Seung-O Park

Unsteady Momentum Source Method for Efficient Simulation of Rotor Aerodynamics

T HE flow field around a rotor is very complicated as we all understand. Prediction of aerodynamic performance of rotor including not only rotor loading but also rotor wake in computational fluid dynamics (CFD) is still an important issue and also very challenging. Various computational methods based on Euler or Navier–Stokes equations with or without wake modeling were developed to predict the flow field around rotor [1–4]. Unsteady rotor-airframe interaction problems requiring both moving and stationary domains was recently simulated by using various techniques such as overset or Chimera grid or sliding mesh scheme [5–7]. Computational burden becomes rather heavy for a full Navier– Stokes simulation of rotor-airframe interaction because rotating blades and stationary body have to be dealt with simultaneously. Some approaches to mitigate this computational burden were suggested that adopted an actuator disk concept for rotating blades. There are two types of actuator disk methods: a pressure-boundary approach and a momentum-source approach [8–13]. O’Brien and Smith [14] discussed rotor-fuselage interaction models of pressureboundary method and momentum-source method using various load distributions. Schweikhard [15] implemented the time-averaged momentum source method in an unstructured flow solver. These actuator-disk models treat the whole disk plane swept by rotors as pressure jump or momentum source plane so that they yield a timeaveraged representation of the flow. Considering that the flow is essentially unsteady, these actuator-disk models are not expected to simulate properly the unsteady-flow features of the rotor. Recently Boyd and Barnwell [16] developed an unsteady rotor/fuselage interactional model that loosely couples a Generalized Dynamic Wake Theory (GDWT) to thin layer Navier–Stokes code with an overset grid. The GDWT estimates the unsteady blade loading by adopting the inflow model of Peters and He [17]. Tadghighi and Anand [18] developed an unsteady rotor source model for the interactional rotor/fuselage aerodynamics. The blade was represented by an unsteadymomentum source distribution in the form of a lifting-line type representing the loading only along a radial line of the blade. Kim and Park [19] tried an unsteady momentum source method by using a Navier–Stokes solver. Themomentum sourcewas evaluated through the blade element theory with inflow model of Peters andHe [17]. Recently, apart from rotor aerodynamics, a simple momentum source method was used to simulate the flowfield around vortex-generator arrays [20,21]. The momentum source magnitude was determined simply by using only the lift force of the vortex generators with a model constant. According to O’Brien and Smith [14], the momentum source method is known to be more stable numerically than the pressure boundary method in the region where the wake is very close to the rotor disk. Moreover the momentum source method is known to represent the rotor aerodynamics better than the pressure boundary method. In the present study, we develop an unsteady momentum source method for unsteady Navier–Stokes solver without employing additionalmodels for induced velocity and tip loss, which is the very first attempt to the authors’ knowledge. The momentum sources are distributed along radial and chord-wise directions of a rotor. The method suggested is validated by several simulation results.

Particle Image Velocimetry Measurement of Laminar Boundary Layer in a Streamwise Corner

Measurement of velocity profiles along a laminar boundary layer in a streamwise corner was carried out by using particle image velocimetry in contrast to conventional hot-wire anemometry. To make the measurement successful, the laser light source for particle image velocimetry was placed at a far-downstream station, and a light sheet was emitted in parallel to the corner intersection line. The platewas tilted at very small angles relative to the freestream to realize the laminar flow starting from the corner leading edge. Discussions on the characteristic features of the velocity profile in the bisector plane are given based on the measurement data. It was found that the measured velocity profiles correspond to the Blasius branch solution of the theoretical corner layer equations.

Stability analysis of a boundary layer over a hump using parabolized stability equations

Parabolized stability equations (PSEs) were used to investigate the stability of boundary layer flows over a small hump. The applicability of PSEs to flows with a small separation bubble was examined by comparing the result with DNS data. It was found that PSEs can efficiently track the disturbance waves with an acceptable accuracy in spite of a small separation bubble. A typical evolution scenario of Tollmien–Schlichting (TS) wave is presented. The adverse pressure gradient and the flow separation due to the hump have a strong effect on the amplification of the disturbances. The effect of hump width and height is also examined. When the width of the hump is reduced, the amplification factor is increased. The height of the hump is found to obviously influence the stability only when it is greater than the critical layer thickness.

