Aeroacoustics is a branch of acoustics that studies noise generation via either turbulent fluid motion or aerodynamic forces interacting with surfaces. Noise generation can also be associated with unsteadily varying flows. Classical aeroacoustic analysis relies upon the so-called Acoustic Analogy which proposed by James Lighthill in the 1950s. Computational Aeroacoustics (CAA) is the application of numerical methods to find approximate solutions of the governing equations for specific aeroacoustic problems. The term of ‘Computational Aeroacoustics’, which can be distinguished from Computational Fluid Mechanics (CFD) and Computational Acoustics, is dated back to just the 1980s. In direct numerical simulation approach to CAA, compressible Navier-Stokes equation which describes both the flow field and the acoustic field is solved directly compared to integral methods such as Lighthill’s acoustic analogy and Kirchhoff integral method. This simulation requires very high numerical resolution due to the large differences in the length scale present between the acoustic variables and the flow variables.
The need of accurate and efficient numerical algorithms with high truncation order and high resolution has been increased for CAA in that these are able to simulate the generation and propagation of high wavenumber or high frequency and small amplitude wave components directly. These are almost non-dissipative and less dispersive than the standard low order ones that have been used widely so far. In our laboratory, the optimized fourth-order penta-diagonal compact scheme is developed for the evaluation of spatial derivatives in the compressible Navier-Stokes equation. The scheme is one of the best spatial derivative schemes in sense of the non-dispersion and non-dissipation property, and it is almost comparable to spectral method.
In the aeroacoustic problems which require strong conservation of flow conserved variables, such as shock problems, cell-centered finite volume methods with upwind biased flux schemes are more suitable than node-based finite difference methods. Inherently upwind schemes are more robust but more dissipative but than central schemes. To overcome the dissipation error, many researchers have designed upwind schemes of high order truncation error. In our laboratory, we focused on this point with compact reconstruction and optimization; so, optimized compact reconstruction(OCR) was developed. With high order, upwind schemes are also experienced with spurius oscillation errors. To impose non-oscillatory properties or nonlinear stability to the schemes, Total Variation Dimishing(TVD), Essentially Non-Oscillatory(ENO) or Monotonicity Preserving(MP) properties can be used.
Generally, computational fluid dynamics (CFD) for rotors requires huge computation resources with numerous grid points to capture tip vortices/wakes and flows around the blade in order to reduce the numerical dissipation of the vortical wake flows in the CFD of the rotor. This paper proposes an Eulerian-Lagrangian coupled method with a freewake method in the rotor wake to describe the vortex behaviors and to provide the boundary condition in the CFD domain around the blade. The coupled method reduces the CFD domain in the wake and around the blade. In addition, the compressible linearized potential panels described by the doublet and source around the blade are applied to include the CFD boundary condition due to the induced flow of the blade circulatory flow and thickness effect. The CFD domain around the blade is reduced further by employing the compressible panel method rather than a lifting line method in providing the CFD boundary condition. The strength of the doublet potential panel is obtained from the surface pressure calculated by CFD with the Bernoulli equation and potential theory. To investigate this method in detail, domain size and linearized compressibility correction studies are performed for a hovering flight condition.Tightly coupled method using Navier-Stokes, Tip-vortex penetration from the freewake
potential panel, and freewake computation into the Navier-Stokes domain
The loss of control effectiveness causing helicopter accidents can be classified into four different types of control: yaw, vertical, roll, and pitch. The accident rate by the loss of yaw control is the highest, and this might be closely related with a tail rotor and an anti-torque issue. The tail rotor produces the thrust force to counterbalance the torque generated by the main rotor. In case of the helicopter control, yawing and rolling controls are coupled because of the characteristics of helicopter control mechanism. This is the reason why flying a helicopter is usually considered to be difficult and dangerous. Understanding a hovering turn flight could be a good first step to analyse and reduce the helicopter accidents caused by the loss of yaw control or tail rotor problems. Even though the hovering turn looks simple, the flow physics during the hovering turn or hovering flight under crosswind is more complex and unsteady than expected. This is because the fuselage is strongly influenced by the main rotor wake as well as the transient crosswind whose magnitude also varies with the distance from the center of rotation. Also, the cross wind makes the tail rotor operate in climb or descent flight. Thus, the tail rotor could enter into vortex ring state. It would cause the loss of yaw control effectiveness and consequently helicopter accidents. The purpose of the present research is to understand the flow characteristics around a helicopter during its hover turn flight.
Helicopter heading change and wake geometry during hover turn
A potential based panel method coupled with advanced time-marching free-wake techniques is developed to achieve fast and accurate prediction for unsteady aerodynamics and wake dynamics of helicopter rotating blades. This coupling analysis is enabled by using the equivalence of the doublet wake panels and the vortex filaments. The coupled panel method allows the inclusion of the self-induced velocity of curved vortex filaments and high order time integration for the computation of wake convection. A parallel computation is applied to the wake convection for fast numerical calculation. The computation of the induced velocity from each vortex filament is parallelized and computed separately. The velocity field integration technique is used to avoid numerical singularity during the interaction between the wake and blades. It is found that blade-pressure predictions and the wake roll-up agree well with the measured data for helicopter rotors, both in hover and forward flight. Tip vortices paring phenomena are also predicted and compared with the measured data.
