Dissertations / Theses on the topic 'High-lift airfoil'

Nowadays, the competition among airlines seriously depend upon the saving operating costs, with the premise that not to degrade its services quality. Especially in the face of increasingly scarce oil resources, reducing fleets operational fuel consumption, is an important means to improve profits. Aircraft fuel economy is determined by operational management strategies and application technologies. The application of technologies mainly refers to airplane’s engine performance, Weight efficiency and aerodynamic characteristics. A market competitive aircraft should thoroughly consider to all of these aspects. Transport aircraft aerodynamic performance mainly is determined by wing’s properties. Wings that are optimized for efficient flight in cruise conditions need to be fitted with powerful high-lift devices to meet lift requirements for safe takeoff and landing. These high-lift devices have a significant impact on the total airplane performance. The aerodynamic characteristics of the wing airfoil will have a direct impact on the aerodynamic characteristics of the wing, and the wing’s effective cruise hand high-lift configuration design has a significant impact on the performance of transport aircraft. Therefore, optimizing the design is a necessary airfoil design process. Nowadays engineering analysis relies heavily on computer-based solution algorithms to investigate the performance of an engineering system. Computational fluid dynamics (CFD) is one of the computer-based solution methods which are more widely employed in aerospace engineering. The computational power and time required to carry out the analysis increases as the fidelity of the analysis increases. Aerodynamic shape optimization has become a vital part of aircraft design in the recent years. Since the aerodynamic shape optimization (ASO) process with CFD solution algorithms requires a huge amount of computational power, there is always some reluctance among the aircraft researchers in employing the ASO approach at the initial stages of the aircraft design. In order to alleviate this problem, statistical approximation models are constructed for actual CFD algorithms. The fidelity of these approximation models are merely based on the fidelity of data used to construct these models. Hence it becomes indispensable to spend more computational power in order to convene more data which are further used for constructing the approximation models. The goal of this thesis is to present a design approach for assumed wing airfoils; it includes the design process, multi-objective design optimization based on surrogate modelling. The optimization design stared from a transonic single-element single-objective optimization design, and then high-lift configurations were two low-speed conditions of multi-objective optimization design, on this basis, further completed a variable camber airfoil at low speed to high-lift configuration to improve aerodynamic performance. Through this study, prove a surrogate based model could be used in the wing airfoil optimization design.

Many flow cases in fluid dynamics face undesirable flow separation due torising static pressure on wall boundaries. This occurs e.g. due to geometry as ina highly curved turbine inlet duct or e.g. on flow control surfaces such as wingtrailing edge flaps within a certain angle of attack range. Here, flow controldevices are often used in order to enhance the flow and delay or even totallyeliminate flow separation. Flow control can e.g. be achieved by using passiveor active vortex generators (VG) that enable momentum mixing in such flows.This thesis focusses on passive VGs, represented by VG vanes that are mountedupright on the surface in wall-bounded flows. They typically have an angle ofincidence to the mean flow and, by that, generate vortex structures that in turnallow for the desired momentum mixing in order to prevent flow separation.A statistical VG model approach, developed by KTH Stockholm and FOI,the Swedish Defence Research Agency, has been evaluated computationally.Such a statistical VG model approach removes the need to build fully resolvedthree-dimensional geometries of VGs in a computational fluid dynamics mesh.Usually, the generation of these fully resolved geometries is rather costly interms of preprocessing and computations. By applying this VG model, thecosts reduce to computations without VG effects included. Nevertheless, theVG model needs to be set up in order to define the modelled VG geometry inan easy and fast preprocessing step. The presented model has shown sensitivityfor parameter variations such as the modelled VG geometry and the VG modellocation in wall-bounded zero pressure gradient and adverse pressure gradientflows on a flat plate, in a diffuser, and on an airfoil with its high-lift systemextracted. It could be proven that the VG model qualitatively describes correcttrends and tendencies for these different applications.

