scholarly journals Morphing Wing Droop Nose with Large Deformation: Ground Tests and Lessons Learned

Aerospace ◽  
2019 ◽  
Vol 6 (10) ◽  
pp. 111 ◽  
Author(s):  
Srinivas Vasista ◽  
Johannes Riemenschneider ◽  
Ralf Keimer ◽  
Hans Peter Monner ◽  
Felix Nolte ◽  
...  
Keyword(s):  
Three Dimensional ◽  
Leading Edge ◽  
Lessons Learned ◽  
Morphing Wing ◽  
Hybrid Layer ◽  
High Lift ◽  
Airframe Noise ◽  
Design Synthesis ◽  
Bending Resistance ◽  
Curvature Distribution

A design for a new high lift system that features a morphing wing leading edge “droop nose” has the potential to generate high lift coefficients whilst mitigating airframe noise emissions. This seamless, continuous, and stepless flexible droop nose potentially offers improvements to stall and compressor requirements for an internally-blown active Coandă trailing edge flap. A full-scale, span-trimmed three-dimensional droop nose was manufactured and ground-tested based on results obtained from new design synthesis tools. A new component of the droop nose is the hybrid fiberglass-elastomeric skin that is tailored in stiffness to meet morphing curvature requirements and spanwise bending resistance. A manufacturing concept of the novel skin was established that led to an adequate manufacturing quality. The skin was driven and supported by two optimized kinematic ribs and conventional actuators and overall shape results show good agreement apart from the region closest to the leading edge. Kinematic trajectory measurements showed that the kinematics met the target trajectories well, with and without the influence of the skin, and it was deemed that the error in curvature is due to a higher than expected skin stiffness in the hybrid layer. Calculated actuator torque levels and strain measurements corroborate this inference. The lessons learned show that means of adjustment post-assembly are needed, and a reduction of torque, energy and a better curvature distribution may be achieved if the skin at the spar junction is allowed to move relative to the main wing. Careful aerodynamic, structural, actuation and manufacturing trade-off studies would be needed to determine the overall performance benefit.

Control of Separated Boundary-Layer on Subsonic Airfoil Using Vortex Generator Jet Devices

2006 ◽  
Author(s):  
A. Kourta ◽  
G. Petit ◽  
J. C. Courty ◽  
J. P. Rosenblum
Keyword(s):  
Boundary Layer ◽  
Friction Coefficient ◽  
Skin Friction ◽  
Three Dimensional ◽  
Vortex Generator ◽  
Leading Edge ◽  
Skin Friction Coefficient ◽  
Synthetic Jets ◽  
High Lift ◽  
Longitudinal Vortices

The control of subsonic high lift induced separation on airfoil may improve the flight envelope of current aircraft or even simplify the complex and heavy high-lift devices on commercial airframes. Until now, synthetic jets have proved a really interesting efficiency to delay or remove even leading-edge located separated areas on high-lift configuration but are not efficient for real scale aircrafts. In case of pressure-like separation (i.e. from trailing-edge), synthetic jets can be replaced by so the called “Vortex Generator Jets” which create strong longitudinal vortices that increase mixing in inner boundary layer and consequently the skin friction coefficient is increased to prevent separation. In this study, numerical simulations were undertaken on a generic three dimensional flat plate in order to quantify the effect of the longitudinal vortices on the natural skin friction coefficient. Both counter and co-rotative devices were tested at different exhaust velocities and distances between each others. Finally co-rotative vortex generators jets were tested on a three dimensional generic airfoil ONERA D. Results show a delay of the separation occurence but this solution does not seem to be as robust as synthetic jets. The study of jets spacing with respect to the efficiency of the devices shows a maximum for a given ratio of spacing to exhaust velocity.

Noise Reduction in a High Lift Wing Using SMAs: Computational Fluid-Structural Analysis

2016 ◽  
Author(s):  
William Scholten ◽  
Ryan Patterson ◽  
Darren Hartl ◽  
Thomas Strganac ◽  
Jeff Volpi ◽  
...  
Keyword(s):  
Structural Model ◽  
Leading Edge ◽  
Fluid Models ◽  
Cfd Analysis ◽  
High Lift ◽  
Airframe Noise ◽  
Weakly Coupled ◽  
Tunnel Model ◽  
Slat Noise ◽  
Flow Speeds

