NASA/TM-2001-209037

Flight Demonstration of X-33 Vehicle Health Management System Components on theF/A-18 Systems Research Aircraft

Keith A. Schweikhard, W. Lance Richards and John TheisenNASA Dryden Flight Research CenterEdwards, California

William Mouyos and Raymond GarbosBAE SystemsNashua, New Hampshire

December 2001

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NASA/TM-2001-209037

Flight Demonstration of X-33 Vehicle Health Management System Components on the F/A-18 Systems Research Aircraft

Keith A. Schweikhard, W. Lance Richards and John TheisenNASA Dryden Flight Research CenterEdwards, California

William Mouyos and Raymond GarboBAE SystemsNashua, New Hampshire

December 2001

National Aeronautics andSpace Administration

Dryden Flight Research CenterEdwards, California 93523-0273

NOTICE

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Available from the following:

NASA Center for AeroSpace Information (CASI) National Technical Information Service (NTIS)7121 Standard Drive 5285 Port Royal RoadHanover, MD 21076-1320 Springfield, VA 22161-2171(301) 621-0390 (703) 487-4650

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FLIGHT DEMONSTRATION OF X-33 VEHICLE HEALTHMANAGEMENT SYSTEM COMPONENTS ON THE F/A-18 SYSTEMS

RESEARCH AIRCRAFT

Keith A. Schweikhard, W. Lance Richards and John Theisen,NASA Dryden Flight Research Center, Edwards, California

William Mouyos and Raymond Garbos,BAE Systems, Nashua, New Hampshire

Abstract

The X-33 reusable launch vehicledemonstrator has identified the need to implement avehicle health monitoring system that can acquiredata that monitors system health and performance.Sanders, a Lockheed Martin Company, hasdesigned and developed a COTS-based openarchitecture system that implements a number oftechnologies that have not been previously used in aflight environment. NASA Dryden Flight ResearchCenter and Sanders teamed to demonstrate that thedistributed remote health nodes, fiber opticdistributed strain sensor, and fiber distributed datainterface communications components of the X-33vehicle health management (VHM) system could besuccessfully integrated and flown on a NASA F-18aircraft. This paper briefly describes components ofX-33 VHM architecture flown at Dryden andsummarizes the integration and flight demonstrationof these X-33 VHM components. Finally, itpresents early results from the integration and flightefforts.

Acronyms and Symbols

BIT built-in test

COTS commercial off-the-shelf

DAS data acquisition system

DHS distributed hydrogen sensor

DLL design limit load

DSS distributed strain system

DTS distributed temperature sensor

FDDI fiber distributed data interface

FTF flight test fixture

GEN II generation II architecture

HOB health optical bus

IVHM integrated vehicle health management

MHz megahertz (million Hertz)

NASA National Aeronautics and SpaceAdministration

RHN remote health node

RS-232 Recommended Standard 232 (serialinterface, IEEE)

RS-422 Differential Serial Interface Standard

SRA Systems Research Aircraft

STE system test equipment

VHM vehicle health management

VME versa module eurocard

b angle of sideslip

µm micrometers (10-6 meters)

Introduction

Integrated vehicle health management (IVHM)is becoming an essential part of aircraft systems forcommercial and military vehicles as well as forspace applications, such as reusable launchvehicles, space stations, and satellites. Thedevelopment of an IVHM involves the integrationof new or existing sensor technology withacquisition and processing devices, implementationof prognostic algorithms, and dissemination ofresults to the correct user in a timely manner. Thedevelopment of such systems can improve vehicle

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safety, reduce maintenance costs, and improvevehicle readiness by identifying potential faults andfailures, taking the proper corrective actions, andinforming the responsible crew member of thehealth and condition of the vehicle.

The X-33 Advanced TechnologyDemonstrator program has taken a first steptowards IVHM by implementing a Generation Ivehicle health management (VHM) system thatmonitors and records all data on the X-33. ThisVHM system transmits critical X-33 data by meansof a radio frequency downlink [1]. Sanders, aLockheed Martin Company (Nashua, NewHampshire), was contracted to develop andintegrate the VHM system on the X-33. This systemconsists of a commercial-off-the-shelf-based(COTS), distributed, open architecture, dataacquisition system, with a ground-based downloadstorage component. This design implemented anumber of new concepts that have beendemonstrated in the laboratory, but have notpreviously been used in critical flight applications.In an effort to reduce the risk to the X-33 program itwas decided that high-risk components of the X-33VHM system should be demonstrated in a flightenvironment. NASA Dryden and Sanders teamed todemonstrate this VHM system modified forinstallation on the NASA Dryden Flight ResearchCenter F/A-18B Systems Research Aircraft (SRA).