A redução do ruído externo gerado por aeronaves operando nas proximidades de grandes centros urbanos é apontada como uma questão vital para a manutenção e expansão sustentável das atividades da aviação civil. Nas últimas décadas, reduções significativas no ruído gerado pelos sistemas de propulsão da aeronave tornaram relevantes as contribuições do trem de pouso e dos dispositivos de hiper-sustentação (flapes e eslates) para o ruído global da aeronave. A caracterização do espectro acústico de cada componente hiper-sustentador é necessária para o desenvolvimento de métodos preditivos de ruído e projetos aerodinâmicos que viabilizem a redução de ruído sem penalizações severas para o desempenho e a segurança da aeronave. Experimentos com modelos em escalas mostram que a contribuição de cada elemento hiper-sustentador para o ruído global é determinada pelo tamanho e modelo da aeronave. Tal fato dificulta a generalização dos resultados experimentais e determina a caracterização do espectro acústico de cada componente de um aerofólio em termos de sua geometria e configuração operacional. Este trabalho tem como objetivo principal a caracterização do ruído do eslate a partir de medições experimentais com um aerofólio hiper-sustentador McDonnell Douglas (30P30N), composto por flape, eslate e elemento principal. Os experimentos foram realizados em túnel de vento de seção fechada, e as medições acústicas contaram com o uso de uma antena composta por 62 microfones. Os dados acústicos foram processados com algoritmos de beam-forming convencional e deconvolução DAMAS (Deconvolution Approach for the Mapping of Acoustic Sources). A aplicação de técnicas de beam-forming permite representar uma distribuição espacial de fontes na forma de um mapa acústico e determinar o nível de ruído gerado por fontes que concorrem de forma independente para o ruído global. A base de dados experimentais permite o estudo do ruído do eslate sob diferentes configurações operacionais e geométricas do aerofólio. A análise do espectro acústico do eslate revela a ocorrência de ruído tonal em baixa e alta frequências, e ruído de banda larga em média frequência. Os mapas de beam-forming obtidos associam o ruído de banda larga com uma distribuição bidimensional de fontes ao longo da envergadura do eslate. O ruído do eslate aumenta com a velocidade de escoamento livre, enquanto que os picos tonais de baixa frequência e o ruído de banda larga decrescem com o aumento do ângulo de ataque do aerofólio de 2° para 10°. Os espectros de ruído do eslate colapsam quando reescalados pelo número de Mach do escoamento livre elevado a uma potência entre 4 e 5, e o ruído tonal colapsa em Strouhal dado pela corda do eslate e pela velocidade do escoamento base. Os resultados mostram que o ruído do eslate é fortemente dependente da geometria do aerofólio, particularmente para variações de overlap. Uma boa correspondência quantitativa foi obtida comparando-se espectros experimentais de ruído do eslate com espectros numéricos, obtidos a partir de um modelo com a mesma geometria e em condições de teste idênticas, o que indica a viabilidade do uso de túneis de vento de secção fechada para a realização de experimentos aeroacústicos.
The reduction in the noise produced by aircraft operating in the vicinity of large urban centers is an important issue for a sustainable growth in the civil aviation activities. Over the last decades, from a signicant reduction achieved in the noise generated by aircraft propulsion systems, the contribution of both landing gears and high-lift devices (flaps and slats) has become important to the aircraft overall noise. The identication of the noise signature of each high-lift component is required for the development of both noise prediction methods and new aerodynamic design concepts toward achieving a noise reduction without severe penalty over the aircraft performance and safety. Scaled model experiments have shown that the importance of each airframe component to the overall noise is determined by particularities in both aircraft geometry and size. Such noise model dependence hampers the generalization of experimental results from a reference testing model and leads to the necessity of assessing noise generation according to the testing model geometry and operational condition. This study focuses mainly on the characterization of slat noise from experimental measurements on a high-lift Mcdonnell Douglas (30P30N) airfoil, composed of a slat, a ap and a main element. Measurements were performed in a closed-section wind tunnel by a 62-microphone array and the acoustic data were processed with in-house codes based on conventional beam-forming and DAMAS (Deconvolution Approach for theMapping of Acoustic Sources) algorithms. Beam-forming techniques potentially enable the representation of a spatial source distribution as an acoustic map, from which the contribution of independent sources to the overall noise can be estimated. The experimental database enables the study of the slat noise from dierent airfoil operational conditions and geometrical settings. The slat noise spectral signature reveals the occurrence of tonal noise over both low- and high-frequency bands and also broadband noise over a mid-frequency range. Beam-forming maps indicate the slat broad-band noise originates from a source spatially distributed along the slat span. The slat noise increases in function of the ow speed, whereas low-frequency tonal peaks and the broadband noise decrease as the airfoil angle of attack increases from 2 to 10. The slat noise spectra scalle when the Mach number is raised to a power between 4 and 5, and the tonal noise collapses with Strouhal based on the slat chord and the ow speed. Results show the slat noise is strongly in uenced by the airfoil geometry, particularly for variations in the overlap. A good quantitative agreement was achieved through the comparison between the experimental and numerical slat noise spectra for the same model geometry and test conditions, which indicates the viability of performing aeroacoustic experiments in closed-section wind tunnels.

The main objective of this diploma thesis is to optimize the high lift device on the wing of the Phoenix Air U-15 ultralight aircraft, so that it complies with the UL-2 regulation regarding the stalling speed – 65 KPH. This is fulfilled by optimization of the slotted flap position. Methods used include the Response Surface Method and the Computational Fluid Dynamics approach – namely Ansys Fluent v6 software package. Furthermore, the paper deals with take-off flap optimization and construction of the flap deflection mechanism.

