NASA/TM-2005-213668

Active Aeroelastic Wing Aerodynamic 

Model Development and Validation for a 

Modified F/A-18A Airplane

Corey G. Diebler

NASA Kennedy Space Center

Kennedy Space Center, Florida

Stephen B. Cumming

NASA Dryden Flight Research Center

Edwards, California

November 2005
NASA/TM-2005-213668

Active Aeroelastic Wing Aerodynamic 

Model Development and Validation for a 

Modified F/A-18A Airplane

Corey G. Diebler

NASA Kennedy Space Center

Kennedy Space Center, Florida

Stephen B. Cumming

NASA Dryden Flight Research Center

Edwards, California

November 2005

National Aeronautics and

Space Administration

Dryden Flight Research Center

Edwards, California 93523-0273
ABSTRACT

A new aerodynamic model has been developed and validated for a modified F/A-18A airplane used

for the Active Aeroelastic Wing (AAW) research program. The goal of the program was to demonstrate

the advantages of using the inherent flexibility of an aircraft to enhance its performance. The research

airplane was an F/A-18A with wings modified to reduce stiffness and a new control system to increase

control authority. There have been two flight phases. Data gathered from the first flight phase were used

to create the new aerodynamic model. A maximum-likelihood output-error parameter estimation

technique was used to obtain stability and control derivatives. The derivatives were incorporated into the

National Aeronautics and Space Administration F-18 simulation, validated, and used to develop new

AAW control laws. The second phase of flights was used to evaluate the handling qualities of the AAW

airplane and the control law design process, and to further test the accuracy of the new model. The flight

test envelope covered Mach numbers between 0.85 and 1.30 and dynamic pressures from 600 to

1250 pound-force per square foot. The results presented in this report demonstrate that a thorough

parameter identification analysis can be used to improve upon models that were developed using other

means. This report describes the parameter estimation technique used, details the validation techniques,

discusses differences between previously existing F/A-18 models, and presents results from the second

phase of research flights.

NOMENCLATURE

AAW Active Aeroelastic Wing

AIL aileron

c mean aerodynamic chord

CG center of gravity

rolling-moment coefficient due to differential TEF deflection

pitching-moment coefficient

pitching-moment coefficient due to angle of attack

pitching-moment coefficient due to symmetric ILEF deflection

pitching-moment coefficient due to dimensionless pitch rate

yawing-moment coefficient due to differential TEF deflection

normal-force coefficient due to angle of attack

d differential surface deflection

g acceleration due to gravity

ILEF inboard leading-edge flap

ITB integrated test block

LEF leading-edge flap

Cl

dTEF

Cm

Cm!

CmILEF

Cmq

CndTEF

CN!
2

NASA National Aeronautics and Space Administration

OBES Onboard Excitation System

OLEF outboard leading-edge flap

PID parameter identification

POPU push-over–pull-up

q pitch rate

RFCS Research Flight Control System

RPO rolling pull-out

RUD rudder

STAB stabilator

SRA Systems Research Aircraft

TEF trailing-edge flap

V true airspeed

!, alpha angle of attack

", delta control surface deflection

INTRODUCTION

In April of 2005, the National Aeronautics and Space Administration (NASA) Dryden Flight

Research Center completed flight tests for the Active Aeroelastic Wing (AAW) program. The program

was a joint research effort involving NASA, the Air Force Research Lab (AFRL), the United States

Navy, and industry. The program looked to develop the multidisciplinary technologies needed to

integrate aircraft aerodynamics, active controls, and structural aeroelastic behavior to maximize vehicle

performance (ref. 1). Control laws were designed for a light, flexible-winged aircraft that would improve

roll performance by exploiting the elastic characteristics of the wing (ref. 2).

At the outset of the flight program, The Boeing Company (Chicago, Illinois) modified a NASA

F/A-18A airplane in preparation for flight testing (ref. 1). In order to develop the integrated control laws,

it was necessary that the designers possess an accurate representation of the stability and control

characteristics of the aircraft. To acquire this representation, a parameter identification (PID) flight phase

was flown to extract the necessary data from flight measurements. This approach had not been pursued

during the development of existing F/A-18 aerodynamic models (ref. 1) and resulted in a more accurate

model of the aircraft than had previously existed.

This report describes the approach taken to develop and validate the AAW aerodynamic model as

well as an informal verification from flight test results. Findings and details about the PID results and

model differences are discussed in addition to data from flight tests.
3

VEHICLE SELECTION

A NASA Dryden F/A-18A airplane, tail number 853, shown in figure 1, was selected as the AAW

research airplane for several reasons. First, NASA already had a full F-18 simulation and the team

already had a wealth of knowledge about the airplane and its systems. Second, the leading-edge flap

(LEF) of the F/A-18 airplane is split into an inboard and an outboard. A standard F/A-18 LEF is operated

as one unit driven by a single actuator so that both sections have the same motion. Additional hardware

was added to the AAW airplane that allowed the inboard section of the LEF to operate independently of

the outboard section. This provided an additional set of surfaces with which to control the twist,

aerodynamics, and load distribution of the wing. Third, the F/A-18 airplane was chosen as the test

airplane because during preproduction performance testing it was discovered that the lighter

preproduction wings were too flexible and did not meet the desired roll performance criteria. For this

reason, the F/A-18 airplane was redesigned with a stronger, heavier wing. This helped to improve the roll

performance, but it resulted in a heavier airplane, reducing range and carrying capability. Since the goal

of the AAW technology was to improve the roll performance of a flexible wing, the preproduction

F/A-18 airplane provided a history in which wing flexibility was detrimental to aircraft performance.

