Study on gust alleviation control and wind tunnel test

SCIENCE CHINA Technological Sciences

© Science China Press and Springer-Verlag Berlin Heidelberg 2013 tech.scichina.com www.springerlink.com

*Corresponding author (email: [email protected])

• RESEARCH PAPER • March 2013 Vol.56 No.3: 762–771

doi: 10.1007/s11431-013-5131-7

Study on gust alleviation control and wind tunnel test

WU ZhiGang1*, CHEN Lei2 & YANG Chao1

1 School of Aeronautic Science and Engineering, Beijing University of Aeronautics and Astronautics, Beijing 100191, China; 2 Division of Structural Strength, Shanghai Aircraft Design and Research Institute, Shanghai 200232, China

Received June 7, 2012; accepted December 24, 2012; published online January 22, 2013

Aircraft are inevitably affected by gust during flight, which disturbs the regular operations of pilots and worsens the ride qual-ity. In more grievous cases, flight mission cannot be completed and the flight safety may be disserved. In order to improve the ride quality and the fatigue life of the plane structure under the affect of gust, it is necessary to explore and validate the gust alleviation schemes. Through the low-speed wind tunnel test, the gust alleviation active control technology applied to elastic aircraft is studied. For a large-type passenger plane configuration with high aspect ratio wing, a test system was designed and three gust alleviation control schemes with PID controllers were proposed. Finally the gust alleviation control low-speed wind tunnel test was carried out in the FD-09 wind tunnel. Test results showed that at certain speed and gust frequency, all of the 3 control schemes can alleviate the acceleration at fuselage and wing-tip to a certain extent, as well as the bending moment of wing-root. The gust alleviation control scheme, which uses aileron, elevator and canard as control surfaces synthetically, gives the most satisfying gust alleviation effect.

aeroelasticity, aeroservoelasticity, gust response, gust alleviation, active control, wind tunnel test

Citation: Wu Z G, Chen L, Yang C. Study on gust alleviation control and wind tunnel test. Sci China Tech Sci, 2013, 56: 762771, doi: 10.1007/s11431-013- 5131-7

1 Introduction

Aircraft are inevitably affected by gust during flight, which disturbs the regular operations of pilots and worsens the ride quality. In more grievous cases, flight mission can not be completed and the flight safety may be disserved [1–4]. Gust can be seen as an external exciter to aircraft, which causes changes not only in rigid motions but also elastic vibrations, and results in additional aeroelastic responses. Gust disturbance mainly affects in two ways: 1) Worsen the ride quality for pilots and passengers. Human body is sensi-tive to low frequency vibrations around 1 Hz, so the accel-eration response caused by gust may worsen the ride quality for pilots and passengers seriously; 2) Bring additional gust load, which may shorten the fatigue life of the structure.

High aspect ratio wings are widely used in large scale transporters, long range bombers and HALE (high-altitude- long-endurance) aircraft, and the features of high flexibility and low nature frequency of structure make these aircraft more sensitive to gust disturbance. It becomes a serious problem for designers to alleviate the gust response. In or-der to improve the ride quality and the fatigue life of aircraft structure under the affect of gust, effectual gust response alleviation approaches are required. Generally speaking, the damping of the elastic modes is so small that it is hardly to achieve gust alleviation only through structure design be-cause of the severe weight punishment. Active control technology (ACT) is an advanced design technology of flight control system, through which active control law is designed to drive the control surfaces to deflect accordingly, so gust alleviation is finally achieved [5].

