Procedia Engineering 67 ( 2013 ) 338 – 346

Available online at www.sciencedirect.com

1877-7058 © 2013 The Authors. Published by Elsevier Ltd. Selection and peer-review under responsibility of the National Chiao Tung Universitydoi: 10.1016/j.proeng.2013.12.033

ScienceDirect

7th Asian-Pacific Conference on Aerospace Technology and Science, 7th APCATS 2013 Effect of Vortex Generator on Lateral and Directional

Aerodynamic Characteristics at Medium Angle of Attack Y. K. Wanga,*, S. Oua, X. Y. Denga, J. Wena

aMinistry-of-Education Key Laboratory of Fluid Mechanics, BeiHang University, Beijing 100191, China

Abstract

A numerical computational investigation of applying vortex generator to conventional figuration aircraft for directional stability improvement at medium angle of attack was conducted in this article.Results showed that vortex generator can put off the critical angle of directional instability from 22°to 29°, in addition, vortex generator also had a very little influence on the lateral aerodynamic characteristic. The flowfield was simulated under the condition of 4° angle of sideslip, Mach number of 0.0882, and Reynolds number of 0.466 million.

© 2013The Authors. Published by Elsevier Ltd. Selection and peer-review under responsibility of theNational Chiao Tung University.

Keywords: vortex generator, lateral stability, directional stability, medium attack angle, aerodynamic characteristics

Nomenclature

CP Surface pressure coefficient, (P-P )/qo

Cy Sectional side force coefficient, 2

0

1 sin2 pC d

Cl Rolling moment coefficient

nC Yawing moment coefficient

nC Yawing moment derivative, nC

Angle of attack (°) Angle of side slip (°)

Circumferential angle(°)

* Corresponding author. Tel: (+86) 010-82317524. E-mail address:wangyankui@buaa.edu.cn

© 2013 The Authors. Published by Elsevier Ltd. Open access under CC BY-NC-ND license.Selection and peer-review under responsibility of the National Chiao Tung University

Open access under CC BY-NC-ND license.

339 Y.K. Wang et al. / Procedia Engineering 67 ( 2013 ) 338 – 346

x Cartesian coordinate

1. Introduction

For the conventional configuration aircraft, the efficiency of directional control rapidly decreases with the increase of angle of attack. According to the investigation of literature, there are two main reasons.

Firstly, the efficiency of vertical tail rapidly declines with the angle of attack increasing. The main reason is that the vertical tail is merged by the upstream separated flow of fuselage and wing. Therefore, the vertical tail cannot provide the required yawing moment to maintain its directional stability. Secondly, the investigation also shows that there is a pair of asymmetric vortexes around its forebody for the conventional configuration aircraft when the angle of attack exceeds 30°[1-2], and these asymmetric vortexes induce a significant amount of side force, which causes directional instability. In order to eliminate the side force, a lot of methods are put forward to control the forebody asymmetric vortex, e.g. forebody strake [3], forebody blowing [4] and forebody disturbance [5]. By changing the state of the asymmetric vortex, the side force is eliminated or used artificially to improve the directional aerodynamic characteristic. However, there is also a great number of aircrafts which become directional instable at medium angle of attack, which is always in the range of 20°to 30°, at this moment, there is no forebody asymmetric vortex to control. Therefore, the conventional methods are useless.

Vortex generator is widely used in the field of aerospace, fluid machinery, automobile and shipping. However, vortex generator is only applied as an engineering method without any explanation for the flow mechanism.In this article, vortex generator is adopted to improve the directional stability of conventional aircraft. It can generate strong vortex which can spread energy to the downstream flowfield.Therefore, vortex generator is used to induce a favorable interference with the downstream flowfield in this article. In addition, the flow mechanism has also been studied.

2. Numerical Computational Method and Model

In thisarticle, the study is conducted by solving the three-dimensional, impressible and steady Navier-Stokes equations, and all the present study make attempt to simulate the flight on the condition of 4° angle of sideslip, a Mach number of 0.0882, and Reynolds number of 0.466 million.

The standard k turbulent model is used in the numerical computational investigation;in addition, the implicit and second-order upwind solver is adopted.

There are two computational models, as shown in Fig.1. Model 1 is a conventional figuration aircraft with wings, horizontal tail, vertical tail, fin and fuselage. Model 2 is created by mounting a vortex generator on the forebody of Model 1. The location and geometry detail of vortex generator are shown as Fig.2 and Fig.3. L represents the length of fuselage.

