NASA/TM-2005-213666

Development and Testing of Control Laws 

for the Active Aeroelastic Wing Program

Ryan P. Dibley, Michael J. Allen, and Robert Clarke

NASA Dryden Flight Research Center

Edwards, California

Joseph Gera

Analytical Services & Materials, Inc.

Edwards, California

John Hodgkinson

Spiral Technology, Inc.

Edwards, California

December 2005
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NASA/TM-2005-213666

Development and Testing of Control Laws 

for the Active Aeroelastic Wing Program

Ryan P. Dibley, Michael J. Allen, and Robert Clarke

NASA Dryden Flight Research Center

Edwards, California

Joseph Gera

Analytical Services & Materials, Inc.

Edwards, California

John Hodgkinson

Spiral Technology, Inc.

Edwards, California

December 2005

National Aeronautics and

Space Administration

Dryden Flight Research Center

Edwards, California 93523-0273
NOTICE

Use of trade names or names of manufacturers in this document does not constitute an official endorsement of such 

products or manufacturers, either expressed or implied, by the National Aeronautics and Space Administration.

Available from the following:

NASA Center for AeroSpace Information National Technical Information Service 

7121 Standard Drive 5285 Port Royal Road

Hanover, MD 21076-1320 Springfield, VA 22161-2171

(301) 621-0390 (703) 605-6000
ABSTRACT

The Active Aeroelastic Wing research program was a joint program between the U.S. Air Force 

Research Laboratory and NASA established to investigate the characteristics of an aeroelastic 

wing and the technique of using wing twist for roll control. The flight test program employed the 

use of an F/A-18 aircraft modified by reducing the wing torsional stiffness and adding a custom 

research flight control system. The research flight control system was optimized to maximize roll 

rate using only wing surfaces to twist the wing while simultaneously maintaining design load 

limits, stability margins, and handling qualities. NASA Dryden Flight Research Center developed 

control laws using the software design tool called CONDUIT, which employs a multi-objective 

function optimization to tune selected control system design parameters. Modifications were made 

to the Active Aeroelastic Wing implementation in this new software design tool to incorporate the 

NASA Dryden Flight Research Center nonlinear F/A-18 simulation for time history analysis. This 

paper describes the design process, including how the control law requirements were incorporated 

into constraints for the optimization of this specific software design tool. Predicted performance is 

also compared to results from flight.

Nomenclature

Ada high-level computer language developed for the U.S. Department of Defense

AAW Active Aeroelastic Wing

AFRL Air Force Research Laboratory

AIL aileron

ASE aeroservoelastic

CONDUIT® Control Designer’s Unified Interface

DFRC Dryden Flight Research Center

FDMS flight deflection measurement system

g acceleration of gravity

GM gain margin

ILEF inboard leading-edge flap

NASA National Aeronautics and Space Administration

OBES onboard excitation system

OLEF outboard leading-edge flap

pmin roll rate at the first minimum of a step response

ppeak first overshoot of a step response

PIO pilot-induced oscillation

PM phase margin


RFCS research flight control system

RPO rolling pullout

STAB stabilator

TEF trailing-edge flap

WUT windup turn

Overview of Active Aeroelastic Wing Project

The Active Aeroelastic Wing (AAW) research program was a joint program between the 

U.S. Air Force Research Laboratory (AFRL) and NASA established to investigate the use 

of a torsionally flexible wing for roll control through and beyond aeroelastic control reversal 

(ref. 1). This overview presents primary and secondary goals of the project, as well as the aircraft 

description and modifications. The following sections also provide details of the Phase I and 

Phase II flights.

Goals

The primary goal of the NASA Dryden Flight Research Center (DFRC, Edwards, California) 

controls effort was to create a design process and use this process to design a controller to maximize 

roll performance at 18 design points, using only wing control surfaces, while maintaining structural 

loads within design limits and maintaining adequate handling qualities. The secondary goal was 

to explore the use of leading-edge and trailing-edge control surfaces of a wing as aerodynamic 

tabs to twist the wing, creating rolling moments through wing twist rather than wing camber. 

In doing so, the amount of required surface deflection could potentially be reduced, allowing 

for hingeless designs that would reduce drag and improve the radar cross section signature 

of an aircraft. Furthermore, this control methodology requires a torsionally flexible wing, 

potentially reducing the amount of structure required in a wing and allowing for the use of higher 

aspect ratios.

Aircraft Description and Modifications

The aircraft selected for this program was an F/A-18A, shown in figure 1, modified by 

replacing the rear wing box skin panels to reduce the torsional stiffness of the wing. Additionally, 

the outboard leading-edge flap (OLEF) surfaces, which are normally driven in unison with the 

inboard leading-edge flaps (ILEFs), were provided with separate actuators to allow independent 

movement. Approximately 200 strain gages were added to the wing structure to measure wing 

loads. An extensive loads calibration was performed to derive the load equations to relate

strain-gage measurements to structural loads (ref. 2). A flight deflection measurement system 

(FDMS) was installed to optically measure the positions of 16 targets on the left wing, thereby 

providing the capability of determining the bending and twisting of the wing during flight (ref. 3). 

