NASA/TM-1999-206585

1998 Research Engineering Annual Report

Compiled byGerald N. MalcolmDryden Flight Research CenterEdwards, California

August 1999

National Aeronautics andSpace Administration

Dryden Flight Research CenterEdwards, California 93523-0273

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Table of Contents

Title First Author Page

Preface Robert Meyer vResearch Engineering Directorate (R)

Organizational Chart viResearch Engineering Directorate Staff viiIn-Flight Observation of Shock Waves David Fisher 1Remote In-Flight Infrared Surface Flow Visualization Daniel W. Banks 2APEX - A High Altitude (Low Reynolds Number)

Flight Experiment Lisa Bjarke 3A Base Drag Reduction Experiment on the X-33 Linear

Aerospike SR-71 (LASRE) Flight Program Stephen A. Whitmore 4Cold Plasma Flight Test Experiment Feasibility Study Tim Moes 6Model X-33 Subscale Flight Research Program Alex G. Sim 7Atmospheric Considerations for Uninhabited

Aerial Vehicles (UAV) Flight Test Planning Edward Teets 8A Stable Algorithm for Estimating Airdata from

Flush Surface Pressure Measurements Stephen A. Whitmore 9X-33 Integration & Real-Time Nonlinear Simulation Cathy Bahm 12X-33 Control Reconfiguration John Burken 13X-33 On-Line Trajectory Design and Control

Reconfiguration John Burken 14U. S. Pilot Evaluation of the Tu-144 Supersonic

Passenger Aircraft Timothy Cox 15F/A-18 Formation Flight Drag Reduction Curtis E. Hanson 16ACTIVE Aeroelastic Wing Controls Design John F. Carter 17Inner Loop Thrust Vectoring Flight Test on the

F-15 ACTIVE Steve Jacobson 18Adaptive Performance Optimization for Transport

Aircraft Glenn B. Gilyard 19Production Support Flight Control Computers John F. Carter 20Separation Analysis of the X-38 Vehicle from the

B-52 Carrier Aircraft Timothy H. Cox 21X-43A Guidance, Navigation and Controls Update Joe Pahle 22Flight Test of an Intelligent Flight Control System

on the F-15 ACTIVE Michael P. Thomson 23ElectroMechanical Actuator (EMA) Validation Program Stephen Jensen 24Flight Termination System Modernization Maria Tobin 25Flight Termination System Transmitter Acquisition Mike Yettaw 26Flight Termination System Telemetry Development Howard Ng 27F-15B Flight Test Fixture Instrumentation

Enhancement James Blanton 28Design of Tu-144LL Airdata Interface Glenn Bever 29Contribution to the Use of an Analog Sweep-

Frequency Spectrum Analyzer for CharacterizingDynamic Data in Flight Phil Hamory 30

Mobile Telemetry Real-Time Instrumentation Support Richard Hang 31Airborne Coherent Lidar for Advanced In-Flight

Measurements (ACLAIM) Rodney Bogue 32Honeywell PPT Smart Sensor Evaluation Gina Branco 33Extended Range Demonstration Flight Test on

the F-15 ACTIVE Bryan Duke 34F-15 ACTIVE Axisymmetric Vectoring Nozzle

Waveform Study John Orme 35

iv

Flight Research of Nozzle Performance for theACTIVE Thrust Vectoring Nozzles John Orme 36

CIAM Scramjet Flight Test Results Trong Bui 37F-15 Skin Friction Flight Test Trong Bui 38Flight Research Plans for the Hyper-X

Research Vehicle Griff Corpening 39Linear Aerospike SR-71 Experiment (LASRE) Stephen Corda 41Evaluation of the Linear Aerospike SR-71

Experiment (LASRE) Oxygen Sensor Kimberly Ennix 42Proposed Vehicle Ground Test Stands Ron Ray 43Laser-Based Mass Flow Sensor on F-18 SRA Kevin R. Walsh 44Simplified Propulsion Controlled Aircraft

(PCA-Ultralite) ACFS B-757 Simulation Frank W. Burcham 45Loading Technique for Light-Weight Structures Stephen V. Thornton 46Selection of Strain Gages for APEX Loads

Measurements: Handbook Values vs. Reality Stephen V. Thornton 47Pegasus Boundary-Layer Glove Experiment

at Mach 8 Leslie Gong 48Fiber Optic Instrumentation Development Lance Richards 49Carbon-Carbon Control Surface Test Program Larry Hudson 50Radiant Heat Flux Gage Calibration System

Characterization Thomas Horn 51On-Line Robust Stability Prediction with

Multiresolution Filtering Marty Brenner 52Applications of Ring Buffered Network Bus Lawrence C. Freudinger 53Ring Buffered Network Bus Technology Maturation Lawrence C. Freudinger 54F-15 ACTIVE Control-Structure Interaction Lawrence C. Freudinger 55Development of an Aircraft Moment of Inertia

Measurement Technique Leonard S. Voelker 56Real-Time Aerothermodynamic Predictions During

X-33 Flight Simulations Craig Stephens 57Acquisition of X-33 Aerothermodynamic and

Thermal Flight Data for RLV Design Tool Validation Craig Stephens 58X-33 Thermal Protection System Durability Studies Craig Stephens 59Transient Thermal Measurements on the Tu-144 at

Supersonic Flight Conditions Craig Stephens 601998 Publications 61Code R Tech Briefs – 1998 66Code R Patents 67Code R Space Act Awards 67

Research Engineering Directorate (R)

Directorate StaffRob Carver, (COTR) (3138)

Kim Flores, (Res Analyst) (3145)

Staff EngineersDr. David Hedgley (3552)

Dr. Kajal Gupta (3710)

Aerodynamics Branch (RA)Bob Curry, Branch Chief (3715)

Flight Instrumentation Branch (RI)Ron Young, Branch Chief (3741)

Dynamics & Controls Branch (RC)Pat Stoliker, Branch Chief (2706)

Propulsion & Performance Branch (RP)Stephen Corda, Branch Chief (Acting) (3687)

Flight Systems Branch (RF)Brad Flick, Branch Chief (3769)

Structural Dynamics Laboratory

Flight Loads LaboratoryKaren Mackall, Manager (3697)

Aerostructures Branch (RS)Mike Kehoe, Branch Chief (3708)

Bob Meyer, Director (3707)Mike DeAngelis, Deputy Director (Acting) (3921)

Gerald Malcolm, Assistant Director (7402)

Everlyn Cruciani, Secretary (3119)Patti Pearson, Office Automation Assistant (3065)

vii

Research Engineering Directorate Staff

Director Robert MeyerDeputy Director (Acting) Michael DeAngelisAssistant Director Gerald MalcolmSecretary Everlyn Cruciani

Branch Codes and Chiefs

RA – Aerodynamics Robert CurryRC – Dynamics and Controls Patrick StolikerRF – Flight Systems Brad Flick (Acting)RI - Flight Instrumentation Ron YoungRP – Propulsion and Performance Stephen Corda (Acting)RS – Aerostructures Mike Kehoe

In-Flight Observation of Shock Waves

Summary

As reported in a 1948 NACA report by Cooperand Rathert, ref. 1, the shadowgraph of shockwaves on aircraft can be observed in-flight if thesun is at a certain angle relative to the shockwave. Cooper and Rathert defined thisrelationship for straight winged aircraft. Thistechnique has been extended for swept wingaircraft . Wing compression shockshadowgraphs were observed on two flightsduring banked turns on an L-1011 aircraft at aMach number of 0.85 and altitude of 35,000 ft(10,700 m). Photos and video recording of theshadowgraphs were taken during the flights todocument the shadowgraphs. GPS and aircraftinstrumentation was used to determine thelocation and attitude of the aircraft relative tothe sun. The shadowgraph was observed forhigh to low elevation angles relative to the wing,but for best results and accurate shock locationrelative to the wing chord, the sun should benearly overhead. Bright sunlight on the aircraftis required.

Objectives

The objective of this experiment was to definethe relationship of the angle of the sun relativeto the wing and shock wave on a swept wingtransport aircraft.

Results

The physics behind this technique are shown infigure 1. The change in density of the air at theshock causes the light rays to be refracted. Sincethe pressure (and density) change is greater nearthe surface, the refraction is greater. This resultsin a dark band immediately behind the shock

Figure 1. Schematic of shadowgraph physics.[1]

Figure 2. Wing compression shock shadowgraphlooking away from sun.

wave followed by a light band. A sampleshadowgraph is shown from the L-1011 isshown in figure 2.

Figure 3. Photo of wingtip shock.

With the proper background, such as scatteredalto-cumulos clouds with a blue ocean below,the shock wave can be observed, fig. 3.

The complete results and the analysis procedureare presented in reference 2.

References1. Cooper, George E. and Rathert, George A.,

Visual Observations of the Shock Wave inFlight. NACA RM A8C25, 1948.

2 . Fisher, David F., Haering, Edward A., Jr.,Noffz, Gregory K., and Aguilar, Juan I.,Determination of Sun Angles for Observationsof Shock Waves on a Transport Aircraft,NASA/TM-1998-206551.

Contact:David F. Fisher, NASA, RA, 805-258-3705Edward A. Haering, Jr., NASA, RA, 805-258-3696

Remote In-Flight Infrared Surface Flow VisualizationSummaryThe feasibility of remotely acquiring infrared images(thermograms) of aircraft surfaces in flight to locatetransition boundaries has been investigated. Previouslythese in-flight thermograms were acquired from acamera located in or on the target aircraft. This methodreduces a number of limitations including limited fieldof view, and the time and expense to instrument eachaircraft on which measurements are desired.