Studies on gust alleviation wind tunnel test and flight test

Wu Z G et al. Sci China Tech Sci March (2013) Vol.56 No.3 763

began in 1960s, and the theoretical methods have been gradually used in engineering. In 1960s–1970s, flight tests to validate the gust alleviation control technology were car-ried out on B-52 [6, 7] and C-5A [8] in America. In 1980s, a series of gust alleviation wind tunnel tests on wings of large-scale transporter were implemented in Japan [9]. In 1980s, flight tests to validate the gust alleviation control system were completed on the civil advanced technology validation aircraft Do128-6 in German [10]. From the mid-dle 1980s to the late 1990s, active flexible wing plan [11] (AFW) and active aeroelastic wing plan [12] (AAW) were proposed successively in America, in both of which gust alleviation was written as a key technology and important content. Since the beginning of the twenty-first century, based on the 3AS Plan of Europe, Israel and Italia have car-ried through a series of wind tunnel test studies on gust al-leviation of the whole plane [13–16]. In 2007, aeroelastic model wind tunnel test of a sensorcraft with fly-wing con-figuration was implemented in America [17], the aim of which was to improve the cruising range and carrying ca-pacity through reducing the structure mass as result of gust alleviation system. Until now, gust alleviation technology has been used in several aircraft in America and Europe, including civil ones such as Airbus-320, Boeing-777 and military ones as B-2.

Studies on gust alleviation control began relatively late in China, and most efforts were made on the design of active control law [18–22]. In the aspect of wind tunnel test, the branch which the author belongs to carried out flutter sup-pression and gust alleviation wind tunnel test for the wing of J8 aircraft in 1990s [23–25]. Since 2008, this branch be-gan the studies on gust alleviation control and wind tunnel test for large-scale aircraft [26, 27]. In February of 2009, gust alleviation wind tunnel test for a single wing with two control surfaces was accomplished with anticipative achiev- ement. This article continues the previous studies to explore the active control scheme of gust alleviation for whole plane configuration through wind tunnel test and to validate the method of gust response analysis and active control design.

2 Mathematic model

2.1 Motion equation of flexible aircraft

Considering a flexible aircraft flying in the longitudinal symmetry plane, the aeroelastic equations of motion in the generalized coordinates can be expressed as [28]

2 2

g g

1 1,

2 2

0 0δ

δV V w

M M C K

δδ δ

Q Q Qδ

(1)

where T1[ ]z y nT R q q is the rigid and elastic

generalized coordinates, is the deflections of control sur-face, wg is gust velocity, is air density, V is airspeed, M and Mare the generalized mass matrices. C is the gener-alized damp matrix. K is the generalized stiff matrix and Q, Q and Qg are the generalized unsteady aerodynamic influence coefficient matrices corresponding to the modal, control surfaces and gust respectively.

2.2 State-space model for gust response analysis

Usually, the aerodynamic influence coefficient in eq. (1) obtained through linear surface element method is described in the frequency domain. In order to set up the state space model of the aeroelastic system, the aerodynamic influence coefficient needs to be handled properly by special tech-nology called rational function approximation, which trans-forms the frequency domain generalized aerodynamic forc-es into the time domain. The unsteady aerodynamic matri-ces can be obtained through rational function approximation by using the minimum state method [29].

T

20 1 2

1

( ) ,N

j j

j j

s s s ss r

D EQ Q Q Q (2)

where Q0, Q1 and Q2 are real matrices, Dj and Ej are real column vectors and rj is a real constant.

By substituting the Q(s) in eq. (2) with Q, Q and Qg in eq. (1), the aerodynamics of modal, control surfaces and gust can be transformed into the Laplace domain. After in-troducing the state term of aerodynamic lag root xa, the aeroelastic equation of flexible aircraft can be written as

ae ae ae aeae ae

ae ae ae ae

,

x A B Ex u w

y C D F (3)

where

T

T T Tae a ,

x ξ x

TT T T

ae , u

T

g g ,w w w

Aae, Bae, Cae, Dae, Eae and Fae are coefficient matrices of the state-space equation. yae is the gust response of flexible air-craft, such as angular velocity, acceleration and inner force of interested structure grid.

The control surface is driven by servo actuator, whose mathematic model can be described by a third-order transfer function [30]. Transform the model of actuator into state- space form as follows:

ac ac acac

ae ac

,0

x A Bx u

u C (4)

764 Wu Z G, et al. Sci China Tech Sci March (2013) Vol.56 No.3

where T T T Tac [ ] x is the state vector of actuator,

u is the control vector, and Aac, Bac, Cac are coefficient ma-trices.