Figure 2 shows the FDMS and target locations.


Figure 2. Flight deflection measurement system target locations (ref. 3).

Figure 1. F/A-18 Active Aeroelastic Wing aircraft (control surfaces labeled).


A research flight control system (RFCS) was interfaced with the F/A-18 flight computer such 

that custom control laws could be flown while still maintaining the ability to revert to the standard 

F/A-18 control laws at any time (ref. 1). Two flight phases were performed in the course of the 

AAW program. Phase I flights used an onboard excitation system (OBES) to add frequency sweep 

and doublet commands to the production F/A-18 control law surface commands (ref. 1), providing 

data for aerodynamic and structural loads model development. The RFCS was reprogrammed for

the Phase II flights to incorporate the AAW control laws, controlling the aircraft independent of the 

production F/A-18 control laws.

Phase I Flights

Phase I of the AAW program was performed between November 2002 through June 2003 

to gather flight data to create aerodynamic and structural loads models of the AAW aircraft for 

control law development. Fifty-one flights were flown using the RFCS with an OBES programmed 

to perform open-loop frequency sweeps and doublets. In addition to the OBES maneuvers, piloted 

maneuvers were flown using the baseline F/A-18 controller. The maneuvers included 5-g windup 

turns (WUTs), 360o rolls, and 4-g rolling pullouts (RPOs) (ref. 1). The data from these maneuvers 

were used to create delta values to the baseline aircraft stability derivatives (ref. 4). Additionally, 

the maneuvers were used to develop a linear structural loads model to predict wing and control 

surface structural loads based on aircraft states and control surface deflections. The structural 

loads model predicts wing-root bending and torsion, wing-fold bending and torsion, and wing 

surface hinge moments, for both wings (ref. 5). A third model was created to predict wing twist in 

a similar manner to the structural loads model; however, this model was not available in time for 

the control law development effort (ref. 3).

Phase II Flights

The 35 AAW Phase II flights began on December 14, 2004, to evaluate the newly developed 

AAW control laws. Maneuvers included stick raps, doublets, WUTs, bank-to-bank and 360o rolls, 

and RPOs. Stick raps were used to check for flutter and aeroservoelastic (ASE) characteristics. 

Piloted doublets were performed to familiarize the pilot with the handling of the aircraft prior 

to more aggressive maneuvering. The 5-g WUTs were used to observe structural loading as a 

result of longitudinal maneuvering. These maneuvers provided a comparison to the production 

F/A-18 controller and gave expectations for how the structure would be loaded in the RPO maneuvers 

to follow. The 360o rolls were executed to demonstrate roll performance and rolling using wing 

twist, while the 4-g RPOs were used to demonstrate the same under loaded conditions.

Design Tools

The design process for the AAW control laws required the integration of both pre-existing 

design tools as well as tools created specifically for the AAW project. The following sections 

give an overview and describe the integration of the primary design tool, CONDUIT, the AAW 

structural loads model, and the DFRC F/A-18 nonlinear simulation.


Overview of CONDUIT Design Tool

The Control Designer’s Unified Interface (CONDUIT®, U.S. Army Aeroflightdynamics 

Directorate, Moffett Field, California) is a commercially available design tool that employs a 

multi-objective function optimization to tune selected control system design parameters (ref. 6).

The software operates within the MATLAB® (The MathWorks, Natick, Massachusetts)

environment. Aircraft dynamics and the control system block diagrams are modeled in Simulink. 

Design parameters, to be optimized by CONDUIT, can be specified as any parameter of any 

Simulink block, but typically are defined as gains. The constraints and objective function of the 

optimization are defined using a number of specifications, which are analogous to two-dimensional 

handling qualities specification plots. Figure 3 illustrates an example of a CONDUIT specification.

Figure 3. CONDUIT specification (blue - level 1; magenta - level 2; red - level 3).

The CONDUIT specifications can be used to represent actual handling qualities specifications, 

or as general constraints or objectives of the optimization. Each specification contains level 1, 

level 2, and level 3 regions, corresponding to blue, magenta, and red regions, respectively. 

Regarding handling qualities specifications, these levels correspond directly to the Cooper-Harper 

scale levels (ref. 7). Concerning all other specifications, the levels are merely used for evaluation. 

All specifications are evaluated graphically by CONDUIT.

A typical CONDUIT problem will contain specifications for basic stability, handling qualities 

requirements, and performance objectives. The AAW program used all of these categories and 

added a fourth category consisting of structural loads. The specifications are given constraint 

priorities, allowing the user to fine tune the optimization process. Listed in order of decreasing 

priority, they are: hard, soft, objective, summed objective, and check only. The system is optimized 

in three phases by first satisfying the hard constraints, then satisfying the soft constraints, and 

finally optimizing the objective and summed objective specifications. Check-only specifications 

are not involved in the optimization but are available to the designer to gain insight about 

the system.


