Aeroservoelastic Testing Of A Sidewall Mounted Free Flying ... · Aeroservoelastic Testing Of A Sidewall Mounted Free Flying Wind-Tunnel Model Robert C. Scott, Travis K. Vetter,yKevin

Aeroservoelastic Testing Of A SidewallMounted Free Flying Wind-Tunnel Model

Robert C. Scott,∗ Travis K. Vetter,† Kevin B. Penning,‡

David A. Coulson,§ and Jennifer Heeg¶

A team comprised of the Air Force Research Laboratory (AFRL), Northrop Grum-man, Lockheed Martin, and the NASA Langley Research Center conducted three aeroser-voelastic wind-tunnel tests in the Transonic Dynamics Tunnel to demonstrate activecontrol technologies relevant to large, flexible vehicles. In the first of these three tests,a semispan, aeroelastically scaled, wind-tunnel model of a flying wing SensorCraft vehi-cle was mounted to a force balance to demonstrate gust load alleviation. In the secondand third tests, the same wing was mated to a new, multi-degree-of-freedom, sidewallmount. This mount allowed the half-span model to translate vertically and pitch at thewing root, allowing better simulation of the full span vehicle’s rigid-body modes. GustLoad Alleviation (GLA) and Body Freedom Flutter (BFF) suppression were successfullydemonstrated. The rigid body degrees-of-freedom required that the model be flown inthe wind tunnel using an active control system. This risky mode of testing necessitatedthat a model arrestment system be integrated into the new mount. The safe and suc-cessful completion of these free-flying tests required the development and integration ofcustom hardware and software. This paper describes the many systems, software, andprocedures that were developed as part of this effort.

Nomenclature

AEI Aerodynamic Efficiency ImprovementAFRL Air Force Research LaboratoryAOS Airstream Oscillating SystemASE AeroservoelasticityBFF Body Freedom FlutterCG Center of GravityDAS Data Acquisition SystemDCS Digital Control SystemDOF Degree of FreedomdSpace1 DCS for Servo PID loops and WatchDogdSpace2 DCS for Trim, GLA, & BFF Suppressionδo Emergency controller command biasFS Flutter SuppressionGLA Gust Load AlleviationGUI Graphical User InterfaceHiLDA High Lift over Drag Active WingLAS Lift Augmentation SystemLE Leading EdgeLM Lockheed MartinLQG Linear Quadratic GaussianNG Northrop Grumman

∗Senior Aerospace Engineer, Aeroelasticity Branch, NASALangley Research Center, Hampton, VA, AIAA Associate Fel-low.†Senior Engineer, VMS & Flight Control, Northrop Grum-

man Corp., El Segundo, CA.‡Senior Staff Aeronautical Engineer, Advanced Development

Programs, Lockheed Martin, Fort Worth, TX.§Senior Research Engineer, Analytical Services & Materials,

Inc., Hampton, VA.¶Senior Aerospace Engineer, Aeroelasticity Branch, NASA

Langley Research Center, Hampton, VA, AIAA Senior Member.

PID Proportional Integral DerivativeΘ Pitch displacement, degΘ̇ Pitch rate, deg/sPZ Vertical position (PZ=0 at centerline), inRVDT Rotary Variable Differential TransducerSnub! Command to engage snubbing systemTDT Transonic Dynamic TunnelT Time, sTE Trailing EdgeVZ Vertical velocity, in/sWatchDog Software for triggering Snub! commandWD Abbreviation for WatchDogWoW Weight on Wheels

IntroductionIn the Fall of 2007, the Air Force Research Labo-

ratory (AFRL), Northrop Grumman, Lockheed Mar-tin, and the NASA Langley Research Center, success-fully completed the third of a series of three wind-tunnel tests of an aeroelastically scaled wind-tunnelmodel of a flying wing SensorCraft vehicle concept.The AFRL SensorCraft is a technology portfolio ofadvanced sensors, communications links, air vehiclecomponents and propulsion elements. AFRL is pur-suing these technology developments for future high-altitude, long-endurance, unmanned surveillance plat-forms. Two research programs intended to developtechnologies relevant to these large, flexible vehiclesare the High-Lift-over-Drag Active (HiLDA) Wing andthe Aerodynamic Efficiency Improvement (AEI) pro-grams.1 The goals of the HiLDA and AEI wind-tunnel

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American Institute of Aeronautics and Astronautics Paper 2008-7186

https://ntrs.nasa.gov/search.jsp?R=20080033678 2020-03-19T23:33:42+00:00Z

HiLDAT574: NASA

T579: Northorp Grumman

AEIT593: Northorp Grumman

T598: Northorp Grumman & Lockheed Martin

T6XX: Boeing

T6XX: Boeing

FY04 FY05 FY06 FY08FY07

Gust Sensor Flying Wing Joined Wing

FY03 FY09

Fig. 1 Completed and planned NASA LangleyTransonic Dynamics Tunnel wind-tunnel tests forthe HiLDA and AEI programs.

tests included the demonstration of Gust Load Allevia-tion (GLA), an enabling technology for a SensorCraftvehicle that will allow it to have reduced structuralweight thereby increasing endurance, range, and pay-load capacity.

The HiLDA and AEI wind-tunnel investigationswere conducted in the NASA Langley Transonic Dy-namics Tunnel (TDT). The completed tests are de-picted in figure 1 along with two planned AEI tests ofthe Boeing joined-wing2,3 SensorCraft vehicle conceptwhere a full span model will be considered. The firsttwo wind-tunnel tests were conducted as part of theHiLDA program. These included a short test of sev-eral candidate sensors for measuring gust flow angle.The HiLDA wing was tested September 2004 with gustload alleviation being demonstrated on a cantileveredmount using the model’s five active control surfaces.4–6

The HiLDA wing was retested in October 2006,where it was mated to a new, multi-degree-of-freedommount. This mount allowed the half-span model totranslate vertically and pitch at the wing root, allowingbetter simulation of the full span vehicle’s rigid-bodymodes. Following some modifications to the mountsystem, the wing was tested on the mount again in thesummer of 2007. Figure 2 depicts the wing and mountsystem in the final configuration used in Test 598. Theupstream flow angle sensor or gust sniffer is also shownin the image. During this test, GLA control systemswere demonstrated in the presence of gusts generatedby the TDT Airstream Oscillating System with peakwing bending loads being reduced by up to 60 percent,and BFF suppression control laws were demonstratedwith flutter dynamic pressure being increased up to 50percent.

The large size of the flying wing model along withthe rigid body degrees-of-freedom afforded by the newmount system, created many unique challenges to suc-cessfully and safely meet the AEI test objectives. Aremotely actuated snubbing system that returned the

Flow Oscillations (“Gusts”)

“Gust Sniffer”

SensorCraft Model

2 DOF Mount

Fig. 2 Illustration of AEI apparatus installed inTransonic Dynamics Tunnel during Test 598.

model to a fixed angle-of-attack was included in themount design. Determining how to effectively use thissystem was critical to the success of the test. Two dig-ital control systems were used, one to implement theflap servo control loops and another to implement thetrim, GLA, and BFF suppression control loops. Thedigital control systems included a safety monitoring(WatchDog) system for automatically triggering thesnubbing system. Along with these systems, a varietyof procedures had to be developed including those formodel launch and recovery.

The purpose of this paper is to provide an overviewof the AEI flying-wing wind-tunnel tests from an op-erational point of view with the focus on systems,procedures, and lessons learned. The paper includesdescriptions of the hardware including the wind tun-nel, the wing model, the mount system, and othersupporting systems developed for this program. Thesoftware developed for the digital control systems willbe described including the control law architecturesand the WatchDog system. The main part of thepaper will conclude with a description of the wind-tunnel testing procedures that were used. Finally, anappendix provides some additional analysis and test-ing details not covered in the main part of the paper.Throughout the paper, wind-tunnel data will be shownwhere applicable or necessary with an emphasis on themost recently completed test, Test 598.

For the purpose of clarity, a brief overview of themodel launch schemes, the flight control architectures,and the WatchDog system is warranted. Two modellaunch schemes were used, the TakeOff and the Releaselaunch schemes developed for use by Northrop Grum-man and Lockheed Martin, respectively. Throughoutthis paper, the TakeOff launch scheme will be asso-ciated with the NG control architecture where trimand suppression controllers are combined into a sin-gle control loop. The Release launch scheme will beassociated with the LM control architecture where

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Fig. 3 NASA Langley Transonic Dynamics Tunnel.

trim and suppression controllers are in separate controlloops. In addition, a separate set of WatchDog moni-tored parameters and associated limits are used witheach launch scheme. These topics will be discussedseparately in the paper.

Transonic Dynamics TunnelThe Langley Transonic Dynamics Tunnel (TDT),

depicted in figure 3, is a unique national facility ded-icated to identifying, understanding, and solving rele-vant aeroelastic and aeroservoelastic problems. TheTDT is a closed-circuit, continuous-flow, variable-pressure, wind tunnel with a 16-foot square test sectionwith cropped corners.7 The tunnel uses either air or aheavy gas as the test medium and can operate at totalpressures from near vacuum to atmospheric. It has aMach number range from near zero to 1.2 and is capa-ble of maximum Reynolds numbers of about 3 millionper foot in air and 10 million per foot in heavy gas.Until 1996, the TDT used dichlorodifluoromethane, R-12, as the heavy gas test medium; since then the TDThas used 1,1,1,2 tetrafluoroethane, R-134a,8,9 an envi-ronmentally acceptable alternative to R-12.

The TDT is specially configured for flutter test-ing, with excellent model visibility from the controlroom and a rapid tunnel shutdown capability for modelsafety. Testing in heavy gas has important advantagesover testing in air: improved model to full-scale simil-itude (which results in heavier, easier to build modelswith lower elastic mode frequencies), higher Reynoldsnumbers, and reduced tunnel power requirements.

Due to the need to perform GLA testing as part ofthe AEI program, the TDT Airstream Oscillation Sys-tem (AOS) was used frequently. The key features ofthe AOS system are shown in the illustration in fig-ure 4. These features include biplane arrangements ofvanes on either side of the entrance to the test section.Each pair of vanes is driven by a separate hydraulicmotor and a flywheel to hold constant vane frequency.

Fig. 4 Illustration of the TDT Air Stream Oscil-lating System (AOS).

FlowOscillation

Vanes

Flying WingSensorcraft

Model

“Gust Sniffer”

Fig. 5 Photo of the HiLDA model taken fromTransonic Dynamics Tunnel settling chamber.

While the two pairs of vanes can be run out of phase,this was not done in the AEI test. Vane frequency wasadjusted from the TDT control room manually or byusing a LabView system. The amplitude of the vanesis manually adjustable from 0◦to 12◦peak-to-peak. Formost of Test 598, the 12◦peak-to-peak setting was usedproviding approximately a ±1◦down wash in the vicin-ity of the model. For the last week of Test 598, thevane amplitude was reduced to 4◦peak-to-peak.

