Dong-Youn Kwak

Numerical Analysis on Flight-Test Results of Supersonic Experimental Airplane NEXST-1

A flight test of a supersonic experimental airplane (National Experimental Supersonic Transport, NEXST-1) was successfully conducted by the Japan Aerospace Exploration Agency in October of 2005. The experimental airplane was designed to reduce drag on supersonic cruise condition by means of some conventional and advanced design concepts: arrow wing planform, warp wing, area-ruled body, and natural laminar flow wing. The purpose of this study was to make clear the verification of those design concepts on the experimental airplane using the computational fluid dynamics analysis. The computational fluid dynamics results were compared with the flight-test results. The aeroelastic deformation and boundary layer transition were taken into account in the computational fluid dynamics analysis. This comparison highlighted the quantitative benefit of the natural laminar flow wing design concept. Furthermore, a real size supersonic transport could be designed using the drag reduction design technologies already validated from the NEXST-1 flight test. The results helped clarify the net drag reduction potentially attainable on a full-scale supersonic transport.

Flight Test Measurements of Surface Pressure on Unmanned Scaled Supersonic Experimental Airplane

Flight test of a supersonic experimental airplane was performed by Japan Aerospace Exploration Agency to improve advanced aerodynamic design technologies for next generation SST. The experimental airplane was designed for reduction of the drag on supersonic cruise condition. A lot of aerodynamic data in the flight test were obtained to validate the aerodynamic design concepts. High quality static pressure measurement systems were constructed to obtain the static pressure distributions. The pressure coefficient distributions in the flight test corresponded with CFD results that were target distributions obtained by optimum design. The pressure results were also consistent with the transition results and aerodynamic force characteristics in the flight test. Aerodynamic design concepts for drag reduction were demonstrated qualitatively and quantitatively by the NEXST-1 flight test.

Transition Measurement of Natural Laminar Flow Wing on Supersonic Experimental Airplane NEXST-1

layer. In this paper, the results of the transition measurement are introduced and compared with the numerically predicted results. The transition locations detected experimentally are in good agreement with the predicted locations, and the natural laminar flow effect is confirmed in the aerodynamic design conditions of the supersonic experimental airplane NEXST-1. Nomenclature C = local chord length CL = lift coefficient of full configuration Cp = surface pressure coefficient Cprms = fluctuation (rms) of surface static pressure of the wind tunnel E = dc output of the hot-film sensor e = ac output of the hot-film sensor H = altitude M = Mach number p = local surface pressure measured by the dynamic pressure transducer PPRT = local total pressure measured by the Preston tube Rec = Reynolds number based on the mean aerodynamic chord S = semispan length Tblow = time from the beginning of the blow in the windtunnel test Tlo = time from the liftoff in the flight test TTC = local temperature measured by the thermocouple X = chordwise position Xtip = axial length from the tip of the wind-tunnel model Y = spanwise position � = angle of attack ’ = circumferential angle from the top line on the nose 0 = fluctuation

Noise Generation Characteristics of a High-lift Swept and Tapered Wing Model

In this paper, noise generation from high-lift devices (HLDs) of a sweptand taperedwing model OTOMO2 measured in low speed wind tunnel experiments are discussed. OTOMO2 has a wing section of simplified high-lift configuration rectangular model OTOMO, and also has a planform of a realistic high-lift configuration aircraft model JSM. Main purpose of the experiment was to answer the questions about the influence of model geometry on HLD noise, which remained in our previous research using the rectangular model OTOMO. Results of the far-field noise measurement in RTRI-Maibara low noise wind tunnel and noise source maps obtained in JAXA-LWT2 low speed wind tunnel are mainly used. Comparison of narrow band spectra of the far-field SPL and noise source maps for the approach, the flap-only-deployed and the slat-only-deployed configurations showed characteristics of the noise generation of the flap-edge and the slat. One of the most remarkable finding was the existence of the multiple tonal component (MTP) in the slat noise spectrum in sweptand tapered-model geometry. The movement of the MTP noise sources along the spanwise location was observed, which was similar phenomenon to the rectangular model geometry. However, for the sweptand tapered-wing model, it is considered that the flow field which is suitable for the MTP generation is produced more locally and it changes continuously along spanwise direction with the variation of the angle of attack. The noise generation of the flap-edge in the sweptand tapered-model geometry was also different from that of the rectangular model geometry because there was the angle of attack dependency. Then it is suggested that there probably existed three different noise sources around the flap-edge.

