Aerofoil Model Research Articles

GESO based RISE Controller for Active Flutter Suppression for Aeroelastic System

Abstract The paper presents a Robust Integral of Signum of Error (RISE) controller for active suppression of flutter of a two dimensional aerofoil and it is integrated with a General Extended State Observer (GESO). Towards this, aerofoil model in state space is first changed into a canonical form. Then, the controller is designed using RISE control technique. As equations of system dynamics are not into integral chain form, a GESO is formulated to observe system states and disturbances. The controller is implemented using estimated states and it is robustified using estimated disturbance. Stability of the proposed GESO-RISE controller is established using the Lyapunov theory. Simulations are performed to check the proposed controller's performance against variations in airspeed, uncertainties in model parameters, external disturbances, time delay and unmodeled dynamics. Comparison of the proposed GESO-RISE controller is carried out with existing controllers using two performance criteria i.e. Control Efforts and Integral of Absolute Error (IAE). It is discovered from simulations that the proposed GESO-RISE controller significantly reduces IAE and control efforts. Additionally, Monte Carlo method is utilized to check robustness of the proposed GESO-RISE controller. The proposed GESO-RISE controller significantly enhances flutter boundary of aerofoil and is completely implementable.

Numerical Studies on Fixed Wing Aircraft Aerodynamic Performance using Injection Suction Mechanism

In this current research work, a NACA 6321 flat bottom aerofoil has been altered to study its aerodynamic performance. A co-flow jet (CFJ) method was used with an injection slot located at 20% of chord length and the suction slot situated at 80% of chord length. Simulations have been carried out on the aerofoil models with various angles of attack (AOA) at a Mach number 0.05 to compute the variation in coefficient of lift (CL) and coefficient of drag (Cd) using a k-omega SST model. From the study, it was observed that CL and stalling AOA of an aircraft are increased by 27% and 33% respectively, with an introduction of the additional mass flow rate of 0.01 kg/sec in comparison with baseline aerofoil. It could be concluded that CFJ method enhanced the aerodynamic performance of fixed-wing aircraft due to boundary layer control over an aerofoil which reduces the overall aircraft weight and maintenance cost.

AERODYNAMIC ANALYSIS IN THE DESIGN OF AN ELECTRIC VEHICLE MODEL TOBACCO STYLE M-164 WITH COMPUTATIONAL FLUID DYNAMIC (CFD) METHOD

The aerodynamic aspect is one of the most important things in the automotive sector which is used to find information on the performance of an aerofoil model design. The performance of an aerofoil through streamflow associated with fuel consumption which means the higher the air speed, the greater the resistance received, so that the fuel consumption will be greater. At this case, fuel consumption can be reduced by creating an aerofoil model design that maintain great aerodynamic to minimize drag forces. The affects of streamflow around the vehicle are discussed in this papper. This research simulated 3D electric vehicle Tobacco Style M-164 in steady condition with various velocities, i.e. 50 km/h, 60 km/h, 70 km/h, and 80 km/h. This simulation use the Tethahedron mesh model and run in SST k-omega turbulence model. The affects can be observed with the quantitative and qualitative data. The quantitative data used as measurable data were Maximum Fluid Pressure, Drag Force, and Coefficient of Drag (CD). The quantitative data is shown to provide a better visual explanation of the streamflow affects. The qualitative data shown in this paper are velocity contours, vectors, and pathlines. The value of the maximum fluid pressure and drag force is directly proportional to the increase in velocity stream. The coefficient of drag decreased as the free stream increased with a percentage decrease of 2.48%. The average value of the coefficient of drag (CD) from this research was 0.318.

Cross validation of the aerodynamic and acoustic measurements in two Kevlar-walled wind tunnels

The wind tunnel with Kevlar walls is becoming more and more popular for aeroacoustic measurements, because the necessary corrections to convert aerodynamic coefficients to free field values are moderate and at the same time the signal to noise ratio of acoustic measurements are in the same order as in an open jet configuration. Yet, the technology is new and the correction methods for aerodynamic and acoustic coefficients are not well established. This paper aims on bench-marking two of the largest Kevlar-walled wind tunnels by cross validation of the measurements performed on a NACA633018 aerofoil model. The measured aerodynamic coefficients showed and excellent agreement in the attached flow regime. Maximum lift was predicted differently in the two facilities. The differences might be due to three-dimensional effects for separated flow and the fact that the aspect ration was different in the two facilities. The difference of the measured overall sound pressure levels was less ± 1 dB for attached flow conditions. However, we observed a difference in the slope of the sound pressure spectra and recommend to further investigate the acoustic correction methods.