碩士
國立成功大學
航空太空工程學系
87
ABSTRACT Subject : A Design Method of High Lift Airfoil Student : Hao-Kun Chu Advisor : Sheng-Jii Hsieh A method of inverse airfoil design for incompressible potential flow presented by Selig and Maughmer is used in this study. The problem, from a given surface velocity distribution determine the corresponding airfoil shape, is solved by conformal mapping method. After determining the relation of mapping, one may compute the airfoil shape from the unit circle. The prescription of upper surface velocity distribution obeys the Liebeck''s high lift laws, including constant-velocity region, followed by a Stratford-type zero-skin-friction portion to ensure the flow unseparate when decelerates, and ensure the average velocity of upper surface as large as possible. By assuming an initial input of lower surface velocity distribution, and then modifying a portion of the lower surface by the use of least squares and Lagrangian multipliers, one can ensure the velocity distribution satisfies the constraints of inverse method, and minimizes the profile of closure condition. Two high lift airfoils are designed respectively for airfoil trailing-edge angle 0°and 16°, with angle of attack α= 8°. In corporated with the vortex panel method, one may obtain the relation of lift coefficient and angle of attack, and get the maximum lift coefficient and maximum lift-to-drag ratio , for each designed high lift airfoil. For the first high lift airfoil (zero trailing-edge angle case), the stall angle of attack is α=20°, and the maximum lift coefficient is 2.05, but the lift-to-drag ratio is only 24.37. However, when α=8°, the lift coefficient is 1.57, but there has maximum lift-to-drag ratio 73.52. For the second high lift airfoil (trailing-edge angle 16°case), the stall angle of attack is α=15°, and the maximum lift coefficient is 1.735, but the lift-to-drag ratio is only 10.73. However, whenα= 10°, the lift coefficient is 1.55, but there has lift-to drag ratio 82.03,. This study provides a practical computer program for high lift airfoil design, but for the calculation of lower surface velocity distribution, one should search better method to deal with the singularity around leading edge, so as to obtain the problem resulting of high lift but low lift-to-drag ratio.

碩士
國立成功大學
航空太空工程學系碩博士班
100
The feasibility design and evaluation of high-lift devices on Ce-71 series UAV is performed by experimental method. With precision model fabrication, the experiments of high-pressure smoke flow visualization is carried out in a suction type low-speed wind tunnel. The controlling factors on AoA of main wing, the flap deflection, the area of control surface and Reynolds number are changed to obtain visualization results. This research starts from wing design and manufacture as the wind tunnel experiment model for wing flow visualization experiment. The high-lift wings are then set on the Tiger II trainer for flight test verifications. Flight test records on relative variation of height and speed are collected for verification. The relationship between control surface to UAV lift provides further development guideline for better flight performance in Ce-71 series.

碩士
國立成功大學
航空太空工程學系
89
ABSTRACT Subject: An Investigation on High-Lift Airfoil Design Method Student: Cheng-Fang Wu Advisor: Sheng-Jii Hsieh A method of inverse airfoil design for incompressible potential flow presented by Selig and Maughmer is applied in this study. This multi- point design method makes use of the conformal mapping to compute the airfoil shape from the unit circle, and one can compute the airfoil with zero or finite trailing-edge angle. The principle of upper-surface velocity distribution employs the Liebeck’s high-lift concept, and uses Stratford type zero-skin-friction requirement at pressure recovery region to assure the boundary layer unseparated flow condition. In order to meet the high-lift demand, it is necessary to ensure the average velocity of upper surface as large as possible, but the average velocity of lower surface is otherwise oppositely. By assuming an initial lower-surface velocity distribution, and then modifying a portion of the lower surface velocity distribution by sine series function where the coefficients are determined by the least squares and the Lagrangian multipliers. One can assure the lower surface velocity distribution satisfies the constraints of inverse airfoil design method, and have the modified distribution as closed to the initial input as possible. In this study the angle of attack(α), the trailing-edge velocity ( ),the length of accelerating region( ), the orthogonal functions numbers(n) and the airfoil design points(N) will be discussed. From this investigation two high-lift airfoils are designed respectively for the trailing-edge angle 0°withα= 8°and the trailing-edge angle 16°withα= 9°. For the first case (zero trailing-edge angle case), whenα= 8°, ＝0.95, ＝0.05, n＝50, N＝60, the designed high-lift airfoil can be obtained with the lift coefficient =1.2264, the drag coefficient =0.03408 and the lift/drag ratio =35.98. For the second case (trailing-edge angle 16°case), whenα= 9°, ＝0.95, ＝0.05, n＝50, N＝86, the designed high-lift airfoil can be obtained with =2.1187, =0.01819 and =116.48.