Objective¥ Remotely acquire infrared thermograms in flight.¥ Process images for motion and spatial corrections, and enhancement.

ApproachBy using a NASA F-18 equipped with a remotelyactuated IR camera and tracking system (FLIR) imageswere obtained from target aircraft (T-34C, Lear 24, andF-15). It was determined from the initial results thatsuch images can be successfully acquired and transitionlocations and patterns can be extracted from thoseimages. While this method alleviated many of thelimitations of the onboard system, it has limitations ofits own. These include distortion due to relative motionand spatial geometry changes between images. Furtherresearch was conducted on grant by UC Davis onmethods to correct for motion, spatial geometry, and ongeneral image enhancement.

Infrared thermograms are measures of surfacetemperature. The higher mixing in a turbulentboundary layer exchanges heat with freestream moreeasily than a laminar boundary layer. Therefore awarmer than freestream surface will show highertemperatures under the laminar boundary layer thanthe turbulent boundary layer. The opposite will occurwith a surface cooler than freestream. Some surfacetreatment (thin black vinyl contact sheets) was used toobtain the current images. The surface treatmentinsulated the surface, and the black vinyl was heated bythe sun which provided a heat source.

ResultsWith optimal geometry between the F-18 and targetaircraft spatial resolution as low as 0.1 inches can berealized. The field of view obtained was significantlybetter than similar images obtained with an onboardsystem. The images obtained were comparable inquality to those obtained with onboard systems.

Status/PlansCurrent research plans includes obtaining images atsupersonic speeds, attempting to visualize shocks andflow separation in addition to transition, and to obtainhigh quality images without surface treatment (vinyl).

Contact: Daniel W. Banks, Code RA

(805) 258-2921 Dan.Banks@dfrc.nasa.gov

Transition pattern on leading edge of T-34C right wing.

F-18 with FLIR system.

Infrared image of right outboard wing of Lear 24,in flight.

Freestream

Trip

TransitionBoundary

FLIR Pod

Leading Edge

Boundary Layer Devices

Test Panel

Contact: Lisa Bjarke, RA (805)258 3706

lisa.bjarke@dfrc.nasa.gov

or Don Greer, RA (805)258 2849

don.greer@dfrc.nasa.gov

Summary:The need for cost effective high altitude vehicles toconduct atmospheric research has created interest inhigh altitude, low Reynolds number, and h ighsubsonic Mach number airfoils. APEX, a remotelypiloted sailplane, is being developed to measure airfoilperformance in high altitude flight. The APEXsailplane will be released from a high altitude balloonfrom approximately 108,000 ft altitude. The first 30seconds after release from the balloon are the mostcritical for the APEX flight control system as transitionto horizontal flight occurs with the assistance of foursmall rockets. The APEX airfoil performance ismeasured as the sailplane descends from 100,000 to70,000 ft, at Mach numbers between 0.5 and 0.65, andReynolds numbers between 100,000 and 700,000.Low Reynolds number airfoils typically exhibitlaminar separation bubbles. These separation bubblesare known to have a significant impact on theperformance of an airfoil. The bubble is formed whenthe laminar flow separates as a result of encounteringthe adverse pressure region of the airfoil. Theseparated free shear layer is unstable which amplifiesthe Tollmien-Schlichting instability waves. The freeshear flow generally transitions rapidly from laminarflow to turbulent flow and then reattaches to the airfoilsurface. The lambda shocks, which occur in thetransonic flight regime, are expected to increase theamplification of the Tollmien-Schlichting instabilitywaves.

Objectives of the APEX flight experiment:The objective is to measure the following airfoilperformance parameters to increase the understandingof low Reynolds number airfoils and validation ofairfoil design codes.¥ The section lift.¥ The section drag.¥ The location of the separation bubble.¥ The vortex shedding characteristics.¥ The Tollmien-Schlichting frequencies.

APEX Description:

The APEX design has been completed and fabricationis near completion. The sailplane is 22.7 ft long with awingspan of 41.2 ft and has a wing aspect ratio of 13.6.The sailplane is fabricated from graphite/epoxy andboron/epoxy composites to minimize weight.

APEX - A High Altitude (Low Reynolds Number) FlightExperiment

Apex Mission Profile

Apex 3-view drawing

Separation bubble schematic

After divergence Mach number is reached,compressibility effects dominate and base dragcoefficient rises rapidly. Beyond Mach 1, base dragdrops-off steadily. In the subsonic flight regime, basedrag accounts for approximately 125% of the overallmodel drag. Approximately 80% of the transonic dragrise can be attributed to compressibility effects on basedrag. Baseline LASRE drag data clearly support theassertion that base drag dominates the overall CDo. Ifone is to reduce the overall drag of the vehicle, then the basearea is clearly the place to start.

Drag Reduction StrategyFor blunt-based objects whose base areas are heavilyseparated, a clear relationship between the base dragand the ÒviscousÓ forebody drag has beendemonstrated (ref. 3, 4). As the forebody drag isincreased; generally the base drag of the projectile tendsto decrease. This base-drag reduction is a result ofboundary layer effects at the base of the vehicle. Theshear layer caused by the free-stream flow rubbingagainst the dead, separated air in the base region acts asa jet pump and serves to reduce the pressure coefficientin the base areas. The surface boundary layer acts as an"insulator" between the external flow and the dead airat the base. As the forebody drag is increased, theboundary layer thickness at the aft end of the forebodyincreases, -- reducing the effectiveness of the pumpingmechanism -- and the base drag is reduced. Because theLASRE drag data lie on the steep vertical portion of thecurve, -- a result of the large base drag -- a smallincrement in the forebody friction drag should result ina relatively large decrease in the base drag.Conceptually, if the added increment in forebody skindrag is optimized with respect to the base dragreduction, then it may be possible to reduce the overalldrag of the configuration.

LASRE Drag Reduction ExperimentThe LASRE drag reduction experiment sought toincrease the forebody skin friction and modify theboundary layer at the back end of the LASRE model.

A Base Drag Reduction Experiment on the X-33 Linear Aerospike SR-71 (LASRE) Flight Program

IntroductionCurrent proposed shapes for single-stage-to-orbitvehicles like the Lockheed-Martin X-33 and "Venture-Star" reusable launch vehicle have extremely large baseareas when compared to previous hypersonic vehicledesigns. As a result, base drag -- especially in thetransonic flight regime -- is expected be very large. Theunique configuration of the X-33, with its very large basearea and relatively low forebody drag, offers thepotential for a very high payoff in overall performance ifthe base drag can be reduced significantly. This briefpresents results of a base drag-reduction experiment,conducted on the X-33 Linear Aerospike SR-71 (LASRE)flight program. Complete results of the experiment arereported in ref. 1.

The LASRE Flight ExperimentThe LASRE experiment (ref. 2) is a flight test of aroughly 20% half-span model of an X-33 forebody with asingle aerospike rocket engine at the rear. As shown infigure 1, the entire test model is mounted on top of anSR-71 aircraft. It was intended that LASRE flight testdata would be used to define the aerospike engineperformance under realistic flight conditions and todetermine plume interactions with the base and enginecowl areas.

In order to measure performance of the LinearAerospike engine under a variety of flight conditions,the model was mounted to the SR-71 with a pylon whichwas instrumented with 8 load-cells oriented to allow asix-degree-of-freedom measurement of the total forcesand moments. The model was also instrumented withsurface pressure ports on the forebody, boat tail, base,engine ramps, and the lower engine fence. These surfacepressure ports allowed the model profile drag to bemeasured by numerically integrating the surfacepressure distributions.

Baseline Drag MeasurementsBaseline drag measurements on the LASREconfiguration demonstrate a large transonic drag risethat is significantly larger than the wind tunnel valuepredicted for the X-33. It is likely that the observedtransonic drag difference is an effect of the sting-mountused to support the X-33 wind tunnel model. In thesubsonic flight regime base drag (referenced to theLASRE base area) remains relatively constant atapproximately 0.38 until the divergence drag rise Machnumber of approximately 0.90 is reached.

Clearly, one of the most convenient methods ofincreasing the forebody skin drag is to add roughnessto the surface. Other methods such as using vortexgenerators to energize the boundary layer wouldprobably work more effectively, but their intrusivenessinto the flow precludes this method for application tothe hypersonic re-entry vehicle problem. Benefits ofusing surface roughness are non-intrusiveness(minimal heating), small weight penalty, mechanicalsimplicity, and low cost.

For the LASRE drag reduction experiment # 24 SiliconCarbide (0.035") grit was glued to the skin using spray-on adhesive and the surface was sealed using a high-tensile strength, heat resistant, white enamel paint. Theresulting surface had an equivalent sand-grainroughness that varied between approximately 0.02 and0.05 inches. In an attempt to avoid inducing additionalflow separation at the boat tail or along the forebody,only the flat sides of the LASRE model were gritted.The gritted area covered approximately 1/3 of theforebody wetted area.

Flight ResultsFlight results verified the effectiveness of the surfaceroughness technique for reducing base drag. Figure 2shows the measured base drag reduction compared tothe predicted base drag reduction(refs. 5,6) assuming asurface roughness of { 0.02Ó, 0.05Ó, and 0.10Ó}. Pre-flightcalculations showed that proposed surface roughnessmodifications would result in base drag reductions of8-14%. The actual flight results showed a peak basedrag reduction of approximately 15%. The base dragreduction also persisted well out into the supersonicflight regime. Since base drag of supersonic projectileshad never been previously correlated to viscousforebody drag, the sizable supersonic base dragreduction was a significant positive result.