EC02-0264-19

Figure 1. The Active Aeroelastic Wing research airplane.
4

VEHICLE MODIFICATIONS

Several modifications were made to the airplane for the AAW program (ref. 1). The wings were

returned to the reduced stiffness levels of the preproduction F/A-18 design to provide greater flexibility

for flight testing (ref. 1). The LEF system was also modified. The F/A-18 LEF system, which operates as

a single unit, was split at the wing fold into an inboard LEF (ILEF) and outboard LEF (OLEF) that can

operate independent of one another. A schematic illustrating the control surfaces can be seen in figure 2.

The LEF drive system was also modified to give the LEFs a greater range of motion as can be seen in

Table 1.

Figure 2. Active Aeroelastic Wing airplane schematic.
5

The AAW flight control computers were modified as well. The standard 701E processor running the

basic F/A-18 flight control laws was augmented with a Research Flight Control System (RFCS). The

RFCS was added to provide a flexible means to run the research software load when desired and revert

back to the standard F/A-18 control laws when needed. The pilot could switch between the standard

control laws and the research control laws while in flight by means of a switch in the cockpit. The

standard controls were flown during take-off, landing, and while transitioning between test points.

The AAW airplane was equipped with a complete set of air data, inertial, and control-surface position

instrumentation. In addition to the F/A-18 standard instrumentation system, a myriad of sensors were

added to the AAW airplane to meet research objectives (ref. 1). Data measurements were telemetered to

the ground control room for real-time monitoring and were also recorded for postflight analysis.

The analysis discussed in this report used measurements from the standard F/A-18 instrumentation

package (obtained through the 1553 data bus) and additional instruments mounted throughout the aircraft

which were telemetered down on a different data stream. Ambient temperature, control surface positions,

and pitch and roll attitudes were obtained from the standard F/A-18 instrumentation package. A research

instrumentation package provided air data measurements (pitot and static pressures, angle of attack, and

angle of sideslip) from a noseboom mounted to the airplane, and angular rates and aircraft accelerations

from rate gyros and accelerometers. Also, a camera pod was mounted on the top of the airplane to record

the amount of wing twist during maneuvering.

FLIGHT TESTS

In order to ensure a safe, thorough evaluation, the flight research program was planned in two phases.

The first phase of AAW flights was dedicated to flutter clearance, loads clearance, and model

Table 1. Control surface rates and limits.

Control Surface

Rate,

deg/sec

Upward travel,

deg

Downward travel,

deg

Aileron 100 25 45

Inboard leading-edge flap 15 3 34

Trailing-edge flaps 18 18 45

Stabilator 40 10.5 24

Rudder 56 30 (left) 30 (right)

Baseline outboard leading-edge flap 15 3 34

Research flight control system 

outboard leading-edge flap

45 10 34
6

development (PID maneuvers). The data gathered from these flights were used to create new

aerodynamic and loads models to be used in the control law design process. The second phase of flights

was dedicated to demonstrating the AAW technology and testing the new flight control laws. Data from

the second phase of flights were also used to further validate, or informally verify, the new models.

The AAW test matrix was made up of 18 separate test points: nine subsonic (Mach 0.85, 0.9, and

0.95) and nine supersonic (Mach 1.1, 1.2, and 1.3). Dynamic pressures ranged from 600 to 1500 psf.

Figure 3 displays the target AAW test matrix and flight envelope. The pilot flew to the conditions

displayed to him by the F/A-18 production air data system. It was found that the production system did

not have a good calibration above Mach 1. More accurate data from the research noseboom were used for

data analysis, but were not displayed to the pilot in the cockpit. Therefore, the true flight conditions did

not exactly match those shown in figure 3.

Figure 3. Active Aeroelastic Wing test matrix.

PHASE I PARAMETER IDENTIFICATION

A total of 51 flights were flown during phase I. Flights consisted of functional checkout, clearance,

air data calibration, PID, and Integrated Test Block (ITB) maneuvers.

Onboard Excitation System Maneuvers

During the first phase flights, the RFCS was used to operate the Onboard Excitation System (OBES).

The OBES consisted of 31 preprogrammed maneuvers which were added to the underlying 701E

commands. The RFCS did not deactivate the standard control laws nor did it eliminate the feedback
7

responses to airplane rates. This method allowed the 701E to retain control of the airplane while the

OBES maneuvers were flown. Six of the 31 OBES maneuvers were designed as PID maneuvers. The

other 25 OBES maneuvers consisted of frequency sweeps to investigate aeroservoelastic effects, LEF

ground test maneuvers and OLEF failure emulations.

Standard F/A-18 control laws use multiple surface inputs in response to pilot stick deflections to

coordinate the motion of the vehicle. The control surface motion is highly correlated and it is not possible

to separate out individual effects. In order to extract accurate stability and control derivatives, it is

necessary to have independent surface motion during the maneuvers. The six OBES PID maneuvers

commanded longitudinal and lateral–directional control surface doublets independent of any piloted stick

motion. Each set of control surfaces was actuated one at a time. This separate actuation provided control

surface inputs that were largely independent of one another. Since the underlying F/A-18 control laws

were still active during these maneuvers, small amounts of control system feedbacks were present, but

surface motion was sufficiently decoupled for PID analysis. A combined inboard and outboard LEF

doublet was also programmed into the OBES commands. This doublet was used to verify superposition

of combined LEF motion and was not used as part of the PID analysis. Tables 2 and 3 show the deflection

magnitudes for each of the OBES PID maneuvers. The columns of the tables are in order of sequence,

from left to right. Surface doublets were separated by two seconds to allow the aircraft motion to dampen

out. Figures 4 and 5 display sample control surface deflections recorded during longitudinal and

lateral–directional OBES maneuvers, respectively.

