Drag Reduction by Laminar Flow Control

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Article Drag Reduction by Laminar Flow Control

Nils Beck 1,*, Tim Landa 2, Arne Seitz 3, Loek Boermans 4, Yaolong Liu 1 ID and Rolf Radespiel 2

1 Aeronautics Research Center Niedersachsen (NFL), TU Braunschweig, Hermann-Blenk-Straße 27, 38108 Braunschweig, Germany; [email protected] 2 Institute of Fluid Mechanics (ISM), TU Braunschweig, Hermann-Blenk-Straße 37, 38108 Braunschweig, Germany; [email protected] (T.L.); [email protected] (R.R.) 3 Institute of Aerodynamics and Flow Technology (DLR-AS), German Aerospace Center, Lilienthalplatz 7, 38108 Braunschweig, Germany; [email protected] 4 Faculty of Aerospace Engineering, Delft University of Technology, Kluyverweg 1, 2629 HS Delft, The Netherlands; [email protected] * Correspondence: [email protected]; Tel.: +49-531-391-66662

Received: 20 December 2017; Accepted: 15 January 2018; Published: 20 January 2018

Abstract: The Energy System Transition in Aviation research project of the Aeronautics Research Center Niedersachsen (NFL) searches for potentially game-changing technologies to reduce the carbon footprint of aviation by promoting and enabling new propulsion and drag reduction technologies. The greatest potential for aerodynamic drag reduction is seen in laminar flow control by boundary layer suction. While most of the research so far has been on partial laminarization by application of Natural Laminar Flow (NLF) and Hybrid Laminar Flow Control (HLFC) to wings, complete laminarization of wings, tails and fuselages promises much higher gains. The potential drag reduction and suction requirements, including the necessary compressor power, are calculated on component level using a flow solver with viscid/inviscid coupling and a 3D Reynolds-Averaged Navier-Stokes (RANS) solver. The effect on total aircraft drag is estimated for a state-of-the-art mid-range aircraft configuration using preliminary aircraft design methods, showing that total cruise drag can be halved compared to today’s turbulent aircraft.

Keywords: drag reduction; laminar ﬂow control; boundary layer suction; transition; aircraft design

1. Introduction and Aim of the Work Since the beginning of aviation, reduction of drag is one of the prime objectives for every aircraft designer. An aircraft with lower drag is not only more economical in every aspect, but also less harmful to the environment, which has become increasingly important in the last decades and will be even more important in the future. In addition to noise, NOX and other pollutants, the primary focus is on the emission of greenhouse gases into the atmosphere, which is directly linked to the combustion of carbon-based fossil fuels. While jet engine technology has provided much of the efficiency improvement in the past, physical and technical limits are reached now which mean that future improvements will be smaller and come at higher costs in terms of weight, size and investment. Understanding in aerodynamics has improved only in small steps since the beginning of the jet age, in part because the subsonic turbulent aircraft was aerodynamically much more mature than the then-new jet engines [1]. Swept wings and supercritical wing profiles have expanded the speed envelope into the transonic region, which again improved engine fidelity and of course travel times, rather than directly reducing drag. However, there is room for substantial drag improvement by laminar flow control. The boundary layer flow on today’s large aircraft is turbulent on almost the entire wetted surface. This results in viscous drag five to ten times larger than that of laminar boundary layers.

Energies 2018, 11, 252; doi:10.3390/en11010252 www.mdpi.com/journal/energies Energies 2018, 11, 252 2 of 28

The first part of this paper, in Section2, provides a review of the research on laminar flow control and transition prediction, with special focus on activities at DLR (Deutsches Zentrum für Luft-und Raumfahrt, German Aerospace Research Center) Braunschweig for partial laminarization of wings. In the second part, beginning in Section3, simple methods are used to assess the potential drag reduction by extending the application of laminar flow control by boundary layer suction to all wetted surfaces of the aircraft. Laminar Flow Control (LFC) is applied to existing airfoils and a generic fuselage geometry, with no shape adaptation taken into account. The combined optimization of shape and suction is outside the scope of this paper, but promises even greater drag reductions than those presented here. The authors’ aim is to make the case for an in-depth investigation of the subject in future research programs, using the sophisticated methods described in the first part of the paper.

2. Review of Recent Research on Laminar Flow Technology in Europe The relevant transition mechanisms that can be found on transonic swept wings of modern transport aircraft are Tollmien-Schlichting Instability (TSI), Attachment Line Transition (ALT), and Crossﬂow Instability (CFI). Basic research conducted throughout the last century (e.g., [2–10]) has led to a good knowledge about the physics of these phenomena and provided principle ideas how to control them. By continuous research work, the German Aerospace Center (DLR) has built up the capabilities for transition prediction as well as for design and testing of wings and empennages following the NLF (Natural Laminar Flow) and HLFC (Hybrid Laminar Flow Control) concepts.

2.1. Transition Prediction A prerequisite for the design of a laminar flow wing is a reliable transition prediction method. At DLR and Airbus, the semi-empirical eN method, established by van Ingen [11,12], is used, which is based on linear stability theory. Velocity profiles of the laminar boundary layer are analyzed with respect to their stability against harmonic oscillations, which are superimposed as small disturbances to an otherwise steady basic flow. If unstable, the downstream amplitude growth of a disturbance can be expressed by the so called N-factor, defined as the natural logarithm of the ratio of disturbance amplitude at a point downstream to its initial value at the so-called neutral point. It is assumed that transition occurs where the N-factor of the most amplified disturbance reaches a limiting value Ncrit. Boundary layer velocity profiles on a swept wing are three-dimensionally warped in regions where a pressure gradient is present. When projected into the direction of the external flow, the velocity profiles are similar to those of two-dimensional boundary layers, while in direction perpendicular to the outer flow, the so-called crossflow profile is present, as illustrated in Figure1. Analogous to 2D flow cases, the profiles parallel to the external flow can become unstable against small travelling disturbances, i.e., Tollmien-Schlichting waves, while the crossflow profiles exhibit at least one inflectional point, making them inherently unstable against disturbances with a wave vector approximately pointing in crossflow direction. Consequently, in the approach followed at DLR [13,14] for transition prediction, chordwise N-factor distributions for two classes of disturbances are calculated:

1. Tollmien Schlichting Instabilities are treated as travelling waves with constant frequency and propagation direction parallel to the ﬂow at the boundary layer edge. 2. Crossﬂow Instabilities are treated as stationary (zero frequency) waves with constant wave length. EnergiesEnergies 2018,, 1111,, 252252 33 ofof 2827 Energies 2018, 11, 252 3 of 27

Figure 1. 3D Boundary Layer on a Swept Wing and related Transition Mechanisms. Figure 1. 3D3D Boundary Layer on a Swept Wing and related Transition Mechanisms.Mechanisms. In order to obtain values of Ncrit for TSI and CFI, in-flight experiments were performed in 1987 In order to obtain values of Ncrit for TSI and CFI, in-flight experiments were performed in 1987 [15] Inwith order the DLR to obtain flying values testbed of ATTAS Ncrit for (Advanced TSI and Technologies CFI, in-flight Testing experiments Aircraft were System) performed shown inin 1987[15]Figure with [15 2a.] withthe A wingDLR the DLR flyingglove flying withtestbed testbedmodified ATTAS ATTAS sections (Advanced (Advanced was Technologiesdesigned Technologies especially Testing Testing for Aircraft the Aircraft purpose System) System) of showna clear shown N-in inFigurefactor Figure 2a.identification.2 Aa. wing A wing glove gloveLinear with with modifiedstability modified sectionsanalysis sections wa delivereds designed was designed amplification especially especially for rates the for purposefor the both purpose of classesa clear of N- aof clearfactordisturbances, N-factor identification. identification.and the Linear envelopes Linearstability of stabilityN TSanalysis and analysisN CFdelivered distributions delivered amplification were amplification correlated rates ratesfor with both for the both classesmeasured classes of disturbances, and the envelopes of NTS and NCF distributions were correlated with the measured oftransition disturbances, locations and from the envelopes the same ofexperiment, NTS and N showCF distributionsn in Figure were2b. During correlated the withATTAS the flight measured tests, transitionthe criterion locations for Attachment from the same Line experiment, Transition (ALTshow shown),n infirst Figure proposed 22b.b. DuringDuring by Pfenninger the ATTASATTAS [5], flightflight was tests,tests, also theconfirmed. criterioncriterion Thefor for Attachmentvalidity Attachment of the Line Linecriteria Transition Transition found (ALT), has (ALT firstbeen), proposed successfullyfirst proposed by Pfenninger proven by Pfenninger during [5], was design also [5], confirmed. andwas flightalso Theconfirmed.testing validity of a The ofNLF thevalidity glove criteria ofon found thea Fokker criteria has been100 found in successfully 1991 has beenwithin successfully proven the frame during provenof the design projectduring and ELFIN flightdesign testing (Europeanand flight of a NLFtestingLaminar glove of Flowa on NLF aInvestigation) Fokker glove 100on a in Fokker[16], 1991 funded within 100 in theby 1991 the frame withinEuropean of the the project Union.frame ELFIN ofIt wasthe project(Europeanassumed ELFIN that Laminar the(European ATTAS Flow Investigation)Laminarcriterion, Flow although [ 16Investigation)], funded evaluated by the[16], from European funded a NLF by Union. experiment, the European It was assumedis Union.also valid that It was thefor ATTAS assumedHLFC criterion,applications that the although ATTAS with evaluatedcriterion,boundary although fromlayer asuction. NLF evaluated experiment, However, from is thea also NLF evaluation valid experiment, for of HLFC the isA320 applications also hybrid valid laminar withfor HLFC boundary fin applicationsflight layer experiment suction. with However,boundaryconducted thelayer in 1998 evaluation suction. [17], where However, of the boundary A320 the hybridevaluation layer laminarsuctio ofn the was fin A320 flight applied hybrid experiment over laminar the front conducted fin 20% flight of inchordexperiment 1998 on [17 an], whereconductedA320 vertical boundary in 1998tail layer plane, [17], suction where showed wasboundary that applied critical layer over N-factors suctio the frontn wasfor 20%TSI applied and of chord CFI over were on the an frontreduced A320 20% vertical compared of chord tail plane, onto thean showedA320NLF case,vertical that as criticalcan tail beplane, seen N-factors showed from forthe that TSIN-factor critical and CFI envelopes N-factors were reduced in for Figure TSI compared and 3. This CFI iswere to caused the reduced NLF by case, the compared inhomogeneity as can be to seen the fromNLFintroduced case, the N-factor as into can thebe envelopes seen flow from by insuction the Figure N-factor through3. This envelopes isdiscrete caused inmicro by Figure the holes inhomogeneity 3. This rather is causedthan introduced through by the inhomogeneityan into ideal the porous flow byintroducedsurface. suction through into the discreteflow by microsuction holes through rather discrete than through micro holes an ideal rather porous than through surface. an ideal porous surface.

(a) (b) (a) (b) Figure 2. (a) ATTAS Flight Experiment and (b) Evalation of NTS-NCF Transition Criterion. Figure 2. (a) ATTAS Flight Experiment and (b) Evalation of NTS-NCF Transition Criterion. Figure 2. (a) ATTAS Flight Experiment and (b) Evalation of NTS-NCF Transition Criterion.

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FigureFigure 3. 3. N-FactorN-Factor Correlations. Correlations.

