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Aerospace Science and Technology 28 (2013) 100132

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Aerospace Science and Technology

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Review

Aerodynamic technologies to improve aircraft performance

A. Abbas a , J. de Vicente b , E. Valero b,

a Airbus Spain, Paseo John Lennon, Getafe, Madrid, Spainb Universidad Politcnica de Madrid, School of Aeronautics, Plaza Cardenal Cisneros, 3, Madrid, Spain

a r t i c l e i n f o a b s t r a c t

Article history:Received 31 July 2012Received in revised form 11 October 2012Accepted 24 October 2012Available online 30 October 2012

Keywords:Aerodynamic efficiencyFlow controlDrag reductionInnovative congurationsFlow separation technologies

An Air Transport System has become an indispensable part of Europes economic infrastructure. The Com-mercial Aeronautics Sector is well aware that it has to nd an acceptable balance between the constanterce competitive pressures upon it and the publics expectations of cheaper fares but reduced environ-mental impact including community noise around airports and global warming. In order to achieve sucha balance in the future, a strategy is required for competitive excellence dedicated to meeting societysneeds.The realization of this vision cannot be achieved without signicant technology breakthroughs in thearea of aerodynamics and other disciplines such as materials and structures. Improved aerodynamic de-signs and the introduction of new aerodynamic technologies should play not only a key role in improvingaircraft performance but, also, contribute strongly to product cost and operability. Substantial R&T explo-ration and development require to be conducted in order to provide the required technologies.In this work, a review of those technologies which show a potential to deliver breakthrough improve-ments in the aerodynamic performance of the aircraft is shown. The focus of this report is on new aircraftcongurations to reduce induced drag and noise, laminar and turbulent drag reduction technologies andow control devices, which aims to improve the performance of the airplane under separated ow con-ditions of unsteady nature, and to reduce the complex high-lift devices. Most of these works have beenexposed in previous KATnet conferences (Key Aerodynamic Technologies for Aircraft Performance Im-provement), although a general overview of the current status of these technologies is included.

2012 Elsevier Masson SAS. All rights reserved.

Contents

1. Introduction. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .2. Aircraft conguration technologies . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1

2.1. Blended wing body (BWB) and boundary layer ingestion (BLI) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1022.2. High aspect ratio . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .2.3. Engine concept . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2.4. Forward swept wings (FSW) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

3. Drag reduction technologies . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3.1. Drag reduction by extended laminar ow . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 107

3.1.1. Laminar ow technologies . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 13.1.2. Laminar ow on nacelles . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1

3.1.3. Hybrid laminar ow . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3.1.4. Alternative laminar ow technology . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 112

3.2. Turbulent skin friction reduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 114. Separation control technologies . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1

4.1. Passive ow control devices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 14.2. Active ow control devices (AFC) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1

4.2.1. Blowing method . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4.2.2. Vortex generator . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

4.3. Other applications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5. System and certication issues . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1

5.1. Safety . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5.2. Performance and ight handling characteristics . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1285.3. Environment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5.4. Industrial issues . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

1270-9638/$ see front matter 2012 Elsevier Masson SAS. All rights reserved.http://dx.doi.org/10.1016/j.ast.2012.10.008
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A. Abbas et al. / Aerospace Science and Technology 28 (2013) 100132 101

6. Implications to aerodynamic tools . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .6.1. Computational uid dynamics . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .6.2. Wind tunnel testing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

7. Conclusions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . References . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Fig. 1. Presentation of FUSIM project (Sept. 2008, Flight Physics, AIRBUS).

1. Introduction

The European societys increasing environmental awareness hasbeen present always in the aeronautical community; industry; andresearch centers. This has had a denite inuence in the way theyforesee the aircraft of the future. In this regard, the ACARE Visionfor 2020, a Group of Renowned Personalities in the aeronauticaleld, formulated a clear set of requirements for civil transportaircraft operation so that the following specic goals could beachieved:

A ve-fold reduction in accidents. Halving perceived aircraft noise. A 50% cut in CO2 emissions per passenger-km. An 80% cut in NOx emissions. An air traffic system capable of handling 16 million ights a

year. 99% of all ights within 15 minutes of timetable.

Although undoubtedly important, an additional secondary objective is the reduction of the enormous dependence of the fuelprices in the nal DOC (Direct Operating Cost) of a typical longrange aircraft. To give an idea, a Brent Crude price rise from 1.5to 4 dollars per gallon supposes a 53% increase in the DOCs fuelprice.

Most of these goals have a direct impact on the aircrafts aero-dynamic performance, mainly on aerodynamic technologies for amore effective, environmentally friendly air transport system. Theaeronautical industry is aware that these objectives can not beachieved only with improvements in the current standards con-gurations ( Fig. 1). A step change in performance is necessary,and this must be accomplished through new breakthrough aero-dynamic technologies.

In this context and nanced by the European Commission, theEuropean Coordination Action KATnet I and II (Key AerodynamicTechnologies for Aircraft Performance Improvement) were con-ceived. KATnets main objective was to support the ACAREs globalstrategic approach. This was achieved by the development of acommon RTD strategy in certain technology areas, and by provid-

* Corresponding author.E-mail address: [email protected] (E. Valero).

ing a communication platform for all aircraft disciplines concerneddirectly or indirectly with aircraft performance improvement.

Within KATnet, the original environmental objectives weretranslated into 10 different aerodynamic objectives ( Table 1 ) whichwere supported by sixteen design concepts ( Table 2 ).

The necessary technology developments to fulll these designconcepts requires considerable changes in new aircraft congura-tions; research in cruise drag; and ow control.

New aircraft congurations.Preliminary design studies of New or Non-Conventional cong-urations have shown that a step change in cruise L/D and noiseare possible but that step is only about half that required.Therefore, it will be necessary to investigate new technolo-

gies which could be applied to such congurations and whichmight deliver the additional improvements in performance re-quired to meet the target.

Cruise drag.To deliver the required step change in cruise drag, the focus of future research and development efforts must be on technolo-gies aimed at delaying laminar to turbulent boundary layertransition and at manipulating turbulent ow structures closeto the aircraft surface. The timescales involved in maturingthese technologies, particularly the ability to manipulate tur-bulent ow structures, are extremely aggressive and carry ahigh risk.

Flow control.Flow control technology describes a variety of techniques by

which aerodynamic performance can be enhanced to levelsbeyond those which can achieved by changes to the exter-nal shape alone. The change in aerodynamic performance maytake the form of enhanced lift; reduced drag; controlled un-steadiness; and reduced noise or delayed transition. Alterna-tively, benets for the same levels of performance may beaccrued through reduced system complexity; less weight; lessmaintenance or reduced life cycle costs.In addition to the work on ow control technologies, work onreducing aircraft weight including high lift and other controlsurfaces is of prime importance since this has a direct ef-fect on fuel burn through reducing overall aircraft weight. Thismeans that technologies to control separation must be devel-oped also.

Besides the environmental goals, other major objectives are thereduction of lead time and the provision of robust solutions withimproved quality. In that context, it is important to exploit any op-portunity provided by enhanced or new classes of tools such ase.g. high delity Computational Fluid Dynamics (CFD) and power-ful High Performance Computing (HPC) capabilities. Also, improvedwind tunnel technology provided by new measurement techniques(e.g. Pressure Sensitive Paint) or model material and manufactur-ing processes must be exploited.

KATnet has served as a forum to expose the technology ad-vances of both industry and academia in the above mentionedareas. A series of conferences and workshop were organized withgreat success. These conferences exposed the most recent advancesof the main aeronautic stakeholders and showed the industrysgrowing interest in these topics. This review aims to describe the
mailto:[email protected]:[email protected]


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102 A. Abbas et al. / Aerospace Science and Technology 28 (2013) 100132

Table 1Aerodynamic objectives.

E n v i r o n m e n t a l g o a l s

Reduce emissions Reduce drag Reduce vortex dragReduce wave dragReduce friction dragReduce pressure drag

Reduce weight Reduce critical loadsReduce community noise Increase structural efficiency

Reduce source noise Reduce airframe source noise contributionReduce engine noise contribution

Increase airport capacity Increase landing & take-off rates Reduce separation distancesImprove affordabil ity Reduce a ir frame cost Reduce complexi ty

Table 2Design concepts.

