Tesis sobre el tema "Aerodynamic interference"

Significant aerodynamic interference can occur between high-speed bodies in close proximity. A complex flowfield develops where shock and expansion waves from a generator body impinge upon the adjacent receiver body. The pressure and flow angularity changes which occur across these disturbances modify the body aerodynamics. The aim of this research is to quantify the aerodynamic interference effects for multi-body configurations and understand the relevant flow physics. The interference aerodynamics for slender bodies in a supersonic flow were investigated through a parametric wind tunnel study. The receiver bodies were finned and un-finned configurations. The effect of lateral and axial body separations, receiver incidence and the strength of the disturbance field were investigated. Measurements included forces and moments, surface pressures and flow visualisations. Supporting computations using steady-state, viscous predictions provided a deeper understanding of the underlying aerodynamics and flow mechanisms. Good agreement was found between the measured and predicted interference loads and surface pressures for all configurations. The interference loads are strongly dependent upon the axial impingement location of the primary shockwave. These induced loads change polarity as the impingement location moves aft over the receiver. The magnitude of the interference loads increase when the receiver is at incidence and are amplified by up to a factor of three when rear fins are attached. In general, the interference loads are larger for a stronger disturbance flowfield. The centre of pressure location is substantially affected and the static stability of the finned receiver changes in some configurations. The effect of the aerodynamic interference on the body trajectories was assessed using an unsteady, Euler prediction in combination with a 6DOF dynamic model. This shows aerodynamic ii interference can cause a collision between the bodies. Moreover, the initial interference loads dominate the subsequent body trajectories and static modelling can be used to evaluate the dynamic trajectories.

In many situations the motion of the fluid and the motion of the body must be determined simultaneously and interactively. One example is the phenomenon of subsonic wing rock. A method has been developed that accurately simulates the pitching and rolling motions and accompanying unsteady flowfield for a slender delta wing. The method uses a predictor-corrector technique in conjunction with the general unsteady vortex-lattice method to compute simultaneously the motion of the wing and the flowfield, fully accounting for the dynamic/aerodynamic interaction. For a single degree of freedom in roll, the method predicts the angle of attack at which the symmetric configuration of the leading-edge vortex system becomes unstable, the amplitude, and the period of the resulting self-sustained limit cycle, in close agreement with two wind-tunnel experiments. With the development of modern aerodynamic configurations employing close-coupled canards, such as the X-29, comes the need to simulate unsteady aerodynamic interference. A versatile method based on the general unsteady vortex-lattice technique has been developed. The method yields the time histories of the pressure distribution on the lifting surfaces, the distribution of vorticity in the wakes, and the position of the wakes simultaneously. As an illustration of the method, the unsteady flowfield for a configuration closely resembling the X-29 is presented. The results show the strong influence of the canards on the main wing, including the time lag between the motions of the canards and the subsequent changes in the vorticity and hence the pressure distributions and loads on the main wing.
Ph. D.
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The flow around a generic car model both in isolation and in proximity to a near side wall has been investigated utilising experimental and computational methods. Phase one of this investigation tested a range of Ahmed generic road vehicle models with varying backlight angles in isolation, employing laser-Doppler anemometry, static pressure and aerodynamic force and moment measurements in the experimental section. Additionally, numerical simulations were conducted using a commercial Reynolds-averaged Navier Stokes (RANS) code with the RNG k-ε turbulence model. This phase served both to extend the previous knowledge of the flow around the Ahmed model, and analyse the effects of both the supporting strut and rolling road. Phase two then used similar methods to investigate the Ahmed model in proximity to a non-moving side wall. Results from phase two are compared with previous near-wall studies in order that an understanding of the effects of wall proximity can be presented, an area lacking in the existing literature. It is found that the flow on the isolated model must be understood before the effects of side wall proximity can be assessed. There is though, in general, a breakdown of any longitudinal vortices on the near-wall side of the model as model-to-wall distance reduces, with an increase in longitudinal vortex strength on the model side away from the wall. There also exists a large pressure drop on the near-wall model side, which increases in magnitude as model-to-wall distance reduces, before dissipating at separations where the boundary layer restricts the flow. Additionally, there is found to be a pressure drop on the top and bottom of the model with decreasing wall distance, with the relative magnitudes of these dependent on model geometry.

