• International Journal of Aeronautical and Space Sciences
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• 제18권3호
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• pp.425-435
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• 2017

In this work, an overview of flow diagnosis in a shock tunnel is made by means of using established techniques that are easy to setup, economical to arrange, and simple to measure. One flow condition was considered having Mach number of 6 at the nozzle-exit, regarded as freestream. Measured aerothermodynamic data such as shock wave speed, wall static and total pressures, surface heat flux, and shock stand-off distance ahead of test model showed good agreement with calculation. This study shows an overall procedure of flow diagnosis in a shock tunnel in a single manuscript. Outcomes are thought to be useful in the field of education and also in a preliminary stage of high-speed vehicle design and tests, that need to be performed within a short time with decent accuracy.

• International Journal of Aeronautical and Space Sciences
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• 제13권4호
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• pp.507-514
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• 2012

The present study is the analyses of what has been attempted and what was understood in terms of improving the regression rate and enlarging the basic understanding of internal flow dynamics. The first part is mainly intended to assess the role of helical grain configuration in the regression rate inside the hybrid rocket motor. To improve the regression rate, a combination of swirl (which is an active method) and helical grain (which is a passive method) was adopted. The second part is devoted to the internal flow dynamics of hybrid rocket combustion. A large eddy simulation was also performed with an objective of understanding the origin of isolated surface roughness patterns seen in several recent experiments. Several turbulent statistics and correlations indicate that the wall injection drastically changes the characteristics of the near-wall turbulence. Contours of instantaneous streamwise velocity in the plane close to the wall clearly show that the structural feature has been significantly altered by the application of wall injection, which is reminiscent of the isolated roughness patterns found in several experiments.

• International Journal of Aeronautical and Space Sciences
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• 제9권2호
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• pp.98-106
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• 2008

The motive gas of a supersonic ejector is supplied from different sources depending on the application. The performance of an ejector that is represented by the secondary flow pressure, starting and unstarting pressures heavily depends on the molecular properties of the motive gas. The effects of specific heat ratio of the motive gas were investigated experimentally for an axi-symmetric annular injection type supersonic ejector. Both the starting pressure and unstarting pressure, however, decreased with the increase of the specific heat ratio of the motive gas. It was discovered that the secondary flow pressure increased as the specific heat ratio of the motive gas decreased even if the stagnation pressure of the motive flow was invariant. However, when the motive gas flow nozzle area ratio is large enough for the motive gas to be condensed, different tendency was observed.

• 한국자동차공학회논문집
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• 제9권1호
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• pp.84-93
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• 2001

The aim of this study is to provide fundamental informations that make it possible to use a cool stream and a hot stream simultaneously. We changed the pressure of compressed air that flows into a tube, the inner diameter of orifice that a cold stream exits, and the mass flow rate ratio. And in each case, we measured the temperature of a cold stream and a hot stream in each exit of a tube. Also we measured the axial temperature distribution and the radial temperature distribution in internal space of a tube. From the study, following conclusive remarks can be made. Average flow rate that flows into a tube is in proportion to square root of inlet pressure. As inlet pressure increases axial and radial temperature distribution in the inner space of vortex-tube increase. As mass flow rate ratio change, separation point moves.

• International Journal of Aeronautical and Space Sciences
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• 제1권2호
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• pp.9-16
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• 2000

For aerodynamic design of missile bodies of non-circular cross-section, the combination of the slender body theory and the cross-flow analogy can hardly be applied owing to the lack of experimental data. An alternative is to utilize the Euler solution in the design stage. For enhanced accuracy, however, an adequate viscous correction is necessary to the Euler solution. In this work, such a procedure is examined to compensate the viscous effect by utilizing the concept of proportionality factor in cross-flow analogy. Predictions of aerodynamic coefficients combining the Euler solution and the viscous correction via proportionality factor are made for a missile body of elliptic cross-section. Results indicate that the present approach can be adopted in designing missile bodies of non-circular cross-sections.

