second is a 

full monitor mode in which a single computer string operates with one touchscreen display I/O. This monitor power of 83 W will most likely be the level 

used during resupply missions at the shuttle or during short standby periods. 

The third occurs during transfer and results in a power consumption of 315 W. 

This estimate includes an assumed user power of 46 W for operation of one 

valve, the user’s computer, and six thermometers; this user estimate is of 

course subject to change. Also shown in Table 2-7 is 375 W for operation of 

the mechanical coolers during the extended standby (storage) at the SS. The 

warning only monitor mode would also be used at this time for a total power 

requirement of 410 W. 

This lowest power mode requires 35 W. 

2.2.2.2 Interconnects 

Two interconnect areas are required. First, an SCS to host interface is 

required to provide an interaction path between the SFHT and its host. Second, the host must provide an interface path between the SCS and the dewar 

mounted avionics to complete SFHT’s control and data loop. 

The SCS-to-host interface, as shown in Figure 2-19, consists of three pairs 

of serial communications lines for downlink data, three pairs of serial communications lines for uplink commands, five wires for critical warning signals to the CWEA, and four power wires rated for at least 3.5 A each. 

Figure 2-20 shows the interconnect requirements between the SCS and the dewar 

mounted avionics. Command and data requires six pairs of serial communications lines per computer for a total of 18 pairs. Discrete wire control requires a total of 18 single wires: 4 for power on/off, 7 for pyro control, and 7 for manual valve safing. Additionally the dewar mounted avionics 

must receive power directly from the host over at least 4 wires capable of 

9.6 A each. User power passes through the dewar mounted avionics and requires 4 wires rated at 0.8 A each. 

2.2.2.3 Data Rates 

All SFHT data rates are relatively modest. We are assuming a maximum available STS data rate of 8 KBPS and SFHT’s uplink/downlink requirements readily 

fall within this constraint. Command and data rates between the SCS and 

F8804112AA8-39 2-39 
Table 2-7 

POWER BUDGET 

F88-04 

WARNING ONLY 

MONITOR MODE MONITOR MODE TRANSFER MODE 

scs 

COMPUTER 9w 

TOUCHSCREEN DISPLAY - 

SWITCH PANEL 

SCS TOTAL 

4 - 

13 

9W 

48 

4 - 

61 

DEWAR MOUNTED AVIONICS 

COMPUTER 16 16 

MAJORITY VOTE - - 

VALVE DRIVE AND PUMP - - 

TEMPERATURE CONTROL 6 

DEWAR TOTAL 

- 

22 22 

6 - 

27 W 

67 

4 - 

98 

48 

12 

105 

6 - 

171 

USER - - 46 

TOTAL 

- - 

35 W 83 W 

- 

315 W 

REFRIGERATOR (SPACE STATION ONLY)375 W 

ASSUMPTIONS: MAXIMUM 2 VALVES + PUMP OPERATE SIMULTANEOUSLY 

MONITOR MODES HAVE SINGLE STRING COMPUTER 

80% POWER CONVERSION EFFICIENCY 

SFHT PROVIDES USER POWER FOR VALVE OPERATION, COMPUTER, AND 

MONITOR 

F8804 11 2AA8-40 2-40 
F88-04 

8 kbps4 

r 

STS Functions 

STSlUpllnk Commands 6 

/ 

MDM 

F8804112AA8-41 

SCS Computers 

(Combinor) 

"i 

Prlmary 

Power 

Figure 2-19 SCS to STS interface 

2-41 

"i 

Auxiliary 

Power 

To Dewar 

% Command 

Data 

> Command 

Data 

> Command 

% Data 

NN 934UMD142.05 
p Irl 

- IIF 

A 

4 

B 

- 

4 

C 

- 

4 

uu 

Switch Panel 

F88-04 

Dlgital 5 5 Command 61 

I /‘ 

5 

5 

5 5 

5 

8 Data 61 

/ 

s 

2 Command 6 

-- e 6 

5 / 

5 5 

5 

s 

5 Command 6/ 

/ 

/ 

Data 61 

s 

s 

Discrotoa 

g 

5 3 

-APH 

4/5 Powor On 

‘5 

EnabloIArmIFiro 7 15 

15 

Valvo Safing 7 - 5 

/5 

AFD 5 Cargo ~ay 

L d s AFD Computers Sorial / 

Powor Distribution DC 

AC 

Dedicated Pins for 

AC or DC Hort Powor 

* 4/ Pyro Clrcuita (Guillotlne) 

Valvo Drive 

~sor powor 

ORIGINAL PAGE IS 

OF POOR QUALITY 

Dewar Computers 

i 

Computers AID 

k Data 

IIF 

- -A -- 

Control - -- 

I I DataITlming - . AID I 

DataITiming 

uu I 

AN 9344ND142.08 

Figure 2-20 Dewar/SCS electrical interface 

F8804112AA8-42 2-42 
F88-04 

SWITCH 

PANEL 

I TOUCHSCREEN 

MANUAL SAFING 

PYA0 CONTROL 

POWER ONIOFF 

POWER ................................................................................. 

b7 

)7 

b4 

L 

x v) 

n 

I 

TOUCHSCREEN I 

e--- 

-\ PRIMARY SELECT ............................................ 

I 

HOST TO SCS I10 d 

I a 

y1 

I3 

0 

0 

5" 

IL , 5 

a0 a! 

0 L 0 

CONDITIONED 

SCS POWER 

I 

BACKUP PRIMARY 

RAW POWER 

Figure 2-21 System control station . 

COMMANDS 

461 

COMMANDS 

-61 

DATA 

-61 

COMMANDS 

-61 

DATA 

-61 

AM 93WMD142.07 

F8804112AA8-44 

PRECEDING PAGE BLANK NOT FILMga) 

2-44 
F88-04 

operator, transfer data to and from the host, and provide only backup limit 

detection to the dewar computers. This limited role of the SCS computers is 

in keeping with a design philosophy which will lead to fully automated 

operation. 

Three microprocessor based computer systems can be housed in one assembly and 

operate as part of the three string computational approach used throughout 

the SFHT avionics system. Use of CMOS microprocessors from the 8086 family 

for these computers will allow the sharing of common software development 

tools with the dewar computers. Computer memory requirements are modest 

compared to 8086 family capabilities. 

Host-to-SCS interface consists of uplink and downlink data, real-time clock, 

and warning lines to the GPC. The downlink data passes information from any 

or all computers. An interleavor in the output interface will allow selection of the desired output format. Uplink and real-time clock inputs will be 

split for application to all three computers simultaneously. The GPC interface will be hardwired for each of five lines. 

2.2.4.2 Touchscreen Display/Input 

Three identical touch panel and display devices are located at the crew control station for the purpose of providing the MMI. The touchscreens enable 

the operator to quickly and readily communicate with the SFHT without the 

need for a keyboard interface. 

Only one of the touchscreens will accept inputs at any given time, with that 

one being the "primary" display. Selection of the primary touchscreen is 

made by switch selection on the SCS switch panel. All computers will receive 

command inputs from the primary touchscreen. 

Each computer communicates with its companion display independently of which 

touchscreen is primary. The primary display provides the operator with a 

continuous operating menu, analysis of the operation in process, and present 

state of the SFHT. The other displays can provide backup or supporting information for the operator upon command. 

2.2.4.3 Switch Panel 

The SFHT switch panel provides a means for the operator to manually control 

several aspects of the SFHT avionics without relying upon the computational 

strings. Functions provided include: 

e Avionics system power on/off control 

Emergency deberth pyrotechnic device control 

Low power "warning only" monitor enable 

e Primary touchscreen selection 

e Manual safing sequence control e 

e 

Figure 2-22 provides a conceptual layout of this uncluttered switch panel. 

One of the three touchscreen displays is colocated with the switch panel but 

has no functional relationship with it. 

F8804112AA8-46 2-45 
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F8804112AA8-46 

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The switch panel has three sets of LED indicators to provide a display of 

first line health and status. These indicators display the source of power 

(primary or auxiliary) and which touchscreen and computer strings have power 

applied. 

The power control switch actuates power relays in the SCS and dewar based 

avionics. This one switch controls power to all SFHT avionics. Primary 

touchscreen selection is provided by a three position switch near the center 

of the control area. A third switch will put the SFHT into its low power 

"warning only" mode. A manual switch is used for this function because all 

touchscreen inputs are disabled during this mode necessitating a manual recovery capability. All toggle switches must have safety flip covers to prevent inadvertent activation. 

Emergency deberthing involves firing of pyrotechnic devices to sever attachment to the user. A set of pushbutton switches on the switch panel must be 

depressed in a unique three step sequence to initiate emergency deberthing. 

A manual safing capability, which closes all SFHT dewar external valves, is 

provided through the use of a rotary selection switch and two push button 

activation switches. This is used as a backup capability to place the dewar 

in a benign mode should two of the three computer chains fail. 

2.2.5 Dewar Monitor and Control Avionics 

The dewar-based avionics includes the following components: dewar computer 

unit (DCU) (three computers in one housing), majority voting unit (MVU), 

valve drive unit (VDU), power distribution unit (PDU), thermal control system 

(TCS), associated interconnect cabling, and emergency deberthing circuits and 

devices. Interconnections and functions of these components are shown in 

Figure 2-23. 

Dewar avionics provides the first line of control over the cryogen system 

sensors and actuators. The sensor processing circuits are located in the 

DCU, and actuator drives are located in the VDU. The design philosophy has 

been to put most of the SFBT controls near the cryogen system because this 

approach will best lead to the autonomous operation that is required of 

SFHT's growth. 

2.2.5.1 Dewar Computer Unit 

The DCU is the heart of the SFHT avionics. Functions within the DCU include 

the basic expert system control and processing computer, sensor excitation 

and readout circuits, and 1/0 interfaces for communication both with the SCS 

and the user. 

Computer - The dewar computer consists of three identical parallel computers 

continuing the three string majority vote computing chain. Each of these 

computers will be based on a microprocessor, preferably from the 8086 family. 

A prime candidate is the 80386 because of its compatibility with the 8086 and 

its growth capabilities for an expanded expert system knowledge base. 

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Figure 2-23 Dewar monitor and control 

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The expert system hardware is configured to support the evolution of the 

system. The kernel of 

the expert system can be contained in W erasable EPROM. Updatable inference 

rules and knowledge base can be stored in an easily programmable, non-volatile memory, such as electrically erasable programmable read only memory 

(EEPROM). Working memory for temporary data, calculations, scratch pad, 

stack, etc. can be handled by a combination of dynamic and static RAM. 

Memory requirements are modest by todays standards. 

Sensor Avionics - Dewar monitoring requires the following sensor types (see 

Table 2-6): 

0 Germanium resistance thermometers (GRT) 0 Silicon diode thermometers (SD) 0 Pressure gage 

0 Flowmeter 

0 Liquid mass gaging 

0 Valve position 

Each of these sensors requires excitation source and readout processing circuits. The result of the basic readout processing is a voltage proportional 

to the measured parameter that can be converted into a digital form for processing by the dewar computer. Analog to digital (A/D) conversion is performed by two A/D converters per computer. A high resolution 16 bit converter is used for sensors requiring high precision, and an 8 bit converter 

is used for less critical sensors. An analog multiplexer, under control of 

the computer, directs the sensor circuits output to the appropriate converter 

at the appropriate time. The redundant and interrelated sensors are distributed amongst the three computer units to 'maintain a high degree of failure tolerance. A summary of the operation, circuit requirements, and status 

of each sensor type follows. 

A GRT is a variable resistor whose resistance depends on temperature. Its 

resistance is measured by passing a known current through it and measuring 

the voltage developed across it. The SHOOT experiment has developed an AC 

excitation and readout circuit that meets SFHT's needs and requirements; the 

SFHT will use the SHOOT GRT circuit. 

The SD relies on the forward voltage drop of a silicon diode which is related 

to both temperature and excitation current. Unlike a resistance thermometer 

the voltage drop across the SD is not linear with current, and accuracy and 

tracking are only guaranteed at fixed currents. SDs suitable for SFHT will 

require a dc current source of 10 mA t0.05 percent for excitation. Current 

sources meeting this value and tolerance are relatively straightforward, and 

applicable circuits exist from other programs. Readout circuits are 

straightforward differential operational amplifier designs suitable for four 

wire resistance measurements and also exist. 

The pressure transducers are to be determined, but the most likely candidates are the widely used Teledyne Taber Model 2215 and Validyne Model AP10. 

The Teledyne Taber is flight qualified, and the Validyne will be flown on the 

SHOOT experiment. Since both of these transducers are familiar, the electronics for their operation should not pose a problem. 

F8804112AA8-49 2-49 
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The SFHT’s standby flowmeter uses a heat input/temperature rise technique. 

The heat input is generated by applying a known electrical power to a resistive heater. The temperature measurements will be made by GRTs. Actual flowrate will be calculated by the computers from the power level, temperature 

data, and catalogued fluid enthalpy. 

Liquid mass gaging will also employ a heat pulse/temperature rise technique. 

The SHOOT program is developing this technique, and we expect the SFHT to be 

able to use the mass gaging technology developed on SHOOT. The SFHT will 

require scaling of the SHOOT mass gaging system to accommodate the increased 

size of the SFHT. Recent SHOOT data indicates that their system may suffer 

from excessive uncertainty in the energy applied due to transient response 

characteristics of the power source. An understanding and solution to this 

problem is expected to exist before the SFHT design is completed. 

The SFHT can use the same valves as SHOOT; therefore we expect the valve 

position sensing circuits to be applicable as well. 

I& -The dewar mounted computers are the heart of the SFHT avionics system 

and must communicate with the operator and host via the SCS and with the 

user. Both of these communication channels carry relatively low data rates 

(1 KBPS) and are best handled by serial interfaces. Numerous serial data 

standards are available, but we feel the best choice for the SFHT is RS422. 

This choice is based on the following factors: 

More than sufficient data rate capability 

Not dependent on close microprocessor support 

Simple in terms of handshaking and protocol 

Architecture is widely used and understood 

Tolerance to EM1 0 

The dewar computer-to-SCS computer interface is made up of three full duplex 

serial communication lines. Each of these lines forms a dedicated interconnect between computers at each end and is part of the total three string 

majority vote computer system. Information that is carried over this 1/0 

path consists of mode initiation commands and limit changes from the SCS to 

dewar computers in one direction and both raw and processed data from the 

dewar to the SCS computers in the other direction. 

Dewar computers each communicate with the berthed user over a serial 110 port 

to send valve commands and to receive instrumentation and status data from 

the user. It is assumed that three serial command channels continue the 

philosophy of three-string majority vote to the user and the user is active 

in this decision process. This is not strictly required, and it may be decided later that it is more appropriate not to extend the three-string majority vote process into the user system. The tradeoffs needed to determine the 

optimum approach are beyond the scope of the current study. Return information from the user can be provided on two serial data lines, which are examined by the three dewar computers. 

F8804112AA8-50 2-50 
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2.2.5.2 Majority Vote Unit 

The Mw is the final link in the three vote computing string where the votes 

are actually tallied for agreement or disagreement. Input to the MW consists of the command output words from the three computers and a system clock 

for synchronization. When at least two of the three simultaneous inputs are 

equal, an output is generated representing the agreeing command. If only two 

computers are in agreement, a message indicating which computer string disagreed is sent back to the SCS via the agreeing computers. 

The MW is a passive comparator, with no computing intelligence, made up of 

many identical sum-of-product (and/or) logic circuits. This circuit can 

readily be implemented in programmable logic arrays (PLA), which can significantly reduce component count. Typically, 15 to 20 discrete ICs can be replaced by a single IC package containing a PLA. 

2.2.5.3 Actuator Power and Control 

The valve drive unit (VDU) is the element of the dewar mounted avionics that 

provides appropriate control power to the cryogenic system actuators. Actuators include motor-operated valves, various heaters, and mechanical coolers. 

Due to the large number of identical valves, there is duplication of a large 

number of circuits. This duplication lends itself to the use of PLA for 

component reduction and increased reliability. 

Command inputs to the VDU come as redundant parallel words, representing 

actuator and action, from the MVU. Continuation of redundancy from the MW 

to the actuator, through the VDU, provides failsafe operation beyond the 

three string computer system. 

Valve Drive - The valves are driven open or closed by a four-phase variable 

reluctance stepper motor. The choice of stepper motors was made rather than 

dc motors because of their much higher efficiency, at the penalty of more 

wires and increased circuit complexity. Stepping sequence and control can 

readily be designed using PLAs followed by power switching devices such as 

DMOS FETs. Drive power and timing constraints will be included in the drive 

circuits to preclude electrical problems from a stuck valve or drive. 

Position feedback will be used to terminate actuation upon completion via the 

dewar based computer. 

TM Pump - The TM pump requires the most complex actuator control found in the 

SFHT. The pumping action is created by an electrical heater and pump 

operation is determined from several GRTs. Electrical power to redundant 

pump heaters is supplied by a linear closed control loop. Loop control is 

performed by a selectable primary dewar based computer using data received 

from flow rate sensing GRTs and a control algorithm that is part of its expert system firmware. The digital result of the computer signal processing 

is applied to an 8 bit digital-to-analog converter (DAC), in the VDU, whose 

output is an analog voltage representing the pump power required. The power 

amplifier supplying the pump heater power (up to 40 W) can best be accomplished with a programmable switchmode dc to dc converter with the DAC 

output being used as the programming control. This type of power amplifier 

can have relatively high efficiency along with a quiet dc output. 

F8804112AA8-51 2-51 
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Mechanical Coolers - The SFHT uses four Stirling cycle refrigerators. These 

refrigerators provide their own electronic control circuits, but the VDU provides power on/off control for them. The high current switching requirements 

for the refrigerators can best be handled by redundant contact relays controlled by simple relay drivers. 

The mechanical cooler subassembly includes five electronics boxes: a control 

unit for each of the four coolers and a power distribution unit. This is the 

same approach that will be used for the AXAF XRS cryogenic subsystem. Details of the control, data, and power interfacing with the refrigerator electronics are not discussed here, but are assumed to be very similar to that to 

be developed for the AXAF instrument. The AXAF XRS program has obtained 

preliminary interface and operations information from the refrigerator manufacturer, British Aerospace. 