H i g h a s p e c t r a t i o w i n g

A d a p t i v e s e c t i o n s

M o r e h i g h l y l o a d e d s e c t i o n s

D i r e c t s h o c k c o n t r o l

L a m i n a r o w p r o m o t i o n

T u r b u l e n t s k i n f r i c t i o n d r a g r e d u c t i o n

M o r e e f f e c t i v e

g u s t a n d m a n e u v e r

l o a d a l l e v i a t i o n

c o n t r o l

U n s t e a d y

o w

s e p a r a t i o n

a n d

t u r b u l e n t n o i s e

s o u r c e

U n s t e a d y l o a d s c o n t r o l ( i

. e .

u t t e r

L C O a n d b u f f e t

o n s e t d e l a y )

M o r e e f f e c t i v e ( s m a l l e r ) c o n t r o l s u r f a c e s

f o r g i v e n a i r c r a f t c o n t r o l r e q u i r e m e n t

M o r e e f f e c t i v e ( s m a l l e r ) h i g h l i f t d e v i c e s

f o r g i v e n l o w s p e e d r e q u i r e m e n t

M a n a g e m e n t / e l i m i n a t i o n o f u n s t e a d y o w

s e p a r a t i o n s +

t u r b u l e n t s o u r c e n o i s e

E n g i n e n o i s e s h i e l d i n g

E n g i n e e x h a u s t

c o n t r o l & t r u s t v e c t o r c o n t r o l

R e d u c e d w a k e

v o r t e x s i g n a t u r e f o r

g i v e n l o w s p e e d r e q u i r e m e n t

S i m p l i e d h i g h

l i f t s y s t e m f o r g i v e n

l o w s p e e d r e q u i r e m e n t

Reduce vortex drag XReduce wave drag X XReduce friction drag X XReduce pressure drag XReduce critical loads X X X XIncrease structural efficiency X X X XReduce airframe source noise X XReduce engine noise X X X XReduce separation distances X XReduce complexity X

main conclusions and works presented in these series of confer-ences. However, the paper tries not to be limited to KATnet and,in taking advantage of this forum, it will give a general overviewof the recent developments in these technologies. In any case, thispaper does not aim to describe separately each of these technolo-gies in detail. Outstanding reviews can already be found in theliterature [24,38,91,98] and, although it is always possible to up-date those works (some of them are more than 10 years old), adetailed and throughout description of all those technologies illus-trated in KATnet is beyond the objectives and the extension of thiswork. Also, the authors wish to point out that this review is bi-ased unavoidably in favor of KATnet activities. As such, we mayhave omitted unintentionally important studies pertaining to owcontrol technologies. It is important, also, to highlight the difficultyassociated with this task. Most of these works were presented onlyin series of KATnet conferences and internal reports, and we foundno further publications in journals or international conferences.Whenever possible, we included the reference to a journal paperor conference proceedings but, many times, the work which is ex-posed here, is unpublished.

2. Aircraft conguration technologies

This general topic comprises of two main different strategies.On the one hand, there are small modications in the geometrywhich produce substantial improvements in drag reduction (fric-tion and wave) and reduce detached areas. Small ow control de-vices as synthetic jets or vortex generators belong to this category.These technologies will be described with detail in Section 4. On

the other hand, there are large geometrical modications over clas-sical aircraft conguration or other innovative congurations whichare aimed at improving aircraft performances substantially by re-ducing drag and/or weight. This section is dedicated mainly to thispoint.

Different lines of action can be addressed.

2.1. Blended wing body (BWB) and boundary layer ingestion (BLI)

In 1994, NASA sponsored one of the rst attempts to studythe feasibility of BWB congurations ( Fig. 2). Liebecks prelimi-nary results [63,64] , showed potential savings in: fuel burn (27%);takeoff weight (15%); operating empty weight (12%); total thrust(27%); and a higher lift/drag (20%). The study was performed ina 800 passenger BWB for a 7000 miles design range compared toa conventional aircraft. Despite these promising results, an impor-tant number of drawbacks had to be solved in order to make thisaircraft technically viable. A new eld of study, related to fuselage-wing integration, is identied: structural integration; aerodynamicstability; the elimination of the conventional empennage; and thepresence of a non-circular fuselage etc.

Additionally, the engines in the considered conguration aremounted near the trailing edge on the upper surface of the wing.Initially, aircraft designers used pylon-mounted nacelles to avoidproblems of surface integration and inlet ow distortion resultingfrom ingesting the incoming boundary layer. However, recent stud-ies indicate that boundary layer ingestion (BLI) offers additionalbenets including reduced ram drag, lower structural weight andless wetted area than a pylon mounted conguration. Because of


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Fig. 2. Blended wind body conguration from [64] .

that boundary layer ingestion, nacelles started to attract the scien-tic communitys attention [91] .

Boundary layer ingestion means taking the fuselage boundarylayer ow through the engines for the purpose of improving fuelefficiency. The idea comes from re-energizing the aircraft wakewhich enables less kinetic energy to be wasted. Comparing the sit-uation with a typical podded engine, it can be proved that, for agiven thrust force, less power needs to be added to a ow enteringthe engine with a lower velocity. Consequently, due to bound-ary layer ingestion, the lower inlet velocity means that the samepropulsive force can be achieved with less power. However, thegain is not without drawbacks, the incoming ow to the engines ishighly non-uniform and produces loss of performance, additionalstress and fatigue to the blades. Moreover, the aerodynamic block-age, associated with the fuselage boundary layer, is much largerthan that due to the duct boundary layer. Therefore, it has a majorrole in the achievable ow rate and the increase in fan pressure fora given thrust. The non-uniformity affects, also, the nozzle exit mo-mentum ux, and the degree to which this occurs in much largerthan typically associated with civil engines. This effect can be alle-viated by careful design of the input S-duct to the compressor andthe use of active or passive ow devices which energize the inputboundary layer in order to avoid detached ow and ow distortionproblems.

In order to reduce the ow pressure distortion at the fan en-trance, the detailed design of the input duct, has been studied by[77] , who considered variations of the inlet duct offset; curvatureof the two bends; area ratio and scalloping of the pre-compressionregion ahead of the intake. It was found that duct offset was themost important parameter governing the strength of the secondaryow and impacted on intake recovery. In these studies, authorsdemonstrated, also, that strong pressure distortion could eliminatethe power saving related to the BLI effect.

The lack of numerical and experimental data available to com-pare and validate new methods and tools is an important obstaclein the development of the BWB. In this line, Carter et al. [19,20]conducted a series of numerical and experimental test cases on

Fig. 3. Example of total drag breakdown (2002 standard).

the Boeings BWB-450-1L model. These focused on the determi-nation of the effectiveness of the trailing edge devices (elevons,drag rudder and winglet rudders), tested at various angles of

attack; sideslip angles; and Mach numbers. The computationalwork focused on particular cruise condition of Mach = 0.85 andRe = 75 millions. Besides, in the range of Mach = 0.2 to 0.88 andRe = 2.4 to 75 millions, four different congurations (no nacelles;pylon-mounted nacelles; BLI nacelles; and redesigned BLI nacelles[19] ) were studied.

2.2. High aspect ratio

In large transport aircrafts, during cruise ight in a still air con-guration, drag is mainly due to friction drag (about 47%) andinduced drag (about 43% see Fig. 3).

Several strategies to reduce the friction drag of conventional air-craft are under examinations and will be reviewed hereafter.

The induced drag which is the other big source of drag de-pends on the span and the lift distribution along the wing span. In


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Fig. 4. Spiroid loop (left) and downward pointing (right) wing tip devices from [42] .

Fig. 5. Design box (hatched zone) from [42] .

conventional large transport aircraft, the lift distribution is so op-timized that no signicant reduction seems possible in the futurewithin the current design approaches. An alternative approach, fora given lift, is obtained through adopting a high aspect ratio wingor a wing tip device. Extensive literature can be found on this topiccovering the theoretical background underlying the induced dragprediction and methods to reduce it [57] ; the link between thewingtip shapes and the induced drag [16] ; and studies of differentwingtip devices [1,27,55,97] .

ONERA, Airbus and Technische Universitat at Braunschweig [42]carried out an interesting study within the M-DAW Project (Mod-eling and Design of Advanced Wing Tip Devices). Two differentand innovative congurations were analyzed and optimized: thespiroid loop; and the downward pointing wingtip devices (Fig. 4).Both concepts can have better structural characteristics related tothe wing root bending moment. For reference, we considered thewing of a generic long haul aircraft. We compared the new solu-tions and the new designs with a standard blended winglet. Themodications were limited to 4% of the wingtip and the height of the ground vehicles limited the downward device extension (seeFig. 5). The wing was designed for cruising at Mach = 0.85.