The leading-edge-slat on an aircraft is a significant contributor to the airframe noise during the low speed maneuvers of approach and landing. It has been shown in previous work that the slat noise may be reduced with a slat-cove filler (SCF). The objective of this current work is to determine how the SMA SCF behaves under steady flow using finite element structural models and finite volume (FV) fluid models based on a scaled wind tunnel model of a newly considered multi-element wing with a SCF. Computational fluid dynamics (CFD) analysis of the wing is conducted at multiple angles of attack, different flow speeds and high lift device deployment states. The FV fluid models make use of overset meshes, which overlap a slave mesh (that can undergo movement and deformation) unto a fixed master mesh, allowing for retraction and deployment of the slat and flap in the CFD analysis. The structural and fluid models are linked using a previously developed framework that permits the use of custom user material subroutines (for superelastic response of the SMA material) in the structural model, allowing for the performance of fluid-structure interaction (FSI) analysis. The fluid and structural solvers are weakly coupled such that the fluid solver transfers pressure data and the structural solver transfers displacements, but the physical quantities of each program are solved independently. FSI results are shown for the cases of the slat/SCF in the fully-deployed configuration as well as for the case of the slat/SCF undergoing retraction in flow.

Three-dimensional design of a large-displacement morphing wing droop nose device

2018 ◽  
Vol 29 (16) ◽  
pp. 3222-3241 ◽  
Author(s):  
Srinivas Vasista ◽  
Felix Nolte ◽  
Hans Peter Monner ◽  
Peter Horst ◽  
Marco Burnazzi
Keyword(s):  
Structural Design ◽  
Three Dimensional ◽  
Leading Edge ◽  
Maximum Error ◽  
Large Displacement ◽  
Lightning Strike ◽  
Morphing Wing ◽  
Optimization Stage ◽  
Kinematic Mechanisms ◽  
Kinematic Optimization

The numerical three-dimensional structural design of a large-displacement flexible morphing wing leading edge, otherwise known as a droop nose, is presented in this article. The droop nose is an essential component of a novel internally blown high-lift system for a transport aircraft to delay stall and reduce internal compressor requirements. A design chain consisting of optimization procedures was used to arrive at the structural design of the droop nose composed of a composite fiberglass skin with integral stringers and supporting kinematic mechanisms. The optimization tools aim to produce a design with minimal error to the critical target shapes. A maximum final error of 10.09 mm between calculated and target trajectories of the stringers was found after the kinematic optimization stage. After inputting the kinematic optimization results into the skin optimization stage and solving, a maximum error in the order of 13 mm and curvature difference 0.0028 1/mm were calculated, occurring in the outboard region. Prior two-dimensional analyses with similar shape deviations showed 0.4% lift reduction though further three-dimensional investigations are required. Concepts for integrating industrial requirements abrasion and lightning strike protection and in-flight de-icing into a multifunctional skin show promise and the resulting aerodynamic surface quality was found to be adequate.

scholarly journals Analisis karakteristik aerodinamika telescopic wing dengan konfigurasi sayap flying wing dan glider menggunakan metode pendekatan simulasi CFD

2021 ◽  
Vol 10 (1) ◽  
Author(s):  
Azhim Asyratul Azmi ◽  
Satriawan Dini Hariyanto ◽  
Arif Hidayat
Keyword(s):  
Cfd Simulation ◽  
Fluid Dynamic ◽  
Angle Of Attack ◽  
Leading Edge ◽  
Aerodynamic Characteristics ◽  
Morphing Wing ◽  
Aircraft Wing ◽  
High Lift ◽  
Tip Vortices ◽  
Wing Tip

A telescopic wing is a shape-changing method of the aircraft wing known as the morphing wing system. Wingspan extends capability on telescopic wing increasing the aspect ratio to get a high lift force. The telescopic wing on a flying wing configuration as an external wing and glider wing as an internal wing can be used to increase the coefficient lift (CL) when carrying out special missions. The aerodynamic characteristics using the Computational Fluid Dynamic (CFD) simulation approach is presented. For the 40% internal wingspan, the highest CL increment was 12.9% at a 10o angle of attack. For the 50% internal wingspan, the highest CL increment was 14.9% at a 10o angle of attack. on the 40% internal wing, the highest coefficient drag (CD) increment was 4.7%, and the largest CD increment on 50% internal was 9.5% at the angle of attack of 20o. The pressure distribution along the internal wingspan was uneven from an angle of attack of 15o due to the wing tip vortices of the external wing. Streamline pattern shown a bubble separation from the leading edge at an internal wing root by external wing tip vortices.Keywords: Morphing wing, telescopic wing, flying wing, glider

scholarly journals The three–dimensional leading–edge vortex of a ‘hovering’ model hawkmoth