This paper gives an overview of the X-33VHM architecture, defines the VHM architecturethat was integrated and flown on the SRA, andsummarizes the VHM system integration efforts onthe aircraft, with an overview of the results of theflight demonstration phase of the experiment.Finally, it summarizes the lessons learned from theexperiment.

Note that use of trade names or names ofmanufacturers in this document does not constitutean official endorsement of such products ormanufacturers, either expressed or implied, by theNational Aeronautics and Space Administration.

Scope and Approach

There were three main purposes forconducting this experiment; to (1) provide a risk

reduction effort for the X-33 program, to (2)demonstrate enabling technologies for futurereusable launch vehicle efforts, and to (3) developgeneric tools and platforms that could be used forfuture VHM research efforts.

The research goals for the X-33 risk reductioneffort included demonstrating that a fiberdistributed data interface (FDDI) architecture couldbe used to distribute data between the remote healthnodes (RHN) and the VHM computer. Proving theaccuracy and capabilities of the RHN was alsoimportant in providing risk reduction for the X-33program. The final area of risk reduction that wasperformed involved the integration and applicationof a limited but representative X-33 VHM systemarchitecture on the aircraft. Data collected using theVHM system could be compared to existing datathat sensors had collected using a separate onboarddata acquisition system (DAS).

Enabling technologies investigated during theexperiment included implementations of fiber opticcable plants that were unique for this experiment.The enabling technologies were (1) a generation II(GEN II) multi-mode fiber optic cable plant, (2) asingle-mode fiber optic cable plant design, and (3) afiber optic demodulation scheme and distributedstrain sensors. None of these had been tested in aflight environment.

The final objective of the flight experimentwas the development of a generic platform forVHM research. The two areas where genericplatforms were implemented during this experimentwere the open architecture VHM system design,which allow COTS hardware to be directlyintegrated with the VHM system; and thedevelopment of a generic structural flight testfixture (FTF) which was designed to generaterepeatable static loads in both a laboratory andflight environment that could be attached to astandard centerline pylon on a fighter class aircraft.Proper design of the fixture would allow rapid,repeatable comparisons of the distributed strainsystem (DSS) strain gage measurements withconventional foil strain in the laboratory as well asin flight.

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X-33 System Architecture

Vehicle ArchitectureThe X-33 VHM architecture (fig. 1) acquires

data from a wide variety of sensors and systems.The system monitors and records all data on thevehicle, including all the flight bus data, all vehiclehealth data (structural, mechanical and system built-in test (BIT) status), and all the flight testinstrumentation sensor data. This system interfacesto a NASA Jet Propulsion Laboratory avionicsflight experiment that communicates over aMIL-STD-1773 fiber optic interface [2].

The VHM system uses a distributed sensingarchitecture to collect data from conventionalsensors. The data are concentrated using 50 RHNsto gather structural, mechanical and environmentalsensor data. Each RHN communicates with theVHM computer over a fiber optic health optical bus(HOB). The VHM computers (1) acquire data fromthe HOB, (2) monitor and record allMIL-STD-1553B data [3], and (3) gather fiber opticdistributed temperature, hydrogen, and strain sensordata from digital signal processors which reside inthe VHM computer. Optical sensors are mounteddirectly on the X-33 propellant tanks, while thedriving optics and processing are housed in theVHM computers.

Figure 1. X-33/SRA configuration.

The X-33 vehicle health manager-A (VHM-A)shown in figure 2 is one of two nearly identicalversa-module-eurocard (VME)-based computershoused in a modified full air transportation rackchassis used to monitor and record data. VHM-Acontains a commercial off-the-shelf or militaryoff-the-shelf standard processor interfaced to a4.5 gigabyte PC-based hard disk drive storagemodule and various input/output modules. VHM-Aconnects to 25 RHNs by means of the fiber opticHOB implementing an FDDI. The distributed fiberoptic temperature sensors and MIL-STD-1553B busmonitors interface directly to the VHM-Acomputer. The VHM-A computer acts as a remoteterminal by using a MIL-STD-1553B interface forVHM command and control information. VHM-Bis nearly identical to VHM-A; the differences arethat the VHM-B line replaceable unit contains thedistributed fiber optic strain and hydrogen sensors.The VHM computer also handles telemetrydownlink, recorder download, and systemdiagnostic functions.