Previous studies of the spiroid winglet [39] showed drag gains,for a slightly lower root bending moment, of the same order of standard wingtip devices. However, it suffers high-transonic inter-actions in the loop. A specic optimization to alleviate this ef-fect was carried out. The optimization was based on 45 designvariables (sweep; maximum thickness; twist; and camber). Someshape design variables were prescribed (spiroid size; minimumchord length; minimum aerofoil relative thickness (5%); and thick-ness law with the same span extension and wetted area of theequivalent blended winglet). Although slightly more efficiency athigh CL in keeping the same root bending moment, the compar-ison showed a discouraging drag reduction of about 86% of theblended winglet drag reduction. On the contrary, a detailed studyof the downward pointing winglet, which considered high and lowspeed, root bending moment and lateral stability, showed somebenets without incurring any major redesign of the wing box. Weobtained promising gains of about 13% in range over a long-ightscenario and an improvement of 46% in low speed L/D.

Another line of research was related to the box-like wing con-cept, which, theoretically, produced the minimum induced drag fora given lift (Prandtl [79] ). The following conditions are satised:same lift distribution and same total lift on each of the horizon-

Fig. 6. Span efficiency for various optimally loaded non-planar systems from [58] .

tal wings; and buttery shaped lift distribution on the vertical tipwings. Following these ideas, Kroo [58] showed a numerical com-parison between the efficiency of different non-planar congura-tion (induced drag of planar wing/induced drag of the non-planarsystem of the same span and lift Fig. 6) for several non-planargeometries. Each of the geometries permitted a vertical extent of 20% of the wing span. Such designs may be of interest because of their potential for lower vortex drag at a xed span which is a key

constraint for many aircraft including very large commercial trans-port concepts.This unconventional non-planar conguration provides the

technology breakthrough necessary to obtain substantial gains indrag vortex reduction. Between these solutions, the wingtip device,already implemented in commercial aircraft, and the joined wingsare probably the ideal Prandtl wing solution.

The practical application of a box-wing can be seen in the Joined-Wing concept (Fig. 7). First proposed by J. Wolkovitch [107] ,this kind of conguration increases substantially the high aspectratio with a theoretical induced drag reduction of up to 40% [35,96] . However, several non-aerodynamics issues must be studied in-cluding the effects on stability and control, characteristics of wakevortices or structural implications. Frediani et al. [35] carried outa practical implementation of this concept on a modied A380.The project focused on stability and structural issues. After an op-timization procedure, they found that the rear wing could not beconnected to the rear fuselage but had to be positioned over thefuselage itself and connected to it by means of two ns ( Fig. 8).This conguration proved to be stable in cruise ight; and, thelift was distributed equally on the front and rear wings giving themaximum L/D improvement. From the structural point of view, thefuselage is equivalent to a doubly supported beam; the supportsbeing the front and the rear wing and, then, bending stresses inthe fuselage are close to zero in the front and rear wing roots.The eigenmodes of the aircraft are completely different from aconventional one; in particular, the lateral bending modes of a con-ventional fuselage are not longer present. The damping momentand the moment of inertia along the pitch axis are higher thanits equivalent conventional aircraft; hence, the ight qualities are


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Fig. 7. Example of non-planar wing conguration, Lockheed box-wing.

Fig. 8. Initial aircraft with joined wing studied by [35] .

considered to be very satisfactory; but this rises problems in theproperly design of the pitch control system [34] .

In the same line, it is worth mentioning the strut-braced wing(Fig. 9) which uses a strut for the wing bending load alleviation.This allows the aspect ratio to be increased and the wing thick-ness to be reduced. A thinner wing has less transonic wave drag;therefore, it is possible to unsweep the wing allowing larger ex-tensions of natural laminar and further structural weight savings.

The optimization of these conguration is concerned usually withstructures, weights, and stability and control [12,41] .

2.3. Engine concept

Increasing the by-pass ratio of turbofans is a successful formulaapplied currently to increase the propulsive efficiency. Geared tur-bofan (GTF) and distributed propulsion are considered to be thetechnology concepts for an increased engine by-pass ratio. How-ever, the growth of the bypass ratio is accompanied by larger andheavier nacelles. At a certain point, the associated nacelle weightincrease and drag penalty outweighs the growth in propulsive ef-ciency. The Contra-Rotating Open-Rotor (CROR) offers a break-through solution. Operating without the drag and weight penalty

of a large nacelle, the by-pass ratio is no longer a limiting factor.However, intrinsically, the open-rotor is noisier than a turbofan atan equivalent thrust setting. To the authors knowledge, very fewstudies of open-rotor congurations can be found in the litera-ture. To give an example, at the 2010 International Congress of the Aeronautical Science meeting (ICAS), there was presented onlyone work on this subject [60] . We could nd, also, no references toopen-rotor at the 46th AIAA Joint Propulsion Conference. This doesnot mean that the scientic community is not increasing their ef-forts in understanding this technology. However, the main effortsare found in an industrial framework. Clearly, different EuropeanProjects are concerned with the investigation and application of this technology, e.g. DREAM, NACRE and, more recently, CleanSky.These are focused on studying tail-mounted CROR conguration on

civil aircrafts (see Fig. 10).The European Project NACRE (New Aircraft Concepts Research)

carried out a comparative study between an Open-Rotor and a con-ventional turbofan. The turbofan was represented by the CFM56-5series with 1995 technology. It had a 1.8 m fan diameter; by-pass ratio of 6.6; and a maximum trust of 1915 daN These gavea typical fuel efficiency at cruise condition of 0 .57 kg / hr / daN. Theequivalent open rotor had a 3.72 m diameter and maximum trustof 1777 daN, giving a typical fuel efficiency of 0 .46 kg / hr / daN,with a 20% of theoretical reduction. The results for a dened civilaircraft mission at 33000 ft altitude with 180 pax of payload and

Fig. 9. Strut-braced wing with tip-mounted engines from [41] .


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Fig. 10. Open rotor conguration studied in NACRE.

Fig. 11. Comparative study between open-rotor and turbofan technologies.

a range of 3000 NM are shown in Fig. 11. These studies obtained apromising 23% reduction in mission fuel burn and 8% reduction inMTOW.

Compared with its equivalent turbofan, one of the main draw-backs of CROR relates to vibration and acoustic problems, in whichthe nacelle acts as an efficient noise shield. Due to the airframesurfaces, the acoustic propagation can be alleviated by adoptingshielding and/or reection. An example is NACRE [60] s study of the U-tail conguration ( Fig. 12) In any case, signicant develop-ments are required to integrate this engine into an aircraft.

Another important issue is the necessity to develop specicdesign tools. A Contra-Rotating Open-Rotor shares some basic fea-tures with the classic propeller, for which various numerical tech-niques exist. However, it is the vortex and viscous wake interactionbetween the pylon and the rotors and between the front-rotor andthe aft-rotor which make the picture more complex. To understandthe underlying physical phenomena specic to a Contra-RotatingOpen-Rotor, it is required to solve accurately the tip vortices andviscous wakes which emanate from upstream stage and interactwith the second.

2.4. Forward swept wings (FSW)

The concept of FSW is not entirely new. In the investigation of the different methods of delaying the onset of the increase of drag

Fig. 12. Open rotor conguration studied in NACRE. Acoustic interaction betweenthe CROR and an U-tail conguration.

for ight in the vicinity of the speed of sound, it was found thatsweeping the wings provided the most effective technique of in-creasing the drag divergence Mach number. Aerodynamically, thesame effect can be obtained regardless of the direction of sweep.However, current aircraft designs favor the use of aft sweepingin order to avoid the phenomenon of structural divergence, in-herent in FSW operating at a high dynamic pressure condition,which cannot be solved through metallic structures without a wingweight penalty. Today, technology advances in composite materialsare providing a promise of eliminating this problem with little orno wing weight penalty.

Some of the benets of FSW include spin resistance; extendedhigh angle of attack and lateral control; and lower transonic ma-neuvering drag. Generally, The FSW separation pattern starts fromthe root and propagates gradually outboard. This allows attachedow to be maintained over the outboard wing, and retains aileroneffectiveness at high angles of attack where, in lateral control, con-ventional (backward) wings may exhibit degradation in lateral con-trol. Other additional benets of FSW are theoretical lower proledrag due to lift and induced drag (for the same lift) [105] . Be-tween the disadvantages, larger trailing edge sweep can strengthenseparation problems at inner wing, (worse with turbulent ow),leading to pitch up, which can be avoided by using vortex gener-ators, wing fence, vortilons (under-wing fences) or slotted airfoil.There is a tendency for static divergence, provoking an unfavor-able gust behavior which can be alleviated by aero-elastic tailoringand unfavorable root mid size effect for laminar pressure distribu-tion.

