• Transactions of the Korean Society of Mechanical Engineers B
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• v.36 no.3
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• pp.233-240
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• 2012

An experimental study to improve the impingement characteristics of the assist gas in laser cutting was carried out. For various assist-gas pressures, and locations and installation angles of the nozzle, the characteristics of the impingement of the jet from a supersonic rectangular nozzle were compared to those previously observed for typical circular nozzles. Schlieren flow visualizations and Pitot pressure measurements downstream of the kerf surface were utilized for this purpose. The present rectangular supersonic nozzle decreased the strength of the Mach disc occurring at the corner of the kerf surface, and thus, could weaken the separation of the assist gas on the kerf surface and increase the Pitot pressures downstream compared to conventional circular nozzles.

• Advances in aircraft and spacecraft science
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• v.7 no.4
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• pp.371-385
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• 2020

A forward-facing aerospike attached to a payload fairing of a satellite launch vehicle significantly alters its flowfield and decreases the aerodynamic drag in transonic and low supersonic speeds. The present payload fairing is an axisymmetric configuration and consists of a blunt-nosed body along with a conical section, payload shroud, boat tail and followed by a booster. The main purpose of the present numerical simulations is to evaluate flowfield and assess the performance of aerodynamic drag coefficient with and without aerospike attached to a payload fairing of a typical satellite launch vehicle in freestream Mach number range 0.8 ≤ M_∞ ≤ 3.0 and freestream Reynolds number range 33.35 × 10⁶/m ≤ Re_∞ ≤ 46.75 × 10⁶/m whichincludes the maximum aerodynamic drag and maximum dynamic conditions during ascent flight trajectory of the satellite launch vehicle. A numerical simulation has been carried out to solve time-dependent compressible turbulent axisymmetric Reynolds-averaged Navier-Stokes equations. The closure of the system of equations is achieved using the Baldwin-Lomax turbulence model. The aerodynamic drag reduction mechanism is analysed employing numerical results such as velocity vector plots, density and Mach contours in conjunction with the experimental flow visualization pictures. The variations of wall pressure coefficient over the payload fairing with and without aerospike are exhibiting different kind of flowfield characteristics in the transonic and low supersonic speeds. The numerically computed results are compared with schlieren pictures, oil flow patterns and measured wall pressure distributions and exhibit good agreement between them.

• Journal of the Korea Academia-Industrial cooperation Society
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• v.17 no.11
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• pp.363-367
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• 2016

A test section of a supersonic wind tunnel was designed for the analysis of flow characteristics over a backward-facing step with Mach 1.0 slot injection in a supersonic flow of Mach 2.5. The cavity flow of a high-speed vehicle is very complex at supersonic speed, so it is necessary to do experiments using supersonic wind tunnels to verify numerical analysis methods. The previous 2D symmetrical nozzle was replaced with an asymmetrical nozzle. The inviscid nozzle contour was designed using Method of Characteristics (MOC), and the boundary layer thickness correction was reflected by experimental data from the wind tunnel. The results were compared with a CFD analysis. The PID control system was changed to be based on the change of tank pressure. This improved the control efficiency, and the run times of supersonic flow increased by about 1 second. The flow characteristics over a backward facing step with slot injection were visualized by a Schlieren device. This equipment will be used for an experimental study of the film cooling effectiveness over a cavity with various velocities, mass flows, and temperatures.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2006.11a
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• pp.377-381
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• 2006

This paper describes experimental work to control supersonic jet noise using a mesh screen that is placed at the nozzle exit plane. The mesh screen is a wire-gauze screen that is made of long stainless wires with a very small diameter. The nozzle pressure ratio is varied to obtain the supersonic jets which are operated in a wide range of over-expanded to moderately under-expanded jets. In order to perturb mainly the initial jet shear layer, the hole is perforated in the central part of the mesh screen. The hole size is varied to investigate the noise control effectiveness of the mesh screen. A schlieren optical system is used to visualize the flow fields of supersonic jet with and without the mesh screen device. Acoustic measurement is performed to obtain the OASPL and noise spectra. The results obtained show that the present mesh screen device leads to a substantial suppression of jet screech tones. The hole size is an important factor in reducing the supersonic jet noise. For over-expanded jets, the noise control effectiveness of the mesh screen appears more significant, compared to correctly and under-expanded jets

• 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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• 2004.03a
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• pp.75-78
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• 2004

Various jet engines (Turbine engine family and RAM Jet engine) have been developed for high speed aircrafts. but their application to hypersonic flight is restricted by principle problems such as increase of total pressure loss and thermal stress. Therefore, the development of next generation propulsion system for hypersonic aircraft is a very important subject in the aerospace engineering field, SCRAM Jet engine based on a key technology, Supersonic Combustion. is supposed as the best choice for the hypersonic flight. Since Supersonic Combustion requires both rapid ignition and stable flame holding within supersonic air stream, much attention have to be given on the mixing state between air stream and fuel flow. However. the wider diffusion of fuel is expected with less total pressure loss in the supersonic air stream. So. in this study the direction of fuel injection is inclined 30 degree to downstream and the total pressure of jet is controlled for lower penetration height than thickness of boundary layer. Under these flow configuration both streams, fuel and supersonic air stream, would not mix enough. To spread fuel wider into supersonic air an aerodynamic force, baroclinic torque, is adopted. Baroclinic torque is generated by a spatial misalignment between pressure gradient (shock wave plane) and density gradient (mixing layer). A wedge is installed in downstream of injector orifice to induce an oblique shock. The schlieren optical visualization from side transparent wall and the total pressure measurement at exit cross section of combustor estimate how mixing is enhanced by the incidence of shock wave into supersonic boundary layer composed by fuel and air. In this study non-combustionable helium gas is injected with total pressure 0.66㎫ instead of flammable fuel to clarify mixing process. Mach number 1.8. total pressure O.5㎫, total temperature 288K are set up for supersonic air stream.

• Transactions of the Korean Society of Mechanical Engineers B
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• v.39 no.3
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• pp.237-243
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• 2015

In this study, the effects of the mixing ratio of emulsified fuel on the droplet evaporation and spray behavior characteristics were analyzed. A surfactant comprising span 80 and tween 80 mixed at a 9:1 ratio was used for the emulsified fuel. The fuel and surfactant were mixed at a ratio of 3:1 for the emulsified fuel. In addition, considering the mixing ratio of the surfactant, the mixing ratio of $H_2O_2$ in the emulsified fuel was set as EF (emulsified fuel)0, EF2, EF12, EF22, EF32, and EF42. To observe the evaporation characteristics, droplets of the emulsified fuel were dropped on a heating plate and observed using scattered light and a Schlieren system. In addition, to analyze the effect of the $H_2O_2$ mixing ratio, the behavior characteristics of the evaporative free spray were investigated in the mixing ratio range of EF0 to EF22 using a constant volume chamber with heaters. Consequentially, it was found that in the case of EF22, the free spray development of the emulsified fuel was faster than that of EF0 (diesel only) because of the promotion of the evaporation due to the phase change in the peroxide contained in the emulsion fuel.