Research and Technology 

2003 Annual Report 

John F. Kennedy Space Center 
Foreword 

When one visits the Kennedy Space Center 

(KSC) Home Page, the words “Launching the 

NASA Vision” immediately appear at the top 

of the screen. This phrase was never more true 

or applicable as the Agency works diligently to 

return to flight safely and implement the new 

vision for space exploration. The KSC team will 

play a major role in these Agency endeavors. 

KSC continues to emphasize programmatic excellence as we work toward becoming a truly great 

spaceport operations and technology Center. We 

remain committed to bringing our customers 

the greatest value hi reachii-ig thcir djectives hy 

working toward a common set of priorities. Our 

technology development activities encompass the 

entire KSC team, consisting of Government and 

contractor personnel working in partnership with 

academic institutions and commercial industry. 

In 2002, KSC celebrated 40 years of pioneering the 

future. This year, NASA exceeded 1,000 days of 

permanent human habitation of the International 

Space Station. To continue our exploration of 

space and the development of the frontiers of the 

universe, we must be true to our vision, inspire 

the next generation of explorers, and partner with 

our customers and suppliers to make space travel 

safer and more accessible. This KSC Research and 

Technology 2003 Annual Report demonstrates 

some of these contributions to NASA’s mission. 

Dr. Dave Bartine, KSC Chief Technologist, 

321-867-7069, is responsible for publication of this 

report and is the point of contact for anv desired 

information. 

James W. Kennedy 

Director 

John F. Kennedy Space Center 

i/ii 
CONTENTS 

Command. Control. and Monitoring Technologies ................................................................................................ 1 

Command. Control. and Monitoring Technologies Roadmap ......................................................................... 3 

Multisensor Array Pressure Transducer ........................................................................................................................... 4 

Color-Indicating Patch for Shuttle Auxiliary Power Unit Hydrazine Transfer Line ....................................................... 6 

A Study of Hail Monitoring Strategies for Shuttle Launch Safety ................................................................................... 8 

Mass Sensor for Cryogenic Fluids ................................................................................................................................... 70 

Shuttle Tile Moisture Sensor ........................................................................................................................................... 12 

New Gas Polarographic Hydrogen Sensor ...................................................................................................................... 14 

Improved Fiber-optic Pressure Block Connector ............................................................................................................ 16 

Common Application Framework (CAF) ........................................................................................................................ 18 

Digital Image Lens Distortion Correction Algorithm ..................................................................................................... 22 

Automated Test Tool ........................................................................................................................................................ 20 

Real-Time Detection and Tracking of Multiple Moving Objects in Complex Image Background .................................. 24 

A Digital Signal Processing (DSP) Filter Tuned To Detect Diurnal Signals ................................................................. 26 

Image Edge Extraction via Fuzzy Reasoning ................................................................................................... ....... 28 

?4Giiit<r&c< 0ptiiiii.i: liiii.ig< Biiiaiktiiioii ~ifi F~AZZ~ Zefisoriing .................................................................................... JV 

Converting Analog Test Data to Audio Wave Files ................ ................................................................................... 32 

Small Mass Spectrometer System Onboard High-Altitude Research Aircraft WB-57 ................................................... 34 

Advanced Data Acquisition System (ADAS) ................................................................................................................. 36 

Visual Anomaly Detection System Prototype ................................................................................................................. 38 

Wireless Inclinometer ...................................................................................................................................................... 40 

Enhanced Wireless Vacuum-Jacketed Line Sensor Network ........................................ ............................................ 42 

A Field Prototype System for Vacuum-Dying Shuttle Tiles .......................................................................................... 44 

Shuttle Tire Pressure Monitor Calibration Station ......................................................................................................... 46 

Vacuum-Jacketed CVJ) Thermocouple Calibration Station ..................................... ... 48 

Fast. Time-of-Flight Mass spectrometer for Rugged Aerospace Applications ................................................................ 50 

Inspection Tool for the Reaction Control System (RCS) Nozzles .................................................................................... 52 

Fluid System Technologies ........................................................................................................................................ 55 

Fluid System Technologies Road map ............................................................................................................................... 57 

Modeling the RLlO Rocket Engine With Dens$ed Liquid Hydrogen and Oxygen Propellants .................................... 58 

Advanced Umbilical Development Automation Testbed ................................................................................................. 62 

Ad~anced Cold-Mate Cryogenic Fluid Quick Disconnect (QD) .................................................................................... 66 

Fiber-Optic Quick Disconnect (QDj ............................................................................................................................... 70 

On the Scaling Laws for Jet Noise in Subsonic and Supersonic Flow ............................................................................ 60 

Advanced Umbilical Location System (AULS) ............................................................................................................... 64 

Collapsible Cryogenic Vacuum-Jacketed (CCVJ) Fluid Transfer Lines ........................................................................... 68 

Ice Suppression Technologies ........................................................................................................................................... 72 

Automated Latching Systems .......................................................................................................................................... 74 

Umbilical Ice Suppression and Mitigation System (ISMS) and Cleanliness Ver$cation System (CVS) ..................... 76 

... 

Ill 
CONTENTS (continued) 

Testing of Space Shuttle Spray-On Foam insulation Under Cryogenic Vacuum Conditions ........................................ 78 

Glass Microspheres for Cryogenic Insulation Applications ............................................................................................ 80 

Standardized Testing of Experimental Thermal Insulation Materials for Low-Temperature Applications. ................... 82 

Thermal Performance Testing of Cryogenic Piping Systems ................... 

Piping System Repair Using Cryogenic Freeze Plug Technology ................................................................................... 86 

Automated Thruster Bench Test (ATBT) System ... 

Venturi Of-Set Technology Cryogenic Valve .................................................................................................................. 90 

.................................................................. 84 

.................................................................................................... 88 

Range Technologies ..................................................................................................................................................... 93 

.................................................................................................. 95 

................................................................. 96 

.............................................................. 98 

Range Technologies Roadmap ............................ 

Space-Based Telemetry and Range Safety (STARS) ......... 

Iridium Flight Modem ....................................................... 

Autonomous Flight Safety System (AFSS) . Phase II ....................................................................................... 

Range lnformation System Management ...................... 

Transient Voltage Recorder .................................................................................................... 108 

Passive Millimeter Wave Imaging ........... 

................. .................................................................................. 102 

.......................................... 104 

..... 106 

................................................... 110 

Spaceport Structures and Materials ....................................................................................................................... 113 

Spaceport Structures and Materials Roadmap .... 

Paschen Breakdown in the Martian Atmosphere 

Wheel Electrometer Technology .............................................. 

Charging of Materials on Earth and on Planetary Surfaces 

Comparison of Surface Resistivity and Triboelectric Charge Generation Characteristics of Materials ........................ 124 

Method for Measuring Charge Buildup on Materials ............. 

Electrostatics Testing of Materials at Low Pressures .................................................................................................... 128 

Spacecraft Charge Monitor .................... 

Antistatic, Field-Emitting, Nanoparticle Coatings .................. 

Corrosion-Resistant Tubing for Space Shuttle Launch Sites ............................................... 

....................................... 115 

.................. 116 

................................................................................................. 118 

................................................... 120 

Electronic Single-Particle Aerodynamic Relaxation Time Analyzer for Mars Robotic Missions . 

Development of a Physical Model of Insulator-lnsulator Triboelectric Charging To Understand Surface 

.............................................. .. 122 

.......................................................... 126 

................................................................................................................. 130 

....................................................... 132 

.................................... 134 

Electrochemical Characterization of Tubing Alloys in Simulated Space Shuttle Launch Pad Conditions ................... 136 

Advanced Multisensor for Electrolytic Characterization .............................................................................................. 138 

Chromate Coating Replacement for Military Aircraft .................................................................................................. 140 

.................................................................... 142 

Electrochemical Analysis of Chloride Rinse Agents on Aircraft Alloys ....................................................................... 144 

............................. 146 

Army Aircraft Alloys - Corrosion-Retardant Additive Testing ... 

Development of Liquid Applied Coatings for Protection of Steel in Concrete ........................... 

Noise Mitigation of Ducted Supersonic Jets for Launch Exhaust Management Systems ............................................ 148 
CONTENTS (continued) 

Deoelopmeizt ofa Robust Finite-Element Method for Rocket Acoustics Problems ....................................................... 1.30 

Tape Adhesion Testing Software .................................................................................................................................... 152 

Process and Human Factors Engineering Technologies ..................................................................................... 155 

Process and Human Factors Engineering Technologies Roadmap ................................................................................ 157 

A Simulator Study of the KSC Slzuttle Landing Facility Air Trafjzc Control Touier Design I58 

Assessing Space Shuttle Launch Dates With Advanced Discrete-Ezieizt Simulation Technologies ............................. 160 

Virtual Reality Simulations of Orbiter Temporary-Access lmprouemeizts ................................................................... 162 

Range Process Simulation Tool (RPST) ......... ........................................................... 164 

Spaceport Safety Modeling and Optimization .................... ........... ........................................... I66 

Advanced Process for Developing and Balancing Quantifatiae Safety, Reliability, 

and Maintainability Requirements ................................................................................................................................ 168 

Human Engineering in Design and Modifications ......................................................................... .. 170 

Task Analysis Technologies ........ ............................. 172 

Human Factors-Process Failure Modes and Effects Analysis (HF-PFMEA) Software Tool ........................................ 174 

Task Analysis of the Payload Test and Checkout System (PTCS) .................................................... 

Actiziity-Based Costing and Simulation for the Operational Assessment of Future Space Transportation Systems ... 180 

Artificial Intelligence Techniques for Payload and Vehicle Processing Scheduling ......................................... 

Toolkit for Enabling Adaptive Modeling and Simulation (TEAMS). ..... 

Design Root-Cause Analysis Knowledge Base.. .............. 

Project Management Integration System (PMIS) .............. 

Portfolio Analysis Tool (PAT) ......................................................................................................................... 

...... 

............................................................ 186 

Biological Sciences ..................................................................................................................................................... 191 

BiologicaI Sciences Roadmap.. ... .................................................................. 

Atmospheric Mariagemen t in Variable-Pressure Enz7iroizments: A Step Toiciard Mortian greenhouse^ .................... 1 96 

Advanced Life Support (ALS) Crop Testing .... ..... .................................................................................... 198 

Plant Lighting System ........................................................................ ...... 

Effect of Microgravity on Raphanus Sativus L. ................... 

Deinoizstratioiz of a Porous Tube Hydroponic System To Control Plant Moisture and Grozofh ..... 

Photosynthesis and Metaboiism of W7zeat in Microgravity .................................................. 

.............................................. 204 

Index ............................................................................................................................................................................. 207 

V 
KSC Technology Development 

.- 

Shuttle Processing 

Launch Services Program 

International Space Station/ 

Payload Processing 
Major Stake ho I de rs 

Robotic/Human Exploration 

Next Generation 

Launch Technologies 

Future Spaceports 

and Ranges 

Orbital Space Plane Concepts 

vii 
Technology Transfer Office 

Introduction 

The John E Kennedy Space Center's (KSC) outstanding record of achievements has earned it an honored place in history and an essential role in 

Space Transportation Systems of today and tomorrow. KSC is increasing the 

momentum in technology development for current and future spaceports. 

KSC's dual mission includes space launch operations and spaceport and range 

technology development. The Spaceport Technology Center (STC) initiative carries out KSC's role within NASA to meet the goals of increased safety, 

reduced cost of space access, and rapid expansion of commercial markets by 

infusing spaceport technologies into all facets of current and future Space 

Transportation Systems. 

KSC's background as the nation's premier launch site creates an ideal environment for the STC. The STC's knowledge, expertise, facilities, and equipment provide technologies and processes to customers who propose to build 

and operate spaceports on Earth, in orbit, and beyond. KSC has three strategic lines of business: Spaceport Operations, Spaceport Design and Systems 

Development, and Spaceport and Range Technology and Science. KSC has 

unparalleled expertise in designing, building, and operating a spaceport with 

all of its complex technologies and systems. 

STC technology development activities are concentrated in six Spaceport 

Technology and Science Product Lines - Fluid System Technologies; Spaceport 

Structures and Materials; Process and Human Factors Engineering Technologies; Command, Control, and Monitoring Technologies; Range Technologies; 

and Biological Sciences. KSC's leadership in producing safer, faster, cheaper, 

and more robust systems by developing and infusing technologies will pave 

the way for the future of the space industry. This report is organized by these 

six Spaceport Technology and Science Product Lines. 

The primary stakeholders and customers for spaceport and range technologies are the programs and initiatives supporting current and future Space 

Transportation Systems and enabling technical programs. KSC aggressively 

seeks industry participation and collaboration in its research and technology 

development initiatives. KSC also seeks to transfer its expertise and technology to the commercial sector and academic community. Programs and 

commercialization opportunities available to American industries and other 

institutional organizations are described in the Technology Transfer Office 

Internet Web site at Izttp://feclzizology.ksc.~zasa.gov. Additional information on 

KSC's Spaceport Technology Center can be found on KSC's home page at 

www.ksc.nasa.gov. 
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Command, Control, and Monitoring 

Technologies 

The vision of Command, Control, and Monitoring (CCM) 

Technologies is to develop technology that enables future affordable, responsive, and safe spaceports. These spaceports can 

be Government-owned, commercial, orbital, or nonterrestrial. 

The goal is to reduce the cost of access to space while increasing safety. CCM Technologies can help in this role through the 

reduction and elimination of unique interface and infrastructure 

requirements. As we move away from a one-for-one cardinality 

between launch vehicle and ground infrastructure toward a truly 

many-to-many paradigm, we must shift our thinking about how 

we process and launch spacecraft. Spacecraft will have to become 

smarter and be in control of their spaceport processing, requesting needed services from the spaceport, as opposed to the smart 

spaceport telling the vehicle that it is now OK to fly. We must 

strive to remove complexity not only in the design but also in 

how information is presented to humans, wherever they may be 

in the loop, and provide decision support. Increases in processing requirements throughout the spaceport will drive advancements in the communication infrastructure, capability, and 

security. Finally, when dealing with human spaceflight, we need 

to support robust, fault-tolerate designs in both the software and 

hardware systems. The technology focus areas that will drive the 

roadmap toward that ultimate vision include the following: 

Sensors and Data Acquisition 

Spaceport Processing and Health Maintenance 

Advanced Software and Computing Ardutectures 

Simulation and Situational Awareness 

for more information regarding Command, Control, and Monitoring 

Technologies, please contact Robert Waterman (RcJ~w~. D. \lhfmiioi!@ 1117517 

.<yoi7),YA, (322) 867-6680; Gregory Clements (C:cgo::;.R.C!~!!!c!!t.."!!r!ia "'5 WT,), 

YA-D8, (321) 867-8992; or David Kruhm (D17;iid.A.Kl.iiliii~~i~~~o.s'~~,), 

YA-D5, (321) 867-6742. 
Command, Control, and 

Monitoring Technologies Roadmap 

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Multisensor Array Pressure Transducer 

Multisensor array (MSA) transducers have been under development for the 

last few years, and recent achievements may make commercialization of such 

devices a reality in the near future. A prototype MSA pressure transducer was 

developed and demonstrated a measurable increase in transducer reliability 

and expanded transducer calibration cycles. 

Common sense suggests that multiple sensors are better than one and, 

if so, how much better? Past experimentation focused on designing and 

implementing tests to answer this question and to quantify the relationship 

between the number of sensors and the associated improvement in sensor 

life and reliability (see KSC Research and Technology 2000/2001 Report, 

"Multisensor Array - Are More Sensors Better Than One?"). 

Multiple sensors can be implemented in numerous ways. The previously 

discussed method took advantage of the reduced size of microelectronic 

sensors and packaged several sensors onto a single circuit board. The result 

was a fully integrated multisensor array pressure transducer. 

The pressure-sensing portion of the device consists of eight individual 

integrated-circuit (IC) pressure sensors mounted on a custom printed circuit 

board. These sensors, coupled with in-house-developed signal conditioning 

electronics, compose the electronics hardware of the device. Using knowledge 

gained from previous sensor characterizations and pressure calibrations, a 

microcontroller within the signal conditioning electronics was programmed 

with the in-house-developed multisensor array algorithm (MSAA). This 

algorithm compares the measurement error of each individual sensor and 

resembles the statistical process used in the Monte Carlo simulation. To 

provide a calculated output, a polling logic, combined with a sensorMultisensor Array Pressure Transducer 
weighting algorithm and a sensor exclusion algorithm (to 

eliminate outliers from the decisionmaking process), is 

embedded in these sensor electronics. 

To calculate the necessary single-pressure output value, 

the embedded software performs several functions and has 

unique characteristics. First, it allows the user to select target 

error /failure criteria. These are values that define acceptable 

overall measurement error and the acceptable error of any one 

sensor element. Once these are defined, the software utilizes 

complex sensor voting and measurement weighting schemes. 

When compared to defined error /failure criteria, this analysis 

provides the determination of valid sensor measurements as 

well as elimination of outlier values from the decisionmaking 

process; the transducer’s single measurement output value is 

based on the voting and weighting results. Last, the availability 

of information pertaining to individual sensor element 

measurements allows for the real-time assessment of individual 

sensor element health. Internal tracking of this information will 

allow for trending and ultimately element and transducer failure 

-..-A: -c: n.. yltulLLlUl1. 

Key accomplishments: 

Paired the developed signal conditioning electronics with a 

commercially developed, eight-element, multiple-strain-gage 

sensor. Successfully demonstrated the coupled system. 

transducer manufacturers. 

Began discussions for commercialization partnerships with 

Contact: J.M. Pevotti ([ow Jl PCI ct:?al ?JP,? ;t’; ), YA-D5-E, (321) 867-6746 

Pnrficipatriig Orgaiirzatroii: USTDC (AV.N. Blalock, R T. Deyoe, A.J. Ecklroff, 

Dr. C. D. Immev, aid 1.7. Raizdazzo) 
Color-Indicating Patch for Shuttle Auxiliary Power Unit 

Hydrazine Transfer Line 

The Auxiliary Power Unit (APU) is a hydrazine-fueled power unit in the 

Shuttle hydraulic system. The APU transfer line consists of 1 /2-inch-outsidediameter stainless-steel tubing encased in a coaxial heater. An insulating glass 

fiber layer wraps around the tube heater assembly and is itself covered with 

LT-100 tape. Kapton tape is the outer layer wrapped around the LT-100 cover 

of the fibrous insulation. During routine maintenance of the APU system, 

evidence of microleaks was observed, as indicated by slight degradation of 

the glass fiber material. Although the leaks in the APU were small, it was 

essential to identify their location. Hydrazine is a highly corrosive and toxic 

substance, and it is a suspected human carcinogen. Currently, hydrazine leaks 

are monitored using the Interscan Hydrazine Monitor and/or Draeger tubes. 

To perform these tests, slits must be cut through the Kapton and LT-100 layers 

into the glass fiber insulation to extract air samples from the annular space 

for analyses. Both methods have major drawbacks. First, both methods are 

intrusive and require compromising the covering and insulation material of 

the APU transfer line. Second, both methods use an air pump to draw the 

sample, thus causing reduction of detection capability due to dilution of the 

hydrazine vapor by surrounding air. Third, it is labor-intensive to measure all 

potential leak locations since each location needs several minutes of analysis to 

accurately determine if there is a leak. Also, because of tight space, it is difficult 

to get to the location for a reliable measurement. A sensitive, easy-to-use, 

nonintrusive, in situ monitoring device of the colorimetric type could alleviate 

all of these difficulties. 

The purpose of the project is to develop a monitoring device for the 

identification of both the presence and location of microleaks of hydrazine in 

the Shuttle APU system. The device is a colorimetric indicator patch that can 

be incorporated into the APU system. The patch changes from pale yellow 

to purplish gray upon exposure to hydrazine liquid or vapor. The active 

ingredient of the indicator, potassium tetrachloroaurate (111), is loaded onto a 

porous 3 / 8-inch-diameter glass fiber or cellulose substrate. Loading levels are 

controlled by pipetting precise volumes of an aqueous solution of the indicator 

reagent onto the substrate. The loaded substrate is dried and attached to the 

I 

Side-by-Side Photographs of 3/8-lnch-Diameter Cellulose Disks Coated 

With the Colormetric Reagent Potassium Tetrachloroaurate (Ill) 

Ul7e left side shozos the color prior to exposure to hydrazine; the right 

side sliozus the color following exposure to hydrazine.) 
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stickv side of a 1-inch-by-3-inch piece of LT-100 adhesive tape and covered by 

a protective laver of waxy paper to form a “band-aid.” After removing the 

protective lcax paper cover, the “band-aid” can be easily applied to the glas 

fiber insulation of the APU system. Once applied, it needs onlv periodic visual 

inspectinn tc! dptprminp if there is a color change. The color chan3e induced 

on the indicator patch is readilv seen, even through the Kapton covering. 

If the color is unchanged, the patch does not need to be replaced. These 

colorimetric patches are a precise monitoring system to identifv the preGence 

and location of hydrazine leaks. In addition, since the patch assembly also 

uses the LT-100 tape, deplovment should not result in significant disruption to 

the integritv of the system or major redesign. To date, tests conducted in the 

lab sho~7ed the color dexreloped is irreversible and stable for at least 2 months. 

Tests also indicated the patch is stable xrhen stored in 40 degrees Celsius and 

under a vacuum of 10 - torr for at least 3 Iveeks. The evaluation of the patch 

for shelf life is ongoing. 

Key accomplisliments: 

Developed the colorimetric chemistry for the detection of hydrazine. 

Verified the color developed is irreversible and stable for at least 2 months. 

L’erified the indicator patch is stable at 30 degrees Celsim and under 

Designed the ”band-aid” for application. 

vacuum of 10.; torr. 

Kev milestones: 

Determine the shelf life of the indicator patch. 

Start using the colorimetric indicator patch in the Shuttle APU slistem. 

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a, IA Study of Hail Monitoring Strategies for Shuttle Launch Safety 

In Florida most precipitation reaches the ground in the form of raindrops, 

with drop diameters ranging from less than 1 millimeter (mm) up to 5 or 6 

mm. Occasionally, these hydrometeors are solid rather than liquid. Solid ice 

or hail occurs in diameters ranging from 0.5 centimeter (cm) or less up to 2 or 

3 cm or more. Hail is a rare event at the Kennedy Space Center, but it does 

occasionally occur. 

The example of hail damage on the surface of the External Tank (ET) shown 

in figure 1 was likely caused by a pea-size (0.5 cm) hailstone. Because of the 

potential of hail damage to the ET while exposed to the weather, it is desirable 

to remotely monitor hailfall in the vicinity of the Shuttle pad. If hail of 

sufficient size and quantity is detected by a hail monitoring system while the 

Shuttle is at the launch complex, the ET is thoroughly inspected for damage. 

For over 40 years, meteorologists have used simple, inexpensive pads of 

Styrofoam for recording hail dents and estimating hailstone sizes. ”Hailpads” 

have a proven record of reliably recording hail impacts down to the smallest 

size, while not recording raindrop impacts. 

A hailpad calibration method developed during the 1978 Alberta Hail 

Project relates the pad dent diameter d to the hailstone diameter D, as an 

empirical second-order polynomial: 

D = 0.38 + l.lld - 0.04d2 (1) 

where D and d are expressed in centimeters. Figure 2 shows a Colorado State 

University hailpad after impact with three 1.6-cm-diameter ice balls, fired by a 

paintball gun at the KSC Fire Training area. 

Even though the hailpad is the “gold standard” in measuring hailfall, it 

was desirable to implement a remote hail monitoring system at the Shuttle 

launch pad. Imaging hailpads with a video camera system was one method 

considered. Another approach investigated by KSC used a single piezoelectric 

ceramic disc mounted under a 1-square-foot flat metal plate. The initial 

concept was subsequently improved by forming a shallow pyramid structure 

(see figure 3) that encouraged hailstones to bounce away from the sensor so 

they would not be counted more than once. In addition, the final prototype 

included a mounting box for the piezoelectric ceramic disc, which was offset 

from the pyramid apex, thus improving response uniformity. 

One notable characteristic of this design is the potential to utilize frequency 

differences between the spectrums created from a raindrop impact and a 

hailstone impact. In other words, the sound of hail hitting a metal plate is 

distinctly different from the sound of rain impacting the same plate. This 

fortuitous behavior of the pyramid sensor may lead to a signal processing 

strategy that is inherently more reliable than one depending on amplitude 

processing only. 
Figure 1. Example of Hail Damage 011 External Tank 

Figure 3. Prototype “Pyramid“ Hail Moiiitor 

Figure 2. Colorado State Hailpad After Ice Ball Test 

Key accomplishments: 

Developed and characterized a hail monitor prototype - the “pyramid” 

sensor. 

Developed a hail simulation and testing procedure based on firing ice balls 

from a paintball gun monitored by a LabVIEW data acquisition system. 

Key milestone: 

Plan to implement a hail monitoring system for Shuttle launch safety based 

on the pyramid sensor with collocated hailpads for in situ calibration and 

verification. 

Coiltact: Dr. R.C. Youlzgquist (i: I i ’ 1 , - . ), YA-C3-E, (321) 867-1829 

Participatiizg Orgaizzzations: USTDC (W. 0. Haskell, Dr. ].E. Lme, Dr. C. D. li?zmer, aid 

R.B. Cox) mid Colorado State Uiiizlersity (Dr. N. Doeskeii) 
10 I 

Mass Sensor for Cryogenic Fluids 

Next-generation launch vehicles will likely utilize densified 

cryogenic propellants so greater amounts of fuel and oxidizer 

can be placed into onboard tanks. This will improve the 

performance of future vehicles by allowing them to fly farther 

or carry heavier payloads. However, much of this enhancement 

will be lost if the total mass of propellant cannot be accurately 

measured. Current limitations on measurement technologies 

require that tanks be filled well over the level calculated as 

necessary, resulting in a loss of payload capacity. Ideally, an 

accurate and robust total mass sensor is needed for future 

launch vehicles so the proper amount of fuel and oxidizer can be 

loaded to match a given mission's needs. 

Figure 1. Cryogenic Fluid Mass 

Sensor 

One mass sensing method uses an extended capacitor 

reaching from the bottom to the top of the tank. Instead 

of determining the liquid height with such a device, as is 

commonly done, a judicious choice of parameters and design is 

made to obtain a total liquid mass measurement. The resultant 

device (shown in figure l), demonstrated experimentally for 

liquid nitrogen and theoretically for liquid nitrogen, hydrogen, 

and oxygen, achieves a mass measurement more accurate than 

k1.0 percent (over a selected temperature and pressure range of 

interest). 

' The sensor is unique in that only a single measurement is 

needed to find the total mass of cryogenic liquid in a tank. Most 

cryogenic tanks do not have uniform cross sections, so normal 

capacitance-level sensors do not provide mass information. As 

illustrated in figure 2, the sensor is designed to compensate 

for noncylindrical tank shapes by varying the capacitance 

as a function of height. This is accomplished by physically 

shaping the inner conductor of a cylindrical capacitor so the 

capacitance is high where the tank area is large and low where 

the tank area is small. Also, the sensor produces a correct mass 

reading even if the cryogenic fluids stratify (i.e., if colder, denser 

fluid drifts to the bottom of the tank). This heavier fluid has a 

correspondingly higher dielectric constant and therefore yields 

a higher capacitance. Finally, if the tank contracts or expands 

thermally, the sensor output can be corrected easily by a simple 

multiplication. While a cylindrical capacitor was used to 

confirm the theory, other capacitive geometries could also be 

used to meet specific application requirements. 
SENSOR LEADS 

INNER CONDUCTOR 

--OUTER CONDUCTOR 

TEST CELL! TANK 

MASS SENSOR 

I 

ACTIVE 

SENSING 

REGION 

1- 

TEFLON SUPPORT 

W I FLUID ACCESS PORTS 

Key accomplishments: Mass vs. Capacitance (711 1/02) 

Designed and executed a proofof-concept experiment where 

an insulated test cell was filled 

with liquid nitrogen and set on 

a scale (figure 3). A LabVIEW 

program was used to acquire 

data from the sensor and from the 

scale at 1-second intervals as the 

liquid nitrogen boiled off (figure 

4). Theoretical development 

is continuing and matches the 

experiment to within 2 percent 

with the expectation of improved 

accuracy. 

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Coiitacts: M.A. Nurge (Mnrk.A.Nurgt~ 

@iz~s~.~~~~~), YA-C3-€, (321) 861-9068; gild 

Dr. R.C. You~~gqrlist, YA-C3-€, (321) 

867-1 829 

11 
Shuttle Tile Moisture Sensor 

12 

Much of the surface of each Space Shuttle is protected from the heat of reentry by a 

high-temperature, reusable insulation commonly known as Shuttle tiles. These tiles are 

composed of 99.7-percent pure silica fibers that form a strong material with extremely 

low thermal conductivity. Each Orbiter has more than 20,000 tiles attached to it forming 

a thermal protection system that can withstand not only temperatures as high as 2,400 

degrees Fahrenheit but also the acoustic levels (up to 166 decibels) common during launch. 

In particular, the tiles across the bottom of the Orbiter experience the greatest temperatures 

and are coated with a thin (about 0.01-inch) layer of borosilicate glass mixed with a silicon 

tetraboride emittance agent (i.e., black glass). This black glass coating raises the emittance of 

the tile above 0.8, allowing the majority of the heat of reentry to be radiated away from rather 

than conducted down to the skin of the Orbiter. 

The tiles are lightweight because the silica fibers only fill 7 percent of the tile’s volume. This 

is an advantage for a launch vehicle but raises an operational difficulty in that the fibers act as 

extremely absorbent sponges. Efforts are made to minimize the amount of water in the tiles 

on an Orbiter because water not only adds to the launch weight but can also freeze in space, 

possibly leading to serious tile damage and Orbiter damage on reentry. 

During fabrication, every tile is treated with the waterproofing agent dimethylethoxsilane 

(DMES). This material coats the silica fibers, making them hydrophobic instead of hydrophilic 

(i.e., causing the fibers to repel water rather than absorb it). This waterproofing system 

is effective, but the DMES burns out of the tile during Orbiter reentry. During Orbiter 

processing, each tile is injected with replacement DMES through a small hole (about 1 

millimeter in diameter) in the black glass coating, effectively rewaterproofing the tiles. 

From reentry until the Orbiter is inside the Orbiter Processing Facility (OPF), the tiles are 

susceptible to water intrusion. Cracks are common in the black glass coating and are believed 

to provide a primary route for water to be drawn into the tiles. Rapidly protecting a recently 

returned Orbiter from adverse weather and other sources of water is a top priority. 

However, it is not always possible to protect the Shuttle from these adverse conditions, 

and water has entered the tiles on several occasions. During vehicle processing, the water 

in the tiles is removed using heat 

lamps. An infrared camera is used 

to determine whether the water 

was sufficiently removed. The 

disadvantage of using an infrared 

camera to evaluate the presence of 

water is that the technicians have 

to wait several hours after the heat 

lamps have been disconnected 

before they can take readings. 

Furthermore, water must be close 

to the tile’s surface for the infrared 

camera to detect it. 

To overcome the limitations of 

the infrared camera, an alternative 

method of water detection based Tile Moisture Sensor 
on capacitance measurements was developed. 

The objective of designing the Shuttle Tile 

Moisture Sensor was to develop a tool capable of 

detecting small amounts of water in a tile. This 

was accomplished by taking advantage of the 

differences in permittivity (dielectric constant) 

of the materials involved. The permittivity of a 

material determines the amount of charge needed 

to create a voltage across the material; therefore, 

the capacitance presented by a material of a certain 

geometrical configuration will increase as the 

permittivity of the dielectric increases. Air has a 

permittivity close to unity, while freshwater has a 

permittivity of about 80, although it depends on 

temperature to some extent. Most nonconducting 

materials have relatively small dielectric 

constants with values typically less than 10. Oils, 

for example, have a value of about 2.0, silica 

glass about 4.0, Teflon about 2.0, and cryogenic 

fluids around 1.5. An accurate measurement of 

capacitance can be used to determine the presence 

of dielectric materials affected by the electric 

field generated by the capacitor’s electrodes. 

The effect that a dielectric has on the measured 

capacitance depends on various factors, such as 

the geometrical configuration of the dielectric, 

its distance from the capacitor’s plates, and the 

presence of other dielectrics or conductors nearby. 

I 

The operation of the new Shuttle Tile Moisture 

Sensor is based on the application of a small and 

fixed current to the electrodes composing the 

capacitive element. The duration of the current 

application is selected to prevent overvoltage 

saturation of the analog circuitry and the analogto-digital converter due to the charge injected into 

the capacitor. Following the application of the 

currrent, the voltage at the capacitor is read. The 

resolution of the voltage measurement is increased 

by taking multiple readings (256 in this case) with 

different durations of the charge current (typically 

125-nanosecond difference between readings) to 

remove the effects of the quantization noise at the 

analog-to-digital converter. This ensures that the 

voltage applied to the analog-to-digital converter 

is not identical for every measurement. Otherwise, 

the quantization noise would still be present and 

would limit the resolution of the capacitance 

measurements. 

1 

The Shuttle Tile Moisture Sensor is 

calibrated to remove the effects of known 

dielectrics, such as the Space Shuttle tiles 

themselves. Since the permittivity of 

the material used in the tiles is not unity, 

the moisture sensor detects a change in 

capacitance as the probe is placed against a 

tile. Preset calibration values are loaded in 

the sensor for various tile depths against the 

metal frame of the Orbiter. 

To achieve the resolution and accuracy 

required, several features were implemented. 

Variable charging times were incorporated 

for successive measurements of the voltage 

developed due to the charging of the 

capacitive elements. An autozero was built 

in to remove the effects of the capacitance 

introduced by the wiring, circuit board, and 

active elements. A very stable active-current 

source capable of sourcing a submicroamp 

curreni was developed. 

The prototype Shuttle Tile Moisture Sensor 

has several features that make it easy to use. 

A liquid crystal display allows the operator 

to select the thickness of the tile being tested 

(adjustable range) as well as to display the 

result of the capacitance measurement. 

Audible and visible alarms with preset 

thresholds make the device easier to operate. 

A 9-volt battery powers the tool. 

Key accomplishments: 

Developed an advanced prototype Shuttle 

Tile Moisture Sensor in conjunction with 

Shuttle Operations. 

OPE 

Currently field-testing the device in the 

Contact: Dr. R.C. Youngquist (R0bert.C. Yoioungquist@ 

~ln~n.goi>), YA-C3-E, (321) 867-1829 

Participating Organization: USTDC (Dr. P.J. 

Mcdclius, J.D. Taylor, J.J. Randnzzo, and J.A. Rees) 
New Gas Polarographic Hydrogen Sensor 

Polarography is the measurement of the current that flows in solution as a 

function of an applied voltage. The actual form of the observed polarographic 

current depends on how the voltage is applied and on the characteristics of 

the working electrode. The new gas polarographic hydrogen (H,) sensor 

shows a current-level increment with concentration of the gaseous H, 

similar to those relating to metal ions in liquid electrolytes in well-known 

polarography. This phenomenon occurs because the diffusion of gaseous H, 

through a gas diffusion hole in the sensor is a rate-determining step in the 

gaseous hydrogen sensing mechanism. The diffusion hole artificially limits 

the diffusion of the gaseous H, toward the electrode located at the sensor 

cavity. 

This gas polarographic H, sensor is actually an electrochemical pumping 

cell since the gaseous H, is pumped via the electrochemical driving force 

generated between the electrodes. Gaseous H, enters the diffusion hole, 

reaches the first electrode (anode) located in the sensor cavity, and is 

transformed into H ions or protons. The H' ions pass through the electrolyte 

and reach the second electrode (cathode) to be reformed into gaseous H,. 

Having a gas binary system and assuming steady-state conditions, no 

reaction, stagnant background gas, and mass transfer in only one direction 

and rate-limited at the diffusion hole, the limiting output current is predicted 

to be a linear function of the logarithm of the background gas concentration. 

This linear relation between the output-limiting current and the logarithm 

of the background molar fraction has a slope that includes the area and the 

length of the gas diffusion hole, the gas pressure, the gas temperature, the 

diffusion coefficient of the gaseous specie in the binary gas mixture, the 

Faraday constant, and the number of electrons per mole of the gaseous specie 

transferred between electrodes (this number is 4 for oxygen [O,] and 2 for H, 

as indicated in figures 1 and 2, respectively). 

Heater Diffusion Hole Diameter: 60 tc 70 prn 

t I 0, + 4e - 202- 

\ 1 Platinum 

YSZ Electrolyte 

202 -02+4e 

Figure 1. Schematic Diagram of Electrovac Gas Polarographic 0, Sensor 
Gas polarographic 0, sensors 

are commercially available. A 

commercial gas polarographic 

0, sensor was used to prove the 

feasibilitv of building a new gas 

polarographic H, sensor. As 

expected, the output-limiting 

current proved to be dependent on 

H, composition; the output-limiting 

current is increased as the gaseous 

H, concentration increases. As in 

the case of gas polarographic 0, 

sensing, a linear relation between 

the output-limiting current and the 

logarithm of the mole fraction of the 

background gas was also observed 

for gas polarographic H, sensing, 

as shown in figure 3. Further test 

runs and precise measurements 

of the diameter and the length of 

the diffusion hole are required to 

cnrrdxrate this !ine.r re!aticx?. 

Key accomplishments: 

Sensor was presented at the 

2nd Institute of Electrical and 

Electronics Engineers (IEEE) 

International Conference on 

Sensors. 

Experimental tests fulfilled 

gaseous hydrogen sensing 

criteria, very fast response, 

good accuracy, and reasonable 

repeatability. 

Sensor's similarity to existing 

and commercially available 

oxygen sensors led to the 

potential use of both sensors to 

enhance performance of both 

sensors when the gas has both 

components - oxygen and 

hydrogen. 

Heater Diffusion Hole Diameter: 60 to 70 prn 

It 

Anode 

fl 

Platinum / 

Electrodes 

2H +2e .H, 

t 

Figure 2. Sclieiiinfic Diagram of a Neir Gas Polaroc~vnpiiic H, Seiisor 

Gas Polarographic H, Sensor: Predicted vs. Experimental Performance 

Gas: H, in N,, 27 "C, 768 Torr. 6000 sccm. May 2002 

000016< 

, 

I OOOO14 

= 000012 

5 

g! 

5 000010' I 6 000008~ 

a 000006! 1 Experimental I 

- 

I 

I 

I I 

0 00004 i 

~ 

~ - Predicted , 

0 0001 0002 0003 0004 0005 0006 0007 0008 0 

-Log (1-YH2) 

9 
Improved Fiber-optic Pressure Block Connector 

When using fiber-optic cable for transmitting video signals from a 

positively pressurized camera housing, it is necessary to use a leaktight fiberoptic connector to prevent loss of pressure from the housing. 

A search for a commercially available connector to meet this requirement 

found very expensive items designed for high-vacuum service and bulkhead 

feed-through items that were not suitable because of the requirement to 

disconnect the fiber-optic cable from the camera housing. The search turned 

toward developing a simple low-cost connector using commercially available 

components that can be easily modified or used as is. 

The main component selected was a $15 standard part, KC-124-KlO 

bulkhead fitting, modified to accept a potting boot on one end and a standard 

FC-type adapter / connector on the other end. To block a leakage path from 

inside the cable jacket, a 3/8-inch section of the jacket is stripped, exposing the 

cladding inside the potting boot before filling the boot with polyurethane. 

An adapter ring is required to mount an FC-type adapter / connector on 

the outside end of the KC-124-K10 bulkhead fitting. A stainless-steel inverted 

nut is welded to the potting boot end and, with a brass ferrule and a 1 / 8-inch 

"B" nut, provides a slight compression to the fiber-optic cable to prevent the 

polyurethane from leaking into the body of the KC-124-K10 bulkhead fitting. 

These fiber-optic pressure block connectors are used in the Operational 

Television (OW) Modernization Camera Housings. The complete design is 

documented on KSC Drawing 79K31859, Fiber-optic Pressure Block. 

Contact: R.A. Stute (Rohc~f.A.Stir te(?~ns~.goi~), YA-E2, (322) 867-3985 

Participating Organizations: USTDC (T. D. Greenfield, C.G. Hallberg, and J.D. Polk) and 

United Space Alliance (R. D. Gibbons) 
I Improved Fiber-optic Pressure Block Connector 

17 
Common Application Framework (CAF) 

The Common Application Framework provides both common services and a layer of 

abstraction that insulates the applications developer from the platform details. The CAF 

accomplished this goal with a set of object-oriented components, base classes, abstract 

patterns, and concrete behaviors / services implemented using C++. The CAF allows the 

application developer to focus on the requirements and business logic necessary for the 

application. In addition, the framework makes extensive use of generic programming 

concepts (i.e., templates), mature open-source software packages, and industry-standard 

interfaces that make the framework easy to use and easy to extend. 

One of the most significant features of the CAF is the embedded concurrency model. The 

CAF concurrency model insulates the application developer from the rigors of making an 

application threadsafe by enforcing serialization at an object level instead of an attribute 

level. This allows an object to be created as either an active or passive object to determine 

the thread of execution associated with the object. To support the concurrency model, the 

framework provides a pool of managed threads and an event-driven model. The combined 

functionality of the managed threads and the event processing means that one thread of 

execution can block waiting on an event, without fragmenting the application code to provide 

the multiplexing capability (i.e., signal handlers and switch statements). The CAF solution 

provides a threadsafe run-time environment and the means to produce software more quickly, 

with fewer experienced developers in a shorter time. 

Half-Sync/ 

Half-Reactive 

I 

I 

vent Demuxing 

nd Dispatching 

-_____ 

I 1 I 

I Queueing Layer 

I Implementation 

I Handler 

1 Types 

I 

I 

I 

I 

I 

I Access 

I Control 

I IntegratedProxy 

Service Handler 

Object Connector 

I 

I 

_ - - _ _ - - 

I I 

I I I 

I 

I 

I 

I I 

I I I 

Synchronization 

Locking I 

I 

I 

_ 

I 

I 

I 

I 

I I Thread ~ Lock 

I Safety 1 Acquisition I 

I I and Release I 

I I I 

I I I 

I 

Locking I 

I 

I 

I Lower-Layer 

1 Abstractions 

I 

I 

I 

I 

I 

i Service Handler 

I Concurrencv 

I 

I Locking Facade ! Mechanism 

I I Encapsulation 

I I 

I 

I i Thread Safety _________- 

Interface 

I Locking Mechanism 

I Encapsulation 

Facade 

Coimnoiz Application Fralnezoork (CAF) Architecture 
The CAF also supports a data-driven initialization model. This means that the object 

composition, initial values, and associations can be derived from instance data when the 

application is initialized as opposed to hard-coded. 

The features of the CAF provide abstraction, commonality, reliability, and increased code 

reuse that lead to enhanced software quality and safety. 

The following are benefits of the Common Application Framework 

Provides an embedded concurrency solution with serialization at the object level and 

Automatically detects deadlock and race conditions. 

Minimizes the time required to design and implement applications. 

Provides the concept of cancellation points (points in the software where a check is 

managed threads of execution. 

performed to determine if the application should continue processing in the current 

thread of execution). 

Provides an event-driven model. 

Supports both client and server applications. 

Provides a number of easy-to-use patterns for both monolithic and distributed 

applications (e.g., publisher subscriber, command, active object, singleton). 

Provides built-in IPC interfaces for CORBA, POSIX, SV5, and Sockets. 

- Uses iildiistr- standards (CGRBA, SOL, ODBC, POSTX, and C++). 

Makes extensive use of open-source packages (ACE, TAO, M4, Flex, and Bison [TAO 

can be replaced with any CORBA-compliant ORB and Name-service]). No licensing is 

required and documentation is readily available on the Web and in many books. 

Supports multiple platforms (e.g., Solaris - SPARC and Intel, Linux) and can be built with 

an ANSI/ ISO-compliant C++ compiler. 

Promotes consistency within a development project. 

Provides a high-level syntax for procedural control within a thread of execution: - 

- 

- 

- 

Read: Reads the value of a specific data point. 

Write: Modifies the value of a data point. 

Verify: Checks a list of data points for a specific value or list of values. 

Verify Within: Checks a list of measurements for a specific value within a specific 

time range. 

Delay: Pauses execution for a specified time. 

Delay Until: Pauses execution until the specified condition is met. 

- 

- 

Minimizes the risk associated with inexperienced developers building concurrent 

applications. 

Key accomplishments: 

Validated Common Application Framework's support of hundreds of requirements. 

Framework used by Hazard Warning System CSCI to validate software that performs the 

Framework software used in test of Flight Controls CSCI to perform end-to-end 

Orbiter Processing Facility checkout against a math model. 

performance tests at Shuttle Avionics Integration Laboratory. 

Contncf: D.A. Nipfell (Dnizicl.A.."\Tictciz~'i7nsn.goi~), YA-D8, (321) 861-2233 

Pnrficipnting Orgalzizatiolzs: USTDC (R. Nieves, A. Crikis, and G. Poueda) and United Space Alliance 

(A.A. Ciazmella) 
Automated Test Tool 

The Automated Test Tool for C++ automates the execution 

and evaluation of software tests, including the logging of pass/ 

fail results. In addition, test driver and script templates are 

automatically generated from C++ header files that are readily 

extensible to other languages. Since the tool encourages a datadriven test design combined with scripting languages, changes 

to test cases can be made easily without the need to recompile 

the application under test (AUT). Also, the AUT can continue 

executing while the test scripts are modified and debugged 

on the fly. If a repository or database is available containing 

information about the data under test (e.g., minimum/ 

maximum values, enumerated values), the Automated Test 

Tool can readily interface with that data store and automatically 

incorporate the data directly into the test cases. 