Aerodynamics analysis of RAF family aerofoil using computational fluid dynamics

Aerofoil is a design that offers a gold standard ratio between the coefficient of drag to the coefficient of lift. Aerofoil are the standard cross-sectional structure of the plane wing. In this paper, a famous aerofoil family which is acknowledged to be the Royal Aircraft Factory (RAF) is regarded with all the accessible aerofoil models in that family. The pinnacle 5 aerofoils are considered and inspected for all drift editions over the viewed models with various flow transforming from 400, 420, 440, 460, 480 and 500 knots. The current research is to determine the high-quality aerofoil sketch over their family, for the quality aerofoil format by using plan amendment, and the float variation is carried out by varying the angle of assault with 5 stages interval from 0-25 ranges and the pressure, velocity, turbulence length, temperature, Mach number and the pressure performing in x & y-direction are recorded. All these parameters in consideration determined the most fulfilling model for the RAF family.

"Influence of Injection on Surface Pressure Coefficients for an Aerofoil Model".(Dept.M)

"Influence of Injection on Surface Pressure Coefficients for an Aerofoil Model".(Dept.M)

Adjoint based optimisation for efficient VAWT blade aerodynamics using CFD

The objective of this work is to demonstrate the viability of applying Adjoint methods to aerodynamic optimisation of VAWTs. Adjoint methods are very powerful optimisation techniques which have been implemented effectively in other fields, yet there is an absence of such work within VAWT literature.A ‘semi-transient’ optimisation process is proposed, using Adjoint optimisation data from single instances in time to improve VAWT performance. This is challenging due to the unsteady nature of VAWT aerodynamics. A pitching aerofoil model approximates the VAWT flow field, drastically reducing computational cost. Details are given on the necessary CFD model(s), Adjoint solver settings, and optimisation philosophy.The optimisation process was applied to a typical VAWT in the commercial CFD software ANSYS Fluent. A high tip-speed-ratio case is chosen to minimise unsteady flow affects. The results show novel blade geometries which improve the VAWT average power coefficient when compared to the original NACA0018 blade.Such a method is novel in the field of VAWTs, and the use of Adjoint methods with low cost CFD models provides an efficient optimisation methodology that can be readily adopted by the VAWT design community. This work sets the foundation for a new and very promising avenue for VAWT research.

Numerical and Experimental Investigation of NACA-0018 Wind Turbine Aerofoil Model Performance for Different Aspect Ratios

Rüzgâr enerjisinin yararlı enerjiye dönüştürülmesinde kullanılan rüzgâr türbinleri, farklı kanat modellerinden oluşmaktadır. Türbin performansını etkileyen en önemli etkenlerden biri kanat modeli aerodinamik performansının değişimidir. Rüzgâr türbin kanatlarında kullanılması muhtemel olan NACA-0018 kanat modelinin aerodinamik performansı bu çalışma kapsamında, sayısal ve deneysel olarak incelenmiştir. Performans analizi için yapılan sayısal çalışmalar hesaplamalı akışkanlar dinamiği (HAD) esasına göre çalışan ANSYS FluentTM 14,5 yazılımında SST (Shear Stress Transport) türbülans modeli altında incelenmiştir. Sayısal çalışmalarda Reynolds (Re) sayısı 5,7x104 kabul edilmiş, 0°’den 60°’ye kadar her 2,5°’lik hücum açısı için analizler tekrarlanmıştır. Deneysel çalışmalar ise açık çevrimli rüzgâr tünelinde her 5° hücum açısı için 0°-60° aralığında gerçekleştirilmiştir. Her iki çalışmada da belirlenen hücum açılarında kanat modelinin kaldırma katsayısı (CL), sürükleme katsayısı (CD) ve aerodinamik verimlilik (CL/CD) değerleri bulunmuştur. Sayısal sonuçlara göre açıklık oranı-1 (AR1) kanat modelinde 32,5°’de irtifa kaybı gözlenirken, açıklık oranı-2 (AR2) kanat modelinde ise 25°’de irtifa kaybı söz konusudur. AR1 ve AR2 kanat modelleri için yapılan deneysel çalışma sonuçlarına göre her iki kanadın CL değeri, sayısal çalışmalar neticesinde elde edilen verilerden sırasıyla %0,41 ve %2,71 oranında daha olumludur. Benzer şekilde deneysel olarak elde edilen CD değerlerinin AR1 ve AR2 kanat modeli için sayısal verilerden sırasıyla %6,35 ve %5,16 kadar daha iyi olduğu sonucuna ulaşılmıştır. Sayısal çalışma sonucu AR1 ve AR2 kanat modelleri için elde edilen CL/CD değerleri aynı yapılar için ölçülen deneysel verilerden her bir kanat modeli için sırasıyla %3,86 ve %12,04 daha yüksektir. Hem deneysel hem de sayısal çalışma sonucunda NACA-0018 kanat modelinin iki farklı açıklık oranına sahip yapılarından AR1 yapısının aerodinamik verimi, irtifa kaybı öncesi ve sonrasında AR2 kanadına göre belirgin bir avantaja sahiptir.