• International Journal of Aeronautical and Space Sciences
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• 제8권2호
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• pp.1-10
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• 2007

A two-dimensional backward facing step flow was comptuted using a Detached Eddy simulation(DES) based on the SST turbulence model. The expansion ratio(ER) was 1.125 and the Reynolds number based on the step height and the mean velocity in the upstream channel was 37,500. The flow condition was the same as with the experimental research[1]. The reattachment length, oscillatory characteristics of the flow and the coherent structures of the present simulation were compared to demonstrate the improtance of spanwise grid spacing.

• International Journal of Aeronautical and Space Sciences
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• 제3권2호
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• pp.67-75
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• 2002

In the course of missile system design, jet plume impingement is encountered in designing airframe as well as launchers, requiring careful investigation of its effect on the system. In the present paper, recent works on such topic are presented to demonstrate usefulness of CFD results in helping design the hardware. The jet impinging flow structure exhibits such complex nature as shock shell, plate shock and Mach disk depending on the flow parameters. The main parameters are the ratio of the jet pressure to the ambient pressure and the distance between the nozzle and the wall. In the current application, the nozzle contour and the pressure ratio are held fixed, but the jet impinging distance is varied to illuminate the characteristics of the jet plume with the distance. The same methodology is then applied to a complex vertical launcher system (VLS), capturing its flow structure and major design parameter. These applications involving jets are thus hoped to demonstrate the usefulness and value of CFD in designing a complex structure in the real engineering environment.

• 한국자동차공학회논문집
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• 제8권3호
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• pp.46-55
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• 2000

Four types of different flow inlet models were tested to improve the flow uniformity at the inlet of the radiator and to prevent thermal damages of auxiliary units from the hot air in the bus engine room. Measurements and numerical calculations were performed and their results were in a good agreement with each other. Simultaneously temperature measurements were carried out under the conditions of actual bus driving. As designing the new flow inlet at the partition board which seperates the engine space and radiator space, flow circulation can be achieved and fresh air comes into the engine room from the bottom. It was proved that new inlet makes the one air temperature cooling down in the engine room, the other uniformity improvement.

• 한국해양공학회지
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• 제25권5호
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• pp.9-14
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• 2011

The divergence-free finite elements introduced in this paper are derived from Hermite functions, which interpolate stream functions. Velocity bases are derived from the curl of the Hermite functions. These velocity basis functions constitute a solenoidal function space, and the gradient of the Hermite functions constitute an irrotational function space. The incompressible Navier-Stokes equation is orthogonally decomposed into its solenoidal and irrotational parts, and the decoupled Navier-Stokes equations are then projected onto their corresponding spaces to form appropriate variational formulations. The degrees of the Hermite functions we introduce in this paper are bi-cubis, quartic, and quintic. To verify the accuracy and convergence of the present method, three well-known benchmark problems are chosen. These are lid-driven cavity flow, flow over a backward facing step, and buoyancy-driven flow within a square enclosure. The numerical results show good agreement with the previously published results in all cases.

• International Journal of Aeronautical and Space Sciences
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• 제17권3호
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• pp.315-323
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• 2016

This paper deals with experimental investigation of active flow control via synthetic jets using an unmanned combat air vehicle (UCAV) planform. Fourteen arrays of synthetic jets, mounted along both leading edges, were fully or partially activated to increase aerodynamic efficiency and reduce pitch-up moment. The measurements were carried out using a six-component external balance, a pressure scanner, and tuft flow visualization. It was observed that aerodynamic efficiency (L/D) and pitching moment were clearly affected by the location of jets. In particular, inboard and outboard actuation could effectively increase L/D. Moreover, inboard actuation showed a reduction in the pitch-up, even more than that generated by the full actuation. These results suggest that inboard actuation not only effectively increases L/D but also reduces the pitch-up using only a few actuators.