A large-eddy simulation (LES) of wake vortices in ground effect is conducted under the influence of airport hangar buildings to investigate the airport-specific behavior of wake vortices. We consider a landing at Tokyo/Haneda International Airport in Japan, where a wake of hangar buildings covers a landing path and touchdown zone under particular wind conditions. We combine an LES code and the Weather Research and Forecasting numerical weather prediction model to retrieve realistic wind conditions around the airport. The behavior of wake vortices is then investigated in the retrieved wind field using the adaptive mesh capability of the LES code for capturing the tiny vortex core of the wake vortex. At the crosswind on the order of a vortex descent speed, the hangar wake reduces the crosswind near the runway, enhancing the windward vortex’s rebound. The effect is alleviated by doubling the crosswind to twice the vortex descent speed. On the other hand, the wind fluctuations caused by the hangar buildings only slightly affect the wake vortex decay.

A conceptual design optimization problem for commercial transport airplanes with turboelectric propulsion, with a reasonable fidelity and comprehensiveness suitable for industrial purposes, is formulated, in order to allow for proper assessment of the benefits of turboelectric propulsion. As a sample problem, we carry out conceptual design optimization of a turboelectric propulsion airplane concept in a conventional tube-and-wing configuration with a turbofan and an associated electric fan on each (i.e., left and right) wing, varying the performance of the turboelectric propulsion devices. The results indicate that proper assessment of the benefits of the turboelectric propulsion can be carried out using the formulated optimization problem. The findings from the sample problem, including notable benefits of the turboelectric propulsion and the performance crossover point where the fuel efficiency of an airplane with conventional propulsion and that of an airplane with turboelectric propulsion cross over, are also presented.

A new model-position-sensing method for the levitation of models with a low fineness ratio (ratio of the longitudinal length to the diameter) in a magnetic suspension and balance system (MSBS) is proposed. MSBS is an ideal model-support device for wind-tunnel testing, which enables the study of flow fields around blunt bodies without flow disturbances introduced by mechanical support devices, with the aerodynamic forces determined from the magnetic forces using a pre-calibrated relationship. The new method allows wind tunnel experiments without mechanical supports with a low fineness ratio model. This method adopts two line sensors placed parallel to the central axis of the model image and measures the position with a resolution finer than 0.06 mm or deg even for thin model geometries. In addition, measurement errors were reduced by correcting a second-order term in the depth direction of the camera. A low fineness ratio circular cylinder model was levitated following sensor calibration. The model was supported in conditions with and without freestream flow. This position measurement method was also applied to a reentry capsule model. The model was levitated while keeping its position and attitude stabilized near the origin.

In this study, an aeroacoustic control mechanism of a plasma jet acting on a high-speed moving wing under a wing-in-ground effect is investigated. Moreover, a novel method is proposed to reduce the aeroacoustics of Aero-train wings. Numerical simulations of the aeroacoustics generated by flow around a National Advisory Committee for Aeronautics 4412 wing are performed under three different plasma excitation modes at four clearances with an incoming flow velocity of 0.3 Ma and an angle of attack of 5°. The results show that different plasma excitation modes interfere with the vortex generation and development in different ways to achieve aeroacoustic reduction. The UP excitation mode delays the airflow separation, delays the vortex generation and development, and reduces the vortex intensity. The BOTH excitation mode forces transverse vortices to transform into streamwise hairpin vortices and reduces the local pressure fluctuation intensity. Hence, plasma jets exhibit a good control effect on the peak aeroacoustics under different clearance conditions but result in the frequency shift effect of acoustic energy transfer to high frequencies. The modal analysis of the flow field of the three excitation conditions via a proper orthogonal decomposition method reveals that the trend of the modal change is similar for the three excitation conditions, and the change in each order of the modal corresponds to the energy decrease at the peak frequency and the energy increase at high frequencies.

This study aims to clarify the flow characteristics and the wake structure of convertible vehicles. Numerical simulations are performed to obtain a preliminary visualization, and the potential vortical motion characteristics are investigated by examining the Q-criterion across multiple cross sections. Comparisons between numerical and experimental results validate the reasonableness of our numerical model. The predominant wake topology of a two-seat convertible is obtained in terms of the location, shape, and spin direction of the vortices. We observe a “nook” vortex that is triggered by the flow acceleration induced by the pressure gradient near the windshield step, provoking undesirable aeroacoustic noise and degrading the cabin comfort. Complicated A-pillar vortex dynamics are revealed, with small vortices that are shed into the cabin and impinge the seats, eventually forming a long tail structure above the back of the vehicle. Moreover, periodic fluctuations of the windshield vortex are induced by the Kelvin–Helmholtz instability, significant impacting the streamwise wake. Ultimately, the combined motion characteristics of the A-pillar and windshield vortices exert undesirable effects on the aeroacoustic noise and drag, suggesting fundamental mechanisms for achieving optimal energy-saving and acoustic convertibles in the future. Based on the wake topology and the vortical generating mechanism, some approaches are proposed to reduce the drag and aeroacoustic noise by impeding the flow over the door into the cabin, modifying the shape of windshield step, and lengthening the windshield in stream direction.

Unsteady drag, unsteady lift, and movement of one or two moving particles caused by the passage of a planar shock wave are investigated using particle-resolved simulations of viscous flows. The particle motion analysis is carried out based on particle-resolved simulations for one or two particles under a shock Mach number of 1.22 and a particle Reynolds number of 49, and the particle migration and fluid forces are investigated. The unsteady drag, unsteady lift, and particle behavior are investigated for different densities and particle configurations. The time evolution of the unsteady drag and lift is changed by interference by the planar shock wave, Mach stem convergence, and the shock wave reflected from the other particle. These two particles become closer after the shock wave passes than in the initial state under most conditions. Two particles placed in an in-line arrangement approach each other very closely due to the passage of a shock wave. On the other hand, two particles placed in a side-by-side arrangement are only slightly closer to each other after the shock wave passes between them. The pressure waves resulting from Mach stem convergence of the upstream particle and the reflected shock waves from the downstream particle are the main factors responsible for the force in the direction that pushes the particles apart. The wide distance between the two particles attenuates these pressure waves, and the particles reduce their motion away from each other.

Multimaterial structures comprising carbon fiber reinforced plastics (CFRPs) and titanium alloys can be widely utilized in the aerospace and automotive industries to achieve optimal weight reduction and reliable structures. In this study, three-dimensional-printed titanium (3DP-Ti) adherends prepared using selective laser melting were co-bonded with woven-fabric carbon fiber reinforced phenolic matrix composite by hot pressing. Upon controlling the scanning speed and path, the 3DP-Ti adherends were prepared with various porosities and macrostructures on their surfaces, and the influences of these surface structures on the lap-shear strengths and fracture morphologies of the 3DP-Ti adherends were investigated. The lap-shear tests of the lap-shear joints with flat 3DP-Ti and pyramidal-macrostructured 3DP-Ti displayed the cohesive failure of the phenolic resin matrix. In contrast, the specimen with cylindrical macrostructure exhibited CFRP fracture and delivered the highest bond strength of 20.6 MPa, which was 18–64% greater than the strength of the lap-shear joint with the conventional Ti alloy plate and 3DP-Ti adherends with flat surface and pyramidal macrostructure. The major possible mechanisms behind the failure mode transition and the highest bond strength included the mechanical interlocking between the cylindrical macrostructure and CFRP.

Characteristics of the generation and propagation of aerodynamic noise are unique to conventional ground vehicles or aircraft due to the wing-in-ground (WIG) phenomenon. This study numerically examines the aeroacoustic characteristics of a multi-directional wing under the WIG effect with different values of clearance. The flow field was simulated by using a large eddy simulation for six groups of wings with different clearances and one group of wings in free space, at a freestream velocity of 0.3 Ma and an angle of attack of 5°. Acoustic simulations were carried out for each condition by using Möhring's analogy, and the acoustic analysis was based on details of the flow field. A multi-vortex system was generated at the trailing edge of the wing when the clearance was lower than 0.2 C. It consisted of an oscillating attached vortex that expanded and shrank with continuous changes in energy, and two periodically shedding vortices were formed that generated a strong wall-bound vortex street downstream of the wing after shedding. An analytical approach was used to simultaneously analyze different dominant frequencies in each region of the sound field. We conclude that different intensities of the WIG effect led to remarkable differences in the characteristic frequency of aerodynamic noise induced by the multi-directional wing. The characteristic frequencies of aerodynamic noise of the multi-directional wing with a low clearance (clearance ≤ 0.2 C) were mainly caused by the wall-bound vortex street while those at a high clearance (clearance ≥ 0.4 C) were mainly caused by vortices on the wingtip.

The accurate measurement of flows and aerodynamic characteristics around a bluff body has been a challenging task due to the existence of interference between the wake and mechanical model supports in wind-tunnel experiments. The present study focuses on a freestream-aligned circular cylinder with an extremely low fineness ratio (the ratio of the axial length to diameter) ranging from 0.30 to 0.50, which has never been investigated without interference from a mechanical model support. We employed a magnetic suspension and balance system to eliminate interference from the model support and measured the drag and velocity fields in the diameter-based Reynolds number between 2.0×104 and 7.7×104. As the fineness ratio decreases below 1.50, the size of the recirculation bubble increases and the velocity distribution on the central axis inside the bubble gradually converges to that of the circular disk. Furthermore, large-eddy simulations were performed in the Reynolds number of 4.0×104, whose drag coefficient agrees well with experiments. Based on those results, it was found that the drag coefficient monotonically converges to that of the circular disk without local maximum. This study revealed that in the low-fineness-ratio regime (0.10-0.50), a critical geometry, at which the drag coefficient shows a local maximum, does not exist in the circular cylinder. Subsequently, unsteady flow analyses were performed, where two characteristic frequencies, i.e., St=0.05 and 0.155, were identified from power spectral densities of the drag coefficient and the pitching moment coefficient. The associated flow structures are then extracted by a phase-averaging procedure, where the phase-averaged flows with St=0.05 represent the recirculation bubble pumping while the phase-averaged flows with St=0.155 show nonaxisymmetric structures inside the recirculation bubble.

The Turbojav is a throwing implement designed for training junior athletes in javelin throwing. However, it apparently exhibits significantly different aerodynamic characteristics compared with a typical javelin. To elucidate the aerodynamic characteristics of the Turbojav, wind tunnel tests were conducted with a 1-m magnetic suspension and balance system to avoid the influence of mechanical supports, which can significantly affect on a thin object. The results reveal that, at small angles of attack, the aerodynamic characteristics of the Turbojav depend on its fins and not on the attitude angle around the airflow axis.

In this study, the influence of the position measurement method of a magnetic suspension and balance system (MSBS) for complex models such as a spaceplane was investigated, and a new position measurement method of MSBS for the spaceplane was developed. The new position measurement method can avoid complex shape parts such as wings by placing the sensor camera at 45 degrees rotated position around the x-axis of 1.0-m MSBS. However, the second-order nonlinear position error was found to occur in this position measurement method. Therefore, it is possible to reduce the error of position by correcting the sensor output value (pixels) in which the second-order nonlinear error occurs. We succeeded in levitating the space return model within the calibration range.

We adopted the Ensemble Kalman Filter (EnKF) methodology in our computational simulation code for pedestrian flows. The EnKF, which is a type of data assimilation methodology, has been developed in the field of weather forecast where the atmospheric condition varies hour by hour. The EnKF estimates the parameters or boundary/initial conditions in the numerical model based on the updated measured data. We considered the EnKF a promising tool for the simulation of pedestrian flows, which are notoriously difficult to predict. In this study, two scenarios were conducted to confirm the usefulness of the EnKF. The first case was unidirectional pedestrian flow in straight corridors, and the second case was Mataf scenario at the Kaaba in Mecca. Needless to say, the second scenario was very challenging because of the number of pilgrims and the degrees of freedom. In each scenario, we conducted the numerical simulation using the original parameter set and then applied the EnKF to improve the accuracy of the simulation.

Abstract. The control technique in a gust alleviation system by using the airborne Doppler Lidar technology is expected to enhance aviation safety to minimize the risks of turbulence-related accidents. Accurate measurement and estimation of the vertical wind velocity are very important in the successful implementation of a gust alleviation system by using the airborne Doppler Lidar technology. An estimation algorithm of airflow vector based on the airborne Lidars is proposed and investigated for preview control to prevent turbulence-induced aircraft accidents in flight. The use of the simple vector conversion method, which is an existing technique, assumes that the wind field between the Lidars is homogeneous. The assumption of a homogeneous field would be wrong when turbulence occurs due to large wind velocity fluctuation. The proposed algorithm stores the line-of-sight (LOS) wind data with each passing moment and uses recent and past LOS wind data in order to estimate the airflow vector. The recent and past LOS wind data are used to extrapolate the wind field between the airborne twin Lidars. Two numerical experiments – ideal vortex model and numerical weather prediction – were conducted to evaluate the estimation performance of the proposed method. The proposed method has much better performance than simple vector conversion in the two numerical experiments, and it can estimate accurate two-dimensional wind field distributions unlike simple vector conversion. The estimation performance and the computational cost of the proposed method can satisfy the performance demand for preview control.

Using the world’s largest magnetic suspension and balance system (MSBS) and a low-turbulence wind tunnel, we successfully measured the aerodynamic forces acting on a non-spinning women’s javelin. It was found that the drag and the lift increased as the angle of attack was increased up to 18°. The pitching moment increased for angles of attack up to about 9°, and then decreased, becoming negative above 12°, indicating nose-down rotation. We used a pseudo supporting rod to simulate a javelin attached to a support, as used in a conventional setup, and confirmed that this interferes with the javelin by creating differences between the aerodynamics forces acting on the javelin with and without the pseudo supporting rod.

The purpose of this study was to investigate the effect of the air permeability of ski jumping suit fabric on aerodynamic characteristics. In the study, four types of fabric with different air permeabilities were installed onto a fabric-clothed elliptic cylinder, and the flow behavior around its surface was investigated through wind tunnel experiments. The stall was delayed by using the fabric with higher air permeability. The air flow permeated the surface of the elliptic cylinder through the fabric for the fabric with high air permeability. This air flowed out into the separation region through the fabric again, which suppress stall. The stall characteristics for the fabric-clothed elliptic cylinder were influenced by the air permeability of the fabric. The boundary layer thickness on the outer surface and the turbulence levels of the boundary layer affected the air permeability of the fabric. The higher air permeability fabrics improved the stall characteristics.

<p>The aerodynamic coefficients of the turbo-jav (a short javelin for junior or disabled athletes) up to 26 degree angle of attack were obtained using Magnetic Suspension and Balance System (MSBS). Ordinary thin models are susceptible to interference caused by mechanical supports. However, MSBS can eliminate the support interference because the model is floating supported by magnetic force instead of mechanical supports. Drag, lift force and pitching moment coefficients were measured in the wind tunnel tests at the wind speed 25 m/s (<i>Re</i> = 6.4 × 10⁴) from 0 to 26 degrees angle of attack. This experiment agreed well with the results of the calculation using ANSYS FLUENT. Especially for pitching moments, the present measurement agreed the calculation better than the previous results using mechanical supports. It suggested that the results using MSBS was more reliable. In addition, the experimental results suggested that the aerodynamic characteristics of the turbo-jav depended on the fins of the turbo-jav but not on the rotation angle around air flow axis of the turbo-jav.</p>

<p>Using the world's largest 1-m MSBS (Magnetic Suspension and Balance System) and low-turbulence wind tunnel at Tohoku University, we successfully measured the aerodynamics force acting on a non-spinning women's javelin with no supporting rod. The primary advantage of the MSBS is the complete elimination of any support, which can interference with the measurements. The advantage of the world's largest 1m-MSBS at Tohoku University is that it is possible to employ a real javelin at relatively high <i>AoA</i>. A women's javelin was used for the wind tunnel tests. It was found that the drag and the Lift increase with increasing the angle of attack up to 18 °. The pitching moment increases up to about 9 °, and it decreases up to 18 °. The pitching moment becomes negative above 12 °. The nose-up rotation was determined as positive. We also confirmed the support interferences with using a pseudo supporting rod. The tip of the pseudo supporting rod was set at the distance of 2 mm away from the surface of the javelin at the angle of attack of 0 °.It was confirmed that there are some differences between the aerodynamics forces with the pseudo supporting rod and those without it. The absolute aerodynamic forces without the pseudo supporting rod become smaller.</p>

&lt;p&gt;The aerodynamic characteristics of sports suits play a crucial role in helping professional athletes gain a competitive edge in their chosen sport. The fluid forces acting on the body of the sports competitor are influenced by the surface topology of the fabric, position and size of the seams, use of fasteners, and air permeability of the fabric. Wind tunnel experiments and numerical simulations provide deep insight on the aerodynamic characteristics of sports suit designs. Ski jumping is a sport in which the competitor is judged based on the flight form and flight distance; therefore, ski jumpers attempt to fly as far as possible upon take-off. The fabric used for ski jumping suits plays a crucial role in achieving favorable aerodynamic characteristics. According to the rules and regulations for ski jumping competitions, the outstretched fabric should have a minimum air permeability of 40 L/m&lt;sup&gt;2&lt;/sup&gt;/s at a water pressure of 98 Pa (10 mmAq). To date, little is known regarding the effects of air permeability on the aerodynamic characteristics of ski jumping suits, which forms the motivation of this study. The purpose of this study is to investigate the effect of air permeabi

&lt;p&gt;In ski jumping competition, there are many regulations to ensure fairness. In the international ski jumping competition rules, suit fabric must show the air permeability of a minimum of 40 liters per m2/sec under 10 mm water pressure. However, the effect of air permeability of suit fabric on aerodynamic characteristics has not been completely clarified yet. The purpose of this study is to investigate the effect of air permeability of suit fabric on aerodynamic characteristics. The fabrics with different air permeability are used and the wind tunnel experiments were carried out using the fabric clothed elliptic cylinder. The experimental results show that the stall characteristics are greatly affected by the air permeability. Furthermore, the flow around the fabric-clothed elliptic cylinder also varies, depending on the position of the fabric&#039;s air permeability on the upper surface of the elliptic cylinder. It is presumed that the aerodynamic performance of ski jumpers during a flight can be improved by adjusting the air permeability of the suit fabric.&lt;/p&gt;

&lt;p&gt;The aerodynamic characteristics of the sports suit play an important role in the victory or defeat of sports. Furthermore, the fluid forces acing on the body of the player are influenced by a surface shape of the cloth for suit, a position or the fastener, and the air permeability. Therefore, a wind tunnel examination and numerical analysis for the design of the suit greatly contributes to these sports. Ski jumping is a sport which competes the flight distance and the flight form, and jumpers attempt to fly as far as possible. The fabric used in ski jumping plays an important role to obtain aerodynamic benefits. In the regulations for ski jumping competition, the outstretched fabric must show a medium air permeability of a minimum of 40 L/m2/sec with 10 mm water pressure. However, the effect of the air permeability of ski jumping suits on aerodynamic characteristics has not been completely clarified yet. The purpose of this study is to investigate the effect of air permeability of jumping suit fabric on aerodynamic characteristics, and four types of fabric with different air permeability are considered from aerodynamic point of view. The wind tunnel experiments were carried out

© 2018, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved. Large eddy simulation of wake vortices in ground effect with realistic terrain around an airport was conducted focusing on the airport-specific behavior of wake vortices. We considered here a landing at Tokyo/Haneda airport in Japan, where a wake of hangar buildings covers landing paths under a certain wind condition. We developed a large eddy simulation code for adaptive mesh refinement capturing fine vortex core of wake vortex. Outer boundary conditions of the simulation were provided by the Weather Research and Forecasting (WRF) numerical prediction model to retrieve the wind condition. The impact of a crosswind over airport buildings on the behavior of wake vortex was evaluated by visualized images so far.

In this study, the aerodynamic analysis of NASA Common Research Model (wing-body configuration) was conducted by a block-structured Cartesian mesh method named Building-Cube method. The building-cube method (BCM) has the advantages of the quick and robust mesh generation and efficient parallel computing, and it is easy to run on a large-scale computing system. However, the BCM suffers a restriction on the resolution of the turbulent boundary layer. The aim of this study is to improve the resolution issue of the turbulent boundary layer at high Reynolds number flows using the immersed boundary method. By utilizing the advantages of the BCM, we challenge to resolve the boundary layer by mesh subdivision. In this paper, the results of the validation before subdivision mesh were shown by using a coarse mesh. The computational results were compared with the transonic wind tunnel test and the results of another flow solver. Although the results showed a similar tendency with the results of another flow solver, the decomposed aerodynamic coefficient appeared with some discrepancy.

High-resolution turbulent-flow simulations using unsteady computational fluid dynamics (CFD) have been widely applied to research and development in the aerospace and mechanical engineering industries. In this study, we propose a data-exploration method to detect anomaly elements between two unsteady simulation data sets which have different structures by sparse structure learning. This method is tested with unsteady pressure-distribution data for two models of an RAE 2822 airfoil with/without the transition trip. The method detects not only the obvious change elements, such as the transition trip, but also small change elements, such as delay of turbulent transition, which are easily overlooked with conventional visualization methods.

The purpose of this study is to measure the aerodynamic forces acting on an aircraft model at high angles of attack with a wind tunnel test using a magnetic suspension and balance system (MSBS). In the magnetic support of the model at high angles of attack, the control characteristics of the model’s motion by MSBS change. The change of the control characteristic might cause an undesirable additional motion in the model’s control. As a result, it is difficult to maintain the evaluation accuracy of a magnetic force. Therefore, the model control system at high angles of attack was proposed in consideration of the change in the magnetic force applied to the magnet in the model. Furthermore, the efficient force evaluation method at high angles of attack excluding rolling moment is presented analytically. As a result, the forces acting on the model in a wide range of angles of attack were evaluated with an error range of about 4%. Finally, three components of the aerodynamic forces acting on the AGARD-B model were acquired at a flow speed of 15 m/s and angles of attack in the range of [0, 30] degrees.

© 2018 The Japan Society for Aeronautical and Space Sciences. The effects of Mach number at Re = 3,000 for different airfoils (NACA0012, NACA0002, NACA4412, NACA4402) with thickness and camber geometries are investigated for the propeller blade design of a Mars airplane. The present study shows that thin and cambered airfoils have larger variations in Cl than symmetric airfoils. As for thin airfoils, Cl at higher ¡ has rapid increases when the M¨ is low. This is because the flow separation occurs at the leading edge, and the flow is reattached on the airfoil surface. However, the rapid increase in Cl disappear as M¨ increases because the flow reattachment does not occurs. As for cambered airfoils, the decrease in Cl becomes larger than that on the symmetric airfoils when M¨ is higher. This is because Cp near the leading edge on the lower surface is smaller than that on the upper surface and the high-speed region on the lower side of the leading edge is enlarged as M¨ increases. Then, the Mcr at Re = 3,000 tends to be larger than that predicted by linear theory.

Computationally expensive multiobjective optimization problems are difficult to solve using solely evolutionary algorithms (EAs) and require surrogate models, such as the Kriging model. To solve such problems efficiently, we propose infill criteria for appropriately selecting multiple additional sample points for updating the Kriging model. These criteria correspond to the expected improvement of the penaltybased boundary intersection (PBI) and the inverted PBI. These PBI-based measures are increasingly applied to EAs due to their ability to explore better nondominated solutions than those that are obtained by the Tchebycheff function. In order to add sample points uniformly in the multiobjective space, we assign territories and niche counts to uniformly distributed weight vectors for evaluating the proposed criteria. We investigate these criteria in various test problems and compare them with established infill criteria for multiobjective surrogate-based optimization. Both proposed criteria yield better diversity and convergence than those obtained with other criteria for most of the test problems.