Combining eqs. (3) and (4), we can get

ae ae ae ac ae

aeac ac ac

acae ae ae ac ae

.

0

0 0

0

x A B C E

xx A B u w

xy C D C F

(5)

Eq. (5) denotes the state space model of the aeroelastic system to be controlled, based on which the gust response of flexible aircraft without control system can be solved.

To relate rigid body modes Tz and Ry to airframe states H, , , y, it is required to derive a transformation matrix that can transform rigid body modes into airframe states, namely [31]

1 0 0 0

0 0 1 0.

0 0

0 0 0 1

z

y

z

yy

T H

R

T V V

R

(6)

According to eq. (6), the flight dynamic parameters, such as AOA and pitching angular velocity, can be obtained from the generalized coordinates of the aircraft.

The state-space equations of the feedback control system, including flight stability augmentation system and gust alle-viation system, can be expressed as

cc cc ac

c

.0

Ax Bx y

Cu (7)

Combining eqs. (5) and (7), we can get the state space model of the aeroelastic closed loop system shown as eq. (8), based on which the gust response of flexible aircraft with control system can be solved.

ae ae ae ac ae

aeac ac ac c

acc c ae c ae ac c c ae

cae ae ae ac ae

.

0

0 0

0

x A B C Ex

x A B Cx w

x B C B D C A B Fx

y C D C F

(8)

3 Design of the test system

3.1 Plane model for wind tunnel test

The aerodynamic configuration of the model for wind tun-nel test in this study is scaled from Boeing-737, as shown in Figure 1. It is a half-model in order to obtain larger structure dimensions and adequate space for devices. This model has a regular configuration with low single wing and high hori-zontal tail. The half span of the wing is 1.8 m, and the sweepback angle of the leading edge of the wing is 25°. The

Figure 1 Plane model for wind tunnel test.

half span of the horizontal tail is 0.64 m, and the sweepback angle of the leading edge of the horizontal tail is 35°. The length of the fuselage is 3.5 m, and the total weight is 22.65 kg.

The structure of the fuselage consists of an elastic metal beam, 14 wooden cabins, and connectors for canard, wing and horizontal tail. The structure of wing is similar to the one in ref. [26], which has two aluminum alloy beams in-side, and wooden frames for shape-keeping outside. The structure of horizontal tail is single beam style.

The aileron is mounted on the tail edge of the wing, and the elevator is mounted on the horizontal tail. Furthermore, a canard is mounted beside the pilot cabin in order to study the active control scheme with multiple control surfaces. Each control surface is driven by an independent Maxon micro direct current servo motor, and the frequency band of each motor is beyond 8 Hz by test.

Linear displacement sensors, accelerometers, angular displacement sensors and angular rate gyros are mounted at the gravity center of the model to measure the plunge dis-placement, vertical acceleration, pitch angular displacement and pitch rate of the fuselage. Accelerometers are mounted at the nose, the rear of the fuselage and the tip of wing to monitor the vibrations of the structure, and strain gauges are mounted at the root of the wing beams to measure the bending moment.

In order to obtain the gust response of the test model through theoretical computation, a finite element model (FEM) is completed as shown in Figure 2. Modal analysis for the FEM is carried out with NASTRAN, and the major analysis results of modal frequency are listed in Table 1 along with the test results for comparison. In order to figure out the safe range of wind speed, flutter analysis for the FEM is carried out beforehand. The result shows that the flutter speed of this model is 31 m/s and the flutter fre-quency is 6.4 Hz. Accordingly, it is safe to test at a wind speed range of 10–24 m/s.

3.2 Support system

Compared to the single wing wind tunnel test, one of the


key technologies of the whole plane wind tunnel test is to release the constraint of the rigid motion degree-of-freedom of the test model so as to simulate the free-free boundary condition. Accordingly, a two degrees-of-freedom support system is designed for the wind tunnel test model as shown in Figure 3.

