Outboard Flap Research Articles

Investigation of Improvement Design on Aileron Surface Flow State of High Lift Configuration in BWB

The aileron is one of the most important tools for adjusting the roll attitude of the aircraft, but the surface flow state of the aileron is likely to be affected by high-lift devices. In this paper, by using computational fluid dynamics (CFD) simulations and wind tunnel tests, the Krueger flap effects on the surface flow state of ailerons in a typical blended wing body civil aircraft were investigated. In order to increase the lift, deflecting the Krueger flap makes the flow separation occur on the aileron surface of BWB civil aircraft. This way of the surface stall that the flow separation in the aileron zone first appears at the wing tip rather than at the wing root is unreasonable for civil aircraft. For the above problem, a sensitivity analysis of the design parameters of the Krueger flaps was carried out. The results indicate that the angle of the outboard Krueger flap mainly affects the flow separation of the ailerons. Its length affects the pitch moment tremendously, while its width slot affects the pitch moment slightly. Finally, the design principles of the BWB Krueger flap for the improvement aileron surface flow state were proposed, and the redesign of the BWB high lift configuration significantly improved the flow state of the aileron zone at a minimal cost of aerodynamic characteristics without losing the existing great aerodynamic performance.

Noise simulations of flap devices for wind turbine rotors

AbstractThe rotor of a large diameter wind turbine experiences more substantial and more dynamic loads due to the fluctuating and heterogeneous wind field. The project SmartBlades 2.0 investigated rotor blade design concepts that alleviate aerodynamic loading using active and passive mechanisms. The present work evaluates the acoustics of the two load alleviating concepts separately, an inboard slat and an outboard flap, using the Fast Random Particle Mesh/Fast Multipole Code for Acoustic Shielding (FRPM/FMCAS) numerical prediction toolchain developed at DLR with input from the averaged flow field from RANS. The numerical tools produce a comparable flap side‐edge noise spectrum with that of the measurement conducted in the Acoustic Wind Tunnel Braunschweig (AWB). The validated FRPM/FMCAS was then used to analyze the self‐noise from a slat at the inboard section of a rotor blade with a 44.45 m radius and compared with that from the outboard trailing edge. Furthermore, the rotational effect of the rotor was included in the post‐processing to emulate the noise observed at ground level. The findings show an increase in the slat's overall sound pressure level and a maximum radiation upwind of the wind turbine for the case with the largest wind speed that represents the off‐design condition. In operational conditions, the slat adds at most 2 dB to the overall sound pressure level. The toolchain evaluates wind turbine noise with conventional or unconventional blade design, and the problem can be scaled up for a full‐scale analysis. As such, the tools presented can be used to design low‐noise wind turbines efficiently.

Yaw-control efficiency analysis for a diamond wing configuration with outboard split flaps

The yaw-control device of a low-aspect ratio flying wing with diamond-shaped wing planform is investigated. Extensive low-speed wind tunnel experiments have been carried out to obtain surface pressure data and the aerodynamic forces and moments of the configuration for six different flap deflection angles at varying angles of attack and sideslip. Complementary unsteady Reynolds-averaged Navier–Stokes simulations are performed for selected configurations. The experimental data is used to examine the validity of the numerical results. The analysis is focused on the aerodynamic coefficients and derivatives. Yaw-control effectiveness, yaw-control efficiency, crosswind landing capabilities and coupling effects are discussed. The results show sufficient yaw-control effectiveness and efficiency for a wide range of considered freestream conditions. The outboard flap exhibits a non-linear characteristic with respect to the flap deflection angle and freestream conditions. The efficiency is considerably reduced at high angles of attack due to large-scale flow separation in the wing outboard section. Non-linear coupling effects with the rolling moment become obvious for moderate to large flap deflections over the whole angle of attack polar. The numerical results show good agreement with the experimental data in the surface pressure distributions and longitudinal aerodynamic coefficients. The yawing moment is overpredicted by numerical simulations for large flap deflection angles.

