1999 Research Engineering Annual Report

Compiled byEdmund Hamlin, Everlyn Cruciani, and Patricia PearsonNASA Dryden Flight Research CenterEdwards, California

August 2000

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1999 Research Engineering Annual Report

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iii

1999 Research Engineering Report

Table of Contents

Title First Author PagePreface Robert Meyer v

Research Engineering Directorate R

Organizational Chart vi

Research Engineering Directorate Staff vii

Supersonic Natural Laminar Flow Flight Research Daniel W. Banks 1

Infrared Flight Test Techniques Daniel W. Banks 2

HAVE RECKON Project Alex Sim 3

Mars Airplane Entry, Descent and Flight

Trajectory Development James Murray 4

Blunt Body Drag Reduction Using Forebody

Surface Roughness Stephen A. Whitmore 5

DC-8 Reduced Vertical Separation Minimum (RVSM)

Certification Edward A. Haering, Jr. 8

F-18 Stability and Control Derivative Estimation

for Active Aeroelastic Wing Risk Reduction Tim Moes 9

SR-71 Testbed Configuration Envelope Expansion Tim Moes 10

Flight Test of a Gripen Ministick Controller in an

F/A-18 Aircraft Patrick C. Stoliker 11

Autopilot Design for Autonomous Recovery of the APEX

Remotely Piloted Vehicle (RPV) Marle Hewett 12

AAW Risk Reduction Flights Using the PSFCC Computers John Carter 13

X-33 Avionics Integration & Real-time Nonlinear Simulation Bob Clarke 14

X-33 Landing Gear Dynamics Model Louis Lintereur 15

X-43A Non-Linear Monte Carlo Analysis Ethan Baumann 16

Unmanned Combat Air Vehicle Neil Matheny 17

X-43A Hypersonic Blended Inertial and Pneumatic

Angle of Attack Joe Pahle 18

F/A-18 Autonomous Formation Flight Curtis E. Hanson 19

F/A-18B Systems Research Aircraft (SRA) X-33 Vehicle

Health Monitoring System Risk Reduction Experiment Keith Schweikhard 20

Flight Test of an Intelligent Flight Control System on the

F-15 ACTIVE Michael P. Thomson 21

X-Actuator Control Test (X-ACT) Program Stephen Jensen 22

X-43A Antenna Patterns Mark W. Hodge 23

Reconfigurable VHF Broadcast Differential GPS Base Station Glenn Bever 24

DSP/GEDAE Signal Processing Algorithms Arild Bertelrud 25

Multiplexed Hot Films Arild Bertelrud 26

LABVIEW Based ESP Pressure Data Acquisition System Alma M. Warner 27

iv

Flight Instrumentation for X-33 Extended Range Test Terry Montgomery 28

Efficient Modulation Techniques

X-43A Flight Research Plans Overview

Don Whiteman

Griff Corpening

29

30

X-43 Fluid Systems Development Masashi Mizukami 31

X-43A Environmental Control Systems Performance Neal Hass 32

Advances in Leak Detection Technologies Neal Hass 33

Hyper-X Propulsion Flight Research Plans Ross Hathaway 34

Development of Scramjet Skin Friction Gages Trong Bui 35

Gust Monitoring and Aeroelasticity Experiment Stephen Corda 36

F-15B Propulsion Flight Test Fixture (P-FTF) Jake Vachon 37

F-15 Skin Friction Flight Test Trong Bui 38

X-43A Propulsion System Controls Update John Orme 39

Development of Remote Site Capability for Flight Research Lawrence C. Freudinger 40

AeroSAPIENT Project Initiated Lawrence C. Freudinger 41

Integration of Finite Element Analysis Predicted Strains with

a Strain Gage Loads Equation Derivation Package Clifford D. Sticht 42

Non-Invasive In-Flight Structural Deflection Measurement William A. Lokos 43

Aerostructures Test Wing David F. Voracek 44

Blackbody Lightpipe Temperature Sensor Development Larry Hudson 45

Spaceliner 100 Carbon-Carbon Control Surface

Test Program Larry Hudson 46

Radiant Heat Flux Gage Calibration System Characterization Thomas J. Horn 47

Thermostructural Analysis of Superalloy Thermal Protection

System (TPS) Dr. William L. Ko 48

Validation of the Baseline F-18 Loads Model Using

F-18/SRA Parameter Identification Flight Data DiDi Olney 49

PC Based Thermal Control System Alphonzo J. Stewart 50

Robust Control Design for the F/A-18 Active

Aeroelastic Wing Rick Lind 51

Application of Gain-Scheduled, Multivariable Control

Techniques to the F/A-18 PSFCC Mark Stephenson 52

Fiber Optic Instrumentation Development Lance Richards 53

1999 Technical Publications 54

Tech Briefs, Patents, and Space Act Awards, FY 99 60

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vii

Research Engineering Directorate Staff

Director Robert R. Meyer, Jr.Deputy Director Victoria A. RegenieAssistant Director Michael DeAngelisAssistant Director Edmund M. HamlinSecretary Everlyn Y. Cruciani

Branch Codes and Chiefs

RA – Aerodynamics Robert E. CurryRC – Dynamics and Controls Patrick C. StolikerRF – Flight Systems Bradley C. FlickRI – Flight Instrumentation Ronald YoungRP – Propulsion and Performance Stephen CordaRS – Aerostructures Michael W. Kehoe

Supersonic Natural Laminar Flow Flight ResearchSummaryAirfoil designs capable of passively maintaining laminarflow at supersonic speeds have been shown by theoryand small scale tests. Recent flight tests now prove thatthese designs can maintain large runs of laminar flow athigher Reynolds numbers in a harsh flight environment.Laminar flow was measured using a commerciallyavailable infrared detector adapted for flight. These testswere performed on the Dryden F-15B aircraft.

Objective¥ Acquire IR images of natural laminar flow up to M=2.0.¥ Obtain large runs of laminar flow at highest unitReynolds number possible with F-15B aircraft.¥ Determine conditions where laminar flow breaksdown.

ApproachUsing an aircraft mounted infrared camera, laminar flowwas measured on a test article mounted on the centerlinepylon of an F-15B aircraft. The IR camera measuressurface temperatures which change with the differentboundary layer states. The surface beneath the turbulentboundary layer will be warmer than that beneath thelaminar layer due to the higher convection of theturbulent layer with the freestream. The laminar flow,for these conditions, will be darker and the turbulentflow lighter.

The test article was fabricated from aluminum with aninsulating layer covering all but the first 1 to 2 inches ofthe leading and trailing edges. A splitter plate wasformed over the test wing to minimize disturbances fromthe bottom of the aircraft effecting the test surface.

ResultsLaminar flow was obtained up to full chord for the outer1/3 span and approximately 80% of the inner 2/3. Thelaminar flow was able to penetrate weak shock waves,but was typically terminated by strong shock waves.The strongest shock wave appeared to emanate from thecamera pod located outboard on the armament rail.

Status/PlansThe next flights will assess cross-flow disturbances byflying with a 30 degree leading edge. Future flights willobtain more detailed data by instrumenting the testarticle with surface pressures and thermocouples. Alsothe data recording system will be updated to record thefull digital images from the camera. A follow onprogram would increase the size of the test article andincorporate potential control surfaces to assess the effectsof higher Reynolds numbers and systems issues.

Contact:

Daniel W. Banks, DFRC, RA

(661) 258-2921 dan.banks@dfrc.nasa.gov

Supersonic Natural Laminar Flow testarticle mounted on F-15B aircraft.

Infrared image of Supersonic NaturalLaminar Flow test fixture.

Predicted transition pattern of test article(cross-hatched area denotes transition).

Freestream

Flow Trips

Transition Boundary

Infrared Flight Test TechniquesSummaryInfrared thermography is a very powerful technique toboth visualize and quantify flow fields over the speedrange. It is both global and non-intrusive. It is mostwidely used in aerodynamic flight test to measurelaminar/turbulent transition, but can also identify shocksand measure aeroheating. Three different techniques ofacquiring the IR thermograms are employed. These are afixed system on the aircraft to be visualized, a remotesystem on a second aircraft, and a ground based remotesystem employing long range optics. These systems andtechniques have been used to visualize flow fields fromsubsonic to hypersonic speeds.

Objectives• Acquire infrared thermograms in flight.• Process images for motion and spatial corrections, and enhancement.• Calibrate and measure surface temperatures• Identify and locate boundary layer transition, shock waves, and other flow phenomena.

ApproachBy using various IR systems and techniques flow fieldinformation has been acquired across the speed range.Remotely mounted aircraft systems have been used toacquire data from other unmodified aircraft in flight fromsubsonic through low supersonic speeds. Fixed systemshave been used to measure test articles at supersonicspeeds but are applicable for a variety of applications andspeed ranges. Ground based remote imaging (Langleylead for this effort) has been developed for measuringtransition and aeroheating of re-entering re-useablelaunch vehicles.

ResultsExcellent results have been obtained from each type ofsystem and at each speed range. The ability to locateboundary layer transition, shock waves, and measuresurface temperatures has been demonstrated.

The top figure shows transition on a subsonic wing. Thesurface is heated by solar radiation so that the laminarregions are lighter in color. The middle figure showstransition on a supersonic laminar flow test article. In thiscase the surface is heated by the freestream, so the laminarareas are darker. The bottom figure shows the spaceshuttle on re-entry. In this case the lighter areas denoteregions of higher surface heating (entire vehicle isturbulent).

Status/PlansContinue developing infrared systems and techniques forflight research.

Contact

Daniel W. Banks, DFRC, RA, (661) 258-2921Robert C. Blanchard, LaRC

T-34 wing imaged remotely by F-18 FLIR.

Infrared image of supersonic naturallaminar flow test fixture imaged with a

fixed IR system.

Freestream

Trip

TransitionBoundary

Transition Boundary

Freestream

Flow Trips

Space shuttle STS-96 imaged withground based IR.

HAVE RECKON ProjectSummaryWhen an Uninhabited Aerial Vehicle, UAV, is piloted from aground control station, it is referred to as a Remotely PilotedVehicle. A delay of typically 0.1 seconds is incurred as thevehicle processes the uplink signal and the vehicle responseis prepared for downlink. When an RPV is flown over thehorizon, the uplink and downlink signals still need to berelayed back to the ground station. A desirable relay is to usea satellite as the link. However a satellite link to the UAVand back often results in an additional signal transport delayof about 0.4 seconds. This 0.5 seconds total delay is enoughto significantly degrade the handling qualities of an RPV.For a high-gain piloting task, this may even be enough lag tomake the vehicle uncontrollable.

The USAF Test Pilot School, TPS, proposed to enhance thehandling qualities by using a model-based predictivealgorithm to compensate for system time delay. Thealgorithm would reside within the ground station andprovide the ground pilot with predictive enhancementsoverlaid on the ground monitor. The compensation wasfunctionally independent of the UAV. To conduct a low-cost flight research experiment, TPS contacted the DrydenSubscale Facility (model shop). Dryden agreed to providethe ÒUAVÓ, safety pilot, ground cockpit, telemetry van andother lesser assets for the flight tests. The scope of theexperiment was limited to the longitudinal axis.

ObjectiveThe USAF TPS general objective was to determine theimprovement in UAV handling qualities when using amodel-based predictive algorithm to compensate for systemtime delay. NASA DrydenÕs primary objectives were toprovide the TPS with an aerodynamic model of the flightvehicle, integrate the vehicle systems, and safely conduct theflight tests.

ResultsDryden chose a large utilitarian model, called theMothership, that was capable of lifting 25 lbs of payload andstill remain within the Dryden definition of a model airplane.A 16-channel data system was carried in a pod under themain fuselage. An airdata noseboom was mounted on thenose. A forward-looking video camera was installed in a podabove the wing. Measurements of elevator, pitch rate, andangle of attack were downlinked on the audio channel of thevideo signal. A telemetry van was used to separate the datafrom the audio channel and put it onto an ethernet line to anUSAF Unix workstation next to the research cockpit. Theworkstation generated the transport lag by passing a delayedvideo frame to the monitor. The workstation was also usedto generate the predictive algorithm and the related monitordisplay. The research cockpit and video monitor was in aseparate van. Altitude and airspeed were merged with thevideo signal onboard the model and overlaid onto the uppercorners of the video image.

An initial longitudinal aerodynamic model of the Mothershipwas calculated using vortex-lattice techniques. Parameterestimation analysis of elevator doublet flight data was thenused to update the update the initial model. A total of 28

flights were flown for the initial functional checks and forparameter estimation.

The NASA model pilot, referred to as the ÒsafetyÓ pilot, stoodoutside of the van containing the ground cockpit. He was pilotin command and insured that the vehicle was kept within hisvisual range. He made all the takeoffs, landings, climbs,descent, and turns. On the straight leg of the flight pattern, hewould give longitudinal stick and throttle control to theresearch pilot within the van. He could take back control atany time.

For the research flying of the program, 62 data sorties wereflown over 16 test days for 20.5 hours of flight time. TPS wasable to show that their model-based predictive algorithmimproved handling qualities of the model flown as an UAV.With the predictor, the time delay tolerated was significantlyincreased. Or at the same time delay, pilot workload wassignificantly decreased. Including the initial checkout flights,a grand total of 90 flights were flown.

Status/PlansDryden started in May, 1999. The research flight testingstarted in mid September and ended in mid October, 1999.USAF TPS technical report was finalized in December 1999.The successful completion of the program has potential to leadto a joint USAF/NASA program using a much larger vehicle.

Mothership Utility Model with 10 ft Wing Span

ContactsAlex Sim, DFRC, RA, (661) 258-3714Tony Frackowiak, DFRC, AS&M, (661) 258-3473Maj. Andrew Thurling, USAF TPS, (661) 277 3131

Mars Airplane Entry, Descent and Flight Trajectory Development

End of pullup,h = 2700 m,

M = 0.76

Airplane unfolds, levels wings,and starts

pullup

Airplane release h = 6400 m,

M = 0.54

Heatshield releaseh = 6900 m, M = 0.9

Parachute deploymenth = 7700 m, M = 1.71

Atmospheric entryh = 136 km,

Va = 5500 m/sec,

γ = – 22°

Mars Airplane Entry, Descent, and Flight Trajectory

SummaryThe Mars Airplane project was to provide the firstopportunity for sustained lifting atmospheric flight on aninterplanetary mission. For the mission, the aircraft iscarried to the planet folded inside a small (75 cmdiameter) aeroshell which makes a direct entry into theMartian atmosphere. Critical to the mission was thetransition from the hypersonic, ballistic aeroshell to thesubsonic lifting airplane. For initial design purposes abaseline Entry, Descent and Flight (EDF) trajectoryprofile was defined to consist of the following elements:ballistic entry of the aeroshell, supersonic deployment ofa decelerator parachute, release of the heatshield, release,unfolding, and orientation of the airplane, and executionof a pullup maneuver to achieve trimmed, horizontalflight.

Given the predicted mass and aerodynamics of theaeroshell and airplane, and the predicted Martianatmospheric and topographical characteristics, asimulation of the EDF trajectory was built using theProgram to Optimize Simulated Trajectories (POST).The trajectory optimization feature of POST was used tofind the control strategy which maximized the surface-relative altitude of the airplane at the end of the pullupmaneuver, subject to a number of design constraints.

ObjectivesThe early development of a baseline (i. e. first-generation)EDF trajectory was important to provide the MarsAirplane design team with essential feedback early in thedesign cycle. The initial objective was to develop an EDFtrajectory consistent with the baseline vehicle design data,and to optimize the trajectory within the allowable designspace. With the optimal trajectory defined, the objectivewas to identify the sensitivity of the trajectory to design

parameter variations. A final objective was to use theresults of sensitivity analysis to develop a second-generation control strategy.

ResultsFor the baseline vehicle design, viable EDF trajectorieswere developed for sites in both the northern and southernhemispheres; altitudes at the end of the pullup maneuverwere above 2 km. However, the high elevation of thedesired science mission site (Parana Valles) precludedcompletion of the pullup maneuver before surface impact.

The performance metric (the surface-relative altitude atthe end of the pullup maneuver) was a most sensitive toairplane mass, airplane lift and drag coefficients, andmaximum Mach number allowed during the pullup. Notsurprisingly, decreasing airplane mass, increasing liftcoefficient, and increasing the allowable maximum Machnumber yielded net altitude gains. Increasing the airplanedrag coefficient to yield a net altitude increment. Theperformance metric showed only small sensitivity to otherdesign parameter variations studied.

With the addition of airplane drag as a design variable, asecond-generation control strategy showed that increasingdrag significantly during the pullup maneuver yielded anet altitude increment of about 1.7 km.Status/PlansThe Mars Airplane project was cancelled near the end ofCY99. However, documentation of current results andsome closeout work continues.

ContactsJames Murray, DFRC, RA, (661) 258-2629Paul Tartabini, LaRC, (757) 864-787

Blunt Body Drag Reduction Using Forebody SurfaceRoughness

IntroductionCurrent proposed shapes for trans-atmospheric launch and crew-returnvehicles like the X-33, X-34, X-38, and"Venture-Star" have extremely large baseareas when compared to previoushypersonic vehicle designs. As a result,base drag - especially in the transonicflight regime -- is expected be very large.The unique configuration of the X-33,with its very large base area and relativelylow forebody drag, offers the potential fora very high payoff in overall performanceif the base drag can be reducedsignificantly.

Drag Reduction StrategyFor blunt-based objects whose base areasare heavily separated, a clear relationshipbetween the base drag and the ÒviscousÓforebody drag has been demonstrated (ref.1, 2). (Figure 1) As the forebody drag isincreased; generally the base drag of theprojectile tends to decrease. This base-dragreduction is a result of boundary layereffects at the base of the vehicle. Theshear layer caused by the free-stream flowrubbing against the dead, separated air inthe base region acts as a jet pump andserves to reduce the pressure coefficient inthe base areas. The surface boundary layeracts as an "insulator" between the externalflow and the dead air at the base. As theforebody drag is increased, the boundarylayer thickness at the aft end of theforebody increases, -- reducing theeffectiveness of the pumping mechanism -- and the base drag is reduced.

Configurations with large base dragcoefficients, necessarily lie on the steepportion of the curve a small increment inthe forebody friction drag should result ina relatively large decrease in the base drag.Conceptually, if the added increment inforebody skin drag is optimized withrespect to the base drag reduction, then itmay be possible to reduce the overall dragof the configuration.

The LASRE Drag Reduction ExperimentThe LASRE experiment (ref. 3) was aflight test of a roughly 20% half-spanmodel of an X-33 forebody with a singleaerospike rocket engine at the rear. Theentire test model is mounted on top of aSR-71 aircraft. In order to measureperformance of the Linear Aerospikeengine under a variety of flight conditions,the model was mounted to the SR-71 witha pylon which was instrumented with 8load-cells oriented to allow a six-degree-of-freedom measurement of the totalforces and moments. The model was alsoinstrumented with surface pressure portsthat allowed the model profile drag to bemeasured by numerically integrating thesurface pressure distributions.

The LASRE drag reduction experimentsought to increase the forebody skinfriction and modify the boundary layer atthe back end of the LASRE model.Clearly, one of the most convenientmethods of increasing the forebody skindrag is to add roughness to the surface.LASRE results verified that as theforebody drag is increased; generally thebase drag of the projectile tends t odecrease. LASRE tests verified that thisdrag reduction persists through transonicand well into the supersonic flight regime.Pre-flight analyses predicted that a tradeoff of viscous forebody drag to base dragmay make it possible to achieve a net dragreduction by adding roughness to theforebody skin.

Unfortunately, even though the LASREexperiments demonstrated a strong basedrag reduction effect; the net drag was notreduced. The roughened forebody causedthe forebody pressures to rise and offsetthe base drag benefits, which were gained.At this time it is unclear whether thisforebody pressure rise was a fault of themanner in which the forebody grit wasapplied, or whether it is even possible t oachieve a net drag reduction. Further tests,under a more controlled flowenvironment, must be conducted t oresolve this uncertainty.