The support system consists of the rotation module, slide module, protection module and mounting base, so it is ca-pable of simulating the pitching motion and plunging mo-tion of the plane. The shaft of the rotation module is con-nected with the plane model, so the model can pitch about the rotation bearing. The slide module is connected with the mounting base by slide rail, so the model can plunge along the slide rail. The pitching motion of the plane model is limited by a disk brake in the protection module, while the plunging motion is limited by a spring damper.

3.3 Gust generator

The major parts of the gust generator are two rectangle blades. The NACA0020 airfoil is used, the span length of

Figure 2 Finite element model of the whole plane.

Table 1 Natural modes and frequencies of the plane model

Mode Description Frequency (Hz)

FEM Test

1 Rigid plunge 0.0 2 Rigid pitch 0.0

3 1st bending of wing 1.72 1.76

6 1st bending of fuselage 5.32 5.33

7 2nd bending of wing 7.65 7.62

9 Pitch of engine 11.0 11.4

10 1st bending of horizontal tail 12.7 12.7

11 Yaw of engine 17.8 17.5

12 1st torsion of wing 18.5 18.9

Figure 3 Support system.

the blades is 2000 mm, and the chord length is 200 mm. The distance between the two blades is 600 mm. The gust gen-erator is placed in front of the test model at a certain posi-tion in order that the wing of the plane model is 1400– 1600 mm downstream the blades of the gust generator. The two blades driven by a direct current motor of 500 W de-flect sinusoidally and the deflecting frequency is 1–6 Hz.

According to the results in ref. [32], when the blades de-flect sinusoidally at certain frequency, the lateral gust which is approximatively sinusoidal can be generated in the test field, and the gust velocity can be written as

g g m( ) sin(2π ),w t A a ft

where wg is the gust velocity; am is the amplitude of the blades deflecting angle; Ag is the gust disturb coefficient which is relevant to wind speed V and blades deflecting frequency f. In this test, am is set to 3°, and Ag is calibrated by test. Figure 4 shows the flow field grid of the gust gener-ator in numerical analysis.

3.4 Stability augmentation system

One of the difficulties in the whole plane wind tunnel test is to stabilize the plane model in the test field at a small angle of attack while the free-free boundary condition is simulated. After preliminary analysis, it is found that the aerodynamic damp and the structure damp of the pitching motion and the plunge motion are both small, and the pitch angle is hardly to be controlled as well as the plunge displacement. Before regular test, it is found that the test model is statically un-stable because of the effect of the fuselage aerodynamics and other unknown reasons. For the sake of safety, an effec-tive stability augmentation system is in badly demand.

In this test, the elevator deflect command is obtained from the proportional feedback signals of pitch angle dis-placement , pitch angle velocity y and plunge displace-ment H.

ELEV .y HK K K H

After preliminary test, the value of each control gain is de-termined as follows: Kθ = 2.3，Kω = 0.4，KH = 40. At a wind speed of 25 m/s, the simulated motion responses of the plane model when the stability augmentation system is open

Figure 4 Flow field grid of the gust generator.


are shown in Figure 5, while the ones when the stability augmentation system is closed are shown in Figure 6. Dur-ing the practical wind tunnel test, this stability augmentation system works well in attitude control and position stabiliza-tion when the wind speed range is 10–30 m/s.

3.5 Measure-control system

The functions of the measure and control module in the gust alleviation test contain stability augmentation control, gust alleviation control, vibration monitor and data recording, as shown in Figure 7. Hardware devices used in the test in-clude USB-6108, PXI-4496 and PXI-6733 data acquisition cards (DAQs) from NI Company in the USA and SDY2101 dynamic strain gauge and SD1476 low-pass anti-mix filter, both of which are homemade. The software module is de-veloped by the branch the author belongs to based on NI Labview software. The stability augmentation control loop and gust alleviation control loop are independent of each other, and the gain of each loop can be adjusted inde-pendently as well as the switches. During the test, the sta-bility augmentation control loop is kept closed all the time, while the gust alleviation control loop can be switched be-

Figure 5 Responses of plane without stability augmentation system.

