• Journal of Korean Tunnelling and Underground Space Association
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• v.23 no.1
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• pp.25-36
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• 2021

Hydrogen in hydrogen-electric vehicles has a wide range of combustion and explosion ranges, and is a combustible gas with a very fast flame propagation speed, so it has the risk of leakage, diffusion, ignition, and explosion. The fuel tank has a Thermally active Pressure Relief Device (TPRD) to reduce the risk of explosion and other explosions, and in the event of an accident, hydrogen inside the tank is released outside before an explosion or fire occurs. However, if an accident occurs in a semi-closed space such as an underground parking lot, the flow of air flow is smaller than the open space, which can cause the concentration of hydrogen gas emitted from the TPRD to accumulate above the explosion limit. Therefore, in this study, the leakage rate and concentration of hydrogen over time were analyzed according to the diameter of the nozzle of the TPRD. The diameter of the nozzle was considered to be 1 mm, 2.5 mm and 5 mm, and ccording to the diameter of the nozzle, the concentration of hydrogen in the underground parking lot increases in a faster time with the diameter of the nozzle, and the maximum value is also analyzed to be larger with the diameter of the nozzle. In underground parking lots where air currents are stagnant, hydrogen concentrations above LFL (Lowe Flammability Limit) were analyzed to be distributed around the nozzle, and it was analyzed that they did not exceed UFL (Upper Flammability Limit).

• 한국전산유체공학회:학술대회논문집
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• 2000.10a
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• pp.54-59
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• 2000

A numerical study of axisymmetric rocket main nozzle flow has been accomplished. The CSCM upwind flux difference splitting method with an iterative time marching scheme having second order accuracy in time and space has been used to simulate unsteady flow characteristics in an axisymmetric rocket main nozzle. Though the pressure vary at nozzle inlet with the lapse of time, Mach No. and the density were not changed significontly compared with the temperature. Specific heat ratio $\gamma$=1.134 predicted higher temperature at nozzle throat and exit and nondimensional thrust coefficients at exit than specific heat ratio $\gamma$=1.4 did.

• Advances in aircraft and spacecraft science
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• v.2 no.2
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• pp.183-198
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• 2015

Scramjets are a class of hypersonic airbreathing engine that are associated with realizing the technology required for economical, reliable access-to-space and high-speed atmospheric transport. After-burning augments the thrust produced by the scramjet nozzle and creates a more robust nozzle design. This paper presents a numerical study of three parameters and the effect that they have on thrust augmentation. These parameters include the injection pressure, injection angle and streamwise injection position. It is shown that significant levels of thrust augmentation are produced based upon contributions from increased pressure, mass flow and energy in the nozzle. Further understanding of the phenomenon by which thrust augmentation is being produced is provided in the form of a force contribution breakdown, analysis of the nozzle flowfields and finally the analysis of the surface pressure and shear stress distributions acting upon the nozzle wall.

• Proceedings of the KSME Conference
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• 2003.11a
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• pp.592-597
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• 2003

Thrust vector control using a coflow-counterflow concept is achieved by suction and blowing through a slot adjacent to a primary jet which is shrouded by a suction collar. In the present study, the flow characteristics of thrust vectoring is investigated using a numerical method. The nozzle has a design Mach number of 2.0, and the operation pressure ratio is varied to obtain various flow features of the nozzle flow. Test conditions are in the range of the nozzle pressure ratio from 6.0 to 10.0, and a suction pressure from 90kPa to 35kPa. Two-dimensional, compressible Navier-Stokes computations are conducted with RNG ${\kappa}-{\varepsilon}$ turbulence model. The computational results provide an understanding of the detailed physics of the thrust vectoring process. It is found that an increase in the nozzle pressure ratio leads to increased thrust efficiency but reduces the thrust vector angle.

• International Journal of Aeronautical and Space Sciences
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• v.18 no.1
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• pp.82-90
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• 2017

The aim of this work is to develop a new computational program to determine the effect of using the gas of propulsion of combustion chamber at high temperature on the shape of the two-dimensional Minimum Length Nozzle giving a uniform and parallel flow at the exit section using the method of characteristics. The selected gases are $H_2$, $O_2$, $N_2$, CO, $CO_2$, $H_2O$, $NH_3$, $CH_4$ and air. All design parameters depend on the stagnation temperature, the exit Mach number and the used gas. The specific heat at constant pressure varies with the temperature and the selected gas. The gas is still considered as perfect. It is calorically imperfect and thermally perfect below the threshold of dissociation of molecules. A error calculation between the parameters of different gases with air is done in this case for purposes of comparison. Endless forms of nozzles may be found based on the choise of $T_0$, $M_E$ and the selected gas. For nozzles delivering same exit Mach number with the same stagnation temperature, we can choose the right gas for aerospace manufacturing rockets, missiles and supersonic aircraft and for supersonic blowers as needed in settings conception.