The Automated Test Tool provides a library of test procedures 

that enable the automatic execution and termination of AUTs 

along with many other predefined methods. Furthermore, the 

Automated Test Tool is able to wrap other test tools, such as 

performance test tools, and launch and drive those tools during 

execution. 

Benefits of the Automated Test Tool include the following: 

Minimizes the time required to test software. 

Allows testing to be run incrementally. 

Removes drudgery of building the test scaffolding associated 

Allows for a rapid buildup of a test suite that can then be used 

Built on top of TCL and Expect - two mature, freely 

with the AUT. 

for regression testing. 

available computer languages. No licensing is necessary and 

documentation is readily available on the Web and in many 

books. 

Eliminates the monotony associated with testing. 

Key accomplishments and milestones: 

Automatic validation of hundreds of requirements for an 

advanced real-time control framework completed in 20 

minutes. 

of the real-time control framework. 

Identification of errors in framework during regression testing 

Contact: G.S. Semmel (Gleii1i.S.Seinmel@nasa.gov), YA-D8, (321) 861-2267 

Participating Organizations: United Space Alliance (K.A. VO and M.C. 

Michel), USTDC (B. Pelakh), IT-D2 (T. W. Merrill), and Aerospace Design 

and Development 
... 

Now testing use case parseStlString ... 

SRS Name --> 

parseStlStringTokenize 

Enter a delimiter character: 

B 

Enter a string that contains no more than one instance 

of delimiter 

character: 

ABC 

AC 

Enter a delimiter string: 

<delimiter> 

Enter a string that contains no more than one instance 

of delimiter string: 

A<delimiter>Z 

Az 

SRS Name -> 

parseStlStringCombine 

Enter the first string: 

CLCS 

Enter the second string: 

World 

CombineStringToken: CLCSXWorld 

SRS Name --> 

PASS: svpParseStlStringTokenize 

PASS: svpParseStlStringCombine 

... 

~ 

Test Run By cma on Wed Jun 12 15:22:56 2002. 

Hostname is ide2ccp2.ksc.nasa.gov. CCWS host 

is ide2hci2. 

PASS: setup spawn-remote-shell. 

PASS: setup launch-application 

PASS: svpLogging 

PASS: setup shutdown-Eim 

Now testing use case publishstate ... 

PASS: svpPublishStatelnitializing 

PASS: svpPublishStateldleAtEndlnit 

PASS: svpPublishStateSupporting 

Now testing use case initiateDisplay ... 

FAIL: SvplnitiateDisplay 

Now testing use case parseStlString ... 

Now testing use case logging ... 

PASS: svpParseStlStringTokenize 

PASS: svpParseStlStringCombine 

PASS: svpConcurrencySingleEntryPoint 

PASS: svpConcurrencyMultipleEntryPoint 

Now testing use case concurrency ... 

... 

_--_ Summary === 

# of expected passes 284 

# of unexpected failures 1 

Screen Shot of Tool Ruiiriing Snmple qfSun2mnvy Log Output 
Digital Image Lens Distortion Correction Algorithm 

Photogrammetry is defined by webster.com as “the science of making reliable 

measurements by the use of photographs and especially aerial photographs (as in surveying).” 

Two or more cameras are generally used to image objects, such that multiple views provide 

information on depth of terrain, for example, or any other three-dimensional (3-D) position 

feature of interest. In these cases, the distance from the cameras to the object (range) is much 

greater than the size of the object (field of view [FOV]). The cameras, which may be flying 

onboard an airplane or satellite, use long-focal-length lenses (or telescopic lenses). Under 

these conditions, the image in each camera appears flat, so all rays connecting the camera 

lens to parts of the imaged object are essentially parallel in the vicinity of the object. This is 

described by the so-called affine camera model. 

Another form of photogrammetry, developed at KSC to make accurate 3-D position 

measurements, uses a single camera and high-contrast target (see ”Surveying With a Digital 

Camera System,” KSC Research and Technology 2002 Annual Report). In this case, the 

rays connecting the target to the camera focal point are not parallel. Complete 3-D position 

coordinates, including depth information, are derived from the imaged target by means of the 

perspective (pin hole) camera model. In this application, the target range is comparable to the 

camera’s FOV. This comes about by a short-focal-length (f> lens. 

Figure 1. Before (top) and After (bottom) 

Application of Lens Correction Algorithm 

Figure 1 (top) shows an outdoor scene 

taken with a commercial digital camera, using 

a short-focal-length setting (wide FOV). On 

close examination, a noticeable curvature can 

be seen along the horizon as well as the pole 

to the left (the pole is actually straight and 

perpendicular to the ground). This “fisheye” effect is a common occurrence in camera 

images and worsens as the FOV increases 

(fdecreases). Note that the pole does not 

appear perpendicular to the ground. This 

can be explained by the perspective camera 

model. The farther a point is from the camera, 

the closer to the center of the image the object 

will appear. In the case of the pole, the top is 

farther from the camera than the bottom, thus 

giving the appearance of a leaning pole. 

The curvature of the pole, as well as the 

curvature of the horizon, is a consequence 

of lens distortion. Lens distortion can be 

thought of as the error produced in a camera 

imaging system that results in deviations 

from that expected from the ideal perspective 

camera model. As can be expected, lens 

distortion causes severe degradation of 

position measurement accuracy in singlecamera photogrammetry. 
A simple empirical lens distortion model considers the distortion effect as proportional to 

the square of the distance from the center of the image, where the image center is coincident 

with the camera optical axis. A lens correction formula is then based on modifying the kth 

image pixel position pk of the original image, using the following formula: 

where a f / d <f is the lens focal length and dc is the charge-coupled device [CCD] pixel 

width), and p:Cis the new pixel coordinate in the corrected image. 

The lens distortion correction factor p can be found by either trial and error or a calibration 

procedure using an xy table. By plotting uncorrected camera data versus the xz-table 

coordinates as shown in figure 2, where the subscript 172 denotes measured coordinates (from 

camera photogrammetry) and the subscript t denotes xz-table coordinates, P is determined by 

the slope of the resulting linear fit. 

Key accomplishments: 

Developed a simple lens distortion correction algorithm for single-camera photogrammetry 

De1.7ised a !ens cdikratim! xr,eth~c! .;sing iiri xy table to deiei-inirle lens correction factor for 

applications. 

any given lens. 

Key milestone: 

Plan to integrate lens correction 

algorithm in Payload Ground 

Handling Mechanism 

(PGHM) position measurment 

photogrammetry system. 

10’ 30 1 I I I I I 

KTGLC 8 inin Tamron _- 

Figure 2. Defevrniiziiig Empirical Len5 Calibratioii 

Factor p From sy-Table Data 
Real-Time Detection and Tracking of Multiple Moving Objects 

in Complex Image Background 

Motion detection is an important process for many machine vision systems. Systems that 

analyze aerial images, design special correlation chips for real-time applications, track moving 

objects, perform robot vision, detect moving targets, analyze traffic flow, or guide autonomous 

vehicles are examples of machine vision systems. The main goal of a motion detection system 

is to detect and track only structural changes in the scene. For machine vision to operate 

successfully, fast, accurate, and robust motion detection algorithms are needed. In the past, little 

work was directed toward solving this problem for time-varying outdoor scenes. 

Motion detection is generally performed at the pixel level, at the edge level, or even at higher 

feature levels. Existing feature-level algorithms for time-varying outdoor backgrounds require 

more computational efforts than those working at the pixel level, but they provide more accurate 

information for higher system modules. We have implemented an innovative motion detection 

system that performs above the pixel level, provides accurate information, and surpasses the 

robustness of the existing high-feature-level algorithms while requiring less computation time. 

The system was originally tested to detect potential foreign object debris (FOD) during the 

Space Shuttle liftoff. The image background during the Space Shuttle liftoff is quite complex 

because moving objects that are not FOD (ice flakes, fume clouds, external fuel line connections, 

etc.) constantly appear in the scene that is taped in black-and-white format. The system was able 

not only to detect two FOD items observed in the tape images but also to keep close track of both 

items (see figures 1 and 2). Figure 3 shows the trajectory for both items; this trajectory was drawn 

in real time as the FOD was detected. In regard to the computation time, the system was also able 

to process and execute 

motion detection 

features on every one 

of the 30 frames per 

second generated by 

the image acquisition 

system. 

This real-time 

moving-object 

detection system 

is currently in its 

preliminary stage. 

The system concept 

and algorithm have 

proven to be robust. 

Implementation and 

validation are still 

needed. The current 

system works using 

a black-and-white 

image format. Images 

acquired in color will 

lead to an even more 

b, ,& - *, 

F 

Figure 1. lmage Sequence Shozoing an Object Moving From Top Center (Close to the 

Shuttle's External Tank) Toward the Bottom Left (Close to the Shuttle's Tail) 
Figure 2. Image Sequence Showing an Object Moaiizg Froin Top Left (Close to the 

Shuttle's Eizgine Exhaust) Toulard the Boffoin Left (Close to the Sl~uttlc's Tail) 

robust system that is not necessarily slower because a multiprocessing CPU 

approach, where individual CPUs would independentlv process each primarv 

color, is quite feasible. 

The potential applications for real-time moving-object detection and 

tracking objects in complex image backgrounds include analysis of aerial 

images, dynamic surveillance, robot vision, detection of moving targets, 

analysis of traffic flow, and guidance of autonomous vehicles. 

Key accomplishments: 

Used as a tool to support the STS-107 investigation. 

Nominated to represent KSC during the 2003 NASA 

Patent submission currently being considered. 

Two of NASA's support contractors (Research 

Triangle Institute and Southeast Regional Tech 

Transfer Center) are showing this software as an 

actual application of image technologies that are 

being currently marketed. 

Software of the Year competition. 

Accepted for publication in NASA Tech Briefs. 

Contact: C.K. Dnws (< 

(321 ) 867-8804 

), YA-€2, 

Pnrt icipat iiig O rgnn izntioi I : US TDC (1. A. Don? iiigu~z nnd 

S.J. Klinko) 

Figure 3. Real-Time Trnckiizg Path of the 

TZOO Moziing Objects (Figures 1 and 2) 
A Digital Signal Processing (DSP) Filter Tuned To Detect Diurnal Signals 

The predominant mesoscale feature across the Florida peninsula during 

the summer months is the land-sea breeze oscillation (LSBO). Intense solar 

heating during the day creates a thermal contrast between the temperature of 

the air over land (Ta) and the temperature of the air over water (Tw). When 

Ta sufficiently exceeds Tw, a thermally induced direct circulation occurs in the 

lowest few kilometers of the atmosphere. Near the surface, air flows from the 

cooler air mass over water to the warmer air mass over land, whereas a return 

circulation aloft flows from land to water. As the day progresses, a boundary 

interface or sea breeze (SB) "front" often develops at the leading edge of this 

circulation and advances inland. Air rises along this leading edge while air 

descends above the water, completing the direct circulation. At night, the land 

breeze (LB) completes this diurnal cycle. 

The processing strategy behind an SB detector is based on the diurnal 

periodicity of the signal. Using a high-order band-pass (BP) filter, based 

on cascading second-order quadratic sections of figure 1, and coefficiencts 

corresponding to a maximally flat Butterworth response, an SB detection time 

filter is constructed as illustrated in figure 2. The BP filter provides an SB 

transition time predictor, which is compared with the low-pass (LP)-filtered 

wind direction signal at every spatial grid point. If the difference between the 

predicted BP-SB time and the LP-SB time exceeds 6 hours, no SB for that day is 

assumed. 

The LP filter is implemented as a moving average 

filter. The BP filter is a zero-phase recursive filter with 

Q = 2, made up of four second-order sections (i.e., 

an eighth-order filter). The center frequencyfi) is set 

to match the 24-hour LSBO period. The zero-phase 

infinite impulse response (IIR) filter is implemented 

by summing the outputs of two filters with identical 

characteristics, where one filter processes a block of 

data forward in time, and the other filter processes 

the data backward in time from the end of the block. 

Figure 3 shows sample output using 5-minutex = 53, y = 47. Note that the SB filter algorithm fills 

in missing data by performing a linear interpolation 

Figure Dsp 'IR Second-order Section duration data samples for July 2000 at grid location Network Diagram 

between end points around the missing data before the LP and BP sections. 

Figure 4 displays the SB filter outputs for all grid points on July 6,2000. 

Figure 5 shows the contour of the transition time plot; its gradient is inversely 

proportional to the SB front-propagation (advection) velocity. 

Key accomplishments: 

Completed processing of 5-minute-duration samples of tower data observed 

Developed an the SB prediction filter using a recursive BP and nonrecursive 

in a continuous 24-hour period to develop a time-filtering technique. 

LP filter. 
- I Comoare 

1.5b 

LP Filter and 

Filter BP Filter 

Signals. 

Output SB 

i I 1 8 I I I 

X-y Grid: {53, 47) 

1 

I Timesand 

Direction Band-Pass 1 Type Codes 

Filter 1 for the Mh 

Figure 2. Sea BreezelLaizd Breeze Detectioiz Algorithm 

I 

Interpolated Sin (Wind Diredon) 

- LP Filer 

0P Filer 

'I SBlimes 

-15c I I I I I 1 I I I* 

300 400 500 600 700 800 900 1000 1100 

Day of July 2000 (UTC) 

Figzdre 3. Sen Breeze Traizsitioii Times, Based oiz Algoritiziiz Froin Figure 2 

Figure 4. SB Transifioii Times Spatial Plot Froin 

Wiiid Tower Data 

Examined output of SB filter 

for all days in July and August 

2000 and compared to analysis 

of daily SB transition times; 

agreement between the SB 

prediction filter and subjective 

SB identification was excellent. 

Coiitact: Dr. E]. Merceret ( 

), YA-D, (321) 

867-0818 

Participatiiig Orgmzizatioiis: USTDC 

fDr. ].E. Lone oizd Dr. C. D. Iinmer) 

aid ENSCO (J.L. Cnse aid Dr. 1.7. 

Manobiaim) 

Figure 5. SB Adzvctioiz Velocity 
Image Edge Extraction via Fuzzy Reasoning 

Edges carry the most important information in images, and accurate 

edge detection is vital to perform advanced image processing and analysis. 

Images of real scenes frequently contain ambiguous and incomplete data. In 

particular, the problem of determining what is and what is not an edge is 

compounded by the fact that edges are often partially hidden or distorted 

by various effects such as uneven lighting and image acquisition noise. 

Furthermore, images frequently contain data with edgelike characteristics, 

but a confident classification of this data can be best solved when high-level 

constraints are imposed on the interpretation of an image. 

The fuzzy reasoning approach we implemented to detect image edges 

significantly surpasses current and widely used edge detection methods. 

An image of a CD was taken as a sample to compare the methods; the fuzzy 

reasoning approach shows edges with tiny details that are not detected or 

properly weighted by other methods. Figure 1 shows the original CD image; 

figure 2 shows the edge detection performed on the CD image (figure 1) using 

this new edge detection technique. The technique clearly detects a scratch in 

the center of the CD and hidden inner circles; it also provides the appropriate 

weight values to the generated edge map. 

Potential uses for this edge detection approach are wide open. It has 

potential to be a key part of many computer vision applications, including 

but not limited to pattern recognition, anomaly or defect detection, tumor 

detection via medical imaging, surveillance, visual motion control, face 

recognition, object tracking, handwriting recognition, and robotic vision. 

Key accomplishments: 

Patent submission being considered. 

Working with two of NASA's support contractors (Research Triangle 

Institute and Southeast Regional Tech Transfer Center) to market this image 

technology and generate leads for licensing. Companies in the medical, 

remote sensing, and biometrics industries will be contacted about the 

technologies. 

Accepted for publication in NASA Tech Briefs. 

Successfully used to preprocess images required by an image analysis 

application, an automated visual anomaly detection system. 

Contact: C.K. Davis (Christopher, K. Davis@nasa.gov), YA-E2, (321) 867-8804 

Participating Organization: USTDC (].A. Dominguez and S.J. Klinko) 

28 
Figure 1. Origiiial CD linage 

Figure 2. Edge Extracfioiz Based 011 Fiizzy Reasotziizg 

Reveals a Hidden Scratch oiz the Ceizfer of the CD 

29 
Noniterative Optimal Image Binarization via Fuzzy Reasoning 

Binary image processing is of special interest because an image with binary 

format can be processed with very fast logical (Boolean) operators. Since 

most image display systems and software assume images of 8 or more bits 

per pixel, the binarization of these images usually takes the two extreme 

gray tones, black and white, which are ordinarily represented by 0 and 255, 

respectively, in an 8-bit gray-scale display environment. 

Usually a binary image is obtained from an 8-bit gray-level image by 

determining a threshold and assigning the low (0) and high (255) binary 

values to all gray levels based on the chosen threshold. Obviously, the 

threshold chosen is critically important since it controls the binary-based 

pattern classification obtained from the gray-level image. The key issue is to 

choose an optimal threshold so the number of misclassified image pixels is 

kept as low as possible. 

A new fast-computational technique based on fuzzy entropy was 

developed to find an optimal binary image threshold. In this method, the 

image pixel membership functions are dependent on the threshold value 

and reflect the distribution of pixel values in two classes; thus, this technique 

minimizes the classification error. The proposed method of choosing a 

threshold surpasses the Huang-Wang and Otsu methods, previously the most 

reliable and fastest thresholding methods, when the image contains textured 

background and poor printing quality (see the figures). The new technique 

provides a much simpler, faster, and more reliable binary segmentation tool 

for optical character recognition (OCR) application and texture analysis. Since 

the proposed method requires relatively little computation time (compared 

with Huang-Wang and Otsu methods), it might be used as a binarization 

operator within real-time image analysis systems. 

Key accomplishments: 

Patent submission currently being considered. 

Currently working with Research Triangle Institute and Southeast Regional 

Tech Transfer Center to market this image technology and generate leads 

for licensing, targeting companies in the medical, remote sensing, and 

biometrics industries. 

Accepted for publication in NASA Tech Briefs. 

Successfully used to preprocess images required by two different image 

analysis applications, a moving object detection/ tracking system and an 

automated visual anomaly detection system. 

Received Space Act Award granted by NASA Inventions and Contributions 

Board. 

Contact: C.K. Davis (Christoplrc7r.K. Dai1is@iznsn,gov), YA-E2, (321) 867-8804 

Participating Organization: USTDC (J.A. Dominguez and S.J. Klinko) 
Original Image I 

Otsii Biizarizatioii 

Hunizg- Wniig Binnrizatioii 

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Fuzzy-RensoiziiigBaserI Biiinrizatioii 

31 
Converting Analog Test Data to Audio Wave Files 

One common method of analyzing any type of data is to first plot the raw data to view its 

general characteristics and prominent features. Following that initial investigation, the data 

analyst may implement standard data evaluation algorithms, such as linear or polynomial 

regressions, data smoothing or filtering, and Fourier transforms to view data features in the 

frequency domain. All of these data analysis methods have one thing in common - they use 

our vision analysis system (eyes) to subjectively evaluate data by examining a graphical plot. 

An automated computer image processing system can be devised to perform a similar type of 

analysis, without actually creating a plot image. 

Another investigative technique that may be equally useful to the data analyst uses our 

audio analysis system (ears). However, most computer utilities (such as those based on PC 

office software and spreadsheets) have little capability to perform audio analysis or to create 

audio from data contained in a data file or spreadsheet. 

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Figure 1. Wind Speed Test Data 

I 

Figure 2. Block Diagram of Data-to-Wave-File Conver 

16x103 

on 

Figure 1 is a plot of static 

pressure data acquired from 

a wind tunnel test of an 

experimental NASA hurricane 

wind sensor (see ”ExtremeVelocity Wind Measuring System: 

3-D Venturi,” KSC Research and 

Technology 2002 Annual Report). 

The vertical axis of the plot shows 

raw pressure sensor output 

voltage (inversely proportional 

to the square of wind velocity), 

while the horizontal axis 

represents the sample index 

corresponding to a sample 

frequency of 100 hertz. 

Figure 2 is a block diagram 

of a system developed at KSC 

that converts any type of data 

into a standard audio wave file 

that can then be played using 

a standard PC sound utility. 

The most important processing 

block in this algorithm is the 

sample rate converter (SRC). The 

primary function of the SRC is 

to generate N output samples for 

every M input samples. Figure 3 

illustrates the effect on a set of 8 

input samples for the particular 

case of M = 1 and N = 20. The 

ideal goal of any SRC is to change 

the number of output samples 

compared to the number of input 

32 
I' I I i 1 1 

71 72 73 74 75 76 77 78 

Wave File Sample Index (Relative to Input Data) 

Figure 3. Sample Rate Conversion Example: 

20 Output Samplesfor Each Input Sample 

I 300 X 1Ok 1 I I I I 

I - 

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UJ 

m 

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20 

, = 10 

I - 

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Q 

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3 

Wave File Sample Index 

Figure 4a. Audio Wave File of Static Pressure Data 

samples without altering the frequency 

characteristics of the data. 

Figure 4a illustrates the result of 

performing the algorithm represented in 

figure 2 on the raw data from figure 1, with 

the gain control suppressed. Figure 4b 

shows similar data converted to wave file 

format where the gain control was used to 

normalize the amplitude as an additional 

investigative tool. Note that gaps of silence 

were inserted to audibly segregate the wind 

speeds. 

Key accomplishments: 

Developed a procedure to convert 

arbitrary data (usually sampled data) to 

standard audio wave file format. 

Can increase or decrease the pitch 

(frequency scaling) by a specified 

amount during the sample rate 

conversion process. 

Accomplished optional amplitude 

normalization by amplitude scaling of 

consecutive frames of data, 

Participating Organization: USTDC (Dr. J.E. Lane) 

F 

I 

800 X IO3 

3 -20- 

-300 x 103 r 

0 200 400 600 

Wave File SamDle Index 

Figure 4h. Wave File of Normalized 

D[ffereiz f id Pressu re Data 

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a, ISmall Mass Spectrometer System Onboard 

High-Altitude Research Aircraft WB-57 

The KSC Hazardous Gas Detection Laboratory is developing small mass spectrometer (MS) 

systems for various applications, including leak detection for the Shuttle, possible air quality 

analysis within the Mini-Payload Logistics Modules from the International Space Station, and 

automated environmental analysis. To meet these goals, it is necessary to research, design, 

and develop mass analyzers, sample transport systems, low- and high-vacuum pumps, and 

control circuitry that are small, lightweight, and rugged and provide rapid, accurate, and 

component-specific information. 

The Aircraft-Based Volcanic Emissions Mass Spectrometer (AVEMS) is the first system 

developed under the Advanced Hazardous Gas Detection System plan. The goal of this 

multiyear plan is to develop small, lightweight, portable, and power-efficient MS systems 

for various types of gas analysis. Developing the AVEMS instrument had two specific goals: 

(1) to determine the Hazardous Gas Detection Laboratory's current ability to build small, 

lightweight systems capable of high-altitude flight and (2) to monitor volcanic emissions. The 

first goal was met by integrating the system into a WB-57 high-altitude research plane (figure 

1). AVEMS has a total weight of 104 pounds and a volume of 3 cubic feet. It requires 350 

watts of power at steady state, can operate at a temperature range of +25 degrees Celsius ("C) 

to -65 "C, and has a pressure range of 760 to 50 torr. Within this range, the system operated 

properly up to an altitude of 41,000 feet above sea level (asl). 

The second goal was met in various ways. First, the system was mounted in the WB-57 

and flown over several Costa Rican volcanoes at a typical altitude of 5,000 feet above ground 

level (agl). At this altitude, it was difficult to quantify components characteristic of volcanic 

emissions, such as helium (He), carbon dioxide (CO,), sulfur dioxide (SO,), hydrogen sulfide 

(H,S), and various organic compounds. However, as the WB-57 flew through the volcanic 

plume, the instrument clearly showed a concentration rise for water, CO,, H,S, and organic 

compounds. 

Fipre 2. Operation of AVEMS at Turrialba 

Figure 1. AVEMS in the WB-57 Payload Bay Volcano in Costa Rica 
It was also desirable to demonstrate the 

portability of the system. For this, two people 

carried the AVEMS into Turrialba volcano 

(10,960 feet asl) where several fumaroles were 

analyzed (figure 2) and compared with the air 

just 30 feet away. Figure 3 shows a significant 

difference between the gas evolved from four 

separate fumaroles and the ambient air around 

the fumaroles. Fumaroles 1 through 4 have 

significant amounts of CO,, SO,, and H,S, with 

minimal amounts of nitrogen (N,) and oxygen 

(0,). A mere 30 feet away, the concentrations 

are similar to normal air. Each compound has 

a different sensitivity. AVEMS will not only 

enable scientists to better understand volcanoes 

but will also help in the development of 

smaller, more robust systems for many different 

applications. 

Key accomplishments: 

?,yluccd ~i-0 LLL n4 VI ~1..-- YUUILtitLILIVe mL.7 I \flC Id e-- 3ys:?l?;. 

Reduced system weight. 

Improved system ruggedness. 

Detected components over / in volcanoes. 

Achieved autonomous operation. 

a, 

Ln 

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0 

a 

v) 

a, 

a: 

Key milestones: 

Further reduce size and weight. 

Improve detection limits. 

Explore other applications such as air 

quality monitoring, hazard warning, 

and mobile leak detection. 

Contacts: EW. Adurns (FI.L'nCrick.~~.Allnlizs 

@ ?insa.goo), YA-D7-E1, (321) 867-6671; Dr. T.P. 

Grifin, YA-F2-C, (321) 867-6755; and D.W. 

Follistein, YA-D7-E1, (321) 867-6747 

Participating Organizations: USTDC (Dr. CR. 

Arkin, G.R. Naylor, W.D. Haskell, D.P. Floyd, 

C. Curley, and N.N. Blalock), University of 

Costa Rica (Dr. J.A. Diaz), Costa Rica Center 

for High Teclznology (CENAT) (Dr. J. Azofefa), 

JSC CBIASPL (Dr. E Chang-Diaz), YA-DZ (P.G. 

Henderson), and PH (1.1. Edelmann, S.K. Czaban, 

and R.J. Beil) 

1200 

Direct Analysis of Turrialba Volcano 

I I 

1000 

800 

600 a, 

> 

ro 

.- 

e 

- 

400 

200 

0 

1 2 3 4 30 fl Away 

Fumarole 

Q 

3 

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Figure 3. Analysis of Fumaroles at Turrialba Volcano in Costa Rica 

35 
Advanced Data Acquisition System (ADAS) 

Current and future requirements of the aerospace sensors and transducers 

field call for the design and development of highly reliable, cost-effective 

data acquisition devices and instrumentation systems. New designs that 

incorporate self-health, self-calibration, and self-repair capabilities allow for 

greater measurement reliability and extended calibration cycles. With the 

addition of power management designs and components, state-of-the-art data 

acquisition systems allow data to be processed and presented with increased 

efficiency and accuracy. 

A smart signal conditioning amplifier was designed incorporating these 

requirements. This device provides increased reliability by automatically 

rerouting signals through different paths when the processor identifies 

a component malfunction. It addresses the issues of signal redundancy 

and signal integrity without taking the traditional and costly approach of 

providing total hardware redundancy for every channel and every function 

block of a data acquisition system. 

In addition to signal redundancy issues, the need for self-calibration 

verification played a defining role in the architectural approach. Data 

acquisition systems such as those used in spacecraft need the ability to 

calibrate automatically without external intervention. Furthermore, the 

quality of the measurement provided by the system is directly related to the 

system’s ability to ensure a proper calibration through the life of the process 

being monitored. Finally, the data acquisition system’s ability to check 

the system’s health, detect and predict failure, and repair itself is critically 

important in systems where operator intervention is not possible. 

Enhancements to the Advanced Data Acquisition System (ADAS) include 

the incorporation of new, commercially available components such as fieldprogrammable gate arrays (FPGAs), field-programmable analog arrays 

(FPAAs), and digital signal processing (DSP) electronic / system controls. The 

ADAS architecture has been divided into three main components: Analog 

Signal Module, Digital Signal / Control Module, and Power Management 

Module. 

The Analog Signal Module provides signal redundancy and signal integrity 

without costly, total hardware redundancy. It also provides self-calibration 

verification and the ability to perform system health checks, failure detection 

and prediction, and automated self-repair that is critical for inaccessible 

systems such as those in deep space. The Analog Signal Module employs an 

innovative concept called ”spare parts - toolbox.” Based on the reliability 

ratings given to the different areas of the system, the design provides the 

toolbox with n number of spare parts (components) necessary to ensure 

continued operation. The number of each type of spare part in the toolbox 

is related to its probability of failure. Areas with higher failure probabilities 

(because of user interaction and/or environment) are stocked with a greater 

number of spare parts than areas well protected by the system (lower 

probability of failure). The number of spare parts contained in the toolbox 

will diminish deeper into the system, where external factors have less effect. 
ADAS Prototype Modides 

The Digital Signal / Control Module controls, monitors, and processes analog signals. It also 

monitors the health of the system bv processing defined measurements and/or trends. The 

Digital Signal /Control Module also implements the ”spare parts - toolbox.” The architecture 

consists of one or more processors and FPGAs. Mean time between failure and redundancy 

requirements determine the final number of processors and FPGAs used in the design. 

The Power Management Module (PMM) monitors, controls, and manages the poIver supplv 

to the different sections of the svstem. To monitor and control the lrarious functions, the PMM 

uses several processors, each with an associated lvatchdog circuitry that resets the processor if 

a problem is detected. Since processors operate in parallel under the same set of rules, control 

is handed off following the failure of one of the processors lvithout polver disruption. The 

processors can also decide whether circuitry must be powered down following the detection of 

an anomaly. 

The ADAS firmware monitors input signals, controls the output lines, and performs signal 

processing. It detects signal path failures and performs self-repairs and periodic c a l’b I rations 

on the Analog Signal Module. 

ADAS provides innovatilre solutions to problems associated lvith traditional data 

acquisition methods. It has been alvarded a U.S. patent, and Circuit Avenue Netrepreneurs, 

LLC, has licensed rights to this technology. 
Visual Anomaly Detection System Prototype 

The Visual Anomaly Detection System finds anomalies or defects in real time under 

normal lighting operating conditions. The application is basically a learning machine that 

integrates the three elements of soft computing (fuzzy logic [FL], artificial neural network 

[ANN], and generic network [GA] schemes) to process the image, run the learning process, 

and finally detect the anomalies or defects. The system acquires the image, performs 

segmentation to separate the object being tested from the background, preprocesses the 

image using fuzzy reasoning, performs the final segmentation using fuzzy reasoning 

techniques to retrieve regions with potential anomalies or defects, and finally retrieves them 

using a learning model built via artificial neural and generic network techniques. 

Fuzzy logic provides a powerful framework for knowledge representation and 

overcomes uncertainty and vagueness typically found on image analysis. ANN provides 

learning capabilities, and the GA approach leads to robust learning results. The system may 

run on either a regular PC under Windows NT or as an embedded version for portability. 

The system acquires the images from a video camera via either an analog or digital 

image grabber. Using fuzzy reasoning, the image undergoes preprocessing including 

primary segmentation, object extraction, enhancement, filtering, and edginess detection. An 

adaptive fuzzy threshold converts the preprocessed image to binary code, and a selective 

blob detector picks image regions with high edginess activity. The system computes 

morphological, geometrical, and fuzzy edginess grading values for each region and uses 

them as inputs of the learning feature via GA-ANN schemes. 

NASA built the Cable and Line Inspection Mechanism (CLIM) to automate image 

acquisition of the fourteen 1,000-foot slidewires used in the emergency egress system 

for the Space Shuttle at KSC. CLIM eliminates the hazardous, manpower-intensive, and 

time-consuming methods previously required to manually maintain the emergency egress 

system at peak performance. Inspecting the emergency egress system using CLIM still 

requires a person to continuously watch cable images for more than 48 hours. 

The intelligent Visual Anomaly Detection System will be integrated into CLIM to 

completely eliminate the need for human involvement in detecting anomalies; the system 

will not only detect the anomaly but also document it. Images acquired by CLIM are being 

used to test the intelligent Visual Anomaly Detection System. Preliminary tests, using 

images recorded from a slide basket cable mounted in the laboratory, show promising 

results. 

Because there is no need for special 

lighting, the potential applications of 

this anomaly detection system in an 

open environment are quite wide. One 

immediate application at KSC is the 

detection of anomalies on the Space 

Shuttle Orbiter’s radiator panels. 

Potential applications include automated 

visual inspection of spacecraft, ground, 

and flight equipment. The embedded 

portable version of this system could 

make visually inspecting this type of CLIM Being Tested at KSC Space Shuttle Launch Pad 
II 

3etrieve/Create 

Session 

Off-Line 1 

On-Line Anomaly 

Detection 

ANNGA Model Generation 

Based on Generated Image 

lnvut Data 

ining and 

Generation of Image 

Input Data 

nsu %.r_ 

rmsissl- cld;wq+laa -- lfqhma34S ~2mPmX4~ 

Softil1arc Maiii Meiiu Holding All tlie Options Uscd To 

Ruii the Aiionialy Detection Sysfciii 

equipment (including spacecraft fuselage) quite 

feasible in 2E opm cnvironmen:. 

The fuzzy operators built to preprocess the 

image can be incorporated into other computer 

vision systems. Adaptive image thresholding 

via fuzzy reasoning, one of the image analysis 

features used to build this anomaly system 

prototype, plays a key role in a system prototype 

currently being implemented at KSC to detect 

foreign object debris (FOD) in real time during 

the Space Shuttle liftoff. 

Key accomplishments: 

Patent submission currently being considered. 

Research Triangle Institute and Southeast 

Regional Tech Transfer Center are 

demonstrating this system as an application of 

currently marketed image technologies. 

Accepted for publication in NASA Tech Briefs. 

Presented at the Institute of Electrical and 

Electronics Engineers (IEEE) International 

Conference on Fuzzy Systems and Soft 

Computing. 

Coiitact: C.K. Dwis (C, 1 -'v' ? . . - ~ ), 

YA-E2, (321) 867-8804 

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Wireless Inclinometer 

The Shuttle Orbiter hang angle operation is performed at every processing flow to ensure 

the Orbiter is within a specified range of inclination prior to transfer and mating with the 

External Tank-Solid Rocket Booster assembly. The current methodology uses a mechanical 

vernier inclinometer and processes that are both time- and labor-intensive. With so few 

cost-effective technological advances available in the field of angular measurement, these 

methods remained the best choice. 

After an Orbiter has been processed for flight but before it leaves the Orbiter Processing 

Facility, the current operations require an angular measurement while the Orbiter is still in 

the horizontal position. Upon the Orbiter’s arrival in the Vehicle Assembly Building and 

prior to being mated with the External Tank, another measurement is taken in the vertical 

position. If the angle is within required specifications, the Orbiter is ready to be mated. The 

present method requires two technicians to carefully maneuver a hydraulically operated 

mobile man-lift into a predetermined location along the payload bay doors and use a 

mechanical vernier measuring device to obtain the required measurement. This method 

not only introduces several potential sources of error but can also put flight hardware and 

personnel in danger. 

Recently, economical high-precision measurement devices were joined with a very 

successful, proven wireless design to provide a complete, all-in-one, dependable 

measurement system. The Wireless Inclinometer System is composed of two main sections: 

the Wireless Inclinometer Module (WIM) and the handheld Central Station Data Collection 

Console (CSDCC). WIM involves the construction of an adapter plate that will mate with 

the Orbiter aft fuselage compartment access door. This section is composed of the following: 

A commercial off-the-shelf inclinometer. 

A wireless radio frequency (RF) transceiver module. 

A power management module. 

A power supply for the unit, currently composed of four C batteries. 

Embedded smart software algorithms. 

Other components necessary for the proper functionality of the system (antenna, etc.). 

The design ensures proper and consistent location of the interface plate and minimizes 

measurement uncertainties caused by the plate-Orbiter interface and plate-inclinometer 

interface. The overall uncertainty of the measurement is anticipated to be 0.1 degree of 

inclination or less to meet the customer’s requirements. 

The handheld CSDCC consists of a portable ac/ dc-powered display console that houses: 

A low-power single-board computer. 

Wireless RF transceiver module. 

Application-specific power management and distribution module. 

Embedded smart software algorithms. 

Other components necessary for the proper functionality of the system (antenna, etc). 

The software algorithm configures the Wireless Inclinometer System for operation, 

acquires the information required by the user, displays and stores the information required 

by the user, and performs RF communication health checks of the system. Both the WIM 

and the handheld CSDCC contain the embedded software of the system. 
General system software requirements include 

the abilitv to calculate and display the inclination 

accurately to within 1.0 degree, calculate and 

display actual inclination and relative (or offset) 

inclination, indicate mode of display (actual / 

relative), and display battery status of both the 

WIM and the handheld CSDCC. 

The CSDCC can retrieve measurements, 

calculate averages, and display the inclinations. 

The CSDCC can also select a measurement to 

be used as a future reference value; check the 

remaining battery life, RF signal strength, and 

remote-unit temperature; and time-stamp and log 

data for later retrieval. 

Incorporated in the WIM is an innovative power 

management system designed to keep power 

consumption to a minimum and extend battery 

life. Communications receivers operate on a low 

percentage of the duty cycle, and the transmitter 

power dries EC)~ px~pec! I! mi!!i~atts. The ivireiess 

communications network has been incorporated 

into numerous recent designs and has a proven 

track record for reliability and robustness. 

The new wireless inclinometer will provide 

an improved method of data acquisition. 

Measurements will be recorded from highaccuracy digital readouts. The equipment can be 

mounted and left in place during Orbiter moves 

and conveniently checked without additional 

manpower and equipment such as the mobile manlift. 

Key accomplishment: 

Developed, fabricated, and preliminarily tested 

the prototype sensing unit and handheld 

CSDCC. 

Key milestones: 

October 2003: Testing of integrated system. 

December 2003: System ready for operational 

testing within an Orbiter process flow. 

Participafiirg Orgnnizafion: USTDC fA.1. Eckhqff, N.h' 

Blalock, B.M. Bzrrizs, Du. C.T. Mota, mid 1.1. Rniidazo) 

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Enhanced Wireless Vacuum - Jacketed Line Sensor Network 

The Space Shuttle launch pads and Mobile Launcher Platforms (MLPs) 

have several hundred feet of transfer lines for liquid oxygen (LO,) and liquid 

hydrogen (LH,) systems. The lines are used to transfer LO, and LH, to the 

Space Shuttle External Tank to be used as propellants for the main engines. 

The transfer lines consist of sections that are kept insulated using vacuumjacketed (VJ) lines. This helps the LO, and LH, keep their liquid form and 

prevents boiloff. 

In the transfer lines, sensors are installed for detecting vacuum within 

the VJ line. When vacuum is lost, this section of the transfer line is pumped 

down to the desired level. Currently, United Space Alliance (USA) personnel 

measure the vacuum level of the transfer lines manually. This requires 

connecting a special gage to a vacuum sensor, recording the reading, and 

disconnecting the gage. This is performed for each sensor in the transfer lines. 

The Wireless Vacuum-Jacketed Sensor Network is designed to eliminate 

manual measurements of all VJ sensors installed in the transfer lines. The 

system will consist of a central station (CS), base station, and numerous 

remote stations. The only hardware interface required for each remote station 

is the connection to the already-installed vacuum 

sensor. Due to restrictions on electromagnetic 

interference (EMC) around some Space Shuttle 

systems, the radio frequency (RF) power output 

was kept to a minimum (10 milliwatts). 

Central Station Front Panel Display 

Remote Station Boards 

The CS is a rack-mountable industrial computer 

with a touch screen. The CS automatically 

performs the following tasks: 

Provides a direct wireless interface to the remote 

stations. 

Polls active remote stations for data and health 

status, coordinates frequency assignments, and 

performs trending analysis. 

sensor measurements. 

selected users via e-mail. 

users. 

Provides data collection and storage for VJ 

Automatically provides system health data to 

Offers remote data access via Ethernet for valid 

Each remote station consists of a battery 

pack, antenna, small enclosure, power board, 

RF / core board, and analog board. The remote 

station boards are integrated into one stack. 

These electronics will be installed in an extreme 

environment and flammable commodities. It 

is important to protect them against the harsh 

environment with an intrinsically safe design. 
The power board manages the use of battery 

power. It cycles the power of the RF/core board 

and provides power to the analog board only 

when required. This minimizes battery pack 

replacement to once every 2 years when acquiring 

and transmitting data twice a day. Unlike the RF/ 

core board, the power board is always on. It has a 

low-power microcontroller that manages the power 

cycles and stores the remote station configuration 

data. It performs a limited health check by tracking 

successful communications between the remotestation and the CS. It automatically reconfigures 

the remote station to use an alternate frequency if 

communication errors persist. 

The RF/core board contains a microcontroller 

and the RF module. Its main functions are to 

control the RF module and acquire data from 

the VJ sensor. It detects incoming transmission 

packets from the CS and validates the remotestation ID number. Once a valid command is 

received, the RF i core board coordinates with the 

power board for power requirements, executes the 

command, and transmits a formatted response to 

the CS. In addition, the RF / core board digitizes 

analog signals, contains calibration data used for 

compensating temperature effects, provides data 

acquisition and control of the analog board, and 

provides the communication protocol. The analog 

board supplies the excitation signal to the sensor, 

performs signal conditioning, and contains an 

analog-to-digital converter. 

Each remote station is designed with a robust 

RF communication scheme. Each remote station 

can operate on alternate multiple preprogrammed 

frequencies assigned by the CS. A remote station 

is also capable of serving as a relay station from 

the CS to another remote station. This may be 

necessary when direct communication between the 

CS and a remote station cannot be achieved due to 

an out-of-range condition or when a structure is in 

the wav. 

Key accomplishments: 

2002: Designed and tested the CS and 

remote station including hardware 

and software. 

monitoring scheme, out-of-bounds 

notification to the user via e-mail, and 

an intrinsically safe design for the 

remote stations. 

2003: Incorporated a health 

Key milestone: 

Install 40 field-rated remote stations 

and a CS for the LO, transfer line. 

Participating Orgaiiization: USTDC (Dr. E‘.]. 

Blalock, A.J. EckhqK arid I.]. Raizdazzo) 

~APJ~~J;~~c, .*.- ET. C.T. MG~G, B.M. uDiti.iis, ;“v’.N. 

Field-Rated Reiiiote Statioiz 
A Field Prototype System for Vacuum-Drying Shuttle Tiles 

Removing water from the thermal protection system tiles is a common 

Orbiter processing problem. The heat of reentry burns out some of the 

waterproofing in these tiles, and if the Orbiter is exposed to water before new 

waterproofing can be applied, the tiles will absorb and hold water, which 

must be removed. This is currently done by heating the tiles with heat lamps 

to boil the water out and using infrared (IR) cameras to determine if moisture 

is still present once the tile returns to thermal equilibrium. A new method of 

vacuuming air and water out of the tile appears to be faster, to remove more 

water, and to have fewer adverse side effects than the heating technique. 

Figure 1. liz-Lab Vacuum System 

(Tke vacuum line comes in from the left fo aflask 

acting as a wafer separator. Attached to theflask 

is a one-to-five spliffer allowing tke vacuum to 

be routed tofve tiles. Connection to fke tiles is 

made via soft, silicon rubber suctioli cups placed 

over the waterproofing holes on each tile.) 

Figure 2. Water Separator Attaclied to tke 

OPF In-House Vacuum S!ystem Through a 

l-l/2-l~zcli-Diatneter Vacuum Liize 

A field prototype is capable of pumping multiple 

tiles at once from a single vacuum source (figure 1). 

Leading from an in-house vacuum system, 1 /4-inch 

Tygon tubes ending in 2-inch-diameter silicon suction 

cups are fitted over the waterproofing holes of five 

Shuttle tiles and the water trapped in the tile is pulled 

out. The lines from the five tiles are combined in a 

manifold and sent into a water separator flask. The 

flask is maintained at vacuum. 

The five-tile test was successful and well received 

by Operations, and the approach was scaled up to a 

full field prototype system. To quiet the operation, the 

vacuum system in the Orbiter Processing Facility (OPF) 

was used. This system pulls up to 750 standard cubic 

feet per minute (scfm) at a vacuum of 8 conventional 

inches of mercury (inHg). In testing, a vacuum 

pressure of 5 to 6 inHg was used and 1 scfm was 

pulled through each tile. Therefore, the OPF system 

may be capable of pulling water out of hundreds of 

tiles at the same time, but this may be limited by in-line 

pressure drops. Measurements were taken in the OPF 

and over 30 scfm could be pulled at about 7.5 inHg 

at any one of the 1-1 /2-inch-diameter access ports 

into the vacuum system. It was decided to develop a 

system capable of drying 25 tiles simultaneously from 

any one access port in the OPE 

Figure 2 shows a water separator attached via a 

grounded vacuum line to one of the OPF vacuum 

ports. The vacuum line from the tank is connected to 

a five-to-one manifold (figure 3) where gray lines can 

be split off from the primary vacuum hose. Each gray 

line splits further into five 1 /4-inch-diameter Tygon 

lines ending in a suction cup. This system can draw 

a vacuum of about 6 to 7 inHg at each of 25 suction 

cups, each with a flow rate of over 1 scfm. Figure 4 

shows all 25 lines attached to a Plexiglas panel with 

holes simulating the tile waterproofing holes. Figure 5 
shows 10 lines drawing water from individual Shuttle 

tileq. Shuttle Operations personnel have been very 

pleased about the prospect of using the new \‘acuurn 

svstem. Benefits of vacuum drying include: 

Time: Running eight vacuum systems simultaneously for 1 hour per tile, about a third of the Shuttle 

tiles can be dried in a standard workweek. Pulling 

the majority of the water from the tile ~IIOWS evaporative drying to complete the moisture removal. 

Electricity Use: The vacuum system uses the OPF 

in-house vacuum that operates continuously, 

requiring no additional electrical power. 

measures the water pulled out of the Orbiter. 

temperature should be faster. 

to any tile on the Orbiter. 

flc.,.,. visibk. 

