• Journal of the Korean Society for Aeronautical & Space Sciences
• /
• v.32 no.3
• /
• pp.10-16
• /
• 2004

In this paper the aeroacoustic characteristics of a helicopter main rotor system is measured by using a pair of scaled rotor blades. A low noise rotor test jig is developed for noise measurement and the rotational speed, thrust and torque are measured simultaneously in order to match the aerodynamic conditions with the full scale rotor. The accuracy of the force measurement device was checked through a calibration procedure. The measured thurst and torque with a 1.2m rotor are compared to the results of analytical prediction and showed that the thrust data at various rotational speed followed the prediction relatively well, but the torque data considered less accurate. It is also found that the background noise level of the test rig is sufficiently low, and the measured noise level from the rotor can be scaled with rotor tip speed. However, the Mach number dependancy and the directivity changes depend on the noise source characteristics.

• Journal of the Korean Society for Aeronautical & Space Sciences
• /
• v.32 no.10
• /
• pp.28-37
• /
• 2004

A two-dimensional direct numerical simulation was performed to investigate the evolution and vortical structure of a single vortex in reacting and non-reacting jet flow fields. A predictor-corrector-type numerical scheme with a low Mach number approximation was used, and a two-step global reaction mechanism was adopted as the combustion model. Through the comparisons of single vortex behaviors in reacting and non-reacting jet flow fields, it was found that the evolution characteristics and vortical structure of the single vortex were significantly influenced by a outer vortex that was generated from the buoyance effect as well as the chemical heat release. Furthermore, it was also identified that the differences of the vortical structure in reacting and non-reacting jet flow fields were mainly attributed to the thermal expansion, Baroclinic torque and buoyance effect.

• The Bulletin of The Korean Astronomical Society
• /
• v.38 no.2
• /
• pp.39.1-39.1
• /
• 2013

We use high-resolution hydrodynamic simulations to study nonlinear gas responses to imposed non-axisymmetric stellar potentials in barred-spiral galaxies. The gas is assumed to be infinitesimally thin, isothermal, and unmagnetized. We consider various spiral-arm models with differing strength and pattern speed, while fixing the bar parameters. We find that the extent and shapes of spiral shocks as well as the related mass drift depend rather sensitively on the pattern speed. In models where the arm pattern is rotating more slowly than the bar, the gaseous arms extend from the bar ends all the way to the outer boundary, with a pitch angle slightly smaller than that of the stellar counterpart. The arms drive mass inflows at a rate of ${\sim}0.5-2.5M{\odot}/yr$ to the bar region to which the shock dissipation, external torque, and self-gravitational torque contribute about 50%, 40%, and 10%, respectively. About 85% of the inflowing mass is added to bar substructures such as an inner ring, dust lanes, and a nuclear ring. while the remaining 15% encircles the bar region. On the other hand, models where the arms corotate with the bar exhibit mass outflows, rather than inflows, over most of the arm region. In these models, spiral shocks are much more tightly wound than the stellar arms and cease to exist in the region where $M{\bot}/sinp*{\geq}25-40$, where $M{\bot}$ denotes the Mach number of a rotating gas perpendicular to the arms with pitch angle p*. We demonstrate that the distributions of line-of-sight velocities and densities can be a useful diagnostic tool to distinguish if the arms and bar corotate or not.

• Transactions of the Korean Society of Mechanical Engineers
• /
• v.11 no.1
• /
• pp.118-123
• /
• 1987

In relation to the dust detonatians which have imposed severe damages on the industry, the ignitability of various dusts has been investigated on a horizontal shock tube in this study. By using a newly designed air injector, very well-distributed clouds could be obtained. The proper reflected shock conditions have been generated by placing a reflector 1.5cm behind the air injector, which reflected the incident shock wave. The incident shock waves in the range of Mach number 2.8-3.3 created the postreflected shock temperature of 1200-1600K. Experimentally the ignition delay was defined as the time interval between the arrival of a reflected shock wave at dusts and the detection of visible light. Measured ignition delays of dusts investigated were located lower than 1msec under the above conditions. These values are one-order higher than those in the incident shock wave condition. In this type of ignitiion process the following three processes are considered to play important roles; heating of a particle, generation of volatile gas by endothermic devolatilization process, and its diffusion from the particle surface and the formation of stoichiometric mixture with oxidizer.

• Journal of the Korean Society for Aeronautical & Space Sciences
• /
• v.32 no.8
• /
• pp.29-36
• /
• 2004

Supersonic inlet buzz can be defined as unstable subcritical operation associated with fluctuating internal pressures and a shock pattern oscillating about the inlet entrance. The flow pulsations could result in flameout in the combustor or even structural damage to the engine. An experimental study was conducted to investigate the phenomenon of supersonic inlet buzz on axisymmetric, external-compression inlet. An inlet model with a cowl lip diameter of 30mm was tested at a free stream Mach number of 2.0. Subcritical instability was investigated by considering the frequency of pressure pulsation and shock wave structure at the inlet entrance. The results obtained show that total pressure recovery ratios were varied from 0.42 to 0.78, and capture area ratio from 0.34 to 0.98. The frequency of the subcritical flow increased with decrease in capture area ratios. Frequency was measured at $224{\sim}240Hz$.