The flow around a generic car model both in isolation and in proximity to a near side wall has been investigated utilising experimental and computational methods. Phase one of this investigation tested a range of Ahmed generic road vehicle models with varying backlight angles in isolation, employing laser-Doppler anemometry, static pressure and aerodynamic force and moment measurements in the experimental section. Additionally, numerical simulations were conducted using a commercial Reynolds-averaged Navier Stokes (RANS) code with the RNG k-ε turbulence model. This phase served both to extend the previous knowledge of the flow around the Ahmed model, and analyse the effects of both the supporting strut and rolling road. Phase two then used similar methods to investigate the Ahmed model in proximity to a non-moving side wall. Results from phase two are compared with previous near-wall studies in order that an understanding of the effects of wall proximity can be presented, an area lacking in the existing literature. It is found that the flow on the isolated model must be understood before the effects of side wall proximity can be assessed. There is though, in general, a breakdown of any longitudinal vortices on the near-wall side of the model as model-to-wall distance reduces, with an increase in longitudinal vortex strength on the model side away from the wall. There also exists a large pressure drop on the near-wall model side, which increases in magnitude as model-to-wall distance reduces, before dissipating at separations where the boundary layer restricts the flow. Additionally, there is found to be a pressure drop on the top and bottom of the model with decreasing wall distance, with the relative magnitudes of these dependent on model geometry.

A new aerodynamic device, based on flow interference effects, is studied in order to significantly improve the cornering performance of racing motorcycles in MotoGP. After a brief overview on why standard downforce devices cannot be used on motorcycles, the new idea is introduced and a simplified mechanic analysis is provided to prove its effectiveness. The concept is based on the use of anhedral wings placed on the front fairing, with the rider acting as an interference device, aiming to reduce the lift generation of one wing. Numerical calculations, based on Reynolds-averaged Navier-Stokes equations, are performed on simplified static 2D and 3D cases, as a proof of concept of the idea and as a preparation for further analysis which may involve experimental wind-tunnel testing. The obtained results show that the flow interference has indeed a significant impact on the lift on a single wing. For some cases the lift can be reduced by 70% to over 90% - which strengthens the possibility of a realistic implementation.
Ett nytt aerodynamisk koncept som nyttjar effekter av flödesinterferenser är utvärderat i syfte att på ett noterbart sätt förbättra en roadracing-motorcykels kurtagningsmöjligheter. Efter en kort genomgång av varför diverse klassiska "downforce" lösningar ej är applicerbara på motorcyklar, presenteras det nya konceptet. Varpå en mekanisk analys genomförs i syfte att se över dess tillämpbarhet. Konceptet bygger på anhedrala vingar som placeras på den främre kåpan, där föraren agerar som ett interferensobjekt, och försöker störa ut lyftkraften som den ena vingen genererar. Numeriska beräkningar baserade på RANS-ekvationer är utförda i förenklade statiska 2D och 3D fall. Som ett vidare steg rekommenderas vindtunneltester. Resultaten visar att flödesinterferenser är ytterst märkbara för vingar och i vissa fall kan lyftkraften reducerats med 70-90%. Detta förstäker möjligheten för en realistisk implementering.

Accurately measuring the total temperature of a high-speed fluid flow is a challenging task that is required in many research areas and industry applications. The difficulty in total temperature measurement generally stems from attempting to minimize measurement error or accurately predict error so it can be accounted for. Conduction error and aerodynamic error are two very common sources of error in total temperature probe measurements. Numerous studies have been performed in prior literature to account for simple cases of both errors. However, the impacts of a mounting strut and freestream flow angle on conduction error and aerodynamic error have not been previously modeled. Both of these effects are very common in gas-turbine applications of total temperature probes. Therefore, a fundamental study was performed to analyze the impact of mount interference and freestream flow angle on a probe's conduction error and aerodynamic error. An experimental study of aerodynamic error was performed using strut-mounted thermocouples in a high-speed jet at Mach numbers ranging from 0.25-0.72. This study showed that a strut stagnation point can provide aerodynamic error reductions and insensitivity to approach Mach number. An off-angle experimental study of conduction error was also performed using strut-mounted thermocouples at pitch angles ranging from -30° to 30°. High-fidelity Computational Fluid Dynamics (CFD) simulations with Conjugate Heat Transfer (CHT) were performed in conjunction with the experiments to provide key heat transfer information and flow visualizations. It was identified that unshielded total temperature probes have reduced conduction error at off-angles, but are sensitive to changes in the freestream flow angle. A low-order method was developed to account for mount interference and flow angle effects. The developed low order method utilizes a local Mach number for aerodynamic error predictions and a local Reynolds number for conduction error predictions. This developed low-order method was validated against experiment and 3D, CFD results, and was shown to accurately capture flow angle trends, mount interference effects, and the impacts of varying probe geometry.
Master of Science
Accurately measuring the total temperature of a high-speed fluid flow is a challenging task that is required in many research areas and industry applications. Many methods exist for measuring total temperature, but the use of thermocouple based probes immersed into a flow remains a common and desirable measurement technique. The difficulty in using thermocouple based probes to acquire total temperature stems from attempting to minimize or accurately predict the probe’s measurement error. Conduction, convection, and radiation heat transfer between the fluid flow and probe create challenges for minimizing measurement error so that the accurate total temperature can be obtained. Numerous studies have been performed in prior literature to account for simple cases of each error source. However, there are many complex, practical applications in which the influence of each error source has not been studied. The impacts of a freestream flow angle and the total temperature probe’s mounting structure have not been previously modeled. Both of these effects are very common in gas-turbine applications of total temperature probes. This Thesis will present a fundamental study analyzing the impact that freestream flow angle and a probe’s mount have on a total temperature probe’s measurement error. The influence of conduction and convection heat transfer was studied experimentally for numerous probe geometries, and the impacts of a mounting strut and freestream flow angle were analyzed. A low-order method was developed to predict conduction error and aerodynamic error for total temperature probes in offangle conditions with the presence of mount interference. The developed low-order method was shown to accurately capture the effects of a mounting strut, varying probe geometry, and varying flow angle. Additionally, the low-order method was validated against experimental and 3D, CFD/CHT results.