We developed a framework for a distributed-memory parallel computer that enables dynamic data management for adaptive mesh refinement and load balancing. We employed simple data structure of the building cube method (BCM) where a computational domain is divided into multi-level cubic domains and each cube has the same number of grid points inside, realising a multi-level block-structured Cartesian mesh. Solution adaptive mesh refinement, which works efficiently with the help of the dynamic load balancing, was implemented by dividing cubes based on mesh refinement criteria. The framework was investigated with the Laplace equation in terms of adaptive mesh refinement, load balancing and the parallel efficiency. It was then applied to the incompressible Navier–Stokes equations to simulate a turbulent flow around a sphere. We considered wall-adaptive cube refinement where a non-dimensional wall distance y+ near the sphere is used for a criterion of mesh refinement. The result showed the load imbalance due to y+ adaptive mesh refinement was corrected by the present approach. To utilise the BCM framework more effectively, we also tested a cube-wise algorithm switching where an explicit and implicit time integration schemes are switched depending on the local Courant-Friedrichs-Lewy (CFL) condition in each cube.

Fluid dynamics simulation is often repeated while changing conditions, and therefore we need to compare a large amount of results. In order to compare results under different conditions, it is effective to overlap the streamlines generated from each condition in a single 3D space. Streamline is a curved line which represents a wind flow. This paper presents a technique to automatically select and visualize important streamlines suitable for comparison of the simulation results. In addition, we present an implementation to observe the flow fields in virtual reality spaces.

The interaction of aircraft exhaust jet and wake vortex generated from a realistic aircraft configuration was numerically studied from the early phase until the complete roll-up. For numerical simulation we employed an unstructured mesh Reynolds-averaged Navier-Stokes (RANS) solver for the flow around an aircraft model, while a Cartesian mesh large-eddy simulation solver implemented on the framework of the building-cube method (BCM) was used for the simulation of the mid- to far-field wake. We especially investigated the feasibility of a coupling approach of the steady RANS solver and the Cartesian LES solver in the simulation of the jet-wake vortex interaction, where unsteady velocity fluctuations are stochastically reproduced based on RANS turbulence variables. The results showed that a multi-level Cartesian mesh of the BCM was effective to capture the evolution of aircraft wake including exhaust jet. The velocity profiles of a wake vortex are compared with several vortex models, which reveals that the velocity profiles fit with different vortex models depending on the vortex ages. The normalized temperature of exhaust jet is compared with the experiment, which confirms that the temperature information is appropriately transferred from the near- to the far-field simulation. (C) 2017 Elsevier Masson SAS. All rights reserved.

The Kriging surrogate model, which is frequently employed to apply evolutionary computation to real-world problems, with a coordinate transformation of the design space is proposed to improve the approximation accuracy of objective functions with correlated design variables. The coordinate transformation is conducted to extract significant trends in the objective function and identify the suitable coordinate system based on either one of two criteria: likelihood function or estimated gradients of the objective function to each design variable. Compared with the ordinary Kriging model, the proposed methods show higher accuracy in the approximation of various test functions. The proposed method based on likelihood shows higher accuracy than that based on gradients when the number of design variables is less than six. The latter method achieves higher accuracy than the ordinary Kriging model even for high-dimensional functions and is applied to an airfoil design problem with spline curves as an example with correlated design variables. This method achieves better performances not only in the approximation accuracy but also in the capability to explore the optimal solution.

<p>Badminton is one of the most popular sports in the world and is famous as the sport having the fastest initial velocity of a batted ball among all ball games. Initial velocity immediately after smashing may reach up to 408 km/h (113 m/s) at maximum. A badminton shuttlecock generates significant aerodynamic drag and it was confirmed that the high deceleration characteristics was related to the slots located at the leg portion of a shuttlecock in the previous study. Turnover refers to the flipping experienced by a shuttlecock when undergoing heading change from nose pointing against the flight path at the moment of impact and a shuttlecock indicates the aerodynamically stable feature for the flip movement just after impact. The purpose of this study is to investigate the effect of gaps on the aerodynamic stability (turnover stability) of a badminton shuttlecock during the flip phenomenon. In the present study, the flow field around the shuttlecock during impulsive change of an angle of attack (flip movement) was measured by using the smoke flow visualization and the behavior of the shuttlecock during the flip movement was evaluated in comparison with that of the conic model (with no gaps). The turnover stability of a badminton shuttlecock is affected by gaps of the shuttlecock skirt.</p>

The vortex shedding in combustion chamber of solid rocket motor may lead the pressure oscillations, which deteriorates the rocket motor performance and could even damage the payload. Cartesian mesh CFD solver is expected to accurately predict such vortex flows using higher-order scheme and to treat complicated configurations easily. Therefore, blockstructured Cartesian mesh solver, Building-Cube Method (BCM), is applied to Taylor flow analysis, which is the simplified combustion chamber model of solid rocket. Immersed boundary method is extended to treat the wall injection from the grain surface. The approach is also applied to real grain configuration, and the usefulness for the complicated geometry is validated.

© 2017, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved. We conducted large eddy simulation of wake vortices in ground effect with realistic terrain around an airport focusing on the airport specific behavior of wake vortices. We considered here a landing aircraft at Tokyo/Haneda Airport in Japan. We developed a framework for a distributed parallel computer that enables dynamic data management for load balancing and flow adaptive mesh refinement for wake vortex. The impact of a crosswind over airport buildings was only qualitatively evaluated so far by visualized images.

This paper proposes a novel approach for fluid topology optimization using genetic algorithm. In this study, the enhancement of mixing in the passive micromixers is considered. The efficient mixing is achieved by the grooves attached on the bottom of the microchannel and the optimal configuration of grooves is investigated. The grooves are represented based on the graph theory. The micromixers are analyzed by a CFD solver and the exploration by genetic algorithm is assisted by the Kriging model in order to reduce the computational cost. Three cases with different constraint and treatment for design variables are considered. In each case, GA found several local optima since the objective function is a multi-modal function and each local optimum revealed the specific characteristic for efficient mixing in micromixers. Moreover, we discuss the validity of the constraint for optimization problems. The results show a novel insight for design of micromixer and fluid topology optimization using genetic algorithm.

In order to achieve sonic boom reduction in real-world atmosphere, we investigate the effects of uncertainties in atmospheric turbulence and numerical weather prediction (NWP) models on sonic boom. We simulate sonic boom propagation of the Drop test for the Simplified Evaluation of Non-symmetrically Distributed sonic boom (D-SEND) #2 flight test, which was conducted by the Japan Aerospace Exploration Agency (JAXA), using the Tohoku University Xnoise (TUXnoise). TUXnoise is a sonic boom analysis tool based on the Khokhlov-Zabolotskaya-Kuznetsov (KZK) equation and capable to consider diffraction, axial convection, and transverse convection effects induced by atmospheric turbulence. First, the sonic boom propagation is simulated with different transverse size of the simulation domain, which is related to the reproducibility of the effects of atmospheric turbulence. The results indicate that the simulation requires a sufficiently large domain size for a good agreement with the measurement in D-SEND#2 flight test. Second, the sonic boom propagation is simulated under the atmospheric uncertainties such as temperature, humidity, and wind velocity uncertainties estimated by ensemble forecast. The results indicate that wind velocity uncertainty can affect sonic boom significantly through a consideration of atmospheric turbulence. They also indicate that the effect of humidity uncertainty on minimum peak pressure is small because the relaxation effects in the KZK equation depend on humidity and inactive in rounded sonic boom signatures, which result in minimum peak pressure. In addition, it is confirmed that the effects of atmospheric turbulence can change significantly even with a slight change of the propagation path. Finally, the sonic boom propagation is simulated with different atmospheric boundary layer (ABL) height (ABLH). Atmospheric turbulence is assumed to exist in ABL. The results indicate that ABLH must be estimated carefully to evaluate sonic boom properly because the sonic boom is affected more significantly as ABL is thicker. They also indicate that wind shear layers over ABL can be other factors to produce atmospheric turbulence and affect the sonic boom signature when D-SEND#2 flight test was conducted. Hence, the uncertainties in atmospheric turbulence and NWP models must be carefully taken into account for more accurate simulation of sonic boom propagation.

The aim of this computational study is to accurately predict complicated flowfields and obtain accurate aerodynamic forces of a thin folding plate under low Reynolds number range. The thin folding plate is a bent flat plate at a point, and is known to cause the non-linear lift curve at the flow regime. One of Cartesian-Mesh CFD solvers, Building-Cube Method (BCM), is used to compute the flows around the thin airfoils. The computed results of a thin flat plate and a thin folding plate are compared with the experimental results. Throughout this study, the effectiveness of BCM against thin airfoils at low Reynolds number flows is discussed. It is expected that the rapid increase of negative pressure in separation bubble at leading-edge causes the increase of Cl, which leads the non-linear lift curve.

Mesh generation is one of the critical problems in numerical simulation of cascade flow, particularly cooled turbine blade. To overcome this problem, Cartesian-mesh CFD solver is focused these days. In this study, block-structured Cartesian-mesh solver, Building-Cube Method (BCM), is applied to cascade flow of T106 low pressure turbine cascade. Ghost-cell approach is employed to represent the curved surface of T106 blade. The result is compared with the result of wall boundary condition with staircase representation. As a result, Ghost-cell approach can mitigate the pressure oscillations and obtain more smooth results. The computational results are also compared with experimental results for the validation, and the similar surface pressure coefficients are obtained. However, the difference is observed at the location of laminar-turbulent transition due to the lack of mesh resolution.

A multi-objective optimization (MOO) approach for multi-point injection of resin-transfer molding (RTM) is proposed to investigate the trade-off relationship between productivity and quality for composite structures. With this approach, the optimum gate positions for their molding properties are evaluated using finite-element analysis (FEA) with a multiple-objective genetic algorithm (MOGA), and the trade-offs are visualized with the combination of a self-organizing map (SOM) and a scatter plot matrix (SPM). We applied this approach to RTM for flat-plate and rib models. For the flat-plate model, we found a negative correlation between fill time and weld line contents, and between fill time and dry spot contents. These results imply difficulty for simultaneous reduction of cycle time and void contents. For the rib model, some tendencies agree with those of the flat-plate model, and others do not. This difference comes from complexity of structure. We also found Pareto solutions that satisfy both productivity and quality for the flat-plate and rib models (i.e., gate positions such as a combination of diagonal and quadrangle positions for the flat-plate model, and aggregation of gate positions at a beam part for the rib model). Furthermore, we conducted simple experiments for the flat-plate model to validate the simulation result. The trends acquired from MOO qualitatively agree with the experiments.

A non-gradient-based approach for topology optimization using a genetic algorithm is proposed in this paper. The genetic algorithm used in this paper is assisted by the Kriging surrogate model to reduce computational cost required for function evaluation. To validate the non-gradient-based topology optimization method in flow problems, this research focuses on two single-objective optimization problems, where the objective functions are to minimize pressure loss and to maximize heat transfer of flow channels, and one multi-objective optimization problem, which combines earlier two single-objective optimization problems. The shape of flow channels is represented by the level set function. The pressure loss and the heat transfer performance of the channels are evaluated by the Building-Cube Method code, which is a Cartesian-mesh CFD solver. The proposed method resulted in an agreement with previous study in the single-objective problems in its topology and achieved global exploration of non-dominated solutions in the multi-objective problems. (c) 2016 The Authors International Journal for Numerical Methods in Engineering Published by John Wiley & Sons Ltd

Understanding of the three-dimensional shock wave-vortex loop interaction phenomena plays a key role in noise reduction. This study focuses on the three-dimensional shock wave distortion and propagation phenomena in a near-field supersonic jet. Shock-square vortex loop interaction was experimentally investigated in a square cross-sectional open-end shock wave generating tube at an incident shock Mach number of 1.39 +/- 0.05. A square vortex loop impinged on a reflected shock wave from a wall located in front of the nozzle end. The planar reflected shock wave transforms into either a concave or convex distorted shape due to the opposing high-speed flow emitted from the nozzle corner. The convex shaped shock wave scatters towards the outside of the vortex loop, whereas the concave one converges towards the centre of the vortex loop. The concave shaped shock wave results in shock wave focusing. In shock square vortex loop interaction, the shock wave is locally focused along the axis of the nozzle corner. (C) 2016 Elsevier Inc. All rights reserved.

A sonic boom signature with a long rise time has the ability to reduce the sonic boom, but it does not necessarily minimize the sonic boom at the ground level because of the real atmospheric turbulence. In this study, an effect of the turbulence on a long rise-time pressure signature was experimentally investigated in a ballistic range facility. To compare the effects of the turbulence on the long and short rise-time pressure signatures, a cone-cylinder projectile that simultaneously produces these pressure signatures was designed. The pressure waves interacted with a turbulent field generated by a circular nozzle. The turbulence effects were evaluated using flow diagnostic techniques: high-speed schlieren photography, a point-diffraction interferometer, and a pressure measurement. In spite of the fact that the long and short rise-time pressure signatures simultaneously travel through the turbulent field, the turbulence effects do not give the same contribution to these overpressures. Regarding the long rise-time pressure signature, the overpressure fluctuation due to the turbulence interaction is almost uniform, and a standard deviation 1.5 times greater than that of the no-turbulence case is observed. By contrast, a short rise-time pressure signature which passed through the same turbulent field is strongly affected by the turbulence. A standard deviation increases by a factor of 14 because of the turbulence interaction. Additionally, there is a non-correlation between the overpressure fluctuations of the long and short rise-time pressure signatures. These results deduce that the length of the rise time is important to the turbulence effects such as the shock focusing/diffracting. (C) 2016 Elsevier Ltd. All rights reserved.

An intrinsic relation is given between the skin-friction vector and the surface pressure gradient through the boundary enstrophy flux (BEF), and it is used to study the possibility to extract some skin-friction structures from a surface pressure field. This attempt contains two related parts. In the first part, when the BEF field is given, a projected skin-friction field in the image plane can be sought from a surface pressure image based on a variational solution of an optical-flow-like equation. This approach is validated in several classical flows. The second part deals with a practical problem in which the BEF field is not known a priori. In this case, a so-called auxiliary skin-friction field is determined from a surface pressure image alone by using the same variational approach. The auxiliary skin-friction field has the magnitude proportional to the skin-friction magnitude and the direction of the negative surface pressure gradient. The physical meaning of the auxiliary skin-friction field and its applicability to global skin-friction diagnostics are discussed.

Multi-objective optimization of vortex generators on a transonic infinite wing is performed using computational fluid dynamics and a multi-objective genetic algorithm coupled with surrogate models. Vortex generator arrangements are defined by five design variables: height, length, incidence angle, chord location, and spacing. The objective functions are to maximize the lift-to-drag ratio at low angle of attack, to maximize lift coefficient at high angle of attack, and to the shift chordwise separation location downstream at high angle of attack. To evaluate these objective functions of each individual in the multi-objective genetic algorithm, the ordinary kriging surrogate model and the radial-basis-function/kriging hybrid surrogate model are employed because numerical analysis of the wing with vortex generators requires a large amount of computational time. Nondominated solutions are classified into five clusters with different aerodynamic characteristics. Comparison of the five clusters revealed that the balance among three objective functions is controlled mainly by vortex generator height, spacing, and their ratio. The solutions in each cluster have specific values of these three parameters, which identify the aerodynamic characteristics. In addition, appropriate values of design variables for generating the vortex most effectively are investigated.

An immersed boundary method (IBM) suitable for aeroacoustic analysis is developed using the block-structured Cartesian mesh method called the building-cube method. The IBM using a ghost cell approach in conjunction with an image point is employed. A hybrid interpolation scheme combining the wavenumber-based method and inverse distance weighting method is introduced to the IBM. This method is evaluated using four test cases. The results of interpolation at an image point indicate that the wavenumber-based method provides more accurate results than standard interpolation methods for aeroacoustic problems. The results of acoustic scattering around a sphere show that the proposed IBM achieves better than second-order accuracy. In the case of noise propagation from a generic inlet, the proposed IBM is shown to yield more accurate predictions for the far-field sound pressure level than the inverse distance weighting-based IBM. In a different application involving the JT15D nacelle, the far-field sound pressure level based on the proposed IBM is in good agreement with the experiment and computational results reported in a previous study. These results validate the conclusion that the proposed IBM can compute noise propagation with a mean flowfield and accurately estimate the far-field sound pressure level.

The magnetic suspension and balance system for the empty set85 mm suction-type supersonic wind tunnel at Tohoku University, which is capable of magnetically suspending the model against the starting and stopping loads without mechanical support, has been successfully developed. To implement fast control system response, a new optical position sensor was built and calibrated. The 1.24 kHz synchronous control system, composed of a proportional-integral controller plus a double-phase advancer, made the rise time of a step response 30% faster than the Japan Aerospace Exploration Agency's existing magnetic suspension and balance system. The frequency characteristics of the magnetic field intensity formed in the test section were measured and improved to achieve faster performance of the control system. The control simulation of the magnetic suspension and balance system for the empty set85 mm suction-type supersonic wind tunnel at Tohoku University indicated the appropriate conditions, such as increasing the moment of inertia of the model and setting a cutoff frequency of 100 Hz, and led to successful magnetic suspension tests.

The effects of plate angle on acoustic waves from a supersonic jet impinging on an inclined flat plate at angles of 30, 45, and 60 deg are numerically investigated. Three-dimensional compressible Navier-Stokes equations are solved using the modified weighted compact nonlinear scheme. Similar to previous studies, the acoustic fields indicate that there are at least three types of acoustic waves in all of the cases considered herein: 1) Mach waves generated from the shear layer of the main jet, 2) acoustic waves generated from the impingement region, and 3) Mach waves generated from the shear layer of the supersonic flow downstream of the jet impingement. Acoustic waves (2) are generated from two different acoustic sources: 1) the interaction between the plate shock wave and the shear layer, and 2) the interaction between the bubble-induced shock waves and the shear layer. The frequency characteristics of acoustic waves are related to the thickness of the shear layer in the impingement region. The results of the present study indicate the source location and the characteristics of acoustic waves (2) for various flat-plate angles.

The purpose of the research is to characterize the ionization process behind a shock wave with precursor photoionization in argon. In this study, the H-β line is observed by spectroscopic measurements using a hypersonic shock tube and the Stark broadening of the H- β line is evaluated to obtain the electron density behind a shock wave. As a result, the electron density is on the order of 10²¹ m^-3 and tends to decrease with increasing the distance from the shock front. The decrease of the measured electron density is caused by the radiative energy loss from the test gas because the radiative energy loss corresponds to a decrease of the temperature or the density of the test gas, both of which can decrease the electron density. Previous study showed that photoionization occurred ahead of the shock wave, making use of the radiation energy emitted from the region behind the shock wave. The fact might be related to the radiative energy loss obtained in the present study. From the present study, it is found that the radiative energy loss is dominant behind the shock wave under the condition that precursor photoionization occurs ahead of the shock wave. In future, we should investigate the radiative transfer phenomena around the shock wave for further clarification.

<p>An experimental study and a computational study are conducted to understand the aerodynamics of the shuttlecock especially about the effect of the chamber of the blade. The lift, drag, pitching moment coefficient is measured in the experiment. It was found that the chamber of the blade can change the aerodynamic properties. The increase in camber causes increase in lift and a decrease in drag. PIV measurements revealed that this decrease in the drag is related to the decrease in the width of wake which is caused by the flow of outside of the shuttle flows along the chamber. A pressure coefficient calculated by CFD reveal that the each blade work as single blade at low angle of attack. On the other hand, at high angle of attack, not so high lift was observed. It is because the front of the outside of the upper blade is inside of the separation region from the cork.</p>

© 2016, National Research Council Canada. In this study, the uncertainty quantification of wake vortex parameters measured by a Doppler lidar is conducted to assess the errors of the lidar-derived parameters due to the measurement process. We employ three methods to estimate vortex parameters from lidar measurements: a simple method to detect velocity peaks on the measurement plane, and a least-square fitting of a vortex model to lidar velocity. In addition, wake vortex parameters along with a surrounding wind field are estimated based on the four-dimensional variational (4D-Var) method assuming a velocity field obtained by the Rosenhead-Burnham-Hallock vortex model. With the last approach, vortex parameters are obtained considering the factors such as insufficient resolution in the line-of-sight direction and vortex movement during a laser scan. It is confirmed from numerical experiments that the average circulation is significantly underestimated and core radius becomes too large with the simple method. The least-square fitting reduces the deviation of output/input parameter ratios. On the other hand, vortex parameters estimated with the help of the 4D-Var method are further improved by compensating the errors due to the measurement process.

Topology optimization is an optimization method which can modify connectivity of an object independently of its predefined topology. In this paper, a global optimization method for topology optimization of flow channels considering fluid and heat transfer using a genetic algorithm is presented. A genetic algorithm (GA) is assisted by the Kriging surrogate model to reduce computational cost required for function evaluation. In the present method, the boundary of a flow channel is represented by a level set function. Topology optimization seldom employs GA since topology optimization requires a large number of design variables for a high degree of freedom for shape and topology representation and GA is not effective to handle such a large scale problem. This paper presents a novel representation method to obtain the distribution of level set function with a reasonable number of design variables. The design variables are given at the scattered control points in the design domain, and the Helmholtz equation is solved in the entire domain. The proposed method is applied to a single-objective optimization problem to maximize heat transfer. As a result, GA found several flow channels, each of which has similar objective function values but with different topology. The result indicates that the objective function is a multi-modal function, which means that a method of population-based multipoint simultaneous exploration such as GA is essential for the present topology optimization problem. Considering minimizing pressure loss of a flow channel as the second objective function, the proposed method is applied to a multi-objective optimization problem. As a result, we confirm that the proposed representation method enables to represent flow channels that balance both objective functions and GA captures the trade-off between two objective functions.

An integrated method consisting of a proper orthogonal decomposition (POD)-based reduced-order model (ROM) and a particle filter (PF) is proposed for real-time prediction of an unsteady flow field. The proposed method is validated using identical twin experiments of an unsteady flow field around a circular cylinder for Reynolds numbers of 100 and 1000. In this study, a PF is employed (ROM-PF) to modify the temporal coefficient of the ROM based on observation data because the prediction capability of the ROM alone is limited due to the stability issue. The proposed method reproduces the unsteady flow field several orders faster than a reference numerical simulation based on Navier-Stokes equations. Furthermore, the effects of parameters, related to observation and simulation, on the prediction accuracy are studied. Most of the energy modes of the unsteady flow field are captured, and it is possible to stably predict the long-term evolution with ROM-PF.

近年災害時における状況確認，沿岸監視，テロ現場での情報収集等を目的として，小型飛行体（MAV : Micro Air Vehicle）の研究開発が盛んに行われている．MAVに要求されている外形サイズ，飛行及び作業形態は，現存の航空機と異なるものであり，革新的な技術開発が必要とされる．本研究では，ステンレス製楕円翼を風洞内でヒービング運動させ，6分力ロードセルを用いて翼厚及びヒービング運動条件が空力特性に与える影響を調査した．さらに模型に組み込まれているエンコーダ出力を基に，特定のヒービング角度で計測を行うPIVトリガー計測システムを構築し，空力特性と前縁剥離渦の挙動との関連性について調査した．その結果，ヒービング振幅，板厚を変化させると，空力特性が変化し，固定翼と比較し良好な特性を示す組合せが存在する事が示された．また，優れた空力特性を示した羽ばたき条件におけるPIVトリガー計測を行った結果，ダウンストローク時において下面からのまき込み渦をきっかけにして，大規模な前縁剥離渦が翼上面近傍に長い間存在している事が確認できた．さらに流体力とPIVの同時計測を実施し，薄翼においてヒービング運動と揚力変動の位相が45°ずれており，前縁剥離渦が翼上面近傍に存在している間，揚力が増加する事が示された．

An integrated method of a proper orthogonal decomposition based reduced-order model (ROM) and data assimilation is proposed for the real-time prediction of an unsteady flow field. In this paper, a particle filter (PF) and an ensemble Kalman filter (EnKF) are compared for data assimilation and the difference in the predicted flow fields is evaluated focusing on the probability density function (PDF) of the model variables. The proposed method is demonstrated using identical twin experiments of an unsteady flow field around a circular cylinder at the Reynolds number of 1000. The PF and EnKF are employed to estimate temporal coefficients of the ROM based on the observed velocity components in the wake of the circular cylinder. The prediction accuracy of ROM-PF is significantly better than that of ROM-EnKF due to the flexibility of PF for representing a PDF compared to EnKF. Furthermore, the proposed method reproduces the unsteady flow field several orders faster than the reference numerical simulation based on the Navier-Stokes equations.