2.2.2.2. NLF NLF and and HLFC HLFC Wing Wing Design Design OnceOnce premature premature transition transition due due to to ALT ALT or or CFI CFI has has been been avoided avoided over over the the first ﬁrst 5–10% 5–10% of of chord, chord, thethe transition transition process process is isdominated dominated by byTollmien-Schlichting Tollmien-Schlichting instabilities. instabilities. It is well It is known well known from two- from dimensionaltwo-dimensional boundary boundary layers layers that thatthe growth the growth of TSI of TSIcan can be belimited limited by bya favorable a favorable pressure pressure gradient. gradient. ForFor a a given given pressure pressure distribution distribution the the initial initial growth growth of of CFI CFI becomes becomes stronger stronger wi withth increasing increasing Reynolds Reynolds numbernumber and sweep, andand ALTALT is is also also more more likely. likely. In conditionsIn conditions where where the suppressionthe suppression of ALT of andALT CFI and in CFIthe frontin the region front ofregion a wing of sectiona wing cansection no longer can no be longer achieved be solelyachieved by tailoring solely by the tailoring pressure the distribution, pressure distribution,boundary layer boundary suction canlayer stabilize suction the can laminar stabilize ﬂow. the Further laminar downstream, flow. Further the development downstream, of the TSI developmentcan also be limited of TSI bycan suction. also be With limited conventional by suctio structuraln. With conventional concepts, the structural installation concepts,of a suction the installationsystem in the of a area suction of the system wing in box the incurs area of a the certain wing penalty box incurs in weight. a certain To penalty avoidthis, in weight. boundary To avoid layer this,suction boundary has so farlayer been suction limited has to so the far wing been nose, limited which to usuallythe wing ends nose, at aboutwhich 20% usually of chord, ends leading at about to 20%the HLFCof chord, concept. leading Aft to of the the HLFC nose suction concept. area, Aft contour of the nose shaping suction was area, used contour to control shaping TSI, and was the used rules tofor control NLF target TSI, and pressure the rules distributions for NLF target were pressure applied. distributions were applied.

2.3.2.3. Design Design Studies Studies and and Demonstration Demonstration Tests Tests IntroductionIntroduction of of laminar laminar technology technology into into series series production production requires requires multidisciplinary multidisciplinary work work on on aerodynamics,aerodynamics, structuralstructural engineering engineering (surface (surface quality quality in terms in ofterms roughness of roughness and waviness), and waviness), production productiontechnology (closertechnology tolerances), (closer systemstolerances), engineering systems (integration engineering of (integration anti- and de-icing), of anti- andand even de-icing), airline andoperations even airline (maintenance, operations damage (maintenance, repair etc.). damage To demonstrate repair etc.). the To feasibilty demonstrate of NLF the technology, feasibilty of a NLF more technology,advanced ﬂight a more test isadvanced currently flight underway test is within currently the framework underway of within the EU-funded the framework Clean Sky of Ithe projects EU- fundedSFWA (SmartClean Sky Fixed I projects Wing Aircraft) SFWA (Smart and BLADE Fixed (BreakthroughWing Aircraft) Laminarand BLADE Aircraft (Breakthrough Demonstrator) Laminar [18]. AircraftFor Demonstrator) the BLADE ﬂight [18]. tests, the outer wing of an Airbus A340 was exchanged for a new NLF wing withFor considerably the BLADE reduced flight sweep,tests, the as outer shown wing in Figure of an4 .Airbus In order A340 to control was exchanged CFI, a leading for a edge new sweep NLF wingof 20 ◦withcould considerably not be exceeded. reduced Practically sweep, as all shown wings in are Figure tapered 4. In for order the reasonto control of a CFI, favorable a leading spanwise edge sweeploading, of and20° could therefore not thebe sweepexceeded. of constant Practically chord all lineswings and, are hence,tapered isobars, for the in reason the shock of a regionfavorable will spanwiseeven be lower loading, than and 20◦ .therefore As a consequence, the sweep with of constant an aft-swept chord wing, lines the and, design hence, Mach isobars, number in the is limitedshock regionto 0.74 will [19] even in order be tolower avoid than high 20°. wave As drag.a cons Thisequence, was thewith motivation an aft-sw toept develop wing, the the design NLF forward Mach numberswept wing is limited concept to 0.74 within [19] the in DLRorder project to avoid LamAiR high wave [20,21 drag.]. The This basic was idea the ismotivation that a forward to develop swept thewing NLF exhibits forward an increaseswept wing of sweep concept in chordwise within the direction DLR project due toLamAiR the taper [20,21]. (see Figure The basic5), thus idea allowing is that afor forward cruise Machswept numberswing exhibits ranging an fromincrease 0.76 of to sw 0.80,eep likein chordwise for the Airbus direction A320. due In anto the aero-structure taper (see Figurecoupled 5), design,thus allowing it could for be cruise shown Mach that numbers the problem ranging of torsional from 0.76 static to 0.80, divergence like for the that Airbus comes A320. along In an aero-structure coupled design, it could be shown that the problem of torsional static divergence

Energies 2018, 11, 252 5 of 28 Energies 2018, 11, 252 5 of 27 Energies 2018, 11, 252 5 of 27 thatwith comes forward along sweep with can forward be resolved sweep by can aeroelastic be resolved tailoring by aeroelastic utilizing tailoring the anisotropic utilizing characteristics the anisotropic of that comes along with forward sweep can be resolved by aeroelastic tailoring utilizing the anisotropic characteristicsCarbon Fiber Reinforcedof Carbon Fiber Plastics Reinforced (CFRP). Plastics (CFRP). characteristics of Carbon Fiber Reinforced Plastics (CFRP).

FigureFigure 4. 4. BLADEBLADE NLF NLF Outer Outer Wing Wing Flight Flight Test Test [18]. [18]. Figure 4. BLADE NLF Outer Wing Flight Test [18].

Figure 5. Effect of Forward Sweep and Taper. Figure 5. Effect of Forward Sweep and Taper. Figure 5. Effect of Forward Sweep and Taper. HLFC with active boundary layer suction is based on the NLF concept because the same design HLFC with active boundary layer suction is based on the NLF concept because the same design rules apply for the required surface pressure distribution. Therefore, the R&D work conducted in rulesHLFC apply withfor the active required boundary surface layer pressure suction dist is basedribution. on the Therefore, NLF concept the R&D because work the conducted same design in Europe since the mid-eighties in this technology path was focused on the suction system. A large Europerules apply since for the the mid-eighties required surface in this pressure technology distribution. path was Therefore,focused on the the R&D suction work system. conducted A large in scale HLFC demonstration was prepared under leadership of Airbus that resulted in a flight test in scaleEurope HLFC since demonstration the mid-eighties was in prepared this technology under leadersh path wasip of focused Airbus on that the resulted suction in system. a flight A test large in 1998 on the vertical tailplane of an Airbus A320 shown in Figure 6. Aerodynamically, the test was a 1998scale on HLFC the vertical demonstration tailplane was of an prepared Airbus underA320 shown leadership in Figure of Airbus 6. Aerodynamically, that resulted in athe flight test testwas in a success, because extensive laminar flow was observed, but the system was much too complex for 1998success, on thebecause vertical extensive tailplane laminar of an Airbusflow was A320 observ showned, but in Figure the system6. Aerodynamically, was much too thecomplex test was for series production. In a subsequent EU funded project, a simplified suction system was elaborated at seriesa success, production. because In extensive a subsequent laminar EU flow funded was project, observed, a simplified but the system suction was system much was too elaborated complex for at DLR, called the ALTTA concept [22]. DLR,series called production. the ALTTA In a subsequent concept [22]. EU funded project, a simplified suction system was elaborated at DLR, called the ALTTA concept [22].

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Figure 6. A320A320 HLFC Fin Flight Test of 1998. Figure 6. A320 HLFC Fin Flight Test of 1998. The ALTTA conceptconcept isis still still state state of of the the art art for for HLFC HLFC system system development development in in Europe. Europe. As As shown shown in inFigure FigureThe7, it ALTTA7, consists it consists concept of: of: is still state of the art for HLFC system development in Europe. As shown in1. FigureA micro-perforated 7, it consists of: 0.6 to 0.8 mm thick metal sheet, preferable titanium, with standard porosity 1. A micro-perforated 0.6 to 0.8 mm thick metal sheet, preferable titanium, with standard porosity (i.e., 50 µm hole diameter, equally spaced hole pitch of 500 µm). 1. A(i.e., micro-perforated 50 µm hole diameter, 0.6 to 0.8 equally mm thick spaced metal hole sheet, pitch preferable of 500 µm). titanium, with standard porosity 2. Stringers parallel to constant chord lines that divide the double skin into chambers. 2. (i.e.,Stringers 50 µm parallel hole diameter, to constant equally chord spaced lines that hole divide pitch of the 500 double µm). skin into chambers. 2.3. StringersAn inner metalparallel sheet to constant with throttle chord orifices. lines that divide the double skin into chambers. 3. An inner metal sheet with throttle oriﬁces. 3.4. AnA plenum inner metal with sheeta constant with throttleunder-pressure. orifices. 4. A plenum with a constant under-pressure. 4. A plenum with a constant under-pressure.

Figure 7. ALTTA Simplified Suction System. FigureFigure 7. 7. ALTTAALTTA Simplified Simplified Suction System. In each chamber an individual under-pressure is adjusted by the throttle orifices, so that the pressureIn each difference chamber between an individual the outside under-pressure and the chamber is adjusted delivers by the the locally throttle desired orifices, orifices, amount so thatof mass the pressureflow through difference the surface. between It has the been outside shown and in the experiments chamberchamber delivers that this the layout locally is self-adaptivedesired amount to a of certain mass flowrangeflow through through of off-design the the surface. surface. conditions. It has It has Theoreticalbeen been shown shown models in experiments in experimentshave provided that this that first layout this insight layout is self-adaptive into is design self-adaptive parameters to a certain to a rangeofcertain the of perforated range off-design of off-design conditions.skin, but conditions. this Theoretical knowledge Theoretical models is noha modelstve yet provided sufficient have first provided to insight guide first into possible insightdesign intolayoutparameters design and ofmanufacturing the perforated processes. skin, but Experience this knowledge shows that is theno tactual yet sufficientpressure lossto guidecharacteristics, possible aslayout a result and of manufacturinga certain laser drilling processes. process, Experience must be shows determined that the expe actualrimentally pressure [23]. loss characteristics, as a result of a certain laser drilling process, must be determined experimentally [23].

Energies 2018, 11, 252 7 of 28 parameters of the perforated skin, but this knowledge is not yet sufficient to guide possible layout and manufacturing processes. Experience shows that the actual pressure loss characteristics, as a result of a certainEnergies 2018 laser, 11 drilling, 252 process, must be determined experimentally [23]. 7 of 27 The ALTTA concept was tested in 2014 within the frame of the LuFo IV (Luftfahrt-Forschungsprogramm,The ALTTA concept was tested German in 2014 Aeronautical within the Research frame of Program, the LuFo part IV IV)(Luftfahrt- project VERForschungsprogramm,2SUS (Verification ofGerman a Simplified Aeronautical Suction System). Research Again, Program, the target part application IV) project was theVER²SUS vertical tailplane(Verification of theof a AirbusSimplified A320. Suction The System). test was Again, conducted the target in the application DNW (Deutsch-Niederländische was the vertical tailplane Windkanäle,of the Airbus German-Dutch A320. The test Windtunnels)was conducted Large in the Low DNW Speed (Deutsch-Niederländische Facility (LLF) in the Netherlands Windkanäle, at flightGerman-Dutch Reynolds Windtunnels) numbersIt successfully Large Low verified Speed theFacility effectiveness (LLF) in ofthe the Netherlands concept as at well flight as the Reynolds design procedures.numbersIt successfully Figure8a showsverified the the model effectiveness mounted of inthe the concept windtunnel as well test as section.the design In procedures. Figure8b, theFigure achieved 8a showsextent the ofmodel laminar mounted flow, as in determined the windtunnel by infrared test section. imaging, In Figure is shown 8b,in the pink. achieved Encouraged extent byof laminar the results flow, of as the determined wind tunnel by test,infrared a flight imaging, demonstration is shown in is pink. currently Encouraged prepared by within the results the EU of projectthe wind AFloNext. tunnel test, a flight demonstration is currently prepared within the EU project AFloNext.