Wake description using equivalence of doublet Vorticity contour in the cross section for
and vortex Caradonna-Tung rotor in hover
Ref. Wie, S. Y., S. K. Lee and D. J. Lee, "Wie, S. Y., S. K. Lee and D. J. Lee, "Potential Panel and Time-Marching Free-Wake Coupling Analysis for Helicopter Rotor," Journal of Aircraft, Vol. 46, No. 3, May-June 2009, pp. 1030-1041.
Aerodynamic force of the rotor is calculated by the panel method and the rotor wake is simulated by the vortex particle method. In case of the vortex method using vortex filaments or doublet panels to describe the wake, the wake components are connected each other so that it might cause a singularity problem as the wake penetrates the body surface. However, in vortex particle method, the wake moves individually without restriction of connectivity and it is treated to reflect on the surface before penetrating the body. Consequently, the vortex particle is easier to treat wake-body interaction and has better performance to expect ground effect.
A viscous vortex particle method solves vorticity transport equation obtained from taking curl of Navier-Stokes equation. There are two parts in the vorticity transport equation: a vortex stretching term and viscous diffusion term. Direct scheme is used for solving the stretching term, which directly calculates velocity gradient from vortex particles. The Particle-Strength-Exchange (PSE) method is used for the viscous diffusion. The characteristic computation speed of the vortex particle method is O(N2), which would be very slow as the number N becomes large. Therefore, the parallel computing is adopted to reduce the running time.
Open suction type
Test section cross size 0.35m X 0.35m
Contraction ratio 21:1
Anechoic room 6.0m X 5.0m X 4.0m
Maximum wind velocity 62.8m/s
An anechoic chamber is a shielded room designed to attenuate waves. Anechoic chambers were originally used in the context of acoustic (sound) echoes caused by reflections from the internal surfaces of the room. Anechoic chambers are commonly used in acoustics to conduct experiments in nominally "free field" conditions. All sound energy will be traveling away from the source with almost none reflected back. Common anechoic chamber experiments include measuring the transfer function of a loudspeaker or the directivity of noise radiation from industrial machinery.
In general, the interior of an anechoic chamber is very quiet, with typical noise levels in the 10-20 dBA range. Full anechoic chambers aim to absorb energy in all directions. Semi-anechoic chambers have a solid floor that acts as a work surface for supporting heavy items, such as cars, washing machines, or industrial machinery, rather than the mesh floor grille found over absorbent tiles present in full anechoic chambers. This floor is damped and floating on absorbent buffers to isolate it from outside vibration or electromagnetic signals. Recording artists recording studio may utilize the semi-anechoic chamber to produce high-quality music free of outside noise and unwanted echoes. Anechoic chambers range from small compartments to chambers as large as aircraft hangars. The size of an anechoic chamber depends on the size of the objects to be tested and the frequency range of the radio or microwave signals used.
A wind tunnel is a research tool developed to assist with studying the effects of air moving over or around solid objects. Air is blown or sucked through a duct equipped with a viewing port and instrumentation where models or geometrical shapes are mounted for study.
Wind tunnel test section is to analyze about the effect of wall and 2 dimensional effect. The viscosity of wall and the interference between wall and body should be considered.
Rotor stand is test equipment for rotor performance and noise measurement. Rotor system is installed on the top of stand with swash plate and we can control collective pitch angle of rotor blade. Using this equipment, we can get the performance of rotor system for hovering case and noise for several collective pitch angles. And partially inclined ground effect is tested using inclined test. When the helicopter land on the roof of building or on the ship, this inclined ground effect can be occurred. So, using this experiment, we can test the stability of rotor system in case of inclined ground effect. High speed camera is used to increase the reliability of pitch angle measurement and several sub-stands are used to avoid vibration problem.
Many smart morphing wing systems rely on smart actuators such as piezoelectric actuators, shape memory alloy (SMA) actuators and pneumatic artificial muscles (PAMs) to change the shape of wings. Among them, SMA actuators have the advantage of the highest power-to-weight ratio at low force ranges (Left figure), and can be used as "direct drive linear actuators" with little or no additional motion reduction or amplification hardware. Hence, the actuating system can be easily simplified and miniaturized in many applications. The SMA-actuating system is implemented to a trailing edge flap system (right figure).
SMAs have non-linear behavior and transform hysteresis come from between the heating and cooling transitions. Than means that the constant inputs involved with the SMA's temperature cannot control the specific deformation of them. The unpredictable factors such as ambient temperatures and external loading conditions, for example, wine tunnel conditions are also the difficult reasons to control the SMAs by a feed-forward control method. For that reason, the proportional-integral-differential (PID) control method, one of the feedback control methods, was implemented to control the SMA-actuated flaps accurately. Left figure shows the block diagram of the control system of the wing model. In case of two-channel system, two independent control systems were constituted in parallel in a system ot synchronize time. Subsequently, wind tunnel experiment was performed for the purpose of evaluating the operational performance of the two flaps with flow condition. Additionally, the unsteady aerodynamic forces were measured in various flap motions and angle of attacks.