Desde o início da história da aviação, a cada novo projeto de aeronave, modificações são realizadas para se obter melhorias nas características aerodinâmicas como, por exemplo, diminuição do arrasto, aumento de sustentação, melhora na manobrabilidade em elevados ângulos de ataque, entre outras. Essas modificações podem ser feitas em diversas partes de uma aeronave como, por exemplo, no perfil aerodinâmico e enflechamento da asa, nas superfícies de controle e até em novas posições da entrada de ar do motor intake. Embora os conceitos de aeronaves com asas de enflechamento negativo, Canard e intake dorsal não sejam novos, ainda não existe uma aeronave que apresente essas três características juntas. Portanto, o presente trabalho teve como objetivo estudar uma aeronave não convencional com essas características. Para isso, testes foram realizados no túnel de vento do Laboratório de Aerodinâmica (LAE), do Departamento de Engenharia Aeronáutica da Escola de Engenharia de São Carlos, da Universidade de São Paulo. Três estudos foram realizados para diferentes ângulos de ataque da aeronave e diferentes ângulos de incidência de três pares de Canard projetados, que foram colocados em duas posições diferentes em relação à asa. No primeiro estudo, as forças aerodinâmicas atuantes na aeronave foram medidas por meio de uma balança aerodinâmica para se obter relações entre seus coeficientes e a eficiência conforme a troca dos pares de Canard. Posteriormente, o estudo do escoamento na entrada do intake dorsal englobou a relação entre a velocidade local no intake, a velocidade do escoamento livre e a recuperação de pressão total no mesmo. Por fim, um mapeamento da vorticidade no campo da asa foi realizado para se analisar a influência do Canard no escoamento raiz da asa. O modelo estudado obteve uma maior eficiência e menor arrasto em elevados ângulos de ataque para ângulos de incidência negativos de Canard. O Canard que se destacou nos dois primeiros estudos foi utilizado no estudo de mapeamento, porém, não apresentou uma influência no escoamento da raiz da asa como desejado. / Since the beginning of aviation history, in each new aircraft design, modifications are made to obtain improvements in aerodynamic characteristics such as drag reduction, increase of lift, improvement in maneuverability at high angles of attack. These modifications can be made in different parts of an aircraft, such as, aerodynamic profile, on the control surfaces and even in new intake positions. Although the concepts of aircraft with forward swept wing, Canard and dorsal intake are not new, there is still no aircraft that presents these three characteristics together. Therefore, the present work had as objective to study an unconventional aircraft with these characteristics. Wind tunnel tests were carried out in order to analyze the aerodynamic characteristics at the Aerodynamics Laboratory of the Department of Aeronautical Engineering of EESC-USP. Three studies were carried out for different angles of attack of the aircraft and different angles of incidence of three pairs of Canard, which were placed in two different positions in relation to the wing. In the first study, the aerodynamic forces acting on the aircraft were measured by an aerodynamic balance to obtain relations between their coefficients and the efficiency according to the change of the Canard pairs. Subsequently, the study of flow at the entrance of the dorsal intake included the relation between the local velocity at the intake and the velocity of the free flow and the total pressure recovery. Finally, a mapping of vorticity in the wing field was performed to analyze the influence of Canard on the boundary layer at the root of the wing. The model studied obtained higher efficiency and lower drag at high angles of attack at negative Canard angles of incidence. The Canard that stood out in the first two studies was used in the mapping study, however, it did not present an influence on the root of the wing as desired.