Wing and Mount SystemThe wing and mount system were designed and

fabricated by NextGen Aeronautics in Torrance, Cal-ifornia. NextGen Aeronautics was a subcontractorto the Northrop Grumman Corporation for the ini-tial development of the wing for Test 579, and later,for the wing modifications and mount system designand fabrication for Test 593. NextGen Aeronauticswas a subcontractor to the Lockheed Martin Companyfor the wing and mount modifications that precededTest 598. This section of the paper will provide a de-scription of the wing and mount system along with adiscussion of the modifications that took place over thelife of the program. A more detailed description of the

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Strain Gage - Bending

Strain Gage - Tension

Strain Gage - Bending (Y-Axis)

Actuator

RVDT

Accelerometers

Servoactuator/Manifold

Pitch Gyro

CTL3RVDT3

RVDT4

RVDT2

STM2

STI2STY

STI2

SBI1

Spar - Cross Section

SBI2

NZ4 SBM1SBM2

NZ5

STM1

RVDT1

CTL4

CTL5

CTL2

CTL1NZ6

NZ7RVDLE

NX2

NZ1

NZ2 NZ3NX1

Hydraulic Lines

Wires

Fig. 8 Wing Instrumentation used for the AEITDT tests (593 and 598). String potentiometersused for measuring pitch angle (Θ), vertical po-sition (PZ), vertical rate (VZ), and moving massposition are not shown.

HiLDA wing can be found in references 4 and 5, anddetails associated with the wing and mount system canbe found in references 10 and 11.

The wing and mount system configuration used inTest 593 is shown in figure 6, and the configurationused in Test 598 is shown in figure 7. These figuresshould be referred to when reading the sections thatfollow.

Wing

The wing is a 12%-length scale, semispan model ofa SensorCraft design concept originally designed foruse in the HiLDA program. The outboard part ofthe wing is a spar-pod design where the scaled stiff-nesses (EI and GJ distributions) are designed into theflanged aluminum spar, and the aerodynamic shapeis provided by discrete fairings, or pods, mounted tothe spar. The original set of pods was made fromglass-filled nylon-12 (Duraform GF) manufactured us-ing a stereolithography system. The inboard wing wasintended to act as a rigid member and consisted ofaluminum spars and internal ribs with fiberglass up-per and lower skins. The closeout ribs for the inboardwing were stainless steel to provide the structural at-tachments for the 5-DOF balance at the root and forthe outboard, flexible spar at the wing break. Leadweights could be attached to the spar to simulate fuelweight at the take off configuration.

The wing has four evenly spaced trailing edge con-trol surfaces and one leading edge control surface onthe outboard, flexible portion of the wing. The five

control surfaces were driven by vane-type hydraulicactuators with position measured by Rotary VariableDifferential Transducers (RVDTs). Two separate hy-draulic systems were included in the model so thatthe trailing edge flaps could be operated at 1,000 psi,while the more heavily loaded leading edge flap couldbe operated at up to 2,000 psi. Other instrumenta-tion included strain gages, accelerometers, a pitch-rategyro, and a gust vane placed in front of the model to al-low lead gust information to be fed into the controller.Figure 5 shows a photo of the HiLDA wing installedin the TDT from the settling chamber.

In preparation for the first AEI test, TDT Test593, several modifications were made to the HiLDAwing. As rigid body degrees-of-freedom were being in-troduced as a result of the new model mount system,overall mass and center of gravity became importantconsiderations. To adjust the model CG, the LE capsof the inboard part of the wing were modified so thatsix blocks of lead could be installed to provide up to30 lbs of forward ballast. In addition, a moving masssystem was included in the rigid root section that pro-vided remote adjustment of the model CG. The deviceconsisted of a compact stepper motor coupled to a ballscrew actuator driving a 32 lb ballast weight. The bal-last assembly was restrained by two linear slides witha string potentiometer for position measurement. Tosimplify the model’s hydraulic requirements, a largerleading edge actuator was built for this test so thatit could operate at the same hydraulic pressure as thetrailing edge flaps.

Modifications to the HiLDA wing’s instrumentationsuite included changes to the accelerometer locationsto better measure the free-free mode shapes and theaddition of a pitch rate gyro near the wing root. Theinstrumentation layout used in both AEI tests is shownin figure 8. Not shown in this figure are leading edgestagnation point sensor arrays developed by Tao Sys-tems that are described in reference 12. Finally, thepods and control surfaces were redesigned to simplifyinstallation and removal, and they were fabricated us-ing a different material (Watershed 11120) to avoidconcerns over moisture absorption. Unfortunately, the11120 material has a low glass transition temperature,and the pods warped due to elevated temperatures intransit to the TDT.

Changes to the wing in preparation for TDT Test598 were limited largely to maintenance items; how-ever, two important modifications were made. First,the pod material was changed back to Duraform toavoid heat related warping. Duraform GF was used forthe new pods, and Duraform AF was used for the newcontrol surfaces. To minimize moisture absorption, allpod surfaces were sealed using primer. The secondmodification was deemed necessary when, during in-

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a) Mount system components.

b) Original fairing assembly with large dynamic plate. Totalvertical travel includes 14.2 inches of free travel plus the 3 inchstrokes of the upper and lower shocks.

c) Back side of pivot carriage showing snubbing mechanism andumbilical.

Fig. 6 Original, 3 degree-of-freedom mount configuration used in Test 593.

spection, it was determined that the inboard wing hadbeen overloaded during Test 593. Wind-tunnel datafrom this event will be shown later in the paper. Asshown in 7, a steel strap was added to the upper andlower surfaces of the inboard wing to strengthen the

center spar by providing an additional load path be-tween the two stainless steel ribs.

Appendix A describes the performance and chal-lenges associated with using and maintaining the wingflap actuators during Test 598.

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a) Revised mount system.

b) Revised fairing assembly with smaller windowand dynamic plate. Total vertical travel includes23.4 inches of free travel plus the 3 inch strokes ofthe upper and lower shocks.

c) Installed mount system showing umbilical ser-vice loop.

Fig. 7 Revised, 2 degree-of-freedom mount configuration used in Test 598.

Mount System

The sidewall multi-degree-of-freedom mount systemevolved considerably over the course of the two TDTtests in which it was used. Initially for Test 593, themount was configured to provide 3 degrees-of-freedom,pitch, plunge, and fore-aft translation. This configu-ration is pictured in figure 6. Based on the opera-tional experience gained in Test 593, several significantchanges were made to the mount system in prepara-tion for Test 598. These changes included the removalof the fore-aft DOF and a complete redesign of the liftaugmentation system. The revised system is shown infigure 7. The key elements of the mount system alongwith the various design changes will be described next.

The pivot carriage assembly was the heart of themount system. It consisted of the pivot bearing andthe hydraulic pitch snubbing mechanism. The snubberconsisted of two linear hydraulic actuators that moveda sliding plate (guillotine) that clamped against al-ternate ends of the half-circle, free-to-pitch structure

that attaches to the wing. The back side of the pivotcarriage assembly showing the pitch snub componentsand the hydraulic/instrumentation umbilical exitingthe model through the center of the bearing assem-bly is shown in figure 6. A string potentiometer wasused to measure model pitch angle during both tests,and another string potentiometer was added prior toTest 598 to provide a direct measurement of the guillo-tine position. For the unsnubbed condition, the totalrotation was controlled by adjusting the retractiondistance of the guillotine by adjusting the hydrauliccylinder mount position. For the snubbed condition,replaceable pads on the free-to-move structure allowedcontrol of the clamped angle pitch angle. Test 593 hada range of motion of -12◦to +16◦with a clamped angleof +2◦, and test 598 had a -6◦to +8◦range of motionwith a clamped angle of +1◦. The range of motion wasrestricted to reduce the risk of damaging the model.Appendix B describes experiments used to assess andimprove the performance of the snubbing system.

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The rail adaptor and vertical rail assemblies pro-vided the translational DOFs via THK 20 mm widthrails and guide blocks. Mil Spec bungee cords, cables,and pulleys were used to restrain for fore-aft DOF, and40” stroke string potentiometers were used to measurethe fore-aft and vertical position of the pivot carriageassembly. To minimize impact loads, the vertical railassembly included dashpot shock absorbers at the topand bottom. During Test 593, it was determined thatthe fore-aft DOF was heavily damped and as a result,no significant fore-aft motion was noted during testing.In preparation for Test 598, the assembly that pro-vided the fore-aft DOF, the rail adaptor, was removed.The pivot carriage assembly was rotated 90◦and at-tached directly to the vertical rails as shown in figure 7.

A fairing assembly enclosed the mount system. Thefairing consisted of fixed exterior fiberglass panels witha cutout or window that allowed for a range of travelof the wing attachment. The window was covered witha dynamic plate that attached directly to and movedwith the pivot carriage assembly. The fairing assemblyused in Test 593 is shown in figure 6. The eliminationof the fore-aft DOF for Test 598 allowed the widthof both the window and dynamic plate to be reducedfrom 23” to 11” as shown in figure 7. As a result,the fixed fairing could be better supported, eliminat-ing contacts between the fixed fairing and the wingroot and the fixed fairing and dynamic plate. A newsmaller, two-piece dynamic plate design was used inTest 598. The allowed the dynamic plate to be re-moved for access mount system components withoutthe need for removing the wing.

A lift augmentation system (LAS) was required tocompensate for the mass of the carriage system andexcess inboard wing weight. The first iteration of theLAS consisted of 8 constant force springs that providedapproximately 40 lbs of force each for a total lift aug-mentation of 320 lbs. The springs could be attachedto the rail adaptor in groups of two via a lanyard andcable assembly. This first system suffered from highfriction associated with the spring tapes twisting andrubbing against the spool frame assembly. The re-designed system used in Test 598 again used 8 constantforce springs, but due to the weight savings associatedwith the elimination of the fore-aft DOF and the re-moval of the rail adaptor assembly, springs providing20 lbs of force each could be used. These springs at-tached to the pivot carriage assembly via pinned balljoint with load cells in series so that LAS force could bemonitored during the test. Other design improvementsincluded spring spool assemblies with an improved setof bearings and a LAS frame design that minimizedpotential spring tape contact points.

The original vertical rail assembly consisted of twoindependent components that bolted directly to the

TDT wall plates as shown in figure 6. As a result ofirregularities in the wind-tunnel wall, shimming wasnecessary to avoid binding the plunge DOF. Also,the hydraulic/instrumentation umbilical had to exitstraight out of the assembly as shown in figure 6.Because there was limited space between the pivot car-riage assembly and the wind-tunnel wall, free verticaltravel was restricted to ±7.1 inches of the designed±12 inches to avoid cutting the umbilical on the open-ing in the wind-tunnel sidewall. The free travel doesnot include the 3 inch strokes of the upper and lowershocks, so total vertical travel is 6 inches greater thanthe free travel. The redesigned mount shown in figure 7included a new standoff support frame that simpli-fied mount installation and provided an additional 4inches of clearance from the wind-tunnel wall. Thesechanges allowed the hydraulic/instrumentation umbil-ical to form a service loop inside the fairing, under thepivot carriage assembly, as shown in figure 7. For Test598 ±11.7 inches of vertical travel was achieved in thisconfiguration.

Appendix C describes the experimental methodsused to estimate the friction in the vertical rails andto monitor the health of the LAS.

