• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.44 no.6
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• pp.477-483
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• 2016

Velocity profiles of laminar, transition and turbulent boundary layers were investigated by using Particle Image Velocimetry(PIV) measurements on the flat plate at Mach 2.96. The Schlieren visualization and PIV measurements are also used to confirm whether the oblique shock wave generated from the leading edge affects the flow field over the flat plate. The laminar velocity profile measured from the experiment was well matched with the compressible Blasius solution. The velocity profile of the transition boundary layer was well correlated with the theoretical turbulent velocity profile from near the wall and the transition began from Re = $1.41{\times}106$. For the turbulent boundary layer, considering compressibility effects, the Van Driest-transformed velocity satisfies the incompressible log-law. It is found that the log region is extended farther in the wall-normal direction compared to the log region in incompressible boundary layer.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.30 no.1
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• pp.28-35
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• 2002

In cavity flow, the rectangular step generates so strong sound that many researchers have investigated method to suppress the nois during interaction between vortical flow and rectangular forward step. In this study the flow noise from the vortex motion in two-dimentional low Mach number flow past a forward step with various chamfering angle is calculated numerically. Inviscid incompressible discrete vortex model and matched asymptotic expansion(MAE) theory are applied to obtain the inner flow field and the outer noise field. Both source acoustic pressure and sound intensity are obtained with various chamfering height, chamfering angle and initial vortex position. The pressure amplitude is most suppressed when the chamfering angle is between $15^{\circ}C$ and $30^{\circC}$ at the chamfering length of 30% of the step height.

• Journal of the Korean Society of Propulsion Engineers
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• v.10 no.3
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• pp.41-47
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• 2006

Effects of mean flow and nozzle damping on acoustic tuning of a gas-liquid scheme coaxial injector are investigated numerically adopting a linear acoustic analysis. The injector plays a role as a half-wave acoustic resonator for acoustic damping in a combustion chamber of a liquid rocket engine. As Mach number of mean flow in a chamber increases, the resonant frequency of the first tangential mode decreases slightly and the optimum injector tuning length varies negligibly. Nozzle damping affects neither the resonant frequency nor the optimum length. From these numerical results, effects of mean flow and nozzle damping on acoustic tuning of a resonator are negligible. As open area of the injectors increases, the acoustic amplitude decreases, but new injector-coupled modes appear.

• Journal of the Korean Society of Propulsion Engineers
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• v.7 no.3
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• pp.15-21
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• 2003

This experimental study describes the propagation characteristics of the impulse noise emitted from the exit of a perforated pipe attached to the open end of a simple shock tube facility. The pressure amplitudes and directivities of the impulse wave propagating outside from the exit of pipe with several different configurations are measured and analyzed for the range of the incident weak shock wave Mach number between 1.02 and 1.2. In the experiments. the impulse waves are visualized by a Schlieren optical system for the purpose of understanding their propagation characteristics. The results obtained show that for the near sound field the impulse noise strongly propagates toward the pipe axis, but for the far sound field the impulse noise uniformly propagates toward the omnidirections, indicating that the directivity pattern is almost same regardless of the pipe type. For this non-directivity in the far sound field, it is shown that the perforated pipe has little performance to suppress the impulse noise.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2007.04a
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• pp.81-86
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• 2007

The air velocity flowing in inner combustion chamber of Scramjet is supersonic and the time of its stay is very short as a few milliseconds. Within this short time, fuel injection, air-fuel mixing, and combustion process should be accomplished. Several methods are suggested for mixing enhancement. Among these, cavity is selected to study for enhancement of mixing. The numerical simulation is performed in the case of freestream Mach number of 2.5 and cavity located in front of fuel jet injection. 8 different sized cavities of length-height ratio were used in order to recognize the effect about cavity size. Also, the case without cavity was analyzed to find the effect of cavity. Used code compared with the result of experiment under identical conditions and it was verified. Through this comparison and verification, mixing enhancement by cavity could be confirmed.

• Journal of the Korean Society of Propulsion Engineers
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• v.18 no.3
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• pp.62-69
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• 2014

This paper presents a preliminary study of a convergent divergent nozzle in an afterburning turbofan engine of a supersonic aircraft engine. In order to design a convergent divergent nozzle, cycle model of a low bypass afterburning turbofan engine of which thrust class is 29,000 lbf at a sea level static condition is established. The cycle analysis at the design point is conducted by Gasturb 12 software and one dimensional gas properties at a downstream direction of the turbine are obtained. The dimension and configuration of an model turbofan engine are derived from take-off operation with wet reheat condition. The off-design cycle calculation is conducted at the all flight envelope on the maximum flight Mach number of 2.0 and maximum flight altitude of 15,000 m.

• Journal of the Korean Society of Propulsion Engineers
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• v.22 no.3
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• pp.119-127
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• 2018

A test stand for an altitude test of reciprocating engine was designed, manufactured and validated by preliminary tests and simple calculations. The test stand was designed to interface with the altitude turbo-shaft engine test facility of Korea Aerospace Research Institute. Many limiting conditions for altitude test of reciprocating engine were assumed and the test stand was developed to satisfy these limitations. The test stand design was focused especially on the altitude, Mach number and fuel temperature control for reciprocating engine altitude tests with smaller air and fuel flow than those of turbo shaft engines.

• Journal of the Korean Society of Propulsion Engineers
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• v.22 no.3
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• pp.53-64
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• 2018

In this study, computational simulation was performed to investigate the characteristics of air/fuel mixing according to the shape of the injector exit when the transverse jet was injected into a supersonic flow. Non-reacting flow simulation was conducted with fixed mass flow rate and the same cross-sectional area. To validate the results, free stream Mach number and jet-to-crossflow memetum ratio are set to 3.38 and 1.4, respectively, which is same as the experimental condition. Further, separation region, structure of the under-expended jet, jet penetration height, and flammable region of hydrogen for five different injectors compared.

• Journal of the Korean Society of Propulsion Engineers
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• v.9 no.1
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• pp.35-45
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• 2005

Numerical study is carried out to investigate the effects of chemistry and thermal radiation on the rocket plume flow field at various altitudes. Navier-Stokes equations for compressible flows were solved by a fully-implicit TVD code based on the finite volume method. An infinitely fast chemistry module for hydrocarbon mixture with detailed thermo-chemical properties and a thermal radiation module for optically thick media were incorporated with the fluid dynamics code. The plume flow fields of a kerosene-fueled rocket flying at Mach number zero at sea-level, 1.16 at altitude of 5.06 km and 2.90 at 17.34 km were numerically analyzed. Results showed the plume structures at different altitude conditions with the effects of chemistry and radiation. It is understood that the excess temperature by the chemical reactions in the exhaust gas may not be ignored in the view point of propulsion performance and thermal protection of the rocket base, especially at higher altitude conditions.

• Journal of the Korean Society of Propulsion Engineers
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• v.9 no.3
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• pp.127-132
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• 2005

In order to investigate transient behavior, of the CRW(Canard Rotor Wing) type UAV(Uninhabited Aerial Vehicle) propulsion system during flight mode transition considering flow control valve operation, the propulsion system was modelled using SIMULINK commercial program. The valve system is to control the gas flow of the rotary duct system and the main duct system, and the analysis was performed with an assumption that the total gas mass flow of the main engine is the same as summation of the rotary duct flow and the main duct flow, and with consideration of valve loss, flow rate and effective area in valve angle variation. The performance analysis was carried out during flight mode transitions from the rotary flight mode to the fixed wing flight mode and vice versa mode at altitude of 1km, flight Mach number 0.1 and maximum engine rpm.