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Flight Tests with a Natural Laminar Flow Glove on a Transport Aircraft.
Abstract
As part of a national laminar low research program, flight tests with a glove on the right-hand wing of the DLR research aircraft VFW 614/ATTAS were carried out. The glove was especially designed to obtain three types of boundary layer instability: Tollmien-Schlichting instability (TSI), crossflow instability (CFI) and attachment line transition (ALT). In all tests, covering the Mach number range from 0.35 to 0.7, the complete transition line has been made visible by infrared image technique. Pressure distributions were measured and and used in stability analyses. Thus prediction by stability calculations at large Reynolds numbers has been obtained. Initial disturbance investigations provided information on the influence of engine noise and geometric steps on transition location.
... Laminar technology, as one of the most promising means, has attracted much attentions and investigations, due to its great potential of reducing friction drag and associated viscous pressure drag. Research activities aiming at delaying laminar-turbulent transition mainly cover natural laminar flow (NLF) [1][2] , laminar flow control [3][4] (LFC), and hybrid laminar flow [5][6] (HLF), among which the NLF technology is the basis of the other laminar flow technologies. ...
... Considering differences in flow quality and geometry models, several wind tunnel campaigns has been executed to determine the N factors for specific wind tunnels. A few flight tests [5][6] were carried out as well to give a reference of critical N factors when stability analysis is performed for configurations in cruise condition. In this paper, two configurations are investigated to validate the current approach with N factors correlated with wind-tunnel tests. ...
... based on mean aerodynamic chord. The thresholds of N factors for TS and CF instabilities in this work are determined as N TS =10.5 and N CF =10 on basis of ATTAS flight tests [5][6] in ideal cruise condition which is on assumptions of low turbulence atmosphere and certain level of surface roughness. ...
This paper aims to develop an efficient global optimization method for design of transonic natural-laminar-flow (NLF) airfoils and wings, based on high-fidelity computational fluid dynamics (CFD) solver. The CFD solver features functionality of automatic transition prediction, by coupling Reynolds-averaged Navier-Stokes (RANS) equations with the linear-stability-theory-based dual e N method for Tollmien-Schlichting and crossflow instabilities. An A320-sized transonic NLF wing with a laminar supercritical airfoil is designed for cruise condition at Mach=0.74, Re=20 million, C L =0.515. In order to further improve the cruise efficiency, this NLF wing is optimized at higher Mach number of 0.75 via an in-house surrogate-based optimizer. The optimization is formulated as a drag minimization problem with constraints on lift, pitching moment and geometric thickness. Through only 130 CFD evaluations, 12.1 counts drag reduction is obtained, while all constraints are strictly satisfied. Further study shows that the drag reduction is contributed by both of shock-wave weakening and laminar-flow extension. On suction side, the favorable pressure gradient is maintained while shock wave is weakened; on pressure side, the cross-flow (CF) instability is effectively suppressed and thereby the laminar flow region is dramatically extended. The improvement of aerodynamic performance is observed not only at design point but also over a certain range of off-design lift coefficients. Nomenclature C = chord length C L = lift coefficient C D = drag coefficient C Df = friction drag coefficient C M = moment coefficient * p C = critical pressure coefficient C p = pressure coefficient D = drag force L = lift force Ma = Mach number N = amplification factor of disturbance Re = Reynolds number t/c = relative thickness W = cross-flow component of velocity inside boundary-layer 1 Professor, P.O. Box 754, 2 X = X coordinate X des = design variables X tr = transition location α = Angle of attack β = wave number Λ = sweep angle δ * = displacement thickness θ = momentum thickness
... One of the techniques of artificial laminarization is to suck the boundary layer off parts of the plane's glider. The effectiveness of a given method was confirmed by experimental studies in wind tunnels [19,20] and flight tests [21][22][23]. Paper [19] shows a decrease in the effectiveness of sucking off the boundary layer with an increase in the sweep of the wing model. Study [20] demonstrates a decrease in the effectiveness of sucking off the boundary layer with an increase in the Mach and Reynolds numbers in wind tunnel tests. ...
... An overview of AC flight tests with the laminar flow on the surface of the glider elements is given in [21]. Work [22] reports a study into the effect of sucking off the boundary layer in the region where the aircraft wing and fuselage are connected, where the flow is three-dimensional. Under certain conditions at sucking off the boundary layer, the authors observed completely laminar zones of the flow, which, under normal conditions, were turbulent. ...
... Over the last decades extensive NLF and HLFC research has been performed worldwide on various aspects, e.g., imperfections of contours, systems integration, and this has resulted in significant progress [12][13][14][15][16]. Several projects have been conducted for example, at DLR (German Aerospace Center), Airbus, and the Technical University of Braunschweig, including theoretical design work as well as flight experiments [17,18]. Currently large scale flight tests are ongoing in Europe [19,20]. ...
... For the first stage system study, we define an LFC factor to reflect the percentage of the total aircraft surfaces that can keep in laminar flow status through laminar flow control technologies. progress [12][13][14][15][16]. Several projects have been conducted for example, at DLR (German Aerospace Center), Airbus, and the Technical University of Braunschweig, including theoretical design work as well as flight experiments [17,18]. Currently large scale flight tests are ongoing in Europe [19,20]. ...