The aerodynamic behavior of wind tunnels with porous, flexible walls formed from tensioned Kevlar has been characterized and new measurement techniques in such wind tunnels explored. The objective is to bring the aerodynamic capabilities of so-called Kevlar-wall test sections in-line with those of traditional solid-wall test sections. The primary facility used for this purpose is the 1.85-m by 1.85-m Stability Wind Tunnel at Virginia Tech, and supporting data is provided by the 2-m by 2-m Low Speed Wind Tunnel at the Japanese Aerospace Exploration Agency, both of which employ Kevlar-wall test sections that can be replaced by solid-wall test sections. The behavior of Kevlar fabric, both aerodynamically and mechanically, is first investigated to provide a foundation for calculations involving wall interference correction and determination of the boundary conditions at the Kevlar wall. Building upon previous advancements in wall interference corrections for Kevlar-wall test sections, panel method codes are then employed to simulate the wind tunnel flow in the presence of porous, flexible Kevlar walls. An existing two-dimensional panel method is refined by examining the dependency of correction performance on key test section modeling assumptions, and a novel three-dimensional method is presented. Validation of the interference corrections, and thus validation of the Kevlar-wall aerodynamic performance, is accomplished by comparing aerodynamic coefficients between back-to-back tests of models carried out in the solid- and Kevlar-wall test sections. Analysis of the test results identified the existence of three new mechanisms by which Kevlar walls cause wall-interference. Additionally, novel measurements of the boundary conditions are made during the Kevlar-wall tests to characterize the flow at the boundary. Specifically, digital image correlation is used to measure the global deformation of the Kevlar walls under wind loading. Such data, when used in conjunction with knowledge of the pre-tension in the Kevlar wall and the material properties of the Kevlar, yields the pressure loading experienced by the wall. The pressure loading problem constitutes an inverse problem, and significant effort is made towards overcoming the ill-posedness of the problem to yield accurate wall pressure distributions, as well as lift measurements from the walls. Taken as a whole, this document offers a comprehensive view of the aerodynamic performance of Kevlar-wall test sections.
Ph. D.