Table 2. Longitudinal Onboard Excitation System maneuvers.

, deg , deg , deg , deg , deg , deg

Large 3 3 5 3 6 0.8

Medium 2 2 4 2 5 0.8

Small 1 1 3 1 4 0.8

Table 3. Lateral–directional Onboard Excitation System maneuvers.

, deg , deg , deg , deg , deg , deg , deg

Large 4 6 6 10 12 6 6

Medium 4 4 4 8 10 6 4

Small 4 2 2 6 8 6 2

"OLEF "ILEF "TEF "LEF "AIL "STAB

"RUD "dOLEF "dILEF "dTEF "dAIL "dSTAB "dLEF
8

Figure 4. Longitudinal Onboard Excitation System maneuver.

Figure 5. Lateral–directional Onboard Excitation System maneuver.

0 5 10 15 20 25 30 35 40 –8

–6

–4

–2

0

2

4

6

8

Surface deflection, deg

Time, s

OLEF

ILEF

TEF

AIL

STAB

050344

0 5 10 15 20 25 30 35 40

0

15

10

5

–5

–10

–15

Surface deflection, deg

Time, s

RUD

OLEF

ILEF

TEF

AIL

STAB

050345
9

Control surface deflections were defined as follows:

Symmetric deflections:

Differential deflections:

Integrated Test Block Maneuvers

In addition to the OBES maneuvers, ITB maneuvers were flown at each test point. The maneuvers

consisted of push-over–pull-ups (POPU), stick rolls, 4-g rolling pull-outs (RPO), and wind-up turns to

5 g’s. The g limits were imposed to give a wide range of data, but stay within safe limits. The pilot flew

POPUs by pushing forward on the stick to –1 g and then pulling back to 3 g’s. The RPOs were flown by

banking the airplane, pulling back on the stick to maintain a level 4-g turn, then quickly rolling the

airplane 180 deg in the opposite direction while maintaining constant longitudinal stick position. The

wind-up turns consisted of the pilot banking the aircraft and pulling back on the stick, slowly increasing

the normal acceleration to 5 g’s while maintaining a constant altitude. These maneuvers were flown using

the standard F/A-18 control laws and were used to validate the new aerodynamic model.

 "AIL 0.5"AILLeft

"AILRight+ # $ % &

"STAB 0.5"STABLeft"STABRight+ # $ % &

"TEF 0.5"TEFLeft"TEFRight+ # $ % &

"ILEF 0.5"ILEFLeft"ILEFRight+ # $ % &

"OLEF 0.5"OLEFLeft"OLEFRight+ # $ % &

"RUD 0.5"RUDLeft"RUDRight+ # $ % & =

=

=

=

=

=

 "dAIL "AILLeft

"AILRight –

"dSTAB "STABLeft"STABRight

 "dTEF "TEFLeft

"TEFRight –

 "dILEF "ILEFLeft

"ILEFRight

"dOLEF "OLEFLeft"OLEFRight = –

= –

=

= –

=
10

PID Technique

A maximum-likelihood output-error PID technique was used to estimate the stability and control

derivatives of the AAW airplane at each test point. Measured airplane responses to the predefined OBES

maneuvers were read into a PID program known as pEst (ref. 3). The pEst program used a mathematical

cost function to quantify the difference between flight-measured airplane responses and the responses

predicted by a vector set of dynamic equations of motion (ref. 4). The program then used an iterative

technique in which the unknown stability and control parameter values were adjusted to reduce the value

of the cost function and thereby increase the agreement of the math model to flight data. This technique

resulted in a set of stability and control parameter values for each maneuver analyzed. Figure 6 shows a

flow diagram of the pEst process. Parameter values from multiple maneuvers at the same test point were

averaged to obtain final values that were then used to create the new AAW aerodynamic model. Figure 7

displays the fit from a sample pEst case. Blue lines display flight data, red lines display the predicted

math model responses using the final parameter values for this particular run. Because of data sensitivity,

all plots in this report showing flight data have the scaling removed.

In order to extract accurate stability and control parameters from the flight data, it was necessary to

have precise knowledge of the weight, center of gravity (CG), and rotational moments of inertia, for each

maneuver of the aircraft. Thus, a weight-and-balance ground test was performed to measure the unfueled

weight and the CG location of the AAW airplane. Capacitance sensors installed in each fuel tank

provided fuel quantities during each maneuver. Fuel data was used postflight along with the tank location

to adjust the CG location and inertias.

Figure 6. The output-error maximum-likelihood parameter identification technique.

Input

Noise

Estimate

Minimization

process

Maximum likelihood

estimate

Mathematical model

of aircraft

Cost

function

Test

aircraft

050346
11

Figure 7. Sample pEst comparison, lateral–directional Onboard Excitation System maneuver at

low-altitude, supersonic conditions.