Blunt Body Drag Reduction Using Forebody SurfaceRoughness

Wind Tunnel Tests A series of low-speed wind tunnel tests is currently beingconducted to prove the existence of thiselusive Drag Bucket. In these tests a two-dimensional cylinder with a blunt afterbody is being tested (Figure 2). For thesimple 2-D tests total body forcemeasurements will be determined bynumerically integrating surface pressuresand by skin drag calculations made usingthe boundary layer velocity profilemeasurements. High-frequency wakepressure (Strouhal number)measurements will also be obtained. Byadding micro-machined surface overlays(Figure 3) to increase the surfaceroughness it is hoped that the overall dragof the configurations can be reduced. Thepredicted drag reduction is shown in figure4. The body flow characteristics predictedby CFD analyses are shown in figure 5.The trailing Von-Karman Òvortex-streetÓwake is clearly visible. Significantinstrumentation system development wasachieved during FY Ô99. Test resultsconfirming these predictions should beavailable by mid year FY Ô00.

References

1) Hoerner, Sighard F., Fluid DynamicDrag, Self-Published Work, Library ofCongress Card Number 64-19666,Washington, D.C., 1965, pp. 3-19, 3-20,15-4, 16-5.

2) Saltzman, Edwin J., Wang, Charles K.,and Iliff, Kenneth W., Flight DeterminedSubsonic Lift and Drag Characteristics ofSeven Blunt-Based Lifting-Body andWing-Body Reentry VehicleConfigurations, AIAA Paper # 99-0383,1999.

1) Whitmore, Stephen A., and Moes,Timothy R., A Base Drag ReductionMethods on the X-33 Linear Aerospike SR-71 Experiment (LASRE) Flight Program,AIAA 99-0277, January 1999.

Figure 1: Visualization of the "Drag Bucket"

1.000.1000.0100.0

0.2

0.4

0.6

0.8

CD, base

CD, total

"Viscous" Forebody Drag Coefficient

X-15

X24-B

M2-F3

X24-A

M2-F1 HL-10

Shuttle

HL-10

M2-F1

X24-AM2-F3

X24-B

ShuttleX-15LASRE

LASRE

Tot

al D

rag

Coe

ffic

ient

5.00"

1.00"

0.75"

0.20"

Top View

Side View

Figure 2: Wind Tunnel Model for Base Drag / Forebody-Roughness Study:

30°

60°

1.50"

3.0" 1.375"

0.250"0.250"

0.433"

0.5" radius

0.75"

3.95"

Etched Surface

h

0.1"

Figure 3: Inset of Bar Grid Pattern

tl

Tunnel Model

Blunt Body Drag Reduction Using Forebody SurfaceRoughness

Table I: "Bar-Grid" Roughness Dimensions (See Inset Diagram) (κs ~ Equivalent Sand Grain Roughness of Plate

0.0020 0.0020 0.0020 0.0064 ± .0005 0.0020 0.0040 0.0020 0.0114 ± .0005 0.0020 0.0060 0.0050 0.0206 ± .0005 0.0050 0.0120 0.0050 0.0330 ± .0010 0.0100 0.0150 0.0100 0.0450 ± .0010 0.0100 0.0200 0.0100 0.0570 ± .0010 0.0200 0.0400 0.0200 0.1140 ± .0015 0.0200 0.0800 0.0200 0.1911 ± .0015 0.0400 0.1000 0.0400 0.2721 ± .0020 0.0400 0.1500 0.0400 0.3660 ± .0020

Forebody Drag Coefficient

Dra

g C

oeff

icie

nt

0.05 0.10 0.15 0.20 0.250.00 0.30

0.20

0.25

0.30

0.35

0.40

0.15

0.45 CD, Base

All laminar

Smooth, transitional

All turbulent

Rough, transitional

CD, Total

All laminar

Smooth, transitional

All turbulent

Rough, transitional

Figure 4: Predicted Drag Reduction for Wind Tunnel Model

Strouhal

Figure 5: CFD Results for Wind Tunnel Model Showing Von-Karman

DC-8 Reduced Vertical Separation Minimum (RVSM) Certification

SummaryThe NASA Dryden DC-8 Airborne Science Laboratory(N817NA) performs research around the globe, recently insupport of the SAGE III Ozone Loss and Validation Experiment(SOLVE). This experiment utilized a large region of airspace,the North Atlantic (NAT) airspace corridor, which is subject toReduced Vertical Separation Minimum (RVSM) requirements.These requirements allow aircraft traffic to be separated by1,000 feet vertically between 29,000 and 41,000 feet MSL, ascompared to the usual 2,000 feet separation. RVSM non-groupaircraft compliance requires ±160 feet pressure altitudeaccuracy. RVSM allows greater traffic density whilemaintaining safe aircraft separation.

A commercial service for RVSM certification was considered,but involved significant modification to the aircraft, high cost,and an unacceptable schedule impact to Airborne Scienceresearch commitments. The approach taken was done internallywith insignificant aircraft modification, minimal cost, and littleschedule impact.

ObjectiveObtain RVSM certification through an airdata calibration of theDC-8 static pressure system achieving ±160 feet accuracy.

ApproachRVSM quality airdata computers were installed in the aircraft,and these data were recorded using the DC-8Õs Data Acquisitionand Distribution System (DADS). These airdata computers havea worst case avionics error of 85 feet after 24 months. For thecalibration flights a carrier-phase differential global positioningsystem (DGPS) receiver and antenna was employed. The DGPSgave geometric altitude of the aircraft to an accuracy better than2 feet. A pressure calibration of the atmosphere on flight testdays was determined by data from a network of rawinsondeweather balloons, synoptic analysis, and surface observations.

By combining DGPS geometric altitude with the pressurecalibration of the atmosphere, the true pressure altitude of theaircraft is determined. This is compared to the airdata computermeasurement with no error corrections to determine the staticsource error correction (SSEC) required to null the pressurealtitude errors. The SSEC for both the pilot and co-pilot systemswere then incorporated into the airdata computers, and checkedon a verification flight. The SSEC is a function only of Machnumber.

The DC-8 was flown near maximum speed (Mach 0.48 to 0.54)at 500 feet above Rogers Dry Lakebed in steady flight. Thesedata gave the SSEC with minimal uncertainty of the atmosphericpressure calibration. Most of the flight data was taken at 29,000to 41,000 feet in stabilized flight between Mach 0.51 and 0.86.The near-ground data were used to adjust the high altitude datafor small temperature biases on the rawinsonde balloons. DGPSdata taken during constant airspeed turns were used to measurewinds independent of the rawinsonde balloons.

Autopilot operation was verified during stabilized flight toconform to the ±65 feet RVSM requirement.

ResultsOne calibration flight was flown to determine the SSECrequired to null pressure altitude errors with the aircraft in aclean configuration. These data yielded error residuals of±15 feet for both the pilot and co-pilot static pressure system.

A verification flight was performed with a variety of airbornescience probes external to the aircraft, including a largenacelle about 5 feet from the static pressure ports. Thisconstituted a worst case scenario of possible SSEC shifts.The maximum residual error on this flight was 73 feet, andwhen combined with the worst case avionics error of 85 feetand the DGPS accuracy of 2 feet gives a total error of 160feet, just meeting the RVSM requirement. The errors wouldbe considerably less if the airborne science probe near thestatic ports is removed or relocated. Autopilot operation wasdemonstrated to be within ±30 feet.

RVSM certification was granted on November 18, 1999.

DC-8 Airborne Science Laboratory (N817NA)

Current RVSM airspace in North Atlantic

ContactsEdward A. Haering, Jr., DFRC, RA, (661) 258-3696Edward H, Teets, Jr., DFRC, RA, (661) 258-2924David A. Webber, DFRC, OE, (661) 258-7541

F-18 Stability and Control Derivative Estimationfor Active Aeroelastic Wing Risk Reduction

SummaryThe NASA Dryden F-18 System Research Aircraft(SRA) was used to obtain stability and controlderivatives from flight data for a baseline F-18 aircraft.This work was done at the higher dynamic pressurerange of the F-18 envelope in support of a future F-18program known as Active Aeroelastic Wing (AAW).The AAW technology integrates vehicle aerodynamics,active controls, and structural aeroelastic behavior tomaximize vehicle performance. In particular, the goalof the AAW project will be to maximize thecontribution of a reduced-stiffness F-18A wing to rollrate performance. In order to support the technology,changes to flight control computers and software will berequired and therefore a good understanding of thebasic F-18 individual control surface effectiveness isessential The SRA was used to provide thatunderstanding.

ObjectiveThe primary objective of this program was to obtaindetailed understanding of the effectiveness of eachcontrol surface at the various test conditions. Theresults can then be used to update the aerodynamicmodel for the F-18 and therefore improve theeffectiveness of the control laws being developed forthe AAW aircraft.

ApproachAn on-board excitation system (OBES) was used toprovide uncorrelated single surface input (SSI) doubletsequences. Longitudinal maneuvers included leadingedge flap (LEF), trailing edge flap (TEF), symmetricaileron, and symmetric horizontal tail SSIs. Lateral-directional maneuvers include rudder, differential LEF,differential TEF, aileron, and differential tail SSIs. Thepilot would initiate the maneuver sequence from thecockpit and the OBES would command the SSIdoublets. In some cases, the surfaces were moved incombinations which are not used by the basic F-18control laws. These included symmetric LEF, TEF, andaileron deflections at high speeds. Data was analyzedpost-flight using an output-error parameter estimationalgorithm.

ResultsA complete set of stability and control derivatives hasbeen obtained for 20 different Mach/Altitude testconditions. The effect of symmetric aileron deflectionon pitching moment in shown in figure 1. As can beseen, the flight determined results (symbols) show thatthere is more pitching moment effectiveness than

StatusThe results of this study are being used to update the F-18aerodynamic model for simulation use AAW control lawdevelopment.ContactsTim Moes, DFRC, RA, (661) 258-3054Greg Noffz, DFRC, RA, (661) 258-2417(NASA TM currently in production)

predicted by the simulation (especially at subsonic Machnumbers). Control surface rolling moment Ò reversalÓ was ofspecial interest to the project. Figures 2 and 3 show TEF andaileron rolling moment derivatives. As seen in figure 2, TEFreversal was measured in flight at Mach 0.95 for altitudesbelow 10,000 ft. Aileron reversal was not measured in flight(figure 3), but reduced aileron effectiveness was clearly seen asthe subsonic Mach number increased and altitude decreased.

Mach number0.9 1.0 1.1 1.2 1.3

Fig. 1 Pitching moment derivative due to symmetric aileron

Fig. 2 Rolling moment derivative due to differential TEF

Fig. 3 Rolling moment derivative due to aileron

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StatusThe SR-71 testbed is capable for flight testing newaerodynamic and propulsion experiments at Machnumbers up to 3.2ContactsTim Moes, DFRC, RA, (661) 258-3054Stephen Corda, DFRC, RP, (661) 258-2103(NASA TM currently in production)

Baseline SR-71

SR-71 Testbed Configuration Envelope ExpansionSummaryA four flight envelope expansion program was conducted onNASA DrydenÕs SR-71A aircraft in the ÒTestbedÓconfiguration. This configuration included a large Òcanoe-likeÓ structure and reflection plane mounted on top of theSR-71 as seen in figure 1. This structure was previouslyused to carry the Linear Aerospike SR-71 Experiment(LASRE). Widespread interest has been expressed in usingthe testbed configuration for flight test of various propulsionand aerodynamic experiments at Mach numbers of up to 3.2.Consequently, NASA conducted the envelope expansion testprogram to look at the configuration performance, thermalenvironment, and stability and control (S&C).

ObjectiveDemonstrate maximum performance and stability andcontrol characteristics of the SR-71 testbed configuration.

ApproachThe approach was to incrementally build up Mach numberon each flight. Longitudinal and lateral-directional doubletswere flown at various Mach/Altitude conditions to obtain theS&C data. A post-flight output-error parameter estimationalgorithm was used to obtain S&C derivatives.

Results

Performance:The added aerodynamic drag of an experiment is mostcritical in the transonic acceleration, where excess thrust isat a minimum. Poor transonic acceleration can affect themaximum Mach number attainable since less fuel isavailable for the acceleration. On ÒhotÓ days, i.e. thetemperature versus altitude profile is well above theStandard Day profile, the SR-71 J58 engine thrust issignificantly reduced, thereby significantly reducing themaximum Mach capability. The SR-71 Test Bedconfiguration successfully reached Mach 3.0 on a ÒhotÓ day.Analysis has shown that Mach 3.2 is attainable, however,flight demonstration of this was not accomplished due to anin-flight system failure. The previous LASRE programreached a maximum Mach of 1.8 in flight. Analysisshowed that the LASRE configuration could have achievedMach 2.5 for a standard day temperature profile. Theadditional drag due to the part of the LASRE experimentlocated above the reflection plane is shown in figure 2.Future experiments desiring Mach numbers greater than 2.5need to have drag increments less than that shown infigureÊ2.

Stability and ControlThe largest S&C concern for the testbed configuration wasreduced direction stability, Cnb. Figure 3 shows the flightdetermined Cnb values for the baseline and testbed SR-71configuration. At Mach 2.1 the testbed configuration showed asteep decline in open-loop directional stability. Cnb did,however, remain positive up through Mach 3.0. The closed-loop directional stability provided by the stability augmentationsystem (SAS) was significantly better due to side accelerationfeedback. Piloted simulations of engine unstart events weredone that showed that the configuration was safe to fly with thereduced stability.

Fig. 1 - SR-71 Testbed configuration.

Fig.2 - Drag coefficient due to LASRE experiment

Ref area = 1605 ft2

Fig. 3 - Directional Stability CoefficientMach number

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Flight Test of a Gripen Ministick Controller in an F/A-18 AircraftSummaryIn March of 1999 five pilots conducted a handlingqualities evaluation in an F/A-18 research aircraft of thesmall displacement center stick controller developed forby Sweden for the JAS-39 Gripen. The ProductionSupport Flight Control Computers (PSFCC) provided theinterface between the controller hardware installed in theaft cockpit and the standard flight control laws for theF/A-18 aircraft.

ObjectiveThe primary objective of the flight research program wasto assess any changes in the handling qualities of the F/A-18 aircraft as a result of the mechanical characteristics ofthe ministick. The secondary objective was ademonstration of the capability of the PSFCC to supportflight test experiments.

ApproachThe ministick controller and demodulator box outputsingle channel pitch and roll stick commands. These DCsignals were input into the PSFCC analog inputs.Software was developed to perform cross-channel datalinks, signal selection and command scaling. The signalswere scaled to be similar to the maximum inputs of thestandard F/A-18 control stick. Because the experimentwas to assess the effects of the mechanical characteristicsof a small displacement center mounted control stick, theoriginal software deadbands and stick shaping were used.

A five flight test program was conducted using fivepilots. General comments and handling qualities ratingswere collected. The flight testing consisted of thefollowing maneuvers: doublets, frequency sweeps, bankattitude captures, pitch attitude captures, echelonformation flight, column formation flight, grossacquisition, and fine tracking. There were three phases tothe echelon formation: gentle maneuvering, vertical

ContactsPatrick C. Stoliker, DFRC, RC, (661) 258-2706John Carter, DFRC, RC, (661) 258-2025

captures and more aggressive maneuvering. The threephases for column formation flight were: gentlemaneuvering, more aggressive maneuvering, and lateralcaptures.

ResultsCooper Harper ratings from the pilots are summarized inthe table below. The pilot comments consistently notedthat the stick was very sensitive in roll with sometendency to ratcheting. This could be mitigated bymodifying the stick shaping and deadband. The generalpilots comments on the stick were favorable. It was notedthat it was extremely easy to put in full amplitude inputs.

StatusData analysis is being completed and a technical reportsare being prepared to document the flight test, compareresults with handling qualities criteria, and to describethe PSFCC implementation and testing process.

Ministick installed in the aft cockpit of an F-18Pilot A B C D EEchelon Formation Phase 1 4 3 2 3 3Echelon Formation Phase 2 4 5 3 2 4Echelon Formation Phase 3 4 to 7 5 4 4 5Column Formation Phase 1 4 3 2 4 4Column Formation Phase 2 3 3 2 4 5Column Formation Phase 3 3 to 4 5 5 4 4Gross Acquisition n / a 6 to 7 2 2 4Longitidinal Fine Tracking n / a 3 2 2 3Lateral Fine Tracking n / a 4 2 3 6

Summary of pilot ratings

Autopilot Design for Autonomous Recovery of the APEXRemotely Piloted Vehicle (RPV)

Summary The APEX lost-link autopilot is used when uplinked pilotcommand signals to the RPV flight controls are lost. Thisresults in the APEX RPV going into autonomous flightcontrolled by the lost-link autopilot. During this lost-linkevent, the autopilot automatically returns the RPV to a safelanding or until the pilot command uplink signals are safelyreestablished.

Objective Develop and demonstrate a reliable lost-link instrumentarrival, approach, and landing procedure that can be flownautonomously by the lost-link autopilot or by the test pilotfrom anywhere on the flight test range, and whichconservatively remains within the RPV flight envelopemaneuver capabilities.

Justification Range safety requires RPVs to remain within the flighttest range confined to RPV airspace boundaries either inremotely piloted or autonomous flight mode or risk flighttermination. The lost-link autopilot would satisfy rangesafety airspace requirements by flying a conservativeinstrument flight procedure that remains well within thoseairspace boundaries during any period when uplink pilotflight command signals are lost.

Approach The lost-link autopilot is programmed to perform apower-off, nonprecision global positioning system(GPS)arrival, approach, and landing procedure. This instrumentapproach procedure is also flown by the test pilot as aconservative set of maneuvers to bring the RPV back afterthe flight test. If lost-link occurs at anytime during theapproach, the lost-link autopilot automatically completes theinstrument approach followed by a safe landing. The stand-alone GPS instrument approach procedure is designed toFederal Aviation Administration specifications resulting inthe required RPV approach maneuvers being well withinflight envelope capabilities.

Results In the illustrated time history, the lost-link autopilot isexecuting the lost-link arrival, approach, and landingprocedures. This set of maneuvers starts with the RPVarriving at the initial approach fix, followed by an entry to aholding pattern with the inbound leg aligned with the finalapproach to the runway used for landing. The unpoweredAPEX RPV remains in holding until the approach decisionaltitude is reached, and then executes a final approach to therunway and lands. The lost-link autopilot can successfullyexecute this GPS instrument approach procedure from anywhere on the flight test range, and remain within the RPVflight envelope.

Time history of lost-link autopilot arrival,approachand landing

Benefits¥Reduced pilot work load during arrival and approachprocedures to landing because the lost-link autopilot isavailable to perform the same instrument approachprocedures.¥ Satisfies range safety requirements to keep the RPVwithin defined flight test airspace boundaries.¥Reduced RPV flight test risk by reliably bringing the RPVback to a safe landing during lost-link.

Status The APEX lost-link autopilot will be reconfigured to beused on the blended wing body flight test program.

ContactsMarle Hewett, DFRC, RC, (661) 258-7445William Figueiredo, DFRC, RC, (661) 258-2001Chong Lee, DFRC, RC, (661) 258-3957

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AAW Risk Reduction flights using the PSFCC Computers

SummaryThe Production Support Flight Control Computers (PSFCC)were used to place aerodynamic parameter identificationinputs on individual surfaces for a standard F/A-18. Theseinputs were executed at high dynamic pressure conditions torefine the aerodynamic and loads models used for the ActiveAeroelastic Wing (AAW) program.

ObjectiveThe objective of this effort was to enhance the quality of theF/A-18 aerodynamic and loads models for use on the F/A-18Active Aeroelastic Wing program.

JustificationThe Active Aeroelastic Wing program will place a moreflexible wing on an F/A-18 to demonstrate increasedperformance using wing flexibility. The load andaerodynamic models used for this program are based on thebest current F/A-18 models with modifications for theincreased flexibility. Data from this flight program is beingused to verify the F/A-18 models within the ActiveAeroelastic Wing test envelopes (see below).

ApproachThe research processor in the PSFCC is programmed to adddoublets and Schreoder frequency sweeps to individualsurface commands at pilot request (right). These doubletsand surface commands were executed at 20 flight conditionswithin 2 regions which will be used by the ActiveAeroelastic Wing program. The test vehicle had standardinstrumentation for rates, accelerations, surface positions,and airdata. In addition, loads instrumentation was installedto measure various loads and hinge moments.

