October 2013

NASA/TP–2013-218051

Aeroservoelastic Testing of Free Flying Wind

Tunnel Models

Part 1: A Sidewall Supported Semispan Model Tested for 

Gust Load Alleviation and Flutter Suppression

Robert C. Scott

Langley Research Center, Hampton, Virginia

Travis K. Vetter

Northrop Grumman Corporation, El Segundo, California

Kevin B. Penning

Lockheed Martin, Fort Worth, Texas

David A. Coulson

Analytical Services and Materials, Inc., Hampton, Virginia

Jennifer Heeg

Langley Research Center, Hampton, Virginia
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October 2013

NASA/TP–2013-218051

Aeroservoelastic Testing of Free Flying Wind

Tunnel Models

Part 1: A Sidewall Supported Semispan Model Tested for 

Gust Load Alleviation and Flutter Suppression

Robert C. Scott

Langley Research Center, Hampton, Virginia

Travis K. Vetter

Northrop Grumman Corporation, El Segundo, California

Kevin B. Penning

Lockheed Martin, Fort Worth, Texas

David A. Coulson

Analytical Services and Materials, Inc., Hampton, Virginia

Jennifer Heeg

Langley Research Center, Hampton, Virginia
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National Aeronautics and Space Administration. 
Abstract

This is part 1 of a two part document. Part 2 is titled: “Aeroservoelastic Testing of Free

Flying Wind Tunnel Models Part 2: A Centerline Supported Fullspan Model Tested for

Gust Load Alleviation.” A team comprised of the Air Force Research Laboratory (AFRL),

Northrop Grumman, Lockheed Martin, and the NASA Langley Research Center conducted

three aeroservoelastic wind tunnel tests in the Transonic Dynamics Tunnel to demonstrate

active control 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 vehicle

was mounted to a force balance to demonstrate gust load alleviation. In the second and

third tests, the same wing was mated to a new, multi-degree of freedom, sidewall mount.

This mount allowed the half-span model to translate vertically and pitch at the wing root,

allowing better simulation of the full span vehicle’s rigid body modes. Gust load alleviation

(GLA) and Body freedom flutter (BFF) suppression were successfully demonstrated. The

rigid body degrees-of-freedom required that the model be flown in the wind tunnel using

an active control system. This risky mode of testing necessitated that a model arrestment

system be integrated into the new mount. The safe and successful completion of these free

flying tests required the development and integration of custom hardware and software.

This paper describes the many systems, software, and procedures that were developed as

part of this effort.

1
Contents

1 Introduction 6

2 Transonic Dynamics Tunnel 8

3 Wing and Mount System 9

3.1 Wing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 10

3.2 Mount System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 14

3.3 Test Configurations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 15

4 Signal Routing, Processing, and Control 16

4.1 DCS Hardware and Development Environment . . . . . . . . . . . . . . . . . 17

4.2 Hydraulics and Snubber Control System . . . . . . . . . . . . . . . . . . . . . 18

5 WatchDog System 18

5.1 Emergency Controller . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 20

5.2 WatchDog Snub Logic . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 21

5.3 NG—Combined Parameters . . . . . . . . . . . . . . . . . . . . . . . . . . . . 21

5.4 LM—Combined Parameters . . . . . . . . . . . . . . . . . . . . . . . . . . . . 22

6 Flight Control Schemes 22

6.1 Takeoff Launch Scheme . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 22

6.2 Release Launch Scheme . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 24

7 Testing Procedures 27

7.1 Open-Loop Flutter Clearance . . . . . . . . . . . . . . . . . . . . . . . . . . . 29

7.2 Snubbed Flight Clearance and Emergency Controller Validation . . . . . . . . 29

7.3 Gust Load Alleviation Testing . . . . . . . . . . . . . . . . . . . . . . . . . . . 30

7.4 Reduced Static Margin Testing . . . . . . . . . . . . . . . . . . . . . . . . . . 32

7.5 BFF Suppression Testing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 32

8 Concluding Remarks 33

9 Acknowledgements 34

A Flap Actuator Response 37

B Snubbing System Performance 38

C Rail Friction and Lift Augmentation System (LAS) 40

D Development of LM Combined Parameters 42

List of Tables

1 Model boundary conditions, mount system states, and 1st bending mode

frequency. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 15

D1 WatchDog system performance assessed using five hard landing cases from

Test 593. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 44

2
List of Figures

1 NASA Langley TDT tests for the HiLDA and AEI programs. . . . . . . . . . 6

2 Illustration of AEI apparatus installed in TDT during Test 598. . . . . . . . . 7

3 NASA Langley TDT. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9

4 Illustration of the TDT Air Stream Oscillating System (AOS). . . . . . . . . 10

5 Original, 3 DOF mount configuration used in Test 593. . . . . . . . . . . . . . 11

6 Revised, 2 DOF mount configuration used in Test 598. . . . . . . . . . . . . . 12

7 Photo of the HiLDA model taken from TDT settling chamber. . . . . . . . . 13

8 Wing instrumentation used for the AEI TDT tests (593 and 598). String

potentiometers used for measuring pitch angle (Θ), vertical position (PZ),

vertical rate (VZ), and moving mass position are not shown. . . . . . . . . . . 13

9 Signal routing used in Tests 593 and 598. . . . . . . . . . . . . . . . . . . . . 16

10 dSpace and snubber control block diagram. . . . . . . . . . . . . . . . . . . . 17

11 Snub control chassis containing the snub control system with manual Snub! switch

and hydraulic pressure displays. . . . . . . . . . . . . . . . . . . . . . . . . . . 19

12 Worst case overload condition from Test 593, point 1833. The bending moment is measured from strain gauge SBI1 shown in figure 8 corrected for

outboard wing weight. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 20

13 WatchDog system snub logic block diagram. . . . . . . . . . . . . . . . . . . . 20

14 Control law and FlightMode logic developed by Northrop Grumman. . . . . . 22

15 Data acquired while using TakeOff launch scheme to achieve trimmed flight.

Test 598, points 1274, 1275, and 1276. . . . . . . . . . . . . . . . . . . . . . . 23

16 Trim and suppression control loops along with FlightMode logic developed

by Lockheed Martin. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 25

17 Data acquired while using Release launch scheme to achieve trimmed flight.

Test 598, point 2384. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 26

18 Portion of the TDT heavy gas operating envelope showing data acquired

during Test 598 where red represents the heavy weight configuration and

green represents the light weight configuration. . . . . . . . . . . . . . . . . . 28

19 Time traces of vertical position, pitch angle, and bending moment while AOS

transitioning from 0.7 to 0.5 Hz. WatchDog Snub! engagement shown along

with subsequent vertical position ramp down. Test 598, point 4891. The

bending moment was measured from strain gauge SBI1 (figure 8) corrected

for outboard wing weight. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 31

20 Time traces of vertical position, pitch angle, and bending moment while AOS

sweeping from 9.5 to 2.0 to 9.5 Hz. Test 598, point 3182. The bending

moment is measured from strain gauge SBI1 (figure 8) corrected for outboard

wing weight. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 32

21 Body freedom flutter time traces for trim only control, light model. WatchDog

Snub! engagement shown along with subsequent vertical position ramp down.

Test 598, point 3525. The bending moment is measured from strain gauge

SBI1 (figure 8) corrected for outboard wing weight. . . . . . . . . . . . . . . . 33

A1 Flap transfer function estimates for beginning (solid), middle (dash), and end

(dot dash) of Test 598, points 216, 1600, and 5409. . . . . . . . . . . . . . . . 38

B1 Snubber system time delay experiments, hydraulics at 2000 psi. . . . . . . . . 39

C1 LAS and wing mounted load cell moving average data acquired during wing

vertical translation. Test 598, points 212 (blue) and 3237 (green). . . . . . . . 41

C2 LAS load cell data acquired during three different launch and flight sequences.

Test 598, points 3245, 3435, and 3455. (Forward load cell data is blue and

the aft load cell data is green). . . . . . . . . . . . . . . . . . . . . . . . . . . 41

D1 Lockheed Martin combined parameter block diagram. . . . . . . . . . . . . . 42

3
D2 Hard landing case plot of vertical position, pitch angle, TE2 command, and

wing root bending moment with Snub! command indicated. Test 593, point

1252. The bending moment is measured from strain gauge SBI1 shown in

figure 8 corrected for outboard wing weight. . . . . . . . . . . . . . . . . . . . 43

D3 Hard landing case comparing nominal WatchDog Snub! command with the

Snub! command generated by the LM combined parameters where two values

of assumed time delay were used in the feed forward estimator. Test 593, point

1252. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 44

4
Nomenclature

AEI Aerodynamic Efficiency Improvement

AFRL Air Force Research Laboratory

AOS airstream oscillating system

ASE aeroservoelasticity

BFF body freedom flutter

CG center of gravity

DAS Data Acquisition System

DCS Digital Control System

DOF degree of freedom

dSpace1 DCS for Servo PID loops and WatchDog

dSpace2 DCS for trim, GLA, & BFF suppression

δo emergency controller command bias

FS flutter suppression

GLA gust load alleviation

GUI Graphical User Interface

HiLDA High Lift over Drag Active Wing

LAS lift augmentation system

LE leading edge flap

LM Lockheed Martin

LQG Linear Quadratic Gaussian

NG Northrop Grumman

PID proportional integral derivative

PZ vertical position (PZ = 0 at centerline), in

RVDT rotary variable differential transducer

Snub! command to engage snubbing system

T time, s

TE trailing edge flap 1, 2, 3, or 4

TDT Transonic Dynamics Tunnel

Θ pitch displacement, deg

Θ pitch rate, deg/s  ̇

V Volts

VZ vertical velocity, in/s

WatchDog software for triggering Snub! command

WD abbreviation for WatchDog

WoW weight on wheels

5
HiLDA

T574: NASA

T579: Northrop Grumman

AEI

T593: Northrop Grumman

T598: Northrop Grumman &

 Lockheed Martin

T606: Boeing

T613: Boeing 

T614: Boeing 

FY04 FY05 FY06 FY08 FY07

Gust Sensor Flying Wing Joined Wing

FY03 FY09 FY10

Figure 1. NASA Langley TDT tests for the HiLDA and AEI programs.

1 Introduction

In the Fall of 2007, the Air Force Research Laboratory (AFRL), Northrop Grumman (NG),

Lockheed Martin (LM), and the NASA Langley Research Center, successfully completed

the third of a series of three wind tunnel tests of an aeroelastically scaled wind tunnel

model of a flying wing SensorCraft vehicle concept. The AFRL SensorCraft is a technology

portfolio of advanced sensors, communications links, air vehicle components and propulsion

elements. AFRL is pursuing these technology developments for future high-altitude, longendurance, unmanned surveillance platforms [1]. Two research programs intended to develop

technologies relevant to these large, flexible vehicles are the High Lift over Drag Active

(HiLDA) Wing and the Aerodynamic Efficiency Improvement (AEI) programs [2]. The

goals of the HiLDA and AEI wind tunnel tests included the demonstration of gust load

alleviation (GLA), an enabling technology for a SensorCraft vehicle that will allow it to

have reduced structural weight thereby increasing endurance, range, and payload capacity.

The HiLDA and AEI wind tunnel investigations were conducted in the NASA Langley

Transonic Dynamics Tunnel (TDT). Three flying wing wind tunnel tests and five joinedwing [3, 4] wind tunnel tests are depicted in figure 1. The first two wind tunnel tests were

conducted as part of the HiLDA program. The first was a short test of several candidate

sensors for measuring gust flow angle. The HiLDA flying wing was tested September 2004

with GLA being demonstrated on a cantilevered mount using the model’s five active control

surfaces [5–7]. This paper will focus on the flying wing tests conducted as part of the AEI

program, while the the joined wing tests will be covered in a separate report.