More recently, one of the main motivations for using FSW re-sides in the fact that transition on swept wings is strongly affectedby leading edge sweep angle. Turbulence transition at lower lead-ing edge angles can be dominated by TollmienSchlichting (TS)waves, whereas higher sweep angles by cross ow instabilities(CF). Different studies have investigated the regimes and effect of sweep on wing turbulent transition. One the main conclusion isthat transonic FSW, because of its lower leading edge sweep an-gle for the same 50% chord line angle (the typical location of theshock (Fig. 13)) presents less CF [84] , which theoretically can de-lay transition until 25% of chord at a Mach = 0.8 conguration,doubling the laminar extension for its equivalent backward sweptwing (Fig. 14).

In this line, an optimization study of forward swept wings(FSW) was performed as part of the novel conguration workpackage in NACRE activities. The design point was dened for acivil aircraft of 180 pax of payload; 3000 nm of range; and cruiseMach number 0.76 at 35.000 ft (Reynolds = 23.7 millions). After


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Fig. 13. Scheme of the main aerodynamic differences between a BSW and a FSW.

a structural and aerodynamic design of the wing (giving a wingaspect ratio of 9.5 for a leading edge sweep of 17.8), a nal in-verse design phase focused on optimization of the laminar region.For a given pressure coefficient distribution taken as the target,the procedure estimated the C p distribution by solving the Eu-ler or NavierStokes equations on the initial geometry. Taking thedifference between the desired and obtained pressure distribution

C p , a geometry correction was computed by solving a potentialequation formulated in inverse design ( z = f ( C p)); the pro-cess was repeated until convergence. Once the nal geometry wascomputed, a second loop could be performed in case the stabil-ity of this wing was not the optima (Fig. 15). The nal designgave an inboard transition moved down-stream (upper and lowerside) with the upper side dominated by TS and a lower side domi-nated by CF with an obtained drag reduction of about 14% of totaldrag.

3. Drag reduction technologies

Viscous drag reduction, which accounts for some 50% of thetotal aircraft drag, shows one of the largest areas of potentialfor improved aircraft efficiency over the next 1020 years. Twomain lines are currently under development. There are investiga-tion of laminar ow (Section 3.1); and turbulent drag reduction(Section 3.2 ).

3.1. Drag reduction by extended laminar ow

Before adopting any resolutions, it is important to clarify thedifferent instability mechanics which may produce turbulent tran-sition in a boundary layer. These are:

TollmienSchlichting (TS) waves are driven by viscous effecton the surface, they occur in two-dimensional ows and themid-chord region of a swept wing.

Attachment-line contamination is provoked by the bound-ary layer of the fuselage which propagates from the wingfuselage junction along the attachment-line and contaminatesthe boundary layer of the leading edge.

Curvature induced instability appears on shear layers over con-cave surfaces.

Cross ow (CF) instabilities occur in regions of pressure gra-dients on swept surfaces. The imbalance between the pres-sure gradient and stream-wise velocity inside and outside theboundary layer provokes a secondary boundary layer ow,called cross ow, which presents a typical infection point in-stability.

Most of the works addressed the need to control TS and CF in-stabilities. Although some examples of attachment-line contamina-tion are shown, also, hereafter, generally, the CF are very sensitiveto free-stream turbulence and to 3D roughness whilst TS are freeof stream sound and 2D roughness. In addition, a negative pressuregradient is favored in dampening the TS while could destabilizeCF. In any case, in a real wing, the transition is triggered by acombination of all these effects, although, for some specic con-gurations, one of them can be the dominant effect.

Fig. 14. Dominan laminar turbulent transition effect for FSW and BSW. Because of its lower leading edge sweep angle FSW conguration presents less CrossFlow instabilities.


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Fig. 15. Algorithms for inverse design of a laminar wing.

In line with these ow instabilities, we considered differentconcepts of Laminar Flow Technologies. These were:

Natural Laminar Flow (NLF) achieved by a favorable pressuregradient which are valid for transition dominated by TS.

Laminar Flow Control (LFC) achieved by Boundary Layer Suc-tion. In highly swept wings t which, usually, are required foright at high subsonic and supersonic speeds, only suctioncan control sweep-induced crossow disturbances which pro-mote boundary-layer transition. Average suction velocity ratiosof 10 310 4 were proven to reduce amplitude growth frome26 to e5 for a at plate boundary layer [87] .

Hybrid Laminar Flow Control (HLFC) is a combination of lead-ing edge suction and pressure gradient/shaping. Generally, suc-tion is applied near the leading edge of a swept wing in orderto control contamination and cross ow instabilities. Appropri-ate shaping of the pressure distribution stabilizes mid-chordTS. Its applicability was demonstrated by a Boeing 757 HLFCFlight Test in 1990 and on an Airbus A320 n in 1998.

More specic approaches, that discussed here also, are activecontrol of transition by wave superposition [86] and span-wiseperiodic distributed roughness elements (DRE) [86,87] .

3.1.1. Laminar ow technologiesDelaying boundary layer transition is a well-known method for

reducing drag. To this end, signicant work has been done on Lam-inar Flow research with different prototypes ying already. Theseare:

The ATTAS (Advanced Technologies Testing Aircraft System)designed by German Aerospace Center (DLR) [85] ;

The Fokker F100 aircraft ight tests within the Europeanproject ELFIN-II [90] ; and

The Piaggio P180 aircraft or the 757 HLFC ight program [40] .

Recently, the need to reduce operating expenses for new com-mercial aircraft has led to a renewed interest in laminar-ow tech-nology to reduce drag in cruising. In this context, the HARLS LowSweep wing conguration optimized for fuel burn rather than op-erating costs, and the idea was discovered that, using a low sweepwing, might unlock the option of laminar ow. Two European FP6projects explored this concept. NACRE (New Aircraft Concepts Re-search) has performed multi-disciplinary assessment of turbulent

Fig. 16. TELFONA PATHFINDER Model in ETW wind tunnel test section from [99] .

HARLS congurations and studied the application of NLF to For-ward Swept Wing conguration (already discussed in the previous

section). TELFONA (Testing for Laminar Flow on New Aircraft) de-veloped the tools to design and test an NFL wing and to be able topredict the in-ight performance standard of an NLF aircraft.

Different activities, carried out in the TELFONA project, in-cluded:

The calibration of transition tools for the ETW (European Tran-sonic Wind Tunnel);

The investigation of the impact of noise and free stream tur-bulence on transition location in a 2D ow in the TsAGi windtunnel;

A receptivity study of traveling CF vortices to free stream tur-bulence and of stationary CF vortices to the surface roughness;and

The windtunnel test of a performance NLF wing in the ETW.

These activities were structured around the design, manufac-turing, testing and analysis of two wing concepts. These werethe pathnder wing, which serves to calibrate transition predic-tion methods for the ETW, and the performance wing, which hasto demonstrate the capacities of the HARLS NLF conguration byReynolds ight number. For the pathnder wing (Fig. 16), the rel-evant test ow conditions were:

Mach = 0.78, 0.02; Re = 15 to 23 millions; T = 117 K; CL = 0.1 to 0.5; and Side slip = 0 and 4.

The wing has been designed by CIRA, DLR and ONERA usinga 3D inverse optimization algorithm with linear stability analy-sis. The Euler equations of gas-dynamics; the laminar boundary-layer equations for compressible ows on innite swept wings;and the linear Parabolized Stability Equations (PSE) were solvedin order to analyze the evolution of convectively unstable distur-bances. Laminar-turbulent transition was assumed to be delayedby minimizing a measure of the disturbance kinetic energy of achosen disturbance which was computed using the PSE. The shapegradients of the disturbance kinetic energy were computed basedon the solutions of the adjoints of the state equations [5]. The de-sign point was chosen at Mach = 0.734; Reynolds number Re = 6.5millions; and angle of attack = 2.1875 . The method showed im-provements in the viscous drag, which was reduced by six dragcounts. The design obtained a pressure distribution which gave


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Fig. 17. TELFONA PATHFINDER Model in ETW wind tunnel test section from [99] .

Fig. 18. ETW and numerical results for pressure distributions and stability analysisat = 0.33 from [99] .

amplication N -factors in the range 5 < N CF < 8, 6 < N TS < 10with linear chord-wise variation. Parallel isobars were obtained fora region which, at least, extended from 30% to 70% of the span,with transition occurring between 30% and 50% of the chord. In theexperiment, pressure taps were located on diagonal sections whichwere located roughly at normalized semi-span positions = 0.33,0.67 (Fig. 17). The results were compared, also, with the Tempera-ture Sensitivity Paint (cryoTSP [33] ) measurements obtained in theETW, showing good agreement with the computations [75,99] ,which proved this methodology to be valid in transition predic-tion (Fig. 18).