Active Aeroelastic Wing Structural Loads Model

A program requirement placed on the AAW control law design was that structural loads had to 

be maintained within their design limits. This requirement initiated the development of a predictive 

structural loads model for use in the control law design. Flight data from Phase I flights were used 

to create a loads model based on 15 aircraft state and surface deflection inputs. The model predicted 

the following loads for the left and right wing: wing-root bending, wing-root torsion, wing-fold 

bending, wing-fold torsion, ILEF hinge moment, OLEF hinge moment, trailing-edge flap (TEF) 

hinge moment, and aileron (AIL) hinge moment. Model coefficients defined unique models at each 

of the 18 design points (ref. 5).

In order to incorporate the structural loads model into the design process, six CONDUIT 

specifications were created to constrain loads during the optimization. Figure 4 shows these 

specifications. A wing-root bending-torque specification and a wing-fold bending-torque 

specification constrained these loads for the left and right wing to remain within a defined envelope. 

Four hinge-moment specifications constrained each of the wing surfaces on the left and right wing. 

For each of these specifications, the borders between the level 2 region (magenta) and level 3 

region (red) represented the 100-percent load limit, while the borders of the level 1 region (blue) 

were based on the predicted load limit uncertainty of the structural loads model for the specific 

load prediction. Thus, designing to these more restrictive borders (although potentially sacrificing 

performance) ensured the greatest probability of maintaining structural design limits in flight.

Figure 4. Load specifications.


Implementation of DFRC F/A-18 Nonlinear Simulation into CONDUIT

A typical application of CONDUIT contains a nonlinear Simulink model, incorporating 

linearized aerodynamics in a state space form. The implementation of AAW in CONDUIT 

initially used this approach. Later, the decision was made to replace the Simulink model 

with a higher fidelity simulation, namely the DFRC nonlinear six-degree-of-freedom 

AAW F/A-18 simulation.

The method by which CONDUIT evaluates the system lent itself to such a modification. 

CONDUIT specifications are self-contained modules of code which, based on their type, either 

linearize the system for frequency-based analysis or run time history simulations for time-based 

analysis. The results are analyzed by each specification, which subsequently displays the results 

graphically. Modifying the specifications to use the nonlinear simulation was simply a matter of 

bypassing the normal routines of calling Simulink, replacing them with calls to run the nonlinear 

simulation. Data from the nonlinear simulation were next formatted to match the expected output 

from Simulink, and the specifications would then continue to run as if nothing had changed. 

Frequency-based specifications of the AAW CONDUIT problem used the standard method of 

evaluation, using a linear Simulink model.

Several benefits could be gained by using the DFRC nonlinear AAW F/A-18 simulation. First, 

the full nonlinear aircraft dynamics could be used. Additionally, the RFCS Ada control laws, used in 

flight, could be used directly, avoiding the problem of creating an identical Simulink representation. 

Finally, because of the complexity of the AAW control laws, the nonlinear simulation with the 

AAW control laws would run faster than real time and approximately two to three times faster than 

the Simulink simulation.

There were two negative aspects to using the nonlinear simulation within CONDUIT, both 

related to the algorithm used to trim the aircraft. Each reset of the nonlinear simulation to the 

nominal conditions required approximately 30 seconds to trim the aircraft. Thus, the time gained 

by the increased run speed was negated by the increased time required to trim the simulation 

before each run. The second problem would occur infrequently. While still untrimmed, the 

simulation would reach a dynamic state that passed the trim criterion. The simulation would run 

prematurely, returning spurious results that could send the optimization in the wrong direction. 

It was discovered, however, that these instances were small enough in number that CONDUIT 

could recover from them. Overall, the benefits of using the nonlinear simulation outweighed the 

problems it introduced.

Controller Design

This section discusses the details of the AAW control law design process. Subsections include 

descriptions of the longitudinal, lateral, and directional control law architectures, the selection of 

design parameters, the design regions and strategies, and the design process.


Control Law Architecture

The AAW control law structure was based on the production F/A-18 controller. Components 

of the system not active in the AAW flight envelope were removed to simplify the system. In 

addition, modifications were made to enable specific AAW control strategy requirements. 

The AFRL contractor defined these strategies, requirements, and the resulting architecture. The 

NASA DFRC designs discussed in this paper were required to fit within this fixed architecture.

Longitudinal Controller

Figure 5 illustrates a simplified representation of the AAW longitudinal control system, showing 

only the basic structure and AAW specific components. Although the basic feedback structure 

remained the same as the production F/A-18 controller, the AAW controller differed in that, rather 

than sending the pilot command error signal to only the stabilator (STAB), the controller branched 

this signal to all of the wing surfaces as well. Each of the five branches (ILEF, OLEF, TEF, aileron, 

and stabilator) had a gain and bias value, allowing the user to define the mixing of the surfaces and 

their trim values.

Figure 5. Simplified longitudinal control laws.

Lateral Controller

Figure 6 shows the AAW lateral-axis controller, simplified to illustrate only the basic structure 

and AAW specific components. As with the longitudinal controller, the architecture was very 

similar to the production F/A-18 control laws but with elements removed that were not required 

within the planned AAW flight conditions. Similar to the longitudinal controller, the pilot command 

error was branched into four signals, each with a gain, for the four wing control surface differential


commands. Additionally, roll-rate feedback was split directly into four signals, each with a gain, 

for a second set of wing control surface differential commands. The four roll-rate feedbacks are 

summed with the four pilot command error signals, resulting in the four wing surface differential 

commands. The differential stabilator was commanded to zero so that wing-only roll control could 

be studied.