ResultsThe flight test results showed significant differencesbetween the existing F/A-18 models for loads andaerodynamics and the flight derived values. The predictivemodels for the F/A-18 are currently being modified to reflectthe flight test performance of the aircraft.

Benefits¥Reduced risk associated with the Active Aeroelastic Wingprogram¥Increased model fidelity within the Active AeroelasticWing regionsStatus The flight test program is over, and the remainder of thedata reduction is being performed. F/A-18 models foraerodynamics and loads are being refined based on theflight test data.

ContactJohn Carter, DFRC, RC, (661) 258-2025

F/A-18 Test Vehicle

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X-33 Avionics Integration & Real-time Nonlinear Simulation

An X-33 Vehicle Mission Computer (VMC).

Comparison of HIL and SIL X-33 six-DOFsimulated flight from liftoff at Edwards AFB tolanding at Michael site in Utah.

ContactsBob Clarke, DFRC, RC, (661) 258-3799,Cathy Bahm, DFRC, RC, (661) 258-3123,Louis Lintereur, DFRC, RC, (661) 258--3307Richard Maine, DFRC, RC, (661) 258-3316,Peggy Williams, DFRC, RC, (661) 258-2508

Summary The X-33 program will prototype new technologiesrequired for design of a Reusable Launch Vehicle (RLV). TheX-33 will be an unmanned vehicle, launched vertically, reachingover 200,000 feet altitude at speeds approaching Mach 10. Thevehicle will operate autonomously from launch to landing. Someof the technologies to be demonstrated are: metallic thermalprotection system, linear aerospike engines, and an integratedvehicle health monitoring system. One of the ways that NASADryden is supporting this activity is through the development ofan X-33 Integration Test Facility (ITF) lab.

Objective The ITF lab provides an essential role in the avionicssystems development of the X-33. The initial phases haveincluded the development and test of a real-time, nonlinear six-DOF simulation, that supports ground operations as well as allphases of flight. Activities are underway in the integration offlight avionics systems into a hardware-in-the-loop (HIL)simulation which will be used for Verification and Validation(V&V) testing of the avionics system.

Approach The development approach begins with simulationmodels for the X-33 vehicle, e.g., aerodynamics, reaction controlsystem, actuators, engine, navigation, guidance, and control.Each model is incorporated into the X-33 batch and real-timesimulations. Six 1553 buses have been integrated into the real-time simulation to provide communication with the avionicssubsystems. Once all of the hardware is fully integrated into thesimulation and the final operational flight program (OFP) isdelivered, formal V&V testing will be performed to certify thesystem is ready for flight.

Status Current work continues to be focused on model updatesand hardware integration to support full, ascent - rollout,hardware-in-the-loop capability. Vehicle System Check Out(SCO) of the avionics subsystems are also getting much attentionin the lab as these subsystems are installed on the vehicle.

Several elements of flight hardware have been and are beingintegrated at the ITF lab. These include Vehicle MissionComputers (VMC’s), Flush Air Data System (FADS) RemotePressure Sensor (RPS) units, INS/GPS units, and Forward, Rear,and Engine Data Interface Units (DIU’s). The real-time VMCHIL simulation is now capable of flying X-33 missions fromlaunch to landing.

Recent work has started in testing system robustness and stabilitymargins of the flight control system. The robustness analysis isconducted through numerous Monte Carlo runs which are usedto determine sensitivity to dispersions in critical simulationmodels, such as aerodynamics, engine, and vehicle massproperties. The stability margin analysis utilizes linear modelsof the vehicle dynamics combined with Matlab models of thecontrol system. The linear models are derived directly from thesix-DOF nonlinear simulation.

StrutFrame

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Landing gear forces and moments.

X-33 Landing Gear Dynamics Model

Summary: One of NASA Dryden’s main roles on theX`33 program is to develop a high fidelity X-33simulation to be used in verification and validationtesting of the avionics subsystem. The most benefit fromV&V testing is gained when the subsystem modelsprovide the most accurate possible representation of theactual subsystems. The landing gear model mustfaithfully portray tire friction and compression as well asstrut loads, compression and dynamics. This will indicateto V&V engineers the potential for gear failure, tailscrape or a dynamic interaction with the automatic flightcontrol system as well as prove the performance of theground steering control laws.

Objective: The objective is to develop an accuratelanding gear model which may be easily integrated intothe X-33 simulation and create the smallest possiblecomputational burden. This will allow for real-time V&Vsimulation testing of the landing system produce highconfidence in the results.

Approach: The core gear dynamics model evolved froma high fidelity landing gear simulation created byLockheed Martin for the F-117. This core was modifiedwith X-33 landing gear parameters and the equations ofmotion were expanded for greater accuracy.

Status: The landing gear model is installed in the X-33six-DOF simulation. Also integrated into the geardynamics model are models for the nosewheel steeringcontroller and brake control unit. Current work is beingperformed to incorporate runway crown effects into themodel and there will be adjustments made on some gearparameters to match drop test results.

ContactLouis Lintereur, DFRC, RC, (661) 276-3307

X-43A Non-Linear Monte Carlo Analysis

SummaryThe Hyper-X program was conceived in order todemonstrate the first ever operation of an integratedairframe-scramjet vehicle in flight. The X-43A is thefirst in a series of such aircraft. It is carried to the testpoint, which is approximately Mach 7 at 95,000 ft, bya modified Pegasus launch vehicle. Followingseparation from the Pegasus, the scramjet engine isfired for a short period of time. Parameteridentification maneuvers are then conducted untilsplashdown in the Pacific Ocean. Since this is thefirst in a series of unique missions, there are manyuncertainties. Aerodynamic performance candiffer from CFD and wind tunnel derived modelsby up to 30%. Actuator, mass properties andsensor uncertainties also have to be taken intoconsideration. The Hyper-X Controls group hasdeveloped a Monte Carlo simulation tool for theuse of the Hyper-X project to demonstrate desiredcontroller performance in the presence of modelingand measurement uncertainties.

ApproachThe Monte Carlo tool consists of a modified non-linear FORTRAN simulation of the X-43A andMATLAB scripts which are used to do the following:drive the non-linear sim with various uncertainties,analyze the results, save the data, and plot the data forfurther analysis. The uncertainties used consist ofuniform or Gaussian distributions for atmosphericproperties, actuator performance, aerodynamiccharacteristics, mass properties, sensor uncertainties,flight condition at separation, and separation forcesand moments.

ResultsThe Monte Carlo tool has proven invaluable inconfirming controller performance when theaforementioned uncertainties are applied. Thecontroller in all cases achieves the requirement for aconstant angle of attack during the scramjet test. Thedata has also been used to develop a trajectoryfootprint and understand the effects of variousuncertainties upon the mission profile. Specific simscripts generated for the Monte Carlo analysis havebeen selected for use during hardware in the loopvalidation testing in order to fully exercise the systemwithin the expected mission profile.

Status/PlansModel updates are added to the Monte Carlo tool, asthey become available. 1,000 runs are generatedevery weekend to provide additional data for analysis.The tools continue to evolve as more uses are foundfor them.

ContactsEthan Baumann, DFRC, RC, (661) 258-3417Michael Richard, DFRC, RC, (661) 258-3543Mark Stephenson, A DFRC, RC, (661) 258-2583

Descent Trajectory

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Artist conception of UCAV

Unmanned Combat Air Vehicle

SummaryThe unmanned combat air vehicle (UCAV) is aBoeing Phantom Works concept for a fullyautonomous airplane designed to perform militarysuppression of enemy air defense (SEAD)missions. Work is currently being done to developtwo UCAV demonstration vehicles, which willserve as a prototypes for the eventual operationalvehicles. The demonstration vehicles will berequired to taxi in formation, takeoff, fly information, execute a collision avoidance maneuverand land in a fully autonomous manner. Operationswill be conducted under the supervision of ahuman operator who will monitor vehicle health,send basic commands and serve as an interface toair traffic control. Dryden will host the flight testprogram, which is expected to begin in December,2000.ObjectiveNASA Dryden’s major roles in the program are todevelop the automatic taxi control laws, andalgorithms for formation taxi and collisionavoidance. These products will be integrated intothe vehicle control system and demonstratedduring flight test.ApproachBoeing has provided Dryden with a detailed, highfidelity Matrix-X 6-DOF simulator. All algorithmsand control laws will be designed and validatedusing this single tool. Ultimately, the flight controlsoftware will be obtained by auto-coding thecontrol law block diagrams directly from thesimulator.The auto-taxi control makes use of the throttle,collective and differential brakes, yaw thrustvectoring, speed brakes and nosewheel steering tosteer the vehicle along a predetermined taxi pathconsisting of surveyed waypoints. The control lawsare designed using nonlinear inversion techniquessimilar to the flight control laws. This will enable asmooth integration into the flight control systemand minimize potential problems during transitionto and from flight.Dryden’s UCAV team is also evaluating conceptsfor the integration of UAV aircraft such as theUCAV into airspace with piloted aircraft. Ofprimary concern is the ability of the UAV to “seeand avoid” other aircraft. The Traffic Alert andCollision Avoidance System (TCAS), currentlyapproved by the FAA for commercial transportaircraft, will be integrated with autonomous

collision avoidance algorithms developed by MIT’sLincoln Laboratories onboard one of Dryden’s high-performance research aircraft to demonstrateautonomous collision avoidance. These algorithms,using simulated TCAS hardware, will also be flownon the UCAV vehicle for a collision avoidancedemonstration.Status The auto-taxi control law architecture isestablished and integrated into a local copy ofBoeing’s 6-DOF simulator. Formation taxi algorithmshave also been developed. Current work is focusedon refining the control system gains and integrationwith the flight control system.

The collision avoidance effort is currently procuringhardware and developing algorithms to be used to testand demonstrate a collision avoidance maneuver forUCAV.

ContactsNeil Matheny, DFRC, RC, (661) 258-3728Louis Lintereur, DFRC, RC, (661) 258-3307Curtis Hanson, DFRC, RC, (661) 258-3966Valerie Gordon, DFRC, RC, (661) 258-2018Peter Urschel, DFRC, RC, (661) 258-2718Jamie Willhite, DFRC, RF, (661) 258-2198

X-43A Hypersonic Blended Inertial and Pneumatic Angleof Attack

SummaryThe Hyper-X research program, conducted jointly by NASADryden and NASA Langley, was conceived to demonstrate ascramjet engine in a flight environment. The X-43AResearch Vehicle, the flight test instrument of the Hyper-Xprogram, will be lofted to its pre-determined research testcondition with the aid of a modified Pegasus first-stagebooster. After separation from the launch vehicle and duringthe engine test phase, the X-43A will be commanded tofollow a nearly ballistic flight path due to mass flow andangle-of-attack requirements imposed by the scramjet enginetest. Key to mission success is the ability to accuratelymeasure and control angle-of-attack.

Angle of Attack derived from inertial sources can provideadequate frequency content for flight control, but absoluteaccuracy is heavily dependent on an accurate atmospheremodel Ð including winds. A Flush Airdata System (FADS)can provide highly accurate flow angles in the presence ofwinds, but significant pneumatic lag often limits theusefulness of this signal for real-time feedback. In order toprovide an accurate, real-time angle-of-attack signal forHyper-X flight control, a FADS derived angle of attack canbe blended with an INU-derived angle of attack.

ObjectiveDryden efforts over the past year have focused on improvingthe FADS algorithm performance with regard to angle ofattack requirements for the scramjet engine test phase. Wind-tunnel data was used to calibrate the FADS angle of attackfrom Mach 3 to Mach 8. Significant effort was put towarddeveloping and validating a simplified pneumatic lag modelin order to compensate for the lag in the flight control laws.For simplicity, a first-order bias filter was selected to blendthe FADS signal with the INU-derived angle of attackfigure(1).

ResultsA dynamic FADS simulation model was developed fromstatic and dynamic wind tunnel test data. The simulationmodel has been used to generate Monte Carlo uncertaintyresults. Two project reviews of the blended angle-of-attackalgorithm were conducted in FY99. Updates to the real-timeangle-of attack estimation algorithm have been incorporatedinto the flight software release.

Figure 2 shows differential FADS pressure results from windtunnel and simulation data for an angle-of-attack sweep fromÐ6.5° to 12° at Mach 8. One key result from this research isthat the pneumatic lag model for a pair of differential portsdoes not change significantly over the desired angle-of-attackrange. This result is evident from the inserts shown in figure2, showing similar differential pneumatic lag response to 1°steps in angle of attack throughout the test range.

Status/PlansHardware-in-the-loop and on- aircraft testing will beused to validate system implementation.First flight of the Hyper-X is scheduled for summer2000.

Figure 1: Block diagram for a blended FADS and inertialangle of attack.

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Figure 2: Mach 8 angle-of-attack sweep (AEDC windtunnel data).

ContactsJoe Pahle, DFRC, RC, (661) 2583185Mark Stephenson, DFRC, RC,(661) 258-2583Mark C. Davis, DFRC, RA (661) 258-2241

Alpha Selection

1alpha_est [deg]

fads_1, deg

fads_2, deg

fads_3, deg

INU alpha, deg

Mach

alphabias _GCA, deg

Lag Models andGood Channel Average

1

s+1

Bias Filter

5Mach

4alpha_INU [deg]

3alpha_fads3 [ deg]

2alpha_fads2 [ deg]

1alpha_fads1 [ deg]

F/A-18 Autonomous Formation FlightSummaryAircraft flying in formation can take advantage of eachother’s wingtip vortices to improve the overallformation efficiency. Two aircraft flying wingtip-to-wingtip can achieve the performance of a single aircraftwith twice the aspect ratio. While wingtip-to-wingtipflight is unrealistic, this effect dissipates slowly withlongitudinal separation. Two aircraft flying withwingtips aligned and within two or three wingspanslongitudinally of each other can achieve significant dragreduction.

ObjectiveDemonstrate through flight test a formation flightautopilot system. This system will include all of thenavigation, guidance and control functionality requiredto accomplish autonomous station keeping of onefollower aircraft to a second leader aircraft.

JustificationFormation flight drag reduction has the potential to savecommercial air cargo companies millions of dollars peryear in fuel costs. This research also has benefits in theareas of Uninhabited Combat Air Vehicle (UCAV)formation flight, swarming and autonomous refueling.

Benefits• Reduced drag and extended range• Autonomous station keeping• Cooperative formation control

ApproachA NASA F/A-18 chase aircraft will be outfitted with aformation flight autopilot system, comprised of a GPSreceiver, an Airborne Research Test System (ARTS)and Production Support Flight Control Computers(PSFCCs). Navigation and guidance algorithms havebeen developed at Dryden along with outer-loop controllaws to complete the formation autopilot system.

This specially modified aircraft will fly in formationunder autopilot control with NASA’s F/A-18 SystemsResearch Aircraft (SRA) in flight tests scheduled for thesummer of 2000. The aircraft will maintain a separationof 250 feet to ensure safe operations.

ResultsThe GPS relative navigation system has undergone riskreduction testing using NASA’s King Air aircraft.These tests have demonstrated that two independentGPS systems on the same aircraft provide a relativeposition accuracy of approximately 1.5 feet. A relativeposition accuracy of 15 feet or better is required forsuccessful formation station keeping.

Contact

Curtis E. Hanson, DFRC, RC, (661) 258-3966

Relative Position Accuracy of Two GPS Antennas inFlight at a Fixed Separation

StatusFinal software and hardware modifications are beingcompleted in preparation for verification and validationtesting of the formation flight autopilot system. Riskreduction flights to further evaluate the GPS relativenavigation system are planned for early 2000. A softwarein the loop simulation that will facilitate control lawevaluation and flight test planning is nearing completion.

Flight test of the Autonomous Formation Flight system isscheduled for June of 2000.

Instrumentation Pallet Containing Components of theFormation Flight Autopilot System

Time (seconds)

Error(ft)

Mean = 1.55 ft

StandardDeviation =

1.61 ft

On FlightCondition

Preflight/Takeoff

Landing/Postflight

GPSReceiver

Airborne ResearchTest System (ARTS)

F/A-18B Systems Research Aircraft (SRA)X-33 Vehicle Health Monitoring System Risk Reduction Experiment

SummaryA VME Based VHM computer, remote health nodes(RHN), and Distributed Strain Sensor (DSS) module wereinstalled and integrated onto the SRA in September of1999. A Structures Flight Test fixture was added to theexperiment in November of 1999 to support the DSSflight research efforts. The fixture was removed in EarlyDecember of 1999 and the aircraft was returned to theVHM computer and health node configuration. Theexperiment has flown for 34 flights, six of which wereflown as dedicated DSS research flights with the flighttest fixture installed on the aircraft. Data reduction andanalysis is ongoing.

ObjectivesThis experiment was designed to demonstrate that theVHM system could be successfully integrated and flownon a high performance aircraft and to validate fiber opticstrain measurements against conventional technologywhen acquired in a realistic flight environment. Thespecific objectives of the experiment included.• Demonstration of the installation and performance of

the Generation II Multi Mode fiber optic cable plant.• Demonstration of the capabilities of the RHN and

validate the accuracy of the measurements.• Verification of the installation and integration

procedures of the VHM system on the aircraft.• Validation of the correlation between the fiber optic

DSS and data recorded using conventional strainsensors.

• Provide a configuration for ongoing VHM researchefforts.

Justification

The realization of integrated health management (IVHM)involves acquisition, distribution, analysis, anddecimation of vehicle information to the user. Thisinvolves fusing new or existing sensor technology withacquisition and distribution devices then implementingprognostic algorithms, and displaying the resultinginformation to the correct person in a timely manner. This experiment applies both new and existing sensortechnologies with a new data distribution network. Components of the system have been tested in alaboratory environment, but have not been demonstratedin a flight environment.

Approach

The VHM computer and one RHN were installed in theGun Bay on the SRA. A second RHN was installed inAvionics Bay 13 Right. These nodes collected analogsensor data and transmitted the information to the VHMcomputer using a multi mode Generation II fiber opticcable plant. The DSS experiment consisted of a tuneablelaser and signal processing module, located in the VHMcomputer. Eight single mode fibers and Bragg Gradingwere installed on the aircraft and flight test fixture (threeon aircraft structure and five on the flight test fixture.) tomeasure strain in a flight environment. Data from overthirty-four flights have been collected using the VHMcomputer and RHN’s. Six flights were dedicated toflying the DSS experiment and flight test fixture. Flighttest fixture flights were flown to 100% of the designlimit load (DLL) on the fixture.

Results

• Installation and integration of a generation I VHMarchitecture consisting of a Distributed COTS Based,Open Architecture design concept.

• Flight demonstration of a Generation II Multi ModeFiber Optic Cable plant transmitting data betweenhealth nodes and the VHM computer.

• Flight demonstration of a network of fiber opticdistributed strain sensors.

• Integration and flight test of a generic flight-testfixture to 100% of DLL.

A NASA paper reporting on the results of the experiment

is being initiated.

Contacts

Keith Schweikhard, DFRC, RF, (661) 258-3411Lance Richards, DFRC, RS, (661) 258-3562

Flight Test of an Intelligent Flight Control System on the F-15ACTIVE

SummaryThe F-15 Advanced Controls Technology forIntegrated Vehicles (ACTIVE) was the testbedfor flight test of the Intelligent Flight ControlSystem (IFCS). The IFCS is a flight controlconcept which utilizes a Neural Network (NN) todetermine critical stability and controlderivatives for a control law that calculates realtime feedback gains with a Ricatti Solver. Thesederivatives are also used for plant identificationin the model following portion of the controllerto insure level one handling qualities. The FlightTest of the IFCS was the culmination of the jointIntelligent Flight Control Advanced ConceptProgram teaming NASA with Boeing PhantomWorks. The goals of the IFCS AdvancedConcept Program (ACP) are:

1. Develop a Flight Control Concept that usesNeural Network Technology to identifyaircraft characteristics to provide optimalaircraft performance.

2. Develop a Self-Learning Neural Network toupdate aircraft properties in flight.

3. Flight demonstrates these concepts on theF-15 ACTIVE Aircraft.

ObjectiveThe IFCS Advanced Concept Programobjectives consist of three phases:

Phase I: Pre-Trained Neural NetworkDevelopment.