Attachment-line contamination can be controlled by an ade-quate design of an anti-contamination device (ACD). ACD is apassive re-laminarization device which is placed at the wingsleading-edge at a close distance to the fuselage which, rstly, aimsat stopping the span-wise propagation of contaminated ow of the fuselage boundary layer since it reaches the wing and, sec-ondly, at initiating a new healthy, laminar attachment line owon the other side of the device. In SUPERTRAC project (Supersonictransition control [7]), a Reynolds-averaged NavierStokes (RANS)-method was considered for the numerical investigation of differentACDs [56] . The change in turbulence state of the ow and the re-laminarization process were monitored through different param-eters. These were: shape-factor; viscosity ratio; streamlines; andpressure distribution. Since leading edge contamination is a localproblem, a simplied geometry of a circular half-cylinder, followedby a thick at plate, was considered to be the leading-edge. It wasxed to a solid cylindrical representing the fuselage at a sweep an-gle of 65 . The ow computations were carried out for the Mach

number 1.7 and a total temperature of 300 K. They considereddifferent values of attachment-line Reynolds number ( R ) rangingfrom 205 to 443. Experimental evidences showed that leading edgecontamination occurred as soon as a critical value around 250 [78]was exceeded.

They investigated two types of standard ACD shapes, namely,two rectangular shapes with different heights (1 mm and 5 mm)and one triangular shape (5 mm height) (Fig. 19). The main con-

clusion was that the rectangular ACD at a height of 1 mm wasincapable of stopping the contaminated ow, i.e., the high viscouslayer crossed over the ACD. The rectangular shape at a height of 5 mm was the most effective one; whilst, possibly due to its spikyedge, the 5 mm-triangular shape was ineffective.

The nal design was a combination of the rectangular and tri-angular shape (trapezoidal shape) with sufficient height (Fig. 20).The rectangular shape with a height of 5 mm was placed on thefront side to stop the contaminated ow and the sloped triangularside faced towards the back in order to initiate a smoother geo-metrical transition. This design was manufactured and tested andshowed that it stopped the contamination effectively and delayedthe critical R to 380, far beyond the theoretical value. Finally, itwas installed on the laminar wings of small supersonic transport

aircraft supplied by Dassault Aviation. The shape of the ACD wasadjusted to the shape of the wings leading-edge for a cruising con-dition at Mach = 1.6 and CL = 0.11.

Also in SUPERTRAC, CIRA (Centro Italiano Ricerca Aeronau-tica) proved the feasibility of a numerical optimization loop inthe design of a laminar supersonic wing. The optimization wasperformed with an evolutionary optimization library, which used3D Euler solver coupled with a full 3D boundary layer, and aboundary layer stability tool which made use of the ONERA-CIRAdatabase stability computations. 25 twist control sections plus 68shapes were dened as design variables functions for each section(68 25 = 1700 variables). The following design points were con-sidered in the computations:

Mach = 1.6; Reynolds = 51.8 millions; Length = 6.27 m; Wing Area 50.0 m 2 ; CL = 0.182; CM= 0.05; and Maximum trailing edge angle 6 and minimum leading edge

radius 0.15 mm.

The optimization loop returns a damping of pressure peaks andmid-chord shocks and an important enhancement of laminar ow(from 2% to about 1012% of wing surface). This proves the feasi-bility of using optimization tools in the design of laminar wings.

TsAGI [23,24] studied experimentally the receptiveness of theboundary layer to free stream turbulence levels and noise studiedat. A LV6 laminarized airfoil with 1 m chord and 35 of sweep wasmeasured at free-stream velocity of around 80 m/s. The experi-ments ( Fig. 21) showed that, for very low free stream turbulencelevel (0.064%), the transition could be delayed as far as 62% of thechord. On the contrary, acoustic perturbation in the range of 2.02.8 Hz and a noise level of 91108 dB reduced the laminar regionto 55%. More dramatic was the effect of the free stream turbulence,which for a 1% level limit, reduced the laminar region to only 10%of the chord. Although, in this case, the acoustic receptivity of theboundary layer was very low, when 3D effects were considered,the cross ow instability mechanism dominated.

3.1.2. Laminar ow on nacellesRecently, the NLF concept focused on nacelle applications. There

were a number of investigations carried out in the past on the


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Fig. 19. Anti-contamination devices studied in SUPERTRAC. Lower, mesh details. Upper, ow details from [56] .

Fig. 20. Anti-contamination device designed in SUPERTRAC from [56] .

aerodynamic interactions of the nacelle, pylon and wing. These canbe broken down into four areas [83] : interference effects of thenacelle, pylon and wing; the effect of different nacelle position;the effect of high by-pass ratio (BPR); ultra-high BPR nacelles; andthe ability of CFD to predict the interference effects. Today, thelarger By-pass Ratio (BPR) engines are receiving renewed atten-tion. Larger BPR engines have better fuel efficiency but larger na-celle diameters. Typically, current nacelles designs feature surfacegaps and steps congurations which can provoke laminar-turbulenttransition:

1. One aluminum lip with anti-icing system and external panelsin composite material.

2. Pneumatic anti-icing system with exhaust panel on nose cowl

external panel.

Fig. 21. The measured transition locations at various experimental conditions LV6airfoil model (semi-bold symbols transition onset; bold symbols end of transition)from [24] .

3. Access panels on external cowls for maintenance purpose(Temperature sensors, Anti-icing system, Oil tank).

4. Junction between xed nose cowl and moveable fan cowls.

It is important to move downstream the junctions between thenose cowl lip and the external panel; the nose cowl and the fancowl; and all access panels. In this context, the nacelles designer,Aircelle, proposed to design cowl concept which integrated nosecowl and fan cowls, allowing an overall improved performance,versus current nacelle design, of about 1% in SFC due to exten-sion of the laminar ow up to the 25% of the wetted area.

Recently, Bombarbier performed a numerical optimization overan original long-cowl nacelle candidate which showed an improve-ment of 7 drag counts in an isolated conguration at ow condi-tion Mach = 0.8 and Re = 16 millions. The optimization was per-formed with Multi Objective Optimization iSIGHT software, which


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Fig. 22. Transition prediction. Isolate nacelle.

Fig. 23. Suction skin connected by perforated honeycomb to structural sandwichfrom [13] .

used an Euler plus boundary layer computations for the externalow and 3D compressible boundary layer stability for the anal-ysis of transition prediction. The nacelle design methodology wasbased on a target pressure optimization. Subsequently, the new de-sign was tested in a wind tunnel and on a test ight and achievedan improvement of more than the 10% in the extension of the lam-inar region ( Fig. 22).

3.1.3. Hybrid laminar ow

The use of boundary layer removal through suction was a wayof maintaining extensive regions of laminar ow which, since thefties, have been the subject of experimental and theoretical inves-tigation [17,67] . Extensive research was undertaken to mature Hy-brid Laminar Flow by suction; the inuence of swept wing effectsand suction ow rate on its efficiency; and the variation of the rel-ative drag co-efficient [9,15,24] . The technology achieved a certainlevel of maturity. Consequently, it is being considered now withinthe industry for future aircraft applications. The technological chal-lenges are mainly non-aerodynamic dealing with the integrationof the complex system solutions into the wing, nacelle, n and hor-izontal tail plane surfaces; manufacturing surface quality includingthat of suction panels; the need for an anti-contamination system,weight of suction system, etc. In this respect, Boermans [13,14]

designed a suction skin which was connected by perforated honey-comb core to structural sandwich (Fig. 23). The typical perforationrequires a diameter of 0.1 mm and a relatively low hole densitywith a porosity 1%. In order to check the maximum suction ve-locity for optimal performance, the solution was tested in a gliderconguration.

Additionally, the intensive use of the DNS (Direct NumericalSimulation) allowed new perspectives of the suction concept to beproposed:

Suction can be combined advantageously with excitation of useful vortices (use of holes/slots to excite and support con-tinuously useful vortices to combine the effect of suction andsuppression of secondary instability), securing and improvingsuction performance;

Fig. 24. Suction test denition in SUPERTRAC.

Fig.25. Stream-wise transition as a function of the suction velocity for different testcases from SUPERTRAC project in [45] .