The HiLDA wing was retested in October 2006 as part of the AEI program, where it was

mated to a new, multi-degree of freedom (DOF) mount. This mount allowed the half span

model to translate vertically and pitch at the wing root, allowing better simulation of the

full span vehicle’s rigid body modes. Following some modifications to the mount system,

the wing was tested on the mount again in the summer of 2007. Figure 2 depicts the wing

and mount system in the final configuration used in Test 598. The upstream flow angle sensor or gust sniffer is also shown in the image. During this test, GLA control systems were

demonstrated in the presence of gusts generated by the TDT airstream oscillating system

(AOS) with peak wing bending loads being reduced by up to 60 percent, and body freedom flutter (BFF) suppression control laws were demonstrated with flutter onset dynamic

pressure being increased up to 50 percent.

6
Flow Oscillations (“Gusts”)

“Gust Sniffer”

SensorCraft Model

2 DOF Mount

Figure 2. Illustration of AEI apparatus installed in TDT during Test 598.

7
The large size of the flying wing model along with the rigid body DOF afforded by the

new mount system, created many unique challenges to successfully and safely meeting the

AEI test objectives. A remotely actuated snubbing system that returned the model to a

fixed pitch angle was included in the mount design. Determining how to effectively use this

system was critical to the success of the test. Two digital control systems were used: one

to implement the actuator servo control loops and the other to implement the trim, GLA,

and BFF suppression control loops. For model safety, a snubbing system was integrated

with mount system, and the digital control systems included a safety monitoring (WatchDog) system for automatically triggering the snubbing system. The WatchDog system was

implemented on the same digital control computer as the servo control loops. Along with

these systems, a variety of procedures had to be developed including those for model launch

and recovery.

This paper will provide an overview of the AEI flying wing wind tunnel tests from an

operational point of view with the focus on systems, procedures, and lessons learned. The

paper includes descriptions of the hardware including the wind tunnel, the wing model,

the mount system, and other supporting systems developed for this program. The software developed for the digital control systems will be described including the control law

architectures and the WatchDog system. The main part of the paper will conclude with

a description of the wind tunnel testing procedures that were used. An appendix provides

some additional analysis and testing details not covered in the main part of the paper.

Throughout the paper, wind tunnel data will be shown where applicable or necessary with

an emphasis on the final test of the flying wing, Test 598.

For the purpose of clarity, a brief overview of the model launch schemes, the flight control

architectures, and the WatchDog system is warranted. Two model launch schemes were

used, the TakeOff and the Release launch schemes developed for use by Northrop Grumman

and Lockheed Martin, respectively. Throughout this paper, the TakeOff launch scheme will

be associated with the NG control architecture where trim and suppression controllers are

combined into a single control loop. The Release launch scheme will be associated with

the LM control architecture where trim and suppression controllers are in separate control

loops. In addition, a separate set of WatchDog monitored parameters and associated limits

are used with each launch scheme. These topics will be discussed separately in the paper.

2 Transonic Dynamics Tunnel

The Langley TDT, depicted in figure 3, is a unique national facility dedicated to identifying,

understanding, and solving relevant aeroelastic and aeroservoelastic problems. The TDT

is a closed circuit, continuous flow, variable pressure, wind tunnel with a 16 ft square test

section with cropped corners [8]. The tunnel uses either air or a heavy gas as the test

medium and can operate at total pressures from near vacuum to atmospheric. It has a

Mach number range from near zero to 1.2 and is capable of maximum Reynolds numbers of

about 3 million per 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 TDT has

used 1,1,1,2 tetrafluoroethane(R-134a) [9, 10], an environmentally acceptable alternative to

R-12.

The TDT is specially configured for flutter testing, with excellent model visibility from

the control room and a rapid tunnel shutdown capability for model safety. Testing in heavy

gas has important advantages over testing in air: improved model to full scale similitude

(which results in heavier, easier to build models with lower elastic mode frequencies), higher

Reynolds numbers, and reduced tunnel power requirements.

Due to the need to perform GLA testing as part of the AEI program, the TDT AOS was

used frequently. Figure 4 shows the key features of the AOS system including the biplane

8
Figure 3. NASA Langley TDT.

arrangement of vanes on either side of the entrance to the test section. Each pair of vanes is

driven by a separate hydraulic motor and a flywheel to hold constant vane frequency. While

the two pairs of vanes can be run out of phase, this was not done in the AEI test. Vane

frequency was adjusted from the TDT control room manually or by using a LabVIEW
R

system. The amplitude of the vanes is manually adjustable from 0◦to 12◦ peak-to-peak.

For most of Test 598, the 12◦ peak-to-peak setting was used providing approximately ±1

◦

variation in the model angle of attack. For the last week of Test 598, the vane amplitude

was reduced to 4◦ peak-to-peak.

3 Wing and Mount System

The wing and mount system were designed and fabricated by NextGen Aeronautics in

Torrance, California. NextGen Aeronautics was a subcontractor to the Northrop Grumman

Corporation for the initial development of the wing for Test 579, and later, for the wing

modifications and mount system design and fabrication for Test 593. NextGen Aeronautics

was a subcontractor to the Lockheed Martin Company for the wing and mount modifications

that preceded Test 598. This section of the paper will provide a description of the wing and

mount system along with a discussion of the modifications that took place over the life of

the program. References 5 and 6 offer a more detailed description of the HiLDA wing, and

references 11 and 12 contain details associated with the wing and mount system.

Figure 5 shows the wing and mount system configuration used in Test 593, and figure 6

shows the configuration used in Test 598. These figures should be referred to when reading

the sections that follow.

9
Figure 4. Illustration of the TDT Air Stream Oscillating System (AOS).

3.1 Wing

The wing is a 12 percent length scale, semispan model of a SensorCraft design concept

originally designed for use in the HiLDA program. The outboard part of the wing is a

spar-pod design where the scaled stiffnesses (EI and GJ distributions) are designed into

the flanged aluminum spar, and the aerodynamic shape is provided by discrete fairings, or

pods, mounted to the spar. The original set of pods was made from glass-filled nylon-12

(Duraform GF) manufactured using a stereolithography system. The inboard wing was

intended to act as a rigid member and consisted of aluminum spars and internal ribs with

fiberglass upper and lower skins. The closeout ribs for the inboard wing were stainless

steel to provide the structural attachments for the 5 DOF balance at the root and for the

outboard, flexible spar at the wing break. Lead weights could be attached to the spar to

simulate fuel weight at the take off configuration.

The wing has four evenly spaced trailing edge control surfaces and one leading edge

control surface on the outboard, flexible portion of the wing. The five control surfaces

were driven by vane-type hydraulic actuators with position measured by rotary variable

differential transducers (RVDTs). Two separate hydraulic systems were included in the

model so that the trailing edge flaps could be operated at 1,000 psi, while the more heavily

loaded leading edge flap could be operated at up to 2,000 psi. Other instrumentation

included strain gauges, accelerometers, a pitch-rate gyro, and a gust vane placed in front

of the model to allow lead gust information to be fed into the controller. Figure 7 shows a

photo of the HiLDA wing installed in the TDT from the settling chamber.

In preparation for the first AEI test, TDT Test 593, several modifications were made to

the HiLDA wing. As a result of the new model mount system with additional DOF, overall

mass and center of gravity (CG) became important considerations. To adjust the model CG,

the LE caps of the inboard part of the wing were modified so that six blocks of lead could

be installed to provide up to 30 lb of forward ballast. In addition, a moving mass system

was included in the rigid root section that provided remote adjustment of the model CG.

The device consisted of a compact stepper motor coupled to a ball screw actuator driving a

32 lb ballast weight. The ballast assembly was restrained by two linear slides with a string

potentiometer for position measurement.

10
(a) Mount system components.

(b) Original fairing assembly with large dynamic

plate. Total vertical travel includes 14.2 in of free

travel plus the 3 in strokes of the upper and lower

shocks.

(c) Back side of pivot carriage showing snubbing

mechanism and umbilical.

Figure 5. Original, 3 DOF mount configuration used in Test 593.

Modifications to the HiLDA wing’s instrumentation suite included changes to the accelerometer locations to better measure the free-free mode shapes and the addition of a

pitch rate gyro near the wing root. Figure 8 shows the instrumentation layout used in both

AEI tests. Leading edge stagnation point sensor arrays developed by Tao Systems are not

depicted in this figure, but they are described in reference 13. To simplify the model’s hydraulic requirements, a larger leading edge actuator was built for this test so that it could

operate at the same hydraulic pressure as the trailing edge flaps.

11
(a) Revised mount system.

(b) Revised fairing assembly with

smaller window and dynamic plate.

Total vertical travel includes 23.4 in of

free travel plus the 3 in strokes of the

upper and lower shocks.

(c) Installed mount system showing umbilical service loop.

Figure 6. Revised, 2 DOF mount configuration used in Test 598.

The pods and control surfaces were redesigned to simplify installation and removal,

and they were fabricated using a different material (Watershed 11120) to avoid concerns

over moisture absorption. Unfortunately, the 11120 material has a low glass transition

temperature, and the pods warped due to elevated temperatures in transit to the TDT.

Changes to the wing in preparation for TDT Test 598 were limited largely to maintenance items; however, two important modifications were made. First, the pod material was

changed back to Duraform to avoid heat related warping. Duraform GF was used for the

new pods, and Duraform AF was used for the new control surfaces. To minimize moisture

absorption, all pod surfaces were sealed using primer. The second modification was deemed

necessary when, during inspection, it was determined that the inboard wing had been overloaded during Test 593. Wind tunnel data from this event will be shown later in the paper.

A steel strap was added to the upper and lower surfaces of the inboard wing to strengthen

the center spar by providing an additional load path between the two stainless steel ribs

(figure 6).

Appendix A describes the performance and challenges associated with using and maintaining the wing flap actuators during Test 598.

12
Flow

Oscillation

Vanes

Flying Wing

Sensorcraft

Model

“Gust 

Sniffer”

Figure 7. Photo of the HiLDA model taken from TDT settling chamber.

Strain Gage - Bending

Strain Gage - Tension

Strain Gage - Bending (Y-Axis)

Actuator

RVDT

Accelerometers

Servoactuator/Manifold

Pitch Gyro

CTL3

RVDT3

RVDT4

RVDT2

STM2

STI2

STY

STI2

SBI1

Spar - Cross Section

SBI2

NZ4 SBM1

SBM2

NZ5

STM1

RVDT1

CTL4

CTL5

CTL2

CTL1 NZ6

NZ7

RVDLE

NX2

NZ1

NZ2 NZ3 NX1

Hydraulic Lines

TE1 Wires

TE3

TE4

TE2

LE

Figure 8. Wing instrumentation used for the AEI TDT tests (593 and 598). String potentiometers used for measuring pitch angle (Θ), vertical position (PZ), vertical rate (VZ), and

moving mass position are not shown.

13
3.2 Mount System

The sidewall multi-DOF mount system evolved considerably over the course of the two TDT

tests in which it was used. Initially for Test 593, the mount was configured to provide 3 DOF;

pitch, plunge, and fore-aft translation (figure 5). Based on the operational experience gained

in Test 593, several significant changes were made to the mount system in preparation for

Test 598. These changes included the removal of the fore-aft DOF and a complete redesign

of the lift augmentation system. Figure 6 shows the revised system. The key elements of

the mount system along with the various design changes will be described next.

