February 2014 

NASA/TP–2014-218170 

Aeroservoelastic Testing of Free Flying Wind 

Tunnel Models 

Part 2: A Centerline Supported Fullspan Model Tested 

for Gust Load Alleviation 

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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February 2014 

NASA/TP–2014-218170 

Aeroservoelastic Testing of Free Flying Wind 

Tunnel Models 

Part 2: A Centerline Supported Fullspan Model Tested 

for Gust Load Alleviation 

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 
Available from: 

NASA Center for AeroSpace Information 

7115 Standard Drive 

Hanover, MD 21076-1320 

443-757-5802 

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National Aeronautics and Space Administration. 
Abstract

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

Flying Wind Tunnel Models Part 1: A Sidewall Supported Semispan Model Tested for Gust

Load Alleviation and Flutter Suppression.” A team comprised of the Air Force Research

Laboratory (AFRL), Boeing, 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

full-span, aeroelastically scaled, wind tunnel model of a joined wing SensorCraft vehicle was

mounted to a force balance to acquire a basic aerodynamic data set. In the second and third

tests, the same wind tunnel model was mated to a new, two degree of freedom, beam mount.

This mount allowed the full-span model to translate vertically and pitch. Trimmed flight at

−10 percent static margin and gust load alleviation 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. The balance and free flying wind tunnel tests will be summarized. The

design of the trim and gust load alleviation control laws along with the associated results

will also be discussed.

1
Contents

1 Introduction 6

2 Transonic Dynamics Tunnel and the Airstream Oscillation System 9

3 Model and Mount System Description 10

4 Model Signals, Conditioning, and Control 14

4.1 dSPACE Digital Control Systems . . . . . . . . . . . . . . . . . . . . . . . . . 16

4.2 Snubber Control System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 17

4.3 RVDT and Servo Valve Signal Conditioning . . . . . . . . . . . . . . . . . . . 18

4.4 TDT Data Acquisition System . . . . . . . . . . . . . . . . . . . . . . . . . . 18

5 WatchDog and Fault Detection Systems 19

6 Model Launch, Test, and Recovery 22

7 Trim Controller Architecture 24

8 TDT Test 613: Free Flying Function Check and ParamID 25

9 TDT Test 614: Free Flight Testing – ParamID and GLA 28

10 Concluding Remarks 34

A Control Surface Actuator Description, Bench Testing, and In-service Performance 37

A.1 Control Surface Actuator Bench Testing . . . . . . . . . . . . . . . . . . . . . 38

A.2 Model Actuator Frequency Response . . . . . . . . . . . . . . . . . . . . . . . 40

B Basic Aerodynamic Testing 45

List of Tables

1 FlightModes, WatchDog parameters, and parameter limits. Limits are in

units of deg, in, micro-strain, and seconds. . . . . . . . . . . . . . . . . . . . 21

2 Fault detection system parameters. Values are in units of deg, in, and seconds.

Values in parenthesis were initial values. . . . . . . . . . . . . . . . . . . . . 22

A1 JWS actuator requirements . . . . . . . . . . . . . . . . . . . . . . . . . . . . 37

List of Figures

1 NASA Langley Transonic Dynamics Tunnel tests for the HiLDA and AEI

programs. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6

2 Flying wing SensorCraft model TDT test configurations. . . . . . . . . . . . . 7

3 Illustration of the AEI Joined Wing SensorCraft model mated to a two-DOF

support system installed in the NASA Langley TDT. . . . . . . . . . . . . . . 8

4 Illustration of the NASA Langley Transonic Dynamics Tunnel (TDT) and

associated Airstream Oscillation System (AOS) used to conduct the JWS tests. 10

5 Close-up views of the TDT air and R134a operation envelopes showing the

conditions where joined wing testing was conducted. . . . . . . . . . . . . . . 11

2
6 Joined Wing SensorCraft model planform view showing onboard instrumentation, actuators, and servo valves. Accelerometer orientation: x = longitudinal, y = lateral, z = vertical. Fuselage gyros measure pitch, roll, and yaw

rates. Wing tip gyros measure local rotation rate. . . . . . . . . . . . . . . . . 12

7 Joined wing, beam support, and carriage assembly installed in the TDT for

Tests 613 and 614. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 13

8 Signal routing used in the joined wing tests. . . . . . . . . . . . . . . . . . . . 15

9 dSpace and snubber control block diagram. . . . . . . . . . . . . . . . . . . . 16

10 Flight control block diagram. . . . . . . . . . . . . . . . . . . . . . . . . . . . 17

11 Snubber control chassis containing the snubber control system with manual

Snub! (chicken) switch and hydraulic pressure displays. . . . . . . . . . . . . 18

12 WatchDog system snub logic block diagram. . . . . . . . . . . . . . . . . . . . 19

13 Description of the launch and test procedures used with the JWS model. . . . 23

14 Data acquired during model launch. Test 614, point 1655. δCmd is command

signal sent to all aft wing control surfaces. . . . . . . . . . . . . . . . . . . . . 24

15 Trim control law diagram. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 25

16 Nominal Test matrix used in TDT Test 613. . . . . . . . . . . . . . . . . . . . 26

17 Nichols stability plot for initial trim controller gains with +5 and −10 percent

static stability. Test 613 data points 412 and 438 are plotted along with

simulation results. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 27

18 Nominal test matrix used in TDT Test 614. . . . . . . . . . . . . . . . . . . . 28

19 Symmetric, all wing control surface sweep 0.1–20 Hz. Amplitude = 2.5◦, Test

614, data point 1019, GLAMode 0, TestMode 1, static margin = −10 percent. 29

20 AOS sweep 012–.1 Hz. Test 614, point 1750, GLAMode 0, TestMode 1, static

margin = −10 percent. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 31

21 AOS sweep 12–0.1 Hz. Test 614, point 1765, GLAMode 5v6, TestMode 1,

static margin = −10 percent. . . . . . . . . . . . . . . . . . . . . . . . . . . . 32

22 GLAMode 5v6 block diagram. . . . . . . . . . . . . . . . . . . . . . . . . . . . 33

23 Tail root and forward wing root strain response to AOS. Test 614, data points

1750, 1769, 1765, and 1770. . . . . . . . . . . . . . . . . . . . . . . . . . . . . 33

24 Nichols plot of trim (GLAMode 0) and trim+GLA (GLAMode 5v6) controllers. Test 614, data points 1774 and 1773. . . . . . . . . . . . . . . . . . . 34

A1 Photographs of actuator bench test rig in two different configurations. . . . . 39

A2 Actuator test bench performance data. Moog Series 30 servo valve. . . . . . . 40

A3 Test bench servo valve manifolds. . . . . . . . . . . . . . . . . . . . . . . . . . 41

A4 JWS model internal hydraulic system. . . . . . . . . . . . . . . . . . . . . . . 42

A5 Forward wing, left-mid-board control surface frequency response with the

single and all control surfaces operating, before and after model plumbing

modifications. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 43

A6 Control surface frequency response function estimates, ±3

◦

linear sine sweep

excitation. Wing control surfaces on the left side of the model are shown using

solid lines, and control surfaces on the right side of the model are shown using

dashed lines. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 43

B1 Joined wing SensorCraft model installed on the TDT sting for Test 606. . . . 46

B2 Normal force and pitching moment associated with aft wing control surface

deflections. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 47

3
Nomenclature

A/D analog to digital

AEI Aerodynamic Efficiency Improvement

AFRL Air Force Research Laboratory

AOS Airstream Oscillation System

ASE aeroservoelasticity

AWI aft wing inboard

AWM aft wing midboard

AWO aft wing outboard

azMode flight control parameter used internally to select which trim control loops operate

BNC Bayonet Neill-Concelman connector

CG center of gravity, referred to throughout the paper in terms of the resulting static margin,

percent mean aerodynamic chord

CS control surface

D/A digital to analog

DAS data acquisition system

DAU data acquisition unit

DCS Digital Control System

DOF degree of freedom

dSpace1 DCS for Servo PID loops and WatchDog

dSpace2 DCS for Trim & GLA

FlightMode flight mode

FWO forward wing outboard

FWM forward wing midboard

FWI forward wing inboard

GLA gust load alleviation

GLAMode user input for GLA control law selection

gpm gallons per minute

GUI Graphical User Interface

HiLDA High Lift over Drag Active Wing

JWS Joined Wing SensorCraft

LESP leading edge stagnation point sensor

MFW O GLA output mix value for forward wing outboard surfaces

MAW I GLA output mix value for aft wing inboard surfaces

OADAS Open Architecture Data Acquisition System

ParamID parameter identification

PID proportional integral derivative

PilotIn user input for initiating model launch or landing

PSNUB snub signal for pitch snub mechanism

Pt TDT total pressure, psf

PTHETA model pitch angle also referred to as Θ, deg

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

PZCenter parameter used in fault detection algorithm, in

PZCmd PZ set point for trim controller, in

PZRange parameter used in fault detection algorithm, in

Reset user input for resetting control systems in preparation for model launch

RUD rudder

RVDT rotary variable differential transducer

SGFWLI strain gauge forward wing left inboard, microstrain

SGFWRI strain gauge forward wing right inboard, microstrain

SGTF strain gauge tail forward (or root), microstrain

4
SGTM strain gauge tail mid, microstrain

SGTA strain gauge tail aft, microstrain

Snub! command to engage snubbing system

SNBXSTAT signal indicating state of snubber control system relays

TestMode user input for selecting the type of ParamID testing

TDT Transonic Dynamic Tunnel

TDTDASD model preparation area OADAS

TDTDASM mobile OADAS

UAV unmanned aerial vehicle

V Volts

VZ vertical velocity, in/s

VZLimit parameter used in fault detection algorithm, in/s

WatchDog Software for triggering Snub! command

ZSNUB Snub signal for brake

δCmd wing control surface and rudder position commands, deg

δCS wing control surface positions, deg

δRudder rudder position, deg

Θ model pitch angle also referred to as PTHETA, deg

ΘCenter parameter used in fault detection algorithm, deg

ΘRange parameter used in fault detection algorithm, deg

ΘCmd Θ set point for trim controller THETA loop, deg

Θ pitch rate, deg/s  ̇

Θ ̇ Limit parameter used in fault detection algorithm, deg/s

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 Transonic Dynamics Tunnel tests for the HiLDA and AEI programs.

1 Introduction

A goal of the Air Force Research Laboratory’s SensorCraft project is to develop technologies

for future high altitude long endurance 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 programs 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.

Several HiLDA and AEI aeroservoelastic wind tunnel tests were conducted in the NASA

Langley Transonic Dynamics Tunnel (TDT) and are depicted in figure 1. The first two

wind tunnel tests were conducted as part of the HiLDA program. These included a short

test of several candidate sensors for measuring gust flow angle. The HiLDA flying wing was

tested in September 2004 on a cantilevered mount with GLA being demonstrated using the

model’s five active control surfaces [3–5]. Figure 2a shows that test configuration.