This thesis deals with receptivity mechanisms of three-dimensional, crossflow-dominated boundary layers. The receptivity of two model problems, a swept-flat-plate and a swept-wing boundary layer, is investigated by solving the parabolised stability equations (PSE) as well as by performing direct numerical simulations (DNS).Both flow cases are known to exhibit strong inflectional instabilities, the crossflow disturbances, whose excitation by external disturbances such as surface roughness or free-stream vorticity is studied. One focus is on worst-case scenarios. This involves the determination of optimal conditions, i.e. those disturbance environments yielding the largest possible response inside the boundary layer. A new method on the basis of the PSE is presented which allows to study optimal disturbances of swept-flat-plate boundary layers. These take the form of tilted streamwise vortices. While convected downstream they develop into streamwise streaks experiencing strong non-modal growth. Eventually, they turn into crossflow disturbances and undergo exponential growth. Non-modal growth is thus found to optimally excite crossflow disturbances and can be related to a receptivity mechanism of three-dimensional boundary layers. Evaluating effects of compressibility reveals that the potential for both non-modal and modal growth increases for higher Mach numbers. It is shown that wall cooling has diverse effects on disturbances of non-modal and modal nature. While destabilising the former it attenuates the growth of modal disturbances. Concave curvature on the other hand is found to be equally destabilising for both types of disturbances. The adjoint of the linearised Navier-Stokes equations is solved for a swept-wing boundary layer by means of DNS. The adjoint solution of a steady crossflow disturbance is computed in the boundary layer as well as in the free-stream upstream of the leading edge. This allows to determine receptivity to incoming free-stream disturbances and surface roughness as well as the corresponding worst-case scenarios. Upstream of a swept wing the optimal initial free-stream disturbance is found to be of streak-type which convects downstream towards the leading edge. It entrains the boundary layer a short distance downstream of the stagnation line. While minor streamwise vorticity is present the streak component is dominant all the way into the boundary layer where the optimal disturbance turns into a crossflow mode. Futher, the worst-case surface roughness is determined. It takes a wavy shape and is distributed in the chordwise direction. It is shown that, under such optimal conditions, the swept-wing boundary layer is more receptive to surface roughness than to free-stream disturbances. Another focus of this work has been the development and evaluation of tools for receptivity prediction. Both DNS and direct and adjoint solutions of the PSE are used to predict the receptivity of a swept-wing boundary layer to localised surface roughness. The configuration conforms to wind tunnel experiments performed by Saric and coworkers at the Arizona State University. Both the DNS and the PSE are found to predict receptivity amplitudes which are in excellent agreement with each other. Though the predicted disturbance amplitudes are slightly lower than experimental measurements the overall agreement with experimental results is very satisfactory. Finally, a DNS of the stabilisation of a transitional swept-wing boundary layer by means of discrete roughness elements is presented. This control approach is found to completely suppress transition to turbulence within the domain studied and confirms experimental results by Saric & coworkers. / QC 20111124

The effect of freestream turbulence on generation of crossflow disturbances over swept wings is investigated through direct numerical simulations.  The set up follows  the  experiments  performed  by Downs  et  al.  in their  TAMU  experi- ment.  In this experiment the authors use ASU(67)-0315 wing geometry which promotes  growth  of crossflow  disturbances.   Distributed  roughness  elements are locally placed near the leading edge with a span-wise wavenumber, to ex- cite the corresponding crossflow vortices.  The response of boundary layer to external disturbances such as roughness heights, span-wise wavenumbers, Rey- nolds numbers and freestream turbulence characteristics are studied.  It must be noted that the experiments were conducted at a very low level of freestream turbulence  intensity  (T u).   In this  study,  we fully  reproduce the  freestream isotropic homogenous turbulence through a DNS code using detailed freestream spectrum data provided by the experiment. The generated freestream fields are then applied as the inflow boundary condition for direct numerical simulation of the wing. The geometrical set up is the same as the experiment along with application of distributed roughness elements near the leading edge to precipi- tate stationary crossflow disturbances.  The effects of the generated freestream turbulence are then studied on the initial amplitudes and growth of the bound- ary layer perturbations.  It appears that the freestream turbulence damps out the dominant stationary crossflow vortices. / <p>QC 20130604</p>