Figure 6 Responses of plane with stability augmentation system.

Figure 7 Measure-control system diagram of the gust alleviation test.

tween on and off so as to validate the effect of the gust alle-viation system.

4 Design of gust alleviation control

The objective of control scheme design for gust alleviation


is to reduce the acceleration at the fuselage and wing-tip of the aircraft and bending-moment at the section of the wing-root. The signals used as feedback to the control sys-tem include accelerations at the center of gravity of the air-craft aF, pitch angle velocity y, pitch angle displacement and wing-tip accelerations aW. The control system for gust alleviation utilizes aileron AILE, elevator ELEV and canard CANA as its control surfaces.

For the purpose of gust alleviation, the general control scheme should be determined qualitatively first: 1) In order to reduce the acceleration of fuselage, it is more effective to choose elevator or canard as control surface, and the motion signals of fuselage should be fed back to control the rigid motion and the elastic mode of the fuselage; 2) In order to reduce the wing-tip acceleration and wing-root bending moment, it is more effective to choose aileron as the control surface, and the difference between wing-tip acceleration and fuselage acceleration should be used as the feedback signal to control the elastic mode of the wing. Figure 8 shows the general gust alleviation control scheme.

In order to identify the alleviation effects of different control surfaces, three control schemes are considered in this test: 1) Only elevator loop and aileron loop are used; 2) Only elevator loop and canard loop are used; 3) All of the three surfaces are used at the same time.

For each loop in the gust alleviation system, the classical PID control theory is used to design the control law in this study. The general form of PID theory is written as follows:

P I0

d ,t

i j j j jj

u K y K y t

where ui is the deflecting command of each control surface;

yj is the feedback signal; KPj and KIj are proportional gain and integral gain, respectively. Optimization methods can be used to obtain the proper gain values. Because there are several goals of the gust alleviation system and the gust responses are relevant to the wind speed and the gust fre-quency, certain tradeoff should be introduced between all the gust alleviation goals according to practical condition. In order to avoid complexity, a set of effective gain values is obtained through an analysis-adjust procedure as shown in Table 2.

5 Gust alleviation wind tunnel test

The gust alleviation wind tunnel test for flexible plane is carried out in FD-09 low speed wind tunnel of the Institute of Astronautics Aerodynamics Technology. The size of the wind tunnel is 3 m×3 m as shown in Figure 9. It is neces-sary to point out that the angle of attack of the plane model is maintained to be small during the test, and the situation in Figure 9 is for static exhibition only.

Figure 8 Gust alleviation control scheme of the flexible plane. Table 2 Control gains of each scheme

Scheme 1

Feedback signals Aileron Elevator Canard

KP KI KP KI KP KI

Acceleration of wing-tip 2.5 25.0

Acceleration of fuselage 2.5 25.0 4.0 40.0

Angular rate of fuselage 0.3 0.0

Angular displacement of fuselage 0.5 0.0

Scheme 2


KP KI KP KI KP KI

Acceleration of wing-tip 2.5 25.0 Acceleration of fuselage 2.5 25.0 0.0 0.0

Angular rate of fuselage 8.0 0.0

Angular displacement of fuselage 40.0 0.0

Scheme 3


KP KI KP KI KP KI

Acceleration of wing-tip 2.5 25.0 Acceleration of fuselage 2.5 25.0 4.0 40.0 0.2 0.5

Angular rate of fuselage 0.1 0.0 0.0 0.0

Angular displacement of fuselage 0.2 0.0 0.0 0.0


Figure 9 Test model in the FD-09 wind tunnel.