It is always a strong motivation for aeronautic researchers and engineers to reduce the aircraft emissions or even to achieve emission-free air transport. In this paper, the impacts of different game-changing technologies together on the reduction of aircraft fuel consumption and emissions are studied. In particular, a general tool has been developed for the technology assessment, integration and also for the overall aircraft multidisciplinary design optimization. The validity and robustness of the tool has been verified through comparative and sensitivity studies. The overall aircraft level technology assessment and optimization showed that promising fuel efficiency improvements are possible. Though, additional strategies are required to reach the aviation emission reduction goals for short and medium range configurations.
... Throughout the past couple of decades, much systematic numerical and experimental research has been performed to tackle the basic problems involved in laminar-turbulent transition on swept wings in transonic flows (e.g. Ref. [1][2][3][4]). This has led to the achievement of significant progress in various passive and active control mechanisms to obtain large portions of laminar boundary-layer flow over the wing surface, like boundary layer suction (Ref. ...
Natural laminar ow technology is a highly promising option to reduce the fuel consumption of next generation "Ultra- Green" commercial transport aircraft. By passively maintaining a laminar ow over the largest possible area of both wings, the overall friction drag of these future commercial transport aircraft is reduced and their aerodynamic performance increased. To further raise the Technology Readiness Level of this laminar technology in aircraft design, joined wind tunnel tests and numerical research has been performed on a newly-designed forward-swept wing con guration in the LuFo VI-1 research initiative ECOWING, which lasted from 2020 to 2021, and its LuFo VI-2 successor project ULTIMATE that commenced in 2022 and continues until 2026. This paper focuses on the location and movement of the boundary-layer transition on the upper surface of the innovative forward-swept wing con guration that were obtained with Temperature- Sensitive Paint (TSP) in the rst performance test of this wing model in the European Transonic Windtunnel (ETW). The test conditions included many combinations of Reynolds and Mach numbers - both design and o -design values - over a wide range of static pitch angles of the wing, as well as various quasi-steady pitch-sweeps at selected Mach and Reynolds numbers. Three di erent methods to achieve a necessary and su ciently large temperature di erence between the model surface, coated with TSP, and the ow have furthermore been tested for comparison of their applicability and e ciency for TSP under cryogenic test conditions. With each of these three methods, the transition pattern could be captured successfully with only minor deviations among the three. In uences of the test parameters on spanwise distribution of the transition location in the surface boundary layer could be clearly visualised.
... For 2 of 17 the particle image velocimetry (PIV) technique [15], Medina et al. [16] investigated the flow mechanism of the pulsating transition jet. Horstmann et al. [17,18] performed transitional measurements on the boundary layer of airfoils with infrared image technology. Holmes et al. [19] confirmed that shear-sensitive liquid crystal technology can be applied well under different wind-tunnel test conditions of airfoils, and Mee et al. [20] accurately measured the boundary layer transition with this technology. ...
Advanced measurement technology on boundary layer transition is an effective means to study the flow mechanism and the performance of the cascade. In this research, to investigate the boundary layer transition of transonic VKI-RG turbine cascade, a noncontact measurement technique named temperature-sensitive paint (TSP) that is capable of quantitatively measuring the surface temperature of a model is used. Under the conditions of outlet Mach number (Ma2) = 0.4, 1.03, and 1.2, the transitional location of a cascade is accurately measured by TSP technique. Moreover, a numerical study on the transitional location of a planar cascade was conducted with the conditions of Mach number (Ma) and attack angle (α). The numerical results show that the transitional location moved forward with the increase in Ma2 and α at Ma2 < 1; otherwise, the transitional location moved backward with increasing Ma2 (Ma2 > 1). Lastly, a strong adverse pressure gradient was formed behind the shock wave, which led to the beginning of the transition.
... Some were also applied to components in flight test aircraft. Flight tests of NLF airfoils have shown laminar flow over 20-45% of the chord [30,31,32,33]. Numerical simulations predict higher achievable ranges up to 60% [34,35,3]. ...
Aircraft designs concerning (hybrid-)electric aircraft often have large discrepancies in the underlying assumptions
regarding highly influential design parameters. These mostly concern assumptions regarding the predicted
performance of the electric energy provider such as batteries and fuel cells, but also for electric motors
and other novel airframe technologies. Having these variations makes it difficult to assess the viability of a proposed
novel aircraft design. Furthermore, it makes a comparison between aircraft for comparable missions or
performance difficult, as differences in the assumed performance of these subsystems can yield a large impact
on the overall design feasibility and performance. The aim of this research is to investigate current trends in
the market for a range of highly influential design parameter for futuristic (hybrid-)electric aircraft designs and
create a forecast for the near future. This will then be used to create a credibility-distribution for performance
assumptions that can be applied to a full aircraft design. Using this criterion, the credibility of the futuristic
aircraft can be assessed. An exemplary application is shown in this paper on a newly developed hybrid-electric
aircraft design.
... The value of N crit is experimentally validated and differs with receptivity factors. N crit value of 13 is used in this study for flight conditions based on the in-flight experiments conducted with DLR flying testbed ATTAS (Advanced Technologies Testing Aircraft System) [30,31]. N-factor identification was performed using a specially designed wing glove with modified sections. ...