A influência da fuselagem sobre a posição do centro aerodinâmico da asa é complexa e deve ser considerada nos cálculos de equilíbrio e estabilidade estática longitudinal da aeronave. Este trabalho apresenta uma análise comparativa para indicar o mais preciso dentre sete métodos teóricos para prever esta influência, em condições de baixa velocidade, utilizando seis configurações de modelos de asa mais fuselagem em escala reduzida, com proporções dimensionais características da aviação leve. Mediram-se os coeficientes de momento e sustentação para cada configuração, através de ensaios em túnel de vento de baixa velocidade, circuito aberto e seção de testes fechada. Calcularam-se as posições experimentais do centro aerodinâmico através da distância do eixo de rotação da balança ao bordo de ataque da asa e derivadas do coeficiente de momento em relação ao coeficiente de sustentação. Aplicaram-se os métodos teóricos às configurações. Os resultados demonstram que a maioria dos métodos prevê comportamentos na variação da posição do centro aerodinâmico semelhantes aos obtidos experimentalmente e apontados na revisão da literatura. A análise dos resultados teóricos ante os experimentais aponta o método descrito em Engineering Sciences Data Unit (1996a) como o mais preciso.
The fuselage influence on the wing aerodynamic center is complex and must be considered within longitudinal static stability and equilibrium calculations of the airplane. This work presents a comparative analysis to indicate the most accurate between seven theoretical methods that predict this influence, at low speed conditions, using six configurations of wing-fuselage reduced scale models, with the dimensional proportions found in light aviation. The moment and lift coefficients have been measured by experiments in a low speed open circuit wind tunnel with a closed test section. The experimental aerodynamic center positions have been found by the distance of the balance trunnion to wing leading edge and the derivation of the moment coefficient relative to the lift coefficient. The theoretical methods have been applied to all configurations. The results show that most of the methods predict variations in aerodynamic center position in the same way as those obtained in experimental results and shown in the literature review. The analysis between theoretical and experimental results indicates the method from Engineering Sciences Data Unit (1996a) as the most accurate.

Kyoto University (京都大学)
0048
新制・課程博士
博士(工学)
甲第12587号
工博第2700号
新制||工||1388(附属図書館)
UT51-2006-S595
京都大学大学院工学研究科社会基盤工学専攻
(主査)教授 松本 勝, 教授 河井 宏允, 助教授 白土 博通, 教授 田村 武
学位規則第4条第1項該当

Esse trabalho descreve uma investigação prática e teórica de características aerodinâmicas de um modelo referente a um avião bimotor leve \"Piper PA-34 SENECA III\" construído em escala 1:6,5. Basicamente, duas configurações fazem parte da avaliação: \"asa limpa\" e Nacele \"Pusher\". O principal objetivo é o de avaliar o comportamento aerodinâmico da configuração \"Pusher\" em três diferentes posições através da obtenção de dados práticos e teóricos de distribuição de pressão estática no aerofólio e no próprio nacele na condição não motorizada. A metodologia emprega testes em túnel de vento de baixa velocidade para a aquisição dos dados experimentais através da variação do ângulo de ataque para as configurações citadas. A obtenção dos dados numéricos, como comparativos aos experimentais, é através da utilização do método dos painéis aplicado à aerodinâmica. Os resultados experimentais e teóricos mostram que a presença do nacele nas posições \"Pusher\" avaliadas gera interferência diminuindo à sustentação do aerofólio.
This work describes a practical and theoretical investigation of aerodynamic characteristics of a 1:6,5 scale model of a light twin-engined Piper Seneca III aircraft. Basically, two configurations are considered in this evaluation: clean wing and pusher nacelle. The principal objective is to evaluate the aerodynamic characteristics of the pusher configuration in three different positions, in order to obtain practical and theoretical data with reference to static pressure distribution on the surface of the aerofoil and of the nacelle in a power-off condition. The methodology employed used low speed wind tunnel test for the acquisition of experimental data for a selection of angles of attach, as cited. The generation of numerical data, to be compared the experimental equivalent, is though the use of a low order panel method. The experimental and theoretical results show that the presence of pusher nacelle in the positions mentioned generated interference, reducing the local lift of the aerofoil.

To explore the influence of wind tunnel test facilities on delta wing aerodynamics, the interference has been separated into two distinct types, wall interference and support structure interference. The wall interference effects have been split into three further components, tunnel blockage, side wall interference, and roof and floor interference. Splitting the tunnel influence in this way allows us to determine the most detrimental interference effects, thus allowing the wind tunnel engineer to design experiments accordingly. Euler and more realistic RANS simulations of tunnel interference have been conducted. To reduce the question of grid dependence when comparing solutions, a common "farfield grid" was created and tunnel grids were extracted. Before doing RANS simulations an analysis of various turbulence models was conducted. It was found that turbulence models have difficulty in predicting turbulence levels in leading edge vortices. As such modifications have been applied to the models which improve predictions. Despite vortex breakdown being widely regarded as an inviscid phenomenon, dependence on turbulence modelling has been exhibited. This is due to the vortex properties being altered with turbulent diffusion of vorticity when turbulence levels are too high. Both 1- and 2-equation models were assessed and it was concluded that a modified 2-equation k-w model was the most suitable of the models available (when compared against experimental results), and was therefore used in all subsequent simulations. From both Euler and RANS simulations it has been concluded that the effect of sidewall proximity significantly promotes vortex breakdown. Side wall induced velocity components increase the mean effective incidence of the wing, the helix angle and the strength of the vortices. The combination of these effects promotes vortex breakdown. Roof and floor proximity has little effect on vortex breakdown as does the frontal area blockage. Pitching simulations have shown that the promotion of vortex breakdown is not consistent on both the upstroke and downstroke. Break-down was observed to be prompted furthest at the higher incidence of the upstroke and on the downstroke. This highlights the dependency of tunnel interference on vortex strength.