Phase II: On-Line Neural Network Development.

Phase III: Neural Network Flight ControllerDevelopment.

ApproachFlight test of the IFCS ACP was performed in 3stages. Phase I & II was flown with systemoutputs provided to instrumentation only andwas not used for aircraft control. Phase IIIutilized the Phase I Pre-Trained NN to providereal-time aircraft stability and control derivativesto a Stochastic Optimal Feedforward andFeedback Technique (SOFFT) controllerdeveloped by NASA Langley. This combinedPhase I/III system was flown utilizing the F-15ACTIVE Research Flight Control System(RFCS). The RFCS allows the pilot to quicklyswitch from the experimental research flightmode back to the safe conventional mode ifrequired.

ResultsIFCS ACP flight test was completed in April,’99. The Phase I/III flight test milestone was todemonstrate across a range of subsonic andsupersonic flight conditions that the Pre-TrainedNeural Network could be used to supply real-time aerodynamic stability and controlderivatives to the closed-loop optimal (SOFFT)flight controller. Additional objectives achievedinclude:1. Flight Qualifying a NN Control System.2. Use a combined Neural Network/closed-

loop Optimal Flight Controller to achieveLevel 1 flying qualities.

3. Demonstrate, via Dial-a-Gain, that differentflying qualities could be achieved by settingnew target parameters ωsp, ζsp, ωdr, ζdr and tr.

In addition, data for the Phase II (On-LineLearning NN) was collected using AdAPT(Stacked Frequency Sweep) Excitation for postflight analysis. Initial analysis of this datashowed good potential for real-timeidentification and online learning.

ReferenceIntelligent Flight Control: Advanced ConceptProgram, Final Report, Boeing-STL 99P0040(NAS2-14181)

ContactsMichael P. Thomson, DFRC, RF,(661) 258-3097mike.thomson@dfrc.nasa.govCharles C. Jorgensen, Ames,(650) 604-6725cjorgensen@mail.arc.nasa.gov

F-15 ACTIVE

X-Actuator Control Test (X-ACT) Program

SummaryFor X-ACT, the X-38 EMA will be installed in place of thestandard hydraulic speedbrake actuator on NASA #837, aheavily modified, preproduction F-15B two seat fighter aircraft.An Actuator Control Unit (ACU) will be used to provideactuator commands normally generated by the X-38 FlightCritical Computers. Actuator commands will be specified by aFlight Test Engineer in the back seat of the aircraft via acockpit interface system. The X-38 Failure Detection,Isolation, and Recovery (FDIR) logic will be hosted in theACU, with full authority to reconfigure the system. Failurescan be injected via the crew interface, either by modifying datapassing between the ACU and actuator controllers, or byinterrupting signals between the actuator and controller.

ObjectiveThe primary technical objectives of the X-ACT program, canbe grouped as follows:Actuator Integration and Flight (JSC)Uncover unanticipated problems in taking the system to flight,validate/develop an actuator model under actual flightconditions, design and integrate a high power, multichannelEMA system onto a testbed aircraft, and evaluate EMAs underspecific simultaneous load, rate, and extension conditions in aflight environment.Redundancy Management System (JSC)Design, code, implement, and test the redundancy managementsoftware for the X-38/CRV EMA system and validateredundancy management software using a fault injectionsystem.Cockpit Interface System (DFRC)Development of a Multipurpose Cockpit Display System forFlight Research., develop graphic display software for onboardmonitoring of research experiments, including mission criticalparameters, develop graphic interface software to controlonboard research experiments, develop a flexible commandprofile generator system, and design, implement and test ahardware and software fault injection system.

JustificationThe research value of the X-ACT flight test program is in theestablishment of a new redundancy configuration, potentiallysuitable for a wide range of new vehicle applications. Theseflight tests will either verify the correctness of the designapproach or identify deficiencies in the configuration to allowcorrection. Lessons learned during this program should beapplicable to other high power EMA and EHA programscurrently in development, both for spacecraft and aircraftapplications. A successful flight test of the actuator could helpfurther the acceptance of Power-By-Wire actuation technologyby industry.

ApproachThe flight test matrix will span the rod end load and rateenvelope of the actuator. The actuator will be testedthroughout the full range of its stroke. Some typical X-38reentry profiles may be flown, from an altitude of 50,000 ft.down to the moment of X-38 parachute deployment. The goalis to have 25 flight hours on the actuator by the time vehicleV-201 ships to KSC for Shuttle Integration. An additional 75flight hours will be accrued prior to first flight of the CRVvehicle.

Results• Currently under development

Benefits• Reduced Risk for X-38.• Validation of new Redundancy Management Schemes for

Power-By-Wire Actuators.• Lessons Learned applicable to multiple EMA development

programs.

ContactStephen Jensen, DFRC, RF, (661) 258-3841Principal Investigator

X-ACT System Locations

X-43A ANTENNA PATTERNS

SummaryThe Hyper-X research program, conducted jointly byNASA Dryden and NASA Langley, was conceived todemonstrate a scramjet engine in a flightenvironment. The X-43A Research Vehicle, theinstrument of the Hyper-X program, will be lofted toits pre-determined research test condition with the aidof a modified air-launched Pegasus booster. Afterseparation from the launch vehicle and during theengine test phase, the X-43A will be commanded tofollow a nearly ballistic flight path - a result ofscramjet engine angle-of-attack requirements. Theengine test phase (which includes post-test vehicleparameter identification maneuvers) is concluded bya recovery to a nominal descent trajectory madepossible by the autonomous controller resident in thevehicleÕs flight control computer. Since the vehicle isautonomous the only link to ground is by transponderand telemetery. The telemetry system is transmitonly. The transponder operates as a beacon. Whenqueried the transponder will transmit. Several factorsdepend upon the reception of data transmission froman aircraft. One is the installed antenna pattern of theaircraft. The X-43A has three antennas, two on eachside and one facing the rear. Knowing the transmitpattern of the antennas will provide the ability toreceive the strongest transmission of radiofrequencies from the aircraft.

ObjectiveThe overall test objective was to gather qualitativeantenna pattern data from installed antennas on theHXRV. This objective was accomplished bycollecting antenna patterns at selected elevations ofthe HXRV. The test data will be used to verifyadequate (radio frequency) RF coverage at possibleranges and altitudes the aircraft will be located duringits flight. The data will identify null areas of the RFcoverage and assist in orienting the flight path andairborne and ground assets for best RF reception.

JustificationAntennas installed on aircraft usually have a muchdifferent pattern compared to uninstalled patterns.The airframe may act as an electrical conductor thatmay transmit electromagnetic energy throughout thestructure. Other aircraft structures may block themagnetic energy. The antennas depending uponlocation may interfere with each other. These andother electromagnetic properties enter into theradiated pattern of the antennas.

ApproachThe X-43A aircraft was placed in the BenefieldAnechoic facility at Edwards AFB, California fortesting. The aircraft exterior was kept as close aspossible to the flight configuration. The internalstructure was the hull without the various flighthardware. The aircraft was lifted by hoist to anelevation that may be seen in flight. A networkanalyzer generated various frequencies to radiate atthe aircraft. The aircraft was tied to a positioner torotate the aircraft 360 degrees. These frequencieswere received by the aircraft antennas and sent to anRF receiver. The receiver was connected to a plotterthat provides antenna polar plots.

ResultsThe antenna plot below provides a tangible picture ofwhat the three antennas would receive or transmitduring flight. The plot is representative of theelectromagnetic environment around the aircraft andthe aircraftÕs transmit and receive capability at thespecific RF frequency and at the elevation of thepattern.

X-43A Antenna Plot

Contacts Mark W Hodge, DFRC, RF, (661) 258-7528 John Kelly, DFRC, RF, (661) 258-2308

Reconfigurable VHF Broadcast Differential GPS Base Station

SummaryDifferential Global Positioning System (DGPS) has beenidentified as a requirement for an increasing number offlight projects. This technology allows positioning ofaircraft in real time to within a few meters or centimeters.Rather than each project putting together their own groundstation, a common resource would provide a standard andsatisfy many project requirements, yet be reconfigurable forcustom requirements.

ObjectiveDesign, develop, and demonstrate a real-time,reconfigurable broadcast DGPS system to achieve sub-meteraircraft positioning in real time.

JustificationWhile various types of DGPS ground stations exist, they arenormally configured in a static configuration and often notoptimized for precision flight research work. Flight researchoften requires a reconfiguration of support systems,including DGPS. Reconfiguration requirements mayinclude changes to frequencies used, data rates, dataformats, message types, transmit power, and date logging.Having a NASA controllable asset that can be scheduledspecifically for particular configurations will help NASADryden to support a variety of GPS activities.

ApproachA dual set of Ashtech Z-12 GPS receivers were purchased,along with sets of error-correcting radio modems andantennas. This equipment was installed in a building nearthe main runway so as to optimize both visibility while inthe pattern and visibility within a 200 mile area surroundingthe airfield. Either standard RTCM SC-104 code phase orDBEN carrier phase (Ashtech proprietary) messages can betransmitted to aircraft containing radio modems and GPSreceivers--allowing aircraft position reporting in the meter(using RTCM) or centimeter (using DBEN) ranges.

ResultsThe system is in place and transmitting. It is being usedinitially to assist in understanding X-33 related GPS systemsand is ready to support autoland and autonomous formationflight (AFF) activities, as well as X-34 activities, if required.

ContactsGlenn Bever, DFRC, RI, 661-258-3747Joseph Collura,DFRC, F, 661-258-2450Becky Flick,DFRC, F, 661-258-3783Ed Haering, DFRC, RA, 661-258-3696Principal Investigators NASA Dryden.

Benefits• Enhances Dryden’s range capabilities.• Improves the accuracy of on-board GPS data .• Provides a basic GPS functionality that can be modified when requirements change.•Increases personnel understanding of DGPS.

Diagram of DGPS ground station and application.

Z-12

RadioModem

Z-12

RadioModem

backup Prime

GPS GPS

VHF

GPS

DSP/GEDAE Signal Processing AlgorithmsSummaryThe understandings of dynamic phenomena are oftenessential in the development of modern aircraft systems. Todescribe these, one may use either a statisticalcharacterization of the signal, such as standard deviation orhigher moments (skewness or flatness) or zerocrossings.These are distinct parameters that fit well in the PCMstream.

However, if spectral information is required it is necessaryto compress the data either accepting some loss ininformation, or if processing speed and memory bufferpermits, do a lossless compression.

The system under development encompasses selectivedigital filtering rather than analog to increase flexibility,signal characterization through statistical parameters toallow parameter identification and spectralanalysis/compression to allow merging with lowerfrequency parameters.

The illustration shows a typical hot-film spectrum taken inflight that would require 1800 bits uncompressed. Theinformation is transferred lossless using approximately 600bits utilizing among other techniques bit-packing, spikeremoval along with inversion. If a small loss is allowable,floating point curve representations can be used, resultingin further reduction in bit transfer requirements.

ObjectivesDFRC needs to improve ground and airborne capabilities toencode, record, process, correlate, and archive highfrequency hot film and microphone research data.Traditionally most flight parameters are made available tothe researcher in the form of time-series in the FDASsystem, while the dynamic data has much higher bandwidthrequirements. The objective of the present work is thedevelopment of a data acquisition subsystem which canprocess high frequency data in real time and output resultsthat can be merged with slower PCM streams for displayand archival.

ResultsGround System Development:

Desktop analysis tools were evaluated and GEDAEwas chosen. An embedded PC board running NTwas chosen as the host system which can thenconfigure the DSP to run the desired algorithm.

Software tool development:Smoothing filters, Integrating filters, DifferentiatingfiltersSpike removal, Curve fitting, deviationscross correlation

Ground system testing:Tools were tested and used against Hot Film data setssupplied by LaRC and data stored on FDAS

Status/PlansDefine a flight system and the associated controlroom part of the system. Acquire flight hardware.

Perform statistical and spectral analysis to reducethe data so that it can be merged with traditionalPCM streams.

Continue to integrate tools onto the embedded DSPboard.

Work directly with Hot Film sensors and theIXTHOS Sharc DSP.

Spectrums from hot film sensor in flight at hypersonic

Mach number. Example of loss less compression.

ContactsArild Bertelrud, AS&M., RI, (661) 258-2989Russell Franz, AS&M., RI, (661) 258-2989Rebecca Flick, DFRC, RI, (661) 258-3783Richard Hang, DFRC, RI, (661) 258-2090

1501005000

50

100

150

FREQUENCY BIN

AMPLITUDE

COM-2

ORIGINALCURVE

TREND

FOLDEDVARIATIONS

Multiplexed Hot FilmsSummarySurface-mounted hot-films are commonly used to detecttransition from laminar to turbulent flow and separated flowregions on wind tunnel models as well as in flight. With theintroduction of high-speed data acquisition and processing itbecomes feasible to document the flow using sheets ofsensors arranged in arrays, which are switched in as needed.

An effort to explore the flow physics of a three-dimensionalhigh-lift configuration in a wind tunnel is planned at NASALangley. Approximately 650 films will be used in variouscombinations depending on real-time decisions when variousflow phenomena are encountered. These can be attainedthrough software-controlled, commercially availableswitches. A total of 24 temperature-compensated constanttemperature bridges are used, organized in three modules of 8channels each. NASA Dryden-developed anemometry is usedto run the films, taking advantage of the size, simplicity,robustness and low noise level. Both power to the bridges aswell as signals are switched.

To evaluate the design of the transition experiment in termsof ability to identify transition/separation features andsupport the software development as well as learning aboutthe flow physics, two risk-reduction experiments wereconducted in Langley's Basic Aerodynamics ResearchTunnel (BART). The system turned out to be very reliable,both in terms of the sheets with sensor arrays, and thefunctionality of the anemometer circuits and the switches.

The utility of the modular design was demonstrated whenhot films were needed for exploration of various models ofthe 2003 Mars Airplane. Depending on the required numberof films, anticipated flow conditions, scheduling etc., thesystem was split into several subsystems, each withanemometer modules and switches. The data acquisitioncodes required minor changes and the self-descriptive filesystem definition originally developed for the high-liftexperiment was used.

ObjectiveDryden efforts are focused on making a system available "onthe shelf", allowing quick response to needs that may arise.This involves mainly hardware and data-acquisition softwareas well as use of film arrays. Utilization of the currentanemometer design together with flight-hardened switchingunits are short-term objectives. A more long-term goal is arevision of the anemometry itself to include analog-digitalconversion, switching and optimization at the board-level.

Real-time processing, compression and merging of the high-frequency data are described in a separate highlight.

ResultsSo far the approach has been used in five successful windtunnel experiments at NASA Langley. Dryden haveparticipated through making the anemometry units availableand working with Langley to develop a better understandingof system performance. Experimental results have been sharedto better guide the post-processing and merging of the datawith low-frequency data.

Status/PlansThe main outstanding issues as far as Dryden is concerned, isthe flight-hardening of the switching units as well as thecontrol of these for flight applications.

PHOTO OF UNITS

SAMPLE SIGNAL

ContactsArild Bertelrud, AS&M, RI, (661) 258-2989Harry Chiles, AS&M, RI, (661) 258-3738Anthony Washburn, LaRC, (757) 864-1290

LABVIEW Based ESP Pressure Data Acquisition System

Summary

A supersonic wind-tunnel test was conductedat Penn State University to studyshock/boundary-layer interactions. Multiplepressure measurements along a flat-plate testarticle were required to adequatelycharacterize the pressure profile in the vicinityof an impinging shock. It was required torecord 32 pressure measurements at 20samples per second to capture the unsteadynature of the shock/boundary-layer interactionin the wind tunnel facility. Since a pressuremeasurement system that could meet theserequirements did not currently exist at thefacility, it was determined that NASA Drydendevelop a low-cost and fairly Ònon-intrusiveÓpressure data acquisition system. This dataacquisition system was to be completelyseparate from the wind tunnel data acquisitionsystem with time tags correlating the twosystems. Utilizing a laptop computercontaining National Instruments hardware,LabVIEW software and an Electronic ScannedPressure (ESP) 32 channel PressureTransducer, a Pressure Data AcquisitionSystem was developed. The test setup isshown in the photograph below.

Lab Test Setup

Objective

Develop a system that would allow theinvestigators to acquire pressure distributionswithout conflicting with the existing wind tunneldata acquisition setup and to provide a Òuser-friendlyÓ interface.

Results

The LabVIEW based ESP Data AcquisitionSystem created for the wind tunnel tests was

LabVIEW program panel

developed at Dryden, shipped to Penn State, andsuccessfully used to collect the desired pressure data.The system obtained data at a rate of 640 scan/secreliably. This resulted in 32 pressure channelmeasurements at a rate of 20 samples per second.This data was then saved to a spreadsheet file.

The program had a pre-test functionality toverify that all 32 ESP channels were i nworking order. Selecting the chart option onthe LabVIEW front panel, shown above gave atime history of the pressure measurementsand could be viewed real-time on the screen.This reduced the sample rate and wastherefore only used for pre-testing the sensor.

Status

This system has been developed and may beutilized in other tests where multiple pressuremeasurements are required and with minimalimpact to the setup. The data acquisitionsystem may also be used to test and verifyESP sensors on an aircraft or to ensure allESP channels are in working order.

ContactsAlma M. Warner, DFRC, RI, (661) 258-3159Tim Moes, DFRC, RA, (661) 258-3054

Flight Instrumentation for X-33 Extended Range TestSummaryThe X-33 project presents unique trackingrequirements because the vehicle trajectory is beyondthe line-of-sight of any single range. The X-33 isplanned to fly between Edwards Air Force Base,California and Dugway Proving Grounds, Utah. Anextended range was established to provide thecontinuous tracking necessary for maintaining publicsafety within the area effected by the X-33 trajectory.High altitude flight tests were required to confirmthat the design, construction, function andperformance of each extended range system satisfiedthe requirements.

ApproachA subset of X-33 avionics was integrated on an ER-2aircraft in addition to a small instrumentationpackage. The X-33 avionics consisted of a C-bandradar transponder and custom high-temperature flush-mount antenna, a Flight Termination System (FTS)receiver and stock antenna, a telemetry transmitterand custom, high-temperature flush-mount antenna, ahybrid RF combiner, and command uplink receiver.The Code RI-implemented instrumentation packageconsisted of DFRC signal conditioning, a PCMencoder, and a DFRC Airborne InstrumentationManagement System (AIMS). The instrumentationpackage monitored avionics temperatures, signalstrengths and FTS commands. The PCM multiplexerencoded these measurements in an IRIG formattelemetry stream. The AIMS added ER-2 avionicsparameters from a serial RS-232 format to thetelemetry stream and output the data stream to thetelemetry transmitter. Power to each subsystem wascontrolled by a series of fuses, circuit breakers, andrelays. The relays allowed the pilot to simulatefailures of various on-board X-33 rangecommunication avionics subsystems.

ResultsAn instrumentation package and a subset of X-33avionics used for range communications weresuccessfully integrated on a NASA Dryden ER-2 forhigh altitude flight test. Uplink receiver signalstrength levels met requirements. The requirement forcontinuous telemetry coverage from takeoff tolanding was met, as was the requirement of at least a6 dB link margin. Continuous FTS coverage andcommand integrity was verified. Integrated rangesystems operations and interfaces were validated, andoperations training and practice for all range systemsthat will be utilized for X-33 operations wereprovided, including investigation of and training forfailure modes and effects.

Status/PlansTwo more ER-2 flights are currently planned to re-validateÊX-33 Extended Test Range assets.Ê Theseflights will take place approximately two monthsprior to the first X-33 launch. Before the next flights,it is planned to upgrade the instrumentation packagewith a higher bit rate PCM encoder. Also, thecapability to wrap-around uplink commands isplanned for the AIMS. Improvement of the wiringharnesses on the avionics pallet associated with bothpackages is planned.