Quantification: The vacuum system captures and 

Use of IR Camera: Recovery to ambient 

Access to Tiles: The \Tacuum system allo~rs access 

Visibility: The clear Tygon tubing makes the water 

CoiitRct: Dr. R.C. Yolrilppist (Robert.C. Yoi~iigpisf@ iinsn.go~), 

YA-U-E, (321 j 867-1819 

Fiprc 5. Teii of flic Possililc 25 Lilies Pirlliiig Wflfer 

Oiif of Shirttle Tiles 

45 
Shuttle Tire Pressure Monitor Calibration Station 

While traditional calibration methods acquire sensor voltages using digital 

multimeters, the new Tire and Strut Pressure Monitor (TPM) calibration 

procedure acquires voltage readings directly from the analog-to-digital (A / D) 

converter of the device. This is achieved after calibrating the A/D converter 

with an in-house-developed linearization algorithm. These voltages are then 

sent to a computer for processing. A multinomial curve-fitting algorithm 

is then applied to generate the calibration coefficients of the TPM. This 

procedure provides better curve fitting and significantly reduces uncertainties 

and the cost of the calibration station since no digital multimeters are needed. 

Furthermore, the calibration station is automated, so highly skilled personnel 

are not needed to perform routine calibrations. 

The TPM is a handheld system to accurately measure Orbiter tire and 

strut pressures and temperatures over wide temperature ranges, as well as 

to improve personnel safety by reducing the pressurized volumes in the 

measuring device. The design places a highly accurate pressure sensor 

(also used as a temperature sensor) as close to the tire or strut measurement 

location as possible, allowing the user to make accurate measurements 

rapidly, minimizing the amount of high-pressure volumes, and allowing a 

reasonable distance between the tire or strut and the operator. The stringent 

requirements of the TPM made calibrating it a challenge. 

The original TPM calibration station consisted of a pressure standard, an 

environmental chamber, a precision voltage source, two digital multimeters, 

and a computer. Reflecting the traditional approach, the calibration algorithm 

created a table of several pressures and temperatures for which sensor 

voltages were recorded using the digital multimeters. These voltages were 

then used to calculate the coefficients of a multinomial that would fit the 

gathered data. The procedure was implemented, and through several data 

verification tests, it was found that the TPM did not meet the desired accuracy 

(although it was within the requirements). Measurement uncertainties 

included nonlinearities and hystersis of the temperature and pressure sensors 

and A / D errors (integral linearity error, differential linearity error, transition 

noise, full-scale error, full-scale error drift, etc.) of the TPM. The calibration 

station also required expensive laboratory equipment and too many manhours of highly skilled personnel, making its implementation very costly. 

To eliminate some uncertainties and to improve the accuracy of the TPM, 

the pressure and temperature sensor voltage readings were taken directly 

from its A/D converter. This procedure was performed by first calibrating 

the analog circuitry of the TPM, which brought the voltages from the A / D 

converter within acceptable error limits. As a result, the digital multimeters 

were no longer needed and were removed from the calibration station. The 

voltages read by the A/D converter were transmitted to a computer and 

used to calculate the calibration coefficients. The new approach resulted in 

approximately half the uncertainties of the traditional approach, and the cost 

of the calibration station was reduced by about half. 
The calibration station now consists of a pressure standard, an 

environmental chamber, a precision voltage source, and a computer with 

custom software to control the calibration process. 

The following are novel features of the calibration station: 

Sensor voltages are taken directly from the TPM A / D converter, avoiding 

Uncertainties of traditional calibration stations are reduced by half. 

Reduction in equipment use results in significant savings on the cost of the 

Automated calibration process requires fewer highly skilled personnel and 

the use of costly multimeters. 

calibration station. 

results in significant man-hour savings. 

Participating Organization: USTDC (Dr. C.T. Mafa, B.M. Bums, aid R.T. Deyoe) 

'I 

Block Diagrain of tlze TPM Calihratioiz Statiori 
Vacuum-Jacketed (VJ) Thermocouple Calibration Station 

A new station was designed to improve the process of calibrating the thermocouple 

gages of the Shuttle launch pads’ VJ cross-country transfer lines. The VJ lines are used to 

pump cryogenic propellants from the liquid hydrogen and liquid oxygen storage tanks 

to the External Tank when fueling the Space Shuttle and to prevent vaporization of the 

cryogenics. Permanently installed thermocouple vacuum gages are used to measure the 

vacuum pressure of these transfer lines. The nature of the installation does not allow 

practical calibration cycles of these thermocouples. 

Before the development of this calibration station, there was no routine calibration 

practice for the thermocouple gages since the only way to calibrate them was to remove 

them from the lines. This impractical procedure was never seriously considered. Readings 

of the thermocouples had to be trusted, and accuracy of the measurements was a real 

concern. Considering that the loss of vacuum in the VJ lines would cause the cryogenic 

propellant to boil off, measurement accuracy had to be known to some degree. Technicians 

had to periodically check the readings of the thermocouples to make certain the vacuum 

pressure was within the required limits. If the pressure read by a VJ thermocouple was 

higher than allowed, that section of the line would be pumped down using the same 

thermocouple readout as the pressure measurement. This meant that degradation of the VJ 

thermocouples could not be identified because there was nothing to compare their readouts 

with. The problem with this practice was obvious because measurements with unknown 

accuracies were used to determine acceptable vacuum pressure levels. 

The VJ calibration station was designed to eliminate the uncertainties of the 

thermocouple gage readings and characterize them in situ. The VJ calibration station 

Y 

... .............. .......... ........ ..... ........ ,.... ....... ... ............... .. ... ....... .......,,. ....... .... ... ......,,..,,... _.,.. .. ...... 

; I 1 

t , ,y SingJe-Board Computer I 

& VJ Signal 

+ Conditioner 

- bratron 0-10Vdc 

I 

Interface Board 

Acopian 

t15 V dc 

RS232 

t6Vdc 120 v 8C 

r? 

Figure 1. V] Calibration Station Block Diagram 
is composed of a vacuum pump, 

a single-board computer, power 

supply modules, a pressure standard, 

communication ports, and a VJ 

thermocouple gage signal conditioner 

(see the block diagram in figure 1). If 

the pressure read by a VJ thermocouple 

is higher than allowed, that section 

of the line is pumped down using 

the VJ calibration station in line with 

the vacuum pump. The pressure of 

the line is measured by a pressure 

standard, and the VJ thermocouple 

output is monitored as the line is 

pumped down. The readouts are 

stored in the single-board computer 

and used to characterize that particular 

VJ thermocouple sensor. As a result, 

a characterization curve is obtained 

and the sensor is calibrated. This in 

situ calibration method is extremely 

practical because the regular 

pumpdown procedure is not altered; 

yet the calibration of the sensors is 

realized. The uncertainties of wrong measurements are now minimized because sensor 

degradation can be identified as long as calibration is performed within certain periods. 

I 

I 

i 

I 

I 

1 

I 

l 

Figure 2. A Viezo ofthe VI Calibratioii Statim in If5 

Coir cep t ual S fage 

The VJ calibration station (except for the vacuum pump) is enclosed in a light, weatherproof 

case that can easily be taken anywhere along the cross-country VJ line (see figure 2). 

The following features are considered novel improvements to the existing system: 

The calibration station is used in-line with the existing vacuum pump without altering the 

I pumpdown procedure. 

I Readout uncertainties due to out-of-calibration or degraded sensors are minimized by 

calibrating the thermocouple sensors when the VJ line sections are pumped down. In 

situ calibration is achieved by characterizing the VJ thermocouple sensor with a pressure 

standard as the line is being pumped down. 

The VJ calibration station is expected to be finished by the end of FY 2003 and to be used 

with the new VJ wireless sensor network during FY 2004 to provide an automated data 

acquisition system with in situ calibration capabilities for VJ thermocouples. 

Pnrficipnting Orgnnization: USTDC (Dr. C.T. Mnfa, A.]. Eckhoff, ._ B.M. Burns, ].I. Rnndnzzo, mid S.]. Stout) 
Fast, Time-of- Flight Mass Spectrometer for 

Rugged Aerospace Applications 

Time-of-flight (TOF) mass spectrometry is designed around the properties of molecular 

matter, Newton’s first law, and the forces imparted to charged particles by electric fields. 

At the beginning of a measurement cycle, pulsed electron beams form ions from neutral 

molecules. Immediately after these ions are created, they are accelerated by pulsed electric 

fields into a drift region where they are separated into groups by their differences in velocity, 

an inverse function of their mass. These groups arrive at the detector inversely according 

to their mass; lightest first, heaviest last. When the ions strike a detector plate, their time of 

arrival and quantity are measured. When the last ion group in the sample has arrived, a new 

cycle can be initiated. For the gas molecules of interest to the aerospace community, the time 

to complete a measurement cycle is about 30 microseconds. At this rate, approximately 30,000 

measurement cycles can be made in 1 second. 

A compact TOF mass spectrometer, developed earlier on a Small Business Innovation 

Research (SBIR) contract and delivered in CY 2000, was modified to incorporate an ionizer that 

does not rely on thermionic emission for operation. The original instrument was designed 

as a fast sensor for use as a propellant (hydrogen, helium, oxygen, and argon) leak detector 

during the first few seconds of a space vehicle launch cycle, after main engine ignition and 

prior to release from the ground. The major NASA-defined benchmarks for this application 

of sensitivity, accuracy, dynamic range, response time, and (to a degree) ruggedness were 

350 7 25 PPm H, , 

8o 7 25 ppm He 

3500 I I I J I 

30 40 50 

Time (s) 

60 

Calibrafioii Gas Data Acquired in 150-Milliseco~zd Sample Tiinefvarrres 
designed into the current prototype using the 

new ionizer technology. The test article is located 

at the vendor’s facility, and NASA personnel 

plan to evaluate the instrument under controlled 

laboratory conditions at KSC. 

A conventional time-to-digital converter was 

used for data acquisition. This has the advantage 

of high sensitivity (low detection limits) but has 

limited dynamic range. The dynamic range of 

a TOF mass spectrometer is principally limited 

by saturation of the microchannel plate detector, 

which can handle only a few million counts per 

second. An approach to overcome that limitation 

for this application is being worked through a third 

SBIR Phase 1 contract with the vendor. A detector 

scheme is under development using a special form 

of the time-to-digital converter, which will increase 

the dynamic range by 2 orders of magnitude to 10j 

counts per second. Patent application PCT/US99 / 

13965 explains this unique approach. 

Low part-per-million detection limits were 

achieved for all species of interest (hydrogen, 

helium, oxygen, and argon) in a nitrogen 

background within an acquisition time of 150 

milliseconds. A fast interface allows response times 

below 1 second for air sampling with an %inch, 

128-micron capillary and a rise time of 200 to 300 

milliseconds. Identical mass spectra were achieved 

by direct comparison of the filament and surface 

desorption (SD) ionizer schemes. 

The ionizer used in this prototype is based on 

a spontaneous SD process and does not require 

a hot-wire filament for operation. Compared 

to thermionic emission, the classical method of 

creating free electrons in a vacuum, this electron 

source provides advantages in efficient use of 

power. Rugged construction leads to a reliable 

instrument that can acquire gas concentration 

data in an extremely harsh operational 

environment, typical of what might be found in 

the aft compartment of a spaceflight vehicle or 

autonomous ground vehicle. 

Electron emission stability and current density 

are similar to that achieved with the electron 

filament used in the first prototype. Adequate 

performance of the SD ionizer in the TOF mass 

spectrometer has been demonstrated. Various 

geometries of the ionizer were tested 

during the development process. 

Electron currents up to 100 nanoamperes 

were registered. This limit is imposed 

by the electrical properties of the 

electron multiplier. Increased currents 

can be obtained by using channel 

electron multipliers that provide 

currents up to several microamperes. 

More current might be possible but has 

been limited by device lifetime. 

The electron current density of this 

ionizer would satisfy most time-offlight mass spectrometer applications 

but at the cost of a bulkier mechanical 

structure. The new design is more 

robust than the filament-based 

version, but it ultimately suffers from 

degradation due to reduced gain in the 

chamel electron multiplier as a function 

of extracted current and time. The 

major limitation to this SD technique 

is identical to that of the channel 

electron-plate-based ion detector - only 

1 coulomb can be extracted over the 

usable lifetime of these channel plate 

structures. 

Work remaining includes the 

miniaturization of control and data 

acquisition electronics, overall 

ruggedization and test, rugged pump 

development, and advanced detector 

and ionizer development to improve 

operational lifetime. 

Key accomplishments: 

Developed a small, fast, aerospacegas-specific, potentially rugged, timeof-flight mass spectrometer. 

Demonstrated filamentless ionizer. 

Con f ac f s : F. W. Adam f Frcdc lick. LV. Adn i i I 5 

Follistein, YA-D7-E1, (321) 867-6747; and Dr. 

T.P. Griffin, YA-F2-C, (321) 867-6755 

Pnrticipating Organizations: USTDC (Dr. CR. 

Arkin, W. D. Hnskell, and G.R. Naylor) mid 

loiizi’erks (Or. [.A. Schultz) 

8;) ~IOSO.;<~?Z~), YA-D7-E1, (321) 867-6671; D. W. 
I 52 

Inspection Tool for the Reaction Control System (RCS) Nozzles 

Figure 1. Version 1 Composed of a11 

Inner Tejlon Cylinder Holding a 

Mirror and an Outer Centering Plug 

Figure 2. Version 2 Slioivn Completed 

(top) arid in Use (17ottorr-r) (The ccwteririg 

plug uias dropped, flit cyliiider ZLVS opened 

up, and a lzaiidle aiid tether we^ .dded.) 

The Space Shuttle Orbiter maneuvers in space using the 

RCS, which is composed of small hypergol thrusters located 

around the vehicle. These thrusters consist of an elaborate 

hydrazine and nitrogen tetroxide delivery system attached to 

the small end of a rocket nozzle. The rocket nozzle contains 

a combustion chamber where the fuel and oxidizer mix and 

explode, followed by a throat that constricts the explosion, 

followed by a parabolic expansion region that funnels the 

expanding gases away from the vehicle. The nozzles are 

made of niobium (also known as columbium) coated with an 

oxide layer designed to survive the high temperatures within 

the combustion chamber. After each mission, the nozzles are 

checked for defects, usually small divots caused by a section 

of the oxide layer coming off. These defects are serious if the 

underlying niobium is exposed because the metal can melt 

from the heat of the hypergol reaction. However, if some of 

the oxide layer remains, then the nozzle may be safe to fly. 

In 2001 a defect was found in a rocket nozzle while the 

Shuttle was at the pad, and it was unclear if the underlying 

metal was exposed. After significant effort, it was 

determined that sufficient oxide was present and that it was 

safe to fly. However, had this defect exposed the niobium, 

not only would the Shuttle have been brought back to the 

Vehicle Assembly Building but the Orbiter would have been 

demated and brought back to the Orbiter Processing Facility 

for repair - an enormous expense of both time and labor. 

Consequently, the decision was made to put more resources 

into defect detection, measurement, and in situ repair. The 

result was the evolution of a rigidly mounted device into a 

handheld, illuminated inspection tool that allows users to 

look inside the throat of a rocket nozzle on the Space Shuttle 

to search for defects. 

The original motivation was to create a tool that could 

evolve into a defect measurement system. Consequently, 

a two-part inspection tool was designed to hold a mirror 

at a known depth and centricity within the nozzle. This 

first-generation tool is shown in figure 1 and consists of a 

Teflon sleeve and a Teflon mirror assembly. Only Teflon was 

permitted to contact the oxide layer within the nozzles. If 

this design proved acceptable to the users, then a microscope 

system could be attached to the mirror assembly, allowing 

a user to scan the throat and combustion regions of the 

nozzle and to measure the size and depth of a defect. Upon 

delivery, though, the inspectors were more interested in just 

the mirror portion of the device. Previously, they had been 

using a rather small, low-quality mirror, and the large mirror 

mounted in Teflon provided better vi6bility of defects within 
Figure 3. Versioii 3 Wifh lllziiniiiafioiz 

Figure 4. Version 4 With Improved Reach, 

Access, mid Opera fioii 

the nozzle. So, the design direction was altered to 

concentrate on a better handheld inspection tool. 

Figure 2 shows the resulting next-generation 

device. The sleeve was removed, the design was 

opened up to make the mirror more visible, and a 

handle and tether were added. This unit was well 

received by Operations and used successfully in 

the field, but it had a significant limitation - the 

inspector had to hold a flashlight to see inside the 

nozzle. 

Figure 3 shows the illuminated version of the 

inspection mirror. Small, high-efficiency, white 

light-emitting diodes (LEDs) were added, and 

circuitry ~7as designed to drive them from a set 

of three AA batteries. These LEDs provided 

significant light as seen in the figure, eliminating 

the need for a separate flashlight. The batteries 

were located in the handle of the unit, which was 

designed to turn on and off by twisting the handle. 

Field testing by Cperations indicated that a shorter 

handle was the only modification necessary. In the 

final design (shown in figure 4), the battery was 

moved back in line with the mirror mount to avoid 

interference with the nozzle. This allowed the 

inspector more reach into the nozzle and permitted 

a thumb-operated on /off switch. 

This iterative design process emphasizing 

user involvement yielded a tool that makes 

rocket nozzle inspections easier to perform while 

minimizing the risk of damaging the nozzle 

coatings. 

Participating Organizntioiz: USTDC (W. D. Haskell arid 

R.B. Cox) 
Fluid System Technologies 

The vision of Fluid System Technologies is to develop technology that enables future affordable spaceports. These spaceports 

can be Government-owned, commercial, orbital, or nonterrestrial. 

The goal is to reduce the cost of access to space and increase safety. 

Fluid System Technologies can help in this role through reducing the 

human oversight of servicing, reducing the waste streams, and reducing the number of hazardous operations and vehicle interfaces. Fluid 

System Technologies will support developing a better understanding 

of the fluid environment in all areas of spaceport activity through 

analysis and simulation. In addition, Fluid System Technologies 

will seek to develop the mears ts reduce thermal iosses associated 

with cryogenic fuels, minimize maintenance and process monitoring 

costs, and provide for safe operation of the spaceport. A major goal 

of this technology product line is to create efficient technologies that 

can be quickly adapted to the changing fluid needs of future space 

vehicle elements and systems residing at the spaceport. The technology focus areas of the roadmap that drive technology development 

toward the ultimate vision include the following: 

Storage, Distribution, and Servicing Systems 

Production, Recovery, and Disposal Management Systems 

Vehicle Interface Systems 

Safety Systems 

Expendable Launch Vehicle (ELV) / Reuseable Launch Vehicle 

(RLV) Thermal / Fluids Environments and Management 

For inore iizforinatioiz regarding Fluid Systeiii Techologies, please contact Jack Fox 

fl17i/i.).Ft~.t~ vw.;.y~), YA-€6, (321) 867-4413; or Russel Rlzodes (Ruwi. €.iu7d~~ 

,J) I~I?X~.><C~C.), YA-D4-A, (321) 867-6298. 

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Fluid System Technologies Roadmap 

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Modeling the RLlO Rocket Engine With Densified Liquid Hydrogen 

and Oxygen Propellants 

There are many benefits to using densified propellants on launch and space vehicles. The 

increase in propellant density translates into smaller propellant tanks that result in lower liftoff weights and larger vehicle payload capacities. The subcooling of densified propellants also 

allows lower system operating pressures in propellant tanks that extend tank life in reusable 

systems. The increased cooling capacity is a potentially vital heat sink for leading edges and 

shock wave regions subjected to aerodynamic heating and for rockets or rocket-based combined-cycle engine combustion chambers and nozzles. 

Analysis of rocket engine performance using densified propellants is a key component of 

the research necessary to optimize engine design and is necessary to advance the maturity of 

densified propellant technology. To further understand the effects of densified propellants on 

rocket engine systems, an RLlO engine was analytically modeled with densified propellants 

using the Rocket Engine Transient Simulation (ROCETS) mathematical modeling code. A validated model of the RL10A-3-3A rocket engine was modified to simulate densified propellant 

inlet conditions and was used to determine the effects on overall system operating conditions. 

The nominal operating conditions of the RL10A-3-3A were compared with operating conditions obtained at various levels of densified propellant inlet conditions. 

The ROCETS code was developed by Pratt & Whitney for NASA in 1990. The program 

simulates a complete rocket engine system and calculates the values for all flow, pressure, 

temperature, and state variables at various locations within the engine system components. 

The code enables a number of study types to be conducted, including parametric, dynamic 

response, component geometry changes, and overall system characterization. The code has 

been used at Pratt & Whitney and several NASA Centers to model rocket engines and rocket 

engine test facilities. 

The ROCETS model was used to study the effects of densified propellants on rocket engine 

performance, which was the ultimate goal of this work. In this study, the model input parameters (such as engine thrust, oxidizer / fuel mixing ratio, combustion chamber pressure, and 

liquid oxygen [LOX] and hydrogen fuel inlet pressures) were fixed. Other inputs, such as 

LOX and hydrogen fuel inlet temperatures, were varied from normal boiling point at ambient 

pressure down to near triple point. The behavior of a number of different engine system components as a function of the thermodynamic state of the LOX and hydrogen propellants was 

predicted. In general, as propellant inlet temperatures decrease, the RL10A-3-3A model shows 

Constant Engine Thrust ~~ ~ ~ 

RLlO DP Study - Fuel Pump Speed 

- 32000 E 

85 105 115 125 135 145 155 

LOX Temperature (R) 

__ 

Figure 1. Fuel Pump Speed Performance 

a significant decrease in fuel and LOX 

pump speeds. This reduces engine system 

failure probability and prolongs engine 

life. Pressure drops across the injectors and 

cooling jacket are smaller due to a smaller 

volume of flow rate for densified fuel and 

LOX. Inlet temperatures at the injectors are 

reduced as engine inlet propellant temperatures decrease. Turbine inlet temperature, 

torque, and power decrease with decreasing 

propellant temperatures. Figures 1 through 

5 illustrate the behavior of several system 

components. 
The research results quantified some of 

the key benefits of densified propellants in 

engine systems, including pressure drop 

reduction and decreased turbine / pump 

speeds. In general, the results show a 

reduction in pressure drop and turbomachinery rotational speed as liquid propellant inlet temperatures decrease and 

fluid densities increase. The advantages 

of these effects on the engine and vehicle 

include increased reliability and engine 

life. The data generated in ths study can be 

used for optimizing current engine system 

configurations for densified propellants or 

for designing new lighter-weight engines 

that take advantage of the increased density 

of the propellants. 

Key accomplishments: 

Modified a ROCETS model of the RLlOA3-3A cryogenic hydrogen and oxygen 

expander-cycle rocket engine to simulate 

the use of densified propellants. 

Performed validation of the mathematical model with existing engine test data. 

Developed methodologies for assessment 

of engine systems optimization methods 

for densified propellants. 

Key milestones: 

Conduct a detailed experimental study 

on engine combustion stability and ignitability. 

lants on increased engine thrust. 

engine components. 

engine with densified liquid hydrogen 

and LOX. 

Research the effects of densified propelConduct further detailed analyses of key 

Conduct testing of an existing RLlO 

Contact: L.K. Walls (Ln1cric.l<.IZ7nl15g7 1~~7in. yz'), 

VA-F3, (321) 476-3637 

Participating Organization: Sierra Lobo, Inc. (M.S. 

Haberbusch, C.T. Nguyen, and A.F. Skafl 

Constant Engine Th~st 

RLlO DP Study - LOX Pump Speed h $ _.__ --__ 12600 x -X-TH2=38R - 12400. X-X-X- 1 0 -x- TH2=34R 

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LOX Temperature (R) 

Figure 2. LOX Pump Speed Performance 

Constant Engine Thrust 

RLlO DP Study - Pressure Drops 

Across Cooling Jacket - 

2 1950 

-X- TH2=38R 

-& TH2=30R 

e TH2=26R 190 0 -_____I 

$5 105 115 125 135 145 155 

LOX Temperature (R) 

Figure 3. Pressure Drop Across Cooling Jacket 

Constant Englno Thrust 

RLlO DP Study - P-u~ Drop. I Across Fwl In- I I 

Figure 4. Pressure Drop Across Fuel lnjector 

Constant Engino Thrust 

RUO DP Study -Turbine Porrar 

-_______ 

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LOX Temperature (R) 

Figure 5. Turbine Pouler 

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60 

On the Scaling Laws for Jet Noise in Subsonic and Supersonic Flow 

Noise generation by turbulent jets is of great practical interest in the 

design of jet engines (subsonic and supersonic civil transport) and the study 

of launch vehicle acoustics. Acoustic loads in a launch vehicle environment 

induce structural vibration of vehicle components, ground support structures, 

and equipment in the immediate vicinity of the launch pad. In designing 

launch vehicles, it is highly desirable to generate data on acoustic loads (nearfield and far-field noise levels) both analytically and from testing small-scale 

and full-scale models. Since full-scale acoustic and vibration testing is often 

cost-prohibitive, the option of small-scale testing combined with analysis 

methods remains a practical alternative. 

Noise from subsonic jets is caused by turbulent mixing, composed of the 

contributions of large-scale and fine-scale structures. The turbulent mixing 

noise is mainly broadband. In perfectly expanded supersonic jets (nozzle 

exit plane pressure equals the ambient pressure), the large-scale mixing 

noise manifests itself primarily as Mach wave radiation caused by the 

supersonic convection of turbulent eddies with respect to the ambient fluid. 

In imperfectly expanded supersonic jets, additional noise is generated by 

broadband shock noise and screech tones. 

In early design stages, scale models are often used to predict the acoustic 

environment associated with flight vehicles. A detailed knowledge of the 

mechanisms of noise generation and noise radiation by jets is essential in 

designing a scale model of the noise source. To ensure complete similarity 

between model and full scale, we need to ensure similarity of flow, noise 

generation, and noise propagation. 

In practice, it is generally difficult to duplicate (simulate) all the 

characteristic parameters in the scale model. Model testing with even small 

rocket engines requires extensive safety precautions. Heated jet facilities also 

involve considerable complexity and cost. The use of less expensive facilities 

or lower gas temperatures, for example, would considerably simplify model 

testing. The ability to conduct a scale-model test with a substitute gas (air, 

nitrogen, helium, etc.) results in considerable savings (reduced costs of test 

facilities, test time) and advantages. For example, helium-air-mixture jets for 

simulating high-temperature effects have been studied. These substitute gas 

tests require some compromise of the actual physics of the hot jet. 

In the absence of an exact match between the dimensionless parameters 

of the scale model and the full scale, a detailed knowledge of the functional 

relationship with respect to the various parameters is essential to aid in the 

interpretation of scale-model data to predict the full-scale environment. 

Key accomplishments: 

Reviewed the scaling and similarity laws applied to jet noises. 

Studied the effect of jet temperature in supersonic jets. 

Proposed a similarity spectrum for sound power level for both subsonic 

and supersonic flows. 
0.01 0.1 1 10 100 1000 

f/fp 

Similarity Spectra for Jef Noise at Mj=2 

Key milestones: 

Perform cold-jet acoustic testing. 

Present the study at the American Institute of 

Aeronautics and Astronautics (AIAA) Aeroacoustic conference. 

Begin the development of hot-jet facility. 

Contact: Dr. B.T. Vu (Bruce. T. Vu@:nasn.ipoz~), YA-C2-T, 

(321) 867-2376 

Participating Organization: USTDC (Dr. M. Kaizdula) 

61 
Advanced Umbilical Development Automation Testbed 

Numerous launch vehicles, planetary exploration systems, and rovers 

require umbilical systems for fluid flow and electrical connections. Simple, 

reliable, autonomous mating is required to make certain missions and systems 

feasible (e.g., Mars methane-fueled rovers) and to provide low-cost and safe 

launch operations. The environment for mating (temperature, humidity, dust, 

cost, weight, and power) varies drastically from system to system. However, 

a common core technology and numerous sensor and mating techniques are 

needed to provide low-cost, common solutions for all systems. 

The Smart Umbilical Mating System (SUMS) is a general-purpose, 

fully automatic, extend-and-retract umbilical system that can be used in a 

Smart Umbilical Hardware (right side) and Simulated 

Flight-Side Plate (left side) 

Control Station Consisting of Redundant 

Control Architecture 

multitude ofapplications. It was built to 

demonstrate the latest in control technology. 

SUMS incorporates a fault-tolerant, failoperational control architecture and utilizes 

linear motors for passive compliance. In 

addition, SUMS will be used as a technology 

testbed to integrate component-level 

mechanical, instrumentation, and other 

umbilical technologies into a working proofof-concept prototype. The intent of this 

system is to provide a platform to allow 

multiple technologies to be evaluated from 

a system perspective. System design is 

generic for use with a wide variety of launch 

vehicles and the development of enabling 

technologies, such as quick disconnect 

connectors, sensors, location / alignment 

systems, latching/ actuation systems, 

deicing / decontamination systems, control 

systems, and verification systems. 

SUMS is capable of extensive 

application of redundancy and alternate 

modes of operation. These capabilities 

are demonstrated in the application of 

techniques to enhance operational reliability. 

For example: 

The control system uses a fault-tolerant, 

voting algorithm for critical controls. 

A redundant vision system that utilizes 

two cameras is employed to align the 

mating cones to receptacles on the flight 

half of the umbilical plate. 

Linear motors are used on all axes, and 

the pitch actuator incorporates a redundant motor to demonstrate the system’s 

fault tolerance. 
Technology areas include redundant 

system control, autonomous operation 

and automated mating, location control 

and reconnect capability, ice suppression 

technology, and man-machine interface. 

SUMS increases the umbilical technology 

readiness level by demonstrating that 

the umbilical connect, disconnect, and 

reconnect capability at any point in the 

countdown process improves safety of 

launch operations; the autonomous mating 

operation reduces costs and improves 

reliability of ground processing operations; 

and the umbilical system integrates system 

health management to verify the system 

operational status. 

Main Operator's Screeiz 

Key accomplishments: 

Developed a remotely controlled graphical user interface for the SUMS hardware. 

Performed liquid nitrogen cryogenic 

mate, demate, and remate testing. 

Demonstrated a potential ice mitigation 

solution. 

Contacts: T.C. Lippitf (- -- ~ ), 

X4-D1, (321) 867-1391; aiid R.l? Mueller, YA-D1, 

(321) 867-2557 

Part icipatiHg Orgal? iza t ion5 : US TDC (R. A. 

Gordie and J. K. Trautiueiiz) mid Rolriiiedder, Iizc. (T. 

Adains) 

63 
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Advanced Umbilical Location System (AULS) 

Automated umbilicals are being developed at KSC that will dramatically 

reduce the time required to mate an umbilical system to the launch vehicle 

interface. To do so, the ground umbilical plate must have a positioning 

mechanism with an associated control system that will locate the ground 

umbilical plate with respect to the flight umbilical plate. The flight umbilical 

plate has six degrees of freedom since it is mounted onto a launch vehicle 

that is positioned on the launch pad and is subject to installation tolerances 

and wind load excursion movements. The ground plate also has six degrees 

of freedom and must be accurately positioned in terms of location and 

orientation to within 0.125 inch, and it must be parallel to the flight plate 

to within +1 degree. The orientation and parallelism can be achieved with 

a mechanical pin/spring mechanism, but the location to within 0.125 inch 

must be provided by a separate system. The location system must be able to 

remotely find a target on the launch vehicle flight plate and then advance and 

position itself over a gap of 6 feet. An AULS proof-of-concept prototype was 

developed to test functionality. 

64 

The AULS uses a modulated light-emitting diode (LED) light source 

mounted on the ground flight plate that shines onto a prismatic reflector 

on the flight-side plate. The light is reflected back to the ground side and is 

detected by a photovoltaic sensor that is divided into four quadrants. The 

electric signal generated is fed into an analog electronics feedback logic 

circuit, which then powers positioning stepper electric motors on lead 

screw mechanisms corresponding to each quadrant on the light sensor. The 

net result is that the ground flight plate travels toward the target location 

on the flight plate represented by the reflector. In this way, the umbilical 

ground plate is able to track the flight plate in real time as the ground plate 

is extended out toward the flight plate in the umbilical mate operation. The 

AULS constantly adjusts the position of the ground plate so it tracks the flight 

plate even if the flight plate is moving or vibrating. The modulated light is 

used to distinguish between the signal light and any stray sources of ambient 

light. The whole assembly has been mounted on a cart. 

AULS Proof-of-Concept Prototype (front view) 
AULS Proofof Concept Prototype (side ziie7ii) 

A UL S Proof- of Co i I cep f Pro to type Ai ialog E1t.c f ro 17 ics 

Coiit rol Sys fein 

Testing shows the AULS is capable of real-time tracking of a moving target 

and is accurate to within 0.125 inch for final location and orientation. 

Key accomplishments: 

Developed AULS prototype concept. 

Designed and fabricated AULS proof-of-concept prototype. 

Tested location capabilities of AULS to within 0.125-inch accuracy. 

Tested real-time response of AULS in ambient light conditions. 

Contact: R.P. Mueller (Rob..ll~~rllc~(ci imw.yx,), YA-D1, (321 i 867-2557 

Parficipafiiig Orgniizatioii: American Reiiiofe Visioii Co. (S. Gleman) 

65 
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Advanced Cold-Mate Cryogenic Fluid Quick Disconnect (OD) 

Loading propellant onto launch vehicles requires using cryogenic fluid 

quick disconnect (QD) fittings. The QD fittings, which are mounted onto 

umbilical systems with related mechanisms, mate or demate the QDs to the 

launch vehicle ground-to-flight interface. Cryogenic propellants such as 

liquid hydrogen and liquid oxygen are then loaded onto the launch vehicle 

through the QDs and related fluid lines. At launch, the umbilical systems 

must be retracted from the launch vehicle and stored in a protective housing 

to avoid damage from the launch blast and to allow reuse of the umbilical 

system. If a launch is scrubbed or aborted, the propellants may be drained 

from the launch vehicle. If the umbilical has been retracted prior to T-0, as 

is anticipated with the use of automated umbilicals, then the QDs must be 

capable of a cold mate while the QDs are still at cryogenic temperatures. 

The problem that arises during cold mates is that the humidity in the air 

condenses onto the cold QD surfaces and forms a layer of frost or ice, which 

compromises the integrity of the sealing surface on the QDs. The frost/ice 

prevents a good mate from taking place between the male and female halves 

of the QD, resulting in a ground-to-flight fluid interface that may leak. Fuel or 

oxidizer leaks at the launch pad may cause dangerous explosive conditions, 

which are a Criticality 1 Hazard that may cause loss of vehicle and crew. 

A specification titled “Operating Parameter Goals for Advanced Cryogenic 

Fluids Disconnect” was developed and released to private industry in a 

solicitation for proposals. NASA selected VACCO Industries, Inc. Space 

Products, of El Monte, California, to develop the Cold-Mate QD. VACCO 

owns a design that was used on the Shuttle Centaur upper stage planned for 

installation in the Space Shuttle Orbiter payload bay. The Centaur design 

has a shroud that encapsulates the QD to prevent the accumulation of any 

propellants leaked in the Orbiter payload bay. The Centaur design was used 

as a baseline to further develop the Cold-Mate QD design (figure 1). 

Ice formation can be prevented by isolating all cold QD surfaces from moist 

atmospheric air. The typical method for achieving this in a spaceport launch 

pad environment is to use a dry, inert-gas purge such as helium or nitrogen. 

Because the respective liquefaction-phase-change temperatures are lower than 

liquid hydrogen and liquid oxygen, the purge gas does not liquefy and is an 

effective isolation medium. 

Figure 1. Cross Sectioiz View of Cold-Mate QD Desigir 
I Figure 2. Mated QD in Test Fixture at LN, Temperature 

1 

Figure 3. Male QD Half Shozi7iii~ Ice Prez7eiitioii 

011 ~ealillg siqfacc 

An inert-gas purge method was selected for ice prevention on 

the QDs. To contain the gas purge, a spring-loaded telescopingtube shroud mechanism was designed and installed around an 

existing VACCO cryogenic fluid QD coupling. A test fixture was 

built (figure 2j to simuiate the ground-to-flight interface mating 

of the two QD halves under liquid-nitrogen (LN,) temperature 

conditions. The LN, test was performed and showed that the 

inert-gas purge was effective in keeping the QD sealing surfaces 

free of frost or ice (figure 3). The QD was pressurized to 50 

pounds per square inch gage (psig) at a temperature of -290 

degrees Fahrenheit (OF) and showed favorable results. 

Tests demonstrated the capability of the advanced Cold-Mate 

QD concepts to mate and demate after a separated cold soak 

simulating a T-0 launch abort. In addition, the prototype QD 

demonstrated the capability to verify leakage at various stages 

of engagement. Further development work is desired to make 

a production version of the prototype QD for use in future 

automated umbilicals. 

1 

Key accomplishments: 

Developed viable inert-gas purge concept for ice prevention 

Designed and fabricated a prototype Cold-Mate QD. 

Successfully tested the Cold-Mate QD prototype at a temperaon a cryogenic QD design. 

ture of -290 OF and a pressure of 50 psig. 

Participatiug Organization: VACCO Iiidustries, lnc. (W. Staliiccker mid D. 

Wicke) 

67 
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Collapsible Cryogenic Vacuum-Jacketed (CCVJ) Fluid Transfer lines 

Launch vehicle umbilical systems require vacuum-jacketed (VJ) fluid 

transfer lines to load cryogenic propellants at the launch pad. Since umbilical 

systems consist of mechanisms that extend and retract the umbilical plates to 

and from the ground to flight interface, these VJ lines must have flexibility or 

a degree of freedom of motion. Typically, this has been accomplished by using 

VJ flex lines. These flex lines may be as large as 8 inches in diameter, and the 

corresponding minimum bend radius is as large as 3 feet with corresponding 

reaction forces that must be controlled. These reaction forces are typically 

restrained by a large energy chain device such as the one shown in figure 1, 

which was used for the Atlas V Autocoupler umbilical system. 

The VJ flex lines are effective but require significant overhead in terms of 

additional mechanisms, structure, and space occupied. Clean launch pad 

concepts and other desired efficiencies require that the automated umbilical 

should have a better packaging ratio. 

To make more compact and efficient umbilical systems, a collapsible 

cryogenic VJ line was developed as a proof-of-concept model. The model 

shows that the VJ flex line can be replaced by a VJ hard line that contains one 

or more cryogenic swivel joints at critical points that then allow the VJ line 

to be folded up or collapsed. A VJ hard line is more thermally efficient than 

a VJ flex line, and it has a far better packaging ratio since the minimum bend 

radius is avoided. This is demonstrated in the proof-of-concept model and is 

shown in figures 2 and 3 in the extended and retracted positions, respectively. 

Key accomplishments: 

Developed CCVJ prototype concept. 

Researched and located unique CCVJ swivel joint. 

Designed and fabricated CCVJ proof-of-concept prototype. 

Contact: R.P. Mueller (Rob.Mueller@izasa.gov), YA-D1, (321) 867-2557 

Participating Organization: Chart Industries, Inc. (B. Appleton) 
Figirre 1. Aflns L7 Aiitocoiqder Uid7ilical With Eiiergy 

Cliniiz for Lal-ge-Bciirl-Radits VJ Flex Lilies 

Fi,yuie 3. CCV] Liiie Proof of Coiicept With Siuiid 

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Fiber-optic Quick Disconnect (QD) 

Launch vehicles require data transmission from the ground to the flight hardware. 

Typically the electrical signals used in the data transmission are conducted across traditional 

shell-and-pin connectors that rely on physical contact of the pin with a corresponding 

receptacle hole. These pins are often subject to corrosion, which causes a malfunction of the 

electrical connector. In addition, there are precise alignment and mating force requirements to 

make the connection. If the alignment is not achieved, then the result is a bent pin condition, 

which also causes a malfunction of the electrical connector. Furthermore, vibrations during a 

T-0 disconnect in the launch environment may cause damage to the pins and prevent reuse of 

the connector. 

To address these issues, an engineering evaluation was performed to identify viable 

solutions and concepts. The results indicated that using fiber-optic light signals to transmit 

data across the ground-to-flight interface would meet all the requirements. 

A fiber-optic QD can transmit data signals across the ground-to-flight interface with no 

physical contact. The light source is transmitted across a fiber-optic interconnect / cable 

assembly using expanded-beam technology, which physically expands and collimates 

transmission signals into an optical beam over 14 times its original diameter. It is then 

refocused back down into the core of the receiving fiber. This concept provides ease of 

alignment and low sensitivity to thermal changes and contamination. High-strength, 

precision connector housings enhance a durable connection, optimizing low loss and 

repeatable performance. Vibration effects are avoided because of the expanded-beam 

footprint. 

Such a fiber-optic QD is manufactured by Amp/Tyco Electronics. It is a Fiber-optic Cable 

Assembly (insert) part number POO-3691-16-1 and an Arinc 404 Single-Cavity Fiber-optic 

Receptacle Assembly (shell) part number POO-3690-26-1. 

Spaceport launch pad environments are unique and not typical of most industrial 

applications. Compatibility with cryogenic temperatures, vibrations, and salt fog is required. 

Such tests were conducted at the KSC Materials Science Laboratory. The fiber-optic QD was 

tested in accordance with the vibration requirements contained in KSC MSL-325-2001. The 

power spectral density (PSD) was defined in 79K31824, Revision A, as 28.2 g,,,. All vibration 

tests showed acceptable levels of transmittance during the vibration application. 

Expanded Beam Principle 

I 

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I 

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From left to nght Pro beam, Pm 

Beam Jr , and Mini-Expanded Beam 

I 

Product Facts 

- Light is expanded, collimated. 

and transmitted across an air gap 

- Ball lens expands cross-sectional 

area of light over 200 times for 

multimode and over 2000 times 

for singlemode 

I 

Seperable Interface 
Cryogenic temperature testing was performed at the Launch Equipment Test Facilitv 

(LETF) Cryogenics Testbed to liquid-nitrogen cryogenic temperatures (-320 degrees Fahrenheit 

[OF]). All cryogenic testing showed that signal loss was within acceptable limits as the 

temperature was lowered. 

The salt fog tests simulated the effects of being in a seacoast environment for 30 daw. The 

testing was conducted in a salt fog test chamber at the Materials Science Laboratory at KSC. 

During the 1,200-hour test, the fiber-optic QD performance did not degrade substantially, and 

the only effect noted was a buildup of salt on the housing. 

The results of environmental testing of the Amp / Tvco Electronics Fiber-optic Cable 

Assembly (insert) part number POO-3691-16-1 and an Arinc 404 Single-Cavity Fiber-optic 

Receptacle Assembly (shell) part number POO-3690-26-1 indicate that this fiber-optic QD is 

acceptable for use in a typical launch pad environment. It is anticipated that this fiber-optic 

QD will be used in future launch vehicle ground-to-flight-interface umbilical systems, and it is 

a candidate for other ground-to-ground interface connections in a launch pad environment as 

well. 

? 

Key accomplishments: 

Performed liquid-nitrogen cryogenic 

t'hermai environment pei-l'orilxiriw testirig. 

Performed launch pad vibrations 

environment performance testing. 

Performed seacoast salt fog environment 

performance testing. 

Identified Tyco / Amp Electronics FiberOptic QD as a viable candidate for use in 

ground-to-flight-interface applications. 

Loss vs Transmittance 

-350 00 -300 00 -250 00 -20000 -15000 -10000 -5000 0 00 50 00 100 00 

Temp (OF) 

-33 00 

-33 20 

-33 40 

-33 60 

-33 80 

-34 00 

-34 20 

-34 40 

-34 60 

-34 80 

-35 00 

-35 20 

71 
Ice Suppression Technologies 

At cryogenic temperatures, ice forms on chilled equipment because of 

the condensation and subsequent solidification of moisture that is present in 

Earth's humid atmosphere. The ice can become foreign object debris (FOD) 

and can interfere with mechanisms and motion or can compromise the sealing 

surfaces of launch pad umbilical systems. An effort was made to identify 

potential solutions to the icing problem that have specific applications to 

spacecraft umbilical systems and generic applications to launch pad cryogenic 

systems. 

A team of NASA and contractor design and operations engineers compiled 

a list of 38 potential solutions whose technology readiness levels (TRLs) 

ranged from very low (such as biological agents) to mid (such as piezo-electric 

vibration methods) to high (such as inert dry gas purges). 

The process yielded the following three methods of ice suppression for 

umbilical systems: 

1. Split Purge Shroud: The shroud encapsulates both halves of the umbilical 

system (ground and flight plates), which mate to form a common purge 

cavity. This approach was pursued with a proof-of-concept prototype and 

is documented separately as the Ice Suppression and Mitigation System 

(ISMS). 

Chamber: This apparatus is designed to encapsulate an individual quick 

disconnect (QD). The EIS-CP chamber is a donut-shaped device designed 

to create a contaminate- and moisture-free cavity around a subject quick 

disconnect. The device is 1-inch thick when compressed. It can be attached 

to an umbilical plate holding a subject QD. As the umbilical plates are 

demated, the EIS-CP chamber will expand and follow the moving umbilical 

plate to a height of approximately 6 inches. At this point, the upper part of 

the EIS-CP chamber will narrow to form a nearly sealed configuration, thus 

reducing the amount of purge gas needed and preventing contaminants 

from reaching the QD sealing surfaces. Once the EIS-CP chamber has 

performed its function and is ready to be retracted into its compressed 

position, the flow pressure to the inner bladder can be stopped. The 

gas pressure keeping the outer bladder expanded can be closed and the 

vacuum pump started, evacuating the air from the inside of the chamber 

causing a negative pressure. This will collapse the chamber and pull it 

down into its housing. The chamber would then be out of the way and 

cause no constraint to the umbilical plate remate operation. 