Flow characteristics and aerodynamic losses of film cooling injection with compound angles of 30 degrees, 60 degrees, and 90 degrees for velocity ratios of 0.5, 1.0, and 2.0 are numerically investigated using the high-quality hybrid meshes comprising tetrahedrons, prisms, and pyramids. The solutions are obtained by solving the compressible Reynolds-averaged Navier-Stokes equations, and the cell-vertex finite volume method is used to discretize these equations. The predicted results are firstly validated against the experimental data. It has been found that the variation tendency of mass average net total pressure loss coefficient with compound angles is reasonably predicted and in many cases well predicted. The analyses based on net total pressure losses show that the aerodynamic losses tend to increase with the increasing of compound angle regardless of the velocity ratio, and the tendency becomes more obvious when the velocity ratio is large. On the other hand, some typical cases are additionally calculated with different turbulence models and flux computation schemes for the comparisons. A latest proposed numerical flux computation method, simple low-dissipative AUSM (SLAU) scheme, is adopted in the predictions. The results indicate that SLAU scheme is slightly superior to Harten-Lax-van Leer-Einfeldt-Wada scheme in the prediction of compound angle film cooling injection.

For high-speed air breathing engines, fuel injection and subsequent mixing with air is paramount for combustion. The high freestream velocity poses a great challenge to efficient mixing both in macroscale and microscale. Utilising cavities downstream of fuel injection locations, as a means to hold the flow and stabilise the combustion, is one mechanism which has attracted much attention, requiring further research to study the unsteady flow features and interactions occurring within the cavity. In this study we combine the transverse jet injection upstream of a cavity with an impinging shock to see how this interaction influences the cavity flow, since impinging shocks have been shown to enhance mixing of transverse jets. Utilising qualitative and quantitative methods: schlieren, oilflow, PIV, and PSP the induced flowfield is analysed. The impinging shock lifts the shear layer over the cavity and combined with the instabilities generated by the transverse jet creates a highly complicated flowfield with numerous vertical structures. The interaction between the oblique shock and the jet leads to a relatively uniform velocity distribution within the cavity.

In this study, experiments were conducted to evaluate the accuracy of the pressure measurement system in a ballistic range at Tohoku University. An effect on the near-field pressure waveform by shock wave interaction with a pressure measurement plate instrumented with pressure transducers was investigated. The pressure waveform changed depending on the measurement positions on the plate. It is important to place pressure measurement positions away from the front and rear edges of the plate. In addition, pressure measurement accuracy was validated by comparison of the pressure waveforms between present and previous results. The pressure was measured at a position away from the edges of the plate. The present data corresponds well with existing experimental data. The results show that the pressure measurement system in the ballistic range at Tohoku University has the ability to accurately measure the near-field pressure waveform. (C) 2015 Elsevier Ltd. All rights reserved.

A computational-fluid-dynamics-based aeroelastic analysis method is proposed to simulate control-surface limitcycle oscillation (LCO) induced by freeplay gap. The present method is based on the previously proposed aeroelastic reduced-order model (ROM), in which an unsteady aerodynamic state-space model is generated from aerodynamic responses to step excitation of individual mode using the eigensystem realization algorithm, and connected to a structural dynamic state-space model within the MATLAB/SIMULINK environment. The aeroelastic ROM is extended to treat structural nonlinearity due to the control-surface freeplay by generating an additional feedback line of generalized residual forces in the SIMULINK model. To reduce the problem size and the computation time, a fictitious-mass modal approach is used, which can afford the possible local change of stiffness. The present method is first validated for its capability to simulate aeroelastic responses of a regional-jet horizontal-tail (HT) model without freeplay, and shows good prediction of flutter characteristics measured in the transonic flutter wind-tunnel testing. The present method is then applied to the HT model with the elevator freeplay gaps. Although there are some differences between the computed and measured limit-cycle-oscillation ranges, the present method well predicts the nonlinear limit-cycle-oscillation behavior in its amplitude and frequency measured in the test, and shows promise toward accurate prediction of control-surface LCO due to its freeplay gap.

This paper discusses aerodynamic optimization of the high-wing configuration to explore fuselage-wing shapes for the high-wing configurations of near-future aircraft, in which it will be possible to install fuel-efficient, ultrahigh-bypass ratio engines, using computational fluid dynamics simulation and the Kriging surrogate-assisted genetic algorithm. First, optimization of the fuselage upper surface is performed, with exploration of the fairing shape suitable for the high-wing configuration. Second, the aircraft nose shape is also optimized, in addition to the fuselage upper surface, to confirm the possibility of generating higher lift by the fuselage itself. Finally, both the fuselage and the wing shape are optimized to improve the lift-to-drag ratio by alleviating the shock wave over the wing, while sustaining the high lift generation of the high-wing configuration. The final optimized configuration achieves not only a lift-to-drag ratio comparable to the DLR-F6, but also a CL approximately 1.5 times higher than the DLR-F6. These results indicate the possibility of producing high-wing aircraft that not only employ fuel-efficient ultrahigh-bypass ratio engines, but also have much better aerodynamic performance than low-wing configurations.

This paper proposes a new approach for the study of complex turbulent flows of aeronautics that integrates experimental fluid dynamics (EFD), employing methods such as wind tunnel experiments, and computational fluid dynamics (CFD) by using a data assimilation technique. The approach aims at representing complex turbulent flows more properly than conventional EFD and CFD approaches by estimating the proper angle of attack, the proper Mach number, and the proper turbulent viscosity, which are the three uncertainty factors in EFD and CFD. To this end, the ensemble transform Kalman filter (ETKF), a sequential advanced data assimilation method, is employed for the estimation and applied to transonic flows around the RAE 2822 airfoil (two-dimensional flow) and transonic flows around the ONERA M6 wing (three-dimensional flow). The results computed using the angles of attack, Mach numbers, and turbulent viscosities estimated by the ETKF diminish the discrepancies between the results of standard computations and experiments. These findings show the effectiveness of this approach, which combines EFD and CFD using data assimilation to represent complex turbulent flows. (C) 2014 Elsevier Inc. Allrightsreserved.

© 2015 AIAA American Institute of Aeronautics and Astronautics. All rights reserved. In this study, the sensitivity analysis of wake vortex parameters measured by a Doppler lidar is conducted to assess the errors due to the measurement process. To reduce the uncertainty caused by the measurement process as possible, wake vortex parameters along with a surrounding wind field are estimated based on the four-dimensional variational (4D-Var) method assuming a velocity field obtained by the Rosenhead-Burnham-Hallock vortex model. With this approach, vortex parameters are obtained considering the factors such as insufficient resolution in the line-of-sight direction and vortex descent during a laser scan. It is confirmed from numerical experiments that the average circulation is significantly underestimated when it is derived from the lidar measurement, while vortex core radius becomes too large in the lidar measurement. On the other hand, vortex parameters estimated with the help of the 4D-Var method are improved by compensating the errors due to the measurement process.

© 2015, American Institute of Aeronautics and Astronautics Inc. All rights reserved. An integrated method consisting of a proper orthogonal decomposition (POD)-based reduced-order model (ROM) and a particle filter (PF) is proposed for real-time prediction of low-level turbulence. The proposed method is validated using identical twin experiments of the low-level turbulence case at Shonai airport in Japan. The results of the experiment show ROM with the estimated temporal coefficients by PF reproduces the wind velocity fluctuations similar to those produced by large eddy simulation with much lower computational cost. In addition, the time evolution of the wind field obtained by ROM-PF show that turbulent spatial structure is expressed accurately in comparison with the reference flow field. It is confirmed that the root mean square error of head and crosswind velocities decreased over time by repeating data assimilation.

In this study, sonic boom propagation is simulated by solving the Khokhlov-Zabolotskaya-Kuznetsov (KZK) equation. The KZK equation accounts for the effects of diffraction, thermoviscous absorption, relaxation, geometrical spreading, inhomogeneity of the medium, axial convection, transverse convection, and nonlinearity. First, the wave propagation prediction based on the KZK equation without diffraction, propagation direction, and transverse terms is compared with that based on the augmented Burgers equation. The result shows good agreement, which proves the validity of the present KZK equation code. Second, the diffraction effect considered in the KZK equation is confirmed by the propagation analyses of sinusoidal planer waves. This result indicates that the proposed tool can simulate actual phenomenon more realistically. Third, propagation of the planer N-wave through turbulence fields is analyzed by the proposed tool. It reveals the capability of the proposed tool to analyze sonic boom propagation considering turbulence effects. Finally, sonic boom propagation in real atmospheric condition is simulated by the proposed tool, and the pressure wave through the turbulence field is estimated.

IEEE CEC 2015 was successfully held during May 24-28, 2015, at Sendai International Center, Sendai, Japan. Technical program consisted of 20 tutorials, 3 workshops, 7 competitions, 68 parallel sessions and 56 special sessions for 403 papers, along with a single poster session for 52 papers. The first joint reception for Women In Computational Intelligence, Young Professionals and Students was held on May 26, where all plenary speakers and TC chairs were invited to this event.

The effects of the number of design variables on the optimization performances in Kriging-model-based many-objective optimizations, which use expected hypervolume improvement (EHVI), expected improvement (EI), and estimation (EST) as the criteria for updating the Kriging model, are investigated based on four independent performance metrics in this paper. Numerical experiments are conducted in 3 to 15-objective DTLZ1 and DTLZ7 problems. The results indicate that the advantages of EHVI over EI and EST are more obvious when the number of design variables increases, and EHVI is more suitable for the problems with a large number of design variables. In addition, the comparison results show that, EHVI obtains faster IGD reduction than EI and EST in most test problems. The advantage of EHVI over EI and EST is mainly shown on the convergence performance. The spread performance is better in both EHVI and EI considering estimation errors than EST without considering estimation errors. However, the uniformity of EHVI is weak, especially for the problems with a large number of objectives.

The Kriging surrogate model, which is frequently employed to apply evolutionary computation to real-world problems, with coordinate transformation of design space is proposed to improve the approximation accuracy of objective functions with correlated design variables. Eigenvalue decomposition is used to extract significant trends in the objective function from its gradients and identify suitable coordinates. Comparing with the ordinary Kriging model, the proposed method shows higher accuracy in the approximation of two-dimensional test functions and reduces the computational cost to achieve the global optimization. In the application to an airfoil design problem with spline curves as correlated design variables, the proposed method achieves better performances not only in the approximation accuracy but also the ability to explore the optimal solution.

A new genetic algorithm to search for Pareto-optimal solutions in multi-objective problems with constraints is proposed. This algorithm employs the parallel evaluation strategy in which feasible and infeasible solutions are preserved in separate populations. Feasible solutions are ranked in accordance with the ordinary non-dominated ranking method. On the other hands, infeasible solutions are ranked based on their objective functions and total constraint violation. The total constraint violation is treated as the (M+1)-th evaluation function in addition to M original objective functions used for ranking infeasible solutions. This non-dominated ranking considering both objective functions and total constraint violation is expected to remove infeasible solutions with large constraint violations and preserve useful solutions. Through the present numerical tests, the proposed algorithm without tunable parameters outperforms the existing genetic algorithms considering either objective functions or constraint violations in multi-objective problems with active constraints. Additionally, the proposed algorithm shows better performance than the genetic algorithm using the penalty approach considering the sum of objective functions and total constraint violation. The improvement of Pareto-optimal solution search capability is accomplished by preserving infeasible solutions near the true Pareto-optimal front restricted by active constraints.

The many-objective optimization performance of the Kriging-surrogate-based evolutionary algorithm (EA), which maximizes expected hypervolume improvement (EHVI) for updating the Kriging model, is investigated and compared with those using expected improvement (EI) and estimation (EST) updating criteria in this paper. Numerical experiments are conducted in 3- to 15-objective DTLZ1-7 problems. In the experiments, an exact hypervolume calculating algorithm is used for the problems with less than six objectives. On the other hand, an approximate hypervolume calculating algorithm based on Monte Carlo sampling is adopted for the problems with more objectives. The results indicate that, in the nonconstrained case, EHVI is a highly competitive updating criterion for the Kriging model and EA based many-objective optimization, especially when the test problem is complex and the number of objectives or design variables is large.

In this paper, we describe a filtering algorithm for removing the error of wind velocity arises in airborne Doppler lidar measurements. The algorithm is based on the Kalman filter with a simplified Kalman gain, which assumes zero variance for correct wind velocity and infinite variance for incorrect wind velocity. The algorithm is applied to 17,487 seconds of airborne Doppler lidar measurements, where a sequence of measurements along the lidar's measurement range is obtained every one second. The reduction of incorrect wind velocity is evaluated at the distance where correct wind velocity exists at least 20-30% of the time out of all measurements. The average standard deviation of filtered wind velocity at the above mentioned distance results in 17.2% of the original value, which is similar magnitude to correct measurements.

A rectangular open cavity with upstream dual injectors at a freestream Mach number of 1.9 was investigated experimentally. To evaluate the effect of the distance between the jets, the flow characteristics were investigated using the high-speed schlieren photography, particle image velocimetry, and surface oil flow techniques. The dual jet distances of 18 and 54 mm were used. Unstable flow occurs over the cavity in all cases and is not improved by changing the distance between the dual jets. Although the distance between the dual jets does not influence the flow stability, the flow field varies decidedly depending on the dual jets distance. The enhancement of air mixing depends on the distance between the jets. A long dual jets distance was found to yield better mixing characteristics within the cavity than a short one. When the jets are further apart, the mainstream between two counter-rotating vortex pairs behind the jets flows strongly into the cavity because of the increased blow-down occurring between the vortex pairs. Additionally, a counterflow with a low velocity magnitude occurs behind the jets. Hence, mixing is enhanced within the cavity by effects of the opposed flow. When the jet pairs are closer to each other, the counter-rotating vortex pairs are in contact; as a result, the blow-down effect does not occur between them. The flow drawn into the cavity from the mainstream is supplied from the sides of the test section into the cavity. (C) 2014 Elsevier Inc. All rights reserved.

Multi-objective optimization tools have been applied in various academic and industry fields. It is often difficult to optimize all the objectives since they often cause trade-offs. It is also difficult to figure what kinds of trade-offs actually cause. We think visualization of multi-objective optimization results assists users to intuitively understand the distributions of their solutions. This paper proposes a visualization of explanatory variable and objective function spaces in the separate views, so that users can easily understand their relevancy. Also, our tool features the linkage mechanism between the two views. When a user selects certain ranges of the values by a mouse click operation, the tool highlights all the corresponding individuals. This mechanism is useful for users to narrow down the results. We expect the tool assists the understanding of the behavior of the optimization processes and improvement of the future processes.

This paper presents electrostriction from the phenomenological perspective, and gives details on two mechanical effects arising from laser-matter interaction. Electrostriction is the tendency of materials to compress in the presence of a varying electric field. In this paper, the investigated materials are polar and nonpolar dielectric liquids. It is stressed that the dominant factor is the time evolution of the laser pulse, which causes tensile stresses and acoustic waves. The study is supported by experimental realization of electrostriction, which can be detected only at favorable conditions (observed in water, but not in castor oil). This study will shed light in developing measurement techniques (e.g., laser-induced grating spectroscopy) and in explaining the onset of cavities and laser-induced liquid breakdown. (C) 2014 The Japan Society of Applied Physics

The purpose of the research is to clarify the generation mechanism of precursor electrons ahead of a hypersonic shock wave in argon. The triple-probe measurements and radiation measurements are carried out using a hypersonic shock tube. Theoretical analysis is also carried out using a one-dimensional photoionization model. The triple-probe measurements show that the electron temperature is almost uniform ahead of the shock wave and ranges from 5000 to 11000 K. In contrast to the electron temperature, the electron density increases exponentially as approaching the shock wave and lies between 10¹⁵ and 10¹⁷ m^-3. These results show that considerable electrons exist in the region ahead of the shock wave. From the radiation measurements, argon ions are found to emit radiation in the region ahead of the shock wave, showing that argon ions are generated from argon atoms by photoionization. The behavior of electron temperature and density is well accounted for by the one-dimensional photoionization model with an assumption of photoionization as the primary ionization process ahead of the shock wave. In conclusion, precursor electrons are mainly generated by photoionization, making use of the radiation energy behind the shock wave.

Mixing between the injected fuel and high speed free stream air is challenging at supersonic speeds. Placing cavities downstream of injection holes or slots addresses the problem of flame holding and stabilisation, however there are still open questions related to mixing enhancement, uniformity and efficiency. The present study examines experimentally the flow field interactions due to a transverse jet - cavity combination with shock impingement at supersonic speeds using PIV, Schlieren photography, and oil flow surface visualisation. The oblique shock lifts the shear layer over the cavity and combined with the instabilities generated by the transverse jet injection creates a highly complicated flowfield with numerous vortical structures. The interaction between the oblique shock and the jet leads to a relatively uniform velocity distribution within the cavity. The lifting of the shear layer is also believed to reduce the drag created by the cavity.

© 2015, American Institute of Aeronautics and Astronautics Inc. All rights reserved. The interaction of aircraft exhaust jet and wake vortex is numerically studied. Following the preceding numerical study of wake vortex evolution from roll-up until vortex decay, we aim to simulate the jet-vortex interaction from the early phase until the complete roll-up. For numerical simulation we employ an unstructured mesh Reynolds-averaged Navier-Stokes solver for a flow around an aircraft model, while a Cartesian mesh large-eddy simulation solver implemented on the framework of the Building-Cube Method (BCM) is used for the mid-to far-field wake. Here we investigate a coupling approach of the above solvers to simulate the roll-up of exhaust jet and aircraft wake. The results show that a multi-level Cartesian mesh of the BCM is effective to capture the evolution of aircraft wake including exhaust jet.

The interaction of aircraft exhaust jet and wake vortex is numerically studied. Following the preceding numerical study of wake vortex evolution from roll-up until vortex decay, we aim to simulate the jet-vortex interaction from the early phase until the complete roll-up. For numerical simulation we employ an unstructured mesh Reynolds-averaged Navier-Stokes solver for a flow around an aircraft model, while a Cartesian mesh large-eddy simulation solver implemented on the framework of the Building-Cube Method (BCM) is used for the mid- to far-field wake. Here we investigate a coupling approach of the above solvers to simulate the roll-up of exhaust jet and aircraft wake. The results show that a multi-level Cartesian mesh of the BCM is effective to capture the evolution of aircraft wake including exhaust jet.

© 2014, American Institute of Aeronautics and Astronautics Inc. All rights reserved. The noise shielding characteristics of the Over-the-Wing Nacelle (OWN) configuration are numerically investigated in the experimental setup. Linearized Euler equations are solved on the block-structured Cartesian mesh method named Building-Cube Method (BCM) to resolve the noise accurately. First, the noise propagation from the isolated nacelle is computed and the noise characteristics are compared with experiment to confirm the characteristics of noise source. Next, the noise shielding metrics in three nacelle positions are computed and the Overall Sound Pressure Levels (OASPLs) of noise shielding metrics are compared with experiment. Computed OASPL shows good agreement with experiment. OASPLs increase as the nacelle position moves backward. In the comparison of noise shielding characteristics of each frequency component, the configuration of forward nacelle position results in lower noise shielding metric in higher frequency. On the other hand, the configurations of backward nacelle position show higher noise shielding metric in higher frequency. Finally, the nondimensional pressure distributions on model surface are compared. It is verified noise from nacelle is shielded by the wing-fuselage configuration. Furthermore, noise distribution of fuselage surface moves backward and forward as nacelle moves.

The noise shielding characteristics of the Over-the-Wing Nacelle (OWN) configuration are numerically investigated in the experimental setup. Linearized Euler equations are solved on the block-structured Cartesian mesh method named Building-Cube Method (BCM) to resolve the noise accurately. First, the noise propagation from the isolated nacelle is computed and the noise characteristics are compared with experiment to confirm the characteristics of noise source. Next, the noise shielding metrics in three nacelle positions are computed and the Overall Sound Pressure Levels (OASPLs) of noise shielding metrics are compared with experiment. Computed OASPL shows good agreement with experiment. OASPLs increase as the nacelle position moves backward. In the comparison of noise shielding characteristics of each frequency component, the configuration of forward nacelle position results in lower noise shielding metric in higher frequency. On the other hand, the configurations of backward nacelle position show higher noise shielding metric in higher frequency. Finally, the nondimensional pressure distributions on model surface are compared. It is verified noise from nacelle is shielded by the wing-fuselage configuration. Furthermore, noise distribution of fuselage surface moves backward and forward as nacelle moves.

The incorporation of cavities within supersonic combustion chambers is an effective means of slowing down the flow for fuel injection and consequent stable combustion. Understanding the flow physics associated with such flows, especially with the injection of a gas to replicate fuel injection, are essential for the optimum design of supersonic propulsion mechanisms. An experimental investigation was performed on a rectangular open cavity with upstream injection model in a Mach number of 1.9 using a trisonic indraft wind tunnel. A rectangular open cavity of dimensions L/D = 5, 100 mm in length (L) and 20 mm deep (D), was adopted, and it was embedded into the lower wall of the test section. An air jet with a jet-to-freestream momentum flux ratio of J = 1.2, 2.7 and 5.3 was injected upstream of the cavity. To evaluate the effect on mixing and flow stability the jet position, measured from the front edge of the cavity, was varied between 0.1L and 1L. The flow field was visualized using schlieren photography, particle image velocimetry, and oil flow measurements. It is found that the mixing characteristic within the cavity when the jet is positioned 0.1L is enhanced independent on the J value because the turbulence intensity of the flow velocity within the cavity is strongly influenced by the jet interaction which lifted the flow from the floor of the cavity compared to the other jet positions. However, the flow over the cavity is unstable at all jet positions. The separation shock formed at the front edge of the cavity oscillates significantly for the case where the jet is located at 0.1L because the separation shock location coincides with the compression shock behind the jet. (C) 2013 Elsevier Inc. All rights reserved.

The many-objective optimization performance of using expected hypervolume improvement (EHVI) as the updating criterion of the Kriging surrogate model is investigated, and compared with those of using expected improvement (EI) and estimation (EST) updating criteria in this paper. An exact algorithm to calculate hypervolume is used for the problems with less than six objectives. On the other hand, in order to improve the efficiency of hypervolume calculation, an approximate algorithm to calculate hypervolume based on Monte Carlo sampling is adopted for the problems with more objectives. Numerical experiments are conducted in 3 to 12-objective DTLZ1, DTLZ2, DTLZ3 and DTLZ4 problems. The results show that, in DTLZ3 problem, EHVI always obtains better convergence and diversity performances than EI and EST for any number of objectives. In DTLZ2 and DTLZ4 problems, the advantage of EHVI is shown gradually as the number of objectives increases. The present results suggest that EHVI will be a highly competitive updating criterion for the many-objective optimization with the Kriging model.