Calculation of Low Aspect Ratio Wing Aerodynamics by Using Nonlinear Vortex Lattice Method

new computational procedure for the Non-Linear Vortex Lattice Method (NLVLM) is suggested in this work. Conventional procedures suggested so far usually involves inner iteration loop to update free vortex shape and an under-relaxation based iteration loop to determine the free vortex shape. In this present work, we suggest a new formula based on quasi-steady concept to fix free vortex shape which eliminates the need for inner iteration loop. Further, the ensemble averaging of the induced velocities for a given free vortex segment evaluated at each iteration significantly improves the convergence property of the algorithm without resorting to the under-relaxation technique. Numerical experiments over several low aspect ratio wings are carried out to obtain optimal empirical parameters such as the length of the free vortex segment, the vortex core radius, and the rolled-up wake length. ĀĀ

Wake Characteristics of Vane-Type Vortex Generators in a Flat Plate Laminar Boundary Layer

Experimental and numerical investigations were conducted to identify the wake characteristics downstream of two vanetype vortex generators over laminar flat plate boundary layer. Experimental study was carried out by using the stereoscopic particle image velocimetry. To describe the flow field around the vortex generator in detail, numerical study was performed. We considered two different planform shapes of vortex generator: triangular and rectangular shape. The height of the generator was chosen to be about the boundary layer thickness at the position of its installation. Two different lengths of the generator were chosen: two and five times the height. Wake measurements were carried out at three angles of attack for each configuration. Wake characteristics for each case such as overall vortical structure, vorticity distribution, and location of vortex center with downstream distance were obtained from the PIV data. Wake characteristics, as expected, were found to vary strongly with the geometry and angle of attack so that no general tendency could be deduced. Causes of this irregular tendency were explained by using the results of the numerical simulation.

Experimental Investigations of Systematic Errors in Wind Tunnel Testing Using Design of Experiments

The variation of systematic bias errors in the wind tunnel testing has been studied. A Design of Experiments(DOE) approach to an experimental study of fuselage drag and stability characteristics of a helicopter configuration was applied. When forces and moments measured in one time block differ significantly from measurements made in another time block under assumption that sample observations can be expected to yield same results within permissible measuring errors. The practical implication of this paper is that the systematic error can not be assumed not to exist. The those error reduction could be achieved through the process of randomization, blocking, and replication of the data points.

PIV Measurement of Laminar Boundary Layer in a Streamwise Corner

Measurement of velocity profiles along a laminar boundary layer in a streamwise corner was carried out by using PIV in contrast to conventional hot-wire anemometry. To make the measurement successful, the laser light source for PIV was placed at a far downstream station and light sheet was emitted in parallel to the corner intersection line. The plate was tilted at a very small angle relative to the free-stream to realize laminar flow from the start of the corner. Based on the measured velocity profiles, effects of plate leading edge shape and angle of incidence of the corner plate were discussed. It was found that the velocity profiles measured correspond to the lower branch solution of the theoretical corner layer equations.

An Experiment Study on Sideslip Angle Effect of Lambda Wing Configuration

ABSTRACT An experimental study on aerodynamic coefficients of a lambda w ing configuration was performed at the low speed wind tunnel of Agency for Defense Development. The main purpose of this study was to investigate the effects of sidesli p angle on various aerodynamic coefficients. In the case of 0 ° sideslip angle, nose-up pitching moment rapidly increases at a specific angle of attack. This unstable pitching moment characteristic is referred to as pitch break or pitch up. As the sideslip angle increases, the pitch break is found to be generated at a higher angle of attack. Rolling mome nt is found to show similar behavior pattern to ‘pitch break’ style with angle of attack at non-zero sideslip angles. This trend gets severer at greater sideslip angles. Yawing moment also shows substantial variation of the slope and the unstable directional stability with sideslip angles at higher angles of attack. These characteristics of the three moments clearly implies the difficulty of the flight control which requires efficient contr ol augmentation system.

Transition Prediction of Boundary Layers over Airfoils based on Boundary Layer Stability Theory

Transition location of boundary layers over airfoils is predicted by using PSE(Parabolized Stability Equations) and -method. Growth rates of disturbances are obtained from the PSE analysis and the N-factor curves are calculated by integrating the growth rates. The computational code developed in the present study is validated by comparing the computed results with the well known data for the cases of flat plate boundary layers and airfoils. Predictions of transition location are made for the boundary layers over NACA0012, NLF(1)-0414F, and NLF(1)-0416 airfoil. Predicted transition locations are found to be in good agreement with the experimental data.

Aerodynamic Simulation of Rotor-Airframe Interaction by the Momentum Source Method

To numerically simulate aerodynamics of rotor-airframe interaction in a rigorous manner, we need to solve the Navier-Stokes system for a rotor-airframe combination in a single computational domain. This imposes a computational burden since rotating blades and a stationary body have to be simultaneously dealt with. An efficient alternative is a momentum source method in which the action of rotor is approximated as momentum source in a stationary mesh system built around the airframe. This makes the simulation much easier. The magnitude of the momentum source is usually evaluated by the blade element theory, which often results in a poor accuracy. In the present work, we evaluate the momentum source from the simulation data by using the Navier-Stokes equations only for a rotor system. Using this data, we simulated the time-averaged steady rotor-airfame interaction and developed the unsteady rotor-airframe interaction. Computations were carried out for the simplified rotor-airframe model (the Georgia Tech configuration) and the results were compared with experimental data. The results were in good agreement with experimental data, suggesting that the present approach is a usefull method for rotor-airframe interaction analysis.