Multi-criteria weighted decision making for operational maintenance processes

This paper proposes an approach towards multi-criteria decision making (MCDM) for operational maintenance processes. It focuses on decision alternative identification and evaluation for short time horizons, thereby addressing problems that need to be resolved in hours or a few days at maximum. This addresses a gap in literature, where MCDM methods are predominantly proposed for strategic maintenance decision making. The proposed approach addresses two distinct steps of decision making: 1) identification of decision alternatives and 2) evaluation of decision alternatives. For identification of decision options, the Boolean Decision Tree (BDT) method is selected to accommodate for the qualitative and discrete operational factors that determine the available, feasible decision alternatives in operational maintenance processes. The feasible alternatives are subsequently evaluated using the weighted sum method (WSM). The approach is applied to a Boeing 777 outboard flap damage case, using real maintenance and operational data. A decision tool has been developed and verified, showing the capability of the approach to systematically identify and evaluate operational maintenance decision making problems in a few minutes. The results suggest that the proposed approach could save in excess of 50% on decision process time, with added benefits in full identification of the available set of decision alternatives at problem onset. In addition, sensitivity analysis on the basis of a global evaluation of the weight space is provided to investigate the impact of weight settings on the decision outcomes.

Analysis of a Multiflap Control System for a Swashplateless Rotor

An analytical study was conducted examining the feasibility of a swashplateless rotor controlled through two trailing-edge flaps (TEFs), where the cyclic and collective controls were provided by separate TEFs. This analysis included a parametric study examining the impact of various design parameters on TEF deflections. Blade pitch bearing stiffness; blade pitch index; and flap chord, span, location, and control function of the inboard and outboard flaps were systematically varied on a utility-class rotorcraft trimmed in steady level flight. Gradient-based optimizations minimizing flap deflections were performed to identify single- and two-TEF swashplateless rotor designs. Steady, forward flight analysis suggest that a two-TEF swashplateless rotor where the outboard flap provides cyclic control and inboard flap provides collective control can reduce TEF deflection requirements without a significant impact on power, compared to a single-TEF swashplateless rotor design.

A self-synchronizing and low-cost structural health monitoring scheme based on zerocrossing detection

Owing to their high specific strength and stiffness, composite materials are increasinglybeing used in aeronautics and astronautics, but such materials are vulnerable to impactdamage and delamination. Structural health monitoring (SHM) techniques have beendeveloped for detecting such defects in recent years. In situ, self-powered and low-cost SHMsystems are a developmental tendency of this technique. This paper introduces theprinciples of a low-cost and self-synchronizing scheme for SHM. Based on the Lamb waveinteractions with the structure, the proposed technique relies on detecting zero crossingtime instants in order to derive an estimation of the structural state. It is shown that sucha method provides a very simple and low-cost way to assess the structural integrity whilebeing computationally efficient. Experimental investigations carried out on a compositeplate with an increasing penetration hold validating the proposed technique show itseffectiveness for detecting the damage. The proposed approach has also been appliedon an aircraft outboard flap to detect the impact damage. The robustness isdiscussed versus time-shift and magnitude jitter assumptions by using the platecase. The temperature effect is also considered by defining a coefficient array inorder to compensate for the material property changes. Finally, an embeddedimplementation of such a SHM technique is presented by using the proposed damage index.

Wake Measurement Downstream of a Hybrid Wing Body Model with Blown Flaps

Flow-field measurements were obtained in the wake of a full-span Hybrid Wing Body model with internally blown flaps. The test was performed at the NASA Langley 14 x 22 Foot Subsonic Tunnel at low speeds. Off-body measurements were obtained with a 7-hole probe rake survey system. Three model configurations were investigated. At 0deg angle of attack the surveys were completed with 0deg and 60deg flap deflections. At 10deg angle of attack the wake surveys were completed with a slat and a 60deg flap deflection. The 7-hole probe results further quantified two known swirling regions (downstream of the outboard flap edge and the inboard/outboard flap juncture) for the 60deg flap cases with blowing. Flowfield results and the general trends are very similar for the two blowing cases at nozzle pressure ratios of 1.37 and 1.56. High downwash velocities correlated with the enhanced lift for the 60deg flap cases with blowing. Jet-induced effects are the largest at the most inboard station for all (three) velocity components due in part to the larger inboard slot height. The experimental data are being used to improve computational tools for high-lift wings with integrated powered-lift technologies.

Optimal Placement of Piezoelectric Actuated Trailing-Edge Flaps for Helicopter Vibration Control

This study aims to determine the optimal locations for dual trailing-edge flaps on a helicopter blade in the presence of actuator hysteresis. An aeroelastic analysis based on a finite element approach in space and time is used in conjunction with an optimal control algorithm to determine the actuator control input for vibration minimization. The reduced hub-vibration level and the flap power are the two optimization indices considered in this study. The location of the flaps along the blade are the design variables. The hysteresis in the piezo-actuators is modeled using a dynamic hysteresis model based on an extension to the classical Preisach model. It is found that second-order polynomial response surfaces based on the central composite design of the theory of design of experiments describe both objectives adequately. Numerical studies for a four-bladed hingeless rotor show that both objectives are more sensitive to outboard flap location compared with inboard flap location. Optimization studies indicate that the dual-flap configuration for the least hub-vibration level is different from the configuration for the least flap power. The Pareto front between the two objectives is found to be discontinuous. However, a reasonable tradeoff configuration is obtained by careful inspection of the Pareto front. This configuration yields about a 70% reduction in hub-vibration levels from the baseline conditions at an advance ratio of 0.30, while requiring about 21% more flap power from the initial configuration of the optimization study.