Test Configurations

In Test 598, the model was tested in two configura-tions: heavy and light. The heavy configuration had 5of 6 inboard leading edge masses, all wing fuel weights,and all 8 LAS springs installed. The light configura-tion had no wing fuel masses, 2 of 6 inboard leadingedge masses, and 4 LAS springs installed. The totalfree flying weight of the heavy and light models was442 lb and 356 lb, respectively.

The boundary conditions and hence, the dynamiccharacteristics of the wing were altered considerably bythe state of the mount system. For instance, when thesnubbing system was engaged and the pivot carriagewas held on the lower stop via gravity, the wing wasessentially cantilevered at the root. The pitch DOFcould be freed by releasing the snubber, and the plungeDOF was only free when the model was in trimmedflight. Thus, the mount system provided four distinctboundary conditions that had to be considered. Thevarious boundary conditions, mount states, and firstmode frequency are shown table 1. For the heavyconfiguration, the frequencies were determined by aground vibration test performed in the TDT modelpreparation area, and for the light configuration, fre-quencies were determined via NASTRAN analysis.11

Signal Routing, Processing, andControl

A high level schematic of the signal routing arrange-ment used in Tests 593 and 598 is shown in figure 9.

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Table 1 Model Boundary Condition and the Fre-quency of the 1st Bending Mode.

Bndry Cond Mount State 1st Mode, HzPitch Plunge Pitch Plunge Heavy LightFixed Fixed Snub On Stop 1.9 2.1Fixed Free Snub Trim 2.0 2.3Free Fixed Trim On Stop 3.0 3.3Free Free Trim Trim 3.5 4.1

All input and output signals were routed from theirsource to the TDT control room via a built-in wiringsystem. Most instrumentation was routed to a PacificInstruments series 6000 chassis that provided instru-ment power (5 or 10 VDC), signal amplification, andanti-alias filtering. The RVDTs and Rate Gyros re-quired ±15 VDC power so a custom power supplywas fabricated with the signals subsequently routedthrough the Pacific Instruments chassis for amplifica-tion and anti-alias filtering. The anti-aliasing filterswere set to 400 Hz for the RVDTs as they were routedonly to a digital control system running at a 1,000Hz frame rate (dSpace1), and all other signals werefiltered at 100 Hz to be compatible with the other dig-ital control system running at a 200 Hz frame rate(dSpace2). Servovalve signals were routed to a Moogvoltage-to-current converter and back to the model asshown in the figure. Model signals were “Teed” androuted to a strip chart for monitoring and to the TDTData Acquisition System (DAS) for recording. TheTDT DAS was set to record data at 500 Hz with itsown anti-aliasing filters set to 200 Hz.

The snubber control system, dSpace1 and dSpace2internal block diagrams, and the external signals com-municating between these systems are depicted in fig-ure 10. All signals external to the dSpace blocksare analog, with dSpace input signals being convertedfrom volts to engineering units prior to processing andoutput signals being converted from engineering unitsback to volts. For signals intended to depict discretesystem states, programming logic was used to decodethe meaning of the signal.

The hydraulic actuators in the snubbing systemwere operated by a solenoid valve that was controlledusing the snubber control system. The heart of thesnubber control system was a latching circuit that canbe tripped by a manual chicken switch or a Snub!command signal from dSpace1. The status of thelatching circuit (snubbed or unsnubbed) was commu-nicated back to the dSpace systems by the SnubStatsignal shown in figure 10. A detailed description ofthe snubber control system and associated hydrauliccomponents is described in a separate section, below.

The trim, GLA, and/or BFF suppression control

laws were implemented on dSpace2. As shown in fig-ure 10, dSpace2 had two internal control blocks and aset of externally generated flap commands that couldbe summed and output as analog flap command signalsto dSpace1. Throughout most of Test 598, only theupper dSpace2 control block was used with two ver-sions being implemented, one developed by NorthropGrumman and the other developed by Lockheed Mar-tin. This upper control block contained GUI interfaceand some programming logic for controlling or initiat-ing certain events like resetting the system or initiatinga launch. Details of the NG and LM flight controlarchitectures including their associated programminglogic will be described later. The FlightMode state iscommunicated to the WatchDog system as shown.

The servo-control loops, the WatchDog system, andthe Weight on Wheels (WoW) switch estimator wereimplemented on dSpace1. The servo-loops were in-dependent PID control loops equipped with outputsaturation blocks to prevent overdriving the actuators.The WoW switch was a hysteresis block intended toprovide debounce or noise tolerance. The WatchDogsystem monitored the model signals and when a faultwas detected, would issue a Snub! command and trans-fer flap control to the emergency control laws via theswitch shown in 10.

Details of the dSpace hardware and software devel-opment environment will be described next, followedby a description of the hydraulics and snubber controlsystem.

DCS Hardware and Development Environment

Each dSPACE DCS consists of a rack containing ahost computer, a target system, a rack mounted key-board and monitor, BNC patch panels for IO, andan uninterruptible power source. The heart of theDCS is the target system that includes a dSPACEDS1006 control processor board utilizing a 2.6GHzAMD Opteron processor connected to three dSPACEDS2002 multi-channel A/D boards and one dSPACEDS2103 multi-channel D/A converter board. The A/Dboards each have 32 channels utilizing 16 bit quantiza-tion with an input range of 10 Volts. The D/A boardcontains 32 channels of 14 quantization bits designedfor 10 Volts and a settling time of 10 µsec.

The controller software is developed within the Mat-Lab Simulink environment, then compiled and down-loaded to the target processor via the dSPACE andMatLab Real-Time Interface. An integral componentof the dSPACE tools is the ControlDesk application.ControlDesk provides the user interface to the targetprocessor for the development and implementation ofthe visual control and indication medium. The targetprocessor is the embedded controller and the host pro-vides the user interface, visual display, and controls all

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HiLDA Wing and AEI Mount

East Wall To Control Room Patch Panel

Gust Sniffer

Test Section

Control Room

Gyros, Strain GaugesAccelerometers,String Potentiometers, &Gust Sniffer

Data Room

Servovalue GustSniffer

Strain Gauges,Accelerometers,

& String-Potentiometers

Rate Gyros

Gyro

Servovalue Commands

RVDT

RVDT

Flap Position Commands

RVDT

Control Room To Data Acquisition Patch Panel

Moog Voltage toCurrent Converter

Custom +/-15 Volt Power Supplies

Pacific InstrumentsSignal Conditioner/Filter

dSpace1Servo LoopWatchDog

Digital Strip Chart dSpace2Trim

GLA / FSS

DAS

Fig. 9 Signal routing used in Test 593 and 598.

δposition

δCmd.

δCmd.

ModelSignals

WoWSwitch

LatchingCircuit

EmergencyControl Law

GLAorFS

External Control Inputs

Manual Chicken Switch

FlightMode

Servo ControlPID

WatchDog

SnubLogic

V Divider

Snubber Control System

To Solenoid Valve:24V=UnsnubbedOV=Snubbed

SnubStat:Hi=UnsnubbedLow=Snubbed

To Model:Servovalves

dSpace1dSpace2

Reset Switch Override Switch

Gnd

Snub!

Trimor

Trim+GLA/FS

See NG or LMBlock Diagrams

Fig. 10 dSpace and snubber control block diagram.

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Fig. 11 Snub control chassis containing the snubcontrol system with manual Snub! switch and hy-draulic pressure displays.

communications between the processors.

Hydraulics and Snubber Control System

Hydraulic control of the pitch snub mechanism wasaccomplished using a custom snub control systemhoused in a chassis located in the TDT control roomand a hydraulic manifold assembly located in the TDTplenum near the wind-tunnel model and mount. Hy-draulic fluid, at 2,000 psi, was delivered to the hy-draulic manifold assembly from the TDT 30 GPMpump. The manifold assembly consisted of an ad-justable pressure switch, a Parker Hannifin solenoidvalve, and two regulators for supplying up to twolower pressures to the wind-tunnel model. Hydraulicpressure transducers were included to monitor the sup-ply/snub pressure as well as the two regulated pres-sures. The solenoid valve was normally open andrequired 24 VDC power to remove pressure from thesnubbing system thereby releasing the model.

The snub control chassis served two purposes, moni-toring of the hydraulic systems and enclosing the snubcontrol system. The chassis housed three hydraulicpressure displays showing supply/snub pressure andup to two lower, regulated model pressures. An amberlight and an audible alarm were wired to the pres-sure switch to provide warning of reduced hydraulicpressure. The snub control system consisted of twomechanically latching relays that were used to send24 VDC to the hydraulic solenoid valve. The relayshad their contacts wired in series so that either re-lay could cut the 24 VDC power to the hydraulicsolenoid and independently snub the model. One relaywas switched directly by the normally open hand-heldchicken switch, and the second relay was switched bya voltage signal from an external source.

The snub control chassis and a handheld chickenswitch are shown in figure 11. Two amber colored

lights on the front of the snub control box indicatedwhich source initiated the Snub! command, chickenswitch or external source (WatchDog system). Reset-ting of the relays and lights was accomplished with amomentary toggle switch (reset switch). There was anindependent toggle (override) switch to hold a snubcondition. This switch was used to keep the model ina snubbed condition independent of the status of therelays allowing systems to be reset with out inadver-tently cycling the snub system.

WatchDog SystemDue to the high risk associated with aeroelastic

wind-tunnel testing, a variety of manual and auto-mated safety systems have been used in the TDT. Themanually operated tunnel bypass valves are generallythe first line of defense for flutter testing as they canrapidly reduce test section dynamic pressure and Machnumber. Depending on the type of model and mountsystem employed, model stabilization or arrestmentmechanisms have also been considered. Previouslyemployed model stabilization devices have included de-coupler pylons that change model dynamics to a morebenign configuration and model arrestment devices in-cluding pneumatic snubber cables for cable mountedmodels. Manual engagement of these devices has attimes been supplemented by automated systems. Suchsystems have previously been employed only for nonflying models with the tunnel bypass valves and/ordecoupler devices being triggered based on thresholdexceedences.13

The range of rigid body motion and the potential forhigh speed impact afforded by the new mount systemmade the AEI flying wing tests among the riskier testsconducted in the TDT. The snubber mechanism wasintended to reduce risk to the model and facility, andcould be triggered manually. However, human reac-tion times for ‘simple’ tasks like the sheep dash gamein reference 14 are only around 0.22 seconds, and moreadvanced tasks, like dodging a baseball, requires a full0.4 seconds.15,16 In the tunnel testing environment,the test engineer must in essence perform system IDin real-time to determine whether a behavior is a be-nign oscillation or a potentially fatal divergence, mixin a little hope, and the reaction time can be signifi-cant. This point is driven home by the worst overloadcase from Test 593 where the automated system wasnot used. Figure 12 shows time histories for verticalposition, pitch angle, and inboard bending momentacquired during a failed model launch attempt. Forthis data point, the peak bending load significantly ex-ceeded the design load of 15,000 in-lbs preceding themanual Snub! command. Clearly, a robust automatedsystem was necessary, and the WatchDog system wasdeveloped for this purpose.