This thesis is a study of the interference forces acting on one cylinder comprising a group of three and four cylinders when situated in a free stream flow. The spacing of the cylinders was such that the lines joining their axes formed an equilateral triangle, and a square respectively. The study is split into two parts, (a) potential flow over the groups of cylinders (b) real flow over the groups of cylinders. Bound with copies of the authors' published papers.~Bibliography: pages 132-137.

Monoposto racecar development is routinely carried out using wheels supported not by the car suspension but by individual, externally-mounted stings. The interference effect of these stings was acknowledged but unquantified in the existing literature. Appraisal of the literature has found that the structure of a wheel wake was not understood, rendering it difficult to assess the support sting interference. These two issues were thus jointly addressed using experimental and computational methods. The two phases of this project each tested a different industrially-representative racecar wheel model. Phase One investigated a single wheel and sting combination, whilst Phase Two extended the work to include two further stings and a model racecar. Non-intrusive velocity measurements were made in the near wakes of the various combinations to extract vertical planes, perpendicular to the tunnel freestream. The measurements made behind the isolated wheels were used to investigate the main flow features of the wake. The flow surrounding an unsupported wheel was established computationally and used to evaluate the interference effects of the support sting. Different wheel support methods (three stings and the car suspension) were used to provide further insight into the sting interference effects and also the impact of sting design on those effects. Testing with and without the model racecar allowed evaluation of its effect on the wheel wake and sting interference. The main characteristics of the near-wake of an isolated wheel rotating in ground contact are proposed from analysis of the data generated in this study. A simplified model of the trailingvortex system induced in the wake of such a wheel is proposed to clarify contradictory literature. The specific interference effects of a wheel support sting are proposed with reference to the main characteristics of the wake. The mechanisms behind these effects are, where possible, identified and presented. The main impact of the support sting, and thus the root of several of the observed effects, is the modification of the axial flow through the wheel. The main effects of the presence of the car on the near-wake are proposed alongside the observation that the wake structure is not fundamentally different to that of an isolated wheel. The proposed sting interference effects are also observed in the presence of the car, albeit at a reduced level.

Monoposto racecar development is routinely carried out using wheels supported not by the car suspension but by individual, externally-mounted stings. The interference effect of these stings was acknowledged but unquantified in the existing literature. Appraisal of the literature has found that the structure of a wheel wake was not understood, rendering it difficult to assess the support sting interference. These two issues were thus jointly addressed using experimental and computational methods. The two phases of this project each tested a different industrially-representative racecar wheel model. Phase One investigated a single wheel and sting combination, whilst Phase Two extended the work to include two further stings and a model racecar. Non-intrusive velocity measurements were made in the near wakes of the various combinations to extract vertical planes, perpendicular to the tunnel freestream. The measurements made behind the isolated wheels were used to investigate the main flow features of the wake. The flow surrounding an unsupported wheel was established computationally and used to evaluate the interference effects of the support sting. Different wheel support methods (three stings and the car suspension) were used to provide further insight into the sting interference effects and also the impact of sting design on those effects. Testing with and without the model racecar allowed evaluation of its effect on the wheel wake and sting interference. The main characteristics of the near-wake of an isolated wheel rotating in ground contact are proposed from analysis of the data generated in this study. A simplified model of the trailingvortex system induced in the wake of such a wheel is proposed to clarify contradictory literature. The specific interference effects of a wheel support sting are proposed with reference to the main characteristics of the wake. The mechanisms behind these effects are, where possible, identified and presented. The main impact of the support sting, and thus the root of several of the observed effects, is the modification of the axial flow through the wheel. The main effects of the presence of the car on the near-wake are proposed alongside the observation that the wake structure is not fundamentally different to that of an isolated wheel. The proposed sting interference effects are also observed in the presence of the car, albeit at a reduced level.