Unfortunately, flight test results for the rough-surfaceconfiguration did not demonstrate an overall net dragreduction. The surface grit caused a rise in forebodypressures. Coupled with increased forebody skin-drag,the forebody pressure rise offset benefits gained bybase drag reduction. Clearly the techniques used toapply the surface grit must be refined.

Conclusions

1) Base drag dominates the overall drag characteristicsof the LASRE/X33/Venture Star configuration.

2) Flight results corroborate the effectiveness ofadding surface forebody roughness to reduce the basedrag.

3) The base drag reduction benefit persists well into thesupersonic flight regime.

4) Because of the mixed results of the experiment --there was no overall net drag reduction -- the existenceof an ÒoptimalÓ viscous forebody drag coefficient muststill be proven.

ContactsStephen A. Whitmore, Timothy R. MoesPrincipal Investigator(s) RA, (805) 258-2002

tony.whitmore@mail.dfrc.nasa.gov,tim.moes@mail.dfrc.nasa.gov

References

1) Whitmore, Stephen A., and Moes, Timothy R., A Base DragReduction Methods on the X-33 Linear Aerospike SR-71Experiment (LASRE) Flight Program, AIAA 99-0277, January1999.

2) Corda, Stephen, Lux, David P., and Meyer, Robert R., Jr.,Blackbird Puts LASRE to the Test, Aerospace America, February1998, pp. 25-29.

3) Hoerner, Sighard F., Fluid Dynamic Drag, Self-PublishedWork, Library of Congress Card Number 64-19666,Washington, D.C., 1965, pp. 3-19, 3-20,15-4, 16-5.

4) Saltzman, Edwin J., Wang, Charles K., and Iliff, KennethW., Flight Determined Subsonic Lift and Drag Characteristics ofSeven Blunt-Based Lifting-Body and Wing-Body Reentry VehicleConfigurations, AIAA Paper # 99-0383, 1999.

5) Mills, Anthony, F., and Hang, Xu, On the Skin FrictionCoefficient for a fully Rough Flat Flate, J. Fluids Engineering, Vol105, 1983, pp. 364-365.

6) Mills, Anthony F., Heat Transfer, Richard D. Irwin, Inc.,Homewood, IL, 1992, pp 282-328.

*Coefficients Referencedto LASRE Base Area

Legend:Flights 46- 49 Base Drag Coefficient (No Grit)

Flight 51 Base Drag Coefficient (With Grit)

Predicted Base DragReduction, κs = 0.02"

Measured Base Drag Reduction (Flight 51 W ith Grit)

a) Base Drag

Mach Number

0.5 0.6 0.7 0.8 0.9 1.0 1.1 1.2 1.3 1.4 1.5

0.010

0.020

0.030

0.040

b) Drag Reduction Increment

0.050

0.060

Predicted Base DragReduction, κs = 0.05"

Predicted Base DragReduction, κs = 0.10"

κs =0.05"

κs =0.02"

κs =0.10"

Flight data

CD, base

∆CD, base

Figure 2: Effect of Forebody Grit on LASRE Base Drag

Cold Plasma Flight Test Experiment Feasibility Study

Summary: A feasibility study is on-going for a flightexperiment to study the effects of a cold plasma fieldon aircraft aerodynamics. Extensive Russian work inshock tubes, ballistic tunnels, and wind tunnels hasshown that a weakly ionized non-equilibrium plasmacan be used to reduce or eliminate a bow shock andconsequently the aerodynamic drag and heating.Aerodynamic performance benefits have also beendocumented for subsonic flow. Other applications ofthis technology include reduced sonic boom,forebody control, scramjet inlet flow control, andengine noise reduction. NASA Dryden has teamed with Boeing NorthAmerican, Rockwell Science Center (RSC), and theMoscow Institute of High Temperature (IVTAN) tocontinue the development of the plasma generators inRussia and to develop a flight test program atDryden using the F-15B aircraft.

Objectives:¥ Design a flight experiment to quantify the effectsof the cold plasma on the aircraft aerodynamics¥ Design plasma generators and power suppliesusing flight qualified components¥ Continue development of the plasma generators inRussia to reduce technical and flight risks¥ Obtain plasma density and temperature data fromflight for theory development and code validation

Results: In March, a 1/3-scale F-15 nosecone wastested in Russia with both erosive jet and highfrequency (HF) discharge plasma generators at Mach1.8. The main objective of the test was to quantifythe aerodynamic effects of the plasma using surfacepressure measurements, Schlieren visualization of theshock, and a force balance to measure drag. Drydenhad provided pressure and thermo-coupleinstrumentation and the data acquisition system forthe test. Dryden also fabricated the plasmagenerator power supplies and controller. A successful wind tunnel test was required beforeadvocacy of a NASA Dryden flight test project on theF-15B. Success was defined as significantquantifiable aerodynamic effects (such as surfacepressure changes) which could then be tested in afull-scale flight experiment. Unfortunately, verylittle quantitative information was obtained duringthe test. The force balance broke early in the test andonly provided information during the low ambientpressure conditions which were not applicable to theF-15 flight envelope. Some drag reduction was

Contact: Tim Moes, NASA Dryden, RA, 805-258-3054, Ron Young, NASA Dryden, RI, 805-258-3741,

and Arild Bertelrud, NASA Langley (AS&M), 757-864-5559

measured and correlated well with base drag reduction.The forebody surface pressure measurements providedalmost no information due to the extreme amounts of EMIcreated by the plasma generators and possibly by theplasma itself. Some boundary layer total pressuremeasurements did indicate some reduced pressure whenaveraging filters were applied to the EMI contaminateddata. Consequently, the decision was made to not proceedwith the F-15 flight experiment at this time. Advocacy offurther cold plasma system development and testing iscontinuing with the goal of a near future F-15 test . The wind tunnel test did provide some positive resultsin the area of operating and controlling plasmas. The testprovided information needed to improve the plasmagenerator configuration so that a more homogeneousplasma could be sustained over the entire nose region.Also, Schlieren visualization did show plasma influenceon the bow shock.

Status/Plans: Boeing and IVTAN have pursued moresmall-scale wind tunnel testing to optimize the plasmagenerator configuration. This included looking at othertypes of generators (e.g. Tesla coil). Following a series oflaboratory development tests, wind tunnel tests of 1/12and 1/6 - scale nosecones were conducted in Russia atMach 1.8. The detailed results of these tests have not yetbeen released. Once an optimized plasma generatorconfiguration is chosen, the 1/3-scale wind tunnel testwill be repeated in Russia with the hopes of achievingquantitative data on the aerodynamic effects of coldplasma. NASA Dryden and RSC personnel haveparticipated in a limited investigation in the U.S. to reducethe EMI noise on the instrumentation. Some approacheshave been identified, however, further work will need tobe done when the plasma generator configuration forflight test is identified.

1/3 scale wind tunnel nosecone model

4 Plasma JetGenerators

9 HFGeneratorElectrodes

InstrumentationPlugs

Model X-33 Subscale FlightResearch Program

Summary: A 4 ft long, instrumented, model of the X-33 in itsgear-down landing configuration has beenfabricated and flown. The model has flown 29flights to date, and the last 20 have been flownwith 16 channels of instrumentation. Its weight is8 lbs empty and 11 lbs with instrumentation. It isvisually controlled from the ground and has nostability augmentation. It is a challengingconfiguration to fly and land due to its limitedperformance, stability, and control. A typicalflight operation starts by launching the X-33model from a larger, powered, model from 1000 ftagl. There is time to perform one flight datamaneuver prior to set up for landing. The modelis back on the ground about 25 sec after launch.The flight data are then downloaded into a laptopcomputer. The flights have been used to maturethe hardware, establish best vehicle trim andcenter of gravity combination, and to gatherlimited flight data. The 3 lbs instrumentationsystem includes power, sensors, and relatedwiring, and was developed for the X-33 model.Limited flight data has been successfully analyzed.The moments of inertia were recentlyexperimentally determined and a detailedcalibration of the airdata parameters will soonfollow. There are plans to evaluate additionallight weight sensors.

Objectives: ¥ To successfully fly a model of the X-33.¥ Develop a small, light weight, instrumentationsystem suitable for model research.¥ Determine limited X-33 aerodynamiccharacteristics from the flight data.¥ To quantify how well parameter estimationtechniques perform on data from this class ofmodel using the lightweight instrumentationsystem.

Justification: ¥ Model flight testing will often highlight anunforeseen characteristic.¥ Development of flight test techniques andinstrumentation applicable to subscale, lightlyloaded, vehicles.

Model with top off showing instrumentationsystem (center) and construction details.

Results: ¥ An X-33 model was fabricated and successfullyflown by Tony Frackowiak.¥ Jim Murray developed a 16 channelinstrumentation system with sensors, noseboom,and batteries that weighs just over 3 lbs.¥ Quality flight data has been gathered foranalysis when the current ground calibration iscomplete.¥ Several maneuvers have been successfullyanalyzed using parameter estimation techniques.

Model near landing touchdown

Side Benefit: ¥ Video from the flight test have been widelyshown throughout NASA and Lockheed Martin.