0 10

Lateral acceleration

Yaw rate Roll rate Beta

20 30

Time, s 050347
12

PID Analysis and Results

Initially, only OBES maneuvers were used to create the new aerodynamic model. However, during

model development, it was discovered that the linear results produced by the pEst analysis were not able

to produce the normal accelerations seen during wind-up turns. One possible explanation is that the

longitudinal OBES maneuvers alone did not produce a wide enough data range to extract values

that could be extrapolated to higher alphas. To alleviate this problem, wind-up turns and POPUs were

used in conjunction with the longitudinal OBES maneuvers in order to widen the angle-of-attack and

normal acceleration ranges used in the analysis. First, the OBES maneuvers were analyzed to obtain the

complete set of stability and control derivatives. Then, while keeping the control derivatives fixed, a

POPU or wind-up turn was analyzed to adjust . While this approach slightly compromised the match

to the OBES data, it worked very well to eliminate the discrepancies seen in high-g maneuvers. Figures 8

and 9, respectively, illustrate the differences between the first version of the model, created using only the

OBES maneuvers, and the second version of the model, which utilized wind-up turns and POPUs along

with the OBES maneuvers in its creation.

Figure 8. Aerodynamic model version 1.0, wind-up turn at Mach 1.2 and an altitude of 20,000 ft.

CN!

CN!

0 5

Normal acceleration

10 15

Time, s 050348

Alpha
13

Figure 9. Aerodynamic model version 2.0, wind-up turn at Mach 1.2 and an altitude of 20,000 ft.

Since it was found that the aerodynamic model did not produce valid longitudinal results outside the

ranges used in the analysis, limits were imposed on the use of the model. The model was considered valid

from 0 to 5 g’s, roughly the range used during the analysis. A similar investigation was conducted on the

lateral–directional results; however, the linearity appeared to hold. Therefore, no limits were imposed on

lateral–directional applications.

After the analysis was completed, it was noted that several substantial differences existed between the

AAW PID results and existing F/A-18 databases. These findings are consistent with the findings of a

previous F/A-18 flight research program conducted at the NASA Dryden Flight Research Center (ref. 5).

In 1999 the NASA Systems Research Aircraft (SRA) completed a series of flight tests that were a

precursor to the AAW flights. The SRA was outfitted with an RFCS and given OBES software nearly

identical to that of the AAW airplane. The major differences between the SRA and the AAW airplane

were that the SRA was an F/A-18B model with unmodified wings and an LEF system that was not

divided into inboard and outboard LEFs.

0 5

Normal acceleration

10 15

Time, s 050349

Alpha
14

The SRA flew similar OBES PID maneuvers at each of the AAW test points. It was also noted during

the SRA program that there were several areas of substantial difference between the flight-derived

stability and control parameters and those of the existing F/A-18 models (ref. 5). The same differences

were seen in the AAW results. The most notable of these differences was in the values. Figure 10

shows some of the TEF results at an altitude of 15,000 ft. While all models show similar results at all

Mach numbers for , the same cannot be said for the values. The AAW and SRA models

have values that remain relatively constant throughout the Mach range while earlier models have

 values that drop off substantially at supersonic Mach numbers.

Figure 10. Trailing-edge flap parameter identification results.

During validation it was found that the AAW model provided a much better match to the measured

response of a TEF doublet than did the baseline model as can clearly be seen in figure 11. This area is one

in which the AAW and the SRA models greatly improved upon the standard models.

CndTEF

CdTEF CndTEF

CndTEF

CndTEF

CndTEF

CldTEF

0.85 0.90 0.95 1.15 1.20 1.25

AAW PID

Baseline sim

SRA model

4

–2

0

2

4

6

8

10

3

2

1

0

–1

Mach number 050350
15

Figure 11. Trailing-edge flap doublet.

15 16 17 18 19

Flight

Baseline sim

AAW sim

20

Time, s

Flight surfaces

Yaw rate

050351

TEF

AIL

STAB

RUD

OLEF

ILEF

Beta
16

Simulation Aerodynamic Model Update

Once the new aerodynamic model was created, it was used to update the NASA F-18 simulation

database. Since the AAW model is only a subset of the much larger F-18 model, it was necessary that

simulation revert back to the baseline F-18 database when outside the AAW test matrix. The baseline

F-18 model was used as a base and linear deltas were added to the baseline database so that the

simulation derivatives were equal to those of the new AAW model. Deltas were ramped to a zero value

outside the AAW test matrix to allow for a smooth transition between the two models.

To create the deltas, baseline parameter values were obtained from the simulator at each of the test

points and subtracted from the PID results to obtain the delta value as shown in equation (1). The delta

values were incorporated into a separate database and added to the simulation. The deltas were then

added to the simulation force and moment equations as shown in equation (2).

(1)

(2)

PHASE I AERODYNAMIC MODEL VALIDATION

The validity of any aerodynamic model is demonstrated by its ability to replicate flight data. The

AAW model underwent a thorough validation process based on that principle before the second phase of

flights began. An accurate aerodynamic model was needed for control law design both to increase the

chances of achieving program goals and to ensure a safe second phase of flights. Since the new control

laws were going to be designed using the newly developed AAW model, any errors in the model would

adversely affect the control laws.

Data from the linear model were compared with flight data from the ITB maneuvers. Then, a

complete set of OBES maneuvers and ITB maneuvers were run through the nonlinear, six

degree-of-freedom simulation at each test point and compared with flight data. Both open-loop and

closed-loop simulation runs were conducted. During open-loop test cases, surface positions measured in

flight were used to drive the simulation. This method ensured that control surface motion was identical to

that seen in flight and provided an isolated aerodynamic model test. For closed-loop cases, the simulation

was driven by the piloted stick and rudder commands measured in flight. This method ensured that the

piloted input was identical to that of flight; however, the control laws were allowed to freely command

the control surfaces in response to simulated aircraft motion. This approach provided a complete

end-to-end test of the entire simulation. Figure 12 shows a comparison of the linear model to flight data.

'Cm! Cm!

 (PID) Cm!