• International Journal of Aeronautical and Space Sciences
• /
• v.13 no.3
• /
• pp.386-397
• /
• 2012

The stability and structure of bluff-body stabilized hydrogen flames were investigated numerically and experimentally. The velocity of coflowing air was varied from subsonic velocity to a supersonic velocity of Mach 1.8. OH PLIF images and Schlieren images were used for analysis. Flame regimes were used to classify the characteristic flame modes according to the variation of the fuel-air velocity ratio, into jet-like flame, central-jet-dominated flame, and recirculation zone flame. Stability curves were drawn to find the blowout regimes and to show the improvement in flame stability with increasing lip thickness of the fuel tube, which acts as a bluff-body. These curves collapse to a single line when the blowout curves are normalized by the size of the bluff-body. The variation of flame length with the increase in air flow rate was also investigated. In the subsonic coflow condition, the flame length decreased significantly, but in the supersonic coflow condition, the flame length increased slowly and finally reached a near-constant value. This phenomenon is attributed to the air-entrainment of subsonic flow and the compressibility effect of supersonic flow. The closed-tip recirculation zone flames in supersonic coflow had a reacting core in the partially premixed zone, where the fuel jet lost its momentum due to the high-pressure zone and followed the recirculation zone; this behavior resulted in the long characteristic time for the fuel-air mixing.

• Journal of the Korean Society for Aeronautical & Space Sciences
• /
• v.35 no.8
• /
• pp.685-692
• /
• 2007

This paper describes the development procedure of GUI Program for the automated generation of airfoil performance table used in helicopter comprehensive code. Considering commercialization, the program is developed based on the Windows operating system. In addition, it is aimed to enhance user's convenience by including embedded postprocessor which enables real-time display of calculation procedure and grid system. Using the validated CFD code, the aerodynamic analyses are automated for a given range of Mach number and angles of attack. The computational grid system is designed to generate automatically once the surface coordinates are given. Mixed-Language scheme is employed in order to combine the CFD code in Fortran with C++ based GUI program, which makes the time-consuming code conversion unnecessary.

• Journal of the Korean Society for Aeronautical & Space Sciences
• /
• v.46 no.6
• /
• pp.445-451
• /
• 2018

This paper describes on introduction of CASE 4(Delta Wing) for EFD-CFD comparison workshop which is in charged of aerodynamic subcommittee of The Korean Society for Aeronautical and Space Science. The wind tunnel test will be performed later, angle of attack is set -5~20deg and mach number is set 0.7, 0.85, 1.2 to solve the transonic flow. The simulation test of grid dependency is conducted to determine the proper grid size of delta wing analysis. The tendency of lift and drag coefficient is determined according to the change of angle of attack based on the selected grid size.

• Journal of the Korean Society for Aeronautical & Space Sciences
• /
• v.46 no.8
• /
• pp.652-661
• /
• 2018

It is necessary to resolve the absolute velocity as well as Mach number to reflect the high temperature effect in high speed flow. So this region is classified as high enthalpy flows distinguished from high speed flows. Many facilities, such as arc-jet, shock tunnel, etc. has been used to obtain the high enthalpy flows at the ground level. However, it is difficult to define the exact test condition in this type of facilities, because some chemical reactions and energy transfer take place during the experiments. In the present study, a quasi 1D code considering the thermochemical non-equilibrium effect is developed to effectively estimate the test condition of a shock tunnel. Results show that the code gives reasonable solution compared with the results from the known experiments and 2D axisymmetric simulations.

• Journal of the Korean Society for Aeronautical & Space Sciences
• /
• v.46 no.10
• /
• pp.791-797
• /
• 2018

In this work, the high-altitude environment simulation study was carried out at an altitude of 65 km exceeding Mach number of 6 after the launch of Korean Space Launch Vehicle using a shock tunnel. To minimize the flow disturbance due to the strut support of test model as much as possible, a few different types of strut configurations were considered. Using the configuration with minimum disturbance, the high-altitude environment simulation experiment including a propulsion system with a single-plume, was conducted. From the thruster test through flow visualization, not only a shockwave pattern, but a general flow-field pattern from the mutual interaction between the exhaust plume and the free-stream undisturbed flow, was experimentally observed. The comparison with the computation fluid dynamic(CFD) results, showed a good agreement in the forebody whereas in the afterbody and the nozzle the disagreement was about ${\pm}7%$ due to unwanted shockwave formation emanated from the nozzle-exit.