Figure 6. Simplified lateral control laws.

Directional Controller

Of the three axes, the directional controller was the least modified. Simplifications were made 

as with the other two axes, and a single gain was added to the rolling surface rudder interconnect 

signal. This gain was seldom used; and for the purpose of this paper, the directional controller can 

be assumed to be the same as in the production F/A-18 control laws.

Design Parameters

Table 1 lists the 23 design parameters that were available for the optimization. In order 

to simplify the design effort, the gains were held constant at each design point rather than 

being scheduled.
10

Longitudinal Gains and Trim Biases

PKSTIK Pilot longitudinal stick gain

PKNZGF Pitch loop gain

PKLEFI ILEF gain

PKLEFO OLEF gain

PKTEF TEF gain

PKAIL AIL gain

PKSTAB STAB gain

NZTRM1 ILEF trim bias

NZTRM2 OLEF trim bias

NZTRM3 TEF trim bias

NZTRM4 AIL trim bias

NZTRM6 STAB trim bias

Lateral Gains

RKPGF Pilot lateral stick gain

RKPD Pilot roll-command error loop gain

RKLEFI ILEF roll-rate gain

RKLEFO OLEF roll-rate gain

RKTEF TEF roll-rate gain

RKAIL AIL roll-rate gain

RKPDG1 ILEF pilot roll-command error gain

RKPDG2 OLEF pilot roll-command error gain

RKPDG3 TEF pilot roll-command error gain

RKPDG4 AIL pilot roll-command error gain

Directional Gain

YKRUD Roll-rate crossfeed gain

Table 1. Design parameter names and descriptions.

Design Regions and Corresponding Strategies

The AAW technology relies on the flexibility of the wing, as well as the ability of the wing 

control surfaces, to effectively twist the wing. In order to twist the wing, the dynamic pressure 

on the control surfaces must be above a critical value. The critical dynamic pressure defines a 

region where control transitions from conventional control to wing twist control. An AAW aircraft 

would likely have to be able to transition from conventional control to wing twist control as the 

aircraft moves through this critical dynamic pressure region. This process of transitioning from 

conventional to wing twist control occurs gradually. Figure 7 shows the effectiveness of the 

wing control surfaces over a range of dynamic pressures for a general AAW aircraft, but does not 

specifically represent the F/A-18 used in the AAW program. Four design regions were defined to 

establish control strategies across the range of dynamic pressures.
11

Figure 7. Wing flexibility design regions.

Region I

Region I defines a conventional wing. Dynamic pressure is too low to produce significant twist 

on the wing. As a result, the leading-edge control surfaces produce very little rolling moment. Twist 

produced by trailing-edge control surface deflection, which normally opposes the lift as a result of 

wing camber, is insignificant. The ideal strategy in this region is to favor the use of trailing-edge 

surfaces, rolling the aircraft primarily using lift derived from wing camber.

Region II

At the higher dynamic pressures of region II, the leading-edge surfaces can twist the wing 

a small amount. The trailing-edge surfaces are still capable of producing rolling moments by 

means of wing camber; however, they impart a twisting on the wing that reduces their overall 

effectiveness. A wing in region II benefits the most by deflecting the control surface positions to an 

optimal combination that maximizes rolling moment using a combination of camber and twist.

Region III

In region III, the leading-edge control surfaces are capable of generating large amounts of wing 

twist, which produces significant rolling moments. The trailing-edge surfaces are at or near the 

point of control reversal, where the opposing wing twist and wing camber effects are balanced. In 

this region, two control strategies could be taken. The first strategy uses the leading-edge surfaces 

to twist the wing, creating rolling moments. Alternately, conventional use of the trailing-edge 

surfaces could be extended into this region by using the leading-edge surfaces to counter the 

adverse twist created by the trailing-edge surfaces.
12

Region IV

Both the leading-edge and trailing-edge control surfaces are capable of significant twist in 

region IV. Thus, only wing twist is used to create rolling moments. Leading-edge surfaces are used 

to twist the wing with greater effectiveness than in the other regions. Trailing-edge surfaces are 

used in control reversal, also creating rolling moments using wing twist.

Predictions and Reality

Early predictions of the performance of the AAW wing showed that control surface reversal 

would occur somewhere within the planned AAW flight test envelope, suggesting the possibility 

of seeing all four design regions. Unfortunately, the increase in wing flexibility was much smaller 

than expected, resulting in observed behavior only in regions I, II, and III (ref. 2). Although 

not a complete picture, these three regions can still demonstrate the AAW concept up to the 

reversal point.