Transport Aircraft Wake Influenced by a Large Winglet and Winglet Flaps

Detailed flowfields of a wind-tunnel investigation are discussed presenting the wake vortex development and evolution in the near field and extended near field behind a four-engined large transport aircraft model fitted with a large winglet. The tests use a half-model of 1:32 scale focusing on the high-lift case of a typical approach configuration at a Reynolds number of 0.5 x 10 6 based on the wing mean aerodynamic chord and at an angle of attack of 7 deg. Flowfields are carefully inspected by advanced hot-wire anemometry at seven crossflow planes up to 5.6 spans downstream of the model. Based on the measured time-dependent velocity components, the wake flowfield is analyzed by distributions of mean velocity and vorticity, turbulence intensities, and spectral densities. The near-field wing vortex sheet is dominated by seven main vortical structures, namely the winglet vortex, the wing tip vortex, the outboard flap vortex, the outboard and inboard nacelle vortices, and the vortices shed at the wing-body junction and the horizontal tail plane. In the extended near field, these vortices roll up and merge to form the remaining rolled-up vortex. The deflection of winglet flaps produce additional vortices influencing the wing tip near field and enhancing the overall merging process. Especially for the cases with asymmetrical flap deflection, the remaining rolled-up vortex shows a significant narrowband concentration of turbulent kinetic energy which may result in an amplification of inherent far-field instabilities.

Optimal placement of trailing-edge flaps for helicopter vibration reduction using response surface methods

This study aims to determine optimal locations of dual trailing-edge flaps to achieve minimum hub vibration levels in a helicopter, while incurring low penalty in terms of required trailing-edge flap control power. An aeroelastic analysis based on finite elements in space and time is used in conjunction with an optimal control algorithm to determine the flap time history for vibration minimization. The reduced hub vibration levels and required flap control power (due to flap motion) are the two objectives considered in this study and the flap locations along the blade are the design variables. It is found that second order polynomial response surfaces based on the central composite design of the theory of design of experiments describe both objectives adequately. Numerical studies for a four-bladed hingeless rotor show that both objectives are more sensitive to outboard flap location compared to the inboard flap location by an order of magnitude. Optimization results show a disjoint Pareto surface between the two objectives. Two interesting design points are obtained. The first design gives 77 percent vibration reduction from baseline conditions (no flap motion) with a 7 percent increase in flap power compared to the initial design. The second design yields 70 percent reduction in hub vibration with a 27 percent reduction in flap power from the initial design.

Vortex Behaviors of Rolled Supersonic Transport Configuration with Leading-Edge Flaps

Wind tunnel measurements were done to investigate the effect of leading-edge flaps on the rolling moment characteristics of the cranked arrow wing for the supersonic transport. Static rolling moment measurements, flow visualization studies, and crossflow velocity measurements by a particle image velocimetry were made at Re = 6.2 x 104. Static rolling moment measurements confirmed linear restoring moment at « < 16deg for the models both with and without flap deflection. When the outboard leading-edge flap is deflected 12 deg, rolling moment hysteresis is observed at the roll angle of 0 as 20 deg and α ≈ 20 deg. The experimental results indicate that the hysteresis is caused by different vortex breakdown behaviors on the inboard wing when the wing is rotated in the clockwise and counterclockwise directions.

Transonic Free-To-Roll Analysis of the F-35 (Joint Strike Fighter) Aircraft

The free-to-roll technique is used as a tool for predicting areas of uncommanded lateral motions. Recently, the NASA/Navy/Air Force Abrupt Wing Stall Program extended the use of this technique to the transonic speed regime. Using this technique, this paper evaluates various wing configurations on the Joint Strike Fighter (F-35) aircraft. The configurations investigated include leading- and trailing-edge flap deflections, leading-edge flap gap seals, and vortex generators. These tests were conducted in the NASA Langley 16-Foot Transonic Tunnel. The analysis used a modification of a figure of merit developed during the Abrupt Wing Stall Program to discern configuration effects. The results showed how the figure of merit can be used to schedule wing flap deflections to avoid areas of uncommanded lateral motion. The analysis also used both static and dynamic wind-tunnel data to provide insight into the uncommanded lateral behavior. The dynamic data were extracted from the time history data using parameter identification techniques. In general, sealing the gap between the inboard and outboard leading-edge flaps on the Navy version of the F-35 eliminated uncommanded lateral activity or delayed the activity to a higher angle of attack.