(a) (b)

Figure 8. Verification of ALTTA concept by Infrared Imaging in Wind Tunnel Figure 8. Veriﬁcation of ALTTA concept by Infrared Imaging in Wind Tunnel.

2.4. Outlook and Next Steps 2.4. Outlook and Next Steps While NLF and HLFC have been examined in great detail, the logical next step is to investigate While NLF and HLFC have been examined in great detail, the logical next step is to investigate full chord laminarization by boundary layer suction. In addition, laminarization of the fuselage full chord laminarization by boundary layer suction. In addition, laminarization of the fuselage should be investigated, as the fuselage is also a major contributor to viscous drag. As described above, should be investigated, as the fuselage is also a major contributor to viscous drag. As described above, detailed analysis of the flow on full configurations including transonic effects and the influence of detailed analysis of the flow on full configurations including transonic effects and the influence of wing sweep induced crossflow on the boundary layer is complex and expensive in terms of scientific wing sweep induced crossflow on the boundary layer is complex and expensive in terms of scientific work, computational effort and experiments. The large effort for taking this approach to the next level work, computational effort and experiments. The large effort for taking this approach to the next must be justified by a thorough study of potentials, using lower order methods. Assuming that cross level must be justified by a thorough study of potentials, using lower order methods. Assuming flow and attachment line instability can be controlled by passive means as shown in the LamAiR that cross flow and attachment line instability can be controlled by passive means as shown in the (Laminar Aircraft Research) project, the second part of this paper focusses on controlling 2D LamAiR (Laminar Aircraft Research) project, the second part of this paper focusses on controlling 2D Tollmien-Schlichting-instabilites by boundary layer suction on wing sections and the fuselage. Tollmien-Schlichting-instabilites by boundary layer suction on wing sections and the fuselage.

3. Low-Order Methods for 2D Suction Design and Airfoil Analysis

3.1. XFOIL and XFOILSUC XFOIL is is a a code code for for the the design design and and analysis analysis of 2D of su 2Dbsonic subsonic airfoils, airfoils, first released ﬁrst released in 1986 in by 1986 Drela by Drela[24]. At [24 the]. AtTU the Delft TU Faculty Delft Faculty of Aerospace of Aerospace Engineer Engineering,ing, several severalstudents, students, in the framework in the framework of their ofmaster their thesis master and thesis supervised and supervised by Boermans, by Boermans, contributed contributed to the implementation to the implementation in XFOIL of in boundary XFOIL of boundarylayer suction layer for suctionlaminarization for laminarization and van Ingen’s and full van e Ingen’s method full foreN themethod calculation for the of transition calculation [12], of transitionwhich is the [12 same], which method is the used same for method the higher-order used for thesimulation higher-order codes simulation developed codesby DLR developed and Airbus. by DLRIn and 2002, Airbus. Ferreira implemented boundary layer suction for laminarization as well as van Ingen’s full eIn 2002, method Ferreira for the implemented calculation of boundary transition layer [25]. suction The latter for laminarizationimplementation as was well needed as van because Ingen’s fullDrela’seN methodenvelope for method the calculation for the calculation of transition of [transition25]. The latter was derived implementation for self-similar was needed Falkner-Skan because laminar boundary velocity profiles and cannot cope with damping of TS instability. In 2004, R.S.W. Broers extended XFOIL for the design of an initial suction distribution, followed by an iterative fine- tuned suction distribution in order to obtain a boundary layer development with prescribed shape factor [26]. Finally, in 2006, Bongers [27] implemented the improved version of the full e method for the calculation of transition, elaborated by van Ingen [28], in XFOIL version 6.93, called from then on XFOILSUC. The improved version offered greater speed and better convergence and calculates damping of the TS waves.

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Drela’s envelope method for the calculation of transition was derived for self-similar Falkner-Skan laminar boundary velocity profiles and cannot cope with damping of TS instability. In 2004, R.S.W. Broers extended XFOIL for the design of an initial suction distribution, followed by an iterative fine-tuned suction distribution in order to obtain a boundary layer development with prescribed shape factor [26]. Finally, in 2006, Bongers [27] implemented the improved version of the full eN method forEnergies the 2018 calculation, 11, 252 of transition, elaborated by van Ingen [28], in XFOIL version 6.93, called from8 then of 27 on XFOILSUC. The improved version offered greater speed and better convergence and calculates dampingXFOILSUC of the TS has waves. been verified by Boermans with the results of detailed boundary layer surveys performedXFOILSUC by van has Ingen been on verified the upper by Boermans surface of with a NACA the results 642-A-215 of detailed airfoil section boundary in the layer low surveys speed performedlow turbulence by van wind Ingen tunnel on the of upperTU Delft, surface Faculty of a of NACA Aerospace 642-A-215 Engineering airfoil section [28]. The in the program low speed has lownot turbulenceyet been released wind tunnel to the of TUscientific Delft, Facultycommunity of Aerospace, but was Engineering kindly provided [28]. The by program Boermans has notto yetTU beenBraunschweig. released to the scientific community, but was kindly provided by Boermans to TU Braunschweig. For the studies presented in this paper, thethe investigatedinvestigated airfoils were discretized with a total of 201 pointspoints,, resultingresulting in in 100 100 panels panels on theon upperthe uppe andr theand bottom the bottom side, respectively, side, respectively, which were which clustered were atclustered the nose at to the account nose to for account the higher for the gradients higher of grad pressureients of distribution pressure distribution and boundary and layer boundary parameters. layer N Theparameters. improved The version improved of the version full e ofmethod the full was e used method for transition was used prediction for transition for all prediction cases, with for and all withoutcases, with suction. and without As there suction. was no As need there to replicatewas no need any windto replicate tunnel any test, wind but rather tunnel provide test, but realistic rather free-flightprovide realistic data, the free-flight critical N-factordata, the was critical set toN-factor 13. was set to 13.

3.2. Auxiliary Software A setset ofof auxiliaryauxiliary programsprograms was was written written in in MatLab MatLab at at Technical Technical University University (TU) (TU) Braunschweig Braunschweig for semi-automaticfor semi-automatic or manual or manual suction suction design, design, data handling data handling and calculation and calculation of properties, of properties, that are notthat part are ofnot the part XFOIL of the dataset, XFOIL such dataset, as the such permeability as the permeability and pressure and loss pressure of the suction loss of skin, the the suction pressure skin, of the internalpressure plenums, of the internal the data plenums, related tothe the data suction relate pump,d to the and suction the calculation pump, and of totalthe calculation drag as described of total indrag Section as described 4.3. Figure in 9Section gives an4.3. overview Figure 9 ofgives the programan overview structure of the andprogram the data structure ﬂow for and automatic the data suctionflow for design. automatic Input suction and output design. data Input is savedand output on disk da forta is every saved iteration on disk of for the every suction iteration distribution, of the allowingsuction distribution, an easy re-start allowing if XFOILSUC an easy re-start gets hung if XFOILSUC up. It also gets allows hung to up. retrieve It also a allows case from to retrieve the data a archivecase from to the use data as a startingarchive to point use foras a further starting design point steps. for further design steps.

Figure 9. XFOILSUC and auxiliary software. Figure 9. XFOILSUC and auxiliary software.

4. Components and Parameters of the LFC System

expelled from a nozzle at the velocity . To improve efficiency, a suction system may have several segments with different plenum pressures. The system shown here has two plenums, one each for the upper and lower side.

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4. Components and Parameters of the LFC System

4.1. General Layout and Components The main components of a suction system are shown in Figure 10. Air is sucked through the porous skin into a plenum. The plenum pressure has to be lower than the external pressure given by the pressure distribution to overcome the ﬂow resistance of the skin. From the plenum, air is sucked via a system of ducts to a pump, which increases the total pressure of the suction air, before it is expelled from a nozzle at the velocity Ujet. To improve efﬁciency, a suction system may have several segments with different plenum pressures. The system shown here has two plenums, one each for the upperEnergies and2018,lower 11, 252 side. 9 of 27

Figure 10.10. Components of Suction System.

The keykey componentcomponent of thethe suctionsuction systemsystem is thethe porousporous wingwing skin.skin. From aa purelypurely aerodynamicaerodynamic pointpoint ofof view, thethe optimumoptimum skinskin wouldwould have a continuouscontinuous and homogenous porosity. Lachmann [[29]29] has usedused sinteredsintered metalmetal andand resin-impregnatedresin-impregnated glass ﬁbrefibre ﬂeecefleece asas surfacesurface materialmaterial inin windwind tunneltunnel experiments, which showed goodgood performance.performance. Ho However,wever, such skin materials cannot carry any structuralstructural loadsloads andand tend tend to to get get clogged clogged with with atmospheric atmospheric dust. dust. The The other other type type of suctionof suction skin skin that that has beenhas been researched researched is sheet is metalsheet with metal a large with number a large discrete number small discrete holes orsmall slots. holes Most or experimenters slots. Most inexperimenters the past have in used the past skins have with used a uniform skins gridwith ofa un suctioniform holes grid of of suction constant hole diameter,s of constant which diameter, results in awhich constant results permeability in a constant of thepermeability suction skin. of the The suction actual skin. suction The rate actual is then suction a function rate is ofthen the a pressurefunction differenceof the pressure across difference the skin, whichacross inthe turn skin, depends which onin turn theouter depends pressure on the distribution outer pressure and thedistribution pressure insideand the the pressure wing. inside In order the to wing. provide In order chordwise to provi tailoringde chordwise of the tailoring suction distribution,of the suction systems distribution, with multiplesystems with plenums multiple have plenums been suggested, have been as wellsuggested, as internal as well structures as internal with structures throttling propertieswith throttling [22]. However,properties with [22]. todays’ However, CNC with machining todays’ possibilities, CNC machining including possibilities, laser beam, including electron laser beam beam, or mechanical electron drilling,beam or itmechanical might even drilling, be possible it might to produce even abe skin possible with a to chordwise produce anda skin spanwise with a variation chordwise of holeand spacingspanwise and variation hole diameter, of hole spacing so that theand suction hole diameter, rate can so be that optimized the suction locally rate with can be just optimized a single or locally very fewwith plenum just a single chambers or very with few a constantplenum chambers inside pressure. with a Ofconstant course, inside changes pressure. in angle Of ofcourse, attack changes change thein angle outer of pressure attack change distribution, the outer so the pressure system mustdistribution, provide so a certainthe system level must of robustness provide anda certain versatility level (thisof robustness is looked and at in versatility more detail (this in Sectionis looked5). at in more detail in Section 5).