While not originally designed for this purpose, the HiLDA flying wing model was mated

to a new, two degree of freedom mount and retested in support of the AEI program in

2007. This mount allowed the half-span model to translate vertically and pitch at the

wing root, allowing the model to “fly” in the tunnel providing a simulation of the full-span

vehicle’s rigid body modes. The large size of the flying wing model along with the rigid

body degrees of freedom afforded by the new mount system, created many unique challenges

to successfully and safely conducting the test. A remotely actuated snubbing system that

returned the model to a fixed pitch angle was included in the mount design. Reference 6

and part 1 of this report both describe applicable systems and procedures including how to

effectively use this system, which is critical to the success of these tests. Figure 2b depicts

the flying wing and mount system in the final configuration used in Test 598 where GLA

and body freedom flutter suppression control laws were successfully demonstrated [6–10].

The Joined Wing SensorCraft (JWS) wind tunnel tests built on the systems, procedures,

and lessons learned from the flying wing tests.

Figure 1 shows the three tests of the JWS that were conducted [11, 12]. During the

first test (TDT Test 606, December 2008) the model was installed on the TDT sting with

6
Flow

Oscillation

Vanes

Flying-Wing

SensorCraft

Model

“Gust 

Sniffer”

(a) Photo of TDT Test 579 configuration showing the AOS vanes and the flying wing attached to the balance.

Flow Oscillations (“Gusts”)

“Gust Sniffer”

Flying-Wing

2 DOF Mount

SensorCraft Model

(b) Illustration of the flying wing SensorCraft model mated to a two-DOF sidewall support system.

Figure 2. Flying wing SensorCraft model TDT test configurations.

7
Figure 3. Illustration of the AEI Joined Wing SensorCraft model mated to a two-DOF

support system installed in the NASA Langley TDT.

8
a balance, and a basic aerodynamic data set was acquired in air and heavy gas. During

the second and third tests a new beam mount, shown in figure 3, was used. The beam

mount along with a carriage assembly installed inside the model provide pitch and plunge

DOF allowing the model to “fly” in the TDT test section. The carriage design incorporates

model arrestment systems that include a plunge brake and a pitch displacement limiter.

During the second test, performance of the trim control law was verified and parameter

identification (ParamID) data were acquired. GLA was demonstrated during the third test.

Reference 13 provides a detailed description of the objectives for the JWS wind tunnel

test program. The primary objectives were to demonstrate stable controlled flight at −10

percent static stability, measure stability margins while flying, and reduce gust loads of

the first symmetric structural mode by 50 percent and of the second symmetric structural

mode by 20 percent. These objectives were achieved except the reduction of the loads of

the second structural mode. This objective was not attempted due to the limited accuracy

of the simulation models at the higher frequencies. A real-time stability margin test mode

was successfully implemented and used throughout testing to determine the phase and gain

margins of several of the key controller loops within the flight control system. These stability

margin measurements were used to verify the modeling used to design the control laws.

Reference 14 provides an overview of the JWS project. References 15 and 16 describe the

wind tunnel model and mount system, respectively, and reference 17 covers some ParamID

research that is not covered here. The purpose of this paper is to provide an overview of the

AEI Joined Wing SensorCraft wind tunnel tests focusing on systems, procedures, control law

development, and results [18]. The main focus of this paper will be on the dynamic testing

conducted during TDT tests 613 and 614. However, a summary of the basic aerodynamic

test (TDT Test 606) is also included in the appendix for completeness.

2 Transonic Dynamics Tunnel and the Airstream Oscillation System

The TDT, depicted in figure 4a, 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 [19]. 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, or R-12, as the heavy gas test medium; since then the

TDT has used 1,1,1,2 tetrafluoroethane, or R-134a [20, 21], 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.

Because of the need to perform GLA testing as part of the AEI program, the TDT

Airstream Oscillation System (AOS) was used frequently. The illustration in figure 4b

shows the key features of the AOS system. These features include biplane arrangements

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. Although

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
c

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

9
(a) TDT. (b) AOS.

Figure 4. Illustration of the NASA Langley Transonic Dynamics Tunnel (TDT) and associated Airstream Oscillation System (AOS) used to conduct the JWS tests.

peak. In order to reduce model risk, the vanes were set to 4◦ peak to peak at the start of

Test 613. Soon after the start of Test 613 it was determined that the JWS model could

accommodate the full 12◦ peak to peak setting. Thus, the vast majority of data acquired

during Test 613 and all of the data acquired during Test 614 used this setting resulting in

a ±1

◦ variation in stream angle in the vicinity of the model for low frequency oscillations;

the amplitude rolls off with increasing frequency.

Figure 5a and b show the close-up views of the TDT air and R134a operation envelopes,

respectively. The red lines represent the regions traversed when conducting the joined wing

tests. The red dots indicate the specific Mach number and dynamic pressure conditions

where virtually all the relevant joined wing data was acquired. In all cases the nominal

dynamic pressure was 51 psf with a Mach number of 0.195 in air and 0.220 in R134a. To

verify model stability, flutter clearance runs were conducted in both test mediums for both

the sting and the beam mount configurations. This testing involved increasing the dynamic

pressure to 55 psf, slightly above the nominal test condition. During these clearance runs,

model response was given a high degree of scrutiny. The red lines extending above the

nominal test value of 51 psf in figure 5 indicate the flutter clearance.

Finally, it should be pointed out that while the model was scaled for testing in R134a,

it is common practice to perform air checkout runs in the TDT prior to filling it with heavy

gas. At the tunnel pressures used, tunnel preparation time for air testing was approximately

30 min, while preparation time for heavy gas was approximately 7 hr. Returning the tunnel

to a condition where personnel could enter the test section took comparable amounts of

time. Thus, basic functionality of all systems was verified in air prior to beginning the

R134a testing.

3 Model and Mount System Description

The model is an eight percent length scale, full-span, aeroelastically scaled model of a Boeing

Joined Wing SensorCraft design concept. The wings and tail boom use a spar pod design

where the scaled stiffnesses (EI and GJ distributions) were achieved by proper geometry

of the metal spars, and the aerodynamic shape was provided by discrete fairings, or pods,

mounted to the spar. Lead weights were attached to the spar to achieve the proper scaled

mass distribution.

The planform drawing of the JWS model in figure 6 shows that each wing has three pairs

of trailing edge control surfaces and the tail boom has a rudder. The 13 control surfaces

10
10

100

Dynamic pressure, psf

0 0.1 0.2 0.3 0.4 0.5 0.6

Mach number

90

80

70

60

50

40

30

20

P =2000 psf

P =2200 psf

P =1800 psf

P =1600 psf

t

t

t

t

(a) Air.

10

100

Dynamic pressure, psf

0 0.1 0.2 0.3 0.4 0.5 0.6

Mach number

90

80

70

60

50

40

30

20

P =2000 psf

P =2200 psf

P =1800 psf

P =1600 psf

t

t

t

t

(b) R134a.

Figure 5. Close-up views of the TDT air and R134a operation envelopes showing the conditions where joined wing testing was conducted.

11
g

g

g

g

g

g

g

b

Control Surfaces

s Servovalve (13)

a Actuator (13)

r RVDT (13)

x Accelerometer (2)

y Accelerometer (1)

z Accelerometer (11)

b Bending Strain Gauge (15)

g LESP Sensor (5)

g Rate Gyro (5)

b

b b

s s

r

r

a a

s

r

a

s

s

r

r

a

a

s

ar

s r

a

s

r

s a

s

r

r

a

a

s

ra

s

s

r

r

a

a

g

z

z

b

b

b

b

b

b

b

g

z

z

z

z

b

b

b

b

z

z

x

z

y

z

z

xg

Figure 6. Joined Wing SensorCraft model planform view showing onboard instrumentation,

actuators, and servo valves. Accelerometer orientation: x = longitudinal, y = lateral, z =

vertical. Fuselage gyros measure pitch, roll, and yaw rates. Wing tip gyros measure local

rotation rate.

12
were driven by vane-type hydraulic actuators operated at 1,000 psi with position measured

by rotary variable differential transducers (RVDTs). Other instrumentation included strain

gauges, accelerometers, and rate gyros for fuselage pitch, roll, and yaw. A movable mass

system was included in the fuselage that provided remote adjustment of the model center

of gravity CG. The device consisted of a compact DC motor coupled via a belt to two ball

screws that each moved 13 lb of lead. Theses masses could move over a range of 26 in. The

position of these ballast weights was mapped to 4 CG locations associated with +5, 0, −5,

and −10 percent static margin. Additional model details are found in references 15 and 13.

(a) Joined wing SensorCraft model installed on the beam support.

(b) Close up photograph of the joined wing model,

carriage, and beam support.

(c) Carriage installed on the beam support without

the joined wing model. The cables were used to support weights to test brake operation.

Figure 7. Joined wing, beam support, and carriage assembly installed in the TDT for Tests

613 and 614.

The beam support system along with a carriage assembly installed inside the model provide pitch and plunge DOF allowing the model to “fly” in the TDT test section. Figures 7a

13
and 7b show the model installed on the carriage and beam support system. As noted in

the upper photo, a string potentiometer was used to measure vertical position (PZ) and

vertical velocity (VZ). The string potentiometer itself was mounted on the upper end of the

beam with the string passing over a small pulley then along the downstream face of the

beam where it attached to the carriage. The carriage design incorporates model arrestment

systems that include a plunge brake and a pitch displacement limiter. Figure 7c shows the

carriage installed on the beam without the model. References 13 and 16 provide additional

details about the design and testing of the beam support system.

Because of the aeroservoelastic nature of these wind tunnel tests, the frequency response

of the control surface actuators is important both in terms of bandwidth and consistency.

During the model design/fabrication process, a bench test rig was used to assess the performance of a single actuator using the same components intended for the wind tunnel model.

Later, the performance of all 13 actuators operating within the model was assessed and a

number of changes had to be made to the model hydraulic system. The appendix discusses

the bench test experiments and subsequent performance in the wind tunnel model.

Reference 13 describes a ground vibration test of the JWS model supported by soft

springs that was performed at the TDT prior to the wind tunnel tests. The first three

symmetric flexible modes had frequencies of 4.57 Hz, 9.05 Hz, and 12.75 Hz. These modes

are dominated by displacement that can best be described as forward wing first bending,

forward wing second bending, and aft wing first bending, respectively.

4 Model Signals, Conditioning, and Control

This section of the paper presents several block diagrams showing increasing levels of detail

on how the model signals were routed, conditioned, and controlled. The first of these

diagrams shows how signals are routed to the model, the various signal conditioners, the

control systems, and the data acquisition system. The next block diagram focuses on the

two digital control systems and the snubber control system, and the third diagram focuses

on the flight control block. Some items in these diagrams will be described in greater detail

in this section of the paper while others will be covered in separate sections. Within these

diagrams, signals depicted in orange are operator inputs and those in red are snubber system

related signals.