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25 26 27 28 29 30

20

15

10

5

0

-5

-10

-15

1.0

0.5

0

-0.5

-1.0

-1.5

-2.0

-2.5

de

g o

r in

x 104

Ro

ot B

en

din

g M

om

en

t, in

−lb

s

Θ, degPZ, in

Snub!

Time, s

Fig. 12 Worst case bending overload conditionfrom Test 593, Point 1833. The bending moment ismeasured from strain gauge SBI1 shown in figure 8corrected for outboard wing weight.

The WatchDog system was developed to keep themodel and mount system from exceeding structuralsafety limits by monitoring signals from the wing andmount system. Difficulty in updating limits and alarge number of false positives limited WatchDog usein Test 593. By Test 598, the need for an improvedWatchDog system was well established and the de-ficiencies of the previous implementation were cor-rected. The system was also expanded to satisfyingthe differing needs of both Northrop Grumman andLockheed Martin.

The WatchDog was implemented on dSpace1 asshown in figure 10. The key features of the Watch-Dog system are the emergency control law and thesnub logic. The snub logic monitors the model sig-nals and issues a Snub! command when a fault isdetected. Issuing the Snub! command engages thepitch snubber and switches model flap control fromdSpace2 to the emergency control law as shown inthe figure. The subsections that follow will describethe emergency controller, the snub logic, and some en-hancements and unique features that were developedfor use by Northrop Grumman and Lockheed Martinfor Test 598.

Emergency Controller

When the model is snubbed, the aerodynamic liftbecomes a function only of the dynamic pressure andcontrol surface position. In the snubbed condition,the flaps can be used as lift only devices, as the pitch-ing moment generated is counteracted by the snubbingmechanism. As a result, a simple emergency controllercould be developed to arrest the vertical motion andsafely land the model after the Snub! command hadbeen issued.

FlightMode

SelectNG or LM

ModelSignals

LimitCheck

Latch

Debounce

NG LimitsWD Enable

ResetValue (=0?)

NG Signals

LM Signals

LM Limits

Snub!

GUIDisplay

Fig. 13 WatchDog system snub logic block dia-gram.

The emergency controller flew the model in thesnubbed condition using vertical position error andvertical velocity feedback. The controller gains weredivided by the measured dynamic pressure to producea controller that worked over the dynamic pressurerange of interest. The controller also had a flap biasterm (δo) consisting of the expected control surface po-sitions that would provide total wing lift equal to netwing weight, actual wing and pivot carriage weightless LAS force. These bias values were also a functionof dynamic pressure, and were initially determined bysimulation and then updated as part of the testing pro-cess. This testing process will be described later in thepaper.

WatchDog Snub Logic

A block diagram of the WatchDog snub logic isshown in figure 13. The system’s primary functionwas to monitor the 32 model signals and issue theSnub! command when parameter limits had been ex-ceeded. The WatchDog could be enabled or disabledby the user from the dSpace1 GUI. While rarely dis-abled during Test 598, the feature was needed to facil-itate testing under unique circumstances.

From the dSpace1 GUI, the operator can select ei-ther the NG or LM limits and signals to be used withthe NG or LM control law architectures, respectively.Both NG and LM signal blocks included the basic setof 32 model signals along with some calculated valuescalled combined parameters. The NG and LM limitsets each included a discrete set of upper and lower pa-rameter limits for each FlightMode. The FlightModewas used to select between the sets of limits, so thata different set could be used for reset, takeoff, flight,or GLA testing. The exact definition of the Flight-Mode parameter was depended on which flight controlarchitecture was being used, NG or LM. In ether case,when the FlightMode associated with model launchand climb out was used, the lower model position lim-its needed to be ignored until a the model cleared acertain height (PZ value). NG accomplished this byhaving a separate FlightMode and WatchDog limit setfor Take off and climb out. LM did not use a separateFlightMode and WatchDog limit set, but instead hadsome coded logic built into the LM Signals block thatignored the lower position parameters until a certainPZ value had been reached. A more detailed descrip-

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tion of these processes will be provided in the flightcontrol schemes section of the paper.

The selected signals and FlightMode dependent lim-its were compared in the Limit Check block. Detectedfaults were passed to the Debounce block where 3consecutive frames of any particular signal fault wererequired to issue the WatchDog Snub! command. Thisprevented a single frame noise event from issuing afalse alarm. A latch was used to quickly diagnose thecause of WatchDog commanded Snub! events. Theparameter that caused the trip would be held anddisplayed to the user via the GUI on dSpace1. TheWatchDog latch was reset by setting the FlightModemode to zero.

Preliminary modeling and review of past tests al-lowed for initial values of the WatchDog limits to beset. As the wind-tunnel test progressed, some limitswere expanded based on operational experience. Forexample, as lower frequency gusts were tested, ver-tical velocity and pitch rate became large exceedinginitial WatchDog limits, but since model response re-mained stable and bounded, the WatchDog limits wereexpanded to permit testing. During testing, the limitvalues could be modified directly in the real time pro-cessor memory from the DSpace1 user interface. Thesemodifications were then recorded back into MatLabscript files for future use.

NG - Combined Parameters

Based on the WatchDog system deficiencies iden-tified during Test 593, Northrop Grumman proposedthat combined parameters be considered in additionto nominal set of 32 model signals. For exam-ple, if the model had negative vertical position (toolow) and positive vertical velocity (moving up), thatwould be acceptable. Alternatively, the combinationof negative position and negative velocity would because for concern. Thus, multiplying vertical posi-tion (PZ) and vertical velocity(VZ) produced a pa-rameter that was acceptable when negative and givena safe upper limit and monitored by the WatchDog(∞ <PZ×VZ<limit). An additional parameter ofPZ×VZ2×sign(VZ) was also used and was a goodindicator of dangerous conditions. These combinedparameters increased model safety and provided rela-tively few false alarms. The NG combined parameterswere implemented within the NG Signal block in fig-ure 13. The NG combined parameter limits were setinitially by simulation and then modified throughoutthe test, as deemed necessary by the test team.

LM - Combined Parameters

A different set of combined parameters was devel-oped by Lockheed Martin based on an estimate ofthe vertical location up and down where the emer-gency controller can bring the model to a stop once

dSpaceGeneratedExcitation

Model Dynamics

ServoloopsWatchDog

Model

+

dSpace2: Upper Control Block

Sensors

SnubStat

Full ControllerΘ-Controller: Ramp or Hold Θ

orZ-Controller: Ramp or Hold PZ

WoW

δCmd.

dSpace1

Model + Controller

GUI Input:SFMode

FlightMode

Stateflow Logic Block

Fig. 14 Control law and FlightMode logic devel-oped by Northrop Grumman.

the snubber had been engaged. These two combinedparameters were generated by a feedforward estimatorin series with a neural network implemented within theLM Signal block in figure 13. Appendix D provides adetailed description of the simulation model, neuralnetwork, and Test 593 data analysis used to developand validate the LM combined parameters.

Flight Control SchemesTwo schemes for achieving trimmed flight were de-

veloped and demonstrated. They were the Take-Off and the Release schemes developed by NorthropGrumman and Lockheed Martin, respectively, withthe assistance of NASA Langley Research Center.The TakeOff scheme simulates a take-off roll, accel-eration, rotation, and lift-off. Alternatively, the Re-lease scheme kept the model snubbed until flight speedwas achieved in the tunnel, then it was unsnubbedand lifted off. This section will describe each launchscheme and its associated control law architecture.Example wind-tunnel data will be provided where ap-propriate.

Takeoff Launch Scheme

The TakeOff launch scheme approximates an air-craft takeoff process beginning with a high speedground roll, followed by rotation, lift off, climb, andfinally, trimmed flight. The control law architectureassociated with the TakeOff launch scheme is shownin figure 14. As shown in the figure, the TakeOfflaunch scheme makes use of the MathWorks State-flow package. Stateflow extends Simulink with a de-sign environment for developing state machines andflow charts. The NG implementation of Stateflow, al-lows most TakeOff launch scheme actions to be eventdriven. The Stateflow Logic Block is controlled by theStateflow Mode (SFMode) parameter via the dSpace2GUI as indicated

The TakeOff launch scheme requires two separatecontrollers. A Θ-controller is used during the pitch-free condition when the model is on the lower vertical

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20

10

0

-10

-20

in o

r deg

60

50

40

30

20

Dyn

amic

Pre

ssur

e, p

sf

0 10 20 30 40 50 60 70 80 90Time, s

SFMode: 0 Taxi

WoW On WoW Off

Snubbed

Fly

Free

Free

Fly

Theta Controller: Hold Rotate

Unsnubbed

Z Controller: Climb

Pitch: Fixed

Plunge: Fixed

Free

Fixed

Hold

User Commanded Event Driven

PZ, inΘ, degTE3POS, deg

FltMode/WD: 0

Cntrl: Open Loop

Taxi T-O/Climb

Fig. 15 Data acquired while using TakeOff launchscheme to achieve trimmed flight. Test 598, Points1274, 1275, and 1276.

stop, and a Z-controller is used when the model takesoff and transitions to the free-free boundary condition.The control laws implemented in the figure were de-veloped using the LQG method and incorporated GLAand trim control into a single controller. This allowedstudies to be performed that directly traded short pe-riod performance with GLA performance. The designof these controllers and their initialization methods arediscussed in more detail in reference 17.

Data acquired during implementation of the Take-Off launch scheme are shown in figure 15. Here, timehistories of PZ, Θ, flap 3 position, and dynamic pres-sure are plotted in the upper part of the figure, andthe timing of boundary condition changes, Stateflowmodes, and other system states are identified in thelower portion of the figure. User and event driven ac-tions are identified. This figure should be consultedwhen reading the following TakeOff scheme steps listedbelow:

• T=0s: Start with model snubbed, dynamic pres-sure set to 20 psf, and SFMode set to 0. Stateflowsets FlightMode to 0 telling the WatchDog to usethe reset limits.

• T=13s: Operator sets SFMode to Taxi. Con-troller is ready and waiting to detect theunsnubbing of the model via the SnubStat signalfrom the snubber control system.

• T=15s: Operator unsnubs the model using thereset switch on snubber control system chassis.Stateflow senses the unsnub condition via theSnubStat signal and initiates the Θ-controller.It also changes FlightMode to Taxi telling the

WatchDog system to use the Taxi limits

• T=24 to 45s: Dynamic pressure is increased tothe launch value of 50 psf

• T=50s: Operator sets SFMode to Fly. Stateflowthen starts the Θ set point ramp and changesFlightMode to T-O/Climb telling the WatchDogsystem to use the T-O/Climb limits.

• T=58s: The model lifts off the lower stop andpasses through the WoW switch causing State-flow to transition from the Θ-controller to theZ-controller. PZ set point for the Z-controlleris ramped up toward tunnel centerline. Also,as WoW switch is cleared, the model boundaryconditions transition to free-free.

• T=83s: Stateflow senses that PZ has reached avalue of -1 inch and changes the PZ set point tothe tunnel centerline value of zero. Stateflow alsochanges FlightMode to Fly telling the WatchDogsystem to use the Fly limits.