(baseline model) = –

'Cm Cm!

! 'Cmq

 

qc

2V

------- 'CmILEF

" = + + ILEF + ...
17

Figure 12. Linear aerodynamic model validation test.

Flight data

Linear model

Time

Lateral acceleration

Yaw rate Roll rate

050352

Beta
18

For comparison purposes, baseline aerodynamic model results were obtained using the same method.

Figure 13 shows an open-loop comparison for a supersonic lateral–directional OBES maneuver. It is

immediately evident that the baseline model produces large oscillations in sideslip angle (beta), yaw rate,

and lateral acceleration. This is not, however, an indication that the baseline model is directionally

neutrally stable. Instead, the oscillations are a result of the test method. Since the simulation is driven by

flight-measured surface positions in open-loop cases, errors in the model responses are not damped out

by control feedbacks. The flight-measured surface motion may even continue to drive the oscillations if

sufficiently out of phase. The AAW model, on the other hand, shows an excellent agreement with flight

data throughout the maneuver. While the open-loop tests provide an excellent isolated test of the AAW

aerodynamic model, it is difficult to compare the AAW model to the baseline because of the oscillations

in the yaw axis.

Figure 14 shows a closed-loop comparison of the same maneuver. Since the simulation is driven by

the measured pilot stick and rudder inputs, the simulation control laws are free to dampen any rates. This

gives a much cleaner baseline model response and it is possible to examine the response for each control

surface doublet. Again the most notable difference between the two models can be seen during the

trailing-edge flap doublet. The baseline model shows no response while the flight data and the AAW

model show a definite response in beta, yaw rate, and lateral acceleration. This response was noted as one

of the largest improvements throughout the AAW test matrix. Other improvements can be seen in roll

rates for the inboard and outboard LEFs and the stabilator.

The OBES maneuvers discussed in the previous paragraphs have shown substantial improvements in

several areas over the baseline model. The ITB comparisons further tested the model to see if these

improvements would carry over to larger coordinated maneuvers. Figure 15 shows a comparison of a

rolling maneuver at the same test point. Again the most obvious differences can be seen in beta and

lateral acceleration. This result is because of the trailing-edge flap discrepancy seen in the OBES

comparisons. The AAW model does a very good job of predicting all the responses for this maneuver.
19

Figure 13. Open-loop lateral–directional Onboard Excitation System validation test at low-altitude,

supersonic conditions.

Time

Surface deflection Lateral acceleration

Yaw rate Roll rate

050353

RUD

OLEF

ILEF

TEF

AIL

STAB

Flight

AAW sim

Baseline sim

Beta
20

Figure 14. Closed-loop lateral–directional Onboard Excitation System validation test at low-altitude,

supersonic conditions.

Time

Flight surfaces Lateral acceleration

Yaw rate Roll rate

050354

RUD

OLEF

ILEF

TEF

AIL

STAB

Flight

AAW sim

Baseline sim

Beta
21

Figure 15. Open-loop integrated test block roll maneuver at low-altitude, supersonic conditions.

Further aerodynamic model checkout was performed using the phase II flights.

Time

Surface deflection Lateral acceleration

Yaw rate Roll rate

050355

RUD

OLEF

ILEF

TEF

AIL

STAB

Flight

AAW sim

Baseline sim

Beta
22

PHASE II FLIGHTS

Phase II consisted of 32 flights, flown with research control laws using RFCS. These control laws

were flown at the 18 flight conditions previously discussed. The main research maneuvers performed

during this phase were bank-to-banks, rolls, wind-up turns, and RPOs. These control laws were designed

to take advantage of the additional control surface actuation available, creating wing twist to assist in

rolling maneuvers. To produce rolling moments, deflections from all four of the wing surfaces could be

used. These surfaces included ILEF, OLEF, TEF, and ailerons. Differential stabilators were not utilized.

Standard F/A-18 control laws use differential stabilators in addition to the other surfaces to provide

sufficient rolling moment. The AAW control laws were designed to create sufficient rolling moment

without this surface deflection. Phase II also served as a means to verify the AAW model, as will be

discussed.

Following the primary phase II flights, the research control laws for several flight conditions were

replaced with new control laws. The flights following this control law change were referred to as

phase IIa flights. These flights allowed researchers to examine different control laws at the same flight

conditions.

AERODYNAMIC MODEL VERIFICATION

Informal verification of the AAW aerodynamic model was performed during phase II of the AAW

project. This verification consisted of comparing flight data from various maneuvers, mainly rolling

maneuvers, with the output from the simulation using the AAW model. Rolling maneuvers were the basis

of the AAW project as well as the focus of the model verification. Results from bank-to-banks,

360-degree rolls, and RPOs were analyzed. Phase II verification was similar to the validation performed

in phase I. In phase II, however, the control laws used for the maneuvers evaluated were designed to

utilize the AAW concept. The phase II maneuvers, therefore, provided the opportunity for a

comprehensive test of the aerodynamic model. Analysis of the results from these flight tests showed that

the AAW model was an adequate aerodynamic model for use with these flight tests.

Lateral–Directional Responses

Roll rates predicted using the AAW aerodynamic model were very good overall when compared with

flight data. The roll rates matched the flight trends throughout the maneuvers and often were nearly

identical to the actual aircraft roll rate. Figure 16 shows an excellent roll rate match throughout the

maneuver. Figure 17 shows a less accurate match. Here, the AAW model roll rate follows the same trend

as the flight data, but with an under-prediction for the steady-state roll rate. For most flight conditions the

steady-state roll rates for 360-degree rolls and bank-to-banks were under-predicted. 