Figure 8 shows a high level schematic of the signal routing arrangement used in joined

wing tests. The input and output signals were routed between the model or mount system

and the TDT control room via the control room-to-sting patch panel or directly to the control

room via the plenum penetration. 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 fuselage mounted 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 DCS running at a 1,000-Hz

frame rate (dSpace1). All other signals were filtered at 100 Hz to be compatible with the

Nyquist frequency for the other DCS running at a 200 Hz frame rate (dSpace2). Figure 8

shows that the servo valve signals were routed to a voltage-to-current converter and back

to the model. 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 9 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; the dSpace input signals were converted from volts to engineering units

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

14
Joined-Wing Model and Beam Support

Test Section

Control Room 

Data Room 

ServoValues

Strain Gauges,

Accelerometers, 

Wing Rate Gyros

& StringPotentiometers

Fuselage

Rate

 Gyros

Fus. Gyros

Servovalue Commands

Unfiltered CS Positions

Filtered CS Positions

CS Position Commands

RVDTs

Voltage to

Current 

Converter

Custom 

+/-15 Volt 

Power Supplies Pacific Instruments

Signal Conditioner/Filter

Digital Strip Chart

DAS

RVDT

Signal 

Conditioner

Strain Gauges,

Accelerometers, 

Wing Rate Gyros

& StringPotentiometers

Snubber

Control

Snub!

Brake

&

Snubber

Solenoids

Control Room

to Plenum Penetration

Moving

Mass

Power

Moving

Mass

Motor

Power

dSpace1

Servo Loop

WatchDog

dSpace2

Trim + GLA 

FlightMode

PSNUB

&

ZSNUB

Control Room to Sting Patch Panel

Control Room To Data Acquisition Patch Panel

Figure 8. Signal routing used in the joined wing tests.

Programming logic was used to decode the meaning of the analog signal for signals intended

to depict discrete system states like FlightMode.

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

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

is a latching circuit that can 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 SNUBXSTAT signal. The PSNUB and

ZSNUB signals communicated the state of the plunge brake and pitch snub mechanisms in

the model. A detailed description of the snubber control system and associated hydraulic

components is described in a separate section, below.

As shown in figure 9, dSpace2 incorporates a flight control block and a set of externally

generated control surface commands that can be summed and sent out as analog control

surface command signals to dSpace1. The excitation signal can be added to various combinations of the control surface commands when conducting ParamID testing in TestMode 1.

The excitation signal is used internally by the flight control block for TestModes of 2 or

greater. The GLA and trim control laws were implemented within the flight control block.

This flight control block includes a Graphical User Interface (GUI) and programming logic

for controlling or initiating certain events like resetting the system or initiating a launch.

These operator inputs are depicted in the figure 9. The FlightMode state is communicated

from dSpace2 to the WatchDog system as shown.

The servo-control loops and the WatchDog system were implemented on dSpace1. The

servo-loops were independent proportional integral derivative (PID) control loops equipped

with output saturation blocks to prevent overdriving the actuators. The WatchDog system

monitored the model signals, and when a fault was detected, it would issue a Snub! command

15
δposition

δCmd.

δCmd

Model Signals

Latching

Circuit

δCmd

TestMode 1 Excitation Input

Manual Chicken Switch

FlightMode

Servo Control

PID

WatchDog

Snub

Logic

V Divider

Snubber Control System

To Solenoid Valves:

24V=Unsnubbed

OV=Snubbed

8V=Unsnubbed

0V=Snubbed

To Model:

Servovalves

dSpace2 dSpace1

Reset 

Switch

Plunge

Override 

Switch

Gnd

Snub!

External Inputs

24 V

Power

Gnd

Pitch

Override 

Switch Plunge

Override 

Power Switch

V Divider

V Divider

Plunge

Brake

Pitch

Snub

= 0

SNBXSTAT

PSNUB

ZSNUB

Excitation

PilotIn

0 or 1

Reset

0 or 1

Test

Mode

0,1,2,3...

Operator

Inputs: GLA

Mode

0,1,2,3...

Flight

Control

Enable

++

Figure 9. dSpace and snubber control block diagram.

and transfer control to the emergency control law via the switch shown in figure 9. For the

joined wing tests, the emergency controller consisted simply of 0◦control surface commands.

In addition to the WatchDog system, the control block on dSpace2 had a fault detection

algorithm and could arrest the model by setting value of FlightMode to 6.

Figure 10 shows key features of the flight control block. The primary components of

this block are the trim and GLA blocks, the FlightMode logic block, and the fault detection

block. Model signals, snubber related signals, operator inputs, and the relevant model

signals are shown. The fault detection algorithm will be discussed in a separate section

along with the WatchDog system, but, when a fault is detected, it is communicated to the

FlightMode logic block whereupon the FlightMode parameter is set to 6. The trim controller

has two modes of operation: Theta hold and altitude (Z) hold. The exact mode of operation

is determined by the user inputs, the model vertical position, and the azMode parameter.

Logic for ramping the vertical position set point from the lower stop to tunnel centerline is

also contained in the trim control block. The user input GLAMode controls the operation

of the GLA control block. When GLAMode is set to 0, the GLA control block simply passes

through the control surface commands. When this parameter is greater than 0, strain gauge

feedback is used to generate GLA control surface commands that are added to the trim

controller outputs. For simplicity, the excitation signal shown in figure 9 is not shown in

figure 10. The logic whereby the TestMode parameter can inject the excitation signal at

various points in the trim or GLA control blocks is also absent. In order to assess controller

performance an additional 17 signals were used to record various parameters within the trim

control and GLA loops. Although they do not appear here, these signals are routed to and

recorded on the DAS.

4.1 dSPACE Digital Control Systems

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

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

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

analog-to-digital (A/D) boards and one dSPACE DS2103 multi-channel digital-to-analog

(D/A) converter board. The A/D boards each have 32 channels using 16 bit quantization

16
GLA Mode

Strain Gages

Fuselage Accelerometers

CG

PZ & VZ

azMode

FlightMode

Trim

Control

Secondary

WatchDog

Θ & Θ

δCmd. δCmd.

.

Θ & Θ

.

Model Signals

Reset

Pilotln

TestMode

PSNUB

ZSNUB

SNBXSTAT

PZ

PZ & VZ

FlightMode

Logic

GLA

Control

Figure 10. Flight control block diagram.

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 is developed within the MATLAB
c Simulink environment, then

compiled and downloaded to the target processor via the dSPACE and MATLAB RealTime Interface. The ControlDesk application is an integral component of the dSpace tool.

ControlDesk 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 Snubber Control System

The system described in reference 6 was modified to support the joined wing tests. The

original system consisted of the snubber control chassis and an external manifold assembly.

It supported only a single solenoid valve located in a manifold assembly external to the

model. In addition to the single solenoid valve, the external manifold assembly consists of

an adjustable pressure switch and two regulators for supplying up to two lower hydraulic

pressures to the wind tunnel model. Hydraulic pressure transducers were included to monitor the supply pressure as well as the two regulated pressures. The pressure switch was

wired to both an indicator light and an audible alarm housed in the snubber control chassis.

For the joined wing tests, a second override switch and alternative 24 VDC power (red

covered toggle switch) were added to the snubber control chassis. Figure 11 shows this

modified version of the system. Although the previously described manifold assembly was

used to monitor and regulate hydraulic pressure to the model, two Parker Hannifin solenoid

valves located within the wind tunnel model were used for model restraint. These valves

provided separate control of the pitch snubber mechanism and the brake. Hydraulic fluid,

at 1,000 psi, from the TDT 30 gpm pump supplied both the joined wing actuators and the

two solenoid valves. The solenoid valves were normally open and required 24 VDC power

to remove pressure from the snubbing system thereby releasing the model.

17
Figure 11. Snubber control chassis containing the snubber control system with manual

Snub! (chicken) switch and hydraulic pressure displays.

Figure 11 shows the snubber control chassis and a handheld chicken switch, while 9

shows key aspects of the snubber control system. Two amber colored lights on the front

of the snubber control box indicate which source initiated the Snub! command, chicken

switch or external source (WatchDog system). The relays and lights were reset with a

momentary toggle switch (reset switch). Two independent lighted switches are used to

hold a braked/snubbed condition. The plunge override power switch allows the operator

to momentarily release the brake to recover (lower) the model after the brake/snubber has

been engaged during flight.

4.3 RVDT and Servo Valve Signal Conditioning

The 13 active control surfaces on the joined wing model exceeded the original TDT signal

conditioning capabilities. As a result, two custom chasses were fabricated using commercially

available signal conditioning components.

The RVDT signal conditioner chasses consisted of 16 LDM-1000 LVDT DIN rail mounted

signal conditioner modules and their DC power source manufactured by Schaevitz Sensors.

These conditioner modules supply an AC signal to the RVDTs and then demodulate and

amplify the output. A number of internal dip switches can be used to configure the units.

For the joined wing tests the RVDTs were driven at 3 V RMS, and the position output

range was set to ±5 VDC. The setting of the internal filter and excitation frequency will be

discussed in a separate section.

The servo valve signal conditioner chassis consists of 8 VC2124 two axis voltage-tocurrent converters manufactured by Delta Computer Systems. Each of these units transforms ±10 V signals into current signals capable of driving the Moog hydraulic servo valves.

Each unit can drive two servo valves, and full scale current output is selectable with the

±10 mA setting used for the joined wing tests.

4.4 TDT Data Acquisition System

The TDT Open Architecture Data Acquisition System consists of three data

acquisition/reduction systems. The real-time systems available to the facility are the tunnel

OADAS (TDTDASE), the model preparation area OADAS (TDTDASD), and the mobile

OADAS (TDTDASM). Together, these computer systems provide the functionality of acquiring, processing, recording and displaying data from the tunnel, the model preparation

18
FlightMode

Model Signals

Limit

Check

Latch

Debounce

WD Enable

==ResetValue (0)

Limit Table File

Snub!

GUI

Display

SNBXSTAT

Figure 12. WatchDog system snub logic block diagram.

area, and other locations where a mobile system may be required. These functions are

largely identical among all three systems. The mobile system was used during model preparation and the tunnel system was using during the three wind tunnel tests of the joined

wing model.

The tunnel data acquisition subsystem is responsible for collecting the data from several

different data acquisition units. The A/D conversions are provided by four NEFF 600 data

acquisition units (DAUs). Each DAU can scan up to 64 analog channels at an aggregate

rate of almost 100,000 samples per second. In addition, the first DAU also contains digital

channels for tunnel parameter and time code data.

5 WatchDog and Fault Detection Systems

Because of 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

including pneumatic snubber cables for cable mounted models [19]. 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 exceedances [22].

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

the new mount system set the AEI free flying tests among the riskier tests conducted in

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

and can be triggered manually. 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. Depending on experience and comfort

level, and the reaction time can be significant. The WatchDog system was developed to

keep the model and mount system from exceeding structural safety limits by monitoring

signals from the model and mount system.