Multi-objective optimization tools have been applied in various academic and industry fields. It is often difficult to optimize all the objectives since they often cause trade-offs. It is also difficult to figure what kinds of trade-offs actually cause. We think visualization of multi-objective optimization results assists users to intuitively understand the distributions of their solutions. This paper proposes a visualization of explanatory variable and objective function spaces in the separate views, so that users can easily understand their relevancy. Also, our tool features the linkage mechanism between the two views. When a user selects certain ranges of the values by a mouse click operation, the tool highlights all the corresponding individuals. This mechanism is useful for users to narrow down the results. We expect the tool assists the understanding of the behavior of the optimization processes and improvement of the future processes.

Difficulty of data assimilation arises from a large difference between the sizes of a state vector to be determined, that is, the number of spatiotemporal mesh points of a discretized numerical model and a measurement vector, that is, the amount of measurement data. Flow variables on a large number of mesh points are hardly defined by spatiotemporally limited measurements, which poses an underdetermined problem. In this study we conduct the sensitivity analysis of two- and three-dimensional vortical flow fields within a framework of data assimilation. The impact of measurement strategy, which is evaluated by the sensitivity of the 4D-Var cost function with respect to measurements, is investigated to effectively determine a flow field by limited measurements. The assimilation experiment shows that the error defined by the difference between the reference and assimilated flow fields is reduced by using the sensitivity information to locate the limited number of measurement points. To conduct data assimilation for a long time period, the 4D-Var data assimilation and the sensitivity analysis are repeated with a short assimilation window.

The applicability of high-performance computing (HPC) to vehicle aerodynamics is presented using a Cartesian grid approach of computational fluid dynamics. Methodology that allows the user to avoid a large amount of manual work in preparing geometry is indispensable in HPC simulation whereas conventional methodologies require much manual work. The new frame work allowing a solver to treat 'dirty' computer-aided-design data directly was developed with a modified immersed boundary method. The efficiency of the calculation of the vehicle aerodynamics using HPC is discussed. The validation case of flow with a high Reynolds number around a sphere is presented. The preparation time for the calculation is approximately 10 minutes. The calculation time for flow computation is approximately one-tenth of that of conventional unstructured code. Results of large eddy simulation with a coarse grid differ greatly from experimental results, but there is an improvement in the prediction of the drag coefficient prediction when using 23 billion cells. A vehicle aerodynamics simulation was conducted using dirty computer-aided-design data and approximately 19 billion cells. The preparation for the calculation can be completed within a couple of hours. The calculation time for flow computation is approximately one-fifth of that of conventional unstructured code. Reasonable flow results around a vehicle were observed, and there is an improvement in the prediction of the drag coefficient prediction when using 19 billion cells. The possibility of the proposed methodology being an innovative scheme in computational fluid dynamics is shown. Copyright © 2014 SAE International.

This paper reports establishment of the simultaneous measurement of the flight attitude and flow field visualization for the near-field pressure measurement experiments in a ballistic range. To measure the flight attitude and flow field visualization for a spherical projectile, the simultaneous visualization method, which is combination of the shadowgraph method and the direct photography, was utilized. It is possible to visualize both the flow field around the projectile and the projectile surface. The angles of pitch, roll and yaw can be calculated from the displacement of line markings on the projectile surface. The image of the projectile by the direct photography is distorted because of the aberration of the optical arrangement, the camera lens and the non-uniform thickness of acrylic windows. Image distortion was calibrated by taking images on actual set-up, so that the accuracy of the angles was improved. In addition, the diffuse reflection light from the projectile surface was confirmed not to influence the flow field visualization. This method is useful for the near-field pressure measurement experiments in a ballistic range. The flight attitude affects the near-field pressure value strongly. Since the near-field pressure with a detailed flight attitude is obtained, its influences can be evaluated using the method which has ability of the flight attitude measurement. In addition, a shock interaction around a projectile can be captured.

A data compression method based on image encoding techniques is presented for a flow simulation data set. An input flow field data set is converted into the octree structure by discrete wavelet transform, and then quantized finely or coarsely depending on its importance in the flow field. Embedded zerotree wavelet encoding as the image encoding technique and entropy encoding reduce the data size by making use of the octree structure created previously. The present compression method is incorporated in a block-structured Cartesian mesh method called Building-Cube method. The Building-Cube method gives not only good performance in the flow simulation but also consistency with the embedded zerotree wavelet encoding in the data compression. Three compression cases for incompressible and compressible flow simulations, including a large-scale simulation with O(10 (8)) mesh points, demonstrate that the present compression method gives both high compression ratios and good qualities of compressed data. Copyright (c) 2013 John Wiley & Sons, Ltd.

Measurement integrated simulation which makes up for uncertainty of measurement techniques in experiment and makes up for uncertainty of modeling in simulation is expected as an innovative technology in future of aerospace science. The measurement integrated simulation has a deep relationship with "data assimilation" that has been developed in the field of earth science. In this paper, value of parameter which is one of uncertainty factors in simulation is investigated by using the ensemble Kalman filter that is an advanced technique of data assimilation. In this research, the values of parameters in the Spalart-Allmaras turbulence model are investigated. And, this paper shows that the values could be determined objectively or statistically by data assimilation techniques and not by subjective "calibration" as performed by the model proposers themselves. (C) 2012 Elsevier Ltd. All rights reserved.

The ejection of material during laser ablation gives rise to the development of discontinuities in the ambient gas. Several of these discontinuities are observed and characterized, including externally and internally propagating shock waves, contact surface, and the ionization front. Qualitative experimental observations and analysis of these discontinuities are presented. Results from shadowgraphy enabled determination of an irradiance threshold between two different ablation mechanisms, and determination of several stages of plasma plume evolution. Consideration of the refractive index as a dynamic sum of the contributions from gas and electrons led to separate identification of ionization front from the contact surface. Furthermore, ionization front was observed to lead the shock wave at the earlier stage of the ablation.

This paper presents a comparison of the criteria for updating the Kriging surrogate models in multi-objective optimization: expected improvement (EI), expected hypervolume improvement (EHVI), estimation (EST), and those in combination (EHVI+EST). EI has been conventionally used as the criterion considering the stochastic improvement of each objective function value individually, while EHVI has recently been proposed as the criterion considering the stochastic improvement of the front of nondominated solutions in multi-objective optimization. EST is the value of each objective function estimated nonstochastically by the Kriging model without considering its uncertainties. Numerical experiments were implemented in the welded beam design problem, and empirically showed that, in an unconstrained case, EHVI maintains a balance between accuracy, spread, and uniformity in nondominated solutions for Kriging-model-based multiobjective optimization. In addition, the present experiments suggested future investigation into techniques for handling constraints with uncertainties to enhance the capability of EHVI in constrained cases.

The wind environment at an airport is affected by terrain features. In Japan, Shonai Airport is known to frequently have wind shear over the runway due to the turbulence induced by the neighboring hills in winter. In this study, large eddy simulation (LES) is performed to investigate the turbulence around Shonai Airport. The initial and boundary conditions are given according to the weather prediction data by the Japan Meteorological Agency non-hydrostatic model (JMANHM). These data are downscaled and transferred to LES domains, which consider actual terrain features as the boundary conditions on the ground, using the two-way nesting method. The present simulations indicate that terrain features may have a significant influence on the turbulence appearing in flight paths; i.e., aircraft safety may depend on wind direction. In addition, it is shown that the present simulation method can predict the turbulence induced by terrain features based on good agreement of results with the turbulence actually observed using a Doppler radar.

Background oriented schlieren (BOS) visualization technique is examined by means of optical geometry. Two most important results are the calculation of the sensitivity and spatial resolution of a BOS system, which allows for the determination of the experiment design space. A set of relations that characterize the performance of a BOS measurement is given, with emphasis on the design of background pattern and spatial placement of optical components. Graphical Abstract: [Figure not available: see fulltext.] © 2013 The Visualization Society of Japan.

A badminton shuttlecock flies in a high-drag, and thus, the sport has been a subject of research from the point of view of aerodynamics. A badminton shuttlecock generates significant aerodynamic drag and has a complex flight trajectory. It also has the smallest ballistic coefficient and exhibits the largest in-flight deceleration of any airborne sporting projectile. The ballistic coefficient of a projectile is a measure of its ability to overcome air resistance in flight and is inversely proportional to deceleration. The primary objectives of this study were to measure the aerodynamic properties of feather shuttlecocks under a range of the wind speed (10-60 m/s) and pitch angle (0°-25°). In particular, measurements of aerodynamic forces were performed at high Reynolds numbers (more than Re = 210,000), and the effect of shuttlecock deformation on aerodynamic properties was also investigated, because it is presumed that the flight dynamics is affected by the deformation of the shuttlecock skirt. A shuttlecock skirt is composed of an array of diverging stems, the ends of which are at the convergent end of the skirt, joined together in an end ring. The shuttlecock rotates about its major axis in actual flight, and thus, the experiments were performed on shuttlecocks with and without rotation (spin). Furthermore, the effect of the flow passing through the gaps between the slots (stiffeners) located at the leg portion of the shuttlecock skirt on aerodynamic characteristics is demonstrated by means of a shuttlecock model without gaps, which was completely covered with cellophane tape. The free rotation rate of a shuttlecock increased with an increase in the Reynolds number, and the drag coefficient gradually decreased above Re = 86,000 for a non-rotating shuttlecock. The reduction of drag can be explained by the deformation of the skirt observed in wind tunnel experiments at high speed. In this study, for a rotating shuttlecock, a reduction of drag was not observed over a whole range of Reynolds numbers, because deformation of the skirt for a rotating shuttlecock becomes smaller than that for a non-rotating shuttlecock. However, there was no significant difference in drag coefficient between rotating and non-rotating shuttlecocks, in contrast to the difference in drag coefficient between shuttlecocks with and without gaps. The drag coefficient for a shuttlecock without gaps was significantly smaller than that for a standard shuttlecock (with gaps). For a standard shuttlecock, the air flowed through the gaps into the shuttlecock skirt, and this flow was related to high aerodynamic drag. © 2013 International Sports Engineering Association.

In this paper, the coupling computation of Computational Fluid Dynamics (CFD) and Computational AeroAcoustics (CAA) is conducted. Three operating conditions of the JT15D nacelle configuration are analyzed. Fan noise from the engine inlet is simulated in each condition. Computation is based on the Cartesian mesh method called Building-Cube Method (BCM). Firstly, the background flow is computed by a compressible Euler solver. And then, the noise propagation from the nacelle is computed by the Linearized Euler Equations (LEE) code. Finally, the far-field noise directivity is estimated by the Ffowcs-Williams and Hawkings (FW-H) method. The low speed flow around the JT15D nacelle was successfully simulated by using Simple Low-dissipative AUSM, which can deal with all speed flows. Sound wave propagation in the radial direction was pronounced at the lower frequency condition. In three conditions, the FW-H estimation of far-field Sound Pressure Level showed the good agreement with the experimental measurements qualitatively.

Quantitative measurements of fluid flow can be achieved by flow visualization techniques, and this paper reviews and outlines background oriented schlieren, with emphasis on its performance: measurement sensitivity and uncertainty. Since the technique depends on cross-correlation, an assessment of image evaluation is also conducted. Background oriented schlieren is applied to two flows: shock reflection from a wedge in a shock tube, and natural cooling byconvection. It is estimated that the technique can be applied to ballistic facilities. © Owned by the authors, published by EDP Sciences, 2013.

Multi-objective optimization and data mining of vortex generators (VGs) on a transonic infinite-wing was performed using computational fluid dynamics (CFD), surrogate models, and a multi-objective genetic algorithm (MOGA). VGs arrangements were defined by five design variables: height, length, incidence angle, spacing, and chord location. The objective functions which should be maximized were three: lift-drag ratio at low angle of attack, lift coefficient at high angle of attack, and chordwise separation location at high angle of attack. In order to evaluate these objective functions of each individual in MOGA, the response surface methodology with Kriging model and the modified version of it was employed because CFD analysis of the wing with VG requires a large computational time. Two types of data mining method: analysis of variance (ANOVA) and self-organizing map (SOM), were applied to the result of the optimization. It was revealed by ANOVA that the ratio of spacing to height and the incidence angle had significant influences to maximizing each objective function. By using SOM, VG designs were split into four types which have different aerodynamic characteristics respectively. The appropriate values of parameters were identified by SOM. © 2013 IEEE.

An multi-objective optimization (MOO) and knowledge extraction were conducted for a structural design of a regional jet horizontal tail using Multi-Objective Design Exploration (MODE). MODE reveals the structure of the design space from the trade-off information and visualizes it as a panorama for a Decision Maker. The present form of MODE consists of Kriging model, Adaptive Range Multi-Objective Genetic Algorithm and Self-Organizing Map (SOM). Combination between the stringer-pitch and the rib-pitch was optimized based on detailed buckling evaluation and MSC/NASTRAN static analysis using realistic aircraft structural model. The resulting Kriging model provided several solutions with improvements, both in the structural weight and the number of structural components, compared with the baseline design. Furthermore, SOM divided the design space into clusters with specific design features. The acquired design knowledge from the present application has been utilized for the horizontal tail design of Mitsubishi Regional Jet (MRJ). © 2013 Springer-Verlag.

Multi-objective genetic algorithms were used to aerodynamically optimize a high-altitude long-endurance aircraft in order to improve its geometry and increase its performance. A conceptual design was created to satisfy a high-altitude long-endurance flight mission. However, to tackle the common problem high-altitude aircrafts have of generating a high CL with low speeds, the airfoil selected for the conceptual design was optimized to maximize its lift capacity at the cruise altitude. The optimization consisted on using adaptive range multi-objective genetic algorithms with 9 design variables represented as the PARSEC definition for an airfoil, and Xfoil was used as the evaluating tool for the generated airfoils. The optimization produced an airfoil that maximizes the required aerodynamic performance for high-altitude flight. This aerodynamic performance improvement eventually leads to the reduction of necessary wing area, which leads to weight and cost reductions which are significant for high-altitude unmanned aircrafts.

This paper presents a comparison of the criteria for updating the Kriging surrogate models in surrogate-based non-constrained many-objective optimization: expected improvement (EI), expected hypervolume improvement (EHVI), and estimate (EST). EI has been conventionally used as the criterion considering the stochastic improvement of each objective function value individually, while EHVI has been proposed as the criterion considering the stochastic improvement of the front of non-dominated solutions in multi-objective optimization. EST is the value of each objective function estimated non-stochastically by the Kriging model without considering its uncertainties. Numerical tests were conducted in the DTLZ test function problems (up to 8 objectives). It empirically showed that, in the DTLZ1 problem, EHVI has greater advantage of convergence and diversity to the true Pareto-optimal over EST and EI as the number of objective functions increases. The present results also suggested the expectation that the Kriging-surrogate-based optimization using EHVI may overcome the direct optimization without using the Kriging models when the number of objective functions becomes more than 10. In the DTLZ2 problem, however, EHVI achieved slower convergence to the true Paretooptimal front than EST and EI. It is due to the complexity of objective function space and the selection of additional sample points.

Multiple-objective optimization problems have multiple objective functions to be optimized simultaneously. It is often difficult to optimize all the objectives since they often cause trade-offs. We produce a visualization tool which displays both explanatory variables and objective functions for multi-objective optimization.

The proper orthogonal decomposition (POD) and the particle filter (PF) are applied to estimate the low-level turbulence around an airport. POD is introduced to represent a wind velocity field by a reduced-order model. PF is employed to estimate coefficients of POD by using observed velocity components from the broadband radar located at an airport. In this study, a numerical experiment was conducted to investigate the effectiveness of the proposed method by comparing the estimated wind velocity field with the original wind velocity field obtained from large eddy simulation (LES). From the comparison results, the variation of turbulence intensity obtained from the present approach was found similar to the original LES result. The main advantage of the present approach is much lower computational cost compared to LES.

This study describes the start/unstart characteristics of a finite and rectangular supersonic biplane wing. Two wing models were tested in wind tunnels with aspect ratios of 0.75 (model A) and 2.5 (model B). The models were composed of a Busemann biplane section. The tests were carried out using supersonic and transonic wind tunnels over a Mach number range of 0.3≤M∞≤2.3 with angles of attack of 0°, 2°, and 4°. The Schlieren system was used to observe the flow characteristics around the models. The experimental results showed that these models had start/unstart characteristics that differed from those of the Busemann biplane (two dimensional) owing to three-dimensional effects. Models A and B started at lower Mach numbers than the Busemann biplane. The characteristics also varied with aspect ratio: model A (1.3<M∞<1.5) started at a lower Mach number than model B (1.6<M∞<1.8) owing to the lower aspect ratio. Model B was located in the double solution domain for the start/unstart characteristics at M∞=1.7, and model B was in either the start or unstart state at M∞=1.7. Once the state was determined, either state was stable. © 2013 Hiroshi Yamashita et al.

Low-level turbulence around airport poses a potential hazard to aircraft operations such as take-off and landing. An advisory system for such atmospheric turbulence is of great help in safe airport operations. Aiming for a part of the advisory system, we investigate terrain-induced low-level turbulence at Shonai airport by an integrated approach of meteorological model prediction and high resolution large eddy simulation. The predicted results by the integrated system are validated against Doppler radar observations focusing on strong-wind events. The results revealed that low-level turbulence is triggered by a hill located on the north side of the runway. Furthermore, the results showed the extent of the low-level turbulence in the wake of the hill, that is, the turbulent area increases vertically as well as horizontally in the downwind direction. The magnitude and the direction of winds vary in short period of time due to the turbulent nature of the terrain-induced turbulence. A simplified model is employed to estimate the vertical load of aircraft along a flight path from the predicted turbulent wind field. These findings would be used to predict the hazardous area to aircraft operations, which is an important part of the advisory system.

The purpose of the research is to develop the method of designing and manufacturing a sensitivity-adjustable three component force balance utilized in a wind tunnel experiment. This paper describes the characteristics of the force balance and its application to wind tunnel testing. In this study, an oval-shape force balance is newly introduced and the characteristics are investigated by FEM analysis and force balance calibration. Also, the validity of the force balance is investigated in a wind tunnel experiment. As a result, the sensitivity of the oval-shape force balance can be adjusted by changing the aspect ratio and aerodynamic forces can be measured in a wind tunnel experiment using it. In conclusion, we can develop the sensitivity-adjustable force balance by applying oval shape and the force balance can be used in a wind tunnel experiment. It is recommended that the strain gauge position and its paste procedure be still improved.

The aerodynamic characteristics and the effects of tip plates of a tapered supersonic biplane wing during the starting process were investigated by experimental and computational fluid dynamics (EFD/CFD). Three types of wing models were used: one with a tip plate covering only the aft-half of the wing tip (type-A); one with a tip plate covering the entire wing tip (type-B); and one without a tip plate (type-N). Experiments were conducted in a supersonic blowdown wind tunnel 600 &times; 600 mm in cross-section located at the high-speed wind tunnel facility of the Institute of Space and Astronautical Science (ISAS/JAXA). The flow conditions ranged from <i>M</i>_&infin; = 1.5 to 1.9 in increments of 0.1. Pressure-sensitive paint was applied to measure the pressure distributions on the wing surface. CFD simulations were conducted for comparison with the experimental results and for detailed investigation of the effects of the Mach number. The tapered biplane wing without a tip plate is found to start between <i>M</i>_&infin; = 1.8 and 1.9. The difference in starting Mach number between type-N and type-A is small. On the other hand, the starting Mach number of type-B is approximately 0.05 higher than that of type-N.

Quantitative measurements of fluid flow properties can be achieved by background-oriented schlieren (BOS). In this paper, it is shown that this depends on several factors. Image-quality index is used to investigate the influence of the image sensor and the quality of its output. Image evaluation is applied to synthetic images, which are treated with a step function, so that they simulate the sharp density jump. The gradual change of the evaluated vector shift revealed the major dependence on the interrogation window, and revealed less of a dependence on background features. BOS applied to shock-wave reflection from a wedge in a shock tube gave qualitative results, due to large uncertainties. But, the application to cooling by natural convection gave satisfactory results, comparable to thermocouple data and theory. (C) 2012 Optical Society of America

A data exploration method based on POD was investigated to get clues to improve actual-product designs from a small set of rough-design candidates. The current method is based on the Oyama method which was used to acquire design knowledge of Pareto-optimal solutions obtained by evolutionary approach with large set of simulation models. The expanded method adopted (1) 2 steps POD analysis, and (2) voxel representation for different topology models. These approaches enabled to evaluate rough-design candidates which had non-parametric structures and those were difficult to handle in the original method. Especially, for the case of handling rough-design candidates, there was no explicit rule for model number ordering at the "first" POD analysis. We found that reordering the model numbers based on the first POD result was important, and rich information was obtained from the expansion coefficient vectors that made it possible to classify the feature regions and to detect feature regions with complicated flow structures at the "second" POD analysis. The effectiveness of the expanded method was verified in a cooling design problem for an optical disc drive. We found that the expanded POD-based data exploration method is a promising method for detecting key factors for actual product designs.

A hybrid method between the Kriging model and the radial basis function (RBF) networks is proposed for robust construction of a response surface of an unknown function. In the hybrid method, RBF approximates the macro trend of the function and the Kriging model estimates the micro trend. Hybrid methods using two types of model selection criteria (MSC), i.e., leave-one-out cross-validation and generalized cross-validation for RBF were applied to three one-dimensional test problems. The results were compared with those of the ordinary Kriging (OK) model and the universal Kriging model. The accuracy of each response surface was compared by function shape and root mean square error. The proposed hybrid models were more accurate than the OK model for highly nonlinear functions because they can capture the macro trend of the function properly by RBF, while the OK model cannot. In addition, the hybrid models can find the global optimum with few sample points using the Kriging model approximation errors.

A weighted compact nonlinear scheme (WCNS) is applied to numerical simulations of compressible multicomponent flows, and four different implementations (fully or quasi-conservative forms and conservative or primitive variables interpolations) are examined in order to investigate numerical oscillation generated in each implementation. The results show that the different types of numerical oscillation in pressure field are generated when fully conservative form or interpolation of conservative variables is selected, while quasi-conservative form generally has poor mass conservation property. The WCNS implementation with quasi-conservative form and interpolation of primitive variables can suppress these oscillations similar to previous finite volume WENO scheme, despite the present scheme is finite difference formulation and computationally cheaper for multi-dimensional problems. Series of analysis conducted in this study show that the numerical oscillation due to fully conservative form is generated only in initial flow fields, while the numerical oscillation due to interpolation of conservative variables exists during the computations, which leads to significant spurious numerical oscillations near interfaces of different component of fluids. The error due to fully conservative form can be greatly reduced by smoothing interface, while the numerical oscillation due to interpolation of conservative variables cannot be significantly reduced. The primitive variable interpolation is, therefore, considered to be better choice for compressible multicomponent flows in the framework of WCNS. Meanwhile better choice of fully or quasi-conservative form depends on a situation because the error due to fully conservative form can be suppressed by smoothed interface and because quasi-conservative form eliminates all the numerical oscillation but has poor mass conservation. (C) 2012 Elsevier Inc. All rights reserved.

The aerodynamic characteristics and the effects of winglets of a tapered supersonic biplane during the starting process have been investigated through Experimental and Computational Fluid Dynamics (EFD/CFD). Three types of the tapered biplane model were used: without winglet (type-N); with the winglet which covers only the aft-half of the wing tip (type-A); with the winglet which covers the entire wing tip (type-B). Experiment was conducted in the supersonic blow-down wind tunnel with 600mm &times; 600mm cross section located at the High-speed Wind Tunnel Facility of Institute of Space and Astronautical Science, JAXA. The flow conditions covered from <i>M</i>_&infin; = 1.5 to 1.9 with increments of 0.1. Pressure-Sensitive Paint was applied to measure pressure distributions on the surface of supersonic biplane models. CFD simulations were conducted to compare with experiment and to investigate effects of the Mach numbers in detail. The tapered biplane without the winglet was found to start between <i>M</i>_&infin; = 1.8 and 1.9. The difference of the starting Mach numbers between type-N and type-A was small. On the other hand, the starting Mach number of type-B was about 0.05 higher than that of type-N.