3. Infrared (IR) Heat and Gaseous Nitrogen (GN,) Air Knife: These two 

methods were used to prevent and mitigate ice buildup and were combined 

in various ways to provide information on the effectiveness of these 

methods. The IR lamps were powered by 480 volts (V) ac and cooled using 

an open-loop water line. They were mounted on a stand that held them 

vertical and parallel with a 1 /2-inch gap between them. The air knife was 

mounted on top of the flight plate so the stream would flow in a laminar 

sheet down the front of the plate. The GN, heaters were mounted on both 

ends of the air knife to provide enough heated gas to the knife. The results 

2. Expandable Ice Suppression and Cleanliness Preservation (EIS-CP) 
-350 

12 14 12 28 12 43 12 57 13 12 13 26 13 40 

Time (rnin) 

~ Baseline Test (Deicing equipment mas Tiof used. If took 30 iiiiniites 

for tl7c ice on the QD to melt aid oiler 2 hours to drtr.) I 

Test 7 

- MOD1 CH4 

- MODI CH5 

- MOD1 CHS 

1243 12 57 13 12 1326 13 40 13 55 14 09 

Time 

Test With Tziio IR Lanips and an Air Knife With Heated GN. 

(It took 0 minutesfur tl7e ice to inelt a?lLjlO iiiiiiufes to dryj 

show that the IR heaters are very effective at deicing the QDs. The 

more difficult problem is removing the water from the QDs. Using 

these two methods, it should be possible to get the melting time 

below 2 minutes and the drying time below 5 minutes. 

Q 

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s 

Participating Organizations: USTDC (1. K. fiautu1ein mid 1.1. Tozc1nsend) and 

American Remote Vision Co. (S. Gleninn) 

73 
Automated Latching Systems 

Automated latching systems are required for remote latching and 

preloading of ground umbilical plates to flight umbilical plates in spacecraft 

vehicle ground-to-flight fluid interface systems. Two such devices were 

developed at KSC. 

The Non-Pyrotechnic Automated Latch/ Lock System (NPALLS) is a highly 

innovative latching mechanism that can be remotely operated. It is based on 

an in-line gripper wire rope device shown in figure 1. 

The in-line gripper was used in a mechanism that pulls on a rod inserted 

into the in-line gripper. As the separation force is increased, then the gripping 

force increases proportionately with a multiplier effect on the force. This 

results in exceptionally high gripping loads for relatively low actuation 

forces. The actuation will be provided by a pneumatically actuated cylinder 

that will surround the in-line gripper mechanism. As a proof of concept, a 

scissors-type mechanism was used to actuate the in-line gripper, and it was 

demonstrated that a separation of 1,000 pound-force could be sustained by the 

in-line gripper device. This proof-of-concept apparatus is shown in figure 2. 

Another automated latching system developed at KSC is the Automated 

Pneumatic Collet Latch (APCL). An existing, manually operated collet 

latch design with flight heritage dating back to the Apollo and Space 

Shuttle programs was used as a baseline for the APCL. The collet latch was 

redesigned so that a pneumatic cylinder mechanism actuates the locking pin, 

which then spreads the fingers of the collet inside a receptacle housing on the 

flight-side plate. This results in a latching of the ground-side umbilical plate 

to the flight-side umbilical plate with the additional advantage of included 

preload forces. The pneumatic cylinder can be actuated by a remote signal 

command to a pneumatic valve, therefore achieving automation capabilities. 

Figure 3 shows an expanded view and an assembly view of the APCL. 

Key accomplishments: 

Developed NPALLS and APCL prototype 

Designed and fabricated NPALLS proof-ofCreated fabrication drawings for APCL. 

concepts. 

concept prototype. 

Contact: R.P. Mueller (Rob.Mueller@nasa.gozl), YAD1, (321) 867-2557 

Participating Organizations: American Remote Vision 

Co. (S. Glernaiz) and YA-D1 (A.C. Littlefield and T. 

Bonner) 

Figure 1. liz-Line Gripper Wire Rope Device 

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Umbilical Ice Suppression and Mitigation System (ISMS) 

and Cleanliness Verification System (CVS) 

Critical automated umbilical systems are retracted from the launch vehicle 

interface shortly before T-0 to ensure reliability of release and to protect the 

umbilical system from launch blast damage. In the event of a launch abort, 

the umbilical system must be reconnected to the launch vehicle interface to 

safe the vehicle and to drain hazardous cryogenic propellants for reuse. A 

chilled umbilical system that is retracted and exposed to humidity in the 

atmospheric air will instantly form a layer of frost, which can turn into 

ice. The frost/ice can compromise sealing surfaces and can inhibit correct 

functioning of the umbilical system’s mechanisms and dynamics. Therefore, 

to enable automated umbilicals, ice must be suppressed from forming, or if 

that is unavoidable, then its presence must be mitigated. 

In addition, the umbilical and all sealing surfaces must be clean and free of 

any contamination. Debris or other contamination in the quick disconnects 

(QDs) can cause improper sealing, resulting in a hazardous propellant leak, or 

may cause impurities to be introduced into the main propulsion system. 

These factors dictate that an automated umbilical system employs an Ice 

Suppression and Mitigation System (ISMS) and a Cleanliness Verification 

System (CVS). A proof-of-concept prototype incorporating an ISMS and a 

CVS was conceptualized, designed, built, and tested. 

The ISMS consists of a spring-loaded, dry, inert-gas purge shroud that 

surrounds the entire QD umbilical plate on both the flight side and the ground 

side. A sliding door keeps the purge contained on each umbilical plate when 

they are separate. The gaseous-helium or gaseous-nitrogen purge prevents ice 

from forming since the humid atmospheric air is kept away by the positivepressure purge gas. When the umbilical plates are mated, the shroud retracts 

with a spring-loaded mechanism, and the purge shroud doors open after the 

two halves of the purge shroud have mated. The umbilical continues to move 

forward until the actual QDs have mated. During the entire operation, the 

QDs are continually purged so there is no chance of any frost / ice formation. 

To verify there is no contamination present on the QDI the CVS employs 

a video camera that transmits a digital image to an electronic image analysis 

device. As long as there is no change in the digital image, the alarm does not 

sound. As soon as a piece of foreign object debris (FOD) lands on the QD, 

there will be a change in the digital image, which causes an alarm to sound 

with a corresponding signal. Once the FOD is identified visually on a display 

monitor, the operator can remotely operate a pulsating pressurized gaseousnitrogen jet nozzle, which is aimed at the FOD on the QD. The pulsating 

gaseous-nitrogen jet blows the FOD off the QDI and a visual inspection 

verifies the QD is clean. 

The purge shroud prevents ice/ frost formation and also maintains a clean 

environment. The CVS allows for a final inspection and corrective action by 

blowing the FOD off the QD. 
The ISMS and CVS prototypes were 

tested at liquid-nitrogen temperatures 

with very positive results. There was 

no ice formation and the CVS was 

demonstrated to be an effective concept, 

I Key accomplishments: 

I Developed ISMS and CVS prototype 

concepts. 

prototypes. 

prototypes at liquid-nitrogen temperar Designed and fabricated ISMS and CVS I 

I Successfully tested the ISMS and CVS 

I ture. 

Partici;~atiizg Orgnzizclfioii: Aii7ericclii Rriiiofe 

Visioii Co. (S. Glcinniz) 

CVS Vision Sysfein Coiztrol Pmel 

Proqf-of- Coizcepf Profotype ISMSICVS 

Mouiited oiz cl Mobile Cart 

77 
.- z Testing of Space Shuttle Spray-On Foam Insulation Under Cryogenic 

Vacuum Conditions 0 

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A key element of space launch vehicles is thermal insulation for cryogenic 

tanks, and the most commonly used thermal insulation material is SprayOn Foam Insulation (SOFI). This lightweight material, a Freon-blown 

polyurethane foam, provides an excellent combination of thermal and 

mechanical properties. However, this type of material is generally good 

only for a one-time-use tank due to problems with multiple thermal cycling 

and environmental degradation. The reusable launch vehicle of the future 

will require a fully reusable high-performance thermal insulation system for 

its cryogenic propellant tanks and feedlines. The cryogenic tank insulation 

must perform well in all phases of ground servicing, ascent, on-orbit, and 

reentry mission operations. A well-established reference point based on 

SOFI is therefore needed for the new materials technology development 

work. Thermal performance testing of SOFI under full-range cryogenic 

vacuum conditions was performed at the KSC Cryogenics Test Laboratory 

in 2002. The SOFI test specimens were sprayed by United Space Alliance 

personnel. 

The steady-state liquid nitrogen boiloff (evaporation rate) calorimeter 

methods established by the Cryogenics Test Laboratory were used to 

determine apparent thermal conductivity (k-value) of the insulation 

material: Cryostat-1 (Cylindrical Test Apparatus) for absolute k-values and 

Cryostat-4 (Flat-Plate Test Apparatus) for comparative k-values. An overall 

view of the Cryostat-1 apparatus is shown in figure 1. A liquid nitrogen cold 

mass maintained the cold-boundary temperature (CBT) at approximately 78 

kelvin (K) (-319 degrees Fahrenheit [OF]). The warm-boundary temperature 

(WBT) was maintained at approximately 293 K (+68 OF) using an external 

heater with electronic controller. The temperature difference was therefore 

215 K (387 OF) while the mean temperature was about 186 K (-125 OF). 

Vacuum environments, or cold-vacuum pressures (CVPs), included the 

following three basic cases: high vacuum (HV) (below l~lO-~ torr), soft 

vacuum (SV) (1 torr), and no vacuum (NV) (760 torr). Additional tests were 

performed at cold-vacuum pressures from 1 x 

portion of a test from Cryostat-1 is shown in figure 2. 

torr to 760 torr. The key 

Thirty-three tests of three different specimens were performed. The 

combined run-time of steady-state liquid nitrogen evaporation was over 

500 hours (not including the vacuum pumping and heating cycles). Each 

of the three test series included at least eight specific vacuum levels (eight 

tests) from high-vacuum to no-vacuum conditions. The test conditions were 

representative of the actual-use conditions for cryogenic insulation. The test 

measurements were made at the full temperature difference and included 

the full-vacuum pressure range. The results, in terms of k-value and 

mean heat flux, are currently being used as a calibration reference and as a 

benchmark of comparison with new materials for next-generation thermal 

insulation systems. Plans for further work include the experimental study 

of environmental effects, such as aging and moisture intrusion, and the 

incorporation of low-temperature compatible coatings. 
2000 

1800 

1600 

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400 

200 

0 

Contacts: J.E. Fesmire (~aiires.E.Fesiirire~ izasa 

.pi), YA-C2-T, (321) 867-7557; and Dr. M.K. 

b’illiains, YA-CZ-T, (321 j 867-4554 

Participating OrRanizatioii: USTDC (K. W. 

Heckle, Sr. and Dr. S.D. Augustynoulicz) 

Fiprc 1. View of Cryostat-1, Cryogenic lnsulation 

Test Apparatus for Cylindrical Specimens, With the 

L’acuuin Can aizd Instrumentation Removed (The 

SOFl test article C136 is shozon mounted on the cold 

mass ofthe cryostat.) 

10.0 

9.0 

8.0 

7.0 

6.0 

5.0 

4.0 

3.0 

2.0 

1 .o 

0.0 

3 4 5 6 7 8 

Time (hours) 

Y 

IFigure 2. Example of Test Results From Cryostat-1, C136, Test 17: Variation of Liquid 

Nitrogen Boiloff Flozo Rate aizd Apparent Tlrermal Conductioity (k-zialue) With Time 

(The cold-zwcwn pressure is q7proximately 0.1 inillitorr.) 

79 
Glass Microspheres for Cryogenic Insulation Applications 

A common and critical element of space launch vehicles and systems is 

thermal insulation for cryogenic tanks and piping. Glass microspheres, also 

known as glass bubbles, represent an alternative insulation material for a 

number of applications. Commonly used materials, such as Spray-On Foam 

Insulation (SOFI) for vehicle tanks and perlite powder for ground storage 

tanks, are targeted for replacement with the new glass bubbles systems. 

Determining a complete thermal characterization of the glass bubbles was 

the first step in producing the engineering solutions required for the energyefficient, low-maintenance cryogenic systems of the future. The testing was 

performed at the Cryogenics Test Laboratory at KSC in 2002. 

The steady-state liquid nitrogen boiloff (evaporation rate) calorimeter 

methods established by the Cryogenics Test Laboratory were used to 

determine apparent thermal conductivity (k-value) of insulation material 

systems. The insulation test material was glass microspheres in bulk form. 

This economical material is manufactured by 3M. These glass bubbles 

are hollow and are formulated for a high strength-to-weight ratio. The 

target isostatic strength, or crush strength, of Type K1 glass bubbles is 1.7 

megapascals (MPa) (250 pounds per square inch [psi]). The particle size 

distribution ranges from 30 to 110 microns with a mean particle size of 65 

microns. The bulk density is typically around 0.060 gram per cubic centimeter 

(g/cm3). 

Fourteen tests were performed using the cryogenic insulation test 

apparatus for cylindrical specimens, Cryostat-1. The test measurements were 

made at the full temperature difference (typical boundary temperatures of 

78 kelvin [K] and 293 K) and included the full vacuum pressure range. The 

results were calculated in terms of k-value and mean heat flux. The test 

conditions were representative of the actual-use conditions for cryogenic 

insulation. This research showed that a number of tests under different 

conditions are needed to understand and characterize the glass microsphere 

material for the full vacuum pressure range. These results will be applied to a 

number of different projects in energy-efficient cryogenics for space vehicles, 

space launch, and industry. 

Different Cryogenic Insulation Materials in the Bulk Fill Form 
LO, Storage Tank at Launch Complex 39B 

Key accomplishments: 

Characterized material thermal performance under cryogenic 

vacuum conditions. 

Tested structural insulation panel system using glass 

microspheres. 

Key milestones: 

Experimental study of vacuum and moisture properties of 

Testing cryogenic tank insulation system. 

material. 

Pal-ficipnfiiig Orgaiiizafions: USTDC (DI.. S.D. Aztgztsfyizmic: mid K. W. 

Heckle, SI..) and Teciinolog~/ Applicnfioiis, liic. (M. Allen) 
Standardized Testing of Experimental Thermal Insulation Materials for 

Low-Temperature Applications 

A major function of the Cryogenics Test Laboratory is to develop a 

world knowledge base for low-temperature thermal insulation systems, 

materials, and testing. The Cryogenics Testbed now has extensive data 

on the performance of a wide range of insulating materials at cryogenic 

temperatures. Ambient temperature data for well-established materials 

are often readily available from the literature, material vendors, or other 

Government agencies. However, many new or experimental insulation 

materials do not have any standard thermal performance characterization 

to compare with other materials. Furthermore, outside data sources are 

not consistent in listing test conditions, boundary temperatures, and other 

important testing details. 

To meet the need for standardized testing of experimental thermal 

insulation materials under ambient conditions, a Netzsch Lambda 2300F 

Heat Flow Meter has been activated at the Cryogenics Test Laboratory. This 

in-house capability, using the American Society for Testing and Materials 

(ASTM) C-518 standard protocol, is being used to complement the extensive 

technology development work being performed using the family of five 

insulation research test cryostats currently in service. While the cryostats 

provide apparent thermal conductivity values for materials under the 

simulated actual-use conditions of cryogenic temperature and full vacuum, 

the heat flow meter provides calibrated thermal conductivity values for a 

material at ambient temperature and ambient pressure only. The new selfcontained unit provides quick consistent results for material screening and 

baseline performance comparisons. 

The operational techniques and procedures of the Heat Flow Meter were 

successfully demonstrated with the proper level of accuracy and consistency 

within the scope of its intended function. The results of the tests conducted 

on a standard reference material, Spray-On Foam Insulation (SOFI), were 

consistent with values obtained from the National Institute for Standards 

and Testing. Important data were also obtained for aerogel beads, aerogel 

blankets, glass microspheres, experimental composites, and a number of 

different types of polyimide foams. In addition, a device and procedure to 

accommodate “loose fill” and uniquely shaped materials in the test apparatus 

were successfully developed. This new capability will continue to support 

research projects for the development of next-generation thermal insulation 

systems. 
Netzsclr Lainbdn 2300F Hent Floii, Meter With Typical Sample 

Key accomplishments: 

I Activation of standardized thermal insulation 

test machine. 

material screening. 

Development of new capability for experimental 

Key milestones: 

Building insulation material thermal 

Testing support of next-generation thermal 

performance database. 

insulation systems research. 

83 
.- G Thermal Performance Testing of Cryogenic Piping Systems 

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From food industry, transportation, energy, and medical applications to the 

International Space Station, cryogenic liquids must be stored, handled, and transferred 

from one point to another. The basic problem for any application of cryogens is 

protection against heat transfer to minimize the evaporation ratio. This energy 

conservation translates into a lower cost of operation for the user. There are a number 

of scenarios for protecting storage tanks and transfer lines from heat. For example, 

short-term storage, using the cryogen to provide a source of gas, does not require a 

tank or piping system with an extremely low thermal conductivity factor. System 

process control is still important, and the thermal insulation systems must be designed 

accordingly. Applications in which long-term storage and maintaining the quality of 

the liquid are critical require high-performance materials that provide the most effective 

level of thermal isolation that can be justified by design and economics. Whatever the 

cryogenic application, knowledge of the system’s operation from the heat transfer point 

of view is essential. That knowledge will prevent the higher-than-expected life cycle 

operating costs that often result from selecting the least expensive thermal insulation 

system, despite its initial cost savings. 

Ia, 

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3 

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The Cryogenics Test Laboratory has developed a new Cryogenic Pipeline Test 

Apparatus for testing pipelines under different temperature, vacuum pressure, and flow 

conditions. With cryogenic use more prevalent and with the requirements for higher 

and higher energy efficiencies, the total system design must be understood in light of its 

heat-load effects. Up to three different pipelines, in lengths of 18 meters or longer, can 

be tested simultaneously. The apparatus gives accurate heat leakage data for the test 

article under simulated conditions of an actual field installation. 

The apparatus includes two cold boxes, one upstream and one downstream, between 

which the test articles are connected. The method of testing can be static (boiloff) 

or dynamic (flow-through). Figure 1 gives a simplified schematic of the apparatus 

configured for the static method of testing. An overall view of the upstream cold box, 

showing three different 18-meter-long cryogenic pipelines under test, is shown in figure 

2. The cold boxes are generally oriented so the downstream end of the pipeline is 

slightly elevated. The cold boxes are mobile such that the apparatus can be reconfigured 

for different test requirements and a variety of test articles. 

The apparatus and method of operation can provide the basis for a standardized test 

of the thermal performance of pipelines or piping systems. Existing standards have 

strict limitations, such as applicability for system operating temperatures above ambient. 

The pipeline test apparatus at the Cryogenics Test Laboratory is U.S. patent-pending 

and currently slated for use in a number of applied research projects supporting the 

development of energy-efficient space launch sites. Preliminary work on the insulation 

envelopes for flexible high-temperature superconducting (HTS) power cables has been 

completed in collaboration with the Department of Energy and Oak Ridge National 

Laboratory. Evacuated microsphere insulation panels for cryogenic piping systems are 

being developed in collaboration with Technology Applications, Inc., Boulder, Colorado. 

Another project, with AMAC International, Newport News, Virginia, is the development 

of a cryogenic transfer line with magnetic suspension. 

This work was supported in part by the Center Director’s Discretionary Fund (CDDF) 

project ”Cryogen c Piping Networks for Energy-Efficient Launch Sites.” 

I 84 
1 I 

+ Line TFturi Sensor iocations + 

+ I 

IilOff 

Meters 

Downstream I L+&o Tank v.1 

W Line Pressure and High-Point Bleed 

Figure I. Simplified Mechanical Schematic of tlie Cryogenic Pipeline Test Apparatus 

Key accomplishments: 

Developing a standardized method 

of static heat leak rate testing. 

U.S. patent application placed 

for test xr.easurement system and 

method. 

commercial vacuum-jacketed 

flexhose. 

Testing of state-of-the-art 

Key milestones: 

Produced prototype experimental 

pipeline for demonstrating new 

insulation technology. 

Generated paper and presentation 

for the International Congress of 

Refrigeration, Washington, D.C., in 

August 2003. 

Provided testing support for Small 

Business Innovation Research 

Figure 2. View of tlie Upstream Cold Box cftlie Cryogenic Pipeline Test Apparatus With Thee 18-MeferLong Test Articles Installed 

(SBIR) projects with Technology 

Applications, Inc., and AMAC 

International. 

Contacts: I. E. Fesmire [\~7i1~5 E.Frt;in i iP(/ ii 

Breakfield, YA-F2-P, (321) 567-1923 

Participating Organizatioiz: USTDC (Dr. 

S.D. Aiigii~tynoz~~icz, Z.F. Nagy, K.R. 

Cummings, and J.E. Gibson) 

i7~ ;;CY), YA-C2-T, (321) 867-7557; al~d R.A 

SI 

3 

n 

I 
Piping System Repair Using Cryogenic Freeze Plug Technology 

A safe and reliable method for isolating leaks in cryogenic systems has been 

developed by the Cryogenics Test Laboratory. In September 2002, a leak caused by 

a cracked gasket upstream of a liquid oxygen storage tank block (shutoff) valve at 

Pad 39A threatened to delay a Space Shuttle launch. The new cryogenic freeze plug 

technology was successfully applied and the cracked gasket was replaced without 

having to drain (and potential waste) many thousands of gallons of liquid oxygen. 

This significantly reduced the cost and amount of time needed for the repair and 

allowed the Shuttle to launch on time. 

Previously, this type of repair required draining the 900,000-gallon liquid oxygen 

storage tank to permit replacement of the pipe sealing gasket. Thermal cycling due 

to draining / warmup induces stresses on the storage vessel and presents a safety 

consideration in that the number of drain/ warmup cycles is limited. 

Creating a freeze plug to isolate the failed gasket and to allow its changeout was 

discussed with lab personnel. Because liquid oxygen in its solid-state is unstable, 

freezing it was ruled out. The team investigated an alternative methodology, and 

an acceptable secondary fluid was introduced into the system downstream of the 

failed gasket. This secondary fluid was frozen within the cryogenic piping to create 

an effective plug in the line. This allowed the gasket to be safely replaced without 

draining the tank. 

To accomplish this repair procedure, several technical issues had to be addressed: 

a dense liquid with acceptable thermal properties had to be found, and it needed to 

be compatible with liquid oxygen. In addition, it had to be easy to remove, down 

to a residual concentration level of less than 10 parts per million. A liquid precision 

cleaning agent was selected and successfully tested on a high-fidelity mockup at 

the Cryogenics Test Laboratory. These successful demonstration tests and the data 

obtained were presented to the Engineering Review Board. The repair procedure 

was then approved for implementation at the launch pad. 

The cryogenic freeze plug technique for piping system repair is being further 

developed for use on larger line sizes and with different fluids. 

Key accomplishments: 

The gasket was changed successfully in a liquid oxygen tank outlet pipe with no 

U.S. patent pending. 

thermal cycle of the tank required and no commodity loss. 

Key milestones: 

Investigate using the technique on larger diameter piping systems. 

Investigate alternative secondary fluids. 

Contact: C.E. McCaskey (~iI~i,i71,[f,Mi.Cf7~/~~,1/-1 t18~1~75~7.,y0~l), PH-G1, (321) 861-3754 

Participating Organizations: USTDC (Z.F. Nagy, K.R. Cummings, and ].E. Gibson) and United 

Space Alliance (P.J. Klonowski) 
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Cryogenic Freeze Plug Tecliizique Beiiig ,4pplied oil Piping at Launch Complex 39A 

Liquid Storage Tniik 

87 
Automated Thruster Bench Test (ATBT) System 

The objective of this effort is to develop a ground test system for 

Shuttle thrusters that will reduce the possibility of human error. In the 

past, test results were often recorded incorrectly due to errors in math, 

data interpretation, and data transcription from the test equipment to the 

procedure. While these errors have been greatly minimized in recent years, a 

backup system to the current test setup was still needed. 

The system developed uses a commercial off-the-shelf data acquisition 

system, a standard office PC, and LabVIEW software to perform all thruster 

tests. Ground testing of thrusters includes solenoid valve timing and valve 

leak rates, heater checkout, and a three-point calibration of the chamber 

pressure transducer. The computer performs all data reduction, eliminating 

mathematical errors. All data is recorded directly onto the computer hard 

drive, eliminating transcription errors as well. The data is stored in a standard 

ASCII text format, allowing easy access by computers connected to the 

network. 

The Shuttle primary thrusters use a solenoid valve with a pressure-assisted 

main stage and a magnetically actuated pilot stage. Requirements for these 

valves call for determining the valve opening and closing times. This can be 

determined for the solenoid primary stage from characteristic "peaks and 

valleys" in the voltage and current during valve opening and closing. These 

peaks and valleys correspond to movement of the magnetic pilot stage. 

Movement of the nonmagnetic primary stage is determined by output from an 

accelerometer. The accelerometer records vibrations through the valve body 

that occur when the primary stage contacts the valve stops. 

The software also monitors the valve timing data for adverse trends and 

calculates statistical control limits from a database of over 100 previous 

thruster bench tests. A warning is sent to the user if any of the timing data 

matches the following criteria: 

Outside of the requirement limits. 

Outside of a 3 sigma control limit (mean 53 standard deviations). 

More than 5 data points trending up or down. 

More than 10 data points that do not vary by more than 1 standard deviation. 

These control limits are particularly important in this application since 

the requirement limits are much larger than the typical spread of the data. 

Without trend monitoring and control limits, abnormal readings could go 

unnoticed. 

Key ATBT features include a software test mode, which allows all the 

functions of the software to be tested; determination of valve timing data by 

the software by detection of peaks and valleys in the solenoid valve voltage, 

current, and accelerometer readings; and application of control limits and 

trend analysis to valve timing data. 

Contact: A. 0. Kelly (A I irlreip. 0. Kell!/in'iznsn.~oii), PH-G2, (321 ,) 862 -392 7 
.. 

3- 

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ATBT Prototype System 

Softicwe Detectioii of Thruster Vnlae Timiiig 

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Venturi Off-Set Technology Cryogenic Valve 

KSC Spaceport Engineering and Technology Directorate is partnering 

with Big Horn Valve, Inc., Sierra Lobo, Inc., and the Florida State University 

(FSU) National High Magnetic-Field Laboratory to prototype and test an 

axially operated, high-efficiency, low-mass cryogenic valve. This valve design 

innovation, termed Venturi Off-Set Technology (VOST), addresses four critical 

needs for effective storage, transfer, and use of cryogens in aerospace: 

Thermal isolation to minimize heat leakage. 

Low-pressure drop to accommodate high flows without cavitation. 

Minimal mass and space requirement due to low profile. 

Adaptability to electromechanical actuation (see the figure). 

Existing cryogenic valves require elaborate insulation measures to mitigate 

heat leakage through actuator stems. They are also characterized by highpressure drop at high flows, large mass-to-envelope ratio, and high actuation 

torques. VOST valves eliminate actuator contact with internal wetted parts, 

potentially improving thermal and fluid containment characteristics necessary 

for cryogenic propellants, while maintaining high flow. The VOST design 

is an axially actuated, high-flow, low-pressure-drop, low-mass, thermally 

isolated valve. The combination of high fluid flow, minimal pressure drop, 

and axial design envelope is unique and significant among cryogenic valves. 

The VOST design will eliminate perpendicular actuation devices that are 

VOST - Electromagnetically Activated Valve - Isolation Characteristics 

Insulation or / 

VOST Valve Cutaway 

90 
known, significant heat leak sources, minimize 

pressure drop under high-flow conditions, 

reduce device mass, and enable a device 

envelope that is suited for superior insulation 

and electromechanical actuation. Currently 

available cryogenic valves require complex and 

bulky insulation strategies to maintain safe and 

predictable fluid conditions, such as warmgas barriers necessary for effective stem seal 

operations. VOST valves are ideally suited for 

electromechanical actuation because only low 

torque is required to open and close them. These 

features combine to improve valve safety and 

reliability necessary for use of cryogens on Earth, 

in space, and in extraterrestrial environments. 

Applications of this technology include storage 

and transfer of cryogenic fuel for testing facilities 

and spacecraft. Non-NASA applications include 

but are not limited to production facilities for 

cryogenic and toxic fluids, short- and long-distance 

transfer of cryogenic and tsxic niiids, and the 

control of cryogenic and toxic fluids within process 

systems (such as refineries and tank farms). 

Participating Organizations: Big Horii Valve, lnc. (J. 

Steider), USTDC (Z.F. Nagy and S.J. Sojourner), and 

Florida Stnfe Uiiioersity National High Magnetic-Field 

Laboratory (5’. Van Scizw) 
Range Technologies 

Range Technologies consist of unique facilities and equipment technologies that provide 

the control and measurement data necessary to ensure the safety of launch and test operations 

for the range / spaceport. Examples of these technologies include (1) data processing equipment for range safety analysis, communications systems for telemetry processing, reception, 

and destruct command systems; (2) sensor and instrumentation systems for vehicle tracking, 

such as radar systems and optical tracking displays; and (3) data communications management, distribution, and recording systems. Significant capital investment is required to build 

and maintain this capability and as a result, these assets have traditionally been slow to modernize or have limited flexibility to accommodate new requirements and vehicles. In addition, 

the cost associated with range system infrastructure is difficult to quantify. 

Future range operations will require flexible range technologies that can incorporate the 

latest technological advances. It is expected to be largely space-based and will be seamlessly augmented with mobile, deployable, and / or ground-based assets. Minimal ground 

infrastructure will be required and will primarily be located at the launch head, reducing the 

operational cost. Self-healing autonomous ground and flight systems will become the norm. 

Range systems will have situational awareness and be intelligent enough to reconfigure 

autonomously when individual systems "drop out" or are unavailable. A National / Global 

Range Network with regional hubs is envisioned, which provides 24 / 7 coverage, is scalable 

to user requirements, and is seamlessly integrated into the national and global air space. The 

modernization activities will also ensure that future technology improvements will be able to 

uration for specific launch vehicles operations will be minimized. Multiple simultaneous test 

I 

1 adapt to different space flight hardware designs, and the degree of customization or reconfigand launch operations by multiple vehicles from multiple spaceports are also envisioned. l 

I 

With the goals and objectives of improving safety, improving range responsiveness (support more launches, reduce lead and turnaround time, and reduce delays), reducing variable 

launch and range costs, reducing cost of fixed infrastructure, and improving capacity and flexibility, the KSC team has stepped up to colead, with the Air Force Space Command, a national 

coalition called Advanced Range Technologies Working Group (ARTWG), which is identifying 

the next-generation range technology roadmaps for the next 25 years. 

The seven range technology areas of focus include: 

Tracking and Surveillance Technologies 

Telemetry Technologies 

Command and Control Technologies 

Communication Architecture Technologies 

Weather Instrumentation and System Technologies 

Decision Modeling and Analysis Technologies 

Planning, Scheduling, and Coordination-of-Assets Technologies 

93 
An overall system capability roadmap has been developed. In additional to the system 

roadmap, individual technology roadmaps have also been developed for each of the seven 

technology focus areas. The complete set of roadmaps is available at the ARTWG Web page 

at /rffLt //Il/U\YC KSC \iASA COV. 

. Demonstrations and Additional operational use 

leveraged use 

GPS, TDRS, etc.) 

mobile/deployable assets space, air, ground assets 

with space and demos with space basea se~essiy augmentea WM 

based constellations (e.g. assets (reduotlm ofgr~d~~~) mobile, deployable, Pmd aSw 

Imulementation with ~Groundassatsminiaudody 

with Improved integration of Autonomous seamless htegfation 

at local departwe point 

of space, air, & ground assets 

mobile/deployable assets Demonstrations 

(UAV, Balloons, etc.) Improved modeling systems, sslfa 

slllmticmlawaaar Improved modeling datahme svstems - 

systems, database systems . Automate brocesses, systems Implementation of autonomous 011 

(knowledge-based, data and functions (semi- board systems (ifdesired) I mining - etc.) autonomous. intelligent) I, * “Realistic” virtual modelitlg 

Semi-automate processes, Demonstration of on-board ‘y-9 database 

Improved interoperability Improved interoperability of 

System and functions With 

FAA and throuhout 

multiple suaceuort & ranges 

Modemized Sensors 

systems, and optimized use of 

autonomy, including &confliction 

National/Global range network 

system providing complete earth 

systems and functions LV autonomy 

and integration of systems 

throughout a single covm@ 

space dran e 24nawSags 

. 

ARTWG Range System Capability Roadmap 

The national Advanced Range Technologies roadmaps and the KSC Range Technologies 

Roadmap are being used by KSC to help define our technology development plans and strategies. 

The overall goals of both of these road-mapping activities are to identify technology 

improvement plans for future launch and test ranges that are more reliable, maintainable, 

available, and operable; more adaptable to fit the mission; more flexible and allow for more 

capacity; more integrated with other systems (range, spacecraft, launch vehicle); more customer- and user-friendly; and more economically viable. 

The technology projects that follow provide a glimpse of some of the current technology 

development activities to allow for the significant improvement desired. 

For more information regarding Range Technologies, please contact Richard Nelson, ~~~~~I~~I~

m ~r~?w.,qi~~~, YA-D7, (321) 867-3332, or Darin Skelly (D(II.~II.,\ 1.Ski~//p irf7ii7.,yoi9, YA-E, (321) 861-3639. 
Range Technologies Roadmap 

f 3s 

wu 

urn 

95 
96 

Space- Based Telemetry and Range Safety (STARS) 

The goal of the Space-Based Telemetry and Range Safety (STARS) project is to 

determine the feasibility of using space-based assets, such as the Tracking and Data 

Relay Satellite System (TDRSS) and Global Positioning System (GPS), to increase flexibility and minimize range operation costs by reducing the need for ground-based 

range assets and infrastructure. 

STARS is divided into the Range Safety (RS) and Range User (RU) subsystems. The 

RS subsystem provides GPS data for vehicle tracking and Flight Termination System 

(FTS) commanding, and the RU subsystem provides vehicle telemetry such as video, 

voice, and data. The core components for the RS subsystem for Flight Demonstration 

1 include a versatile low-power transceiver (LPT) (figure 1) for bidirectional communications with TDRSS and ground-based facilities near the Launch Head, a state-ofthe-art Command and Data Handler (C&DH) (figure 2) to process the FTS commands 

and format the return link telemetry, and a commercial high-dynamics GPS receiver. 

The RU subsystem uses a commercial mulitplexer S-band transmitter and S-band 

antennas. 

Data from the flight demonstrations will be distributed in real time to Dryden 

Flight Research Center (DFRC), KSC, and Wallops Flight Facility (WFF) via White 

Sands Complex (WSC) (see figure 2). KSC has integrated a ”Range Simulator” using 

commercial off-the-shelf software and hardware to support STARS and other Range 

Technology projects. The simulator was used to support acceptance testing of STARS 

hardware and will be used to graphically display flight parameters and telemetry data 

in real time during flight and to aid in postflight analysis. 

For Flight Demonstration 2, the LPT, GPS receiver, and the C&DH will be combined into a single box that will perform nearly all the Range Safety functions, and the 

FTS command link will be encrypted and encoded. The RU subsystem will use a Kuband transmitter and a Ku-band phased-array antenna that is under development to 

support increased communications bandwidth requirements. 

Figure 1. Low-Power Transceiver (LPT) 

Figure 2. Command and Data 

Handler (C&DH) 
TDRS \ 

1- 

STARS Flight Demonstration Data Flow 

- 400-bps Forward Link -%-I 

I I L 400-bps Forward LinkJ 

STARS is truly a One NASA project. Eleven NASA facilities, as well as the Eastern and 

Western Ranges, are involved. 

I 

Key accomplishments: 

Developed and delivered RS and RU hardware to DFRC. 

Integrated STARS hardware on F-15B. 

Successfully completed Flight Demonstration 1, which consisted of a series of seven flight 

I 

tests to prove the concept of a spaced-based range. 

Key milestones: 

2003: Flight Demonstration 1 on an F-15B at DFRC. 

2004: Flight Demonstration 2 on an F-15B at DFRC. 

2005: Flight Demonstration 3 on a hypersonic vehicle. 

Contacts: R.A. h'ek0Jl (Riciiniil.A.S~I~l~iz~ )ins~7.~0?), YA-D7, (321) 867-3332; L.M. VnIeiicia, YA-E6, (322) 

851-7682; ni~d E.C. DCllSO71, Y,4-D7, (321) 857-6537 

Pnvticipntiiig Ovgantatioiis: USTDC (R. B. Bin), GSFC (T.C. Sobclink), DFRC (R.D. Snkalmra md D.E. 

Whiteinan) WFF (S.N. Buiidick), WSC (1.M. Gni)uvrr), n17d ITT (M. Hrrrlncher and S. Casfed) 
Iridium Flight Modem 

This has been a very successful second 

year for Iridium Flight Modem, a joint project at KSC and Wallops Flight Facility (WFF) 

to investigate the iridium satellite system 

for low-rate, full-duplex aerospace-related 

communications (Global Positioning System 

[GPS] tracking data and commands). Flight 

Modem uses commercial off-the-shelf communications equipment, a high-dynamics Coarse Acquisition (C/ A-Code) GPS 

receiver, and a PC-104 computer. The target 

applications for this project are tracking and 

telemetry for weather balloons, manned and 

unmanned aircraft tracking and commanding, and possibly two-way backup communications systems for rockets. 

Iridium is a commercial global satellite cell phone system for voice and data. 

Worldwide coverage is provided by 66 low 

Earth orbit satellites equally spaced in 6 

orbital planes at an altitude of 780 kilometers (km). Each satellite has 48 spot beams, 

and each spot beam has a diameter of 300 

to 400 km on the Earth's surface. Because 

these beams overlap, each satellite can effectively cover a circular region on the ground 

with a diameter of 3,000 to 4,000 km. 

L-band frequencies are used for the uplink/ 

downlink segments, and the signaling is frequency 

division multiple access / time division multiple 

access. The guaranteed data rate is 2,400 bits 

per second (bps) with a 12-decibel link margin, 

although this can be increased to 3,800 bps with 

a smaller link margin. The primary gateway is 

in Tempe, Arizona, and a Department of Defense 

gateway is in Hawaii. The main satellite ground 

control station is in Leesburg, Virginia, and a 

backup facility is being added in Chandler, Arizona. The current cost of modems is about $1,200 

and is expected to drop to $200. Airtime currently 

costs about $0.80 per minute and may also decrease 

in the future. 

Various types of connections are available, 

including mobile-to-PSTN (public switched 

telephone network), PSTN-to-mobile, mobile-tomobile, and direct Internet, where Iridium is the 

Internet service provider. Short-burst messaging 

capability is in beta testing and mobile-to-direct 

Internet protocol is planned for the future. 

Flight tests indicate that Iridium Flight Modem 

operates reasonably well for relatively slow 

moving aircraft, although the currently attainable 

reconnect times of 20 to 30 seconds may preclude 

GPS Satellite , Iridium Satellite Iridium Satellite 

/ 

Combined 

Iridium and 

GPS Antenna A // A /' 

Iridium Flight Modem 

PC-104 Computer 

Configuration of Iridium Flight Modern During the First TWO Test Flights 
Integrated lridium Flight ModemlComputer Box 

its use as a primary Range Safety asset. Flight 

Modem’s performance during high-dynamic 

and high-altitude (greater than 100,000 feet) 

maneuvers with large Doppler shifts and 

decreasing spot size wiii be tested on an upcoming sounding rocket flight. 

A future weather-related project under consideration is using Iridium Flight Modem to 

transmit the current position of surveillance 

planes to an integrated weather display for prelaunch operations. 

Key accomplishments: 

P November 2001: Developed flight-ready 

prototype of Iridium Flight Modem. 

December 2001: Performed ground-based 

demonstration with GPS data sent from KSC 

to WFF via iridium satellite system. WFF 

verified data originated from KSC with 

mapping software. 

March 2002: Conducted first aircraft flight 

test on a Piper Cherokee 140 flying in the 

KSC vicinity. 

May 2002: Conducted second flight test 

on a Piper Cherokee 140 flying a series of 

banked turns and spirals off the coast of KSC. 

Changed to a Linux operating system. 

June 2002: Conducted third and fourth flight 

tests on a P-3 Orion. These were flights of 

opportunity and consisted of a 3-hour roundtrip training flight from WFF to Greensboro, 

North Carolina, and a 3-1 /2-hour geophysical 

mapping operation from WFF to Iowa. 

July 2002: Developed prototype multiplexed 

dual modems. 

February 2003: Designed and built an 

integrated computer/modem box for quick 

installation on aerospace vehicles. 

March 2003: Conducted first unmanned aerial 

vehicle test on a large-model aircraft at KSC. 

Showed that antennas and modems could be 

mounted on such a vehicle and initial link 

could be established. 

Key milestone: 

June 2003: Sounding rocket flight scheduled 

at WFF. This test will help determine the 

limits of this technology. 

Contacts: Dr. J.C. Sin?pson f 

- , ), YA-D7, (321) 867-6937; E.C. Denson, YA07, (321) 867-6537; and A.J. Mirrray, YA-D7-€1, (321) 

867-6673 

72.70 

Lap 

FIM 

0.00 

T.N.T 

I 

11 - Asbestos Phenolic Nose Cone Adapter 

Participating Orgaiiizations: USTDC (R.B. Birr and Dr. 

C.T. Mota), NAL Rescarch Corporation (Dr. N. Honn<q), 

niid WFF (A.M. Colcinaii mid R. L. Wessclls) 

72 70 89 70 106.20 126.20 13745 142.95 

Lap RADAX RADAX RADAX RADAX RADAX 

0 00 

T N.T FIM FA4 FIM FIM FIM 

11 - Asbestos Phenolic Nose Cone 

89.70 106.20 126.20 137.45 142.95 

RADAX RADAX RADAX RADAX RADAX 

F/M FIM FIM FIM j lridium Telemetry 1 NMACS 1 AD9 

12.052 After Motor Separation 

Placciiicnt of Iridium Hariizarc 017 a npical Soiiiiding Rocket 

99 
Autonomous Flight Safety System (AFSS) - Phase II 

The Autonomous Flight Safety System (AFSS) is a real-time onboard hardware and software system for tracking and possible flight termination. AFSS 

is designed primarily for small expendable vehicles at remote launch sites 

where providing traditional ground-based range safety infrastructure, including radio frequency (RF) communication and command links, radar stations, 

data processing, display facilities, and trained operators, would be extremely 

expensive. Advantages of using AFSS include global coverage and decreased 

costs during remote launch site operations. The system may also be useful as 

a training aid and as a backup system for other types of vehicles. It may also 

serve as a flight crew advisory system for manned vehicles. Global Positioning System (GPS) receivers provide tracking data. An onboard processor compares the current position, velocity, and attitude with the nominal trajectory 

and predetermined flight safety rules to autonomously terminate flights when 

appropriate. A feasibility study (Phase I) was completed in 2000, and a proofof-concept hardware system (Phase 11) was successfully tested in 2002. 

This report covers AFSS Phase 11 testing and test results (see Research 

and Technology 2002 Annual Report for design information). A small team 

of NASA and contractor employees as identified below, along with major 

responsibilities, performed the Phase I1 work: 

NASA - KSC: flight hardware, real-time operating system (RTOS), and 

NASA - Wallops Flight Facility (WFF): trajectory simulation. 

Lockheed Martin, prime contractor, supported by subcontractor KT 

Engineering: algorithm, decision logic, and display software. 

integration. 

The primary objective of the laboratory demonstration test was to evaluate 

the AFSS decision logic software installed in a fully integrated AFSS system, 

using WFF-generated flight trajectory simulations. Testing was conducted 

in the WFF GPS Simulation Laboratory in accordance with the AFSS Test 

Scenario Procedures document (developed by the WFF Guidance, Navigation, Control, and Systems Engineering Branch). Two baseline missions were 

utilized; the first was based on Kodiak Star, and the second was a customdesigned trajectory referred to as Wallops Express. Various failure scenarios 

created by WFF personnel based on these two missions were run and graded. 

The AFSS Phase 11 flight configuration used for demonstration testing 

included Ashtech G12 GPS receivers in conjunction with a VME Radstone 

PPC4A-750 single-board computer (SBC) running VxWorks RTOS. 

Although there were several failures of the AFSS system (both decision 

logic and SBC RTOS input / output [I / 01-based), the primary test objective 

of verifying this type of system’s feasibility was met during the testing of 28 

unique scenarios. 

Many of the decision logic problems can be traced back to an attempt to 

emulate the human decision process exactly. Among the unique problems 
encountered with the decision logic are “first motion of the flight 

vehicle” and divergence of redundant sensors. In addition, more 

robust 1/0 routines are needed on the SBC RTOS to ensure reliability of the AFSS. 

Currently the potential of a Phase I11 to produce an integrated, flight-ready, and flight-tested system with improved 

nonmap-based algorithms is being studied, and use of onboard 

inertial measurement units is being examined. 

Contacts: Dr. J.C. Sirnpsoiz (li7incs.C.Siin??soi~~ iim?.p3), YA-D7, (321) 

867-6937; B.A. Ferrell, YA-D7, (321) 867-6678; aiid R.A. Nelson, YA-D7, 

(321) 867-3332 

Participating Organizations: USTDC (J.A. Marin, R.D. Zoeriier, and T. 

Erdogaiz); WFF (1. B. Bull, J.C. Hickman, R.J. Laiizi, M.S. Patterson, W.R. 