The RHN shown in figure 3 acts as a dataconcentrator for analog sensor information. EachRHN communicates with its respective VHMcomputer over two independent HOBs using FDDIas the network protocol. The FDDI uses a fullduplex, dual counter-rotating token ring topology toprovide reliable communications in the event of afailure.

Figure 2. VHM-A/B.

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Figure 3. Figure 3 RHN.

Each individual RHN can collect data from upto 40 sensors located in close proximity of the RHNmounting location. This configuration reducessensor wire weight and the chance of lost data thatcould result from electromagnetic interference andother outside phenomena. The RHN interfaces withseveral types of sensors including accelerometers,strain gages, thermocouples, resistive temperaturedevices, and sensors that measure pressure, rate,angular position, voltage, current, angular rate, andlinear position. The RHN also provides theexcitation to the sensor, if any is required. Thesensor data are collected, conditioned, digitized,and time-tagged in the RHN. The data are sent tothe associated VHM computer by means of theHOB to be recorded on a mass storage device in theVHM computer. After the vehicle lands, the dataare downloaded by an optical connection from theVHM-A and VHM-B computers to the ground-based processing and storage facilities.

An optical network of distributed strain,temperature, and hydrogen sensors make up thefiber optic distributed sensor system on the X-33.The goal of this network of fiber sensors is to createa global strain, temperature, and hydrogen leak mapfor monitoring the health of the tank structure andcryogenic insulation. The fiber optic sensorsemploy two unique sensing techniques. The firsttechnique used for strain and hydrogen leakdetection utilizes a wavelength tunable narrowlinewidth laser, a fiber optic network containing asensing fiber with Bragg grating sensors, lightdetection photodiodes, signal conditioningelectronics, and a digital signal processor. Thesecond technique used a multi-mode fiber with abroadband laser source as a temperature sensor.

Distributed Strain SystemThe DSS is a fiber-optic-based sensing

network that uses a NASA Langley patenteddemodulation technique to infer local information(strain, temperature, gas concentration, etc.) atmultiple vehicle locations [4]. Sanders modified theLangley demodulation architecture [5]. ThroughX-33 and internal research and developmentfunding the design further evolved over the courseof the flight demonstration experiment. The DSS isdesigned to collect information from eight DSSfibers with up to 20 discrete fiber Bragg gratingsmultiplexed along a single fiber. Bragg gratings arelocalized, photo-induced, periodic perturbations ofthe refractive index in the core of single-modeoptical fiber. These gratings have the uniquecharacteristic of reflecting back a narrowbandwavelength, called the Bragg wavelength, whenilluminated by either a broadband light source, or inthe case of the DSS, a tunable laser diode. Thedemodulation technique begins by firsttransforming the reflecting signals from each of thefiber Bragg gratings from the wavelength to thelength domain by performing fast Fouriertransforms. This yields the location of each fiberBragg grating from the reference reflector along thesensing fiber. Each grating is then windowed, aninverse fast Fourier transform is performed, and thecenter wavelength of each grating is calculatedduring loading, and then subtracted from the initialwavelength at zero load [6].

SRA Experiment Architecture

SRA Aircraft ConfigurationThe SRA is a multi-purpose facility with the

primary goals of identification and flight test oftechnologies beneficial to subsonic, supersonic,hypersonic, and space applications [7]. The aircraftis configured to fly multiple independentexperiments during any given flight block, allowingfor maximum utilization of vehicle assets and flighttime by sharing flight time and costs for any givenmission. Experiments flown with the VHMexperiment included: a parameter identificationexperiment, a conformal ultrahigh frequency andvery high frequency antenna experiment, and aflush air data system.

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SRA VHM System ArchitectureHigh-risk components of the X-33 VHM

system were identified for the SRA flightdemonstration. These components comprised arepresentative subset of the overall X-33 VHMsystem architecture (fig. 1). Slight modificationswere made to the X-33 VHM components to eithervalidate technologies that would eventually befielded in a VHM system or that were necessary tointerface into the F/A-18 SRA. A diagram of theSRA experiment configuration is shown in figure 4.

Figure 4. VHM experiment configuration.