Turbulent Triggers and the Model Quality Influence on Aerodynamic Characteristics of the Laminar Aerofoil in Transonic Flow Regime

Abstract The test with a roughness application on the laminar aerofoil has been conducted in the N-3 trisonic wind tunnel of the Institute of Aviation in Warsaw. The main goal of tests was to investigate the influence of the boundary layer transition triggers on a laminar profile aerodynamic characteristic. For baseline configuration, the natural transition was applied. As a local roughness on the upper model surface, the carborundum strips with different heights were applied. These were positioned on the upper model surface in the front of the shock position occurrence. The Mach number during test was equal Ma = 0.7 and Reynolds number was about 2.85·106. Tests have been conducted for different model incidence in range 0°-7°. Current article refers partially to the previous study, where aerofoil model with lower quality of surface had been tested. Investigation results from previous work indicated that some of transition positions improved an aerodynamic characteristic by reducing the drag coefficient value and decreasing shock wave unsteadiness in the transonic regime. However, current article indicates that beneficial effects in respect to the baseline configuration are also strictly dependent on the model quality and turbulent triggers size. Improved surface quality of the laminar aerofoil model affected on aerodynamic characteristics with and without turbulent triggers. Resultant aerodynamic coefficients of all tested cases i.e. drag, lift and lift to drag ratio were compared.

An experimental investigation of a laminar separation bubble on the leading-edge of a modelled aerofoil for different Reynolds numbers

This paper describes the dynamics of a laminar separation bubble formed on the semi-circular leading edge of constant thickness aerofoil model. Detailed experimental studies are carried out in a low-speed wind tunnel, where surface pressure and time-averaged velocity in the separated region and as well as in the downstream are presented along with flow field visualisations through PIV for various Reynolds numbers ranging from 25,000 to 75,000 (based on the leading edge diameter). The results illustrate that the separated shear layer is laminar up to 20% of separation length and then the perturbations are amplified in the second half attributing to breakdown and reattachment. The bubble length is highly susceptible to change in Reynolds number and plays an important role in outer layer activities. Further, the transition of a separated shear layer is studied through variation of intermittency factor and comparing with existing correlations available in the literature for attached flow and as well as separated flow. Transition of the separated shear layer occurs through formation of K-H rolls, where the intermittency following spot propagation theory appears valid. The predominant shedding frequency when normalised with respect to the momentum thickness at separation remains almost constant with change in Reynolds number. The relaxation is slow after reattachment and the flow takes about five bubble lengths to approach a canonical layer.

A multi-physics computational model of fuel sloshing effects on aeroelastic behaviour

A multi-physics computational method is presented to model the effect of internally and externally-carried fuel on aeroelastic behaviour of a pitch–plunge aerofoil model through the transonic regime. The model comprises three strongly coupled solvers: a compressible finite-volume Euler code for the external flow, a two-degree of freedom spring model and a smoothed particle hydrodynamics solver for the fuel. The smoothed particle hydrodynamics technique was selected as this brings the benefit that nonlinear behaviour such as wave breaking and tank wall impacts may be included. Coupling is accomplished using an iterative method with subcycling of the fuel solver to resolve the differing timestep requirements. Results from the fuel-structural system are validated experimentally, and internally and externally-carried fuel is considered using time marching analysis. Results show that the influence of the fuel, ignoring the added mass effect, is to raise the flutter boundary at transonic speeds, but that this effect is less pronounced at lower Mach numbers. The stability boundary crossing is also found to be less abrupt when the effect of fuel is included and limit cycles often appear. An external fuel tank is seen to exhibit a lower stability boundary, while the response shows a beating effect symptomatic of two similar frequency components, potentially due to interaction between vertical and horizontal fuel motion.