Pozpell, S.E. Kreiner, aiid G.F. Robinson); aid Locklieed Martin (S. Halei/) 

1 GPS 1- Flight Computer Receiver 1 RS-232 

Decision Logic 

r---------_---------- 

I Flight System I 

GPS-1 Flight I 

I 

I 

GPS Simulator B GPS-2 RS-232 I 

I 

i 

Ethernet Ethernet 

support 

Computer 

Simulation 

suppori 

Computer 

Data Capture 

Lahorn tory Denioiis t ra f ion Test Coi$g~t ra t ion 

101 
Range Simulator 

The KSC Advanced Range Technologies Laboratory has successfully integrated several different 

hardware and software packages into a versatile support environment that is used daily for a variety of 

projects. The core packages include a CAST Global 

Positioning System (GPS) simulator, high-dynamic 

GPS receivers, Satellite Tool Kit (STK), and Veridian telemetry decommutation boards and telemetry 

software. There are also real-time operating system 

development environments such as VxWorks and 

Linux for embedded processors; digital signal processors; and MATLAB, LabVIEW, and many custom programs for technical computing and visualization. 

The GPS simulator produces a subset of the signals 

created by the GPS satellite constellation at specific 

places and times determined by either internal test 

scenarios or external trajectory files. These signals are 

then connected to a GPS receiver whose output can 

be used to test the GPS coverage and the receiver’s 

performance. External files for antenna patterns can 

also be included. The simulator has been used for 

pre- and postflight analysis and acceptance testing. 

STK is the leading commercial off-the-shelf software solution for the aerospace industry. Some of 

STKs basic applications include calculating and 

visualizing a vehicle’s position and attitude in two or 

three dimensions using either internal test scenarios 

or external data files, determining 

acquisition times from space- or 

ground-based sites, analyzing the 

vehicle’s field of view, and determining communications link margins. 

A large database of satellite orbits is 

also available. 

Veridian hardware and software 

are widely used in the aerospace 

industry for real-time and posttest data processing and telemetry 

signal decommutation and reproduction. Data from a wide variety of 

archived and real-time sources can be 

displayed, manipulated, and easily 

reformatted for further analysis. 

The support provided for the 

Space-Based Telemetry and Range 

Safety (STARS) project is an excellent example of the capabilities of 

the Range Simulator. STARS is a 

multifaceted and multicenter NASA 

project to determine the feasibility of 

using the NASA Tracking and Data 

Relay Satellite System (TDRSS) and 

GPS to provide reliable communication, telemetry, and tracking for 

launch vehicles. The GPS simulator was used to flow data through 

the custom-built Command and 

Data Handler (C&DH) flight processor with a real-time operating 

system during acceptance testing 

and was used to build a scenario to 

test predicted link margins. During 

the first series of flight tests with 

an F-15B at Dryden Flight Research 

Center (DFRC) in May 2003, KSC 

received the real-time Range Safety 

data over the NASA data network. 

The Veridian hardware decommutated and reformatted this data for 

display on a large video screen in a 

conference room at KSC. LabVIEW 

was used to display various health 

and status information, and STK/ 

MATLAB showed the aircraft’s posiLabVlEW Telemetry Interface Showing STARS Internal 

Processor Health and Status 
STK Displny of the F15 During 

tlie First STARS Test Flight Using 

Actrinl Flight Data 

tion and heading on a 3-D animated map. Data 

was recorded and archived at DFRC for postflight 

analysis using Veridian’s database, playback, and 

alization tools; and many of the customized analysis programs. 

doli1-Laiting ~iipddities; STK’S afi.!ysi~ ad U~SLIKey accomplishments: 

STARS Support. 

- VxWorks real-time operating system on 

embedded computer. 

I - Trajectory and link margin analysis. 

I - Acceptance testing of the C&DH processor. 

- Real-time telemetry decommutation. 

- Data extraction/analysis for flight tests using 

- Real-time data displays and 3-D visualization 

during flight tests. 

1 archived files. 

I 

Columbia Accident Investigation 

Support. 

- Trajectory and vehicle attitude 

analysis. 

Key milestones: 

2001: Conceptual design of the simulator. 

2002: Integrated STK, MATLAB, and 

GPS simulator. 

2003: Added k’eridian hardware / 

software and LabVIEW capabilities. 

An upgrade is planned to the GPS 

simulator to include simulated inertial 

measurements capabilities. 

Autonomous Flight Safety System Support. 

- VxWorks real-time operating system on 

embedded computer. 

- Interface development and testing using GPS 

receivers and simulated GPS data. 

Pnrticiptiizg Orgaizizntioiz: USTDC fT. Erdogni~, 

R.B. Birr, J.A. Mnriii, R. D. Zocnicr, Dr. C.T. Mntn, 

aid Dr. C.D. I171117er) 

Iridium Flight Modem Support. 

- Signal analysis. 

- Linux and VxWorks real-time operating 

system on embedded computers. 

- Trajectory and link margin analysis including 

use of actual antenna patterns. 

- Acceptance testing of miniaturized 

integrated modems / GPS receivers. 

103 
Range Information System Management 

Current Range Technology systems rely heavily on expensive networks of 

unique and costly ground-based communication assets to interconnect tracking, communications, and flight termination functions. Much of this aging 

infrastructure is interconnected through legacy-installed, nonmovable cables. 

A future vision of a more flexible Spaceport and Range is emerging, based on 

high-speed first mile / last mile communication technology arising from recently 

developed commercial technologies. These new technologies, seamlessly 

extending existing legacy Range infrastructure without wires and cables, hold 

great promise for preserving the utility of existing Range infrastructures while 

providing life cycle cost savings and increased operational flexibility. 

The common theme needed to accomplish this future vision is a coherent 

Range Information System Management (RISM) plan. In FY 2002, the RISM project began work on this plan by surveying the Range and identifying the existing key communication functions. In addition, the RISM project investigated 

the leading candidate commercial technologies in which early involvement 

would likely provide the greatest returns on investment toward accomplishing 

future Range-related communication tasks. In FY 2003, the follow-on project, 

Emerging Communication Technology (ECT), continues this effort through a 

detailed investigation of the leading candidate technologies identified by RISM. 

Together, RISM and ECT provide a course of action for achieving a coherent 

vision for meeting future Range communication technology needs. 

During FY 2002, the RISM project identified three key emerging technology 

candidates to be further investigated on ECT. These three technologies were 

Ultra-Wideband (UWB), Wireless Ethernet (Wi-Fi), and Free-Space Optical (FSO). 

The RISM project also released a report identifying the major communication 

trends and techniques being developed today in the private sector, as well as 

their likely contributions to meeting future Spaceport and Range needs. 

The following are key aspects of the Range communication vision: 

Mobile, portable, fixed Range / Spaceport requirements. 

High bandwidth, high data-rate capability over three decades of data rates (10 

Seamless connections to today’s wired communication infrastructure and 

Selectable features versus communication capability to best match future 

megabits per second [Mb/s] to 10+ gigabits per second [Gb/s]). 

future systems. 

needs and flexible provisioning for quick turnaround. 

The ECT project will procure UWB, Wi-Fi, and FSO exemplars and will collect 

measured results on various performance parameters using a specially instrumented advanced networking laboratory at KSC to determine the limits of these 

technologies. With this knowledge, the shortcomings of the commercial products will be identified, and it will be possible to influence the developments of 

the next generation of commercial products to meet specialized Range needs. 
Coiizparisoiz of Leading Candidate Cominunicatioiz Tecliiiologies 

Key Range 

Communication Attributes 

Candid ate Technolow 

Maximum Range 

Typical Range 

Minimum Range 

Selectable Data Security 

Portable, Mobile, Fixed 

Regulatory Impacts 

Medium Data Rates 

(10Mb/s- lOO+Mb/S) ~ 

Wireless Ethernet (Wi-Fi) 

Yes 

Yes 

None 

High Data Rates Highest Data Rates 

(100 Mbis - 1,00O+Mb/s) (1,000 Mbis- 1O,OOO+Mb/s) 

Ultra-Wideband (UWB) Free-Space Optical (FSO) 

> 10 km > 1.5 km 

>5km >I km 

<2m < 100 m 

Yes Yes 

< 50 m 

Maximum Range at 

Present for UWB 

The ultimate outcome will be to establish a set of criteria for best fitting the appropriate 

technology to the Range communications needs while achieving connectivity over at least 

three decades of data rates. Once the performance limits of these technologies are determined on ECT, it will be possible to enhance future Range capabilities while maintaining or 

even increasing Range safety and lowering life cycle costs. 

Key accomplishments: 

Attended Advanced Range Technology Working Group (ARTWG) and Advanced 

Spaceport Technology Working Group (ASTWG) conferences at KSC and Colorado 

Springs, Colorado. 

Established the Future Integrated Range and Spaceport Technology Working Group 

(FIRSTWG) comprising more than 80 senior aerospace leaders and managers from 

Government, industry, and academia. 

biweekly through FY 2003.) 

Arranged technical seminars for FIRSTWG presentations in FY 2002 (continuing 

Arranged site briefings and familiarization visits for NASA and contractor personnel to 

Supported ARTWG and ASTWG development efforts through participation in the 

I 

, the Eastern Range facilities. 

t following subgroup teams: Vision, Communication, Telemetry, Command and Control, 

and Tracking and Surveillance. 1 

Key milestone: 

September 2002: Released RISM Phase I final report identifying Range communication 

infrastructure sites and links, existing commercial communications technologies, and 

three key technologies for further investigation. 

Coiztact: R.A. Nelsoi? (~ic~i~i.~,A.~~,l~@ii~ iraw.go~~), YA-D7, (321) 867-3332 

Participating Organizations: USTDC (Dr. G.L. Bastiiz, W.C. Harris, C. Kerios, T. Erdogan, and E.H. 

Williams). More than 80 peoplefroin at least 14 orgaizizatioizs are also participating iiz this project in 

FIRSTWG telepltorie conferences. Major scieiitific coiitributors include Wallops Flight Faciliti, 45th Space 

Wing, KSC, FAA, NOTU, CCAFS, VAFB, SMCICWP, ENSCO, DOC, DOD, Boeiizg, CSR, LocklieedMartiiz, Harris Corporation, Uiiiuersity of Ceiitral Florida, Florida liistitute of Techizology, and Uiiiwrsity of 

Florida. 

105 
106 

Accurate Location of Lightning Strikes 

Lightning strikes generate large electric and magnetic fields. These fields, 

when introduced into electrical conductors, can result in large voltage or current spikes. These large induced voltages can result in immediate damage to 

sensitive components, especially those 

containing semiconductor devices. 

Following a lightning strike at the 

Space Shuttle launch pads, equipment 

and instrumentation are tested according to the test plan in KSC SWE-0018, 

Volume 2. Complete testing involves 

turning all the equipment on and is 

conducted when either of the following 

conditions is met: 

The voltages or currents recorded 

by the Lightning-Induced Voltage 

Instrumentation System (LIVIS) at 

the launch pads exceed the limits 

established by the operation and 

maintenance requirements. 

The lightning channel traveled past 

the mast and the catenary lightning 

protection wire and struck either the 

pad structure or the vehicle. 

There have been instances when it 

was impossible to see the attachment 

SOLLO 

point of the lightning strike due to heavy rain. 

Without a clear attachment point, it is hard to 

know whether to conduct a complete test or waive 

the requirement. The accurate lightning location 

system can assist in determining the exact attachment point and can help assess whether additional 

testing is required. 

The fast-varying electric current associated with 

lightning discharges generates large electric field 

variations, which propagate at the speed of light in 

a radial direction away from the striking point. At 

the same time, thunder is generated as a result of 

the sudden heating of the air caused by the large 

currents associated with the lightning discharge. 

For an observer located away from the striking 

point, the electric field waveform arrives earlier 

because it travels at a speed of 300 million meters 

per second (m / s), while the sonic wave travels at 

about 350 m / s depending on air temperature and 

humidity. 

The Sonic Lightning Locator (SOLLO) consists 

of an electric field sensor and four sonic (thunder) 

recorders. Three of these sonic sensors are located 

in the perimeter of a 4-m-radius circle, 120 degrees 

apart from each other, with the electric field sensor 

placed in the center of the circle. The fourth sonic 

sensor is located 2 or 4 m above the plane formed 

by the other three sensors. Based on the differences 

in the time of arrival of the sonic wave at each 

sensor, a set of equations is solved to obtain the 

direction to the lightning strike. The differences in 

the times of arrival can be measured to better than 

10 microseconds by performing real-time digital 

cross-correlations among the signals received by 

the four sonic sensors. The result consists of direction (azimuth and elevation) and distance to the 

lightning strike. The location accuracy is better 

than 5 meters for a strike within a 0.5-kilometer 

radius. Digital signal processing techniques are 

used to discriminate between thunder waveforms 

from a close lightning strike (within approximately 

a 1-mile radius) and thunder waveforms from a 

preceding but more distant strike. This process 

is important because during the active phase of a 

thunderstorm lightning activity could occur at a 

rate exceeding 10 strikes per minute. 
Key accomplishment: 

The Sonic Lightning Locator was successfully tested during the 

thunderstorm season of 2002, recording 7 triggered lightning strikes 

and 29 natural lightning flashes. The sensitivity of the electric field 

sensor was adjusted for it to trigger on lightning strikes occurring 

within 1.5 miles of the recording station. The sensitivity calibration 

is based on the typical strength of the electric field generated by a 

lightning strike in Florida. Different calibration parameters can be 

applied when using the SOLLO at other geographical locations. 

Key milestone: 

To measure the magnitude of the electric field and to provide an 

estimate of the lightning current, circuitry is being added to enable 

recording of the following main parameters of the electric field 

waveform: 

- Peak magnitude. 

- 10- to 90-percent rise time. 

- Mean duration. 

- Number of strokes. 

Con fact: J. T. Mad u ra (/ti ;I 1. T.,Z I r7d I i i.17 ((7 I IGW .:q:*), YA- D, (32 1 1 86 7-0814 

Participating Organization: USTDC (Dr. P.J. Medelius) 

Thunder Sensors 

Timing, 

Digital Cross-Correlation, 

Digital Filtering, 

Location of Strike 

Digital Signal of Location 

Processing of Lightning Strike 

Threshold 

Detector 

I 

SOLLO Block Dingmm 

107 
Transient Voltage Recorder 

The Transient Voltage Recorder was designed for monitoring the effects of lightninginduced voltages on communication and power lines associated with Expendable Launch 

Vehicles at Cape Canaveral Air Force Station. Although electronic equipment and instrumentation usually have some form of transient protection, limitations exist on the maximum 

voltage and maximum duration of the transient voltages to which they can be exposed before 

permanent damage occurs. Electronic components and equipment can fail immediately when 

subjected to voltages larger than their maximum-rated values. Transient voltages can also 

alter the characteristics of an electronic component without any immediate sign of damage. 

Unless there is a way to determine transient voltages are present, such anomalies can remain 

undetected and eventually lead to a premature and unexpected failure of an electronic component or an instrument. 

Transient voltage recorders have been used for many years to monitor voltage spikes in 

cables. Common sources of transient voltages include large electric motors, switching of 

power equipment, induced voltages from nearby lightning, and direct lightning strikes. The 

transients generated by sources other than lightning tend to have frequency components well 

below 1 megahertz (MHz), which most commercial transient recorders can detect and record. 

On the other hand, lightning-induced transients contain voltage peaks whose amplitude can 

rise from 10 to 90 percent in a fraction of a microsecond and can generate multiple transients 

within an few milliseconds of each other. It is therefore important to be able to record and 

characterize the impulsive nature of transients caused by lightning strikes. 

Commercially available transient voltage recorders can sample multiple signals at fast rates 

(e.g., 20 megasamples per second [MS/s]) and measure voltage transients up to several hundred volts. However, the main disadvantage of many recorders is that data must be retrieved 

through a serial communication link, and depending on the amount of data being recorded, 

the onboard storage can be filled before data retrieval is completed. Large amounts of data are 

transmitted, although the users usually require only a relatively small part of the information. 

The technology implemented for the new Transient Voltage Recorder overcomes the limitations of existing transient recorders and provides essential information regarding the transient 

voltages without the overhead associated 

with the recording, storage, and transmission of a complete transient waveform. 

The technology consists of a transient voltage recorder that characterizes the transient 

waveform in real time. A fast-sampling 

analog-to-digital converter, operating at 

speeds up to 100 MS/s is used to sample 

the transient waveforms. Compared to 

traditional recorders that sample and store 

complete waveforms, the new recorder 

analyzes the data as it is being sampled 

and stores only the parameters of interest. 

A real-time comparator and peak detection circuit was implemented with the use 

of a fast Field-Programmable Gate Array 

(FPGA). Compared to the traditional Transient Voltage Recorder 
approach that requires thousands of bytes to store 

each waveform, the present technology requires 

less than 20 bytes of storage per waveform. In 

addition, the typical dead time between subsequent triggers is now reduced from hundreds of 

microseconds to about one-tenth of a microsecond. 

This recorder calculates and stores the following 

parameters: 

Peak voltage value. 

Duration of transient at 25,50, and 75 percent of 

10- to 90-percent rise time. 

90- to 10-percent decay time. 

peak (in 10-nanosecond increments). 

The Transient Voltage Recorder can be configured for different input ranges, depending on the 

expected magnitude of the transient voltage being 

monitored. Typical input ranges include *lo, 50, 

and 100 volts. The termination can be either singleended or differential and selectable among %, ?2G, 

or 10 kilohms. A bipolar scheme is used so either 

a positive or negative transient can trigger the 

system and initiate the real-time waveform characterization. Depending on the implementation, 

multiple data channels could be analyzed simultaneously, with a trigger signal from any channel 

determining the start of the waveform characterization process. 

A real-time clock is included to accurately 

time-stamp any recorded waveform. This is 

important if it is necessary to correlate a transient with other measurements, such as those 

provided by a lightning location system. The 

power supply includes backup batteries that 

can maintain the transient voltage recorder 

operational for up to 15 days in case of main 

power loss. The batteries are automatically 

charged during normal operation. 

Key milestone: 

The Transient Voltage Recorder was 

used to monitor transient voltages in 

support of the GALEX mission. Current 

measurements were conducted at four 

different power, communication, and 

ground iines. 

Participating Orgaizizafion: USTDC (Dr. P.J. 

Medelius) 

Real-Time 

AmDlifier and 

TimeStamp 

Antialiasing 

Filters 

Channel 1 

I I 

4-b Ethernet - Micro- - controller RS-232 

Nonvolatile 

Detector Trigger Memory 

: 1- 

I 

-15 V 

Gain 110Vac 

Selection 4 

Traiisienf Voltage Recorder Block Diagram 

109 
Passive Millimeter Wave Imaging 

Current Range Safety launch requirements for missions carrying nuclear 

payload materials call for continuous observation by Range Safety officers 

from T-0 through tower clear. Historically, optical and laser scanning imaging systems have provided the bulk of the imaging equipment needed to meet 

this Range Safety observation requirement. However, with the advent of 

larger launch vehicles with ever increasing numbers of strap-on solid rocket 

motors (e.g., graphite epoxy motors), more capable water deluge systems 

are needed on launch pads to reduce potentially damaging acoustic energies 

during launch. The unintended consequence of building more capable pad 

deluge systems is an increase of obscuring clouds of plume exhaust and steam 

during launches. An updated method of viewing launch vehicles through 

these larger clouds during launch is needed. 

The fundamental issue, relative to legacy optical and laser scanning systems, is the attenuation introduced by the clouds of plume exhaust at the 

wavelengths of both the visible imaging systems and the scanning laser 

beams. To increase the efficiency of solid rocket motors, exhaust particle size 

must be kept very small. Unfortunately, current particle sizes are now of such 

a small diameter that high attenuation occurs at both visible light wavelengths 

and scanning laser beam wavelengths. In short, the old viewing methods no 

longer can meet the imaging needs for maintaining Range Safety for larger 

launch vehicles. 

One solution investigated this year on the Millimeter Wave Imaging project 

was passive millimeter wave imaging systems. These work because, at all 

temperatures above absolute zero, physical objects naturally emit millimeter 

wave energies passively. No millimeter wave energy illumination is required, 

and unlike radar beams, no energy must be radiated at the launch vehicle. 

Passive millimeter wave imaging systems therefore hold great promise for 

providing enhanced observation capabilities for the current generation of 

heavy launch vehicles while remaining inherently safe for use with launch 

ordnance devices and pyrotechnics. 

The Millimeter Wave Imaging project extended the theoretical understanding of applying passive millimeter wave imaging technology to launch situations. The tasks accomplished on the project included modeling of expected 

attenuations versus wavelength at several millimeter wavelengths, resulting 

in the identification of optimal wavelengths that can best enhance Range 

Safety. This theoretical effort included developing models for estimating 

steam production, particle sizes, and total attenuation through plume exhaust 

launch clouds. The results of this theoretical investigation were released in a 

final report identifying the major limits of this technology. 

Follow-on tasks anticipated for the coming year are to coordinate a vendor 

demonstration of a military millimeter wave imaging system to implement 

an initial proof of concept, followed by additional theoretical investigations to 

further determine the limits of applying this technology for enhancing Range 

and Spaceport Safety. A wide range of applications exists for passive millime- 
ter wave imaging technology, including the passive 

inspection of external Orbiter tanks during fueling 

on the launch pad and the passive inspection of 

cargo and container trucks at guard station checkpoints. 

Key accomplishments: 

Arranged site briefings and familiarization visits 

to Innovative Science and Technology Experimentation Facility (ISTEF), Launch Complexes 

17A and 17B. 

Researched key technical journals and papers. 

Developed theoretical models for estimating the 

performance limits of passive millimeter wave 

imaging in a launch environment. 

Key milestones: 

October 2002: Commenced research. 

December 2002: Released final report to the 

Evolved Expendable Launch Vehicle (EELV) 

program identifying the major technical limits 

to applying passive millimeter wave imaging 

technology to launch environments. 

Contact: Dr. R.C. Youngquist (,?, c ‘ $1 , I . 11.1’ 1;’’ ), YA-C3-E, (321) 867-1829 

Participating Organizations: LISTDC (Dr. G.L. Bastin and T. 

Erdogan) and VA-C (A. Sierra) 

Vehicle Obscurafion During Launch 

11 1/112 
Spaceport Structures and Materials 

Advanced structures and materials are critical to acheving the 

goals of reduced costs, increased reliability, and higher flight rates 

for future spaceports. Spaceport structures and materials technology 

development areas include corrosion abatement, static charge dissipation, nondestructive evaluation (NDE), reduction of vibroacoustic loads during launch, and nonflammability. Sources of corrosion 

and material degradation include humid saltwater environments 

surrounding launch structures and aggressive ozone and hard radiation environments of Earth orbit. Static charge buildup on payloads, 

spacecraft, and launch structures can present significant safety 

issues for personnel ad equipment. NEE tecb,-,dogies arc criticd 

to obtaining rapid status analysis of flight hardware. Passive sound 

mitigation techruques to replace active noise reduction methods such 

as water suppression systems are desirable for future launch vehicle 

pads. Advanced nonflammable materials are desirable to enhance 

the safety of crew and ground personnel during operations. Since 

materials are ubiquitous in spaceflight, the reach of this technology 

thrust area will encompass launch structures, payload processing, 

spacecraft design, and vehicle maintenance. 

Technology focus areas include the following: 

Launch Structures and Mechanisms 

Corrosion Science and Technology 

Electromagnetic Physics 

Materials Science and Technology 

Nondestructive Evaluation 

The goals and objectives of Spaceport Structures and Materials 

include the following: 

Ensure safe, efficient, and reliable structures techniques 

Enhance reliability and reduce maintenance cost of infrastructure 

Improve safety and reliability of operations through detection, 

mitigation, and prevention of electrostatic generation on equipment 

Develop specialty materials in support of future structures and 

materials initiatives 

For inore iizforrnatioiz regarding Spaceport Structures and Materials, please confact 

Karen Thompson, (321) 867-7555, Kniwz. L. Tiiompsoiz@ ii~~o.p~; or Dr. Carlos 

Calle, (321) 867-3274, Carlnd.Cnllr@ iim?.goi’. 

113/114 
Spaceport Structures and Materials Roadmap 

115 
116 

0l;tics 

Electronic Single- Particle Aerodynamic Relaxation Time Analyzer 

for Mars Robotic Missions 

With funding from NASA's Office of Space Science, under its Planetary 

Instrument Definition Program, KSC's Electrostatics and Surface Physics 

Laboratory and the University of Arkansas at Little Rock are currently designing and building an Electronic Single-Particle Aerodynamic Relaxation Time 

(E-SPART) analyzer that will perform real-time, simultaneous measurements 

of the aerodynamic diameter and the electrostatic charge distributions of dust 

particles in the Martian atmosphere (figure 1). To date, there have been no 

direct measurements on the surface of any planet (other than Earth) characterizing the concentration, distribution, size, or shape of suspended dust particles or the charge deposited on them. 

1 

15 cm 

I Insulator II 

II 

I1 

I' 

Integrated : I I, 

Figure 1. Schematic of the Electronic Single-Particle 

Aerodynamic Relaxation Time (E-SPART) Analyzer 

The E-SPART analyzer consists of 

a dual-beam, frequency-biased laser 

Doppler velocimeter (LDV); a relaxation 

cell; and an electronic signal and data 

processing system. In the E-SPART 

analyzer, the LDV measures, in a noninvasive way, the particle velocity in its 

sensing volume. The sensing volume 

is located at the center of a relaxation 

chamber. The Mars dust particles flow 

vertically downward through the sensing volume. As the particles move 

through the sensing volume, in the 

direction normal to the plane containing 

the two converging laser beams, they 

experience an alternating current (ac) 

electric excitation that causes the particles to oscillate in a horizontal direction, 

parallel to the direction of LDV velocity 

measurements. The ac electric excitation 

is generated using a pair of electrodes located inside the relaxation cell and 

positioned symmetrically across the LDV sensing volume. 

The electronic signal and data processor analyzes the phase lag of the particle motion with respect to the ac electric field driving the particle. The gas 

dynamic diameter is derived from the phase lag value. The measurements of 

diameter and the direction and amplitude of the electrical migration velocity 

of the particles with respect to the electric field provide the polarity and magnitude of electrical charge of a particle. 

Results of an example of particle sampling of a Martian regolith simulant 

charged against stainless-steel beads for 10 minutes at room temperature with 

44-percent relative humidity are shown in figures 2 and 3. The simulant was 

ground down using a ball mill and measured with the E-SPART. The average 

particle diameter was 6.63 millimeters (mm) with a standard deviation of 1.91 

mm. The total particle count was 5,039 with a total mass of 3,901.3 nanograms 
I 

and total charge of -311.02 femtocoulombs 

(fC). This gives a charge-to-mass ratio of 

-0.13 millicoulomb per gram (mC/g). The 

results are given as a three-dimensional 

view showing the number of particles 

versus the diameter and the percentage 

of maximum charge (figure 2) and as the 

charge-to-mass ratio (Q/ M) of positively 

and negatively charged particles as a 

function of particle size (figure 3). 

While the net charge-to-mass ratio 

(positive plus negative) for the Martian 

simulant dust in the size range 1 to 30 

mm is -0.13 mC/g against stainless steel, 

the individual particles were found to be 

highly charged. For example, 1.06-mmdiameter particles showed Q/M of -27.83 

mC / g for negatively charged particles 

and +20.16 mC / g for positively charged 

particles. 

This technology will allow us to find 

the charge-to-mass ratio and concentration of particles suspended in the Martian 

atmosphere. The technology will also 

provide KSC with a future flight instrument for inclusion in the Mars 2009 mission. 

Coil fact: Dr. C.1. Calle (Cni-lo5 I.Cdlc@ iinsa.p~), 

YA-C2-T, (321) 867-3274 

Participating Organizations: Uiiiuersity of 

Arkansas at Little Rock (Dr. M.K. Mazumderi arid 

Szilales Aerospace (Dr. C.R. Bidder) 

144 

z f33 

22 

11 

0 

-20 -15 -10 -5 0 5 10 15 20 

PercantMaxchsrge 

Figure 2. E-SPART Results for n Martian Soil Simulant 

Electrostnfically Charged Agaii-1st Staiizless-Steel Beadsfor 2 0 

Humidity 

3.83 

-3.68 

2-4m 

-I" 

Paltick? Diameter da (pn) 

Figirre 3. Charge-to-Mass Ratio Plotted as II Firiictioii of the 

Equiilalent Aerodyiiamic Diameter of Dust Particles Charged 

Against Stainless-Steel Beads for 10 Minutes 

117 
m 

U .- 

c 

ti 

0 

Paschen Breakdown in the Martian Atmosphere 

The high probability of dust interactions during Martian 

dust storms and dust devils, combined with the cold, dry 

climate of Mars, will most likely result in airborne dust that is 

highly charged. On Earth, potential gradients up to 5 kilovolts 

per meter have been recorded and in some cases resulted in 

lightning. Although the Martian atmosphere is not conducive 

to lightning generation, it is widely believed that an electrical 

discharge, in the form of a corona, occurs. 

To understand the breakdown of gases, Paschen 

measurements are taken, which relate the minimum potential 

required to spark across a gap between two electrodes. The 

minimum potential is plotted versus the pressure-distance value 

for electrodes of a given geometry. For most gases, the potential 

decreases as the pressure decreases. For carbon dioxide (CO,), 

the minimum in the curve happens to be at Mars atmospheric 

pressures (5 to 7 conventional millimeters of mercury [mm Hg]) 

for many distances and geometries. However, a very small 

amount (less than 0.1 percent) of mixing gases radically changes 

the curve, as noted by Leach (figure 1). Here, we present 

experimental results of a Paschen curve for a Mars gas mixture 

compared with 100-percent pure CO,. 

Voltages of either positive or negative polarity were applied 

at one custom-made brass electrode, while the other was 

grounded. The cylindrical electrodes have a diameter of 5.059 

centimeters (cm) with curved edges to eliminate edge effects 

that create strong electric fields. Thus the electrodes fixed 0.1 cm 

apart provided the parallel-plate geometry. 

The chamber was pumped down to 104 mm Hg and 

backfilled to approximately 50 mm Hg with the atmospheric 

gas. It was pumped down again to remove residual gases and 

contaminants. The gases that were used are 100-percent CO, 

and a Mars gas mixture provided by Praxair, 

Inc., that consisted of 95.5-percent CO,, 2.7- 

percent nitrogen, 1.6-percent argon, 0.13-percent 

oxygen, and 0.07-percent carbon monoxide. 

Measurements were taken by simply increasing 

the voltage until the power supply shut off 

as a result of the current divergence during 

breakdown. Figure 2 shows the minimum voltage 

required for electrical breakdown of the two gases 

for both positive and negative polarities. 

j 

V 

Clearly, as the pressure is lowered, less voltage 

is required for breakdown to occur for both gases. 

At the higher pressures, the spark potential does 

not seem to depend on the makeup of the gases. 

Pressure - Distance 

Figure 1. Leach Shows That Paschen 

Measurements D#ev Greatly with Minimal 

Change in Atmospheric Gas Content 
At lower pressures, a clear difference in 

breakdown potentials can be seen. Figure 3 shows 

how the small concentrations of nitrogen, oxygen, 

argon, and carbon monoxide greatly alter the 

Paschen curve for pressures ranging between 1 and 

10 mm Hg. We can see that the minimum in the 

CO, curve at around 4 to 5 mm Hg becomes shifted 

to 6 to 7 mm Hg once the other gases are added. 

This suggests that, even though breakdown occurs 

easily at Mars pressures, it may not occur at the 

minimum in the Paschen curve for CO,. Instead, 

the results seem to indicate that there may actually 

be a local maximum near the minimum in the 

CO, curve. Once the pressure is lowered below 1 

mm Hg (not shown), both the CO, and Mars gas 

mixture rise again since there are no molecules to 

ionize in a hard vacuum. 

It is interesting to note that around Mars 

pressures the blue spark seen at higher pressures 

turns into a purple-blue corona. At pressures less 

than 2 mm Hg, there is no longer a well-defined 

spark between the cylinders; instead fhere is 

only the corona breakdown to the gas around the 

charged electrode. 

This research yielded the following major 

conclusions: 

Paschen curves for a Mars gas mixture are 

not too different from pure CO,, although 

differences can be seen. 

Knowledge of the electrical breakdown of the 

Martian atmosphere is crucial for future human 

and robotic exploration of Mars. 

Contact: Dr. C.I. Calle fCnrlod.C~lle@ 77nsn.goi1), 

YA-C2-T, (321) 867-3274 

Participating Organization: Su?ales Aerospace 

(Dr. C.R. Bulzler) 

Figure 3. Minimuin Spark Potential iiz the Range 

1 to 10 inin Hg for Pure CO, aiid tlze Mars Gas 

Mixture Usiizg Brass Electrodes at 0.1 cm Apart 

za 

P 

2 

3 

n 

3 

n 

0 10 20 30 40 50 60 

Pressure - Distance (mm Hg * cm) 

Figure 2. Minimuin Voltage Required for Breakdomn 

of CO, and Mars Gas Mixture as a Fuiictioii of the 

Pressii re- Dis tan ce Relatioiislz ip 

(The error bars are smaller tlraiz the points.) 

I - -+ Mars mixture 

420-- 1 I I 1 

0.2 0.4 0.6 0.8 1.0 

Pressure - Distance (mm Hg cm) 

119 
120 

Wheel Electrometer Technology 

Surface sand and dust electrification is a highly probable phenomenon in the low-pressure, lowhumidity environment of Mars. Surface soil and dust particles on Mars may become electrostatically 

charged due to incident ultraviolet (UV) radiation reaching the surface. Recent experiments have 

shown that this process can electrify the soil and provide a photoelectron sheath that levitates 

charged particles about 1 meter above the surface, a phenomenon that was possibly seen on the 

Moon and that may also exist on Mars. Contact charging may also occur because of collisions 

between wind-blown dust particles and stationary surface particulate matter. It has been suggested 

that the formation of soil agglomerates and sand dunes may be attributed to the electrostatic 

properties of the Martian soil. To date, no experiment on the surface of Mars has directly measured 

the amount of charge on the surface dust and soil particles. 

A system of embedded sensors that can be incorporated into the wheel of any future mission rover 

would provide for a simple and fairly unobtrusive way to measure the distribution of electrostatic 

fields on the Martian surface and to measure variations in soil electrostatic response. The Wheel 

Electrometer Sensor (WES) consists of two types of sensor modules. These sensors will be attached 

just beneath the rover wheel so each sensor will be exposed to the Martian regolith either by line 

of sight through a small amount of Martian atmosphere or by direct contact with the regolith. A 

description of how the WES might be incorporated into one of the wheels of the Field-Integrated 

Design and Operations (FIDO) helps to illustrate the technology. FIDO is an advanced vehicle used 

in technology definition and field tests for future NASA Mars programs. 

Figure 1 shows the FIDO wheel with both the electric field sensor (ELF) and triboelectric (TRIBO) 

sensor modules. The five TRIBO sensors use a different insulator material and are backed by 

independent miniature electrometers, while the ELF is a bare sensor recessed from the wheel surface. 

The ELF and TRIBO sensors are actually based on the same technology. Each type of sensor simply 

measures the amount of charge induced on a metal electrode that has been exposed to some external 

distribution of electrostatic charge, and both have sensitivities that are achieved by adjusting circuit 

component values and the sensor area. 

Figure 1. Wheel Electrometer System (WES) Shown on the FIDO 

Wheel (left) and Cross-sectional View of the ELF and TRIBO Module: 

(right). (The ELF is recessed to allow measurement of an undisturbed 

soil plug that has riot beeii compressed or rubbed by the wheel rim.) 

In the case of the ELF sensor, the 

source of the charge would be any 

charged soil particles present on the 

Martian surface at the time the rover 

wheel rolls over it. The ELF sensor 

electrode will be recessed several 

centimeters radially inward from the 

outer surface of the wheel through a 

hole in the wheel. This will ensure the 

ELF directly measures any naturally 

occurring charge that may be present 

on a small patch of undisturbed 

Martian regolith as the wheel rolls 

forward. The ELF will provide 

an output voltage that is directly 

proportional to the amount of charged 

regolith that the sensor ”sees” through 

the hole. The regolith’s surface charge 

density will be determined using the 

charge measurement and the known 

hole area. As the rover travels across 
the Martian surface, the local surface charge density 

will be mapped using the ELF measurements. 

These data will provide direct measurements of 

the presence of electrically charged particles on the 

Martian surface. 

The TIUBO sensor module will have five 

independent sensors. The electronic circuitry for 

each sensor is identical, but a different insulator 

material will cover the electrometer sensor electrode 

of each sensor. As the rover wheel rolls over the 

Martian regolith, each of the five different insulators 

will make contact with the surface. The electrostatic 

response to contact charging of each insulator with 

the regolith will provide data regarding how the 

regolith fits into the triboelectric series. 

A prototype of the WES with four TRIBO 

sensors was built in the KSC Electrostatics and 

Surface Physics Laboratory to test the concept in 

a simulated Martian environment using Johnson 

Space Center USC) Mars-1 simulant sd (figui-t. 

2). The prototype wheel is 12.7 centimeters (cm) in 

diameter and 10.3 cm long. The four TRIBO sensors 

have a diameter of 1.84 cm with a concentric guard 

and a shield. The sensors are capped with Teflon, 

Lucite, fiberglass-epoxy G-10, and Lexan disks 2.0 

cm in diameter and 0.7 cm thick. Figure 3 shows 

preliminary data obtained with the prototype WES 

in dry air at 9-percent relative humidity and at 

atmospheric pressure. The prototype wheel was 

rolled along a 60-cm tray containing JSC Mars-1 

simulant. The four insulators acquire different 

electrostatic charges when in contact with this 

taken prior to the one generating the data 

presented here. The insulators and simulant 

were exposed to an ionizer to neutralize their 

surface charges before this run but were not 

cleaned. Thus, this procedure is fairly close 

to an actual procedure that could be used on 

a flight instrument. 

This research yielded the following major 

conclusions: 

A wheel electrometer could easily detect 

surface changes covered with fine dust 

that would be undetected visually. 

electrostatic fields while traversing 

surfaces. 

properties of the Martian regolith could 

help us understand several unsolved 

problems rdatcc! to sand transport and 

the low erosion rate on Mars. 

This wheel electrometer will map 

Determination of the electrostatic 

Coli fact: Dr. C.1. Calk (Cnrlos.l.CallrirLiinsa.~~oi~), 

YA-C2-T, (321) 867-3274 

Participfing Orgaiitafioizs: USTDC (Dr. C.R. 

Buliler), Florida lizstitufc of Techiiology, (Dr. J.G. 

Mnntoaniii), YA-C2-T fE.E. Groop), Jet Propulsiniz 

Lahoratoni (Dr. M.G. Buehler), aizd USTDC (A.W. 

Noiiiicki) 

simulant. The sharp peaks observed in the graph 

are due to the initial contact with the soil. Repeated 20 Id' 

contacts show an increase in the charge exchanged 

between simulant and insulator. Several runs were 15 

-10 

0 

1 - kite 

~ - Fiberglass 

20 40 sb ai7 100 

?me (seconds) 

Figure 3. Charge Generated on the Four Polymers 

Figure 2. Prototype Wlzeel Electrometer System Cappiiig the Electrometer Sensors oii the Prototype WES 
Development of a Physical Model of Insulator- Insulator Triboelectric 

Charging To Understand Surface Charging of Materials on Earth and 

on Planetary Surfaces 

The electrostatic charging of materials by contact or rubbing is called triboelectric charging. 

It is a serious problem for materials on Earth and on dusty planetary surfaces, such as that 

of Mars. The least understood type of triboelectric charging occurs when insulators contact 

each other. A hypothesis being investigated to explain insulator-insulator charging is that 

an exchange of surface ions takes place upon material contact or rubbing. A physical model 

being developed is that ions are weakly attached to an insulator’s surface either in a thin 

(about 100-angstrom) water layer and/ or directly attached to the surface by intermolecular 

forces such as the van der Waals force. When the surfaces are in contact, there is ion exchange 

resulting (upon separation) in one surface being negatively charged and the other positively 

charged. 

The surface charge can be viewed as n ions occupying n sites of N total sites on the surface. 

Using a thermodynamic relation for n and Gauss’s law, an equation for the surface charge 

density can be found: 

where P, is the vapor pressure, A is the surface area, q, is the electron charge, rn is the ion 

mass, T is temperature, k, is Boltzmann’s constant, A is Planck‘s constant divided by 2n, and 

E,, is the ion surface adsorption energy. Experiments were performed where surface charge 

was measured for insulator materials (polymers) that were charged up triboelectrically and 

then had the atmospheric pressure decreased. The curve fits to experimental data in the figure 

are from (1). The fitting parameters used in (1) provide values for the adsorption energies 

and number of occupiable charge sites per unit area at room temperature (300 kelvin [K]). 

Thus, it may be possible to determine the amount of charge deposited on insulators after 

insulator-insulator contact if these two parameters are known. The atomic species chosen was 

the sodium ion (Na+, m = 3.82 x 

common and easily solvated ionic compounds. 

kilogram) because sodium chloride is one of the most 

The values of 5, developed by the curve fit correspond to the difference of the work 

functions for material A and material B, (@A - @,), measured for metal-insulator contact. 

Conventionally, insulators are assigned “effective work functions” based on their charging 

properties against metals of known work function. Therefore, it is believed that charging of 

two insulators could be determined by comparing their effective work functions. 

Thus, an alternative method for determining the amount of charge transferred between 

insulators is presented. Here the mechanism is that of ion transfer as opposed to electron 

transfer. Experiments indicate that surface properties, such as hydrophobicity, play a key 

role. The model focuses on identifying the source and meaning of the adsorption energies for 

insulators. Determination of insulator work functions by contact with a metal appears to be 

equivalent to our determination of the adsorption energy. Thus, charging an insulator and 
t 

removing the vapor pressure may provide an alternative method for determining (@A - @J, 

the difference of work functions. This interpretation may provide a clearer picture of the 

phenomenon. 

Key milestones: 

ri: 

c) 

2002 / 2003: Development of a pressure-dependent physical model of insulator-insulator 

triboelectric charging. 

2003: Refine physical model by developing realistic values for N/A, and Eo and perform 

additional triboelectric experiments where surface charge is measured as a function of 

pressure for various insulator materials. 

will mitigate triboelectric surface charging. 

2003 and beyond: Use model and experimental data to develop methods or materials that 

Contacts: M. D. Hogue (Michn~l.n.Hoc~zic~?in~~.~@~~), YA-C2-T, (321) 867-7549; and Dr. C.1. Calle, 

YA-C2-T, (321) 867-3274 

Participating Organizations: Swales Aeyospace (Dr. C.R. Bulzler) and University of Central Florida 

(Dr. W. Luo) 

Surface Charge vs. Pressure Stair Step Maximums I- 

> 

1.5- 

d / 100 200 300 400 500 600 

-0.5 1 

Pressure (Torr) 

+Fiberglass -Lexan +Teflon -Rulon J -Luclte I 

123 

I 
Comparison of Surface Resistivity and Triboelectric Charge Generation 

Characteristics of Materials 

Material 

The Electrostatics and Surface Physics Laboratory at KSC has 

recently focused on providing test data that demonstrates the 

lack of correlation between the surface resistivity of materials 

and the materials’ ability to generate charge due to contact. This 

research was undertaken to emphasize the need for contact or 

triboelectric charge generation testing when using materials in 

sensitive or hazardous environments. 

Surface Resistivity 

(Ohms / Square) 

When the threat of electrostatic discharge is a concern during 

an operation or application, the users need to be aware of the 

material’s tendency to acquire large voltages that could lead to a 

spark. The majority of material testing concentrates on surface 

resistivity measurements. These measurements determine the 

material’s ability to bleed off accumulated charge. However, 

resistivity measurements do not give an indication of the initial 

quantity of charge and therefore on the voltage that the material 

has a tendency to develop. This initial charge accumulation can 

pose a threat in hazardous environments. 

Copper 

Aluminum 

Dissipative or conductive materials can retain large amounts 

of charge and hence maintain hazardous voltages if the materials 

are not connected to a sink to which they can bleed the charge. 

This research concentrated on the importance of triboelectric 

charge generation tests for hazardous areas. Triboelectric testing 

can identify safer materials that have a low tendency to retain 

charge and therefore a low tendency to develop high voltages. A 

spark can pose a serious threat in areas with sensitive electronics 

or flammable materials. Areas where conductive and dissipative 

materials cannot be adequately 

List of Tested Materials in Order of Increasing Surface Resistivity grounded or where insulating materials 

<1 

<1 

Polyethylene Film 

Antistatic Nylon Film 

PTFE 

lo1* 

1013 - 

I PVF Film I 107 I 

Polycarbonate 

HDPE 

1 Electronics BaE ~~ I 1010 I 

~~ 

>1016 

1 AN100 I loll I 

must be used pose an increased risk for 

dangerous charge accumulation. 

Analysis of the data emphasizes that 

there is no correlation between surface 

resistivity and charge accumulation 

tendencies of materials. The lack of this 

correlation points to the importance 

of developing a better understanding 

of the triboelectric charging tendency 

of materials and to the importance of 

developing a test method that can be 

used during operations to monitor 

charge accumulation. 
20 

~~ Materials Listed in Order of Increasing Surface Resistivrty 

80 

Induced Charge per Unit Area on a Copper Plate Wien Contacfed by the Listed Materals 

Wzose Surface Resistivity Increases From Left fo Right 

I I I * 

a 

U 

K -50- 

0 

m 

r 

- 

F -50- 

0 -70- 

Bo' 

I 

t 

Materials Listed in Order of Increasing Surface Resistivity 

Induced Charge per Unit Area on a PTFE Sheet When Contacted by the Listed Materals 

Wzose Surface Resistizlity Increases From Left to Riglit 

Contacts: E.E. Groop (Elleii.E.G~.oc)~lii( ~zasn.goi)), YA0-7'' (321) 867-4861; and Dr. C.I. Calle, YA-C2-T, 

(321) 867-3274 

Participating Organization: USTDC (A. W. Nuwicki) 

125 
Method for Measuring Charge Buildup on Materials 

0.30 

al 

m 2 0.25 

0 

B 

2 0.20 

- 2 

T 

0 

3 

- - 

0 

9 0.15 

s 

lJl 

m 

- 

0.10 - 

0 

c 

c 

8 & 0.05 

a 

0.00 

The methods developed to test for the charge generation propensity of 

materials due to contact or triboelectric charging are not very reproducible. 