2.2.5.4 Power Distribution Unit 

The PDU provides all power conditioning required by the dewar based avionics 

as well as power on/off control for the user. Our concept has the S HT asFa 

28 V dc system, and use with a host providing different power will require a 

power adapter. The PDU receives both primary and auxiliary dc power from the 

host with automatic switching to auxiliary power in the case of failure of 

the primary power. Dewar avionics power is connected directly at the PDU and 

is on/off controlled by redundant contact relays powered from the power 

switch located on the SCS switch panel. 

2.2.5.5 Thermal Control Unit 

The SFHT thermal control is the same as that used with OSCRS. This semiautonomous unit controls the avionics base plate temperature by controlling 

the opening of a metal foil shade covering the dewar mounted avionics and 

avionics baseplate. "'his temperature control is closed-loop in that the 

baseplate temperature is sensed by silicon diodes whose output, after processing, drives a stepping motor to position the shade. This simple control 

loop is independent of the dewar computers, except that the computers can 

command the shade to close in preparation for shutdown and transportation. 

2.2.6 Operations 

The SFHT flight avionics is designed to provide closed-loop control over as 

many aspects of operation as is applicable. In particular, tending and 

transfer of the cryogen is carefully controlled to provide reliability and 

efficiency. Each valve actuation is closely followed by observing the open/ 

close position sensors associated with that valve. Other sensors such as 

temperature, pressure, and flow are observed for their expected changes in 

response to the valve movement in accordance with the overall control scheme. 

The reading from a single sensor is considered not only on its own, but more 

importantly by what it indicates about the overall state of the system when 

considered with evidence from other sensors. The construction of a knowledge 

base and model such as this serves to provide operational reliability and 

efficiency as well as providing a means of contingency control should sensors 

fail. 

F8804112AA8-52 2-52 
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Major operational modes can be broken into four categories: normal operation, 

intialize/finalize, power save monitor, and emergency. Each major mode can 

be broken down into functional submodes that perform the specific tasks 

required. 

2.2.6.1 Normal Operation Mode 

operation 

0 

0 

0 

0 

0 

The normal operating, full power mode provides capability to conduct normal 

cryogen resupply operations. Functional submodes conducted from the normal 

mode include: 

Orbital SFHT pumpdown 

Standby (short term or long term) 

User berthing 

User cooldown and transfer 

Preparation for descent 

During normal operating mode, all computers and sensors will be operating. 

The primary SCS display/touchpanel will show the top level operating selections that the operator can make, plus the status of key parameters including key predictions and trends. The secondary displays can provide supporting information. Any anomolous condition alerts will be displayed in 

real-time at the appropriate location. 

2.2.6.2 Initialize/Finalize 

Initialization occurs upon application of power to the SFHT avionics. When 

power is applied, a self-test and configuration determination routine is performed. During this period, all sensors will be examined to determine cryogen and avionics conditions. Interfaces will be exercised and verified, and 

any changes in system status will be conveyed to the operator and/or ground 

control along with operating status and applicable operational recommendations. In the event that no operator interchange occurs within a certain 

period of time and normal dewar conditions exist, the system will revert to a 

power saving posture. 

Finalization is an operator initiated functional mode in which the system 

prepares for removal of power. During this operation, the system assures 

that all cryogen valves and actuators are properly positioned. Given the 

plans for shutdown duration and other environmental aspects, the expert system can formulate predictions of cryogen loss during the unmonitored state. 

If conditions are detected during finalization that could create difficulties 

during the unpowered state, the operator will be notified of the situation 

and options. 

2.2.6.3 Power Save 

Two power saving modes are available. The first is a low power operational 

monitoring mode. In this mode the dewar avionics continue to operate but at 

a much reduced level. Power reduction methods include: only a single channel computer is operational, only critical sensors are excited and monitored, 

sensor reading rates are reduced, display memory updating is slowed, and 

F8804112AA8-53 2-53 
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communications are limited. Interfaces with the operator and emergency channels are still operative, but the avionics activities are restricted to only 

critical functions. 

The second, even lower power, critical warning monitor mode uses a single 

computer string, monitors only critical parameters, and provides no touchscreen interface. This mode is initiated and terminated by a switch on the 

SCS switch panel. If an anomolous condition is detected, the operator is 

notified by an alert dispalyed on the CWEA. This is a request for the operator to switch back to the normal operation mode so that the nature of the 

anomolous condition and recommended corrective action can be displayed. 

Corrective action would usually be implemented automatically, but the avionics must be switched back to the normal operation mode so that it can command 

corrective action. An example of this would be a plumbing failure resulting 

in helium leakage into the guard vacuum and large heat input to the cryogen 

tank. This occurrence would be detected as a fast and steady rise in flowrate and bath temperature and pressure. The corrective action would be to 

open the high-flow vent in hopes that the tank pressure could be controlled 

without blowing the burst discs. This would simplify the eventual repair 

process. 

2.2.6.4 Emergency 

This mode is an automated quick safing method for the SFHT and its attached 

user. It does not remove the manual safing controls from the operator but 

only automates the procedure. The emergency mode is available from all other 

modes except the warning only monitor mode. The following functional submodes are accessible in the emergency mode. 

Shutdown - This functional submode shuts down the SFHT and/or user dewars and 

co?trols venting or cryogen flow into benign plumbing conduits. 

Jettison of the User from the SFHT - This action will require firing of pyrotechnic devices for activation of the decouplinn -- nuillotine that severs the 

fluid line and electrical cables. This capability may not be needed for the 

follow-on concept where couplings are mated and demated automatically. Actual firing of these devices must be done by the operator; and the expert system only makes preparation for the firing and displays instructions for the 

operator. Simultaneous with instructing the operator, the SFHT will put the 

user into as favorable a state as possible before jettison. 

Jettison of the SFHT from the Cargo Bay - This functional submode sets both 

the SFHT and the user into states which are the most safe to the orbiter, and 

secondly to states that will prevent or minimize damage to either the user or 

the SFHT. 

Emergency Checkout - This functional submode provides a means by which the 

operator can rapidly determine the status of all critical functions within 

the SFHT and its external interfaces. It presents both critical limits and 

actual measurements along with questionable or anomolous trends. Operational 

status of both the cryogenic and avionic subsystems will also be displayed. 

Ongoing activity will not be altered by this checkout mode, and it is directly accessible from other modes. 

F8804112AA8-54 2-54 
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2.2.6.5 Data Storage 

There are two considerations to storage of cryogen and avionics sensor data. 

First, if TDRSS is used in conjunction with the downlink data, a 40 to 

90 percent daily coverage must be considered. Based on the moderate data 

rates expected from the SFHT, a relatively small amount of temporary storage 

would be required. This could readily be formed with available static RAM. 

Secondly, it may be desirable to store a complete mission data set for reconstruction on the ground near or after mission end. 'Assuming a 12 month 

mission with two 10 hour resupplies, the data could be stored in 90 K bytes 

of memory. Incorporating this memory in solid state non-volatile form, such 

as EEPROM, can save significant volume and power. Storing the data in three 

stores, one for each computer, will provide significant redundancy. 

2.2.7 Redundancy 

The only safety related function of the avionics is the command firing of the 

pyrotechnically operated guillotine for emergency jettison of the user, as 

discussed in 2.2.6.4. In all other respects, the SFHT concept satisfies STS 

and SS safety requirements with passive design features that require no power 

or command. During times when the avionics is powered, unusual conditions 

will be detected by the avionics and displayed at the SCS, downlinked to MCC, 

and if necessary displayed on the CWEA. 

Two fault tolerance for safety and single fault tolerance for mission success 

are maintained in the avionics by using a three string majority vote computer, redundancy, and the ability to interpolate around failed sensors. 

This is followed by hardwired manual controls to critical operations. 

A three string computational control approach (Figure 2-24) has been baselined. Three computational strings operate independently in the process of 

making a command decision; then each outputs its vote to be processed by a 

majority voting unit. If at least two out of three strings agree, the vote 

passes and becomes a command to operate an actuator. Three string computation with majority voting is employed in all operations where cryogen fluid 

control is involved. 

All dewar sensors are redundant, and each member of a pair is controlled and 

measured from a different dewar computer. If both sensors of a pair should 

fail, sufficient data will be supplied from other sensors to allow mission 

completion. 

Cryogen valves are generally series-parallel redundant for opening and closing cryogen and vent lines. This provides sufficient failure tolerance to 

meet all SFHT safety and reliability objectives. 

Failure of any two computational strings can be handled by operator activation of the manual emergency safing sequence from the SCS switch panel. The 

remaining computational string will function to provide monitoring of the 

safing process. The SFHT avionics will attempt to control the user and render it usable and intact under all conditions. 

F8804112AA8-66 2-55 
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2.2.8 Adapting to Different Host Vehicles 

Our baseline concept has been developed primarily to perform resupply at the 

shuttle, with resupply at the SS as an alternative. Resupply at the user 

orbit with the OMV as host has been given low priority in the concept definition effort. 

2.2.8.1 Space Station 

We envision installing a SCS in the SS that is the same as used in the STS. 

Major differences between SS and STS are the internal communications and 

power. Figure 2-25 is a block diagram showing the electrical interface between the SFHT and SS. Communication between the SS LAN and SFHT, and also 

between the SCS and dewar-based avionics, will require converters to simulate 

the STS system. Power for the SFHT will require ac-to-dc converters to 

change the SS 440 Vac at 20 kHz to 28 Vdc. 

2.2.8.2 OMV 

Figure 2-26 shows the envisioned interface between the OW and SFHT. Special 

interface adapters exist between each of ON'S command/telemetry busses and 

the SFHT. Two interfaces are required to achieve mission success redundancy. 

A video signal interface to route berthing and boresight camera signals 

through the OMV downlink to the MCC is required. 

2.2.9 Deviations from the OSCRS concept 

At the direction of JSC, the Martin OSCRS design concept has been used as a 

starting point for our SFHT avionics design concept. Commonality between the 

OSCRS and SFHT avionics designs, especially regarding operator interfacing, 

would have obvious benefits. The following changes to the Martin OSCRS 

design are recommended for the baseline SFHT concept. 

e Three flat panel displays instead of four 

e True three independent string majority vote design rather than 

"vote of confidencen design 

0 Take advantage of modern technology for improved reliability, 

reduced power, and lower parts count 

0 Simplify switch panel 

0 Use solid state nonvolatile memory (e.g., EPROM) for data storage 

instead of optical disc 

The rationales for these changes are given below. 

Displays - The OSCRS design uses four flat panel displays at the system control station. Three displays, including the display collocated with the 

switch panel, are more than adequate to provide the MMI for the SFHT. The 

SFHT data rates and quantity of sensors are less than with OSCRS; hence less 

total display capability is required. 

F8804112AA8-57 2-57 
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2-58 
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2-59 
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Switch Panel - The OSCRS switch panel includes a joy stick and calculatorlike pushbuttons as inputs. The SFHT will utilize a touchscreen as the primary input, which will provide a simpler and more convenient MMI. Toggle 

switches will have protective covers to prevent inadvertent actuation. The 

resulting switch panel will be less cluttered and easier to operate. 

Modern Technology - The OSCRS design makes use of many time tested technologies, which unfortunately are not as power or space efficient as current 

technology will provide. The SFHT avionics concept employs CMOS microprocessors from the 8086 family for all computers. Considering the rapid 

advancement in microprocessors and the growth path for the SFHT, an 80386 is 

the microprocessor of choice for the dewar based computers. 

The many redundant and repetitive circuits in the SFHT avionics affords an 

excellent opportunity to reduce component count, reduce volume, reduce cost, 

increase reliability, and reduce testing through the use of programmable 

logic devices. A single programmable device can conservatively replace 15 to 

20 IC packages. 

Data Storage - The OSCRS design includes two write-once-read-many optical 

disk drives at the control station. Each drive can store 200 M bytes per 

side. The SFHT does not require anywhere near this much data storage. It is 

convenient and efficient to provide solid state non-volatile data storage in 

either the SCS or dewar mounted computers. No development or procurement 

costs are involved in this approach, and higher reliability is expected than 

with a rotating device. 

Computer Redundancy - The SFHT three string majority vote computer system is 

configured differently than that used in the OSCRS. The SFHT system is simpler and provides the redundancy required to meet all operational goals. An 

added advantage of the SFHT approach is that the three strings can be extended into the user with the user making a majority vote decision for commands it receives. 

Cargo Bay Payload Retention Latches - The OSCRS design provides monitoring of 

the payload bay retention latches including opening, closing, and pin insertion. For a two trunion configuration this means twelve position sensors, 

including redundancy. The SFHT will not provide this monitoring because 

monitoring of cargo in the cargo bay is best left to the STS. It would not 

be expeditious or efficient to rely on the SFHT to provide this data since 

SFHT must be powered to provide such monitoring. SFHT does provide monitoring of the coupling mechanisms for resupply, such as the berthing latches and 

fluid coupling. 

2.2.10 Development Issues 

Expert System - Many reasons have been presented in various sections for the 

rationale behind incorporating an expert system design into the SFHT avionics. The SFHT is a very good application for an expert system not only from 

the standpoint of having an automated advisor available but also when the 

evolution to remote autonomous resupply operation is considered. The key 

development effort in a system such as the SFHT involves designing the knowledge base required. No general methodology formula yet applies because of 

the nature of knowledge base characteristics. Fortunately, both tools and 

F8804112AA8-60 2-60 
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techniques for doing this are rapidly evolving and fit the SFHT timeline. We 

envision the knowledge base and inference rule set used in early SFHT missions to be relatively simplistic, but as the system becomes better understood through use, the expert system can become "smarter." Initial development and continued evolution of the expert system must be planned for. 

TMP-Flowmeter-Mass Gaging - The SHOOT program will prove the in-flight viability of the TM pump and pulsed heat mass gaging techniques. The SFHT can 

benefit from, and depends upon, the SHOOT progress to make SFHT design objectives attainable at low risk. 

The SFHT uses a closed-loop control of the TM pump. The control input is 

pressure downstream of the pump and temperatures on each side of the pump, 

and the output variable is pump flow rate. The TM pump flow rate is a function of power supplied to its heater and is determined by a calibration curve 

for the porous plug and heater set. Complete ground characterization of the 

control loop must be made before stable operation can be assured. 

The SFHT mass gaging technique is the same as the SHOOT method but must be 

scaled for larger helium quantity. Continued work must be done on 

characterization of the temperature response profile vs. input power pulse 

for the SFHT. Impact on dewar performance of the heater lead wire size must 

also be considered. 

Flow rate through the standby (low-flow) vent will be determined by applying 

a small, known heating rate to the flow stream and measuring the resultant 

temperature rise, as described in 2.1.6. The electrical heater power and 

temperature rise must be accurately measured. Development of the electronic 

requirements for operation of such a flow meter will be required. 

2.3 GROUND SUPPORT EQUIPMENT 

Ground support equipment needed to prepare the SFHT for launch will be a 

significant part of the program development effort. As determined in the 

first phase of the study, it is advisable for several reasons, system mass 

penalty being the main driver, to launch the SFHT with SfHe in the dewar. 

This has been done; the IRAS dewar was 93 percent full at a temperature of 

1.7 K when launched. However, the fluid capacity of the SFHT is about 20 

times greater than IRAS. Filling a dewar of this size with SfHe has not been 

done, and the process and equipment involved should be carefully planned. 

Selection of the fill procedure must consider development cost, development 

and operational risks, operational cost, facility constraints, and 

operational requirements. 

Selection of the GSE design is a five-step process: 

1. 

2. Identify all possible approaches. 

3. 

4. 

5. Select the most desirable solution. 

Understand the operational requirements and design constraints. 

Define viable approaches to the extent needed to evaluate their 

merits. 

Evaluate and compare the viable alternatives based on the criteria 

listed above. 

F8804112AA8-61 2-61 
F88-04 

Our choice is the same approach used for IRAS and COBE; i.e., successive 

topoffs at a pressure just above the lambda point. The rationale for this 

selection is given in the remainder of the section. 

2.3.1 GSE Design Concept 

GSE consists of three major categories: 

1. Mechanical - This is equipment needed to move the SFHT around 

within a given facility and to ship it long distances. It 

consists mainly of a handling dolly and a shipping container. The 

design of this equipment is straightforward and won’t be discussed 

further. It is assumed to be similar to that used in support of 

the IRAS and COBE dewars. 

2. Cryogenic - This includes the supply dewars (custom made), pumps, 

heat exchangers, transfer lines, refrigerators, etc., needed to 

accomplish fill of the SFHT dewar and maintenance of the dewar 

guard vacuum. 

3. Monitor/Control - This includes the computers and electronic 

equipment needed to monitor and control the dewar during all 

ground operations (up until payload integration). This equipment 

is not the flight avionics, although it will take advantage of 

applicable software and hardware designs created for the flight 

avionics. The monitoring system will likely evolve toward greater 

automation as SFHT missions are accomplished. The monitoring GSE 

is discussed further in 2.3.1.2. 

The remainder of this section is concerned with the cryogenic GSE. 

Requirements and Constraints 

Operational requirements and design constraints for the cryogenic GSE are 

relatively simple; this is largely due to the great size and excellent thermal performance of the SFHT dewar. These two features provide an extremely 

long lock-up capability; i.e., the time it takes for the bath to heat up from 

its nominal ground temperature of 1.7 K to the lambda point (2.18 K) with no 

venting. This time is predicted at 4 months with the flight mechanical coolers running and 25 days with the coolers turned off. This superior lock-up 

capability provides flexibility in planning and performing the prelaunch 

operations. Our recommended facility usage and schedule of operations is 

presented in 2.3.2. 

The GSE must interface with the SFHT fluid connections and fill the dewar to 

nearly full with superfluid helium. The fuller and colder the cryogen tank 

is, the more efficient the system is from the perspective of launch mass. 

The IRAS dewar was launched 93 percent full at a temperature of 1.7 K. There 

are methods that will improve this fill efficiency, as will be discussed. 

There is no hard constraint on the time required for the SFHT fill process; 

1 month is realistic. The faster the dewar must be filled, the larger and 

more expensive the equipment involved, but the manhours required may be 

fewer. There are GSE approaches that minimize the man attendance required, as 

will be discussed. 