The German research Cluster of Excellence SE2A (Sustainable and Energy Efficient Aviation) is investigating different technologies to be implemented in the following decades, to achieve more efficient air transportation. This paper studies the Hybrid Laminar Flow Control (HLFC) using boundary layer suction for drag reduction, combined with other technologies for load and structural weight reduction and a novel full-electric propulsion system. A multidisciplinary design optimisation framework is presented, enabling physics-based analysis and optimisation of a fully electric aircraft wing equipped with HLFC technologies and load alleviation, and new structures and materials. The main focus is on simulation and optimisation of the boundary layer suction and its influence on wing design and optimisation. A quasi three-dimensional aerodynamic analysis is used for drag estimation of the wing. The tool executes the aerofoil analysis using XFOILSUC, which provides accurate drag estimation through boundary layer suction. The optimisation is based on a genetic algorithm for maximum take-off weight (MTOW) minimisation. The optimisation results show that the active flow control applied on the optimised geometry results in more than 45% reduction in aircraft drag coefficient, compared to the same geometry without HLFC technology. The power absorbed for the HLFC suction system implies a battery mass variation lower than 2%, considering the designed range as top-level requirement (TLR).
... Aerodynamics is inherently the key driver for development of laminar technology on transport aircraft. Consequently, throughout the past decades the basic problems regarding laminar to turbulent transition on transonic swept wings have been investigated thoroughly and flow control methodologies capable of delivering large portions of laminar boundary layer flow are well understood [2]. However, so far there are only a few fully developed laminar aircraft design studies available [3] that exceed conceptual or preliminary design level. ...
In order to further raise the Technology Readiness Level (TRL) of laminar technologies in aircraft design the German Aerospace Center DLR conducted an internal project called TuLam (Toughen up Laminar Technology), which lasted from 2014 to 2017. In the course of the project two technology paths were pursued, namely Natural Laminar Flow (NLF) and Hybrid Laminar Flow Control (HLFC). Within the frame of the NLF path a short and medium range transport aircraft with forward swept laminar wing was designed. The present paper is focused on the aerodynamic design of the forward swept wing in cruise flight. As a special feature in comparison with previous designs of transonic laminar flow wings a trailing edge flap of 10% chord depth is employed to allow for an adaptation of the laminar bucket to off-design conditions. The resulting wing was assessed on overall aircraft level with respect to its fuel reduction potential, whereby the CSR-01 configuration, essentially a re-design of the Airbus A320-200, was used as a reference.
... Atkin and Courtenay [3] found that HLFC decreased the total drag by 1.6%, whereas it dropped to 1% when taking the pumping system into consideration. The research of NLF technology dates back to the 1940s, when NACA 6-series airfoils [4] were first developed, and the feasibility of it has been revealed by some early works [5][6][7][8]. At present, NLF technology is still an active research topic. ...
This paper presents a combined experimental and computational study of a low-sweep transonic natural laminar flow (NLF) wing with shock-control bumps (SCBs). A transonic NLF wing with a relatively low sweep angle of 20 deg was chosen for this study. To avoid the complexity of the flow introduced by perforated/slotted walls commonly used for transonic wind-tunnel tests for reducing the wall interference, both experimental tests and computational simulations were conducted with solid wind-tunnel wall conditions. This allows for like-to-like validation of the computational simulation. Optimization of the shock-control bumps was first conducted to design the wind-tunnel test model with bumps. Two critical parameters of the three-dimensional SCBs for shock control (i.e., bump crest position and bump height) were optimized in terms of total drag reduction at the given design point in the wind tunnel. We show that the strong shock wave on the low-sweep NLF wing can be effective controlled by well-designed SCBs deployed along the wing span. The optimized SCBs result in 18.5% pressure drag reduction with 5% viscous drag penalty, and the SCBs also bring some benefits at off-design conditions. The wind-tunnel tests include pressure measurement, particle image velocimetry, and temperature-sensitive paint to provide detailed insight into the shock-control flowfield and to validate the computational simulations. Comparisons include surface pressure profile, velocity distribution, and transition location.
... The studies associated with NLF technologies find its origin from the 1940s when the NACA 6-series airfoils [1] were first developed. Thereafter, a number of representative design works [2][3][4][5] and experimental works [6][7][8][9] have revealed its feasibility and potential. A few real-world applications to commercial aircrafts, such as the NLF airfoils of Honda's HA-420 business jet [10,11], the NLF nacelles of the Boeing 787 aircraft [12], and the NLF winglets of the Boeing 737 MAX aircraft [13], successfully demonstrated the benefits of NLF technology and greatly inspired further research and development in this area. ...
Aerodynamic shape optimization of a swept natural-laminar-flow wing in the transonic regime is still challenging. The difficulty is associated with reliable prediction of laminar–turbulence transition and reasonable compromise of viscous and wave drags. This paper proposes to use efficient global optimization based on surrogate models to address this problem. The Reynolds-averaged Navier–Stokes flow solver features automatic transition prediction via a full eN method, in which dual N factors are used for Tollmien–Schlichting and crossflow instabilities, respectively. The optimizer is based on the kriging surrogate model and parallel infill-sampling method. The baseline natural-laminar-flow wing for short- and medium-range transport aircraft is designed at a cruise Mach number 0.75. Then, drag minimization with up to 42 design variables is carried out, and significant drag reduction (8.79%) has been achieved. A close examination of the optimal wing shows that the drag reduction mainly comes from shock-wave weakening on the upper surface and laminar flow extending via suppression of crossflow instability on the lower surface. Robustness of the optimal wing is investigated, and multipoint optimization is further exercised to improve the robustness to the Mach number variation. It is demonstrated that surrogate-based optimization is feasible and effective for aerodynamic shape optimization of transonic natural-laminar-flow wings.