Yaw rates matched flight trends well overall, but not as well as roll rates. Figure 18 is an example of

the AAW model matching the flight yaw rate throughout the course of the maneuver. Figure 19 shows a

maneuver in which the AAW model simulation does not match the flight yaw rate closely. The matching

is reasonable at the start of the maneuver, but as the maneuver progresses, the yaw magnitudes seen in

flight are not well-predicted by the AAW model.
23

Figure 16. Lateral–directional responses from a 75-percent stick left roll at low-speed subsonic

conditions.

Time

Surface deflection Lateral acceleration

Yaw rate Roll rate

050356

RUD

OLEF

ILEF

TEF

AIL

STAB

Flight

AAW sim

Beta
24

Figure 17. Lateral–directional responses from a 75-percent stick left roll at high-speed subsonic

conditions.

Time

Surface deflection Lateral acceleration

Yaw rate Roll rate

050357

RUD

OLEF

ILEF

TEF

AIL

STAB

Flight

AAW sim

Beta
25

Figure 18. Lateral–directional responses from a full-stick left roll at high-speed subsonic conditions.

Time

Surface deflection Lateral acceleration

Yaw rate Roll rate

050358

RUD

OLEF

ILEF

TEF

AIL

STAB

Flight

AAW sim

Beta
26

Figure 19. Lateral–directional responses from a full-stick left roll at low-speed subsonic conditions.

Time

Surface deflection Lateral acceleration

Yaw rate Roll rate

050359

RUD

OLEF

ILEF

TEF

AIL

STAB

Flight

AAW sim

Beta
27

Lateral acceleration outputs from the AAW model were often offset from the lateral acceleration seen

in flight. Though it matched closely for some maneuvers, such as in figure 18, often the lateral

acceleration comparisons more closely resemble figures 16 and 20. Simulation trends match flight data

very well, but the offset is persistent. At times the AAW simulation would not trim to the initial flight

lateral acceleration. The resulting offset would continue, and in some cases, increase throughout the

maneuver, as in figure 16. In other instances, the simulation would trim well at a steady-state condition,

but would be offset from the flight data throughout the actual maneuver, as in figure 20. This offset,

while often present, is generally of a magnitude small enough to be insignificant.

Beta was a particularly important parameter in control law design for AAW as well as during the

flights. Excessive sideslip during rolls could result in a departure from controlled flight. Beta prediction

by the AAW model was generally adequate. Beta did not match trends as well as some of the other

lateral–directional responses, but still matched general trends reasonably well. Figure 17 is a good

example of a reasonable match without near-perfect trend matching. Figure 21 is a good example of beta

from the AAW simulation matching the flight trends well throughout the maneuver.
28

Figure 20. Lateral–directional responses from a full-stick right roll at low-altitude, supersonic conditions.

Time

Surface deflection Lateral acceleration

Yaw rate Roll rate

050360

RUD

OLEF

ILEF

TEF

AIL

STAB

Flight

AAW sim

Beta
29

Figure 21. Lateral–directional responses from a 75-percent stick left roll at high-altitude, supersonic conditions.

Longitudinal Responses

The longitudinal responses from the AAW aerodynamic model matched the flight data reasonably

well for rolling maneuvers. While the simulation did not always trim to the flight values, the outputs

generally followed the flight data trends. Figure 22 shows flight data and AAW model simulation data for

the longitudinal responses from a subsonic roll maneuver. As can be seen in the figure, the longitudinal

Time

Surface deflection Lateral acceleration

Yaw rate Roll rate

050361

RUD

OLEF

ILEF

TEF

AIL

STAB

Flight

AAW sim

Beta
30

time history from the simulation matches the trends of the flight data. Figure 23 shows a supersonic roll

maneuver. Again, the AAW simulation follows the flight data trends reasonably well considering many

of the responses result from cross-coupling. Angle of attack is under-predicted for a portion of the

maneuver because of the slight under-prediction in the negative pitch rate spike that occurs during the

roll. Considering that cross-coupling factors were not included in the aerodynamic model, the

longitudinal responses for lateral–directional maneuvers matched flight trends reasonably well.

Figure 22. Longitudinal responses from a 75-percent stick left roll at low-speed subsonic conditions.

Time

Surface deflection

Normal acceleration Pitch rate

050362

OLEF

ILEF

TEF

AIL

STAB

Flight

AAW sim

Alpha
31

Figure 23. Longitudinal responses from a 75-percent stick left roll at high-altitude, supersonic conditions.

Time

Surface deflection

Normal acceleration Pitch rate

050363

OLEF

ILEF

TEF

AIL

STAB

Flight

AAW sim

Alpha
32

Effects of Maneuver Variations

The agreement between the flight data and the simulation results using the AAW model

remained reasonable regardless of the magnitude or direction of stick input. Trends were followed

reasonably for the various stick inputs, including rolls initiated by various magnitude lateral stick

deflections. Figures 24, 17, and 18 show half-stick, 75-percent stick, and full-stick left rolls. While the

AAW simulation results in all three figures do not match flight data identically, the predictions do match

flight data trends reasonably. For this specific flight condition, the roll rate matching appears best for the

full-stick maneuver. While this result is found in some flight conditions, it is not a consistent trend.

Figure 25 shows a full-stick right roll at the same test conditions. The AAW simulation does not match

the flight data identically for the two maneuvers, but trends are followed reasonably for both. While rolls

in different directions did not result in identical matching, the AAW aerodynamic model proved to be

adequate regardless of roll direction.