Modifications to the SRA VHM experimentincluded: a single VHM computer with solid-statememory replacing the disk drive, a single FDDImodule, a DSS module, and GEN II opticalinterconnects. Two RHNs were configured in ahalf-FDDI ring configuration.

HOB communication was accomplished usinga prototype GEN II fiber optic cable plantconsisting of a five-fiber, 62.5 mm multi-moderibbon cable that had five fibers interwoven withthe strength member and clear rubberized jacketand a multi-fiber termination on each end. Thetermination of the fiber could only be accomplishedat the manufacturer facility. As a result, themulti-mode cable plant was delivered cut to lengthwith termini and connectors on both ends asdescribed above.

The ground-based system test equipment(STE) was based on a standard PC architecture.This equipment consists of a 166 MHz PC, a2-gigabyte hard drive, an Ethernet module, and theFDDI module. The STE software was developed bySanders using a PC-based operating systemand C++.

Flight Test FixtureThe flight test fixture was designed to address

three specific issues. First, eliminating errorsresulting from composite-based materials used inthe F-18 fuselage panels. A large percentage ofF-18 wing and fuselage panels are made of quasi-isotropic graphite-epoxy composite material. Thefixture was fabricated out of 2024 aluminum withthe goal of separating sensor from substratephenomenon during thermal and mechanicalin-flight loading. If the aircraft composite structuralcomponents were relied upon to validate the fiberoptic sensors, then significant strain measurementerrors, such as apparent strain, would be much moredifficult to eliminate. Second, it is difficult to flyrepeatable steady-state maneuvers which load theprimary aircraft structure sufficiently to generatelevels of strain necessary to validate the DSS. Theability to generate well-controlled, highly loadedand repeatable maneuvers was important inachieving both research quality and flight safetygoals. Finally, the development of a “loads” fixtureprovides a configuration that could be easily testedin both a flight environment and a laboratoryenvironment.

The FTF provided an extremely usefulconfiguration for verifying the accuracy of the DSSsensors. The FTF was designed to maximize theapplied strains within a limited flight envelope,therefore, it was to be installed and flown within itsdesigned flight envelope for a limited number offlights. Instrumentation attached to the FTF wouldbe active when the FTF was installed on theaircraft.

Installation and Integration of theSystem on the SRA

The installation and integration of the VHMsystem on the aircraft posed a number ofchallenges. The entire system contained newcomponents or configurations. This forced some

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unique aircraft installations. Both X-33 flighthardware and the configuration of the SRA aircraftwere modified to support this experiment.

VHM ComputerThe modified chassis design for the VHM

computer (fig. 2) resulted in a box that would not fitin a standard avionics rack. As a result the computerwas installed vertically on an instrumentation racklocated in the gun bay on the SRA. This installationsubjected the VHM computer to “off axis” loadingfor a significant portion of the test program.

VHM computer modifications includedincorporation of GEN II fiber optic connectors anda multi-channel single-mode fiber optic connectorfor the DSS. Software modifications were made tointegrate the system onto the SRA. Thesemodifications changed the computer operatingcommands and time synchronization from aMIL-STD-1553B interface to an analog switch tocommand the recorder and an RS-422 interface toprovide time synchronization. Additional changesincluded; changing to a single FDDI module forboth the RHN network and the recorder download,and the use of solid-state memory replacing thehard disk drive.

The original DSS laser module integrated intothe VHM computer exhibited thermal stabilityproblems. A redesigned optics module wasdeveloped and installed in the computer before thefirst dedicated DSS flights. Part of the redesignincluded software modifications that disabledautomatic thermal protection logic in the DSS lasermodule. This resulted in a requirement for groundcooling to be supplied to the gun bay during aircraftground operations to prevent the DSS laser fromoverheating while operating the system on theground. Airflow through the gun bay was sufficientto cool the VHM computer during flight. The othermajor DSS software modification involvedchanging the digital signal processor samplingfrequency in an attempt to improve the coherencyof the sampled data. This modification effectivelydecreased the sampling frequency of the DSSmodule to approximately one sample every90 seconds.

Remote Health Nodes and Fiber OpticInterconnects

The RHNs were distributed in two locations inthe aircraft. These components were designed forinstallation in locations in the X-33 that are exposedto extreme temperature ranges and highvibro-acoustic levels [8]. Their compact size andlarge thermal envelope made the choice of RHNlocation dependent on sensor location, not theenvironment of the RHN module. The RHNs weremodified slightly from the X-33 configuration. Themajor change to both RHNs was the change toGEN II optical interconnects. One RHN had theouter chassis changed to a composite chassis.