F8804112AA8-62 2-62 
F88-04 

Figure 2-27 is an options tree for filling the dewar. The option of simply 

filling and launching the dewar with normal helium is shown for completeness, 

but discarded because of the severe system mass penalty (approximately 50 

percent) involved. The next major division is to produce the superfluid helium either by an evaporative process, as done for IRAS, or by an active cooling approach (i. e. , Joule-Thomson expansion) . Approaches using active cooling are more complex in concept and require more complex SFHT dewar plumbing. 

They are discussed separately in 2.3.1.1. 

The evaporative branch is divided according to the condition of the fluid 

transferred from the GSE dewars to the SFHT dewar: 15 psi (normal), 1.0 psi 

(just above the lambda point, as done for IRAS and COBE), and 0.2 psi (temperature approximately 1.7 K) . This third possibility requires a GSE dewar 

with a TM pump and can theoretically provide a 100 percent full tank at 

1.7 K. Practical operational considerations, like vacuum pump size, might 

result in a temperature somewhat higher than 1.7 K. 

There are two major distinctions between the first and second evaporative 

methods. The second method can provide a much smaller ullage at launch, but 

it requires a supply dewar capable of operating at negative gage pressures to 

allow pumpdown of the supply bath. Given that the supply dewar must be 

custom designed and built, this is not a significant cost impact. Caution 

must of course be exercised though operational procedures when handling a 

negative-pressure dewar to insure that air ingestion and ice blocking don't 

occur. During SFHT verification testing, there will be extended periods of 

operation at negative pressure. The system must therefore be designed for 

negative-pressure operation regardless of how the prelaunch fill is accomplished. Since the alternative of transferring from a normal supply dewar 

offers no significant advantage and incurs a large launch mass penalty, it is 

eliminated from further consideration. ' 

Figure 2-28 is a schematic of the IUS and COBE cryogenic GSE; the complexity 

of the equipment and operations is evident. This basic equipment would be 

required regardless of whether fluid is transferred just above the lambda 

point or below the lambda point. There is one major distinction between the 

two approaches. When transferring SfHe, the supply tank bath must be vacuum 

pumped to remove the heat produced by the TM pump heater. It would be necessary to transfer at a relatively high rate (approximately 1,000 L/hr) to 

minimize the degrading effect of parasitic heat input through the transfer 

system, just as is the case during on-orbit transfer. A rather large 

(approximately 3,000 CFM) vacuum pump would be required. When transferring 

above the lambda point, the flow can be driven by pressurization. However, 

the receiver bath must be vacuum pumped to remove the flashoff produced by 

the transfer system parasitic heat leaks. This can be done successfully with 

a relatively small vacuum pump; a 300-CFM pump was used for IRAS and COBE. 

While it is desirable to minimize the vacuum pump size, there are no facility 

constraints that prohibit use of a 3,000-CFM pump. 

The tradeoff between transferring above or below the lambda point comes down 

to cost and development risk: the cost of slightly increased launch mass 

versus the cost and risk of developing a GSE dewar that uses a TM pump. The 

TM pump method will also save time and money in the prelaunch operation. 

F8804112AA8-63 2-63 
p Irl F88-04 

TRANSFER 

AT 15 PSI 

MUST LAUNCH 60% 

FULL SINCE NO 

TOPOFF POSSIBLE. 

UNACCEPTABLE 

MASS PENALTY 

I TAKE LlOUlD 

HELIUM I INTO ORBIT 

CONVERT BY 

EVAPORATION 

TRANSFER 

AT 1 PSI 

7% ULLAGE AT 1.7K 

ACHEIVAELE (IRAS). 

TM PUMP N,OT 

REQUIRED IN GSE 

DEWAR. 

TRANSFER 

AT 0.2 PSI 

ZERO ULLAGE AT 1.7K 

THEORETICALLY 

ACHIEVABLE. GSE 

DEWAR MUST HAVE 

TM PUMP. 

ZERO ULLAGE AT 1.7K MAY BE ACHIEVLABLE: PERFORMANCE UNCERTAIN WITI(OUTlK0EPlH 

NauAnm RISK LEVEL mm WITHOUT 

ZERO ULLAGE AT 1.7K MAY BE ACHIEVLABLE: PERFORMANCE UNCERTAIN WITI(OUTlK0EPlH 

I COOLINO I NauAnm RISK LEVEL mm WITHOUT 

2K PRODUCED 

AT EXTERNAL 

2K PRODUCED 

BY J-T VALVE 

Figure 2-27 Alternative fill process 

F8804112AA8-64 2-64 
~ 

F88-04 

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F8804112AA8-65 2-65 
F88-04 

If the launch cost is mass driven, as with the oblate spheroid dewar configuration, the launch cost savings of the third approach is estimated at %1.5M. 

If launch cost is driven by shuttle bay space, as is the case with the 

spherical configuration, there is no launch cost savings. Based on these 

consideration and our recommended spherical configuration, our choice is the 

second approach, the same as used for IRAS and COBE. However, the use of a 

supply dewar with a TY pump is a very tempting alternative that should be 

revisited as SfHe transfer technology continues to mature. 

2.3.1.1 Production and Maintenance of Stored Superfluid by Means of Active 

Refrigeration 

There are at least four active cooling schemes that can be used to produce 

and maintain the superfluid within the SFHT as shown in Figure 2-27. In all 

of these systems it is assumed that the SFHT has been filled with normal 

helium from a separate source, and that makeup or topping will be done 

through that source. These active cooling systems are closed loop in the 

sense that there is no fluid communication between the cooling system and the 

stored helium. The true advantages and disadvantages of these arrangements 

compared with the direct evaporation method will only become clear through 

much more detailed study. Some potential advantages of active cooling are 

reduction of GSE cost and lower chance of line ice blockage due to "negative" 

pressure in the tanker. 

Figure 2-29 shows a completely external mechanical cooler that produces superfluid temperature across the J-T valve from A to B. The internal heat 

exchanger removes heat from the helium bath. The heat exchanger, shown 

mounted internally, could also be mounted to the outside of the cryogen tank 

if carefully located. To maintain the proper low temperature in this process, the compressor must have a rather large vacuum pump at its first stage. 

An advantage of this approach is that almost all of the refrigeration equipment is external in the GSE and accessible. A disadvantage is that superfluid must be transferred through the GSE connection and through warmer regions of the insulation before it can be used to remove heat from the tanker. 

This is a highly inefficient way to transfer the SfHe and will lead to a 

large and costly refrigeration system. 

Figure 2-30 shows the J-T valve moved to the inside of the tanker. Superfluid will be produced by expansion from B to C so the difficulty with pressurized superfluid transfer of 2-29 will be eliminated. The compressor will 

still require a vacuum pump for the first stage to produce the required temperature in the heat exchanger, but this pump is expected to be much smaller 

than for 2-29. The disadvantage is that the J-T valve is buried within the 

tanker where it cannot be easily accessed and maintained. 

Figure 2-31 shows another version of 2-30, which employs an internal ejector. 

The function of the ejector is to produce a vacuum inside the tank by using 

part of the helium flow to produce vacuum on the other portion of the flow by 

the Bernoulli effect. This scheme has been used in low temperature coolers 

both to avoid the thermal penalty of the large vent tube passing through the 

insulation and to eliminate the vacuum pump first stage of the compressor. 

(The helium will be returned to the compressor at about 1 atmosphere.) The 

F8804112AA8-68 2-66 
 

F88-04 

COMPRESSOR 

INTERMEDIATE 

COOLING STAGES 

(NOT ALL SHOWN) f 

JOULE-THOMSON 

VALVE 

GSE CONNECTIONS THERMAL STANDOFF 

DEWAR FILL DEWAR VENT LINE 

HEAT EXCHANGER 

AM 9344MD14222 

Figure 2-29 Active cooling using external J-T valve and internal heat 

exchanger 

F8884112AA8-67 2-67 
F88-04 

GSE CONNECTION 

COMPRESSOR 

INTERMEDIATE PRECOOLING 

THERMAL STANDOFF 

DEWAR VENT LINE 

HEAT EXCHANGER 

STAGES 

NN 9344iMD14224 

Figure 2-30 Active cooling using internal J-T valve and heat exchanger 

F8804112AA8-68 2-68 
F88-04 

JOUL 

COMPRESSOR 

INTERMEDIATE PRECOOLING STAGES 

GSE CONNECTIONS 

THERMAL STANDOFF LENGTH 

DEWAR FILL LINE DEWAR VENT LINE 

HELIUM DRIVEN EJECTOR PRODUCES 

1.6 K IN HEAT EXCHANGER 

BY LOWERING THE PRESSURE 

.E THOMSON VALVE 

HEAT EXCHANGER 

TO REMOVE HEAT 

FROM TANKER 

STORAGE AT 1.6 K 

SFH 

INSULATED LINE 

NN 9344iMQ142.23 

Figure 2-31 Active cooling using internal J-T valve, heat exchanger, and 

e j ector 

F8864112AA8-69 2-69 
F88-04 

disadvantage of 'this system is that the ejector and J-T valve both are in the 

cryogen tank where they are difficult to maintain. 

Thermodynamic and Thermal Considerations 

We now examine the penalties and sizing of the various active cooling 

techniques. A first step in doing this can be taken by studying Figure 2-32. 

In this scheme, a pressurized tanker supplies liquid helium through the 

connection to point B, which has a condition of elevated pressure and 

temperature, i.e., 4 atmospheres and 5 K. Four atmospheres is about the 

highest pressure available from a transport dewar, and 5 K supercritical 

helium is produced from the 4.2 K liquid by the heat leak of the 

interconnecting plumbing. The enthalpy of this fluid is 14.32 J/g. An 

isenthalpic expansion takes place across the J-T valve as shown in Figure 

2-33. If the vacuum pump is able to maintain a pressure of 5.6 torr at C, 

then a temperature of 1.6 K will be produced to remove heat from the tanker. 

The amount of heat that can be removed is proportional to the increase in 

enthalpy of the helium in the heat exchanger as the helium boils. If all of 

the superfluid helium boils, but the temperature at D is not allowed to rise, 

then the amount of heat that can be absorbed at 1.6 K is 23.09 - 14.32 = 8.77 

J/g . 

The total amount of heat that must be removed to convsrt 10,000 Q. of 4.5 K 

normal helium to 1.6 K superfluid helium is 14.2 x 10 Therefore, :he 

total amount of.helium that must flow through the cooling loop is 1.6 x 10 g 

or about 13,000 e of normal liquid helium from the supply, assuming there are 

no other losses. Note that if the temperature at B is allowed to rise to 

6.9 K at 4 atmosphere, the cooling.available will have decreased to zero. 

This temperature sensitivity of the cooling effect makes the cooling system 

approach a very imposing design problem requiring significant study.' Simply 

increasing or decreasing the supply pressure will not be an adequate solution 

as can be seen by study of the thermodynamic properties of helium in this 

region.. 

J. 

The closed-loop compression driven systems do not waste helium, but the same 

inlet line temperature sensitivity problem exists. In all of the active 

cooling approaches a very important consideration is the size of the return 

line that must pass through the SFHT insulation space. The density of the 

return vapor is extremely low, and the pressure losses due to flow must be 

held to less than about 1 torr to avoid a significant rise in the desired low 

storage temperature (approximately 1.6 K) . 

This leads to trade-offs involving vent line heat leak, vent line size, tanker thermal efficiency, and cool down time. The use of the internal ejector 

could largely desensitize this trade-off. The feasibility and risks of the 

active cooling methods cannot be well understood without a considerably more 

detailed evaluation than possible within the scope of this study. 

Conclusion 

Without a more detailed evaluation, it is not clear whether an active cooling 

approach can be made to work well for the SFHT or what the advantages might 

F8804112AA8-70 2-70 
F88-04 

VACUUM PUMP 

GSE CONNE 

THERMAL STANDOFF 

DEWAR VENT LINE DEWAR FILL LINE 

JOULE THOMSON VALVE 

HEAT EXCHANGER 

A/N 9344iMD142.25 

Figure 2-32 Internally produced refrigeration using pressurized external 

storage tanker 

F8804112AA8-71 2-71 
TEMPERATURE P I 5.6 TORR 

DECREASE DUE TO 

INTERMEDIATE HEATING UP AT 

COOLING STAGES 

PRESSURE) 

A 

s 

k 

P I 4 ATM 

T = 5K A h I 23.09-14.32 

I 8.77 JIG 

CONSTANT ENTHALPY 

SATURATION CURVE 

SUPERFLUID = 5.6 TORR 

EVAPORATING 

PROVIDING COOLING 

IN TANKER 

S (ENTROPY) JIG-K i 

AJN 9344lMD142.26 

Figure 2-33 Temperature-entropy diagram of helium showing the cooling 

process . 

F8804112AA8-72 2-72 
F88-04 

be compared to an evaporative technique. Active cooling approaches all 

result in design impacts to the SFHT and associated developmental and 

operational risks. The use of a GSE dewar with a TM pump will provide lower 

launch mass than any active cooling method, although it also involves 

developmental risk compared to the IRAS/COBE approach. Considering all this, 

we baseline the proven, low-risk procedure used for filling the IUS and COBE 

dewars. 

2.3.1.2 Monitor/Control Equipment (MCE) 

The MCE provides monitor and control of the dewar during the final stages of 

in-process testing, during system testing, during maintenance and prelaunch 

fill operations, and for post integration checkout. Its functions include 

monitoring of all dewar sensors, control of motor-operated valves and 

heaters, and data reduction and storage. 

If the SFHT uses mechanical coolers, as currently baselined for SS-based 

resupply, the MCE must also provide control commands and monitoring of the 

coolers and related sensors. This capability will be needed during system 

testing and possibly during ground lock-up, if it is decided to use the 

coolers at that time. 

Monitor Functions 

Dewar sensors fq11 into two basic categories: 1) those, such as temperature 

sensors, which indicate the general condition or health of the dewar, and 2) 

those, such as valve position indicators, which give the status of active 

devices. Table 2-6 is a complete list of sensors and electrical devices in 

and on the dewar. Only one item on that list, the EVA coupling position 

sensor, will not be monitored by the MCE. 

Monitoring serves several functions. It provides characterization of the 

dewar performance so that changes in behavioral patterns can be detected. 

These changes can then be related to quickly or slowly developing 

malfunctions that might be correctable. Based on the dewar performance 

observed during system testing, regions of abnormal readings will be 

established for various critical sensors. Out-of -limit warnings, both visual 

and audible, will be provided. The audible warning will be used to announce 

impending burst disc rupture or that a guard vacuum failure is in progress; 

these would be. indicated by the bath pressure. No action would be required 

other than to stand clear of the emergency vent. During dewar fill and 

topoff operations, certain sensors will be used as the signal to proceed to 

the next step of the operation. The MCE will provide continuous indication 

of all valve positions and heater power. 

Control Function 

The MCE will provide commands via mechanical switches and 28 V power for 

actuation of valves. It will also provide constant or variable 28 V and 

100 V power to the various heaters within the dewar. If mechanical coolers 

are used in the design, the MCE will provide command and power to the 

coolers. 

F8804112AA8-73 2-73 
F88-04 

Data Reduction and Storage 

In some cases, raw data from more than one sensor is combined according to an 

algorithm to provide the desired information. This is true, for example, of 

the mass quantity gage that determines mass quantity based on heater current 

and voltage drop, fluid temperature rise, and a curvefit of fluid specific 

heat versus temperature. In other cases, such as thermometers, the 

engineering data is produced from the sensor voltage output by a simple 

curvefit. All data generated during the test program and during prelaunch 

fill operations will be recorded for future reference and troubleshooting. 

Hardware Description 

There are two separate sets of MCE: 1) the full-up MCE that is used for dewar 

testing before the dewar-based avionics package is installed and 2) the 

partial MCE used after the dewar-based avionics is installed. The full-up 

set essentially duplicates the control and monitor functions of the’flight 

avionics, although aspects such as packaging, display, redundancy, and use of 

high-reliability parts will be different from the flight system. The partial 

MCE performs the functions of the remote segment of the flight avionics. The 

partial MCE is a segment of the full-up MCE; the interfacing would have to be 

somewhat modified to be compatible with the dewar-based flight avionics. The 

partial MCE will also be able to monitor selected interfaces between SFHT 

functional blocks, SFHT and host, and SFHT and users to verify correct 

interface functions. 

The MCE consists of the following major components: 

0 

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AT-sized personal computer with interface buses, printer, and 

plotter 

Digital voltmeter 

A variety of power sources for valves and heaters 

Constant current sources for various sensors 

Controller/readout for liquid level sensor and Vac-ion gage 

Calve drive electronics 

Digital-analog converter 

Scanner 

Front control panel 

Connector panel 

Sensor stimulation board 

Strip chart recorder 

Audible warning system 

2.3.2 Ground Operations/Facilities 

K-STSM-14. 122 prescribes that the filling of cryogens be accomplished on vertically integrated payloads. Payloads involving hazardous operations are 

normally serviced in the PHSF23 or the SAEF 2. The SFHT will be serviced and 

installed into the orbiter in a shorter time before launch when filled and 

installed in the sequence dictated by vertical, rather than horizontal, processing. There is almost continuous access to the SFHT during the vertical 

processing flow until it is transferred to the vertical canister for 

F8804112AA8-74 2-74 
F88-04 

transporting to the RSS. Vertical integration is accomplished in the VPF. 

Post-flight operations are performed in preparation for storage and the next 

flight. 

Horizontal processing is also described so that a comparison can be made to 

the vertical processing time frame and operations. 

2.3.2.1 Vertical Processing 

Figure 2-34 summarizes the flow of operations to install the SFHT in the 

orbiter when it is in the vertical position. The time to process the SFHT 

from receipt at KSC to launch is shown by the schedule in Figure 2-35 to be 

1,041 hr (43.4 days). 

The SFHT is transported to the PPF in its shipping container upon receipt at 

KSC. The shipping container provides and maintains a clean environment for 

the SFHT during the handling at KSC. The shipping container cover is removed 

at the PPF for damage inspection, functional tests, and checkout. 

After completion of the functional tests and checkout, the shipping container 

cover is reinstalled, and the SFHT is moved to a PHSF or SAEF 2 area to load 

it with SfHe. The SFHT will have 

some final tests and checkout made in preparation for its next move to the 

VPF. 

The fill operation is discussed in 2.3.1. 

The fill operation will take a maximum of 30 days. 

