The effects of the sidewall boundary layers in transonic shock tube airfoil flow were investigated. We attempted to correct the effects of the sidewall boundary layers using the Barnwell-Sewall and Murthy methods for shock tube boundary layers. Petersen’s boundary layer theory, which evaluates the modern wall-skin friction coefficients for shock tubes, was used in this analysis, and the results showed that the Mach number correction ΔM (the difference between the free stream Mach number (hot gas Mach number) and the corrected Mach number) increases as the hot gas Mach number M₂ increases under the condition of fixed time for the shock tube. This is caused by the boundary layer development, which grows thicker as the hot gas Mach number increases. Furthermore, when analysis is performed under the condition of constant displacement thickness 2δ^*⁄b, the Mach number correction ΔM gradually increases with an increase in the hot gas Mach number. This trend becomes very pronounced with increasing displacement thickness. In addition, after performing a comparison, we found that the correction of the shock wave location is in the direction of the improved agreement with the 2D CFD results when applied to the shock tube experiment.

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A resonance type flapping wing, which is suitable for an insect-like Micro Air Vehicle, is presented. It utilizes the resonance phenomenon of a two-degrees-of-freedom elastic system, that is, the wing is supported by the springs for flapping and feathering motions, and is oscillated at the resonance frequency of the system only by the torque in flapping motion. The amplitudes of flapping and feathering motions and the phase angle between them are controlled by changing the amount of damping. An optimum design method is applied to determine the structural parameters of the wing, the amount of applied torque and damping. Using these determined optimum quantities, arbitrary control of the flapping and feathering motions is possible. This paper presents the theoretical bases of the concept. The results of the numerical simulation using a Navier-Stokes code that takes into account the aeroelastic effects are also presented to verify the concept.

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Trailing edge vortex shedding from a turbine cascade was numerically simulated for an exit isentropic Mach number of 0.79 and a Reynolds number of 2.8×10⁶. The objective of this study is to clarify the time evolution of the vortex shedding process from a turbine blade and the mechanism of energy separation appearing in the wake. Calculations used a locally developed numerical code, employing a second-order AUSM scheme for inviscid numerical fluxes, a second-order implicit dual time method for time integration, and Detached Eddy Simulation for turbulence. Calculated results confirmed a non-uniform pressure distribution along the trailing edge, which was observed experimentally and different from a uniform distribution at low subsonic Mach number. The energy separation where instantaneous total temperature splits into hot and cold spots in the wake is caused by convection and vortex rotation. In addition, the formation and dissipation phases of vortices affect the energy separation.

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The Busemann biplane airfoil is considered one of the candidates for reducing sonic boom. In aircraft designs utilizing the biplane concept, high-lift devices must be used for takeoff and landing in low-speed conditions. In this work, flow visualizations were performed around a Busemann biplane airfoil equipped with leading and trailing edge flaps in a smoke wind tunnel. The lift coefficient of the biplane airfoil was estimated by utilizing a method based on measurements of smoke line patterns. The aspect ratio of the baseline Busemann biplane model was 0.75, the thickness ratio of the single element was 5%, and the wave cancellation condition was designed for Mach number 1.7. The length of each of the flap chords was 30% of the baseline. The Reynolds number, which is based on the chord length of the airfoil, is about 2.8×10⁵. The results of the study are summarized as follows. For the baseline Busemann airfoil without flaps, the lift coefficient increases linearly as the angle of attack increases. The slope of the lift coefficient c_l is 0.062 (1/deg.), which is in good agreement with reference data. This indicates that measuring smoke line patterns is a valid method for estimating the lift coefficient of biplane airfoils. Based on the visualization of the flow around the biplane model equipped with deflected leading and trailing edge flaps, confirmed that the separation bubble is smaller than in the baseline model due to the effective increase in camber. When the deflection angle of the trailing edge flap is increased, the lift coefficient also increases. The trend of the increasing c_l is similar to that of conventional monoplane airfoil models with trailing edge flaps. Therefore, such flaps can be considered effective high-lift devices for Busemann biplane airfoils.

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Nonlinear aeroelastic phenomena of a two-dimensional system in the transonic regime were examined numerically. Investigation was made of the unsteady aerodynamics excited by pitching and heaving oscillations, based on the first harmonic component of the aerodynamics computed by Navier-Stokes code. The results indicate that nonlinearity exists not only in oscillations accompanying massive flow separation but also in pitching oscillations with small amplitudes of less than 0.5°. Stability analysis of the first harmonic component of the unsteady aerodynamics resulted in a bifurcation diagram from which an estimate of the Limit Cycle Oscillation (LCO) amplitude could be obtained. The estimated LCO amplitude was confirmed by numerical simulation. Nonlinearity of the low-amplitude unsteady aerodynamics, which affects the stability boundary, was found to be highly correlated with boundary layer transition.

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Jet commercial aircraft in flight are frequently subject to atmospheric turbulence resulting in rapidly varying aerodynamic and flight dynamic characteristics. These varying characteristics not only pose threats to flight safety, but also may cause structural damages and reduce fatigue life. The sudden plunging motion in severe turbulence is the major reason to cause the flight injuries. To express the turbulence intensities as the hazardous levels, the root-mean-square g-load, usually for structural load, has been examined in the past twenty years. The present study is to examine alternative ideas to express the hazardous levels of sudden plunging motion in atmospheric turbulence. The flying-quality parameters of the plunging mode and integrated turbulence-induced downwash will be proposed as more reliable parameters for the turbulence hazard levels.

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The optimum aeroelastic design method for a resonance-type flapping wing for a Micro Air Vehicle (MAV) is presented. It uses Complex Method and 3D Navier-Stokes code to determine the optimum structural and aerodynamic parameters of a 2 DOF flapping wing system. The method is used to design a dragonfly-type MAV, and numerical simulation shows that the designed flapping wings can generate sufficient lift to sustain the weight and sufficient thrust to overcome the body drag.

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An efficient Multidisciplinary Design and Optimization (MDO) framework for an aerospace engineering system should use and integrate distributed resources such as various analysis codes, optimization codes, Computer Aided Design (CAD) tools, Data Base Management Systems (DBMS), etc. in a heterogeneous environment, and need to provide user-friendly graphical user interfaces. In this paper, we propose a systematic approach for determining a reference MDO framework and for evaluating MDO frameworks. The proposed approach incorporates two well-known methods, Analytic Hierarchy Process (AHP) and Quality Function Deployment (QFD), in order to provide a quantitative analysis of the qualitative criteria of MDO frameworks. Identification and hierarchy of the framework requirements and the corresponding solutions for the reference MDO frameworks, the general one and the aircraft oriented one were carefully investigated. The reference frameworks were also quantitatively identified using AHP and QFD. An assessment of three in-house frameworks was then performed. The results produced clear and useful guidelines for improvement of the in-house MDO frameworks and showed the feasibility of the proposed approach for evaluating an MDO framework without a human interference.

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