Longitudinal Design

The design of the longitudinal system was not the primary goal of the AAW project. Despite the 

fact that the system was designed to provide active load alleviation, the DFRC design team decided 

that this concept had been explored in a previous research project and did not contribute directly 

to the wing twist methodology (ref. 8). Thus, the DFRC requirements on the longitudinal control 

system were that this system had to satisfy structural load limits as well as basic handling and 

stability criteria. These criteria included gain margin (GM) and phase margin (PM), short period 

damping ratio, short period frequency and acceleration sensitivity, stick force per g, equivalent 

time delay, and pilot-induced oscillations (PIOs). In order to focus the design effort on the lateral 

controller, the DFRC team chose to design the AAW controller to mimic the F/A-18 longitudinal 

controller, which already satisfied these requirements. CONDUIT was used only to verify that the 

longitudinal control law requirements were met.

Lateral Design

The goal of the lateral control system design was to create a controller that would roll the aircraft 

according to the four design regions to maximize roll rate, while simultaneously maintaining 

structural loads within their design limits and adhering to handling qualities requirements. The 

ten gains listed in Table 1 were defined for the lateral control laws. The pilot stick gain was set with 

RKPGF, while the inner loop roll-rate error gain was set with RKPD. The remaining eight gains 

were assigned to each of the four wing surfaces, two per surface.

The first set of four gains was applied to the roll-rate feedback signal, one for each of the wing 

control surfaces. These gains could be used individually to shape the response of a particular 

control surface or collectively to change the effective roll damping of the aircraft. The natural 

roll damping of the aircraft was seen to be sufficient and did not require modification. Shaping 

of the control surface response was also not necessarily required, but was an available option
13

to smooth out transient hinge-moment loads. Thus, the approach to the lateral design was to 

optimize the system without these four gains, applying nonzero values to them only if transient

hinge-moment loads were observed. Throughout the process of designing the control laws, however, 

these four gains were never required.

The other set of four gains was applied to the pilot roll-rate command error signal, split to each 

of the four wing control surfaces. These gains were given to CONDUIT as design parameters for 

the lateral control system optimization, and control the mixing and magnitude of the wing control 

surface positions as a result of roll commands.

The AAW lateral design was largely a problem of determining a mixing function for the 

four wing surfaces for rolling maneuvers. Afterwards, the inner loop gain (RKPD) and pilot stick 

gain (RKPGF) were adjusted to disable the feedback loop, since additional disturbance rejection 

was not required.

Specifications and Constraints

Table 2 shows a list of the AAW specifications, grouped by constraint type, in order of relative 

importance to the optimization. The first phase of a CONDUIT optimization starts with the hard 

constraints, where CONDUIT will make any changes necessary to move these specifications into 

level 1, even if that means violating the other constraint groups. Once this has been accomplished, 

CONDUIT moves the optimization into the second phase, in which the soft and summed objective 

specifications are moved into level 1 while maintaining the level 1 rating of the hard constraints. At 

this point, the third optimization phase begins. In this third phase, the optimization works to move 

the individual objective and summed objective specifications as far into level 1 as possible, again 

without violating the hard or soft constraints. Additionally, the optimization works to achieve 

the same result for the sum of the summed objective specifications. This summation constrains 

the possible see-saw effect, in which one objective specification is improved at the expense of 

another, and vice versa. Check-only specifications are only for reference and are not included in 

the optimization. They are included in a design to provide further insight about the optimization to 

the designer. They also provide a convenient way of temporarily removing a specification from the 

optimization without having to delete the specification permanently.

The hard constraints of the AAW problem were reserved primarily for stability. Since the 

results of many of the specifications are invalid for an unstable aircraft, designating only stability 

specifications as hard constraints ensures that the optimization will ignore any of these invalid 

results until stability is achieved. At that point, the optimization will evaluate the remaining 

specifications, which would then be returning valid results. The EigLcG1 specification ensures 

that all poles are stable, while the StbMgG1 specification maintains gain and phase margins of 

6 dB and 45° respectively. This specification was set to evaluate all three axes. The RFCSEng 

specification monitors the status of the RFCS engage status to prevent undesired disengagements 

from RFCS during maneuvering.
14

Table 2. Active Aeroelastic Wing lateral design CONDUIT specifications.

Specification Constraint Description

EigLcG1 Hard Eigenvalues

StbMgG1 Hard GM/PM

RFCSEng Hard RFCS engage enforcement

LdFdEnv Soft Wing-fold loads envelope

LdRtEnv Soft Wing-root loads envelope

HngILEF Soft ILEF hinge moment

HngOLEF Soft OLEF hinge moment

HngTEF1 Soft TEF hinge moment

HngAil1 Soft AIL hinge moment

MinBta1 Soft Sideslip excursions during roll

RollOsc Soft Roll-rate oscillations

LatAcc1 Soft Lateral acceleration at pilot station

MinBta2 Soft Constraining of beta excursions

MaxRoR1 Summed objective Maximization of roll-rate

PrfRoD1 Summed objective Time to bank

SMinAil Check-only Visualization of AIL deflection only

SMinILF Check-only Visualization of ILEF deflection only

SMinOLF Check-only Visualization of OLEF deflection only

SMinTEF Check-only Visualization of TEF deflection only

RollTC1 Check-only Roll mode time constant

RollED1 Check-only Roll equivalent delay

MinBeta Check-only Visualization of sideslip only

LdBtaQb Check-only Vertical tail load constraint

LatPIO1 Check-only Lateral-directional PIO

RollTC2 Check-only Roll mode time constant on centering stick

RollED2 Check-only Roll equivalent delay on centering stick

The soft constraints comprised the largest number of specifications. The first six specifications 

constrain loads to remain within limits during maneuvering. These include the wing-root and 

wing-fold loads, and the four wing surface hinge moments. Each of these specifications contained 

evaluations for both the left and right wings and surfaces. The MinBta1 specification constrained 

sideslip due to rolling maneuvers to within the MIL-F-8785C (ref. 9) level 1 boundaries of 