Upon arrival at the VPF, the shipping container will undergo cleaning operations in the airlock. Then the unit will be moved to the bay area where the 

shipping container cover will be removed, and the SFHT will be hoisted into 

the VPHD. The SFHT will be connected to the CITE simulated orbiter 

interfaces to verify all hardware interfaces including redundant paths. 

These tests will include power turn-on, health and status checks, power 

control tests, and SFHT command, monitor, and control checks. All data and 

command paths, all formats, and the data stream will be verified in mission 

simulation tests. 

GSE will be disconnected after all tests are completed in preparation for 

installation into the canister and transporting to the RSS. Any GSE required 

to support the RSS and launch pad SFHT operations will be installed and 

validated in the RSS prior to arrival of the SFHT. The SFHT must be 

installed into the RSS prior to the shuttle transfer to the launch pad. 

After the canister is installed in the RSS, the SFHT will be removed by the 

PGHM, which will retract into the RSS. The canister is then removed after 

the RSS doors have been closed. When the RSS is moved into position to 

enclose the orbiter, the PGHM will install the SFHT into the cargo bay. 

SFHT-to-orbiter interfaces are connected and verified, and then the SFHT is 

ready for launch. Launch can take place a minimum of 247.5 hr after the SFHT 

is installed in the orbiter or 309.5 hr (12.9 days) after the SFHT is filled. 

The predicted dewar lock-up capability of 25 days, without the mechanical 

coolers running, therefore provides 12 days of margin to accommodate 

integration and launch delays. 

F8804112AA8-75 2-75 
F88-04 

yef ess -+ VERIFY + TRANSFER SFHT + MOVE TO LAUNCH+ ~~~G\~~DED 

RETRACTED 

INTO RSS 

INTERFACE TO CANNISTER PAD RSS 

TESTS 

PERFORMCITE 

STORAGE 

AT KSC 

+ 

-b MOVE TO 

PPF 

I PJ3 inn 

PHYSICAL SEtS AEF .I . H pZZAZ: H ::.,SFHE 

CHECKOUT 

I 'I I-' 

4T MOVE TO 

9RBITFq 

INSTALLSFHT 

INTO ORBITER 

CONNECTSFHT 

TO ORBITER 

INTERFACES 

LAUNCH 

SUPPORT 

LAUNCH 

OPERATIONS 

NN 9344ND142.13 

Figure 2-34 Preparation of SFHT for launch through the vertical processing 

facility 

F8804112AA8-70 2-76 
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F8804112AA8-77 2-77 
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2.3.2.2 Horizontal Processing 

Figure 2-36 summarizes the flow of operations to install the SFHT in the 

orbiter when it is in the horizontal position. The time to process the SFHT 

from the receipt at KSC to launch is depicted by the schedule shown in Figure 

2-37 to be 1,480 hr (61.7 days). 

The SFHT is transported to the PPF in its shipping container upon receipt at 

KSC. The shipping container cover is removed at the PPF for damage 

inspection, functional tests, and check out. 

After completion of the functional tests and checkout, the shipping container 

cover is reinstalled, and the SFHT is moved to a PHSF or SAEF 2 area to load 

it with SfHe. The SFHT 

will have some final tests and checkout made in preparation for its next move 

to the O&C building. 

The fill operation will take a maximum of 30 days. 

Upon arrival at the OM, the shipping container will undergo cleaning operations. Then the unit will be moved to the bay area where the SFHT will be 

removed from the shipping container and installed in the CITE-simulated 

orbiter interfaces to verify all hardware interfaces including redundant 

paths. These tests will include power turn-on, health and status checks, 

power control tests, and SFHT command, monitor, and control checks. All data 

and command paths, all formats, and the data stream will be verified in 

mission simulation tests. 

GSE will be disconnected after all tests are completed in preparation for 

installation into the canister and then transporting to the OPF. The SFHT 

will be hoisted from the canister in a horizontal attitude and positioned 

over the orbiter and then lowered and secured into the cargo bay. SFHT-toorbiter interfaces are connected and verified, and then the SFHT is ready for 

launch. Launch can take place a minimum of 710 hr after the SFHT is 

installed in the orbiter or 748.0 hr (31.2 days) after the SFHT is filled. 

Compared to the vertical processing schedule, an additional 18 days of dewar 

lockup is therefore required. 

F8804112AA8-78 2-78 
F88-04 

pgc. Bss 

STORAGE 

SFHT AT KSC 

-b MOVE TO + 

oac 

MOVE TO 

PPF 

INSTALLIN + TRANSFER SFHT + MOVE TO LAUNCH + SHFT + 

PERFORMCITE VERIFY CITE STAND TO CANNISTER PAD RSS 

TESTS INTERFACES 

PHYSICAL 

INSPECTION MOVE TO 

FUNCTIONAL PHSFlSAEF 2 LOAD 

WITH SFHe TESTlCHECKFigure 2-36 

F8804112AA8-79 

t TRANS FER 

SHUTTLE 

TO LAUNCH PAD 

LAUNCH 

SUPPORT 

LAUNCH 

OPERATIONS 

I 

Preparation of SFHT for launch through the horisontal 

processing facility 

2-79 
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F8804112AA8-80 2-80 
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Section 3 

MISSION OPERATIONS 

3.1 HOST/USER INTERFACES AND OPTIONS 

The SFHT must interface with the following: 

e Launch vehicle 

e The user with the shuttle as host 

e The user with the SS as host 

e OW 

e GSE 

These interfaces require the mating of docking/berthing hardware and electrical and fluid connectors. With the exception of the docking/berthing interfaces, the initial connections of the fluid and electrical connectors will be 

accomplished by EVA. Eventually all interfaces are to connect automatically. 

The shuttle is to be the baseline launch vehicle and host for the initial 

SfHe resupply customers. The docking/berthing mechanism recommended for 

initial mission is the grapple and the SEE because of its successful use on 

the shuttle program to remove and handle payloads. It has been designed to 

remove 32,000-lb payloads from the cargo bay. Allowable grapple fixture 

loads are specified in NSTS 07700, Volume XIV, Appendix 8, Section 4.2. A 

special purpose end effector may be used to provide an electrical copgnection 

across the interface for control power and/or signals to payloads The 

electrical connector is located on the outside of the SEE at the SEE/payload 

interface. 

. 

It is recommended that the TPDM be used when considering the automatic connection of the electrical and fluid lines. Considerable force is required to 

make these connections, which can be accommodated by the TPDM. The TPDM has 

two 56 pin electrical connectors incorporated in the design that provide 

sufficient pins for the SFHT power, control, and monitor requirements for 

transfer of SfHe to a user. The fluid coupling will require additional design effort to incorporate into the TPDM configuraiion. Moog is currently 

under contract to NASA/JSC to design a SfHe coupling . This coupling will be 

used on the SHOOT program for the transfer of SfHe in space, and may be 

extended to robotic. 

3.1.1 Initial Capabilities 

Figure 3-1 depicts the initial concepts and EVA connections for servicing of 

users with the shuttle or the SS as host and shows the SFHT in storage at the 

SS. The electrical and fluid EVA connections that have to be made are shown. 

An electrical cable connected to the shuttle or to the SS provides power and 

the capability to monitor and control the activities of the SFHT. The power, 

monitor, and control hook-up to the shuttle, Figure 3-1(s), is assumed to use 

the shuttle umbilical retraction system (SURS) so that EVA is not required 

for disconnect of this cable when the SFHT is removed for transport to the 

SS. This requires a separate, parallel power, monitor and control cable 

F8804112AAS-1 3- 1 
F88-04 

@hf CABLE 

ORBITER 

(LOOKING 

FORWARD) 

A) RESUPPLY USER AT ORBITER 

SPACE STATION 

CONTROL 

CABLE 

SIRTFAJSER I B) RESUPPLY USER AT SPACE STATION 

SPACE STATION 

I 

ATTACHED 

AND 

CONTROL CABLE 5 sm~YLoAD 

C) RESUPPLY ATTACHED PAYLOAD 

(SM ATTACHED) 

SPACE STATION 

0) RESUPPLY AITACHED PAYLOAD 

(SFHT ALONGSIDE ARACHED PAYLOAD) 

LEGEND 0 GRAPPLE EVA ELECTRICAL CONNECTlON 

0 SEE EVA FLUID CONNECTION 

AUTOMATIC ELECTRICAL CONNECTION 

ku 9ooJ(u013s 020 

Figure 3-1 Initial capability - EVA required to connect avionics and fluid 

1' ines 

F8884112AA9-2 3-2 
F88-04 

SPACE STATION 

L 

Snir 

E) RESUPPLY LAB MODULE EXPERIMENTS 

(SFHT ALONGSIDE LAB MODULE) 

SPACE STATION 

L 

- 

/POWER. 

MONITOR 

SFHT CONTROL 

CABLE 

PROTECTIVE 4 .-. ENVELOPE l.ll--l-.... 

G) SRlT IN STORAGE AT SPACE STATION 

SPACE STATION 

L . 

CONTROL 

CABLE 

F) RESUPPLY LAB MODULE EXPERIMENTS 

(SFHT ATTACHED) 

LEGEND: 

0 GRAPPLE 

0 SEE 

EVA ELECTRICAL CONNECTION 

a EVA FLUID CONNECTION 

Figure 3-1 Initial capability - EVA required to connect avionics and fluid 

lines (concluded) 

F8804112AA9-3 3-3 
F88-04 

which will then be connected to the SS. The electrical connection made between the SFHT and the user, AP or LM, provides the capability to monitor and 

control the components of the user, AP or LM, during the fill operation. Two 

fluid couplers are provided on the SFHT for redundancy to connect to the 

user, AP or LM, for transfer of the SfHe. A grapple is shown on the port 

side (see Figure 2-3) so that it will be located on the same side as the 

shuttle RMS. The grapple is used to remove the SFHT from the shuttle cargo 

bay and to transport on the truss system of the SS using the mobile RMS of 

the SS. A grapple is also shown on the lower end of the SFHT, which is used 

to attach the SFHT to the SS truss system for storage and to locate near 

users for servicing. The options shown in Figure 3-1 for resupplying are as 

follows : 

a. 

b. 

c. 

d. SFHT docked to SS truss with electrical cables and fluid lines 

e. SFHT docked to SS truss with electrical cables and fluid lines 

f. 

g. 

SFHT services user while located in the shuttle cargo bay 

SFHT docked to SS truss and user docked to SFHT 

SFHT docked directly to AP such as ASTROMAG 

going to AP such as ASTROMAG 

going to LM 

SFHT docked directly to LM 

SFHT in storage at SS 

3.1.1.1 Resupply at Space Station 

The grapple at the top of the SFHT will be used to remove the SFHT from the 

orbiter and to transport it around the SS from storage to the user, AP and 

LM, with the SS mobile RMS. The electrical and fluid connections must be 

made by EVA. The required connections are power, monitor, and control cable 

connection to the SS SCS, two 52 pin electrical connectors to the user, and 

the fluid connection to the user. 

Two different configurations are shown for servicing the AP, Figure 3-1 (c 

and d) and the LM, Figure 3-1 (e and f). In the configurations shown in 

Figure 3-1 (c and f) the SFHT is mated directly to the AP or the LM. The 

configurations shown in Figure 3-1 (d and e) have the SFHT docked to the SS 

truss system next to the AP or the LM. The actual configuration to be used 

must be determined at a later date when the SS configuration becomes more 

solidified and the relative advantages and disadvantages of each approach can 

be evaluated. 

The SFHT is shown stored in Figure 3-1 (g) on the SS truss system in a protective envelope for thermal and meteoroid protection. When the SS servicing 

facility becomes available, the SFHT may be stored in that facility using the 

two sill and one keel trunnion fittings as the attachment method. 

3.1.2 Automatic Resupply Capability 

Figure 3-2 depicts the follow-on concept with automatic connections for servicing of users with the shuttle or the SS as host, and shows the SFHT in 

storage at SS. The electrical and fluid connections required and whether 

these connections are to be made by EVA or automatically are shown. The 

shuttle power, monitor, and control cable will use the SURS, as in the 

F8804112AA9-4 3-4 
A) RESUPPLY USER AT ORBITER 

SPACE STATION 

A 

AND 

B) RESUPPLY USER AT 

FW -

SPACE STATION 

F88-04 

SPACE STATION 

ASTROMAG 

AlTACHED 

CONTROL 

CABLE 

C) RESUPPLY ARACHED PAYLOAD 

SPACE STATION 

4 

AND ~ -1- I rp CONTROL 

CABLE 

sF&r - 

D) RESUPPLY AlTACHED PAYLOAD 

(OPTION) 

LEGEND 0 GRAPPLE 0 EVA ELECTRICAL CONNECTION 

a TPW LATCH PIN 

0 TPOY LATCH A EVA FLUID CONNECTION 

0 AUTO ELECTRICAL CONNECTION 

A AUTO FLUID CONNECTION 

Am~0135op 

Figure 3-2 Automatic connection of avionics and fluid lines . r-a 

2. 

F8804112AA9-5 3-5 
F88-04 

SPACE STATION 

f 

E) RESUPPLY LAB MODULE EXPERIMENTS 

(SFHT ALONGSIDE LA8 MODULE) 

SPAMONITOR 

AND 

CONTROL 

CABLE 

F) RESUPPLY LAB MODULE EXPERIMENTS 

(SM ATTACHED) 

SFHT 

AND 

CONTROL 

CABLE 

G) RESUPPLY OF USER ON-OABJT 

SPACE TATION 

CABLE ;# - PROTECTIVE 

ENVELOPE 

H) SFHT IN STORAGE AT SPACE STATION 

LEGEND 0 GRAPPLE 

TPDMIATWPIN 

0 TPDMLATW 

o EVA ELECT. CONNE~N 

a AUTO ELECT. CONNECTION 

A EVA FLUID CONNECTlON 

A AUTO FLUID CONNECTION 

5M = 196.85 In I 16.404 ft 

Figure 3-2 Automatic connection of avionics and fluid lines (concluded) 

F8804112AAQ-6 3-6 
F88-04 

initial concept. However, for automatic operation, the parallel cable may 

connect either to the TPDM or directly to the SS, as shown in Figure 3-2. A 

grapple is shown in the port side of the shuttle so that it will be located 

on the same side as the RMS. The TPDM latch pin is shown on the lower end of 

the SFHT which is used to attach the SFHT, to the SS truss system for storage 

and to locate near users for servicing. The options shown in Figure 3-2 for 

resupplying are as follows: 

a. 

b. 

d. 

C. 

e. 

f. 

g. 

h. 

SFHT services user while located in the shuttle cargo bay 

SFHT docked to SS truss and user docked to SFHT 

SFHT docked directly to attached payload such as ASTROMAG 

SFHT docked to SS truss with electrical cables and fluid lines 

connected by EVA to the attached payload such as ASTROMAG. An 

option is shown for connecting the control and power cable 

directly to SS or to SFHT 

SFHT docked to SS truss with electrical cables and fluid lines 

connected by EVA to the lab module 

SFHT docked directly to LM. An option is shown for connecting the 

control and power cable directly to SS or to SFHT. 

SFHT delivered to user orbit by OW. All operations are performed 

in the automatic mode 

Storage of SFHT on SS truss system. An option is shown for 

connecting the control and power cable directly to SS or to SFHT 

3.1.2.1 Resupply at the Orbiter 

The SFHT is attached to the orbiter in the cargo bay with two sill and one 

Power, monitor, and control electrical cable is 

connected to the orbiter using the SURS. This connection is made on the 

ground and no EVA will be required for servicing users at the orbiter. 

. keel trunnion fitting. 

The user will be docked to SFHT using the RMS. As the user is mated to SFHT, 

the two electrical connectors and the fluid connector will be automatically 

connected to the user. 

3.1.2.2 Resupply at Space Station 

The grapple at the top of the SFHT will be used to remove the SFHT from the 

orbiter and to transport it around the SS from storage to the user, AP and 

LY, with the SS mobile RMS. EVA will not be required to disconnect the 

power, monitor and control cable from the orbiter to connect it to the base 

of the TPDY (latch pin side). The electrical and fluid connections are made 

automatically to the user, AP or LM. 

Two different configurations are shown for servicing the AP, Figure 3-2 (c 

and d) and the LM, Figure 3-2 (e and f). In the configuration shown in 

Figure 3-2 (c and f) the SFHT is mated directly to the .AP or the LM. The 

configurations shown in Figure 3-2 (d and e) have the SFHT docked to the SS 

truss system next to the AP or the LM. The actual configuration to be used 

must be determined at a later date when the SS configuration becomes more 

solidified and the relative advantages and disadvantages of each approach can 

be evaluated. 

FEE041 12AA9-7 3-7 
F88-04 

The SFHT is shown stored on the SS truss system in a protective envelope in 

Figure 3-2 (h) for thermal and meteoroid protection. When the SS servicing 

facility becomes available the SFHT may be stored in that facility using the 

two sill and one keel trunnion fittings as the attachment method. For 

storage outside the servicing facility, where the SS heat rejection system 

may not be available, handling the 350 W heat load from the mechanical 

coolers will be more difficult. 

3.1.2.3 Resupply at User Orbit 

Users may be resupplied on-orbit with the SFHT based either at the orbiter or 

SS. The OW will be required to transfer the SFHT to the user orbit. All 

connections and the berthing/docking maneuvers must be made automatically. 

Figure 3-2 (g) displays the configuration required for the resupply on-orbit 

of the user. 

3.2 RESUPPLY OPERATIONS 

The step-by-step process of resupply is shown in Figure 3-3. It accounts for 

resupply based at either the shuttle or the SS and the user being either warm 

or cold. The user is not vented after liquid begins collecting in the user 

cryoge tank. Figure 3-5 shows a plumbing schematic for the AXAF XRS 

dewar, which is presented here as an example user system. User valve 

numbers referred to in Figure 3-3 are those shown in Figure 3-5. Notice that 

the XRS fluid management concept does not incorporate all the redundant 

valves required of SFHT to meet the requirement of fault tolerance for 

mission success. User fault tolerance is a topic that should be evaluated 

further. Figure 3-4 shows the operation timeline for resupply at the 

shuttle, where resupply must be accomplished within 7 days after launch (or 

3-4 weeks with the extended duration orbiter). The total operation duration 

shown in Figure 3-4 is 4.4 days for a warm SIRTF. 

fl 

The time required to pump down the SFHT bath in preparation for transfer is 

not shown in Figure 3-4. As shown in Figure 4-4, this operation will take 12 

hours, or less. It can, therefore, be performed during the time that the OW 

is retrieving the SIRTF. 