Stereoscopic PIV measurements of leading edge separation vortices on a cranked arrow wing

Three-component velocity measurement results via stereoscopic PIV for leeward flow of a cranked arrow wing with a centre body are presented in order to understand leeward flow phenomena, emphasizing angle-of-attack effects and Reynolds number effects. The tests are conducted in a low-speed wind tunnel with the Reynolds number range from 0.15 × 106 to 1.5 × 106 at angles of attack of 8° to 20°. Based on instantaneous results, statistical properties such as turbulence kinetic energy are derived. The test results indicate that with increase in angle of attack, interaction between two leading edge separation vortices and a vortex on the body becomes strong, merging into a single vortex. In results on angle of attack effects, vorticity magnitude and kinetic energy have a positive correlation in the cores of the separation vortices. Results of Reynolds number effects show that with increasing Reynolds number, the secondary vortex structure becomes small with the two primary separation vortices moving outward, and the core diameter and vorticity of the vortices become larger and smaller, respectively, due to boundary layer transition on the inboard wing. In general, these Reynolds number effects are similar to the effects of decrease in angle of attack.

Rolling Moment Characteristics of Supersonic Transport Configurations at High Incidence Angles

Wind-tunnel tests were performed to investigate rolling moment characteristics of a cranked arrow wing supersonic transport configuration at high incidence angles. Force measurements and surface static pressure measurements were conducted when the wing model was rolled statically at a Reynolds number of 9.21 x 10 3 based on the mean aerodynamic chord. To understand the behavior of the leading-edge separation vortices formed both on the inboard and outboard of the cranked arrow wing, a stereoscopic particle image velocimetry survey was performed at 55 and 83% root chord locations. As the wing roll angle was increased, linear stable rolling moments were observed at low incidence angles, whereas abrupt changes from stable to unstable rolling moments were observed at incidence angle of 20 deg. Variations in suction peaks of pressure distributions at the rear part of the wing contribute to destabilization of the rolling moment component. These variations in suction peaks that induce the unstable rolling moment are related to the vortex breakdown characteristics. Chordwise breakdown locations of the inboard and outboard vortices are different between the windward and leeward wings. This difference causes the sudden change of rolling moments.

Investigation of Eddy Viscosity Turbulence Models on the Vortical Flow over the Highly Swept and Low Aspect Ratio Wing

A numerical investigation on a supersonic aircraft configuration at high alpha and low speed flight conditions has been performed using the Reynolds-Averaged Navier-Stokes Equations. The aerodynamic characteristics of the supersonic aircraft were strongly influenced by the vortical flow behaviors that formed over the wing and leading edge flaps. The paper is focused on the rotation curvature (RC) correction which can play an important role in the accurate simulation of the vortex dominant flow field. The RC correction effects were investigated using two eddy viscosity turbulence models (SA and SST). Moreover, these results were compared with simulations using the EARSM and experimental results. In results, it is found that small scaled vortices were detected by the RC correction. These trends are similar with the EARSM results. Furthermore, from the adjustment of the strength of the RC effects, the computational solutions were in good agreement with experimental results.

Investigation of Turbulence Models for the Supersonic Transport Configuration at Low-Speed and High Alpha Flight Condition

A numerical study on a supersonic transport at low speed and high alpha condition has been performed using Reynolds-Averaged Navier-Stokes equations. The aerodynamic performance is strongly influenced by the separated vortices formed over the wing. In this vortex dominant flow field, the characteristics of several turbulence models (SA, SA-RC, EARSM) were investigated in comparison with the experimental results. The vortex behaviors strongly depended on the turbulence models. These tendencies were clearly observed at the leading edge flaps and at the tail plane. It was found that the Cm of the tail plane is strongly influenced by the strake vortex. On the tail plane flow field, the numerical result by the SA-RC shows good agreement with the experimental result.

Comparison of CFD solvers for low speed vortex dominated flows

CFD computations on a supersonic aircraft configuration at low speed and high angle of attack have been performed using three different software packages with the aim of studying the capability of the code to simulate the flow and to check the solver dependency. The solvers used are ADCS, elsA, and TAU. A single hexahedral mesh is used throughout the study. Results show a good correlation between the solvers with some discrepancy in terms of vortical flow for the high angle of attack, high Reynolds number case.

Rolling Moment Characteristics at High Alpha on Several Planforms of Cranked Arrow Wing Configuration

Rolling moment characteristics at high alpha on several cranked arrow wings were investigated by low speed wind tunnel tests. Aerodynamic forces measurement and oil flow visualization were conducted for cranked arrow wing models with different leading edge kink locations and different inboard leading edge sweep back angles. Averaged velocity distributions at some chordwise locations were obtained by means of the stereoscopic PIV measurements. Interactions of the inboard and outboard vortices are observed at aft part of the kink locations. Behaviors of the inboard vortex breakdown are strongly affected by the difference of the inboard and outboard sweep back angle and the chord length at the kink.