5.1 Open-loop gust response

Open-loop gust response is defined as the response of the plane when the stability augmentation system is closed while the gust alleviation system is opened. For the safety of the test model, the wind speed range in test is 12–24 m/s, and the gust frequency range is 1.0–5.0 Hz. The plane mod-el is forced to vibrate sinusoidally under the perturbation of gust. Open loop gust responses at different wing speeds and gust frequencies are shown in Figure 10, and Figure 10(a), (b) and (c) represent the fuselage acceleration, wing-tip ac-celeration and wing-root bending moment, respectively. It can be seen that the responses of fuselage acceleration are excited by the first bending mode of the wing which is about 2.0 Hz. With the wind speeding up, the acceleration at the wing-tip has its peak value for low frequency changed from 2 to 3 Hz. The major reason for the entire phenomenon is that the first bending mode of the wing increase as wind speeds up. As for fact that the bending moment at the wing-root sections is more obvious at frequency of 1–3 Hz than 4–6 Hz, that is mainly because the first bending mode of the wing happens around frequency of 2 Hz.

5.2 Analysis of gust alleviation efficacy

In this paper, the gust alleviation efficacy is defined as

open close

open

( , ) 100%,A A

V fA

where Aopen and Aclose are the amplitudes of gust responses when the gust alleviation system is opened and closed re-spectively; alleviation efficacy is relevant to wind speed V and gust frequency f.

The theoretical results and experiment results corre-sponding to the three control schemes at the wind speed of 18 m/s are shown in Figures 11–13. It can be seen from Figure 11 that the three schemes are not able to alleviate the fuselage acceleration at the frequency of 1–1.5 Hz, but they work equivalently at the frequency of 1.5–5.0 Hz. It can be

Figure 10 Results of the open loop gust response wind tunnel test. (a) Acceleration of the fuselage; (b) acceleration of wing-tip; (c) bending moment of the wing-root.

seen from Figure 12 that the three schemes cannot alleviate the wing-tip acceleration at the gust frequency of 1–1.5 Hz, but they function nearly the same at frequency of 1.5– 5.0 Hz. It can be seen from Figure 12 that the responses of wing-root bending moment are much stronger at low fre-quency, and all of the three control schemes work well at gust frequency of 1–3.5 Hz; the responses of wing-root bending moment are smaller at 4–6 Hz, and the gust allevia-tion efficacies of the three control schemes are less.

According to the results above, it can be said that the ai-


Figure 11 Gust response of plane model (acceleration of fuselage). (a) Numerical results; (b) test results.

leron loop in which the difference between the fuselage acceleration and the wing-tip acceleration is used as the feedback signal is capable of alleviating the wing-tip accel-eration and the wing-root bending moment caused by gust. The alleviation efficacy of the aileron loop is not affected by the elevator loop or canard loop. The general control scheme that the aileron loop is used to alleviate the wing-tip acceleration and the elevator loop or canard loop is used to alleviate the fuselage acceleration is proved feasible by the wind tunnel test.

The alleviation efficacies of each control scheme acting on the fuselage acceleration are shown in Table 3. It can be seen that control scheme 1 works better at the wind speed of 16–18 m/s than at 20 m/s; control scheme 2 works better at the wind speed of 18–20 m/s than at 22 m/s; control scheme 3 is more effective among much wider range of wind speeds and gust frequencies. Therefore, the conclusion is drawn out that scheme 3 is the most efficient means to alleviate fuse-lage accelerations, and scheme 1 takes the second place, leaving scheme 2 the most inefficient.

The alleviation efficacies of each control scheme acting on the wing-tip acceleration are shown in Table 4. It can be

Figure 12 Gust response of plane model (acceleration of wing-tip). (a) Numerical results; (b) test results.

seen that scheme 1 can gain better effect at the wind speed of 18 and 20 m/s than at 16 m/s, and scheme 2 functions the same at all wind speeds with a good effect, while by com-paring status at 18 and 20 m/s, scheme 3 can be rated as the best means from both scope of frequency and effectiveness. Moreover, scheme 2 takes the second place, and scheme 1 is the most inefficient in this respect.