3.3 USER IMPACTS 

Interfacing between the SFHT and user has been discussed in 2.2, 3.1, and 

3.2. This section summarizes the user design impacts that specifically 

result from the on-orbit resupply. 

The following SFHT-to-user interface features are required: 

e Berthing - The recommended berthing concept is to use the SEE and 

grapple for the initial EVA configuration and the three point 

docking mechanism for the automated configuration. The user would 

therefore need to have either the grapple or the mating half of 

the TPDM. 

e - Fluid - The superfluid helium coupling being developed for the 

SHOOT experiment is baselined for the fluid connection. It would 

have to be somewhat modified for incorporation into the TPDM in 

F8804112AA9-8 3-8 
ORGINAL PAGE IS 

OF POOR QUALITY 

Remove SFHT from Cugo Bay 

Dock to OMV 

Go to SS Rmove SFHT from OMV 

Dock to SS 

F88-04 

I Launch 

& I 

I 

c 

On Orbit 1 

c 

Moasun Fluid Ouantities 

in SFHT and Uaer 

and Warm 

c c 

Op.n U.ar Vdwo I,& s and 7 Open U*N Valve. 1. 2. and 6 

ClOU Uaer Valw 10 

Low Raw (SO Uhr) I Power TMP at 

Low Rate I 

Temperature 

Open User Valw 5 Close Uur Valves 7, 10. 

Clou Valve 6 

Figure 3-3 SFHT resupply process flow chart 

3-9 

. 

F8804112AA9-9 
F88-04 

- 

Close User Valves 1 and 5 

Open 10 

+ 

Remove Power from User 

Measure Fluid Quantity 

in SFHT and User 

Increase TMP 

Flow to Optimum 

+ 

Open SFHT Valves 7, 9, 

and 17 

Le User Full 

Prepare SFHT to Wait 

for Next Transfer 

v 

Turn Off TMP 

Undock from SS 

Mate with OMV 

Go to STS 

Demate from OMV 

Stow SFHT in Cargo Bay 

I 

3. 7, 9. 15, and 17 

Descend 1 

I Power TMP I at Low Level 

Yes 

6 

I Power Down SFHT 1 

AiN 9344MD14Z11 

Figure 3-3 SFHT resupply process flow chart (concluded) 

F8804112AA9-10 3-10 
F88-04 

SlRTF 

Electrical 

SIRTF 

Mounted in 

Seprrated Shuttle 

OMV RMS 

Crpturos Sunshrdo 

2 OMV 1 OMV 

4 SiRTF -+ Covor -+ 

Interface 

Structure 

rnd 

Stowed rnd A ttrchod 

OMV .I - 1 

24 OMV 0.5 

Connection - 

Made 

I OMV 0.5 EVA 0.5 OMV 0.5 0.5 EVA 0.5 

- .. 

Initial 

Transfer 

Cooldown 

Linos 

Conflgurrtion Attrch brk Transfer Chock 4 for -+ Transfor -+ Transfer 

Llnos Ho Transfor Llnos 

. , 

1 RIA 1 EVA 1 l(30) 5 

-+ -+ 

rnd 

SlRTF 

at 

1000 UHr 

- _i . 

O(12) O(2) EVA 1 1 OMV 1 

7 

Fin81 In-Bay 

Chockout Attach Dowar . Topoff Romove 

-a Thormrl -+ SiRTF -0 Trrnrfer - 

Strbillzitlon Dewrr Lines 

, 

VJ r OMV 

Tolomotry 

i 

Legend 

EVA 0.5 OMV 1.5 OMV 16 OYV 1 1 .s . OMVlSiRTF OMV Prepare for 

Dopioy Boost Rendezvous Deorbit and Romovo 

4 Sunrhrdo -+ 

A 

I AM 920YMD135.064 

A: Roaourco Roquirod 

E: Hours Roqulrod with SlRTF Cryogm Tank 81 2K 

C: Hours Roqulrd with SlRlT Cryogen Tank rt lSOK 

Actlvity 

Figure 3-4 Operations for shuttle-based resupply 

OMVlSiRTF * to with 

Orbit Shuttlo Covor 

L I 1 

F8804112AAQ-11 

Return to 

Earth 

3-11 
F88-04 

+ 

@ 

3- 

f 

PP 

PS 

0 

EV 

I 

-IF 

Relief Valve 

0.5 in. Motor Valve 

Commercial Bayonet 

Burst Disk 

Porous Plug 

Phase Separator 

Pressure Transducer 

Emergency Vent 

EL Connector 

EL Feedthrough 

EVA Coupling 

High 

Normal Flow 

Vent Vent Fill 

EC @ 

142.12 

F8804112AAQ-12 

Figure 3-5 AUF XRS fluid management schematic 

3-12 
. 

F88-04 

the automatic configuration. The baseline has two of these 

coupling for redundancy; this is because a single failure in the 

coupling could prevent mating and mission completion. This 

requirement and the coupling failure modes should be examined 

further to determine if two of these costly couplings, or if two 

of the user half, are absolutely required. 

0 Electrical - Command, data, and power hookup to the user requires 

two 56-pin connectors. These are existing connectors for the 

initial EVA approach or connectors mounted in the TPDM for the 

automatic approach. The SFHT avionics provides power to the user 

for operation of its computer, valves, and sensors. 

The SFHT must be able to command all user valves except the phase separator 

bypass valves. It is probably just as easy to plan on giving the SFHT control of all user valves; a reason for operating the bypass valves may be 

determined later. The SFHT must be able to monitor user thermometers at 

approximately four locations: 1) inside the tank, 2) the downstream side of 

the low-flowrate phase separator, 3) the downstream side of the high-flowrate 

phase separator, and 4) the inner VCS. It is assumed that the user is responsible for monitoring the sensors needed to verify the health of its system and issue alerts of abnormal conditions to the SFHT. 

It is assumed that the three-string majority vote feature of the avionics 

extends into the user and that the user participates in the voting process 

for command of its valves. This approach is not absolutely required, but it 

does provide for more flexible contingency action if abnormal conditions 

occur in the user. 

There are two potential impacts to the user fluid management system. First, 

a high conductance vent line must exist for removal of flashoff gas without 

excessive back pressure during cobldown. This is an insignificant impact if 

the user already has a high conductance line as part of its emergency vent 

system. If the user is vented through this line during the fill (to minimize 

the post-transf er bath temperature), then' a high-f lowrate phase separator is 

required. Second, a valve is required on the upstream side of the standby 

phase separator. Our baselined procedure is to close the user tank when it 

approaches full (or throughout the transfer process). Thus, when the user 

reaches full, it can easily be detected by a sharp pressure rise in the 

transfer system. In the absence of the aforementioned valve, SfHe would be 

forced through the porous plug phase separator. The large amount of heat 

conducted to the porous plug through the downstream column of SfHe would then 

prevent establishing the temperature drop through the plug, which is needed 

for it to act as a phase separator. SfHe would continue to flow unconstrained through the plug with no means of recovery. This is the "Castles' catastrophe." The only failsafe method of preventing this situation is to place a 

valve upstream of the plug. 

3.4 SAFETY 

The SFHT design and operations must comply with the safety requirements of 

NEB 1700.7A. The underlying requirement is that no credible SFHT failure 

shall result in personnel injury or damage to any part of the STS or other 

payloads. Four hazards are identified that require specific safety control 

F8804112AAS-13 3-13 
F88-04 

measures: 1) overpressurization of the cryogen tank, 2) overpressurization 

of the vacuum shell, 3) overpressurization of plumbing, and 4) emergency 

jettison. 

3.4.1 Cryogen Tank Pressure Relief 

Cryogen tank overpressurization could occur if the vent system cannot remove 

fluid from the tank fast enough to keep up with the heat input to the tank. 

A failure resulting in rapid leakage of helium or air into the guard vacuum 

is the circumstance during which this becomes an issue. If helium floods the 

guard vacuum, the heat input to the tank is estimated at 30,000 times greater 

than during normal operation. This estimate is based on the assumption of 

continuum gaseous conduction, which will occur if the gas pressure is above 

about 0.001 torr. Higher pressure does not result in increased heat conduction. This heating rate is equivalent to the condensation of air leaking 

through a 1 cm diameter hole in the vacuum shell. Occurrence of a hole 

larger than this is not considered a credible failure, and flooding of the 

guard vacuum with helium is therefore considered the driving case for design 

of the cryogen tank emergency vent system. 

The SFHT fluid management system provides redundant vent paths for preventing 

cryogen tank overpressurization in the case of a guard vacuum failure. (See 

2.1.3.) This provides two-fault tolerance, the first fault being the guard 

vacuum failure and the second being loss of one of the vent paths. Safety is 

thus provided by totally passive design features (burst discs) and does not 

rely in any way on the avionics. Analysis of the emergency vent system is 

discussed in 4.3.3. 

3.4.2 Vacuum Shell Pressure Relief 

The vacuum shell is equipped with a 3 psi relief device like that used in the 

IRAS and COBE dewars. This is a 3-in. diameter piston in a cylinder with two 

"0" rings for sealing; it also doubles as the vacuum pumpout port. It is 

available off-the-shelf from Cryolab. The internal pressure capability of 

the vacuum shell is several atmospheres, providing considerable margin. 

3.4.3 Plumbing Pressure Relief 

Any section of plumbing in which cold fluid could be trapped 

ed by a relief device to prevent overpressurization as the 

This is done with off-the-shelf relief valves where possible 

fied burst discs where the risk of relief valve back leakage 

unacceptable. (See 2.3.1.) 

must be protectfluid warms up. 

and with qualior freeze out is 

A pyrotechnically actuated guillotine is provided to severe the fluid and 

electrical lines to the user so that jettison of the user can be accomplished 

without EVA. Redundant guillotines and firing circuits are provided. Firing 

of the umbilical cutter is commanded from the GPC and requires the SFHT avionics to be in the normal operating mode. This feature is probably not required if the SFHT is designed for fully automatic mating/demating with the 

user. 

F8804112AAQ-14 3-14 
F88-04 

No EVA is required in the event the SFHT dewar must be emergency jettisoned 

from the shuttle. Electrical hookup from the dewar-based avionics to the 

shuttle uses the shuttle umbilical retraction system for automatic 

disconnection. 

F8804112AA9-15 3-15 
Several types of analyses have 

trades and concept development 

baseline design. These include 

fluid transfer analyses. 

4.1 STRUCTURAL ANALYSIS 

F88-04 

Section 4 

ANALYSIS 

been performed in support of configuration 

and to characterize the performance of the 

structural, thermal/cryogenic, venting, and 

Calculations have been made to determine thicknesses and masses of the cryogen tank, vacuum shell, and vapor-cooled shields and to size the cryogen tank 

support straps. The assumed operating conditions and structural design criteria are given in Table 4-1. 

4.1.1 Sizing of Shells 

The vacuum shell is sized for 1 atm external pressure and launch loads. It 

consists of two shells welded to a girth ring, which also supports the cryogen tank, insulation system, and plumbing. The required shell thickness is 

0.65 cm for the spherical shape and 0.98 cm for the 42 oblate spheroid 

shape. 

The cryogen tank also must withstand 1 atm of external pressure for the inprocess leak test. It must also be designed for the maximum operating pressure (internal) and for launch loads with essentially no pressure. The external and internal pressure conditions drive the design about equally; buckling strength is usually the primary driver. The required wall thickness is 

0.55 cm for the spherical tank and 0.92 cm for the 42 oblate spheroid tank. 

The cryogen tank could be lightened somewhat if a smaller factor of safety is 

allowed for the leak check condition. This was done for the IUS and COBE 

tanks. 

The thickness of the vapor-cooled shields is estimated at 0.10 cm. 

provides adequate buckling strength for launch loads. 

This 

4.1.2 Sizing of Cryogen Tank Support Straps 

The support straps are sized by fatigue; 100 cycles of limit load per flight 

is assumed. Considering the relatively large mass of the SFET and results of 

the coupled loads analysis for systems of similar size, such as the COBE, 

this assumption is realistic. With 50 missions and a fatigue factor of safety of four, the total fatigue life requirement is 20,000 cycles of limit 

load. Measured fatigue strength of straps the size of those used in the COBE 

dewar is shown in Figure 4-1. Straps made of fiberglass/epoxy (i.e., the 

COBE straps) and alumina/epoxy were tested by the National Bureau of Standards in Boulder, Colorado, as part of a BASG internal research and development program. The figure shows that for 20,000 cycles, the fiberglass/epoxy straps have a fatigue strength of 65 ksi, while the alumina/epoxy 

straps provide 100 ksi. These values were used for the strap sizings. 

For a given dewar geometry, strap angles are arranged to provide minimum 

strap area. The strap-supported 

mass is estimated at 2,070 kg for the 10,000-e spherical 

The SFHT dewar is assumed to use 12 straps. 

F8804112AB9-1 4- 1 
F88-04 

Table 4-1 

STRUCTURAL DESIGN CONDITIONS AND CRITERIA 

Operating Conditions 

- Vacuum Shell 

-- 1 ATM External Pressure with Launch Loads 

- Cryogen Tank 

-- 1 ATM External Pressure (During Leak Check) 

-- 54 psi Maximum Operating Pressure (Internal) 

-- Launch Loads with No Pressure 

- Vapor-cooled Shields 

-- Launch Loads with No Pressure 

-- Thermal Contraction Effects 

Design Criteria 

- log Limit Load (STS or Titan IV Launch) 

- Yield Factor of Safety: 

-- 1.6 (Internal Pressure) 

-- 1.4 (Other) 

- Ultimate Factor of Safety: 

-- 2.0 (Internal Pressure) 

-- 1.4 (Other) 

. 

F8804112AB9-2 4-2 
F88-04 

140 

100 

60 

20 

A FIBERGLASS/EPOXY 

0 ALUMINA/EPOXY - 

1 o2 1 o3 1 o4 1 os 1 o6 I o7 

CYCLES TO FAILURE A/N 92031MD135.063 

Figure 4-1 Strap fatigue test results at P=O.OS and room temperature 

F8804112ABQ-3 4-3 
. F88-04 

tank and 2,370 kg for the oblate spheroid. The calculated strap size is 

approximately the same for both the sphere and the oblate spheroid. The 

greater supported mass of the oblate spheroid is compensated for by a mors 

favorable strap angle. 

for alumina/epoxy and 25.5 cm for f iberglass/epoxy . 

The measured tensile mgdulus of elasticity is 8.7~10' psi for the fiberglass/epoxy and 19.7~10 Based on an estimated strap 

length of 100 cm and the strap-supported masses, the natural frequencies for 

the various combinations of strap material and dewar shape vary from 23 to 30 

Hz. These are acceptable levels. 

The cross-sectional area of all 12 straps is 16.0 cm 

for the alumina/epoxy. 

4.2 THERMAL ANALYSIS 

The thermal analysis indicates that a hybrid system with four Oxford Sterling 

cycle refrigerators offers a significant mass savings and extended dewar 

lock-up capability while meeting the thermal performance requirement of delivering 8,900 & of SfHe at the end of a 270-day on-orbitlgtandby. This 

amount of SfHe is required to cool and fill a 150 K SIRTF. The 270-day 

period represents a 180-day servicing interval with a 50 percent margin 

added. The objective of this analysis was to define a thermally optimum 

dewar subsystem while meeting cost and payload constraints. 

4.2.1 Model Description 

The SFHT dewar thermal anaiysis was conducted using a SINDA computer model 

(Figure 4-2). The model uses single nodes to represent the CT, IVCS, OVCS, 

and the VS. Heat transfer paths include conduc$ion through the supports and 

&I, and radiation through the MLI. Also incorporated in the model are secondary effects including heat transfer through the plumbing and wiring and 

parasitics at the refrigerator thermal switch. The ground rules for this 

modeling effort were: 

8,900 & of SfHe at 1.6 K are available for a warm SIRTF resupply 

at the end of a 270-day on-orbit standby period. 

The baseline design includes four Oxford Stirling cycle refrigerators providing OVCS cooling, each with a capacity of 3 W at 93 K. 

The nominal conductance of both the refrigerator cold-finger and 

thermal switch is 500 mW/K. 

mal switch is 500 mW. 

The parasitic heat leak at the therThe nominal on-orbit and ground hold vacuum shell temperature is 

300 K. 

Dewar exit wiring consists of 130 40-gage and 102 28-gage manganin 

wires exiting along a support strap. 

is inherently vapor cooled. 

Wiring inside the vent line 

12 support straps are of alumina/epoxy construction. 

Vent and fill lines are 3/4-in. diameter, 0.016-in. wall thickness, stainless steel. 

F8804112ABQ-4 4-4 
MC 

COOLING 

F88-04 

f 

PARASITIC 

THERMAL 

SWITCH 

CONDUCTION 

0- - - - -0 RADIATION 

F8804112ABQ-6 

Figure 4-2 Schematic of thermal math model 

4-5 

I vs 

I 

ovcs 

IVCS 

Lo CT 

A/N 9203lMD135.044 
F88-04 

0 A 5 percent ullage exists when the dewar is locked up for prelaunch hold. 

The model was designed to run either a steady-state or transient case. 

Steady-state runs were used to determine on-orbit standby losses. Transient 

runs were used to determine the dewar lock-up capability during prelaunch 

ground hold. 

The model incorporates refrigerator cooling capacity values as a function of 

OVCS temperature. Integrated values of the alumina/epoxy support thermal 

conductivity and temperature dependent values of helium vent gas enthalpy 

were also included in the modal. Heat transfer through the MLI was calculated using conduction and radiation correlations validated by test data from 

the IRAS and COBE programs. 

4.2.2 Baseline System Performance 

System performance was calculated for the dewar with the mechanical coolers 

(SS-based resupply) and without the mechanical coolers (shuttle-based resupply). The heat load breakdown for the two cases based on a 300 K vacuum 

shell temperature is as follows: 

WITH COOLERS WITHOUT COOLERS 

Total Heat Leak, mW 62 163 

Support Straps, % 

m1, I 

Plumbing, X 

Dewar Exit Cables, I 

52 

32 

14 

2 

43 

43 

12 

2 

Boil-Off Rate, g/day 238 630 

Bath temperature rise during the prelaunch lockup is shown in Figure 4-3. 