... s one of the effective methods to decrease the drag of aircraft, the concept of laminar flow has been proposed in 30s last century 1 . Laminar flow technology, including natural-laminar-flow(NLF) 2,3 , laminar-flowcontrol(LFC) 4,5 and hybrid-laminar-flow(HLF) 6,7 , commits to delaying laminar-turbulent transition. Although laminar flow technology could bring great benefits in reducing drag and the associated fuel burn, it is only recently that these benefits have begun to be realized on commercial transport aircrafts. ...
... In this paper, we consider the results of five large-scale laminar flow tests: three wind tunnel experiments in the ONERA S1MA wind tunnel, flight tests with the ATTAS VFW614 NLF glove [2,3] and with the Fokker F100 NLF glove. ...
The e
N
-method is applied to three large-scale experiments performed in the ONERA S1MA wind tunnel and two flight experiments with the ATTAS VFW614 and a Fokker F100 aircraft. N-factor correlations using incompressible as well compressible linear stability theory obtained for natural and hybrid laminar flow surfaces are discussed.
... During the flight test, a database was obtained for variations in Mach numbers from 0.35 to 0.7, Reynolds numbers from 12 × 10 6 to 30 × 10 6 , and sweep angles from 18° to 24° (obtained with sideslip). For a Mach number of 0.35, the transition front ranged from 8 to 50 percent chord dependent on flap and yaw settings (Horstmann et al. 1990). For TS-disturbance-dominated transition, the transition front was at nearly the same chordwise location across the span, whereas for CF-disturbancedominated transition, a distinct sawtooth pattern arose (reminiscent of CF transition). ...
The history of Laminar Flow Control (LFC) from the 1930s through the 1990s is reviewed and the current status of the technology is assessed. Early studies related to the natural laminar boundary-layer flow physics, manufacturing tolerances for laminar flow, and insect-contamination avoidance are discussed. Although most of this publication is about slot-, porous-, and perforated-suction LFC concept studies in wind tunnel and flight experiments, some mention is made of thermal LFC. Theoretical and computational tools to describe the LFC aerodynamics are included for completeness. 1. Introduction This overview reviews Laminar Flow Control (LFC) research that began in the 1930s and flourished through the early 1960s until it was de-emphasized because of a change in national priorities. During the 1970s when the oil embargo by OPEC led to a fuel shortage and high-cost fuel, LFC research became important again because of the aerodynamic performance benefits it could potentially prod...
Local stability theory (LST) combined with an \(e^N\) method is a well-suited approach for predicting transition. However, simpler or surrogate LST models are often used in an automated transition prediction framework for robustness and efficiency reasons. There are, however, few surrogate-based stability methods for the instabilities encountered in three-dimensional compressible flows. To address this problem, this paper proposes a radial-basis-function-based surrogate model capable of reproducing the instability characteristics of two-dimensional Tollmien-Schlichting waves and stationary cross-flow instabilities. The basic flows necessary to set up the model stem from compressible local Falkner-Skan-Cook similarity solutions. The suitability of the proposed method is demonstrated by N-factor computations for three ATTAS flight test points at transonic conditions.
The principal source of difficulty in the application of gradient-based optimizers is the requirement for having a non-discontinuous and mathematically predictable design space. Non-gradient based methods can prove more complex to implement than Gradient Based Methods (GBM), but they do not require continuity or predictability over the design space, and usually increase the likelihood of finding a global optimum. Methods of optimization known as metaheuristics can offer robust methods of finding a solution, and increases the likelihood of converging onto a solution at the global optimum. These gradient-free methods are known to be capable of engaging with numerically noisy optimization problems that be difficult for GBM. This is because metaheuristics operate from a completely different paradigms usually based on the some naturally occurring phenomenon. Unlike gradient methods, derivatives of the cost functions are not necessary, allowing metaheuristics to easily cope with non-continuous or numerically noisy cost functions. Furthermore, no pre-defined baseline design or knowledge of the design space is required and gradient-free methods typically optimize several solutions in parallel.