4.2. Suction & Skin Parameters Within XFOILSUC, the the suction suction design design for for a asingle single operating operating point point defined deﬁned by bylift liftcoefficient, coefﬁcient, Re ⁄ Reand and MachMach number number involves involves only only the the non-dimensional non-dimensional suction suction velocity velocity v0/U∞ asas the designdesign parameterparameter and requires no consideration of the suctionsuction skin properties or any internal pressures. However, these areare relevantrelevant forfor thethe assessmentassessment ofof thethe suctionsuction compressorcompressor system,system, it’s powerpower consumption,consumption, andand thethe thrust produced byby the dischargeddischarged suctionsuction air.air. InIn addition,addition, thethe permeability ofof the skin is a mechanical feature that does not change, which means that instead the suction

distribution changes with the pressure distribution, when or are varied. To calculate such off- design cases, it is also necessary to include the skin properties in the design chain. The porosity is the fraction of flow-through area, i.e., the cross-sectional area of holes or

slits per unit of skin area. If is the mean flow velocity inside the suction holes, then the nominal suction velocity becomes = ∙, which is identical to the volume flow per unit skin area. The direction of is defined positive in direction of the surface normal vector and consequently, is negative when suction is applied. The nondimensional volumetric suction coefficient is the total volume flow through the skin, , divided by reference area and free stream velocity. It can also be

obtained by integrating the nondimensional suction velocity ⁄ along the 2D profile contour (Equation (1). The contour coordinate runs from the trailing edge along the upper side of the airfoil to the nose and then back to the rear end along the bottom side (Figure 11):

Energies 2018, 11, 252 10 of 28

the skin is a mechanical feature that does not change, which means that instead the suction distribution changes with the pressure distribution, when CL or α are varied. To calculate such off-design cases, it is also necessary to include the skin properties in the design chain. The porosity por is the fraction of flow-through area, i.e., the cross-sectional area of holes or slits per unit of skin area. If vh is the mean flow velocity inside the suction holes, then the nominal suction velocity becomes v0 = vh·por, which is identical to the volume flow per unit skin area. The direction of v0 is defined positive in direction of the surface normal vector and consequently, v0 is negative when suction is applied. The nondimensional volumetric suction coefficient is the total volume flow through the skin, Q, divided by reference area and free stream velocity. It can also be obtained by integrating the nondimensional suction velocity v0/U∞ along the 2D profile contour (Equation (1). The contour coordinate s runs from the trailing edge along the upper side of the airfoil to the nose and then back to Energiesthe rear 2018 end, 11, 252 along the bottom side (Figure 11): 10 of 27 Energies 2018, 11, 252 10 of 27 Q 1 I v C = = − 0 ds (1) Q · 1 = Sre f U∞=− l s U∞ (1) 1 ∙ = =− (1) ∙

Figure 11. Airfoil Contour Cordinate System. Figure 11. Airfoil Contour Cordinate System Figure 11. Airfoil Contour Cordinate System

FigureFigure 12 12 gives gives an an overview overview of ofthe the most most important important pressure pressure definitions. definitions. c isp isthe the wall wall pressure pressure distributiondistributionFigure created created12 gives by by thean the overview external external flow, of flow, the nondim most nondimensionalized importantensionalized pressure by by the the definitions. free free stream stream dynamic dynamic is the wall pressure: pressure: pressure distribution created by the external flow, nondimensionalized by the free stream dynamic pressure: − − pw − p∞ pw − p∞ = = (2) cp = −= ∙∙− (2) = q = 1· · 2 ∞ 2 $∞ U∞ (2) ∙∙ To obtain the skin pressure coefficient ,, the pressure drop across the suction skin is also To obtain the skin pressure coefficient c , the pressure drop across the suction skin is also made dimensionlessTo obtain the with skin freepressure stream coefficient dynamic pressure:p,,skin , the pressure drop across the suction skin is also mademade dimensionlessdimensionless withwith freefree stream stream dynamic dynamic pressure: pressure: Δ − , = Δ = −− (3) ∆pskin ppl pw cp,skin,== == (3)(3) q∞ q∞

Figure 12. Cp, Cp,pl & Cp,skin. Figure 12. Cp, Cp,pl & Cp,skin. Figure 12. cp, cp,pl & cp,skin. The plenum pressure coefficient is defined similarly to the wall pressure coefficient, except that the pressureThe plenum inside pressurethe suction coefficient plenum is isdefined used. similarlyThe pressure to the coefficientwall pressure coefficient,can be positive except or that the Thepressure plenum inside pressure the suction coefficient plenum is defined is used. similarly The pressure to the wall coefficient pressure coefficient, can be exceptpositive that or negative, depending on the pressure distribution on the profile. When suction is applied, , is thenegative, pressure depending inside the on suction the pressu plenumre isdistribution used. The pressureon the profil coefficiente. Whencp cansuction be positive is applied, or negative, is always less than the minimum , and in consequence, , , though not constant, is always, negative:always less than the minimum , and in consequence, , , though not constant, is always negative: − , = −=, + (4) , = =, + (4) The plenum pressure coefficient , for each plenum segment is found by inserting the The plenum pressure coefficient for each plenum segment is found by inserting the maximum , (the least negative) and the, minimum for that segment into Equation (4). The maximum (the least negative) and the minimum for that segment into Equation (4). The minimum is, taken from the pressure distribution, while the maximum , must be defined by settingminimum the minimum is taken absolute from the skin pressure pressure distribution, loss. Several while researchers the maximum have attempted, must to bederive defined the by flowsetting resistance the minimum of a suction absolute skin as skin a function pressure of loss. geometrical Several researchersand flow parameters, have attempted either tofrom derive first the principlesflow resistance or experimental of a suction data, skin but asinvariably a function without of geometrical success [23]. and While flow parameters, the main driver either is knownfrom first to beprinciples the porosity, or experimental too many other data, parametersbut invariably were without found successto be of [23]. relevance, While somethe main of which driver can is known not be toproperly be the porosity,measured too or manyreproduc othered parametersduring manufacturing. were found Forto be the of purpose relevance, of somethis paper, of which a simple can not relationshipbe properly typical measured for problemsor reproduc ofed internal during manufacturing.flow is assumed, For inthe which purpose the of pressure this paper, drop a simple is proportionalrelationship to thetypical dynamic for problems pressure ofof the internal suction fl velocity,ow is assumed, , and a inpressure which lossthe coefficientpressure drop: is proportional to the dynamic pressure of the suction velocity, , and a pressure loss coefficient :

Energies 2018, 11, 252 11 of 28

depending on the pressure distribution on the proﬁle. When suction is applied, cp,pl is always less than the minimum cp, and in consequence, cp,skin , though not constant, is always negative:

ppl − p∞ cp,pl = = cp,skin + cp (4) q∞

The plenum pressure coefficient cp,pl for each plenum segment is found by inserting the maximum cp,skin (the least negative) and the minimum cp for that segment into Equation (4). The minimum cp is taken from the pressure distribution, while the maximum cp,skin must be defined by setting the minimum absolute skin pressure loss. Several researchers have attempted to derive the flow resistance of a suction skin as a function of geometrical and flow parameters, either from first principles or experimental data, but invariably without success [23]. While the main driver is known to be the porosity, too many other parameters were found to be of relevance, some of which can not be properly measured or reproduced during manufacturing. For the purpose of this paper, a simple relationship typical for problems of internal flow is assumed, in which the pressure drop is proportional to the dynamic pressure of the suction velocity, qs, and a pressure loss coefficient ζ:

1 ∆p = −ζ·q = −ζ· ·$ ·v2 s s 2 w 0 By deﬁning the skin permeability as the inverse of the pressure loss coefﬁcient, we can express the suction velocity as function of the skin pressure drop, the wall density, and the permeability:

1 ϕ = ζ

2 −∆ps v0 = 2·ϕ· (5) $w The pressure differential is expressed by the skin pressure coefﬁcient and free stream dynamic pressure: $ ∆p = c · ∞ ·U2 (5) s p,skin 2 ∞ Combining Equations (5) and (6) and some rearrangement results in the nondimensional suction velocity v0/U∞ and eliminates U∞ from the right hand side of the equation:

v2 $ 0 = − · · ∞ 2 cp,skin ϕ U∞ $w

The wall density ratio $w/$∞ is equal to the pressure ratio multiplied by the inverse of the temperature ratio, which can both be expressed as functions of Mach number and wall pressure coefﬁcient using basic gas dynamic relationships for boundary layers [3,30]:

$ p T 0.16055·M2 + 1 w = w · ∞ = ∞ (6) $ p T 1 2 ∞ ∞ w 2 ·cp·κ·M∞ + 1

Inserting the densitiy ratio $w/$∞ given by Equation (7) leads to an expression for the nondimensional suction velocity containing only quantities known from XFOIL output, and the skin permeability distribution:

v2 1 ·c ·κ·M2 + 1 0 = − · · 2 p ∞ 2 cp,skin ϕ 2 (7) U∞ 0.16055·M∞ + 1

By solving Equation (8) for ϕ, we can extract the permeability distribution of the skin from a given case, which in turn allows us to compute the suction velocity distribution for other operating points. Energies 2018, 11, 252 12 of 28

4.3. Drag and Thrust Bookkeeping The 2D profile drag is the sum of all forces parallel to free stream velocity caused by external pressure and wall shear stress. It can be determined by integrating pressure and shear stress over the profile contour. Another classical approach to profile drag identification is to integrate the momentum deficit in the wake of the profile. The underlying assumption is that any portion of air influenced by the airfoil also passes through the wake plane, which is not true if the boundary layer is partially sucked into the wing. For airfoils with suction, the remaining wake drag is only a small portion of the total drag. It can be shown by momentum analysis that the profile drag is equal to the sum of the wake drag and the momentum intake of the suction system [26], which in turn can be shown to be exactly twice the suction coefficient CQ:

CDp f = CDp + CD f = CDw + 2·CQ

As can be seen from this equation, the amount of suction should be kept as low as possible, just enough to ensure laminar flow. Any additional suction will not reduce drag further, but leads to increased profile drag. The drag that has to be overcome by engine thrust is the profile drag minus the thrust produced by the jet thrust of the expelled suction air. In terms of total energy, the aircraft power system also has to deliver the power to drive the suction compressor. We therefore define a total drag coefficient CDtot, on which calculations of drag savings in this paper will be based:

CDtot = CDp f − CTjet + CDC

The pressure difference that must be overcome by the suction pump has four components:

1. Difference of local static pressure on outer contour to p∞; 2. Flow resistance through porous skin; 3. Flow resistance through ducting; 4. Dynamic pressure of discharge velocity.