The pivot carriage assembly was the heart of the mount system. It consisted of the pivot

bearing and the hydraulic pitch snubbing mechanism. The snubber consisted of two linear

hydraulic actuators that moved a sliding plate (guillotine) that clamped against alternate

ends of the half-circle, free-to-pitch structure that attaches to the wing. Figure 5 depicts

the back side of the pivot carriage assembly showing the pitch snub components and the

hydraulic/instrumentation umbilical exiting the model through the center of the bearing

assembly. A string potentiometer was used to measure model pitch angle during both

tests, and another string potentiometer was added prior to Test 598 to provide a direct

measurement of the guillotine position. For the unsnubbed condition, the total rotation

was controlled by setting the retraction distance of the guillotine by adjusting the hydraulic

cylinder mount position. For the snubbed condition, replaceable pads on the free-to-move

structure allowed control of the clamped pitch angle. Test 593 had a range of motion of

−12◦to +16◦ with a clamped angle of 2◦, and test 598 had a −6

◦

to 8◦range of motion with

a clamped angle of 1◦. The range of motion was restricted to reduce the risk of damaging

the model. Appendix B describes experiments used to assess and improve the performance

of the snubbing system.

The rail adaptor and vertical rail assemblies provided the translational DOFs via THK

20 mm width rails and guide blocks with vertical and fore-aft position measured using string

potentiometers. To minimize impact loads, the vertical rail assembly included dashpot shock

absorbers at the top and bottom. During Test 593, Mil-Spec bungee cords, cables, and

pulleys were used to restrain the fore-aft DOF. It was determined that the 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 provided the fore-aft DOF, the rail adaptor,

was removed. The pivot carriage assembly was rotated 90◦ and attached directly to the

vertical rails (figure 6).

A fairing assembly enclosed the mount system. The fairing consisted of fixed exterior

fiberglass panels with a cutout or window that allowed for a range of travel of the wing

attachment. The window was covered with a dynamic plate that attached directly to and

moved with the pivot carriage assembly. Figure 5 shows the fairing assembly used in Test

593. The elimination of the fore-aft DOF for Test 598 allowed the width of the window

to be reduced from 23 in to 11 in (figure 6). As a result, the fixed fairing could be better

supported, eliminating contacts between the fixed fairing and the wing root and the fixed

fairing and dynamic plate. A smaller, two-piece dynamic plate was used in Test 598. The

two-piece design allowed the dynamic plate to be removed for access to the mount system

components without removing the wing.

A lift augmentation system (LAS) was required to compensate for the mass of the carriage system and excess inboard wing weight. The first iteration of the LAS consisted of

eight constant force springs that provided approximately 40 lb of force each for a total lift

augmentation of 320 lb. The springs could be attached to the rail adaptor in groups of

two via a lanyard and cable assembly. This first system suffered from high friction associated with the spring tapes twisting and rubbing against the spool frame assembly. The

redesigned system used in Test 598 again used eight constant force springs, but because

of the weight savings associated with the elimination of the fore-aft DOF and the removal

14
of the rail adaptor assembly, springs providing 20 lb of force each could be used. These

springs attached to the pivot carriage assembly via pinned ball joints with load cells in series so that the LAS force could be monitored during the test. Other design improvements

included spring spool assemblies with an improved set of bearings and a LAS frame design

that minimized potential spring tape contact points.

The original vertical rail assembly consisted of two independent components that bolted

directly to the TDT wall plates as shown in figure 5. As a result of irregularities in the

test section wall, shimming was necessary to avoid binding the plunge DOF. Also, the

hydraulic/instrumentation umbilical had to exit straight out of the assembly as shown in

figure 5. Because there was limited space between the pivot carriage assembly and the test

section wall, free vertical travel was restricted to ±7.1 in of the designed ±12 in to avoid

cutting the umbilical on the opening in the test section wall. The free travel does not include

the 3 in strokes of the upper and lower shocks, so total vertical travel is 6 in greater than the

free travel. The redesigned mount shown in figure 6 included a new standoff support frame

that simplified mount installation and provided an additional 4 in of clearance from the

wind tunnel wall. These changes allowed the hydraulic/instrumentation umbilical to form a

service loop inside the fairing, under the pivot carriage assembly, as shown in figure 6. For

the Test 598 configuration ±11.7 in of free vertical travel was achieved.

Appendix C describes the experimental methods used to estimate the friction in the

vertical rails and to monitor the health of the LAS.

3.3 Test Configurations

In Test 598, the model was tested in two configurations: heavy and light. The heavy

configuration had five of six inboard leading edge masses, all wing fuel weights, and all eight

LAS springs installed. The light configuration had no wing fuel masses, two of six inboard

leading edge masses, and four LAS springs installed. The total free flying weight of the

heavy and light models were 442 lb and 356 lb, respectively.

Table 1. Model boundary conditions, mount system states, and 1st bending mode frequency.

Bndry Cond Mount State 1st Mode, Hz

Pitch Plunge Pitch Plunge Heavy Light

Fixed Fixed Snub On Stop 1.9 2.1

Fixed Free Snub Trim 2.0 2.3

Free Fixed Trim On Stop 3.0 3.3

Free Free Trim Trim 3.5 4.1

The boundary conditions and hence, the dynamic characteristics of the wing were altered

considerably by the state of the mount system. For instance, when the snubbing system

was engaged and the pivot carriage was held on the lower stop via gravity, the wing was

essentially cantilevered at the root. The pitch DOF could be freed by releasing the snubber,

and the plunge DOF was only free when the model was in trimmed flight. Thus, the mount

system provided four distinct boundary conditions that had to be considered. Table 1 shows

the various boundary conditions, mount states, and first mode frequency. For the heavy

configuration, the frequencies were determined by a ground vibration test performed in the

TDT model preparation area, and for the light configuration, frequencies were determined

via NASTRAN analysis [12].

15
HiLDA Wing and AEI Mount

East Wall To Control Room Patch Panel

Gust Sniffer

Test Section

Control Room 

Gyros, Strain Gauges

Accelerometers,

String Potentiometers, &

Gust Sniffer

Data Room 

Servovalue Gust

Sniffer

Strain Gauges,

Accelerometers,

& StringPotentiometers

Rate

 Gyros

Gyro

Servovalue Commands

RVDT

RVDT

Flap Position Commands

RVDT

Control Room To Data Acquisition Patch Panel

Moog Voltage to

Current Converter

Custom +/-15 Volt 

Power Supplies

Pacific Instruments

Signal Conditioner/Filter

dSpace1

Servo Loop

WatchDog

Digital Strip Chart dSpace2

Trim

GLA / FSS

DAS

Figure 9. Signal routing used in Tests 593 and 598.

4 Signal Routing, Processing, and Control

Figure 9 shows a high level schematic of the signal routing arrangement used in Tests 593

and 598. All input and output signals were routed from their source to the TDT control

room via a built-in wiring system. Most instrumentation was routed to a Pacific Instruments

series 6000 chassis that provided instrument power (5 or 10 VDC), signal amplification, and

anti-alias filtering. The RVDTs and rate gyros required ±15 VDC power so a custom power

supply was fabricated with the signals subsequently routed through the Pacific Instruments

chassis for amplification and anti-alias filtering. The anti-aliasing filters were set to 400 Hz

for the RVDTs as they were routed only to a digital control system running at a 1,000 Hz

frame rate (dSpace1), and all other signals were filtered at 100 Hz to be compatible with

the Nyquist frequency for the other digital control system running at a 200 Hz frame rate

(dSpace2). Servovalve signals were routed to a Moog voltage-to-current converter and back

to the model as shown in the figure. Model signals were “Teed” and routed to a strip chart

for monitoring and to the TDT Data Acquisition System (DAS) for recording. The TDT

DAS was set to record data at 500 Hz with its own anti-aliasing filters set to 200 Hz.

Figure 10 depicts the snubber control system, dSpace1 and dSpace2 internal block diagrams, and the external signals connecting these systems. All signals external to the dSpace

blocks are analog, with dSpace input signals being converted from volts to engineering units

prior to processing and output signals being converted from engineering units back to volts.

For signals intended to depict discrete system states, programming logic was used to decode

the meaning of the signal.

The hydraulic actuators in the snubbing system were operated by a solenoid valve that

16
δposition

δCmd.

δCmd.

Model

Signals

WoW

Switch

Latching

Circuit

Emergency

Control Law GLA

or

FS

External Control Inputs

Manual Chicken Switch

FlightMode

Servo Control

PID

WatchDog

Snub

Logic

V Divider

Snubber Control System

To Solenoid Valve:

24V=Unsnubbed

OV=Snubbed

SnubStat:

Hi=Unsnubbed

Low=Snubbed

To Model:

Servovalves

dSpace2 dSpace1

Reset Switch Override 

Switch

Gnd

Snub!

Trim

or

Trim+GLA/FS

See NG or LM

Block Diagrams

(RVDTs)

Figure 10. dSpace and snubber control block diagram.

was controlled using the snubber control system. The heart of the snubber control system

was a latching circuit that could be tripped by a manual chicken switch or a Snub! command signal from dSpace1. The status of the latching circuit (snubbed or unsnubbed) was

communicated back to the dSpace systems by the SnubStat signal in figure 10. A detailed

description of the snubber control system and associated hydraulic components is described

in a separate section, below.

The trim, GLA, and/or BFF suppression control laws were implemented on dSpace2.

Figure 10 shows that dSpace2 had two internal control blocks and a set of externally generated flap commands that could be summed and sent out as analog flap command signals

to dSpace1. Throughout most of Test 598, only the upper dSpace2 control block was used

with two versions being implemented, one developed by NG and the other developed by LM.

This upper control block included a GUI interface and programming logic for controlling

or initiating certain events like resetting the system or initiating a launch. Details of the

NG and LM flight control architectures, including their associated programming logic, will

be described later. The FlightMode state is communicated from dSpace2 to the WatchDog

system as shown.

The servo-control loops, the WatchDog system, and the Weight on Wheels (WoW) switch

estimator were implemented on dSpace1. The servo-loops were independent PID control

loops equipped with output saturation blocks to prevent overdriving the actuators. The

WoW switch was a hysteresis block intended to provide debounce or noise tolerance. The

WatchDog system monitored the model signals, and when a fault was detected, it would

issue a Snub! command and transfer flap control to the emergency control law via the switch

shown in figure 10.

Details of the dSpace hardware and software development environment will be described

next, followed by a description of the hydraulics and snubber control system.

4.1 DCS Hardware and Development Environment

Each dSPACE DCS consists of a rack containing a host computer, a target system, a keyboard, a monitor, BNC patch panels for IO, and an uninterruptible power source. The heart

17
of the DCS is the target system that includes a dSPACE DS1006 control processor board

using a 2.6 GHz AMD Opteron processor connected to three dSPACE DS2002 multi-channel

A/D boards and one dSPACE DS2sec103 multi-channel D/A converter board. The A/D

boards each have 32 channels using 16 bit quantization with an input range of ±10 V. The

D/A board contains 32 channels of 14 quantization bits designed for ±10 V and a settling

time of 10 μs.

The controller software was developed within the MATLAB
R Simulink environment,

then compiled and downloaded to the target processor via the dSPACE and MATLAB
R

Real-Time Interface. An integral component of the dSPACE tools is the ControlDesk application which provides the user interface to the target processor for the development and

implementation of the GUI. The host computer runs the GUI and controls all communications between the processors.

4.2 Hydraulics and Snubber Control System

Hydraulic control of the pitch snub mechanism was accomplished using a custom snub

control system housed in a chassis located in the TDT control room and a hydraulic manifold

assembly located in the TDT plenum near the wind tunnel model and mount. Hydraulic

fluid, at 2,000 psi, was delivered to the hydraulic manifold assembly from the TDT 30 GPM

pump. The manifold assembly consisted of an adjustable pressure switch, a Parker Hannifin

solenoid valve, and two regulators for supplying up to two lower pressures to the wind tunnel

model. Hydraulic pressure transducers were included to monitor the supply/snub pressure

as well as the two regulated pressures. The solenoid valve was normally open and required

24 VDC power to remove pressure from the snubbing system thereby releasing the model.

The snub control chassis served two purposes, monitoring of the hydraulic systems and

enclosing the snub control system. The chassis housed three hydraulic pressure displays

showing supply/snub pressure and up to two lower, regulated model pressures. An amber

light and an audible alarm were wired to the pressure switch to provide warning of reduced

hydraulic pressure. The snub control system consisted of two mechanically latching relays

that were used to send 24 VDC to the hydraulic solenoid valve. The relays had their contacts

wired in series so that either relay could cut the 24 VDC power to the hydraulic solenoid

and independently snub the model. One relay was switched directly by the normally open

hand-held chicken switch, and the second relay was switched by a voltage signal from an

external source.