Aircraft drag can be drastically reduced by maintaining laminar flow over the wing. Laminar flow is maintained by reducing the amplification of instability mechanisms which exist in the boundary layer. Accounting for these instability mechanisms by optimizing airfoil design is referred to as the Natural-Laminar-Flow concept. To stabilize the Tollmien-Schlichting (TS) mechanism, a favorable pressure gradient can be promoted along the majority of the chord, and the crossflow instability can be minimized by reducing the sweep. The main subject of this paper is the S207 airfoil, which is a slotted, NLF (SNLF) airfoil. The main goal of this paper is to determine the critical sweep angle beyond which the crossflow instability negates the benefit of the SNLF design of the S207 airfoil, over a range of angles of attack. The amplification of all anticipated unstable disturbances is calculated by means of the Linear Parabolized Stability Equations (LPSE) in tandem with Linear Stability Theory (LST). Base-flow data was obtained using the Navier-Stokes solver, OVERFLOW, and the boundary-layer equation solver, DEKAF. These solutions were used for cross-verification. At the reference angle of attack -1.520 degrees, for sweep angles greater than or equal to 25 degrees, the stationary crossflow (SCF) N-factors exceed 5 units; the estimated critical value for SCF. For a sweep angle of 20 degrees, although SCF N-factors remain below 5, traveling crossflow (TCF) N-factors do. Lowering the sweep angle to 15 degrees, the TS and CF N-factors do not exceed their critical values for the reference angle of attack, but when the angle of attack is varied, the TCF N-factors exceed 5 units. Studying a slightly lower sweep angle of 12.5 degrees, and considering the full range of angles of attack, the crossflow N-factors remain below 5, while the Tollmien-Schlichting N-factors reach a maximum value of 4.6 units. Assigning a critical N-factor for TS induced transition of 9, the boundary layer on both sides of the airfoil are not expected to transition for all considered angles of attack at a 12.5 degree sweep. This statement appears to hold when the interaction between stationary crossflow and TS is taken into account in a qualitative manner. For this reason, the crossflow instability does not appear to negate the SNLF design of the S207 airfoil at design flight conditions for the sweep angles up to and including 12.5 degrees.
A multi-sensor hot-film concept was developed which meets the requirements when the hot-film technique is used under flight conditions. Experimental and numerical investigations were performed for the development of this measurement technique to show that the coated hot films maintain their sensitivity to detect transition on airfoils in free flight for several months. After the introduction of an improved measurement system, several coating procedures, materials, and coating thicknesses to improve mechanical durability and signal quality of the hot-film sensors were analysed. The paper shows that several solid sensors with protective coatings exist such that multi-sensor arrays are on the one hand, sufficiently robust to stand the environmental conditions and on the other hand, still possess the sensitivity to locate transitional areas.
Already in the early phase of the first axial flow compressors for jet engines some aspects of this turbomachinery have been of importance in having exerted a substantial influence on the whole development of jet engines. To these aspects belong the off-design conditions and the aerodynamics of the blading of the axial flow compressor as well as the necessary performance improvement of the compressors in order to produce larger propulsive forces (thrust). These influences will be dealt with here, where also the findings of the early phase will be addressed because these topics refer to typical contributions made in Germany. From that it follows that the subject is not new.
The aerodynamic performance of a multi-element high-lift system has a critical influence on the direct operating cost of a subsonic civil transport aircraft. A thorough understanding of the aerodynamic characteristics of these multi-element aerofoils and wings allows aircraft companies to design and build more competitive aircraft with high-lift systems that are less complex and lighter for given high-lift performance or that have improved lift and drag characteristics for given system complexity and weight. Flight experiments on NASA Langley's B737-100 aircraft have been conducted to further enhance the understanding of the complex flows about multi-element high-lift systems at full-scale flight conditions. In this paper, an overview of the flight program is provided, followed by high-lights of experimental results and computational analysis. Measurements included surface pressures on the slats, main element and flap elements using flush pressure ports and pressure belts, surface shear stresses using Preston tubes, off-surface velocity distributions using boundary layer/wake rakes, aeroelastic deformations of the flap elements using an optical positioning system, and boundary layer transition detection using hot-film anemometers and an infrared imaging system. Boundary layer transition measurements on the slat using hot-film sensors are correlated with the flow visualisation results from an infrared imaging technique. Extensive application of several computational techniques and comparisons with flight measurements are shown for a limited number of cases. This program has generated an extensive set of data, much of which are still being analysed.
Since the classical experiments performed by Osborne Reynolds (1883), the instability of laminar flow and the transition to turbulence have maintained a constant interest in fluid mechanics problems. This interest results from the fact that transition controls important aerodynamic quantities such as drag or heat transfer. For example, the heating rates generated by a turbulent boundary layer may be several times higher than those for a laminar boundary layer, so that the prediction of transition is of great importance for hypersonic reentry spacecraft, because the thickness of the thermal protection is strongly dependent upon the altitude where transition occurs. For high-subsonic speed, commercial transport aircraft, the achievement of laminar flow can significantly reduce the drag on the wings and hence the specific consumption of the aircraft. The potential benefits are important, because transition separates the laminar flow region with low drag from the turbulent region where skin friction dramatically increases, see figure 3.1. A good knowledge of the transition mechanisms is also required to reduce hydrodynamically generated noise beneath transitional flows.
The objective is to discuss the possibilities of skin friction drag reduction in two- and three-dimensional flows. The discussion is restricted to low speed and transonic problems. The stabilizing or destabilizing effects of pressure gradient, heating/cooling and suction are explained and illustrated by experimental results obtained in wind tunnel or in free flight conditions. Wave cancellation techniques are also presented. An important part of the paper is devoted to the use of the linear stability theory as a practical tool for transition prediction.
DLR’s concept for a natural laminar flow transonic transport aircraft with forward swept wings is presented. Giving an overview on the aircraft’s configurational layout first,
the focus is on the multidisciplinary design of NLF wing and its aerodynamic performance.
Results from high-fidelity coupled aero-structural simulations show that a significant extent of laminar flow is achievable. Torsional divergence of the wing is successfully suppressed by
aeroelastic tailoring. The impact of elastic wing deformation on boundary layer stability and NLF performance is studied. Finally, results of the aircraft’s cruise performance and
expected fuel savings are provided and compared to results from preliminary design.