Contact: Alex G. Sim, 661 258 3714Jim Murray, 661 258 2629Tony Frackowiak, 661 258 3473

Atmospheric Considerations for Uninhabited Aerial Vehicles (UAV) Flight TestPlanning

Summary:A process to evaluate the atmospheric behaviorin support of UAV planning and operationshas been developed. By examining the hour tohour, day to day, and month to monthvariations in the atmosphere, a picture as to thefeasibility to conducting flight test operationsat any site becomes apparent. Evaluating theforecast output to the real time observedchanges in the atmosphere the meteorologistcan produce a Nowcast that will providevaluable new data to mission planners. Thedesired goal of updating of the mission plan inis to reduce the amount of distance and time theaircraft needs to travel in conducting themission. Using climatology to determinefavorable locations and seasons to fly,supported by real time forecasting andobservations, had made flight operations saferand more repeatable.

Objective:Atmospheric considerations in support ofuninhabited aerial vehicle (UAV) flight testinginvolve characterizing and understanding thelocal atmospheric environment (winds, windshear, temperature, precipitation, andturbulence) in preparation and support ofaircraft operations. A primary objective of thisprocess is to ensure vehicle, test range, andground safety. The generalized atmosphericbehavior of any potential operation sites is bestdescribed by combining the local seasonalclimatology, daily upper atmospheric wind andtemperature profiles, and hourly surface andlow level wind observations. For the supportof UAV flights, a continuous forecast updateprocess based on atmospheric turbulence withsurface and low level wind monitoring isdescribed. Updates ensure mission plannersthe most current available data needed forplanning. Each mission plan is developed tonot exceed operational limits due to adverseweather.

Process Components:¥ Understanding the surface and upper airclimatology¥ Evaluating the model output forecasts andexpected variability¥ Obtaining observation data to furtherevaluate the forecasts¥ Real time measurements and Nowcasts in atimely manner for flight planners

Process Flowchart:

Results:On July 7, 1997 the Pathfinder solar poweredaircraft reached an altitude of 71,500 ft abovethe Pacific ocean off the coast of Hawaii. Daysearlier, the flight process described was usedand because of bad weather a 4 day delay in theflight occurred. In addition, the use of theSODAR enabled the meteorologists to monitorthe development of a low level gravity wavewhich caused the winds to change directionmany times. The relaying of this Nowcasts tothe planners followed by recommended changespermitted the pilots to change from normalprocedures delaying the landing for nearly anhour and to land from the opposite directionadverting a possible accident. When themission was completed a new record altitudefor a propeller driven vehicle was achieved.

Contact:Edward Teets, NASA, RA, (661)258-2924Ed.Teets@dfrc.nasa.gov

A Stable Algorithm for EstimatingAirdata from Flush Surface

Pressure Measurements

Stephen A. WhitmoreBrent R. Cobleigh

Edward A. Haering, Jr.NASA Dryden Flight Research

Center.

IntroductionThe design of a series of

algorithms and the correspondingsoftware used to derive airdata from aFlush Airdata Sensing System (FADS)are presented in this text. The FADSconcept, where air data are inferred fromnon-intrusive surface pressuremeasurements, does not require probingof the local flow-field to compute airdataparameters. This innovation allows theextreme hypersonic heating caused by thesmall radius of a flow sensing probe to beavoided, and extends the useful range ofthe airdata measurement system to theHypersonic flow regime. The FADSalgorithms presented here are used as aflight critical part of the real-time avionicssystems for the Lockheed-Martin X-33,the Orbital Sciences X-34 advancedlaunch-systems technologydemonstrators, and the X-38 assuredcrew recovery vehicle (ACRV). It isanticipated that the algorithms will beused by LMA for the full scale "Venture-Star" Re-usable launch Vehicle (RLV)program. These trans-atmosphericvehicles generally require airdatameasurements for flight-critical sub-systems such as inertial guidance, innerand outer-loop flight control, and forterminal area energy management.

Since all of the above vehiclesmust perform an unpowered landing,knowledge of the dynamic pressure,Mach number, angle-of-attack, andsurface winds is critical for terminal areaenergy management (TAEM) to insurethat the target runway can be reachedunder a wide variety of atmosphericconditions. To determine the best meansof meeting the airdata requirements,

NASA Dryden Flight Research Center(DFRC) performed a feasibility study tocompare the performance and cost of aFlush Airdata Sensing (FADS) system toa set of deployable probes similar to thesystem installed on the space shuttle.(The hostility of the heating environmentsprecluded the use of permanentlydeployed probes.) The study concludedthat a FADS system was moreeconomical by a factor of approximately2. Two issues made the probe-basedsystem prohibitively expensive: 1)integration onto the vehicle airframe, and2) system calibration. The FADS systemrequires no deployment mechanisms andcan be integrated directly onto the vehiclenose cap with no movable parts. Becausethe FADS system does not “probe” theflow field, but instead uses the naturalcontours of the forebody, the flow field ismuch cleaner and is easier to calibrate. Anadditional advantage of the FADS systemis that it offered the potential to senseairdata on ascent, an option not availableto the probe-based system. Based on theresults of the study, the FADS systemwas selected in favor of the deployableprobes.

BackgroundThe DFRC FADS design builds

on work which originated in the early1960's with X-15 program1, continued atNASA Langley2,3, and Dryden Flight

1 Cary, John P., and Keener, Earl P., FlightEvaluation of the X-15 Ball-Nose Flow DirectionSensor as an Air Data System, NASA TN D-2923

2 Siemers, Paul M., III, Wolf, Henry, andHenry, Martin W., Shuttle Entry Air DataSystem (SEADS)-Flight Verification of anAdvanced Air Data System Concept, AIAA Paper88-2104

3 While, D.M., Shuttle Entry Air DataSystem(SEADS) Hardware Development, Vol 1,Summary, NASA CR 166044, January, 1983.

Research Centers 4 ,5 in the 1970's and1980's, and recently concluded flighttesting of an onboard real-time system inthe early 1990's 6, 7. For early real-timeFADS algorithms developed byWhitmore, et. al,6,7, Surface pressuremeasurements were related to the desiredairdata parameters using a calibratedaerodynamic model derived from themodified Newtonian flow theory8. Themodel captures the salient features of thelocal flow, but is simple enough to beinvertable in real-time. Non-linearregression9 was used to invert theaerodynamic model. In this algorithm allsurface pressure measurements were usedsimultaneously to infer the airdata bylinearizing the equations around the resultof the previous data frame. These

4 Larson, Terry J., Whitmore, Stephen A.,Ehernberger, L. J., Johnson, J. Blair, andSiemers, Paul M., III, Qualitative Evaluation ofa Flush Air data System at Transonic Speeds andHigh Angles of Attack, NASA TP-2716, 1987

5 Larson, Terry J., Moes, Timothy R., andSiemers, Paul M., III, Wind TunnelInvestigation of a Flush Air Data System atMach Numbers From 0.7 to 1.4, NASA TM-101697, 1990

6 Whitmore, Stephen A., Moes, Timothy R.,and Larson, Terry J., Preliminary Results From aSubsonic High Angle-of-Attack Flush AirDataSensing (Hi-FADS) System: Design,Calibration, and Flight Test Evaluation, NASATM-101713, 1990.

7 Whitmore, Stephen A., Davis, R.J., and Fife,J. M., In-flight Demonstration of a Real-TimeFlush Air Data Sensing System, AIAA Journalof Aircraft, Vol. 33, Number 5, September-October, 1996, pp. 970-977.

8 Anderson, John D., Jr., Hypersonic and HighTemperature Gas Dynamics, McGraw-Hill BookCompany, New York, 1989.

9 Bendat, Julius S. and Piersol, Allan G., RandomData: Analysis and Measurement Procedures, Wiley-Interscience, New York, 1971.

algorithms were successfully flight testedby Whitmore et. al. on the DrydenSystems Research Aircraft7.Unfortunately, the non-linear regressionalgorithm exhibited problems withalgorithm stability in the transonic andsupersonic flight regimes or in thepresence of undetected sensor failures.These stability problems required ad-hocsoftware patches to artificially aid thestability by throwing “Bad ports” out ofthe estimation algorithm. These ad-hocadditions to the code werecomputationally cumbersome, and did notuniversally stabilize the algorithm for allflight regimes. Because the FADS is to beused for closed-loop flight control, thenon-linear regression algorithm wasdetermined to be too risky and wasabandoned.

The FADS “Triples” AlgorithmTo avoid problems encountered

with the non-linear regression algorithm,a new solution algorithm was developedfor the X-33, X-34 and X-38 spacevehicles. A better solution algorithm isoffered by taking strategic combinationsof three sensor readings to analyticallyde-couple the angles of attack and -sideslip from Mach number, dynamicpressure, and static pressure. Thisinnovation allows for the development ofan estimation algorithm whose solutionspeed is superior to the non-linearregression algorithm, and whose stabilitycharacteristics can be analytically pre-determined for a given port arrangement.

Detailed disclosures of the FADSestimating algorithms, and the associatedsoftware are presented by Whitmore, et.al. In ref 10. A brief derivation of thesolution algorithms are also presented in

10 Whitmore, Stephen A. Cobleigh, Brent R.,and Haering, Edward, A., Design and Calibrationof the X-33 Flush Airdata Sensing (FADS)System, AIAA - 98 - 0201, Prepared forPublication at the 36th AIAA AerospaceSciences Meeting and Exhibit, January 12-15,1998 , Reno Hilton, Reno, NV, Also Publishedas NASA TM: 1998 - 206540

the attached appendix. A design criterionfor insuring algorithm stability has beendeveloped; This analysis-- although easyto apply -- is extremely complex and willnot be presented here. The reader isreferred to reference 10 for this analysis.