Figure 11 shows the snub control chassis and a handheld chicken switch. Two amber

colored lights on the front of the snub control box indicated which source initiated the

Snub! command, either the chicken switch or the external source (WatchDog system). Resetting of the relays and lights was accomplished with a momentary toggle switch (reset

switch). There was an independent toggle (override) switch to hold a snub condition which

was used to keep the model in a snubbed condition independent of the status of the relays

allowing systems to be reset without inadvertently cycling the snub system.

5 WatchDog System

Due to the high risk associated with aeroelastic wind tunnel testing, a variety of manual

and automated safety systems have been used in the TDT. The manually operated tunnel

bypass valves are generally the first line of defense for flutter testing as they can rapidly

reduce test section dynamic pressure and Mach number. Depending on the type of model

and mount system employed, model stabilization or arrestment mechanisms have also been

considered. Previously employed model stabilization devices have included decoupler pylons

that change model dynamics to a more benign configuration and model arrestment devices

18
Figure 11. Snub control chassis containing the snub control system with manual Snub! switch

and hydraulic pressure displays.

including pneumatic snubber cables for cable mounted models. Manual engagement of

these devices has, at times, been supplemented by automated systems. Such systems have

previously been employed only for non flying models with the tunnel bypass valves and/or

decoupler devices being triggered based on threshold exceedences [14].

The range of rigid body motion and the potential for high speed impact afforded by the

new mount system made the AEI flying wing tests among the riskier tests conducted in

the TDT. The snubber mechanism was intended to reduce risk to the model and facility,

and could be triggered manually. However, human reaction times for “simple” tasks like

the sheep dash game in reference 15 are only around 0.22 s, and more advanced tasks, like

dodging a baseball, requires a full 0.4 s [16, 17].

In the tunnel testing environment, the test engineer must in essence perform system

identification in real time to determine whether a behavior is a benign oscillation or a

potentially fatal divergence, mix in a little hope, and the reaction time can be significant.

This point is driven home by the worst overload case from Test 593 where the automated

system was not used. Figure 12 shows the time histories for vertical position, pitch angle,

and inboard bending moment acquired during a failed model launch attempt. For this data

point, the peak bending moment significantly exceeded the design load of 15,000 in-lbs with

the peak load preceding the manual Snub! command. Clearly, a robust automated system

was necessary, so the WatchDog system was developed.

The WatchDog system was developed to keep the model and mount system from exceeding structural safety limits by monitoring signals from the wing and mount system.

Difficulty in updating limits and a large number of false positives limited WatchDog use in

Test 593. By Test 598, the need for an improved WatchDog system was well established

and the deficiencies of the previous implementation were corrected. The system was also

expanded to satisfying the differing needs of both Northrop Grumman and Lockheed Martin.

The WatchDog was implemented on dSpace1 as shown in figure 10. The key features

of the WatchDog system are the emergency control law and the snub logic. The snub logic

monitors the model signals and issues a Snub! command when a fault is detected. Issuing the

Snub! command engages the pitch snubber and switches model flap control from dSpace2

to the emergency control law. The subsections that follow will describe the emergency

19
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

deg or in

x 104

Root Bending Moment, in−lbs

Θ, deg

PZ, in

Snub!

Time, s

Figure 12. Worst case overload condition from Test 593, point 1833. The bending moment

is measured from strain gauge SBI1 shown in figure 8 corrected for outboard wing weight.

FlightMode

Select

NG or LM

Model

Signals

Limit

Check

Latch

Debounce

NG Limits

WD Enable

==ResetValue (0)

NG Signals

LM Signals

LM Limits

Snub!

GUI

Display

Figure 13. WatchDog system snub logic block diagram.

controller, the snub logic, and some enhancements and unique features that were developed

for use by Northrop Grumman and Lockheed Martin for Test 598.

5.1 Emergency Controller

When the model is snubbed, the aerodynamic lift becomes a function only of the dynamic

pressure and control surface position, as the model is held at constant angle of attack by the

snubbing mechanism. In the snubbed condition, the flaps can be used as lift-only devices.

As a result, a simple emergency controller could be developed to arrest the vertical motion

and safely land the model after the Snub! command had been issued.

The emergency controller flew the model in the snubbed condition using vertical position

error and vertical velocity feedback. The controller gains were divided by the measured

dynamic pressure to produce a controller that worked over the dynamic pressure range of

interest. The controller also had a flap bias term (δo) consisting of the expected control

surface positions that would provide total wing lift equal to net wing weight which was the

actual wing and pivot carriage weight less LAS force. These bias values were also a function

of dynamic pressure, and were initially determined by simulation and then updated as part

of the testing process. This testing process will be described later in the paper.

20
5.2 WatchDog Snub Logic

Figure 13 shows a block diagram of the WatchDog snub logic. The system’s primary function

was to monitor the 32 model signals and issue the Snub! command when parameter limits

had been exceeded. The WatchDog could be enabled or disabled by the user from the

dSpace1 GUI. While rarely disabled during Test 598, the feature was needed to facilitate

testing under unique circumstances.

From the dSpace1 GUI, the operator can select either the NG or LM limits and signals to

be used with the NG or LM control law architectures, respectively. Both NG and LM signal

blocks included the basic set of 32 model signals along with some calculated values called

combined parameters. The NG and LM limit sets each included a discrete set of upper and

lower parameter limits for each FlightMode. The FlightMode was used to select between the

sets of limits, so that a different set could be used for reset, takeoff, flight, or GLA testing.

The exact definition of the FlightMode parameter was dependent on which flight control

architecture was being used, NG or LM. In either case, when the FlightMode associated

with model launch and climb-out was used, the lower model position limits needed to be

ignored until the model cleared a certain height (PZ value). NG accomplished this by having

a separate FlightMode and WatchDog limit set for take off and climb out. LM did not use

a separate FlightMode and WatchDog limit set, but instead had some coded logic built into

the LM Signals block that ignored the lower position parameters until a certain PZ value

had been reached. A more detailed description of these processes will be provided in the

flight control schemes section of the paper.

The selected signals and FlightMode dependent limits were compared in the Limit Check

block in figure 13. Detected faults were passed to the Debounce block where three consecutive frames of any particular signal fault were required to issue the WatchDog Snub! command. This prevented a single frame noise event from issuing a false alarm. A latch was

used to quickly diagnose the cause of WatchDog-commanded Snub! events. The parameter

that caused the trip would be held and displayed to the user via the GUI on dSpace1. The

WatchDog latch was reset by setting the FlightMode to zero.

Preliminary modeling and review of past tests allowed for initial values of the WatchDog

limits to be set. As the wind tunnel test progressed, some limits were expanded based on

operational experience. For example, as lower frequency gusts were tested, vertical velocity

and pitch rate became large exceeding initial WatchDog limits, but since model response

remained stable and bounded, the WatchDog limits were expanded to permit testing. During

testing, the limit values could be modified directly in the real time processor memory from

the dSpace1 user interface. These modifications were then recorded back into MATLAB
R

script files for future use.

5.3 NG—Combined Parameters

Based on the WatchDog system deficiencies identified during Test 593, Northrop Grumman

proposed that combined parameters be considered in addition to nominal set of 32 model

signals. For example, if the model had negative vertical position (below tunnel centerline)

and positive vertical velocity (moving up), that would be acceptable. Alternatively, the

combination of negative position and negative velocity would be cause for concern. Thus,

multiplying vertical position (PZ) and vertical velocity(VZ) produced a parameter that was

acceptable when negative and given a safe upper limit and monitored by the WatchDog,

i.e. −∞ < PZ × VZ < limit. An additional parameter of PZ × VZ2 × sign(VZ) was

also used and was a good indicator of dangerous conditions. These combined parameters

increased model safety and provided relatively few false alarms. The NG combined parameters were implemented within the NG Signal block in figure 13. The NG combined

parameter limits were set initially by simulation and then modified throughout the test as

21
dSpace

Generated

Excitation

Model Dynamics

Servoloops

WatchDog

Model

+

dSpace2: Upper Control Block

Sensors

SnubStat

Θ-Controller: Ramp or Hold Θ

or

Z-Controller: Ramp or Hold PZ

WoW

δCmd.

dSpace1

Model + Controller

GUI Input:

SFMode

FlightMode

Stateflow Logic Block

Figure 14. Control law and FlightMode logic developed by Northrop Grumman.

deemed necessary by the test team.

5.4 LM—Combined Parameters

A different set of combined parameters was developed by Lockheed Martin based on an

estimate of the vertical location up and down where the emergency controller can bring

the model to a stop once the snubber had been engaged for a given vertical speed. These

two combined parameters were generated by a feedforward estimator in series with a neural

network implemented within the LM Signal block in figure 13. Appendix D provides a

detailed description of the simulation model, neural network, and Test 593 data analysis

used to develop and validate the LM combined parameters.

6 Flight Control Schemes

Two schemes for achieving trimmed flight were developed and demonstrated: the TakeOff

and the Release schemes developed by Northrop Grumman and Lockheed Martin, respectively, with the assistance of NASA Langley Research Center. The TakeOff scheme simulates

a take-off roll, acceleration, rotation, and lift-off. Alternatively, the Release scheme kept the

model snubbed until flight speed was achieved in the tunnel, then it was unsnubbed and

lifted off. This section will describe each launch scheme and its associated control law

architecture. Example wind tunnel data will be provided where appropriate.

6.1 Takeoff Launch Scheme

The TakeOff launch scheme approximates an aircraft takeoff process beginning with a high

speed ground roll, followed by rotation, lift-off, climb, and finally, trimmed flight. Figure 14

shows the control law architecture associated with the TakeOff launch scheme. As shown

in the figure, the TakeOff launch scheme makes use of the MathWorks Stateflow package.

Stateflow extends Simulink with a design environment for developing state machines and flow

charts. The NG implementation of Stateflow allows most TakeOff launch scheme actions to

be event-driven. The Stateflow Logic Block is controlled by the Stateflow Mode (SFMode)

parameter via the dSpace2 GUI as indicated.

22
20

10

0

-10

-20

in or deg

60

50

40

30

20

Dynamic Pressure, psf

0 10 20 30 40 50 60 70 80 90

Time, s

SFMode: 0 Taxi

WoW On WoW Off

Snubbed

Fly

Free

Free

Fly

Θ-Controller: Hold Rotate

Unsnubbed

Z-Controller: Climb

Pitch: Fixed

Plunge: Fixed

Free

Fixed

Hold

User Commanded Event Driven

PZ, in

Θ, deg

TE3POS, deg

 FltMode/WD: 0

Cntrl: Open Loop

Taxi T-O/Climb

Figure 15. Data acquired while using TakeOff launch scheme to achieve trimmed flight.

Test 598, points 1274, 1275, and 1276.

The TakeOff launch scheme requires two separate controllers. A Θ-controller is used during the pitch-free condition when the model is on the lower vertical stop, and a Z-controller

is used when the model takes off and transitions to the free-free boundary condition. These

control laws were developed using the LQG method and incorporated GLA and trim control into a single controller. This allowed studies to be performed that directly traded short

period performance with GLA performance. Reference 18 discusses the design of these

controllers and their initialization methods in more detail.

Figure 15 shows data acquired during implementation of the TakeOff launch scheme.

Here, time histories of PZ, Θ, TE3 position, and dynamic pressure are plotted in the upper

part of the figure, and the timing of boundary condition changes, Stateflow modes, and

other system states are identified in the lower portion of the figure. User- and event-driven

actions are identified. This figure should be consulted when reading the following TakeOff

launch scheme steps:

T = 0 s: Start with model snubbed, dynamic pressure set to 20 psf, and SFMode set to 0.