Figure 9 shows the implementation of the WatchDog on dSpace1. 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 brake and pitch snubber and switches the control of model

control surfaces from dSpace2 to the emergency control law as shown in figure 9. The

system was originally developed for the flying wing tests [6] and was modified for use here.

As previously mentioned, the emergency controller used on the joined wing tests consisted

of simply commanding the control surfaces to 0◦.

19
Figure 12 shows a block diagram of the WatchDog snub logic. The system’s primary

function is to monitor a select set of model signals and issue the Snub! command when

parameter limits were exceeded. In addition to the model signals exceeding limits, the

Snub! command would be issued if a FlightMode value of 6 was detected. The FlightMode

values were designed to facilitate launch and recovery, incremental testing, stability margin

data collection and GLA testing. Based on the specific value of FlightMode, different sets

of WatchDog limits could be used for reset, takeoff, flight, or GLA testing. For instance,

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). Table 1 shows the FlightMode designations and integer values used in the JWS

tests. As indicated previously, the FlightMode parameter was not a direct user input. The

FlightMode logic block in figure 10 determined the appropriate value based on user inputs

and the vertical position of the model. Additional details will be covered in the model

launch, test, and recovery procedures section of the paper.

The selected signals and FlightMode dependent limits were compared in the limit check

block in figure 12. Detected faults were passed to the debounce block where 3 consecutive

frames of any particular signal fault were required to issue the WatchDog Snub! command.

The debounce 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 previous 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
cscript files for future use. Table 1 defines the various FlightMode limits and

lists the monitored parameters along with the final (end of Test 614) set of limits. As can

be seen from the table a total of 22 model signals were monitored directly: 12 wing control

surfaces, 1 rudder, 4 rigid body measurements, and 5 strain gauges, as well as, three ad hoc

or combined parameters. These parameters were an attempt to improve safety by applying

some engineering judgement. For instance, if the model had negative vertical position (below tunnel centerline) and positive vertical velocity (moving up), that combination 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. Similar logic was used to construct a parameter consisting of

Θ and PZ.

20
Table 1. FlightModes, WatchDog parameters, and parameter limits. Limits are in units of

deg, in, micro-strain, and seconds.

FlightMode Reset Taxi 1 Taxi 2 Transition Fly 1 Fly 2 Snub

0 1 2 3 4 5 6

Parameter Min Max Min Max Min Max Min Max Min Max

δRudder ±∞ −10 10 −10 10 −10 10 −10 10 −10 10 ±∞

δCS ±∞ −13.5 13.5 −13.5 13.5 −13.5 13.5 −13.5 13.5 −13.5 13.5 ±∞

V Z × P Z ±∞ −150 50 −150 150 −350 350 −400 200 −400 250 ±∞

V Z2 × P Z ±∞ −2000 1600 −2000 1600 −2000 1600 −2000 1600 −8000 8000 ±∞

(Θ − 1

o

)P Z ±∞ −40 40 −40 40 −40 20 −40 20 −40 25 ±∞

VZ ±∞ −20 20 −20 20 −20 20 −20 20 −40 40 ±∞

PZ ±∞ −22 −18.5 −22 −18.5 −22 0 −16 16 −16 16 ±∞

Θ ±∞ −1 3 −1 3 −0.5 3 0 3 0 3.25 ±∞

Θ ̇ ±∞ −15 15 −15 15 −15 15 −15 15 −30 30 ±∞

SGFW LI ±∞ −150 150 −150 150 −150 150 −150 150 −150 150 ±∞

SGFW RI ±∞ −150 150 −150 150 −150 150 −150 150 −150 150 ±∞

SGT F ±∞ −450 450 −450 450 −450 450 −450 450 −450 450 ±∞

SGTM ±∞ −200 200 −200 200 −200 200 −200 200 −200 200 ±∞

SGT A ±∞ −150 150 −150 150 −150 150 −150 150 −150 150 ±∞

21
An additional fault detection system was implemented along with the control laws on

dSpace2 to provide an alternative method of detecting an unstable or runaway condition.

This system monitors a combination of PZ and VZ, and a combination of the pitch angle

and rate. The state of the FlightMode parameter controlled the limits used. The position

and velocity terms were scaled by the limits so that ±1 indicated that the particular DOF

was at its limit. The scaled terms were squared and added to form circular error terms as

shown in equations 1 and 2. If the circular error term exceeded 1, the FlightMode was set to

the Snub! value of 6. When this was detected by the primary WatchDog, the Snub! signal

was sent to Snubber Control System and the model brake and pitch snubber were activated.

The values of the FlightMode dependent terms are listed in table 2.



P Z − P ZCenter

P ZRange 2

+



V Z

V ZLimit 2

≤ 1.0 (1)



Θ − ΘCenter

ΘRange 2

+

 

Θ ̇

Θ ̇ Limit !2

≤ 1.0 (2)

Table 2. Fault detection system parameters. Values are in units of deg, in, and seconds.

Values in parenthesis were initial values.

FlightMode Reset Taxi 1 Taxi 2 Transition Fly 1 Fly 2 Snub!

0 1 2 3 4 5 6

Parameter

ΘCenter 0.5 0.5 0.0 0.9 0.9 0.9 0.5

ΘRange 2.0 2.0 3.0 2.0 2.0 (0.5) 3.5 (2.5) 2.0

Θ ̇ Limit 15.0 15.0 15.0 15.0 (5.0) 15.0 (5.0) 50.0 15.0

P ZCenter −21.6 −21.6 −21.6 −8.4 0.0 0.0 −21.6

P ZRange 4.2 4.2 4.2 12.0 12.0 (8.0) 18.0 4.2

V ZLimit 18.0 18.0 24.0 30.0 18.0 60.0 18.0

6 Model Launch, Test, and Recovery

This section describes how the joined wing model was launched, tested, and recovered.

Figure 13 shows the nominal procedures and depicts a notional sequence where the dynamic

pressure is ramped up to 51 psf, tests are performed, the model is secured, and dynamic

pressure is returned to zero. The status of the snubber control system, the dSpace2 control

system, and the vertical position of the model (PZ) are shown. When parameters change

they are depicted in red. The yellow highlighted changes require manual inputs while the

others are automatic. Figure 13 shows the five groupings of events. These groupings are

described next.

Reset and PSNUB Release. Start with model snubbed. Tunnel dynamic pressure is set

to 51 psf, the relays (SNBXSTAT) and flight controller are reset, and the pitch snubber

(PSNUB) is released. At this point dSpace2 engages the THETA control loop, which

tracks a Θ set point value. Generally, the set point was around 0.5◦.

Theta Loop ParamID Testing. The operator sets TestMode to a value greater than 0.

Depending on the value of TestMode, excitation signals are added to various signals

within the Theta control loop and ParamID data are acquired. When ParamID data

22
PilotIn Trim CL FlightMode Reset TestMode GLAMode CG PZ Notes

0 TRIP SNUB ON 0 OPEN 6 0 0 0 5% < ZLT Model Secure, Tunnel Off

51 TRIP SNUB ON 0 OPEN 6 0 0 0 5% < ZLT Tunnel Conditions Set

51 RESET SNUB ON 0 OPEN 0 1 0 0 5% < ZLT Reset Systems

51 RESET SNUB ON 0 OPEN 1 0 0 0 5% < ZLT

51 RESET FREE ON 0 THETA 1 0 0 0 5% < ZLT Release Pitch Snubber

51 RESET FREE ON 0 THETA 2 0 >0 0 5% < ZLT Engage ParamID Testing Mode

51 RESET FREE ON 0 THETA 2 0 >0 0 5% < ZLT Acquire ParamID Data

51 RESET FREE ON 0 THETA 1 0 0 0 5% < ZLT Disengage from ParamID Testing Mode

51 RESET FREE ON 1 THETA 1 0 0 0 5% < ZLT

51 RESET FREE FREE 1 Z 1 0 0 0 5% < ZLT PZ Setpoint Starts Ramp to O"

51 RESET FREE FREE 1 Z 3 0 0 0 5% > ZLT PZ Based Transition

51 RESET FREE FREE 1 Z 4 0 0 0 5% > ZFT PZ Based Transition

51 RESET FREE FREE 1 Z 5 0 0 0 5, 0, -5,-10% ~ 0.0' Adjust CG to Provide Desired Stability

51 RESET FREE FREE 1 Z 5 0 ≥0 0 5, 0, -5,-10% ~ 0.0' Engage ParamID Testing Mode

51 RESET FREE FREE 1 Z 5 0 ≥0 ≥0 5, 0, -5,-10% ~ 0.0' Select GLA Control Law

51 RESET FREE FREE 1 Z 5 0 ≥0 ≥0 5, 0, -5,-10% ~ 0.0' Acquire ParamID/GLA Data

51 RESET FREE FREE 1 Z 4 0 0 0 5, 0, -5,-10% ~ 0.0' Disengage from Test/GLA Modes

51 RESET FREE FREE 1 Z 4 0 0 0 5% ~ 0.0' Set CG to 5% Static Margin

51 RESET FREE FREE 0 Z 4 0 0 0 5% ~ 0.0' PZ Setpoint Starts Ramping Down

51 RESET FREE FREE 0 Z 3 0 0 0 5% < ZFT PZ Based Transition

51 RESET FREE FREE 0 THETA 1 0 0 0 5% < ZLT PZ Based Transition

51 TRIP FREE FREE 0 OPEN 6 0 0 0 5% < ZLT Engage Brake and Pitch Snubber

51 TRIP SNUB ON 0 OPEN 6 0 0 0 5% < ZLT Engage Overide Switches

0 TRIP SNUB ON 0 OPEN 6 0 0 0 5% < ZLT

ZLT=Z-Landing-Threshold: PZ=-18.25" ZFT=Z-Fly-Threshold: PZ=-6.0" Tunnel Center Line: PZ=0.0" Red= Automatic Change Red = Manual Change

Free Flying 

ParamID 

and/or GLA 

Testing

Land Model 

and Secure

dSpace Controls/States

Reset & 

PSNUB 

Release

Theta Loop 

ParamID 

Testing

Launch

Snub/Brake Control

SNBXSTAT PSNUB ZSNUB

Test

Group Q

Figure 13. Description of the launch and test procedures used with the JWS model.

acquisition is complete, TestMode is set back to zero. This set of sequences was used

early in the test program, but it was generally skipped.

Launch. The operator changes the PilotIn parameter from 0 to 1 and then releases the

brake (ZSNUB). The control system engages the Z control loop and ramps up the PZ

set point to a centerline value of 0. Along the way, FlightMode changes automatically

as the model passes through certain PZ thresholds.

Free Flying ParamID and/or GLA Testing. Using the movable mass system, the CG

is remotely adjusted to provide the desired static margin. The operator selects a

TestMode value to determine the type of ParamID testing to be conducted and selects

the GLAMode to determine which, if any, GLA control law is to be tested. When

ParamID and/or GLA testing is complete, TestMode and GLAMode are returned to

0.