This study is undertaken to characterize the aerodynamic behavior of Kevlar-wall test sections and specifically those containing two-dimensional, lifting models. The performance of the Kevlar-wall test section can be evaluated against the standard of the hard-wall test section, which in the case of the Stability Wind Tunnel (SWT) at Virginia Tech can be alternately installed or replaced by the Kevlar-wall test section. As a first step towards the evaluation of the Kevlar-wall test section aerodynamics, a validation of the hard-wall test section at the SWT is performed, in part by comparing data from NACA 0012 airfoil sections tested at the SWT with those tested at several other reliable facilities. The hard-wall test section showing good merit, back-to-back tests with three different airfoils are carried out in the SWT's hard-wall and Kevlar-wall test sections. Kevlar-wall data is corrected for wall interference with a panel method simulation that simulates the unique boundary conditions of Kevlar-wall test sections including the Kevlar porosity, wall deflection, and presence of the anechoic chambers on either side of the walls. Novel measurements of the boundary conditions are made during the Kevlar-wall tests to validate the panel method simulation. Finally, sensitivity studies on the input parameters of the panel method simulation are conducted. The work included in this study encompasses a wide range of issues related to Kevlar-wall as well as hard-wall tunnels and brings to light many details of the performance of such test sections.
Master of Science

This investigation sought first to determine the feasibility of generating a surrogate model of the interference drag between nacelles and wing-fuselage systems suitable for the inclusion in a multidisciplinary design optimization (MDO) framework. The target aircraft was a single-aisle, transonic aircraft with a freestream Mach number of 0.8 at 35,000 feet and a design lift coefficient of 0.5. Using an MDO framework is necessary for placing the nacelle because of the competing objectives of the disciplines involved in aircraft design including structures, acoustics, and aerodynamics. A secondary goal was to determine what tools are necessary for accurately capturing interference drag effects on the system. This research used both Euler computational fluid dynamics (CFD) with a coupled viscous drag estimation tool and Reynolds Averaged Navier-Stokes (RANS) CFD to estimate the system drag. The initial trade space exploration that varied the nacelle location across a baseline airframe configuration was completed with the Euler solver, and it showed that appreciable overlap between the wing and nacelle led to large increases in interference drag. A follow-on study was conducted with RANS CFD where the wing shape was tailored for each unique nacelle position. In comparing the results of the Euler and the RANS CFD, it was determined that RANS is required to accurately capture the flow features. Euler solvers can create artifacts due to the lack of viscous effects within the model. Wing tailoring is necessary because of the sensitivity of transonic flows to geometric changes and the addition of neighboring components, such as a nacelle. The research showed that for above and aft wing locations, a nacelle can overlap the trailing edge without incurring a drag penalty. Nacelles placed in the conventional location, forward and beneath the wing, displayed low interference drag effects, as the nacelle had a small and local impact on the wing's aerodynamics. Given the high cost of computing a RANS solution with wing tailoring, and the large design space for nacelle locations, building a surrogate model for interference drag was found to be prohibitive at this time. As the cost of computing and mesh generation decreases, collecting the data for building a surrogate model may become tractable.
Doctor of Philosophy
Engine placement on an aircraft is dependent on multiple disciplines. Engine placement affects the noise of the aircraft because the wing can shield or reflect the engine noise. Engine placement impacts the structural loads of an aircraft, with some positions requiring more reinforcement that adds to the cost and weight of the aircraft. Aerodynamically, the engine placement impacts the vehicle's drag. Taken together, the only means of trading the different disciplines' needs is through a multidisciplinary design optimization (MDO) framework. The challenge of MDO frameworks is that they require numerous solutions to effectively explore the trade space. Thus, MDO frameworks employ fast, low-order tools to compute hundreds or thousands of different combinations of features. A common approach to make running MDO analysis feasible is to develop surrogate models of the key considerations. Current aerodynamic drag build-ups for aircraft do not consider the interference drag associated with engine placement. The first goal of this research was to determine the feasibility of generating a surrogate model for inclusion in an MDO framework. In order to collect the data required for the surrogate, appropriate tools to capture the interference drag are required. Building a surrogate requires a large number of samples, thus the aerodynamic solver must be fast, robust, and accurate. An Euler (inviscid) computational fluid dynamics (CFD) was used do explore the engine placement design space to test the feasibility of building the surrogate model. The target aircraft was a single-aisle, transonic aircraft with a freestream Mach number of 0.8, flying at an altitude of 35,000 feet and a design lift coefficient of 0.5. The initial vehicle used a baseline wing, and the engine placement was varied across the wing span and fuselage. The results showed that the conventional location, where the engine is forward and beneath the wing, had the a modestly beneficial interference drag, though positions near the trailing edge and above the wing also showed neutral interference drag. In general, if the engine overlapped the wing, the interference drag increased dramatically. % A follow-on study used Reynolds Averaged Navier-Stokes (RANS) CFD to investigate seven engine placements above and aft of the wing. Each of these positions had the wing tailored such that the wing performance would be typical of a good transonic wing. The results showed that with wing tailoring, a moderate amount of overlap between the wing and nacelle results in reduced or neutral interference drag. This is in contrast with the baseline wing results that showed moderate overlap led to large increases in interference drag. % The results from this research suggest that building a surrogate model of interference drag for transonic aircraft is not feasible given today's computational resources. In order to accurately model the interference drag, one must use a RANS CFD solver and tailor the wing. These requirements increase the cost of evaluating an engine position such that collecting enough for a surrogate model is prohibitively expensive. As computational speeds increase, and the ability to automate CFD mesh generation becomes less time intensive, the feasibility may increase. Using an Euler solver is insufficient because of the lack of viscous effects in the flow. The lack of a boundary layer leads to artifacts appearing in the flow when the nacelle and wing are in close proximity.