Global variation of sonic boom overpressures with the realistic atmospheric gradients was discussed. The atmospheric gradients were estimated by upper-air observational radiosonde data and a simple N-wave was extrapolated through all seasonal atmospheric gradients without winds around the world. Results demonstrated that sonic boom overpressure varies widely with season and geographic position compared to that of the standard atmospheric condition. The results also showed the tendencies of the global variation in overpressure.

This paper compares the criteria for updating the Kriging response surface models in multi-objective optimization: expected improvement (EI), expected hypervolume improvement (EHVI), estimation (EST), and those combination (EHVI+EST). EI has been conventionally used as the criterion considering the stochastic improvement of each objective function value individually, while EHVI has been recently proposed as the criterion considering the stochastic improvement of the front of non-dominated solutions in multi-objective optimization. EST is the value of each objective function, which is estimated non-stochastically by the Kriging model without considering its uncertainties. Numerical experiments were implemented in the welded beam design problem, and empirically showed that, in a non-constrained case, EHVI keeps a balance between accurate and wide search for non-dominated solutions on the Kriging models in multi-objective optimization. In addition, the present experiments suggested future investigation into the techniques for handling uncertain constraints to enhance the capability of EHVI in a constrained case.

The shuttlecock used in the game of badminton has astonishing aerodynamic characteristics. The shuttlecock which has an open conical shape is the smallest ballistic coefficient and exhibits the largest in-flight deceleration of any airborne sporting implement. The ballistic coefficient of a body is a measure of its ability to overcome air resistance in flight and is inversely proportional to the deceleration. To evaluate correctly the forces acting the shuttlecock, it is necessary to know the in-flight properties determined from the aerodynamics. The purpose of this study is to investigate the relationship between aerodynamic properties and flow fields for a shuttlecock with spin rotation at high Reynolds numbers. Particle image velocimetry (PIV) was used to aid in understanding the flow field in the near field of the shuttlecock edge for the wind tunnel test. The effect of shuttlecock deformation on aerodynamic properties was also investigated because it is presumed that aerodynamic forces are affected by the deformation of the shuttlecock skirt. The drag coefficient for the shuttlecock with no gap is significantly smaller than that for the ordinary shuttlecock (with gap). For the ordinary shuttlecock, the air flows through the gap in the shuttlecock skirt, and this flow is related to high aerodynamic drag. That is, the high aerodynamic drag of a badminton shuttlecock is causes by the gap in the shuttlecock skirt. (C) 2012 Published by Elsevier Ltd.

In this study, the computational accuracy of BCM compressible Euler solver was first validated with NACA0012 airfoil. The result of validation showed that BCM was in good agreement with the TAS code. Next, the BCM solver was applied to the aerodynamic design optimization problem of shock free airfoil. The airfoil without a shock wave was successfully designed. Finally, data mining was performed using POD in order to understand the characteristics of configuration for the obtained airfoils by optimization. As a result, it is revealed that the acceleration of flow on upper surface is reduced by the flattened upper surface of airfoil to weaken a shock wave.

Experimental and numerical methods were performed to investigate the influence of engine-cooling exit flow on aerodynamic performance. The engine-cooling channel was constructed at a simplified notchback model with various outlet layouts. The changes in the main aerodynamic coefficients with different cooling exit flow inclination angles and cooling outlet locations were measured and evaluated, and the reasons and mechanisms were analysed and discussed. The results indicate that reducing the inclination angle is an effective method to improve the cruise performance of a passenger vehicle, and wheels and wheel-housings have significant effects on interference between the underflow and the cooling exit flow.

Several data mining techniques for the acoustic sources of computational data of a supersonic jet impinging on an inclined flat plate with its angle of 30, 45 and 60 degrees are applied and the results are discussed. Cluster analysis clearly shows us the classification of three kinds of acoustic waves without complicated analyses based on try-and-error process. The correlation analysis shows that the acoustic wave radiated from around the impingement is generated from the shear layer-shock interaction and has strong relationship with the fluctuation of shear layer. The POD analysis clearly shows the acoustic wave radiation pattern and illustrates that there are two types of acoustic waves from around impinging region for large plate angle case.

This study was performed to determine the relations between the features of wall shear stress and aneurysm rupture. For this purpose, visual data mining was performed in unsteady blood flow simulation data for an aortic aneurysm. The time-series data of wall shear stress given at each grid point were converted to spatial and temporal indices, and the grid points were sorted using a self-organizing map based on the similarity of these indices. Next, the results of cluster analysis were mapped onto the real space of the aortic aneurysm to specify the regions that may lead to aneurysm rupture. With reference to previous reports regarding aneurysm rupture, the visual data mining suggested specific hemodynamic features that cause aneurysm rupture.

This paper describes the design optimization of a sport shoe sole structure by evolutionary computation coupled with eigenmode analysis based on the finite element method. A genetic algorithm assisted by the Kriging response surface model was used for global and efficient optimization. The present study implemented two optimization cases based on equivalent design problem formulations: the single-objective constrained problem and multi-objective non-constrained problem. The optimization described here provided a midsole structure with optimal material properties, in which the elastic modulus should be reduced in rearfoot -&gt; midfoot -&gt; forefoot-centre domains sequentially as the midsole weight becomes lighter. In addition, a comparison of the results between the two cases yielded insight into constraint handling for faster convergence in midsole design optimization by evolutionary computation.

Multi-objective design optimization for a steam turbine stator blade was implemented using three-dimensional large eddy simulation (LES) and a genetic algorithm (GA). The GA used here was assisted by the Kriging response surface model for global and efficient optimization. The aim of the optimization described here was to reduce overall pressure loss and local pressure loss due to end walls simultaneously. The optimization results revealed the blade design candidates that overcame the baseline design in terms of overall loss and local loss, and the trade-off relation between them. In addition, these results provided a specific design concept and corresponding flow mechanism to realize a highly efficient stator blade.

The shock stand-off distance of a spherical model flying with transonic speeds is determined through numerical simulations. The model decelerates due to drag forces caused by the pressure and viscous shear stress at the model surface. Two-dimensional axisymmetric numerical codes with numerical grids fixed to the flying spherical model are used in the simulations. Numerically determined shock stand-off distances are compared with experimental data obtained in a previous study as well as with those obtained in our ballistic-range experiments. The numerical results and the experimental data are found to be in good agreement. In addition, the time-dependent shock stand-off distance of a decelerating model is investigated.

Multi-objective design exploration (MODE) and its application for design rule extraction are presented. MODE reveals the structure of design space from the trade-off information. The self-organizing map (SOM) is incorporated into MODE as a visual data-mining tool for design space. SOM divides the design space into clusters with specific design features. The sufficient conditions for belonging to a cluster of interest are extracted using rough set theory. The resulting MODE was applied to the multidisciplinary wing design problem, which revealed a cluster of good designs, and we extracted the design rules of such designs successfully.

In this research, "Data Assimilation" was applied to a wind tunnel experiment. A square cylinder wind tunnel experiment and the ensemble Kalman filter were used to integrate surface pressures of a square cylinder obtained from computation and experiment. The results indicated that surface pressures of the square cylinder and flow field can be replicated by the present data assimilation method, while they could not be replicated by numerical simulation alone. The results of the ensemble Kalman filter were compared with those of "Measurement-Integrated Simulation," which is another integrated method between computation and experiment for fluids.

The drag reduction of a pickup truck by a rear flap add-on was examined through CFD simulations and wind tunnel experiments. When installed at the rear edge of the roof, the flap increased the cabin back surface pressure coefficient, causing the downwash of the bed flow to be inclined on the tailgate. Thus, the attachment of the bed flow to the tailgate was eliminated; consequently, the drag coefficient was reduced with increasing flap length and downward angle despite the enlarged reverse flow in the wake. However, the drag coefficient did not decrease any further after a specific downward angle was reached because the bed flow increased the drag force at the tailgate and the flap lowered the pressure field above the flap. To maximize the drag reduction effect, the rear downward flap should be designed to have an optimum downward angle.

To examine the possibility of sonic boom mitigation by the supersonic biplane with a fuselage, two kinds of wing-fuselage models for the ballistic range experiment were designed by using CFD simulations. One was the low-boom model with a blunt nose and the other was the high-boom model with a spike nose. From the CFD result, the near-field pressure strength from the low-boom model becomes 70% of that of the high-boom model. The far-field pressure peak of the low-boom model becomes 85% of that of the high-boom model. To verify the near-field pressure signature calculated by CFD, the experimental model using the supersonic biplane with a finite-thickness leading edge was designed and launched with the ballistic range. As a result, the reliability of the CFD result was confirmed.

In order to construct a response surface of an unknown function robustly, a hybrid method between the Kriging model and the radial basis function (RBF) networks is proposed in this paper. In the hybrid method, RBF approximates the macro trend of the function and the Kriging model estimates the micro trend. Then, hybrid methods using two types of model selection criteria (MSC): leave-one-out cross-validation and generalized cross-validation for RBF and the ordinary Kriging (OK) model for comparison are applied to three types of one-dimensional test problems, in which the accuracy of each response surface is compared by shapes and root mean square errors. As a result, the hybrid models are more accurate than the OK model for highly nonlinear functions because the hybrid models can capture the macro trend of the function properly by RBF, but the OK model cannot. However, because the accuracy of the hybrid method turns down significantly when RBF causes overfitting, stable MSC is required. In addition, the hybrid models can find out the global optimum with a few sample points by using the Kriging model's approximation errors effectively.

Flow and acoustic fields of a supersonic jet impinging on an inclined flat plate are investigated by applying data mining techniques. The data of flow and acoustic fields obtained in the previous study are used. The self-organizing map (SOM) and k-means method are applied to this dataset based on the normalized sound pressure level spectra. The results of SOM and k-means method show the clear characterization of the regions based on the frequency characteristics of acoustic waves. Some of clustered regions correspond to the region at which three kinds of aeroacoustics wave are (i) Mach wave from the main jet, (ii) acoustics waves from impinging, and (iii) Mach waves from the supersonic flow downstream of impinging region. The results of SOM and k-means method validate the relationship among these three kind of aeroacoustic waves which is clarified in the previous study, i.e. lower frequency components are stronger for Mach/acoustic waves with the higher index. Copyright © (2011) by the Institute of Noise Control Engineering.

Experimental and numerical studies are performed to evaluate and analyze the influence of the notchback rear diffuser angle on aerodynamic drag and wake structure. The relationship between aerodynamic drag and rear diffuser angle is summarized, and the flow mechanism are analyzed and discussed. A speculation regarding lower trailing vortices is proposed, and is verified in the present research model. Rear diffuser angle is an important factor influencing the wake structure, and optimizing the vehicle rear diffuser at a favorable angle can contribute to reduce the drag force and improve the wake structure.

The aerodynamic characteristics and the effects of tip plates of a tapered supersonic biplane wing during the starting process have been investigated through Experimental and Computational Fluid Dynamics (EFD/CFD). Three types of the wing model were used: without tip plate (type-N); with the tip plate which covers only the aft-half of the wing tip (type-A); with the tip plate which covers the entire wing tip (type-B). Experiment was conducted in the supersonic blowdown wind tunnel with 600 mm × 600 mm cross section located at the High-speed Wind Tunnel Facility of Institute of Space and Astronautical Science (ISAS/JAXA). The flow conditions covered from M∞ = 1.5 to 1.9 with increments of 0.1. Pressure-Sensitive Paint was applied to measure pressure distributions on the surface of the wing. CFD simulations were conducted to compare with experiments and to investigate effects of the Mach numbers in detail. The tapered biplane wing without the tip plate was found to start between M∞ =1.8 and 1.9. The difference of the starting Mach numbers between type-N and type-A was small. On the other hand, the starting Mach number of type-B was about 0.05 higher than that of type-N. Copyright © 2011 by ASME.

The badminton shuttlecock has the smallest ballistic coefficient and exhibits the largest in-flight deceleration of any airborne sporting implement. In the present study, measurements of aerodynamic forces are performed at high Reynolds numbers, and the effect of shuttlecock deformation on aerodynamic properties was also investigated. A shuttlecock skirt has an array of diverging stems, whose ends are at the convergent end of the skirt, joined together in an end-ring. Furthermore, the shuttlecock rotates about shuttlecock's major axis in actual flight, and the experiments are performed by the shuttlecocks with and without rotation (spin). The effect of the flow passing through the gap between slots (stiffeners) located at leg portion of the shuttlecock skirt on aerodynamic characteristics is also demonstrated by using no-gap shuttlecock model, which is completely covered with cellophane tape. The free rotation rate of a shuttlecock increases with increasing Reynolds number, and the drag coefficient gradually decreases over Re = 86000 for the shuttlecock with no rotation. On the other hand, the drag coefficient for the shuttlecock with rotation is almost constant over the whole range of Reynolds number in the present study because the deformation of shuttlecock skirt for the shuttlecock with rotation becomes smaller than that for the shuttlecock with no rotation. However, there is no significant difference in drag coefficient between the shuttlecocks with and without rotation in contrast to the difference in drag coefficient between the shuttlecocks with and without gap. The drag coefficient for the shuttlecock with no gap is significantly smaller than that for the ordinary shuttlecock (with gap). For an ordinary shuttlecock, the air flows through the gap in the shuttlecock skirt, and this flow is related to high aerodynamic drag. (C) 2011 Published by Elsevier Ltd. Selection and peer-review under responsibility of RMIT University

Research of Long-Distance Human-Powered Flight has been performed through the following four items: Aircraft design, Risk management, Pilot performance, Weather prediction. Actual flight took place on August 12, 2009. The flight distance was 20.72km. In this study, results of weather prediction using 1-way downscaling technique are validated by surface observation data and feasibility of weather prediction for Long-Distance Human-Powered Flight is discussed. The weather prediction with 1-km mesh decreases RMSE of wind speed by 0.1--0.2m/s compared with that with 5-km mesh. The weather prediction with 1-km mesh also has a potential to reproduce nonstationary wind. The RMSE gradually increase with time, which is mainly caused by initial and boundary data given from coarse mesh model. To reduce the RMSE, it is desirable to use newest analysis data as possible for initial and boundary data.

A helicopter rotor is optimally designed for aeroacoustic performance improvement. As shown in previous reports, the blade shapes can be designed to minimize high-speed impulsive noise but tend to have excessively high tapers and be swept back. Since an overly short chord length around the blade-tip region may cause structural problems and safety issues in autorotation, an autorotation index has been introduced to keep the tip region from having excessive taper ratios. In addition, the changes in thickness and camber of the airfoils can also be taken into account to better represent realistic rotor shapes. Aeroacoustic analysis is performed using Kirchhoff&apos;s method coupled with computational fluid dynamics analysis, and the optimization is performed using the kriging-model-based genetic algorithm method. Optimization results are presented that show that the designed blades have improved aerodynamic performance and reduced high-speed impulsive noise characteristics. It is found that a more practical blade shape can be obtained by using airfoil transitions and an autorotation index. The results of the analysis of variance and self-organization map indicate that the taper ratios, the swept back, the tip chord length, the protrusion shape, the camber, and the thickness of the root airfoil are the prominent features affecting the aeroacoustic performance of the rotor.

This work attempts to optimize stems that are implanted at the neck of coronary or cerebral aneurysms. to effect a flow diversion A two dimensional version of the stent which is a series of struts and gaps placed at the neck is considered as the first step Optimization is carried out based on the principles of exploration of design space using reductions in velocity and vorticity in the aneurysm dome as the objective functions Latin hypercube sampling first develops 30-60 samples of a strut gap arrangement Flow past an aneurysm with each of these samples is computed using the commercial software FLUENT and the objective functions evaluated This is followed by a Kriging procedure that identifies the nondominated solutions to the system which are the optimized candidates Three different cases of stents with rectangular or circular struts are considered It is found that placing struts in the proximal region of the neck gives the best flow diversion [DOI 10 1115/1 4001861]

This paper investigates the aerodynamic performance of the three-dimensional lifting supersonic biplane and its sonic boom. Although the Busemann biplane is known to cancel the wave drag, it does not produce lift, either. A few decades later, the supersonic biplane airfoils with lift were reported. This paper extends their ideas to the three-dimensional biplane. The aerodynamic performance was revealed by using computational fluid dynamics. The possibility of sonic boom mitigation due to shock wave interaction was demonstrated.

This paper proposes and demonstrates the construction of performance maps, which represent relations between various performance and geometry parameters, with the aid of data mining techniques. Data mining can reveal characteristic patterns in high-dimensional data with performance and geometry parameters. Therefore, the data mining results make it easy to interpret complex features of performance vs. geometry relations, and help engineers to discover new knowledge for engineering design through interpretation. The present demonstration of a centrifugal diffuser demonstrated that the data mining techniques are suitable and applicable to high-dimensional performance map construction, together with actual acquisition of new knowledge for diffuser design that was unknown from conventional quasi-one-dimensional nozzle theory.

Computation Fluid Dynamics (CFD) has become an important tool in optimization and has seen successful in many real world applications, Most important among these is in the optimisation of aerodynamic surfaces which has become Multi-Objective (MO) and Multidisciplinary (MDO) in nature. Most of these have been carried out for a given set of input parameters such as free stream Mach number and angle of attack. One cannot ignore the fact that in aerospace engineering one frequently deals with situations where the design input parameters and flight/flow conditions have some amount of uncertainty attached to them. When the optimisation is carried out for fixed values of design variables and parameters however, one arrives at an optimised solution that results in good performance at design condition but poor drag or lift to drag ratio at slightly off-design conditions. The challenge is still to develop a robust design that accounts for uncertainty in the design in aerospace applications. In this paper this issue is taken up and an attempt is made to prevent the fluctuation of objective performance by using robust design technique or Uncertainty.

Advection database of wake vortices were constructed based on the Doppler lidar measurements at Sendai Airport to understand the relations among measurement, weather factors and behavior of wake vortex. The advection database of wake vortex consists of three measurement factors and fifteen weather factors as variables to determine the wake vortex behavior. However, the extraction of correlations from the high-dimensional data is difficult in general. Therefore, in this paper, data mining was employed as an efficient approach to extract relations among the variables from the advection database of wake vortex. As a first step of the data mining, redundant measurement, weather factors were removed through Spearman rank method, then Self-Organizing Map (SOM) was created to find correlations among the measurement, weather factors and wake vortex behavior. Next, the measurement, weather factors which have larger influences to the behavior of the wake vortex were identified by Analysis of Variance (ANOVA). Finally, rough set theory revealed the specific rules of measurement, weather factors to determine wake vortex behavior. © 2010 by Her Majesty the Queen in Right of Canada.

The purpose of this research is to understand the combined flow of underneath flow and engine-cooling exit flow, as well as the interaction between the rear diffuser angle and the cooling airflow outlet on aerodynamic performance. A 1/8 scale simplified notchback model with an unrestricted internal cooling airflow path was used in this study. And a series of experimental investigations were conducted into aerodynamic performance on different rear diffuser angles in the conditions of closing and opening the cooling airflow path, and on different combinations of the rear diffuser angle and the layout of cooling airflow outlet. The test results show that, after opening the cooling airflow path, the minimum drag can be obtained at a smaller rear diffuser angle and the lift doesn't continuously decrease with the increase of rear diffuser angle, but decreases first and then increases at an angle. To gain a further understanding of the flow mechanism of the combined flow and the interaction, CFD was also employed to obtain the flow details and the distributions of static pressure on models surfaces. The simulation results indicate that, when the upper and the lower vortices on the wake region achieve a relative equilibrium, the drag reaches the minimum value. After opening the cooling airflow path, the separation region of the wake is increased, and the relative equilibrium of upper and lower vortices can be achieved at a smaller rear diffuser angle. In addition, the pressure distribution of the model under surface changes dramatically after opening the airflow path, particularly the variation tendency of pressure distribution on the surface of rear diffuser takes a marked change on different rear diffuser angles. Copyright © 2010 SAE International.

A surrogate model approach has attracted a lot of attention in the field of engineering design, due to its significant reduction in the computational time. The kriging model, one of the most widely used surrogate models, provides efficient tools for global optimization because of its stochastic characteristics. Thus, authors had developed helpful design tools using the kriging model, such as the kriging model based GA (Genetic Algorithm), and had developed ANOVA (ANalysis Of VAriance), and SOM (Self Organization Map) for data mining. However, the developed tools have inconvenience and inefficiency on the whole design works, because of text user interface and separated tools. Thus, the integrated GUI platform MEDOC (Multi-purpose Engineering Design and Optimization Code) has been developed. The main features of MEDOC are supporting multi-OS (Operating System), easy GUI to use, and high calculation efficiency by multi-threading and JNI (Java Native Interface : multi-language coupling). As an example of MD() ( Multidisciplinary Design Optimization), the design of the helicopter rotor blade shape for aero-acoustics performance improvements has been conducted by MEDOC. This application demonstrates the efficiency and the usability of MEDOC.

This paper describes the material design optimization for a sport shoe sole by the evolutionary computation coupled with the eigenmode analysis based on finite element method. In the present evolutionary computation, a genetic algorithm assisted by the Kriging response surface model was used for global and efficient optimization. The present study implemented two optimization cases based on equivalent design problem formulations: single-objective constrained problem and multi-objective non-constrained problem. The present optimization provided optimal sole material properties, in which an elastic modulus should be reduced in rear -&gt; mid -&gt; front-center foot domains sequentially as the sole weight becomes lighter. In addition, a comparison of the results between the two cases showed a suggestion about constraint handling for faster convergence in the sole design optimization by evolutionary computation.

This study was performed to examine the mechanism of flow-induced vibration (FIV) in a hard disk drive (HDD). For this purpose, data mining using self-organizing map (SOM) and Bayesian network was applied to unsteady computational fluid dynamics (CFD) simulation data for a hard disk drive. The present data mining started from the extraction of temporal indices from the time series data of fluid properties given at each grid point. Then, a set of grid points was divided into several clusters based on the similarity of the temporal indices by using SOM, and the clustered data were mapped onto a real space of HDD. Through this process, characteristic phenomena latent in the unsteady flow field were classified and identified. Finally, the relations between temporal indices and FIV were constructed by using Bayesian network. The resulting network structure revealed a possible mechanism ofFIV that originates from a temporal sequence of flow energy dissipation and production. © 2010 IEEE.

This article presents a combined use of multi-objective optimization and quantitative design rule mining methods to improve the aerodynamic efficiency and stability of a centrifugal impeller with a vaned diffuser. A time-averaged but spatially distributed flow is considered at the mixing plane to evaluate the flow uniformity, which affects aerodynamic stability. First, the impeller's shape has been optimized using a multi-objective genetic algorithm to improve efficiency and flow uniformity. It was found that the trade-off among non-dominated solutions can be controlled by the vane-less diffuser's dimensions and the aerodynamic load distribution. A compromise solution has been experimentally shown to improve both objectives. Second, decision tree analysis and rough set theory have been applied to extract design rules for improving each objective. Although the design rules derived from both methods are consistent with regard to the main effects of design variables, some differences are found regarding the interaction effects.