Land Model and Secure. The CG is remotely adjusted to the full forward position providing +5% static margin. The operator then sets PilotIn to 0 telling the controller

to land the model. The PZ set point is ramped down. At certain PZ thresholds,

FlightMode transitions to 3, then to 1 where the Z control loop is disengaged. Finally,

the operator trips the snubber relays (SNBXSTAT) and engages the override switches

(PSNUB and ZSNUB) to secure the model.

Figure 14 provides an example set of data acquired during model launch. The upper plot

shows time histories of PZ, Θ, and aft wing control surface commands. The lower plot shows

the three snubber related values along with FlightMode. Snubber values of 0 VDC indicate

a snubbed or tripped state while a value of approximately 8 VDC indicates an unsnubbed

or reset condition. The launch procedure described above and in figure 13 is followed except

that ParamID testing is not shown and the time histories end once the model has achieved

flight.

It should be pointed out that the nominal sequence shown in figure 13 was almost

never followed exactly. Typically, the WatchDog system would arrest the model during

ParamID or GLA testing. When this happened the pitch snub (PSNUB) and brake (ZSNUB)

override switches were set to the snubbed/brake-on positions and tunnel dynamic pressure

23
PZ, in. 

or

δCmd, deg.

20 25 30 35 40 45 50 55 60

3

2

1

0

-1

Θ, deg.

PZ

δCmd

Θ

10

8

6

4

2

0

25

0

-25

Time, sec.

Voltage

or 

Mode

SNBXSTAT

PSNUB

ZSNUB

FightMode

Figure 14. Data acquired during model launch. Test 614, point 1655. δCmd is command

signal sent to all aft wing control surfaces.

was reduced to about 42 pfs. The plunge override power switch and the plunge override

switch were then used to momentarily release the brake to allow the model to settle to the

bottom stop. At that point, tunnel dynamic pressure would be returned to 51 psf, and the

sequence in figure 13 would resume from the top.

7 Trim Controller Architecture

The trim controller was designed to launch, fly and land the model in the wind tunnel and

to serve as the reference for GLA performance. The trim controller consisted of two main

elements, a vertical (Z) feedback loop and a pitch (Θ) feedback loop. The vertical loop

consisted of a simple PID controller that generated an acceleration command. The pitch

loop consisted of a PID plus acceleration feed-forward controller to generate a control surface

command from the acceleration command. These two loops used gains that were scheduled

based on CG. The single control surface command was passed through a third-order low

pass filter to attenuate the response at higher frequencies. The initial set of gains were

determined using an analytically derived aeroservoelastic model.

Figure 15 shows a diagram of the trim controller. As indicated, the trim controller has

two modes of operation controlled by the azMode parameter: Pitch control only or altitude

(vertical position) control. Figure 15 also shows the routing of the excitation signal to the

control surface command that is used in TestMode 3. The excitation signal along with the

unaugmented control surface command were routed to the DAS for near real time assessment

of controller stability. Other test modes add the excitation to other signals within the trim

control block diagram or the flight control block (figure 10). For clarity the details on these

additional test modes and the associated signals that were also routed to the DAS are not

shown. The single control surface command signal shown in figure 15 was sent to all six aft

wing control surfaces.

24
+ +

+

+

Integrator zigain

Altitude Control

Fuselage

Accelerations

Pitch Control

Cmd

PZ

VZ

θ

Forward

Aft

kzi 1

- s

zpgain signchange

Acc Blend

kzp

zdgain

kzd

k*u k*u

k*u

+

-

+

+

+

deg per Integrator

ft/sec

pitchgain

kpi

pitchpgain

kpp

pitchdgain

kpd

+

pitchd1gain

kpff

1

s

trimfiltnum(z)

trimfiltden(z)

CmdFilter

ActCmd

Low Pass

Filter

Az Feed-forward

PZ

+

-

.

θ

θ

Cmd Mode Select

(azMode)

+

+

Excitation

TestMode 3

Break Point

Parameters

Zeroed or Added 

Based on PSNUB 

To DAS

To DAS

Parameters

Zeroed or Added 

Based on ZSNUB 

Figure 15. Trim control law diagram.

8 TDT Test 613: Free Flying Function Check and ParamID

The primary objectives of the first free flying test of the joined wing model were to (1) assess

the functionality of the model mated to the carriage and beam structure, (2) demonstrate

and verify the model launch, test, and recovery procedures, and (3) acquire parameter

identification (ParamID) data. Figure 16 shows the nominal test matrix used for Test

613. As indicated, the model could be tested in one of three different boundary conditions:

restrained (pitch snubber and brake engaged), free-in-pitch (brake engaged), and flying.

The data acquired in the restrained and free-in-pitch configurations were used largely for

system evaluation in preparation for flying the model. The free flying data were the highest

priority, and the majority of data was acquired in this configuration.

Most of the data acquired in Test 613 used GLAMode 0, in which only the basic trim

control law was used. GLAMode values of 1 and greater denote the use of various GLA

control laws. TestMode refers to the type of ParamID testing performed. TestMode 1 was

used with the AOS and control surface excitation signals. TestMode values greater than

1 indicated that the excitation signal was being injected at some point within the flight

control system. Most of the data acquired during Test 613 consisted of linear sine sweeps

as shown in figure 16. In general, the amplitudes of excitation were maximized based on

trial and error. In each case, the goal was to obtain the maximum model response without

exceeding WatchDog limits. The amplitude of the AOS vanes was fixed at their maximum

value of 12◦ peak-to-peak. The typical values of control surface excitation are shown, but

sometimes larger or smaller amplitudes were used. Specific amplitudes for TestModes of 2

through 11 are not listed because they varied greatly depending on CG location and other

parameters. The block of data on the right-hand side of figure 16 was repeated for each

movable mass location or static margin value. Acquiring data at all the combinations of

parameters indicated in the test matrix was in fact not done. Model configurations and

data types were prioritized. Some sample wind tunnel data will be shown in the Test 614

section of the paper.

An analytical aeroservoelasticity (ASE) model was developed for use in determining

the trim controller gains. The initial modeling was based purely on analysis updated with

25
!"#$%&'( ")*(+, "&- .,(/0(*123$45 #678)90'( .,(/0(*123$45 #678)90'(

!"#$%&'"()*

**!+,,"-*!+./+-",0" 1 1 1 2

**345*6+#$ 1 2 78"--# 9:

**345*6+#$* 1 1 1 2 92# 2;(<(= 9:

**>?4*30$+@$&.* 1 1 1 2 92# 2;(<=2 A:

**>?%*30$+@$&.* 1 1 1 2 92# 2;(<=2 A:

**>?B*30$+@$&.* 1 1 1 2 92# 2;(<=2 A:

**>?*3--*30$+@$&. 1 1 1 2 92# 2;(<=2 =;C:

**3?4*30$+@$&.* 1 1 1 2 92# 2;(<=2 A:

**3?%*30$+@$&. 1 1 1 2 92# 2;(<=2 A:

**3?B*30$+@$&. 1 1 1 2 92# 2;(<=2 A:

**3?*3--*30$+@$&. 1 1 1 2 92# 2;(<=2 =;C:

**3--*30$+@$&. 1 1 1 2 92# 2;(<=2 =;C:

!"#$%&'"*= 1 2 92# 2;(<=2D*2;(<A E@1FEFG"'

!"#$%&'"*H 1 1 2 92# 2;(<=2D*2;(<A E@1FEFG"'

!"#$%&'"*A 1 2 92# 2;(<=2D*2;(<A E@1FEFG"'

!"#$%&'"*C 1 2 92# 2;(<=2D*2;(<A E@1FEFG"'

!"#$%&'"*9 1 2 92# 2;(<=2D*2;(<A E@1FEFG"'

!"#$%&'"*I 1 2 92# 2;(<=2D*2;(<A E@1FEFG"'

!"#$%&'"*J 1 2 92# 2;(<=2D*2;(<A E@1FEFG"'

!"#$%&'"*K 1 2 92# 2;(<=2D*2;(<A E@1FEFG"'

!"#$%&'"*(2 1 2 92# 2;(<=2D*2;(<A E@1FEFG"'

!"#$%&'"*(( 1 2 92# 2;(<=2D*2;(<A E@1FEFG"'

:&0*'+,2$;&*')<&*

=9+<1$%+,-)*$>$?@3$A3$B@3$C$BDAE

=266(9,)1 #*<F266(9,)1

G(F9,+)*('

.,((B)*BH)91I

.82)*-

=J((7

Figure 16. Nominal Test matrix used in TDT Test 613.

static aerodynamic terms determined from the JWS balance test, TDT Test 606. This ASE

model, along with the previously described flight control block, support system dynamics,

and a model for the AOS-generated gust input were part of a Simulink model. Reference 13

describes this model in more detail. This model was used to develop the initial gains for

the trim controller. With those gains, the controller was determined to be stabilizing at all

model configurations (CG locations). The original gains were designed with 9 dB of gain

margin and 45◦ of phase margin. The actual margins obtained from testing were lower than

the predicted margins. Data acquired using TestMode 3 were used to estimate margins of

6 dB and 18◦ at the -10 percent static stability condition. Figure 17 shows the Nichols plot

for +5 and −10 percent static stability from both the simulation model and test data.

There were several false starts in developing the best approach to using the wind tunnel

data to refine the ASE model. The technique that was ultimately used involved the use

of an optimizer to tune the simulation model to match the test data. This method used

nine sets of actuator sweep time histories (six sets of symmetric actuator pairs, all forward

actuators, all aft actuators. and all forward and aft surfaces) and one set of gust sweep time

histories. These data were acquired using TestMode 1 with the model flying via the trim

controller. The optimization technique required a large amount of memory and processor

time to complete. A typical solution for one CG condition would take approximately 4 CPU

days. The solutions obtained from this method produced models with generally very good

agreement with the stability margins produced from test data and an accurate match of the

first symmetric structural mode. The final linear model contained 18 modes and accurately

reproduced the measurements required for the trim and GLA controllers.

26
−360 −300 −240 −180 −120 −60

20

15

10

5

0

-5

-10

-15

-20

-25

Phase, deg

Magnitude, db

 0.12 Hz

 0.52 Hz

 1.01 Hz

2.01 Hz

 3.02 Hz

0.21 Hz

0.21 Hz

 2.01 Hz

0.12 Hz 3.02 Hz

 2.01 Hz

 0.21 Hz

Nominal Margins

Robust Margins

Simulation +5% CG

Simulation -10% CG

Data Point 412 +5% CG

Data Point 438 -10% CG

 0.12 Hz

 1.01 Hz

 1.01 Hz

 0.21 Hz

 1.01 Hz

 0.52 Hz

 0.52 Hz

 0.52 Hz

Figure 17. Nichols stability plot for initial trim controller gains with +5 and −10 percent

static stability. Test 613 data points 412 and 438 are plotted along with simulation results.