Stateflow sets FlightMode to 0 telling the WatchDog to use the reset limits.

T = 13 s: The operator sets SFMode to Taxi. Controller is ready and waiting to detect the

unsnubbing of the model via the SnubStat signal from the snubber control system.

T = 15 s: The operator unsnubs the model using the reset switch on the snubber control

system chassis. Stateflow senses the unsnub condition via the SnubStat signal and

initiates the Θ-controller holding a constant 1◦set point. It also changes FlightMode

23
to Taxi telling the WatchDog system to use the Taxi limits.

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

T = 50 s: The operator sets SFMode to Fly. Stateflow then starts the Θ set-point ramp,

i.e. rotation, and changes FlightMode to T-O/Climb telling the WatchDog system to

use the T-O/Climb limits.

T = 58 s: The model lifts off the lower stop and passes through the WoW switch causing

Stateflow to transition from the Θ-controller to the Z-controller. The PZ set point for

the Z-controller is ramped up causing the model to climb. Also, as the WoW switch

is cleared, the model boundary conditions transition to free-free.

T = 83 s: Stateflow senses that PZ has reached a value of −1 in and changes the PZ set

point to the tunnel centerline value of 0. Stateflow also changes FlightMode to Fly

telling the WatchDog system to use the Fly limits.

T > 90 s: Dynamic pressure is increased, as necessary, to the desired test condition. Operator can set SFMode to Test. Stateflow then changes FlightMode to Test which tells

the WatchDog system to use the Test limits. When testing is complete, the snubber

chicken switch is used to manually engage the snubbing system and land the model

using the emergency control law.

This method was shown to be successful and led to a better understanding of the model

in early testing. However, the large dynamic pressure changes that were required for each

launch made it very time consuming. In addition, the pitch-free boundary condition is not

very stable and is significantly more difficult to control than the free-free boundary condition.

As a result, engineering effort had to be used to develop stable Θ-controllers. Also, the

transition from the Θ-controller to the Z-controller must have a small enough transient not

to overly perturb the model. Thus, the Z-controller bandwidth and performance was being

driven by the need to mitigate the takeoff transition. While this challenge is understood

and manageable in conventional flight control architectures, it presents particular challenges

when investigating large state-space controllers.

6.2 Release Launch Scheme

For the Release launch scheme, the control surfaces were prepositioned for flight prior to

releasing the pitch snub mechanism so that when the snubber was released the model flies

off the bottom stop. The model transitions from the stable fixed-fixed boundary condition

to the stable free-free conditions without dwelling at the unstable pitch-free condition. This

quick transition does not allow the unstable pitch-free configuration to build up any response

amplitude. Therefore, there was no need to control the pitch-free configuration using a Θcontroller.

The control architecture used with the Release method was a two-loop system as shown

in figure 16. The inner loop contained the trim controller and was responsible for flying

the model, and the outer loop contained the suppression controller and was responsible for

reducing the dynamic response. The suppression controller could be switched on and off

with little or no effect on the average control surface trim position. As a result, switching

between controllers caused little trim/elevation change and was done frequently throughout

the test to acquire GLA data with as many controllers as possible without having to repeat

the launch procedure.

The trim controller consisted of vertical position PID gains plus a pitch rate gain and

a static (bias) term. The proportional and integral terms were ignored until the model

was unsnubbed. The output of the trim controller was a single command that was sent to

all four trailing edge flap surfaces. The trim controller gains were initially established from

analysis, and a single set of trim controller gains were identified that were applicable to both

the heavy and light model. The bias term and the dynamic trim gains could be adjusted

24
LabView

Generated

Excitations

+

dSpace2: Upper Control Block

Snub!

δCmd.

Trim: T10 or T11

Ramp or Hold PZ

FlightMode

FlightMode Logic Block

Model Dynamics

Servoloops

WatchDog

Model Sensors

dSpace1

Model + Trim

GUI Input:

 FlightMode

 Suppression On/Off

Suppression

Figure 16. Trim and suppression control loops along with FlightMode logic developed by

Lockheed Martin.

from the dSpace2 GUI. Two versions of the trim controller were used during Test 598; T10

and T11. Trim controller T11 was better suited to TDT testing with the AOS system

operating.

The suppression controllers were developed utilizing the system identification and

LQR/LQG techniques described in reference 19. System identification data were acquired

around the inner loop, as indicated by the lines in figure 16, providing the model and model

plus trim controller system dynamics. These system identification data were then used with

the LQR/LQG methods to build the suppression controller. Therefore, the suppression controllers were aware of and expected the trim controller to be part of the system. Reference 20

covers this subject in more detail.

As with the TakeOff launch scheme, the WatchDog limits associated with the Release

launch scheme had to be set to ignore the lower position limits during the launch and

climb out phase of flight. For the TakeOff launch scheme, the logic for ignoring lower

vertical limits was contained entirely within the Stateflow logic block, and a separate set

of WatchDog limits was selected for this phase of flight. For the Release launch scheme, a

separate FlightMode for launch and climb out was not used. Instead, logic was built into the

LM Signal block shown in figure 13 that ignored certain parameters until PZ had reached a

value of −7 in. Another difference between the two launch schemes is that in the TakeOff

launch scheme the operator sets the SFMode parameter via the GUI, but FlightMode is

event-driven via the Stateflow logic block. When using the Release launch scheme, the

operator sets the FlightMode directly via the GUI.

For the Release launch scheme, FlightMode 0 was a reset mode, FlightMode 1 was a

trim control only mode, and FlightMode 2 allowed the use of the suppression loop. Each

FlightMode had a different set of WatchDog limits. FlightMode 1 was used for launch

and climb out, so parameters that would snub the model when it was on the lower stop

were ignored until PZ reached a value of −7 in. To protect the model during takeoff, in

FlightMode 1 the WatchDog Θ lower limit was set to 0◦

to minimize the possibility of a

pitch over and subsequent hard landing event.

Figure 17 shows data acquired during implementation of the Release launch scheme.

Here, time histories of PZ, Θ, TE3 position, and tunnel dynamic pressure are plotted in

25
in or deg

Dynamic Pressure, psf

0 5 10 15 20 25 30 35 40

Time, 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, in

Θ, deg

TE3POS, deg

 WD: Flight Limits 2

Lwr Prms Disabled All Paramerters Enabled

 WD: 0

Figure 17. Data acquired while using Release launch scheme to achieve trimmed flight. Test

598, point 2384.

26
the upper part of the figure, and the timing of boundary condition changes, FlightMode

changes, and other system states are identified in the lower portion of the figure. Userand event-driven actions are identified. This figure should be consulted when reading the

following Release launch scheme steps:

T = 0 s: The model is snubbed. Tunnel dynamic pressure is set to the launch value of

50 or 60 psf. FlightMode is set to zero by the operator causing the control surfaces

to be pre-positioned by the static part of the trim controller (δo). This also puts the

WatchDog in reset mode (0) which resets the latching logic associated with the lower

position limits in WatchDog limit set 1.

T = 5.5 s: The operator unsnubs the model using the reset switch on snubber control system

chassis. The FlightMode logic block within the LM flight control architecture senses

the unsnubbed condition via the Snub! command signal and initiates the dynamic

part of the trim controller. The PZ set point for the trim controller begins ramping to

tunnel centerline. The model starts to rotate but does not immediately lift off. This

is an unstable condition, and the model starts to oscillate in pitch.

T = 7 s: The model lifts off the bottom hard stop but is supported on the extended

shocks. The integral part of the trim controller continues to increase the control

surface deflection and Θ.

T = 12 s: The model clears the WoW switch indicating that it is off the shocks and

has transitioned to the free-free boundary condition. This event does not trigger any

action within the Release scheme and associated LM control architecture.

T = 13 s: The vertical position passes through −7 in and the WatchDog begins monitoring

all signals including the lower vertical limits.

T = 17.5 s: The PZ set point for the trim controller reaches and holds the tunnel centerline

value (PZ = 0).

T = 27 s: FlightMode is manually changed to FlightMode 2 which allows the use of the

suppression control law but does not close the suppression loop.

T = 31 s: A suppression controller is engaged by the operator via the dSpace2 GUI. The

higher frequency control surface activity is associated with the suppression loop being

closed.

T>40 s: Dynamic pressure is increased, as necessary, to the desired test condition. When

testing is complete, the snubber chicken switch is used to manually engage the snubbing system and land the model using the emergency control law.

The Release launch scheme proved to be easy, effective, and insensitive to dynamic

pressure and model configuration.

7 Testing Procedures

With the exception of several air checkout runs, all Test 598 data was acquired in heavy gas

with a tunnel total pressure between 1,670 and 1,880 psf. This maps to a fairly tight band

of Mach number and dynamic pressure combinations as shown in figure 18 where a portion

of the TDT heavy gas operating envelope is shown along with the Test 598 data points. At

any given tunnel total pressure, Mach number and dynamic pressure are not independent of

each other and vary in unison with changes in the RPM of the TDT main drive motor. For

the test procedure of Test 598, increasing RPM varied tunnel conditions along the plotted

data points starting with the lower left. As Mach number is relatively insensitive to tunnel

RPM in this region, dynamic pressure was used to identify all tunnel conditions.

As previously mentioned, the model and mount system had four testable boundary

conditions. Data presented in table 1 shows that the first bending mode frequency changed

significantly based on these conditions, and as a result, model flutter speeds varied with

27
Mach Number, M

Dynamic Pressure, PSF

20

30

40

50

60

70

80

90

200

H=50

PSF

75

100

150

200

300

400

500

600

700

800

10

900

00

1200

1400

1600

1800

2202000

0

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

Figure 18. Portion of the TDT heavy gas operating envelope showing data acquired during

Test 598 where red represents the heavy weight configuration and green represents the light

weight configuration.

28
snub and flight status. Due to the possibility for high speed vertical impact, the riskiest

configuration was with the model in a free-free, flying configuration. The safest configuration

was with the pitch DOF snubbed and the model sitting on the lower stop. For both the

heavy and light configurations, the relatively safe configurations were cleared for flutter

and the functionality of the emergency controller verified prior to proceeding with free-free

flight at a given dynamic pressure. The subsections that follow will describe the various test

procedures in general order of increasing complexity.

7.1 Open-Loop Flutter Clearance

The model and mount system could be tested open-loop when the pivot carriage was sitting

on the lower mount with and without the pitch snub mechanism engaged. Flutter clearance

testing of these configurations was performed in a manner typical of TDT flutter testing.

Dynamic pressure was increased slowly while the model was monitored visually and its

signals monitored via strip chart.

For the fixed-fixed case, flutter clearance was somewhat pro forma, as this configuration

had the least modeling unknowns and the highest analytical flutter speeds [5]. Also, it

had previously been tested in this configuration (Test 579 and Test 593) up to a dynamic

pressure of at least 80 psf. During Test 598, the fixed-fixed configuration was only cleared

for flutter to a dynamic pressure of 60 psf. The high degree of confidence in the analysis

and the concern that the model would not remain on the lower stop during this testing due

to the fixed, positive, snub pitch angle precluded flutter clearance testing above 60 psf.

The other open-loop configuration that was tested for flutter was pitch-free with the

model on the lower stop. This configuration required that the model be manually trimmed

to keep pitch angle from getting too large causing the model to lift off the lower stop.

This configuration was of interest because it was the model boundary condition state for

the Taxi mode in the TakeOff launch scheme. The exact flutter onset dynamic pressure

was somewhat ambiguous for this configuration as this flutter mechanism appeared to be

a hump mode with a very shallow flutter crossing. Experimentally, this was demonstrated

by an inconsistent flutter dynamic pressure and a relatively benign flutter mechanism. This

behavior occurred at dynamic pressures between 30 and 50 psf.