1997 ◽  
Vol 352 (1351) ◽  
pp. 329-340 ◽  
Author(s):  
Coen van den Berg ◽  
Charles P. Ellington
Keyword(s):  
Wing Length ◽  
Three Dimensional ◽  
Axial Flow ◽  
Leading Edge ◽  
Tip Vortex ◽  
Flow Visualisation ◽  
High Lift ◽  
Leading Edge Vortex ◽  
Edge Vortex ◽  
Wing Chord

Recent flow visualisation experiments with the hawkmoth, Manduca sexta , revealed small but clear leading–edge vortex and a pronounced three–dimensional flow. Details of this flow pattern were studied with a scaled–up, robotic insect (‘the flapper’) that accurately mimicked the wing movements of a hovering hawkmoth. Smoke released from the leading edge of the flapper wing confirmed the existence of a small, strong and stable leading–edge vortex, increasing in size from wingbase to wingtip. Between 25 and 75 % of the wing length, its diameter increased approximately from 10 to 50 % of the wing chord. The leading–edge vortex had a strong axial flow veolocity, which stabilized it and reduced its diamater. The vortex separated from the wing at approximately 75 % of the wing length and thus fed vorticity into a large, tangled tip vortex. If the circulation of the leading–edge vortex were fully used for lift generation, it could support up to two–thirds of the hawkmoth's weight during the downstroke. The growth of this circulation with time and spanwise position clearly identify dynamic stall as the unsteady aerodynamic mechanism responsible for high lift production by hovering hawkmoths and possibly also by many other insect species.

scholarly journals Advanced Design of a Full-Scale Active Morphing Droop Nose

2020 ◽  
Vol 2020 ◽  
pp. 1-19
Author(s):  
Alessandro De Gaspari ◽  
Vittorio Cavalieri ◽  
Sergio Ricci
Keyword(s):  
Complete Solution ◽  
Three Dimensional ◽  
Leading Edge ◽  
Element Model ◽  
High Lift ◽  
Bird Impact ◽  
Dimensional Shape ◽  
Definition Of ◽  
Three Dimensional Shape ◽  
Previous Phase

This paper describes the design of a morphing droop nose conceived to increase the performance in high-lift conditions of a twin-prop regional aircraft, while ensuring the natural flow laminarity of the wing. Starting from the results obtained in a previous phase, mainly concerned with the performance augmentation, a detailed structural design is conducted. The main aim is the achievement of a feasible solution based on the use of conventional materials, such as aluminium alloy for the internal structure and glass-fibre for the skin. A finite element model of the complete device is generated for the three-dimensional shape quality evaluation and for the skin structural verification. Stress analyses on high-fidelity models of the single adaptive ribs are also performed. At the same time, various design aspects are evaluated, such as installation and inspection issues, actuation power, and weight considerations. All these requirements contribute to the definition of an advanced and complete solution for the device, up to the realization of a detailed CAD model. Final verification on the virtual prototype assesses the functionality of the device when attached to the wing-box. Moreover, the bird impact safety of the leading edge is demonstrated according to the certification rules.

In-Flight Deployable Micro Devices in Noise Control: Design and Evaluation

2011 ◽  
Author(s):  
Nesrin Sarigul-Klijn ◽  
Brian C. Kuo
Keyword(s):  
Noise Reduction ◽  
Sound Pressure Level ◽  
Sound Pressure ◽  
Aerodynamic Force ◽  
Three Dimensional ◽  
Pressure Level ◽  
Frequency Range ◽  
High Lift ◽  
Airframe Noise ◽  
Study Results

In this paper, time-accurate RANS simulations and FWH acoustic analogy were carried to study the three-dimensional unsteady flowfield and acoustic components around a three-element high-lift wing with and without micro devices. Micro devices are designed to be attached to the pressure side of the high lift surface near its trailing-edge to help reduce the noise generated. The analysis revealed that with the deployment of the micro device, along with reduced high-lift device setting angles, an overall airframe noise reduction of 2–5 dB is obtained over the entire frequency range. Noise reduction in the mid-frequency range, where human hearing is the most sensitive to, was particularly evident. As seen in an earlier 2D study by the authors, the application of the micro device caused strong aerodynamic force oscillations, resulting in a tone spike at a very low frequency. However, looking at the A-weighted scale sound pressure level spectrum, noise sources from the high-lift devices still dominated and it was the slat noise which dominated the overall 1/3 octave band sound pressure level. Through the reduced high-lift setting angles and the micro device application, an overall 2.3 dB noise reduction was achieved. Based on the current three-dimensional and the previous two-dimensional acoustic study results, micro devices designed by the authors demonstrated its potential to be applied onto commercial airliners as well as any aerial platforms for the use in airframe noise reduction during approach to landing phase of flight.