• T¿90s: Dynamic pressure increased, as necessary,to the desired test condition. Operator can setSFMode to Test. Stateflow then changes Flight-Mode to Test which tells the WatchDog systemto use the Test limits. When testing is complete,snubber chicken switch is used to manually engagethe snubbing system and land the model using theemergency control law.

This method was shown to be successful and leda better understanding of the model in early test-ing. However, the large dynamic pressure changes thatwere required for each launch made it very time con-suming. In addition, the pitch-free boundary conditionis not very stable and is significantly more difficult tocontrol than the free-free boundary condition. As aresult, engineering effort had to be used to developstable Θ controllers. Also, the transition from the Θ-controller to the Z-controller must have a small enoughtransient not to overly perturb the model. Thus, the Z-controller controller bandwidth and performance wasbeing driven by the need to mitigate the takeoff tran-sition. While this challenge is understood and man-ageable in conventional flight control architectures, itpresents particular challenges when investigating largestate-space controllers.

Release Launch Scheme

For the release launch scheme, the control surfaceswere prepositioned for flight prior to releasing the

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LabViewGeneratedExcitations

+

dSpace2: Upper Control Block

Snub!

δCmd.

Trim: T10 or T11Ramp or Hold PZ

FlightMode

FlightMode Logic Block

Model Dynamics

ServoloopsWatchDog

Model Sensors

dSpace1

Model + Trim

GUI Input: FlightMode Supression On/Off

Suppression

Fig. 16 Trim and suppression control loops alongwith FlightMode logic developed by Lockheed Mar-tin.

pitch snub mechanism so that when the snubber wasreleased the model flies off the bottom stop. Themodel transitions from the stable Fixed-Fixed bound-ary condition to the stable free-free conditions withoutdwelling at the unstable Free-Fixed condition. Thisquick transition does not allow the unstable pitch-free configuration to build up any response amplitude.Therefore, there was no need to control the pitch-freeconfiguration using a Θ-controller.

The control architecture used with the releasemethod was a two-loop system as shown in figure 16.The inner loop contained the trim controller and wasresponsible for flying the model, and the outer loopcontained the suppression controller and was respon-sible for reducing the dynamic response. The suppres-sion controller could be switched on and off with littleor no effect on the average control surface trim posi-tion. As a result, switching between controllers causedlittle trim/elevation change and was done frequentlythroughout the test to acquire GLA data with as manycontrollers as possible without having to repeat thelaunch procedure.

The trim controller consisted of vertical positionPID gains a plus pitch rate gain and a static (bias)term. The proportional and integral terms were ig-nored until the model was unsnubbed. The output ofthe trim controller was a single command that was sentto all four trailing edge flap surfaces. The trim con-troller gains were initially established from analysis,and a single set of trim controller gains were identi-fied that were applicable to both the heavy and lightmodel. The bias term and the dynamic trim gainscould be adjusted from the dSpace2 GUI. Two ver-sions of the trim controller were used during Test 598,T10 and T11. Trim controller T11 was better suitedto TDT operations with the AOS system operating.

The suppression controllers were developed utilizingthe system identification and LQR/LQG techniquesdescribed in reference 18. System identification datawas acquired around the inner loop as indicated by the

thin lines in figure 16 providing the model+trim sys-tem dynamics. These system identification data werethen used with the LQR/LQG methods to build thesuppression controller. Therefore, the suppression con-trollers were aware of and expected the trim controllerto be part of the system. This subject is covered inmore detail in reference 19.

As with the TakeOff launch scheme, the WatchDoglimits associated with the Release launch scheme hadto be set to ignore the lower position limits during thelaunch and climb out phase of flight. For the Take-Off launch scheme, the logic for ignoring lower verticallimits was contained entirely within the Stateflow logicblock, and a separate set of WatchDog limits was se-lected for this phase of flight. For the Release launchscheme, a separate FlightMode for launch and climbout was not used. Instead, logic was built into theLM Signal block shown in figure 13 that ignored cer-tain parameters until PZ had reached a value of -5 in.Another difference between the two launch schemes isthat in the TakeOff launch scheme the operator setsthe SFMode parameter via the GUI, but FlightModeis event driven via the Stateflow logic block. Here, theoperator sets the FlightMode via the GUI when usingthe Release launch scheme.

For the Release launch scheme, FlightMode 0 wasa reset mode, FlightMode 1 was a trim control onlymode, and FlightMode 2 allowed the use of the sup-pression loop. Each FlightMode had a different set ofWatchDog limits. FlightMode 1 was used for launchand climb out, so parameters that would snub themodel when it was on the lower stop were ignored un-til PZ reached a value of -5 in. To protect the modelduring takeoff, in FlightMode 1 the WatchDog Θ lowerlimit was set to 0◦to minimize the possibility of a pitchover and subsequent hard landing event.

Data acquired during implementation of the Releaselaunch scheme are shown in figure 17. Here, time his-tories of PZ, Θ, flap 3 position, and tunnel dynamicpressure are plotted in the upper part of the figure,and the timing of boundary condition changes, Flight-Mode changes, and other system states are identifiedin the lower portion of the figure. User and eventdriven actions are identified. This figure should beconsulted when studying the following TakeOff schemesteps listed below:

• T=0s: The model is snubbed. Tunnel dynamicpressure is set to the launch value of 50 or 60 psf.FlightMode is set to zero by the operator causingthe control surfaces to be pre-positioned by thestatic part of the trim controller (δo). This alsoputs the WatchDog in reset mode, 0. The resetmode gets the latching logic associated with thelower position limits in WatchDog limit set 1ready for launch.

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in o

r deg

Dyn

amic

Pre

ssur

e, p

sf

0 5 10 15 20 25 30 35 40Time, s

FlightMode 2

Snubbed

Free

Fixed

Pitch: Fixed

Plunge: Fixed

FlightMode 0

WD: Flight Limits 1

Trim + GLA

FlightMode 1

WoW On

Trim: PZ Ramp

Unsnubbed

Free

Free

WoW Off

Trim: δo

User Commanded Event Driven

Trim: PZ Hold

0

-10

70

60

50

40

PZ, inΘ, degTE3POS, deg

WD: Flight Limits 2

Lwr Params Disabled All Paramerters Enabled

WD: 0

Fig. 17 Data acquired while using Release launchscheme to achieve trimmed flight. Test 598, Point2384.

• T=5.5s: Operator unsnubs the model usingreset switch on snubber control system chassis.The FlightMode logic block within the LMflight control architecture senses the unsnubbedcondition via the Snub! command signal andinitiates the dynamic part of the trim controller.The PZ setpoint for the trim controller beginsramping to tunnel centerline. The model startsto rotate but does not immediately lift off. Thisis an unstable condition and the model starts tooscillate in pitch.

• T=7s: The model lifts off the bottom stop andis supported on the damper. The integral partof the trim controller continues to increase thecontrol surface deflection and Θ.

• T=12s: The model clears the WoW switchindicating that it is off the dampers and hastransitioned to the free-free boundary condition.This event does not trigger any action withinthe Release scheme and associated LM controlarchitecture.

• T14s: The vertical position passes through -5inches and the WatchDog begins monitoring allsignals including the lower vertical limits.

• T=17s: The model vertical position passesthrough tunnel centerline (PZ=0), and the PZset point for the trim controller is set to zeroby the FlightMode logic block in the LM flight

Mach Number, M

Dyn

amic

Pre

ssur

e, P

SF

20

30

40

50

60

70

80

90

200

H=50PSF

75

100

150

200

300

400

500

600

700

800

900

1000

1200

1400

1600

1800

2000

2200

10

100

0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8

BFF Trim Only, Q=82 psf

BFF Trim Only, Q=100 psf

TakeOff Launch Scheme: Unsnub

TakeOff Launch Scheme: Rotate

Release Launch Scheme: Unsnub

GLA Data Acquired Here

BFF, Q=120 psf

BFF, Q=125 psf

Pitch-Free High Response

Fig. 18 Portion of the TDT heavy gas operat-ing envelope showing data acquired during Test598 where red represents the heavy weight con-figuration and green represents the light weightconfiguration.

control architecture.

• T=27s: FlightMode is manually changed toFlightMode 2 which allows the use of the suppres-sion control law but does not close the suppressionloop.

• T=31s: A suppression controller is engaged bythe operator via the dSpace2 GUI. The higherfrequency control surface activity is associatedwith the suppression loop being closed.

The Release launch scheme proved to be easy, effec-tive, and insensitive to dynamic pressure and modelconfiguration.

Testing ProceduresWith the exception of several air checkout runs, all

Test 598 data was acquired in heavy gas with a tunneltotal pressure between 1,670 and 1,880 psf. This mapsto a fairly tight band of Mach number and dynamicpressure combinations as shown in figure 18 where aportion of the TDT heavy gas operating envelope isshown along with the Test 598 data points. At anygiven tunnel total pressure, Mach number and dy-namic pressure are controlled by varying the RPM ofthe TDT main drive motor. For the test procedure

15 of 24


of Test 598, increasing RPM varied tunnel conditionsalong the plotted data points starting with the lowerleft. As Mach number is relatively insensitive to tun-nel RPM in this region, dynamic pressure was used toidentify all tunnel conditions.

As previously mentioned, the model and mount sys-tem had four testable boundary conditions. Data pre-sented in table 1 showed that the first bending modefrequency changed significantly based on these condi-tions, and as a result, model flutter speeds varied withsnub and flight status. Due to the possibility for highspeed vertical impact, the riskiest configuration waswith the model in a free-free, flying configuration. Thesafest configuration was with the pitch DOF snubbedand the model sitting on the lower stop. For boththe heavy and light configurations, the relatively safeconfigurations were cleared for flutter and the func-tionality of the emergency controller verified prior toproceeding with free-free flight at a give dynamic pres-sure. The subsections that follow will describe thevarious test procedures in general order of increasingcomplexity

Open-Loop Flutter Clearance

The model/mount system could be tested open-loopwhen the pivot carriage was sitting on the lower mountwith and without the pitch snub mechanism engaged.Flutter clearance testing of these configurations wasperformed in a manner typical of TDT flutter test-ing. Dynamic pressure was increased slowly while themodel was monitored visually and its signals moni-tored via strip chart.

For the Fixed-Fixed case, flutter clearance wassomewhat pro forma, as this configuration had theleast modeling unknowns, the highest analytical flut-ter speeds.4 Also, it had previously been tested inthis configuration (Test 579 and Test 593) up to a dy-namic pressure of at least 80 psf. During Test 598, thefree-free configuration was only cleared for flutter to adynamic pressure of 60 psf. The high degree of confi-dence in the analysis and the concern that the modelwould not remain on the lower stop during this testingdue to the fixed, positive, snub pitch angle precludedflutter clearance testing above 60 psf.

The other open-loop configuration that was testedfor flutter was pitch-free with the model on the lowerstop. This configuration required that the model bemanually trimmed to keep pitch angle from gettingtoo large causing the model to lift off the lower stop.This configuration was of interest because it was themodel boundary condition state for the Taxi mode inthe TakeOff launch scheme. The exact flutter onsetdynamic pressure was somewhat ambiguous for thisconfiguration as this flutter mechanism appeared tobe a hump mode with a very shallow flutter crossing.