Active and passive vortex wake mitigation using control surfaces

Aircraft trailing vortices constitute a hazard to following aircraft and are therefore one of the main concerns for airport capacity constraints. At the Institute of Aerospace Engineering (ILR) experiments on wake vortices are conducted in a towing tank using particle image velocimetry. The motivation behind the presented investigations is the alleviation of the rolling moment induced on following aircraft using control surfaces of the preceding aircraft. An unstable vortex system is generated by the use of control surfaces. The instabilities lead to a mitigation of the vortex system. A variety of outboard flaps are used in preliminary studies to estimate the sizing and deflection of ailerons that lead to a vortex system with strong cooperative instabilities and thus to an alleviation of the rolling moment induced on following aircraft. In subsequent experiments a rectangular wing is equipped with preset ailerons, producing an unstable vortex system in its wake. The instabilities are excited by oscillating the ailerons. Each half-wing is equipped with two adjacent ailerons that oscillate in counter-phase in order to minimise fluctuations of the lift coefficient.

遷音速領域での超音速航空機の前縁フラップによる揚抗比の改善について

Wind tunnel tests were conducted to investigate lift to drag ratio improvement by the leading-edge flap the outboard wing on an SST model at transonic regions. Force measurements and surface pressure measurements were performed for the SST model with and without outboard leading-edge flaps of 5 and 12.2 degrees deflection angles. The lift to drag ratio was improved due to a reduction in the drag component when flaps deflected, because the flow was attached to the leading edge surface of the wing. The optimum flap deflection angles to attain the maximum lift to drag ratio at a fixed lift coefficient were estimated using experimental results.

Noise Characteristics of Aircraft High Lift Systems

The characteristics of far-field noise from aircraft high lift systems are discussed. Analyses are made of the data from an airframe noise test conducted in the 40 by 80 ft wind tunnel at NASA Ames Research Center, using a 4.7% DC-10 aircraft model. Discussions are given on both far-field free microphone data and measurements from a phased microphone array. Major trends are revealed and discussed from the free microphone data, which include far-field frequency characteristics, dependence of the acoustic radiation on flow Mach numbers, effects of flap and/or slat deployment, and far-field directivity. Data from the phased microphone array are used to identify locations of major noise sources. Four subregions on the wing are identified as important source locations, namely, the leading-edge slat region, the inboard and outboard side edges of the outboard flap, and the inboard flap region close to the hub of the wing. The source strength distributions in these subregions are integrated to reveal dependencies of various noise sources on flow conditions and high lift system configurations. The effects of flap side edge fences on far-field noise are also discussed, which shows a reduction of a few decibels in flap-related noise.

Studies on the influence of outboard flaps on the vortex wake of a rectangular wing

Aircraft trailing vortices constitute a hazard to following aircraft, and are therefore one of the main concerns for airport capacity constraints. At the Institute of Aerospace Engineering (ILR) experiments on wake vortices up to a distance of 60 spans behind the model of a rectangular wing are conducted in a towing tank. The motivation behind the presented experiments is the alleviation of the rolling moment induced on following aircraft by constructive means at wings and flaps of the preceeding aircraft. Following the approach of Ortega et al., the influence of the geometry of outboard flap extensions on the wake of a rectangular wing is investigated. Towing tank experiments using Particle Image Velocimetry are conducted, and the induced rolling moment is calculated from the experimentally determined downwash. In addition a stability analysis is performed for the experimentally investigated four-vortex configurations utilizing the linearized Biot–Savart law as proposed by Crow. It is assumed that the vortex system attunes to the most unstable eigenform. The theoretical results help giving an interpretation of the experimental data. It is shown that a significant reduction of the induced rolling moment by more than 50% can be achieved by the use of outboard flap extensions. This is due to an increase of the effective vortex size resulting from the counter-rotating vortex pairs produced by the flaps.