C-BandTransponder

FTSReceiver

Combiner

Temperatures,Signal Strengths

FTS Status

PCM Encoder

ER-2 AvionicsParameters

AIMSUplinkReceiver

Transmitter

Block Diagram

ContactsTerry Montgomery, DFRC, RI, (661) 258-3188Glenn Bever, DFRC, RI, (661) 258-3747Keith Krake, Sparta, Inc., RI, (661) 258-2147

Efficient Modulation Techniques

SummaryNASA Dryden and UCLA are investigating and testingefficient modulation techniques for use during flight- testing.The efforts include flight testing of Feher Quadrature PhaseShift Keying (F-QPSK)* modulation, research into otherefficient modulation techniques, investigation of methods toimprove the performance of F-QPSK as well as othermodulation systems and research into the use of cellularphone Space-Time Coding methods for efficient datatransmission.

F-QPSK modulation has been adopted by the RangeCommanders Council for high data-rate telemetry systemsand is capable of transmitting twice the data of aconventional Pulse Code Modulated/Frequency Modulation(PCM/FM) system in the same bandwidth, figure 1. Drydenis investigating the performance of sample F-QPSK systemsthrough flight testing aboard the NASA King Aire aircraft.Dryden is also working with UCLA to investigatemodulation techniques that may yield efficiencies greaterthan that of F-QPSK.

UCLA is currently investigating equalization methods toimprove the current performance of QPSK systems, whichcould also be applicable to F-QPSK. The equalizer providesamplitude and phase compensation to restore the receivedsignal to itÕs original transmitted form and reduce the effectsof fading on the quality of received data.

UCLA has proposed the investigation of Space-Time Codingto increase the efficiency of QPSK, Quadrature AmplitudeModulation (QAM) and other modulation techniques. TheSpace-Time Coding technique utilizes multiple transmit andreceive systems, figure 2, to improve bandwidth efficiencyand Bell Laboratories has reported laboratory test results forcellular phone applications.

ObjectiveCurrent and proposed instrumentation systems require greaterbandwidth for the transmission of test data than those used isthe past. The bandwidth available for the transmission offlight test data is limited and methods need to be investigatedto alleviate restrictions that limit the quantity and quality ofdata that can be obtained during flight testing. This researchwill allow for the transmission of instrumentation data atgreater rates resulting in improved accuracy andperformance.

ResultsUCLA has completed an initial search of existing modulationtechniques which have the possibility of yielding greaterbandwidth efficiency.

UCLA has also completed the initial modeling andsimulation of equalizers for QPSK systems as well asmodeling and simulation of a Space-Time coding systemusing QAM and a Rayleigh fading channel model.

Status/PlansFlight-testing and evaluation of F-QPSK system performance isplanned for FY00. A PCM/FM system will be flown at thesame time for comparison. Bit error rates and multipathperformance will be evaluated and compared. Compatibilitywith existing ground station equipment will also be evaluatedduring integration.

Modeling and simulation of equalizers for various efficientmodulation systems will be pursued with future efforts focusedon equalizer optimization for DrydenÕs transmission channel.

Space-Time Coding system performance, for DrydenÕs fadingchannel, will be evaluated in future simulations to determinethe applicability of this technique to DrydenÕs fading channelmodel. The use of Space-Time Coding for alternative efficientmodulation techniques will also be the focus of futureinvestigations. If Space-Time coding proves to be effective forDrydenÕs channel model it could be applied to improve theperformance and efficiency of F-QPSK and other efficientmodulation techniques.

Figure 1: Bandwidth of FQPSK vs PCM/FM

Figure 2: Space-Time Coding System

ContactsDon Whiteman, DFRC, RI, (661) 258-3385Kung Yao, Elec. Engr. Dept., UCLA* F-QPSK is Feher patented QPSK.

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X-43A Flight Research Plans Overview

Summary The X-43A Flight Project includesthe development and flight test of threehypersonic aircraft. Each fully autonomousvehicle is nominally 12 feet long, has a 5 footwingspan, and is powered by a single airframe-integrated scramjet propulsion system. The firsttwo vehicles will be flown at Mach 7 and thethird at Mach 10. The research vehicle (HXRV)will be boosted to the desired flight testconditions by a modified Pegasus booster(HXLV). The HXRV will then separate from theHXLV, quickly stabilize, open the engine cowl,and perform a series of pre-programmedmaneuvers to evaluate vehicle performance.

Objectives The overall program objectivesinclude: (1) evaluation of the in-flightperformance of an airframe integrated hydrogenfueled, dual-mode scramjet powered researchvehicle at Mach 7 and 10; (2) demonstration ofcontrolled powered and unpowered hypersonicaircraft flight; and (3) obtaining ground andflight data to validate computational methods,prediction analyses, test techniques, andoperability that comprise a set of designmethodologies applicable to future hypersoniccruise and space access vehicles.

Specific flight research flight research objectivesinclude development of in-flight performanceanalysis techniques, characterization andcomparison of the predicted and flight aero-thermodynamic and propulsion data bases,validating the flight control law design andpropulsion control law design, evaluation of aflush air data system, characterizing theperformance of the vehicle systems, andcharacterizing stresses on key vehicle structures.

Justification The potential payoffs in a scramjetpowered vehicle for access to space and/or highspeed cruise has been well documented. Theflight demonstration of an aircraft powered by anairframe integrated scramjet has never beenaccomplished and will add greatly to theconfidence and credibility of airbreathingpropulsion powered vehicle concepts.

Approach The approach from the beginning ofthe project has been to balance technicalobjectives, safety, and mission risk. To this endcareful attention has been paid to hardware andsoftware qualification and testing. A very largenumber of ground tests and analyses have also

been conducted to better understand everythingfrom engine performance and operability to theeffects of side loads on the pistons used toseparate the research vehicle from the launchvehicle.

Monte-Carlo analyses have played a critical rolein understanding mission risk and maximizingmission success. At the backbone of theseanalyses are three highly sophisticated simulatorscorresponding to the three mission phases,namely launch, separation, and research vehiclefree flight. Significant project resources havegone into the development of these simulatorsand the models and databases used by them.

Status Several significant milestones wereachieved this year. Of particular note is thedelivery of the first research vehicle and launchvehicle to Dryden, verification testing of allsystems aboard the research vehicle, andcompletion of validation testing of the launchvehicle.

The Hyper-X Stack

Contacts Griff Corpening, DFRC, RP, (661) 258-2497Bob Sherrill, LaRC, (757) 864-7085

X-43 Fluid Systems Development

SummaryThe X-43 autonomous, free flying, expendableresearch vehicles will conduct short duration scramjetflight tests at Mach 7 and 10. Each vehicle will bedelivered to the test condition by modified Pegasusbooster launched from a B-52 aircraft. The X-43propulsion system consists of the scramjet engineitself, as well as the associated fluid systems. Thefluid systems consist of the following. A highpressure gaseous hydrogen (GH2) blowdown systemfuels the engine. A pyrophoric 20% silane - 80%GH2 igniter mixture from a blowdown systeminitiates combustion. Gaseous nitrogen (GN2)systems in the X-43, adapter and B-52 are used forcompartment purge, line purge, and coolantpressurant. Pressure-fed water-glycol coolantsystems in the research vehicle and adapter, withtanks incorporating bladder positive expulsiondevices, provide cooling for the engine during highspeed flight.

ObjectiveThe goal of the X-43 fluid system development effortis to ensure their proper operation, for flight safetyand mission success.

JustificationThe fluid systems contain a large amount of energy,both in the form of combustible chemicals andpressure, and therefore constitute a major potentialsource of hazards in the flight program. Also,propulsion systems have historically been a majorcause of failures for expendable launch vehicles.

ApproachFirst, appropriate design and testing requirementswere established. Performance requirements werebased on needs of the scramjet engine and itsresearch objectives. Hardware design and testingrequirements were based on guidance of establishedmilitary standards, and to meet particular NASArequirements. A mindset was taken like that ofexpendable launch vehicles and spacecraft, where themission is a one shot deal and the systems must workcorrectly the first time, with little or no humanintervention.

Vehicle space limitations led to the use of higherpressures, custom designed components, anddemanding performance requirements, therebyincreasing risk and developmental effort. Thecontractor designers implemented palletized

subsystems, which allowed for subsystem testingindependently of the vehicle.

Component and subsystem testing were performed bycontractors and vendors. Qualification testing oncomponents determine if the design is sufficientlyrobust for the operating conditions. Verificationtesting on components and subsystems determine ifthey meet the performance requirements. Acceptancetesting at all levels check the integrity andfunctionality of the hardware.

NASA will perform validation testing on the installedsystems in the vehicle, to ascertain the systems canperform the mission. Captive carry flight will be thefinal test of system integrity prior to the scramjet testflight.

StatusTesting to date have been highly successful introubleshooting, understanding and maturing thecomponents and systems, as intended. Adjustmentsand minor changes were made to the hardware, asissues were identified by testing. Subsystemperformance verification testing was indispensablefor designing the system control logic. Validationtests in the vehicle are expected to commence shortly.All these tests build confidence for safe and correctoperation of the fluid systems in the vehicle.

ContactsMasashi Mizukami, RP, (661) 258-3096

masashi.mizukami@dfrc.nasa.govNeal Hass, RP (661) 258-2641

neal.hass@dfrc.nasa.govJohn Orme, RP (661) 258-3683

john.orme@dfrc.nasa.gov

X-43A Environmental Control Systems Performance

SummaryThe Hyper-X research program, conducted jointly byNASA Dryden and NASA Langley, was conceived todemonstrate a scramjet engine in a flight environment.The X-43A Research Vehicle, the instrument of theHyper-X program, will be lofted to its pre-determinedresearch test condition with the aid of a modified air-launched Pegasus booster. After separation from thelaunch vehicle, the research vehicle will conduct engineoperation tests, and then perform PID maneuvers atselected points in the deceleration phase.

One of the many facets critical to achieving missionsuccess of the operation is control of the vehicleenvironment. Deflagration and detonation hazardsexist due to the nature of the ignitor and fuel that areused for the SCRAM jet engine experiment. In additionto this hazard, there are thermal issues affectingguidance and navigation, instrumentation and controlsystems equipment, and propellant feed systemperformance. Therefore, effective mitigation is critical tokeep both vehicle and crew safe during the missionphases leading up to drop from the B-52 and ensuringmission success.

ObjectiveDryden’s Propulsion efforts have, in part, focused uponmitigation of the fire hazards and assurance that theoperating environment will be maintained untilcompletion of primary mission objectives. Considerableeffort has been expended on analyzing; leak detectioncapability and sensitivity, deflagration and detonationcharacteristics of fuel and ignitor, gas species detectorperformance, purge capability to transport leaking fueland oxidizer, inert vehicle cavities, and cool equipment,and nitrogen system performance. Validation testing onenvironmental control system performance will beconducted for all flight phases.

ResultsDynamic calibrations to temperature and pressurecompensated species detectors demonstrate goodperformance for entire flight phase. Analysis of the firstpurge system tests have indicated that further effortsmust be made to seal the panel seams of the hull toincrease pressure differential between inside andoutside. This will insure that the purge operates asdesigned (see Figure 1) and can overcomeaerodynamic ram air and maintain an inert internalenvironment. Environmental testing to date hasindicated that the hazard of exceeding operationaltemperature limits due to internal equipment heatgeneration does not appear to be a problem, but morecomprehensive tests are pending. The results from thefirst heat test are shown in Figure 2. The nitrogensystem model indicates that the system will be depletedapproximately 1 minute into its flight. This data isshown in Figure 3.

Status/PlansAnalysis of pressure differential and hull leakagecontinues. Further testing with ultrasonic leak detectorsalong panel seams will continue to identify areas ofinadequate seals. Fully integrated vehicle thermal andinert validation tests are scheduled for early March of2000. The true measure of purge effectiveness will bedetermined from the first captive carry flight on the B-52with inerts in the ignitor and fuel tanks.

ContactNeal Hass, DFRC, RP, (661) 258-2641

Hyper-X Research Vehicle Nitrogen Tank Blowdown in Free Flight

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Advances in Leak Detection TechnologiesSummaryLASRE, X-33, Hyper-X, and X-34 are research programs meant,in part, to demonstrate advanced propulsion systems. In orderto demonstrate the performance of these systems, high-energyfuels and oxidizers such as gaseous hydrogen, liquid hydrogen,and liquid oxygen are employed. These types of fuels andoxidizers are inherently hazardous, posing a deflagration, orworse, detonation threat to mission success of the programs.

One of the many facets critical to the mission success of theseflight test operations is the flight safety of the vehicle. Beingable to detect leakage of particular propellant species before theymanifest into a hazard gives the program a greater likelihood ofmission success. This requires a clear understanding of thelimits detectable by species sensors and algorithms used toenhance species detectors.

ObjectiveDrydenÕs Propulsion group is analyzing the effectiveness ofspecies sensors to detect and assess leakage, and understandingthe leak rate relationship between inert and high-energypropellants when inerts are substituted into the propellant feedsystem. Several commercial sensors and a prototype sensorwill, or are in the process of, being tested for their ability toaccurately, and quickly, detect their respective species. Analysisof the data will determine if effective detection capability over awide range of temperature and pressure is possible. If suitable,the instruments can be employed on various flight researchexperiments requiring this capability. Analysis of the inert andhigh energy propellant leak testing will help to scale leakagefrom gaseous propellant feed systems without the use of actualpropellants.

ResultsCommercial-off-the-shelf (COTS) oxygen sensors that wereintended for automobile and medical use have proven quiteeffective for monitoring oxygenated environments. Thecalibrations demonstrated an accuracy of 1% at pressure altitudesup to 50 kFt. The uncertainty analysis results for theseinstruments are shown in Figure 1. A fully pressure andtemperature compensated oxygen sensor design that wascalibrated and dynamically tested, demonstrated accuracy towithin 0.7% while operating at a 60 kFt/min ascent rate from 18kFt to 105 kFt. The results of this dynamical response test areshown in Figure 2. A prototype sensor that measures speed ofsound changes has had significant success in detecting hydrogenconcentrations in nitrogen. The results of this sensor are shownin Figure 3.

Status/PlansThe COTS oxygen sensor and the fully compensated oxygensensor have finished calibrations and are integrated into theHyper-X for flight testing. Validation testing will beginsometime in mid-March. A fully compensated hydrogendetection sensor is scheduled for calibration and dynamic testingin mid-February. If testing meets requirements, it too will beintegrated into the Hyper-X flight test vehicle. Negotiations arepresently under way to procure another prototype speed of soundsensor for characterization and calibration by Dryden. A researchgrant with UCLA is in progress to determine a leak rate

correlation between gaseous hydrogen and gaseous heliumleaking through orifices simulating leaks.

ContactNeal Hass, DFRC, RP, (661) 258-2641

Hyper-X Propulsion Flight Research Plans

SummaryThe Hyper-X Flight Project, conducted jointlyby NASA Dryden and NASA Langley, willdevelop and flight test three hypersonic vehiclesthat are designated X-43A. The X-43A ResearchVehicles will demonstrate the performance of ascramjet engine in a flight environment. Eachfully autonomous vehicle is 12 feet long, has a 5foot wingspan, and is powered by a singleairframe-integrated scramjet propulsion system.The initial X-43A Research Vehicle will bedropped from the NASA Dryden B-52 and loftedto its pre-determined research test condition(Mach 7, 100,000 feet) with the aid of a modifiedPegasus booster. After separation from thePegasus booster, the X-43A will follow a nearlyballistic flight path during the engine test phase.The sequence of events during the engine testare: 1) open inlet cowl, 2) initial engine tare, 3)engine ignition, 4) engine burn, 5) post enginetare, 6) parameter identification maneuvers and7) close inlet cowl. The vehicle will then splashdown in the Pacific Test Range with no plan torecover it.

ObjectivesThe overall program objectives include: (1)evaluation of the in-flight performance of anairframe integrated hydrogen fueled, dual-modescramjet powered research vehicle at Mach 7 and10; (2) demonstration of controlled powered andunpowered hypersonic aircraft flight; and (3)obtaining ground and flight data to validatecomputational methods, prediction analyses, testtechniques, and operability that comprise a set ofdesign methodologies for future hypersoniccruise and space access vehicles.

JustificationSupersonic combustion has been demonstratedsince the 1950Õs in wind tunnel tests. However,the uncertainties in the measurements and testtechniques have made actual scramjet engineperformance open to debate. The X-43A willprovide the first free flying scramjet engineperformance data.

ApproachBecause the engine is so highly integrated into

the hypersonic vehicle the predicted poweredand unpowered vehicle aero-propulsive forcesand moments and the resulting predicted vehicledynamics will be compared with the actualvehicle dynamics through trajectoryreconstruction. A simplified first order successcriteria from the overall vehicle standpoint is thatthe vehicle accelerate under power. It is alsodesirable that engine and engine componentperformances be determined from the flight dataand compared to predictions and ground testdata. Three separate but complimentaryapproaches are planned for post-flight dataanalysis. These are: 1) integrated vehicleperformance through trajectory reconstruction todetermine vehicle acceleration, 2) integratedvehicle performance using the engine test phasesequence to identify force increments, and 3)using pressure and temperatures to determineengine component performance.

Key to the success of each of these methods is arealistic stackup of uncertainties and a verystrong effort to reduce these prior to flight.

Hyper-X Research Vehicle

StatusMonte Carlo simulation runs are being made toidentify the major parameters affecting theengine test phase. Wind tunnel testing continuesat NASA Langley to help identify the scramjetengine performance.

ContactsRoss Hathaway, DFRC, RP, (661) 258-3618Randy Voland, LaRC, (757) 864-6275

SummaryNASA Dryden is working with Virginia PolytechnicInstitute and State University (Virginia Tech) to develop askin friction gage for scramjet flight test application onthe X-43. Although the Virginia Tech skin friction gagehas been used extensively in scramjet wind tunnel tests, ithas never been used in flight. As the results, an extensivedevelopment, testing, and validation program is beingpursued to help develop the best possible gage technologyfor the X-43 scramjet flight application.

ObjectiveDevelop and validate sensor technologies for skin frictionmeasurement inside scramjet engines of hypersonic flightvehicles such as the X-43.

JustificationSurface skin friction force is one of the important forcesaffecting hypersonic flight vehicles and their propulsionsystems. For scramjet engines such as those used in theX-43, scramjet skin friction drag is a significant portion ofthe net engine thrust. Accurate measurement of the skinfriction force is important to validate analysis tools,provide wind tunnel to flight correlation, and determinescramjet engine efficiency.

ApproachThe Virginia Tech scramjet skin friction gage uses thecantilever-beam design. In this design, the shear stressproduced by the flow above displaces the sensor head atthe top of the beam and produces strain at the base of thebeam. Semiconductor strain gages installed at the base ofthe beam measure the strain. The strain gage output isproportional to the wall shear stress.

Although the gage concept described above is simple, it isquite a challenging task to build a flight gage that wouldwork inside a scramjet combustor. The gage will have tooperate in an environment of extreme pressures,temperatures, and heat transfer rates. Also, effects suchas in-flight g-loading and vibration have to be considered.Currently, the following tests are being planned:

• The F-15B/FTF flight test.• The NASA Glenn Trailblazer Mach 6.4 scramjet test

at GASL.• The Hyper-X engine model (HXEM) Mach 7 scramjet

test at NASA Langley.• The Trailblazer Mach 3 direct-connect combustor test

at NASA Glenn.Plans are being made to test skin friction gages in theX-43 flight no. 2 (Mach 7) and flight no. 3 (Mach 10).

ResultsIn July 1999, the scramjet skin friction gage design wastested successfully in the NASA Glenn Mach 6.4Trailblazer scramjet test at GASL. Trailblazer was anRBCC launch vehicle concept at NASA Glenn. It used ascramjet as part of its operating cycle, and the GASL testprovided an excellent opportunity to validate the X-43gage concept in a relevant environment. Good skin

friction data were obtained for three hydrogen−fueledscramjet tests. The skin friction data for one of those testsis shown in figure 2 below. The skin friction gageperformed well during the GASL test with no majorproblems.

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Work is currently underway to build and install scramjetskin friction gages in the NASA Langley Mach 7 HXEMArc-Heated Scramjet Test Facility (AHSTF) test. A skinfriction gage prototype for the HXEM test is shown inFigure 2 below.

Figure 2. HXEM Skin Friction Gage Prototype.