Triboelectric charge testing is highly dependent on the environment in which 

the testing is conducted and the type of contact that is made (rubbing, rolling, 

etc.). Research has been completed in the KSC Electrostatics and Surface 

Physics Laboratory that compares the current triboelectric testing method for 

electronic components developed by the Electrostatic Discharge Association to 

a new triboelectric testing method based on capacitive sensors for monitoring 

charge generation throughout a process or activity. 

For the most part, triboelectric testing of materials has consisted of using 

the material or device as it is used in an application and then depositing the 

device into a Faraday pail. Figure 1 shows a histogram of 25 data points taken 

using the Faraday pail method. This test method gives only a limited amount 

of information since it only measures the charge that remains on the device 

after it is used. 

The new device, shown in figure 2, is a rough prototype intended to 

be used to simulate a microchip. The charge that is generated on the pins 

of the chip is read as a voltage across a known capacitor (figure 3). The 

advantage of using this device to monitor charging is that the charge can be 

monitored throughout the process and not just at one point. This allows for 

the identification of electrostatic discharge hot-spots where high voltages are 

developed during an application or operation. 

i 

2 3 4 5 6 7 8 9 

Charge per Unit Area Induced on Sensor Heads (nC/cmz) 

Figure 1. Histogram of the Charge Developed on the Pins of the Device During Testing Using a Faraday Pail 
Coiztacts: E.E. Groop (€!lt’ll.€.G~.~’(’~”(’ 11~x7 

.pw), YA-C2-T, (321) 867-4861; aiid Dr. C.1. 

Calle, YA-C2-T, (321) 867-4861 

Participating Orgai~ization: LlSTDC (A. W. 

Nowicki) 

, 

Figure 2. Microchip Simulator Device for Measuring 

the Charge Induced on the Pins During ai2 Operation 

Q 

3 

Q 

_. 

I 

Figure 3. Charge Deueloped on Two Pins of the Microchip Simulator Wliile 

Being Rocked Six Times in a Magazine 

127 
128 

Electrostatics Testing of Materials at low Pressures 

The Electrostatics and Surface Physics Laboratory at KSC recently 

developed an Automated Electrostatics Test Device (figure 1) that operates in a 

chamber at KSC. With this device, it is now possible to perform electrostatics 

testing at atmospheric pressures down to 0.4 torr and at temperature ranges 

from -120 to 200 degrees Celsius ("C). The gas environment in the chamber 

can also be selected to reproduce a variety of environments found not only on 

Earth but also on Mars. 

Our laboratory has been testing the electrostatic charging and decay 

properties of materials for 40 years. Tests on thousands of materials during 

this time show the effects of relative humidity on the charging and decay 

properties. The new device allows us to test and identify the effects on 

electrostatic charging and decay under controlled humidity, temperature, low 

atmospheric pressures, and selected gas environments. Figures 2,3, and 4 

show typical test reports. 

This laboratory can duplicate a wide variety of environments where 

materials are used. Materials can be tested in these simulated environments to 

select the best one for a given application. The Automated Electrostatics Test 

Device was designed to measure both the electrostatic-generating potential 

and the discharge time of the material under test. It can test films, clothing 

materials, space suits, solar panels, solid foams, gloves, paints, and coatings. 

The device operates in a laboratory test chamber (after proper material 

acclimation time) under the following environmental conditions: 

Atmospheric pressure: 0.4 to 760 torr (0.533 to 1010 millibars). 

Temperature: -123 to 200 "C. 

Humidity: 0.5 to 100 percent. 

Various atmospheric gas combinations. 

Static fl 

Detector Head .c 

Figure 1. Automated Electrostatic Test Device 

Sample Carrier / R~EIC with Samples 

Buffer MotorMead 
ESD Test Report 

Ti. 

v-0 

Figurc 2. Material Test at Ambient Atinospheric Coizditioizs 

Coiztacts: Dr. C.i. Caiie i(--arios.i.eaiie~lzasa 

.pzi), YA-C2-Tf (321) 867-3274; and Dr. R.H. 

Gomyf, YA-C2-T, (321) 867-3267 

Pnrticipatiiig Organization: YA-F1 (PF. 

Riclziuso and J.A. Bayliss) 

Figure 3. Material SIzomiz iiz Figure 2 at 100 Torr 
Spacecraft Charge Monitor 

Electrostatic charging can damage spacecraft, cause instrument biases that limit the 

accuracy of scientific measurements, and even pose a direct hazard to personnel during 

extravehicular activity. Currently, there are few options available for monitoring spacecraft 

charge. An inexpensive, compact, reliable, easily deployed spacecraft charge monitor is 

needed. Funded by a NASA Small Business Innovation Research (SBIR) Phase I1 grant, 

Goembel Instruments is developing the Spacecraft Charge Monitor (SCM) for KSC’s 

Electrostatics and Surface Physics Laboratory. 

Goembel Instruments has advanced the SCM from concept to the initial stages of a 

complete flight instrument (figures 1 and 2). An SCM test apparatus was built to test the 

different prototypes of the SCM (figure 3). The SCM will bring a substantial improvement to 

the field of spacecraft charge detection and will find a ready market in aerospace. 

The SCM is based on electron spectroscopy. The method is dependent on a phenomenon 

known as photoionization. When sunlight strikes the Earth‘s upper atmosphere the 

molecules of nitrogen and oxygen eject electrons that travel with kinetic energies of only a few 

electronvolts (ev). A significant portion of the electrons produced in photoionizaton travel at 

four well-defined energies between 20 and 30 eV 22.2,23.9,25.2, and 27.2. These electrons 

travel away from the lower altitudes where they are produced (approximately 150 to 250 

kilometers [km]) to higher altitudes where they can be detected by spacecraft. 

The SCM measures the spacecraft charge by measuring the energy of electrons produced 

by the photoionization of the Earths upper atmosphere. The SCM’s reference potential is 

spacecraft ground. Spacecraft charging is apparent if we see a shift in the energy location of 

the peaks that would appear at 22.2, 23.9,25.2, and 27.2 eV on an uncharged spacecraft. If it 

Electrostatic 

Figure 1. Cutaway Drawing of the Flight SCM 

appears that the electrons have been slowed 

down as they approach the spacecraft, it is 

because the spacecraft is charged negatively, 

and the negative charge of the spacecraft 

repulses the negatively charged electrons. If 

the spacecraft has a negative floating potential 

of 10 volts relative to the space plasma where 

the photoelectrons were produced, the four 

photoionization peaks will appear at 12.2, 13.9, 

15.2, and 17.2 eV. By the same mechanism, 

positively charged spacecraft will accelerate 

the electrons to higher speeds, and the peaks 

will appear shifted to higher energies. One 

of the outstanding features of this method of 

determining spacecraft floating potential is 

that there is a one-to-one correlation between 

the shift in the location of the photoionization 

peaks and the spacecraft floating potential. 

Thus, the data analysis for floating potential 

promises to be direct and accurate. The 

diagram in figure 4 illustrates this one-toone correlation between spacecraft floating 

potential and a shift in the energy location 

I 130 
of the photoionization peaks in an electron 

energy spectrum. 

In designing and building the flight SCM, 

the device will be configured so it can be 

easily accommodated on a wide variety of 

spacecraft. The SCM will weigh 500 grams 

and consume 1 watt of power. The device will 

fit within a space of 6.5 by 3 by 3 inches. In 

addition, the SCM will use an RS-422 interface, 

and it is likely that we will be able to build the 

flight prototype SCM with radiation-tolerant 

parts (rated to 12 kilorads). Because the Air 

Force Research Laboratory (Hanscom AFB, 

Massachusetts) has offered to recommend 

the flight of the SCM through the Air Force 

Space Test Program (STP), the device is being 

designed to meet or exceed standards for STP 

flights. Their standards include operation 

at a temperature range of -24 to +61 degrees 

Celsius, vibration tolerance to 9 g's, and the 

ability to operate with power from a 28 26-volt 

spacecraft power system. 

The SCM has been included, for a second 

time, in a proposal to fly as a component of 

a suite of science instruments on a series of 

satellites. More information about the SCM 

can be found at MiMr~2r.booembel.hiz. 

The following tasks were completed under 

the SBIR Phase I1 grant: 

Converted an empty room into a state-ofthe-art electron spectroscopy laboratory. 

Completely designed a flight-quality SCM 

sensor head. 

Completed a preliminary design for the 

flight electronics. 

Contracted Payload Systems, Inc. and Elf 

Electronics, Inc., to deliver flight-quality 

boards for the SCM. 

Coittact: Dr. C.1. Calk (Cados.l.Callc@ ~ma.pl), 

YA-C2-T, (321) 867-3274 

Participating 0 rga IZ iza t ion: Goem bel 171s t ru ineiz ts 

(Dr. L. Goentbel) 

Figure 2. SCM (liicomplete Assembly) and 

Standard Business-Size Emelope for Scale 

Figure 3. The SCM Test Apparatus iiz tlze Goembel 

ltistrziineizts Laboratory 

Floating Floating 

I Potential I Potential 

I =V, -23.9 I =V, -23.9 ' 

I I I 

I I I: I I 

15 v, 23.9 v, 35 

Energy (eV) 

Figure 4. Spectra (Simulated) That Exhibit 

n Sh<ft in the Energy Locatioii of the Peaks 

131 
132 

Antistatic, Field-Emitting, Nanoparticle Coatings 

Production of antistatic, field-emitting coatings that can prevent 

destructive electric discharges is currently underway at Eltron 

Research Inc., funded by a Small Business Innovation Research 

(SBIR) Phase I1 contract with NASA KSC. The antistatic coatings are 

being designed to survive the atomic oxygen and intense ultraviolet 

light of low Earth orbit by using refractory electron-conducting 

ceramic oxides. To silently discharge a surface through electron 

field emission, it is necessary to produce local electrostatic fields 

in the range of lo9 to 1O1O volts per meter (V m-I). To achieve such 

high fields at modest applied voltage, small, nano-sized particles 

are used in the coatings. Assuming as a first approximation that 

the electrostatic field E of an isolated spherical particle is given by 

E = V/r, where V is the applied potential in volts and Y is the radius 

in meters, then the estimated required radius of the field emitters is 

given by Y = V/E. For a modest applied potential of 500 V, the radii 

of the particles need to lie in the range between Y = 500 V/109 V m-I= 

5 x m and 500 V/lOIO V 5 x m. In nanometers (nm), the 

particles must have radii between 50 and 500. Particle diameters, 

therefore, must be in the range of 100 to 1,000 nm. 

Because of the extremely high surface free energy of finely 

divided matter, most base metals and semiconductors of nanometer 

size would rapidly oxidize in Earth's atmosphere or would react 

with atomic oxygen in orbit to form insulating surface layers. 

Therefore, electron-conducting ceramic oxides were chosen for the 

new coating materials. Since these materials are already stable 

oxides, there is little tendency for them to further oxidize. All but 

one of the compounds synthesized over the past year are stable to 

at least 900 degrees Celsius ("C) (1173 kelvin [K]) in air. The figure 

is a scanning electron microscopy image of an electron-conducting 

coating showing the typical submicron particle size of the materials. 

The new materials compare favorably to conductive graphite and 

some of the conductive oxides used previously by NASA. 

By varying the stoichiometry of the compounds, materials can 

be made that are either metallic conductors or semiconductors. 

Metallic conductors increase in resistivity with increase in 

temperature, whereas semiconductors increase in resistivity with 

decrease in temperature. Metallic conductors are preferred to avoid 

possible formation of insulators from semiconductors if coatings 

experience cryogenic temperatures in space. 

Tests are currently under way to optimize the use of the new 

materials with inorganic binders that have been used successfully 

by NASA as binders for zinc orthotitanate and other coatings. 

Inorganic binders are preferred because, unlike organic binders, 

they are not degraded by atomic oxygen and vacuum ultraviolet 

light. Preparations have also been made for testing the stability 
of materials in vacuum while exposed to atomic 

oxygen and ultraviolet radiation from a deuterium 

lamp equipped with magnesium fluoride windows 

Mrith transmission down to 115 nm. Samples of 

three materials were accepted by NASA for the 

next long-duration exposure facility flight. 

Coizfact: Dr. C.I. Calk (Cdos.l.C(~lle@ ium.gw), YAC2-T, (322) 867-3274 

Part icipa t iiig 0 rgan iza t ion : Elf YO 11 Research Iiz c. (D r. 

M.V. Murzdsclzau, Dr. X. Xie, Dr. 1.S. TIroinpsoi7, D.A. 

Gribble, and M.A. TayloY) 

72) 

CD 

?2 

0 

ri 

I 

Q 

Is 

a 

133 
Corrosion-Resistant Tubing for Space Shuttle Launch Sites 

The existing 304 and 316 stainless-steel (SS) tubing and fittings 

at Launch Pads 39A and 39B are susceptible to pitting corrosion. 

This pitting corrosion can cause cracking and rupture of both highpressure gas and fluid systems. The failures can be life-threatening 

to launch pad personnel and can influence schedules. Outages 

in the systems where the failures occur can affect the safety of 

Shuttle launches. Improved corrosion-resistant tubing systems 

will greatly enhance the safety of personnel and the Shuttle. These 

new materials will require less maintenance over their lifetime and 

significantly reduce costs associated with these systems. 

A range of materials was selected for testing that includes 

stainless steels (austenitic, low-carbon, Mo-alloy, superaustenitic, 

duplex, and superferritic), Ni-Cr-Mo alloy, Ni-Cr-Mo-Fe-W alloy, 

and austenitic Ni-base superalloy. Each tubing test article contains 

a series of 90-degree bends, 37-degree tube flairs, and an orbital 

weld. Four environmental conditions are being tested on the same 

beach where the launch pads are located: seacoast (unsheltered 

and sheltered) and seacoast with acid rinse (unsheltered and 

sheltered). 

Based on the 1-year exposure data of the materials tested, 

austenitic stainless steel AL6XN was found to be the most costeffective tubing material to replace the existing 300-series stainless 

steel. AL6XN welded tube was determined to be in the same price 

range as seamless 300-series SS while offering much higher pitting 

corrosion resistance. 

The 304L SS tubing is in the process of being replaced with 

AL6XN at the launch pad. The tubes are being installed using 

orbital welding. This welding process did not compromise the 

mechanical properties of AL6XN SS. During a pull test, a welded 

AL6XN tube failed far away from the orbital weld, proving the 

strength of the weld joint. Several tests were conducted to assess 

the integrity of the material’s properties in the postwelded tube. 

X-ray dot mapping was conducted to test for segregation 

of alloying elements. This test showed no anomalies with the 

microstructure of the material. X-ray photoelectron spectroscopy 

(XPS) was used to test the integrity of the protective oxide layer. 

The surface of the postwelded AL6XN tubing was restored by 

passivating it with nitric acid, which is the preferred passivating 

method for SS in industry. 

Metallography of representative samples was performed to 

check for sensitization issues using American Society for Testing 

Materials (ASTM) standards as a guideline. Figure 1 shows a 

micrograph of an AL6XN sample that has been etched with oxalic 

acid to test for sensitization. Sensitization is the formation of 
chromium carbides along the grain 

boundaries, lvhich weakens the 

material’s ability to withstand certain 

corrosive environments and eventuallv 

leads to failure. When sensitation is 

present, it is revealed by the formation 

of a ditch structure as shown in Figure 

2. The AL6XN micrograph in Figure 1 

does not show the ditch structure that is 

represe1~tatix.e of a sensitized sample. 

(Reprinted. with permission from A 262-02a Standard Practices for 

Detecting Susceptibilitv io lntergranuiar Attack in Austenitic Stainless 

Steels copyright ASTM International) 
Electrochemical Characterization of Tubing Alloys in Simulated 

Space Shuttle Launch Pad Conditions 

The 304L stainless-steel (304L SS) tubing is used in various supply lines that 

service the Orbiter at the KSC launch pads. The atmosphere at the launch site 

has a very high chloride content caused by its proximity to the Atlantic Ocean. 

During a launch, hydrochloric acid is produced as a by-product of the Solid 

Rocket Booster exhaust. The acidic chloride environment is aggressive to 

most metals and causes severe pitting in many common stainless-steel alloys. 

The 304L SS tubing is susceptible to pitting corrosion that can cause 

cracking and rupture of both high-pressure gas and fluid systems. These 

failures can be life-threatening to launch pad personnel in the immediate 

vicinity. Furthermore, outages in the systems where the failure occurs can 

result in expensive delays in Shuttle operations. The implementation of a 

new tubing alloy will greatly improve safety and reduce maintenance costs, 

downtime losses, and the probability of failure. 

The objective of this investigation is to study the electrochemical behavior 

of 10 corrosion-resistant stainless-steel alloys to replace the 304L SS tubing at 

the Space Shuttle launch sites. The alloys included 317L, 316L, 2205, 

C-276,625,254 SMO, C-2000, AL-6XN, AL29-4C, and 2507. As a reference for 

the candidate materials, 304L SS was included in the study. The specimens 

were obtained from tubing that was cut and flattened prior to testing. Three 

electrochemical techniques were used to evaluate the performance of the 

materials: corrosion potential, polarization resistance, and cyclic polarization. 

The samples were placed in an electrochemical cell designed to expose a 

surface area of 1 square centimeter to the electrolytic solution. Electrolytic 

solutions were chosen to emulate conditions that are less aggressive than, 

similar to, and more aggressive than those found at the launch pads at KSC. 

The figure shows the results from the cyclic polarization measurements 

for 304L SS alloy exposed to neutral 3.557% NaCl and 3.5551 NaC1-1.ON HC1 

solution. The addition of hydrochloric acid results in a more aggressive 

environment, and the increased acidity of the solution is detrimental to the 

performance of the material. The position of the scans in the potential-current 

diagram shows that the corrosion performance of 304L deteriorated in the 

3.55% NaC1-1.ON HC1 solution. 

A complete set of measurements is being collected using standard coupon 

samples under the same electrolytic conditions. Preliminary data suggest a 

strong correlation between the performance of the samples in the laboratory 

and similar specimens that were atmospherically exposed at the KSC Beach 

Corrosion Test Site. 

Contacts: Dr. L.M. Calk (Luz.M.Calle@nasa.gov), YA-C2-T, (322) 867-3278; and 

L.G. MacDozuell, YA-C2-T, (321) 867-4550 

Parficiyafiiicy Orgaizizatioiz: USTDC (R. D. Vinje) 

136 
300 . 200 

1 1 N-HCL.DAT 100 

0 

1 NEUTRAL.DAT 

-400 

-500 

-600 

-700 

-1 0 

rurrL 

-9 

Cyclic Polal-izatioizs for 303L Stniiiless Steel iii Neiitial3.55' r NaCl mid 3.55'( h'aC1-1 .Oh' HCl 

137 
The final element of the sensor is a galvanic cell array. The galvanic cell 

methodology is unique (developed at KSC) in that it attempts to analyze 

electrolytes based upon the corrosion rate and corrosion potentials exhibited 

by the metals. In essence, the galvanic cell concept (differing metals) is 

used to drive the system at a potential indicative of the corrosive processes 

taking place. This allows the electrochemical processes to occur without the 

application of external power, except for that needed to take measurements. 

While these open-circuit potential values indicate the corrosivity of the 

Advanced Multisensor for Electrolytic Characterization 

This project investigated the feasibility of developing a sensing device for 

electrolyte characterization using an array of galvanic cells. The methodology 

needed to analyze and interpret the data is also being developed. The 

ultimate goal is to integrate the sensor into a subsurface explorer probe to 

determine the feasibility of grounding experimental devices on planetary soil, 

utilizing the soil as an electrical return path, and using the experimental data 

to select materials that will not corrode when in contact with the soil. While 

the focus of the research centers upon developing a device to study planetary 

surfaces, commercial applications are also feasible. These may include but are 

not limited to agricultural, biological, and geological analyses. The selection 

and optimization of galvanic cell electrodes to determine the corrosivity of 

solutions and soils are important to the investigation. 

The first component of the multisensor (shown in figure 1) is an ion 

selective electrode (ISE) array for the qualitative and quantitative detection 

of individual ions. The second constituent of the device complements the 

first technique through the use of anodic stripping voltammetry (ASV) and 

cyclic voltammetry (CV). ASV is an analytical method in which analytes are 

preconcentrated on the surface of a working electrode for a specific duration. 

Through an anodic potential scan, the analyte is stripped from the working 

electrode and oxidized back to its original form, with a voltammetric signal 

indicating the ions present in solution. CV is a similar technique, except that 

it does not include the preconcentration step of the anodic stripping analysis. 

These components are being investigated with standard, well-characterized 

redox couples to determine the sensitivity of the electrodes and resistance to 

interfering signals. 

system, they fail to provide information on 

the specific corrosion rates. These are often 

obtained from potentiodynamic polarization 

measurements. 

For potentiodynamic determinations, the 

potential is scanned through a preset range 

of voltages near the open-circuit potential. 

Straight lines can be fit to the anodic and 

cathodic areas of the potentiodynamic 

polarization scans, which result in a Tafel plot. 

This plot gives the open-circuit potential of the 

svstem as well as the corrosion current. 

Figure 1. Multiseiisiiig Ceramic Substrate J 
- -0.25 

(D 

0 

m 

C 

0) 'j -0.15 

c 

0 

- c 

L - -0.2 

- 

L 

n 

(D -0.1 

9 

b 

- 

0 -0.05 

0 

DI H,O Tap Water 0.25% NaCl 0.25% NaCl + Sat. KCI 

1% NaN03 

1.8 E-07 

1.6 E-07 

- 

e 1.4 E-07 4 

1.2 E-07 2 

1.0 E-07 5 

0 

8.0 E-08 s 

- 

E 

.- 6.0 E-08 

5 4.0 E-08 0 

2.0 E-08 

0.0 E+OO 

Figure 2 shows the corrosion potentials and corrosion currents for a set of 

aqueous solutions. These were specifically chosen to illustrate a wide range 

of corrosivities. As shown, deionized (DI) water and tap water were the least 

corrosive. An addition of 0.25-percent sodium chloride (NaCl) to the DI water 

increases the corrosivity of the system as shown by the corrosion potential and 

corrosion current values. In contrast, an addition of 1-percent sodium nitrate 

(NaNOJ to this solution shows a decrease in both values. This is indicative 

of the corrosion-inhibiting properties of the NaNO;, which is a well-known 

corrosion inhibitor. Finally, the corrosion current of the saturated potassium 

chloride (KCI) solution clearly shows the greatly increased corrosivity of the 

system. These results are consistent with the known corrosivities of these 

solutions. 

Kev accomplishments: 

Completed cyclic voltammetry experiments with standard redox couples to 

investigate the response of the electronics and CV electrodes of the sensor. 

Integrated the sensing ceramic with commercial instrumentation to validate 

the response of the sensor. 

Gathered open-circuit potentials for the galvanic couples for a wide range 

of solutions varying in corrosivity. 

Obtained potentiodynamic polarization measurements for the individual 

galvanic couples exposed to solutions with a wide range of corroqivities. 

139 
Chromate Coating Replacement for Military Aircraft 

Alloys 

2024-T3 

2219-T87 

5083, H131 

7075-T6 

To enhance corrosion resistance of aluminum alloys, hexavalent chromiumbased products are presently used. Unfortunately, tighter restrictions are 

being placed on the use of these coatings because of their carcinogenic 

effects and endangerment to the environment. The objective of this project 

is to expose 960 nonchromate-coated coupons to a harsh, outdoor marine 

environment at the KSC Beach Corrosion Test Site (BCTS) while tracking 

the deterioration of the surface pretreatments and coatings. This study is 

to determine the effectiveness of protection of environmentally compliant 

chromate replacement technology for chromate-inhibited coatings for 

aluminum materials on military aircraft. 

The 960 coupons form a matrix of four aluminum alloys having nine 

pretreatments, each of which has five paint systems, all in sets of three or 

five coupons (see the table). Each coupon is scribed and was evaluated in 

accordance with ASTM D 1654, Standard Test Method for Evaluation of 

Painted or Coated Specimens Subjected to Corrosive Environments. This 

ASTM standard uses a rating system of 0 (worst) to 10 (best) depending on 

the distance of creepage from the scribe. The coupons were observed on a 6- 

week schedule and intensively evaluated on a 4-, 8-, and 12-month schedule 

using ASTM D 1654 ratings for blister sizes. All blisters were measured onesided, that is from the scribe to the creepage front. All blister measurements 

were recorded for each coupon in millimeters. The maximum, minimum, and 

representative means were recorded along with each blister measured. 

Alloys, Pretreatiizcizts, and Coating Systems 

Pretreatments 

Alodine 1200s (immersion) 

TCP 10 (immersion) 

Alodine 5200 (immersion) 

Sanchem (spray) 

X-IT Prekote (wipe) 

Oxsilan AL-0500 (immersion) 

Chemdize 727ND (wipe) 

Bi-K Alkimate (immersion) 

Boegel (spray) 

Primers 

53022, carc solvent 

53030, carc water 

85582nc, nonchromatedwater 

85582~1, chromated-water 

23377, chromate-solvent 

Topcoats 

53039, carcsolvent 

82585, high 

solids 

After 1 year, all coupons with an ASTM rating of 2 or less were considered 

failures, removed from the site, and sent back to the customer for evaluation. 

The remaining coupons that showed little or no degradation at the end of the 

1-year period are being exposed for a longer period. There were 194 coupons 

that had an ASTM rating of 2 or below. All of the failures came from the 

coatings applied to the 2024-T3, 7075-T6, and 2219-T87 alloys. These failures 

consistently came from the 53030 / 53039,53022 / 53039, and 85582nc / 85285 

paint systems. There were no fail rankings from the same coating systems 

applied to the 5083-Hl31 alloy. 

140 
Coupoiz Racks ut tl7e Beach Corrosioii Test Site 

Blistered Coupoiz, 8-Moi~th Exposure 

Key accomplishments: 

Blistered aiid Contiizg Delniniiintioii 

Coupoii, 8-Moiith Espsz4re 

Built panels to hold 960 nonstandard-sized coupons. 

Installed approximately 960 coupons at the BCTS. 

Completed 1-year evaluations and photographic documentation. 

Developed a portable camera stand for taking consistent coupon photographs. 

Compiled a 6-month summary of results. 

Compiled a 12-month report of the test evaluation. 
Army Aircraft Alloys - Corrosion-Retardant Additive Testing 

This project tested and evaluated commercially available chloride rinse 

agents (CRAs) for effectiveness of corrosion reduction by rinsing typical 

aviation metals. These metals are used in Army aircraft, missile, and ground 

vehicle systems and components. Four separate CRAs are being tested for 

their effectiveness in preventing or reducing corrosion of metals exposed to 

seacoast environments. Nine different metals specified by the U.S. Army have 

experienced a harsh, outdoor marine environment at the KSC Beach Corrosion 

Test Site (BCTS). Replicate alloy samples were exposed and rinsed weekly for 

periods of 1 and 2 years. CRAs and control solutions were sprayed onto metal 

coupons each week. Comparison testing of products and recommendations 

were reported at the 1-year and 2-year points. CRAs, numbered 1 through 4, 

were compared to seawater, demineralized water, and no-rinse as controls. 

The following materials / coatings were selected for testing: 

Aluminum 2024-T3 with 8625 coating. 

Aluminum 2024-T3 with 5541 coating. 

Aluminum 7075-T6, uncoated. 

Magnesium 4377 with 3171 coating. 

Steel 4340 carbon steel, uncoated. 

I" 13-8 molybdenum (Mo) high-strength steel, uncoated. 

AM-355 CRT high-strength steel, uncoated. 

C-250 maraging high-strength steel, uncoated. 

Titanium 6A1-4V, uncoated. 

To establish CRA ranking, pitting characteristic ratings were applied to 

the aluminum and magnesium alloys. The results of this study showed 

that two of the chemical rinse agents tested, CRA 1 and CRA 3, were more 

effective in reducing corrosion on both aluminum and magnesium than the 

same exposure and rinsing with either demineralized water or tap water. 

These results contrast with the results for the other two commercial rinse 

agents tested. Rinsing with either CRA 2 or CRA 4 resulted in even more 

corrosion on aluminum or magnesium than rinsing with demineralized water 

or even with seawater. Because of the detrimental effects on the aluminum 

alloys, these two products are not recommended for use on military aircraft. 

Typical Pit, 7075T6 Aluminum Alloy - 

Magizificafion x 100 

Typical Pit, Magnesium Alloy - 

Magnification x ZOO 
Corrosion ratings were based on weight loss for the 4340 steel and the C-250 maraging 

steel. The results of testing with the four rinse agents on these steels produced mixed 

results. Rinsing with CRA 3 produced the loivest corrosion rate, while rinsing with CRA 1 

resulted in the highest corrosion rate. Rinsing with CRAs 2 and 4 produced similar results. 

The weight loss results are unclear due to the limited number of exposure samples for each 

alloy /exposure combination in this extensive studv. The corrosion-resistant stainless steels 

and titanium alloys only showed superficial tarnishing and had no beneficial or detrimental 

effects. 

Key accomplishments: 

Completed 1-year and 2-year exposures of coupons. 

Cleaned and evaluated all nine coupon material sets at the 1-year and 2-year points. 

Completed and delivered 1-year progress report. 

Presented I-year data at Department of Defense Tri-Services conference. 

Presented 2-vear data at National Association of Corrosion Engineers (NACE) 2003 

Developed and instituted a method to rate CRA effectiveness using ultrasound technologv 

corrosion conference. 

and image analysis software. 

Key milestone: 

Prepare and deliver final report and 

associated coupons. 

:<C~I,), YA-C2-T, (321) 867-4550 

Pm?icipf ir ig Organ iza fioiz: US TDC (1.1. 

Curran ailn 1.P C211.ra12~ 
Electrochemical Analysis of Chloride Rinse Agents on Aircraft Alloys 

NASA’s Corrosion Technology Testbed is investigating the effectiveness 

of four chloride rinse agents (CRAs) to prevent or reduce corrosion of 

alloys commonly used in aircraft. The investigation included atmospheric 

exposure of rinsed samples of the alloys at the Beach Corrosion Test Site and 

electrochemical measurements at the Corrosion Laboratory. Both facilities are 

located at the Kennedy Space Center. Results from the atmospheric exposure 

investigation are the subject of ”Army Corrosion-Retardant Additive Testing” 

(KSC Research and Technology 2002 Annual Report). The current report 

presents the results obtained using electrochemical techniques. 

The alloys employed in this investigation are listed in the table. 

Electrochemical measurements are performed using an electrochemical cell 

to expose the aircraft alloys to four commercially available CRAs (CRA 1, 

CRA 2, CRA 3, and CRA 4), seawater, and deionized (DI) water. Seawater 

and DI water are included as reference electrolytes. The rinse agents are 

diluted according to the manufacturers’ specifications. The electrochemical 

techniques being used include corrosion potential measurement, linear 

polarization, potentiondynamic scans, and cyclic polarization. The techniques 

used depend on the alloy being investigated. 

Metal Alloys Used in the Electrochemical Analysis of Chloride Rinse Agents 

Metal UNS Designation 

Alloys Alloy Designation This Study 

J9100 AI 2024-T3 / 5541 

Aluminum 

A97075 AI 7075-T6 

Steel Alloy G43400 4340 

S13800 PH13-8Mo 

1 S45850 I C-250 High-Strength 

Steel 1 S35500 1 AM-355 I I 

Titanium Alloy 1 R65400 Ti-6A1-4V 

I M11311 1 AZ31B-H24 Magnesium 

Allov 

Percent Composition 

Al, 4.5 Cu, 1.5 Mg, 

0.6 Mn 

AI, Zn 5.6, 2.5 Mg, 

1.6 Cu, 0.3 Cr 

Fe, 0.4 C, 1.8 Ni, 0.8 

Cr, 0.25 Mo 

Fe, 13 Cr, 8 Ni, 

2 Mo 

Fe, 18 Ni, 7.5 Co, 

5 Mo, Ti, A1 

Fe, 15.5 Cr, 4.5 Ni, 3 

MO 

Ti, 6 AI, 4 V 

Mg, 3 AI, 1 Zn 
To date, data has been collected on 

AI 2024-T3 and A1 7075-T6 aluminum 

alloys. Open-circuit potential and cyclic 

polarization techniques were used to test 

the samples. Sample preparation involved 

cleaning the samples with ethyl alcohol, 

which served as a degreasing agent. This 

method was utilized to preserve the 

samples in a condition similar to the initial 

state of the panels placed at the Beach 

Corrosion Test Site. 

Open-circuit potential measurements 

were monitored until a stable potential was 

achieved. Measurements were determined 

to be stable when the values of potential 

varied by no more than 2 5 millivolts 

(mV) for a period of 10 minutes. Once a 

stable potential was achieved, the cyclic 

polarization technique was used. 

P,na!ysis of he data shows that, of the 

four rinse agents, CRA 3 is the best and 

CRA 4 the worst for AI 2024-T3 / 5541. 

Seawater proved to be the most damaging 

electrolyte to this aluminum alloy. Figure 

1 shows the results of cyclic polarization 

scans for A1 2024-T3 in CRA 3 and DI water. 

Plot 1 of figure 1 has a positive hysterisis 

loop, which is indicative of pitting, and 

a smaller corrosion current than plot 

2, indicating that CRA 3 is better at 

reducing corrosion than DI water. Similar 

performance was revealed in the case of 

A1 7075-T6. CRA 3 performed best while 

CRA 4 proved to be the most aggressive 

rinse agent. Interestingly, CRA 4 proved 

to be more detrimental to the sample than 

seawater. Figure 2 shows the results of 

cyclic polarization for A1 7075-T6 in CRA 4 

and DI water. 

I 

I 

Contacts: Dr. L.M. Calk 

iLuz.M.Cnlle@?zasa.gou) YA-C2-T, (321) 867- 

3278; and L.G. MacDowell, YA-C2-T, (321) 

867-4550 

Parficipatiizg Orgaizizatioiz: USTDC (R. D. 

Vi71 je) 

le00 1 . . .."..I . , . . "., , . , ....'I . . , ..."I * , , 'ml 

lsoo - 

200 - 

0- 

-1 3 -1 1 4 -7 -3 

log (l)(log (A)) 

Figure 1. Cyclic Polarization Plots for A1 2024-T3 in Dl 

Water and CRA 3 

Figure 2. Cyclic Polarization Plots for AI 7075-T6 iiz DI 

Wafer and CRA 4 
F? 

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Development of Liquid Applied Coatings for Protection of Steel in Concrete 

Corrosion of reinforcing steel in concrete is an insidious problem facing 

the NASA facilities at KSC, other Government agencies, and the general 

public. These problems include damage to KSC launch support structures, 

including transportation and marine infrastructures, as well as building 

structures. Because of these problems, the development of a galvanic 

liquid applied coating system would be a breakthrough technology having 

great commercial value for the following industries: transportation 

infrastructure, marine infrastructure, civil engineering, and construction. 

Developing this type of system had several goals: 

Phase I concentrated on the formulation of coatings with easy application 

characteristics, predictable galvanic activity, long-term protection, and 

minimum environmental impact. These new coating features, along with 

the electrical connection system, will successfully protect the embedded 

reinforcing steel through the sacrificial cathodic protection action of the 

coating. 

Phase I1 improved on the formulations and included optimizing metallic 

loading as well as incorporating a humectant for continuous activation. 

In addition, development of optimum electrical connections continued. 

Phase TI1 incorporated improvements from the previous phases to the test 

blocks. 

Phase IV incorporated the final upgrades onto large structures that are 

heavily instrumented (see figure 1). 

Laboratory testing has demonstrated the commercial potential of this 

technology. Presently, testing is being conducted at the NASA/ KSC 

Corrosion Technology Testbed Beach Corrosion Test Site with positive 

preliminary results. Figure 2 shows the correlation between weather and 

protection of steel in reinforced concrete. This demonstrates that protective 

current is available as needs arise. In addition, the data is being collected 

and remotely accessed from offsite locations. 

Successful development and continued optimization of this breakthrough 

system would stimulate great interest in the NASA/KSC Corrosion 

Technology Testbed. Commercial patents on this technology would 

enhance KSC’s ability to attract industry partners for similar corrosion 

control applications and establish the Corrosion Technology Testbed as a 

leader in solving corrosion problems. 

Key accomplishments: 

Proved the feasibility of using liquid applied coatings for protection of 

embedded reinforcing steel in concrete. 

Determined the optimum mix ratio for specific liquid applied coatings. 

Determined that the addition of moisture attractors adds little or no 

benefit to coating performance. 

Developed improved electrical connectivity between the coating and the 

iTiternal reinforcing steel. 
Added intensive embedded 

instrumentation in new simulated 

concrete structures for monitoring 

corrosion protection. 

Obtained patent. 

Successfully licensed product / 

Coated Phase IV structures. 

technology to commercial partner. 

Key milestones: 

Analyze metallic loading. 

Install and evaluate embedded 

Ag / AgCl reference electrodes. 

Investigate coating-concrete 

interfacial resistance. 

Realkalinize carbonated concrete. 

Pnrticipating Orgaizizatiolz: USTDC 

(1.1. Curraiz, 1.P Curran, and Dr. R.G. 

Barile) 

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DAYS 
Noise Mitigation of Ducted Supersonic Jets for Launch Exhaust 

Management Systems 

The Launch Systems Testbed (LST) at KSC was created to support the development of clean 

launch pads for future launch vehicles, incorporating passive sound mitigation techniques to 

replace active noise reduction methods, such as water suppression systems. These launch pad 

designs focus on reducing cost while enhancing safety and quick turnaround times. The overall 

goal of the LST is to establish a capability to simulate a small-scale launch vehicle environment for 

use in testing and evaluation of launch pad designs for future space vehicles. 

One of the LST's specific objectives is to develop launch exhaust management systems (LEMS) 

that effectively reduce the generated acoustic loads with innovative duct designs devoid of water 

injection systems. Scale-model testing, along with analytical methods, is helpful as a means of 

predicting the full-scale acoustic environment in designing such systems. A preliminary series 

of tests was designed to define the jet acoustic load environment for closed-duct and partially 

open-duct (upper duct wall removed) configurations. The LST has conducted cold-jet tests with 

nitrogen gas issuing from an ideally expanded small-scale nozzle of 1-inch exit diameter with an 

exit Mach number of 2.5. Tests were also conducted with free jets, jets passing through closed 

ducts, and jets passing through partially open ducts with an upstream J-type deflector. 

With the use of a closed duct, the overall sound power of a Mach 2.5 supersonic jet is reduced 

by about 3 decibels (dB). The peak frequency is found to increase above the free-jet value as 

the angle from the jet axis is increased. These results also suggest that there is an optimum 

distance between the nozzle exit plane and the duct inlet for minimizing the sound power. The 

partially open duct results in increased sound levels near the duct axis relative to the free-jet case. 

Regarding the closed duct, larger reductions in sound power may be realized by increasing the 

-= 

660 

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Pressure 

144 Regulator 

, N2 Bottles dd,iiill (6000 psi) 

r' 

Oxrall Test Setup 

duct length, widening the duct 

cross section (adding a diffuser), 

and including absorbing liners 

on the duct walls. 

Key accomplishments: 

Characterized the flow and 

acoustic environment of a free 

supersonic jet. 

environment of a supersonic 

jet flowing through a covered 

duct with a J-deflector. 

Characterized the acoustic 

environment of a supersonic 

jet flowing through an open 

duct. 

Characterized the acoustic 

Key milestones: 

Completed Phases I and I1 of 

Air Force Research Laboratory 

(AFRL) LEMS. 
130 

120 

110 - 

F 

5 100 0 

c? c 

90 - 

i (I) 

80 

70 

fin 

Run#2 

Mach 2 5 Free Jet 

Nozzle Exit Diameter 1 inch 

Nozzle Exit Plane 73 inches above gound 

Acoustic Sensors 80-inch arc radius 54 inches above ground 

~~- 

- Mic 1 

Mic 2 

__ Mic 3 

MlC 4 

Mic 5 

+ MI~ 6 

Mic 7 

-C- Mic 8 

- -Mic9 

-- 

- 

JetlDuct Configuration 

-- 102 103 10' 1 05 

Frequency, kHz 

Spectral Sound Power for the Free Jef 

130 

120 

110 

$ 100 

8 

2 90 Navy-NASA-Air Force (JANNAF) 

Conference - c and International - 

Presented the study at Joint Army- - 

J 

a 80 Congress in Sound and Vibration v) RuMl 

- Mic 1 

4 Mic 2 

Mic 3 

--t ?k 4 

Mic 5 

--F MIC 6 

Mic 7 

--It Mic 8 

-+ Mic 9 

__ 

- 

Closed Duct Mach 2 5 Jet 

Nozzle Exit Diameter 1 inch 

Nozzle Exit Plane 73 inches above gound 

Duct Inlet Plane 10 inches below nozzle exit plane 

Acoustic Sensors 80-inch arc radius 54 inches above ground 

70 (ICSV). 

Delivered test results to the AFRL 

(RMLS) team. 

Contact: Dr. B.T. Vu 

Reusable Military Launch Systems 60 

50 1 02 103 104 10' 

Frequency, kHz 

Spectral Sound Power for the ]et Flo-iuing in a Closed Duct (E!*I/Lc' 7 \ ~/~(J~Io<P YC~Y~,), YA-C2-T, (321) 

867-23 76 

Par ticipa f iizg Organization: US TDC 

(Dr. M. Kaizdula) 

140 

130 

- 5 110 

rn 100 r 

1 6' E RuM10 

Open Dud Mach 2 5 Jet Nozzle Exit Diameter 1 inch 

Nouie Exit Plane 73 inches above aound 

Dun inlet Plane 10 incnes miow nonie exit pian 

ACG~S:,; Sensors 8Ginch ax rad,~s 

54 inches above ground 

60 .~. 

102 103 

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Frequency, kHz 

Spectral Sound Power for the Jet Flowing in a Partially 

Ope?? Duct 

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149 
Development of a Robust Finite-Element Method for Rocket 

Acoustics Problems 

Launching space vehicles generates extreme vibration and acoustic 

conditions that can affect the launch pad, launch vehicles, and payloads. 

These acoustic loads result from the acoustic field generated by the rocketengine exhaust stream with the ambient atmosphere and are the primary 

source of structural vibrations and internal loads during launch or static 

firings. Thorough knowledge and elimination of acoustic loads are critical to 

proper functioning of vehicle components and their supporting structures. 

Problems that might arise from acoustic loading include (1) malfunction 

of electrical and mechanical components because of structural vibrations 

and internal acoustic loading, (2) fatigue failure of internal components and 

supporting hardware, (3) fatigue of lightweight exterior structures, such 

as antenna panels, (4) fatigue of lightweight spacecraft structures, and (5) 

unsuitable environmental conditions for the vehicle occupant. To manage and 

suppress the undesirable conditions, we must be able to accurately predict the 

acoustic and vibration levels by either analysis or testing. 

The objective of the current project is to develop a user-friendly program 

to solve acoustic equations using the finite-element method (FEM). The 

method used is mixed-explicit implicit algorithm. This program will utilize 

the object-oriented program languages to create a unified environment for the 

preprocessor, the main solver, and the postprocessor. With this approach, the 

user can obtain the solution to the compressible Navier Stokes equations all 

in one environment, from the very raw data all the way to postprocessing and 

analysis of the results. 

Key accomplishments: 

Analyzed flow around complex geometry such as ramjets, scramjets, 

internal flow in a Space Shuttle main engine, and flow past cone-shaped 

objects. 

Developed a mixed-explicit implicit algorithm in finite elements. 

Conducted a seminar on FEM for rocket acoustics at NASA KSC. 

Key milestones: 

Acoustic computation using Kirchhoff and Flowc-Williams Hawkings 

Presentation of study at the Thermal-Fluid Analysis Workshop (TFAWS), 

Software patent submission. 

(FWH) codes. 

NASA-Langley Research Center. 

Con tact: Dr. B. T. Vu (R 1.1 i ctl . T. 1'1 1 i( 1 i 117517, go ii ), YA-C2 - T, (32 1 ) 86 7-23 76 

Pavticipating Organization: Alabama A&M University (Dr. A. Mobasher) 
Compufatioiial Grid for Ramjet 

Close-up View of the Stream Liiie Coiitoiirs Year 

the Posifioiz Wliere Shock Wme Iiiteracfs Wit17 the 

Boundary Layer (Tlze recirculatioli zone is i%ible.) 

Presswe Coilfours for Ramjet 

Coinpi~tatioizal Grid for Scramjet 
Tape Adhesion Testing Software 

The Materials and Processes (M&P) Laboratory, located in the Vehicle 

Assembly Building (VAB), supports External Tank (ET), Solid Rocket Booster 

(SRB), and Orbiter operations with in-process testing for a variety of materials, 

including adhesives, foams, ablatives, and films. 

The M&P Laboratory uses a Tinius-Olsen (T-0) 10,000-pound Universal 

Testing Machine to evaluate tape adhesion strength. Tape adhesion testing 

verifies tape for shelf life extension. The tape specimen is placed in the jaws 

of the T-0 machine, which pulls 2 inches of tape from a steel test panel. The 

tape adhesion testing software makes over 500 measurements during the 2- 

inch pull. The software combines this data to determine the force required 

to remove tape from the panel. It then automatically reports the results on 

the tape adhesion testing input screen for inspection and storage in the tape 

adhesion database. 

Previously, United Space Alliance (USA) laboratories at KSC did not have 

the ability to perform tape adhesion testing onsite. Performing this test onsite 

provides an obvious increase in processing efficiency. 