7.2 Snubbed Flight Clearance and Emergency Controller Validation

The next stop on the way toward achieving trimmed, free-free flight was to verify and tune

the emergency controller. The purpose of the emergency controller was to gently return

the model to the lower stop using gains and biases scheduled with dynamic pressure. The

bias values had to be adjusted based on experimental data. An additional purpose of this

testing was to simultaneously verify that the model was flutter free in the snubbed flying

configuration up to the maximum expected dynamic pressure of interest.

The process for clearing the model for flutter was to fly the model in the snubbed configuration at tunnel center line (PZ = 0) and incrementally increase dynamic pressure stopping approximately every 10 psf. Once aeroelastic stability had been established, controller

validation data was acquired using the emergency controller to track a sawtooth position

command between −10 in and +10 in with a ramp rate of ±1 in/s. Data was acquired

for three complete saw tooth cycles. The average control surface position needed to fly the

model at each dynamic pressure was determined and used as the static component (δo) of

the emergency controller.

In preparation for GLA testing this clearance/validation process was performed from

the takeoff dynamic pressure of 50 psf to 70 psf, the maximum dynamic pressure where

GLA data would be acquired. In preparation for BFF testing, the process was repeated for

29
dynamic pressures above 70 psf to a maximum dynamic pressure of 130 psf. The testing

process had to be repeated for the light model.

7.3 Gust Load Alleviation Testing

So far, this paper has described the apparatus, procedures, and preliminary testing that

were necessary to achieve trimmed, free-free flight in a safe and predictable manner. Demonstrating that a control law reduced loads due to gusts required that data be acquired using

different control laws all in the same gust environment. Here, the acquisition of the GLA

data with the TDT AOS operating will be discussed.

The biggest problem encountered with AOS testing was simply to get the model in

a trimmed flying state with the AOS operating. Initial attempts failed 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 vane frequency, and therefore, changes in

frequency occur relatively slowly. As a result, when the AOS system was engaged, the model

encountered a low frequency sinusoidal 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 to fly through this gust field exceeded the vertical travel available on the mount

system, and the WatchDog system or test engineer would snub the model. The solution to

this problem was to bring the AOS on-line prior to launching the model. An AOS frequency

well above the rigid body modes and between the resonant frequencies of the flexible modes

(firstin-plane and second out-of-plane bending) was used, typically 6.5 Hz. This technique

allowed both the TakeOff and Release launch schemes to be successfully used.

An interesting aside regarding the AOS is that on several occasions launch sequences

that had previously been successfully employed failed to work. In these instances, the model

would fly to the top of its vertical travel and trip the WatchDog system. It was determined

that the AOS vanes had been parked at a nonzero angle, and this flow angularity change

was enough to cause the model to fly to its upper travel limit during launch attempts.

Re-zeroing the AOS vanes solved the problem.

For the Release scheme, it was experimentally determined that trim controller T10 could

only be used during model launch with the AOS off. Thus, the trim controller gains had to

be adjusted slightly in order to successfully launch the model when the AOS was operating.

Two versions of the Release scheme trim controller were ultimately used, the AOS and nonAOS versions or T11 and T10 trim controllers, respectively. The higher gains and initial

control surface position in T11 provided a snappier launch and better position tracking.

Although T11 could be used to launch the model with and without the AOS operating, this

was generally not done as T10 was the preferred baseline controller for the AOS off cases.

In the case of the TakeOff scheme, the same control laws could be used to launch with and

without the AOS operating. Since trim and GLA were combined for the TakeOff scheme,

the inherent disturbance rejection associated with GLA may have allowed these control laws

to achieve this without any special AOS related design considerations.

At this point, the differences between the control law design methods associated with

the TakeOff and Release schemes dictated how the testing for each was conducted. For

the TakeOff scheme, the trim/GLA controllers were designed using state-space analytical

models. LQG methods were employed where the weightings were varied to produce different

controllers that provided various levels of GLA performance, and many such controllers were

evaluated. The GLA evaluation procedure consisted of acquiring model response data for

each control law subject to AOS dwells in 0.5 to 0.1 Hz frequency increments. Generally,

this type of testing would start at a relatively high AOS frequency and proceed to lower

frequencies until the WatchDog or test engineer snubbed the model. When the model was

snubbed, the AOS operator would return to the system to 6.5 Hz to reduce the large amplitude wing bending oscillations associated with peak response frequencies and to prepare

30
0 2 4 6 8 10

10

0

-10deg or in

6000

4000

2000

0

Root Bending Moment, in−lbs

Θ, deg

PZ, in

Snub!

Time, s

Figure 19. Time traces of vertical position, pitch angle, and bending moment while AOS

transitioning from 0.7 to 0.5 Hz. WatchDog Snub! engagement shown along with subsequent

vertical position ramp down. Test 598, point 4891. The bending moment was measured

from strain gauge SBI1 (figure 8) corrected for outboard wing weight.

for another launch using a different control law. Figure 19 provides some example data

where the AOS frequency was about 0.6Hz. Here, the vertical displacement is quite large

approaching ±10 in. At T = 4.9 s, a WatchDog limit was exceeded and the model was

snubbed and returned to the lower stop by the emergency controller.

When using the Release launch scheme and associated LM controller architecture, a

system ID approach was used to generate the GLA controllers. The system ID data was

acquired with uncorrelated random excitations sent simultaneously to each flap. These

excitations were generated externally using a LabVIEW
Rsystem with the excitations added

to the flap commands within dSpace2 as shown in figure 16. As discussed earlier, the system

ID data needed to capture the model with the appropriate trim control loop engaged.

Acquisition of system ID data for use when the AOS was not engaged was relatively

straight forward. However, a new procedure had to be developed to generate system ID

data for generating GLA controllers for use with the AOS engaged. Here, the external

frequency command feature of AOS was used with the LabVIEW
Rsystem to sweep through

the frequency range of interest. Typically, this was from an upper frequency of 9.5 to 11

Hz down to a frequency slightly above where the model was expected to exceed WatchDog

limits then back to the upper frequency. The random excitations were simultaneously sent

to the flaps while the AOS system was sweeping.

Suppression controllers generated using these data could be evaluated using the AOS

dwells, described earlier, or the AOS sweep excitation without the random flap inputs

included. Figure 20 shows time traces acquired during an AOS sweep. Here, the AOS

frequency is varied from 9.5 to 2.0 back to 9.5 Hz during a 100 second data record. Large

bending moment and plunge displacement responses were obtained at the low AOS frequencies.

The lowest frequency that could be obtained using either the AOS dwell or sweep excitations was a function of dynamic pressure. For the heavy model and a dynamic pressure

of 70 psf, the WatchDog system would generally trip between 3.0 and 2.5 Hz, and at 60 psf

31
5

0

-5deg or in

10

5

0

AOS Frequency, Hz

0 10 20 30 40 50 60 70 80 90 100

4000

2000

0

-2000

Bending Moment, in-lb

Time, s

Θ, deg

PZ, in

Figure 20. Time traces of vertical position, pitch angle, and bending moment while AOS

sweeping from 9.5 to 2.0 to 9.5 Hz. Test 598, point 3182. The bending moment is measured

from strain gauge SBI1 (figure 8) corrected for outboard wing weight.

it would typically trip between 1.7 and 2.5 Hz. The actual lower limit was a function of

the specific controller engaged at the time. Similar trends were noted for the model in the

light configuration. For the final week of testing, the peak-to-peak amplitude of the AOS

vanes was reduced from 12◦to 4◦. Following this change, data could be acquired with AOS

frequencies below 1.0 Hz for dynamic pressures of 60 psf as previously shown in figure 19.

References 18 and 20 show more details on the GLA control laws and associated experimental results.

7.4 Reduced Static Margin Testing

The movable mass could be remotely adjusted from the TDT control room during wind

tunnel testing. For model launch and most subsequent testing, the nominal location of this

mass was full forward for the heavy configuration and full aft for the light configuration. In

the heavy configuration, the movable mass and model’s CG could be moved aft to investigate

reduced static margins. The procedure was to establish trimmed flight with the mass in the

nominal (full forward) location and set the dynamic pressure to the desired value. Then the

mass was moved aft, incrementally, until the model went unstable and was snubbed or the

mass made it to the full aft position. Data was acquired periodically during this process.

7.5 BFF Suppression Testing

For flutter testing, the goal was to establish the experimental flutter onset dynamic pressure associated with a given control law. As with the flutter clearance testing, dynamic

pressure was increased incrementally; however, since flutter was expected, the increments

were reduced as dynamic pressure was increased into uncharted territory. For this testing,

naturally occurring tunnel turbulence was generally deemed adequate to perturb the model,

but control surface doublets were occasionally used to free stiction in the mount system.

Ultimately, flutter onset was determined by the test engineer’s assessment that the model

was unstable or when the WatchDog limits had been exceeded. In either case, the model

32
deg or in

84

82

80

Dynamic Pressure, PSF

20 22 24 26 28 30 32 34 36

10000

5000

0

-5000

Bending Momeing, in−lbs

Time, s

5

0

-5

Θ, deg

PZ, in

Snub!

Figure 21. Body freedom flutter time traces for trim only control, light model. WatchDog

Snub! engagement shown along with subsequent vertical position ramp down. Test 598,

point 3525. The bending moment is measured from strain gauge SBI1 (figure 8) corrected

for outboard wing weight.

would be snubbed. Figure 21 shows data from one such BFF encounter where the WatchDog system engaged the snubber and the model was subsequently landed via the emergency

controller. Note that Θ and PZ have the same frequency and are growing in amplitude.

Controllers associated with the TakeOff launch scheme were tested past their designed

dynamic pressure to determine the onset of BFF. While not a specific test objective of

Northrop Grumman, the demonstration of BFF suppression and the system ID data acquired

provided a measure of controller robustness and validated flutter onset predictions.

Controllers associated with the Release launch scheme were tested extensively for flutter

in the heavy and light configurations. As the objective of this testing was to generate BFF

suppression controllers that would 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-controlleronly (baseline) flutter onset dynamic pressure, acquiring system ID data incrementally as

dynamic pressure was increased. These data were then used, off-line, to generate BFF

suppression control laws. When testing resumed with the combined trim/BFF-suppression

controller engaged, system ID data would again be acquired as dynamic pressure was increased until a new flutter onset value was established. The system ID and control law

design procedure was repeated until the maximum value of flutter onset dynamic pressure

was established. References 12 and 20 contain more details on the BFF suppression control

laws and a discussion of the experimental results.

8 Concluding Remarks

In the Fall of 2007, the Air Force Research Laboratory (AFRL), Northrop Grumman, Lockheed Martin, and NASA Langley Research Center successfully completed the third of a

series of three wind tunnel tests of an aeroelastically scaled wind tunnel model of a flying

wing SensorCraft vehicle concept. The first of these tests was conducted on a cantilevered,

33
sidewall mount. The second and third tests used a new, multi-DOF mount. This mount

allowed the semi-span model to translate vertically and rotate in pitch at the wing root,

allowing better simulation of the full span vehicle’s rigid-body modes. The large size of

the flying wing model along with the rigid body DOF afforded by the new mount system,

created many unique challenges to successfully and safely flying the model in the tunnel

and meeting the AEI test objectives. This paper has provided an overview of the AEI flying wing wind tunnel tests from an operational point of view. It has focused on systems,

procedures, and lessons learned that enabled the test to be successful in demonstrating

Gust Load Alleviation (GLA) and Body Freedom Flutter (BFF) suppression. Descriptions

of the hardware including the wind tunnel, the wing model, the mount system, and other

supporting systems developed for this program were provided.

9 Acknowledgements

From Northrop Grumman the authors would like to acknowledge Tony Shimko for his expertise and guidance, Elaine Shaw for her tireless control design, test engineers, Lou Scherer,

Chris Curnes and Kian Tehrani, and finally, Ted Schulman for much appreciated experience

and insight.