Dynamic and flutter suppression analysis of 2D aerofoil by receptance method

A dynamic characteristic analysis is firstly performed for the purpose of identifying the eigenvalues problem under different velocities of a pitch and plunge degree-of-freedom aerofoil system. Then the open-loop and closed-loop for aeroelastic system is derived. The receptance method is described and shown to be very suitable for aeroelastic system. Finally, a numerical example of 2D aerofoil model is demonstrated successfully by simulating aeroelastic system with the control force excitation being provided by a control surface. The results show that the velocity of the flutter speed for the open-loop aeroelastic system is very limited, but it could be extended by active control.

Flutter control using resistively shunted piezoceramics

AbstractThis paper investigates the feasibility of flutter control using resistively shunted piezoceramics. 2D aerofoil model is elastically restrained in heave and pitch by a set of leaf springs modeled as Bernoulli-Euler beams. The piezoceramic transducers are bonded to the leaf springs. Unsteady aerodynamic theory is used to solve the flutter determinant incorporating the additive damping terms due to resistive shunting. Tuned dampers realised by resistive shunted piezoceramics in heave and pitch springs could enhance flutter speed in the range 25-50 % based on additive damping of 10 % achieved due to resistive shunting of piezoceramics in either pitch spring, heave spring or both.

A Virtual Inspection Technique for Assessing the Dimensional Accuracy of Forged Compressor Blades Using FE Modeling and CMM Inspection

This paper presents research for developing a virtual inspection system that evaluates the dimensional tolerance of forged aerofoil blades formed using the finite element (FE) method. Conventional algorithms adopted by modern coordinate measurement processes have been incorporated with the latest free-form surface evaluation techniques to provide a robust framework for the dimensional inspection of FE aerofoil models. The accuracy of the approach had been verified with a strong correlation obtained between the virtual inspection data and coordinate measurement data from corresponding aerofoil components.

A changeable aerofoil actuated by shape memory alloy springs

An aerofoil has great influences on the aerodynamic quality of aerial vehicles and efficiency of flight can be greatly improved by adopting the optimum aerofoil under different flying conditions. A changeable aerofoil model is designed and manufactured. By changing the constraint condition of the skins, they can achieve large deformation without overstepping their strain allowance. Shape memory alloy springs with the help of stop structures are used to actuate accurately certain points on the skins to approach the target aerofoil. A finite element model of the skins is built and deformation of the skins under the control of the discrete points is simulated. Deformation of the skins actuated by SMA springs is measured and the results of the experiment and simulation are compared and analyzed. A method of designing a changeable aerofoil is provided.

On wake response to asymmetric blowing in spurts from a blunt trailing edge

In the present investigation, a technique has experimentally been evolved with the aim to control the unsteady wake dominated by periodic vortex-shedding. The technique uses pulsating surface blowing that is applied asymmetrically just before separation from the blunt trailing edge of a thick aerofoil model. Thus, momentum is injected in spurts in only one of the two separated shear layers. The control parameters are the mass flow rate and the forcing frequency of the pulsating blowing. The technique is particularly effective in suppressing the vortex-shedding when the blowing is moderate and the forcing frequency is twice the natural vortex-shedding frequency. As a consequence of injecting momentum in spurts, there is a significant saving in mass flow required to achieve a condition of the momentumless wake. In spite of the initial conditions being strongly asymmetric, the flow pattern in the wake was observed to quickly assume a remarkable symmetry.

A comparison of full non‐linear and reduced order aerodynamic models in control law design using a two‐dimensional aerofoil model

AbstractThe prediction of the flutter boundary of an aircraft is a necessary but time consuming process, particularly as for the most realistic results a time accurate simulation of the interaction between the non‐linear aerodynamic and structural forces is required. Extension of the flight envelope by the design of active control laws to suppress flutter further increases the demands on computational time, to presently unrealistic levels. Use of a reduced order model (ROM) derived from, and in place of, the full non‐linear aerodynamics greatly reduces the time required for calculation of aerodynamic forces. However, this is necessarily accompanied by some loss in accuracy, and hence the method must be verified by comparison with results obtained by the full aerodynamic model before it may be used with confidence.Such a comparison is presented here, using a two‐dimensional aerofoil and control surface combination as a test case. Active control of the deflectable surface is used to attempt to increase the flutter speed across the complete Mach range, feedback control being achieved by gains acting on heave and pitch proportional and differential signals, interpreted as a hinge moment demand. Full non‐linear and reduced order aerodynamic models are then used to obtain optimum control law gain for flutter suppression. The results demonstrate that the ROM accurately predicts the open loop flutter boundary, gives a good approximation to the increase in flutter speed that may be produced by gain optimization, and produces a similar response given the identical gain values in each system for a significantly reduced cost. Copyright © 2005 John Wiley & Sons, Ltd.