Another benefit is cost savings. Even though the primary use of this 

software is to test flight tapes, it also can be used to test ground support 

equipment (GSE) / nonflight tapes. Presently, we discard thousands of dollars’ 

worth of tape with expired shelf life each year. This software gives USA the 

onsite ability to test tapes in the VAB M&P Laboratory and to create cost 

savings. 

The software team and design team chose Oracle Forms and Reports 

running on a Microsoft Windows platform for the user interface. This 

combination of software application and operating system provided current 

database technology. However, Oracle Forms did not provide the specialpurpose communication functions this effort required. In response, the 

software team embedded C++ function calls within the form. This allows the 

user to click an on-screen button to 

initiate two-way communication with 

the T-0 via a serial port. This digital 

interface to an analog device allows 

the software to automatically direct the 

operation of the T-0 machine, receive 

over 500 test values, average these 

values, store the data in a database, and 

report the results. 

The C++ functions significantly 

extended tracking capabilities. These 

on-screen objects provide real-time 

display of T-0 operational values and 

display dialog boxes for direct user 

interaction as required. Although it 

automatically initiates the T-o and Tinius-Olsen 10,000-Pound Universal Testing Machine 

Performing an Adhesion Test 
Tqw Adhesion Software biterface 

starts testing, the software still preserves all 

manual backup capabilities. 

Advances in communication technology 

make the direct interface of a test process 

and its tracking software both convenient 

and practical. The ability to take multiple 

measurements and mathematically manipulate 

the results prior to automatic storage in a 

database adds processing flexibility. It avoids 

the duplication of effort (and errors) involved 

when using a human being / keyboard to 

interface an analog device to a database. In 

addition, it performs hundreds of calculations 

to provide the most accurate average possible, 

a process that would take hours to perform 

manually. 

This contribution has been used in the Space 

Shuttle program. The tape adhesion testing 

software provides the M&P Laboratory with a 

convenient and accurate way to automatically 

record and report its tests. It provides better 

support to SRB /ET processing with automatic 

recording and mathematical manipulation of 

multiple values from the test-generating device 

directly to the tracking and reporting 

software. 

The concepts developed in this tracking 

system can be used in any manufacturing 

process requiring automated multiple 

recording, manipulation, and storage of 

analog information to a database. 

Key accomplishment: 

Integration of C++ communications 

modules into Oracle Forms. 

Key milestones: 

2003: Program designed, coded, and 

fully implemented. 

Co7ZfaCt: PI. BfJOkHZa71 (p(?l?IC’l(7./. B00/<ilI~? I I 

((1 I/o~L~.:<oT~), YA-C1, (321) 867-6381 

Participfiiig Organizatiotz: United Space 

Allimce (S.M. Schizeider, A.M. Fodemusky, 

G. Lebovifz, and A. Beckus) 

153/154 
I 

Process and Human Factors Engineering 

Technologies 

iu 

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2 

: 

3- 

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Process and Human Factors Engineering Technologies include 

research and development of innovative tools and technologies 

to improve process safety and efficiency. Spaceport and range 

systems have many unique aspects that require development of 

advanced process, human factors, and industrial engineering 

technologies. Process and Human Factors Engineering emphasizes the interfaces among people, processes, and hardware / 

software systems in specific work environments. Process and 

Human Factors Engineering focuses on the science of process 

improvement and optimization of the operational phases of 

complex systems, including current and future space transportation systems. The overall goal of Process and Human Factors 

Engineering Technologies is to develop highly effective technologies for designing, implementing, improving, and managing 

safe and efficient processes, systems, and work environments 

that can be quickly adapted to the changing needs of current and 

future spaceports and ranges. 

Process and Human Factors Engineering Technologies 

directly support NASA's goals of achieving safe, reliable, and 

low-cost space access and exploration. Technology projects 

develop new concepts, methodologies, processes, and /or systems that advance the state of the art in one or any combination 

of the following technology focus areas: 

Modeling and Simulation 

Human Factors and Ergonomics 

Task Analysis Technologies 

Process and Operations Analysis 

Life Cycle Systems Engineering Tools 

Scheduling and Risk Assessment Technologies 

Management Support System Technologies 

For iizorc iizforinafioii reprdiiig Process arid Huii7nii Factors Eizgiuceriizg 

Tcchiiologies, please coiztacf Tiiir Barfh ( 

(327) 867-6230; or Dr. Gciia Bnkcr ( 

86 7-42 61. 

155/156 
, 

Process and Human Factors Engineering 

Technologies Roadmap 

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A Simulator Study of the KSC Shuttle Landing Facility Air Traffic Control 

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The design of the new Shuttle Landing Facility (SLF) tower was well under 

way when personnel needed assurance that known problems would be mitigated by the new design. The best way possible was to ”experience” the SLF 

from the new tower - a feat that could easily be accomplished by a high-fidelity, life-sized simulation. Fortunately, this technology was already in place at 

NASA’s Ames Research Center (ARC). KSC conducted the study in collaboration with Langley Research Center (LaRC) and ARC. 

.- 

.- 

2 

0 

U 

U 

c Modeling the KSC airspace required close interaction with the Master 

Planning Office to acquire specifications of land mass, latitude and longitude 

of the runway, existing structures, antennas, and other geographical mapping information. In addition, aerial photographs were taken of the terrain 

to model the topography of the airport and surrounding airspace. This task 

was accomplished by flying two helicopter missions over the SLF. Operations 

data were collected next. Through interviews with SLF operations personnel, 

documented procedures, and direct observations of landings, we gained a 

good understanding of the way business is done. Finally, all the data collected 

was modeled in FutureFlight Central - ARC’S 3-D Air Traffic Control (ATC) 

tower simulator. Its twelve 10-by-10-foot displays provided a 360-degree view 

of the SLF from the proposed tower, all before breaking ground. 

: 

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New SlF Air Traffic Control Tower 

158 

The goal was to perform a human factors assessment of the proposed tower design. There were two 

main areas of concern - the cab interior layout and the 

location of the tower. This study was conducted to 

ensure that pertinent ergonomics considerations were 

taken to reduce fatigue while maintaining engagement. Special attention was paid to reducing errors 

due to misperceived affordances. The chosen site 

was evaluated to ensure that the controller in the new 

tower would no longer encounter the glare from the 

xenon lights at the ends of the runway experienced at 

the old tower. The height of the tower was also varied 

in the study to verify that controllers had an unobstructed view of both ends of the runway. 

To date, all study objectives have been realized and 

construction of the new tower is ongoing. One of the 

biggest achievements is the opportunity SLF personnel had to operate in the tower environment and make 

recommendations before construction began. The 

study was completed in FY 2002. 

In the future, with the model of KSC now available, a training environment will be developed where 

controllers at the Dryden Flight Test Facility and other 

alternate landing sites can practice. Other landing 
Coiitrol Roniii Being Placed oiz Top of New SLF Air Paffic Coiifrol Tower 

by a Cruize 

operations such as convoy, security, and emergency 

medical will also benefit from this training environment. This model will be an invaluable resource 

to maintain the level of proficiency for all support 

personnel. 

Key accomplishments: 

Modeled the KSC airspace. 

Started a library of landing support aircraft and 

ground vehicles. 

Determined the optimum tower height for unobstructed views of runway, taxiway, and apron. 

Key milestones: 

Verified the adequacy of the tower design. 

Created the capability to simulate landing operations. 

Participa f iizg 0 rga i i iza f io i i s : NASA A i es Research Cei i fer (D r. 

B.G. Kunki aiid K.L. Clzrisfeizseii) aiid NASA Laiigle!y Research 

Ceizter (G. W. Lohr) 
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Assessing Space Shuttle Launch Dates With Advanced Discrete- Event 

Simulation Technologies 

This report documents the efforts to quantitatively assess the likelihood 

that the NODE-2 Space Shuttle mission (U.S. Core Complete) would launch 

on February 19,2004. The Shuttle Program sought a tool for assessing how 

proposed changes might affect the chances of meeting the NODE-2 launch 

date and for identifying ways to increase the probability of on-time launch. 

In response, a discrete-event simulation model was developed of the 

Space Shuttle manifest from STS-114 through STS-120 (NODE-2). The model 

was populated with planned processing times and the order of events from 

the baseline Space Shuttle manifest. Figure 1 shows the manifest that was 

modeled. Note that this was the manifest prior to the Columbia accident. 

Probability distributions for delays to processing activities and mission 

durations were developed using empirical data. Such data was also used 

to assign event probabilities to launch scrubs and diverted landings to 

California. The model was further populated with these distributions and 

probabilities, but the occurrence of a fleetwide grounding (for example, 

because of a hydrogen leak or Orbiter wiring problem) or a major mishap was 

not factored. 

Based on 1,000 replications of the discrete-event simulation model, it was 

determined that the NODE-2 mission had a .15 probability of launching on 

February 19,2004 (see figure 2). The 95-percent confidence interval for that 

date ranged from .134 to .165. As time progresses, the probability of launching 

increases, as one would expect. For example, by March 25,2004, there is 

greater than .8 probability that launch will have occurred. 

Designating holidays as available processing days increased the probability 

of launch to .245. It was also theorized that the launch probability could be 

further improved by shifting margins within the manifest. When 6 days of 

reserve were shifted from the Orbiter Processing Facility to the launch pad for 

each of the seven Shuttle processing flows, the launch probability for February 

19 increased to .31. 

The discrete-event simulation model is ready for analysis of future Space 

Shuttle manifests. 

Contacts: G.R. Cafes (Grant.R.Cates@nasa.,yov), PH-M3, (321) 861-9216; and A.L. Warren, 

l"-M3, (321) 861-9212 
Figure I. Manifest Used for Discrete-Eoent Simulation Modeling 

Kl 

0 

0 

0 

Baseline Arscsrm ent 

NODE-2 Launch Date Cumulative Percentaae 

100% 

90 % 

80% 

70% 

60% 

SO % 

40% 

30% 

20% 

10% 

Figure 2. One-Tliousand Replications of NODE-2 Mission Discrete-Event Simulation Modeling 

161 
.- 3 Virtual Reality Simulations of Orbiter Temporary-Access Improvements 

UJ 

162 

Various temporary platforms are used to gain access to the Orbiter and 

Orbiter Maneuvering Subsystem (OMS) pods during preflight processing 

through the Orbiter Processing Facility (OPF) and the Hypergolic 

Maintenance Facility (HMF). Temporary platforms consist of narrow, 

aluminum picboards installed by the High Crew using C-clamps and nylon 

roping. To improve preflight processing safety and efficiency, a crossfunctional team was established to identify picboard temporary installations 

occurring for every processing flow. United Space Alliance (USA) Industrial 

Engineering and Shop personnel evaluated the process by comparing factors 

such as risk reduction, frequency of installation / removal, ease of completion, 

and project cost. Requests were then prioritized to have the top seven 

“temporary picboards” replaced with permanent, portable platforms designed 

to improve safety and access to the Orbiter / OMS pods. 

To further enhance design efforts and aid future design reviews, NASA 

Industrial Engineering for Safety (IES) partnered with USA to simulate 

concepts in their respective processing environments. The virtual reality (VR) 

simulations were created by USA KSC Advanced Visualization Environment 

(KAVE) and Boeing KSC Design Visualization personnel using 3-D models 

generated by USA Design Engineering. The VR simulations provided the 

following benefits: 

Provided rapid prototyping without physical field modifications or models. 

Demonstrated feasibility of implementation. 

Provided easily grasped presentation of the concept to laymen (3-D versus 

Refined design based on input from individuals with vested interest in the 

2-D). 

actual use of finished hardware. 

Figure 1 shows the current temporary picboard setup for Engineering 

Support Request (ESR) K16997 (OMS Pod Leading Edge Access When Vertical 

at the 35-Foot Level) at the HMF east and west cells. The VR simulation 

depicted deployment, operation, and stowage of the conceptualized platform 

(figure 2). An avatar, a humanlike form, was also included to simulate and 

demonstrate a typical technician-to-platform and OMS pod relationship. 

The following ESRs were also selected for 

VR simulation: 

Figure 1. Temporary Picboard 

K17000, Elevon Fall Protection Barrier: 

Simulation included deployment and 

retraction motions of the proposed railing 

for OPF 3. An avatar was included in the 

simulation to demonstrate typical technician-to-railing relationship. 

K16964, Access to Engine Cavity 1: Simulation included installation and removal 

of the proposed access platform. An 

avatar was included in the simulation to 
Figure 2. VR Siiniilntioiz 

demonstrate typical technician-to-platform relationship. 

Access at the Y-Web Platform: Simulation 

suitable for OPFs 1 and 2 depicts deployment, operation, and stowage of the conceptualized platform. A 3-D model of the 

OMS pod and an avatar were included to 

show how a human operates and utilizes 

the conceptualized platform. 

K16994, Access Between the OMS Pod and 

Vertical Stabilizer: Simulation suitable 

for OPF 2 conceptualized platform was 

completed by Boeing. 

K16996, Access to Orbiter Forward Electrical Manifold in OPF 3: Simulation with 

an avatar included deployment, operation, stowage, and utilization of the conceptualized platform. 

K17001, OMS Pod Access to Flat Side 

When Vertical at 19-Foot Level (HMF): 

Simulation included deployment, operation, and stowage of conceptualized platform. 

K16993, Stinger /Up-Firing Thruster 

Key accomplishments: 

March 29,2002: Received approval for 

Change of Partnering Agreement (CPA). 

May 22,2002: Completed task list, platform requirements, and tools used. 

July 8,2002: Created 3-D models for VR 

simulation development. 

September 18,2002: Completed NASA 

deliverables (i.e., video and CDs). 

October 4,2002: Presented final report to 

IES. 

Key milestones: 

October 25,2002: Control Configuration 

Board Directive (CCBD) issued to complete design, fabrication, and installation 

of K16964 and K16996. 

December 12,- 2002: Desip csmpkted foz 

K16997. 

February 28,2003: Partnered fabrication/ 

installation of K16997 and design of 

K17001 with NASA. 

Participating Organizntions: KSA MA (J.M. 

Roxmki); PH-M3-A (D. B. Stanibolian and M.M. 

Groiz-Hainniond); United Spacc Alliance (M.A. 

Sortinan, E.]. Dereiner-Cook, 1.0. Moist, B.A. 

Lail?reiicc, D.C. Larseii, W.C. Maizley, C.J. Brozivz, 

B.X. Taylor, G.C. AtiLlood, B.C. Burns, K.P. Warren, 

P.A. Tlzoinas, G.A. Farmer, E.E. Brindley, L.E. 

Hmiltoii. J.T. Noble, T.R. 14%ite, D.]. Luth, R.A. 

Birflcr, I. Lnckascyck, PA. Peabody, E.G. Rieck, S.G 

Waclrosrski, and A.M. Rakofsky); and Boeing (R.P. 

Hiiiireiziuk) 

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Range Process Simulation Tool (RPST) 

Robust and practical tools to quantitatively assess the impact of technology 

insertion on operational performance of space transportation systems are 

not currently available. In particular, new methods and tools are needed 

to quantitatively assess the impact of planned technology modernization / 

technology insertion on the operational performance of current and future 

spaceports and ranges. 

This project designed, developed, and demonstrated the Range Process 

Simulation Tool (RPST), a decision support tool that facilitates quantitative 

change impact assessment. The technical approach involves defining change 

impact assessment requirements, designing and developing RPST software, 

developing baseline Range process models, defining focused test application 

scenarios, acquiring test application data, and testing and validating RPST. 

The RPST is intended for use by Range Planners and Range Technology 

Portfolio managers to quantitatively assess Range operational performance 

and assess the impact of Range technology changes / upgrades on operational 

performance. The RPST is an important first step toward deploying a realistic 

and robust Range operations change impact assessment capability. RPST 

also provides key enabling technology for modeling and analyzing future 

spaceports. The key RPST innovations are a novel, model-based approach 

for change impact assessment and a component-based, scaleable, openarchitecture approach that facilitates rapid tailorability and cost-effective 

deployment to varying spaceport application situations. The main RPST 

benefits are the ability to accurately and reliably predict impacts of technology 

changes on space transportation systems and the ability to reduce the Range 

operational and maintenance costs through improved change management 

decision support. 

Key accomplishments: 

2001 : Defined RPST requirements. Developed the preliminary RPST prototype. Identified RPST test scenarios. Defined a data collection and software 

testing approach. 

lection. Completed initial RPST testing and validation and Phase I1 Final 

Report. 

2002: Completed a scalable RPST prototype and test scenario data colKey milestones: 

2003: Continue RPST testing and validation. Transfer RPST to other NASA, 

Air Force, and contractor organizations. Generalize RET for commercial 

use. 

Contacts: T.S. Bartk (~j~~r~~~li1/,S.B~7~~lr~~~~~~~~~7,~~~~~~), YA-C, (321) 867-6230; G.D. Phillips, TAF-A, (321) 494-9167; and W.H. Haase, TA-F-A, (321) 494-9163 

Participating Organization: Knowledge Based Systems, Inc. (Dr. P.C. Benjamin and M. 

Grad) 
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Many current and future NASA programs have invested heavily in simulation models 

to identify cost and schedule drivers. However, despite safety goals that are orders-ofmagnitude more aggressive than cost performance goals, these programs have not invested 

heavily in simulation models to identify operations and maintenance safety drivers. Existence 

of a simulation modeling technology gap is part of the reason for the deficiency in safety 

modeling and optimization tools. 

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This research project explores the development of revolutionary new tools and technologies 

for macrolevel safety modeling and optimization. The goal is to develop a discrete-event 

simulation model that can be applied to any process to determine which tasks, systems, 

and maintenance actions have the greatest potential for causing personnel injury, hardware 

damage, and efficiency bottlenecks. With this information, a program can exercise control 

to mitigate these risk factors and thereby improve its operations. Developing a prototype 

tool will enhance the capability of preexisting (e.g., Shuttle, Atlas) or new (e.g., Orbital Space 

Plane) discrete-event simulation models already populated with typical task information 

(resource requirements and duration) by adding human factors and safety data. 

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The basic concept involves developing innovative analytical tools to help identify 

and quantify the critical, high-level safety drivers of a system, process, or product line. 

NASA / Kennedy Space Center’s dual mission of space launch operations and spaceport 

and range technology development uniquely positions KSC to perform this research. KSC 

has established technology development capabilities and technology core competencies in 

industrial engineering. KSC’s spacecraft processing facilities provide an ideal testbed for 

process and human factors engineering technology development. 

The research involves identifying relevant human factors safety data to accurately simulate 

the tasks affiliated with a process flow. The human factors safety data sample will include 

historical data (past health and safety incidents), human factors workforce safety data, 

human-error risk factors, etc. Developing the prototype tool will show proof of concept that 

human factors safety data can be combined with scheduling/ resource data to more accurately 

determine safety and efficiency bottlenecks affiliated with a process flow. 

Key accomplishments: 

2002: Basic research was completed on schedule and within budget to explore the concept of using discrete-event simulation to model and optimize macrolevel safety drivers in 

spacecraft processing. An initial prototype was developed and tested with KSC data. The 

research results are encouraging. 

Key milestones: 

2003: Additional refinement of the concept based on testing with current Shuttle, Expendable Launch Vehicle, and potential future Reuseable Launch Vehicle spacecraft processing 

flows. Based on the results of these tests, funding for developing this core enabling technology will be pursued from the Launch Services Program, Space Shuttle Program, and/or the 

Next-Generation Launch Technologies Program. 
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Participating Organizafioirs: Lockliced Marfiii Space Syfclizs Company (K. Bri.~glielli) md Floridn 

Ii?sfitiife of Tec111iolog~y (DI: D.S. Carstens) 

167 
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Because of the interrelationships between system maintainability and reliability, it is 

important to consider their integrated effects on system-, element-, and component-level 

design, development, test, and evaluation. Tradeoffs must be evaluated in establishing 

system reliability and maintainability requirements because of their impact on nonrecurring 

and recurring costs and the minimization of system life cycle cost. System dependability 

and responsiveness attributes are both heavily influenced by achieving an effective balance 

between reliability and maintainability design characteristics. These attributes then drive the 

safety, cost, and operation of the system. 

.- 

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(S), reliability (R), and maintainability (M) derives and integrates Level I requirements and the 

controls needed to obtain program key objectives for safety and recurring cost (see the figure). 

Being quantitative, the process conveniently uses common mathematical models. Even 

though the process is shown as being worked from the top down, it can also be worked from 

the bottom up. Two illustrations using this process are provided. 

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Number of 

Missions (n) 

Probability (A) of x or More 

Successes Out of n Missions 

Note: 1 -A = Program Risk 

Number of Successful 

Missions (x) 

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Via Math Model 

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Number of Serial Minimum Mission 

System Elements (e) Reliability (Ps) 

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for Any Element (Psi) 

Maximum Failure Rate 

for Any Element (Pfi) 

Note: 1 - Psi = Pfi 

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I f Maximum Number of 

System Elements Expecting 

Repair (r) per Mission 

.; 'I 

Repair Rate (h), Repairable 

Jse e When e = N 

Elements (N)' Use pfi When e = N 

Via Math Model 

Probability (M) of No More Than r 

Number of Repairs per Mission 

A Process for Developing and Balancing the Quantitative S, R, and M Requirements 
This process uses three math models: (1) the 

binomial distribution (greater-than-or-equal-to 

case), (2) reliability for a series system, and (3) the 

Poisson distribution (less-than-or-equal-to case). 

The zero-fail case for the binomial distribution 

approximates the commonly known exponential 

distribution. Either model can be used. The 

binomial distribution was selected for modeling 

flexibility because it conveniently addresses both 

the zero-fail and failure cases. The failure case is 

typically used for nonhuman spacecraft as with 

missiles. 

As the first step of the process, the Systems 

Engineering designer begins with three inputs: 

(1) the desired number of missions the program 

is planning (n); (2) the minimum number of 

successful missions for the duration of the program 

(x); and (3) the assurance (A) of obtaining x or more 

successes out of the n missions. In risk terms, 1 -A 

is the probability or likelihood of not obtaining x or 

more successes niut nf II r?~mber ef attempt; or not 

obtaining the desired level of safety and reliability 

over the life of the system’s program. When these 

three inputs are used in the binomial distribution, 

the minimum mission reliability (Ps) is calculated. 

At this point of the process, the Level I safety 

requirement has been established. 

The second step uses the minimum mission 

reliability (Ps) and an estimate of the number of 

serial line replaceable unit (LRU) elements (e) as 

inputs into the formula for reliability of a series 

svstem to calculate minimum element reliability 

(Psi). Maximum element failure rate (Pfi) is equal 

to 1 - Psi. Without considering the maintainability 

burden, which has a very large influence on 

recurring cost, including the system’s acquisition 

(fleet) size, the process at this point has established 

the safety and reliability requirements for the 

program. 

The last step addresses the maintainability 

parameter, the parameter that provides a control 

for recurring costs resulting from maintenance and 

repair. Similar to assurance or program reliability 

(A), program maintainability (M) is a probability. 

The probability M is determined bv the Poisson 

distribution and uses the following inputs: (1) the 

number of missions (n), (2) the number of elements 

(N, where e 5 N), (3) the LRU failure rate (Pfi or h, 

where h 5 Pfi), and (4) the maximum number of 

LRU repairs (r). Technically, M is the probability of 

no more than r number of repairs occurring 

at a particular mission using e number of 

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the enabling requirements to balance and 

i$ and the desired A, adjustments in e, Pfi, N, 

achieve the desired key objectives of the 2 

program. 0 

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Illustration 1 (e = N case): If A = 0.99 

for 100 successes out of 100 attempts is 

required by program management and 

the current design concept calls for 100 

serial systems (e = loo), then Pfi = 1 x 

will satisfy the Assurance requirement. 

Additionally, if N = e = 100 and Pfi = h 

= 1 x 

no more than one repair per mission is 

0.?????9??5. This, 3 ?V:ciintciinabiliky 

requirement desired at virtually any 

level for these management and system 

conditions is forecasted to be satisfied. 

then the probability of having 

Illustration 2 (e < N case): If A = 0.99 

for 100 successes out of 100 attempts 

is required by program management 

and the current design concept uses 100 

serial systems (e = loo), then Pfi = 1 x 

will satisfy the Assurance requirement. 

L4dditionally, if each of the 100 serial 

systems contains an average of 1,000 

subelements and each of the 100,000 

subelements (N = 1,000 x e = 1,000 x 100 = 

100,000) has an average repair rate of i, = 

1 x lO-j, then the probability of having no 

more than one repair per mission is 0.7356 

or about 74 percent. In other words, this 

design concept under a 99-percent level of 

assurance indicates there is a 26-percent 

chance of having two or more (up to N) 

repairs for each of the 100 missions. Thus, 

if the Maintainability requirement was 

targeted to be no less than 90 percent, the 

Maintainability requirement is forecasted 

not to be satisfied. Adjustments in the 

design parameters are required. 

169 
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Human Engineering in Design and Modifications 

Space Shuttle processing at KSC is envisioned to continue during the next 20 years and 

beyond in support of International Space Station activities, satellite repair and/or launches, 

and other key NASA objectives. During this timeframe, the Space Shuttle and ground support 

systems and hardware will continue to age. In an effort to maintain operational status, new 

engineering designs will be required to maintain, modify, upgrade, and incorporate changes. 

These design changes require the integration of human engineering principles as described 

in MIL-STD-1472, Department of Defense Design Criteria Standard Human Engineering, the 

Department of Transportation/ Federal Aviation Administration (DOT/ FAA) Human Factors 

Design Guide, and other related standards and guidelines. 

Results of KSC research and human factors investigations demonstrate the occurrence 

of undesirable events stemming from human-system interactions. These incidents were 

attributed to human factors issues such as industrial design and workplace design. The 

combination of these two contributing cause categories represents 28 percent of all 

contributing causes identified in Shuttle Processing human factors investigations. These 

contributing causes identified systemic issues that included misapplication or omission of 

basic human engineering principles identified in accepted standards and guidelines. 

To ensure the human engineering principles are implemented by the design entities, 

training to understand the requirements in these standards is required. The primary training 

tasks include (1) a human factors awareness class, (2) use of a human factors design review 

checklist, (3) use of a human factors design quick-reference application guideline, and (4) 

a training module on the application of human engineering principles into the KSC design 

processes. 

The training module consists of the following four main objectives: 

1. Motivate the design community to consider human engineering principles early in 

the design development and modification stage. (The figures illustrate the practical 

improvements that can be realized when deliberate and early focus is placed on human 

engineering principles.) 

and detailed scientific texts and scholarly research into human engineering principles. 

2. Serve as an intermediate step or bridge between introductory human engineering principles 

3. Serve as an understandable source of human factors design information. 

4. Provide current information on human-oriented design requirements. 

Key accomplishments (2001 - 2003): 

Developed human factors and MIL-STD-1472 awareness overview. 

Developed the human factors design review checklist and quick-reference guideline. 

Developed the detailed MIL-STD-1472 training module concept. 

Delivered the human factors awareness overview and MIL-STD-1472 familiarization. 

Key milestones: 

Complete development of detailed training course material on application of MIL-STDDeliver course to designers, design engineers, and ”Train the Trainer.” 

1472. 
Lifl Configuration - 1" Vertical Lip at Ramp 

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171 
172 

Task Analysis Technologies 

This research effort focused on conducting task analyses on potential 

bottlenecks and hazardous processes identified at KSC. A task analysis is 

performed to identify specific improvement opportunities to mitigate or 

eliminate the risks associated with the potential bottlenecks. Task analysis is 

a generic term including work methods analysis, process failure modes and 

effects analysis (PFMEA), and other tools and techniques. The benefits of task 

analysis to spaceport operations on safety and cost performance have been 

demonstrated on specific tasks, and KSC seeks to perform many additional 

detailed task analyses. However, the relative complexity and infrequency of 

spaceport processing tasks make existing task analysis technologies expensive 

and time-consuming, require extensive training to adapt to the spaceport 

environment, and limit their use. This research produced a task analysis tool 

that enables a human factors analyst to perform a task analysis in days versus 

weeks because of the automated human-error taxonomies. 

The steps involved in the Task Analysis Technologies project are as follows: 

Identification of hazardous tasks performed at KSC: Two current 

hazardous operations and two hazardous operations that will exist in nearterm future operations were selected for analysis. Hazardous tasks are 

classified as any tasks that have resulted in close calls or actual incidents 

affecting (safety of personnel, collateral damage, etc.) or future tasks that 

have the potential to result in close calls or actual incidents. 

Task analysis training: Task analysis training was given to the team 

members by NASA Headquarters to ensure consistency. 

Performance of task analyses and development of recommendations for 

improvement: Different task analysis methodologies were performed 

based on the task being analyzed. A PFMEA was performed on the current 

operation of the Solid Rocket Booster blast shield installation (figure 1). A 

PFMEA is a procedure used to analyze the system’s processes rather than a 

specific piece of equipment (NASA Human Factors [HFI-PFMEA Training, 

2002). An HF-PFMEA was performed on the current operation of the Cargo 

Late-Access Task (figure 2). An HF-PFMEA is a systematic method to 

analyze each task in a process and identify human errors (failures) and the 

worst-case effect on the system (NASA HF-PFMEA Training, 2002). Both 

of the current operations analyses resulted in several recommendations to 

mitigate human error in task performance, enhancing the safety of both 

technicians and hardware. An HF-PFMEA was also performed for the 

future operation of the Advanced Technology Development Center LOX 

Pump Acceptance Test Procedure. This enabled the team to design the 

work procedure for the LOX Pump Acceptance Test with safety and human 

error in mind. A general task analysis was performed for the Payload Test 

and Verification System that resulted in the creation of the ideal process 

flow for future payloads that mitigates human-error-related risk factors and 

eliminates bottlenecks associated with current payload processing. 

Identification of technology barriers and challenges to performing better 

task analyses at KSC: It is important to identify current task analysis 

barriers and challenges to enable researchers to identify current technology 

gaps in this area, therebv revealing additional research opportunities. This 
segment of the research focused on addressing 

such issues as how we become more efficient 

and effective in performing task analyses. 

Task Analysis Tool Development: The prototype 

tool (figure 3) was developed by automating 

the manual HF-PFMEA process, allowing an 

analyst to quickly identify actual improvement 

recommendations. This enables an analysis 

to be performed within days versus perhaps 

months, depending on the size of the task being 

analyzed. 

Key accomplishments: 

2002: Identified hazardous tasks performed at 

KSC, received task analysis training, conducted 

task analyses, developed recommendations 

for improvement, and identified technology 

barriers and challenges to performing better task 

analyses at KSC. Completed Task Analysis Tool 

and reports. 

Key milestones: 

W 

Figure 1. Technician Working on the 500- 

P01171d Solid Rocket Booster Blast Shield 

Liffiiig and bistnllation Process 

. 2003: Conduct additional task analyses on KSC 

operations. Make task analysis tool available to 

end users. 

Con facts: M.M. Groii-Hammond (;I .In!~-in.M. G vbHL? 111 vL1 T! .> 

1I~?~i?.~~~T'), PH-M3-A, (321) 867-0572; and T.S. Bartii, YAC, (321) 867-6230 

Participating Organizations: NASA YA, PH, aizd QA; 

Bociizg; Uirifrd Space Alliaizcc; Florida Iiisfitiitc of 

Teci7n010gy (Dr. D.S. Carstens); nizd USTDC 

Figirre 2. Cargo Late-Access Task Wllere 

Cargo Is Placed on the Payload at the 

SIriittle Latiiich Pad 

Figure 3. Task Aiialysis Protot!ype 

173 
Human Factors-Process Failure Modes and Effects Analysis (HF-PFMEA) ln 

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The goals of HF-PFMEA are to identify factors that contribute to human 

error and either eliminate them or control them so their effects are manageable 

or acceptable. To that end, a new technology was sought to enable quick, 

user-friendly, and affordable HF-PFMEA; task-level analysis focusing on 

safety and performance issues; and the capability to analyze human factors 

issues such as health and safety risks. The HF-PFMEA Software Tool was 

developed to systematically analyze each task in a process to identify potential 

human errors, their respective worst-case effects on a system, and the factors 

that increase the likelihood of the human error. The HF-PFMEA Software Tool 

helps the user identify: 

Potential individual or team human errors. 

Factors contributing to or affecting the potential for human error 

occurrence. 

Barriers to prevent errors or inhibit the effect of errors. 

Risks associated with the human errors. 

Recommendations to reduce errors or mitigate their effects. 

The tool consists of modules that allow the user to create a functional 

analysis and a hierarchical task analysis. With this information, the software 

then generates a dataset that interactively relates task action data to potential 

human errors, factors that can affect those errors, and barriers and controls 

that can prevent the errors or mitigate their effects. The database functions 

incorporated into the software develop HF-PFMEAs faster than the timeconsuming manual method and create, print, and store reports, including 

HF-PFMEA tables that contain most of the analysis data. In addition, the 

tool provides the user a means to perform a risk assessment for the potential 

errors and their effects by determining the severity of the errors identified, the 

likelihood of their occurrence, and the probability of error detection prior to 

consequence. The software even provides suggestions for the development of 

recommendations to eliminate or reduce the risk of errors. 

The HF-PFMEA Software Tool was developed from specific HF-PFMEA 

requirements and HF graphical user interface requirements. The code is 

written in Visual Basic and constructed with a program called J-Builder. 

The tool has three main interrelated components: analysis, task tree, and 

reporting. The application is very dynamic. All data entered and selected by 

the user is automatically saved and may be edited at any time. It allows users 

to build their own lists of data and can accept new missions, functions, tasks, 

subtasks (to any level), errors, performance-shaping factors, and barriers 

when necessary. 

The program itself is written in Java, runs on a PC under the Windows 

operating environment (98, NT, or 2000), and can be accessed from the 

Internet. The program is compatible for data exchange with Microsoft Word, 

Excel, and Access. 
The HF-PFMEA Software Tool is flexible enough 

to analyze a variety of operations, including 

maintenance. It can also be used to collect data 

on maintenance processing scenarios from a crew 

implementation and human factors perspective to 

enhance human performance and well-being. By 

doing so, the HF-PFMEA Software Tool promotes 

increased human capabilities to perform a variety 

of tasks while maintaining effectiveness and 

physical health. 

Coiztacts: M.M. Groh-Hammond ( ' I ,- 1 

QA-C, (321) 867-4261 

* -' , ), PH-M3-A, (321) 861-0572; and Dr. G.H. Baker, 

Participatiilg Organizations: The Boeing Conzpany (M. 

Philippart, Dr. K. Relvini, W.D. Valeiztino, and D.E Baska) 

and Florida Institute of Technology (Dr. D.S. Carstens) 

mu I' 

HF-PFMEA Graphical User lnterface 
.- G Task Analysis of the Payload Test and Checkout System (PTCS) 

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In response to a request from the International Space Station Utilization 

Division, the KSC Process and Human Factors Engineering Working Group 

(PHFEWG) and The Boeing Company Human Factors / Industrial Engineers 

(HF/IE) performed a complete task analysis of the Payload Test and Checkout 

System (PTCS). The purpose was to use the insight gained to develop a 

Payload Test and Verification System (PTVS) for payload developers (PDs) 

and investigation developers (IDS) to use to test and check out their racks 

prior to delivery to KSC. Knowing how the racks are processed at KSC 

will help PDs develop racks that integrate more smoothly into the handling 

process. The working group selected the Microgravity Science Glovebox 

(MSG) rack as the test article and analyzed its progress from conception of 

requirements and documentation to testing and associated experiments in the 

PTCS at KSC. 

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The following seven steps were used to develop the MSG PTCS task 

analysis flow: m 

V 

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2 1. 

2. 

3. 

4. 

5. 

6. 

7. 

Collect and review documentation: general documents, mission-specific 

documents, lessons learned, mishap reports, and questionnaires and 

interviews with process owners. 

Compile tasks and constraints: categorize in three areas - rack planning 

activities prior to delivery at KSC, customer off-line activities at KSC 

delivery, and on-line activities after delivery. 

Design task analysis flow storyboard. 

Convert storyboard to computer model. 

Review automated task analysis flow. 

Revise task analysis flow. 

Complete deliverables: recommendations, lessons learned, and final report. 

The MSG PTCS Task Analysis Flow identified potential and existing 

bottlenecks, interdependencies between processes and documents, and gaps 

in the process. The flow describes each task and document and illustrates its 

potential interaction with other processes by color and pattern. It clarified 

the complete process and identified the constraints that each participant 

represents in the process and the steps the hardware needs to follow for 

efficient end-to-end processing. 

The PTCS task analysis of the MSG rack yielded the following 

recommendations that have broad application for rack processing: 

Require PDs to submit electronic drawings to KSC so equipment 

interference issues can be identified and prevented. 

Develop standard sets of common test equipment capable of supporting all 

types of racks to use when PD missions slip and the originally earmarked 

ground support equipment is obligated to other missions. 

Involve Human Factors / Industrial Engineering specialists at the beginning 

of design and throughout procedural reviews of processes requiring 

human-machine interface. 

176 
Micrograuify Science Glovebox Payload 

Bridge documentation gaps to increase process understanding, eliminate redundant 

tests, track completion of tasks, and mitigate undocumented open items after turnover. 

Work toward an electronic hardware transfer log to reflect real-time rack configuration. 

Achieve more comprehensive and frequent communication between PDs /IDS and KSC. 

Refine definitions of effects on KSC integration and constraints on KSC operations and 

strengthen requirements for shipping instructions and for problem reports on latestowage hardware to reduce open verification items after turnover. 

The product of this endeavor was a Payload Test and Verification System that enables 

PDs and IDS to design and configure racks that integrate better into the Payload Test and 

Checkout System at KSC. The task flow analysis yielded a tool that can serve as a template 

for rack processors performing PTCS operations. 

Contact: Dr. G.H. Baker (Gc~a.H.Rak~~~ci~ i7ml7,p;l), QA-C, (321) 867-4261 

Participatiilg Organization: Tlie Boeiizg Coinpaizy (Dr. K. Reloini and J.G. Posada) and Florida Institute of 

Technology (Dr. D.S. Carstens) 

177 
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Toolkit for Enabling Adaptive Modeling and Simulation (TEAMS) 

As systems have become more complex, using quantitative operations 

analysis decision support tools, such as simulation, scheduling, activity-based 

costing, and optimization, has become increasingly important. These decision 

support tools, however, have several technical and pragmatic limitations. For 

example, systems simulation practice is afforded little automated support 

for the initial analysis, problem solving, and design tasks, which are largely 

qualitative in nature, and uses too much of both the domain expert’s and the 

simulation analyst’s time. It also requires significant investment of time and 

money to deploy and maintain simulations over extended periods of time and 

is not widely accepted by decisionmakers. Likewise, there is limited support 

for activity-based cost (ABC) model development and model maintenance. 

Similar problems exist for optimization modeling and scheduling. TEAMS 

targets these technical challenges associated with operations analysis, 

modeling, simulation, and cost analysis and will provide current and future 

spaceport designers with a knowledge-based infrastructure for quickly and 

easily developing, maintaining, and reconfiguring operations analysis models. 

This project will design, develop, and demonstrate TEAMS, a decision 

support tool that promotes the quantitative process analysis of space 

transportation operations and maintenance. The technical approach involves 

the following tasks: (1) identify focused TEAMS application areas, (2) 

develop TEAMS applications, (3) test and validate TEAMS, and (4) generalize 

TEAMS for commercialization and technology transfer. Anticipated TEAMS 

products include a validated TEAMS software application, a reconfigurable 

spaceport operations modeling framework, and a scalable operations analysis 

knowledge repository. Key innovations include (1) a process-centered 

approach maximizing reuse of domain knowledge for rapid operations 

analysis model development; (2) open architecture and distributed plug-andplay architecture allowing for mass customization and rapid deployment; and 

(3) novel, simulation-based optimization mechanisms facilitating affordable 

risk minimization through exploration of a large number of spaceport designs. 

Key accomplishments: 

2001: Start of Phase I project and definition of TEAMS requirements. 

2002: Completion of TEAMS prototype software and start of Phase I1 

project. 

Key milestones: 

2003: Enhancement of TEAMS framework, refinement and hardening of 

TEAMS software, selection of KSC application areas, and development of 

KSC applications. 

2004: Refinement of TEAMS framework, completion of TEAMS 

applications, and technology generalization for commercial use. 

Contact: G.R. Rhodeside ( . 9 5 , ), QA-C, (321)867-7910 

Participating Organization: Knowledge Based Systems, Inc. (Dr. P.C. Benjamin) 
Spaceport Operations Needs Mission Constraints 

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Activity- Based Costing and Simulation for the Operational Assessment of 

Future Space Transportation Systems 

NASA’s long-term goals include dramatically reducing the cost of space 

transportation. Achieving this goal requires innovative technologies and new 

approaches to vehicle design, development, manufacturing, and operations. 

Developing models that analyze all life cycle activities for space transportation 

system concepts, from development to operations, is one of the needed 

innovations. 

Operations models (ground operations or spaceport operations) are an 

important part of the assessment of new reusable launch system architectures. 

Models are required to estimate significant portions of the system’s life cycle 

cost, such as recurring and nonrecurring operations costs. Overall ground 

cycle time is also an important parameter to estimate because this will 

determine the system’s yearly flight-rate capability per set of assets (flight 

and ground elements). This operational factor of turn-time will affect other 

up-front, nonrecurring costs, such as the number of vehicles, facilities, and 

equipment to acquire to meet a desired flight rate per year. Thus, downstream 

operations that seem far away actually affect up-front and immediate costs as 

well. The recurring costs and the flight rate are the result of tasks or activities 

that are required during ground operation, for example, the preparation of a 

payload for integration with the vehicle. Typically, the cost and task duration 

assessment of these processes is performed by experienced engineers who 

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lnput Srrwnfor Eqtimating TPS 1,ahor Bawd on Surh Drin~rq a5 Arwap 1 oration, 

Material, Accessllnterfaces, and Design Approach 

employ their knowledge of production and 

operations technology, methods analysis, and 

engineering economics to predict the probable 

cost and production time of a product - in this 

case a ground operation activity. 

This research focuses on developing and 

evaluating new techniques for automating 

the operations assessment of future space 

transportation systems by using a combination 

of activity-based costing and simulation 

modeling. The approach translates vehicle 

design parameters into a set of activities 

and a related process mapped in a domain 

(operations characteristics of future space 

vehicle concepts) where there is limited 

knowledge. This approach is innovative 

because it will, for the first time, combine 

activity-based cost modeling (which is known 

to work well in well-defined environments) 

with expert knowledge to estimate the 

activities, cost, and time characterizations 

associated with proposed space transportation 

concepts. A critical element of this research is 

gathering data and expert opinions to develop 

“knowledge engines” for major vehicle 

subsystems. The knowledge engines 

capture existing and future approaches 

to a vehicle subsystem and link them 

to an activity set, time, cost, failure 

characterizations, and process map. 

Key accomplishments: 

Created the initial framework for 

reusable launch system knowledge 

engines. 

Developed a preliminary knowledge 

engine (a spreadsheet, with a graphic 

user interface) for estimating operations 

factors, such as labor hours, for the 

thermal protection systems of future 

reusable launch systems. 

activity-based cost modeling for future 

reusable launch systems. 

Documented the results of using 

Participating Organization: Blue Frog 

Technologies, lnc. (Dr. A. Ruiz-Torres) 
.- M Artificial Intelligence Techniques for Payload and Vehicle 

Processing Scheduling 0 

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Preparing vehicles and payloads for launch is an extremely complex process 

involving thousands of operations for each mission. Each operation requires a 

variety of equipment and workspaces. Since several missions are in preparation 

simultaneously, they all compete for scarce resources. Furthermore, since many of 

these resources are extremely expensive and limited in number (often operating 

at or beyond their capacity), optimal assignment and efficient use are critically 

important. Ground rules, safety requirements, and the unique needs of processing 

vehicles and payloads destined for space impose additional constraints. These 

challenges are compounded by the endless changes to the schedule caused by late 

deliveries, delayed flights, and malfunctioning equipment. To resolve the many 

conflicts and to predict possible problem areas, operators must use a number of 

rules of thumb specifying where things should happen, whether they will happen 

on time, and whether the requested resources are actually necessary. 

Although there are a number of commercially available scheduling systems, 

the degree of domain knowledge required for decisions and the unusual set of 

constraints make these of limited use. Stottler Henke solved these problems 

by applying a combination of artificial intelligence (AI) techniques to produce 

Aurora, a system capable of rapidly completing a near-optimal schedule. Aurora 

is unusual in that it combines sophisticated scheduling mechanisms with domain 

knowledge to solve the scheduling problem. Aurora’s initial scheduling domain 

was Space Station Processing Facility (SSPF) scheduling support. This domain 

includes a number of problems unique to KSC, such as the need to schedule 

floorspace and maintain certain spatial relationships among the tasks and 

components. This complexity proved a good test of Aurora’s scheduling potential. 

Aurora Flozo Editirzg mid DisplaJy Window 

182 
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Aurora allows a range of informational 

access, graphically displaying resource usage, 

floorspace usage, and the spatial relationships 

among different activities. It also permits the 

user to create reports that can be analyzed in 

Excel. Scheduling experts can interactively 

modifv and update the schedule and can 

request more information about specific 

scheduling decisions. This allows them to 

verify the system's decisions and override 

them to resolve novel conflicts. 

Stottler Henke has taken advantage of 

Aurora's built-in flexibility to apply it to such 

disparate problems as Orbiter scheduling for 

USA and classroom and teacher scheduling 

for GE Medical Systems. The version of 

Aurora in use by KSC will be able to respond 

to changing scheduling needs in coming 

years by tracking personnel requirements and 

temporal relations, in addition to the more 

static resource requirements that are currently 

of central concern. 

Key accomplishments and milestones: 

June 2001: Proof-of-concept prototype 

completed. 