Redundancy ManagementThe number of measurements in

the pressure matrix was selected as acompromise between the need toaccurately measure the flow conditions atthe nose, and the cost of locating ports onthe vehicle. At least 5 independentpressure measurements must be availableto derive the entire airdata state. Usingfive sensors to estimate the airdata isequivalent to a higher order spline fit andwill result in an estimating algorithmwhich is sensitive to noise in themeasured pressures. Providing anadditional 6th sensing location mitigatesthe noise sensitivity, increasesredundancy options, and results in asystem which gives overall superiorperformance. Each flight-criticalmeasurement subsystem must have a fail-operational (Fail-Op) capability. That is,the subsystem can tolerate one failureanywhere in the system software orhardware and still produce a usableresult. Typically, this Fail-Op capabilityis achieved by installing tri-redundantsystems, each operating alongindependent paths. The results of thethree systems are then "voted" and themedian value signal is selected as themost reliable. The FADS design exploitsthe built in redundancy of the pressureport matrix, to achieve fail-op capabilitywith dual redundant system hardware.This innovation allows for considerablesavings in up-front costs.

Dual redundancy is achieved ateach surface measurement location byinstalling a plug with two surface ports.This dual-port design for the FADSsystem provides a total of 12 surfacepressure measurements; however, thepressures from the dual redundantpressure ports are always analyzedindependently. Defining data flow path Ias the set of computations which use the

grouping of the six upper and outboardpressure measurements on each plug, anddata flow path II as the set ofcomputations which uses the 6 lower andinboard pressures, then the Fail-Opcapability of the system is always beinsured by selecting the computationalpath with the minimum mean-square fiterror (MSE).

This redundancy management schemeselects the system with the best overall fitconsistency, and allows for a soft sensor failure -- one which is not detected by the hardwarediagnostics -- to occur without degrading theperformance of the system. If the MSE of theoutput flow path is normalized by an expectedpopulation variance -- that is, by the expectedrange of fit error that is allowable for a systemwith no failures -- then the MSE becomesdistributed as χ2 and is a good indicator of theabsolute system health.

SummaryThe design of a series of algorithms and thecorresponding software used to derive airdatausing a Flush Airdata Sensing System (FADS)are presented in this text. The FADS algorithmspresented here are used as a flight critical part ofthe real-time avionics systems for the Lockheed-Martin X-33, the Orbital Sciences X-34advanced launch-systems technologydemonstrators, and the X-38 assured crewrecovery vehicle (ACRV). It is anticipated thatthe algorithms will be used by LMA for the fullscale "Venture-Star" Re-usable Launch Vehicle(RLV) program. The FADS design utilizes amatrix of pressure orifices on the vehicle nose toestimate air data parameters. A fail-op capabilityis achieved with dual-redundant measurementhardware, giving two independent measurementpaths. The airdata parameters corresponding tothe measurement path with the minimum fit-errorare selected as the output values. This methodallows for a single sensor failure to occur withminimal degrading of the system performance. Inprevious flight-critical sub-systems, this fail-oplevel of redundancy has been achieved only byusing tri-redundant hardware and software paths.Thus the current FADS design offersconsiderable savings.

X-33 Integration & Real-time Nonlinear Simulation

SummaryThe X-33 program will demonstrate new technologiesrequired for a Reusable Launch Vehicle (RLV) using a half-scale prototype. The X-33 will be an unmanned vehicle,launched vertically, reaching an altitude of over 200,000feet at speeds approaching Mach 10. The vehicle willoperate autonomously from launch to landing. Some ofthe technologies to be demonstrated are: metallic thermalprotection system, linear aerospike engines, and anintegrated vehicle health monitoring system. NASADryden is supporting this activity through thedevelopment of an X-33 Integrated Test Facility (ITF).

Objective: The ITF provides an essential role in the overallsystems development of the X-33. The initial phase will bethe development and test of a real-time, nonlinear 6-DOFsimulation, that supports all phases of flight. Activitieswill culminate with the integration of flight avionicssystems into a hardware-in-the-loop (HIL) simulation forverification and validation testing.

Approach: The development approach begins withsimulation models for the X-33 vehicle, e.g., aerodynamics,reaction control system, actuators, engine, navigation,guidance, and control. Each model is incorporated into theX-33 batch and real-time simulations. The real-timesimulation has been integrated with the 1553 buses usedfor communication with the avionics hardware. When allof the hardware are fully integrated into the simulationand the final operational flight program (OFP) is delivered,formal verification and validation testing will beperformed to certify the system readiness.

Status: Current work is focused on model updates andhardware and software integration to support full, ascent -rollout, hardware-in-the-loop capability by late 1999.Formal Verification and Validation is scheduled to beginthe fall of 1999.

The first X-33 batch end-to-end simulation was completedin May 1998. Integration of the triplex VMC system forAscent flight was achieved in September 1998. A single setof FADS Hardware was integrated and INSGPS integrationand test was performed at MSFC using the ITF simulation.Hardware Forward, Rear and Engine Data Interface Units(DIU) and Engineering Test Stations (ETS) were deliveredto the lab and integration with the simulation has begun.

Contact: Cathy Bahm, NASA Dryden, RC, x-3123, BobClarke, NASA Dryden, RC, x-3799, Louis Lintereur, NASADryden, RC, x-3307

Longitude deg

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36

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38

39

40

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The X-33 Vehicle in simulated flight.

The X-33 6DOF simulated flight from liftoff atEdwards AFB to landing at Michael site inUtah.

X-33 Control Reconfiguration

SummaryThe X-33 Vehicle will include theautomatic control system reconfigurationcapability in the event of an actuatorfailure. The reconfigurable system willincrease the possibility of landing theÒcrippledÓ X-33 at a landing site with ajammed or floating actuator. Nonlinearsimulation results show a definiteimprovement with reconfiguration ascompared to the nominal control systemwith the same failure.

Objective: Increase the likelihood oflanding the X-33 with a failed actuator ordelay flight termination to a less severelocation. Other objectives includemaintaining stability, rejecting gust andperform maneuvers while havingacceptable stability margins.

Justification: The X-33 is an all electricactuated vehicle (all the control surfacesare powered by electric actuators rated at270 Volts/50 Amps). This is a relativitynew actuation method and hence has ahigher risk involved. To mitigate the riskof a failure, the reconfigurable controllaws were developed.

Approach: The constrained controlallocation approach was taken for thereconfigurable design. The X-33 has 8control surfaces and in the event of onefailed surface the other 7 healthy surfacesare used to control the vehicle. The off-line sequential quadratic programmingmethod was used because rate saturationand rate limiting can be accounted for inthe design. The ability to incorporate thenonlinear surface rate limiting andposition limiting was very important inthe success of the controller.

Nonlinear simulation comparison of EntryNominal controller Vs Reconfigurablecontroller for outboard elevon jammed at 25degrees. Reconfigurable controller is stable.

X-33 Vehicle and Aerosurfaces

Benefits¥ Land a ÒcrippledÓ X-33.¥ Delay flight termination¥ Improve performance

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Contacts: John J. Burken, Lead X-33 ControlsEngineer, NASA Dryden, RC, X-3726 ValerieGordon, Flight Controls Engineer NASADryden, RC, X-2018

X-33 On-Line Trajectory Design and Control Reconfiguration

SummaryThe X-33 Vehicle will fly a pre-definedtrajectory; these trajectories have beendeveloped in such away that thermo or aeroload limits are not exceeded. The trajectorydesign process takes a lot of computerpower to do an exact trajectory. The intentof this study is to design an approximatetrajectory when there has been a failure onboard, and land with minimal damage. Byusing some approximations, the trajectorycan be re-designed on-board and therefore ifa failure occurs, the guidance system couldhelp command the vehicle to a benign flightpath. We are investigating whether it ispossible to redesign the trajectory toaccommodate the reduced controleffectiveness when the failure occurs.

Objective: The objective is to use theSequential Quadratic Programming (SQP)method for combined guidance (trajectorydesign) and investigate differentreconfigurable control law methodologies.From nonlinear simulation studies,determine the ÒbestÓ method.

Justification: The X-33 will havereconfigurable controls on-board but thedesign method may change depending onthe simulation trade studies.

Approach: Use SQP for the trajectory re-design when a failure occurs and use twomethods for control system reconfigurationfor the simulation trade-off study. The twomethods for control reconfiguration are; 1)robust servomechanism design and 2)control-allocation approach based on aquadratic programming formulation.

Note: The two control system methods arebeing designed for flight test on DrydenÕs F-18 PSFCC aircraft, and hope to fly in the fallof 1999.

X-33 Vehicle and Aerosurfaces

Benefits¥ Land a ÒcrippledÓ X-33.¥ Build new trajectory due to a failure.¥ Design a simpler reconfigurable controller¥ Impact RLV reconfigurable control laws.

Contacts: John J. Burken, Lead X-33 ControlsEngineer, NASA Dryden, RC, X-3726

Dr. Ping Lu, Iowa State University

TrajectoryGeneration

PlantAircraft

ControllerK

UncertaintyDelta

Ud

Xd

On-line Trajectory Generation to construct feasible trajectories/controller.