Components #2 and #3 can be small for a well-designed suction system, component #1 depends on the body geometry and the angle of attack. If the suction air is simply spilled overboard at a negligible discharge velocity, the #4 component is zero. The pressure components can be made dimensionless by introducing pressure coefficients. The first component is simply the wall pressure coefficient defined in Equation (2), the second is the skin pressure coefficient defined in Equation (3), and their sum is the plenum pressure coefficient defined in Equation (4). This simplifies the calculation of pump power, as cp,pl is constant per segment of the suction system. The flow resistance of the ducting from the plenum to the pump is expressed by the duct pressure loss coefficient cp,duct, which depends on the design of the ducting and the mass flow. As the internal design of the LFC system is beyond the scope of this paper, cp,duct is also assumed to be a constant. The pressure at the pump inlet then becomes:

ppump,in = (cp,pl + cp, duct)·q∞ + p∞

The temperature at the pump inlet is identical to the outer skin temperature, as the skin and the duct can be considered adiabatic throttling devices. Two outlet properties, the static pressure p∞ and the jet velocity Ujet, are given. For the dynamic pressure component, the density must be determined from static pressure and the pump outlet temperature, which is a function of the pressure ratio and the pump efficiency. The correct values are determined by iterative calculations, and the final compressor power is equal to the change in specific enthalpy of the suction air multiplied by the mass flow. In incompressible flow, the mass flow could be calculated from CQ and free stream quantities. . However, for compressible flows, m is not proportional to volumetric CQ, and we need to define a Energies 2018, 11, 252 13 of 28

mass flow dependent suction coefficient. Using the previously defined wall densitiy ratio, $w/$∞, this can also be calculated from non-dimensional 2D quantities, analogous to Equation (1):

. I m 1 v0 $w CQm = = − · ds Sre f ·U∞·$∞ l s U∞ $∞

Dividing pump power by ﬂight speed results in the equivalent compressor drag DC, which can be converted to a nondimensional compressor drag coefﬁcient:

D P C = C = C DC 1 2 1 3 2 ·$∞·U∞·Sre f 2 ·$∞·U∞·Sre f

Finally, the jet of expelled suction air produces thrust, which is expressed by the thrust coefﬁcient:

. m ·Ujet C = Tjet 1 2 2 ·ρ·U∞·Sre f

Obviously, compressor power and jet thrust both increase with jet velocity. If the design goal is to minimize pump power, the jet velocity should be kept as low as possible, while the optimum from a total energy point of view is found by minimizing CDC − CTjet. It can be shown that the optimum discharge velocity is equal to flight velocity for ηC = 1 and decreases with compressor efficiency. The jet thrust is generated with very high propulsive efficiencies (ηprop > 1), because the inlet momentum or sink drag is already contained in the profile drag. Correct matching of the jet velocity to the compressor efficiency is important for the overall system efficiency. For the calculation of pump drag and suction air jet thrust, a compressor efficiency of ηC = 0.7 was assumed and the jet velocity set to Ujet = 0.7·U∞. These rather conservative values are similar to those reported for the Northrop X-21, the only large jet plane with full chord LFC flown [31]. The minimum skin pressure loss coeffcient and the duct pressure loss coefficient were both set to a value of 0.1.

4.4. Strategies for Suction Design For very simple cases such as the flat plate at zero incidence, theoretical solutions exist for suction velocities and shape factors that will ensure laminar flow in incompressible flow. For real world cases with varying pressure gradients, high Reynolds numbers and compressability effects, suction designs must be made by iteratively running the boundary layer solver, checking the solution and adaptation of the suction where certain target values are not met. As the primary objective is the transition location, one could start out with a small constant suction and simply scale it up until the target transition location is achieved. However, this would lead to oversucking on much of the wing surface, which results in unwanted additional skin friction and can even cause early transition rather than delaying it. As the local shape factor H12, which is the ratio of the displacement thickness to the momentum thickness, has been shown to be a good design parameter for suction design, the method used for automatic suction design was to increase or decrease suction locally, depending on wether the shape factor was under or over target. Because the boundary layer flow at a given location is a result of the flow “history”, rather than just the suction at that point, back stepping of the suction adaptation is necessary. Applying moving average smoothing to both the shape factor and the adapted suction distribution provided sufficient chordwise propagation and also proved helpful in damping out numerical oscillations. For flat plates in incompressible natural laminar flow, H12 is known to be 2.6, and this has been successfully used as the target value for the suction design of sailplane wing sections, which operate at low speeds. H12 must be kept below 2.6 for higher Reynolds numbers in incompressible flow, while higher Mach numbers provide damping of the TS waves, which permits running higher shape factors, in excess of 3, without unwanted transition. For cases with both, high Reynolds and Mach numbers, a varying target shape factor was used for some cases, decreasing Energies 2018, 11, 252 14 of 28 linearly from LE to TE. It was also tried to use the N-factor calculated by XFOILSUCs transition prediction module as target parameter, the idea being that a steady increase towards the TE might lead to a good suction design. However, the N-factor turned out to be too sensitive to small changes in suction for use in an automated design routine. For some cases, where no good solution could be found by automatic design, the suction was adapted manually. It must be pointed out that neither the perfect target criterion nor its optimum value have been reliably identified and that further research on this subject is necessary.

5. Results for a Supercritical Airfoil The effect of boundary layer suction shall be demonstrated for the example of the DLR F15 (see Figure 13), which is a generic supercritical airfoil developed during the FNG project [32]. Although this airfoil was not designed for laminar flow, and no modifications to the shape were made, it shows a certain extent of natural laminar flow in 2D calculations, as can be seen from the jump in the skin friction coefficient Cf shown in Figure 14. At Mach 0.7, a Reynolds number of 30 million and CL = 0.5, natural transition (NT) occurs around 5% chord length on the upper side, while the lower surface stays laminar almost up to 50%. With suction (LFC) up to 80% of chord length, the transition is delayed to 85% on both sides. Also shown in Figure 14 are the amplification factor, N, and the shape factor, H12, together with the applied non-dimensional suction velocity, v0/U∞, and the necessary permeability, ϕ, for the same operating point, which is typical of cruise conditions for a midrange jet aircraft or Energies 2018, 11, 252 14 of 27 regional turboprop.

FigureFigure 13. 13. InvestigatedInvestigated Airfoils. Airfoils.

Energies 2018, 11, 252 15 of 28 Energies 2018, 11, 252 15 of 27

Figure 14. Boundary Layer and Suction Parameters; DLR-F15; =30∙106 ; = 0.7; =0.5. Figure 14. Boundary Layer and Suction Parameters; DLR-F15; Re = 30·10 ; Ma = 0.7; CL = 0.5. For off-design operating points, the performance and boundary layer data must be computed in three Forsteps. off-design First, the operating skin permeability, points, the given performance implicitly and by the boundary suction layer and pressure data must distribution, be computed is extractedin three steps. from First,the initial the skindesign permeability, case. The pressure given implicitly distribution by the is obtained suction and by running pressure the distribution, new case withis extracted the oldfrom suction the initial velocity design distribution case. The pressurethrough distribution XFOILSUC. is obtainedThe upper by runningand lower the newplenum case with the old suction velocity distribution through XFOILSUC. The upper and lower plenum pressures pressures are adjusted to comply with the maximum , —condition, and the new suction c velocitiesare adjusted are todetermined comply with using the Equation maximum (8), pbefo,skin—condition,re the case is and re-run the newin XFOIL suction to velocitiesget the final are solution.determined When using this Equation is done (8),for beforean angle-of-attack-sequence, the case is re-run in XFOIL a =to get the-polar ﬁnal can solution. be plotted.TwoWhen this is done for an angle-of-attack-sequence, a ϕ = const-polar can be plotted.Two examples are shown in examples are shown in Figure 15, for permeabilites designed at = 0.4 and 0.6. The polars have the C typicalFigure 15shape, for permeabilitesof NLF profiles designed at lower at ReynoldsL = 0.4 and numbers, 0.6. The polarswith a havepronounced the typical laminar shape ofbucket NLF covering a certain span of lift coefficients around the design point. The laminar bucket can be

Energies 2018, 11, 252 16 of 28

Energies 2018, 11, 252 16 of 27 profiles at lower Reynolds numbers, with a pronounced laminar bucket covering a certain span of lift extendedcoefficients by around lowering the the design plenum point. pressure The laminar and thus bucket increasing can be extendedthe suction, by as lowering can be theseen plenum in the pressure and thus increasing the suction, as can be seen in the dashed curve, for which the minimum dashed curve, for which the minimum skin pressure loss was increased from , =−0.1 to −0.2. Itskin can pressure also be seen loss wasthat increasedfor lift coefficients from cp,skin abov= −e 0.1the tolaminar−0.2. Itbucket, can also drag be seenincreases that forquickly lift coefficients and even surpassesabove the laminarthat of the bucket, unsucked drag base increases airfoil, quickly which and is due even tosurpasses the increased that skin of the friction unsucked induced base by airfoil, the whichnow obsolete is due suction. to the increased The data skinfor the friction natural induced transition by reference the now obsoleteare repres suction.ented by The the dash-dotted data for the line,natural while transition the solid reference black line are is representedthe minimum by drag the dash-dottedenvelope obtained line, while by adjusting the solid the black permeability line is the forminimum each individual drag envelope operating obtained point. by adjusting the permeability for each individual operating point.

FigureFigure 15. 15. PolarsPolars with with fixed ﬁxed and and adjusted adjusted permeability; permeability; DLR-F15; DLR-F15; Re = 3030· ∙10 106; ; Ma = =0.7. 0.7.

One of the questions that arise when designing suction systems is what portion of the wing One of the questions that arise when designing suction systems is what portion of the wing chord chord should be sucked. Extension of the suction area will of course increase the technical complexity should be sucked. Extension of the suction area will of course increase the technical complexity and and the weight and power consumption of the suction system, which have to be compensated by the the weight and power consumption of the suction system, which have to be compensated by the drag saved. drag saved. In Figure 16, a variation of the suction length for the DLR F15 and the resulting drag components In Figure 16, a variation of the suction length for the DLR F15 and the resulting drag components are shown. As can be seen in Figure 16a, both the profile and the total drag decrease over the entire are shown. As can be seen in Figure 16a, both the proﬁle and the total drag decrease over the entire interval, which means the increase in pump drag is more than compensated by the decrease in interval, which means the increase in pump drag C is more than compensated by the decrease in profile drag up to the maximum suction length. In FigureDc 16b, the relative drag reduction is shown, proﬁle drag up to the maximum suction length. In Figure 16b, the relative drag reduction is shown, along with the ratio of drag reduction and resulting pump drag. In conclusion, the suction length is along with the ratio of drag reduction and resulting pump drag. In conclusion, the suction length is only limited by technical feasibility of the suction system integration. As provision must be made for only limited by technical feasibility of the suction system integration. As provision must be made trailing edge movables for lateral control and high lift, it was decided to limit suction to 80% of chord for trailing edge movables for lateral control and high lift, it was decided to limit suction to 80% of length for the investigations in this paper, which results in transition at approximately 85% on both chord length for the investigations in this paper, which results in transition at approximately 85% on sides. both sides.

Energies 20182018, ,1111, ,252 252 1717 of of 27 28 Energies 2018, 11, 252 17 of 27

(a) (b) (a) (b) = 30 ∙ 10; =0.7 FigureFigure 16. 16. VariationVariation of of suction suction length; length; DLR-F15; DLR-F15; Re = 30·106; Ma = 0.7.. Figure 16. Variation of suction length; DLR-F15; = 30 ∙ 10; =0.7. Profile drag in general depends on Reynolds and Mach number, which remains true for airfoils Profile drag in general depends on Reynolds and Mach number, which remains true for airfoils with Profilelaminar drag flow in control, general where depends also on the Reynolds necessary an suctiond Mach to number, prevent which transition remains changes. true forAs airfoilscan be with laminar flow control, where also the necessary suction to prevent transition changes. As can be seenwith inlaminar the left flow graph control, in Figure where 17, also in the the interval necessary between suction Re to = prevent10–40 million, transition the changes.profile drag As canof the be Re unsuckedseen in the supercritical left graph in airfoil Figure stays 1717,, inalmost the interval constant. between The reason Re = 10–40 for this million, is the mutualthe profile profile compensation drag of the ofunsucked two effects: supercritical Skin friction airfoilairfoil for staysstays either almost almost laminar constant. constant. or turbulent The The reason reason boundary for for this this islayer is the the mutualdecreases, mutual compensation compensation while at the of sameoftwo two effects: time, effects: the Skin Skintransition friction friction for location eitherfor either laminaron thelaminar lower or turbulent or si turbulentde moves boundary boundaryforward, layer increasinglayer decreases, decreases, the while portion while at the ofat same the airfoilsametime, the time,that transition has the a transition turbulent location locationboundary on the loweron layer. the side lower For moves the si deLFC forward, moves cases, forward, increasingthe transition increasing the portionposition the of does portion the airfoilnot move,of thatthe andairfoilhas ahence, turbulent that thehas profilea boundary turbulent drag layer. boundarydecreases For the withlayer. LFC Re For. cases, The the amount LFC the transition cases, of suction the position transition air, as does well position not as the move, does pump andnot power, move, hence, Re decreaseandthe profilehence, as dragthewell. profile decreases The rightdrag withdecreaseshand graph. The with in amount ReFigure. The of 17amount suction shows of air,the suction asinfluence well air, as as theof well Mach pump as numberthe power, pump for decrease power, Re = = constdecreaseas well.. The Theas profile well. right dragThe hand right stays graph hand nearly in graph Figureconstant in 17Figure up shows to 17a Mach theshows influence number the influence of 0.6, Mach beyond of numberMach which number for aRe significant for constRe = . dragTheconst profilerise. The can profiledrag be observed, stays dragnearly stays both nearly constant with constant and up towithou aup Mach tot asuction. Mach number number The of 0.6,suction of beyond 0.6, requirements beyond which which a significant and a significantthe pump drag dragrise can alsorise be increasecan observed, be observed, with both Mach withboth numbers andwith without and above withou suction.0.6. t suction. The suction The suction requirements requirements and theand pump the pump drag dragalso increasealso increase with Machwith Mach numbers numbers above above 0.6. 0.6.