The alleviation efficacies of each control scheme acting on the wing-root bending moment are shown in Table 5. It can be seen that at gust frequency of 1.5–3.0 Hz, control scheme 1 can alleviate the wing-root blending moment by 15.0%–59.0%, and control scheme 2 can achieve an allevia-tion efficacy of 18.0%–56.5%, while alleviation efficacy of control scheme 3 is 11.5%–55.6%. According to the results above, the three schemes have similar effects to alleviate the wing-root blending moment at the gust frequency of 1.5– 3.0 Hz.

According to the comparison above, it comes to the con-clusion that all of the three gust alleviation control schemes only work well at certain range of wind speed and gust fre-quency in alleviating fuselage acceleration, wing-tip accel-eration and wing-root bending moment. Utilizing aileron,


Figure 13 Gust response of plane model (bending moment of wing-root). (a) Numerical results; (b) test results.

Table 3 The alleviation efficacies of the three control schemes (acceleration of fuselage)

Scheme 1

Wind speed (m/s) 16 18 20

Frequency (Hz) 1.5–3.5 4.5–5.0 1.0–5.0 1.5–2.0 3.5

Alleviation efficacy 12.0%–69.0% 62.0%–91.2% 18.8%–52.7% 10.0%–30.3% 14.1%

Scheme 2


Frequency (Hz) 1.0–3.0 4.5–5.0 1.5–2.5 4.5 3.5 4.5

Alleviation efficacy 24.8%–63.9% 28.1%–52.8% 28.6%–57.3% 27.5% 28.6% 16.1%

Scheme 3


Frequency (Hz) 2.0–3.0 4.0–4.5 1.0–3.0 4.0–4.5 1.0–2.5 3.5–4.5

Alleviation efficacy 32.4%–90.4% 22.6%–34.2% 30.8%–55.5% 27.3%–66.5% 10.3%–47.1% 13.0%–18.1%

Table 4 The alleviation efficacies of the three control schemes (acceleration of wing-tip)

Scheme 1


Frequency (Hz) 2.0–3.5 2.0–4.5 1.5–4.0

Alleviation efficacy 18.5%–51.9% 20.0%–49.5% 13.8%–44.3%

Scheme 2


Frequency (Hz) 1.5–4.5 1.5–4.5 1.5–4.0


Scheme 3


Frequency (Hz) 1.5–4.0 1.5–4.5 2.0–4.0


Table 5 The alleviation efficacies of the three control schemes (bending moment of wing-root)

Scheme 1


Frequency (Hz) 1.5–3.0 5.0 1.0–3.5 1.0–3.0

Alleviation efficacy 16.7%–47.2% 11.2% 15.8%–59.5% 16.9%–54.6%

Scheme 2


Frequency (Hz) 1.5–3.5 1.5–3.0 1.5–3.0


Scheme 3


Frequency (Hz) 1.5–3.0 1.0–3.5 1.0–3.0



elevator and canard as control surfaces, scheme 3 outstands among all the schemes for its widest range of frequency and best alleviation effect.

6 Conclusion

Focusing on a large-type passenger plane configuration with high aspect ratio wing, we designed the test system includ-ing test model, support system, gust generator, stability augmentation system and measure-control system, and car-ried out the gust alleviation control low speed wind tunnel test in the in FD-09 wind tunnel.

Numerical gust responses were obtained through time domain method, based on which three gust alleviation con-trol schemes with PID controllers for the whole plane were proposed. Test results have shown that at certain speed and gust frequency, all of the three control schemes can alleviate the acceleration at fuselage and wing-tip to a certain extent as well as the bending moment of wing-root. The gust alle-viation control scheme 3 which uses aileron, elevator and canard as control surfaces synthetically gives the most sat-isfying gust alleviation effect.

The effectivity of the gust response analysis method and the gust alleviation control method were verified by the wind tunnel test. The relevant work in this study is valuable for engineering application.

This work was supported by the National Natural Science Foundation of China (Grant Nos. 91116005, 10902006).

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Study on gust alleviation control and wind tunnel test