High-lift Devices Research Articles

Aerodynamic Shape Design and Validation of an Advanced High-Lift Device for a Regional Aircraft with Morphing Droop Nose

In the present work, the aerodynamic shape design of an advanced high-lift system for a natural laminar flow (NLF) wing, based on the combination of a morphing droop nose and a single slot trailing edge flap, is presented. The paper presents both the aerodynamic design and optimization of the NLF wing and the high-lift configuration considering the mutual effects of both flap devices. Concerning the morphing droop nose (DN), after defining the parameterization techniques adopted to describe the geometry in terms of morphing shape and flap settings, the external configuration is obtained by an aerodynamic shape optimization procedure able to meet geometrical constraints and the skin structural requirements due to the morphing. The final performance assessment of the three-dimensional high-lift configurations is performed by high-fidelity aerodynamic analyses. The design procedure is applied to a twin-prop regional aircraft equipped with a natural laminar flow wing. The morphing droop nose is compatible with an NLF wing that requires the continuity of the skin and, at the same time, extends the possibilities to improve the performances of the class of regional aircraft which usually are not equipped with conventional leading edge devices. Additionally, the morphing technology applied to the flap allows the design of a tracking system fully integrated inside the airfoil geometry, leading to a solution without external fairings and so with no extra friction drag penalty for the aircraft.

Toward Noise Certification during Design: Airframe Noise Simulations for Full-Scale, Complete Aircraft.

An overview of a recent, NASA-sponsored effort to substantially advance simulation-based airframe noise prediction is presented. An accurate characterization of this component of aircraft noise requires a high-fidelity representation of the finer geometrical details associated with the landing gear and wing high-lift devices, such as slats and flaps, which constitute major noise sources. To achieve this ambitious goal, a systematic approach was followed to extend our current state-of-the-art computational tools to a full-scale, complete aircraft in landing configuration within a realistic flight environment. The work involved several phases: high-fidelity, large-scale, unsteady flow simulations; model-scale experiments in ground-based facilities; and farfield noise prediction for a full-scale, complete aircraft. The comprehensive aeroacoustic database generated during the course of the effort provided a wealth of relevant information for full validation and benchmarking of the advanced computational tools used in the present work. The database will also foster the development of simulation methodologies with improved predictive capabilities.

A flight demonstration for airframe noise reduction technology

This article overviews activities and results for an initial flight demonstration of a project, FQUROH, to develop and mature airframe noise reduction technologies. The project aims to demonstrate technologies for reducing airframe noise through three flight tests using aircraft with modified high-lift devices and landing gear. It intends to verify the design methods that utilize advanced computational fluid dynamics (CFD) and the feasibility of practical noise reduction concepts. The main purposes of the initial demonstration using a JAXA’s research aircraft in 2016 were to establish the processes for flight testing, including modification of the aircraft and experimental procedures, as well as to evaluate preliminary noise reduction designs. In the design process, noise reduction concepts for the flap and main landing gear were successfully applied to the actual complex geometries of the aircraft using advanced CFD, and the noise reduction effects were confirmed with noise measurements in wind tunnel tests. Based on the designs, the flap and main landing gear modifications were manufactured and installed on the aircraft. The flight demonstration campaign was successfully carried out with 177 noise source measurements, and encouraging results were obtained that show the feasibility of the noise reduction design method.

Numerical studies of active flow control on wing tip extension

PurposeIn the European project AFLoNext, active flow control (AFC) measures were adopted in the wing tip extension leading edge to suppress flow separation. It is expected that the designed wing tip extension may improve aerodynamic efficiency by about 2 per cent in terms of fuel consumption and emissions. As the leading edge of the wing tip is not protected with high-lift device, flow separation occurs earlier than over the inboard wing in the take-off/landing configuration. The aim of this study is the adoption of AFC to delay wing tip stall and to improve lift-to-drag ratio.Design/methodology/approachSeveral actuator locations and AFC strategies were tested with computational fluid dynamics. The first approach was “standard” one with physical modeling of the actuators, and the second one was focused on the volume forcing method. The actuators location and the forcing plane close to separation line of the reference configuration were chose to enhance the flow with steady and pulsed jet blowing. Dependence of the lift-to-drag benefit with respect to injected mass flow is investigated.FindingsThe mechanism of flow separation onset is identified as the interaction of slat-end and wing tip vortices. These vortices moving toward each other with increasing angle of attack (AoA) interact and cause the flow separation. AFC is applied to control the slat-end vortex and the inboard movement of the wing tip vortex to suppress their interaction. The separation onset has been postponed by about 2° of AoA; the value of ift-to-drag (L/D) was improved up to 22 per cent for the most beneficial cases.Practical implicationsThe AFC using the steady or pulsed blowing (PB) was proved to be an effective tool for delaying the flow separation. Although better values of L/D have been reached using steady blowing, it is also shown that PB case with a duty cycle of 0.5 needs only one half of the mass flow.Originality/valueTwo approaches of different levels of complexity are studied and compared. The first is based on physical modeling of actuator cavities, while the second relies on volume forcing method which does not require detailed actuator modeling. Both approaches give consistent results.