U. S. Pilot Evaluation of the Tu-144 Supersonic Passenger AircraftSummaryIn September of 1998 two U. S. pilots conducted ahandling qualities evaluation of the Tu-144supersonic passenger aircraft. Three flights wereflown to evaluate the take-off, climb to cruise,cruise, descent from cruise, and approach andlanding characteristics. The results indicated heavycontrol forces, sensitive pitch dynamics, goodlateral/directional dynamics, and good approachand landing characteristics.

ObjectiveCollect U. S. pilot comment data on the Tu-144flying qualities. Of special interest was anevaluation of the Mach 2 flight regime and severalevaluations of the approach characteristics.

JustificationApproximately 5 years ago the Tupolev AircraftCorporation, through funding by the High SpeedResearch program, re-engined, refurbished andinstrumented a Tu-144 aircraft to support the HighSpeed Civil Transport (HSCT) program as anexperimental supersonic flying laboratory.Nineteen flights of the modified aircraft were flownfrom the fall of 1996 to Feb. 1998, and data for sixflight experiments were acquired, one of which wasin the field of handling qualities. One of the primeobjectives of the handling qualities experiments wasto collect data to validate handling qualities criteriabeing used in the design of the HSCT. However, thehandling qualities data collected during thisprogram was quantitative in nature. The data,although useful in the application of handlingqualities criteria, did not include pilot derivedqualitative data. Thus, no comparisons betweencriteria predictions and pilot commentary could bemade.

ApproachOf the three evaluation flights flown by theAmerican pilots, the first was restricted to besubsonic. The last two were supersonic flights.withapproximately 20 minutes of Mach 2.0 time each. Aset of maneuvers was defined to aid the pilot inevaluating the aircraft. They were pitch attitudecaptures, bank angle captures, heading captures,steady heading sideslips, airspeed captures,simulated engine failures, and slow speed flight.With these maneuvers evaluations of the subsonicand supersonic cruise condition were conducted aswell as the nominal take-off and approachconfigurations. In addition the pilots evaluatedapproaches with both manual throttle and autopilotoperation, both canard extended and retractedconfigurations, and a simulated engine failure.

Contact:Timothy H. Cox, NASA Dryden, RC, X2126

ResultsThe pilot comment data are summarized below:

¥ The wheel and column control forces are large,especially noticed in the roll axis when effectingreasonable roll rates.¥ The acceleration and climb to Mach 2 is a highworkload task. This workload can be attributed toinherent pitch sensitivity in the aircraft, which maybe a function of the aggressiveness of themaneuvering, and a sensitive pitch attitude scalecoupled with poor out of the cockpit visualreferences.¥ Good, predictable lateral/directional dynamicsexist throughout the envelope.¥ Although the aircraft lands on the Ôback sideÕ of thepower curve, using throttle manually to maintainairspeed during approaches was not difficult.¥ In crosswinds the aircraft landed in a crab tends toalign itself with the runway and is easy to perform.¥ Using the ground effect while maintaining aconstant pitch attitude is an effective, easy techniquefor soft landings.

Pilot ratings, without considering the high degree ofcontrol and column forces, were level 1 in thelateral/directional axis and level 2 in the pitch axisfor all configurations and flight regimes tested.

StatusComparison of these pilot evaluations to such criteriaas CAP, short period and dutch roll damping, timedelay, Neal/Smith, bandwidth, flightpathbandwidth, and time-to-bank is currently on-going.

Tu-144LL landing at Zhukovsky Air DevelopmentCenter near Moscow, Russia (EC98-44749-25)

F/A-18 Formation Flight Drag ReductionSummaryAircraft flying in formation can take advantage ofeach otherÕs wingtip vortices to improve theoverall formation efficiency. Two aircraft flyingwingtip-to-wingtip can achieve the performance ofa single aircraft with twice the aspect ratio. Whilewingtip-to-wingtip flight is unrealistic, this effectdissipates slowly with longitudinal separation.Two aircraft flying with wingtips aligned andwithin two or three wingspans longitudinally ofeach other can achieve significant drag reduction.

ObjectiveDemonstrate through flight test the potential ofachieving enhancements in performance throughclose formation flight. Measure the drag reductionand required control compensation associatedwith flight within the wingtip vortex of anotheraircraft.

JustificationDrag reduction through formation flight cansignificantly reduce fuel costs for passenger andtransport aircraft. Additionally, high-altitude,indefinite-duration missions may be performed byformations of smaller, more durable solar poweredaircraft, allowing maintenance rotation withoutloss of mission. This research also has benefits inthe areas of Uninhabited Combat Air Vehicles(UCAVÕs), autonomous refueling and vortexcancellation.

ApproachThe F/A-18 Systems Research Aircraft (SRA) wasflown in left echelon formation with a chase F/A-18 aircraft. Maintaining constant throttle, the pilotplaced the SRAÕs right wingtip into the vortex shedby the leading aircraft. The pilot applied rightlateral stick to maintain wings-level and reducedpower as required to match speed with the leadaircraft. Once again maintaining constant throttle,the pilot exited the vortex and observed any loss inspeed.

ResultsA 10-15% drag reduction associated with flying inthe wingtip vortex of the lead aircraft wasmeasured. This result was repeated on multipleflights. Pilot compensation was required primarilyin the roll axis and amounted to approximately20% of the available roll authority. Pilot commentsindicated a noticable increase in speed associatedwith flying in the vortex and estimated the widthof the vortexÕs outer half to be on the order of fivefeet.

Contact:

Curtis E. Hanson, Principle InvestigatorNASA Dryden, RC, X-3966

Lateral Compensation and Drag CoefficientAssociated with Formation Flight

StatusThe degree of pilot workload required for formationdrag reduction on long flights and the application ofthis approach to formations of UAVÕs point to theneed for an Autonomous Formation Flight (AFF)system. Efforts are currently underway at DFRC todevelop a prototype AFF system. GPS receivers willbe used to provide position estimates accurateenough for autonomous outer-loop control in closeformation flight. Initial station keeping tests of thissystem using the SRA are scheduled for early 1999.

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

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NASAÕs Systems Research Aircraft (SRA)and F/A-18 Chase

ACTIVE Aeroelastic Wing Controls Design

SummaryThe ACTIVE Aeroelastic wing program will usean F/A-18 with more flexible wings to flightdemonstrate aeroelastic wing technology.Precise roll control of the aircraft will beperformed using only wing and rudder surfacesbeyond aileron reversal flight conditions.

Objective: To demonstrate precise aircraftcontrol with good handling qualities using wingroot deflection. To perform this task, structuresand aerodynamics and controls need to bedesigned concurrently in order to satisfyprogram constraints. Modern controls methodsinvolving optimization will be needed toaccount for handling qualities, load constraints,actuator performance, and stability.

Justification: Aircraft of the future have a needfor lighter, more flexible structures. In order tominimize loads and eliminate flutter in wings,the flexible structure needs to be used as asurface to provide controlling moments to theaircraft.

Status: Two flight control design methodologiesare being pursued. One is based on directstructural optimization, producing surfacedeflections which will provide for absoluteminimum wing loading. The other method,based on H infinity methods, keeps the wingloading within specified bounds whileattempting to maintain handling qualties withobtainable surface rates and accelerations. In1999, one methodology will be selected andused for the AAW control law.

Contact: John F. Carter , PrincipalInvestigator NASA Dryden, RC, X-2025

F/A - 18 During Flight Research

Benefits¥ Reduced weight and cost for aircraft wings

¥ Development of aeroservoelastic, finite element, and aerodynamic prediction tools

¥ Development of control law design strategies for flexible wings

wing rotation beyond aileron reversal

Airstream std wingAAW wing

Inner Loop Thrust Vectoring Flight Test on the F-15 ACTIVE

SummaryThe Inner Loop Thrust Vectoring (ILTV) control lawsutilize the Pratt and Whitney Pitch/Yaw Balance BeamNozzles (P/Y-BBN) as a primary control effector on theF-15 ACTIVE. The nominal configuration uses the pitchnozzles to produce 50% of the commanded pitchingmoment and also uses yaw nozzles to augment yaw control.

F-15 ACTIVE during a research flight

This enhanced (research) control system resides in anadditional processor within the flight controller. The pilotmay switch between the conventional mode and enhancedmode via the trigger switch. The control law, which wasdesigned by Boeing Phantom Works, uses a pseudo-dynamic inversion process that cancels out the naturalvehicle dynamics and commands desired dynamics.

This control law was auto-coded from graphical blockdiagrams that were created in Matrix-X System build. Thesignificance of this is that ACTIVE is the first pilotedvehicle to use this auto-coded control laws.

In addition to the baseline system there are 15 additionalpreset Dial-A-Gain (DAG) options. These DAGÕs allowthe pilot to select configurations that vary control ratios andsystem handling qualities. Of particular interest is DAG 25which is referred to as the longitudinal thrust vectoring onlyDAG. DAG 25 schedules the horizontal stabilator toprovide longitudinal trim, and uses the pitch vectoringnozzles for stabilization and dynamic pitching maneuvers.

ObjectiveThe objectives of the program are to:1) Demonstrate the P/Y-BBN thrust vectoring nozzles in anoperational environment. 2) Evaluate the process forimplementing software that has been auto-coded fromblock diagrams. 3) Demonstrate that thrust vectoring maybe used as the primary control effector.4) Demonstrate the use of thrust vectoring for powerapproach and landing. 5) Evaluate high angle of attackflight in an operationally representative vehicle.