The main obstacle to integrating the RHN wasconnector placement on both ends of the RHN, asseen in figure 3. Fiber optic and power connectorswere located on one end of each RHN. The sensorinputs on two connectors were located on theopposite end of the RHN. This design made itdifficult to install and remove RHNs from theaircraft because it either required access to bothends of the RHN, or that RHN connectors be matedprior to mounting the RHN to the aircraft structure.

Routing of the HOB required that thebackshell and part of the connector be removedfrom the GEN II cable plant and a routing andprotection device supplied by the manufacturer beinstalled over the exposed multi-fiber termini. Thecable plant was then routed as part of a larger cableharness using a convoluted tubing cable guide. Afiber optic termini was damaged during this routingprocess. Routing of the cable plant continued afterthe termini was damaged. The damaged termini andapproximately 10 ft of excess ribbon cable wereremoved from the end of the cable plant and a newsection of cable plant, with an undamaged termini,was mechanically spliced to the installed cableplant. Excess fiber optic cable was coiled intoservice loops and stowed near the RHN or theVHM computer. A second multi-fiber termini wasdamaged during routine connector mate and demateoperations. The probable cause of the damage wasthe result of a misalignment between the connectorand multi-fiber termini when it was being mated tothe RHN. The connector with the damaged terminirequired that a “blind” mate of the cable plant to theRHN be accomplished. This damaged cable plantwas removed from the aircraft and returned to themanufacturer for repair. The cable plant was

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reinstalled in the aircraft upon its return from themanufacturer.

The single-mode fiber optic cable plant wasrouted from the VHM computer to the aircraftsensors and the FTF using convoluted tubing cableguides. The fiber was fusion-spliced after the fiber,sensors, and FTF were installed on the aircraft.

DSS Sensor LocationConventional and fiber optic strain sensors

needed to support the DSS experiment were routedto four separate aircraft locations: at the wing root,wing leading edge, wing skin, and FTF (fig. 4).Fiber optic sensors were installed at the wing rootto compare with strain gages already existing aspart of the conventional F-18 health monitoringsystem. Both fiber optic strain sensors and straingages were installed on the outboard leading edgetransmission to demonstrate that fiber could beapplied on geometrically challenging aircraftcomponents that were normally instrumented withstrain gages. The underside of the right wing wasinstrumented with 16 fiber optic strain sensorsspaced 1 ft apart on a single fiber. The fiber wasrouted from the inboard wing root area for 8 sensorsin the outboard direction, and returned inboard withthe remaining 8 sensors bonded at the samelocations of the first 8. The goal for thisconfiguration was to address surface attachmentissues on realistic aircraft composite materials, aswell as to establish what sort of precision colocatedfiber optic sensors had when compared with eachother at the same location. These three locationswere permanently attached to the aircraft. Thefourth location was the removable FTF mounted onthe centerline pylon. The design andinstrumentation philosophy of the fixture wasdescribed in the SRA Experiment Architecturesection of this paper.

DSS Sensor AttachmentThe fiber optic DSS sensors were bonded

directly to the aircraft (and FTF) structure usingmethods and materials developed specifically forthis experiment. Limited information was availablein the literature for bonding fiber optic sensors.Procedures and techniques were developed basedon strain gage technology and extensive flight andground test experience. Lessons learned from the

attachment of the fiber optic sensors for thisprogram are being documented.

RHN Sensor InstallationIntegration of analog sensors with the RHN

was relatively straightforward. A combination ofstrain gages, accelerometers, thermocouples, andpressure sensors were integrated with the RHNs.Sensor selection was based primarily on using thesame sensors specified in the X-33 mastermeasurement list. However, in some cases, COTStransducers that were not specified in the X-33master measurement list were used. The RHNsignal conditioning, sample rate, and gainconfigurations were loaded into each RHN prior tothe installation on the aircraft. These configurationswere modified periodically during the flightprogram as needed for data quality purposes. Therewere no serious problems integrating the selectedsensors with either RHN.

Once the sensors connected to the RHN bridgecompletion, their circuits were zeroed in order toremove bridge offsets from each channel. This wasaccomplished using ground system test equipment.Care had to be taken during this process because aninvalid system command issued to the RHN wouldresult in the loss of sensor load information.