Albi II – a new generation development

The UL-39 Albi university's all-composite ultralight aircraft project, powered by a piston engine and ducted fan, continues at the Department of Aerospace Engineering at Faculty of Mechanical Engineering CTU in Prague and its partners ZALL JIHLAVAN airplanes, s.r.o. and LA composite, s.r.o. by developing its new generation. The article is a follow-up to a contribution from 2017, where the entire genesis of the first prototype was described. The introduction summarizes the experience of the prototype operation and analyzes the deficiencies that required a major redesign of the propulsion unit. Aspects leading to the choice of another propulsion unit arrangement and changes in the ducted fan, airframe and systems are described. The fuselage of the airplane has undergone a dominant change. The paper describes not only structural changes leading to the reduction of the width of the fuselage and its wetted area, but also the changes in manufacturing process of composite parts leading to weight reduction. Following the changes in the fuselage design modifications of the wing (mainly high lift devices) and modification of the horizontal tail plane are described. At the end there is a plan of further development described, which should ultimately lead to the commercialization of the project.

Reliability analysis for high-lift device based on Copula function and Bayesian updating

In order to improve the reliability analysis accuracy of the aircraft high-lift, an approach based on the Copula function theory and Bayesian updating is proposed. Considering the influence of the random variables’ correlation in the process of updating, choosing the reasonable prior joint distribution and likelihood function is crucial. Under the condition of the incomplete probability information, the analytic expressions of the prior joint distribution and likelihood function of the correlated random variables are derived through the Copula function. Then, the posterior joint distribution is obtained by Bayesian updating. The reliability of the lifting device is calculated based on the posterior distribution. The case analysis shows that the reliability results based on the proposed approach are more accurate and more coincident with the factual situation than the reliability analysis results based on the independence assumption of random variables.

An improved high-lift aerodynamic prediction method for transport aircraft

The aim of this work is the development of a methodology to predict lift characteristics for transport aircraft in the whole flight envelope, useful in the preliminary aircraft design stage. The purpose is an attempt to improve the classical methodologies for wing load distribution and lift prediction, applicable to both clean and flapped configuration. This has been obtained considering the airfoils’ aerodynamic characteristics until stall and post-stall conditions during the process, and modifying 2D characteristics in the case of high-lift devices to take into account 3D effects introduced by the devices themselves. The method is a modification of standard vortex-lattice procedures which are capable of predicting wing aerodynamic characteristics. As regards the clean configuration, the enhanced method works by integrating airfoil characteristics, whereas as far as the high-lift devices’ effect is concerned, the improved method works by substituting clean airfoil aerodynamic characteristics for the flapped aerodynamics ones, and introducing a correction to evaluate the 3D effects induced by the high-lift devices’ geometrical discontinuities. The methodology is explained separately for these two configurations. The results of the developed method have been compared with CFD and experimental data showing good agreement, making available a fast and reliable method useful in preliminary aircraft design phase.

Mechanism/structure/aerodynamic multidisciplinary optimization of flexible high-lift devices for transport aircraft

Mission adaptive variable camber wing in both chord-wise and span-wise directions that can improve the aerodynamic performance during takeoff, landing and cruising flight, will be the state-of-the-art high-lift system for next generation airliners. Based on NASA TrapWing model released on the 1st AIAA CFD High-lift Prediction Workshop, a smart high-lift system with “Flexible Droop Nose & Single Slotted Hinge Flap combined with spoiler deflection & Flexible Trailing Edge Flap” is proposed in this paper. The Flexible Droop Nose is actuated by kinematic chains mechanism, the Single Slotted Hinge Flap is actuated by simple hinge mechanism and the Flexible Trailing Edge Flap is actuated by link/track mechanism.A mechanism/structure/aerodynamic multidisciplinary optimization platform based on iSIGHT software is constructed for this smart high-lift system. This platform consists of stress analysis, high-lift configuration generation, high-lift configuration structure grid generation, computational fluid dynamics and optimization algorithm modules. The optimal takeoff and landing configurations with comprehensive performance of mechanism, aerodynamic and structure is then obtained after multidisciplinary optimization. Finally, the CFD results show that the aerodynamics performance of this smart high-lift system is more effective than the original NASA TrapWing model.