The HXEM skin friction test is scheduled to take place inMarch 2000. HXEM is a wind tunnel model of the X-43scramjet engine. This test will allow the evaluation of theskin friction gage under the realistic operating conditionsof the X-43.

ContactTrong Bui, DFRC, RP, (661) 258-2645, trong.bui@dfrc.nasa.gov

Development of Scramjet Skin Friction Gages

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Gust Monitoring and Aeroelasticity Experiment

SummaryThe Gust Monitoring and Aeroelasticity(GMA) experiment was initiated byUniversity of California, Los Angeles(UCLA) graduate students under theUCLA/NASA Center for Flight SystemsResearch. A small test wing was flownon the NASA Dryden F-15B Flight TestFixture II (FTF-II) flight data on gustmonitoring and aeroelasticity.

ObjectiveThe primary purpose of the GMA flightexperiment was to measure theaeroelastic response of a small test wing,and to monitor in-flight gust intensityusing a customized laser-based device.

ApproachThe small test wing and a low-powerlaser were flown attached to the aft/leftpanel of the FTF-II test article. Thesmall, flexible 18-inch long, 4-inchchord, test wing was instrumented withaccelerometers and strain gages tomeasure the dynamic response of thewing and FTF-II. Video camerasmounted on the underside of the aircraftrecorded the movement of the test wing.

ResultsThe GMA experiment was flown on theF-15B / FTF-II at speeds of up to Mach0.80 and altitudes up to 20,000 feet overa series of 4 flights. Flight data wasused to investigate the relationship of thedynamic response of a flexible wing toin-flight maneuvers and atmosphericwind gusts. In addition, the experimenttested a new gust-monitoring devicebased on measurements of the forwardscattering of a laser beam. Newaeroelastic modeling techniques are

being developed and will be tested usingdata collected during this experiment.These modeling techniques are beingdeveloped to aid in the synthesis ofactive gust alleviation and fluttersuppression control algorithms.

Fig. 1. F-15B with FTF-II

Fig. 2. GMA Experiment on FTF-II

Contact

Oscar Alvarez-Sanchez, TRWStephen Corda, DFRC, RP(661) 258-2103stephen.corda@dfrc.nasa.gov

Contact

Jake Vachon, DFRC, RP, (661) 258-3450

jake.vachon@dfrc.nasa.gov

Summary

Many new, experimental propulsion systems areapproaching Flight Test, including Pulse Detonation Engines(PDE) and components, Rocket-Based Combined-CycleEngines (RBCC) and components, Advanced Inlets, andAerodynamic Models. NASAÕs F-15B provides a unique,low-cost opportunity to develop analysis techniques foradvanced propulsion concepts: (1) Subsonic/Transonic/LowSupersonic data points, (2) Relatively simple flowpaths(PDE, RBCC), (3) Analysis tool validation, (4) Gainvaluable experience in PDE / RBCC in-flight forcemeasurement & flight test techniques prior to full-scaletesting. The Propulsion Flight Test Fixture (P-FTF) isdesigned to support these efforts.

In the past half-year, the P-FTF has evolved from anintangible concept into a design ready for manufacture. Thecurrent design has two main external features: a pylon and aflight experiment (initially a simple ÒtubeÓ for flow qualityanalysis). The pylon is designed for mounting the entirefixture to the F-15B centerline pylon, transfer all loads intothe F-15B suspension lugs, and to house instrumentation andtankage for the propulsion experiment.

The flight experiment is attached to the pylon by a uniquetwo-point force balance. The in-flight force balanceseparates the flight experiment from the pylon and measuresthe forces encountered or generated by the experiment whilein flight. The balance is the result of a novel approach tomeeting several stringent design criteria, includinggeometric constraints within the pylon, providing a methodof mating the pylon to an experiment, tolerance of maximumforces and moments expected in-flight, catastrophic failureprotection, and the versatility to accommodate futurerequirements that are currently unknown.

There are three main bays within the P-FTF pylon. Each baycontains a specially designed rack. The racks are easilyremovable from the lower cavity of the pylon, which allowsfor simple manipulation and setup of the experimentcomponents. The forward bay will notionally contain theinstrumentation for the experiment, while the other two bays(mid and aft) will contain additional instrumentation asneeded and the main tankage: one insulated LOX tank, oneN2 or He pressurant/purge tank, and one fuel tank (JP, etc.).

The P-FTF is expected to be manufactured and flying withinthe next year.

F-15B Propulsion Flight Test Fixture (P-FTF)

NASA 836 In-Flight with P-FTF Concept Mounted

Internal Configuration

The Propulsion Flight Test Fixture

SummaryNASA Dryden is working with Virginia PolytechnicInstitute and State University (Virginia Tech) to develop askin friction gage for scramjet flight test application onthe X-43. Although the Virginia Tech skin friction gagehas been used extensively in scramjet wind tunnel tests, ithas never been used on a flight vehicle. The F-15B/flighttest fixture (FTF) provides an excellent opportunity fortesting the Virginia Tech gage in the flight environment.

ObjectiveEvaluate and validate the Virginia Tech skin friction gageconcept in flight using the F-15B/FTF.

JustificationSurface skin friction force is one of the important forcesaffecting hypersonic flight vehicles and their propulsionsystems. For scramjet engines such as those used in theX-43, skin friction drag is a significant portion of the netengine thrust. Accurate measurement of the skin frictionforce is important to validate analysis tools, provide windtunnel to flight correlation, and determine scramjet engineefficiency.

ApproachA diagram of the Virginia Tech skin friction gage conceptis shown in figure 1 below.

Figure 1. The Virginia Tech Skin Friction Gage Concept.

The non-intrusive floating head is mounted flush with thesolid surface. The shear stress produced by the flow fromabove displaces the floating head slightly. This producesstrain at the base of the cantilever beam. Extremelysensitive displacement sensors measure the strain at thebase of the beam, and the voltage output of the sensorswill indicate the actual skin friction shear stress.

In the F-15B/FTF flight experiment, the skin friction at alocation on the surface of the FTF II will be measured inflight using the Virginia Tech skin friction gage, thePreston tube method, and the Clauser plot method. Theflight conditions will be chosen so that the flow over theFTF II approximates the simple flat plate flow, and agood estimate of the skin friction can be obtained usingthe Preston tube and the Clauser plot methods. Theseestimates can then be used to evaluate the accuracy of theVirginia Tech gage.

ResultsTo date, an F-15B/FTF skin friction sensor complex hasbeen built. This complex was designed to fit into existing8-inch hatches on either side of the FTF, facilitating jointflight testing with other FTF experiments. Shown infigure 2 below, this skin friction sensor complex includesthe boundary layer rake, RTD temperature sensors, heatflux gages, and a Preston tube. The skin friction gage willbe mounted on the plate between the rake and the Prestontube.

Figure 2. The F-15B/FTF Skin Friction Sensor Complex.

In March, 1999, the boundary layer rake and the Preston tubewere tested in the NASA Glenn Research Center (GRC) 8 ×6 Supersonic Wind Tunnel at Mach numbers 0 to 2 withexcellent results. As can be seen in figure 3, skin frictionvalues computed from the rake and the Preston tube dataagree with theory to within ±5 percent.

Transformed Reynolds number based on momentum thickness, Reθ

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Figure 3. The Wind Tunnel Skin Friction Results.

The skin friction sensor complex (without the skin frictiongage) was flown together with the F-15 Gust Monitoring& Aeroelasticity (GMA) experiment in September, 1999 .The skin friction gage is planned to be installed in theskin friction sensor complex and flown together with theF-15 Hot-Wire Anemometry experiment in February,2000.

ContactTrong Bui, Code RP, (805) 258-2645,trong_bui@mail.dfrc.nasa.gov

F-15 Skin Friction Flight Test

B.L. rake

Preston tube

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Temp. sensorFlow

X-43A Propulsion System Controls UpdateSummaryThe Hyper-X research program, conducted jointly by NASADryden and NASA Langley, was conceived to demonstrate ascramjet engine in a flight environment. The X-43AResearch Vehicle, the instrument of the Hyper-X program,will be lofted to its pre-determined research test conditionwith the aid of a modified air-launched Pegasus booster.After separation from the launch vehicle the propulsionsystem will be commanded to follow a prescribed ignitionand fueling sequence. Key to mission success is the ability toaccurately measure and control fuel flow rates, as well asavoiding inlet unstart and flameout stability limits. Currentpropulsion system control (PSC) design features activestability control and fault accommodation.

The propulsion control laws rely on adequate engineperformance and fuel system operability knowledge. Langleyfull-scale engine wind tunnel and GASL fuel system testinghave proven invaluable to algorithm development. Note thatthe design, analysis, integration and implementation ofpropulsion control software are the responsibility of a jointindustry and NASA team of which Dryden is only a part.

ObjectiveDryden Propulsion Control efforts over the past year havefocused on improving the PSC algorithm performance in adetailed design phase and planning for pre-flight validationtests of the Research Vehicle. Of prime consideration hasbeen ensuring a fueling sequence that maximizes thelikelihood for a successful scramjet engine ignition andoperation. Considerable effort has been expended ingathering and reducing GASL fuel system data. The data isbeing utilized to verify fuel system performance with flighthardware. Fault accommodation has been developed toimprove control law robustness to single point failures ofcritical engine and fuel system sensors.

ResultsFuel system and engine performance simulation modeldevelopment is nearing completion. The mass flowestimation and valve timing sequence are being finalized foran updated release of the PSC flight software. Fuel systemoperation and performance has been improved by accountingfor test hardware and fluid characteristics observed duringGASL fuel systems performance tests. Langley wind tunneldata was used to devise and demonstrate logic that will detectand respond to the onset of inlet unstart. Current fuel systemand wind tunnel results predict successful engine operationwith adequate fuel supply and ample stability margins.

Status/PlansThe control system has matured considerably over the past yearand final flight software expected to be delivered by mid-February of 2000. Hardware-in-the-loop and aircraft-in-the-loop tests are scheduled during the first half of CY99 wherevehicle system validation tests are expected to increaseconfidence in hardware integration - including control laws.

X-43A Flight Research Vehicle being prepared in Drydenhangar.

Hyper-X Flight Engine Simulator in Langley 8Õ hypersonicwind tunnel

ContactsJohn Orme, DFRC, RP, (661) 258-3683Ken Rock, LaRC, (757) 864-6265Mark Nugent, Boeing, (562) 593-2256

Development of Remote Site Capability forFlight Research

SummaryLower costs and higher productivity for flight research andflight test activities can be achieved through effective use ofnetwork communications. There are two characteristics offlight test environments that impact how effectivecommunications should operate. First, it is recognized thatdata analyses and signal processing is often required toextract useful information and knowledge from the data.Effective communications over a network must thus enablesignal processing to be conducted at the downstream end ofthe network. Second, flight projects often have a substantialnumber of test engineers and researchers that cannot or neednot be in the control room, but still require rapid access todata. This fact suggests that communications must alsoenable a scalable network of remote sites. Thus, a genericremote site capability for flight research enablesgeographically distributed teams to rapidly assimilate andprocess flight test data.

As an outgrowth of efforts to automate flight flutter testing, anetwork caching data server has been created that manageslive data streams in manner suitable for use in a remote sitenetwork infrastructure. RBNB/DataTurbine servertechnology has been introduced as a commercial product andreceived recognition in NASAÕs 1999 Software of the Yearcompetition.

Dryden engineers have integrated telemetry streams as datasources this past year, developed software tools to managemultiple channel groups, designed and implemented networkdistribution to control room and selected Dryden remotesites, implemented prototype tools for flight flutter testingand more general data visualization, and have initiated designand deployment of remote site nodes that extend DrydenÕsability to support remote sites for flight operations across theAgency.

ObjectiveDryden remote site efforts over the past year have focused ongetting telemetry data flowing through the RBNB, applyingthe baseline functionality in support of flight test operations,extending the capability to meet the demands of supportingnon-Dryden remote sites, and leveraging the AeroSAPIENTproject to actually move data to non-Dryden remote sites.

ResultsA software package called REDDS (RBNB Executive DataDistribution System) establishes baseline capability to extractgroups of tag/data pairs, create engineering units (EU) dataand feed those channels into the RBNB/DataTurbine. Designof a second-generation version has been initiated that permitssophisticated parsing of the tag/data pairs and remote EUprocessing. This version splits REDDS functionality intotwo modules called DAPCapture and tmSplit.

Connectivity to control room and remote sites in PAO andB4800 was established. Matlab-based flutterometer softwarefor improved flutter testing diagnostics began on-line testing.A number of tools were created including:

§ A MATLAB-based data acquisition front-end permitscommanded or data-driven triggering during flight

§ A java plug-in module to RBNB enables on-demandextraction of discrete bits from packed integers

§ A utility to extract time histories from RBNB and writethem to disk in GETDATA and MATLAB file fomats

§ An on-line version of DrydenÕs in-house time history toolQuickPlot was created

§ An NT-based data visualization package calledDataWorks provides high quality stripchart, 2D, and 3Ddata visualization tools

Status/PlansInfrastructure buildup and deployment continues.AeroSAPIENT, F-18SRA/AAW and X-43 dominant users nearterm CY00. Initiate development of improved multi-servermanagement and native ATM support. Validate and/or

evaluate existing and in-progress security enhancements suchas Entrust/PKI encryption module.

MATLAB Data Acquisition Utilities

Data Visualization Client (DataWorks)ContactsLawrence C. Freudinger, DFRC, RS, (661 258-3542Jim Harris, DFRC, FE, (661) 258-2431

AeroSAPIENT Project Initiated

SummaryNASAÕs Aviation Safety Program is part of a nationaleffort to reduce the number of aircraft accidents andincrease passenger safety. The AgencyÕs role withinthe national effort is to accelerate relevant technologydevelopment. There does not yet exist an aviationsystem concept of operations for the year 2020, but allcandidate concepts share fundamental requirements forthe following:

§ Information is available to all users in near real time§ Up to date knowledge of current state of the

national airspace system, airport configurations,terminal airspace, vehicle health, etc.

§ Integration of information and services such thatintuitive intelligent systems exist in the air and onthe ground.

These requirements dictate a comprehensive air andground infrastructure that accommodates all types ofinformation Ð including communications, navigation,surveillance, maintenance or health monitoring, etc. - ina timely manner that maintains integrity and security ofthe overall system. In an effort to examine thefeasibility of bringing the aeronautical community intothe global information network, the IT R&T BaseProgram has created and is sponsoring theAeroSAPIENT project (Aeronautical Satellite AssistedProcess for Information Exchange through NetworkTechnology) in support of the Aviation Safety Program.AeroSAPIENT is a collaborative project involving Ames,Glenn, Langley and Dryden Research Centers.

ObjectiveThe Phase 1 objectives of AeroSAPIENT includeestablishing network connectivity between NASAÕs DC-8 Airborne Laboratory and the ground via commercialKu-band geosynchronous satellite relay. Connectivity isenabled with new electronically steered phased arrayantennas developed under NASAÕs Aviation SystemsCapacity Program. The wireless link will supportmultiple protocols and prioritization of data channels.

Several applications will run on board the aircraft thatare representative of applications and data typesrelevant to aviation safety. These include

§ Weather data into and out of the aircraft, includinglidar-derived turbulence information (via ACCLAIMproject)

§ Digital voice and text messaging§ Engine monitoring and other information

representative of that recorded on digital flight datarecorders (DFDR).

§ Link performance metrics

The aircraft will have an on board data managementsystem based on the RBNB/DataTurbine software.

When the satcom link is active, information can flowto/from the aircraft through a ground station at GlennResearch Center, where data propagates throughAeroSAPIENT remote site server nodes (anRBNB/DataTurbine server network) to researchersacross the Agency.

Status/PlansA prototype on board data management node (COACTrack) was designed and evaluated as a piggybackexperiment during the KWAJEX deployment July Ô99.Modifications to the COACT rack are in progress. Buildupof the satcom rack has been initiated. AeroSAPIENTremote site nodes have been designed and are in a build-up phase. Dryden has contributed resources and isleveraging remote site nodes for general flight testapplication. Most tasks for integrating applications havebeen defined. Plans for CY00 include deployment andcheckout of remote site nodes during 1st and 2nd QtrCY00. Installation of onboard systems scheduled tocommence late July. Flight test over continental U.S.scheduled for August through October.

On-board Data Management Node and Data Visualization

AeroSAPIENT Remote Site Server Node

ContactLawrence C. Freudinger, DFRC, RS, (661) 258-3542l.freudinger@dfrc.nasa.gov

Integration of Finite Element Analysis Predicted Strainswith a Strain Gage Loads Equation Derivation Package

SummaryIt is desired to create strain calibration files that aredirectly usable by EQDE from finite element resultsfiles. This will require development of methods, whichwill convert selected finite elements into strain gageand bridge outputs and arrange them into a formatexpected by EQDE. The equation derivation program(EQDE) main functions are calculation of strain gageinfluence coefficients and derivation of equationcoefficients for the load equations. Demonstration ofthis approach will be with the F18 AAW for theanalytical load calibration of aircraft structures.

Loads calibrations methods used in industry are notsignificantly different than those used by NASA.Currently, both NASA and industry calibrate aircraftstructures by loading the actual structure with hydraulicactuators, recording the output of pre-installed straingages, and then attempting to choose the bestcombination of available gages. This process isdeficient in several ways. It requires engineeringjudgment for placement of a master group of straingages and access to the airframe during a time, whichis usually schedule critical.

ObjectiveThe objective is to increase productivity of strain gagecalibration and reduce time associated with aircraftloads calibration.

JustificationThe increased cost associated with conducting loadcalibrations on aircraft has impacted aircraft programs.Aircraft industry is faced with economic constraints tofind less expensive methods for strain gage loadcalibrations. Future applications could be used todevelop analytical loads calibration for surface areasthat cannot be loaded.

Approach1. Create a finite element model of the structure with

detail in the area of interest.2. Identify strain gage locations to be examined by an

analytical code.3. Generate finite element strain output for analytical

strain calibration.4. Develop a code to translate analytically predicted

strains by NASTRAN into EQDE5. Use the NASA developed equation derivation

(EQDE) with search algorithm to locate minimumnumber of gages and determine error and standarddeviation for shear, bending and torque.

6. Through applied loads in NASTRAN validatecalibration

Future workThis analytical procedure will then be validated againsta calibration performed on the F-18 AAW through afull-scale test in the Flight Loads Laboratory.Comparison of the resulting loads equations will bemade with benefits from the analytical method.

F18 Active Aeroelastic Wing

ContactClifford D. Sticht, DFRC, RS, (661) 258-2538

Non-Invasive In-Flight Structural Deflection Measurement

SummaryAs air vehicle design and predictive toolshave matured and increased in precisionthere has been increased interest in not onlypredicting the airframe elastic shape undervarious flight conditions but also measuringit as well. Early in-flight deflectionmeasurement efforts involved the use ofphotographic film as the sensing medium.Later efforts have involved the use of singlecolumn photo diode array sensors andinfrared light emitting diode target markers.Recent improvements in the resolution andsampling rates of two-dimensional videosensors combined with new softwaretechniques and micro computerdevelopments have provided the potentialfor a new approach to this measurementtask. Video motion analysis has alreadyfound application in athletic sports trainingfeedback. Among other features, thistechnology potentially does away with theneed for wired active targets. For this andother reasons its application to airbornestructural measurement is desirable.

ObjectiveTo begin to mature this new technology forflight use, a considerable amount oflaboratory work must be accomplished first.The following objectives are planned:¥ Demonstrate a ten-fold increase in targetcount (relative to current Flight DeflectionMeasurement System, FDMS).¥ Demonstrate a two and a half-foldincrease in sampling rate (relative to currentFDMS).¥ Demonstrate a two-fold increase inmeasurement resolution (relative tocurrent FDMS)¥ Characterize data stream handlingrequirements for typical flight application.¥ Identify required hardening/ruggedizing ofhardware for flight environment

Approach¥ Procure required hardware and softwareincluding a high speed video camera, on-board computer and storage unit, datareduction computer, and data reductionsoftware.¥ Build up a functioning ground test motionanalysis system.¥ Perform measurement tests with typicaltarget ranges.¥ Generate and record representative datastream.¥ Supply sample data recording to TestInstrumentation Engineers to begindeveloping on-board data recording and TMtechniques.¥ Document results.