(a) (b)

Fig.1 Computational models of conventional configuration aircraft, (a) Model 1and (b) Model 2

Y

X

Y

X

340 Y.K. Wang et al. / Procedia Engineering 67 ( 2013 ) 338 – 346

Fig.2 Location detail of vortex generator

Fig3. Geometry detail of vortex generator

3. Boundary Conditions

In the simulated flowfield, the incoming and outgoing flow boundariesare respectively placed at 2 and 2.8 aircraft lengths upstream and downstream from the aircraft apex. In the vertical direction, the computational domain spanned 10mean aerodynamic chords. In the lateral space, the outer boundary is placed at 6 half spans from the plane of symmetry.All these boundaries are adopted pressure far-field condition, and the no-slip wall boundary condition is adopted on the surface of the aircraft.

4. Computational Method Validation

The numerical computational method in this article is verified by comparing the simulation result of Model 1 with the experimental data. It can be concluded from Fig.4 that the aircraft becomes directional instable at 22° angle of attack, which matches experimental result well. Therefore, conclusion can be drawn that the computational method used in this article is valid and reliable.

0.05L

A

A 0

90

180

270 60

Section A-A0.

006L

0.09L

0.02

L

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Fig.4 Yawing moment derivative comparison between computational result and experimental data( 30 /V m s )

5. Results and Discussions

5.1. Effects of Vortex Generator on Aerodynamic Characteristics

The numerical computational results are presented as follow. It can be concluded from Fig.5 that the critical attack angle of directional instability is put off from 22°to 29°when the vortex generator is adopted on model2 and there is almost no influence on lateral aerodynamics characteristic from the vortex generator, as shown in Fig.6.

In order to investigate the flow mechanism improving the directional instability by vortex generator, we can compare the contributions to the yawing moment of whole model from different component and analyze flow over them.

Fig.5 The variation of yawing moment derivative with angle of attack( 30 /V m s )

0 5 10 15 20 25 30

-0.002

-0.001

0.000

0.001

0.002

0.003

Cn

Cn -EXP Cn -CFD

0 5 10 15 20 25 30

-0.0015

-0.0010

-0.0005

0.0000

0.0005

0.0010

0.0015

0.0020

0.0025

Cn

Cn of Model 1 Cn of Model 2

UNSTABLE

STABLE

342 Y.K. Wang et al. / Procedia Engineering 67 ( 2013 ) 338 – 346

Fig.6 The variation of rolling moment with angle of attack( 30 /V m s )

5.2. Contribution to Yawing Moment of Different Component of Aircraft

The yawing moments of different components are extracted from the yawing moment of entire aircraft. By comparing the component yawing moments of Model 1 and Model 2, it can be found that the yawing moment from fuselage of Model 2 greatly increases when angle of attack exceeds 10°compared with that of Model 1 and maintain a small change with the increasing of attack angle, as shown in Fig.7, however, in Model 1, the yawing moment of fuselage rapidly decreases when the angle of attack exceeds 10°. In addition, the yawing moment of vertical and fin are almost unchanged. Therefore, the conclusion can be drawn that the vortex generator mainly has an influence on the flowfield around fuselage.

Fig.7 The variation of yawing moments from different component with angle of attack( 30 /V m s )

5.3. Effect of Vortex Generator on Flow over Fuselage

Fig.8 gives out distribution of pressure-checking sectionson the fuselage.The pressure distribution of sections is extracted from the surface of fuselage, and the side force and yawing moment of each section can be obtained by integrating the pressure distribution along the circumferential angle, the direction of circumferential angle is defined as shown Fig.2.

0 5 10 15 20 25 30

-0.014

-0.012

-0.010

-0.008

-0.006

-0.004

-0.002

0.000

Cl

Cl of Model 1 Cl of Model 2

0 5 10 15 20 25 30

-0.015

-0.010

-0.005

0.000

0.005

0.010

0.015

0.020

Cn

Fuselage of Model 1 Fuselage of Mode 2 Vertical tail of Model 1 Vertical tail of Model 2 Fin of Mode1 Fin of Model 2

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By comparing the variation of section side force and yawing moment of each section with angle of attack in Fig.9 and Fig.10, we can find that there exists obvious influence of vortex generator on flow over section X2, X3 and X4 and little on that over section X1, X5, X6 and X7. From it, we canknow that the side forces of section X2, X3 and X4 on Model 2 increase compared with that on Model 1, which is beneficialto directional stability of the aircraft, so does the yawing moment of section X2, X3 and X4.

Fig.11 gives out the pressure distribution over section X2, X3 and X4 at 20°angle of attack respectively, the x-coordinatestands for circumferential angle, which is defined as shown in Fig.2. Along the clockwise direction, the x-coordinate ranges from 0° to360°.

From the pressure distribution over section X2, X3 and X4,it showsthat the vortex generator mainly has a remarkable influence on flow over the leeward side of section X2 and X3, the pressure of leeward side increases and the pressure of windward side remains a little change. Therefore, the side force provided by forebody of fuselage decreases, which results in the yawing moment of aircraft increasing. As a result the directional instability has been improved and the critical directional instable attack angle has been put off 7 degrees.