Experimentally, this was demonstrated by an inconsis-tent flutter dynamic pressure and a relatively benignflutter mechanism. This behavior occurred at dynamicpressures between 30 and 50 psf.

Snubbed Flight Clearance and EmergencyController Validation

The next stop on the way toward achieving trimmed,free-free flight was to verify and tune the emergencycontroller. The purpose of the emergency controllerwas to gently return the model to the lower stop us-ing gains and biases scheduled with dynamic pressure.The bias values had to be adjusted based on exper-imental data. An additional purpose of this testingwas to simultaneously verify that the model was flut-ter free in the snubbed flying configuration up to themaximum expected dynamic pressure of interest.

The process for clearing the model for flutter wasto fly the model in the snubbed configuration at tun-nel center line (PZ=0) and incrementally increase dy-namic pressure stopping approximately every 10 psf.Once aeroelastic stability had been established, con-troller validation data was acquired using the emer-gency controller to track a sawtooth position commandbetween -10 inches and +10 inches with a ramp rate of±1 in/sec. Data was acquired for 3 complete saw toothcycles. The average control surface position neededto fly the model at each dynamic pressure was deter-mined and used as the static component (δo) of theemergency controller.

In preparation for GLA testing this clear-ance/validation process was performed from the take-off dynamic pressure of 50 psf to 70 psf, the maximumdynamic pressure where GLA data would be acquired.In preparation for BFF testing, the process was re-peated for dynamic pressures above 70 psf to a maxi-mum dynamic pressure of 130 psf. The testing processhad to be repeated for the light model.

Gust Load Alleviation Testing

So far, this paper has described the apparatus, pro-cedures, and preliminary testing that were necessaryto achieve trimmed, free-free flight in a safe and pre-dictable manner. Demonstrating that a control lawreduced loads due to gusts required that data be ac-quired using different control laws all in the same gustenvironment. Here, the acquisition of the GLA datawith the TDT AOS system operating will be discussed.

The biggest problem encountered with AOS test-ing was simply to get the model in a trimmed fly-ing state with the AOS operating. Initial attemptsfailed when the model was first launched into a stable,trimmed flight condition prior to turning on the AOS.The AOS uses a large flywheel to hold constant vanefrequency, and therefore, changes in frequency occurrelatively slowly. As a result, when the AOS system

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was engaged, the model encountered a low frequencysinusoidal gust field with flow angles of approximately±1◦(at the initial AOS amplitude setting of 12◦peak-to-peak). The vertical travel required for the wing tofly through this gust field exceeded the vertical travelavailable on the mount system, and the WatchDogsystem or test engineer would snub the model. Thesolution to this problem was to bring the AOS on-lineprior to launching the model. An AOS frequency wellabove the rigid body modes and between the resonantfrequencies of the flexible modes (1st in-plane and 2ndout-of-plane bending) was used, typically 6.5 or 9.5 Hz.This technique allowed both the TakeOff and Releaselaunch schemes to be successfully used.

An interesting aside regarding the AOS is that onseveral occasions launch sequences that had previouslybeen successfully employed failed to work. In theseinstances, the model would fly to the top of its ver-tical travel and trip the WatchDog system. It wasdetermined that the AOS vanes had been parked ata nonzero angle, and this flow angularity change wasenough to cause the model to fly to its upper travellimit during launch attempts. Re-zeroing the AOSvanes solved the problem.

For the Release scheme, it was experimentally de-termined that trim controller T10 could only be usedduring model launch with the AOS off. Thus, the trimcontroller gains had to be adjusted slightly in orderto successfully launch the model when the AOS wasoperating. Two versions of the Release scheme trimcontroller were ultimately used, the AOS and non-AOSversions or T11 and T10 trim controllers, respectively.The higher gains in T11 provided a snappier launchand better position tracking. Although T11 could beused to launch the model with and without the AOSoperating, this was generally not done as T10 was thepreferred baseline controller for the AOS off cases.

In the case of the TakeOff scheme, the same controllaws could be used to launch with and without theAOS operational. Since trim and GLA were combinedfor the TakeOff scheme, the inherent disturbance re-jection associated with GLA may have allowed thesecontrol laws to achieve this without any special AOSrelated design considerations.

At this point, the differences between the controllaw design methods associated with the TakeOff andRelease schemes dictated how the testing for each wasconducted. For the TakeOff scheme, the trim/GLAcontrollers were designed using state-space analyti-cal models. LQG methods were employed where theweightings were varied to produce different controllersthat provided various levels of GLA performance, andmany such controllers were evaluated. The GLA evalu-ation procedure consisted of acquiring model responsedata for each control law subject to AOS dwells in 0.5

to 0.1 Hz frequency increments. Generally, this type oftesting would start at a relatively high AOS frequencyand proceed to lower frequencies until the WatchDogor test engineer snubbed the model. When the modelwas snubbed, the AOS operator would return to thesystem to 6.5 Hz to reduce the large amplitude wingbending oscillations associated with peak response fre-quencies and to prepare for another launch using adifferent control law. Figure 19 provides some ex-ample data where the AOS frequency was about 0.6Hz. Here, the vertical displacement is quite large ap-proaching ±10 in. At T=4.9s, a WatchDog limit wasexceeded and the model was snubbed and returned tothe lower stop by the emergency controller.

When using the Release launch scheme and associ-ated LM controller architecture, a system ID approachwas used to generate the GLA controllers. The sys-tem ID data was acquired with uncorrelated randomexcitations sent simultaneously to each flap. These ex-citations were generated externally using a LabViewsystem with the excitations added to the flap com-mands within dSpace2 as shown in figure 16. Asdiscussed earlier, the system ID data needed to cap-ture the model with the appropriate trim control loopengaged.

Acquisition of system ID data for use when the AOSwas not engaged was relatively straight forward. How-ever, a new procedure had to be developed to generatesystem ID data for generating GLA controllers for usewith the AOS system engaged. Here, the external fre-quency command feature of AOS was used with theLabView system to sweep through the frequency rangeof interest. Typically, this was from an upper fre-quency of 9.5 to 11 Hz down to a frequency slightlyabove where the model was expected to exceed Watch-Dog limits then back to the upper frequency. Therandom excitations were simultaneously sent to theflaps while the AOS system was sweeping.

Suppression controllers generated using these datacould be evaluated using the AOS dwells, describedabove, or the AOS sweep excitation without the ran-dom flap inputs included. Time traces acquired duringan AOS sweep are shown in figure 20. Here, the AOSfrequency is varied from 9.5 to 2.0 back to 9.5 Hzduring a 100 second data record. Large bending andplunge displacement responses were obtained at thelow AOS frequencies.

The lowest frequency that could be obtained usingeither the AOS dwell or sweep excitations was a func-tion of dynamic pressure. For the heavy model anda dynamic pressure of 70 psf, the WatchDog systemwould generally trip between 3.0 and 2.5 Hz, and at 60psf it would typically trip between 1.7 and 2.5 Hz. Theactual lower limit was a function of the specific con-troller engaged at the time. Similar trends were noted

17 of 24


0 2 4 6 8 10

10

0

-10

deg o

r in

6000

4000

2000

0

Root B

endin

g M

om

ent, in−

lbs

Θ, degPZ, inSnub!

Time, s

Fig. 19 AOS transitioning from 0.7 to 0.5 Hz,WatchDog engaged Snub! command. Test 598,Point 4891. The bending moment is measured fromstrain gauge SBI1 shown in figure 8 corrected foroutboard wing weight.

5

0

-5

de

g o

r in

10

5

0 AO

S F

req

ue

ncy, H

z

0 10 20 30 40 50 60 70 80 90 100

4000

2000

0

-2000

Be

nd

ing

Mo

me

nt, in

-lb

Time, s

Θ, deg

PZ, in

Fig. 20 Time traces of vertical position, pitch an-gle, and bending moment while AOS sweeping from9.5 to 2.0 to 9.5 Hz. Test 598, Point 3182. Thebending moment is measured from strain gaugeSBI1 shown in figure 8 corrected for outboard wingweight.

for the model in the light configuration; however, asthe light model had a higher first bending mode fre-quency, the lower frequency limit at which data couldbe acquired was higher. For the final week of testing,the peak-to-peak amplitude of the AOS vanes was ad-justed 12◦to 4◦. Following this change, data couldbe acquired with AOS frequencies below 1.0 Hz fordynamic pressures of 60 psf as previously shown infigure 19.

More details on the GLA control laws and the exper-imental results can be found in references 17 and 19.

Reduced Static Margin Testing

The moving mass could be remotely adjusted fromthe TDT control room during wind-tunnel testing. For

model launch and most subsequent testing, the nomi-nal location of this mass was full forward for the heavyconfiguration and full aft for the light configuration. Inthe heavy configuration, the moving mass and model’sCG could be moved aft to investigate reduced staticmargins. The procedure was to establish trimmedflight with the mass in the nominal (full forward) loca-tion and set the dynamic pressure to the desired value.Then the mass was moved aft incrementally until themodel went unstable and was snubbed or the massmade it to the full aft position. Data was acquiredperiodically during this process.

BFF Suppression Testing

For flutter testing, the goal was to establish the ex-perimental flutter onset dynamic pressure associatedwith a given control law. As with the flutter clearancetesting, dynamic pressure was increased incrementally;however, since flutter was expected, the incrementswere reduced as dynamic pressure was increased intouncharted territory. For this testing, naturally occur-ring tunnel turbulence was generally deemed adequateto perturb the model, but control surface doubletswere occasionally used to free stiction in the mountsystem. Ultimately, flutter onset was determined bythe test engineer’s assessment that the model wasunstable or when the WatchDog limits had been ex-ceeded. In either case, the model would be snubbed.Figure 21 shows data from one such BFF encounterwhere the WatchDog system engaged the snubber andthe model was subsequently landed via the emergencycontroller. Note that Θ and PZ have the same fre-quency and are growing in amplitude.

Controllers associated with the TakeOff launchscheme were tested for flutter. However, BFF suppres-sion was not a test objective of Northrop Grumman,and BFF testing was limited to experimentally deter-mining the flutter onset dynamic pressure using themanner described above. System ID data was not ac-quired.

Controllers associated with the Release launchscheme were tested extensively for flutter in the heavyand light configurations. As the objective of this test-ing was to generate BFF suppression controllers thatwould increase flutter onset above the trim-controller-only value, a boot strap method had to be employed.The first step was to establish the trim-controller-only(baseline) flutter onset dynamic pressure, acquiringsystem ID data incrementally as dynamic pressure wasincreased. This data was then used, off-line, to gen-erate BFF suppression control laws. When testingresumed with the Trim+BFF-suppression controllerengaged, system ID data would again be acquiredas dynamic pressure was increased until a new flut-ter onset value was established. The system ID and

18 of 24


de

g o

r in

84

82

80

Dyn

am

ic P

ressu

re, P

SF

20 22 24 26 28 30 32 34 36

10000

5000

0

-5000

Be

nd

ing

Mo

me

ing

, in

−lb

s

Time, s

5

0

-5

Θ, degPZ, inSnub!