Direct stabilization of secondary instability of crossow vor-tices by pin-point suction reveals itself successfully with rela-tively low suction rate (optimal for xed vortices).

Very few studies of supersonic velocities are available. In SU-PERTRAC the laminar control by suction of CF-dominated laminar-turbulent transition was studied at Mach 2. After a preliminarynumerical analysis, a windtunnel model was built based on asymmetric arc-shaped airfoil with relative thickness of t / c = 0.13;a sharp leading edge; and a chord length of c = 300 mm. The suc-tion panel was located between 5% and 20% of the chord. The holediameter was 17 m. The model was mounted in the windtunneltest section at zero angle of attack and a sweep angle of 20 and

30 (Fig. 24). The suction pressure ranged between 1.2 and 4.8 barfor Reynolds numbers between 6 and 24 millions. The experimentswere conducted in the Ludwieg Tube Facility (RWG) at DLR. Asshown in Fig. 25, a signicant delay of laminar turbulent-transitioncan be obtained (typically from 20% to 60% of the chord). It wasobserved, also, that, beyond a certain level of suction velocity, nofurther improvement was obtained.

The HISAC project investigated the application of laminar owtechnology for a business supersonic jet. The aircraft was a sweptwing monoplane. The wing and horizontal tail were tailored in or-der to keep the ow laminar over a large portion of the wing/tailarea. The wing design beneted from an imposed negative pres-sure gradient over the largest possible length of local chords. Thelaminar wings were tested in the ONERA-S2MA wind tunnel atMach = 1.6 and Re = 7 millions. By using suction and cooling,


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Fig. 26. Applications of NLF & HLFC concept on a supersonic test (HISAC project).

the laminar studies focused on high-sweep wings. The main re-sults showed that Natural Laminar Flow should be achievable onthe outboard wing, which could be enhanced by using laminar

ow cooling, whilst suction and the use of anti-contamination de-vices were needed to maintain laminar ow on the inboard wing(Fig. 26). The orders of magnitude of the suction rate were be-tween 0.0005 and 0.001 at the upper side and 0.001 at the lowerside, giving a 50% upper surface transition location and an esti-mated viscous drag reduction of 28.5% (7.2 dc) [70] . The maindifficulties, in maturing the concept, were the design of high-liftdevices and the development of de-icing system compatible withthe laminar ow concept.

The City University London [8] employed Parabolized StabilityEquations (PSE) to consider the effect of Mach and Reynolds num-bers on the stability of boundary layers ( Fig. 27). They showedthat the critical N -factors decreased faster as the Mach numberand Reynolds number increased. This corrected the previous sta-bility analysis which did not consider the compressibility. To ob-tain a comprehensive idea, a typical decrease of N -factor from5.8 to 4 suggested an increase in the suction mass ow rate of 9% and an increment in the pump pressure ratio from 2.0 to 2.5which decreased by 9% the expected benets in drag reduction of HLFC (after deducting system weight and power penalties). It was

shown, also, that, although critical N -factors for linear PSE anal-ysis differed little from classical theory for standard experimentaldatasets, when the effect of the leading edge curvature was takeninto account, a different picture emerged. In this case, the leadingedge modes were inuenced positively by the curvature and, com-pared to the HLFC for the same design, resulted in a reduction of 24 suction mass ow rate and an increased benet of 9% withoutcurvature effects.

As a nal remark concerning the Hybrid Laminar Flow Con-trol (HLFC) technology demonstration, a test ight was planned inthe framework of the JTI-SFWA project. The new outer wing ele-ments of the A340 wing incorporating HLFC could be ight testedin 2014.

3.1.4. Alternative laminar ow technologyThe current advances in micro and nano machining and electro-

mechanical fabrication have facilitated the emergence of alter-native methods in promoting laminar ow. Identied researchlines include the polishing of the surface; application of dis-tributed roughness to delay CF; active control of TS by mass-less jets/surface actuation; and laminar ow control by heat transfer orplasma. Such technologies may offer a more lightweight solution

than the conventional hybrid laminar ow system. However, moreresearch is required to understand the viability of such approachesboth in terms of the fundamental ow physics of how transitioncan be delayed and the resultant control system requirements.

Polishing the leading edge is one way to extend the laminarow region. Test ights conducted by the Texas A&M Flight Re-search Lab in a Cessna O-2A Skymaster showed 80% laminar owat Re = 8 millions, = 30 , obtaining N factor > 16. The advan-tage was that it was necessary to polish only 10% of the chord tobe efficient. However, typical of ight conditions, adequate eval-uation requires very low free-stream turbulence, which makes itimpossible to calibrate in wind tunnel testing. Other disadvantagesare associated to the manufacturing process like the necessity touse harder surfaces to reduce roughness or the use of leading edgeKrueger aps to protect during landing and takeoff.

The Micron Sized Roughness (MSR) concept (or periodic Dis-crete Roughness Elements, DRE) can control the stationary crossow vortices. The idea is to create an articial surface roughnesswhich introduces weakly growing wavelength, s , which generatesa modied mean ow which is stable to all wave-lengths greaterthan s . This concept was checked by Saric [87] who demon-

Fig. 27. Critical N -factor for TS wave as a function of the Mach and Reynolds numbers from [8] .


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Fig. 28. Swept wing in ight tests (SWIFT).

Fig. 29. Flight measurements carried by W. Saric with painted LE et DRE concept.

strated that the DRE could increase laminar from 30% to 60% of the chord at Re = 8.1 millions; sweep angle of = 30 and ve-locity of 92 m/s (Mach 0.28). The experiments were performedon the Swept Wing In Flight Tests (SWIFT) (Fig. 28) mounted ina Cessna O-2A Skymaster at the Texas A&M Flight Research Labo-ratory. A detailed computational study was performed in advance.The design obtained a pressure minimum between x/ c = 0.7 to

0.8, making the boundary layer subcritical to TS instability whilstdestabilizing crossow waves. The nose radius was restricted toRe < 100, making an attachment-line subcritical to instabilitiesand contamination. Finally, the Cp distribution was optimized bycrossow control. The computation was performed on the Eulerand NavierStokes equations for Cp and boundary layer calcula-tions; OrrSommerfeld for stability; and Parabolized NavierStokesfor nal assessment [44] .

Stability calculations veried that the 4.5 mm wavelength wasextremely un-stable and that s = 2.25 mm was the candidate tocontrol crossow. Two layers of DRE were placed at 1% x/ c on theinboard pressure row, and 1.3% x/ c on the outboard pressure row.The DREs were 2.25 mm spacing, 1 mm of diameter and 30 mi-crons high. For this conguration the transition of the chord movesfrom 30% to 60% (Fig. 29).

SUPERTRAC and HISAC projects and, within the UK nationalAERAST project for transonic congurations, studied, also, micron-sized roughness elements for supersonic congurations. In SUPER-TRAC [6], in a swept wing with a strong ow acceleration, thetransition was triggered by CF instabilities. Controlling this typeof instability required the knowledge of the target mode of themost amplied natural vortices. Then, theoretically, a killer modegenerated articially by MSR and arranged along the leading edgecould interact with the unsteady mode resulting in a more stableevolution. The success of this approach depended on the pressuregradient and on the choice of the killer wavelength [88] . In SU-PERTRAC, a row of MSR were implemented on the wing leadingedge, using rows of roughness elements of 10 m height and 0.2or 0.15 mm in diameter; Reynolds from 3 to 7 millions; and sweepangle from 15 to 30 (Fig. 30). Unfortunately, no clear positive ef-

Fig. 30. MSR concept tested in the S2MA wind tunnel (SUPERTRAC project).

Fig.31. Visualization of boundary layer transition for the NLF(2)-0415 airfoil at Re =2.2 millions. Left uncontrolled, right controlled.

fect was observed, according to receptivity computations, the usedMSR lay at the limit of efficiency in terms of height and diame-ter. Probably, new manufacturing processes and/or tests at a largerscale are needed.

More recently, Sui-han et al. [100] , in collaboration with Air-bus, studied the NLF(2)-0415 airfoil with a 45-degree sweep wing.The experiments were performed at Reynolds 2.2 millions and50 m / s in the Northwestern Polytechnical Universitys Low Tur-bulence Wind Tunnel (0.05%). One line of DREs were disposed at1.7 mm spacing roughness located at x/ c = 3.5%, with a diameter

of 0.7 mm and a mean height of 19.25 m. The separation was se-lected based on the most unstable wavelength of 3 3.3 mm. Asobserved in Fig. 31, the region downstream of the roughness, thetransition was suppressed successfully until the trailing-edge.