Future Work¥ The longer term intention is to fly a videomotion analysis system on the ActiveAeroelastic Wing F-18 aircraft side-by-sidewith our current electro-optical FDMS andcompare results. Following this, expansionto three-dimensional measurementcapability could be pursued.

Active Aeroelastic Wing F-18 With FDMSContactWilliam A Lokos, DFRC, RS,(661) 258-3924

Aerostructures Test Wing

Background/ObjectivesThe AeroStructures Test Wing (ASTW) willprovide a functional flight test planform tovalidate stability prediction algorithms and flighttest techniques. The next generation tools forrobust stability predictions will reduce thecurrently high levels of risk and cost associatedwith flight flutter testing. The ASTW will beflown to the point of instability while real-timestability prediction algorithms will be estimatingthe flutter margins. The design flutter point is0.80 Mach at 10,000 feet. The primaryobjective of this project is to investigate pre-flight and on-line analysis tools for loadpredictions and stability estimation. In-flightflutter estimation algorithms will be evaluated todetermine their accuracy in safely predicting theonset of flutter instabilities and applicabilitytowards an on-line flutterometer concept. Asecondary objective is to investigate open andclosed-loop control strategies for loadsalleviation and flutter suppression.

ApproachThe research approach of the ASTW consists offour main components: analytical modeling,ground tests, flight tests, and technology transfer.The analytical approach will consist of usingSTARS and ASTROS to create structural modelsand predict flight loads and flutter speeds.STARS analysis will also provide a state spacemodel required for the µ flutter predictionalgorithms. The ground tests will consist of aload test to measure the static stiffness of the testwing. A ground vibration test will be done toobtain the structural frequencies and modeshapes. The ground tests will be used to updatethe models and stability estimations. The flighttest of the ASTW will take place on the F-15B/FTF-II test bed. The main objective of theflight test will be to gather static and dynamicload data for load predictions and stabilityestimations. A build up approach from 25% to99% of the flutter speed will be completed on theASTW.

ResultsA design was completed and the wing wasmanufactured this last year. A finite elementanalysis of the completed design was completed.The first structural mode for wing bending wasfound to be 18 Hz and for wing torsion thefrequency was 23.6 Hz. A flutter analysis of thewing estimated the flutter speed to be 437 Keasat Mach 0.80 with a frequency of 22.4 Hz.

Status/PlansThe wing is currently being mounted to the flighttest fixture panel. The panel with the test wingwill undergo a series of ground tests. Theground tests include a wing stiffness and straingage calibration, a ground vibration test, and atest of the excitation system. The results of thetests will be used to update the finite elementanalysis and flutter predictions. TheAerostructures Test Wing should complete itsflight testing during this next fiscal year.

F-15B with Flight Test Fixture

Aerostructures Test Wing Flutter AnalysisResults

Flutterometer Concept, µ-Method

ContactDavid F. Voracek, DFRC, RS, (661) 258-2463Rick Lind, DFRC, RS, (661) 258-3075

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Blackbody Lightpipe Temperature Sensor Development

Summary

A contact temperature measurement technique using ablackbody lightpipe temperature sensor is being developedto control surface temperatures on a carbon-carbon controlsurface above 2000¡ F. A blackbody lightpipe sensorcompleted verification tests in the FLL blackbodycalibration furnace at a peak temperature of 2732¡ F. Thesensor was calibrated to within 2¡ F of a NIST calibratedoptical pyrometer over a temperature range of 1472¡ F to2732¡ F. The sensor is awaiting validation testing in anactual test application.

Objectives

¥Develop a temperature measurement technique to controltemperatures on a carbon-carbon control surface above2000¡ F.

¥Develop a technique to calibrate a blackbody lightpipetemperature sensor up to 2732¡ F.

¥Perform V&V tests of a blackbody lightpipe temperaturesensor.

¥Calibrate the blackbody lightpipe temperature sensorsrequired to support the Carbon-Carbon Control Surface testprogram

Benefit

¥Establishes a method of controlling temperaturesaccurately and repeatedly on carbon-carbon structures athigh temperatures without disturbing the substrate withbonded thermocouples

Contacts

Larry Hudson, DFRC, RS, (661) 258-3925

Tom Horn, DFRC, RS, (661) 258-2232

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Spaceliner 100 Carbon-Carbon Control Surface TestProgram

SummaryWork continued in the Flight Loads Laboratory (FLL) toestablish a 3000¡ F test capability in support of theSpaceliner 100 Carbon-Carbon Control Surface (CCCS)test program. The three main development areas include alarge aluminum test chamber, a heating system, and a waterand gas cooling system. Construction of all major testsystems is complete.

System checkout and integration tests are underway. Theprimary focus to date has been on the purge/gas-coolingsystem. This system utilizes a high output blower to bothpurge the chamber and provide cooling to the radiantheaters. The blower delivers approximately 4000 SCFM ofgas at 14 to 15 psig. Liquid nitrogen is sprayed into theflow to cool the gas, which has been heated by thecompression process, and displace the oxygen in thechamber atmosphere.

Checkout testing of the purge/gas-cooling system has beencompleted and chamber purge levels of 140 ppm oxygenhas been demonstrated. Gas temperature entering theradiant heaters has been maintained at or below 70¡ F.

Objectives¥ Establish a 3000¡ F nitrogen atmosphere test capability.¥ Perform thermal/mechanical test up to 3000¡ F.¥ Support the Spaceliner 100 carbon-carbon development

work through test and analysis of the CCCS.

Benefits¥ Provides valuable test data to the technical community on a

flight-weight carbon-carbon control surface.¥ Establishes methods of testing oxidation sensitive structures

at high temperatures.¥ Serves as a testbed to test innovative sensors.

ContactsLarry Hudson, DFRC, RS, (661) 276-3925Tom Horn, DFRC, RS, (661) 276-2232Dr. Wayne Sawyer, LaRC, (757) 864-5432

Blower(≈ 4000SCFM)

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system

Radiant Heat Flux Gage Calibration System Characterization

Summary Heat flux measurements typically have large uncertainties(typically 10% to 20%) associated with them. Heat fluxcalibration facilities being developed at the NationalInstitute of Standards and Technology (NIST) operate atheat fluxes well below (<20 Btu/ft2/sec) the levelsachieved during high speed flight. This leaves the heatflux gage user interested in these higher heat flux levelsonly two options: 1) take the gage manufacturerscalibration on faith or 2) develop and understand your owncalibration process.

An effort is underway at NASA Dryden’s Flight LoadsLaboratory (FLL) to reduce the uncertainty of heat fluxmeasurements taken at Dryden. The first phase of thiseffort is to thoroughly characterize the radiant heat fluxgage calibration system located in the FLL. Future phasesof this project will develop methods for using gagescalibrated in this system in radiant heating tests performedin the FLL and flight.

Objectives: • Characterize the radiant heat flux gage calibration

system (both blackbody and flat graphite plate heaters)in the Flight Loads Laboratory in order to quantify andreduce, if possible, calibration uncertainty.

• Be able to demonstrate to customers that weunderstand our heat flux gage calibrations.

Challenges: • Quantify errors associated with calibrating a water

cooled heat flux gage inside a cylindrical blackbody• Determine effect of graphite plate erosion• Determine effect of natural convection around flat

graphite plate and gage• Determine effect of distance between flat graphite

plate and gage on absorbed heat flux distributionRecent Results: • Steady state numerical thermal models of the

blackbody have been developed and tuned toexperimental data @ 2012° F (see figures at right)

• Analysis of data from the steady state numerical modelsand experiments show:- negligible convection in the blackbody- effective emissivity of the cylinder end wall is approx.

0.98 for insulated and uninsulated configurationsFuture Work: • Develop 3-D model of the flat plate heater for use in

evaluating calibration configurations• Develop transient numerical models of blackbody/heat

flux gage calibration process• Determine experimental heat flux using different

principals of physics.

Contact: Thomas J. Horn, DFRC, RS, (661)276-2232

Blackbody Source with Wall Temperature MeasurementProbe

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Reference:Jiang, S., Horn, T. J., and Dhir, V. K., 1998, “Numerical

Analysis of a Radiant Heat Flux Calibration System,”HTD-Vol. 361, Proceedings of the ASME HeatTransfer Division, V. 5, pp. 609 – 616.

Summary

Superalloy honeycomb thermal protection system (TPS) is anew candidate TPS for the future space transportation system.The light weight TPS must provide high temperature gradientacross its core depth, stiff enough to resist bulging-outthermal bending, and its cell walls must be stiff enough towithstand buckling under external air pressure. Therefore,thermo-structural analysis of TPS is needed to obtain basicknowledge for the design of low-density honeycomb TPSpanels with high heat-shielding performance, and ample stiffness to resist thermal bending and cell-wall buckling.

Objective

To fully understand the heat-shielding performance andthermostructural characteristics of superalloy TPS panels withvarying honeycomb cell geometry, and to generate a set oftheoretical curves for the designers to determine the mostefficient superalloy honeycomb TPS panel structures.

Approach

For a given superalloy, the heat shielding and thermostructuralperformance of the honeycomb TPS panels vary with thethickness of the face sheets, honeycomb-core depth,honeycomb cell wall thickness, and the size and shape of thehoneycomb cell, etc. Therefore, thermostructural analysis of honeycomb TPS must cover the following key domains: (1)Geometrical analysis of different types of candidate honeycombcells with the same effective density but different crosssectional shapes. (2) Heat transfer analysis of TPS with different honeycomb cell geometry. (3) Thermal bendinganalysis of the TPS panels subjected to one-sided heating andunder different support conditions. (4) Comparative bucklinganalysis of different types of honeycomb cell walls under axialcompression.

Results

1. The heat shielding performance of honeycomb TPS panelis very sensitive to the change of honeycomb core depth, butinsensitive to the change of honeycomb cell cross sectionalshape.2. The thermal deformations and thermal stresses in the TPSpanel are very sensitive to the edge support conditions.3. A slight corrugation of the honeycomb cell walls cangreatly enhance the buckling strength of the honeycomb cellwalls.

Contact

Dr. William L. Ko, NASA Dryden, RS, (661) 258-3581

Thermostructural Analysis of Superalloy Thermal Protection System (TPS)

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Figure 1. Honeycomb thermal protection system (TPS) under surface heating and pressure loading.

Figure 2. Effect of cell size and core depth on the heat shielding performance or superalloy TPS.

Figure 3. Plots of buckling pressure and core density of TPS cell as functions of corrugation depth.

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Validation of the Baseline F-18 Loads ModelUsing F-18/SRA Parameter Identification Flight Data

SummaryThe F-18 baseline loads model validation wasperformed as part of the Active Aeroelastic Wing(AAW) Risk Reduction Experiment on the SRAaircraft. The F-18 loads model was developedby Boeing for use with the flight simulator todetermine loads at discrete locations on theaircraft. It was then modified to reflect theincreased flexibility of the AAW aircraft so thatit could be used in the development of the AAWcontrol laws which are being optimized tomaximize roll performance while maintainingload limits. Currently, the loads models for boththe AAW aircraft and the baseline F-18 arepredicting higher than expected loads especiallyfor the Outboard Leading Edge Flap which hasbecome the limiting factor in the AAW controllaw development. In order to ensure the successof the AAW program, a reasonably accurateloads model must be developed. The purpose ofthis experiment was to determine the level ofconservatism inherent in the baseline loadsmodel.

ApproachThree control surface hinge moments were usedin this experiment: the right outboard leadingedge flap, the left aileron, and the left stabilator.These measured control surface hinge momentswere monitored in flight along with loads modelpredicted values. The loads model values wereobtained by incorporating the loads model codeand database into the real time fortran and usingthe aircraft measured flight data including:Mach, altitude, and control surface deflections asinputs to the model.

Longitudinal and lateral directional doubletswere performed during which each individualcontrol surface was deflected. Measured hingemoments were compared to loads modelpredicted hinge moments at twenty flightconditions ranging from Mach .85 to 1.3 andaltitude from 5000 to 25,000 feet.

ResultsFlight test has been completed on the F-18 SRA.The first stage of data analysis has beencompleted in which measured hinge momentsand predicted hinge moments at each test pointwere compared directly. The measured leadingedge flap hinge moment was approximately 75%

of the loads model predicted value as seen in thegraph. The aileron hinge moment measuredapproximately 40% of the predicted hingemoment while the stabilator hinge momentexceeded the predicted by approximately 10% ofthe flight limit load.

Future PlansCurrently, analysis is being performed in which alinear regression technique is being used todetermine the loadsÕ derivatives for each of thehinge moments, e.g. the load on the leading edgeflap due to control surface deflection. Thesevalues will be directly compared to thederivatives in the loads model database at eachflight condition to determine the existence of anytrends, which can be used to evaluate thedatabase.

F-18 Component Loads Diagram(All Loads Predicted by Loads Model)

Measured and Predicted Leading EdgeHinge Moments vs Time

ContactsDiDi Olney, DFRC, RS, 661-258-3988Heather Hillebrandt, DFRC, RS, 661-258-2843

Measured

SummaryThermal testing in the Flight Loads Laboratory (FLL) hasformerly been accomplished with the use of a mini-computer/array processor combination. The FLL's third generation dataacquisition and control system (DACS III) thermal control systemnow under development, is being designed based on a Sparc10processor. The DACS III thermal control system is at least a yearfrom completion. There exists an immediate need for a smallchannel count thermal control system that can be implementedusing personal computers.Objectives - Develop a PC based control system with thefollowing capabilitiesUse present FLL power controller cabinets for temperaturecontrol functions.• Follow a flight profile with a rise rate Å5¼ F/sec• Provides limit checking per control channel that produces

warnings and automatic shutdown when fatal limits arereached

• Permits maximum power level protection (ability to limitmaximum obtainable power level for a given command)

• Power level recording• Control switching with bumpless transfer from power level

control to closed loop temperature control• Ability to slave control zones• Display and record all data• View data in the control room as well as at the test siteChallenges• Compensating for time lags between the data acquiring

computer and the control computer such that the controlcomputer is processing the correct data set

• Developing a self-tuning algorithm that automatically adjustsas it encounters various temperature ranges, control rates andmaterial thickness

Recent Results• Demonstrated the ability to fire heat lamps on command

using the FLL's present controller cabinet receiving signalsfrom the PC based system

• Demonstrated the ability to track a user generated thermalprofile

Future Work• Implement the remaining system requirements• Investigate using a 'model free' adaptive control algorithm for

this application

ContactAlphonzo J. Stewart, DFRC, RS, (661) 276-3579

PC Based Thermal Control System

ControlComputer

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Test Specimen

TemperatureSignals

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PowerPulses toHeat LampsLAMPS

PowerControllerCabinet

Data Displays

The control system consists of two personalcomputers linked via Ethernet. The acquisitioncomputer receives data from the test article andconverts this data to engineering units. Theconverted data are transmitted to the controlcomputer. The control computer produces firingcommands for the power controller cabinets basedon the desired temperature profile.The control and display features are implementedusing National Instruments Labview software.

Robust Control Design

for the F/A-18 Active Aeroelastic Wing

Summary:

The Active Aeroelastic Wing (AAW) is being developed asa research testbed to demonstrate technologies related toaeroservoelastic e�ects such as wing twist and load mini-mization. Currently, the F/A-18 uses a strong structurewith large stabilators to perform rolling maneuvers at highdynamic pressures. The AAW will demonstrate that wingtwist can be used as an eÆcient tool for maneuvering andthat the induced loads can be reduced beyond current lev-els. A robust control design is formulated for the AAW us-ing a multi-objective approach that considers all the closed-loop goals. Simulations indicate the controller simultane-ously achieves maneuvering and load performance and isrobust to small levels of modeling uncertainty.

Objective:

The objective of the robust control design is to synthesize a ight controller that enables the AAW to eÆciently demon-strate its key technologies. In particular, the controllermust induce wing twist to generate rolling maneuvers. An-other objective is to minimize the loads throughout thestructure during these twisting maneuvers. Furthermore,these objectives must be met without violating constraintsassociated with actuate position and rate limiting that mayseverely restrict the closed-loop performance.

Justi�cation:

The AAW presents an inherently multi-objective designproblem with equally important criteria such as maximiz-ing roll performance while minimizing the load on the struc-ture. The design is further complicated by varying levelsof �delity in elements of the analytical models. Traditionaland optimal control strategies are unable to eÆciently con-sider this type of problem whereas a robust design is inher-ently well-suited for the AAW application.

Approach:

The controller synthesis uses a model that describes themultiple closed-loop objectives. The roll performance andhandling qualities are included by formulating an error sig-nal that describes the di�erence between the responses ofthe AAW and a current F/A-18. The load minimizationis included by formulating error signals that relate the dif-ference between the AAW loads and a set of desired loads.Also, a simple uncertainty description is associated withthe model to re ect potential errors in the structural andloads models. The controller is designed using a worst-caseapproach that attempts to trade-o� objective constraintsand result in reasonable performance for both maneuveringand loads characteristics for the uncertain model.

Results:

Initial controllers have been designed for the AAW thatdemonstrate the concept is feasible. Linear simulations in-dicate the design objectives are satis�ed for maneuveringand loads. Also, the controller can be realized by a simplearchitecture that enables eÆcient implementation for the ight vehicle.

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Bene�ts:� Robust control design provides multi-objective ap-proach to achieve several closed-loop goals

� Load minimization may result in weight savings forfuture aircraft designs by lightening structures

� Roll maneuvering using wing twist may reduce needfor ineÆcient and high-maintainence control surfaces

References:� AIAA Guidance, Navigation, and Control Conference,Portland OR, August 1999

� ICSE International Forum on Aeroelasticity andStructural Dynamics, Williamsburg VA, June 1999

Contact:

Rick Lind, NASA Dryden, RS, x3075

Application of Gain-Scheduled, Multivariable ControlTechniques to the F/A-18 PSFCC

SummaryThe current F/A-18 flight controller was developedfrom a collection of independent linear controllerscreated at specific flight conditions. These linearcontrollers were combined, through gain schedulingtechniques, to create a full envelope controller.Traditional gain-scheduling methods, like thoseemployed on the F/A-18, are inherently ad hoc andthe resulting scheduled controller provides nostability or performance guarantee for rapid changesin the scheduling variables. The lack of a systematicapproach to gain scheduling and an absence ofstability and performance guarantees are among themain motivators for current research intomultivariable gain-scheduled control techniques.

The research area attempting to address the issueslisted above has become known as control of linear,parameter varying (LPV) systems. The LPVmethodology provides for a systematic approach thatexplicitly designs a global controller for all values ofthe scheduling parameters throughout the intendedoperational envelope. The resultant controllerguarantees stability and performance in the presenceof time-varying parameters. In addition, thesecontrollers are robust to errors in model estimationand sensor noise while simultaneously achievingdesired handling quality characteristics throughout theflight envelope.

ObjectiveThe objectives of this research are a demonstrationand evaluation of the design methodology, itsexposure and suitability to a flight environment, andits viability as a design approach for futureapplications. To achieve these objectives, thecontroller will be implemented on the F/A-18 PSFCC.After the controller is verified and validated in ahardware environment it will be flight-tested within alimited, low-energy, relatively safe region of the F/A-18 envelope.

ApproachController point designs are created within thePSFCC Class-B envelope using robust controltechniques. By virtue of the robust design, all pointcontrollers are linear, time-invariant (LTI) and are

represented as state space systems. LPV techniquesproceed to combine all independent point designs tocreate a single, composite controller valid over theentire design envelope. The resultant compositecontroller, however, loses the LTI nature of the pointdesigns since the elements of its state spacerepresentation are now scheduled, through LPVtechniques, as a function of flight condition.Fortunately, the LPV generated schedules can beimplemented in a similar manner to traditionalmethods tending to reduce concerns of excessivememory and computational resource utilization.

Status/PlansLPV designs for both sets of aircraft axes(longitudinal and lateral/directional) are complete andhave been implemented in a Mathworks Simulinkenvironment as well as in a Dryden F/A-18, sixdegree of freedom simulation. Piloted simulationevaluations are expected during the summer of 2000and a flight test is tentatively scheduled for fall of2000.