碩士
淡江大學
土木工程學系碩士班
106
There is no standard recommendation for the interference effect of buildings wind-resistance design specifications on neighboring buildings, and generally, the surrounding environment of urban terrain is very complicated. Therefore, wind tunnel tests must be used as the basis for wind design. This study investigates the effect of wind forces on the mutual interference caused by two square pillar buildings, and compares the effects of aerodynamic damping under the influence of disturbance effects and without interference effects. In addition, the calculation of the displacement of high-rise buildings due to wind forces, the transverse wind aerodynamic behavior is complex, especially when the low-scruton number (low-quality and low-damping) high-rise building structure, due to wind caused by excessive displacement response, resulting in aerodynamic. The phenomenon of unstable force has great harm to the structure. Therefore, this study specifically discusses the structural system of the low Scruton number. The wind tunnel experiment about the aero-elastic vibration test is conducted in the 18.0 × 1.8 × 2.2 m boundary layer wind tunnel which is conducted by the guidance professor Professor Lo Yuanlong at Wind Engineering Research Center at Tokyo Polytechnic University in August 2017. A 1/400 scale turbulent flow over a sub-urban terrain with a power law index exponent for mean velocity profile of 0.2 is simulated with properly equipped spires, saw barriers, and roughness blocks. Changing 12 reduced wind speeds (6.5, 7.5, 8.3, 9.0, 9.7, 10.5, 11.0, 11.5, 12.2, 13.1, 13.8, and 14.6). As a study of the aerodynamic characteristics of across-wind vibratory behavior under different reduced wind speeds, the square prism model is 0.07 m in both width (B) and depth (D) and 0.56 m in height (H), which make the aspect ratio (H/B) 8. The tapered model is 0.04 m in width on the roof-top and 0.10 m in width on the bottom. The height is the same as the square one and the aspect ratio (height to the averaged width) is also 8. Both the two principal building models are manufactured in the same volume in order to have a basic comparison level. Addiction the interfering model which is made rigid-pivoted aero-elastic and tuned to vibrate in the same fundamental frequency as the principal building models. There are 20 interference locations. In this study, experimental results are used to investigate the aerodynamic damping behavior of high-rise buildings under disturbance effects. Different experimental settings include changes in different interference positions; structural responses caused by different major buildings; different reduced wind speeds and different interferences Four possible factors such as the impact of buildings (rigid interference and vibration). Two methods for calculating aerodynamic damping are used: random decay method, wavelet theory. In this study, firstly, the comparison of the calculation results of the two methodologies was performed. After the determination, the aerodynamic damping value was discussed using one of them. It was expected that the aerodynamic forces could not be estimated in the aeroelastic experiment. The interaction between the two buildings when vibration occurs due to wind force is compared with the distribution of the aerodynamic damping value and the disturbance displacement response. It is shown that changes in the aerodynamic damping will lead to changes in the disturbance displacement.