2° proverse and 6° adverse yaw. RollOsc enforced the requirement that the roll rate at the 

first minimum ( pmin) following the first peak (ppeak) should be at least 60 percent of the initial peak, 

illustrated in figure 8. This requirement was restricted even further to 90 percent, being both more 

desirable and attainable.
15

Figure 8. Roll oscillations after first peak.

LatAcc1 evaluated the lateral acceleration at the pilot’s head, as defined by MIL-STD-1797 

(ref. 10). This guideline was developed for a class III aircraft with very low roll rates. It was 

found that the guideline did not scale well to the roll rates of the AAW aircraft. Discussions with 

the aircraft manufacturer resulted in an acceptable level 2 value of less than ±0.75 g. MinBta2 

was included to ensure sideslip remained within the MinBta1 sideslip limits for a full stick, 

1-second pulse. 

There were two summed objective specifications that form the focus of the optimization. 

The first, MaxRoR1, determines the steady-state roll rate due to a step input, plotting it onto a 

one-dimensional specification, shown in figure 9. This type of specification has only one active 

axis ( y‐axis), while the other axis is held at a constant value, centered in the plot. This type of 

specification is useful for evaluating a single parameter. As a summed objective specification, the 

optimization will attempt to push the evaluated point as far into level 1 as possible, maximizing 

roll rate. The second summed objective specification, PrfRoD1, calculates the time to bank to 90°, 

plotting the value on a one-dimensional specification. Again, the optimization routine will try to 

move this value into level 1 as far as possible, minimizing the time to bank.

Figure 9. One-dimensional CONDUIT specifications, MaxRoR1 and PrfRoD1.
16

In addition to optimizing the individual summed objectives, the optimization would also try to 

minimize the sum of the two summed objectives. Since the MaxRoR1 and PrfRoD1 specifications 

were mutually beneficial, this part of the optimization task was simplified.

There were many check-only specifications in the AAW CONDUIT problem, included primarily 

to provide further information about the system during the design process. These included 

specifications to visualize the wing control surface deflections and aircraft sideslip. Additionally, a 

number of soft constraints were converted to check-only specifications to simplify the optimization 

cost function. Some of these were converted because they did not drive the optimization, but 

were easily satisfied by achieving level 1 ratings in a few critical specifications. The remaining 

specifications were converted to check only because their algorithms were unreliable or inconsistent 

and caused problems with the optimization. After completion of the optimization, the specifications 

that had been converted from soft constraints were checked for violations.

Modifications to Optimization Process

Modifications were made to the AAW implementation in CONDUIT to improve the efficiency 

and accuracy of the optimization. The implementation of the AAW constraints in CONDUIT 

required 30 time history simulations for a complete evaluation, but only 2 of these were unique. 

Rather than running all 30 simulations, the specifications were modified so that only the 

2 unique simulations were run; and the remaining specifications were modified to reuse the data. 

This modification reduced the required run time by a factor of 15. Since a typical CONDUIT 

problem might require several hundred evaluations for a complete optimization, the time savings 

of this modification was significant. The accuracy was improved through the incorporation of 

the DFRC nonlinear AAW F/A-18 simulation, as described earlier in this paper, resulting in time 

histories created with the most accurate models available. Both of these modifications are not 

typical, require detailed knowledge of CONDUIT, and demonstrate possible improvements for 

future versions of this design tool.

Optimization

A typical optimization would start with the four roll-rate command error gains set to 1.0, the 

four roll-rate feedback gains set to zero, RKPD set to 1.0, and RKPGF set to a value based on 

flight condition. All gains except the four wing surfaces gains would be frozen, that is, removed 

from the active set of design parameters. The optimization usually reached Phase II immediately 

or after a few evaluations of the system. If after several iterations, the roll rate or time-to-bank 

specifications could not be moved into level 1, the borders of these specifications were adjusted to 

allow the optimization to move to Phase III. CONDUIT would then work to improve the roll rate 

and time to bank as much as possible.