ApproachMost of the flight envelope will be cleared using thereversionary ILTV control laws. Automatic reversion

occurs when the vehicle exceeds flight envelope boundaries(speed, g, angle of attack, etc.) or encounters systemfailures. The reversionary phase will be followed by flighttest of the non-reversionary ILTV control laws. Envelopeexpansion will be accomplished at discrete points in theenvelope and will include handling qualities evaluations.

StatusFlight tests of the Reversionary ILTV control laws wereconducted during November and December of 1998. Thecontrol laws were cleared for basic maneuvering at flightconditions ranging from 20 kft/0.6 Mach to 30 kft/ 0.9Mach. 1 g expansion up to 25° angle of attack and out toMach 1.2 was also completed.

Handling qualities evaluations were accomplished at fourflight conditions. The two tasks performed were formationflight, and 3 g tracking. The ILTV enhanced modeconsistently performed better than the conventional modecontrol laws during these handling qualities tasks. Thetable below summarizes some of the Cooper-Harper ratings(CHRÕs) awarded. Note that lower numbers indicate betterhandling qualities.

Cooper-Harper ratings forLongitudinal tasks

Task Mode Long gross acqu. Long fine trackAve CHR Std. dev. Ave CHR Std. dev.

Formation Conv 3.5 1.7 2.1 0.4Formation Enh DAG 0 2.4 1.1 1.9 0.93 g tracking Conv 2.9 1.1 2.7 1.03 g tracking Enh DAG 0 2.4 1.0 1.6 0.73 g tracking Enh DAG 25 1.8 0.8 2.0 0.7

Cooper-Harper ratings forLateral directional tasks

Task Mode Lat-dir gross acqu. Lat-dir fine trackingAve CHR Std. dev. Ave CHR Std. dev.

3 g tracking Conv 2.8 0.8 2.9 1.23 g tracking Enh DAG 0 2.5 0.5 2.4 0.5

In general the pilots commented that the enhanced modecontrol laws were very solid, performed very well and werean improvement over the conventional mode in the twotasks. In fact, four out of four pilots awarded CHRÕs of 1during the 3 g tracking task. One important result from theevaluation was that the thrust vectoring only mode (DAG25) performed as well as the baseline enhanced mode,including several CHRÕs of 1.

Contact: Steve JacobsonF-15 ACTIVE Lead Flight Controls Engineer,NASA Dryden Flight Research Center, RC X-7423

Adaptive Performance Optimization for Transport AircraftSummaryAdaptive Performance Optimization (APO) is anapproach for improving transport aircraft performanceusing existing control surfaces. This approach exploitsexisting redundant control effector capability byautomatically reconfiguring control surface deflectionsto achieve a minimum drag trim condition.Implementation on a fly-by-wire transport can be assimple as a new Flight Management System softwareload; transports with mechanical systems requireadditional control system hardware modifications.

Objective: Design, develop, and demonstrate real-timedrag minimization techniques using symmetricdefections of wing control surfaces and, indirectly,horizontal stabilator. The algorithm will identify theminimum drag control surface configuration for thatcombination of aircraft configuration and flightcondition.

Justification: A 1-percent reduction in drag can saveoperators of long-range wide-body transports $120,000per aircraft per year in reduced fuel costs. For long-range aircraft at maximum takeoff weight (which isalso affected by both temperature and airport altitude),the benefit of a 1-percent drag reduction can increaserevenue by $4,000,000 per aircraft per year and withthe fuel tanks full, the benefit of a 1-percent dragreduction can increase revenue by $12,000,000 peraircraft per year. These revenue enhancements arehuge and benefits of more than 1-percent dragreduction would be proportionately larger.

Approach: The outboard ailerons of the modifiedL-1011 aircraft are commanded symmetrically tochange the lift distribution of the entire wing. Anexcitation command function is applied to thesesurfaces while Mach and altitude are being controlled.The resulting aircraft responses are analyzed on-boardthe aircraft to identify the minimum dragconfiguration for the entire aircraft. The out-boardailerons are then optimally repositioned.

Results: In the time history illustrated, the autopilotcontrolled altitude while the pilot controlled Machnumber during the optimization maneuver. Themaneuver was a two-sided raised-cosine, duringwhich the data was analyzed real-time. The thrust andstabilator indirectely indicate the effect the maneuveris having on trim conditions. At the completion of themaneuver, the symmetric aileron was commanded toits optimal (minimum drag) position. The analysisresults indicate a minimum drag condition at -3.4 deg.(trailing edge down) symmetric aileron deflection.

Contact: Glenn B. Gilyard, Principal InvestigatorNASA Dryden, RC, X-3724

Reference : AIAA 99-0831

Benefits¥ Reduced drag reduces fuel expense¥ Reduced drag provides for large increases in revenue for aircraft at maximum takeoff weight or with full fuel tanks¥ Optimization utilizes available control surfaces

Time history of a real-time drag minimization maneuver.

Variation of incremental drag with symmetricoutboard aileron deflection.

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deg

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Machnumber

980532

20 x 10– 4

15

10

5

0

– 5840– 4– 8

Symmetric aileron, deg

∆CD

980533

Best quadratic fit

Production Support Flight Control Computers

SummaryThe Production Support Flight ControlComputers (PSFCC) were developed inconjunction with the United States Navy andwill provide any of the F/A-18 aircraft at NASADryden with a flight control law or sytemsresearch capability. A research control lawprocessor imbedded in the flight controlcomputer avionics box provides a capability forpilot-selectable research control laws. Throughthe use of analog inputs there is capability tointerface hardware with the control system.

Objective: Develop a facility for fast andefficient flight test of advanced control laws,flight systems, and handling qualitiesexperiments, which can be performed on a timeavailable basis.

Justification: Design, implementation, and testof experimental control laws are currentlyextremely expensive, and connected directlywith specific flight programs under tightschedules. This facility would reduce the timeand cost associated with implementation, andallow researchers more time for investigationand discovery in control law and control systemresearch in the flight environment.

Status: Initial flight test of the PSFCC wasperformed in March/April of 1998 using theF/A-18 Systems Research Aircraft (SRA). Fourflights demonstrating the operationalengagement and disengagement of the systemwere conducted. Constraints imposed duringthe software verification and validation testingresulted in the aircraft being cleared to fly in arestricted envelope where failures within theresearch system would not result in exceedanceof the aircraft structural limits. Currently asoftware load is being developed and tested to aflight critical level which will not have as manyenvelope restrictions for flight test.

Contact: John F. Carter , PrincipalInvestigator NASA Dryden, RC, X-2025

PSFCC integration into baseline F/A - 18flight control computers

Input signalmanagement

F/A-18BaselineControl laws

Outputsignalselection

Research Processorand Control Laws

Memory Int erface

Basic F/A - 18Flight Cont rol System

F/ A - 18SensorInput sand1 5 5 3bus data

Surf aceCommands

The first experiment underdevelopment is to integrate the controlstick from the JAS-39 Gripen aircraftwith the standard F-18 control laws toassess any handling qualities issuesassociated with the stick. The softwarefor this experiment was developed andtested in house. Aircraft testing hasresulted in the requirement forhardware filtering of the control signals.The second experiment is to use thePSFCC to perform aerodynamicparameter identification maneuvers forthe Active Aeroelastic Wing program bycommanding single surface excitations.This will reduce risk for the AAWprogram by collecting requiredaerodynamic data for the the controlsystem design and establish theprocedures that will be required duringAAW flight operations.

ReferencesNASA TM-1997-206233

Separation Analysis of the X-38 Vehicle From the B-52 Carrier Aircraft

Contact:Timothy H. Cox or Martin Martinez-LavinNASA Dryden, RC, X2126

SummaryIn March of 1999 an X-38 research vehicle with anactive control system will begin test flights. Theresearch vehicle will be launched from a B-52aircraft at various flight conditions. Analysis of theseparation dynamics was conducted to investigatewhether the X-38 would re-contact with the B-52.Results indicate even in the event of a worse casecontrol system failure at the moment of launch theX-38 would clear the B-52.

ObjectiveTo perform analysis of separation dynamics and topresent this analysis visually to insure that the X-38will not re-contact the B-52 during launch.

JustificationThe NASA Johnson Space Center is currentlydeveloping the technology to design and build alifeboat for the space station. The program,designated as the X-38, is investigating the use of anautomated parafoil landing system. To successfullydevelop this landing technology two experimentalflight vehicles were designed to investigate thetransition of controlled aircraft-like flight throughparafoil deployment to pure, parafoil flight totouchdown. The experimental vehicles, one with anactive control system and one with fixed controlsurfaces, were designed to be carried under a B-52aircraft to a pre-determined flight condition andthen launched. One flight with fixed controlsurfaces has already been successfully conducted.Separation analysis for this flight showed no threatof B-52 re-contact. Flights with an active controlsystem are scheduled to begin in March of 1999.Recontact of the research vehicle with the B-52 inthe event of a control system failure, not covered inthe analysis for the first flight, must be addressed.

ApproachThe worse case scenario envisioned, full controlsurface hardovers in the worst direction (nose upfor pitch, rudder for lateral/directional) at theinstant of launch, was incorporated into an X-38simulation at the highest feasible dynamic pressure.Analysis for the longitudinal axis included themaximum amount of positive pitching momentuncertainty and a constrained lateral/directionalaxis. The lateral/directional axis was constrained toadd conservativism by not allowing the lift vector toroll off. Graphics models of the X-38 and the B-52were developed and driven by simulation data toillustrate the motion of the X-38 relative to the B-52.The graphic presentation uses a red line to illustratethe trajectory of the X-38 nose and blue lines to trackthe trajectory of the right and left fintips.