April 2002: Released a scheduling 

system that provides support for spatial 

requirements and constraints as well as 

traditional scheduling needs for SSPF 

floorspace scheduling. 

December 2002: Released a refined 

version of Aurora with expanded support 

for the SSPF scheduling domain. 

March 2003: Completed development of 

temporally oriented Aurora for Orbiter 

scheduling. 

May 2003: Delivered completed Aurora 

scheduling system, featuring spatial and 

standard resource displays, temporal and 

resource-based scheduling, and import / 

export capabilities for interfacing with 

other planning systems. 
.- G Design Root-Cause Analysis Knowledge Base 

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In pursuing NASA's goal to advance space transportation for anyone, anytime, and 

anywhere, knowledge of specific design shortfalls is greatly needed to more effectively 

develop launch vehicles and advanced spaceports. Knowing where the processing time is 

spent and where the operational infrastructure costs are incurred is a critical need in modeling 

the operations and the infrastructure of future concepts for both flight and ground systems. 

Acquiring information on specific operations functions and infrastructure elements will help 

in developing higher-fidelity operations models, as well as more effective flight and ground 

systems designs that are also more affordable, more responsive, and safer. 

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The Next-Generation Launch Technologies (NGLT) program is investing in the 

development of a design root-cause knowledge base to help meet this need. The design RootCause Analysis (RCA) Knowledge Base is made up of thousands of Space Shuttle maintenance 

tasks organized into three generic operations classes: operations modules, functions, and 

subfunctions. Further, the maintenance tasks are classified by the generic design discipline 

(flight and ground) that generates the work. In this manner, top-level knowledge of the 

experience gained by the Shuttle Program can be extracted to determine where the overall 

work content is concentrated. For example, the overall amount of unplanned troubleshooting 

and repair activity, versus servicing, versus assembly and handling, versus inspections and 

testing, can be determined and quantified. The RCA Knowledge Base then expands on this 

information by organizing design cause statements associated with why a specific operations 

subfunction exists, or why it takes a certain amount of time to perform. Currently (April 2003) 

the RCA Knowledge Base contains over 200 such cause statements. Associated with each 

design cause statement is at least one operational design need statement. It is expected that 

designers and technologists will quickly discover high-priority needs and more effectively 

develop new technologies and systems for the space transportation operator from a database 

of such statements. 

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Ongoing research and development include the following: 

RCA Knowledge Base support to development of the Schedule and Activity Generator / 

Estimator (SAGe). Specifically, identification of key system characteristics for SAGe concept 

definition and input, as well as report generation. 

architecture characteristics effort. 

RCA Knowledge Base support to a Design for Operations (D40ps) space transportation 

Key accomplishments: 

Captured actual operations for eight Space Shuttle on-vehicle direct maintenance actions. 

Created an operations Functional Breakdown Structure (FBS) to three levels. 

Defined a set of generic design disciplines. 

Assigned a set of operations functions, subfunctions, and design disciplines for each 

maintenance task in a complete Space Shuttle processing flow (Space Shuttle Atlantis /STS81). 

Developed algorithms for automatically mapping STS-81 applied knowledge to seven other 

processing flows. 

Analyzed and extracted information from the STS-81 analysis for the NGLT program and 

ongoing technology work in the field of space transportation operations modeling. 
Key milestones: 

Completion of RCA Knowledge Base design. 

Completion of report generation capability. 

Engineering release of RCA Knowledge Base 

to the Space Transportation Information 

Network (STIN) for use by NASA and its NGLT 

contractors. 

Participating Organizafioiis: Locklzeed Martin Space 

Systeins (K.A. Ingoldsby) mid USTDC (D.E. loliizson, Jr.) 

Design Root-Cause Aiialysis Knowledge Base 

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Project Management Integration System (PMIS) 

A commonly lacking yet critical component of successful application of project 

management concepts is the seamless integration of project management tools being used 

throughout the organization. Development and implementation of an integrated framework 

and system architecture for project management tools will result in optimized planning, cost 

estimating, risk management, and project control. NASA and KSC have designed a thorough 

system of well-defined policies and procedures based on traditional project management 

techniques. Investigation and analysis of project management strategies revealed that many 

different software applications are being used as project management tools. However, 

there was no integrated on-line system to assist project managers to efficiently use these 

applications to complete compliance requirements. Such a system would allow project teams 

to collaborate on project responsibilities while permitting project managers to spend more 

time managing project requirements rather than continually learning new project management 

tools. Therefore, the purpose of this work was to develop a framework for integrating project 

management tools used at KSC so they might be easily accessed and efficiently used from a 

central system architecture, administered through KSC’s Systems Management Office (SMO). 

An additional goal of the project management framework was to reveal proven principles that 

establish the basis for project management success while integrating those principles to form a 

reliable system. 

Prior to deciding on a particular project management tool, organizations must carefully 

evaluate the reliability of internal policies, procedures, and practices. Thus, initial 

developmental tasks for this architecture included investigating applicable policies, 

procedures, tools, and techniques. While many project management tools and techniques 

are integrated into project management software applications, many are also developed 

using other software such as Microsoft Access, Excel, Outlook, PowerPoint, and Word. 

The advantage of project management software is that it generally provides tools such as 

work breakdown structure (WBS), task list, Gantt chart, critical path analysis, and earned 

value analysis that help project managers and project team members accomplish scope, 

time, cost, and quality management. However, according to leading project management 

experts, quality is actually the fundamental goal of project success in accomplishing efficient 

integration of cost and schedule. Investigation of quality issues revealed that KSC SMO uses a 

diverse network of rationale-based metrics with specific goals developed around its mission to 

control quality. 

Project management methodologies commonly used by industry experts were used to 

design the proposed interactive system. A well-integrated project management system 

with quick and easy Internet access was one of the primary objectives of PMIS. We used 

the experience of industry and academia to create an architecture for integrating the project 

management software applications in use at KSC. Research and analysis revealed that project 

managers at NASA KSC need the capability to (1) use project management software and 

systems without being an expert in the framework of the project management system and 

how it works, (2) access the project management system at any time from any place while 

maintaining organizational security, and (3) instantaneously and automatically notify team 

members that task requirements or action items have changed. 

Furthermore, the desired architecture should eliminate redundancy by automating 

project team tasks to allow the transfer of data from forms, templates, and other required 

documentation into a knowledge base for integrated and distributed project management. 

Review of both Government and private-sector experiences resulted in a list of desirable 
features for integrated project management in a 

Web-based environment that includes (1) simple, 

powerful processes, (2) form-based interfaces 

for easy connection to existing data swrces and 

input /modification by nonproject management 

personnel, (3) well-organized, consistent data, (4) 

quick-scan dashboards, (5) in-depth reports, and 

(6) consistent and reliable security. 

For large, complex organizations such as 

KSC, a system based on distributed project 

management concepts is the most feasible of 

the methods reviewed. Distributed project 

management offers the most versatile and effective 

solution to integration and automation for project 

management. This technology, while continuing to 

advance as a result of research and development in 

lraried u7ork environments, has shown significant 

benefits. 

The proposed PMIS may be deploved using 

a Web-based graphical interface to integrate 

commerciai off-the-sheii (COTS) software, 

application service provider (ASP) software, 

custom-designed software, or a combination 

of these. The figure shows the proposed 

architecture of such an application. Review 

of the literature shows that Microsoft 

Project’s custom programming language, 

Visual Basic for Applications (VBA), can be 

used to automate recurrent tasks and interact 

with other Windows applications. Java, 

XML, and VBA programming tools combine 

effectively to integrate these components 

into an interactive, on-line, distributed 

project management system. Unfortunately, 

research revealed that one of the most 

prevalent problems faced by organizations 

attempting to institute more efficient systems 

is the lack of implementation. Development 

of a KSC-wide project group to complete a 

prototype based on the PMIS architecture 

will facilitate deployment of the proposed 

solution. 

Coirfnct: T.E. Fredricksoii (~nce!i.E.Fii,iIi.iCX~(111 

(il ~insn.goi~), QA-C, (321) 867-2770 

Pnrficipntiirg Orpiiznfioir: Miles College (C.D. 

Birtlfr) 

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Distributed PAT End Users 

Portfolio Analysis Tool (PAT) 

PAT Functions 

A key characteristic distinguishing successful enterprises of the 

next millennium will be the ability to respond quickly, proactively, and 

aggressively to unpredictable change. Robust methods and tools to facilitate 

change management are therefore essential for the modern enterprise. This 

project targets the following key technology gaps and challenges associated 

with strategic change management: 

Lack of scientific methods for strategic change management. 

Lack of information-integrated software tools to support the change 

Lack of knowledge management mechanisms that capture, store, and 

management process. 

leverage corporate knowledge for strategic planning and control. 

This project developed and demonstrated a Portfolio Analysis Tool (PAT), 

a Web-based software tool that facilitates strategic change management. Key 

functional requirements addressed by PAT include the following: 

Provide technology investment decision information: PAT provides 

Facilitate identification of technology gaps: A direct result of the portfolio 

Promote technology roadmap development: The portfolio gap analysis 

decision-making information to facilitate strategic planning. 

analysis is to assist the precise identification of technology gaps. 

results provide information to construct a time-phased technology 

development approach / roadmap. 

Facilitate decision making over distributed teams: PAT facilitates the 

collaborative, distributed decision-making processes associated with 

strategic planning and change management. 

Intended PAT end users are strategic planners and technology portfolio 

managers. The PAT architecture includes a database server that provides 

storage of the information managed by PAT. PAT also provides a Web-based 

user interface that facilitates distributed capture of key information required 

to perform portfolio analysis. A central function of PAT is to facilitate sound 

project / technology investment decisionmaking as part of strategic change 

management. Different types of outputs and visualizations are generated 

by PAT to accomplish this function, including (1) Risk-Return Portfolio 

Analysis, (2) Gap Analysis, (3) Sensitivity Analysis, (4) Roadmap Generation, 

Portfolio Analysis Tool Functional Architecture 

188 
5 

High Risk. 

Low Return H@ <- Risk -> Lo*r LorrRislr. 

Low Return 

Example Risk Returti Aizalysis Output 

and (4) On-Demand Reports. Key innovations include novel 

adaptation and application of portfolio theory for strategic 

decision support and a structured, knowledge-based method 

for strategic planning and change management. 

Key accomplishments: 

2001: Developed a prototype Web-based portfolio analysis 

tool. Completed change management method and initial test 

data collection and tool testing. 

2002: Completed Phase I1 Final Report and PAT testing and 

validation. Completed PAT deployment at KSC and enduser training. 

Key milestones: 

2003: Expand PAT user base at KSC. Develop spreadsheet 

representations for Technology roadmapping. Refine PAT 

- roadmap linkages. Transfer technology to other NASA 

organizations and the U.S. Air Force. 

Coiztacts: T.S. Bartk (T~iiioriii/ S P17i iiw iir7~7 +': ), YA-C, (321) 867-6230; 

aiid E.Z. Zapata, YA-D4-A, (321) 867-6234 

Parficipafing Orgaizizatioiz: Kiioiidedge Based Systems, lilt. (Dr. P.C. 

Berzjamiiz) 
Biological Sciences 

The Biological Sciences program at KSC conducts research and technology development in 

four basic life sciences areas: 

Bioregenerative life support in closed biological systems. 

Metabolic response of plants to the space environment. 

Plant flight hardware development, design, construction, and validation. 

Application of results from this research and technology development to the management 

of agricultural and natural ecosystems on Earth. 

Efforts in bioregenerative life support systems research for over 10 years have focused 

~n cr=p gr~wth and deve!~pment vzith the pri~ar:. emphasis ~n bi~mzss pr~ductix under 

closed conditions. In recent years, the focus has shifted toward the potential use of plants in 

recycling solid and liquid waste. Another goal for these studies was to determine the stability 

of microbial communities associated with the plant and recycling svstems. The Biological 

Sciences program initiated significant efforts in understanding photosynthetic processes 

under spaceflight conditions. A series of studies was designed and initiated to determine the 

response of higher plants to microgravity and to determine if they would remain viable and 

useful components of a life support system. These experiments required adding new plant 

flight hardware and refurbishing older flight hardware that had been used for previous plant 

experiments. Plant flight hardware technology development efforts include investigating the 

use of light-emitting diodes and their effects on plant growth, using porous tube svstems to 

deliver water and nutrients to the roots of plants, and developing both physical and biologica 

sensors that could give a more complete understanding of the hardware operation and 

environment during spaceflight testing. 

Studies to determine the stability of microbial communities with physiological profiling 

continued during the past year, and the use of molecular approaches to characterize these 

communities was expanded. Molecular and biological technologies that assist in determining 

the stability and functional; ty of microbial communities are essential for managing biological 

systems during long-duration spaceflights. The ultimate goal is to test the capacity of 

microorganisms to recycle solid and liquid wastes, including graywater and trash. The 

successful demonstration of bioregenerative life support technologies and the ability to 

support them in the spaceflight environment ill permit NASA to estimate the resources 

needed to design, construct, and test life support systems for future missions. Ultimatelv, 

biological life support systems will be required to sustain autonomous human colonies in 

space. Understanding and optimizing both the biomass production and biological waste 

recycling concepts for life support, along with their associated reliabilities and costs, will be 

essential for future exploration of space. 
Biological Sciences Roadmap 70 

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Demonstration of a Porous Tu be Hydroponic System 

To Control Plant Moisture and Growth 

Many areas of plant research can benefit from developing technologies to precisely control, 

alter, and monitor plant growth and physiological processes. The ideal test plant growth 

system would allow for control of (1) the spectral (the particular wavelengths of light) quality 

and intensity of the light environment, (2) the water and nutrient contact with the root system, 

(3) the temperature and humidity of the growth area, and (4) the concentration of gases in the 

atmosphere. Such a system would provide tremendous opportunities for plant research in 

areas such as remote sensing of plant stress, toxicology testing, tissue culture of plants, and 

plant physiology. Today's high-quality plant growth chambers can be used to address the 

atmospheric gas, temperature, and humidity concerns. Advances in lighting technology, such 

as light-pipe, light-emitting diodes, and other innovative light sources, are addressing issues 

associated with fine-control lighting. By accurately controlling the plant growth environment 

to minimize variance resulting from environmental responses, we can begin to accurately 

define treatment effects and minimum detectable differences in physiological characteristics 

as they relate to sampling strategies and procedures. In our work, we are exploring the 

possibility of creating a reliable and simple system for controlling the water and nutrients that 

are available to plant roots. This is the one characteristic of plant growth chambers that has 

not been adequately controlled. Specifically, we are evaluating configurations of the Porous 

Tube Plant Nutrient Delivery System (PTPNDS) (figure 1) as well assessing our ability to 

detect plant responses to the various configurations. 

WhiteiBlack 

Polyethylene Outer Open Tubular 

Support With Slot 

A Porous Ceramic 

Filter Tube 

WhiteiBlack Polyethylene ~ Space qc& for Seeds ..::' 

Outer Open Tubular 

Support With Slot Nutrient Supply 

Fitting to Pump and 

Porous Ceramic 

Filter Tube 

Airspace 

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23crnOD ' Nutrient Solution 

Fipre 1. A Diagram of the Porous Tube Coin,uoizenf offlie 

PTPNDS 

Eight porous ceramic tubes, 2 

centimeters (cm) in diameter and 80 

cm in length, were placed on each of 

three laboratory carts. These tubes 

were connected to a standpipe using 

a manifold at the upstream end and a 

peristaltic pump at the downstream 

end. The nutrient solution for each 

cart was contained in an 80-liter 

cooler and pumped to the standpipe 

using a submersible (magnetically 

coupled impeller) pump. Each cart 

can be moved independent of the 

others, allowing ready access to any 

of the plants being grown on the cart 

(figure 2). 

Superdwarf wheat plants (fiiticu777 aestiuum, cv. Perigee) were grown for a complete life 

cycle. Planting was conducted on June 6,2002, with solution flowing through each tube. The 

three variable nutrient solution pressures, henceforth referred to as applied water potentials, 

were established at the initiation of the study: -0.1 cm of water (-0.8 pascal [Pa]), -16 cm of 

water (-1,570 Pa), and -32 cm of water (-3,140 Pa). 

It is important to note that these values are negative; the pressure is lower inside the tubes 

than outside. This is the pressure that controls the amount of water available to the plant 

roots. The main characteristics monitored at this stage are the progression of plant height, 

spectral reflectance, and fresh and dry weights of the plants. 
The PTPNDS performed well, and wheat plants were 

cultured using the three applied water potentials in the 

rhizosphere. It was expected that the -0.1-cm treatment 

would exhibit superior growth compared to the other 

two applied water potentials. This was not the case. 

Instead these plants appeared to be oversaturated with 

nutrient solution early in their life cycles, and until the 

amount of leaf transpiration was sufficient to overcome 

this condition, the plant roots were oxygen-stressed. 

When the leaf transpiration rates were sufficient, the 

growth rate of the plants accelerated and matched or 

exceeded the growth rate of the plants grown with the 

other h70 applied water potentials. As expected, the 

percent plant moisture in the -32.0-cm treatment was 

significantly lower than the percent plant moisture of 

the other tu70 treatments, which were not significantly 

different from each other. In addition, we found a strong 

correlation between percent leaf moisture and light 

reflectance signatures. While there was no significant 

difference found in the light reflectance <ignatiirw 

between the -0.1-cm and the -16.0-cm treatments, there 

was a significant difference between this pair and the 

-32.0-cm treatment. 

Under water-stressed conditions, plants are expected 

to exhibit a decrease in photosynthesis and conductance 

while experiencing an increase in fluorescence at 

higher light intensities. No distinguishable trends 

were found for the parameters, with the exception of 

conductance. The -0.1-cm tray consistently had the 

highest conductance of the three treatments. Since 

photosynthesis was similar among all treatments, this 

indicates that plants on the -0.1-cm trap were the least 

water-use-efficient. 

Key accomplishments: 

Developed hydraulic control systems for laboratoryCollected experimental data on dwarf wheat. 

scale crop tests. 

- Leaf moisture content with root water potentials 

controlled at -0.1, -1.6, and -3.2 kilopascals. 

- Collected percent moisture measurements along 

with spectral analysis for the three treatments. 

Key milestones: 

Refine models defining relationships between 

moisture in components of the remote sensing scene 

and moisture in the plant canopy. 

responses. 

Characterize and address individual species 

I I 

Figure 2. One of Three PTPNDS Units Used 

for flze Study 

Incorporate the LI-COR 

photosynthesis system for 

measuring plant physiological 

responses. 

195 

resistances to gas transfer (e.g., CO, or 

H,O) from leaves to air. 

Related tests of water evaporation 

showed that saturation pressures at a 

constant temperature were independent 

of pressure (figure 2) and that 

evaporation rates at a constant vapor 

pressure deficit increase as pressures 

decrease (figure 3). The increased 

evaporation rates can be explained by 

increased gas diffusion rates at lower 

pressures. 

Improvements to the deployable 

greenhouse prototype continued over 

the past year and the system is nearly 

complete (figure 4). Load cells for 

tracking the water used by individual 

plants and the condensate collection 

reservoir were instaiied and tested, 

and a series of solenoids and water 

pumps was installed for recycling the 

condensate to the plants. A chilledwater circulation system with a 

variable-speed pump was connected 

and tested for holding temperature and 

humidity set points. Two additional 

hemispherical Lexan domes were 

manufactured, and reinforcement ribs 

were added to one to accommodate 

low-pressure tests directly in the 

laboratory. Software routines were 

refined for controlling temperature, 

humidity, and CO, concentration 

and for rewatering the plants with 

condensate. 

Participating Organizations: Dyiianiac 

Corporation, (Dr. PA. Fozuler) and Uni~iersity 

of Florida (Dr. V. Y. Rygnlozi) 

+- - - + 15°C - -- -*- -- - 

0 20 40 60 80 100 

Total Pressure (kPa) 

Figure 2. Test of Saturation Vapor Pressure for Water 

af Different Temperatures and Pressiires 

16 

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Pressure (kPa) 

Figure 3. Water Ei~aporation Rates at D$f&wit 

Huinidities and Pressures 

Figure 4. Reillforced Lexan Dome for Groioing Plants at 

~ozo Pressure in tlie LaEoratory 
Advanced Life Support (ALS) Crop Testing 

Through the processes of photosynthesis, plants (crops) could be used 

to provide oxygen (0,) and food, while removing carbon dioxide (CO,) in 

closed life support habitats. In addition, plants and their associated rootzone microflora can be used to purify wastewater where the transpiration 

is condensed as clean water. To effectively use plants for life support 

and understand the associated costs, information on their environmental 

requirements and responses is needed. The earliest opportunity for testing life 

support contributions of plants in space might involve small growth chambers 

flown on the International Space Station (ISS) or some related flight system. 

ALS crop research activities at KSC over the past year involved (1) cultivar 

trials of radish (Raphanus sativus) and bunching onion (Allium fistulosum), 

(2) light spectral studies with radish and onion and photoperiod studies 

with potato (Solanum tuberosum), and (3) CO, studies with bean (Phaseolus 

vulgaris). Radish cultivars Sora and Cherry Bomb I1 performed well in 

hydroponic systems, even at 25 to 28 degrees Celsius (“C), temperatures that 

might be encountered in spaceflight. Yields of eight radish cultivars (grown 

for 21 days) and eight onion cultivars (grown for 42 days) were similar 

within a species under either cool white fluorescent or high-pressure sodium 

lighting at light levels of about 300 pmol rn-, s-* photosynthetic photon flux 

(PPF), which should be achievable on the ISS. Studies with potato showed 

that light levels as low as 2 to 4 pmol rn-, s-l PPF during dark cycles delayed 

tuberization. This suggests that ”short-day” events in some plants, such as 

flowering and tuberization, may be perturbed by low levels of light during 

“night” cycles, as might occur in open- or mixed-crop production systems 

in space. CO, studies with bean showed that yields and water use efficiency 

were highest at CO, levels of 1,200 parts per million (ppm) (0.12 kilopascal 

[kPa]) compared to 400, 4,000, 8,000, and 16,000 ppm (0.04, 0.4, 0.8, and 

1.6 kPa). CO, levels in space habitats with human crews (such as the ISS) 

commonly reach levels of 4,000 ppm, suggesting that CO, concentrations in 

spaceflight environments may be somewhat higher than optimal for certain 

plants. 

Twen ty-One-Day-Old Hydroponically Grown 

Radishes (Raphanus sativus L. cv. Cherry Bomb 11) 

Resulting From Cultivar Evaluation Tests 
Forty-Two-Day-Old Bunching Onions 

(Alliumfisfulosum) Growing in 

Hydroponic Systems Under Cool 

Wliite Fluorescent Lamps 

Key accomplishments: 

Completed radish and onion cultivar evaluations 

under high-pressure sodium and cool white 

fluorescent lamps. 

Completed “light pollution” studies with potato. 

Completed super-elevated CO, studies with bean. 

Key milestones: 

Initiate environmental tests with radish, onion, aid 

lettuce to include typical ISS environments. 

Initiate mixed-cropping tests with radish, onion, 

and lettuce grown with solid media (arcillite) and 

hydroponics. 

Initiate cultivar evaluation tests with lettuce, 

tomato, and pepper. 

Contacts: Dr. R.M. Wlreeler (Ri~.~/mo~~ii.M. ~~~~eelrr((~’rr~~n.~i~~~), 

YA-E4-C, (321) 476-4273; aizd Dr. J.C. Sager, YA-E4-C, (3211 

476-42 70 

Participating Organization: Dyizamlzc Corporution (DE G. D. 

Goins, N.C. Yorio, aizd Dr. G.W. Sttrtte) 
Photosynthesis and Metabolism of Wheat in Microgravity 

The photosynthetic rate of higher plants is a critical component of plantbased atmospheric regeneration systems being proposed for long-duration 

space missions. The Photosynthesis Experiment and System Testing and 

Operations (PESTO) experiment is designed to directly measure photosynthesis 

in microgravity so an informed decision on the feasibility and design of these 

systems can be made. The overall objective of this research was to determine 

the effect of microgravity on photosynthetic response of plant tissues developed 

in either gravity or microgravity. A specific objective of this research is to 

determine the effect of microgravity on (1) carbon dioxide (CO,) and light 

response curves of wheat, (2) metabolism and electron transport processes 

associated with photosynthetic and respiratory gas exchange, (3) carbohydrate 

partitioning in wheat, and (4) gas exchange, including water, over a range of 

atmospheric vapor pressure deficits. Another objective is to use the knowledge 

gained to understand the response of plants grown under the elevated CO, 

conditions of commercial, controlled-environment crop production systems. 

The PESTO experiment, which was conducted during Increment IV of the 

International Space Station (ISS), was one of the most complex plant experiments 

ever performed by NASA. The PESTO experiment was launched to the ISS 

onboard STS-110 (Atlantis) on April 8,2002, and was returned to Earth June 

19,2002, onboard STS-111 (Endeavour), spending a total of 73 days in space. A 

ground-based control experiment was conducted on a 14-day delay and was 

harvested on July 3,2002. During the flight period, there were seven harvests of 

PESTO wheat and priming of six wheat root modules on-orbit. Eighteen plants 

were fixed on-orbit for morphological and structural analysis, 4 plants were 

fixed for RNA and molecular genetic analysis, and more than 280 plants were 

frozen for biochemical analysis. The Biomass Production System (BPS) cameras 

collected 3,018 images during flight, and 3,640 images were obtained from the 

Summary of Gas Exchange Characteristics, Including CO, and Light 

Response Curves Obtained Between Stands of Wheat Grown for 20 

Days Onboard the ISS and on Earth 

ground-based controls. Sixtyfive digital images of the wheat 

harvest were collected during 

flight, with 75 being collected 

from ground-based controls. 

In addition, crew members 

recorded approximately 50 

images and over 3 hours of video 

of flight operations. Over 2 

gigabytes of environmental data 

were obtained from the flight 

and ground-based control units. 

Activities have been directed to 

the preparation, maintenance, 

and return of the experiment 

from the ISS. The ISS experiment 

has produced a wealth of data 

that is currently being analyzed. 

There was excellent germination 

(greater than 95 percent) in each 

of the six root modules primed 
on-orbit, and growth rates between each replicate 

were comparable. Initial analysis of the data reveals 

no significant differences in either germination or 

growth rate between the flight and ground-based 

control treatments. 

A primary objective of the flight experiment was 

to characterize gas exchange characteristics of the 

developing wheat canopy to differences in CO, 

and light. These CO, and light response curves 

were conducted at three stages of development on 

each chamber in orbit. The treatments consisted 

of obtaining CO, response curves at five different 

photosynthetic photonflux intensities. There were no 

statistically significant differences between the flight 

and ground-based control treatments in CO, response 

characteristics (including maximal photosynthetic 

rate and CO, compensation point) and in the light 

response characteristics of the developing wheat 

stands. Analysis of the morphological, structural, 

and biochemical processes associated with stand gas 

exchange measurements during the E-day n;issioii is 

ongoing. 

A number of long-duration, preflight experiments 

were conducted in the BPS during FY 2002, including 

a high-fidelity ground-based control test and 

experiments to determine the effects of elevated 

root zone temperatures on wheat growth and 

development. 

Key accomplishments: 

Completed 73-day mission during Increment IV 

onboard the ISS. 

Completed high-fidelity ground control 

experiment in the BPS. 

Completed five 24-day temperature experiments 

using ground control root modules. 

Published two manuscripts in peer-reviewed 

journals. 

Presented five talks at international scientific 

meetings. 

Images of Wlieat Grouifli at FD1 nnd FD8 

Key milestones: 

Complete biochemical and biomolecular 

Complete ground control temperature 

Present results at international meetings 

analysis of plant tissue. 

experiments. 

and submit results for publication. 

Part iciyat iiig Organ izat io12 : Dytzainac Corprn t ion 

(Dr. G. W. Stictte) 

~~ 

al 

(D 

2 

: 

3 

Q 

3 

Q 

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3 

3 

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(0 

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h) 

0 

0 

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57 

0 

20 1 
Plant Lighting System 

Plants can produce food, regenerate oxygen, and purify water for space 

inhabitants. A major challenge to growing plants in space will be controlling 

and supplying sufficient quantity and quality of light. Conventional lighting 

technologies would prohibit growing plants on a large scale in space due to low 

electric power conversion efficiencies. Current lighting research for space-based 

plant culture is focused on innovative lighting technologies that demonstrate 

high electrical efficiency and reduced mass and volume. Accordingly, lightemitting diodes (LEDs) are promising technologies being developed to 

efficiently generate photosynthetic radiation. LEDs can illuminate near the 

peak light absorption regions of chlorophyll while producing virtually no nearinfrared radiation (which is not utilized in photosynthesis). The work in the 

KSC Life Sciences Support Facility, in association with NASA's Advanced Life 

Support (ALS) /Advanced Human Support Technology Project, is studying the 

use of LEDs as alternative light sources for plant production. 

Within the ALS Project, salad-type plants represent crops that could provide a 

portion of fresh food as well as psychological benefits to the crew aboard future 

space transportation vehicles. Laboratory data generated with salad-type crops 

grown under various lighting sources could provide important information for 

modeling and developing future missions. 

Work has been completed with Swiss chard, spinach, radish, and lettuce 

plants grown under different lighting sources for 28 days. Three lamp banks 

represented broad-spectrum white light sources (microwave, high-pressure 

sodium, and cool white fluorescent). Past tests also included separate LED 

arrays filled with a given peak wavelength of red (664, 666, 676, 688 nanometers) 

LEDs. Each LED array contained single rows of blue LEDs (474 nanometers) 

evenly distributed within the multiple rows of red LEDs. The relative 

stoichiometry of photosystem I and I1 reaction centers (which are responsible for 

the initial conversion of light energy in plants) was measured for each treatment. 

Additonal measurements throughout growth, including plant chlorophyll 

fluorescence, accessory pigment concentrations, and responses of photosynthetic 

rate to light and carbon dioxide levels, were incorporated. 

Arrays of Red and Blue LEDs With Green (left) and Without Green (right) Supplementation 

I 202 
ldeiitical Lettuce Plants Under Arrays of Red aid Blue LEDs With Greeir (left) mid Without Green fright) 

Supplenientatioiz 

The combination of red and blue light 

proved to be an effective lighting source 

for growing spinach, radish, and lettuce. 

However, the addiiioii of green light may 

promote increased plant growth since 

green light penetrates the plant canopy 

better than red or blue light. Leaves in the 

lower canopy could use the transmitted 

green light in photosynthesis. For a space 

mission, highly optimized lighting systems 

are necessary to conserve power and 

maximize plant growth, so supplemental 

green light needs to be evaluated. Green 

light may also have additional psychological 

benefits for the crew. Plant leaves readily 

absorb red and blue light, so absorptance 

is high and reflectance is low. Therefore, 

even healthy plants grown under red and 

blue LEDs alone appear gray or black to 

humans. Hence, the addition of green LEDs 

e 

e 

to red and blue LED arrays makes plants 

appear green to the crew. Research has 

begun to evaluate the beneficial aspects of 

supplemental green light on salad crops in 

terms of increasing light penetration in the 

canopy and enhancing the aesthetic appeal 

of plants grown under LEDs. 

Key accomplishments: 

1999: Began experiments with growing 

salad-type plants with LEDs and 

microwave lamps. 

2000: Completed initial salad-type plant 

growth studies with LEDs and microwave 

lamps. NASA NRA Solicitation 98-HEDSextended through 2001. 

2001: Began experiments with mixtures 

of salad plant species to compare the 

growth of multiple crops in a common 

environment / hydroponics system. 

Testing included characterization 

of photosynthetic reaction center 

stoichiometry (photosystem I versus 11) in 

response to light quality. 

2002: Began experiments to evaluate the 

beneficial aspects of supplemental green 

light on salad crops in terms of increasing 

light penetration into the canopy and 

enhancing the aesthetic appeal of plants 

grown under LEDs. 

01 gar;: to 5~;namac Csrnnratinn IAJ~C r UL*--- 

Key milestone: 

2003: Continue the evaluation of 

supplemental green light at different 

levels to determine the photobiological 

effects on salad crops. 

Coiztacts: Dr. J.C. Sager (Jnliiz.C.Srr~ci~iinsn.por~), 

YA-E4-C, (321) 476-4270; aiid Dr. R.M. Mreeler, 

YA-E4-C, (321) 476-4273 

Paificiyatiiig Organizations: Nafioiial Research 

Coziiicil (Dr. H.H. Kim) aid Dyizainac Corporation 

(Dr. G.D. Goiiis) 
Effect of Microgravity on Raphanus Sativus L. 

The objective of this experiment is to determine the effects of microgravity 

on the growth and development of radish (Raphanus sativus L.). Salad-type 

crops that have a short planting-to-harvest cycle may be used to supplement 

the primary diet of crews on short-to-medium-duration space missions. In 

addition to the life support applications, radish serves as a model system for 

evaluating the effect of microgravity on carbohydrate partitioning to the root. 

Ethylene is a volatile plant hormone that has wide-ranging effects on radish 

growth and development. Accumulation of ethylene produced from both 

biogenic and anthrogenic sources may reduce plant productivity and alter 

growth. Experiments to determine the threshold of ethylene sensitivity in 

radish were conducted. Significant physiological and morphological changes 

occur with chronic exposure of 30-to-1,000-part-per-billion (ppb) ethylene. 

These concentrations are 7,500 times lower than the current guidelines 

for allowable spacecraft concentrations of ethylene. Research to develop 

techniques for removing ethylene is ongoing. 

The Radish Assimilation in Spacecraft Testbed Atmospheres (RASTA) 

experiment will be conducted in space in a specialized plant growth chamber, 

the Plant Growth Facility-Split Plenum (PGF-SP) that is being built by 

Bionetics Corporation (Kennedy Space Center, Florida). Numerous tests were 

performed to optimize the conditions for plant growth in preprototype 

PGF-SP chambers, including types of watering systems, lighting systems, and 

rooting systems. Radishes grown in all tests undergo detailed morphological, 

developmental, and physiological analysis. 

Key accomplishments: 

Completed experiments to determine dose response curve of radish to 

Continued experiments to quantify production of volatile organic 

Designed preprototype plant growth chambers for testing. 

Presented papers at two international conferences. 

Published three peer-reviewed manuscripts. 

ethylene. 

compounds from radish. 

Key milestones: 

Complete automated control system for continuous monitoring of 

Determine developmental effects of starch partitioning in radish. 

Characterize morphological and physiological adaptations to light quality. 

Continue testing experimental protocols in prototype flight hardware. 

Present results at international scientific conferences. 

biologically active volatile compounds. 

Participating Organization: Dynamac Corporation (Dr. G.W. Stutte) 

204 
Radish Plants Being Grown in Preprototype Flight Ckambers Using 

Light-Emitting Diodes To Validate Methodsfor Growing Piam in Space 

Closed Exposure Ckambers Alloil1 the Eflects of Common Spacecraft 

Atmospheric Components on Plaiit Grouith To Be Determined 

Q 

3 

z1 

2051206 
INDEX 

A 

Adams, E W., 35,51 

Adams, T., 63 

Adams, T.C., 169 

Allen, M., 81 

Appleton, B., 68 

Arkin, Dr. CR., 35,51 

Atuwod, G.C., 163 

Augustynowicz, Dr. S.D., 79, 81, 85 

Azofcfa, Dr. J., 35 

B 

Baker, Dr. G.H., 155, 175,177 

Barile, Dr. R.G., 15, 147 

Bartlz, TS., 155, 164, 167, 173, 189 

Baska, D.F., 175 

Bastiiz, Dr. G.L., 105, 111 

Bayliss, J.A., 129 

Beckus, A., 153 

Beil, R.J., 35 

Benjamin, Dr. PC., 164, 178,189 

Bigos, S., 183 

Birr, R.B., 97, 99, 103 

Blalock, N.N., 5, 35,41, 43 

Boiiizer, T,, 74 

Bookman, PI., 153 

Boyd, S.W., 103 

Breakfeld, R.A., 81, 85 

Brindley, E.E., 163 

Brown, C.J., 163 

Brughelli, K., 167 

Buehler, Dr. M.G., 121, 139 

Buhler, Dr. C.R., 117,119,121, 123 

Bull, J.B., 101 

Bundick, S.N., 97 

Burns, B.C., 163 

Bums, B.M., 23,37,41,43,47,49 

,aut!er, C.D., 187 

Butler, R.A., 163 

Buftner, Dr. W.J., 7 

C 

Calle, Dr. C.I., 113, 117, 119, 121, 

123,125,127,129,131,133 

Calle, Dr. L.M., 135, 136, 139, 145, 

147 

Carstens, Dr. D.S., 167, 173, 175, 177 

Case, J.L., 27 

Casteel, S., 97 

Cafes, G.R., 160 

Chamberland, D.W., 197 

Chang-Diaz, Dr. E, 35 

Christensen, K.L., 159 

Ciauarella, A.A., 19 

Clements, G.R., 1 

Coleman, A.M., 99 

Cox, R.B., 9,53 

Crikis, A., 19 

Cummings, K.R., 85,86 

Curley, C., 35 

Curran,J.J., 71, 135, 141, 143, 147 

Curran, J.l?, 141, 143, 147 

Czaban, S.K., 35 

D 

Dazis, C.K., 25,2t?, x,39 

Denson, E.C., 97, 99, 103 

Deremer-Cook, E.]., 7 63 

Deyoe, R.T., 5,47 

Diaz, Dr. J.A., 35 

Doctor, M.A., 83 

Doesken, Dr. N., 9 

Dominguez, J.A., 15,25,28,30, 39 

Dreschel, Dr. T.W., 195 

E 

Eckhof, A.J., 5,37,41,43,49 

Edelmann, J.J., 35 

Elliott, Dr. D.M., 159 

Erdogan, T., 101,103,105,111 

Ewald, J.R., 171 

INDEX 207 
F 

Farmer, G.A., 163 

Ferrell, B.A., 101 

Fesmire, J.E., 79, 81, 83, 85 

Floyd, D.P., 35 

Foderousky, A.M., 153 

Follistein, D.W., 35,51 

Foster, T.E., 195 

Fowler, Dr. P.A., 197 

Fox, I.]., 55 

Fredrickson, T.E., 187 

C 

Gavura, J.M., 97 

Gibbons, R.D., 16 

Gibson, J.E., 85,86 

Gleman, S., 65, 73, 74, 77 

Goembel, Dr. L., 131 

Goins, Dr. G.D., 199,203 

Gompf, Dr. R.H., 129 

Gordie, R.A., 63 

Graul, M., 164 

Greenfield, T.D., 16 

Gribble, D.A., 133 

Grier, R., 71 

Grifjin, Dr. T.P., 35, 51 

Groh-Hammond, M.M., 163,171, 

173,175 

Groop, E.E., 121,125,127 

H 

Haase, W.H., 164 

Haberbusch, M.S., 59 

Haley, S., 101 

Hall, C.R., 195 

Hallberg, C.G., 16 

Hamilton; L.E., 163 

Harlacher, M., 97 

Harris, W.G., 105 

Haskell, W.D., 9,35,51, 53 

Heckle, Sr., K. W., 79, 81 

Henderson, P.G., 35 

Hickman, J.C., 101 

Hoang, Dr. N., 99 

Hodge, T.R., 45 

Hogue, M.D., 123 

Humeniuk, R.P., 163 

I 

lmmer, Dr. C.D., 5, 9,23,27,103 

Ingoldsby, K.A., 185 

J 

Johnson, Jr., D.E., 185 

K 

Kandula, Dr. M., 61, 149 

Kanki, Dr. B.G., 159 

Kelly, A.O., 88 

Kerios, C., 105 

Kim, Dr. H.H., 203 

Klinko, S.J., 25, 28,30, 39 

Klonowski, P.J., 86 

Kolody, Dr. M.R., 139 

Kounaves, Dr. S.P., 139 

Kremer, S.E., 101 

Kruhm, D.A., 1 

1 

Lackascyck, J., 163 

Lane, Dr. J.E., 9, 23, 27, 33 

Lanzi, R.J., 101 

Larsen, D.C., 163 

Lawrence, B.A., 163 

Lebovitz, G., 153 

Lippitt, T.C., 63, 73 

Littlefield, A.C., 74 

Lohr, G.W., 159 

Lucena, A.R., 37,41,43 

Luo, Dr. W., 123 

Luth, D.J., 163 

208 INDEX 
INDEX 

M 

MacDowell, L.G., 71, 135, 136, 141, 

143,145,147 

Madura, J.T., 107 

Maiiley, W.C., 163 

Maizobianco, Dr. J.T., 27 

Manfoz~ani, Dr. J.G., 121 

Mariii, J.A., 101, 103 

Mata, Dr. C.T., 37,41, 43,47, 49, 99, 

103 

Mazumder, Dr. M.K., 117 

McCasicey, C.E., 86 

McCleskey, C.M., 185 

Medelius, Dr. P.J., 13, 37, 43, 107, 109 

Merceret, Dr. F.J., 27 

Merrill, T. W., 20 

Michel, M.C., 20 

Mobasher, Dr. A., 150 

Moist, J.O., 163 

Mueller, R.P, 15, 63, 65, 67, 68, 71, 

73, 74, 77 

Mundsclrau, Dr. M.V, 133 

Murray, A.J., 99 

N 

Nagy, Z.F., 71, 85, 86, 91 

Naylor, G.R., 35,51 

Nelson, R.A., 94, 97, 101, 105 

Nguyen, C.T., 59 

Nieten, D.A., 19 

Nieves, R., 19 

Noble, J.T., 163 

Notardonato, W.U., 91 

Nowicki, A.W., 121,125, 127 

Nurge, M.A., 11 

0 

O'Brien, T.L., 171 

P 

Parks, S.L., 45 

Patterson, M.S., 101 

Peabody, PA., 163 

Pelakli, B., 20 

Peroffi, J.M., 5, 37, 41, 43, 47, 49 

Peters, J.W., 53 

Philippart, M., 175 

Plrillips, G. D., 164 

Polk, I. D., 16 

Posada, J.G., 177 

Poveda, G., 19 

Poiiiell, W.R., 101 

Q 

Quiiicy, C.D., 191 

Quiiiiz, S.M., 201, 204 

R 

Rakofsky, A.M., 163 

Randazzo, ].I., 5, 13,41,43,49 

Rees, J.A., 13 

Reluini, Dr. K., 175, 177 

Rhodes, R.E., 55, 169 

Rhodeside, G.R., 178 

Riclziuso, PF., 129 

Rieck, E.G., 163 

Riley, Jr., H. W., 171 

Robinson, G.F., 101 

Rozewski, J.M., 163 

Ruiz-Torres, Dr. A., 181 

Rygalou, Dr. VY, 197 

S 

Sager, Dr. J.C., 199,203 

Sakahara, R.D., 97 

Schneider, S.M., 153 

Scholfens, B.E., 81, 83 

Schultz, Dr. J.A., 51 

Semmel, G.S., 20 

Sierra, A., 111 

Simpson, Dr. J.C., 99, 101, 103 

SkafiA.F., 59 

Skelly, D.M., 94 

Smitlz, E.C., 71 

Sobchak, T.C., 97 

INDEX 209 
INDEX 

Sojourner, S.J., 91 

Sortman, M.A., 163 

Stalnecker, W., 67 

Stambolian, D.B., 163 

Stender, J., 91 

Stottler, R., 183 

stout, S.J., 49 

Stute, R.A., 16 

Stutte, Dr. G.W., 199,201,204 

Surma, J.M., 45 

Sweeney, G.P., 171 

T 

Taylor, B.R., 163 

Taylor, J. D., 13 

Taylor, M.A., 133 

Thomas, PA., 163 

Thompson, Dr. J.S., 133 

Thompson, K.L., 113 

Townsend, I.I., 73 

Trautwein, J.K., 63, 73 

Trout, D.H., 109 

U 

v 

Valencia, L.M., 97 

Valentino, W. D., 175 

Van Sciver, S., 91 

Vinje, R.D., 135,136,145 

Vo, K.A., 20 

Vu, Dr. B.T., 61,149,150 

w 

Wachowski, S.G., 163 

Walls, L.K., 59 

Warren, A.L., 160 

Warren, K.P, 163 

Waterman, R.D., 1 

Wessells, R.L., 99 

Wheeler, Dr, R.M., 197, 199, 203 

White, T.R., 163 

Whiteman, D.E., 97 

Wicke, D., 67 

Wilkhu, S.S., 163 

Willard, D.E., 53 

Williams, E.H., 105 

Williams, Dr. M.K., 79, 83 

X 

Xie, Dr. X., 133 

Y 

Yorio, N.C., 199 

Young, R.C., 7 

Youngquist, Dr. R.C., 9,11,13,23, 

45,53,111 

Z 

Zapata, E.Z., 181,189 

Zeitlin, N.P., 139 

Zoerner, R.D., 101,103 

Zysko, J.A., 33 

210 INDEX 
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Research and Technology 

2003 Annual Report of the John F. Kennedy Space Center 

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NASA, John F. Kennedy Space Center 

Kennedy Space Center, Florida 32899 

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13. ABSTRACT (Maximum 200 words) 

The John F. Kennedy Space Center (KSC) is America's Spaceport Technology Center. The KSC 

technology development program encompasses the efforts of the entire KSC team, consisting of 

Government and contractor personnel, working in partnership with academic institutions and 

commercial industry. KSC's assigned mission areas are space launch operations and spaceport 

and range technologies. KSC's technology development customers include current space transportation programs, future space transportation programs / initiatives, and enabling technical 

programs. The KSC Research and Technology 2003 Annual Report encompasses the efforts of 

contributors to the KSC advanced technology development program and KSC technology 

transfer activities. 

Dr. Dave Bartine, KSC Chief Technologist, (321) 867-7069, is responsible for publication of this 

report and should be contacted for any desired information regarding KSC's research and technology development activities. 

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