The use of steady-state aerodynamic optimization methods in the computational fluid dynamic (CFD) community is fairly well established. In particular, the use of adjoint methods has proven to be very beneficial because their cost is independent of the number of design variables. The application of numerical optimization to airframe-generated noise, however, has not received as much attention, but with the significant quieting of modern engines, airframe noise now competes with engine noise. Optimal control techniques for unsteady flows are needed in order to be able to reduce airframe-generated noise. In this thesis, a general framework is formulated to calculate the gradient of a cost function in a nonlinear unsteady flow environment via the discrete adjoint method. The unsteady optimization algorithm developed in this work utilizes a Newton-Krylov approach since the gradient-based optimizer uses the quasi-Newton method BFGS, Newton's method is applied to the nonlinear flow problem, GMRES is used to solve the resulting linear problem inexactly, and last but not least the linear adjoint problem is solved using Bi-CGSTAB. The flow is governed by the unsteady two-dimensional compressible Navier-Stokes equations in conjunction with a one-equation turbulence model, which are discretized using structured grids and a finite difference approach. The effectiveness of the unsteady optimization algorithm is demonstrated by applying it to several problems of interest including shocktubes, pulses in converging-diverging nozzles, rotating cylinders, transonic buffeting, and an unsteady trailing-edge flow. In order to address radiated far-field noise, an acoustic wave propagation program based on the Ffowcs Williams and Hawkings (FW-H) formulation is implemented and validated. The general framework is then used to derive the adjoint equations for a novel hybrid URANS/FW-H optimization algorithm in order to be able to optimize the shape of airfoils based on their calculated far-field pressure fluctuations. Validation and application results for this novel hybrid URANS/FW-H optimization algorithm show that it is possible to optimize the shape of an airfoil in an unsteady flow environment to minimize its radiated far-field noise while maintaining good aerodynamic performance.

碩士
淡江大學
航空太空工程學系碩士班
99
Global warming has led to extreme weather around the world frequently such as low level wind shear, typhoon, ice/snow etc. If aircraft taking-off and landing will unavoidably meet with the heavy rain, then aircraft designer must put these severe weather influences into considerations in the conceptual design phase. Aerodynamic influences due to heavy rain are still the on-going research subject, and needs further investigation. But for the past decade there are neither experimental nor numerical researches about heavy rain except our research team conducted in recent years. In this thesis, we first review some research finding of heavy rain effects on the aerodynamic performance degradation. Secondly, commercial CFD package FLUENT and preprocessing tool Gambit is used as our main analysis tools, and the simulation of rain is accomplished by using Two-Phase Flow approach’s Discrete Phase Model (DPM) and surface roughness provided by FLUENT. The results show that this research successfully simulates the aerodynamic investigation of high-lift airfoil under the influence of heavy rain effects, the doubts or errors in the previous numerical and experimental works are also revealed. The degradation rate increases with the rain rate, and the premature stall phenomenon is also discovered. It is expected that the quantitative information gained in this thesis could be useful to the operational airline industry, and greater effort should put in this direction to further improve modern transport aircrafts safety.

碩士
國立臺北科技大學
冷凍空調工程系所
105
The demands of an efficient high-lift device have been one of major driving forces in aerodynamic research for many years. The fan-wing was described as a simple, stable, quiet, short takeoff and landing, no stall at large angle of attack, stability in flight and very efficient high-lift device. Having these advantage, it recently becomes a hotspot in unmanned aerial vehic1e and exhibits great potential for exploitation a wider range of application for either military or civilian purpose. Foreign research in this area has been for years, but there are few domestic research literature published today. To catch up, it is worthy for people to participate in research and promotion. Therefore, the purpose of this program is to develop an innovative unmanned aerial vehicle with cross-flow fan propulsive wings. The numerical simulation and experimental measurement are used to analyze the effects of different flight conditions on the distribution characteristics of streamline, the static pressure, the lift and drag forces along surfaces of thrust wing and the actual aerodynamic performance (e.g. flying takeoff and landing, flight attitude, flight in high angle of attack, monitoring system, etc.), propulsion efficiency and so on. The research results can be used as a reference for future commercialization. The results show that (1) an wireless remote control UAV using the cross-flow fan wing as thrust system is completed and verified that it can fly; (2) the maximum thrust created by this CFF-UAV reach 9.1kgf and the shortest distance for taking off/landing is 10m; (3) when the speed reach 6000rpm, the average wind speed generated by the outlet reach 18.65m/s.