The flow detail can also be got from the vorticity field and streamline, as shown in Fig.12 and Fig.13. The vortices coming from the vortex generators have strong effect of flow over the two sides on section X2 and X3, and the flow on the leeward side is completely changed, but the flow on the windward side remains a little difference. This difference between flow over leeward side and windward side on these two sections maybe result from the angle of sideslip.

In addition, it can also be concluded by comparing Fig.12 and Fig.13 that the vorticity on both sides rapidly weaken and their locations gradually lift. Therefore, the vortex can’t affect the flowfield far away from the vortex generator, this is the reason thatwhy the flowfieldaround fin and vertical tail almost remain unchanged.

Fig.8 The position detail of each section of fuselage

X 1 X 2 X 3 X 4 X 5

X 6X 7

0 . 1 L 0 . 1 L 0 . 1 L 0 . 1 L 0 . 1 L

0 . 1 L

0 . 1 L

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Fig.9 The variation of side force of each section with angle of attack( 30 /V m s )

Fig.10 The variation of yawing moment of each section with angle of attack( 30 /V m s )

(a) (b)

0.1 0.2 0.3 0.4 0.5 0.6 0.7

-0.16-0.14-0.12-0.10-0.08-0.06-0.04-0.020.000.020.04

Cy

x

Model 2, a =10o

Model 2, a =15o

Model 2, a =20o

Model 2, a =24o

Model 2, a =30o

Model 1, a =10o

Model 1, a =15o

Model 1, a =20o

Model 1, a =24o

Model 1, a =30o

0.1 0.2 0.3 0.4 0.5 0.6 0.7

-0.040-0.035-0.030-0.025-0.020-0.015-0.010-0.0050.0000.0050.010

Cn

x

Model 1, a =10o Model 1, a =15o

Model 1, a =20o

Model 1, a =24o

Model 1, a =30o

Model 2, a =10o

Model 2, a =15o

Model 2, a =20o

Model 2, a =24o

Model 2, a =30o

0 50 100 150 200 250 300 350

-0.3

-0.2

-0.1

0.0

0.1

0.2

0.3

Cp

Cp-Model 2 Cp-Mode 1

0 50 100 150 200 250 300 350

-1.0-0.8-0.6

-0.4-0.2

0.00.20.4

0.60.8

Cp

Cp-Model 2 Cp-Model 1

345 Y.K. Wang et al. / Procedia Engineering 67 ( 2013 ) 338 – 346

(c)

Fig.11Pressure distributions of section (a)X2, (b)X3 and (c)X4 with the circumferential angle at 20°angle of attack( 30 /V m s )

Fig.12 Vorticity field and streamline of section X2 at 20°angle of attack( 30 /V m s )

Fig.13 Vorticity field and streamline of section X3 at 20°angle of attack( 30 /V m s )

0 50 100 150 200 250 300 350

-1.0

-0.8

-0.6

-0.4

-0.2

0.0

0.2

0.4

Cp

Cp-Model 2 Cp-Model 1

x-vorticity2000180016001400120010008006004002000

-200-400-600-800-1000-1200-1400-1600-1800-2000

x-vorticity2000180016001400120010008006004002000

-200-400-600-800-1000-1200-1400-1600-1800-2000

x-vorticity2000180016001400120010008006004002000

-200-400-600-800-1000-1200-1400-1600-1800-2000

x-vorticity2000180016001400120010008006004002000

-200-400-600-800-1000-1200-1400-1600-1800-2000

Model 1 Model 2

Model 2 Model 1

346 Y.K. Wang et al. / Procedia Engineering 67 ( 2013 ) 338 – 346

6. Conclusion

Several conclusions can be drawn from the above discussion: (1)Vortex generators have a noticeable favorable effect on directional aerodynamic characteristic and the

critical directional instable attack angle of this configuration has been put off 7 degrees. (2)The vortices coming from the vortex generators have strong effect of flow over the two sides on fore

fuselage and especially reduce the flow suction on leeward side obviously. This is the main reason that the yawing moment of aircraft has been improved and directional instable attack angle is put off.

Acknowledgements

The project is supported by the National Natural Science Foundation of China (11272035)and Aeronautical Science Foundation of China(2011ZA51003)and Specialized Research Fund for the Doctoral Program of Higher Education of China (20101102110015).

References

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[3] Daniel G. Murri,Gautam H. Shah,t and Daniel J. DiCarlo NASA Langley Research Center, Hampton, Virginia 23681 Actuated Forebody Strake Controls for the F-18 High-Alpha Research Vehicle JOURNAL OF AIRCRAFT Vol. 32, No. 3, May-June 1995

[4] Ng T T, Suarez C J, Malcolm G N.Forebody Vortex Control Using Slot Blowing[J]. AIAA Paper, 1991: 91-3254. [5] L.E.Ericsson, CONTROL OF FOREBODY FLOW ASYMMETRIY A CRITICAL REVIEW.AIAA-90-2833.