Experimental study of the instability of unequal-strength counter-rotating vortex pairs

A rapidly growing instability is observed to develop between unequal-strength counter- rotating vortex pairs. The vortex pairs are generated in a towing tank in the wakes of wings with outboard triangular flaps. The vortices from the wing tip and the inboard tip of the flap form the counter-rotating vortex pair on each side of the wing. The flow fields are studied using flow visualization and particle image velocimetry. Both chord- based and circulation-based Reynolds numbers are of O(105). The circulation strength ratios of the flap- to tip-vortex pairs range from −0.4 to −0.7. The initial sinuous stage of the instability of the weaker flap vortex has a wavelength of order one wing span and becomes observable in about 15 wing spans downstream of the wing. The nearly straight vortex filaments first form loops around the stronger wing-tip vortices. The loops soon detach and form rings and move in the wake under self-induction. These vortex rings can move to the other side of the wake. The subsequent development of the instability makes the nearly quasi-steady and two-dimensional wakes unsteady and three-dimensional over a distance of 50 to 100 wing spans. A rectangular wing is also used to generate the classical wake vortex pair with the circulation ratio of −1.0, which serves as a reference flow. This counter-rotating vortex pair, under similar experimental conditions, takes over 200 spans to develop visible deformations. Velocity, vorticity and enstrophy measurements in a fixed plane, in conjuction with the flow observations, are used to quantify the behaviour of the vortex pairs. The vortices in a pair initially orbit around their vorticity centroid, which takes the pair out of the path of the wing. Once the three-dimensional interactions develop, two-dimensional kinetic energy and enstrophy drop, and enstrophy dispersion radius increases sharply. This rapid transformation of the wake into a highly three-dimensional one offers a possible way of alleviating the hazard posed by the vortex wake of transport aircraft.

Computation of Viscous Flow for a Boeing 777 Aircraft in Landing Configuration

A series of Navier-Stokes simulations of a complete Boeing 777-200 aircraft configured for landing is obtained using a structured overset grid process and the OVERFLOW CFD code. At approach conditions, the computed forces for the 777 computation are within 1.5% of experimental data for lift, and within 4% for drag. The computed lift is lower than the experiment at maximum-lift conditions, but shows closer agreement at post-stall conditions. The effect of sealing a span-wise gap between leading edge elements, and adding a chine onto the nacelle is computed at a high angle of attack. These additions make a significant difference in the flow over the wing near these elements. Detailed comparisons between computed and experimental surface pressures are shown. Good agreement is demonstrated at lower angles of attack, including a prediction of separated flow on the outboard flap.

Evaluation of the In-Service Performance Behavior of Honeycomb Composite Sandwich Structures

When honeycomb composite structures are fabricated for the aerospace industry, they are designed to be closed to their operating environment for the life of the composite structure. However, once in service, this design can break down. Damage can set in motion a chain reaction of events that will ultimately degrade the mechanical integrity of the composite structure. Through thermographic analysis, the tendency of honeycomb composite structures to absorb and retain water was investigated, and an attempt was made to quantify the extent of water ingression in the Boeing 767 aircraft. Through thermographic analysis, the exterior honeycomb composite structures were found to contain less than 50 kg of water per plane. On average, over 90% of the water found on an aircraft was contained in five problematic parts, which included the outboard flap wedge, the nose landing gear doors, the main landing gear doors, the fixed upper wing panels, and the escape slide door. Kevlar lamina induced microcracking, skin porosity problems, and cracked potting compound were the root causes of water ingression and migration in these structures. Ultimately, this research will aid in the fundamental understanding and design of future honeycomb composite sandwich structures.

Effects of structural nonlinearities on flutter characteristics of the CF-18 aircraft

The describing function method is used to analyze the flutter characteristics of the CF-18 aircraft with structural nonlinearities. The first nonlinearity studied is located at the CF-18 wing-fold hinge. From ground test data, this hinge can be represented by a bilinear spring. A flutter sensitivity study is carried out, which shows that when the hinge stiffness is reduced, divergent flutter involving the wing bending and torsion modes is replaced by limit-cycle flutter of the wing torsion and outer wing rotation modes. Another form of nonlinearity at the outboard leading-edge flap is also studied by treating it as a spring with free-play. Limit-cycle oscillations are possible only within a small range of velocities in the vicinity of its linear flutter velocity. The flutter modes remain unchanged and their frequencies are practically constant for hinge-stiffness values ranging from a few percent to its nominal value. Positive aileron angles are found to be more effective in alleviating limit-cycle flutter at the wing-fold than negative angles. Similar observations are made at the outboard leading-edge flap hinge where downward deflection of the aileron gives larger values of the preload and, hence, the equivalent stiffness than upward deflection.