The bed is one of the most important parts of a pickup truck for aerodynamic performance. The flow characteristics of a pickup truck were examined in a series of wind tunnel experiments and numerical simulations with regard to the bed geometry variation, the bed length, and the bed height. The drag coefficient was changed in accordance with the bed geometry variation so that the bed length and bed height had a significant interaction effect. The main factors that affected the drag coefficient were the bed recirculation flow over the bed and the reverse flow in the wake. The drag coefficient increased when the downwash of the bed flow was not recirculated into the bed but was attached to the upper part of the tailgate. The larger reverse flow in the wake and the enlarged adverse pressure area inside the bed influenced the drag increment when there was no attachment of the bed flow. For a low-drag pickup truck, the bed should be designed such that the bed flow is not attached to the tailgate and the reverse flow in the wake is small.

This study focuses on the aerodynamic performance of the supersonic biplane at the low-speed region. The performance was evaluated and discussed through Computational Fluid Dynamics (CFD) and Experimental Fluid Dynamics (EFD). The result of the CFD simulation was compared with the experimental result to validate the simulation and confirmed to be reliable. Therefore, the CFD results were employed to derive the aerodynamic performance coupled with the theoretical equations. In the wind tunnel experiment, the three-component force measurement was conducted to obtain lift, drag and pitching moment coefficients. The wake survey was conducted to measure the drag in detail. The results proved the low-speed aerodynamic performance of the supersonic biplane can be described by the classical ``general biplane theory'' reasonably well.

The two-dimensional (2D) supersonic biplane is well known for its shock wave cancellation effect and zero wave drag at supersonic speed. Nonetheless, in a three-dimensional (3D) setting, this favourable shock wave interaction becomes disturbed in the Mach cone regions at the wing tips of the supersonic biplane, which results in a severe drag penalty near the wing tips. However, this can be alleviated by appropriately designing the planform of the biplane. This study was performed to investigate the patterns of the 3D shock wave interaction using computational fluid dynamics by considering the 3D supersonic biplane with different planforms. The drag was shown to become large when the wing has a high sweepback angle, and it became small when the wing has a small taper ratio. However, an excessively low taper ratio could also lead to a drag greater than that of the rectangular supersonic biplane. It was found that the low drag 3D supersonic biplane has a small sweepback angle and an adequate taper ratio, which results in the vertex line becoming almost perpendicular to the freestream. Thus, in the design of the low drag 3D supersonic biplane, it is important to take into consideration the trade-off between the drag reduction at the mid-span section and the drag penalty near the wing tip and the wing root.

The two-dimensional (2D) supersonic biplane is well known for its shock wave cancellation effect and zero wave drag at supersonic speed. Nonetheless, in a three-dimensional (3D) setting, this favourable shock wave interaction becomes disturbed in the Mach cone regions at the wing tips of the supersonic biplane, which results in a severe drag penalty near the wing tips. However, this can be alleviated by appropriately designing the planform of the biplane. This study was performed to investigate the patterns of the 3D shock wave interaction using computational fluid dynamics by considering the 3D supersonic biplane with different planforms. The drag was shown to become large when the wing has a high sweepback angle, and it became small when the wing has a small taper ratio. However, an excessively low taper ratio could also lead to a drag greater than that of the rectangular supersonic biplane. It was found that the low drag 3D supersonic biplane has a small sweepback angle and an adequate taper ratio, which results in the vertex line becoming almost perpendicular to the freestream. Thus, in the design of the low drag 3D supersonic biplane, it is important to take into consideration the trade-off between the drag reduction at the mid-span section and the drag penalty near the wing tip and the wing root.

This study describes the nondeterministic effect of homogeneous atmospheric turbulence on sonic boom propagation. The Sears-Haack body was calculated in three dimensions by computational fluid dynamics in inviscid flow (Euler) mode to create the near-field pressure wave. The homogeneous atmospheric turbulence field was represented by the finite sum of discrete Fourier modes based on the von Karman and Pao energy spectrum. The sonic boom signature was then calculated by the modified waveform parameter method, considering a random velocity of homogeneous atmospheric turbulence. The results of this study indicated that atmospheric turbulence has ked influence oil sonic boom overpressure during its propagation to the ground. In comparison to the no-turbulence condition. we found that the sonic boom decreased in 59% of the cases and increased in 41% of the cases. Thus, homogeneneous atmospheric turbulence seems to favor a decrease, rather than an increase, in boom overpressure. In addition, we found that turbulence has a small effect on the propagation path front the flight attitude to the ground. Nonetheless, this small change in the propagation path may result in a variability, of the point at which the sonic boom reaches the ground, of tip to 1820 ft in the north-south direction (flight direction) and 115 ft in the east-west direction.

One of the most critical technical issues with regard to supersonic commercial transportation is the sonic boom that occurs during supersonic cruising flight, which causes impulsive noise on the ground. The "supersonic biplane theory" has been proposed to reduce the sonic boom. Shock wave interaction and cancellation between the wings of a supersonic biplane can be realized at a specific design Mach number, but does not work at off-design values. Here, the low-speed aerodynamic performance, as off-design performance, of a baseline supersonic biplane was investigated and discussed using experimental and computational fluid dynamics approaches. The thin airfoil stall characteristics of a supersonic biplane were shown to be caused by the stall of both upper and lower wings at an angle of attack of 20 degrees. Although there was leading flow separation of the upper wing at lower angles of attack, the stall of the lower wing was suppressed by interference with the upper wing. The lift of the lower wing was almost dominant to produce the lift of the supersonic biplane in the low-speed range. However, the lower wing caused greater drag than the upper wing at higher angles of attack.

The drag reduction system which consisted of the rear underbody slant and the rear flat flaps of a bluff-body was optimized by using design of experiments with robustness for cross wind. Drag force measurements accompanied by the orthogonal array were carried out at the low turbulence wind tunnel varying in the yaw angle. As a result of the analysis of variance, it was identified that the rear underbody slant angle, the lower and upper flaps' lengths were significant to the drag force while the side flap's length and interaction effects were insignificant. The slant angle, the most significant factor, should be set very carefully in designing the drag reduction device.

This study describes a method to overcome the choked-flow and flow-hysteresis problems of the Busemann biplane at off-design conditions, by using plain flaps at the leading and trailing edges. Two-dimensional analyses of four different biplanes were addressed in order to examine the effect of the flaps, using Computational Fluid Dynamics (CFD) in inviscid flow (Euler) mode. Results show that the leading-edge flaps can alleviate the drag increase due to the choked-flow, and reduce the area of flow-hysteresis. In contrast, the trailing-edge flaps are not effective in overcoming these problems. However, the trailing-edge flaps can reduce the wave drag near to the speed of sound, and also shift the drag-divergence Mach number to a higher one. Thus, the combined effect of both flaps allows us to smoothly achieve the design point (M∞ = 1.7) from subsonic regime, and to avoid the severe drag penalty due to the choked-flow and flow-hysteresis problems.

A new approach for multi-objective robust design optimization was proposed and applied to a practical design problem with a large number of objective functions. The present approach is assisted by response surface approximation and visual data-mining, and resulted in two major gains regarding computational time and data interpretation. The Kriging model for response surface approximation can markedly reduce the computational time for predictions of robustness. In addition, the use of self-organizing maps as a data-mining technique allows visualization of complicated design information between optimality and robustness in a comprehensible two-dimensional form. Therefore, the extraction and interpretation of trade-off relationships between optimality and robustness of design, and also the location of sweet spots in the design space, can be performed in a comprehensive manner.

The Busemann biplane is well known as the airfoil that has zero wave drag at the supersonic flight in the linear theory. It is found that this airfoil has a hysteresis in drag values from the transonic speeds through the low supersonic speeds based on Computational Fluid Dynamics (CFD) analysis. This paper shows that this hysteresis is explained by the Kantrowitz-Donaldson Criteria that usually defines the start and unstart of the supersonic intake.

The nonlinear lateral characteristics of a lifting-body type reentry vehicle were improved by applying efficient optimization techniques to determine the optimum configuration. First, the causes for the nonlinear lateral characteristics were investigated by computational fluid dynamics analysis. The results indicated that fins mounted on the Japan Aerospace Exploration Agency's baseline lifting-body configuration cause unsymmetrical development of vortices and result in nonlinear lateral characteristics. An efficient optimization algorithm, called efficient global optimization for multiobjective problems, was adopted to improve these nonlinear lateral characteristics. Efficient global optimization for multiobjective problems predicts potential optimum solutions based on the probability estimated by the Kriging model. The computational time for optimization was markedly reduced owing to use of the Kriging model. Further investigations were also performed to determine the effects of the upswept upper-aft and swept-back fin angles. The results indicated that the swept-back fin angle is an important factor controlling the lateral characteristics of the lifting-body type reentry vehicle.

We propose a new method of design called MORDE (multi-objective robust design exploration) that combines a multi-objective robust optimization approach and data-mining techniques for analyzing trade-offs. The probabilistic representation of design parameters, which is compatible with the Taguchi method, is incorporated into the optimization system we previously developed that uses a multi-objective genetic algorithm. The means and standard deviations of responses of evaluation functions against uncertainties in design variables are evaluated by descriptive Latin hypercube sampling using Kriging surrogate models. Design space is visualized by Self-organizing map (SOM). To extract design rules further, a new approach that adopts the association rule with an "aspiration vector" is proposed. MORDE is then applied to an industrial design problem with a centrifugal fan for a washer-dryer. Taking dimensional uncertainty into account, we optimize the means and standard deviations of the resulting distributions of the fan efficiency and turbulent noise level. Steady Reynolds-averaged Navier Stokes simulations are carried out to collect the necessary dataset for Kriging models. We demonstrate the advantages of the proposed method of multi-objective robust optimization over traditional non-robust ones in that the solutions are diverse. We clarify that the association rule can extract both sufficient and necessary conditions as design rules to achieve trade-off balances. The association rule is also more beneficial than SOM in finding quantitative relations, particularly those that are concerned with more than three design parameters.

The supersonic biplane is well known as the airfoil that has zero wave drag at the supersonic speed. This paper investigates aerodynamic characteristics of the supersonic biplane with the finite span length and effects of the aspect and taper ratios by comparing with the two-dimensional supersonic biplane which has the infinite span length using the computational fluid dynamics. The rectangular wing was employed to examine the effect of the aspect ratio and tapered wing was employed to inspect the effect of the taper ratio. Both the rectangular wing and the two-dimensional supersonic biplane choke at the high angle of attack. The drag of the three-dimensional wing becomes smaller when it has a higher aspect ratio. The drag polar curve of the three-dimensional wing approaches asymptotically to the drag polar curve of the two-dimensional supersonic biplane when it has a higher taper ratio. The drag of the tapered wing becomes smaller when it has an adequate taper ratio.

A hard disk drive (HDD) involves high-speed trailing flow on the rotating disk, which may cause vibration force. In order to reveal the information of flow fields related to the vibration force, this study performed the data mining in the unsteady computational fluid dynamics (CFD) simulation data for HDD flow fields. The present data mining started from the extraction of temporal characteristics from the time-series data of fluid properties given at each grid point, and then moved to the clustering analysis of the grid points based on the similarity of the temporal characteristics. Finally, the clustering results were projected onto a real space of the HDD for discussion. The present results indicated that a significant sequence of flow energy production and dissipation, which leads the vibration force, occurs near the filter entrance, arm root, and shroud wall between them.

Transition criteria for crossflow instability are investigated in order to apply the γ-Reθ transition model to boundary layer flows with strong three-dimensionality. The criteria are evaluated by comparing their predictions with experimental results for the flowfields around a cylinder at various yaw angles, and the flowfields around a swept wing. It is found that the crossflow criterion proposed by Kohama et al. is a promising one because it can be evaluated by using only local flow variables. Transition prediction of a yawed cylinder has been performed by the γ-Reθ transition model with the proposed crossflow criterion. The production of eddy viscosity due to the present criteria was confirmed. A modification of the criterion was suggested to treat its dependency on flow conditions. Copyright © 2009 by the American Institute of Aeronautics and Astronautics, Inc.

Measurement of wake vortices has been conducted at Sendai airport in Japan since April 2006. The measurements were obtained from 620 departures using a Doppler lidar, where measured aircrafts were Boeing 767, Airbus 320, Boeing 737 and MD-81. The collected measurements were classified by weather conditions such as a wind direction and wind strength to construct an advection database of wake vortices at Sendai airport. In this paper, a detection algorithm of wake vortices which was used to process all the lidar measurements automatically is firstly described. Then, details of the advection database are presented. Representative results from the advection database are shown in terms of a wind direction, a wind strength and static stability, which were dominant factors to affect the advection of wake vortices. Copyright © 2009 by the American Institute of Aeronautics and Astronautics, Inc.

A new approach for multi-objective robust design optimization was proposed and applied to a real-world design problem with a large number of objective functions. The present approach is assisted by response surface approximation and visual data-mining, and resulted in two major gains regarding computational time and data interpretation. The Kriging model for response surface approximation can markedly reduce the computational time for predictions of robustness. In addition, the use of self-organizing maps as a data- mining technique allows visualization of complicated design information between optimality and robustness in a comprehensible two- dimensional form. Therefore, the extraction and interpretation of trade-off relations between optimality and robustness of design, and also the location of sweet spots in the design space, can be performed in a comprehensive manner. © 2009 Springer-Verlag Berlin Heidelberg.

空港における実運航機の後方乱気流の解析は大気境界層や飛行経路を考慮する必要から一般に困難である.本研究では仙台空港において離陸機の後流をライダ計測し,その計測値を数値流体力学シミュレーションに融合する手法を開発した.ライダ計測値は4次元変分法により実際の計測プロセスを模擬しながら同化した.本論文では手法を説明した後,ライダ計測値の同化過程と再現された後方乱気流の挙動に関して議論する.

We propose a new multi-objective parameter design method that uses the combination of the following data mining techniques: analysis of variance, self-organizing map, decision tree analysis, rough set theory, and association rule. This method first aims to improve multiple objective functions simultaneously using as much predominant main effects of different design variables as possible. Then it resolves the remaining conflictions between the objective functions using predominant interaction effects of design variables. The key to realizing this method is the obtaining of various design rules that quantitatively relate levels of design variables to levels of objective functions. Based on comparative studies of data mining techniques, the systematic processes for obtaining these design rules have been clarified, and the points of combining data mining techniques have also been summarized. This method has been applied to a multi-objective robust optimization problem of an industrial fan, and the results show its superior capabilities for controlling parameters to traditional single-objective parameter design methods like the Taguchi method.

The nonlinear lateral characteristics of a lifting-body type reentry vehicle were improved by applying efficient optimization techniques to determine the optimum configuration. First, the causes for the nonlinear lateral characteristics were investigated by computational fluid dynamics analysis. The results indicated that fins mounted on the Japan Aerospace Exploration Agency&apos;s baseline lifting-body configuration cause unsymmetrical development of vortices and result in nonlinear lateral characteristics. An efficient optimization algorithm, called efficient global optimization for multiobjective problems, was adopted to improve these nonlinear lateral characteristics. Efficient global optimization for multiobjective problems predicts potential optimum solutions based on the probability estimated by the Kriging model. The computational time for optimization was markedly reduced owing to use of the Kriging model. Further investigations were also performed to determine the effects of the upswept upper-aft and swept-back fin angles. The results indicated that the swept-back fin angle is an important factor controlling the lateral characteristics of the lifting-body type reentry vehicle.

Data mining has been performed on the result,; of an aerodynamic design optimization of a two-stage-to-orbit reusable launch vehicle flyback-booster wing. Three data mining techniques were compared, including selforganizing map, functional analysis of variance, and the rough set theory. The optimization problem had four aerodynamic objective functions and 71 wing shape design variables. The hypothetical design database resulting from the optimization contained a total of 302 solutions which included 102 nondominated solutions. Consequently, the acquired knowledge of the design space consisted of general design characteristics, correlation between objective functions, and the effects of these design variables on the objective functions, for both nondominated as well as all solutions. The comparison also revealed the similarities and differences among the three data mining techniques used in this study. Even though all three techniques discovered detailed design knowledge and the results produced by the combination of all three methods compensated disadvantages of each method when applied individually, it was discovered that the self-organizing map produced the overall best results. Moreover, this study has also shown that the knowledge acquired from both nondominated solutions and from all solutions found was consistent despite the differences between the design spaces. Furthermore, it was shown that data mining is essential for visualizing results of an evolutionary multi-objective optimization problem and extracting useful design knowledge from these results.

Multi-Objective Optimization has been applied to a design problem of the twin engine concept for Silent Supersonic Business Jet (SSBJ). This problem aims to find main wing, body, tail wing and engine nacelle configurations, which can minimize both sonic boom and drag in a supersonic cruising flight. The multi-objective genetic algorithm (MOGA) coupled with the Kriging model has been used to globally and effectively search for optimal design candidates in the multi-objective problem. The drag and the sonic boom have been evaluated by the computational fluid dynamics (CFD) simulation and the waveform parameter method. As a result, the present optimization has successfully obtained low-boom and low-drag design candidates, which are better than the baseline design by more than 40% regarding each performance. Moreover, the structure of design space has been visualized by the self-organizing map (SOM).

An unsteady flow phenomenon of clear air turbulence was investigated by numerical simulation integrating flight data of encounters with severe turbulence. To determine the initial condition of the numerical simulation based on the flight data, the four-dimensional variational method was employed. The validation of the, present method was conducted by using pseudo flight data that are obtained from a reference flowfield produced by the numerical simulation. Then the four-dimensional variational method was applied to two actual severe turbulence cases to reproduce the flowfield of these turbulent flows. The validation results show reproducibility of the flowfield, even with spatially limited measurement data such as flight data, as long as the data involve primary flow features. Unsteady flowfields of turbulence were successfully simulated by using the four-dimensional variational method based on two cases of actual flight data. The results show the existence of vortices that correspond to clear air turbulence. In addition, the existing hazard indices F-factor and root-mean-square normal load were applied to the reproduced flowfield to obtain a spatial distribution of these hazard indices.

Numerical simulation of wake turbulence was performed by integrating the lidar measurements using four-dimensional variational method. A bogus vortex technique was adopted to ensure the existence of wake vortices in the flow field. The validation of the method was performed by an idealized test case using virtual lidar measurement which was produced by the reference simulation of a vortex pair. The results of the validation showed the convergence of a vortex parameter such as a circulation to the parameter of reference simulation case. It was also confirmed that the velocity distribution on a measurement plane agreed with reference one.

In this research, the wake turbulence from actual passenger airplanes taking off from Sendai airport was measured with Electronic Navigation Research Institute's Doppler laser radar (lidar). First, the influence of the surrounding wind on the behavior of wake vortices was investigated. The wake vortices in the crosswind case disappear more quickly from the runway than those of the low surrounding wind and head wind cases. In addition, the wake vortices in the case of large crosswind move faster than those in small crosswind. Next, the correction factor was estimated by using the pseudo lidar measurements based on Computational Fluid Dynamics (CFD). The corrected data for weak surrounding wind case agreed with the existing wake vortex model.

The stent design itself seems to be one of the factors responsible for restenosis. As a remedy, the present work attempts to perform a design optimization of coronary stents from a hemodynamic point of view. For the purpose, we have applied the principles of modern exploration of design space restricting ourselves to two-dimensional considerations. Width, thickness, and spacing of the struts of the stent formed the design variables. The objectives chosen for optimization were the vorticity generated, length of recirculation zone, and the reattachment distance in between the struts. Both semicircular and rectangular cross sections of stents were included. Starting with the range of design variables, sample stent cases were generated using Latin hypercube sampling. Objective functions were calculated for each of these by computing the two-dimensional flow using software FLUENT under the assumption of a steady, Newtonian flow considering a model stent with three struts. This was followed by Kriging to construct a response surface, which gives the relationship between the objectives and the design variables. The procedure gave nondominated fronts, which consist of optimized designs. Stents with minimum vorticity, with minimum recirculation distance, and the ones with maximum reattachment length in between struts were generated. The procedure is capable of producing the optimum set of design variables to achieve the prescribed objectives. [DOI: 10.1115/1.2885145]

In conventional windmills, the high tip speed creates aerodynamic noise, and when they are used at very low Reynolds numbers, their performance deteriorates due to laminar separation. These are important issues in modern windmills. Present study deals with a new windmill concept, the "flapping wing power generator," which would solve such problems. The concept is to extract energy via the flutter phenomena and the concept has been developed by some researchers. In 2003, Isogai 2003, "Design Study of Elastically Supported Flapping Wing Power Generator," International Forum on Aeroelasticity and Structural Dynamics, Amsterdam) proposed a new system. The system utilizes dynamic stall vortices efficiently and generates high power. The dynamic stall vortex is something that should be avoided in conventional windmills. They optimized the system to maximize the efficiency and obtained the set of design parameters, which achieved best efficiency. The system works at low frequencies and it enables high efficiency. To realize the system, it is necessary to consider the power and the efficiency. Thus, the present study optimized the system to maximize both the power and the efficiency. To obtain nondominated solutions, which are widely distributed in the design space, adaptive neighboring search, which is one of evolutionary algorithms, has been extended to handle multiple objectives and was used in the present study. Self-organizing map was used for the data mining. The trade-off between the power and the efficiency has been visualized. The trade-off curve was shaped by the constraints on the reduced frequency and the phase delay angle, which were imposed so that the dynamic stall phenomenon gives favorable effects on the power generation. The heaving amplitude was a parameter correlated to the objective functions. The reduced frequency and the phase delay angle change to control the heaving amplitude. Consequently, when the power is emphasized, the system undergoes a large heaving motion with a low frequency. On the other hand, when the efficiency is emphasized, the system undergoes a small heaving motion with a high frequency. Multiobjective optimization and data mining revealed the trade-off of the objective functions and the parameters correlated to the objective functions. The power obtained was comparable to that of present windmills at low tip-speed ratio region.

Numerical simulation of aircraft's wake turbulence was carried out by integrating lidar measurements. Four-dimensional variational method (4D-Var) was applied to integrate the lidar measurements into three-dimensional large eddy simulation (LES). This paper describes the results using lidar measurements at Sendai airport.

Large eddy simulation of wake turbulence was performed by integrating the lidar measurements. Four-dimensional variational method was applied to integrate lidar measurements with the three-dimensional large eddy simulation. In addition, a bogus vortex technique was adopted to ensure the existence of wake vortices in the flow field. The validation of the method was performed by idealized test cases using virtual lidar measurement which is produced by the reference simulation of a vortex pair. The results of the validation cases showed the convergence of vortex parameters such as a circulation to that of reference simulation case. It is also confirmed that the velocity distribution on a measurement plane agrees with reference one. Copyright © 2008 by the American Institute of Aeronautics and Astronautics, Inc.

Data mining is an important facet of solving multi-objective optimization problem. Because it is one of the effective manner to discover the design knowledge in the multi-objective optimization problem which obtains large data. In the present study, data mining has been performed for a large-scale and real-world multidisciplinary design optimization (MDO) to provide knowledge regarding the design space. The MDO among aerodynamics, structures, and aeroelasticity of the regional-jet wing was carried out using high-fidelity evaluation models on the adaptive range multi-objective genetic algorithm. As a result, nine non-dominated solutions were generated and used for tradeoff analysis among three objectives. All solutions evaluated during the evolution were analyzed for the tradeoffs and influence of design variables using a self-organizing map to extract key features of the design space. Although the MDO results showed the inverted gull-wings as non-dominated solutions, one of the key features found by data mining was the non-gull wing geometry. When this knowledge was applied to one optimum solution, the resulting design was found to have better performance compared with the original geometry designed in the conventional manner.