From Lockheed Martin the authors would like to acknowledge Wink Baker and David

White for their contributions to Test 598 that including model preparation, calibration, and

wind tunnel testing.

From NASA Langley Research Center the authors would like to acknowledge Chuck McClish for building the hydraulic manifold assembly, Patrice Quander-Haas and Gary Koepple

for fabricating the snub control chassis, the moving mass controller, and their patience in

dealing with ever changing requirements, Vic Tumwa for his assistance in conducting ground

vibration tests, William Johnston and John Newman for performing the LAS spring fatigue

test on such short notice, and Mike Ramey, Jeff Wilfong, and Anthony Pototzky for their

assistance in conducting the wind tunnel test.

Finally, we would like to acknowledge the design, fabrication, and test support efforts

of NextGen Aeronautics including David Cowan, Matthew Scott, Dana Howard, Robert

Alphenaar, Robert Myers, Jonathan Prouty, and Richard Guhr.

References

1. Lucia, D., “The SensorCraft Configurations: A Non-Linear AerServoElastic Challenge

for Aviation,” 46nd AIAA/ASME/ASCE/ASC Structures, Structural Dynamics and

Materials Conference, Austin Texas, April 2005.

2. Martinez, J., “An overview of SensorCraft capabilities and key enabling technologies,”

26th AIAA Applied Aerodynamics Conference, No. AIAA-2008-7185, Honolulu, Hawaii,

Aug. 2008.

3. Reichenbach, E., “Aeroservoelastic Design and Test Validation of the Joined Wing Sensorcraft,” 26th AIAA Applied Aerodynamics Conference, No. AIAA-2008-7189, Honolulu, Hawaii, Aug. 2008.

4. LeDoux, S., Vassberg, J., and Fatta, G., “Aerodynamic Cruise Design of a Joined

Wing SensorCraft,” 26th AIAA Applied Aerodynamics Conference, No. AIAA-2008-

7190, Honolulu, Hawaii, Aug. 2008.

5. Lockyer, A. J., Drake, A., Bartley-Cho, J., Vartio, E., Solomon, D., and Shimko, T.,

“HIGH LIFT OVER DRAG ACTIVE (HiLDA) WING; Delivery Order 0007: HiLDA

34
Wing Program,” Tech. Rep. AFRL-VA-WP-TR-2005-3066, Northrop Grumman Corporation, 2005.

6. Vartio, E., Shimko, A., Tilmann, C. P., and Flick, P. M., “Structural

Modal Control and Gust Load Alleviation for a SensorCraft Concept,” 46th

AIAA/ASME/ASCE/AHS/ASC Structures, Structural Dynamics and Materials Conference, No. AIAA-2008-7185, Austin, Texas, April 2005.

7. Silva, W., Vartio, E., Shimko, A., Kvaternik, R. G., Eure, K. W., and Scott,

R. C., “Development of Aeroservoelastic Analytical Models and Gust Load Alleviation Control Laws of a SensorCraft Wind-Tunnel Model Using Measured Data,” 47th

AIAA/ASME/ASCE/AHS/ASC Structures, Structural Dynamics and Materials Conference, No. AIAA-2006-1935, Newport, Rhode Island, May 2006.

8. Staff of the Aeroelasticity Branch, “The Langley Transonic Dynamics Tunnel,” Langley

Working Paper LWP-799, Sept. 1969.

9. Corliss, J. M. and Cole, S. R., “Heavy Gas Conversion of the NASA Langley Transonic

Dynamics Tunnel,” Proceedings of the 20th Advanced Measurements and Ground Testing

Technology Conference, No. 98-2710, Albuquerque, NM, June 1998.

10. Cole, S. R. and Rivera Jr, J. A., “The New Heavy Gas Testing Capability in the NASA

Langley Transonic Dynamics Tunnel,” Royal Aeronautical Society Wind Tunnels and

Wind Tunnel Test Techniques Forum, No. 4, Cambridge, UK, April 1997.

11. Bartley-Cho, J. and Henderson, J., “Design and Analysis of HiLDA/AEI Aeroelastic

Wind Tunnel Model,” 26th AIAA Applied Aerodynamics Conference, No. AIAA-2008-

7191, Honolulu, Hawaii, Aug. 2008.

12. Love, M. H. and et al., “Final Report for the Aerodynamic Efficiency Improvement Contract No.: FA8650-05-C-3501,” Tech. Rep. FMZ-9454, Lockheed Martin Corporation,

2008.

13. Mangalam, S. and Mangalam, A., “Unsteady Aerodynamic Observable for Gust Load

Alleviation and Flutter Suppression,” 26th AIAA Applied Aerodynamics Conference,

No. AIAA-2008-7187, Honolulu, Hawaii, Aug. 2008.

14. Perry, B., Cole, S. R., and Miller, G. D., “Summary of an Active Flexible Wing Program,” AIAA Journal of Aircraft, Vol. 32, No. 1, 1995, pp. 10–15.

15. “http://www.bbc.co.uk/science/humanbody/sleep/sheep/,” BBC Homepage, Science

& Nature: Human Body & Mind, Sheep Dash! A simple on-line game testing human

reaction time.

16. McDowell, M. M., “A Controlled Study on Batted Ball Speed and Available Pitcher

Reaction Time in Slowpitch Softball,” British Journal of Sports Medicine, Vol. 39,

2005, pp. 223–225.

17. “Sanchez v. Hillerich & Bradsby Co.” Court of Appeal, Second District, Division 4,

California, 2002, http://www.napil.com/PersonalInjuryCaseLawDetail31772.htm.

18. Vartio, E. and Shaw, E., “GLA Flight Control System Design for a SensorCraft Vehicle,”

26th AIAA Applied Aerodynamics Conference, No. AIAA-2008-7192, Honolulu, Hawaii,

Aug. 2008.

19. Larimore, W. E., “Optimal Order Selection and Efficiency of Canonical Variate Analysis

System Identification,” Proc. 13 IFAC World Congress, San Francisco, California, July

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20. Penning, K., Zink, P. S., Wei, P., De La Garza, A. P., and Love, M. H., “GLA and

Flutter Suppression for a SensorCraft Class Concept Using System Identification,” 26th

AIAA Applied Aerodynamics Conference, No. AIAA-2008-7188, Honolulu, Hawaii, Aug.

2008.

36
Appendix A

Flap Actuator Response

Due to the aeroservoelastic nature of these wind tunnel tests, the frequency responses

of the five flap actuators were important. The PID gains on the five control loops were

established by examining one actuator loop at a time while tracking a square wave command.

Command and position are plotted in near real time on the dSpace1 host system. The

procedure for setting the gains is to start with zero integral and derivative gains, and

increase the proportional gain until actuator instability, then back off approximately 20

percent. The derivative gain is then adjusted to improve stability. This process is repeated

until satisfactory response characteristics are obtained. Integral gain was not found to be

useful and was generally set to zero.

Given the fact that GLA testing is quite demanding in terms of the actuator stroke

requirements and the number of cycles, the flap actuators’ performance was satisfactory.

This was especially true for Test 598 where many hours of trimmed flight were successfully

demonstrated with the AOS operating. The actuators were, however, the highest maintenance items for Test 598, and three main problems were encountered. The first of these

were leaks in the o-rings that seal the actuator shafts. They were a known wear item, and

periodic replacement was expected. The other two problems were more troublesome. There

was an issue with the end cap o-ring seal where slight variations in the installation procedure

and/or lot number variations in the o-rings could lead to part of the o-ring binding against

the actuator shaft. This binding would limit actuator bandwidth and linearity, and could be

identified by the flap position time trace having a distinctive clipped appearance. Finally, in

terms of actuator performance, free play was the biggest problem. Since, the flap position

sensor (RVDT) was not attached directly to the actuator shaft, but to the flap, any free

play between the actuator shaft and the flap would interfere with the servo control loop.

As part of the start-of-day operations, data was acquired with a 30 s sine-sweep position

command being applied to all five flaps. Command and position time histories were acquired,

and the actuator transfer functions were estimated using the MATLAB
R TFESTIMATE

command. Figure A1 shows transfer function estimates for the beginning, the middle, and

the end of Test 598. As can be seen, there was some variation in actuator transfer functions

throughout the test with flaps TE2 and TE4 having chronic phase issues compared with the

other surfaces due the inability to remove the free play.

Cross resonance in the model could also excite the actuator free play in other actuators

causing sympathetic limit cycles. In these cases, a reduced proportional gain was needed to

maintain stability. By the end of the test, the LE actuator would enter a limit cycle above

a dynamic pressure of 60 psf, and it had to be substantially detuned in order to maintain

stability as indicated by the large phase lag in figure A1.

37
1.5

1.0

0.5

0

0

-20

-40

-60

-80

-100 Magnitude

0 4 8 12 16 20

Frequency, Hz

Phase, Deg

TE1

TE2

TE3

TE4

LE

Figure A1. Flap transfer function estimates for beginning (solid), middle (dash), and end

(dot dash) of Test 598, points 216, 1600, and 5409.

Appendix B

Snubbing System Performance

This section discusses the performance of the snubbing system as an important consideration for the AEI wind tunnel tests and snubber performance data acquired during Test 593

and Test 598. Figure B1 shows the evolution of the snubber system performance where pitch

angle is plotted versus time. These time traces are lined up such that the Snub! command

for each was issued at T = 0 s.

The upper part of figure B1 shows the initial testing of the snubber system performance

during Test 593. An adjustable flow rate valve was added to the hydraulic line feeding the

snubber system to provide some control over the snub rate, and three flow rate settings

were considered. Here, with the least restrictive valve setting, the time required to snub the

model from the initiation of a Snub! command exceeded 0.5 s, and this setting was used in

Test 593. It is interesting to note that these data were acquired when the snubber hydraulics

were first installed, and no further timing studies were performed. However, subsequent and

frequent operational usage of the snubber system may have resulted in some break-in as the

operational snubs indicated a shorter time delay of about 0.3 s. For example, in data from

the end of Test 593, plotted in figure 12, there appears to be about 0.3 s of delay between

the Snub! command and the pitch angle being clamped.

In preparation for Test 598, the time delay between issuing a Snub! command and

the pitch angle being limited and subsequently clamped was a concern. The center plot

in figure B1 shows the results of the snubber system timing experiments conducted prior

to wing installation. The initial data set (blue) indicated a time delay similar to what

was observed in Test 593. As a result, three possible fixes were considered. They were

the removal of the flow rate valve, a better method of bleeding the air out, and a diode

installed across the solenoid electrical leads. As it turned out, the rotation of the pivot

carriage assembly, discussed earlier, ended up being serendipitous as it facilitated bleeding

38
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 Restrictive

T598 Setting

Θ, deg.

Θ, deg.

Most Restrictive

Figure B1. Snubber system time delay experiments, hydraulics at 2000 psi.

the hydraulics. With the actuator in a vertical stroke position, air in the actuator would rise

to the top. If the actuator was repeatedly cycled and bled, all the air from the system could

be removed. Using this procedure, a satisfactory result was finally obtained as indicated in

the center figure.

The lower plot of figure B1 shows snubber timing studies of the system in the final Test

598 configuration. Here, the wing has been installed, the system has been thoroughly bled,

and the flow rate valve was used. Several flow rate settings were investigated with the final

Test 598 setting identified in the plot where the net time delay was estimated to be 0.15 s.

The reason the flow rate valve was included was that engaging the snubber mechanism

could, under some circumstances, induce wing bending loads near limit values. The final

flow rate setting was a compromise between snubber speed and induced loads. One known

circumstance worth mentioning was air-off snubber engagement where the wing was at

maximum Θ. The large difference between the initial and snubbed values of Θ could result

in large bending moments. To avoid this problem, the model was electrically snubbed

(solenoid valve open) prior to bringing the hydraulic pump on-line. As the initial set point

for the pump was 200 psi, the model could be brought to a snubbed state relatively slowly

with minimal bending moments. The hydraulic pump would then be brought up to its

operational pressure of 2,000 psi.