In this paper results of wind-tunnel experiments and flight tests that demonstrate the effectiveness of coated multisensor hot-film arrays using constant current anemometry to determine boundary-layer transition under cruise conditions are presented. The study focuses on the analysis of several coating procedures, materials, and coating thicknesses to improve mechanical durability and signal quality. After several numerical simulations and environmental resistance tests, for example, in climate chambers and low-speed flight experiments, wind-tunnel experiments and flight tests at transonic speeds are carried out. The numerous experimental and numerical investigations and flight tests in cooperation with the Deutsche Airbus GmbH on a Beluga airplane (300-600ST) show a 1-mu m parylene layer covering the extremely sensitive hot-film sensors and conducting paths to be sufficiently robust to stand the environmental conditions in cruise flight for several months. Furthermore, flight tests at transonic speeds on the Mystere Falcon 20 E5 at an altitude up to 30,000 ft and Mach numbers up to 0.79 are conducted in a joint project with DLR Oberpfaffenhofen. Modified wing models are mounted under both wings of the Falcon. The surfaces are instrumented with coated hot-film arrays using 128 sensors. The tests on the Falcon prove parylene coated hot-film sensors to still possess the sensitivity to locate transition under real flight conditions.
The flow around infinite swept wings is computed using a Reynolds-averaged Navier-Stokes method coupled to a. boundary layer and a transition prediction method based on the e(N) approach. This approach applies the two N factor e(N) strategy, which allows the differentiation between excited Tollmien-Schlichfing waves of constant frequency and stationary crossflow waves of constant spanwise wave number. The applicability is documented by comparing computations with infinite swept-wing experiments on four configurations, where one configuration was tested in two different wind tunnels. The limiting N factors for both types of waves are a priori unknown. Hence, firstly the limiting N factors, which represent the stability limit, are deduced from the measured transition locations. Subsequently, iterative computations are executed to determine the transition location using the coupled system, Navier-Stokes, boundary-layer, and two N factore(N) method incombination with the stability limit. The correlation between experimentally observed and computed transition locations is shown to be good for all considered wind tunnel studies.
The flow around two infinite swept wings and an infinite swept cylinder model is computed using a Reynolds averaged Navier-Stokes method coupled to a boundary-layer and transition prediction method based on the e(N), approach. The configurations were wind-tunnel tested, including transition location measurements. In all test cases, transition is provoked purely by crossflow instabilities. The inifinite swept wing experiments represent critical test cases because they contradict two assumptions made in the conventional e(N) approach, where the limiting crossflow N factor is supposed to be independent of the surface curvature and the quality of the wing surface finish. It will be shown that the extended version of the e(N) method can resolve this problem and produce acceptable results. The infinite swept cylinder model tested at high sweep angles between 53 and 71 deg is excellently suited to demonstrate that the e(N) method is capable of describing the leading-edge contamination problem. For sweep angles between 53 and 58 deg, a relative constant limiting crossflow N factor exists, whereas for higher sweep angles the crossflow N factors at the measured transition location deacrease gradually. Applying the leading-edge contamination criterion of Pfenninger, the e(N) method predicts the sweep angle, where leading-edge contamination starts to influence the laminar boundary-layer flow. Furthermore, the sweep angle, which produces, according to. an empirical correlation, fully turbulent leading-edge flows, is well predicted by the e(N) method.
The capability of the e(N) method to predict transition in complex three-dimensional. laminar boundary layers is demonstrated. Flows around inclined prolate spheroids are ideally suited for investigation because they exhibit highly divergent and convergent three-dimensional viscous flows on curved surfaces. The inviscid flowfield is described by the potential theory and the viscous layer by a three-dimensional laminar boundary-layer method. This approach is applied in weak viscous/inviscid interaction regions only, that is, in attached laminar flow regions, and validated by comparison with measured wall pressure and skin friction in magnitude and direction. For the determination of the transition location, the laminar boundary layer is analyzed by the two N factor e(N) method. Excited Tollmien-Schlichting waves of constant frequencies and stationary crossflow waves of constant wavelength are computed by the local, linear stability theory. Both N factor integrations are executed along 21 streamlines. which regularly cover the surface of the prolate spheroid. First, the values of both N factors at the measured transition locations are calculated, which deliver the stability limit of the prolate spheroid in the considered wind tunnel. Second, based on the knowledge of the stability limit, the transition locations are evaluated. The prolate spheroid with an aspect ratio of 6:1 was tested in two wind tunnels at angles of attack from 0 to 30 deg and Reynolds numbers from 1.5 x 10(6) to 43 x 10(6). The measured transition locations compare remarkably well with computations for all investigated flow cases.
Flight experiments on NASA Langley Research Center's B737-100 airplane were conducted to document flow characteristics for further understanding of high-lift now physics. The measurements included surface pressure distributions measured using flush pressure taps and pressure belts on the slats, main element, and flap elements, and boundary-layer state changes measured using hot-film anemometry and infrared thermography. In this paper, results obtained in the final phase of flight experiments are presented and analyzed. The analysis primarily focuses on changes in the boundary-layer state measured on the slat as a result of changes in nap setting and/or flight condition. The measurements show that extended runs of laminar now exist on the slat at relevant angles of attack. Flow mechanisms that affect the extent of laminar flow include attachment-line contamination, crossflow instability, and relaminarization.