27
!"#$%&' "()'*+ "%, -+'./')012345 #678(9/&' -+'./')012345 #678(9/&'

!"#$%&'"()*

**!+,,"-*!+./+-",0" 1 1 1 23*(3*4555* 6 6

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**789*:+#$* 1 1 1 23*(3*4555* (42# 25(6(4 =>

**?@8*70$+A$&.* 1 1 1 23*(3*4555* =2B(42# 25(6423*25(6C C> 25(6423*25(6C C>

**?@%*70$+A$&.* 1 1 1 23*(3*4555* =2B(42# 25(6423*25(6C C> 25(6423*25(6C C>

**?@D*70$+A$&.* 1 1 1 23*(3*4555* =2B(42# 25(6423*25(6C C> 25(6423*25(6C C>

**?@*7--*70$+A$&. 1 1 1 23*(3*4555* =2B(42# 25(6423*25(6C 45E>

**7@8*70$+A$&.* 1 1 1 23*(3*4555* =2B(42# 25(6423*25(6C C> 25(6423*25(6C C>

**7@%*70$+A$&. 1 1 1 23*(3*4555* =2B(42# 25(6423*25(6C C> 25(6423*25(6C C>

**7@D*70$+A$&. 1 1 1 23*(3*4555* =2B(42# 25(6423*25(6C C> 25(6423*25(6C C>

**7@*7--*70$+A$&. 1 1 1 23*(3*4555* =2B(42# 25(6423*25(6C 45E>

**7--*70$+A$&. 1 1 1 23*(3*4555* =2B(42# 25(6423*25(6C 45E>

!"#$%&'"*4 1 23*(3*4555* =2B(42# 25(6423*25(6C FA1GFGH"'

!"#$%&'"*I 1 1 23*(3*4555* =2B(42# 25(6423*25(6C FA1GFGH"'

!"#$%&'"*C 1 23*(3*4555* =2B(42# 25(6423*25(6C FA1GFGH"'

!"#$%&'"*E 1 23*(3*4555* =2B(42# 25(6423*25(6C FA1GFGH"'

!"#$%&'"*= 1 23*(3*4555* =2B(42# 25(6423*25(6C FA1GFGH"'

!"#$%&'"*J 1 23*(3*4555* =2B(42# 25(6423*25(6C FA1GFGH"'

!"#$%&'"*K 1 2

!"#$%&'"*L 1 23*(3*4555* =2B(42# 25(6423*25(6C FA1GFGH"'

!"#$%&'"*(2 1 43*I3*CM =2B(42# 25(6423*25(6C FA1GFGH"'

!"#$%&'"*(( 1 2

:%/)&*+13;%)&(<%)

=9*<03$*+,()3>3?@23A23B@23C3BDAE

=166'9+(0 #)<F166'9+(0

G'F9+*()'&

-+''B()BH(90I

-81(),

=J''7

Figure 18. Nominal test matrix used in TDT Test 614.

9 TDT Test 614: Free Flight Testing – ParamID and

GLA

The primary objective of the second free flying joined wing test was to demonstrate GLA.

Based on the lessons learned and a review of the data acquired in Test 613, the nominal test

matrix used in Test 613 was modified as shown in figure 18. In general, logarithmic sweeps of

longer duration were used. The sweeps were better suited to the previously described ASE

model optimization procedure. In addition, some antisymmetric control surface data were

acquired. The most important change to the test matrix was the inclusion of GLAMode

values of 1 and greater indicating that GLA control laws were evaluated.

Figure 19 shows a sample ParamID data set used for ASE model refinement.

GLAMode 0 and TestMode 1 were used for this data set. All control surfaces received

the same 2.5◦excitation signal. Figure 19 shows time histories of the model rigid body

positions (PZ and Θ), a select set of strain gauges and accelerometers, and the wing control

surface positions. The fuselage accelerometers are included as they are used as inputs to

the trim control law, and the forward wing inboard and tail root strain gauges are shown as

these signals were used as inputs to the most successful version of the GLA control law. The

aft wing control surface positions are a sum of the trim controller and the excitation signal.

The trim controller did not use the forward wing surfaces, so what is plotted is simply the

sum of the bias values and the excitation signal. It should be pointed out that the forward

wing control surfaces include the biases to help trim the model and bring the mean aft wing

control surface commands closer to zero. Using the forward wing surfaces to help trim the

model was an important part of testing in general, with different biases being used with

different CG locations (static margins). This trimming procedure was necessary to keep the

aft wing control surface commands from exceeding their physical limits.

Figures 20 and 21 show model responses to AOS sweeps. The format of these plots is

essentially the same as what was previously described in figure 19 except that the frequency

28
Θ, deg.

10 20 30 40 50 60 70

Strain Gauges: Blue=Forward Wing Left Inboard, Green=Forward Wing Right Inboard, & Red=Tail Root

1.0

0.5

0

-0.5

400

200

0

15

10

5

0

-5

-10

-15

15

10

5

0

-5

-10

-15

15

10

5

0

-5

-10

-15

4

3

2

1

0

Fuselage Accelerometers: Blue=Aft, Green=Forward

Aft Wing Control Surfaces : Blue=Outboard, Green=Midboard, Red=Inboard

0

Forward Wing Control Surfaces: Blue=Outboard, Green=Midboard, Red=Inboard

Time, sec.

PZ, inches Microstrain Acceleration, G’s Position, deg. Position, deg.

Figure 19. Symmetric, all wing control surface sweep 0.1–20 Hz. Amplitude = 2.5◦, Test

614, data point 1019, GLAMode 0, TestMode 1, static margin = −10 percent.

29
of the AOS vanes is also shown in the upper plots. The data plotted in figure 20 was

acquired using GLAMode 0 and TestMode 1. The data acquired in figure 21 was acquired

using GLAMode 5v6 and TestMode 1. Forward wing outboard control surface motion can

be seen in figure 21 where GLA was engaged. These data points were selected because they

are among the last sets of data acquired during Test 614. These same points will be further

analyzed in the frequency domain below.

Test 614 was conducted over a period of approximately six weeks. Throughout this test,

both the trim controller and the GLA control law were continuously refined. The general

process was to design, implement, test, and evaluate the trim and GLA controllers. This

process was repeated multiple times during the testing period. As the testing progressed,

improved testing procedures were developed including more optimal trim values for the

control surfaces and expanded WatchDog parameters, which allowed better ParamID data

sets to be acquired. These data sets were used to further refine the ASE model helping to

improve the trim and GLA controller designs.

The trim controller was tuned to improve the stability margins and general performance

based on the refined simulation model. The stability margin improvements were designed

to achieve a nominal stability margin target of ±6 dB of gain margin and ±30◦ of phase

margin at all four of the target CG positions. The design margins were generally achieved,

although the nonlinear break in the slope of the pitching moment associated with the aft

wing control surfaces tended to shift the pitch loop gain up and down depending on the

required trim deflections. The nonlinear responses are shown in the appendix.

The development of the GLA controller used a classical design approach. This process

provided a more direct way to address this specific GLA problem. This approach also

facilitated the controller being tuned based on actual test results that were not captured in

the simulation modeling. Fourteen GLA control laws were designed, tested, and evaluated.

The first eight control laws were implemented as separate modes as they were developed

and could be selected by modifying the GLAMode input. The final six GLAModes altered

versions of GLAModes 4 and 5. The GLAModes were known as: 2, 6, 7, 8, 9, 3, 4, 5, 4v2,

5v2, 5v3, 5v4, 5v5, and 5v6; and were developed and tested in the order listed. Reference 13

includes detailed descriptions of the various GLA control law designs. The GLA controllers

included alterations to the mix of control surfaces used by the trim controller and the

inclusion of forward wing inboard and tail boom root strain gauge feedback, along with

some basic first and second order filters.

The GLA controller with the best performance in terms of minimizing strain response

was GLAMode 5v6. This controller used feedback from the averaged forward wing inboard

strain gauges along with the tail boom root strain gauge. Figure 22 shows a block diagram

of this control law. It used three inputs that were combined, filtered and then distributed

to four control surfaces. The inputs were mixed together using weights that enhanced the

observability of the first forward wing bending mode while reducing the observability of the

second forward wing bending mode. The result was then passed through a series of filters.

The first filter was a second order high pass filter with a frequency of 1 Hz and a damping

ratio of 0.5. The second filter consisted of a first order lead with a frequency of 3 Hz and a

first order lag with a frequency of 3.57 Hz. The third filter consisted of a first-order lag with

a frequency of 20.23 Hz. This filtered result was then scaled and distributed to the forward

wing outboard surfaces and the aft wing inboard surfaces.

Figure 23 shows plots of the frequency response of the average forward wing strain and

the tail root strain acquired during AOS sweeps with GLAMode 0 (trim control only) and

GLAMode 5v6 (trim control + GLA). Each plotted data set is composed of two data records

that contained a sweep from 0.1 to 12 Hz and a sweep from 3 to 12 Hz. Here, figures 23a

and b show a 5 db reduction in the first flexible mode forward wing inboard strain response

and a 6 dB of reduction in the first flexible mode tail strain response. These reductions were

achieved without degrading (increasing response) in other modes.

30
Inches or Hz

PZ

AOS Freq

Microstrain

Strain Gauges: Blue=Forward Wing Left Inboard, Green=Forward Wing Right Inboard, & Red=Tail Root

Fuselage Accelerometers: Blue=Aft, Green=Forward

Acceleration, G’s

Aft Wing Control Surfaces : Blue=Outboard, Green=Midboard, Red=Inboard

Position, deg.

0 20 40 60 80 100 120

Forward Wing Control Surfaces: Blue=Outboard, Green=Midboard, Red=Inboard

Time, sec.

Position, deg.

Θ, deg.

15

10

5

0

-5

-10

-15

15

10

5

0

-5

-10

-15

1.0

0.5

0

-0.5

-1.0

15

10

5

0

-5

-10

-15

5

0

-5

400

300

200

100

0

Figure 20. AOS sweep 012–.1 Hz. Test 614, point 1750, GLAMode 0, TestMode 1, static

margin = −10 percent.

31
400

300

200

100

0

Microstrain Acceleration, G’s

1

0.5

0

-0.5

-1

15

10

5

0

-5

-10

-15

15

10

5

0

-5

-10

-15

Position, deg.

0 20 40 60 80 100 120

Time, sec.

Position, deg.

Strain Gauges: Blue=Forward Wing Left Inboard, Green=Forward Wing Right Inboard, & Red=Tail Root

Aft Wing Control Surfaces : Blue=Outboard, Green=Midboard, Red=Inboard

Forward Wing Control Surfaces: Blue=Outboard, Green=Midboard, Red=Inboard

Inches or Hz

PZ

AOS Freq

15

10

5

0

-5

-10

-15

Θ, deg.

5

0

-5

Fuselage Accelerometers: Blue=Aft, Green=Forward

Figure 21. AOS sweep 12–0.1 Hz. Test 614, point 1765, GLAMode 5v6, TestMode 1, static

margin = −10 percent.