Time to reach the lambda point, temperature after the required 18-day lockup, 

and mass loss after orbital pumpdown may be seen for the cooler and no-cooler 

configuration. Figure 4-4 shows orbital pumpdown time based on pumping 

through the high-flow vent. For the predicted fluid exit temperature of 

2.05 K, pumpdown time is 12 hours without use of coolers during ground lockup 

and 4 hours with coolers used during ground lockup. Pumpdown would be at 

least 10 times slower if it were done through the low-thrust vent alone. 

Pumpdown time is critical for only shuttle-based resupply, in which case 

coolers are not used. For SS-based resupply the major advantage of using the 

coolers during lockup is not shortened pumpdown time, but rather a system 

launch mass reduction of about 75 kg. 

Table 4-2 gives the breakdown of helium usage throughout the mission for the 

two cases. It is assumed that the transfer rate is throttled back near the 

end of the resupply operation so that a 5 percent residual is achievable. 

4.2.3 Sensitivity Studies 

Several cases have been considered concerning uncertainties in vacuum shell 

temperature, the effect of larger support straps, and a 360-day standby period. The results are presented in Table 4-3 as a dewar mass penalty. The 

F8804112ABQ-6 4-6 
2.2 

2.1 

n 

s 2.0 

W 

K 

3 

c 

a 

K 

W 1.9 

n 

2. 

w 

c 

I 

1.8 

m 

1.7 

1.6 

LAMBDA 

WITHOUT COOLERS 

(10,000 LITERS) 

WITH COOLERS 

(1 0,000 LITERS) 

4% ORBITAL PUMPDOWN 

1.5% ORBITAL PUMPDOWN LOSS 

18 DAYS REQUIRED 

I I I I I I I 1 I I I 

VACUUM SHELL 

TEMPERATURE = 300 K 

0 20 40 60 80 100 120 

CUMULATIVE LOCK-UP PERIOD (DAYS) 

F8884112ABQ-7 

AiN 9203lMD135.060 

Figure 4-3 Ground lock-up capability 

4-7 
F88-04 

n 

U 

U 

0 

t 

<n 

W 

U 

P 

Y 

z 

F 

a 

28 

26 

24 

22 

20 

18 

16 

14 

12 

10 

8 

6 

HIGH-FLOW VENT OPEN 

T = FLUID EXIT TEMPERATURE 

-1.7 K 

y1.6 K 

T = 2.05 K 

(PREDICTED) \ 

4 I 

0 4 8 12 16 20 24 

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

TIME (HOURS) 

A/N 9203iMD135.058 

F8804 11 2AB9-8 

Figure 4-4 Orbital pumpdown time 

4-a 
' * 

F88-04 

Table 4-2 

BREAKDOWN OF HELIUM USAGE 

Final Ground Pumpdown 

Orbital Pumpdown to 1.6 K 

Orbital Standby 

Transfer (Warm SIRTF) 

Residual (5%) 

Total 

WITH 

co OLERS * 

500 

150 

440 

(9 Months) 

8,400 

500 

WITHOUT 

COOLERS* * 

500 

400 

30 

(1 Week) 

8,400 

500 

9,990 9,830 Liters 

*Resupply at Space Station 

** Resupply at Shuttle 

differences are slight since nearly 90 percent of the total mass can be attributed to the 8,900 liter resupply requirement which is not subject to 

standby losses. 

Also, the failure of mechanical coolers on orbit and the resulting reduction 

in standby lifetime is presented in Table 4-4. The standby lifetime presented represents the maximum period of time SFHT can remain in the standby mode 

with an adequate helium volume available for a warm SIRTF resupply. 

4.2.4 Configuration Tradeoffs 

Three dewar design tradeoffs have been assessed including a 42 oblate spheroid shaped dewar with four Oxford coolers, a spherical dewar with no coolers, 

and the substitution of fiberglass support straps for alumina/epoxy straps on 

the baseline dewar. Table 4-6 shows the mass penalty associated with each 

alternative for SS-based resupply. 

The oblate spheroid dewar with four Oxford coolers has nearly the same thermal performance as the baseline dewar with a 118 day ground lock-up capability and a standby boiloff rate of 243 g or 0.016 percent per day. Therefore, 

nearly all of the 720 kg mass penalty associated with this dewar configuration can be attributed to the dry weight of the dewar. 

The spherical dewar without the mechanical coolers yields a 26 day lock-up 

period, before the helium bath goes normal, or only 22 percent of the capability of the baseline system. The heat leak to the cryogen tank during 

orbital standby is nearly 2.4 times greater than the baseline design. The 

F8804112AB9-9 4-9 
1 

F88-04 

EXAMINED 

PARAMETER 

Base1 ine 

360 Day Standby 

(No Margin) 

Orbital Vacuum Shell 

Temperature (K) 

320 

310 

290 

280 

Table 4-3 

SENSITIVITY ANALYSES 

CRYOGEN TANK DEWAR MASS X OVER 

VOLUME (L) PENALTY (kg) BASELINE 

10,Ooo 

10,150 

10,080 

10,030 

9,950 

9,900 

Support Strap Size 

1.25 X Baseline 10,070 

1.50 X Baseline 10,170 

27 

12 

-11 

-24 

24 

50 

--- 

1.5 

0.9 

0.4 

-0.3 

-0.8 

0.8 

1.6 

Table 4-4 

EFFECT OF MECHANICAL COOLER FAILURE ON DEWAR PERFORMANCE 

0 Reduction in Standby Lifetime Calculated - 

- 

8,900 Liters Required at End of Standby 

Complete Failure of 1,2,3 or 4 Coolers at Start of Standby 

Considered 

STANDBY LIFETIME 

NUMBER OF LIFETIME REDUCTION 

COOLERS FAILED (DAYS) m 

270 

250 

180 

125 

100 

- 

7 

33 

54 

63 

F8804112AB9-10 4-10 
F88-04 

Table 4-5 

MASS PENALTIES ASSOCIATED WITH CONFIGURATION TRADEOFFS 

DEWAR DEWAR MASS X OVER 

CONFIGURATION VOLUME (11 PENALTY (ke;l BASELINE 

--- --- Baseline 10,Ooo 

Oblate Spheroid 10,Ooo 720 23 

Sphere/No Coolers 10,800 

Fiberglass Straps 10,200 

250 

60 

8 

2 

8 percent mass penalty is a result of the larger helium and tank volume necessary to meet the 270-day standby requirement, less the weight of the mechanical coolers. 

Replacement of the alumina/epoxy support straps with fiberglass straps results in a total mass penalty of 2.4 percent. This entire mass penalty is in 

the form of excess helium necessary to offset the degradation of the lock-up 

and standby thermal performance resulting from the necessarily larger fiberglass support straps. The alumina/epoxy material provides 40 percent higher 

frequency and 10 percent lower thermal conductance than f iberglass/epoxy for 

the SFHT. 

4.3 VENTING ANALYSIS 

The SFHT venting subsystem is composed of three separate venting paths: the 

low-flow vent to provide venting in the standby mode, the high-flow vent to 

provide initial orbital pumpdown and venting throughout the resupply mode, 

and the emergency vent to relieve uncontrolled tank pressurization due to 

loss of guard vacuum. Both the low-flow and high-flow vents function by 

removing heat from the cryogen tank in the form of vapor. The design goal 

for both of these vent systems is to allow adequate flow conductance to maintain the lowest feasible bath temperature without significantly impacting the 

thermal performance of the dewar. The emergency vent is a straight exit path 

specifically designed to maintain tank pressure below yield for a mishap 

involving guard vacuum loss as a result of helium or air intrusion into the 

annulus. 

4.3.1 Low-Flow (Standby) Vent 

The low-flow vent (Figure 4-6) is opened subsequent to launch by a baro 

switch and remains open until descent. The venting system incorporates a 

passive phase separator, or porous plug, with roughly four times the surface 

area of the plug developed for COBE to remove parasitic heat leaks to the 

cryogen tank during standby. This plug is designed to the higher vapor flowrates associated with a shuttle-based resupply mission. For the SFHT baselined 3/4-in. low-flow vent system, analysis indicates a bath temperature of 

1.6 K is attainable with a cryogen tank heat input of 163 mW for a shuttlebased resupply. A bath temperature of 1.5 K is predicted for the lower 

standby venting rates associated with space station-based resupply with refrigerators installed. 

F8804112ABQ-11 4-11 
STANDBY 

LTVl 

BC - Commercial Bayonet Coupling BD - Burst DISC 

EC - Electrical Connector 

EF - Electrical Feedthru 

EVAC - EVA Coupling 

HFPP - Hlgh-Flow Porous Plug IVCS - Inner Vapor-cooled Shield 

LAD - Liquid Acqulstlon Device 

LFPP - Low-Flow Porous Plug OVCS - Outer Vapor-cooled Shield @ - Motor-operated Value 

RD - Relief Device/Pump-out Port TM - Thermomechanical 

PV - Purge Valve (Manual) 

VACUUM SHELL 

F88-04 

EMERGENCY HIGH-FLOW 

VENT 

- LTVI 

VENT LTVI IIv 

DCI (20 p.1) BC4 

- 

T 

IkW vCRYOGEN TANK 

F8804112AB9-12 

Figure 4-6 Standby (low-flow) vent 

4-12 
F88-04 

The low-flow vent plumbing system includes nearly 19 ft of line length internal to the vacuum shell and 25 ft total. A strong thermal attachment at both 

the IVCS and OVCS provides 100 percent vapor cooling efficiency. External to 

the vacuum shell are redundant flow paths each with an in-line 3/4-in. valve 

making the system 1 fault tolerant to mission success. Finally, the vapor 

exits the system through a low-thrust vent. 

The analysis was conducted using the continuum flow equation to predict sectional pressure drops in the line based up?? fluid temperatures equivalent to 

the CT, IVCS, OVCS, and VS temperatures. A 1/4-in. wire bundle entering 

the line at the vacuum shell and extending into the cryogen tank is accounted 

for in the analysis. 

The predicted bath operating temperature in the standby mode is 1.6 K for the 

system without coolers and 1.5 K for the system with coolers incorporated. 

The majority of the flow impedance can be attributed to line length, with 

valves and bends making an insignificant contribution. This is a result of 

the laminar flow condition where viscosity effects are dominant. A venting 

mode with both valves open offers little advantage with respect to flow 

conductance. 

4.3.2 High-Flow Vent 

The high-flow vent (Figure 4-6) is opened just prior to resupply and remains 

open throughout thz resupply mode. The venting system includes a large POrous plug (200 cm , 10 Fm) to remove a greatly escalated heat input to the 

cryogen tank during transfer while maintaining small pressure drops across 

the porous medium. This is the same plug as baselined for the SHOOT experiment. The high-flow vent system has been designed to extract large volumes 

of cold gaseous helium from the cryogen tank while sustaining the bath at the 

baselined temperature of 1.6 K. 

The high-flow vent plumbing is comprised of nearly 6 ft of line internal of 

the vacuum shell exiting in a nearly straight line. There is a relatively 

weak thermal attachment at the IVCS. An attachment clip with a small conductance is used to intercept heat conducted toward the cryogen tank during 

standby while allowing very little temperature rise in the effluent gas during transfer to minimize flow impedance. The remainder of the line is isolated from any heat sources except a 1 W heat input at the external valves. 

There is a redundant flow path associated with an internal set of valves and 

a separate redundant flow path associated with two pairs of external valves. 

A low thrust vent at the outlet completes the system. 

The high-flow vent analysis was conducted using the Darcy equation for turbulent flow to predict sectional pressure drops in the line. The design is 

based upon all valves being open to flow. The vapor exits the tank slightly 

above the bath temperature, with heating at the IVCS and external valve set 

only. 

In order to maintain a tank temperature of 1.6 K during resupply, the flow 

conductance through the baselined 3/4-in. high-f low vent must be minimized by 

. 

F8804112ABQ-13 4-13 
F88-04 

EVAE 1 WAC 2 

STANDBY VENT E 

I 

LTVI 

IID L 

(5 Pd 

BC - Commercial Bayonet Coupling BD -Burst Disc 

EC - Electrical Connector 

EF - Electrical Feedthru 

EVAC - EVA Coupling 

HFPP - High-Flow Porous Plug 

IVCS - Inner Vapor-cooled Shield 

LAD - Liquid Acquistion Device 

LFPP - Low-Flow Porous Plug OVCS - Outer Vapor-cooled Shield @ - Motor-operated Value 

RD - Relief Device/Pump-out Port 

TM - Thermomechanicai 

PV - Purge Valve (Manual) 

VACUUM SHELL 

VI 

0 

CRYOGEN TANU 

Figure 4-6 Orbital pumpdown (high-flow vent) 

F8814112ABQ-14 4-14 
F88-04 

restricting the exit fluid temperature to 2.05 K or less. This is accomplished by design of the vent line thermal attachment at the IVCS. An attachment clip with a nominal conductance of 25 mW/K achieves the desired result without significant impact to the standby thermal performance. For this 

high-flow situation, the flow impedance associated with valves and fittings 

dominates the overall system. 

4.3.3 Emergency Vent 

The emergency vent (Figure 4-7) is used only during a situation of uncontrolled tank pressurization due to loss of guard vacuum and is specifically 

designed for the scenario of helium gas leakage into the annulus. For such 

an event, it is necessary to prevent the tank from pressurizing beyond yield 

(68 psi) and sustaining irreparable structural damage. The SFHT baselined 

3/4-in. emergency vent system is capable of such a task. Heat input to the 

cryogen due to helium leakage is equivalent to air leakage through a hole of 

about 1-cm diameter; this size of hole is not considered a credible failure. 

Two completely redundant flow paths both with two in-line burst discs set at 

30 psi on the warm and cold end define the emergency vent system. One of the 

lines doubles as the high flow vent line; the other is a dedicated emergency 

vent line. Each line has a total path length of nearly 17 ft. Both lines 

have a weak thermal connection to the IVCS. The same attachment clip designed for the high-flow vent is used to minimize exit fluid temperature and 

flow impedance. At the outlet of the line there is a low thrust vent. 

The analysis was conducted using the Darcy equation for turbulent flow. 

Three18cases were considered to study the effect of vapor cooling at the 

VCSs. The first case considers no effective vapor cooling, resulting in a 

maximum cryogen tank heat leak. For this situation, the fluid exits at the 

bath temperature. The second case accounts for a 100 percent heat transfer 

efficiency at the vapor-cooled shield;, with a greatly reduced heat leak to 

the cryogen tank. The exit fluid temperature is taken sectionally as the CT, 

IVCS, and OVCS temperatures predicted by a SINDA thermal model. The final 

case accounts for the heat transfer effects at the IVCS with the designed 

clip in place. A SINDA thermal model was used to account for the attachment 

clip conductance at the IYCS. The clip modeled is identical to the clip used 

in the high-flow venting system. 

From these three models, it has been determined that the worst case with 

respect to.pressure drop across the emergency vent line exists when the heat 

leak to the tank is a maximum and heat transfer to the fluid at the VCSs is 

negligible. From the third case it can be concluded that this is very near 

the true situation. However, the important conclusion is that the baselined 

3/4-in. emergency vent system will provide adequate flow conductance to maintain a below-yield tank pressure in an emergency venting situation. 

4.4 SUPERFLUID HELIUM TRANSFER ANALYSIS 

A computer mode12bhat incorporates the physics involved in the transfer of 

superfluid helium has been applied to verify the baselined transfer system 

design. The governing equations are shown in Table 2-4; it is the GorterMellink equation and the internal heat flux that differentiate this set of 

F8804112ABQ-15 4-15 
F88-04 

STANDBY 

LTVI 

E VENT - 

BC - Commercial Bayonet Coupling BD -Burst Disc 

EC - Electrical Connector 

EF - Electrical Feedthru 

EVAC - EVA Coupling 

HFPP - High-Flow Porous Plug IVCS - Inner Vapor-cooled Shield 

LAD - Liquid Acquistion Device 

LFPP - Low-Flow Porous Plug OVCS - Outer Vapor-cooled Shield @ - Motor-operated Value 

RD - Relief DevicelPump-out Port 

TM - Thermomechanical 

PV - Purge Valve (Manual) 

VACUUM SHELL 

F8804112A89-16 

t 

FLEXIBLE 

TRANSFER 

LIME 

EMERGENCY HIGH-FLOW 

VENT VENT 

J 

CRYOGEN TANK 

Figure 4-7 Emergency vent 

4-16 
F88-04 

equations from that governing the behavior of normal fluids. The GorterMellink equation expresses the relationship among the temperature gradient, 

the pressure gradient, and the mutual friction. The result of this 

relationship is a heat flux going upstream in the SfHe flow. At very high 

flow velocities, mutual friction couples the normal and superfluid components 

together, and the flow behavior approaches that of a normal fluid. At the 

low flowrates that will exist during start-up of the SFHT transfer process 

this is not the case, and it is important to use a model that properly 

accounts for the unique behavior of SfHe. 

The transfer system analyzed includes the TM pump, the SFHT plumbing, a flexible transfer line, and the user plumbing. The model accounts for flow constrictions due to valves and bends and for heat inputs along the line. The 

thermodynamic properties of SfHe are contained in the program. The line 

configuration is input as a number of elements. Each element specifies a 

portion of the line through inputs of length, diameter, and heating. Equivalent lengths are used in the traditional manner to account for constrictions 

and changes of direction. These configurational and heating data, the saturation temperature of the source tank, the heat flux at the receiver tank, and 

the porous plug properties define the transfer system. The model uses a 

finite difference approximation to predict temperature, pressure, velocity, 

and heat flux along the line length. The model has been validated by correlation with the results of SfHe transfer testing performed by BASG. 

4.4.1 Modeling Ground Rules 

The ground rules used for modeling the SFHT transfer system are: 

e The supply dewar operates at a temperature of 1.6 K. 

e Transfer takes place into a non-venting receiver. The only heat 

transfgr taking place at the receiver tank is due to condensation 

of the gas within the tank as the ullage shrinks. 