Fig. 21 Body freedom flutter time trace for trimonly control, light model. WatchDog Snub! en-gagement shown along with vertical position rampdown. Test 598, Point 3525. The bending momentis measured from strain gauge SBI1 shown in fig-ure 8 corrected for outboard wing weight.

control law design procedure was repeated until themaximum value of flutter onset dynamic pressure wasestablished. More details on the BFF suppression con-trol laws and a discussion of the experimental resultscan be found in references 11 and 19

Concluding Remarks

In the Fall of 2007, the Air Force Research Labora-tory (AFRL), Northrop Grumman, Lockheed Martin,and NASA Langley Research Center successfully com-pleted the third of a series of three wind-tunnel tests ofan aeroelastically scaled wind-tunnel model of a flyingwing SensorCraft vehicle concept. The first of thesetests was conducted on a cantilevered, sidewall mount.The second and third tests (2006 and 2007) used anew, multi-degree-of-freedom mount. This mount al-lowed the semi-span model to translate vertically androtate in pitch at the wing root, allowing better simu-lation of the full span vehicle’s rigid-body modes. Thelarge size of the flying wing model along with the rigidbody degrees-of-freedom afforded by the new mountsystem, created many unique challenges to successfullyand safely flying the model in the tunnel and meet-ing the AEI test objectives. This paper has providedan overview of the AEI flying-wing wind-tunnel testsfrom an operational point of view. It has focused onsystems, procedures, and lessons learned that enabledthe test to be successful in demonstrating Gust LoadAlleviation (GLA) and Body Freedom Flutter (BFF)suppression. Descriptions of the hardware includingthe wind-tunnel, the wing model, the mount system,and other supporting systems developed for this pro-gram were provided.

1.5

1.0

0.5

0

0

-20

-40

-60

-80

-100

Ma

gn

itu

de

0 4 8 12 16 20Frequency, Hz

Ph

ase

, D

eg

TE1TE2TE3TE4LE

Fig. 22 Flap transfer function estimates for be-ginning (solid), middle (dash), and end (dot dash)of Test 598, Points 216, 1600, and 5409.

AcknowledgementsFrom Northrop Grumman the authors would like

to acknowledge Tony Shimko for his expertise andguidance; Elaine Shaw for her tireless control design;test engineers, Lou Scherer, Chris Curnes and KianTehrani; and finally, Ted Schulman for much appreci-ated experience and insight.

From Lockheed Martin the authors would like toacknowledge Wink Baker and David White for theircontributions to Test 598 that including model prepa-ration, calibration, and wind-tunnel testing.

From NASA Langley Research Center the authorswould like to acknowledge Chuck McClish for buildingthe hydraulic manifold assembly; Patrice Quander-Haas and Gary Koepple for fabricating the snub con-trol chassis, the moving mass controller, and theirpatience in dealing with ever changing requirements;Vic Tumwa for his assistance in conducting ground vi-bration tests; William Johnston and John Newman forperforming the LAS spring fatigue test on such shortnotice; and Mike Ramey, Jeff Wilfong, and AnthonyPototzky for their assistance in conducting the wind-tunnel test.

Finally, we would like to acknowledge the design,fabrication, and test support efforts of NextGen Aero-nautics including David Cowan, Matthew Scott, DanaHoward, Robert Alphenaar, Robert Myers, JonathanProuty, and Richard Guhr.

Appendix A: Flap Actuator ResponseDue to the aeroservoelastic nature of these wind-

tunnel tests, the frequency responses of the five flapactuators was important. The PID gains on thefive control loops were established by examining oneactuator-loop at a time while tracking a square wavecommand. Command and position are plotted in nearreal time on the dSpace1 host system. The procedure

19 of 24


for setting the gains is to start with zero integral andderivative gains, and increase the proportional gainuntil actuator instability then back off approximately20%. The derivative gain is then adjusted to improvestability. This process is repeated until satisfactory re-sponse characteristics are obtained. Integral gain wasnot found to be effective and was generally set to zero.

Given the fact that GLA testing is quite demand-ing in terms of the actuator stroke requirements andthe number of cycles, the flap actuators performancewas satisfactory. This was especially true for Test 598where many hours of trimmed flight was successfullydemonstrated with the AOS operating. The actua-tors were, however, the highest maintenance items forTest 598, and three main problems were encountered.The first of these were leaks in the o-rings that sealthe actuator shafts. They were a known wear item,and periodic replacement was expected. The othertwo problems were more troublesome. There was anissue with the end-cap o-ring seal where slight varia-tions in the installation procedure and/or lot numbervariations in the o-rings could lead to part of the o-ringbinding against the actuator shaft. This binding wouldlimit actuator bandwidth and linearity, and could beidentified by the flap position time trace having adistinctive clipped appearance. Finally, in terms ofactuator performance, free play was the biggest prob-lem. Since, the flap position sensor (RVDT) was notattached directly to the actuator shaft, but to the flap,any free play between the actuator shaft and the flapwould interfere with the servo control loop.

As part of the start-of-day operations, data was ac-quired with a 30 second sine-sweep position commandbeing applied to all five flaps. Command and positiontime histories were acquired, and the actuator transferfunctions were estimated using the MatLab TFESTI-MATE command. Transfer function estimates for thebeginning, the middle, and the end of Test 598 areshown in figure 22. As can be seen, there was somevariation in actuator transfer functions throughout thetest with flaps TE2 and TE4 having chronic phase is-sues compared with the other surfaces due the inabilityto remove the free play.

Cross resonance in the model could also excite theactuator free-play in other actuators causing sympa-thetic limit cycles. In these cases, a reduced propor-tional gain was needed to maintain stability. By theend of the test, the LE actuator would enter a limitcycle above a dynamic pressure of 60 psf, and it had tobe substantially detuned in order to maintain stabilityas indicated by the large phase lag.

8

6

4

2

0

Test 593

Θ, deg.

8

6

4

2

0

3

2

1

Most Restrictive

Least Restrircted(T593 Setting)

Test 598 Without Wing Installed

Valve Removed,Thorough Bleed

Valve Installed,Least Restrictive

0 0.2 0.4 0.6 0.8 1.0

Test 598

Time, s

Least RestrictiveT598 Setting

Θ, deg.

Θ, deg.

Most Restrictive

Fig. 23 Snubber system time delay experiments,hydraulics at 2000 psi.

Appendix B: Snubbing SystemPerformance

The performance of the snubber system was an im-portant consideration for the AEI wind-tunnel tests.Snubber performance data acquired during Test 593and Test 598 will be discussed here. Figure 23 showsthe evolution of the snubber system performance start-ing with data from Test 593, followed by some experi-ments from Test 598 buildup where the wing was notinstalled, and finally, data acquired at the start of Test598. In all cases, pitch angle is plotted versus time, andthese time traces are lined up such that the Snub! com-mand for each was issued at T=0s.

Initial testing of the snubber system performance forTest 593 is shown in the upper part of figure 23. Anadjustable flow rate valve was added to the hydraulicline feeding the snubber system to provide some con-trol over the snub rate, and three flow rate settingswere considered. Here, with the least restrictive valvesetting, the time required to snub the model from theinitiation of a Snub! command exceeded 0.5 seconds,and this setting was used in Test 593. It is interest-ing to note that these data were acquired when thesnubber hydraulics were first installed, and no furthertiming studies were performed. However, subsequentand frequent operational usage of the snubber systemmay have resulted in some break-in as the operationalsnubs indicated a shorter time delay of about 0.3 sec-onds. For example, data from the end of Test 593 isplotted in figure 12 where there appears to be about0.3 seconds of delay between the Snub! command andtheta being clamped.

In preparation for Test 598, the time delay betweenissuing a Snub! command and the pitch angle beinglimited and subsequently clamped was a concern. Thecenter plot in figure 23 shows the results of some stud-

20 of 24


ies of the snubber system prior to wing installation.The initial data set (blue) indicated a time delay sim-ilar to what was observed in Test 593. As a result,three possible fixes were considered, and they werethe removal of the flow rate valve, a better methodof bleeding the air out, and a diode installed acrossthe solenoid electrical leads. As it turned out, the ro-tation of the pivot carriage assembly, discussed earlier,ended up being serendipitous as it facilitated bleedingthe hydraulics. With the actuator in a vertical strokeposition, air in the actuator would rise to the top. Ifthe actuator was repeatedly cycled and bled, all theair from the system could be removed. A satisfactoryresult was finally obtained as indicated in the centerfigure.

Snubber timing studies of the system in the finalTest 598 configuration are shown in the lower plot offigure 23. Here, the wing has been installed, the sys-tem has been thoroughly bled, and the flow rate valvewas used. Several flow rate settings were investigatedwith the final Test 598 setting identified in the plot.The net time delay used in subsequent analysis and inthe WatchDog system was estimated to be 0.15 sec-onds. The reason the flow rate valve was includedwas that engaging the snubber mechanism could, un-der some circumstances, induce wing bending loadsnear limit values. The final flow rate setting was acompromise between snubber speed and induced loads.One known circumstance worth mentioning was air-offsnubber engagement. If the model was snubbed air-off with full hydraulic pressure, large bending loadswould result. To avoid this problem, at the start ofTDT operations, the model was electrically snubbed(solenoid valve open) prior to bringing the hydraulicpump on-line. As the initial set point for the pumpwas around 200 psf, the model could be brought to asnubbed state relatively slowly with minimal bendingloads. The hydraulic pump would then be brought upto its operational pressure of 2,000 psi.

The snubbing system performed quite well through-out Test 598 with one minor exception. The solenoidvalve must be energized to unsnub the system. There-fore, removing power to the solenoid valve snubs themodel. There were several instances during both Test593 and Test 598 when the snubber mechanism couldnot be unsnubbed as the valve had apparently stuckin the snubbed position. If the solenoid was left de-energized for approximately 30 minutes, it would startfunctioning properly again. Installation of a pancakefan to cool the solenoid coil and valve solved the prob-lem during Test 598.

Appendix C: Rail Friction and LASThe model and mount system contained a variety

of unusual or high maintenance systems that were all

critical to the success of Test 598. One critical areawas the performance and longevity of the LAS andthe friction associated with the mount system rails.These items will be discussed here.

Prior to Test 598, one area of concern was the po-tential for frequent replacement of the LAS springs.The time involved in clearing the R134a test mediumfrom the TDT along with the time required to changethe LAS springs would result in the loss of at leastone day of test time. As the model would be flyingas close to tunnel center line as possible, the centerof the LAS springs would receive the most cycles, anddata from Test 593 indicated that the vendor ratedlife of 13,000 cycles could be reached several times aweek. To assess the level of conservatism in the ven-dors rated life, a fatigue test of the springs mounted inthe LAS assembly was conducted using an Instron ma-chine at NASA Langley Research Center. The springswere pulled out a distance of 12 in to be consistentwith the configuration expected with the model flyingat tunnel centerline, and the springs were cycled at ±1inch at 1 Hz. These tests indicated that the springshad a fatigue life of at least 98,000 cycles.