32
SGFWLI

SGTF

Trim Control Inputs

Trim + GLA 

Control Inputs

+

+

SGFWRI Input

Mixer

Output

Mixer

MFWO

0

0

0

0

MFWO

0

0

MAWI

MAWI

0

0

0

s

2

s

2 + 2ζω1

s + ω1

2 s + ω3

s + ω2

s + ω4

k

Figure 22. GLAMode 5v6 block diagram.

10-1 100101

40

35

30

25

20

15

10

5

0

AOS Excitation Frequency, Hz

Trim+GLA (GLAMode 5 v6)

Trim (GLAMode 0)

Magnitude, db 

(a) Average forward wing inboard strain.

10-1 100101

60

55

50

45

40

35

30

25

20

15

10

Magnitude, db 

AOS Excitation Frequency, Hz

Trim+GLA (GLAMode 5 v6)

Trim (GLAMode 0)

(b) Tail root strain.

Figure 23. Tail root and forward wing root strain response to AOS. Test 614, data points

1750, 1769, 1765, and 1770.

33
−540 −480 −420 −360 −300 −240 −180 −120 −60

10

0

-10

-20

-30

-40

-50

Phase, Deg.

Magnitude, db

0.06 Hz

0.52 Hz

5.00 Hz

5.00 Hz

10.01 Hz

15.01 Hz

0.12 Hz 0.21 Hz

1.01 Hz

2.01 Hz

3.02 Hz

10.01 Hz

Nominal Margins

Robust Margins

GLA 0 (pt. 1774)

GLA 5 v6 (pt. 1773)

0.03 Hz

(a) Nominal view.

−240 −210 −120 −180 −150

10

8

6

4

2

0

-2

-4

-6

-8

-10

Phase, deg

Magnitude, db

0.03 Hz

0.52 Hz

1.01 Hz

2.01 Hz

0.06 Hz

0.06 Hz

3.02 Hz

3.02 Hz

(b) Detail.

Figure 24. Nichols plot of trim (GLAMode 0) and trim+GLA (GLAMode 5v6) controllers.

Test 614, data points 1774 and 1773.

A Nichols plot for the −10 percent static stability margin case where TestMode 3 was

used to acquire the data is shown in figure 24a. This plot includes both data from the trim

controller only and for the trim controller with GLAMode 5v6 active. The plot shows that

there is little impact to the pitch loop due to the addition of the GLA controller. A more

enhanced view of the stability margins is shown in figure 24b, which shows the desired 6 dB

of gain margin and 30◦ of phase margin.

10 Concluding Remarks

In the summer of 2010, the Air Force Research Laboratory (AFRL), Boeing, 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 joined wing SensorCraft vehicle

concept. The first of these tests was conducted on sting and balance. The second and

third tests used a new, two-DOF support system. This model support allowed the full-span

model to translate vertically and rotate in pitch, allowing better simulation of the vehicle’s

rigid-body modes. The large size of the joined 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 test objectives. This paper has provided an

overview of the AEI joined wing wind tunnel tests. Descriptions of the hardware including

the wind tunnel, the wing model, the model support system, and other supporting systems

developed for this program were provided. The design of the trim and gust load alleviation

control systems were also described. Trimmed flight at −10 percent static margin and a

reduction of structural response of at least 50 percent were successfully demonstrated.

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.

34
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. Lockyer, A. J., Drake, A., Bartley-Cho, J., Vartio, E., Solomon, D., and Shimko, T.,

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

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Wind Tunnel Model,” 26th AIAA Applied Aerodynamics Conference, No. AIAA-2008-

7191, Honolulu, Hawaii, Aug. 2008.

8. 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.

9. 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.

10. 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.

11. 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.

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

7190, Honolulu, Hawaii, Aug. 2008.

13. Reichenbach, E., Castelluccio, M., Sexton, B., and Scott, R., “Aerodynamic Efficiency

Improvement Final Report Volume 2: Aeroservoelastic Test Program,” AFRL Contractor Report FA8650-05-C-3500, Boeing Research and Technology Computational Sciences, 2010.

35
14. Reichenbach, E., Castellucio, M., and Sexton, B., “Joined-Wing SensorCraft Aeroservoelastic Wind Tunnel Test Program,” 52nd AIAA/ASME/ASCE/ASC Structures,

Structural Dynamics and Materials Conference, No. AIAA-2011-1956, Denver Colorado,

April 2011.

15. Scott, M. J., Enke, A., and Flanagan, J., “SensorCraft Free-Flying Aeroservoelastic

Model: Design and Fabrication,” 52nd AIAA/ASME/ASCE/ASC Structures, Structural Dynamics and Materials Conference, No. AIAA-2011-1957, Denver Colorado,

April 2011.

16. Sharma, V. and Reichenbach, E., “Development of an Innovative Support System for

SensorCraft Model,” 52nd AIAA/ASME/ASCE/ASC Structures, Structural Dynamics

and Materials Conference, No. AIAA-2011-1958, Denver Colorado, April 2011.

17. Heeg, J. and Morelli, G., “Evaluation of Simultaneous Multi-sine Excitation of the

Joined Wing Aeroelastic Wind Tunnel Model,” 52nd AIAA/ASME/ASCE/ASC Structures, Structural Dynamics and Materials Conference, No. AIAA-2011-1959, Denver

Colorado, April 2011.

18. Scott, R., Castelluccio, M., Coulson, D., and Heeg, J., “Aeroservoelastic Wind-Tunnel

Tests of a Free-Flying, Joined-Wing SensorCraft Model for Gust Load Alleviation,” 52nd

AIAA/ASME/ASCE/ASC Structures, Structural Dynamics and Materials Conference,

No. AIAA-2011-1960, Denver Colorado, April 2011.

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

Working Paper LWP-799, Sept. 1969.

20. 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.

21. 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.

22. 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.

36
Appendix A

Control Surface Actuator Description, Bench Testing,

and In-service Performance

The flight control surface actuators are key elements in the overall flight control system.

They need to be fast, reliable, and compact enough to fit within the wings. The control

surface actuator requirements were established based on demonstrated capabilities in similar

applications, the specific requirements of the JWS model, and preliminary estimates of the

requirements to achieve the GLA objectives. Table A1 shows the actuator requirements.

Table A1. JWS actuator requirements

Metric Value

Travel limit ≥ ±15 ◦

Phase shift at 20 Hz with ± 2

◦

command ≤ 30◦

Phase shift at 30 Hz with ± 2

◦

command ≤ 60◦

Gain at ≤ 25 Hz with ± 2

◦

command ≤ −2 dB

Hysteresis ≤ 0.1◦

Max torque ≥ 120 in-oz

The travel limits were set based primarily on the limitations imposed by the available

thickness of the wing and supported by preliminary GLA simulations of the model response

to the gust input. The phase shifts and gain (referred to together as bandwidth) were chosen

so that the control surfaces would be effective at the first mode frequency of approximately

5 Hz. The hysteresis requirement was based on past operational experience with similar

actuator designs along with the need to maintain linear behavior for the trim and GLA

control laws. The torque requirement was specified to account for the maximum hinge

moment on the control surface and to provide the required acceleration to achieve the

bandwidth.

Electric and hydraulic actuation methods were investigated. The performance of each

type was evaluated against the actuator requirements to identify a suitable candidate. Electric actuation was the initially favored method because of its perceived simplicity and recent

advances in the electric actuation technology. A number of electric actuator types were reviewed including hobby servos, military unmanned aerial vehicle (UAV) servos, and voice

coil actuators. The hobby servos had enough torque, but were too slow. The UAV servos

had a lot of torque and were faster than the hobby servos, but were still too slow, approximately 12-Hz bandwidth. The voice coil actuators had enough potential bandwidth and

could be ganged together for enough torque, but the size of the actuators would have made

packaging difficult to implement in the model wings. Hydraulic actuation had the advantage of historical demonstration of the required performance, recent operational experience

of the TDT, and design and manufacturing experience of the model subcontractor, NextGen

Aeronautics.

In the end, the electric actuation option was determined to be too risky to pursue,

and hydraulic actuation was selected for the JWS control surface actuators. The hydraulic

actuation system consisted of a rotary vane actuator, a servo valve, a position transducer,

and the associated electronics and plumbing. The operating pressure was selected to be

1,000 psi. The rotary vane actuator was designed and built by NextGen with a 0.5-in

square vane and a ±15◦travel. The servo valve was initially baselined as a Moog Series 30

37
servo valve with the Moog E024 servo valve later replacing 10 of the 13 Series 30 servo valves

in the design because of the E024s smaller size and lighter weight. The position transducer

selected was a RSYN-8-30 RVDT with the associated signal conditioner, LDM-1000. After

a significant design effort combined with much trial and error, the previously described

performance objectives were largely met.

A.1 Control Surface Actuator Bench Testing

A test bench was created to evaluate the actuation system. The test bench can be configured

to test both the Moog Series 30 and the EO24 servo valves with the actuator connected to

the servo valve by a single manifold (0-in hydraulic tubes) or using separate manifolds for

the actuator and servo valve using 12-in or 36-in of hydraulic tubes. The goal was to ensure

that the actuator performance would be sufficient at the extremes of the potential wind

tunnel model configurations.

The actuator test bench consists of several components. These include the integrated

manifold, a remote servo valve manifold, and a number of adaptors and tubes. In all

configurations, the actuator mounts directly to the integrated manifold while the servo

valve can attach directly to the integrated manifold or to the remote servo valve manifold.

Figure A1 shows these components in several different configurations. Figure A1a shows

an actuator with the end-caps removed and the shaft and vane visible. In the background

of this photo, a series 30 servo valve is shown mounted to the integrated manifold with

hydraulic supply and return tubes attached. Figure A1b shows the actuator attached to the

integrated manifold. An EO24 servo valve is shown in figure A1c connected via the 12-in

tubes and remote servo valve manifold.

For the bench tests, the PID gains on the control loop were established by tracking

a square wave command. Command and position are plotted in near real time on the

dSpace1 host system, so the procedure for setting the gains is to start with zero integral

and derivative gains, increase the proportional gain until actuator instability and then back

off approximately 20 percent. The derivative gain is then adjusted to improve stability.

Integral gain was not found to be necessary and was set to zero. The same proportional

and derivative gains were used for all the Series 30 and EO24 servo valve experiments.

The initial series of experiments were designed to determine the optimal settings for

the LDM 1000 signal conditioner. Reference 13 describes these experiments in detail. Ultimately, it was determined that the best actuator response, in terms of phase lag, was

obtained with the LDM 1000 filter set to 1000 Hz and the excitation frequency set to

2.5 KHz. Reference 13 also documents the results of experiments with simulated control

surface inertia. Control surface inertia was found to have a small effect on actuator performance.

