24TH INTERNATIONAL CONGRESS OF THE AERONAUTICAL SCIENCES AERODYNAMIC EFFICIENCY OF HIGH MANEUVERABLE AIRCRAFT APPLYING ADAPTIVE WING TRAILING EDGE SECTION Christian Breitsamter Institute for Fluid Mechanics, Aerodynamics Division, Technische Universität München Boltzmannstrasse 15, 85748 Garching, Germany Keywords: Adaptive wing, high agility aircraft, high angle of attack, force measurements Abstract environmental impacts enforce continuous im- The aerodynamic characteristics of a high- provement of the performance of both civil and agility aircraft of canard-delta wing type are military aircraft. A key issue enhancing the air- analyzed in detail comparing the performance craft efficiency is the use of advanced and adap- obtained by smooth, variable camber wing trail- tive wing technologies (Fig. 1) [1-4]. ing-edge sections versus conventional trailing- The aerodynamic characteristics of adap- edge flaps. Wind tunnel tests are conducted on a tive and form-variable wing elements and multi- detailed model fitted with discrete elements rep- functional control surfaces have been inten- resenting the adaptive wing trailing-edge sec- sively studied in several German and European tions. Force and flow field measurements are research programs [5-8]. Results show increased carried out to study the aerodynamic properties. glide number, reduced drag and improved high Lift, drag and pitching moment coefficients are lift performance. A main focus is on structural evaluated to assess aerodynamic efficiency and concepts for adaptive flap systems as well as for maneuver capabilities. Comparing the data for local surface deformations and the integration of adaptive and conventional flap configurations, related sensors, actuators and controllers in the it is shown that smooth, variable camber results aircraft [9]. Local surface thickening known as in an increase in lift at given angle of attack and ‘Shock Control Bump’ (SCB) is used to modify in a decrease in drag at given lift coefficient. the shock development at transonic flight condi- These benefits hold also for trimmed conditions. tions [10]. An appropriate position and shape of They are demonstrated for an angle-of-attack the SCB may lead to a nearly isentropic recom- range of −10 to +30 deg, substantiating a sig- pression along with a reduced pre-shock Mach nificant improvement in flap efficiency for the number. This reduction results in a lower shock adaptive wing concept. Thus, the form-variable strength and, therefore, in a decrease in wave wing demonstrates a remarkable potential to drag. Extensive studies are conducted on the enhance flight and maneuver performance and aerodynamic and structural design of SCBs to to alleviate maneuver loads, respectively. make them applicable for large transport aircraft wings [11, 12]. 1 Introduction Also, advanced structural concepts for variable camber wing trailing-edge sections The forthcoming demands on safe and efficient have been developed. The structural solutions flight operations in context of society needs and were proven by full scale ground demonstrators applying the so called ‘horn concept’ [13], the Copyright © 2004 by C. Breitsamter. Published by the ‘girdle concept’ [14] and the ‘finger concept’ International Council of the Aeronautical Sciences, with [9]. Further studies concentrate on multifunc- permission. tional control surfaces [15]. For example, small 1 CHRISTIAN BREITSAMTER 2 Experiment and Test Program Laminar Shock Turbulent 2.1 Model and Facility flow control control flow control Perturbation Experiments are performed on an 1:15 scaled damping detailed steel model of a modern high-agility Boundary turbulent aircraft (Fig. 2). Major parts of the model are layer suction Transition laminar Load nose section, front fuselage with rotatable can- control ards and a single place canopy, center fuselage with delta wing and a through-flow double air Variable intake underneath, and rear fuselage including camber nozzle section and empennage (fin). The base- line wing configuration is equipped with con- Fig. 1 Adaptive wing technologies. ventional inboard and outboard trailing-edge control surfaces mounted at the trailing-edge of flaps. The model elements designed to represent conventional flaps show an increase in maxi- the smooth, variable camber of the adaptive mum lift of about 7% [16]. Recently, investiga- wing trailing-edge section have the same span- tions on ‘Miniature Trailing Edge Devices’ wise and chordwise dimensions as the conven- (Mini-TEDs) have been performed in wind tun- tional flaps. The form-variable section is com- nel and flight tests within the technology plat- posed by six discrete wing elements which are form AWIATOR (‘Aircraft WIng with Ad- exchanged for the conventional flaps to run the vanced Technology OpeRation’) funded by the adaptive wing tests. The corresponding de- European Commission [8]. flections are defined by the tangent on the Concerning military aircraft, comprehen- trailing-edge including values of ηFT = 0°, 5°, sive flight tests were carried out on the F-111 10°, 15°, 20°, and 25° (Fig. 3). To