One of the last remaining challenges preventing the laminarization of sweptwings is the control of unstable crossflow vortices. In low-disturbance environments the transition from laminar to turbulent flow on the swept-wing initially takes the path of receptivity, where surface roughness or disturbances in the environment introduce shortwavelength disturbances into the boundary layer. This is followed by development and linear growth of stationary crossflow vortices that modify the mean flow, changing the stability characteristics of the boundary layer. Finally, breakdown to turbulence occurs over a short length scale due to the high-frequency secondary instability. The receptivity mechanism is the least understood, yet holds the most promise for providing a laminar flow control strategy. Results of a 3-year flight test program focused on receptivity measurements and laminar flow control on a 30-degree swept-wing are presented. A swept-wing test article was mounted on the port wing of a Cessna O-2A aircraft and operated at a chord Reynolds number of 6.5 to 7.5 million. Spanwise-periodic, micronsized discrete roughness elements were applied at the leading edge of the swept-wing in order to excite the most unstable crossflow wavelength and promote early boundary layer transition. An infrared camera was used to detect boundary-layer transition due to changes in leading-edge roughness. Combined with the IR camera, a new technique of calibrating surface-mounted hotfilms was developed for making disturbance-amplitude measurements downstream of modulated roughness heights. This technique proved to be effective at measuring disturbance amplitudes and can be applied in future tests where instrumentation is limited. Furthermore, laminar flow control was performed with subcritically-spaced roughness. A 100% increase in the region of laminar flow was achieved for some of the conditions tested here.

A series of flight tests was performed using a swept-wing model mounted on a Cessna O-2 aircraft. The crossflow waves on the airfoil were excited by pneumatic spanwise-periodic distributed roughness elements (DREs). The objective of the experiment was to determine the roughness receptivity i.e. the relationship between roughness height and the amplitude of the unstable crossflow wave. The local skin-friction variation was measured using an array of calibrated and temperature-compensated hotfilm sensors. The amplitudes of the disturbance shear stress were compared to the amplitudes of the DREs. It was found that there is a relationship between the shear stress and DRE amplitude that needs to be studied more before any definitely conclusions can be made. It was also found that the sensitivity of the crossflow to DREs is highly dependent on the freestream turbulence levels.

No description available.

Many futuristic aircraft such as the Hybrid Wing Body have numerous control surfaces that can result in large hinge moments, high actuation power demands, and large actuator forces/moments. Also, there is no unique relationship between control inputs and the aircraft response. Distinct sets of control surface deflections may result in the same aircraft response, but with large differences in actuation power. An Artificial Neural Network and a Genetic Algorithm were used here for the control allocation optimization problem of a Hybrid Wing Body to minimize the Sum of Absolute Values of Hinge Moments for a 2.5-G pull-up maneuver. To test the versatility of the same optimization process for different aircraft configurations, the present work also investigates its application on the Forward Swept Wing aircraft. A method to improve the robustness of the process is also presented. Constraints on the load factor and longitudinal pitch rate were added to the optimization to preserve the trim constraints on the control deflections. Another method was developed using stability derivatives. This new method provided better results, and the computational time was reduced by two orders of magnitude. A hybrid scheme combining both methods was also developed to provide a real-time estimate of the optimum control deflection schedules to trim the airplane and minimize the actuation power for changing flight conditions (Mach number, altitude and load factor) in a pull-up maneuver. Finally, the stability derivatives method and the hybrid scheme were applied for an antisymmetric, steady roll maneuver. / Master of Science / Many futuristic aircraft such as the Hybrid Wing Body have numerous control surfaces that can result in large actuation power. An Artificial Neural Network and a Genetic Algorithm were used here to minimize the actuation power on the Hybrid Wing Body. To test the versatility of the same optimization process for different aircraft configurations, the present work also investigates its application on the Forward Swept Wing aircraft. A method to improve the robustness of the process is also presented. A completely different method was developed, and it provided better results with the computational time reduced by two orders of magnitude. A hybrid scheme combining both methods was also developed to provide a real-time estimate of the optimum control deflection schedules to trim the airplane and minimize the actuation power for changing flight conditions (Mach number, altitude and load factor) in a pull-up maneuver.