The bench testing showed that the actuator performance was sufficient in each of the

hydraulic tube length configurations. Figure A2 shows the phase and gain responses at four

conditions. The reference black line shows the response with the integrated manifold with

a 1◦ amplitude excitation. The blue line shows the response with the integrated manifold

with a 3◦ amplitude excitation. The green line shows the response with the 12-in long

hydraulic tubes with a 3◦ amplitude excitation. The red line shows the response with 36-in

long hydraulic tubes with the 3◦excitation. The sharp breaks in the phase plots are an

indication of the servo valve hitting its flow rate limit. The responses for the longer hydraulic

tubes show that the flow rate limit is hit earlier indicating that there are additional pressure

drops associated with the use of the longer tubes although there is little difference between

the 12-in and 36-in cases. At the time, this lack of difference was deemed satisfactory. Later

review revealed that the servo valve manifold used for the 12-in and 36-in test cases differed

from the integrated (0-in) manifold in a critical way. The pictures of the two manifolds

shown in figure A3 show that the manifold used for the 12-in and 36-in test cases (left)

38
(a) Disassembled rotary vane actuator. Integrated manifold with series 30 servo valve installed shown in

the background.

(b) Integrated manifold with actuator and RVDT

installed. The remote servo valve adaptor with 12-

in tubes is also attached to the actuator manifold.

(c) Remote servo valve manifold with EO24 servo

valve and 12-in tubes installed.

Figure A1. Photographs of actuator bench test rig in two different configurations.

39
20

15

10

5

0

-5

-10

0

-20

-40

-60

-80

-100

-120

-140

-160

-180 Magnitude, db

101 102

Frequency, Hz.

Flow Rate Limited

Phase, Deg.

0" Hydraulic Tubing, ±1°

0" Hydraulic Tubing, ±3°

12" Hydraulic Tubing, ±3°

36" Hydraulic Tubing, ±3°

Flow Rate Limited

Figure A2. Actuator test bench performance data. Moog Series 30 servo valve.

had very small ports compared to those on the integrated manifold (right). The much

smaller ports add a significant amount of pressure drop, thereby decreasing the peak flow

rate through the servo valve and providing an explanation for the sharp drop in performance

when switching to the longer hydraulic tubes.

A.2 Model Actuator Frequency Response

The joined wing model arrived at the TDT in September 2008. Getting the actuators to

perform reliably and consistently proved to be a challenge. Eventually, the control surface

actuators reached a minimal state of functionality such that they could be used for static

testing, but, when surfaces were operated simultaneously, their frequency response was well

below what was observed in the bench test experiments. This performance was not a major

problem for Test 606 as the primary purpose of this was to acquire steady data. After the

completion of Test 606, in December 2008, the model was returned to the lab where actuator

performance was reassessed.

Analyses of the pressure drops throughout the hydraulic system revealed the need for

some modifications to the internal hydraulic system and a minor change to the actuators.

The actuators were initially built with 0.030-in diameter ports for the hydraulic fluid. The

pressure drop through these ports was excessive. Analysis showed that increasing the port

diameter to 0.045 in would significantly reduce the pressure drop. The external hydraulic

flex tubing was determined to be too long and too small of a diameter, producing significant

pressure drops when all of the actuators were operating. The external hydraulic tubing was

replaced with shorter, larger diameter flex tubing. In addition, a number of hydraulic tubes

within the model were replaced with tubes of larger diameter where feasible.

40
(a) Remote manifold. (b) Integrated manifold.

Figure A3. Test bench servo valve manifolds.

Figure A4 shows photographs of the model internal hydraulic system used in Test 606

and in Tests 613 and 614. One modification involved installing two hydraulic valves in the

model. These valves could be configured in one of three ways: 1) both off; 2) by-pass mode

where the external supply and return are connected with the model isolated; and 3) normal

operation where the external supply and return are appropriately connected to the model

hydraulic system. The valves minimized the chances of contaminating the model hydraulic

system. Another modification was to replace some of the 1/8-in ID hydraulic tubing in the

model. Larger diameter tubing (1/4-in ID) was installed in the fuselage, the forward wing,

and the tail. In addition to the larger diameter tubing, an extra set of supply and return

tubes was also added to the forward wing to increase flow capacity. The other modification

was to plumb in the brake and snubber solenoid valves. The final modification was to add

three 10-micron filters to the model to help minimize the risk of servo valve contamination.

Figure A4 shows the final configuration.

After these modifications were made, the performance of all of the actuators improved

dramatically. Figure A5 shows the frequency response of a single actuator when excited

individually and with all the actuators both before and after the modifications to the model

hydraulic system. The bandwidth requirements described in table A1 had been met or

exceeded.

Unfortunately, during the preparations for Test 613, the servo valves failed twice. Both

failures were due to contamination, and the valves had to be sent back to Moog for repair.

The initial failure was thought to be associated with fabrication related contamination.

After the second failure, it was determined that the TDT hydraulic system was the source

of the contamination. The TDT 30-gpm hydraulic system was subsequently cleaned, two

3-micron filters were added in series to the model external supply tube, and the three 10-

micron filters, noted earlier, were added to the model itself. These steps combined, with a

very thorough model flush procedure, eliminated the problem of servo valve contamination.

The improvements to the model hydraulic system along with the lack of contamination

resulted in much improved actuator performance and eliminated the need to individually

tune the servo loops. For TDT tests 613 and 614 only proportional gains were needed with

two different gains used, one for the series 30 valves and another for the EO24s. Figure A6

41
(a) Initial configuration used for Test 606.

(b) Final configuration: 3-way valves, larger diameter tubing, second set of tubes to forward wings, solenoid

valves plumbed, and 10-micron filters added. Used in Tests 613 and 614.

Figure A4. JWS model internal hydraulic system.

42
20

10

0

-10

-20 Magnitude, db

150

100

50

0

-50

-100

-150

100 101

Frequency, Hz

Phase, deg

Initial Model, Single CS

Initial Model, All CS

Modified Model, Single CS

Modified Model, All CS

Figure A5. Forward wing, left-mid-board control surface frequency response with the single

and all control surfaces operating, before and after model plumbing modifications.

1.5

1.0

0.5

0

Magnitude

0 5 10 15 20 25 30 35 40 45 50

0

-20

-40

-60

-80

-100

Frequency, Hz

Phase, deg

FWO

FRM

FWI

AWO

AWM

AWI

RUD

(a) Early in Test 613, data point 246.

1.5

1.0

0.5

0

Magnitude

0 5 10 15 20 25 30 35 40 45 50

0

-20

-40

-60

-80

-100

Frequency, Hz

Phase, deg

FWO

FRM

FWI

AWO

AWM

AWI

RUD

(b) End of Test 614, data point 1881.

Figure A6. Control surface frequency response function estimates, ±3

◦

linear sine sweep

excitation. Wing control surfaces on the left side of the model are shown using solid lines,

and control surfaces on the right side of the model are shown using dashed lines.

43
shows frequency response function estimates for all actuators when operated simultaneously.

These data sets show that the control surface frequency response remained very consistent

throughout Tests 613 and 614. It should also be pointed out that the three surfaces that

use the series 30 servo valves (rudder and forward wing inboard control surfaces) tend to be

grouped together especially in terms of phase. The other surfaces all use the EO24 valves

but have varying tubing length between the servo valve and the actuator.

44
Appendix B

Basic Aerodynamic Testing

The initial wind tunnel test of the JWS aeroelastic model was conducted with the model

supported on a sting and force balance so that a full range of aerodynamic data could be

collected. This data was used to refine the analytical models developed for flight control

design.

Figure B1 shows the joined wing model attached to NASA balance 1630 among other

sting components. Except for the balance cables which run inside the wing, all other instrumentation wiring and hydraulic flex tubing were taped to the outside of the sting. In order

to allow the roll coupler to operate, the wires and hydraulic flex lines were grouped into two

service loops, as can be seen in Figure B1a. The black wires running through the center

slot in the test section floor run directly into the control room. While not shown in the

photograph, these wires were grouped into a large service loop along the test section spitter

plate allowing for sting pitch changes. These same wires were used with the beam mount

in Tests 613 and 614. The sting pitch and roll coupler DOF enabled testing over a range of

angles of attack (alpha) and sideslip (beta). The model was fully powered both electrically

and hydraulically so that a multitude of control surface positions could be efficiently tested.

Data from the on-board instrumentation system was recorded in addition to the balance

forces.

The bulk of the aerodynamic data acquired during Test 606 was steady and fell into

three categories: 1) Alpha sweeps with various combinations of symmetric control surface

deflections; 2) Beta sweeps at various alpha values with control surfaces fixed at zero; and

3) Symmetric and antisymmetric control surface deflection sweeps with model held at its

expected trim attitude. Reference 13 describes these data sets in detail. Overall, there

were few surprises in the data, with lift and drag being well behaved linear and quadratic

functions, respectively, of alpha. The forward wing control surfaces were effective at generating lift, but provided limited pitching moment. These forward wing control surface loads

were also found to be relatively linear. The only unusual finding was the lift and pitching

moment associated with aft wing control surfaces. As shown in figure B2, the normal force

and pitching moment are nonlinear, especially between deflections of −10◦ and 0◦. This

behavior is likely due to the aft wing’s relatively thick airfoil. It is also believed that this

nonlinear behavior had a slight destabilizing effect on the trim controller stability margins.

45
(a) Joined wing SensorCraft model installed on the

TDT sting.

(b) Side view of TDT test section showing sting components and joined wing model.

Figure B1. Joined wing SensorCraft model installed on the TDT sting for Test 606.

46
252

250

248

246

244

242

240

238

236

234

Normal

force, lb

AWO

AWM

AWI

Control surface deflection, deg

-15 -10 -5 0 5 10 15

(a) Normal force.

1200

1100

1000

900

800

700

600

500

400

Pitch

moment,

in-lb

Control surface deflection, deg

AWO

AWM

AWI

-15 -10 -5 0 5 10 15

(b) Pitching moment.

Figure B2. Normal force and pitching moment associated with aft wing control surface

deflections.

47
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Aeroservoelastic Testing of Free Flying Wind Tunnel Models-Part 2: A 

Centerline Supported Fullspan Model Tested for Gust Load Alleviation 

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Scott, Robert C.; Vetter, Travis, K.; Penning, Kevin B.; Coulson, David A.; 

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This is part 2 of a two part document. Part 1 is titled: "Aeroservoelastic Testing of Free Flying Wind Tunnel Models Part 1: A Sidewall Supported Semispan 

Model Tested for Gust Load Alleviation and Flutter Suppression." A team comprised of the Air Force Research Laboratory (AFRL), Boeing, 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 full-span, aeroelastically scaled, wind tunnel model of a joined wing SensorCraft vehicle 

was mounted to a force balance to acquire a basic aerodynamic data set. In the second and third tests, the same wind tunnel model was mated to a new, two 

degree of freedom, beam mount. This mount allowed the full-span model to translate vertically and pitch. Trimmed flight at10 percent static margin and gust 

load alleviation 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. The balance and free flying wind tunnel tests will be summarized. The design of the trim and gust load alleviation 

control laws along with the associated results will also be discussed.

15. SUBJECT TERMS

Active controls; Aeroservoelasticity; Body freedom flutter; Flutter supression; Gust load alleviation; Wind tunnel tests

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