0 The desired maximum transfer rate is 1,OOO &/hr. 

e The user transfer system configuration and heat inputs are identical to those of the SFHT, except for the TM pump. 

4.4.2 Baseline System Description 

Figure 4-8 is a plumbing schematic of the transfer system; major features are 

internal and external 3/4 in. valves in both the SFHT and user, a bayonet 

coupling, an EVA coupling, and a flexible transfer line. The inside diameter 

of the SFHT plumbing, user plumbing, and transfer line are assumed to be 

1.8 cm. The total length of the flow path is just under 11 m; Figure 4-9 

shows the diameter profile over the flow path. Flow impedances of the 

valves, couplings, and corrugated flexible line are represented in the model 

by equivalent lengths of 1.8 cm pipe. 

Based on the measured valve flow coefficient of 5.3", an equivalent length 

of 1.8 cm tubing was calculated for input to the model. The flexible transfer line is assumed to be 3 m long with the minor diameter of the inner corrugated line being 1.8 cm. To account for the increased flow impedance due 

F8804112A89-17 4-17 
F88-04 

10 FEET LONG 

TRANSFER LINE 

_----. -- 

EVA 

COUPLING 

DISC 

VACUUM SHELL 

BURST 

EXTERNAL 

VALVE SET 

I 

I,, I TM PUMP 

1 

SFHT CRYOGEN TANK 

(1.6 K) 

RELIEF VALVE I-+ 

EXTERNAL 

VALVE SET 

- - - - - IVCS 

INTERNAL 

VALVE SET 

USER CRYOGEN TANK NN 9203lMD135.43 

Figure 4-8 Transfer system plumbing schematic 

4-18 
F88-04 

Q) 

C 

i' 

8 

!! 

). 

(R 

C 

c 

a 

> 

W 

m Q) rn > 

m 

> 

- 

I 

8 

s. 

% 

W 

Q) 

> 

m 

- 

mi I L 

U 

Q) 

C 

I I I I I I I I I I I I I I I I I I I 

- 

P 

aD 

CD 

d 

cu 

0 

I 

4 

II) 

x 

II) 

Q) 

bl 

1 

M 

.rl 

&I 

F8804112ABQ-19 4-19 
F88-04 

to the corrugation, th;e&J.xible line was modeled as being 10 times longer 

than its actual length . The impedance of the flexible line could be 

lowered by incorporating a liner inside the corrugated line. As results of 

the analysis show, however, the pressure drop through the flexible line is a 

small part of the total, and this may not be necessary. Inside diameters of 

the EVA coupling and bayonets are 1.12 cm, with appropriate equivalent 

lengths of smooth 1.8 cm tubing used in the model. 

The porous plug modeled is identical to that baselined for the SHOOT experiment with a nominal pore size of 0.5 p. The plug surface area is 200 cm , 

and the thickness is taken to be 1.2 cm. The plug is modeled as a 16 cm 

diameter disk, although a cylinder with identical surface area would perform 

essentially the same. The heater section succeeds the porous plug with the 

electrical element 0.5 mm off the plug face (Figure 4-10). The heater is 

modeled as a 16 cm diameter section of line tapered down to the 1.8 cm line 

diameter of the SFHT plumbing with electrical heating over a length of 5 mm. 

There are several sources of heat input to the transfer system, as shown 

be low. 

0.8 W - EVA coupling 

0.5 W - Bayonet coupling 

0.5 W/m - Flexible transfer line (3 m) 

0.5 W - External valve (each) 

0.5 

- 

W - Inner VCS (each) 

0.15 W - Warm relief line (eachj 

5.1 W Total heat input 

The levels of heat input shown are low, but achievable. The 0.8 W for the 

EVA coupling is the goal of the on-going coupling development program under 

contract from NASA/JSC. The 0.5 W for the bayonet is about twice as good as 

currently available commercial units; this level of heat leak is achievable 

with careful design and construction. The transfer line heat leak of 0.5 W/m 

is about twice as good as currently available vacuum- jacketed flexible lines. 

A heat leak of 0.5 W/m is achievable for a carefully fabricated vacuumjacketed line or for a flexible line that is foam-insulated rather than 

vacuum-jacketed. With ground and flight operations as we envision, the SFHT 

transfer line need work only in a vacuum environment and that a heavy vacuum 

jacket is therefore not required. The last three heat inputs shown depend on 

configurational details of the plumbing system. 

4.4.3 Transfer Analysis Results 

The transfer model was run for two operating conditions: full flowrate 

(1,000 &/hr) and minimum flowrate (70 t./hr). A maximum flowrate of 1,000 

&/hr was established as a study goal by the statement of work. Our analysis 

indicates that even higher flowrates are achievable with no deleterious 

impacts on the SFHT cost or development risk. The minimum flowrate is 

defined to be either (1) that existing with no pump heater power, as long as 

the flow remains superfluid or (2) that required to prevent the flow from 

going normal, using pump heater power. The minimum flowrate level is 

dictated by the overall flow impedance and heat input of the transfer system. 

Although transition to normal flow within the transfer system is physically 

F8804112ABQ-20 4-20 
b1.2 cm4 0.8 cm b1.2 cm-b 

HELIUM II 16 

(1.6 K) 

F8804112AB9-21 

F88-04 

AN 9203lMD135.025 

Figure 4-10 Thermomechanical pump configuration 

4-21 
F88-04 

possible, it is difficult to analyze and may be an unstable condition. The 

minimum achievable f lowrate is of interest because low f lowrate is desirable 

for efficient user cooldown and for fill of small users. 

Full-Flow Condition - A summary of the predicted full-flow condition is given 

in Table 4-6. The most direct indicator of transfer efficiency is the fluid 

temperature rise, which is 0.20 K; this is adequate considering that the user 

fluid mass loss during the subsequent bath pumpdown is about 1 percent for 

each 0.1 K of bath temperature reduction. For example, if helium is transferred to the user at a temperature of 1.8 K, then 2 percent of the bath is 

lost in pumping down to 1.6 K. The pump heater power required to generate 

the 1,OOO t/hr flowrate is 28.5 W. 

Profiles of temperature, pressure, and conducted heat flux along the transfer 

system are shown in Figures 4-11, 4-12, and 4-13. Slightly more than half of 

the total temperature rise takes place in the pump plug. Most of the remainder takes place from the SFHT external valve to the user external valve; 

almost all the transfer system heat input is along this section. The sharp 

temperature rise at the exit to the user vessel reflects heat input to the 

transfer system due to condensation of gas in the user vessel as the liquid 

volume increases and the ullage shrinks. 

The total transfer system pressure drop is about 34 torr; 24 torr of this 

occurs in the four valves, and most of the remainder is in the corrugated 

transfer line. Since a TM pump can produce a driving pressure of several 

hundred torr, pressure drop is an issue only because of the corresponding 

temperature rise. 

The heat flux shown is that described by the Gorter-Mellink relationship; it 

is gero for a normal fluid. The heat flux to the pump is about 0.5 torr, 

which is small compared to the electrical heater power. At lower flowrates 

its effect on the pump operation and the heater power required to produce a 

given flowrate becomes greater. As may be seen, the heat flux into the 

transfer system from condensation of vapor in the user vessel is about 2.5 W, 

corresponding to a flowrate of 1,000 k/hr. This is actually a specified 

boundary condition of the model. 

Minimum-Flow Condition - Heat conducted to the TM pump through the fluid from 

the warm transfer system is sufficient to initiate flow: this has been observed by experimenters at BASG and by others during -superfluid helium 

transfer testing. If the impedance and heat input of the transfer system are 

small enough, superfluid helium can be transferred to a receiver vessel 

without applying electrical heater power to the pump. This might be thought 

of as a "self starting" capability, which is desirable, but probably not 

necessary. Self starting will occur if the heat conducted upstream to the 

pump produces a flowrate high enough to prevent the parasitic heat input from 

raising the temperature of the fluid in the transfer system above the lambda 

point. 

The model was run with successively lower heater power until the fluid temperature at the user came up to the lambda point. The condition at which this 

occurred is shown in the heat flux, temperature, and pressure profiles of 

Figures 4-14, 4-15, and 4-16. A summary of pump operating parameters is 

given in Table 4-6. The resultant electrical power required is 1.1 W, and 

F8804112AB9-22 4-22 
Table 4-6 

RESULTS SUMMARY 

FULL-FLOW CASE 

F88-04 

MINIMUM-FLOW CASE 

70t/ hr 

1.6 K 

2.177 K 

1.1 w 

1.9 w 

0.4 W 

29.4 torr 

0.1 g/s 

FLOWRATE 

SUPPLY TEMPERATURE (SFHT) 

RECEIVER TEMPERATURE 

PUMP HEATER POWER 

HEAT INPUT TO PUMP 

PRESSURE RISE AT PUMP 

BACKFLOW FROM PUMP 

SFHT VENT RATE 

1 

1.6 

1.80 

28.5 

28.9 

41.5 

4.1 

1.4 

the flowrate is 70 &/hr. 

relatively small 0.16 torr, corresponding to the low flowrate. 

The transfer system pressure drop in this case is a 

Results indicate that it may be difficult to design a system of this size to 

be self starting. When operation evolves to automatic, the commercial 

bayonet and transfer line would be eliminated. This would result in reduced 

heat input to the transfer system and secondarily in reduced flow impedance. 

These changes would promote self starting, which would simplify the startup 

operation somewhat. The alternative is to power the electrical heater to a 

predetermined low level at some appropriate point of the start-up process. 

Without further analysis, the difficulties involved in the start-up process, 

especially regarding the control system, are not well understood. 

4.4.4 Conclusions from Transfer Analysis 

Although a major objective of the modeling effort has. been accomplished with 

the verification of the recommended transfer system, it should in no way be 

construed as an optimum design. It is, in fact, a step on the way to understanding the behavior of superfluid helium during the course of the transfer 

operation. The modeling effort has provided considerable insight into the 

parameters that impact system performance. It has also opened the door to 

some unanswered questions that are beyond the scope of this particular task. 

The following conclusions can be drawn: 

0 

0 

0 

It is feasible to establish a 1,000 &/hr flowrate (and 

greater) in a system of this size. It would actually be 

easier if the system were smaller. 

Minimizing parasitic heat input to the transfer system is 

important. It not only minimizes the fluid temperature at 

the receiver; it also sets the lower limit on superfluid 

flowrate. This second point is important because a low 

flowrate is desirable to achieve high efficiency during user 

cooldown. If a high flowrate must be used during cooldown, 

there will likely be a resultant fluid mass penalty. 

The start-up process must be better understood. It appears 

that if the parasitic heat input to the transfer system is 

F8804112AB9-23 4-23 
Y 

1.8 

1.78 

1.76 

1.74 

1.72 

1.7 

1.68 

1.66 

1.64 

1.62 

1.6 

0 

TEMPERRTURE 

System 

F88-04 

2 4 6 

Length (m) 

8 18 

Figure 4-11 Full-flow (1,OOO &/hr) temperature profile 

F8804112ABQ-24 4-24 
40 

30 

20 

10 

0 E 

0 

p Ir. 

PRESSUR E 

System 

~ 

F88-04 

2 4 6 

Length (m) 

8 18 

Figure 4-12 Full-f low (1 ,OOO &/hr) pressure profile 

F8804112A89-26 4-25 
. 

Q 2 

HEAT FLUX 

System 

4 6 

Length (m) 

8 18 

Figure 4-13 Full-flow (1,000 t/hr) heat flux profile 

F8804112A89-26 4-26 
F88-04 

F8804112ABQ-27 

Length (m) 

Figure 4-14 Minimum-flow (70 E/hr) heat flux profile 

4-27 
TEMP ERRTURE 

System 

2.2 

2.1 

2 

1.9 

1.8 

1.7 

1.6 

2 4 6 

Length (m) 

F88-04 

8 18 

Figure 4-15 Minimum-flow (70 &/hr) temperature profile 

F8804112ABQ-28 4-28 
c t 

F88-04 

Ll 

0 

IF8804112ABQ-29 

Figure 4-16 Minimum-flow (70 t./hr) pressure profile . 

4-29 
F88-04 

below a certain threshold, then start-up will happen without pump heater 

power. However, for greater parasitic heat input, the pump heater must 

be turned on at some point of the start-up process. Exactly when this 

should be done and how critical the timing is remains uncertain without 

further analysis. 

4-30 
F88-04 

Section 5 

COMMONALITY ISSUES 

Table 5-1 lists all the components used in the SFHT. An evaluation was made 

to determine if these components could be used as is or modified for use in 

tankers for other cryogens. Many components would not be compatible without 

modification or unless a course of action was taken to accommodate the other 

cryogens in the initial design of the SFHT. 

Table 5-1 

ASSESSMENT OF USE OF SFHT ELEMENTS ON SUBCRITICAL/SUPERCRITICAL TANKERS 

Applicable 

To Other Technical Difficulties With 

- Item Component Cryogens Using With Other Cryogen 

1. 

2. 

3. 

4. 

5. 

6. 

7. 

8. 

9. 

10. 

11. 

Thermomechanical No 

Pump 

Motor operated Yes 

Valve 

Transfer Lines Yes 

Bayonets Yes 

Mass Quantity Sensor No 

- Heat Pulse 

Flowmeter-Standby Vent No 

Flowmeter-Transfer No 

Liquid Level Sensor Yes 

-Superconducting.Wire 

Liquid Acquisition Yes 

Device 

Porous Plug Phase No 

Separators 

Avionics Yes 

Relief Valve Yes 

Will only work with SfHe 

Acceptable for use, operating pressure 

may be an issue 

Acceptable for use, operating pressure 

may be an issue 

Acceptable for use, operating pressure 

may be an issue 

Will only work with SfHe 

Measurement of flow heat capacitance 

works best at very low temperature. 

Method might work for low pressure 

systems such as solid hydrogen or neon. 

Will only work with SfHe 

Redesign required for use with other 

cryogens (use of higher temperature 

superconductor required) 

Redesign required for use with other 

cryogens because of different capillary 

action 

Use TVS with other cryogens 

Acceptable for use with software 

modification 

Acceptable for use, operating pressure 

may be an issue 

F8804112AD0-1 5- 1 
1 t 

F88-04 

Table 5-1 

ASSESSMENT OF USE OF SFHT ELEMENTS ON 

SUBCRITICAL/SUPERCRITICAL TANKERS (Continued) 

Applicable 

To Other Technical Difficulties With 

- Item Component Cryogens Using With Other Cryogen 

12. Burst Disc Yes Acceptable for use, operating pressure 

may be an issue 

13. Electrical Connector Yes Acceptable for use, operating pressure 

may be an issue 

14. Electrical Feedthrough Yes Acceptable for use, operating pressure 

may be an issue 

15. Low Thrust Vent Yes Acceptable for use provided size is 

right for flow rate 

16. Plumbing Yes Acceptable for use, operating pressure 

may be an issue 

17. EVA Coupler Yes Acceptable for use, operating pressure 

may be an issue 

18. Germanium Resistance Yes Acceptable for use up to a temperature 

Thermometer of 40 K (hydrogen and neon) 

19. Silicon Diode Yes Operating temperature range is probably 

Thermometer acceptable for most applications with 

other cryogens 

20. Pressure Transducer Yes Acceptable for use, operating pressure 

may be an issue 

21. Triaxial Accelerometer Yes Acceptable for use, operating range may 

be an issue 

22. Strain Gauges Yes Acceptable for use, operating range may 

be an issue 

23. Ion Gauge Yes Acceptable for use 

24. 

25. 

Cryogen Tank Yes Other cryogens, except hydrogen, are 

Support Straps more dense requiring stronger support 

straps 

Cryogen Tank (CT) Yes Supercritical cryogens may require a 

CT with a much heavier wall because 

of higher operating pressures. 

Required quantities of other cryogens 

may be much different 

F8804112AD0-2 5-2 
F88-04 

Table 5-1 

ASSESSMENT OF USE OF SFHT ELEMENTS ON 

SUBCRITICAL/SUPERCRITICAL TANKERS (Concluded) 

Applicable 

To Other Technical Difficulties With 

Item Component Cryogens Using With Other Cryopen 

26. Vacuum Shell Yes Must be beefed up for heavier CT 

- 

27. Mechanical cooler Yes Potential system mass benefit unknown 

without thermal analysis. Ground 

lock-up not a major issue with other 

cryogens 

28. Pyrotechnic Umbilical Yes Acceptable for use providing line sizes 

Cutter are the same 

5.1 LIQUID ACQUISITION DEVICE 

As part of the commonalty investigation, the LAD configuration appropriate 

for a 10,000 liter spherical tank containing either liquid oxygen or hydrogen 

has been determined. This analysis was performed for the study by McDonnell 

Douglas Astronautics Division, who have extensive experience in the design 

and use of channel capillary LADS. 

The LAD configuration consists of four channel legs, the same as recommended 

for our baseline SFHT concept, which i,s discussed in 2.1.4. The legs are 

joined by a mainfold to the pump inlet. The channel cross-section forms an 

isosceles triangle, 7.5 cm in height and 7.5 cm in basewidth. The base of 

the triangle, which faces the wall, is made of standard 325x2300 DDTW screen. 

The screen is 2 cm from the tank wall at all locations. 

Figure 5-1 shows the expulsion efficiency of the LAD as a function of adverse 

acceleration level for storage of oxygen or hydrogen. Performance with both 

fluids is insensitive to flowrate over the range from 10 to 100 percent of 

tank fluid capacity per hour. A residual of 2 percent is achieved for acceleration levels up to 0.1 g, well above the expected frequent disturbances of 

the resupply environment. The slightly lower performance with oxygen is 

attributable !f the lower ratio of surface tension to density of oxygen 

(i .e. , 1.27~10 compared to 1.95~10-') . 

In summary, the overall configuration of the LAD for use with liquid oxygen 

or hydrogen can be very similar to the device needed for SfHe. Therefore, 

with careful thought the mounting bracketry and overall structural design of 

the LAD could be identical for the three fluids. However, as discussed in 

2.1.4, the physics behind operation of the LAD with SfHe, and therefore the 

details of the screen (or porous material), are unique compared to normal 

fluids. 