Rolling pull-outs were performed to examine the effects of loading on roll performance. The AAW

model did a reasonable job predicting performance for these maneuvers at supersonic flight conditions.

Subsonic RPOs were not analyzed for this report. Figure 26 shows a supersonic RPO with the same stick

deflection as figure 20 (full-stick right). The matching is similar in this maneuver as it was for the

360-degree roll. The RPOs more consistently had simulation results offset from the flight data for

steady-state roll rate than did unloaded rolls. While the consistency of this offset is notable, the

magnitude is not significantly larger than what was seen for 360-degree rolls. Figure 27 shows another

example of a supersonic RPO. Again, the matching is similar to that found in figure 21, a roll at the same

flight conditions and stick magnitude, but in the opposite direction (left). Longitudinal responses from the

simulation for the RPOs generally matched flight data well. Figure 28 shows the longitudinal responses

for an RPO. The simulation outputs match flight data reasonably well.
33

Figure 24. Lateral–directional responses from a half-stick left roll at high-speed subsonic conditions.

Time

Surface deflection Lateral acceleration

Yaw rate Roll rate

050364

RUD

OLEF

ILEF

TEF

AIL

STAB

Flight

AAW sim

Beta
34

Figure 25. Lateral–directional responses from a full-stick right roll at high-speed subsonic conditions.

Time

Surface deflection Lateral acceleration

Yaw rate Roll rate

050365

RUD

OLEF

ILEF

TEF

AIL

STAB

Flight

AAW sim

Beta
35

Figure 26. Lateral–directional responses from a full-stick right rolling pull-out at low-altitude, supersonic

conditions.

Time

Surface deflection Lateral acceleration

Yaw rate Roll rate

050366

RUD

OLEF

ILEF

TEF

AIL

STAB

Flight

AAW sim

Beta
36

Figure 27. Lateral–directional responses from a 75-percent stick right rolling pull-out at high-altitude,

supersonic conditions.

Time

Surface deflection Lateral acceleration

Yaw rate Roll rate

050367

RUD

OLEF

ILEF

TEF

AIL

STAB

Flight

AAW sim

Beta
37

Figure 28. Longitudinal responses from a full-stick right rolling pull-out at low-altitude, supersonic

conditions.

Time

Surface deflection

Normal acceleration Pitch rate

050368

OLEF

ILEF

TEF

AIL

STAB

Flight

AAW sim

Alpha
38

Phase IIa

During the portion of the project informally known as phase IIa, new control laws were implemented

for several of the flight conditions, and maneuvers were performed utilizing these control laws. These

control laws used the available surfaces differently than did the phase II control laws. Examples of

maneuvers using the phase IIa control laws can be seen in figures 29 and 30. Figures 17 and 21,

respectively, are roll maneuvers flown with the same stick deflection using the phase II control laws.

A comparison of the differential surfaces between the two figures from the same flight conditions

illustrates the differences in control surface allocation. The degree to which the simulation matches flight

data does not vary significantly for these maneuvers. This result demonstrates that the accuracy of the

aerodynamic model is not dependent upon the control laws used or surfaces actuated.
39

Figure 29. Lateral–directional responses from a 75-percent stick left roll at high-speed subsonic

conditions (phase IIa control laws).

Time

Surface deflection Lateral acceleration

Yaw rate Roll rate

050369

RUD

OLEF

ILEF

TEF

AIL

STAB

Flight

AAW sim

Beta
40

Figure 30. Lateral–directional responses from a 75-percent stick left roll at high-altitude, supersonic

conditions (phase IIa control laws).

Time

Surface deflection Lateral acceleration

Yaw rate Roll rate

050370

RUD

OLEF

ILEF

TEF

AIL

STAB

Flight

AAW sim

Beta
41

Comparison with Baseline

A comparison of the AAW aerodynamic model with the baseline F-18 model clearly shows that the

AAW model is superior for the AAW aircraft flying at applicable flight conditions. While the baseline

model did a reasonable job for some flight conditions, it rarely outperformed the AAW model.

Maneuvers previously discussed solely in the context of the AAW model will now be shown with both

the AAW and the baseline model. 

Figures 31 through 34 show comparisons between flight data, the AAW model and the baseline

model. Figure 31 is an example of the baseline model matching flight beta reasonably, but poorly

matching the roll rate from flight. Yaw rate and lateral acceleration from the baseline model do not match

flight data particularly well. Figure 32 shows another subsonic roll maneuver. In this instance, beta and

lateral acceleration outputs from the baseline simulation trend the opposite direction from flight and the

AAW simulation outputs. The yaw rate from the baseline model is reasonable and the roll rate is nearly

as accurate as the AAW model. Figure 33 shows a supersonic maneuver. Similarly to the subsonic

maneuver, beta and lateral acceleration trend in opposite directions from the flight data and AAW model.

Again, roll rate and yaw rate from the baseline model match flight data reasonably, but do not match as

closely as those from the AAW model. Figure 34 shows a supersonic RPO. Here, the baseline simulation

trends similarly to the AAW simulation, but is farther away from the flight data for every

lateral–directional response. Figure 35 illustrates a typical set of longitudinal responses from a roll for the

baseline model. The baseline model does a reasonable job, but does not match flight data as well as the

AAW model. From these figures, it is clear that the AAW aerodynamic model is an improvement over

the baseline aerodynamic model. 
42

Figure 31. Lateral–directional responses from a 75-percent stick left roll at low-speed subsonic

conditions.