With the age of hypersonic flight imminent just beyond the horizon, researchers are working hard at designing work-arounds for all the major problems as well as the minor quirks associated with it. One such issue, seemingly innocuous but one that could be potentially deadly, is the problem of interference heating due to surface protuberances. Although an ideal design of the external surfaces of a high-speed aircraft dictates complete smoothness to reduce drag, this is not always possible in reality. Control surfaces, sheet joints, cable protection pads etc. generate surface discontinuities of varying geometries, in the form of both protrusions as well as cavities. These discontinuities are most often small in dimension, comparable to the local boundary layer thickness at that location. Such protuberances always experience high rates of heat transfer, and therefore should be appropriately shielded. However, thermal shielding of the protrusions alone is not a full solution to the problem at hand. The interference caused to the boundary layer by the flow causes the generation of local hot spots in the vicinity of the protuberances, which should be properly mapped and adequately addressed. The work presented in this thesis aims at locating and measuring the heat flux values at these hot spots near the protrusions, and possibly formulating empirical correlations to predict the hot spot heat flux for a given set of flow conditions and protrusion geometry. Experimental investigations were conducted on a flat plate model and a cone model, with interchangeable sharp and blunt nose tips, with attached 3D protuberances. Platinum thin-film sensors were placed around the protrusion so that the heat fluxes could be measured in its vicinity and the hottest spot located. These experiments were carried out at five different hypersonic free stream flow conditions generated using two shock tunnels, one of the conventional type, and the other of the free-piston driven type. The geometry of the protrusions, i.e., the height and the deflection angle, was also parametrically varied to study its effect on the hot spot heat flux. The results thus obtained for the flat plate case were compared to existing correlations in the open literature from a similar previous study at a much higher Reynolds number range. Since a mismatch was observed between the results of the current experiments and the existing correlations, a new empirical correlation has been developed to predict the hot spot heat flux, that is valid within the range of flow conditions studied here. A similar attempt was made for the case of the cone model, for which no previous correlations exist in the open literature. However, a global correlation covering the entire range of flow conditions used here could not be formed. A correlation that is valid for just one out of the five flow conditions used here is presented for the cones with sharp and blunt nose tips separately. Schlieren flow visualization was carried out to obtain a better understanding of the shock structures near the protuberances on both models. For most cases, where the protrusion height and deflection angle were large enough to cause flow separation immediately upstream of the protuberance, a separation shock was manifested which deflected some part of the boundary layer above the protuberance, while the rest of the fluid in the boundary layer entered a recirculating region in the separated zone before escaping to the side. Some preliminary computational analysis was conducted which confirmed this qualitatively. However, the quantitative match of surface heat flux between the simulations and experiments were not encouraging. Schlieren visualization revealed that for the flat plate case, the foot of the separation shock was located at a distance of 10.5 to 12 times the protrusion height ahead of it, whereas in the case of the sharp cone, it was at a distance of 9 to 10.5 times the protrusion height. The unsteady nature of the separation shock was also captured and addressed. Some preliminary experiments on boundary layer tripping were also conducted, the results of which have been presented here. From this analysis, it has become evident that a single global correlation cannot be formed which could be used for a wide range of flow conditions to predict the hot spot heat flux in interference interactions. The entire range of conditions that may be encountered during hypersonic flight has to be broken down into sections, and the interference heating pattern should be studied in each of these sections individually. By doing so, a series of different correlations can be formed at the varying flow conditions which will then be available for high-speed aircraft designers.

碩士
國立中興大學
土木工程學系所
96
In order to be done further understanding the tall buildings by the aerodynamic phenomenon of wind. This research carries on the experiment of wind-tunnel under the circumstances that different intervals arrange the position to twin cylinders with a square close section. To confer the downstream cylinder to incur wind loading to produce vertically response. Hope to understanding aerodynamic data of the single square cylinders, by way of set up experience formula. Demonstrate of the aerodynamic interference effects of twin cylinders with a square close section, and hope tocompare each other with the experiment achievement of 3-D cylinder in the future. Different intervals position of twin square cylinders, will be produced strengthened or suppressed vertically response and will influence us to predict the position of wind speed of resonance. Find that because flow field influences each other among the twin square cylinder. Then analyse that find out relevant experience formula of twin square cylinder in the permutation position. By the experience formula to calculate aerodynamic damping and vertically response, find the same as experiment value.

碩士
逢甲大學
機械工程研究所
80
Combat aircraft carry external stores (bombs, rockets, fuel tanks, instrument pods, or containers) at various station on the wing. There are mutual aerodynamic interferences between the aircraft and the external stores. This condition can change the characteristics of aircraftoflow field that may result in a wing/store flutter phenomena, an instability of flight control, and external stores departing from the predicted separation trajectory behavior and possibly colliding with the aircraft。 　　An investigation is made, simplifying the aerodynamic interference to two-dimensions and using the hydraulic analogy analysis technique to simulate the flow field. Experiments in a water table provide qualitative observation and quantitative analysis of interference flow fields at subsonic and supersonic speed. The results show that different wing/store configurations and flight conditions have different aerodynanmic interferences. Thus, interference flow fields can analyze the force applied to the external store and lift distribution of the wing before store separating from the aircraft.