Drag reduction, by means of active control of TS, could allowthe laminar ow to extend in difficult conditions (adverse pressuregradient, high Reynolds number). The control of boundary layertransition by wave-superposition is not new [102] , but putting thisidea into practice is not an easy task. The principle of an ActiveWave Control System (Fig. 32) considers two steps. These are:

The TS waves measurement with a reference-sensor; The counter-wave actuation and the monitoring of results with

an error sensor which allows the control to be adapted.

The successful application of the sensoractuator system in therange of Ma = 0.20 .5 has been carried out by TU Berlin [3032] .The TS wave amplitude has been reduced by 90% (at the error sen-sor) in the wind tunnel and around 50% in ight (with a glider),yielding to substantial delay of laminar-turbulent transition. Theexperiments were performed in a two-dimensional NACA0004 pro-le of 750 mm of chord length. The wind tunnel was providedwith an adaptive test section so that the pressure prole could bechanged to force the transition in the planar region of the NACAprole where the experimental facilities were disposed. The freeight test was done in an unswept Grob G103 Twin II Glider at22.5 m / s with the actuator located at 47% of the chords length.The difficulties for using this technique at higher Mach numberswere linked to the development of oblique TS. In order to developthese technologies, it is important to acquire more information


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Fig. 32. Principle of an active wave control system (TU Berlin).

Fig. 33. Dielectric barrier discharge actuator.

about the properties of TollmienSchlichting waves. Another im-portant issue is the design of a robust control system capable of managing more realistic congurations.

Other alternatives for laminar ow control come from heattransfer and plasma actuation. In the past, controlling the tem-perature of the wing skin was found to be relatively inefficient.However. with micro fabricated lms and sensors, it may be pos-sible to control the heat transfer using relatively small amountsof energy which could make this solution more attractive. TheHISAC project considered the new approaches and could be ap-

plied to nacelles in the future.The application of a near wall direct current to create a corona

discharge can delay the laminar-turbulent transition by generationa body force [52] which accelerates the boundary layer ow in thestream-wise direction. Theoretical estimations [49] revealed somebenecial effects of this method in drag reduction if the boundarylayer included both laminar and turbulent parts [53,54,59] . A rel-atively simple device employs two thin electrodes, separated by adielectric barrier (Fig. 33). Alternating voltage is applied betweentwo electrodes ionizing the air and creating plasma. The movementof the ions transfers global momentum to the neutral air which isperceived macroscopically as a body force on the ow. However,the design of the experiments is not trivial, and it is important toknow the parameters of the plasma actuators (dimensions; currentstrength; applied voltage; etc.), which can inuence the boundary

Fig. 34. Typical riblet geometry from [36] .

layer stability. The experiments performed by [59] in a at plate

at u = 30 m / s, T = 290 K, and the Reynolds number, at the anodeposition of Re 1 millions concluded that the plasma ow pro-duced a signicant increase in the value of the critical Reynoldsnumber and constrained the range of unsteady TS wavenumbers.In any case, further experimental and theoretical validations mustbe performed to evaluate the real performance of these methods.

3.2. Turbulent skin friction reduction

Even if laminar ow control is successful, a signicant propor-tion of skin friction drag due to turbulence remains. This sectiondiscusses the different technologies: riblets and dimples or surfaceactuation which aim to reduce the turbulent drag.

Riblets are small surface protrusions, aligned with the direction

of ow, which confer an anisotropic roughness to a surface. Theyare one of the few techniques which have been applied success-fully to the reduction of the skin friction in turbulent boundarylayers, both in the laboratory and in full aerodynamic congura-tions. From an aerodynamic perspective riblets were identied asa mature technology which could provide modest reductions (78% skin friction drag reduction for riblet spacing of approximately15 wall units) in aircraft drag. Flight tests conrmed total drag re-duction of 1.6% (5% of Cd on 66% of the wetted surface) for anA320 model (proved in the S1MA Onera Wind Tunnel). The cur-rent development of riblets is in line with a better understandingof the physics and the application of surface technologies like theaerodynamic evaluation of riblet material for turbulent skin fric-tion reduction. A current analysis of the status of this technologycan be found in [18,25,104] . The physical mechanism of the ribletdrag reduction effect is caused by a protrusion height between thevirtual origin, seen by the stream-wise shear ow and some meansurface location. This offset would result in a greater separationbetween the wall and the turbulent stream-wise vortices, reducingthe exchange of momentum at the wall [10,46] . The correct ex-planation of the underlying physics is still a topic of research. Inthis line Garcia-Mayoral and Jimenez [36] found that the groovecross section A+ g , expressed in wall units, was a better characteri-zation of this breakdown than the riblet spacing, with an optimum( A+ g )

1/ 2 11 (Fig. 34). However, the drag reduction was affectedgreatly by the riblet spacing and size or orientation which, in somecases, could produce drag increase. Other non-aerodynamic issueswere maintenance of riblet shape and adhesive over operationallife (hydraulic uid, dirt, deformation by hail and maintenance); vi-sual appearance, and time required to install, remove and re-applyriblets.

Dimples are regular arrangements of surface depressions dis-tributed along the wall. Dimples are a well-known measure toincrease the heat-transfer from a wall [21] ; however, they can beuseful for drag reduction. Compared to riblets, they can be advan-tageous since they are composed of macroscopic structures whichare less sensitive to dirt and mechanical degradation. Despite itsexpected interest, very few works could be found in the literature,e.g. only two articles about dimples were presented at the last49th Aerospace Sciences Meeting of 2011, and some preliminarystudies [65] were somewhat discouraging by showing very com-plex structures inside the dimple and little or no improvements inturbulent drag reduction. These made them unattractive due to thecosts associated with their operation.


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Since the potential drag reduction could be large, the possibilityof using nano or micro scale electro-mechanical systems to controlthe development of turbulent structures in the boundary layer wasinvestigated also. One interesting drag reduction technique, whichis conceptually simple and has the potential for a positive ener-getic budget, is the cyclic span-wise movement of the wall [29,81,82] . In this active technique, the wall moves sinusoidally creat-ing a traveling wave which alters substantially the near-wall ow

structure. The main features of the oscillating wall technique arethe existence of an optimal frequency to maximize signicantlydrag reduction, which can be as high as 45% when the amplitudeof the oscillation is comparable with the ow center line veloc-ity. Following this idea, Tardu and Doche [101] showed how thereaction of the near wall turbulence and the drag were sensitiveto the temporal waveform of the localized time periodical blow-ing. The injection velocity is periodical and asymmetric in time,with a rapid acceleration phase followed by a slow decelerationone. Mainly in the deceleration phase, the ow is re-laminarizedduring 70% of the oscillation period. Induced by the blowing, thelatter maintains the stability pf the vorticity layer and prevents itsrollup contrary to a sinusoidal time periodical blowing. Therefore,a time mean drag reduction of 50% is obtained in the region recov-

ering 200 wall units downstream of the blowing slot. This is 40%greater than the drag reduction obtained by a steady blowing withthe same time mean severity parameter.

Unfortunately, the effectiveness of these method decreases asthe Reynolds number increases [22] . To overcome this constraint,Pamies et al. [74] proposed the use of a simple modication of Op-position Control (OC) to increase the performance at high Reynoldsnumbers and, thereby, showing a signicant improvement on dragreduction when performing a blowing-only opposition control. Theanalysis was performed by a LES numerical simulation in a spa-tially developing at plate (fully turbulent ow) at Mach = 0.1,Re = 1100 and Re = 3300. This showed a local drag reductionof up to 61% in the area of control and, in the analysis, approxi-mately ve times the boundary layer thickness.

The physical realization of this surface waves can be done byusing Electro-Active Polymer (EAP) [28] . EAP, also referred to asMaxwell Stress, is based on the compression, by electrostatic pres-sure, of an elastomer lm, which is applied by compliant elec-trodes upon the application of an electric eld. By virtue of in-compressibility, the compression of the elastomer results in anelongation of the membrane which is used for actuation. Thisallows continuous surface deformation, both in-plane and out-of-plane. With this material, it is possible to create an array of activedimples which act as time dependent depressions and which al-low changes in frequency and amplitude. These inject local owdisturbances capable of interacting with structures existing in theturbulent boundary layer. A similar effect can be obtained by usingenergy deposition (plasmas).