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PPPPSSSSFFFFCCCCCCCC CCCCllllaaaassssssss BBBB EEEEnnnnvvvveeeellllooooppppeeee LLLLiiiimmmmiiiittttssss

Class BEnvelope

ContactsMark Stephenson, DFRC, RC, (661) 258-2583Rick Lind, DFRC, RS, (661) 258-3075

Fiber Optic Instrumentation Development

SummaryFiber optic sensor technology has been the subject ofconsiderable research interest over the past few years.This emerging technology potentially offers numerousadvantages over conventional aircraft sensors. Acomprehensive research effort has been underway atDryden to develop fiber optic systems, evaluate sensorattachment performance, and evaluate this new technologyin the laboratory, and through ground- and flight-testing.

ObjectiveSystems Development - Develop a self-contained,portable fiber optic measurement system for in-flight andground base testing. Sensor Evaluation - Developattachment techniques and evaluate the accuracy of fiberoptic sensors in controlled laboratory conditions. Groundtest evaluation - Evaluate the feasibility of using fiberoptic sensors in large-scale ground-based test applications.Flight test evaluation - Demonstrate in-flight fiber opticsensor measurements on simple flight structures made ofwell-characterized materials.

JustificationA comprehensive fiber optic research program involvinglaboratory and flight research must be performed first tosuccessfully implement fiber optic technology at DFRC.

ApproachAcquire off-the-self measurement systems and sensors forevaluation. Modify and develop fiber optic digital signalprocessing hardware and sensor attachment techniques.Evaluate sensors and systems in both laboratory and flightenvironments.

ResultsA comparison between four fiber optic strain sensors andconventional strain gages showed agreement to within 2%after sensors were calibrated. The sensors were attached toa cantilever beam and subjected to a concentrated load atthe beam tip. A vacuum/inert atmosphere furnace was alsopurchased to perform high-temperature evaluation of fiberoptic strain sensors on a variety of material systems. Thefurnace completed qualification testing at 3000¡ F. Inaddition, a flight test fixture was designed and fabricatedand instrumented with both fiber optic and conventionalfoil strain gages. The fixture was as a flight testexperiment on the F-18 SRA. Test data from the flightexperiment are currently being analyzed.

ReferenceFiber Optic Instrumentation Development at NASADryden, Proceedings of the Western Regional Strain GageCommittee, SEM; Edwards, CA; February 1-3, 1999.

ContactsLance Richards, Lead & Flight Test PI, (661) 276-3562Allen Parker, Systems Development PI, (661) 276-2407Anthony Piazza, Sensor Evaluation PI, (661) 276-2714Larry Hudson, Ground Test PI, (661) 276-3925

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ain

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)

Comparison Between Fiber Opticand Conventional Foil Strain Gages

Overview ofFiber OpticResearch Program

SensorEvaluation

Ground testEvaluation

Flight testEvaluation

SystemsDevelopment

F-18 SRAFlight Test Fixture

Vacuum/Inert Atmosphere Furnace

54

1999 Technical Publications

1. Gupta, K.K. and Lawson, C.L., Structural Vibration Analysis by a Progressive Simultaneous Iteration Method, Proc. Roy. Soc. A (1999) 455, 3415–3423.

2. Cowan, T.J, Arena, A.S. and Gupta, K.K., Development of a Discrete-TimeAerodynamic Model for CFD-Based Aeroelastic Analysis, Paper No. AIAA-99-0766,37th AIAA Aerospace Sciences meeting and exhibit, January 11–14,1999.

3. Stephens, C.H., Arena, A.S. and Gupta, K.K., CFD Based AeroservoelasticPredictions with Comparisons to Benchmark Experimental Data, AIAA-99-0764.Presented at the 37th AIAA Aerospace Sciences Meeting and Exhibit, Reno, Nevada,January 11–14, 1999.

4. Murray, James E., “The Bungee Mode in Towed Sailplane Flight,” Technical Soaring,January 1999.

5. Gilyard, Glenn B., Jennifer Georgie, and Joseph S. Barnicki, Flight Test of an AdaptiveConfiguration Optimization System for Transport Aircraft, AIAA-99-0831. Presentedat the 37th AIAA Aerospace Sciences Meeting and Exhibit, Reno, Nevada,January 11–14, 1999.

6. Bui, Trong T., A Parallel Finite Volume Algorithm for Large-Eddy Simulation ofTurbulent Flows, AIAA-99-0789. Presented at the AIAA 37th Aerospace SciencesConference, Reno, Nevada, January 11–14, 1999.

7. Teets, Edward H., Jr., and Natalie Salazar, Wind and Mountain Wave Observationsfor the Pathfinder Hawaiian Flight Test Operation, NASA CR-1999-206571,January 1999.

8. Saltzman, Edwin J., K. Charles Wang and Kenneth W. Iliff, Flight-DeterminedSubsonic Lift and Drag Characteristics of Seven Lifting-Body and Wing-Body ReentryVehicle Configurations With Truncated Bases, AIAA-99-0383. Presented at the 37thAIAA Aerospace Sciences Meeting and Exhibit, Reno, Nevada, January 11–14, 1999.

9. Gilyard, Glenn B., Jennifer Georgie, and Joseph S. Barnicki, Flight Test of an AdaptiveConfiguration Optimization System for Transport Aircraft, NASA TM-1999-206569,January 1999.

10. Bui, Trong T., A Parallel Finite Volume Algorithm for Large-Eddy Simulation ofTurbulent Flows, NASA TM-1999-206570, January 1999.

11. Whitmore, Stephen A. and Timothy R. Moes, A Base Drag Reduction Experiment onthe X-33 Linear Aerospike SR-71 Experiment (LASRE) Flight Program, AIAA-99-0277. Presented at the 37th Aerospace Sciences Meeting and Exhibit, Reno, Nevada,January 11–14, 1999.

55

12. Iliff, Kenneth W., Kon-Sheng Charles Wang, Flight-Determined, Subsonic, Lateral-Directional Stability and Control Derivatives of the Thrust-Vectoring F-18 High Angleof Attack Research Vehicle (HARV), and Comparison to the Basic F-18 and PredictedDerivatives, NASA/TP-1999-206573, January 1999.

13. Steenken, W. G., Williams, J. G., Yuhas, A. J., and Walsh, K. R., Factors AffectingInlet-Engine Compatibility During Aircraft Departures at High Angle of Attack for anF/A-18A Aircraft, NASA/TM-1999-206572, February 1999.

14. Whitmore, Stephen A. and Petersen, Brian J. “Dynamic Response of Pressure SensingSystems in Slip-Flow with Temperature Gradient,” AIAA Journal, February 1999.

15. Whitmore, Stephen A. and Timothy R. Moes, A Base Drag Reduction Experiment onthe X-33 Linear Aerospike SR-71 Experiment (LASRE) Flight Program, NASA TM-1999-206575, March 1999.

16. Hunley, J. D., The History of Solid-Propellant Rocketry: What We Do and Do NotKnow, (Invited Paper), AIAA-99-2925. Presented at the 35th AIAA, ASME, SAE,ASEE Joint Propulsion Conference and Exhibit, Los Angeles, California, June 20–24,1999.

17. Greer, Donald, Phil Hamory, Keith Krake, and Mark Drela, Design and Predictionsfor a High-Altitude (Low-Reynolds-Number) Aerodynamic Flight Experiment,AIAA-99-3183. Presented at the 17th Applied Aerodynamics Conference, NorfolkVirginia, June 28–July 1, 1999.

18. Fisher, David F., Fifty Years of Flight Research: An Annotated Bibliography ofTechnical Publications of NASA Dryden Flight Research Center, 1946–1996, NASATP-1999-206568, May 1999.

19. Burcham, Frank W., Jr., Trindel A. Maine, John Kaneshige, and John Bull, SimulatorEvaluation of Simplified Propulsion-Only Emergency Flight Control Systems onTransport Aircraft, NASA TM-1999-206578, June 1999.

20. Saltzman, Edwin J. and Robert R. Meyer, Jr., A Reassessment of Heavy-Duty TruckAerodynamic Design Features and Priorities, NASA TP-1999-206574, June 1999.

21. Hamory, Philip J., John K. Diamond and Arild Bertelrud, Design and Calibration ofan Airborne Multi-Channel Swept-Tuned Spectrum Analyzer. ITC Conference Paper.Presented at the 35th Annual International Telemetering Conference (ITC),“Telemetry: Meeting the 21st Century Challenge,” Riviera Hotel and ConventionCenter, Las Vegas, Nevada, October 25–28, 1999.

22. Ko, William L., Open-Mode Debonding Analysis of Curved Sandwich PanelsSubjected to Heating and Cryogenic Cooling on Opposite Faces. NASA-TP-1999-206580, June 1999.

56

23. Greer, Donald, Phil Hamory, Keith Krake, and Mark Drela, Design and Predictionsfor a High-Altitude (Low-Reynolds-Number) Aerodynamic Flight Experiment. NASATM-1999-206579, July 1999.

24. Carter, John, and Mark Stephenson, Initial Flight Test of the Production SupportFlight Control Computers at NASA Dryden Flight Research Center, NASA TM-1999-206581, August 1999.

25. Burken, John, Ping Lu, and Zhenglu Wu, Reconfigurable Flight Control Designs WithApplication to the X-33 Vehicle, NASA TM-1999-206582, August 1999.

26. Balakrishnan, A. V., Unsteady Aerodynamics – Subsonic Compressible Inviscid Case.NASA/CR-1999-206583, August 1999.

27. Ennix, Kimberly, Griffin Corpening, Michele Jarvis and Harry Chiles, Evaluation ofthe Linear Aerospike SR-71 Experiment (LASRE) Oxygen Sensor, AIAA-99-4815.Presented at the AIAA 9th International Space Planes and Hypersonic Systems andTechnologies Conference, Norfolk, Virginia, November 1–4, 1999.

28. Greer, Donald, Cryogenic Fuel Tank Draining Analysis Model, Paper CLB-3.Presented at the CEC/ICMC 1999 Cryogenic Engineering and International CryogenicMaterials Conference, July 12–16, 1999, Montreal, Quebec, Canada.

29. Orme, John and Robert L. Sims, Selected Performance Measurements of the F-15ACTIVE Axisymmetric Thrust-Vectoring Nozzle, ISABE Paper No. IS 166. Presentedat the 14th ISABE (International Society for Airbreathing Engines) AnnualSymposium, September 5–10, 1999, Florence, Italy.

30. 1998 Research Engineering Annual Report, Compiled by Gerald N. Malcolm,NASA/TM-1999-206585, August 1999.

31. Larson, Richard R., Automated Testing Experience of the Linear Aerospike SR-71Experiment (LASRE) Controller, ITEA Meeting Paper. Presented at the 1999 ITEAInternational Symposium, September 21–24, 1999, Atlanta, Georgia.

32. Larson, Richard R., Automated Testing Experience of the Linear Aerospike SR-71Experiment (LASRE) Controller, NASA/TM-1999-206588, September 1999.

33. Bolonkin, Alexander and Glenn B. Gilyard, Estimated Benefits of Variable-GeometryWing Camber Control for Transport Aircraft, NASA/TM-1999-206586, October1999.

34. Hamory, Philip J., John K. Diamond, and Arild Bertelrud, Design and Calibration ofan Airborne Multichannel Swept-Tuned Spectrum Analyzer, NASA/TM-1999-206584,October 1999.

57

35. Hass, Neal, Masashi Mizukami, Brad Neal, Clinton St. John, Bob Beil, Tim Griffin,Propellant Feed System Leak Detection and Assessment—Lessons Learned From theLinear Aerospike SR-71 Experiment (LASRE), AIAA-99-4805. Presented at theAIAA 9th International Space Planes and Hypersonic Systems and TechnologiesConference, Norfolk, Virginia, November 1–5, 1999.

36. Cobleigh, Brent R., Stephen A. Whitmore, and Edward A. Haering, Jerry Borrer, andV. Eric Roback, Flush Airdata Sensing (FADS) System Calibration Procedures andResults for Blunt Forebodies, AIAA-99-4816. Presented at 9th International SpacePlanes and Hypersonic Systems and Technologies Conference, November 1–5, 1999,Norfolk, Virginia.

37. Ennix, Kimberly, Griffin Corpening, Michele Jarvis and Harry Chiles, Evaluation ofthe Linear Aerospike SR-71 Experiment (LASRE) Oxygen Sensor, NASA/TM-1999-206589, November 1999.

38. Hass, Neal, Masashi Mizukami, Brad Neal, Clinton St. John, Bob Beil, Tim Griffin,Propellant Feed System Leak Detection and Assessment—Lessons Learned From theLinear Aerospike SR-71 Experiment (LASRE), NASA/TM-1999-206590, November1999.

39. Cobleigh, Brent R., Stephen A. Whitmore, Edward A. Haering, Jerry Borrer andV.Eric Roback, Flush Airdata Sensing (FADS) System Calibration Procedures andResults for Blunt Forebodies, NASA/TP-1999-209012, November 1999.

40. Marshall, Laurie A., Boundary-Layer Transition Results from the F-16XL-2Supersonic Laminar Flow Control Experiment, NASA/TM-1999-209013, December1999.

NASA-Sponsored UCLA DocumentsNASA Grant NCC 2-374

(Submitted to CASI in July 1999)

41. Balakrishnan, J. D., “Measures and Interpretations of Vigilance Performance:Evidence Against the Detection Criterion,” Human Factors, Vol. 40, No. 4, December1998, pp. 601–623.

42. Balakrishnan, J. D., “Decision Processes in Discrimination: FundamentalMisrepresentations of Signal Detection Theory,” Journal of Experimental Psychology:Human Perception and Performance/(In press).

43. Balakrishnan, J. D., “Some More Sensitive Measures of Sensitivity and ResponseBias,” Psychological Methods, Vol. 3, No.1, March 1998, pp. 68–90.

44. Balakrishnan, J. D., “Form and Objective of the Decision Rule in AbsoluteIdentification,” Perception and Psychophysics, Vol. 59, No.7, 1997, pp. 1049–1058.

58

45. Tatineni, Mahidhar and Xiaolin Zhong, Numerical Simulations of Unsteady Low-Reynolds-Number Flows Over the APEX Airfoil, AIAA-98-0412. Presented at the36th Aerospace Sciences Meeting & Exhibit, Reno, Nevada, January 12–15, 1998.

46. Lee, Theodore K. and Xiaolin Zhong, Spurious Numerical Oscillations in NumericalSimulation of Supersonic Flows Using Shock Capturing Schemes, AIAA-98-0115.Presented at the 36th Aerospace Sciences Meeting & Exhibit, Reno, Nevada,January 12–15, 1998.

47. Tatineni, Mahidhar and Xiaolin Zhong, A Numerical Study of Low-Reynolds-NumberSeparation Bubbles, AIAA-99-0623. Presented at the 37th Aerospace SciencesMeeting and Exhibit, Reno, Nevada, January 11–14, 1999.

48. Selerland, T. and A. R. Karagozian, “Ignition, Burning and Extinction of a StrainedFuel Strip with Complex Kinetics,” Combustion Science and Technology, Vol. 131,1998, pp. 251–276.

49. Mitchell, M. G., L. L. Smith, A. R. Karagozian, and O. I. Smith, “Burner EmissionsAssociated with Lobed and Non-Lobed Fuel Injectors,” Twenty-Seventh Symposium(International) on Combustion/The Combustion Institute, 1998, pp. 1825–1831.

50. Strickland, J. H., T. Selerland, and A. R. Karagozian, “Numerical Simulations of aLobed Fuel Injector,” Physics of Fluids, Vol. 10, No. 11, November 1998,pp. 2950–2964.

51. Mitchell, M. G., L. L. Smith, A. R. Karagozian, and O. I. Smith, EmissionsMeasurements from a Lobed Fuel Injector/Burner, AIAA-98-0802. Presented at the36th Aerospace Sciences Meeting & Exhibit, Reno, Nevada, January 12–15, 1998.

52. Karagozian, A. R., Control of Mixing and Reactive Flow Processes (Invited), AIAA-99-3572. Presented at the 30th AIAA Fluid Dynamics Conference, Norfolk, Virginia,June 28 – July 1, 1999.

53. Mitchell, M. G., O. I. Smith, and A. R. Karagozian, Effects of Passive Fuel-Air MixingControl on Burner Emissions Via Lobed Fuel Injectors, AIAA-99-2400. Presented atthe 35th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit, LosAngeles, California, June 20–24, 1999.

NASA-Sponsored UCLA DocumentsNASA Grant NCC 2-374

(Submitted to CASI in November 1999)

54. Jiang, Shanjuan, V. K. Dhir, and Thomas J. Horn, “Numerical Analysis of a RadiantHeat Flux Calibration System,” published in Proceedings of International MechanicalEngineering Conference and Exposition (IMECE), November 1998.

55. Friedmann, P. P., “Renaissance of Aeroelasticity and Its Future,” published inJournal of Aircraft, Vol. 36, No. 1, January–February 1998, pp.105–121.

59

NASA-Sponsored DocumentsWest Virginia University NASA Grant NCC 2-759

(Submitted to CASI in November 1999)

56. Napolitano, Marcello R., Carla Iorio, and Brad Seanor, Parameter Estimation fromFlight Data for a Flexible Aircraft, Progress Report, January 1998.

60

Tech Briefs, Patents, & Space Act Awards, FY 99

NASA Tech Briefs Articles

Tim Conners, Robert SimsDirect Thrust Measurement Technique Applied to F-15 ACTIVEPublished May 1999

Ed Teets, Natalie SalazarWind and Mountain Wave Considerations for the Pathfinder Hawaiian Flight Test

Operations

John Muratore (JSC), Chris Nagy

X-38 Crew Return VehiclePublished July 1999

Tony Whitmore, Tim MoesExperiment On Reducing Drag On an Aerospace Launch VehiclePublished August 1999

Sangmon Lee, Linda Kelly, Art Lavoie, Kiefer, ChampaigneRF Wireless Flight Control System

Al Stewart, Allen Parker, Knut Roepel, Van Tran, Manny Castro, Ron Rough, ChrisTalley

Third Generation Data Acquisition Now in OperationPublished October 1999

Dave Richwine, Craig Stephens, Kirsten Carpenter, Michelle Greslik, David McAllisterDurability Studies on the F-15B Flight Test FixturePublished July 1999

Alex Sim, Jim Murray, Tony FrackowiakModel X-33 Subscale Flight Research Program

Published October 1999

61

Klaasen (SVT Assoc/SBIR company)

UV Dual and Photodiode Sensors for Dynamic Combustion Control

Patents

Brent Cobleigh, Tony Whitmore, Ed HaeringFlush Airdata Sensing (FADS) System Calibration Procedures and Results for Blunt

Forebodies

Bryan Duke

Algorithm and Test Technique for a Frequency Excitation Using an Optimized MultipleFrequency Sweep Waveform (Sweepstack)

Paul Curto, Gerry Brown, and Jan Zysko (NASA HQ)Device for Providing AC Cockpit Display of Runway Status, Local Wind Velocity &Direction, and Global Position

Mark NunneleeForce Measuring C-Clamp

Anthony PiazzaAlumina Encapsulated Strain Gauge, Not Mechanically Attached To Substrate, Used to

Temp Compensate an Active High Temp Gauge in Half-Bridge Configurations

Space Act Awards

David HedgleyA Formal Algorithm for Routing Traces on A Printed Circuit Board

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NSN 7540-01-280-5500 Standard Form 298 (Rev. 2-89)Prescribed by ANSI Std. Z39-18298-102

1999 Research Engineering Annual Report

WU 529-50-04-MI-00-RR-0-005-000

Compiled by Edmund Hamlin, Everlyn Cruciani, and Patricia Pearson

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

H-2413

National Aeronautics and Space AdministrationWashington, DC 20546-0001 NASA/TM-2000-209026

Selected research and technology activities at Dryden Flight Research Center are summarized. These activitiesexemplify the Center’s varied and productive research efforts.

Aerodynamics, Flight, Flight controls, Flight systems, Flight test,Instrumentation, Propulsion, Structures, Structural dynamics

A05

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August 2000 Technical Memorandum

Unclassified—UnlimitedSubject Category 99

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