A hybrid evolutionary algorithm, consisting of a genetic algorithm (GA) and particle swarm optimization (PSO), is proposed. Generally, GAs maintain diverse solutions of good quality in multi-objective problems, while PSO shows fast convergence to the optimum solution. By coupling these algorithms, GA will compensate for the low diversity of PSO, while PSO will compensate for the high computational costs of GA. The hybrid algorithm was validated using standard test functions. The results showed that the hybrid algorithm has better performance than either a pure GA or pure PSO. The method was applied to an engineering design problem-the geometry of diesel engine combustion chamber reducing exhaust emissions such as NO., soot and CO was optimized. The results demonstrated the usefulness of the present method to this engineering design problem. To identify the relation between exhaust emissions and combustion chamber geometry, data mining was performed with a self-organising map (SOM). The results indicate that the volume near the lower central part of the combustion chamber has a large effect on exhaust emissions and the optimum chamber geometry will vary depending on fuel injection angle.

Two data mining techniques have been performed for a large-scale and real-world multidisciplinary design optimization (MDO) results to provide knowledge in design space. The MDO among aerodynamics, structures, and aeroelasticity for a regional-jet wing was carried out using high-fidelity evaluation tools on adaptive range multi-objective genetic algorithm. MDO generated 130 solutions included in nine non-dominated solutions. All solutions were investigated regarding tradeoffs among three objective functions and effects of design variables on objective functions using a self-organizing map (SOM) and a functional analysis of variance (ANOVA) to extract key features of the design space. Consequently, as SOM and ANOVA compensated with respective disadvantages, the design knowledge could be obtained more clearly by the combination between them. Although the MDO result showed the inverted gull-wings as non-dominated solutions, one of the key features revealed by data mining was the non-gull wing geometry. When this knowledge was applied to one optimum solution, the resulting design was found to have better performance compared with the original geometry designed in the conventional manner. Data mining can discover better design due to the salvage of information from design space even when optimization itself does not converge. Copyright © 2007 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

Data mining is an important facet of solving multi-objective optimization problems. In the present study, two data mining techniques were applied to a large-scale, real-world multidisciplinary design optimization (MDO) problem to provide knowledge regarding the design space. The use of MDO in the aerodynamics, structure, and aeroelasticity of a regional-jet wing was carried out using high-fidelity evaluation models with an adaptive range multi-objective genetic algorithm. As a result, nine non-dominated solutions were generated and used for tradeoff analysis of three objectives. All solutions evaluated during the evolution were analyzed for the influence of design variables using a self-organizing map (SOM) and a functional analysis of variance (ANOVA) to extract key features of the design space. As SOM and ANOVA compensate for respective disadvantages, the design knowledge could be obtained more clearly by combinating them. Although the MDO results showed inverted gull-wings as non-dominated solutions, one of the key features found by data mining was a non-gull wing geometry. When this knowledge was applied to one optimum solution, the resulting design was found to have better performance compared with the original geometry designed in the conventional manner.

A 5-year R&D project has been in progress in Japan toward the development of an environmentally friendly high performance small jet aircraft since 2003. Authors have been trying to apply a MDO tool to the aircraft design. In this MDO tool, SOM was adopted to extract the essential design knowledge from the optimization results. This paper presents the example of the application of the SOM to an engine-airframe integration design.

A numerical simulation was carried out corresponding to recent experiments using delta wings with sharp and blunt leading-edges, which indicates the second primary vortex, at NASA Langley Research Center. However, the experimental data did not reveal the detailed physical phenomena regarding the second primary vortex, because the experiment used only on-the-body-surface data. In the present study, the physical phenomena were revealed using Reynolds-averaged Navier-Stokes computations with three one-equation turbulence models on an unstructured hybrid mesh. The adaptive mesh refinement method in the vicinity of the vortex center was also applied to have more mesh resolution. Consequently, the result quantitatively revealed that appropriate modeling regarding turbulent kinematic viscosity was significant. Moreover, the three-dimensional visualization of the computational fluid dynamics results suggested that the second primary vortex was a developing shear layer merging into an open-type separation generated late by the primary vortex.

A new approach, Multi-Objective Design Exploration (MODE), is presented to address multidisciplinary design optimization (MDO) problems using computational fluid dynamics-computational structural dynamics (CFD-CSD) coupling. MODE reveals the structure of the design space from the trade-off information and visualizes it as a panorama for Decision Maker. The present form of MODE consists of the Kriging Model, adaptive range multi objective genetic algorithms, analysis of variance and a self-organizing map. The main emphasis of this approach is visual data mining. An MDO system using high-fidelity simulation codes, a Navier-Stokes solver and NASTRAN has been developed and applied to a regional-jet wing design. Because the optimization system becomes very expensive cornputationally, only brief exploration of the design space has been performed. However, visual data mining results demonstrate that design knowledge can produce a good design even after brief design exploration.

Performance of a fish caudal fin is brought out from many factors, such as the shape, the movement and the elasticity. The present study treats all of these factors simultaneously and attempts to visualize the complex design space using Kriging and SOM. As a result, the present study succeeded in visualizing the complex structure of the design space of the oscillating wing (caudal fin), and the combined effects of the design variables are shown. This data will become extremely useful for practical design of fish robots and other nautical machines.

We applied data mining techniques (Decision Tree Analysis and Rough Set Theory) to an optimal design database in order to extract useful design rules. The database was created using data obtained by solving of a four-objective aerodynamic optimization problem for a centrifugal fan, in which three types of aerodynamic efficiencies and the turbulence noise level were optimized. The extracted rules are expressed in "if…, then…" form and can be used by designers to make decisions. We demonstrated that these data mining techniques can be used to select important design variables applicable to various design objectives. While Decision Tree Analysis derives a single statistical rule, to which the most sensitive design variables are related, Rough Set Theory derives multiple rules, with which even non-linear interactions among design variables can be taken into account. We confirmed that using several data mining techniques is necessary for correct understanding of the design problem.

A new approach for multi-objective robust design optimization has been proposed and applied to a real-world design problem with a large number of objective functions. The present approach is assisted by response surface approximation and visual data-mining, which results in two major gains regarding computational time and data interpretation. The Kriging model for response surface approximation can realize accurate predictions of robustness measures, and dramatically reduces the computational time for objective function evaluation. In addition, the use of self-organizing maps as a data-mining technique allows visualization of complicated design information between optimality and robustness of design in a comprehensible two-dimensional form. Therefore, the extraction and interpretation of trade-off relations between optimality and robustness of design, and also the location of sweet-spots in the design space, can be performed in a comprehensive manner.

Data mining technique based on Rough Set theory has been applied to solutions obtained from three-objective optimization for regional jet wing design with engine nacelle. To reveal design tradeoffs, multiobjective optimization was performed by using Evolutionary Algorithms coupled with the Kriging model. High dimensional data (design parameters and the corresponding objective function values) are mapped onto the two-dimensional Self-Organizing Map (SOM) where global tradeoffs are visualized. The rule sets are derived by Rough Set theory so as to determine the importance of design parameters corresponding to the clusters obtained from SOM.

Large amounts of data regarding various physical phenomena can be acquired from CFD analysis of unsteady flow field. However, the post-process of the unsteady CFD data is usually performed by averaging each time-step data or by searching eventual correlations among points in the flow field. In those cases, most of the unsteady CFD data end up being unused. The objective of this work is to construct a Data Mining method for the post-process of unsteady CFD data. The Data Mining method in this study consists of a time-series analysis and a cluster analysis. The time dependent flow characteristics are extracted by the time-series analysis, while the features regarding space are extracted by the cluster analysis. The Data Mining method was applied to unsteady CFD data of base flow at transonic speed and it successfully captured instability of shear layer, shock wave and vortex shedding.

An unsteady flow phenomenon of clear air turbulence (CAT) was investigated by large eddy simulation (LES) using flight data available from airlines encounter with severe turbulence. To incorporate the flight data into the LES, the four-dimensional variational method (4D-Var) was employed. This paper presents formulation and numerical method of the 4D-Var applied to the flight data. The validation of the present method was conducted using pseudo-flight data obtained from reference flow field. Then, the 4D-Var was applied to two severe turbulence cases. Existing hazard indices, F-factor and root mean square (RMS) normal load were evaluated on the reproduced flow field. The validation results show reproducibility of flow field using limited measurements if the measurements include primary flow features. An unsteady flow of CAT was simulated by using the 4D-Var and the results show the existence of characteristic vortices which correspond to CAT.

This paper presents results of extensive computational experiments in which evolutionary multiobjective algorithms were used to find Pareto-optimal solutions to a complex structural design problem. In particular, Strength-Pareto Evolutionary Algorithm 2 (SPEA2) was combined with a mathematical programming method to find optimal designs of steel structural systems in tall buildings with respect to two objectives (both minimized): the total weight and the maximum horizontal displacement of a tall building. SPEA2 was employed to determine Pareto-optimal topologies of structural members (topology optimization) whose cross-sections were subsequently optimized by the mathematical programming method (sizing optimization). The paper also presents the shape of the Pareto front in this two-dimensional objective space and discusses its dependence on the building's aspect ratio. The results reported provide both qualitative and quantitative knowledge regarding the relationship between the two objectives. They also show the trade-offs involved in the process of conceptual and detailed design of complex structural systems in tall buildings.

The problem of designing a centrifugal blower was explored using multi-objective genetic algorithm and data mining techniques. Blade-to-blade regions of an impeller and a diffuser were modeled and time-averaged non-uniform inflow to the diffuser was considered. The design objectives were blower efficiency and uniformity of the inflow to the diffuser. The impeller's shape was represented by NURBS curves and then optimized. The obtained non-dominated solutions showed a trade-off relationship and the design variables controlling the trade-off were found to be related to the dimensions of the vane-less diffuser and the load balance of the impeller. We also applied Decision Tree Analysis and Rough Set Theory to reveal design rules, which led to good performance. Although the design rules derived from the optimization result and data mining results partly agreed with each other, we also clarified that there were some differences due to the characteristics of the data mining methods we used.

A new approach for multi-objective robust design optimization has been proposed and applied to a real-world design problem with a large number of objective functions. The present approach is assisted by response surface approximation and visual data-mining, which results in two major gains regarding computational time and data interpretation. The Kriging model for response surface approximation can realize accurate predictions of robustness measures, and dramatically reduces the computational time for objective function evaluation. In addition, the use of self-organizing maps as a data-mining technique allows visualization of complicated design information between optimality and robustness of design in a comprehensible two-dimensional form. Therefore, the extraction and interpretation of trade-off relations between optimality and robustness of design, and also the location of sweet-spots in the design space, can be performed in a comprehensive manner.

航空宇宙数値シミュレーション技術シンポジウム2006論文集

航空宇宙数値シミュレーション技術シンポジウム2006論文集

Diesel engine combustion chamber which reduces exhaust emission has been designed using CFD analysis and optimization techniques. In order to save computational time for design, the Kriging model, one of the response surface models, is adopted here. For a robust exploration, both the estimated function value of the model and its uncertainty are considered at the same time. In the present problem, the k-means method is used to limit the number of additional sample points to a reasonable level. Among the additional sample points, two combustion chamber shapes dominate the baseline configuration in terms of all objective functions. Compared with the previous optimization with the evolutionary algorithm, its computational time for design was cut by 95%. The results indicate that the present method is a practical approach for real-world applications.

An adaptive search region method for design optimization is proposed. The validity of the current search region is checked by investigating the probabilistic distribution of the design variable, and if the search region is inadequate, it is changed adaptively. To ease the computational burden, evaluation of the objective function, in both probabilistic analysis of the distribution of design variables and optimization, is performed using a Kriging model. The present method was validated by application to transonic airfoil design. Even when starting with an inadequate initial search region, it was possible to obtain an airfoil with good aerodynamic performance using the present method. Functional analysis of variance (ANOVA) was performed to identify the effects of each design variable on the objective function and constraints. The results indicated that the relevant definition of the search region is essential to obtain correct information from ANOVA.

The overset unstructured grid method is extended to viscous flow computations of multiple-bodies in contact. Overset treatment of hybrid meshes composed of tetrahedral, prismatic, and pyramidal elements is discussed in detail when those elements are embedded in or in contact with the wall boundary of another grid. This method is applied to the ONERA M5 wing-body model and computed by the Navier-Stokes code. The computational results show good agreement with those obtained by a single grid and the experimental data. The boundary layer in the intersection region between two grids is well captured without a collar grid.

The wing shape of flyback booster for a Two-Stage-To-Orbit reusable launch vehicle has been optimized considering four objectives. The objectives are to minimize the shift of aerodynamic center between supersonic and transonic conditions, transonic pitching moment and transonic drag coefficient, as well as to maximize subsonic lift coefficient. The three-dimensional Reynolds-averaged Navier-Stokes computation using the modified Spalart-Allmaras one-equation model is used in aerodynamic evaluation accounting for possible flow separations. Adaptive range multi-objective genetic algorithm is used for the present study because tradeoff can be obtained using a smaller number of individuals than conventional multi-objective genetic algorithms. Consequently, four-objective optimization has produced 102 non-dominated solutions, which represent tradeoff information among four objective functions. Moreover, Self-Organizing Maps have been used to analyze the present non-dominated solutions and to visualize tradeoffs and influence of design variables to the four objectives. Self-Organizing Maps contoured by the four objective functions and design variables are found to visualize tradeoffs and effects of each design variable.

One of the fundamental problems preventing commercial transport from going supersonic is the creation of Sonic Boom due to strong shock waves. Although it is nearly impossible to sufficiently reduce the shock waves generated by the aircraft designed conventionally to meet acceptable Sonic Boom levels at ground, it has been shown by the present authors that Sonic Boom can be reduced by using biplane. Here, the Busemann biplane is used as the basic shape of biplane configuration. Even though the Busemann biplane has a wide range of low wave drag, it has a choked flow phenomenon at off-design condition. Hence, we study methods for avoiding choked flow phenomenon for biplane by changing wing configuration. It was found that the configuration used in this research can alleviate the problem of sudden increase of the C_d value in choked flow phenomenon.

In this paper, Genetic Algorithm (GA) is adopted as the searching algorithm, and the Kriging model is used as the approximation model, for the study of the influence of the elasticity of the car tires on its noise performance. Finally, Self-Organizing Map (SOM) is used to study trade-off relationships between the objective functions.

A numerical simulation of Kelvin-Helmholtz (K-H) instability which is considered the major cause of clear air turbulence (CAT) is carried out. The Fourier analyses of velocity components are applied to extract a dominant vortex structure. A main vortex structure of K-H instability is reflected as peaks of spectrum for both horizontal and vertical velocities. Therefore, it is expected that the Fourier analysis of a horizontal velocity component is used to predict CAT.

The γ-Reθ transition model with a newly developed transition correlation set is applied to the stall angle prediction of thin airfoil, NACA 64A006. Empirical correlations used in the present model are developed based on literatures and numerical experiments and validated by skin friction coefficients on flat plate for several free stream turbulent intensities: T3 test cases for bypass transition regime and Schubauer& Klebanof for natural transition regime. For validation of separated flow transition, T3LC test case which uses coupland flat plate are computed and have a good agreement with experimental results of skin friction coefficient. The present γ-Reθ transition model is applied to stall angle prediction of NACA 64A006 airfoil to investigate the effect of boundary layer transition on massively separated flow problem within the framework of Reynolds-averaged Navier-Stokes computation. The results show that the γ-Reθ transition model improves the prediction capability of stall angle.

The overset unstructured grid method is extended to multiple bodies that are in contact with each other, and the treatment of nodes in the overlapping regions of two bodies in contact is discussed in detail. The developed overset unstructured grid method is applied to the following cases: ONERA M5 body/wing, ONERA M6 store separation, and the National Aeronautical Laboratory (NAL) experimental supersonic airplane/booster separation. In each case, there are two bodies in contact, and the calculation is performed with the Euler code. The calculation results clearly show smooth transitions of the contour lines between two bodies. In the ONERA M6 and NAL airplane/booster cases, a simulation of separation from the connecting configuration is performed.

This study applied a Kriging model to optimization of a constrained aerodynamic design problem. The objective function and all constraint functions are considered statistically independent to avoid treating the complicated multivariate normal distribution in the constrained optimization problem. In the Kriging model of objective functions, expected improvements (EI) is calculated and in the Kriging model of constraint function, the probabilities of satisfying the constraints are calculated. Based on these values, an efficient exploration of the global optimum is performed. Two data mining techniques are used to investigate the information of design space such as the relationship between objective function and design variables: Functional Analysis of Variance (ANOVA), and Self-Organizing Map (SOM). ANOVA shows information quantitatively, while SOM shows it qualitatively. Based on the information, elimination of design variables with little effect on objective function is performed. The present method is applied to two-dimensional (2D) transonic airfoil design. The results showed the validity of the present method.

In this paper we illustrate a new method for design of wings based on the optimization of the performance stability. This technique is an alternative approach to the traditional one which considers just a design point and tends to "over-optimize", producing solutions that perform well at the design point but have poor off-design characteristics. This method has been implemented for wing design in transonic field, searching for solutions that are stable in terms of performance with the fluctuations of Mach number and angle of attack. The high fidelity simulation has been performed with a Navier-Stokes code and a multi objective genetic algorithm has been used to optimize; in fact it has been proved that the design of airfoils, where the performance (lift and drag) and the stability are both objectives, is implemented more efficiently using a multi objective approach. In the optimization phase in order to reduce the number of needed simulations a response surface based on statistics (DACE) will be used.

A large-scale, real-world application of Evolutionary Multi-Objective Optimization is reported. The Multidisciplinary Design Optimization among aerodynamics, structures, and aeroelasticity of the transonic regional-jet wing was performed using high-fidelity evaluation models. Euler and Navier-Stokes solvers were employed for aerodynamic evaluation. The commercial software NASTRAN was coupled with a Computational Fluid Dynamics solver for the structural and aeroelastic evaluations. Adaptive Range Multi-Objective Genetic Algorithm was employed as an optimizer. The objective functions were minimizations of block fuel and maximum takeoff weight in addition to drag divergence between transonic and subsonic flight conditions. As a result, nine non-dominated solutions were generated and tradeoff information among three objectives was revealed.

In this paper, the effectiveness of the Fuzzy Decision Tree as a data mining technique for datasets obtained by Evolutionary Algorithms, which is a preferred optimization technique for multi-objective problems, is investigated. The Fuzzy Decision Tree is constructed to evaluate the influence of design variables on the objective functions. Based on the results of the Fuzzy Decision Tree, designers can reduce the number of design variables and focus on the important design variables. The technique is applied to the aerodynamic optimization of a supersonic wing design which has four objective functions and 72 design variables.

The wing shape of flyback booster for a Two-Stage-To-Orbit reusable launch vehicle has been optimized considering four objectives. The objectives are to minimize the shift of aerodynamic center between supersonic and transonic conditions, transonic pitching moment. and transonic drag coefficient, as well as to maximize subsonic lift coefficient. The three-dimensional Reynolds-averaged Navier-Stokes computation using the modified Spalart-Allmaras one-equation model is used in aerodynamic evaluation accounting for possible flow separations. Adaptive range multi-objective genetic algorithm is used for the present study because tradeoff can be obtained using a smaller number of individuals than conventional multi-objective genetic algorithms. Consequently, four-objective optimization has produced 102 non-dominated solutions, which represent tradeoff information among four objective functions. Moreover, Self-Organizing Maps have been used to analyze the present non-dominated solutions and to visualize tradeoffs and influence of design variables to the four objectives. Self-Organizing Maps contoured by the foul. objective functions and design variables are found to visualize tradeoffs and effects of each design variable.

A large-scale, real-world application of Evolutionary Multi-Criterion Optimization (EMO) is reported in this paper. The Multidisciplinary Design Optimization among aerodynamics, structures and aeroelasticity for the wing of a transonic regional jet aircraft has been performed using high-fidelity models. An Euler/Navier-Stokes (N-S) Computational Fluid Dynamics (CFD) solver is employed for the aerodynamic evaluation. The NASTRAN, a commercial software, is coupled with a CFD solver for the structural and aeroelastic evaluations. Adaptive Range Multi-Objective Genetic Algorithm is employed as an optimizer. The objective functions are minimizations of block fuel and maximum takeoff weight in addition to difference in the drag between transonic and subsonic flight conditions. As a result, nine non-dominated solutions have been generated. They are used for tradeoff analysis among three objectives. One solution is found to have one percent improvement in the block fuel compared to the original geometry designed in the conventional manner. All the solutions evaluated during the evolution are analyzed by Self-Organizing Map to extract key features of the design space.

Analysis of variance (ANOVA) and self-organizing map (SOM) were applied to data mining for aerodynamic design space. These methods make it possible to identify the effect of each design variable on objective functions. ANOVA shows the information quantitatively, while SOM shows it qualitatively. Furthermore, ANOVA can show the effects of interaction between design variables on objective functions and SOM can visualize the trade-offs among objective functions. This information will be helpful for designers to determine the final design from non-dominated solutions of multi-objective problems. These methods were applied to two design results: a fly-back booster in reusable launch vehicle design, which has 4 objective functions and 71 design variables, and a transonic airfoil design performed with the adaptive search region method. Copyright © 2005 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

A large-scale, real-world application of Evolutionary Multi-Criterion Optimization (EMO) is reported in this paper. The Multidisciplinary Design Optimization among aerodynamics, structures and aeroelasticity for the wing of a transonic regional jet aircraft has been performed using high-fidelity models. An Euler/Navier-Stokes (N-S) Computational Fluid Dynamics (CFD) solver is employed for the aerodynamic evaluation. The NASTRAN, a commercial software, is coupled with a CFD solver for the structural and aeroelastic evaluations. Adaptive Range Multi-Objective Genetic Algorithm is employed as an optimizer. The objective functions are minimizations of block fuel and maximum takeoff weight in addition to difference in the drag between transonic and subsonic flight conditions. As a result, nine non-dominated solutions have been generated. They are used for tradeoff analysis among three objectives. One solution is found to have one percent improvement in the block fuel compared to the original geometry designed in the conventional manner. All the solutions evaluated during the evolution are analyzed by Self-Organizing Map to extract key features of the design space.

Numerical simulation method for the prediction of supersonic intake performance is discussed. To perform a practical simulation, it is necessary to consider not only the nacelle-only configuration but also the engine/airframe configuration. In this study, the unstructured mesh, known to have advantages with regard to adaptation to complex geometry, was generated around aircraft geometry and the Euler equation was solved. To obtain higher numerical accuracy for the shock wave/boundary layer interaction, which appears at the inlet, the structured Navier-Stokes solver was applied to solve flows inside the flow-through nacelle. These regions were coupled with using the unstructured-structured zonal approach. The nacelle-only configuration and the engine/airframe configuration were calculated using this method and the effects of the airframe existence on supersonic intake performance were investigated.

The objective of the present study is to demonstrate performances of Evolutionary Algorithms (EAs) and conventional gradient-based methods for finding Pareto fronts. The multiobjective optimization algorithms are applied to analytical test problems as well as to the real-world problems of a compressor design. The comparison results clearly indicate the superiority of EAs in finding tradeoffs.

A multiobjective design optimization system of exhaust manifold shapes with tapered pipes for a car engine has been developed by using divided range multiobjective genetic Algorithm (DRMOGA) to obtain more engine power as well as to produce less pollutant. Although the present design problem is known to be highly non-linear, the exhaust manifold has been successfully designed to improve both objectives. A comparison of the results obtained by DRMOGA and MOGA was performed and DRMOGA was demonstrated to find better solutions than MOGA.

In this paper, wing-body configurations for a next generation Supersonic Transport are designed by means of Multiobjective Evolutionary Algorithms. SST wing-body configurations are designed to reduce the aerodynamic drag and the sonic boom for supersonic flight. To lower the sonic boom intensity, the present objective function is to satisfy the equivalent area distribution for low sonic boom proposed by Darden. Wing and fuselage is defined by 131 design variables and optimized at the same time. Structured multiblock grids around SST wing-body configuration are generated automatically and an Euler solver is used to evaluate the aerodynamic performance of SST wing-body configuration. Compromised solutions are found as Pareto solutions. Although they have a variety of fuselage configurations, all of them have a similar wing planform due to the imposed constraints. The present results imply that a lifting surface should be distributed innovatively to match Darden's distribution for low boom.