Experimental investigation of trailing-edge devices at transonic speeds

AbstractThe influence of trailing-edge devices such as Gurney flaps and divergent trailing edges of different height on the aerodynamic performance of an aerofoil at transonic speeds has been investigated experimentally. The investigation has been carried out in the Transonic Wind Tunnel Göttingen (TWG) using the two-dimensional aerofoil model VC-Opt at freestream Mach numbers of M ε [0.755, 0.775, 0.790] and a Reynolds number of Re = 5.0 x 106.The results have shown that the trailing-edge devices increase the circulation of the aerofoil leading to a lift enhancement and pitching-moment decrease as well as an increase in minimum drag compared to the baseline configuration. The maximum lift-to-drag ratio is considerably improved and the onset of trailing-edge flow separation is shifted to higher lift. Besides the increased rear-loading, a downstream displacement of the shock provides the main lift enhancement in transonic flow.The simple Gurney flap provides the largest additional circulation of all geometries tested. The smoother turning of the flow due to the additional ramp of the divergent trailing edge leads to a smaller increase of circulation. Slightly less lift but considerably less viscous (pressure) drag is generated enhancing the maximum lift-to-drag ratio compared to the Gurney flap. The negative affect of the Gurney flap on the pitching moment is also reduced.For the high divergent trailing edges, different ramp slopes have a significant influence on the aerodynamic performance whereas at low device heights the influence is considerably diminished.The results show that the divergent trailing edge proves to be the better trailing-edge device at transonic speeds. The application as an element for an adaptive wing is generally possible.

Investigation of slat heel effect on the flow field over multi-element aerofoils

This paper presents results obtained from an experimental study of the flow field over a multi-element aerofoil which incorporates either a conventional or an advanced slat. Detailed measurements of the mean flow and turbulent quantities over a multi-element aerofoil model equipped with either type of slat have been made in a wind tunnel using stationary and flying hot-wire (FHW) probes. The performance of the two slats at two angles of attack, alpha = 10degrees and 20degrees, were investigated and compared with each other. The results showed a better performance for the advanced slat in terms of the mean velocity field and hence an increase in the lift performance. The advantage of the advanced slat was more pronounced for the multi-element aerofoil placed at the higher angle of attack, i.e., 20degrees. These findings were substantiated by the Reynolds stresses measured over the multi-element aerofoil, with the conventional slat exhibiting higher values compared with its advanced slat counterpart. Both the mean velocity and Reynolds stress results clearly demonstrated that the conventional slat had a lower stall margin than the advanced slat. (C) 2002 Elsevier Science Inc. All rights reserved.

An Investigation of Flow Fields Over Multi-Element Aerofoils

This paper presents results obtained from a combined experimental and computational study of the flow field over a multi-element aerofoil with and without an advanced slat. Detailed measurements of the mean flow and turbulent quantities over a multi-element aerofoil model in a wind tunnel have been carried out using stationary and flying hot-wire (FHW) probes. The model configuration which spans the test section 600mm×600mm, is made of three parts: 1) an advanced (heel-less) slat, 2) a NACA 4412 main aerofoil and 3) a NACA 4415 flap. The chord lengths of the elements were 38, 250 and 83 mm, respectively. The results were obtained at a chord Reynolds number of 3×105 and a free Mach number of less than 0.1. The variations in the flow field are explained with reference to three distinct flow field regimes: attached flow, intermittent separated flow, and separated flow. Initial comparative results are presented for the single main aerofoil and the main aerofoil with a nondeflected flap at angles of attacks of 5, 10, and 15 deg. This is followed by the results for the three-element aerofoil with emphasis on the slat performance at angles of attack α=10, 15, 20, and 25 deg. Results are discussed both for a nondeflected flap δf=0deg and a deflected flap δf=25deg. The measurements presented are combined with other related aerofoil measurements to explain the main interaction of the slat/main aerofoil and main aerofoil/flap both for nondeflected and deflected flap conditions. These results are linked to numerically calculated variations in lift and drag coefficients with angle of attack and flap deflection angle.