Time

Surface deflection Lateral acceleration

Yaw rate Roll rate

050371

RUD

OLEF

ILEF

TEF

AIL

STAB

Flight

AAW sim

Baseline sim

Beta
43

Figure 32. Lateral–directional responses from a full-stick left roll at high-speed subsonic conditions.

Time

Surface deflection Lateral acceleration

Yaw rate Roll rate

050372

RUD

OLEF

ILEF

TEF

AIL

STAB

Flight

AAW sim

Baseline sim

Beta
44

Figure 33. Lateral–directional responses from a 75-percent stick left roll at high-altitude, supersonic

conditions.

Time

Surface deflection Lateral acceleration

Yaw rate Roll rate

050373

RUD

OLEF

ILEF

TEF

AIL

STAB

Flight

AAW sim

Baseline sim

Beta
45

Figure 34. Lateral–directional responses from a full-stick right rolling pull-out at low-altitude, supersonic

conditions.

Time

Surface deflection Lateral acceleration

Yaw rate Roll rate

050374

RUD

OLEF

ILEF

TEF

AIL

STAB

Flight

AAW sim

Baseline sim

Beta
46

Figure 35. Longitudinal responses from a 75-percent stick left roll at low-speed subsonic conditions.

Time

Surface deflection

Normal acceleration Pitch rate

050375

OLEF

ILEF

TEF

AIL

STAB

Flight

AAW sim

Baseline sim

Alpha
47

CONCLUSION

A new aerodynamic model has been developed and validated for a modified F/A-18A airplane used

for the Active Aeroelastic Wing (AAW) research program. Results from the parameter identification

analysis conducted for the AAW program have shown that existing F/A-18 aerodynamic models have

several flaws and could be substantially improved through extensive parameter identification testing. The

output-error technique discussed in this report has proven to be effective in providing linear stability and

control estimates from flight data. It is recommended, however, that large coordinated maneuvers be used

in conjunction with individual control surface inputs to ensure that estimates are applicable to a wide

range of maneuvers. It is also necessary to thoroughly validate the model to identify any limitations.

The AAW aerodynamic model produced simulation data that compared well with flight data.

Significant improvement over the baseline model was shown. The accuracy of the model did not vary

significantly with control surfaces used. The model also performed well for loaded maneuvers, such as

rolling pull-outs. Sideslip excursions not predicted by the baseline model were evident in flight and were

predicted adequately by the AAW model. Overall, the aerodynamic model created for the AAW program

proved to be adequate and was an improvement over the existing models for the particular vehicle used.

Dryden Flight Research Center

National Aeronautics and Space Administration

Edwards, California, July 15, 2005

REFERENCES

1. Voracek, David, Ed Pendleton, Dr. Kenneth Griffin, Eric Reichenbach, and Leslie Welch, The Active

Aeroelastic Wing Phase I Flight Research Through January 2003, NASA/TM-2003-210741, 2003.

2. Dibley, Ryan P., Michael J. Allen, and Robert Clarke, Development and Testing of Control Laws for

the Active Aeroelastic Wing Program, NASA/TM-2005-213666, 2005.

3. Murray, James E. and Richard E. Maine, pEst Version 2.1 User’s Manual, NASA Technical

Memorandum 88280, 1987.

4. Maine, Richard E. and Kenneth W. Iliff, Application of Parameter Estimation to Aircraft Stability

and Control: the Output-Error Approach, NASA Reference Publication 1168, 1986.

5. Moes, Timothy R., Gregory K. Noffz, and Kenneth W. Iliff, Results From F-18B Stability and

Control Parameter Estimation Flight Tests at High Dynamic Pressures, NASA/TP-2000-209033,

2000.
REPORT DOCUMENTATION PAGE Form Approved

OMB No. 0704-0188

1. REPORT DATE (DD-MM-YYYY)

04-11-2005

2. REPORT TYPE 

Technical Memorandum

 4. TITLE AND SUBTITLE

Active Aeroelastic Wing Aerodynamic Model Development and Validation for 

a Modified F/A-18A Airplane

5a. CONTRACT NUMBER

 6. AUTHOR(S)

Corey G. Diebler and Stephen B. Cumming 

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NASA Dryden Flight Research Center

P.O. Box 273

Edwards, California 93523-0273

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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.

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14. ABSTRACT

A new aerodynamic model has been developed and validated for a modified F/A-18A airplane used for the Active Aeroelastic Wing (AAW) 

research program. The goal of the program was to demonstrate the advantages of using the inherent flexibility of an aircraft to enhance its 

performance. The research airplane was an F/A-18A with wings modified to reduce stiffness and a new control system to increase control authority. 

There have been two flight phases. Data gathered from the first flight phase were used to create the new aerodynamic model. A maximumlikelihood output-error parameter estimation technique was used to obtain stability and control derivatives. The derivatives were incorporated into 

the National Aeronautics and Space Administration F-18 simulation, validated, and used to develop new AAW control laws. The second phase of 

flights was used to evaluate the handling qualities of the AAW airplane and the control law design process, and to further test the accuracy of the 

new model. The flight test envelope covered Mach numbers between 0.85 and 1.30 and dynamic pressures from 600 to 1250 pound-force per 

square foot. The results presented in this report demonstrate that a thorough parameter identification analysis can be used to improve upon models 

that were developed using other means. This report describes the parameter estimation technique used, details the validation techniques, discusses 

differences between previously existing F/A-18 models, and presents results from the second phase of research flights.

15. SUBJECT TERMS

Aerodynamic characteristics, Aeroelastic research wings, Aircraft models, F-18 aircraft, Flight simulation, Flight test, 

Maximum likelihood estimates, Parameter identification

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