evaluate the research aircraft to develop and assess the tech- aerodynamic performance obtained by adaptive nology enhancements of the ‘Mission Adaptive wing versus conventional flaps, the comparison Wing’ (MAW), employing smooth contour, is based on a measure of deflection angles with variable camber leading- and trailing-edge sur- respect to the skeleton line. Fig. 4 depicts the faces [17]. Flight tests were conducted in frame corresponding definitions used for conventional of the joint U.S. Air Force/NASA/ Boeing Ad- flap deflection ηK and deflection of the form- vanced Fighter Technology Integration (AFTI/ variable trailing-edge section ηFS. F-111) program. Improvements in glide number, The experiments have been carried out in large reductions in drag at off-design conditions the large low-speed facility A of the Institute for and alleviation of maneuver loads have been Fluid Mechanics (Aerodynamics Division) of successfully demonstrated [18-20]. Technische Universität München. This closed- However, there is a lack of investigations return wind tunnel can be operated with both applying adaptive leading- and trailing-edge open and closed test section at maximum usable surfaces on high-maneuverable aircraft fitted velocities of 75 m/s and 65 m/s, respectively. with slender wing geometry [21]. Research ac- Test section dimensions are 1.8 m in height, 2.4 tivities concentrate mainly on related structural m in width and 4.8 m in length. Because high concepts but there is also a need for systematic angle-of-attack experiments are conducted the aerodynamic investigations. Therefore, the pre- open test section is used. The test section flow sent study uses an aircraft model of canard-delta was carefully inspected and calibrated docu- wing type to assess the aerodynamic perform- menting a turbulence level less than 0.4% and ance of adaptive wing sections versus conven- uncertainties in the spatial and temporal mean tional trailing-edge flaps. velocity distributions of less than 0.067%. 2 AERODYNAMIC EFFICIENCY OF HIGH MANEUVERABLE AIRCRAFT APPLYING ADAPTIVE WING TRAILING EDGE SECTION 2s = 0.740 m ϕW = 50° Index: lµ = 0.360 m ϕW = 45° W: Wing ARW = 2.45 λW = 0.14 C: Canard ARF = 1.38 ϕF = 54° F: Fin Skeleton line a) Conventional trailing-edge flap. Skeleton line b) Form-variable trailing-edge section. Fig. 4 Definition of trailing-edge deflections. Fig. 2 Model geometry of delta-canard fighter aircraft. a) Front view. η = 0° FT ηFT = 5° ηFT = 10° ηFT = 15° ηFT = 15° ηFT = 20° ηFT = 25° ηFT = 25° b) Rear views. Fig. 3 Wing trailing-edge elements representing the smooth contour, variable camber trailing-edge section. Fig. 5 Model mounted in test section. 3 CHRISTIAN BREITSAMTER gauge bridges. The balance is designed to sus- Wiring to connect strain gauges tain maximum loads of 900 N, 450 N, and 2500 with amplifier modules N for axial, lateral and normal forces, respec- tively, and maximum moments of 120 Nm, 160 Components for strain gauge balance integration Nm, and 120 Nm for rolling, pitching, and yaw- ing moments, respectively. The accuracy based Strain gauge balance on maximum loads ranges from 0.05% to 0.1% for the force components and from 0.8% to 1.2% for the moment components. The strain Tail sting and Z-adapter to link gauge bridges are connected with six-wire con- balance with support sting ductors to the carrier frequency amplifier mod- ules of a specific measurement system (HBM Fig. 6 Components for strain gauge balance integration MGC plus). This system is fully computer con- and connection with tail sting adapter. trolled providing also the excitation for the strain gauge bridges. The calculation of forces and moments from the balance signals employs a comprehen- sive balance calibration. The mathematical rela- tionship assigning balance signals to given loads uses a third order approximation of the balance behavior. Pure single loads and combinations of Fig. 7 Six-component strain gauge balance. single loads are stepwise applied to the balance during the calibration. The evaluation of each The model is sting mounted using a spe- loading sequence is based on a least square error cific tail sting with its left cylinder attached to second order polynomial approximation. The the internal strain gauge balance and its right complete calibration coefficient matrix is de- conical part connected with the horizontal sting rived from those evaluations resulting in a sys- of the 3-axis model support (Figs. 5 and 6). The tem of equations. The related solution applies a computer controlled model support provides an least square error method to obtain the aerody- angle-of-attack range of −10° ≤ α ≤ +30°, and namic loads from the balance signals [22]. models may be yawed and rolled 360°. The un- certainty in angle setting is less than 0.05°. 2.3 Test Conditions and Data Processing Force measurements have been made at free 2.2 Force Measurements stream reference velocities of U∞ = 40 m/s and An internal six-component strain gauge balance U∞ = 60 m/s at ambient pressure p∞ and ambient is used to measure aerodynamic forces and mo- temperature T∞ .