Post Flight Data Management SystemThe final part of the VHM system to be

integrated was the interface between the VHMcomputer and the ground STE. This interface wasused after each flight to download data recorded onthe VHM computer and to initialize the disk afterthe data were downloaded. The interface proved tobe a challenge for the program. The FDDI recorderdump capability between the VHM computer andthe ground STE worked well for most of theprogram. There was one incident of the systemfailing to work properly. This problem was latertraced to a faulty processor in the VHM computer,not the FDDI link. Transfer of data between theVHM computer and the ground STE exhibitedcommunications problems related to timing anddata buffering. As a result of problems with theground STE about 20 percent of the data recordedby the VHM computer was lost. This problem wastraced to software problems with the ground STE.The flight program was completed before thatproblem could be fully identified and corrected.

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Flight Experiment Phase

Flight SummaryThe VHM experiment began flying on the

SRA at the end of September 1999. Flight testsconcluded in early April 2000. The experiment wasinstalled on the aircraft for 53 flights and flew forover 46 hours. These missions covered a largeportion of the SRA flight envelope. Figure 5 showsthe flight conditions accomplished during the testprogram. Thirteen of the 53 flights were dedicatedto testing the VHM system. Eleven of the thirteendedicated VHM flights were used to characterizethe DSS using the FTF. The remaining twodedicated flights were used to gather RHN data inregions of the aircraft flight envelope not coveredduring the nondedicated flights. The remaining40 flights were dedicated to testing otherexperiments being flown on the aircraft. Data werecollected during these 40 missions, but there wereno test maneuvers specifically dedicated to theVHM experiment.

Figure 5. Flight envelope.

Two of the fifteen dedicated flights were usedto gather RHN data in regions of the flight envelopethat were not reached during nondedicated flighttest, nor during FTF flights. Constant speed climbsand descents, and dynamic maneuvers wereperformed at eighteen flight conditions highlightedin figure 5. The aircraft was dynamically loaded to6 g and subjected to high pitch- and roll-ratemaneuvers during these flights. The aircraft reachedaltitudes of over 40,000 ft and speeds of Mach 1.6during these test maneuvers.

Eleven of the dedicated flights were flownwith the FTF installed on the aircraft. Of these

eleven flights, seven were dedicated to validatingthe DSS sensor measurements. The other fourflights were used to gather analog instrumentationdata from the conventional strain gages beingrecorded by the RHNs and the NASA DAS. TheFTF flights consisted of flying to a specific airspeedand altitude.

Flight Loading Approach—FTF FlightsThe fixture was loaded by applying an angle of

sideslip (b) to the fixture. The use of rudder trimand a fixed offset angle set at the FTF allowed b tobe added to flight loads. This loading conditioncould be held until a representative data samplecould be taken.

The FTF test points were flown using a buildup approach starting from 10 percent of the DLL ofthe flight test fixture and continuing in anincremental fashion until 100 percent DLL wasachieved. The percentage of DLL was determinedprior to flight through a series of ground tests on thefixture. These tests were performed to gather abaseline strain gage survey based on loadingconditions up to 115 percent of the DLL. Thesesame strain gages were monitored in flight todetermine the actual loading condition in real-time.

The FTF flights showed that repeatableloading conditions could be obtained consistentlyusing the FTF offset angle and rudder trim. Both theDSS fibers and analog RHN strain gages were flighttested using the FTF configuration and flightapproach.

Results and Discussion

The VHM system gathered data during theentire flight demonstration phase. The overallsystem performance suffered from a combination offaulty hardware, system integration, and systemprogramming problems. The system performance isnow discussed in detail.

The RHNs performed without failure duringthe entire flight program. The data correlated wellwith the NASA DAS but exhibited higher levels ofnoise on the individual channels. Figure 6 showselevated noise levels for colocated thermocouplesrecorded using the VHM System and the NASADAS. RHN-B data were lost for the first 21 flights

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as a result of integration problems with the FDDIring communications between RHN-A and RHN-B.This problem was tracked to a problem with themulti-mode cable plant between the two RHNs. Thecable plant was repaired and the system performedwell for the remaining 32 flights. RHN-A data werelost for nine flights because of a corrupt sensorsystem load in the RHN. The RHN continued toreport valid BIT status to the VHM during theseflights, but did not produce the requested data. Thiswas traced to the BIT logic in the RHN whichchecks the health and status of the individual RHNchannels, but does not check for valid system load.