Global trade-offs for aerodynamic design of Supersonic Transport (SST) have been investigated by Multi-Objective Evolutionary Algorithms (MOEAs). The objectives are to reduce both drag and sonic boom to make next-generation SST more feasible. Adaptive Range Multi-Objective Genetic Algorithms (ARMOGAs) are utilized for the efficient search. The trade-offs are analysed by Self-Organizing Map (SOM), which provides a topology preserving mapping from the high dimensional space to two dimensions. ARMOGAs and SOM can be good design tools to conduct aerodynamic design optimizations and analyse the results.

Multidiciplinary optimization is a key element of the design process. To date multidisciplinary optimization methods that use low-fidelity methods are well developed. Gradient-based optimization methods that use data from three-dimensional linear aerodynamic solvers and two-dimensional structural solvers have been applied to complex aerospace configurations. However, use of high-fidelity methods such as Euler/Navier-Stokes methods for fluids and three-dimensional finite element method for structures is not as well developed. As an activity of the Multidisciplinary Design Optimization Technical Committee (MDO TC) of AIAA, an effort was initiated to assess the status and use of high-fidelity methods in multidisciplinary optimization. Contributions were solicited through the members of the MDO TC. This paper provides a summary of that effort.

The development of a high-fidelity aerodynamic design optimization tool based on evolutionary algorithms for turbomachinery is attempted. A three-dimensional Navier-Stokes solver was used for aerodynamic analysis, so that flowfields would be represented accurately and so that realistic and reliable designs would be produced. For efficient and robust design optimization, the real-coded adaptive range genetic algorithm was adopted, and the computation was parallelized and performed on an SGI Origin 2000 cluster to reduce turnaround time. The aerodynamic redesign of the NASA rotor 67 blade demonstrated the superiority of the present method over the conventional design approach, increasing adiabatic efficiency by 2% over the original design. This increase is achieved not only at the design condition, but over the entire operating range. This design optimization method has proven to be suitable for parallel computing. This promising tool is shown to help turbomachinery designers to design higher-performance machines while shortening the design cycle and reducing design costs.

Based on Navier-Stokes equations and structural and flight dynamic equations of motion, dynamic responses in vertical discrete gust flow perturbation are investigated for a supersonic transport model. A tightly coupled method was developed by subiterations between aerodynamic equations and dynamic equations of motion. First, under the assumption of rigid-body and single freedom of motion in the vertical plunging, the results of a direct-coupling method are compared with the results of quasi-steady model method. Then, gust responses for the one-minus-cosine gust profile are analyzed with two freedoms of motion in plunging and pitching for the airplane configurations with and without the consideration of structural deformation.

A fully implicit multiblock aeroelastic solve; which coupled thin-layer Navier-Stokes equations with structural equations of motion, has been developed for the flutter simulation on complex aerodynamic configurations. Navier-Stokes equations are solved with a lower-upper symmetric Gauss-Seidel subiteration algorithm and a modified Harten-Lax-van Leer-Einfeldt-Wada scheme. Structural equations of motion are discretized by a direct second-order differential method with subiteration in generalized coordinates. The transfinite interpolation is used for the grid deformation of the blocks neighboring the flexible surface. To evaluate the effectiveness of the grid deformation, a comparison is first done for the oscillating LANN wing with a rigidly. attached grid and a deforming grid. Then the method is applied to predict the flutter speed and frequency of the AGARD 445.6 standard aeroelastic wing. Finally, the aeroelastic instability referred to as aileron buzz is simulated for the supersonic transport model of the National Aerospace Laboratory of Japan.

Unstructured hybrid grids with 3 turbulence models are applied to simulate experiment about flows around delta wings with a blunt leading edge at Reynolds number of 6 millions. The modified S-A turbulence model is found effective to capture complex vortex structure. The second primary vortex at the inboard of the first primary vortex is captured numerically as reported in the experiment. Its structure is investigated by using CFD visualization.

In this research, multiobjective optimization system has been developed to reduce the aerodynamic drag and the sonic boom for SST wing-body configuration. Wing and fuselage are defined by 131 design variables and optimized at the same time. Multiblock grids around SST wing-body configuration are generated automatically and an Euler solver is used to evaluate the aerodynamic performance of SST wing-body configuration. Sonic boom is evaluated based on the equivalent area distribution by Darden. MOGA is used as an optimizer of the present optimization. To reduce the large computation time, CFD computations are parallelized by a simple master-slave concept on SGI ORIGIN 2000 at The Institute of Fluid Science at Tohoku University.

Multiobjective design optimization system of exhaust manifold shapes for a car engine has been developed using Divided Range Multiobjective Genetic Algorithm (DRMOGA) to obtain more engine power as well as to achieve less environmental impact. The three-dimensional manifold shapes are evaluated by the unstructured, unsteady Euler code coupled with the empirical engine cycle simulation code. This automated design system using DRMOGA was confirmed to find Pareto solutions for the highly nonlinear problems.

This paper presents numerical simulation of flowfields around a supersonic transport aircraft with integrated engine nacelles. In this study, flowfields were simulated by solving the Euler equations with the unstructured grid method for handling the complex geometry. To simulate intake flows at actual flight conditions, a bump was introduced inside the nacelle. The effect of nacelle mass flow ratios on overall aerodynamic performance was investigated in detail by changing bump heights. The spillage drag was calculated and found to have a large impact on the total drag. Computed results showed excellent agreements with wind tunnel data obtained at National Aerospace Laboratory of Japan.

Self-Organizing Maps (SOMs) have been used to visualize tradeoffs of Pareto solutions in the objective function space for engineering design obtained by Evolutionary Computation. Furthermore, based on the codebook vectors of cluster-averaged values of respective design variables obtained from the SOM, the design variable space is mapped onto another SOM. The resulting SOM generates clusters of design variables, which indicate roles of the design variables for design improvements and tradeoffs. These processes can be considered as data mining of the engineering design. Data mining examples are given for supersonic wing design and supersonic wing-fuselage design.

The design optimization of wings for supersonic transport by means of multiobjective evolutionary algorithms is presented. Three objective functions are first considered to minimize the drag for transonic cruise, the drag for supersonic cruise, and the bending moment at the wing root at the supersonic condition. The wing shape is defined by planform, thickness distributions, and warp shapes in total of 66 design variables. A Navier-Stokes code is used to evaluate the aerodynamic performance at both cruise conditions. Based on the results, the optimization problem is further revised. The definition of the thickness distributions is given more precisely by adding control points. In total, 72 design variables are used. The fourth objective function to minimize the pitching moment is added. The results of the revised optimization are compared with the three-objective optimization results as well as National Aerospace Laboratory's (NAL's) design. Two pareto solutions are found superior to NAL's design for all four objective functions. The planform shapes of those solutions are "arrow-wing" type.

Two-dimensional Navier-Stokes simulations have been performed for three types of flows at stall conditions about NACA63<SUB>3</SUB>-018, NACA63<SUB>1</SUB>-012 and NACA64A-006 airfoils. The Baldwin-Lomax algebraic eddy-viscosity model was applied in three different manners. First, a fully turbulent flow was assumed. Second, a transition point from laminar to turbulent was set manually. Third, the Degani-Schiff modification was applied. The fully turbulent computations predicted maximum lift coefficients much higher than experiment in all the cases, while the computations agreed well with experiment when specifying transition points properly. The computations with the Degani-Schiff modification predicted the stall angle close to experiment, but showed different behaviors at post-stall conditions. It is important to capture a laminar separation bubble at the leading-edge for predicting stall angles and post-stall behaviors correctly.

An implicit multiblock Navier-Stokes solver, which contains the LU-SGS subiteration method and the HLLEW scheme, has been developed for numerical simulations on complex and realistic aerodynamic configurations. Two-level halo cells are used to communicate data between abutting blocks. The transfinite interpolation (TFI) and the elliptic method with boundary control are employed to generate the initial multiblock grid and to smooth the grid distribution in each block. A comparison is first done for the AGARD supercritical LANN-wing with single- and multi-block grids. Then the present method is applied to a NASA transport wing/fuselage configuration and the NAL supersonic transport (SST) model. The choice of multiblock grid topology from the view of aeroelastic calculation and the comparison of Euler and Navier-Stokes solutions are investigated.

Self-Organizing Map (SOM) has been applied to analyze exhaust manifold design solutions obtained during multiobjective optimization using Evolutionary Algorithms. All the numerical data (design variables and the corresponding objective function values) are mapped onto the two-dimensional SOM where global tradeoffs are successfully visualized. SOM is confirmed to be a versatile datamining tool for fluid engineering.

Evolutionary Algorithm (EA) is applied to a practical three-dimensional shape optimization for aerodynamic design of an aircraft wing. Aerodynamic performances of the design candidates are evaluated by using the three-dimensional compressive Navier-Stokes equations. A structural constraint is introduced to avoid an apparent solution of zero thickness wing for low drag in high speeds. To overcome enormous computational time necessary for the optimization, the computation is parallelized on Numerical Wind Tunnel at National Aerospace Laboratory in Japan, a parallel computer with 166 vector-processing elements. The results ensure the capability of the EA in handling large-scale design optimizations.

The design optimization of wings for supersonic transport by means of multiobjective evolutionary algorithms is presented. Three objective functions are first considered to minimize the drag for transonic cruise, the drag for supersonic cruise, and the bending moment at the wing root at the supersonic condition. The wing shape is defined by planform, thickness distributions, and warp shapes in total of 66 design variables. A Navier-Stokes code is used to evaluate the aerodynamic performance at both cruise conditions. Based on the results, the optimization problem is further revised. The definition of the thickness distributions is given more precisely by adding control points. In total, 72 design variables are used. The fourth objective function to minimize the pitching moment is added. The results of the revised optimization are compared with the three-objective optimization results as well as National Aerospace Laboratory's (NAL's) design. Two pareto solutions are found superior to NAL's design for all four objective functions. The planform shapes of those solutions are "arrow-wing" type.

Wing flap-deflection angles of a supersonic transport are optimized to improve transonic cruise performance. Toward this end, a numerical optimization method is adopted using a three-dimensional unstructured Euler code and a discrete adjoint code. Deflection angles of 10 flaps, five for leading edge and five for trailing edge, are employed as design variables. An elliptic equation method is adopted to enable interior grid modification during the design process. Interior grid sensitivities are neglected for efficiency. Also tested is the validity of the approximate gradient evaluation method for the present design problem; it is found to be applicable for leading-edge flap design in cases where there are no shock waves on the wing surface. The Broydon-Fletcher-Goldfarb-Shanno method is used to minimize the drag with constraints on the lift and upper surface Mach numbers. Two design problems are considered; one involves a leading-edge flap design, and the other involves simultaneous design of leading-edge and trailing-edge flaps. The latter gave a smaller drag than the former by about two counts. Successful design results suggest that the present design method is valid and efficient.

An aerodynamic design method has been developed by using a three-dimensional Euler code and an adjoint code with a discrete approach. The method is applied to wing design problems of supersonic transports with wing-body-nacelle and wing-body configurations. Hicks-Henne shape functions are adopted for the surface geometry perturbation, and the elliptic equation method is employed for the interior grid modification during the design process. Interior grid sensitivities are neglected except those for design parameters associated with a body translation such as nacelles. The sequential quadratic programming method is used to minimize the drag with constraints on the lift and airfoil thickness. Successful design results confirm validity and efficiency of the present design method.

This paper describes an application of Adaptive Range Multiobjective Genetic Algorithms (ARMOGAs) to aerodynamic wing optimization. The objectives are to minimize transonic and supersonic drag coefficients, as well as the bending and twisting moments of the wings for the supersonic airplane. A total of 72 design variables are categorized to describe the wing's planform, thickness distribution, and warp shape. ARMOGAs are an extension of MOGAs with the range adaptation. Four-objective optimization was successfully performed. Pareto solutions are compared with Pareto optimal wings obtained by the previous three-objective optimization and a wing designed by National Aerospace Laboratory (NAL).

Real-coded adaptive range genetic algorithm (ARGA) has been applied to a practical three-dimensional shape optimization for aerodynamic design of an aircraft wing. The real-coded ARGA possesses both advantages of the binary-coded ARGA and the floating-point representation to overcome the problems of having a large search space that requires continuous sampling. The results confirm that the real-coded ARGA consistently finds better solutions than the conventional real-coded genetic algorithms do. © 2001 Elsevier Science B.V. All rights reserved.

In this paper, a study on designing a thick supercritical airfoil by utilizing Takanashi's inverse design method is discussed. One of the problems to design a thick supercritical airfoil by Takanashi's method is that an oscillation of the geometry may occur during the iteration process. To reduce the oscillation, an airfoil parameterization method is utilized as the smoothing procedure. A guideline to determine the target pressure distribution to realize the thick airfoil is also discussed.

This paper discusses the design optimization of a wing for supersonic transport (SST) using a multiple-objective genetic algorithm (MOGA), Three objective functions are used to minimize the drag for supersonic cruise, the drag for transonic cruise, and the bending moment at the wing root for supersonic cruise. The wing shape is defined by 66 design variables, A Euler flow code is used to evaluate supersonic performance, and a potential flow code is used to evaluate transonic performance, To reduce the total computational time, flow calculations are parallelized on an NEC SX-4 computer using 32 processing elements. The detailed analysis of the resulting Pareto front suggests a renewed interest in the arrow wing planform for the supersonic wing.

Real-coded Adaptive Range Genetic Algorithms (ARGAs) have been developed. The real-coded ARGAs possess both advantages of the binary-coded ARGAs and the use of the floating point representation to overcome the problems of having a large search space that requires continuous sampling. First, the efficiency and the robustness of the proposed approach are demonstrated by test functions. Then the proposed approach is applied to an aerodynamic airfoil shape optimization problem. The results confirm that the real-coded ARGAs consistently find better solutions than the conventional real-coded Genetic Algorithms do. The designed airfoil shape is considered to be the global optimal and thus ensures the feasibility of the real-coded ARGAs in aerodynamic designs.

A multiobjective genetic algorithm (GA) based on Fonseca-Fleming's Pareto-based ranking and fitness-sharing techniques has been applied to aerodynamic shape optimization of cascade airfoil design. Airfoil performance is evaluated by a Navier-Stokes code. Evaluation of GA population is parallelized on the Numerical Wind Tunnel, a parallel vector machine. The present multiobjective design seeks high pressure rise, high flow turning angle, and low total pressure loss at a low Mach number. Pareto solutions that perform better than existing control diffusion airfoils were obtained.

A practical three-dimensional shape optimization for aerodynamic design of a transonic wing has been performed using Evolutionary Algorithms (EAs). Because EAs coupled with aerodynamic function evaluations require enormous computational tune, Numerical Wind Tunnel (NWT) located at National Aerospace Laboratory in Japan has been utilized based on the simple master-slave concept. Parallel processing makes EAs a very promising approach for practical aerodynamic design.

Multi-objective genetic algorithms (MOGAs) have been applied to optimize an inverse design of a transonic wing shape. First, the wing planform is optimized by solving a multidisciplinary optimization problem based on aerodynamic, structural, and fuel storing objectives and constraints. Second, three-dimensional target pressure distribution is optimized for the aerodynamic inverse design with the previously designed planform, Minimization of the profile drag and the induced drag is performed under constraints on lift and other design principles. Applying these two preprocessing procedures by using MOGAs, Pareto surfaces can be studied for tradeoffs among multiple objectives. Thus, a designer is able to choose a good compromise for wing planform and target pressures for the inverse design. Corresponding for wing surface geometry is obtained by Takanashi's inverse method, and a sample design result is given.

A design optimization system for a supersonic wing has been developed. The present design system consists of the supersonic inverse design method and the Multiple-Objective Genetic Algorithm (MOGA) code for target pressure optimization. Furthermore, a wing twist specification technique is introduced to the present inverse design method by taking advantage of the non-uniqueness of three-dimensional aerodynamic inverse problems. With this method, a designer can select a wing that is easy to be manufactured and aerodynamically optimal. The designed wing shows better performance than the second baseline wing of National Aerospace Laboratory's (NAL) scaled supersonic experimental airplane program.

An inverse method for supersonic wing design has been developed. The present method is based on iterative "residual-correction" concept. The geometry correction is calculated by solving linearized small perturbation (LSP) equation with a difference of surface pressure distributions as a boundary condition. In the present method, LSP equation is analytically transformed to integrodifferential equations by using Green's theorem. Design results of an isolated wing and wing-fuselage configuration are presented here.

This paper examines several niching and elitist models applied to Multiple-Objective Genetic Algorithms (MOGAs). Test cases consider a simple problem as well as multidisciplinary design optimization of wing planform shape. Numerical results suggest that the combination of the fitness sharing and the best-N selection leads to the best performance.

Flowfields about a scramjet inlet model were numerically simulated by the Thin-Layer Navier-Stokes equations, and effect of inflow boundary layer thickness was investigated. The computations were conducted for inflow Mach number of 3.4. Although pressure at the throat region is relatively constant for various boundary layer thickness, corresponding flow pattern near the top wall was very different to each other, and the total pressure recovery decreased with the boundary layer thickness. The spilled impulse function increased with the thickness of inflow boundary layer, and stream thrust function of the spilled air slightly increased with the boundary layer thickness. The tendency was due to the increase of the separating region with the boundary layer thickness.

A genetic algorithm has been applied to optimize target pressure distributions for inverse design methods. Pressure distributions around airfoils are parameterized by B-spline polygons, and the airfoil drag is minimized under constraints on lift, airfoil thickness, and other design principles. Once target pressure distributions are obtained, corresponding airfoil/wing geometries can be computed by an inverse design code coupled with a Navier-Stokes solver. Successful design results were obtained for transonic cases with and without a shock wave.

A genetic algorithm has been applied to optimize target pressure distributions for inverse design methods. Pressure distributions around airfoils are parameterized by B-spline polygons, and the airfoil drag is minimized under constraints on lift, airfoil thickness, and other design principles. Once target pressure distributions are obtained, corresponding airfoil/wing geometries can be computed by an inverse design code coupled with a Navier-Stokes solver. Successful design results were obtained for transonic cases with and without a shock wave.

A Genetic Algorithm has been applied to optimize a wing shape for generic subsonic transport aircraft by using Navier-Stokes computations. To overcome enormous computational time necessary for this optimization, Numerical Wind Tunnel at National Aerospace Laboratory, a parallel vector machine with 166 processing elements, was used. Design results indicate feasibility of the: present approach for the aerodynamic optimization in advanced computational environments.

Unsteady Navier-Stokes simulations have been performed for vortical flows over an ''arrow-wing'' configuration of a supersonic transport in the transonic regime. Computed steady pressures and integrated force coefficients with and without control surface deflection at a moderate angle of attack are compared with experiment. For unsteady cases, oscillating trailing-edge control surfaces are modeled by using moving grids. Response characteristics between symmetric and antisymmetric oscillatory motions of the control surfaces on the left and right wings are studied. The antisymmetric case produces higher lift than the steady case with no deflection and the unsteady symmetric case produces higher lift than the antisymmetric case. The detailed analysis of the wake structure revealed a strong interaction between the primary vortex and the wake vortex sheet from the nap region when the flap is deflected up.

New capabilities have been developed for a Navier-Stokes solver to perform steady-state simulations more efficiently. The how solver for solving the Navier-Stokes equations is based on a combination of the lower-upper factored symmetric Gauss-Seidel implicit method and the modified Harten-Lax-van Leer-Einfeldt upwind scheme. A numerically stable and efficient pseudo-time-marching method is also developed for computing steady flows over flexible wings. Results are demonstrated for transonic flaws over rigid and flexible wings.

Unsteady Navier-Stokes computations have been performed for simulating transonic flows over wings with oscillating control surfaces using a locally moving grid and a stationary-mismatched zoning scheme. An F-5 wing and a clipped delta wing are chosen for the present study. The computed unsteady pressures and the response characteristics to the control surface motions are compared with experimental data. The results successfully predict main features of the unsteady pressure profiles, such as the double peaks at the shock wave and at the hinge line.

A multiobjective Genetic Algorithm based on Fonseca-Fleming's Pareto ranking method has been applied to optimize the three-dimensional target pressures for the aerodynamic inverse design of wing shape. The optimization problem was formulated to minimize the induced drag for wings as well as to minimize the viscous drag for airfoil sections. Performances of both the simple Genetic Algorithm and Vector Evaluated Genetic Algorithm were found unsatisfactory to the present optimization problem. The present design procedure was successfully applied to transonic wing design.

A new algorithm has been developed for obtaining higher-order accuracy in upwind schemes for the Euler and Navier-Stokes equations. In this method, the compatibility relations for the Euler equations are used to construct formulas for the higher-order interpolates. By using these formulas, computed results are obtained for steady, inviscid flow through a nozzle and also steady, inviscid, and viscous flow over an airfoil. This approach provides an alternative to total-variation-diminishing and essentially nonoscillatory schemes.

The viscous, three-dimensional flowfield of a helicopter rotor in hover is calculated by using an implicit, upwind, finite difference numerical method for the solution of thin-layer Navier-Stokes equations. The induced effects of the wake, including the interaction of tip vortices with successive blades, are calculated as a part of the overall flowfield solution without specifying any wake models. Comparison of the numerical results at subsonic and transonic flow conditions show good agreement with the experimental data for the surface pressures and the nearfield vortex trajectory. However, the captured vortex structure is diffused due to the coarse grids, but this appears to have minimal influence on the prediction of surface pressures. The paper also presents limited comparisons of the Navier-Stokes results with Euler results along with some discussion on grid refinement studies.

Unsteady Navier-Stokes computations have been carried out for simulating transonic flows over a clipped delta wing undergoing oscillatory and ramp motions, including flexibility. The implicit upwind algorithm has been validated by comparing the solutions with experimental data for the oscillatory pitching motion cases. The numerical and experimental results agree well at moderate angles of attack, where a leading-edge vortex develops. The ramp motion cases have demonstrated the effects of unsteadiness of the flowfield and structural flexibility on the wing responses. For the 10-deg ramp motion, a vortex breakdown is observed. The inviscid interaction with the shock wave plays an essential role in the process of the breakdown observed in the present calculation.

A streamwise upwind algorithm has been extended to compute unsteady flows with a moving grid system and applied to compute flows over oscillating wings at transonic Mach numbers. Comparisons have been made between results obtained from this upwind algorithm, using both temporally nonconservative- and conservative-implicit methods, with the results obtained from a central-difference method, and also with experimental data. The efficiency and practicality of the temporally nonconservative implicit solver are illustrated. Results to show the robustness and accuracy of the upwind method for unsteady computations compared to the central-difference method are presented.

A finite-volume upwind algorithm for solving the three-dimensional Euler equations with a moving grid has been developed for computing helicopter forward-flight rotor flows. The computed pressure distributions and shock positions of high-speed rotor flow are compared with various experimental data as well as with other numerical results, and the agreement is encouraging. A comparison of quasisteady solutions with unsteady solutions reveals that when a shock occurs in the flowfield, the assumption of quasisteady flow may fail due to the time lag of the shock motion. Similarly, three-dimensional effects cannot be neglected. Sufficient subiterations for each time step are required to avoid numerical lag effects in using the present method. The redistribution of the residual due to the coordinate transformation is discussed. For high-order Monotone Upstream-Centered Conservation Law (MUSCL)-type schemes, a coordinate-independent solution can be obtained by interpolating primitive variables.