Two- and Three-Dimensional Prescribed Surface Curvature Distribution Blade Design (CIRCLE) Method for the Design of High Efficiency Turbines, Compressors, and Isolated Airfoils

2013 ◽  
Vol 135 (4) ◽  
Author(s):  
T. Korakianitis ◽  
M. A. Rezaienia ◽  
I. A. Hamakhan ◽  
A. P. S. Wheeler
Keyword(s):  
Turbine Blade ◽  
High Efficiency ◽  
Turbine Blades ◽  
Circle Method ◽  
Three Dimensional ◽  
Leading Edge ◽  
Surface Curvature ◽  
Blade Design ◽  
Compressor Blades ◽  
Curvature Distribution

The prescribed surface curvature distribution blade design (CIRCLE) method is presented for the design of two-dimensional (2D) and three-dimensional (3D) blades for axial compressors and turbines, and isolated blades or airfoils. The original axial turbine blade design method is improved, allowing it to use any leading-edge (LE) and trailing-edge (TE) shapes, such as circles and ellipses. The method to connect these LE and TE shapes to the remaining blade surfaces with curvature and slope of curvature continuity everywhere along the streamwise blade length, while concurrently overcoming the “wiggle” problems of higher-order polynomials is presented. This allows smooth surface pressure distributions, and easy integration of the CIRCLE method in heuristic blade-optimization methods. The method is further extended to 2D and 3D compressor blades and isolated airfoil geometries providing smooth variation of key blade parameters such as inlet and outlet flow angles, stagger angle, throat diameter, LE and TE radii, etc. from hub to tip. One sample 3D turbine blade geometry is presented. The efficacy of the method is examined by redesigning select blade geometries and numerically evaluating pressure-loss reduction at design and off-design conditions from the original blades: two typical 2D turbine blades; two typical 2D compressor blades; and one typical 2D isolated airfoil blade geometries are redesigned and evaluated with this method. Further extension of the method for centrifugal or mixed-flow impeller geometries is a coordinate transformation. It is concluded that the CIRCLE method is a robust tool for the design of high-efficiency turbomachinery blades.

Assessment of the Academic CFD Code Galatea-I With the DLR-F11 Model in High Lift Configuration

2016 ◽  
Author(s):  
Sotirios S. Sarakinos ◽  
Georgios N. Lygidakis ◽  
Ioannis K. Nikolos
Keyword(s):  
Mach Number ◽  
Incidence Angle ◽  
Three Dimensional ◽  
Leading Edge ◽  
Navier Stokes ◽  
Test Case ◽  
High Lift ◽  
Low Mach Number ◽  
Flow Phenomena ◽  
Number Flow

In this study the development and assessment of an academic CFD (Computational Fluid Dynamics) code, named Galatea-I, is reported. The proposed solver employs the RANS (Reynolds-Averaged Navier-Stokes) approach, modified by the artificial compressibility method, along with the SST (Shear Stress Transport) turbulence model to predict steady or unsteady turbulent incompressible flow phenomena on three-dimensional unstructured hybrid grids, composed of prismatic, tetrahedral and pyramidal elements. Parallel processing and an agglomeration multigrid method have been included for the acceleration of the solver’s methodologies. Galatea-I is evaluated against a test case of the HiLiftPW-2 (Second High Lift Prediction Workshop). In particular, the low Mach number flow at 7° incidence angle over the DLR-F11 aircraft configuration of Case 1 of the aforementioned workshop was examined; it considers a three-element wing with a leading edge slat and a trailing edge flap attached on a body pod, without including though any of the support brackets used in the wind tunnel experiments. The obtained results are close to the available experimental data, as well as the numerical results of other reference solvers, indicating the proposed methodology’s potential to predict accurately such low Mach number flows over complex geometries.

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

2019 ◽  
Vol 2019 ◽  
pp. 1-21 ◽  
Author(s):  
Alessandro De Gaspari ◽  
Frédéric Moens
Keyword(s):  
Laminar Flow ◽  
Tracking System ◽  
Three Dimensional ◽  
Design Procedure ◽  
Leading Edge ◽  
Optimization Procedure ◽  
Shape Design ◽  
High Lift ◽  
Aerodynamic Shape ◽  
Natural Laminar Flow

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.

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