Specifically, test points within design regions I and II were optimized using the previous 

procedure on all four wing surface gains. This optimization allowed the largest roll rates 

within the conventional and transitional regions, a design approach that would be typical for an 

AAW aircraft in these regions.
17

For test points that fall within design regions III and IV, where wing twist becomes 

effective, an alternate approach was taken. The trailing-edge surface positions were zeroed and 

frozen at the start of the optimization to allow CONDUIT to maximize roll rate using only the 

leading-edge surfaces and, in turn, wing twist. After the optimization was allowed to progress 

as much as possible, the use of the trailing-edge surfaces was explored. Since region IV was not 

achieved and few test conditions met the criteria for region III, this approach was applied to some 

of the region II test points as well. This artificially restricted the trailing-edge surfaces to create 

additional “region III-like” designs that used primarily wing twist to roll the aircraft.

Typically, the optimization would end with either the sideslip or wing control surface hinge 

moments at their limits. Since an optimization of this complexity will often have multiple local 

minimums, the designer would then investigate the history of the optimization, supporting 

information about the current system, and sensitivity analysis of the active design parameters 

to determine if further improvement were possible. From this, the number of active design 

parameters could be changed to refine the optimization at the current minimum. Alternately, the 

design parameter values could be changed manually to provide a new direction to the optimization 

or to break out of a local minimum. The degree of user intervention required varied at each flight 

condition from zero intervention to several redirections of the optimization.

Phase II Flight Data

Figure 10 through figure 14 show results from left and right 360° rolls at a region III design 

point. Left and right full-stick rolls are shown in each plot along with the predicted response from 

a right full-stick roll from the AAW F/A-18 simulation. Although the aircraft loads and handling 

qualities requirements are not shown in the following plots, these requirements remained within 

limits for both predicted and actual maneuvering. All of the plots within this section share common 

time scales.

Figure 10 shows the roll rates from the predicted simulation and the two 360° rolls from flight. 

The predicted and flight right rolls show similar roll rates, with the flight results being slightly 

higher than that of the simulation. The roll rate for a left roll was slightly higher than for the right 

roll; a trend consistent with Phase I results and most likely caused by asymmetry in the aircraft 

(ref. 1). Figure 11 shows that sideslip remains within the level 1 region for proverse sideslip for 

the duration of each roll. Predicted sideslip matched the flight data closely throughout the roll. 

Differences between the predicted and flight responses at the end of the roll were a result of 

differences in the stick commands.

Figure 12 shows surface deflections for the full-stick rolls. The rudder plot shows the average 

collective rudder deflection, while the remaining plots show surface differential deflections; 

representing the included angle between the left and right surfaces. From these plots, it is apparent 

that the only wing surfaces that moved appreciably were the OLEFs. The rudders moved to 

counter the sideslip, and the stabilator differential command was essentially zero. Differences 

between the surface deflections from flight and those from the simulation were due to inaccuracies 

in the aerodynamic hinge-moment predictions in the AAW F/A-18 nonlinear simulation. At this 

particular flight condition, the aileron was at the reversal point, making this a region III test point. 
18

The OLEFs were found to be the most effective surface; more effective than any combination 

of the other surfaces. Additionally, other surface combinations with OLEF were found to violate 

hinge-moment or sideslip limits.

As stated earlier in this paper, figure 2 shows the locations of the FDMS targets. For the 

two rolling maneuvers discussed in this section, however, only targets 10, 12, 13, and 15 were 

active; providing twist at spanwise locations 193.8 in. and 222.1 in. Figure 13 and figure 14 show 

the left wing twist at the two outboard twist stations for a right and a left roll, respectively. The 

dark blue (solid line) segments indicate valid data whereas the light blue (dashed line) segments 

indicate a span of time in which the optical sensor data was invalid. The right roll shows an increase 

in twist (leading edge up) of the left wing, the correct direction for creating a roll through wing 

twist. Likewise, the left roll shows a decrease in wing twist, also in the correct direction. Thus, 

the controller was using the wing surfaces in such a way to roll the aircraft through wing twist. 

To achieve the same amount of twist with a reduction in the amount of required surface motion, 

additional reduction in the wing structure would be required to increase torsional flexibility. In 

such a situation, the aircraft performance would improve since the same amount of twist would be 

produced with less adverse camber.
19

Figure 10. Roll rate: predicted left, flight left and flight right 360° rolls.

Figure 11. Sideslip: predicted left, flight left and flight right 360° rolls.
20

Figure 12. Surface deflections: predicted left, flight left and flight right 360° rolls.
21

Figure 13. Wing twist: right 360° roll, positive leading edge up.

Figure 14. Wing twist: left 360° roll, positive leading edge up.
22

Conclusions

A design process has been developed and proven in flight to produce control laws to roll the 

Active Aeroelastic Wing F/A-18 aircraft at the maximum possible roll rate, using wing twist 

created by using the wing control surfaces as aerodynamic tabs. Results for a representative flight 

condition were shown. Both loads and handling qualities requirements were maintained within 

acceptable levels. The roll rate for a left 360° roll was greater than the same maneuver to the right 

due to aircraft asymmetries and both were slightly larger than the predicted simulation roll rate. 

Sideslip was constrained within the 2° proverse limit as defined in MIL-F-8785C, matching the 

predicted response from simulation. Roll control was allocated almost entirely to the outboard 

leading-edge flaps, imparting twist on the wing to roll the aircraft. This flight condition demonstrated 

that effective roll control can be achieved through the use of wing twist only.