(a) (b) (a) (b) Figure 17. DLR F15, Sensitivity to (a) Reynolds and (b) Mach number. Figure 17. DLRDLR F15, F15, Sensitivity to ( a)) Reynolds Reynolds and ( b) Mach number. 6. Results for NLF- and BWB-Airfoils 6. Results for NLF- and BWB-Airfoils In the previous section, total drag reduction by a factor of three was shown on a supercritical airfoilIn whichthe previous normally section, is mostly total dragturbulent. reduction Airf byoils a designedfactor of threefor natural was shown laminar on aflow supercritical employ extendedairfoil which areas normally with negative is mostly pressure turbulent. gradients Airf inoils the designed front part for to keepnatural the laminarflow laminar flow onemploy both extended areas with negative pressure gradients in the front part to keep the flow laminar on both

Energies 2018, 11, 252 18 of 28

6. Results for NLF- and BWB-Airfoils In the previous section, total drag reduction by a factor of three was shown on a supercritical airfoil whichEnergies normally2018, 11, 252 is mostly turbulent. Airfoils designed for natural laminar flow employ extended18 areas of 27 with negative pressure gradients in the front part to keep the flow laminar on both sides without suction, followedsides without by a suction, short and followed strong by pressure a short rise and on strong the rear pressure part, whichrise on causes the rearimmediate part, which transition causes. Theimmediate low drag transition. characteristics The low are drag only characteristics achieved for a limitedare only window achieved around for a thelimited design window lift coefficient. around Belowthe design or above lift coefficient. the so-called Below laminar or above bucket, the so transition-called laminar moves bucket, forward transition on at least moves one forward side of theon airfoil,at least causingone side drag of the coefficients airfoil, causing similar drag to thosecoeffici ofents fully similar turbulent to those wing of section.fully turbulent wing section. The DLR-LC2B also shown in Figure 13 is a NLFNLF airfoilairfoil originallyoriginally developed for use on commuter planesplanes.. To To investigate investigate the the potential potential for for further further drag drag reduction, reduction, suction suction was was applied applied to the to theregion region behind behind the thepressu pressurere minimum. minimum. The The laminar laminar flow flow length length could could be be extended extended almost almost to to the trailing edge, inducing a relative drag reduction on thethe orderorder ofof 40%.40%. Compared to thethe supercriticalsupercritical airfoil with with suction, suction, the the resulting resulting total total drag drag is lower is lower by 16%. by 16%.The lowest The lowestvalues observed values observed were were= C0.0015Dtot = for0.0015 the fortotal the drag total coefficient. drag coefficient. Outside Outside the lami thenar laminar bucket, bucket, a sharp a sharp suction suction peak peak builds builds up upat atthe the leading leading edge, edge, causing causing instant instant transition transition that that requ requiresires additional additional suction suction at at the the nose nose to to suppress, suppress, which explains the jump in CDtot near near the theupper upperand and lowerlower limitslimits ofof thethe laminarlaminar region,region, asas shown in Figure 1818..

= 30 ∙ 10; = 0.7 FigureFigure 18. 18. EffectEffect of of additional additional suction suction on on a a NLF NLF airfoil; airfoil; DLR DLR LC2B; LC2B; Re = 30·106; Ma = 0.7. .

While the minimum thickness of wing sections is limited by strength and weight consideration, While the minimum thickness of wing sections is limited by strength and weight consideration, blended wing body (BWB) configurations require airfoils also optimized for internal volume. One blended wing body (BWB) configurations require airfoils also optimized for internal volume. example is the NASA SC(2)-0518, also shown in Figure 13, that was used by Lyu and Martins [33] for One example is the NASA SC(2)-0518, also shown in Figure 13, that was used by Lyu and Martins [33] the center section in their study of a BWB configuration. It features a very blunt nose and a high for the center section in their study of a BWB configuration. It features a very blunt nose and a high thickness-to-cord ratio of 18%, resulting in a fairly high profile drag coefficient between 0.007 and thickness-to-cord ratio of 18%, resulting in a fairly high profile drag coefficient between 0.007 and 0.008 for Re = 30 million and Ma = 0.7, shown by the dash-dotted curve in Figure 19. Application of 0.008 for Re = 30 million and Ma = 0.7, shown by the dash-dotted curve in Figure 19. Application of suction can reduce the total drag coefficient to just above 0.003, which is larger than the minimum suctiondrag coefficient can reduce found the totalfor the drag much coefficient thinner to supercri just abovetical 0.003, and NLF which airfoils, is larger but than represents the minimum a reduction drag coefficientof 60% compared found forto the muchturbulent thinner base supercritical airfoil. These and numbers NLF airfoils, must be but treated represents with asome reduction care, ofas 60%most compared BWB configurations to the turbulent presented base airfoil. today Thesehave st numbersrong leading must beedge treated sweep with and some higher care, Reynolds as most BWBnumbers configurations than those examined presented here. today Nevertheless, have strong leading the application edge sweep of BL and suction higher to Reynoldsvery thick numbers profiles thanpromises those significant examined here.drag Nevertheless,reductions and the is application surely worth of BLa closer suction look, to very using thick more profiles sophisticated promises significantsimulation dragtools. reductions and is surely worth a closer look, using more sophisticated simulation tools.

Energies 2018, 11, 252 19 of 28 EnergiesEnergies 20182018,, 1111,, 252252 1919 ofof 2727

FigureFigure 19. 19. EffectEffect ofof of SuctionSuction Suction onon on aaa high-thickhigh-thick high-thicknessnessness BWBBWB BWB airfoil;airfoil; airfoil; NASANASA NASA SC(2)-0518;SC(2)-0518; SC(2)-0518; Re = == 30 3030 ∙ ∙· 10 106;;; Ma = = 0.7 0.7..

7. Application of Suction to Fuselages 7. Application of Suction to Fuselages

7.1.7.1. Design Design of of Fuselage Fuselage Suction Suction NaturalNatural laminarlaminar laminar flowflow flow onon on thethe the frontfront front 30%30% 30% ofof aa ofbusinessbusiness a business jetjet fuselagefuselage jet fuselage hashas beenbeen has investigatedinvestigated been investigated byby HolmesHolmes by etHolmeset al.al. [34],[34], et al.andand [34 potentialpotential], and potential ACAC levellevel AC dragdrag level savingssavings drag savings ofof 7%7% were ofwere 7% reported.reported. were reported. However,However, However, thethe muchmuch the larger muchlarger ReynoldslargerReynolds Reynolds numbersnumbers numbers ofof passengerpassenger of passenger planesplanes planes makemake make significantsignificant significant NLFNLF NLF impossible.impossible. impossible. ActiveActive Active fuselagefuselage laminarizationlaminarization has has been been suggestedsuggested by by Lachmann [29], [[29],29], including re-lam re-laminarizationre-laminarizationinarization behind behind a a turbulent turbulentturbulent cockpitcockpit section section byby completecomplete complete boundaryboundary boundary layerlayer layer removal,removal, removal, asas as earlyearly early asas as 1962,1962, 1962, andand and laterlater later byby by PfenningerPfenninger Pfenninger [35][35] [35 inin] 1987.in1987. 1987 OfOf. Of course,course, course, suctionsuction suction onon on aa fuselage afuselage fuselage skinskin skin posesposes poses aa a numbernumber number ofof of technicaltechnical technical challenges. challenges. WhileWhile windows cancan inin theorytheorytheory be bebe replaced replacedreplaced by byby artificial artificialartificial vision visionvision using usingusing cameras camerascameras and andand monitors, monitors,monitors, some somesome doors doorsdoors and andand hatches hatcheshatches will willalwayswill alwaysalways be required. bebe required.required. However, However,However, these thesethese problems problemsproblems may maymay be overcome bebe overcomeovercome by minimizing byby minimizingminimizing gaps gapsgaps and and stepsand stepssteps and andinstallationand installationinstallation of suction ofof suctionsuction skin on skinskin door onon panels.doordoor panels.panels. As the AsAs aim thethe of aimaim this of paperof thisthis ispaperpaper an all-out isis anan effortall-outall-out to efforteffort reduce toto drag reducereduce by draglaminardrag byby laminar flowlaminar control, flowflow thecontrol,control, potential thethe potentialpotential of fuselage ofof fuselagefuselage suction is suctionsuction investigated isis investigatedinvestigated on an isolated onon anan simplifiedisolatedisolated simplifiedsimplified fuselage fuselagebody.fuselageThe body.body. geometry TheThe geometrygeometrywas derived waswas from derivedderived the LamAiR fromfrom thethe [ LamAiR20LamAiR] configuration [20][20] configurationconfiguration mentioned mentionedmentioned in Section 2.3inin Section,Section which 2.3,is2.3, very whichwhich similar isis veryvery to similarsimilar the fuselage toto thethe fuselagefuselage shape of shapeshape the Airbus ofof thethe AirbusAirbus A320. A320.A320. The waist TheThe waistwaist line was lineline was projectedwas projectedprojected onto ontoonto the theXY-Plane,the XY-Plane,XY-Plane, and and anand axisymmetric anan axisymmetricaxisymmetric body bodybody of revolution ofof revolurevolu wastiontion generated waswas generatedgenerated fromthe fromfrom resulting thethe resultingresulting 2D-curve, 2D-curve,2D-curve, shown showninshown Figure inin 20FigureFigure. 20.20.

Figure 20. LamAiR Fuselage Geometry. Figure 20. LamAiRLamAiR Fuselage Fuselage Geometry. Geometry.