For Test 598, Lockheed Martin introduced the use ofload cells to measure the LAS force in the lab, duringTDT buildup, and during wind-tunnel testing. An ad-ditional load cell could be temporarily attached to thewing to measure the force required to lift the wingthrough its vertical range of motion. Reference 11describes the extensive rail and LAS friction experi-ments performed during build-up in the TDT modelpreparation area. Based on these experiments, theLAS friction was determined to be about ±2.5 lbs perspring set (forward and aft) and the rail assembly fric-tion was between ±11 and ±17 lbs. Once the wing,mount, and fairing were assembled in the final testconfiguration an abbreviated friction experiment wasperformed periodically to assess LAS and rail friction.This experiment consisted of acquiring load cell andvertical position data while lifting the wing via thethird load cell through its vertical range of motion upand down. Data acquired at the start of Test 598 andafter about 4 weeks of wind-tunnel testing is plottedin figure 24. These data indicated a hysteresis loopwhere force is dependent on the direction of motion.For the purposes of health monitoring, these data wereexamined for consistency. The data acquired at point3237 was deemed satisfactory and the wind-tunnel testwas continued.

The LAS load cell data was acquired by the DAS andmonitored by the test engineers during wind-tunneltesting. LAS load cell data plotted versus wing ver-tical position is shown for three different take-off andsubsequent trimmed flights in figure 25. The data frompoint 3245 indicates a healthy LAS system. Closer ex-

21 of 24


80

70

60

80

70

60

Forward Load Cell

Aft Load Cell

−15 −10 −5 0 5 10 15

350

300

250

PZ, in/sec

Fo

rce

, lb

sF

orc

e, lb

sF

orc

e, lb

s

Wing Mounted Load Cell

Fig. 24 LAS and wing mounted load cell movingaverage data acquired during wing vertical transla-tion. Test 598, Points 212 (blue) and 3237 (green).

100

60

20

Fo

rce

, lb

sF

orc

e, lb

sF

orc

e, lb

s

Point 3245

Point 3435

−15 −10 −5 0 5PZ, in

Point 3455

100

60

20

100

60

20

Fig. 25 LAS load cell data acquired during threedifferent takeoff and flight sequences. Test 598,Points 3245, 3435, and 3455. (Forward load celldata is blue and the aft load cell data is green).

amination of this data indicates that during flight, theLAS friction is no more than ±1.5 lbs, less than whatwas measured air-off. The higher vibration environ-ment associated with wind-on operations may improveLAS performance in this regard. The next two datapoints indicated a weakening and failed, respectively,forward LAS spring assembly. The LAS system failedtwice during Test 598. The first failure was at thepoint shown here near the middle of the test. Duringrepair of the LAS system, the model was changed tothe light weight configuration. The LAS also failed onthe final day of testing.

Appendix D: Development of LMCombined Parameters

A significant analysis effort was undertaken byLockheed Martin Aeronautics to determine additionalWatchDog parameters so that a Snub! command wouldbe issued in time to snub the model and subsequentlycontrol plunge displacement using the emergency con-troller before the pivot carriage hit the dampers.These combined parameters needed to take into ac-count the time delay in physically snubbing the modeland vertical travel needed by the emergency controllerto obtain a zero sink rate. This appendix will describehow the LM combined parameters were developed, ver-ified, and implemented.

A simulation model of the wing, mount system, andcontrol systems was developed in MatLab Simulink.The wing, including the mass of the pivot carriageassembly, was modeled using a series of state-spacemodels for various combinations of Mach number anddynamic pressure with two-dimensional interpolationused to generate the ASE model used at a given flightcondition. The ASE model contained inputs for 5 con-trol surface position commands, and a vertical gustvelocity input. The physical constraints and forcesassociated with the rest of the mount system were im-posed on the wing root using 6 force/moment inputs(x, y, z, roll, pitch, yaw). The effects of mount flexibil-ity were modeled using these force terms. In the caseof the vertical force, terms for friction as a function ofbending moment, the LAS, the damper, and the hardstop were included. The pitch moment force includedterms for the pitch dampers, hard stop, and the pitchclamp actuator. The outputs of the ASE model werethe 23 aircraft sensors.

A series of simulations were run to determine therelationship between the initial condition of the modeland vertical travel needed to obtain a zero sink rate.The simulation started with an initial vertical rate,pitch angle, pitch rate, and dynamic pressure for boththe heavy and light model configurations The initialconditions for all other parameters were assumed to bezero. The time delay between Snub! command and apitch restoring force being applied to the wing root wasalso assumed to be zero. The simulation studies indi-cated that if the model was not snubbed at the timeof pivot carriage to damper contact, large pitch excur-sions and wing bending moments would result. Thus,wing bending moments were within allowable limitsas long as the model was snubbed prior to verticaldamper contact. The simulation data was then used toconstruct a new WatchDog parameter. The parameterwas generated by determining the relationship betweenvertical velocity, pitch angle, and dynamic pressureand the vertical distances up or down needed to stopthe model once the snubber had been engaged. This

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PZ(t)

VZ(t)

Θ(t)

Θ(t)

Feed ForwardEstimator:

Assumed SnubTime Delay = ∆t

10x10Neural

Network

DynamicPressure

Up Dist

COMBLO

COMBHI

Down Dist

VZ(t+∆t)

PZ(t+∆t)

+

+.Θ(t+∆t)

Fig. 26 Lockheed Martin combined parameterblock diagram.

mathematical relationship was represented by a 10 by10 neural network.

In order to account for any time delay between theSnub! command and pitch snub mechanism physicalengagement, some parameters were extrapolated intothe future. A simple feed forward estimation routineusing an assumed snubber time delay was applied tovertical position, vertical rate, pitch angle, and pitchrate. The forward estimator and neural network werecombined in series in the WatchDog system as shownin figure 26 to provide a real-time estimate of the ver-tical position, up or down, where the model would bebrought under control. These values were monitoredby the WatchDog system, and as with the other moni-tored signals, the WatchDog would issue a Snub! com-mand when thresholds were exceeded.

To assess snubber time delay requirements, 5 hardlanding cases from Test 593 were examined. For oneof these data points, time histories of vertical position,pitch angle, Snub! command, trailing edge surface 2command, and wing root bending moment are plottedin figure 27. Here, it can be seen that the model startsto climb then falls back, lands hard, and bounces.Pitch angle and bending moment decrease rapidly af-ter contacting the damper just before T=42s whereslope discontinuities are observed in the vertical posi-tion, pitch angle, and bending moment time histories.These data are consistent with the simulation resultsindicating the need to have the model clamped in pitchprior to damper contact to minimize subsequent pitchangle excursions and high wing bending loads.

Table 2 WatchDog performance assessed using 5hard landing cases from Test 593.

Assumed RequiredSnub TimeTime Delay WD PerformanceDelay (pt. 1252) In-Time False Alarm

0.065 2/5 1/50.10 0.113 5/5 0/50.25 0.165 2/5 1/5

For the 5 hard landing cases, the performance ofthe WatchDog system monitoring only the originalmeasured parameters was compared with the Watch-Dog system configured to monitor the LM combined

40.0 40.5

Time, s

41.0 41.5 42.0 42.5 43.0 43.5 44.0

15

10

5

0

-5

-10

-15

-20

-25

PZ, inΘ, degTE2 Command, degBending Moment, 1000 In−lb

Snub!

Fig. 27 Hard landing case plot of vertical po-sition and model angle with damper contact andSnub! command. Test 593, Point 1252. The bend-ing moment is measured from strain gauge SBI1shown in figure 8 corrected for outboard wingweight.

41.2 41.4 41.6Time, s

41.8 42.0 42.2

10

8

6

4

2

0

-2

-4

-6

-8

-10

PZ, in

Θ, deg

Bottom Damper Contact

Snub! Nominal Parameters

Snub! LM Comb, delay=0.10 s

Snub! LM Comb, delay=0.25 s

in o

r deg

Fig. 28 Hard landing case comparing nominalWatchDog performance with and without com-bined parameters. Test 593, Point 1252.

parameter using two different assumed snubber timedelays of 0.1 and 0.25 seconds. One of these hardlanding cases is shown in figure 28 where the modelvertical position and pitch angle are plotted along withthe Snub! command associated with the three Watch-Dog configurations. Damper contact time is indicatedby the red line. For this data, the original Watch-Dog system would have snubbed the model as it wasdescending through a -10 inch/sec velocity threshold.The combined parameter cases caught the model ear-lier, as shown. The figure of merit for these cases is thetime difference between damper contact and when thecommand was issued. If the actual, physical snubbertime delay is less than the required time delay, then a

23 of 24


given WatchDog parameter would have been success-ful in avoiding model and damper contact and highloads. Required time delay for the data shown in fig-ure 28 along with a summary of the performance of theWatchDog parameters for all 5 Test 593 cases exam-ined is shown in table 2. A false alarm is defined as acase where a Snub! command is issued, but the bend-ing load limits were not subsequently exceeded in thedata set. The combined parameter with an assumed0.1 second snubber time delay was able to snub themodel before contacting the damper in all 5 cases withno false alarms. As previously discussed, the physicalsnubber system was ultimately configured to providea delay of 0.15 seconds.

References1Martinez, J., “An overview of SensorCraft capabilities and

key enabling technologies,” 26th AIAA Applied AerodynamicsConference, No. AIAA-2008-7185, Honolulu, Hawaii, Aug. 2008.

2Reichenbach, E., “Aeroservoelastic Design and Test Valida-tion of the Joined Wing Sensorcraft,” 26th AIAA Applied Aero-dynamics Conference, No. AIAA-2008-7189, Honolulu, Hawaii,Aug. 2008.

3LeDoux, S., Vassberg, J., and Fatta, G., “AerodynamicCruise Design of a Joined Wing SensorCraft,” 26th AIAAApplied Aerodynamics Conference, No. AIAA-2008-7190, Hon-olulu, Hawaii, Aug. 2008.

4Lockyer, A. J., Drake, A., Bartley-Cho, J., Vartio, E.,Solomon, D., and Shimko, T., “HIGH LIFT OVER DRAGACTIVE (HiLDA) WING; Delivery Order 0007: HiLDA WingProgram,” Tech. Rep. AFRL-VA-WP-TR-2005-3066, NorthropGrumman Corporation, 2005.

5Vartio, E., Shimko, A., Tilmann, C. P., and Flick, P. M.,“Structural Modal Control and Gust Load Alleviation for aSensorCraft Concept,” 46th AIAA/ASME/ASCE/AHS/ASCStructures, Structural Dynamics and Materials Conference, No.AIAA-2008-7185, Austin, Texas, April 2005.

6Silva, W., Vartio, E., Shimko, A., Kvaternik, R. G.,Eure, K. W., and Scott, R. C., “Development of Aeroser-voelastic Analytical Models and Gust Load Alleviation Con-trol Laws of a SensorCraft Wind-Tunnel Model Using Mea-sured Data,” 47th AIAA/ASME/ASCE/AHS/ASC Structures,Structural Dynamics and Materials Conference, No. AIAA-2006-1935, Newport, Rhode Island, May 2006.

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