This problem was repeated during groundoperations when the sensor system loads for bothRHNs were inadvertently corrupted during acalibration procedure. Both RHNs reported validBIT status without valid sensor system loads beingpresent.

Validation of the DSS experiment was onlypartially successful. There were a number of factorsthat contributed to this lack of success. The firstproblem was attributed to poor performance of theDSS laser and overall system. The DSS did notgenerate meaningful fiber optic strain data as aresult of the high technological risk associated withfielding a flight-rugged tunable laser source. Theexperiment objective to validate fiber optic sensingtechnology in flight was, therefore, not achieved.Thermal stability issues with the laser also resultedin a very low sample rate (~0.01 Hz) during theflight experiment, which required that the testcondition be maintained for a longer duration. Thisincreased the total number of flights required tovalidate the DSS. The second factor that contributedto the lack of success in validating the DSS was acombination of a compressed flight schedule andintermittent operation of the DSS module. This

Figure 6. Flight data.

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intermittent operation was not identified throughanalysis until flight testing of the DSS wascomplete. As a result, DSS data were only collectedfor 25–60 percent of the DLL cases.

The second series of four FTF flights werededicated to validating analog sensor data. The DSSfibers mounted on the FTF were not connectedduring the flight phase. These four flights werededicated to validating analog strain sensor data,after changes were made to sensor gains and bridgeoffsets in the RHN. The DSS was not connected,and the sample rates of the analog gages were muchfaster, therefore the duration of each test pointcould be reduced. Data were gathered for FTFloading cases up to 100 percent of the DLL.

The flight demonstration program succeededin collecting over 25 gigabytes of data from 53 testmissions. Data from these flights are being analyzedto fully assess the system performance andcapabilities.

Conclusions

The basic VHM system architecture was welldesigned and implemented. Experiment integrationand flight test showed the flexibility of theCOTS-based open architecture design. Aredesigned DSS module was integrated into theVHM computer after system integration began onthe aircraft. The COTS-based VME boards, FDDIinterface, and RHNs performed nominally after theinitial integration problems were solved.Acquisition of analog sensor data was successfulafter the RHN channels were correctly configuredwith the correct system gains and bridge offsets.Data from the RHN channels show a tendency forlevels of noise that are higher than noise levels ofdata recorded using the onboard DAS. Likelycauses for this noise include differences in samplerates, gains and signal conditioning. As analysis ofthe flight data continues the accuracy andrepeatability of individual sensor data will be morefully assessed.

There were some areas where designimprovements would have aided the systemintegration and operations efforts. Hardwaremodifications that should be considered include

implementation of a standard size air transportationrack chassis for the VHM computer and placementof RHN connectors at only one end of the chassis.Software issues that need to be addressed includemodifications to the RHN BIT logic to test thevalidity of sensor system loads. A rework of thedata transfer software in the ground STE is neededto avoid losing critical flight data during grounddata transfers.

The integration of the fiber optic cable plantsand DSS sensors on the aircraft presented their ownchallenges that included routing, splicing,termination, and attachment issues. The cable plantsworked well after they were installed on theaircraft, demonstrating that both single- and multi-mode fiber can be used to gather and transmitinformation in a flight environment. Follow-onwork is needed to address the installation andintegration issues that can make the technologycommercially feasible.

The DSS module did not meet expectations.Problems integrating the tunable laser into aflightworthy piece of hardware resulted in a systemthat performed intermittently. As a result, it was notpossible to fully assess the performance of the DSSin a flight environment. Follow-on research isneeded to develop a flightworthy tunable lasermodule that will meet the accuracy and sample raterequirements needed to implement structural healthmonitoring functions for future airframe andcryogenic tank applications.

Implementation and demonstration of the FTFshowed that the fixture could be integrated with theVHM experiment. It could also produce predictable,stable, repeatable loading conditions in a flightenvironment. The FTF can be used in the future toaid in the development and demonstration of newstructural sensors in a flight environment.

In general, the integration and flight programreduced the risk to the X-33 program throughdemonstrated system capabilities and identifiedareas where improvements or follow-on researchare needed to meet system requirements. Theprogram also demonstrated enabling technologiesthrough the flight demonstration of the single- andmulti-mode fiber optic cable plants. Finally it

11

developed tools, such as the FTF, that can be usedfor follow-on VHM research.