Experimental study on aerodynamic noise characteristics of high-lift configuration with variable gap leading-edge slat

The aircraft high-lift devices are important components to ensure the safety of take-off and landing, and it is also the main source of airframe noise in the process of take-off and landing, especially the leading-edge slat. The slat that reduces gap by rotating with the fixed axis is proposed in this paper, and experimental study of far-field aerodynamic noise characteristics of high-lift configuration with the slat is also carried out. The experimental results show that discrete tones in the low-middle frequency range disappear and the broadband frequency amplitude also slightly decline, resulting in a significant decrease in the overall sound pressure level of the high-lift configuration with slat gap of 0, compared with the configuration with conventional leading-edge slat. However, in the process of rotation from conventional slat to seamless slat, more prominent discrete tone in low-middle frequency range or humps in mid-high frequency range appear at angles of attack below 6°, which give rise to the overall increase in the sound pressure level. At angles of attack above 6°, these aero-acoustic phenomena are not obvious. Hence, the overall sound pressure levels of high-lift configuration gradually decrease with the rotation process in larger angles of attack.

A PCA–ANN-based inverse design model of stall lift robustness for high-lift device

The concept of stall lift robustness for high-lift device (HLD) measures the stability of lift values under a series of angles of attack (AoA) around stall, which plays a significant role in flight safety. In this article, a stall lift robustness design with the consideration of aerodynamic constraints (stall AoA, average lift, etc.) is carried out, where design targets are set as a series of lift values on Lift-AoA curve. A Principle Component Analysis (PCA)–Artificial Neutral Network (ANN)-based inverse design model is introduced. The design targets are transformed by PCA for data dimension reduction. Then, the new set of design targets are input into the surrogate model of ANN, and corresponding geometry of new HLD is predicted. The ANN is constructed through database and sample points are screened considering lift unsteadiness. The design procedure is iterated to meet the design accuracy. The process of stall lift robustness design with the proposed model is discussed in this article, and the design results are validated by Detached-Eddy Simulation (DES).

Dielectric Barrier Discharge Actuators Employed as Alternative to Conventional High-Lift Devices

The performance and effectiveness of serpentine dielectric barrier discharge plasma actuators as hingeless high-lift devices on a three-dimensional airplane model were investigated. Attempts were made to design the actuators in the most optimized way, based on the reported results of the latest dielectric barrier discharge optimization efforts. Four actuators were mounted onto four different locations of the wings as the leading-edge slat, spoiler, flap, and leading-edge aileron. Operation of the actuators on one wing increased and/or decreased the lift/drag and generated a gradual rolling moment. The experiments were performed separately for each actuator at three different Reynolds numbers, seven angles of attack, three peak-to-peak voltages, and three wave frequencies. The Reynolds numbers were approximately , , and . The evaluation based on actuators effects on rolling moment generation was performed, and the results indicate that employing the serpentine dielectric barrier discharge actuators as high-lift devices for low Reynolds number applications is possible, especially as DBD slat and DBD spoiler, and they can have the same effect of a conventional aileron only for normal flight maneuvering, with low power consumption.

Experimentaland Numerical Analysis over the Truncated Airfoil with Slotted Flap Configuration

Objectives: The objective of this work is to investigate the flow over the truncated NACA2414 airfoil with slotted flap configuration. Investigation is carried using wind tunnel and commercial CFD tools with the consequences of the variation of speed, angle of attack of control surface and effective angle of attack. Methods/Statistical Analysis: Wind tunnel tests are performed on a reduced scale model (3:10) in wind tunnel at Manipal University. Experimentally determined static pressure distribution data at equal interval stations, placed over the wing are used to generate pressure data. The results obtained are used in turn as a benchmark to validate the CFD simulations. Findings: Findings clearly states that the blunt trailing edge has high drag coefficient. Although off set cavity and cavity has proven to be the ideal modification for trailing edge shape in a plane wing from the study. It can be concluded that the same is not true when such modifications are applied to slotted flap configuration. Application/Improvements: Further studies can be carried out with the alternative designs and a rigorous research can contribute in efficient designs in high lift devices. Keywords: High-lift Devices, Numerical Analysis, Slotted Flaps, Truncated Airfoil, Wind Tunnel

Leading-edge vortex development on a pitching flat plate with multiple leading edge geometries