ResultsThe longitudinal analysis indicates that the X-38immediately pitches up to a high angle-of-attackafter launch and then translates past the tail. Evendespite the conservative nature of the assumptionsthe vehicle easily clears the B-5 2 wing andhorizontal tail.

The lateral/directional analysis for the rudderhardover indicates that soon after launch the X-38 rollsoff and begins to tumble. Although the fin tips swingin the direction of the engine nacelles, the visualgraphics show that plenty of margin exists. Althoughnot shown here, a rudder hardover in the oppositedirection likewise clears the B-52 fuselage.

Visual graphics illustrating the worse case longitudinalanalysis of the X-38 launch.

Visual graphics illustrating a rudder hardover duringan X-38 launch.

X-43A Guidance, Navigation and Controls Update

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 and during the engine test phase, the X-43A will be commanded to follow a nearly ballistic flightpath - a result of scramjet engine angle-of-attackrequirements. The engine test phase (which includespost-test vehicle parameter identification maneuvers) isconcluded by a recovery to a nominal descent trajectorymade possible by the autonomous controller resident inthe vehicleÕs flight control computer. Key to missionsuccess is the ability to accurately measure and controlangle-of-attack. Current plans include the integration ofa flush airdata system (FADS) into the flight controlsystem, which provides corrections to inertially-derivedalpha. Blending methods under investigation includecomplementary and Kalman filtering.

Critical to mission completion is the design andimplementation of inner-loop control laws and an outer-loop guidance algorithm. Both, of course, rely heavilyon adequate aerodynamic knowledge. The guidancealgorithm, however, has also proven sensitive toseparation condition uncertainties. Note that thedesign, analysis, integration and implementation offlight control software is the responsibility of a jointindustry and NASA team of which Dryden is only a part.

Objective:Dryden GNC efforts over the past year have focused onimproving the GNC algorithm performance in a detaileddesign phase. Of prime consideration has beenensuring compliance with scramjet engine experimentphase success criteria. Considerable effort has beenexpended in gathering and reducing FADS wind-tunneldata. The data is being utilized to create a dynamicFADS model for controller design and simulationimplementation. The same GNC efforts have centeredon improving control law robustness to aerodynamicand separation condition uncertainties.

Results:FADS simulation model development is nearingcompletion. The angle-of-attack estimation and sensorselection algorithms are being finalized for a projectreview to be conducted in CY99. A decision regardingappropriate estimation technique and fault detection willbe made in time for flight software delivery. Thelongitudinal guidance algorithmÕs performance hasbeen improved by accounting for uncertainties invehicle kinetic and potential energy immediatelyfollowing the separation event. Current CFD and windtunnel results predict a strong influence on lateral-directional aerodynamics with elevator position.

The lateral-directional controller has been redesigned toaccount for the aerodynamic dependency on elevatorposition.

Status/Plans:The control system has matured considerably over thepast year and is expected to be complete by the secondquarter of CY99. Final flight software delivery is expectedshortly after GNC design efforts have ceased. Hardware-in-the-loop and aircraft-in-the-loop tests are scheduledduring the summer of CY99 where vehicle systemvalidation tests are expected to increase confidence inhardware integration - including control laws.

X-43A Research Vehicle in free-flight

X-43A Representative Mission Trajectory

Contacts:Joe Pahle, NASA Dryden, RC, X-3185Mike Richard, NASA Dryden, RC, X-3543Mark Stephenson, NASA Dryden, RC, X-2583

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Mach 7 engine test

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

Summary:The F-15 Advanced Controls Technology forIntegrated Vehicles (ACTIVE) will be thetestbed for flight test of the Intelligent FlightControl System (IFCS). The IFCS is a flightcontrol concept which uses a Neural Network(NN) to determine 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 will be the culmination of thejoint Intelligent Flight Control AdvancedConcept Program teaming NASA Ames withBoeing Phantom Works. The goals of the IFCSAdvanced Concept Program (ACP) are to:

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 demonstrate these concepts in the F-15 ACTIVE Aircraft.

Objective:The 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.

Approach:Flight test of the IFCS ACP will be performed in3 stages. Phase I & II will be flown with systemoutputs provided to instrumentation only andwill not be used for aircraft control. Phase IIIwill use 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 will be flown utilizing the F-15 ACTIVE Research Flight Control System(RFCS). The RFCS allows the pilot to quicklyswitch from the experimental research flightmode back to the safe conventional mode ifrequired.

Status:Phase I development and flight test wascompleted in 1996. Phase II development iscomplete and has been ground tested on aSHARC Digital Signal Processor (DSP).Integration of the SHARC DSP into the F-15ACTIVE Vehicle Management SystemComputer (VMSC) is in progress. Flight test isscheduled for summer of Õ99. Phase IIIdevelopment is complete and the Verification,Validation and Flight Qualification of the PhaseI/III Control System is in progress. Flight test isscheduled for winter of Õ99.

Reference:Intelligent Flight Control: Advanced ConceptProgram, Annual Report, MDC 97M0004(NAS2-14181)Intelligent Flight Control: Advanced ConceptProgram, Annual Report, MDC 98P0026(NAS2-14181)

Contact:Michael P. Thomson, F-15 ACTIVE Lead FlightSystems EngineerCode RF, (805) 258-3097mike.thomson@dfrc.nasa.gov

Charles C. Jorgensen, Program ManagerNASA Ames, (650) 604-6725cjorgensen@mail.arc.nasa.gov

F-15 ACTIVE

ElectroMechanical Actuator (EMA) Validation Program

SummaryThe ElectroMechanical Actuator (EMA) is the third and finalresearch actuator developed and tested as part of the AirForce/Navy/NASA Electrically Powered ActuationDevelopment Program (EPAD). The EMA system replaces thestandard hydraulic aileron actuator on the left side of theF/A18B Systems Research Aircraft (SRA) with a 5HP, all-electric, dual-motor EMA actuator. The EMA actuator wasdesigned and built by MPC to be a one-to-one replacement forthe standard actuator, and its performance will be comparedwith concurrent data collected from the standard actuator on theright wing of the SRA. The Power Control and ManagementElectronics unit (PCME) for the actuator was built byLockheed Martin Control Systems. It combines both the lowpower control electronics and high power switching circuitryinto a box small enough to be mounted into the wing of theF-18. The ±135 VDC power for the actuator was supplied by aPower Conversion Unit (PCU), built by Dynamic Controls,Inc. Two Interface Boxes (IBOXs), also supplied by DynamicControls, allowed for installation of a research actuatorwithout modification of the standard F-18 Flight ControlComputers.

ObjectiveThe objective of the EMA experiment, and the EPAD programin general, is to establish the credibility of electric actuation asthe primary method of control for flight critical controlsurfaces on tactical aircraft and spacecraft.

JustificationAll-electric actuation of aircraft and spacecraft control surfaceshas long been seen as a way of increasing efficiency, loweringweight, greatly reducing maintenance, lowering cost, andreducing the amount of required support equipment andpersonnel for these vehicles. However, it is still considered“high risk” technology, and is usually abandoned in favor ofbetter understood actuators powered by central hydraulicsystems. The EPAD program will help answer many of thequestions with Power By Wire technology, and spur the use ofthis technology both in new aircraft and spacecraft designs, aswell as upgrades and retrofits of current models.

ApproachAfter an extensive Verification and Validation phase on the F–18 Iron Bird test bench, the EMA system was installed onthe aircraft for 25 hours of flight testing at representativeportions of the F-18 flight envelope.

Results• Flight Test completed successfully after 25 flight hours.• Data reduction is continuing.• EMA design was used to develop EMA actuators on X-38

Vehicle 132, currently undergoing flight test at Dryden.

Benefits• Lower A/C maintenance costs.• Lower A/C weight.• Greater efficiency.• Reduced requirements for ground support equipment.• Lower operating costs.• Increased Safety and Reliability

ContactStephen JensenPrincipal InvestigatorCode RF, (805) 258-3841NASA Dryden Flight Research Center

EMA ACTUATOR

55 56 57 58 59 60 61-30

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ght)

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Flight Data Comparison of Aileron Position,Left (EMA) vs. Right (Std.)

Flight Termination System Modernization

Summary: NASA Dryden operates experimental aircraft in controlledairspace over sparsely populated terrain and is responsiblefor public safety. To prevent unmanned aircraft from leavingsafe areas, Dryden requires the use of a Flight TerminationSystem (FTS), an independent system capable of endingflight. A typical FTS deploys a parachute and turns offaircraft engines upon activation.

The standardized airborne parts of a FTS include an antennasystem and receiver/decoders (as shown in Figure 1) whichterminate upon reception of a termination command from aground transmitter. The parts used by Dryden before 1998(for almost 20 years) weighed over 6.5 lbs. and occupiedover 147 cubic inches. They were expensive to buy/buildand did not fully comply with existing standards. Smallweight-sensitive unmanned aircraft motivated developmentof an improved FTS. The older FTS limited Dryden tooperation of only one unmanned aircraft at any given time.

Figure 1: Flight Termination System Architecture

Objective: Develop a FTS that is compact, light, inexpensive, andreliable and me