ForFor thethe designdesign ofof thethe fuselagefuselage suctionsuction distribudistribution,tion, aa sliceslice throughthrough thethe fuselagefuselage geometrygeometry waswas For the design of the fuselage suction distribution, a slice through the fuselage geometry was importedimported inin XFOILSUC.XFOILSUC. TheThe thicknessthickness waswas scaledscaled downdown toto obtainobtain aa pressurepressure distributiondistribution withwith similarsimilar imported in XFOILSUC. The thickness was scaled down to obtain a pressure distribution with similar gradientsgradients toto thosethose obtainedobtained fromfrom 3D3D RANSRANS computcomputations.ations. MachMach andand ReynoldsReynolds numbernumber werewere setset toto gradients to those obtained from 3D RANS computations. Mach and Reynolds number were set to cruisecruise conditionsconditions forfor anan A320A320 classclass aircraft,aircraft, whichwhich resultedresulted inin aa ReynoldsReynolds numbernumber basedbased onon fuselagefuselage cruise conditions for an A320 class aircraft, which resulted in a Reynolds number based on fuselage lengthlength ofof 166166 millionmillion andand aa MachMach numbernumber ofof 0.8.0.8. WiWithoutthout suction,suction, naturalnatural trantransitionsition waswas predictedpredicted byby XFOILSUCXFOILSUC atat 10%10% relativerelative length.length. SuctionSuction waswas appliedapplied fromfrom 7%7% toto 70%70% fuselagefuselage length,length, upup toto thethe endend

Energies 2018, 11, 252 20 of 28

Energieslength 2018 of 166, 11, million 252 and a Mach number of 0.8. Without suction, natural transition was predicted20 of by27 XFOILSUC at 10% relative length. Suction was applied from 7% to 70% fuselage length, up to the end of the cylindrical part, and adjusted manually until the transition location was significantly signiﬁcantly aft of the end of the suction region, beyond 80%.

7.2. Numerical Setup To analyze the influence influence of suction on the drag drag co coefficientefficient of a a fuselage, fuselage, three-dimensional three-dimensional RANS simulations were performed on a quarter section of the fuselage body geometry described above, above, using the DLR TAU-code TAU-code [36,37], [36,37], release 2016.2.0. Tw Twoo computational grids of different different fineness fineness were ◦ ◦ created, bothboth withwith symmetrysymmetry planes planes at at 0 0°and and 90 90°in in circumferential circumferential direction. direction. The The boundary boundary layer layer and andthe wakethe wake downstream downstream of the of fuselage the fuselage were resolvedwere reso withlved structured with structured hexahedral hexahedral cells, while cells, the while farfield the farfieldgrid was grid made was up made from up unstructured from unstructured prism and prism tetrahedral and tetrahedral cells. Within cells. theWithin structured the structured part, the part, fine thegrid fine has 1005grid pointshas 1005 in streamwisepoints in streamwise direction, 45 direction, points in 45 circumferential points in circumferential direction and direction81 points andin wall 81 + pointsnormal in direction. wall normal The direction. y -values The are y below+-values 1 along are below the nose 1 along and the the nose cylindrical and the part cylindrical of the fuselage part of. + theOnly fuselage. at the conicalOnly at partthe conical at the end,part at y the-values end, slightlyy+-values larger slightly than larger 1 occur. than 1The occur. boundary The boundarylayer is layerresolved is resolved with more with than more 30 than cells 30 globally. cells globally. In total, In the total, fine the grid fine consists grid consists of 7.75 of million 7.75 milliongrid points. grid Topoints. check Tothe check influence the influence of the grid of resolution,the grid resolution, a coarse grida coarse was grid derived was from derived the finefrom grid, the withfine grid, grid pointwith grid spacing point doubled spacing in doubled all three in dimensions all three dimensions in the structured in the partstructured of the gridpart (500of the× grid23 × (50041 points × 23 ×). The41 points). unstructured The unstructured part was adapted part was accordingly, adapted resultingaccordingly, in 1.24 resulting million in grid 1.24 points million in total.grid Figurepoints 21in total.gives anFigure impression 21 gives of thean structureimpression of theof the computational structure of grids. the computational For a better illustration, grids. For the a coarsebetter illustration,grid is shown. the coarse grid is shown.

Figure 21. Details of the grid structure (coarse grid displayed) at the nose (left) and at the rear end Figure 21. Details of the grid structure (coarse grid displayed) at the nose (left) and at the rear (right). end (right).

The TAU-code solves the Reynolds-averaged Navier-Stokes equations. Turbulence was modeledThe TAU-codeby the Menter-SST solves the Reynolds-averagededdy viscosity model Navier-Stokes [38,39], which equations. is widely-used Turbulence for was industrial modeled applications.by the Menter-SST For time eddy stepping, viscosity the modelimplicit [38 lower-upper,39], which issymmetric widely-used Gauss-Seidel for industrial (LU-SGS) applications. scheme wasFor timeused.stepping, First, fully the turbulent implicit lower-upper simulations symmetricwere performed, Gauss-Seidel aimed (LU-SGS)at the identification scheme was of used. a suitable First, numericalfully turbulent setup simulations for the simulations were performed, including aimed su atction the identificationand providing of athe suitable drag numericalcoefficient setupof a turbulentfor the simulations fuselage as including a baseline suction for the andevaluation providing of LFC. the dragFinally, coefficient two simulations of a turbulent including fuselage suction as showa baseline the potential for the evaluation of drag reduction of LFC. due Finally, to the two delay simulations of transition, including preservi suctionng the showlaminar the boundary potential layer.of drag For reduction these simulations, due to the the delay suction of transition, velocity distribution preserving thedesigned laminar in boundaryXFOILSUC layer. was projected For these ontosimulations, the RANS the suctionsurface velocitygrid. For distribution the simulations designed presented in XFOILSUC here, the was transition projected position onto the was RANS set manuallysurface grid. slightly For thedownstream simulations of the presented area where here, suction the transition was applied, position assuming was set the manually boundary slightly layer todownstream be laminar offrom the the area nose where of the suction fuselage was up applied, to the assumingtransition theposition. boundary Table layer 1 gives to be an laminar overview from of the cases nose ofsimulated, the fuselage including up to the the transition results of position.the drag calculations. Table1 gives For an overview the drag coefficients, of the cases simulated,the wetted surfaceincluding was the used results as ofthe the reference drag calculations. area. Drag For is predicted the drag coefficients, similarly on the both wetted computational surface was grids, used indicatingas the reference that the area. boundary Drag is predictedlayer development similarly onis bothwell computationalresolved on the grids, fine indicatinggrid. To ensure that the a sufficientboundary resolution layer development of the suction is well distribution, resolved the on thesimulations fine grid. including To ensure suction a sufficient were performed resolution on of the suctionfine grid. distribution, For the fully the turbulent simulations case, including a drag estimation suction were as performedin Subsection the7.4, fineusing grid. handbook For the methods for preliminary aircraft design, confirmed the drag results for the turbulent case.

Energies 2018, 11, 252 21 of 27

Table 1. RANS Calculations for Fuselage Drag.

Suction Transition Drag Area Flow Condition Grid , (x/c)Start–(x/c)End (x/c)Transition (m²) Fine - 0.0 0.845 0.00204 Turbulent Coarse - 0.0 0.849 0.00205 Energies 2018, 11, 252 21 of 28 Suction 53% Fine 0.07–0.53 0.550 0.449 0.00108 Suction 70% Fine 0.07–0.70 0.776 0.235 0.00057 fully turbulent case, a drag estimation as in Subsection 7.4, using handbook methods for preliminary 7.3. Resultsaircraft design, conﬁrmed the drag results for the turbulent case.

In Figure 22, the boundaryTable layer 1. parametersRANS Calculations of the for 2D Fuselage XFOILSUC Drag. and 3D TAU simulations are shown in comparison. Note that the skin friction coefficient is identical for the part were suction is applied, whichFlow is in accordance with boundarySuction layer Transitiontheory and provesDrag that the data import from Grid 2 CDFF, swet Condition (x/c)Start–(x/c)End (x/c)Transition Area (m ) XFOIL to TAU was implemented correctly. Also shown is the development of the shape factor , Fine - 0.0 0.845 0.00204 which is consistentlyTurbulent smaller for the 3D case than for the 2D data. At the same time, transition is Coarse - 0.0 0.849 0.00205 assumed earlier than predicted by XFOILSUC, which means that the calculated drag reduction can Suction 53% Fine 0.07–0.53 0.550 0.449 0.00108 be considered conservative. Suction 70% Fine 0.07–0.70 0.776 0.235 0.00057 To show the effect of suction length, a case with reduced fuselage suction length was also calculated, with the suction velocity set to zero aft of 53% relative fuselage length. The reduction of 7.3. Results total drag for both suction lengths is shown in Figure 23a, while the relative drag reduction and the ratio of pumpIn Figure drag 22 ,to the drag boundary reduction layer parametersare shown of in the Figure 2D XFOILSUC 23b. Note and that 3D the TAU difference simulations between are shown in comparison. Note that the skin friction coefﬁcient c f is identical for the part were suction pressure and friction drag () and total drag () is almost neglibile. The fuselage pressure is applied, which is in accordance with boundary layer theory and proves that the data import from distribution is very flat, and only very little suction mass flow is needed to ensure a laminar boundary XFOIL to TAU was implemented correctly. Also shown is the development of the shape factor H , layer. In addition, the compressor power for the static pressure increase is also small, and 12so the which is consistently smaller for the 3D case than for the 2D data. At the same time, transition is pumpassumed drag is earlier almost than completely predicted compensated by XFOILSUC, by which the meansjet thrust. that theThe calculated total fuselage drag reduction drag reduces can be by a factorconsidered of four, and conservative. the necessary compressor power is only around 7% of the saved drag.

FigureFigure 22. 22. BLBL and and Suction Suction Parameters Parameters for for Fuselage. Fuselage.

Energies 2018, 11, 252 22 of 28

To show the effect of suction length, a case with reduced fuselage suction length was also calculated, with the suction velocity set to zero aft of 53% relative fuselage length. The reduction of total drag for both suction lengths is shown in Figure 23a, while the relative drag reduction and the ratio of pump drag to drag reduction are shown in Figure 23b. Note that the difference between pressure and friction drag (CDp f ) and total drag (CDtot) is almost neglibile. The fuselage pressure distribution is very ﬂat, and only very little suction mass ﬂow is needed to ensure a laminar boundary layer. In addition, the compressor power for the static pressure increase is also small, and so the pump drag is almost completely compensated by the jet thrust. The total fuselage drag reduces by a factor of four, and the necessary compressor power is only around 7% of the saved drag. Energies 2018, 11, 252 22 of 27

(a) (b)

Figure 23. Results for Fuselage Drag with and without suction; DLR LamAiR; = 1666 ∙ Figure 23. Results for Fuselage Drag with and without suction; DLR LamAiR; Re f us = 166·10 ; 10Ma; = 0.8; = 0.8;α = 0◦ =. 0°.

7.4. Estimation of Fuselage Drag by Handbook Methods 7.4. Estimation of Fuselage Drag by Handbook Methods To validate the RANS results for the turbulent reference case, the drag was was estimated using To validate the RANS results for the turbulent reference case, the drag was was estimated using handbook methods. Gur [40] presents a methodology commonly used for preliminary aircraft design, handbook methods. Gur [40] presents a methodology commonly used for preliminary aircraft design, were the friction and form drag coefficient of each component is the product of the equivalent flat were the friction and form drag coefficient of each component is the product of the equivalent flat plate skin friction coefficient, a form factor, and the ratio of the components’ wetted surface area to plate skin friction coefficient, a form factor, and the ratio of the components’ wetted surface area to the the reference area, which is usually the wing area: reference area, which is usually the wing area: = ∙∙ Swet CDFF = CFe·FF· Sre f The form factor accounts for the pressure drag due to boundary layer effects. Because the wing area isThe undefined form factor for accountsan isolated for fuselage, the pressure we use drag the due wetted to boundary surface as layer reference effects. area. Because This results the wing in

area⁄ is undeﬁned=1, and for consequently: an isolated fuselage, we use the wetted surface as reference area. This results in Swet/Sre f = 1, and consequently: , = ∙ CDFF, swet = CFe·FF For the fuselage form factor, an estimate is given by Gur and Raymer [40,41] as a function of the For the fuselage form factor, an estimate is given by Gur and Raymer [40,41] as a function of the fuselage slenderness: fuselage slenderness: L f us 60 60 Df us ==1++ + FFf us 1 3 400 L f us 400 D f us For the equivalent skin friction coefficent, Raymer [41] recommends a formula taken from DATCOM [42]: 0.455 = . . () ∙ 1 + 0.144 ∙

The main parameters and the results of the handbook and numerical calculations are summarized in Table 2. The drag area values estimated by handbook methods and integrated from CFD data differ by less than 0.5%, which indicates that the CFD method is sufficiently accurate to predict the potential drag reduction by laminar flow control.