References

[1] Garbos, Ray, Leland Childers, Bruno Jambor,Oct. 1997, “System Health Management/ VehicleHealth Management for Future Manned SpaceSystems,” AIAA DASC Conference, SHM/VHM.

[2] Military Standard Fiber Optics Mechanizationof a Digital Time Division Command/ResponseMultiplex Data Bus MIL-STD-1773 (DOD), May20, 1988.

[3] Military Standard—Aircraft Internal TimeDivision Command/Response Multiplex Data Bus.MIL-STD-1553B (DOD), Sept. 21, 1978.

[4] Froggatt, Mark, September 1996, “DistributedMeasurement of the Complex Modulation of aPhotoinduced Bragg Grating in an Optical Fiber,”Appl. Opt., vol. 35 no. 25, pp. 5162–5164.

[5] Bodan, Patricia, Carl Bouvier, April 20–23,1998, “X-33/RLV Reusable Cryogenic Tank VHMUsing Fiber Optic Distributed SensingTechnology.” Proceedings of theIAA/ASME/ASCE/ASC Thirty-ninth Structures,Structural Dynamics, and Materials Conferenceand Exhibit, Long Beach, California,AIAA-98-1929.

[6] Melvin, L., B. Childers, R. Rogowski,W. Prosser, J. Moore, M. Froggatt, S. Allison, M.C. Wu, J. Bly, C. Aude, C. Bouvier, E. Zisk, E.Enright, Z. Cassadaban, R. Reightler, J. Sirkis, I.Tang, T. Peng, R. Wegreich, R. Garbos, W.Mouyos, D. Aibel, and P. Bodan, September 18–20,1997, “Integrated Vehicle Health Monitoring(IVHM) for Aerospace Vehicles,” Structural HealthMonitoring-Current Status and Perspectives,Technomic Publishing Co. Inc., edited by Fu-KuoChang, proceedings of the International Workshopon Structural Health Monitoring, StanfordUniversity, Stanford, California.

[7] Sitz, Joel, December 1992, The F-18 SystemsResearch Aircraft Facility, NASA TM-4433.

[8] Garbos, Raymond, William Mouyos,April 20–23, 1998, “X-33/RLV System HealthManagement/ Vehicle Health Management,”Proceedings of the IAA/ASME/ASCE/ASC Thirty-ninth Structures, Structural Dynamics, andMaterials Conference and Exhibit, Long Beach,California, AIAA-98-1928.

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NSN 7540-01-280-5500 Standard Form 298 (Rev. 2-89)

Prescribed by ANSI Std. Z39-18298-102

Flight Demonstration of X-33 Vehicle Health Management SystemComponents on the F/A-18 Systems Research Aircraft

WU 529-35-34-E8-RR-00-000

Keith A. Schweikhard, W. Lance Richards, John Theisen, WilliamMouyos, and Raymond Garbos

NASA Dryden Flight Research CenterP.O. Box 273Edwards, California 93523-0273

H-2435

National Aeronautics and Space AdministrationWashington, DC 20546-0001 NASA/TM-2001-209037

The X-33 reusable launch vehicle demonstrator has identified the need to implement a vehicle healthmonitoring system that can acquire data that monitors system health and performance. Sanders, a LockheedMartin Company, has designed and developed a COTS-based open architecture system that implements anumber of technologies that have not been previously used in a flight environment. NASA Dryden FlightResearch Center and Sanders teamed to demonstrate that the distributed remote health nodes, fiber opticdistributed strain sensor, and fiber distributed data interface communications components of the X-33 vehiclehealth management (VHM) system could be successfully integrated and flown on a NASA F-18 aircraft. Thispaper briefly describes components of X-33 VHM architecture flown at Dryden and summarizes the integrationand flight demonstration of these X-33 VHM components. Finally, it presents early results from the integrationand flight efforts.

Aircraft systems, Bragg gratings, Fiber optic sensing, Flight testing, Integratedvehicle health management, Intelligent systems, IVHM, Structural healthmonitoring

16

Unclassified Unclassified Unclassified Unlimited

December 2001 Technical Memorandum

Presented at the 19th Digital Avionics Systems Conference, Philadelphia, Pennsylvania on October 7-11,2000.

Unclassified—UnlimitedSubject Category—06

This report is available at http://www.dfrc.nasa.gov/DTRS/