A short overview of the different ways to apply linear stability theory in the form of the e(N) method for transition prediction in boundary-layer flow is given. The ATTAS/VFW614, the Fokker F100, and the Airbus A320 flight tests are evaluated and compared to two hybrid laminar-flow tests in the SIMA tunnel. N-factor correlations are given for all five tests. It is shown that transition prediction based on two N-factors, one to model the crossflow instability and a second for Tollmien-Schlichting instability, is superior to the envelope method for hybrid laminar flow applications. Results concerning the influence of surface roughness on the crossflow instability are discussed. Furthermore, the evaluation shows that in hybrid laminar flow the internal noise coming from the suction system cannot be neglected for the Tollmien-Schlichting amplification.
The LamAiR (Laminar Aircraft Research) project deals with the design of a laminar wing for short and medium range transport aircraft operated in the transonic regime. It is well known that extensive laminar flow on wings of such aircraft still can be achieved by natural means, i.e. solely by contour shaping of the airfoil sections. But with Reynolds numbers being in the order of 25 millions in cruise condition the leading edge sweep of the wing should not be higher than approximately 20deg in order to limit the growth of cross-flow instabilities and, hence, prevent early transition. Consequently, the design cruise Mach number for laminar wings of conventionally aft-swept configurations cannot exceed values of about 0.75 and it is expected that the high-speed off-design performance is rather poor. Within the DLR project LamAiR it is therefore investigated if these aerodynamic shortcomings can be overcome by employing forward sweep in combination with aeroelas-tic tailoring using CFRP (Carbon Fiber Reinforced Plastics) materials. In particular the goal is to design a forward swept laminar wing having a design Mach number of 0.78 and the capability of reaching Mach 0.80 in high-speed off-design. The present paper gives an overview on the current status of the project as well as prospects for future work.
The requirement to reduce the specific fuel consumption of future advanced transport aircraft, in order to achieve definite improvements in the areas of economy and ecology, has resulted in dedicated activities among others in two fields of aeronautics. In the area of propulsion the current trend in the development of turbofan engines is directed towards very high and ultra high bypass ratios with the objective to lower the specific fuel consumption and the noise level. In the area of aerodynamics the laminarization of extensive surface areas of the airframe wetted by the external flow aims at a reduction of the skin friction drag. Since the higher bypass ratios of the advanced powerplants are associated with a growth in the engine nacelle dimensions, not only the skin friction drag rises, but the problem of engine-airframe integration is also intensified due to negative engine-airframe interactions. Therefore there is a challenge to an efficient integration of.
The laminar flow technology is one of the key technologies in aeronautics offering substantial improvements in the fields of economy and ecology. Owing to the sensitivity of laminar boundary layer flow over aerodynamically smooth surfaces to effects of disturbances as for example insect contamination, noise and vibration as well as the atmospheric conditions, all these aspects of the receptivity problem have to be taken into account in the application of the laminar flow technolog. In this context one important element is the atmospheric turbulence. In-flight measurements of the atmospheric turbulence intensity have been undertaken within the framework of a European collabrative programme to flight test a natural laminar flow nacelle. The measurement of the atmospheric turbulence intensity was one of the tasks of the University of Oxford. The knowledge gained concerning the turbulence intensity has been compared with the results of in-flight measurements in Europe from other sources, most of which also stem from investigations related to the laminar flow technology. In contrast investigations specifically addressing the atmospheric turbulence ranging from still air and cumulus cloud to thunderstorm have been conducted in the United States of America by NASA. A comparison is made of the freestream turbulence intensities obtained in flight with results for some wind tunnels. The in-flight measurements have yielded turbulence intensities for the still air atmosphere which are less than 0.05%, a figure that for the wind tunnels examined is only reached in the low speed regime.
Based on a Navier-Stokes approach, the flow around a lamiar airfoil is described together with transition prediction based on the e*-method. The main emphasis is placed on the computation of the transitional flow region, i.e. the range between fully laminar and fully turbulent flow. The extension and intermittency of the transition region are predicted on the basis of available models originating from the turbulent spot theory. A comprehensive computational study is performed as the transition length differs considerably depending on the model used. The influence of the transition length, varying from zero to 20 per cent chord, is investigated in a strong adverse and in a zero pressure gradient airfoil flow region. Based on the outcome of the study a conventional transition length model is proposed, which is applicable in flow regions where transition is predicted well upstream of the laminar separation location. Furthermore, a special transition length model is given, typically for low Reynolds number flows, where the boundary layer stays laminar up to separation. In this case transition is fixed right upstream of the laminar separation point, as the problem of laminar separation bubbles is not considered in the present context. The present investigation concerns the flow over the DoAL3 laminar airfoil at 2° incidence. For that particular case, the e*-method predicts transition at 8 per cent chord on the upper surface. Along the lower surface, the computed N-factors are lower than the limiting N-factor up to laminar flow separation.