4. Separation control technologies

Separation is a continual source of nuisance for the aircraft.At low speed, maximum lift, it is related directly to the exis-tence of ow separation. For highly deployed high-lift devices, themaximum lift is obtained immediately before an extended owseparation region appears. Additionally, highly detached and un-steady ows in gaps, slots and cavities in landing gear are the mainsources of aero-acoustic noise. In a high speed regime, the causesof concern come from reducing the inuence of the shock withthe upper surface boundary layer and delaying the Mach numberat which this interaction leads to ow separation and buffet.

The control of high lift induced separation on an airfoil mayimprove the ight envelope of current aircraft or even simplifythe complex and heavy high-lift devices on commercial airframes.

Now, traditional high-lift systems have evolved greatly with onlysmall improvements in current aircraft performance. An innovativestep is necessary to obtain the desired improvement in L/D andCLmax. This level of improvement allows an increase in payloadand reduced landing speed and, therefore, reduces noise emissionsand increases aircraft safety.

Separated ows can appear in at least two different conditions.These are progressive boundary layer separation (PBLS) such as

trailing edge separation, which displaces progressively upstreamwhen the angle of attack is increased, and very localized bound-ary layer separation (LBLS) such as those generated by geometricaldiscontinuities. These two ow separation types can be found on awing in high lift conguration. For instance, the PBLS type can beobserved either on a slotted ap leeward side or on the leewardside of the main body trailing edge near the pre-stall angles of at-tack. LBLS type ow separation can be observed near the leewardleading edge of the main body at stall and post-stall angles of at-tack, and, also, near the hinge line of a strongly deected aileron.

The use of ow control devices to control ow separation hasvery attractive properties; at high speed, they can delay the onsetof buffet, and enable more aggressive low-drag designs or lighterstructures with the same aero-performance. At low speed, they can

enable simpler congurations and mechanisms, e.g. xed leadingedge; short chord slat; simple hinged ap; and improve take-off lift to drag ratio or recover performance for alternative platforms e.g. increased sweep or taper. In term of ow control actuators,a PBLS type separated ow can be controlled with sufficiently up-stream mechanical or pneumatic discrete vortex generators (VG).A LBLS type separated ow can be controlled by a pulsed synthetic jet slot (located very near to the separation onset) (Kroo [58] ;Wygnanski [108] ; McCormick [71] ; and Courty et al. [26] ), by apulsed blowing jet slot or pulsed pneumatic VGs (Kilbens andBower [50] ).

Numerous research activities on Separated Flow Control wereundertaken in the USA and Europe, with many wind tunneldemonstrations. These works allows the assessment of the effec-tiveness of various actuators and various kinds of actuation, aswell as the optimal parameters of such devices (geometry; size;orientation; ow rate; and frequency). A particular problem withthis approach is that the use of wind tunnel models requires theuse of smaller actuators than needed for in ight demonstrations.Moreover, in pneumatic pulsed actuators, for instance, actuationfrequencies are usually an order of magnitude which are higher inwind tunnel demonstrations than during ight demonstrations.

Flow control technologies can be divided into passive, activeand reactive. Passive systems do not require a power input for theiroperation but have, in general, a drag penalty during cruise oper-ation. Examples are vortex generator (VG) or sub-boundary layervortex generator (SBVG). An active system requires some power in-put to work. Between them, we can nd uidic vortex generators,pneumatic synthetic jets, electrical or plasma actuator or mag-netic. Finally, a reactive system possesses some kind of intelligenceand actuates according to the information supplied by the sensor.As described in previous sections, these kinds of systems are usu-ally more suitable for boundary layer stabilization processes.

4.1. Passive ow control devices

Passive VG are simply small aspect ratio airfoils mounted nor-mally to the lifting surfaces ahead of the ow separation point inorder to energize the boundary layer and to prevent separation.The only difference with SBVG, also known as low-prole vor-tex generators, is that SBVG are submerged below the boundarylayer to reduce the drag penalty. VG and SBVG can be classied asPBLS ow control actuators. Typical applications of these devicesare the control of low-speed separated ows in adverse pressure


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Fig. 35. Numerical surface streamlines of a backward-facing wedge VG at Mach 2.5from [61] .

gradient and supersonic shock-induced separation. Although it hasbeen proven that the VG reduces the separation zone, there is noexisting satisfactory explanation of how they act. Lu et al. [68] de-scribed a detailed review of the underlying physics and ow topo-logical structures, presented around these devices. This was mainlyfor SBVG at high speed, where it was shown that, although conven-tional, vane-shaped vortex generators were well established (theyproduce a par tip vortices which entrains free-stream uid, and,thereby, energizing the boundary layer), other congurations, asforward and backward-facing wedge or wishbone and doublet typeWheeler vane (better suited for high-speed ows in view of its ro-bustness and produced typically a dominant horseshoe vortices)showed an extremely complex ow topology ( Fig. 35). This behav-ior can be explain because of the presence of a tiny, sub-boundarylayer protuberance which arises from the ow separating off theslant sides, showing that understanding the unsteadiness of theow caused by these devices and its effects on the external owwas not as easy as expected. A second debatable approach is thepresence of an instability mechanism triggered by the VG whichis thought to produce a train of either hairpin vortices or vortexrings which entrain high momentum of the free-stream uid.

VG are found to be best suited for applications where, rela-tively, the ow-separation locations are xed and the generatorscan be placed close upstream of the separation. The effect of dif-ferent parameters (device height; length; spacing; and stream-wisedistance) involved in its design was discussed in [66] . The resultsshowed the effectiveness of the low-prole VG h/ 1, ; theboundary layer thickness; and h the height of the device in elim-inating the inection point in the pressure distribution, thus allow-ing a signicant reduction of separated area. Taller VGs ( / 1)reduce, also, the detached area but at the cost of causing a strongmodication of the input ow, creating 3D ow structures (Fig. 36)with a subsequent drag penalty. In any case, it is important toemphasize the strong inuence and sensitivity of the different pa-

rameters involved in their congurations in order to obtain theoptimum results.

More recent applications of these devices are found within theAWIATOR project, where SBVGs were mounted on the ap up-per surface producing the delay of ap boundary layer separationonly when the ap was deployed ( Fig. 37). When the ap wasstowed, the SBVG were contained in the cove region under thewing shroud. The design was tested on an A340. Computational

analysis; wind tunnel testing; and ight test validation were car-ried out to obtain an optimized design. The addition of SBVGsto the ap at = 35 deection increased the lift coefficient byup to 2.2% over a wide incidence range compared to the base-line conguration of = 32 without SBVGs. The effect of SBVGsat a xed landing ap deection increased the lift coefficient CL from 0.01 to 0.04 over part of the incidence range. As expected,this indicated that the vast majority of the lift increase was dueto an increase in ap angle. A trend to a small increase in dragof 4 counts (Cd = 0.0004) in takeoff conguration was within therepeatability of the wind tunnel balance and, therefore, deemedinsignicant. Although the stall was produced 0.3 early, the ighttest conrmed this improvement in the behavior of the ap at 35 ,with an increment of CL of approximately 2.5% at the reference an-

gle of attack.QinetiQ studied experimentally, in the DERA (Defense Evalua-

tion and Research Agency) high speed wind tunnel, different SBVGarrangements to prevent the onset of buffet. The SBVG were dis-posed at 55% of the chord, ahead of the shock. The results showedthat split vane SBVG performed better than wedges in delayingseparation in these congurations. However, the ideal location of the VG was not obvious. Additionally, the effect of the wing sweep,which was not considered in basic studies, could be the cause of discrepancies between the expected rate of boundary layer growth(much higher) and the optimum VG height. Another important ap-plication was found in improving attached ows in pylon-wing in-terference or internal ows S-duct engine air-intake congurations(Fig. 38). Studied by Onera, QinetiQ or numerically by KTH [103] ,

vane and air-jet VG arrays were effective when located ahead of the separation.A more energetic approach is the passive air jet vortex (AJVG)

generator (Fig. 39). Studied by City University [80] , they can beused at low speed to prevent detaching or, at high speed, to con-trol buffet. The idea of the slots is to inject fresh momentum tothe boundary layer particles which have been slowed down by theaction of viscosity. The analysis at Mach = 0.1, Re = 1 millions and

Fig. 36. 0 .2 (0.8 right)-high vane-type counter-rotating VGs at 10 h (6 h-right) upstream of baseline separation from [66] .

Fig. 37. Flap separation controlled by VG. Stow inside the ap well.


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Fig. 38. Flow control application to S-duct engine air-intake.

Fig. 39. Passive air jet vortex system.

an angle of attack of 18 was performed on a NACA23012C aero-foil a modication of the NACA 23012. The model was equippedwith a span-wise array of 15 air jets of 4.8 mm di

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