Dryden Flight Research Center

National Aeronautics and Space Administration

Edwards, California, June 29, 2005
23

REFERENCES

1. Voracek, David, Ed Pendleton, Eric Reichenbach, Dr. Kenneth Griffin, and 

Leslie Welch, The Active Aeroelastic Wing Phase I Flight Research Through January 2003, 

NASA/TM-2003-210741, 2003.

2. Lokos, William A., Candida D. Olney, Natalie D. Crawford, Rick Stauf, and 

Eric Y. Reichenbach, Wing Torsional Stiffness Tests of the Active Aeroelastic Wing 

F/A-18 Airplane, NASA/TM-2002-210723, 2002.

3. Lizotte, Andrew M. and Michael J. Allen, Twist Model Development and Results From 

the Active Aeroelastic Wing F/A-18 Aircraft, NASA/TM-2005-212861, 2005.

4. Cumming, Stephen B. and Corey G. Diebler, “Active Aeroelastic Wing Aerodynamic 

Model Development and Validation for a Modified F/A-18A,” AIAA-2005-6312, 

AIAA Atmospheric Flight Mechanics Conference and Exhibit, San Francisco, CA, 

August 15–18, 2005.

5. Allen, Michael J., Andrew M. Lizotte, Ryan P. Dibley, and Robert Clarke, “Loads

Model Development and Analysis for the F/A-18 Active Aeroelastic Wing Airplane,” 

AIAA-2005-6313, AIAA Atmospheric Flight Mechanics Conference and Exhibit, 

San Francisco, CA, August 15–18, 2005.

6. Tischler, Mark B., et al., “CONDUIT–A New Multidisciplinary Integration Environment

for Flight Control Development,” AIAA-1997-3773, AIAA Guidance, Navigation, and

Control Conference, New Orleans, LA, 1997.

7. Cooper, George E. and Robert P. Harper, Jr., The Use of Pilot Rating in the Evaluation of 

Aircraft Handling Qualities, NASA TN D-5153, 1969.

8. Thornton, Stephen V., Reduction of Structural Loads Using Maneuver Load Control on 

the Advanced Fighter Technology Integration (AFTI)/F-111 Mission Adaptive Wing, 

NASA TM 4526, 1993.

9. MIL-F-8785C, “Flying Qualities of Piloted Airplanes,” November 5, 1980, available

from ASD/ENESS, Wright-Patterson AFB, OH.

10. MIL-STD-1797 (USAF), “Flying Qualities of Piloted Vehicles,” March 31, 1987,

available from ASD/ENES, Wright-Patterson AFB, OH.
REPORT DOCUMENTATION PAGE Form Approved

OMB No. 0704-0188

1. REPORT DATE (DD-MM-YYYY)

28-12-2005

2. REPORT TYPE 

Technical Memorandum

 4. TITLE AND SUBTITLE

Development and Testing of Control Laws for the Active 

Aeroelastic Wing Program

5a. CONTRACT NUMBER

 6. AUTHOR(S)

Ryan P. Dibley, Michael J. Allen, Robert Clarke, Joseph Gera, 

and John Hodgkinson 

 7. PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES)

NASA Dryden Flight Research Center

P.O. Box 273

Edwards, California 93523-0273

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National Aeronautics and Space Administration

Washington, DC 20546-0001

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H-2615

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13. SUPPLEMENTARY NOTES

Also presented at the AIAA Atmospheric Flight Mechanics Conference and Exhibit, San Francisco, CA, August 15–18, 2005. 

An electronic version can be found at the NASA Dryden Flight Research Center Web site, under Technical Reports.

12. DISTRIBUTION/AVAILABILITY STATEMENT

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Subject Category 01, 05, 08 Availability: NASA CASI (301) 621-0390 Distribution: Standard 

19a. NAME OF RESPONSIBLE PERSON

STI Help Desk (email: help@sti.nasa.gov)

14. ABSTRACT

The Active Aeroelastic Wing research program was a joint program between the U.S. Air Force Research Laboratory and NASA established to 

investigate the characteristics of an aeroelastic wing and the technique of using wing twist for roll control. The flight test program employed the 

use of an F/A-18 aircraft modified by reducing the wing torsional stiffness and adding a custom research flight control system. The research flight 

control system was optimized to maximize roll rate using only wing surfaces to twist the wing while simultaneously maintaining design load 

limits, stability margins, and handling qualities. NASA Dryden Flight Research Center developed control laws using the software design tool 

called CONDUIT, which employs a multi-objective function optimization to tune selected control system design parameters. Modifications were 

made to the Active Aeroelastic Wing implementation in this new software design tool to incorporate the NASA Dryden Flight Research Center 

nonlinear F/A-18 simulation for time history analysis. This paper describes the design process, including how the control law requirements were 

incorporated into constraints for the optimization of this specific software design tool. Predicted performance is also compared to results from flight.

15. SUBJECT TERMS

AAW, Active aeroelastic wing, CONDUIT, Control reversal, F/A-18, Handling qualities, Structural loads 

model, Wing twist

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