This article investigates the dynamic stall of a pitching flat plate in order to facilitate lift control in applications such as flapping wings and micro-aerial vehicles. Constant-pitch-rate rotation, similar to ramp-and-hold, of a flat plate was investigated in an open-jet wind tunnel using embedded surface pressure sensors to assess vortex formation and its impact on aerodynamic characteristics. Three leading edge and trailing edge geometries were investigated, with all demonstrating minor effects on the overall aerodynamic performance for the angles of attack presented. For all leading and trailing edge combinations, increasing the rotation rate increased the magnitude of the localised low pressure on the upper surface the flat plate resulting in increased lift. Separation of the leading-edge vortex was characterised by the motion of a half-saddle flow feature on the upper surface of the flat plate, which in turn was linked to the rotation rate. A critical angle of attack was found whereby further increases in the angle of attack provided no benefit to overall force generation. The surface pressures shown in the current investigation can be utilised for the development of adaptive boundary layer control devices to either limit or promote the dynamic stall process on a flat plates. This study confirms that vortex lift, which is critical for flight of micro-aerial vehicles and flapping wing devices, is limited by the maximum angle of attack and rotation rate. This is significant for high-lift devices where limited power is available for lift generation.

Aeroacoustic analysis of an airfoil with Gurney flap based on time-resolved particle image velocimetry measurements

Particle image velocimetry for the experimental assessment of trailing edge noise sources has become focus of research in recent years. The present study investigates the feasibility of the noise prediction for high-lift devices based on time-resolved particle image velocimetry (PIV). The model under investigation is a NACA 0015 airfoil with a Gurney flap with a height of 6% of the chord length. The velocity fields around and downstream of the Gurney flap were measured by PIV and used to compute the corresponding pressure fields by solving the Poisson equation for incompressible flows. The reconstructed pressure fluctuations on the airfoil surface constitute the source term for Curle's aeroacoustic analogy, which was employed in both the distributed and compact formulation to estimate the noise emission from PIV. The results of the two formulations are compared with the simultaneous far-field microphone measurements in the temporal and spectral domains. Both formulations of Curle's analogy yield acoustic sound pressure levels in good agreement with the simultaneous microphone measurements for the tonal component. The estimated far-field sound power spectra (SPL) from the PIV measurements reproduce the peak at the vortex shedding frequency, which also agrees well with the acoustic measurements.

A Study of Multi-Objective Aerodynamic Optimization Design for Variable Camber Airfoils and High Lift Devices

Aiming at the synthetical optimization of the aerodynamic performance between the low-speed condition of two-dimensional high lift devices during take-off and landing phase and the high-speed condition of variable camber airfoil during cruise phase, an aerodynamic optimization design method for high lift device based on Kriging based surrogate model and multi-objective genetic algorithm has been developed. With the application of Adaptive Dropped Hinge Flap mechanism, the low-speed take-off and landing performance and high-speed cruise performance of the aircraft is improved by coupling deflection of the flap and spoiler. The position of flap hinge, deflection angle of spoiler and deflection angle of flap are taken as design variables; The Navier-Stokes equations are used to predict the aerodynamic forces of initial samples; The Kriging based surrogate model is employed to establish the algebraic relation between design variables and aerodynamic forces at take off, landing and cruise, obtaining four efficient prediction models for aerodynamic forces; Multi-objective optimization design with multi-objective genetic algorithm is conducted on the basis of surrogate models. The automatic generation of computational grid is achieved by the mesh deformation method based on RBF (Radial Basis Function) when the design variables change. On the basis of efficient global multi-objective optimization design platform, the synthetical optimization of high-speed and low-speed aerodynamic performance is conducted; The multi-objective solution set of the Pareto frontier is verified and analyzed, and the optimal solution with well matched high and low speed performance is selected.

AIRCRAFT ENVIRONMENTAL NOISE LEVEL ESTIMATION AT LANDING APPROACH

Three approximate methods are proposed for the estimated noise level of an aircraft performing approach. The noise level from the power plant is determined using existing software packages. In the first two methods, the estimation of the airframe noise level is based on modern concepts of the ratio of the acoustic noise emission from the power plant and that of the aircraft airframe, or on the known data on the relationship between the noise emissions from the chassis and the high lift devices, and the chassis noise level is calculated using known semi-empirical methods. In the third method, to estimate aircraft noise it is proposed to use experimental data on the statistically minimum environmental noise level at landing approach of an airplane with a conventional airframe layout.