Some lessons learned as regards flow transition detection in the course of flight-testing carbon fiber composite conventional and laminar flow nacelles are presented. The key objective of the flight trials was to examine the influences of environmental phenomena and of the engine nacelle structure regarding the interpretation of the nacelle infrared images and of the nacelle surface temperature measurements, both with reference to determining the location of the flow transition front on the nacelles. As regards the application of the infrared thermography, the study has yielded the result that the quality of the infrared images is influenced by heat radiation to space and from Earth, as well as by reflection of sunlight from the airframe. For good quality infrared images, only carbon fiber composites should be used as structural material for the nacelle fan cowls. In addition, the mechanical design of the fan cowls should not contain any stiffening frames in the substructure, but the should be designed as a uniform honeycomb skin structure. Concerning the surface temperature measurements, for good-quality results,a uniform honeycomb carbon fiber composite skin structure is just as essential as for the infrared imaging system.
Navier-Stokes airfoil computations coupled to e* transition prediction are feasible, provided the limiting N factor is known beforehand. In the present study a procedure is outlined to validate free-transition airfoil experiments in wind tunnels for which the limiting N factor is not known a priori. The approach does not rely on mesh adaption procedures to obtain adequate laminar viscous layer data from Navier-Stokes computations for the stability analysis. To the contrary, the laminar viscous layer is computed by a boundary-layer method applying as input the pressure distribution from Navier-Stokes computations on initial meshes. Two measurement campaigns are validated: the NLF(1)-0416 laminar airfoil in the low-speed NASA Langley Low-Turbulence Pressure Tunnel LTPT and the NACA 64*A015 airfoil in the NASA Ames 12-Foot Pressure Tunnel. The free-transition measurements in both tunnels include pressure distributions and transition locations and, in supplement, for the NLF(1)-0416 laminar airfoil lift and drag measurements. The computational results document the validity of the present approach, the existence of a constant limiting N factor for a specific wind tunnel, and an excellent agreement with the experimental findings.
The laminar flow technology is one of the key technologies in aeronautics offering substantial improvements in the areas of economy and ecology. This paper describes the aerodynamic design methodology for a natural laminar flow (NLF) nacelle and the subsequent verification of the design quality by flight tests with a subsonic transport aircraft. The aerodynamic design was a contribution within the framework of a European collaborative programme to flight test a natural laminar flow nacelle. The participants in this programme were the aero-engine manufacturers Rolls-Royce and MTU as well as the Institute for Design Aerodynamics of the DLR. The nacelle manufacture was undertaken by Hurel-Dubois (UK) Ltd. and some components of instrumentation were provided and operated by the University of Oxford. The flight trials, apart from having the goal of validating the aerodynamic design of the nacelle, had the major objective of obtaining realistic experience with the operational aspects of the NLF nacelle. In the first series of flight trials, which began in September 1992, the flight test results confirmed the suitability of the aerodynamic design tools used.
The DoAL3 airfoil was designed for a laminar wing of a commuter aircraft with small leading edge sweep flying at moderate Mach numbers, M = 0.45-0.6, and Reynolds numbers up to Re = 14 x 106. Free transition measurements were executed in the Transonic Wind Tunnel Braunschweig (TWB) of the DLR, German Aerospace Center, at a Mach number of 0.48 and a Reynolds number of 3.2 x 106. To quantify the airfoil performance for flight Reynolds numbers a Navier-Stokes method coupled to eN transition prediction is applied. The limiting N factor for the TWB facility has been shown beforehand to achieve a value of six, in flight a mean value of twelve can be expected. The extrapolation to flight Reynolds numbers considers the aircraft flying in adiabatic and heating flow conditions. When exposed to extensive sunshine on the airfield the fuel contained in the wet wings will be heated and the aircraft is then flying a considerable time in climb and cruise with wings exposed to non negligible heat transfer. The effects of Reynolds number and heat transfer on the extent of laminar wing flow are investigated separately. The important findings of the numerical study are, firstly, that in adiabatic conditions the wing will fly with the same extent of the laminar bucket as observed in the wind tunnel. Secondly, in heating conditions laminarity is completely lost, the drag is almost doubled with respect to the minimum drag for adiabatic flows. © 2002 Éditions scientifiques et médicales Elsevier SAS. All rights reserved.
Summary An earlier paper described a method of calculating the turbulent boundary layer flow over the rear of an infinite swept wing. It made use of an entrainment equation and momentum integral equations in streamwise and cross-flow directions, together with several auxiliary assumptions. Here the method is adapted to the calculation of the turbulent boundary layer flow along the attachment line of an infinite swept wing. In this case the cross-flow momentum integral equation reduces to the identity 0 = 0 and must be replaced by its differentiated form. Two alternative approaches are also adopted and give very similar results, in good agreement with the limited experimental data available. It is found that results can be expressed as functions of a single parameter C *, which is evidently the criterion of similarity for attachment-line flows.
The transition behaviour of the boundary layer which is formed along an infinite swept attachment line has been studied experimentally. Circular trip wires and turbulent flat plate boundary layers have been used as sources of disturbance and the range of parameters covered has been such that the reuslts are directly applicable to full scale flight conditions. Simple criteria have been deduced which allow the state of the boundary layer to be determined for given geometric and free stream properties.
The use of infrared imaging to detect transition lines, where the
laminar boundary layer changes to the turbulent state, is described. The
basic physical principles as well as the system and examples are
discussed.
Flight and wind-tunnel experiments have been carried out to measure the pressure distribution and the transition location on a special wing glove. By means of the linear stability theory of laminar boundary layers limiting N values of Tollmien-Schlichting waves at the transition location have been evaluated. The values of N ≈ 13.5 are nearly independent of Reynolds number and are the same in flight and wind-tunnel tests.