F8804112AD0-3 5-3 
F88-04 

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F8804112AD0-4 5-4 
F88-04 

Section 6 

TECHNOLOGY DEVELOPMENT 

An assessment was made to determine the technology development status of all 

the components used on SFHT. The 

development status is based on the criteria for Development Status and Criticality shown at the bottom of the table. Criticality is a judgment based on 

the need for that component to complete the mission. The cost and schedule 

has been noted for those hardware items requiring further development. Comments have been included to explain why development is required and to give a 

history of successful use of the proposed components. 

This assessment is presented in Table 6-1. 

F8804112ADl-1 6- 1 
L 

F88-04 

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F88-07 

Section 7 

PROGRAM PLAN 

This section defines a preliminary program plan for development and certif ication of the SFHT system and associated GSE described in Section 2. The 

plan assumes one protoflight SFHT system with spares only at the component 

level and one set of GSE, also with only component spares. 

7.1 HARDWARE TO BE DEVELOPED 

In addition to the flight hardware and its GSE, special test equipment, 

training equipment, and production tooling must be developed. The items of 

development hardware assumed for construction of the program plan are summarized below. 

SFHT Flight System - Two major subsystems, the dewar and the avionics, comprise the SFHT, as shown in Figure 7-1. The subassemblies of the two are 

shown in the figure. Almost all of the technology development risk is 

associated with the dewar and its operations. 

Ground Support Equipment - As described in Section 2.3, the GSE is comprised 

of three categories of equipment: mechanical (handling), (cryogenic operations), and monitor/control. Development of this complex set of equipment is 

an important part of the SFHT program. The most significant items of GSE are 

the 15,000 e supply dewars. Two are needed, and they must operate at 

negative pressure to perform the low pressure topoff of the SFHT. 

Special Test Equipment (STE) - There are four items of STE. First is a test 

dewar to be used for verifying the SFHT transfer operations. This dewar may 

be relatively small (e.g.-, COO e) and must interface with the SFHT fluid 

subsystem. It must have appropriate internal valving and a high conductance 

vent similar to the orbital resupply users. It is not necessary to include 

the EVA coupling in the transfer system performance verification, and the 

fluid couplings of the test dewar will therefore be standard bayonets. GSE 

transfer lines can probably be used for the test. 

The other three items of STE are electronic simulators of 1) the SFHT dewar, 

2) the user, and 3) the host vehicle. These will be used for verification of 

the avionics subsystem before it is mated to the dewar and for final 

certification of the avionics. These simulators.wil1 allow the two SFHT 

major subsystems, the dewar and the avionics, to be fabricated and certified 

somewhat independently of each other. The two will be mated and verified as 

a total system only after they have each been fully verified as a subsystem. 

This approach will minimize schedule risk factors related to the complex enditem testing required for the two subsystems, especially that of the dewar. 

Training Equipment - An avionics simulator will be provided for astronaut 

training that duplicates human interfaces and control functions of the flight 

system. 

F8804112AD4-1 7- 1 
. 

ORIGINAL PAGE IS 

OF POOR QUALITY 

fi I I DEWAR I I AVIONICS 

INSULATION SENSORS 1 ::$’ 1 INTERNAL CABLING 

E XTE RN AL 

E XTERN AL 

CONTROL 

EXTERNAL CABLING 

t sm OEWAn sENson DATA 

INTERFACES OPERATINO INTERFACES COYPONENTS 

INFORYATION 

EMERGENCY VENT VENTINO conTRoL STRUCTURE 

POWER 

Figure 7-1 SFHT flight hardware breakdown 

F8804112AD4-2 7-2 
F88-07 

Production Tooling - During fabrication and assembly of the dewar, carefully 

designed production tooling is required. These include holding fixtures for 

automatic welding, holding fixtures for fabrication of subassemblies such as 

plumbing manifolds, templates for MLI cutting, fixtures for assembly of 

vapor-cooled shields, a cryogen tank holding fixture for MLI layup, proof 

test fixtures, etc. 

7.2 PROGRAM SCHEDULE 

The development program schedule is shown in Figure 7-2. The 5.5 year phase 

C/D is preceeded by a phase B study. The preliminary design review is shown 

1 year after contract start. To maintain the planned schedule, it is essential that critical long-lead procurements, such as spinnings for the cryogen 

tank, be initiated immediately after the PDR. The extensive dewar subsystem 

testing is reflected by the estimated 9 month test duration. 

7.3 RISK 

Critical technology development issues are discussed in Section 6. The most 

critical area is the transfer system, especially the liquid acquisition device and understanding of the transient aspects of the transfer operation, 

particularly the startup process. The SHOOT program will provide many of the 

answers, but, because the SFHT is considerably larger, it is likely that 

further development testing guided by the results of the SHOOT experiment 

will be needed. The schedule of Figure 7-2 shows component development 

starting during the phase C/D preliminary design stage. 

7.4 MAKE-OR-BUY PLAN 

Preliminary make-or-buy decisions for the major components are shown below: 

BUY 

BUY 

EITHER 

BUY 

BUY 

BUY 

BUY 

BUY 

MAKE 

MAKE 

MAKE 

MAKE 

MAKE 

EITHER 

BUY 

Spinnings for the cryogen tank, vacuum shell, and VCSs 

Forging for the vacuum shell girth ring 

Machining of cryogen tank and girth ring 

Support straps 

Motor-operated valves 

Mechanical cooler and electronics 

EVA coupling 

Electronic parts/instrumentation 

Thermal switch 

Low flowrate flowmeter 

Avionics electronics boxes 

Dewar exit cable assemblies 

Liquid acquisition device 

15,000 k?. supply dewars 

Transfer lines, bayonet couplings 

7.5 LONG-LEAD PROCUREMENTS 

Critical long-lead procurements are: 

0 Spinnings for the cryogen tank, vacuum shell, and VCSs 

0 Girth ring forgings/machinings 

0 Support 0s traps 

F8804112AD4-3 7-3 
F88-07 

ORIGINAL PAGE IS 

OF POOR QUALITY 

N?4 03uIUD142 27 

Figure 7-2 SFHT program schedule 

F8804112AD4-4 7-4 
0 Motor-operated valves 

0 Electronic parts 

0 EVA coupling 

0 Mechanical cooler 

0 15,000 E supply dewars 

0 Liquid level sensors 

F88-07 

Criticality is based on a combination of estimated delivery schedule and need 

date in the fabrication schedule. For example, items associated with the 

cryogen tank, which is the first major subassembly on the program critical 

path, have high criticality. 

7.6 FACILITIES 

Figure 7-3 shows the flow of the SFHT dewar fabrication process and the 

facilities required. BASG has all the facilities needed for fabrication and 

test of the SFHT system. 

F8804112AD4-S 7-5 
F88-07 

ASSEMBLE 

FABRICATE - CLEAN - CRYOQEN 

PARTS PARTS TANKS 

ORIGINAL PAGE IS 

OF POOR QUALITY 

WELD X-RAY AND 

TANK 

CRYOQEN - LEAK - INSPECT 

TEST WELDS 

- 

ASSEMBLE WELD PIPES LEAK X-RAY AND 

ASSEMBLE MLI X-RAY AND AND VCS TO EACH VCS TEST INSPECT 

INSPECT - ONTO CRYO TANK WELDS 

CRYO TANK TEST WELDS 

ASSEMBLE 

CRYO TANK AND WELD - LEAK - AND STRAPS - 

INTO DOLLY TO MANIFOLD 

- 

WELD GIRTH X-RAY AND RINQS. CYLINDER - LEAK - INSPECT 

APERTURE CONE 

AND TEST WELDS 

FACILITY 

- FINAL 

CLEAN 

WELD 

LOWER H 1 ASSEMBLE 4 AND VACUUM 

CRYO TANK H DOME 

I SHELL IUU 

CLEAN WELD CLEAN 

ROOM SHOP ROOM 

X-RAY 

EXTERNAL 

INSPECT MANIFOLD 

X-RAY AND WELD 

DYE-PENETRANT SHOP 

X-RAY 

TESTS INSPECT 

CLEAN X-RAY AND 

ROOM DYE-PENETRANT 

FACILITY FACILITY 

UN 9344IMOlU 29 

Figure 7-3 SFHT dewar manufacturing process and facility usage 

F8804112AD4-6 7-6 
i 

Section 8 

CONCLUSIONS AND RECOMMENDATIONS 

F88-04 

A SFHT design concept has been recommended that meets all mission requirements using low-cost and low-risk approaches. The all-aluminum, 10,000 f?. 

spherical dewar will perform a warm SIRTF resupply at either the shuttle or 

the space station (after a 9-month standby). The insulation system design is 

optimized to provide minimum launch mass based on the 9-month standby 

requirement; mechanical coolers are used to cool the dewar outer VCS during 

the extended SS standby. 

An oblate spheroid configuration was investigated to see if it would reduce 

shuttle launch cost; it has essentially the same launch cost and thermal 

performance as the spherical dewar, which is 1,600 lb lighter. It requires 

about 19 in. less shuttle bay length, providing better manifesting flexibility. Although the spherical shape has been selected as the baseline, there 

is not a strong preference. 

The study has emphasized the areas of technology risk, which center around 

the fluid subsystem design and operations. Technology development is needed 

for the follow-on automatic system, but only a cursory consideration of those 

items was appropriate within the scope of this study. When the SFHT hardware 

is developed, state-of-the-art, space-qualified technology will be used for 

the avionics. Therefore, little study effort was spent on this topic. Ref garding avionics, the study was more concerned with assuring that SFHT power 

and data requirements are compatible with shuttle and SS resources. Considerable software development will be required for the avionics "smart" 

computer system. , 

BASG has a proven capability for designing and fabrication spaceborne dewars 

with high thermal performance. Although this is not a technology risk area, 

considerable study effort was spent on providing valid mass and thermal performance estimates for the dewar. This was necessary to make legitimate 

configuration tradeoffs and plan operations, both ground and flight. Also, 

minimizing the parasitic heat input to the cryogen tank will minimize the 

SfHe bath temperature and maximize the transfer efficiency. 

The primary technology development area for the SFHT is the transfer system, 

consisting of the liquid acquisition device, thermomechanical pump, SFHT 

plumbing, flexible transfer line, and user plumbing. Demonstration of this 

system is the main function of the SHOOT experiment. However, there are 

important scaling factors and detailed design impacts that can only be determined through a rigorous analytical understanding of the system. Regardless 

of ground component testing and SHOOT flight testing, a valid analytical 

model of the entire transfer system is required to establish the SFHT design 

with any confidence. That analytic capability does not yet exist. 

Our approach has been to take a very careful end-to-end look at a point 

transfer system design based on the fundamental physics involved, therefore 

verifying the feasibility of performing on-orbit superfluid helium resupply. 

We conclude that it can be done within the imposed design constraints, but 

the system design and operation must be very carefully planned and verified. 

F8804112AD3-1 8- 1 
F88-04 

The next step is to construct a comprehensive analytic tool for understanding 

the system better, particularly the transient aspects of the resupply 

process. 

F8804112AD3-2 8-2 
F88-04 

REFERENCES 

1. Personal communication with Tom Bradshaw of Rutherford Appleton 

Laboratory. 

2. Memo from Bill Boyd (JSC) dated June 13, 1988. 

3. Multimission Modular Spacecraft Flight Support System Description, Satellite Servicing Workshop 111, June 1987. 

4. NASA/JSC Contract Number NAS9-17872 to Moog. 

5. Verbal information from Bob Brown of JSC. 

6. Carrington, H., "A Review of the Low Gravity Environment of the Shuttle 

Orbiter," BASG Report 4828-640-HGC-86.001, December 18, 1986. 

7. Study of Helium Transfer Technology for STICCR-Fluid Management, NASA 

Contract Report 177463. 

8. 

9. Daney, D.E., "Cavitation in Flowing Superfluid Helium,"Cryogenics, FebSHOOT Program Testing, CDR Handout, August 1988. 

ruary 1988, page 132. 

Hopkins, R.A., and Payne, D.A., "Optimized Support System for Spaceborne 

Dewars," Cryogenics, April 1987. 

10. 

11. Hopkins, R.A., and Payne, D.A;, "Next-Generation Tension Strap Supports 

for Spaceborne Dewars," presented at AIM Thermophysics Conference, June 

1987. (Preprints available) 

12. Space Shuttle System Payload Accommodations, NSTS 07700, Volume XIV, 

Revision J; Appendix 2, Thermal. 

13. Space Station Program Definition and Requirements, JSC 30,000. 

14. Space Station Program Definition and Requirements, JSC 30,000. 

15. Final Report, Contract NAS9-17585, Orbital Spacecraft Consumables Resupply System Study, MCR-87-1339, September 1987. 

NASA Contractor Report 177381, SIRTF Telescope Instrument Changeout and 

Cryogen Replacement Study, November 1985. 

16. 

17. BASG System Engineering Report SFHT.002, Standby Vent Flow Conductance 

Analysis, May 2, 1988. 

BASG System Engineering Report SFHT.003, Emergency Vent Flow Conductance 

Analysis, April 19, 1988. 

18. 

F8804112AB7-1 1 
J 

F88-04 

19. 

20. 

21. 

22. 

23. 

24. 

25. 

26. 

Space Shuttle System Payload Accommodations, NSTS 07700, Volume XIV, 

Revision I, Section 8.1.1.1(b). 

Snyder, H.A., "Dewar to Dewar Model for Superfluid Helium Transfer," 

Cryogenics, February 1988. 

AXAF XRS Cryogenic Subsystem Definition Study, Contract NAS5-30186, 

Final Technical Report, December 1987. 

Launch Site Accommodation Handbook for STS Payloads, K-STSM-14.1, Rev. 

B. 

Telecon with Roger Wall, NASA/KSC, July 1,1988 and Jim Weir, NASA/KSC, 

July 5, 1988. 

SHOOT Preliminary Design Review, Volume I, September 29, 1987. 

Hawthorne, R.C. and Helms, H.C., "Flow in Corrugated Hose," Product 

Engineering, June 10, 1963. 

Walstrom, P.L., Weisend, J.G., Moddocks, J.R., and Van Sciver, S.W., 

"Turbulent Flow Pressure Drop in Various He I1 Transfer System 

Components," Cryogenics, February 1988. 

F8804112AB7-2 

NASA-JSC 

2 
1 NASA JSC Tc2 J. Williams 

1 NASA JSC K2 K. Havens 

1 NASA JSC DF6 T. White 

1 NASA JSC DF4 R. Trevino 

1 NASA JSC EA4 Technology Utilization Officer 

4 NASA JSC JM6 Technical Library 

1 NASA MSFC KA21 L. Hartings 

1 NASA MSFC KA4 1 U. Hueter 

1 NASA MSFC PD2 2 S. Tucker 

1 NASA MSFC EP53 L. Jones 

1 NASA MSFC EP53 G. Langford 

1 NASA MFC EB24 E. Hinman 

1 NASA MSFC KA51 R. Hoodless 

2 NASA STIF Accessing Dept. 

P. 0. Box 8757 

Baltimore/Washington International 

Airport, M> 21240 

1 USAF Space Division XR 

P. 0. Box 92960 

Loa Angeles, CA 90009 

Maj. L. Jackson 

1 Lockheed Missiles and Space Co. J.Woh1, Dept 5751 

1111 Lockheed Way 

Sunnyvale, CA 94086 

1 McDonnell Douglas Astronautics Co. K. box, h4!S 11-3 

5301 Bolsa Avenue 

Huntington Beach, CA 92647 

1730 NASA Rd. 1, Suite 203 

Houston, TX 77058 

1 McDonnell Douglas Corp. A. Eisenberg 

1 ILC Space Systeme 

16665 Space Center Blvd 

Houston, TX 77058 

M. Wi they 

PRECEDiNG PAGE BLANK NOT FlLMED 
DRL Item +6 -- Final Report (cont) 

Qry 

--- 

1 

1 

1 

1 

1 

1 

1 

1 

1 

1 

1 

1 

1 

General Dynamics Corp. 

1300 Hercules, Suite 204 

Houston, TX 77058 

TRW 

2525 Bay Area Blvd., Suite 620 

Houston, TX 77058 

Teledyne Brown 

300 Sparkman Drive 

Huntsville, Alabama 35807 

Hughes Aircraft Co. 

Space & Camnunications Group 

P.O. Box 92919 

Ball Aerospace Systems Division 

P.O. Box 1062 

Boulder, 00 80306 

LO8 hgele8, 90009 

Astronautics Corporation 

Technology Center 

5800 Cottage Grove Rd 

Madison, Wisconsin 53716-1387 

Eagle Engineering 

711West Bay Area Blvd, Suite 315 

Webster, TX 77598 

Boeing Company 

P.O..Box 58747 

Houston, ?x 77258 

Grunmlln Corp. 

2200 Space Park Drive 

Houston, TX 77058-3690 

SAIC 

1701 E. Woodfield, Suite 819 

Schaumburg, IL 60195 

Rockwell International 

12214 Lakewood Blvd 

Downey, CA 90241 

Rockwell International 

555 Gemini Ave. D, Suite 301 

Houston, TX 77058 

Fairchild Space Co. 

Ge rmantown, Mary1 and 20874- 11 81 

J. Crenshaw 

C. Peterson 

W. 'Ifrvichell/MS-52 

J. Juraco/h4S C350 

D. Hopkins/LTl 

E. Rice 

L. Fisher 

G. Bottomley 

G. Stone 

S. Hoffman 

T. ThonetIAD38 

D. Finley 

M. Levy/D-8 
. . 

DRL Itemr6 -- Final Report (cont) 

QTY ORGANIZATION MAIL CODE TO THE A"TI0N OF 

--- ------------ --------- ------------------- 

1 Fairchild Control Systems S. Mu 

1800 Rosencrans Ave. 

Manhattan Beach, CA 90266-3797 

1 Martin Marietta Corp. 

P.O. Box 179 

Denver, CO 80201 

1 Martin Marietta Corp. 

P.O. Box 179 

Denver, CO 80201 

1 PRC Systems Services 

7911 Charlotte Drive 

Huntsville, AL 35802 

1 Aerospace Corporation 

P.O. Box 92957 

LO8 Angeles, CA 90009 

T. Tracy/S8082 

R. EberhardtlS8082 

A. Prince 

B. Cheney/M4-922 

1 Ccunputer Technology Associates D. Beaudet 

7501 Forbes Blvd 

Suite 201 

Lanham, MI 20706 

.