Particle Image Velocimetry Investigation of the Coherent Structures in a Leading-Edge Slat Flow

Air traffic volume is expected to triple in the U.S. and Europe by 2025, and as a result, the aerospace industry is facing stricter noise regulations. Apart from the engines, one of the significant contributors of aircraft noise is the deployment of high-lift devices, like leading-edge slats. The unsteady turbulent flow over a leading-edge slat is studied herein. In particular, particle image velocimetry (PIV) measurements were performed on a scale-model wing equipped with a leading-edge slat in the H.J. Irving–J.C.C. Picot Wind Tunnel. Two Reynolds numbers based on wing chord were studied: Re = 6 × 105 and 1.3 × 106. A snapshot proper orthogonal decomposition (POD) analysis indicated that differences in the time-averaged statistics between the two Reynolds numbers were tied to differences in the coherent structures formed in the slat cove shear layer. In particular, the lower Reynolds number flow seemed to be dominated by a large-scale vortex formed in the slat cove that was related to the unsteady flapping and subsequent impingement of the shear layer onto the underside of the slat. A train of smaller, more regular vortices was detected for the larger Reynolds number case, which seemed to cause the shear layer to be less curved and impinge closer to the tail of the slat than for the lower Reynolds number case. The smaller structures are consistent with Rossiter modes being excited within the slat cove. The impingement of the shear layers on and the proximity of the vortices to the slat and the main wing are expected to be strong acoustic dipoles in both cases.

System-Level Assessment of Active Flow Control for Commercial Aircraft High-Lift Devices

The use of active flow control technology for augmentation of aircraft aerodynamic performance is of considerable interest to the aerospace industry. Research to date has included subscale and full-scale wind-tunnel experiments as well as limited flight-test demonstrations. Out of several possible applications of active flow control to commercial aircraft, this paper focuses on its application to high-lift devices, which may permit weight and drag reduction through simplification of the flap mechanism. These, however, must be traded off against the additional weight of the active flow control system, its power requirements, and redundancy considerations to assess net system-level impacts. In this work, the Integrated Subsystem Sizing and Architecture Assessment Capability is used to assess the system-level impact of multiple high-lift device active flow control architectures for a small twin-aisle aircraft. A comparative assessment of these architectures reveals a strong dependency of fuel burn performance and the identity of the best-performing architecture on the system total mass flow requirements. For a selected architecture, the impact of the system mass flow requirements on the off-design performance is investigated. Despite an increase in vehicle operating empty weight, fuel consumption still decreases for most missions in the payload–range envelope. Finally, the sensitivity of the architecture’s performance to several sources of uncertainty is ascertained through a sensitivity analysis, and it is observed that the uncertainty in estimation of flap track fairing drag has the most significant impact on the estimation of mission fuel burn.

Turbulent planar wakes under pressure gradient conditions

Accurate prediction of the spatial evolution of turbulent wake flows under pressure gradient conditions is required in some engineering applications such as the design of high-lift devices and wind farms over topography. In this paper, we aim to develop an analytical model to predict the evolution of a turbulent planar wake under an arbitrary pressure gradient condition. The model is based on the cross-stream integration of the streamwise momentum equation and uses the self-similarity of the mean flow. We have also made an experimentally supported assumption that the ratio of the maximum velocity deficit to the wake width is independent of the imposed pressure gradient. The asymptotic response of the wake to the pressure gradient is also investigated. After its derivation, the model is successfully validated against experimental data by comparing the evolution of the wake width and maximum velocity deficit. The inputs of the model are the imposed pressure gradient and the wake width under zero pressure gradient. The model does not require any parameter tuning and is deemed to be practical, computationally fast, accurate enough, and therefore useful for the scientific and engineering communities.

Concurrent wing and high-lift system aerostructural optimization

A method is presented for concurrent aerostructural optimization of wing planform, airfoil and high lift devices. The optimization is defined to minimize the aircraft fuel consumption for cruise, while satisfying the field performance requirements. A coupled adjoint aerostructural tool, that couples a quasi-three-dimensional aerodynamic analysis method with a finite beam element structural analysis is used for this optimization. The Pressure Difference Rule is implemented in the quasi-three-dimensional analysis and is coupled to the aerostructural analysis tool in order to compute the maximum lift coefficient of an elastic wing. The proposed method is able to compute the maximum wing lift coefficient with reasonable accuracy compared to high-fidelity CFD tools that require much higher computational cost. The coupled aerostructural system is solved using the Newton method. The sensitivities of the outputs of the developed tool with respect to the input variables are computed through combined use of the chain rule of differentiation, automatic differentiation and coupled-adjoint method. The results of a sequential optimization, where the wing shape and high lift device shape are optimized sequentially, is compared to the results of simultaneous wing and high lift device optimization.