The snubbing system performed quite well with one minor exception. As the solenoid

valve must be energized to unsnub the system, the solenoid coil would heat up during

prolonged test sessions. There were several instances during both Test 593 and Test 598

when the snubber mechanism could not be unsnubbed as the valve had apparently stuck in

the snubbed position. If the solenoid was left de-energized for approximately 30 minutes,

it would start functioning properly again. Installation of a pancake fan to cool the solenoid

coil and valve solved the problem during Test 598.

39
Appendix C

Rail Friction and Lift Augmentation System (LAS)

The consistent behavior and longevity of the LAS and mount system rails were critical

to the success of Test 598. These items will be discussed here.

Prior to Test 598, one area of concern was the potential for frequent replacement of the

LAS springs. The time involved in clearing the R134a test medium from the TDT along

with the time required to change the LAS springs would result in the loss of at least one

day of test time. As the model would be flying as close to the vertical center of travel as

possible, the center of the LAS springs would receive the most cycles, and data from Test

593 indicated that the vendor-rated life of 13,000 cycles could be reached several times a

week. To assess the level of conservatism in the vendors-rated life, a fatigue test of the

springs mounted in the LAS assembly was conducted using an Instron machine at NASA

Langley Research Center. The springs were pulled out a distance of 12 in to be consistent

with the configuration expected with the model flying at tunnel centerline, and the springs

were cycled at ±1 in at 1 Hz. These tests indicated that the springs had a fatigue life of at

least 98,000 cycles.

For Test 598, Lockheed Martin introduced the use of load cells to measure the LAS

force in the lab, during TDT buildup, and during wind tunnel testing. An additional

load cell could be temporarily attached to the wing to measure the force required to lift

the wing through its vertical range of motion. Reference 12 describes the extensive rail

and LAS friction experiments performed during build-up in the TDT model preparation

area. Once the wing, mount, and fairing were assembled in the final test configuration, an

abbreviated friction experiment was performed periodically to assess LAS and rail friction.

This experiment consisted of acquiring load cell and vertical position data while lifting the

wing via the third load cell through its vertical range of motion. Figure C1 shows data

acquired at the start of Test 598 and after about 4 weeks of wind tunnel testing. These data

indicated a hysteresis loop where force is dependent on the direction of motion. For the

purposes of health monitoring, these data were examined for consistency. The data acquired

at point 3237 was deemed satisfactory, and the wind tunnel test was continued.

The LAS load cell data was acquired by the DAS and monitored by the test engineers

during wind tunnel testing. Figure C2 shows LAS load cell data plotted versus wing vertical

position for three different launch and flight sequences. The data from point 3245 indicates

a healthy LAS system. Closer examination of these data indicates that during the flight,

the LAS friction is no more than ±1.5 lbs, less than what was measured air-off. The higher

vibration environment associated with wind-on operations may improve LAS performance

in this regard. The next two data points indicated a weakening and failed forward LAS

spring assembly. The LAS system failed twice during Test 598. The first failure was at the

point shown here near the middle of the test. To minimize tunnel entries, the model was

changed to the light weight configuration when the LAS was repaired. The LAS also failed

on the final day of testing.

40
80

70

60

80

70

60

Forward Load Cell

Aft Load Cell

−15 −10 −5 0 5 10 15

350

300

250

PZ, in

Force, lbs Force, lbs Force, lbs

Wing Mounted Load Cell

Figure C1. LAS and wing mounted load cell moving average data acquired during wing

vertical translation. Test 598, points 212 (blue) and 3237 (green).

100

60

20

Force, lbs Force, lbs Force, lbs

Point 3245

Point 3435

−15 −10 −5 0 5

PZ, in

Point 3455

100

60

20

100

60

20

Figure C2. LAS load cell data acquired during three different launch and flight sequences.

Test 598, points 3245, 3435, and 3455. (Forward load cell data is blue and the aft load cell

data is green).

41
PZ(t)

VZ(t)

Θ(t)

Θ(t)

Feed Forward

Estimator:

Assumed Snub

Time Delay = ∆t

10x10

Neural

Network

Dynamic

Pressure

Up Dist

COMBLO

COMBHI

Down Dist

VZ(t+∆t)

PZ(t+∆t)

+

. +

Θ(t+∆t)

Figure D1. Lockheed Martin combined parameter block diagram.

Appendix D

Development of LM Combined Parameters

A significant analysis effort was undertaken by Lockheed Martin Aeronautics to determine additional WatchDog parameters so that a Snub! command would be issued in time

to snub the model and subsequently control plunge displacement using the emergency controller before the pivot carriage hit the shocks. These combined parameters needed to take

into account the time delay in physically snubbing the model and the vertical travel needed

by the emergency controller to obtain a zero sink rate. This appendix will describe how the

LM combined parameters were developed, verified, and implemented.

A simulation model of the wing, mount system, and control systems was developed in

MATLAB
R Simulink. The wing, including the mass of the pivot carriage assembly, was

modeled using a series of state-space models for various combinations of Mach number and

dynamic pressure with two-dimensional interpolation used to generate the ASE model for

at a given flight condition. The ASE model contained inputs for five control surface position

commands and a vertical gust velocity input. The physical constraints and forces associated

with the rest of the mount system were imposed on the wing root using six force/moment

inputs (x, y, z, roll, pitch, yaw). The effects of mount flexibility were modeled using these

force terms. In the case of the vertical force, terms for friction as a function of bending

moment, the LAS, the shocks, and the hard stop were included. The pitch moment force

included terms for the pitch shock absorbers, hard stop, and the pitch clamp actuator. The

outputs of the ASE model were the 23 aircraft sensors.

A series of simulations were run to determine the relationship between the initial condition of the model and 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 both

the heavy and light model configurations. The initial conditions for all other parameters

were assumed to be zero. The time delay between Snub! command and a pitch restoring

force being applied to the wing root was also assumed to be zero. The simulation studies

indicated that if the model was not snubbed at the time of pivot carriage to shock contact,

large pitch excursions and wing bending moments would result. Thus, wing bending moments were within allowable limits as long as the model was snubbed prior to vertical shock

contact. These simulation data were then used to construct a new WatchDog parameter.

The parameter was generated by determining the relationship between vertical velocity,

pitch angle, and dynamic pressure and the vertical distances up or down needed to stop the

model once the snubber had been engaged. This mathematical relationship was represented

by a 10 by 10 neural network.

In order to account for any time delay between the Snub! command and pitch snub

mechanism physical engagement, some parameters were extrapolated into the future. A

42
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, in

Θ, deg

TE2 Command, deg

Bending Moment, 

1000 In−lb

Snub!

Figure D2. Hard landing case plot of vertical position, pitch angle, TE2 command, and wing

root bending moment with Snub! command indicated. Test 593, point 1252. The bending

moment is measured from strain gauge SBI1 shown in figure 8 corrected for outboard wing

weight.

simple feed forward estimation routine using an assumed snubber time delay was applied

to vertical position, vertical rate, pitch angle, and pitch rate. The forward estimator and

neural network were combined in series in the WatchDog system as shown in figure D1 to

provide a real-time estimate of the vertical position, up or down, where the model would be

brought under control. These values were monitored by the WatchDog system, and as with

the other monitored signals, the WatchDog would issue a Snub! command when thresholds

were exceeded.

To assess snubber time delay requirements, five hard landing cases from Test 593 were

examined. For one of these data points, figure D2 plots the time histories of vertical position,

pitch angle, Snub! command, TE2 command, and wing root bending moment. The figure

shows that the model starts to climb then falls back, lands hard, and bounces. Pitch angle

and bending moment decrease rapidly after contacting the shocks just before T = 42 s where

slope discontinuities are observed in the vertical position, pitch angle, and bending moment

time histories. These data are consistent with the simulation results indicating the need to

have the model clamped in pitch prior to shock contact to minimize subsequent pitch angle

excursions and high wing bending moments.

For the five hard landing cases, the performance of the WatchDog system monitoring

only the original measured parameters was compared with the WatchDog system configured

to monitor the LM combined parameters using two different assumed snubber time delays

of 0.1 and 0.25 s. Figure D3 shows one of these hard landing cases where the model vertical

position and pitch angle are plotted along with the Snub! command associated with the

three WatchDog configurations. Shock contact is indicated by the red line. For this data

point, the original WatchDog system would have snubbed the model as it was descending

through a −10 in/s velocity threshold. The combined parameter cases caught the model

earlier, as shown. The figure of merit for these cases is the time difference between shock

contact and when the command was issued. If the actual, physical snubber time delay is less

than the required time delay, then a given WatchDog parameter would have been successful

43
41.2 41.4 41.6

Time, s

41.8 42.0 42.2

10

8

6

4

2

0

-2

-4

-6

-8

-10

PZ, in

Θ, deg

Bottom Shock Contact

Snub! Nominal Parameters

Snub! LM Comb, delay = 0.10 s

Snub! LM Comb, delay = 0.25 s

in or deg

Figure D3. Hard landing case comparing nominal WatchDog Snub! command with the

Snub! command generated by the LM combined parameters where two values of assumed

time delay were used in the feed forward estimator. Test 593, point 1252.

Table D1. WatchDog system performance assessed using five hard landing cases from Test

593.

Assumed Required

Snub Time

Time Delay, s WD Performance

Delay, s (pt. 1252) In-Time False Alarm

0.065 2/5 1/5

0.10 0.113 5/5 0/5

0.25 0.165 2/5 1/5

44
in avoiding model and shock contact and high loads.

Table D1 shows the required time delay for the data point in figure D3 along with a

summary of the performance of the WatchDog system for all five Test 593 cases examined.

A false alarm is defined as a case where a Snub! command is issued, but the bending

load limits were not subsequently exceeded in the data set. The combined parameter with

an assumed 0.1 s snubber time delay was able to snub the model before contacting the

shock in all five cases with no false alarms. As previously discussed, the physical snubber

system was ultimately configured to provide a delay of 0.15 s, and this value was used in

feed forward estimator to compute the LM combined parameters during Test 598. The

WatchDog performance in Test 598 was satisfactory.

45
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Aeroservoelastic Testing of Free Flying Wind Tunnel Models Part 1: A 

Sidewall Supported Semispan Model Tested for Gust Load Alleviation and 

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Scott, Robert C.; Vetter, Travis K.; Penning, Kevin B.; Coulson, David A.; 

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This is part 1 of a two part document. Part 2 is titled: "Aeroservoelastic Testing of Free Flying Wind Tunnel Models, Part 2: A Centerline Supported Fullspan 

Model Tested for Gust Load Alleviation." A team comprised of the Air Force Research Laboratory (AFRL), Northrop Grumman, Lockheed Martin, and the 

NASA Langley Research Center conducted three aeroservoelastic wind tunnel tests in the Transonic Dynamics Tunnel to demonstrate active control 

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 vehicle was mounted to a force balance to demonstrate gust load alleviation. In the second and third tests, the same wing was mated to a new, 

multi-degree of freedom, sidewall mount. This mount allowed the half-span model to translate vertically and pitch at the wing root, allowing better simulation 

of the full span vehicle's rigid body modes. Gust load alleviation (GLA) and Body freedom flutter (BFF) suppression were successfully demonstrated. The 

rigid body degrees-of-freedom required that the model be flown in the wind tunnel using an active control system. This risky mode of testing necessitated that 

a model arrestment system be integrated into the new mount. The safe and successful completion of these free flying tests required the development and 

integration of custom hardware and software. This paper describes the many systems, software, and procedures that were developed as part of this effort. 

15. SUBJECT TERMS

Active control; Aeroservoelasticity; Body freedom flutter; Flutter suppression; Gust load alleviation; Wind tunnel tests

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