Energies 2018, 11, 252 23 of 28

For the equivalent skin friction coefficent, Raymer [41] recommends a formula taken from DATCOM [42]: 0.455 CFe = 2.58 2 0.65 [log10(Re)] ·[1 + 0.144·Ma ] The main parameters and the results of the handbook and numerical calculations are summarized in Table2. The drag area values estimated by handbook methods and integrated from CFD data differ by less than 0.5%, which indicates that the CFD method is sufficiently accurate to predict the potential drag reduction by laminar flow control.

Table 2. Comparison of Drag Results from RANS and Handbook Methods.

Geometry

Lfus 38 m Dfus 4 m 2 Swet 414.5 m Flight Conditions

HCR 13,000 m MaCR 0.8 ReFUS 166,240,000 Handbook

FFfus 1.094 CFe 0.00187 CDFF,swet,HB 0.00205 RANS Data

CDFF,swet,CFD 0.00204 (Upwind, ﬁne grid)

8. Drag Reduction on AC Level For a medium-range jet with top level aircraft requirements similar to an A320, a reference configuration and an optimized design, taking into account LFC on wings, tails and fuselage, were evaluated using methods for preliminary aircraft design [43]. To simplify the process, the suction system was not modeled in detail. The power of the suction pump and the jet thrust of the expelled suction air were included in the total drag coefficients, as explained in Section 4.3. No weight penalty for the suction system is considered, assuming that for future aircraft, new weight-saving structural technologies will compensate for the additional components. The relative drag reductions for wing sections were determined for 2D profiles at Ma = 0.7, which corresponds to a medium range jet aircraft flying at Ma = 0.8, with a sweep angle in the order of 20◦ to 30◦. In Section 2.3, it was shown that sweep induced transition can be avoided with forward swept wings and moderate sweep angles, and the suction system only has to suppress the amplification of TS instabilities on the rear part of the wing. The relative effect of viscous drag reduction on absolute drag is illustrated in Figure 24. The first column shows the drag breakdown for the reference aircraft. The major contributors are induced drag and viscous drag on wings, tail surfaces and the fuselage. Smaller portions such as nacelle drag, interference and wave drag are summed up in the miscellaneous fraction. For wings and tails, a drag reduction of 56% is applied. The assumptions are a relative reduction of profile drag by 68%, as shown in Section5, from which a deduction was made to account for junction areas which can not be laminarized. For the fuselage, a reduction of 65% was assumed, based on the 72% reduction for the isolated fuselage shown in Section7, and again making a deduction for junction areas. The immediate result is a reduction in viscous drag by 60%, and in consequence a reduction of absolute AC drag by 30%. This changes the balance of viscous and induced drag, recommending a re-design of the wing layout, with the optimum aspect ratio increased from 9.5 to 15. The intermediate result is an aircraft of roughly the same size, but with significantly increased range. Resizing for the original mission leads Energies 2018, 11, 252 24 of 28 to a reduction in wing area from 124 to 109 m2, an according change of tail plane area, plus a reduction in MTOW, from 75 down to 66 tons. This results in further reduction of total drag, even for fractions of the drag not directly affected by laminar flow control or wing design. Due to the reduced mass, the energy required for climb to altitude and take-off acceleration is also reduced. The resulting cruise drag, represented by the right column in Figure 24, is only 50% of the reference value. Mission fuel

Energiesburn is 2018 reduced, 11, 252 from 15.9 to 8.3 tons, a remarkable reduction by 47%. 24 of 27

Figure 24. Viscous Drag Reduction and Effect on Total Aircraft Drag. Figure 24. Viscous Drag Reduction and Effect on Total Aircraft Drag. 9. Conclusions and Outlook 9. Conclusions and Outlook Laminar flow control by boundary layer suction holds potential for great reductions in drag. For Laminar flow control by boundary layer suction holds potential for great reductions in drag. wing sections, the lowest drag values were achieved by applying LFC on the rear part of airfoils For wing sections, the lowest drag values were achieved by applying LFC on the rear part of airfoils originally designed for NLF, to further delay the TS-induced transition towards the trailing edge. originally designed for NLF, to further delay the TS-induced transition towards the trailing edge. Compared with today’s turbulent supercritical airfoils, profile drag can be reduced by more than Compared with today’s turbulent supercritical airfoils, profile drag can be reduced by more than 70%. 70%. Application of suction to supercritical airfoils results in slightly higher total drag values, but Application of suction to supercritical airfoils results in slightly higher total drag values, but with with better adaptability to changes in lift coefficient. It seems reasonable to assume the same drag better adaptability to changes in lift coefficient. It seems reasonable to assume the same drag reduction reduction potential also for tail surfaces. Wing sections with increased thickness, designed for use on potential also for tail surfaces. Wing sections with increased thickness, designed for use on future future blended wing body configurations, also showed significant potential for drag reduction with blended wing body configurations, also showed significant potential for drag reduction with the the application of BL suction. In the future, integrated design of airfoil shape and suction distribution application of BL suction. In the future, integrated design of airfoil shape and suction distribution may may lead to even lower total drag coefficients. lead to even lower total drag coefficients. Next to wings and tails, the fuselage is a great contributor to viscous drag. For a simplified Next to wings and tails, the fuselage is a great contributor to viscous drag. For a simplified generic generic contour, drag could be reduced to one fourth of the fully turbulent fuselage by applying contour, drag could be reduced to one fourth of the fully turbulent fuselage by applying suction from suction from near the nose to the beginning of the tailcone. near the nose to the beginning of the tailcone. The technical assumptions made in this study are believed to be rather conservative. Chordwise The technical assumptions made in this study are believed to be rather conservative. Chordwise extent of suction on wing sections was limited to 80%, leaving room for full span trailing edge extent of suction on wing sections was limited to 80%, leaving room for full span trailing edge movables. For the fuselage, suction up to 70% of fuselage length was investigated. The efficiency of movables. For the fuselage, suction up to 70% of fuselage length was investigated. The efficiency of the suction compressor was set to = 70%, based on flight tests performed in the 1960s, and none the suction compressor was set to η = 70%, based on flight tests performed in the 1960s, and none of of the investigated geometries was coptimized or modified for the application of suction. the investigated geometries was optimized or modified for the application of suction. By using the knowledge gained by the DLR during research on NLF and HLFC for the integrated By using the knowledge gained by the DLR during research on NLF and HLFC for the integrated design of wing and fuselage contours and suction systems, and application of state-of-the art suction design of wing and fuselage contours and suction systems, and application of state-of-the art suction compressor technology including electric drive systems, fuel burn reductions of 47%, as claimed in compressor technology including electric drive systems, fuel burn reductions of 47%, as claimed in Section 8, appear feasible. Section8, appear feasible. Apart from further development of methods and numerical tools for aerodynamic design and Apart from further development of methods and numerical tools for aerodynamic design and analysis, a number of additional issues need to be addressed and solved in order to achieve the analysis, a number of additional issues need to be addressed and solved in order to achieve the necessary technological maturity: necessary technological maturity: • Structural concepts for porous skins with high surface quality, good mechanical robustness, low weight and long term stability, possibly employing multiple-shell designs with function integration into the sandwich core, • Integration of BL suction design concepts with design of propulsion integration, high-lift and anti-ice system, • Treatment of surface imperfections such as door gaps, possibly by gap suction or downstream re-laminarization.

Energies 2018, 11, 252 25 of 28

• Structural concepts for porous skins with high surface quality, good mechanical robustness, low weight and long term stability, possibly employing multiple-shell designs with function integration into the sandwich core, • Integration of BL suction design concepts with design of propulsion integration, high-lift and anti-ice system, • Treatment of surface imperfections such as door gaps, possibly by gap suction or downstream re-laminarization.

Acknowledgments: We would like to acknowledge the support of the Ministry for Science and Culture of Lower Saxony, Germany (Grant No. VWZN3177) for funding the research project “Energy System Transformation in Aviation” in the initiative “Niedersächsisches Vorab”. We also thank the North-German Supercomputing Alliance (HLRN) for providing computational resources. Finally, the authors are grateful for the many helpful and productive discussions and the cooperation with other members within the Energy System Transition in Aviation project team. Special thanks go to the project leader, Jens Friedrichs, for defining and clarifying the concepts. Author Contributions: Nils Beck performed the XFOILSUC calculations, coded the auxiliary MatLab software, analyzed the airfoil and fuselage data with respect to suction requirements and total drag, estimated the fuselage drag by handbook methods, and was responsible for the overall editing of the paper. Tim Landa prepared and executed the RANS simulations of the isolated fuselage and wrote the respective part of the paper. Arne Seitz compiled the review of laminar flow research presented in chapter 2. Loek Boermans contributed the XFOILSUC software, gave valuable advice on it's application, and reviewed the chapters on XFOIL, suction design and total drag. Yaolong Liu analyzed the effect of the drag reduction on overall aircraft level presented in chapter 8. Rolf Radespiel initiated and coordinated the work, gave valuable guidance, reviewed the complete paper and contributed to the introductory and outlook chapters. Conflicts of Interest: The authors declare no conflict of interest.

Nomenclature

CD; CDtot Drag coefficient; total drag coefficient including suction system CDp; CD f ; CDp f ; CDc; CDi CD due to: pressure; friction; profile drag; compressor power; lift induced CDw Wake drag coefficient, calculated from momentum deficit CL; α Lift coefficient; angle of attack CP Pressure coefficient CP,pl; CP,skin; CP,duct CP of: plenum pressure; skin pressure loss; duct pressure loss Ma Mach number U∞; p∞; $∞; T∞; q∞ Free stream velocity; pressure; density; temperature; dynamic pressure pw; $w; Tw Wall pressure; density; temperature ppl Suction plenum pressure Ujet Jet velocity of suction air exhaust Sre f ; Swet Reference areas; wetted surface area Re; Re f us Reynolds number based on: airfoil chord length; fuselage length v0 v0; U Absolut; nondimensional suction velocity . ∞ Q; m Volume flow; mass flow of suction system CQ; CQm Suction coefficient based on volume flow; mass flow x, z; s; l Cartesian airfoil coordinates; airfoil contour coordinate; chord length ϕ; ζ Permeability; flow resistance of suction skin H12; N Shape factor; growth factor NT; NLF Natural transition; natural laminar flow LFC; HLFC Laminar flow control; hybrid laminar flow control TS; CF; ATL Tollmien-Schlichting; cross flow; attachment line TLAR Top Level Aircraft Requirements DLR Deutsches Zentrum für Luft- und Raumfahr (German Aerospace Research Center) TU BS Technische Universität Braunschweig (Technical University Braunschweig) TU Delft Delft University of Technology Programming language for technical and mathematical problems by MathWorks, MatLab 3 Apple Hill Drive, Natick, Massachusetts, 01760 USA Energies 2018, 11, 252 26 of 28

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