This work is concerned with stability and receptivity analysis as well as studies on control of the laminar-turbulent transition in boundary-layer flows through direct numerical simulations. Various flow configurations are considered to address flow around straight and swept wings. The aim of this study is to contribute to a better understanding of stability characteristics and different means of transition control of such flows which are of great interest in aeronautical applications. Acoustic receptivity of flow over a finite-thickness flat plate with elliptic leading edge is considered. The objective is to compute receptivity coefficient defined as the relative amplitude of acoustic disturbances and TS wave. The existing results in the literature for this flow case plot a scattered image and are inconclusive. We have approached this problem in both compressible and incompressible frameworks and used high-order numerical methods. Our results have shown that the generally-accepted level of acoustic receptivity coefficient for this flow case is one order of magnitude too high. The continuous increase of computational power has enabled us to perform global stability analysis of three-dimensional boundary layers. A swept flat plate of FSC type boundary layer with surface roughness is considered. The aim is to determine the critical roughness height for which the flow becomes turbulent. Global stability characteristics of this flow have been addressed and sensitivity of such analysis to domain size and numerical parameters have been discussed. The last flow configuration studied here is infinite swept-wing flow. Two numerical set ups are considered which conform to wind-tunnel experiments where passive control of crossflow instabilities is investigated. Robustness of distributed roughness elements in the presence of acoustic waves have been studied. Moreover, ring-type plasma actuators are employed as virtual roughness elements to delay laminar-turbulent transition. / <p>QC 20161124</p>

In this study, the non-equilibrium Johnson and King Turbulence Model (JK model) is implemented in a three-dimensional, Navier-Stokes flow solver. The main program is a structured Euler/Navier-Stokes flow solver in which spatial discretization is accomplished by a finite volume formulation and a multigrid technique is used as a convergence accelerator. The aim is the validation of this in-house developed CFD (Computational Fluid Dynamics) tool with this enhanced enlarged capability in order to obtain a reliable flow solver that can solve flows over swept wings accurately. Various test cases were evaluated against reference solutions in order to demonstrate the accuracy of the newly implemented JK turbulence model. The selected test cases are NACA 0012 airfoil, ONERA M6 wing, DLR-F4 wing and two wings taken from the 3rd Drag Prediction Workshop. The solutions were analyzed and discussed in detail. The results show appreciably good agreement with the experimental data including force coefficients and surface pressure distributions.

In the military arena there is an increase in demand for Low Observable (LO) flight vehicles. This drive for low observability imposes limits on Leading Edge (LE) sweep angles and prohibits the use of a tailplane/fin resulting in unconventional configurations; a typical example of which are Unmanned Combat Aerial Vehicles (UCAVs). This class of aircraft poses stability and control problems due to the early onset of flow separation. The focus of this project is on the on the use of Tangential Leading Edge Blowing (TLEB) as a means of separation suppression on such vehicles. This project is unique in that the TLEB slot is positioned on the wing lower surface facing the oncoming freestream. Also, the model in this project is representative of the outboard panel of a UCAV wing, a geometry on which TLEB has not been explored in the past. A swept wing model (LE sweep = 47 degrees, AR = 3) was designed. The model has a TLEB nozzle with a slot on the lower surface at approx. 1% yawed chord that spans 0.58 m (approx. 70% LE length). Baseline wing characteristics were obtained with the full slot exposed. The wing showed a variation in pitch between CL = 0 and 0.6 which from oil flow visualisation is believed to be due to laminar separation. At CL = 0.6 there is a positive pitch break which flow visualisation suggests is due to the occurrence of a LE vortex. Sensitivity studies for slot configuration, Re number and transition fixing were carried out. The blowing rates 0.0025, 0.005, 0.025, 0.05 were tested for two slot lengths; one full span (0.58 m) and another third span positioned at the midpoint of the full slot. All blowing rates show some suppression of the LE vortex and therefore reduction in severity of the pitch break at CL = 0.6. High blowing rates produce a negative shift in CM, which CFD suggests is due to a large amount of suction produced on the lower wing surface adjacent to the slot exit. This means the available trim power is less than for the lower blowing rates. Wool tuft results for high blowing rates from the middle slot show an increase in streamwise flow at the TE suggesting TLEB is capable of improving the effectiveness of TE devices. The effectiveness of TLEB at low blowing rates has been shown to be high compared to that found in literature. A 1st order analysis of the impact of TLEB on a full scale system shows realistic options.