• 한국전산유체공학회:학술대회논문집
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• 2005.10a
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• pp.293-298
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• 2005

We calculate the coordinates of an axisymmetric nozzle with a central body. This nozzle ensures a transonic flow with a plane sound surface, which is orthogonal to the symmetry axis and has a wall kink at the sonic point, The Chaplygin transformation in the subsonic part of the flow leads the Dirichlet problem for a system of nonlinear equations. The definition domain of the solution in the velocity-hodograph plane is taken as a rectangle. This enables one to obtain the nozzle with a monotonic distribution of velocity along its subsonic part. In the nonlinear differential equation, the linear Chaplygin operator for plane flows is separated, which allows the iterative calculation of the solution. The supersonic part of the nozzle is calculated under the assumption that the flow at the nozzle exit is uniform and parallel to the symmetry axis; i.e., the supersonic jet outflows to the submerged space with the same pressure. The calculation is performed by the characteristic method. The exact solution of Tricomi equation for near-sonic flows with the straight sonic line is used to 'move away' the sound plane. The velocity distribution alone the supersonic part of the nozzle is also monotonic, which ensures the absence of the boundary-layer separation and, therefore, the adequacy of the ideal-gas model. calculations show that the flow in the supersonic part of the nozzle is continuous (compression shocks are absent)

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.11a
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• pp.919-921
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• 2011

Variable nozzles are used to enhance the effectiveness of aircraft engines at various altitudes. Unsynchronized movements of variable nozzle flaps affect the direction of thrust in case the variable nozzle consists of many flaps. A synchronization test system was developed to verify the synchronization characteristics of variable nozzle mechanism including flaps. The test system has a capability to simulate combustion pressure in variable nozzle space. The test system was used to qualify the synchronization characteristics of a variable nozzle flaps affected by magnitude and uniformity of simulated combustion pressure, and time delay of each nozzle actuators.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.43 no.4
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• pp.304-310
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• 2015

Pintle thrusters use pintle stroke to change nozzle throat area, and this controls thrust. Using MATLAB, one-dimensional simulation has been investigated and the results are compared to those of cold flow tests and computational fluid dynamics for the pintle thruster of Chungnam National University. The prediction based on one-dimensional flow theory shows good agreement with measurements for chamber pressure, but deviates for thrust, partly because of nozzle wall separation. Computational results show that nozzle wall separation occurs at an early stage of nozzle expansion, near the design nozzle throat, for the course of pintle strokes. Empirical thrust prediction incorporates nozzle wall separation, and thus 1-D simulation using empirical thrust prediction showed good results for an early stage of pintle stroke.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.48 no.3
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• pp.207-215
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• 2020

Design of a shape transition nozzle is carried out as a part of building a lab-scale supersonic combustion experimental equipment. In order to connect directly the circular shaped vitiation air heater to the square shaped scramjet combustor, area change is evaluated by using the method of characteristics. Shape transition function is introduced to control the transition rate. Boundary layer correction was made through the three-dimensional computational fluid dynamics with the assessment on the several shape transition functions. The shape transition nozzle is proved minimizing the growth of boundary layer at the center of the rectangular nozzle surfaces that caused by the pressure gradient at the corners of the rectangular nozzle and the following recirculation regions.

• Proceedings of the KSME Conference
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• 2003.11a
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• pp.508-513
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• 2003

Experiments are performed to investigate the detailed structure of underexpanded twin jet impinging on a perpendicular flat plate. The major parameters, such as nozzle operating pressure and nozzle spacing, are varied to create different jet flow fields resulted from the complicated interactions of the twin jets. From the surface pressure measurements and shadowgraphs taken by schlieren optical system, the jet structure is strongly dependent on the nozzle operation pressure and the spacing. The results obtained show that the closer nozzle spacing may induce to decrease the diameter of the Mach disk within the first shock cell in the underexpanded twin jet. With the increasing nozzle operating pressure and decreasing the nozzle spacing, a new shock wave appears at the entrainment region between the two jets, due to the enhancement of mixing effect of the both jets. The closer nozzle spacing makes the overall impinging pressure level higher, while severe pressure oscillation along the axis of symmetry. Furthermore it is recommended the wider spacing to obtain higher thrust under the present experimental conditions.