• Journal of computational fluids engineering
• /
• v.6 no.2
• /
• pp.32-39
• /
• 2001

Hypersonic analysis on the KSR-Ⅲ payload configuration has been performed using an axisymmetric Navier-Stokes code. A numerical code based on the Harten and Yee's upwind TVD scheme with simplified curve fits in the chemically reacting equilibrium air was developed. The carbuncle phenomenon on detached shock in front of the payload is controlled by using pressure gradients to tune the dissipation. Chemically reacting equilibrium computations for the reentry flight conditions of Mach No. 10.2, 8, 4.9 are presented and compared with the results of calorically perfect gas.

• 한국전산유체공학회:학술대회논문집
• /
• 1999.05a
• /
• pp.29-34
• /
• 1999

A two-dimensional Navier-Stokes code based on AUSMPW+ scheme has been developed to simulate the hypersonic flowfield of hypersonic shock-shock interaction. AUSMPW+ scheme is a new hybrid flux splitting scheme, which is improved by introducing pressure-based weight functions to eliminate the typical drawbacks of AUSM-type schemes, such as non-monotone pressure solutions. To study the real gas effects, three different gas models are taken into account in this paper: perfect gas, equilibrium flow and nonequilibrium flow. It has been investigated how each gas model influences on the peak surface loading, such as wall pressure and wall heat transfer, and unsteady flowfield structure in the region of shock-shock interaction. With the results, the value of peak pressure is not sensitive to the real gas effects nor to the wall catalyticity. However, the value of peak heat transfer rates is affected by the real gas effects and the wall catalyticity. The structure of the flowfield also changes drastically in the presence of real gas effects.

• International Journal of Aerospace System Engineering
• /
• v.2 no.2
• /
• pp.62-66
• /
• 2015

The starting processes of the Ludwieg tube hypersonic quiet tunnel plays very important role in the achievement of the quiet flow in the test section, which could affect the confidence coefficient of the data in the hypersonic transition experimental investigations. Thus, numerical analysis on that processes could help to understanding the running mode of the Ludwieg tube quiet tunnel and the propagation principle of the expansion wave series. To verify our computational method, the same parameter of the BAM6QT (the Boeing/AFOSR Mach-6 quiet tunnel at Purdue University) is used to compute, and it is agrees with our computational results.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2004.03a
• /
• pp.649-656
• /
• 2004

Hypersonic wind tunnel test of the rectangular variable geometry intake is performed. For realization of a Precooled turbojet engine, development of a hypersonic ramjet engine is planned. To investigate performance of the intake of the hypersonic ramjet engine, wind tunnel test is done with freestream Mach number of 5.1. The total pressure recovery was 18 % with 12.9 % of ramp bleed. Several reasons for low total pressure recovery are shown. Supersonic internal compression is not enough. Then, the throat Mach number is high (M2.61) and total pressure losses at the terminal shock is large. Supersonic flow at the throat and position of the terminal shock is sensitive to a difference of the second ramp's throat height and the third ramp's throat height. Flow separations at the second ramp's trailing edge and the third ramp's leading edge are seen those could result in the trigger of unstart. The seal mechanism between the ramps and the sidewalls is important.

• Communications of the Korean Mathematical Society
• /
• v.10 no.3
• /
• pp.767-781
• /
• 1995

In this paper we present computation of a reflected shock in the hypersonic flow of air with chemical reactions. We consider two dimensional steady inviscid hypersonic flow of air around bodies including chemical reaction effects. At a high Mach number, a strong shock is formed in front of the body when a wedge is placed against the flow. In front of the shock, temperature and pressure increase greatly and the flow is in nonequilibrium state. If the shock hits a wall, then a reflected shock is formed in the nonequilibrium flow region. Behind this reflected shock, the temperature and pressure are very high. We carry out the computation of the reflected shock and the flow behind it. The jump conditions at the reflected shock are presented. A technique combining smooth transforms of domain and implicit difference methods is used to overcome numerical difficulties associated with the lack of resolution behind the shock and near the body.

• 한국전산유체공학회:학술대회논문집
• /
• 2009.04a
• /
• pp.1-8
• /
• 2009

This is a short review of how CFD contributed to hypersonic flights in the past 50 years. Two unexpected phenomena that occurred in the entry flights of the Apollo and Space Shuttle made us aware of the impact of the high temperature real-gas effects on hypersonic flights: pitching moment anomaly of up to 4 degrees, and radiation overshoot behind a shock wave. The so-called two-temperature nonequilibrium model was introduced to explain these phenomena. CFD techniques were developed to accommodate the two-temperature model. Presently, CFD can predict trim angle of attack to an accuracy of about 1 degree. A concerted effort was made to numerically reproduce the experimentally measured flow-field over a double-cone. As yet, perfect agreement between the experimental data and computation is not achieved. Scramjet technology development is disappointingly slow. The phenomenon of ablation during planetary entries is not yet predicted satisfactorily. In the future, one expects to see more research carried out on planetary entries and space tourism.

• Journal of computational fluids engineering
• /
• v.20 no.1
• /
• pp.99-104
• /
• 2015

Aero-optic phenomena cause the image position displacement on an imaging plane of the airborne optical/IR systems. Particularly, the aero-optic boresight error(BSE) is important factor for homing, positioning and aiming applications of hypersonic flight interceptor missile. In this paper, an estimating method of aero-optic BSE for a hypersonic flight vehicle is studied. A ray tracing method and a transform method of refractive index fields from flow density fields are combined with computational fluid dynamics(CFD) method.

• Journal of the Korea Institute of Military Science and Technology
• /
• v.4 no.2
• /
• pp.202-214
• /
• 2001

In this paper the upwind flux difference splitting Navier-Stokes method has been applied to study quasi one-dimensional nozzle flow and axisymmetric sphere-cone($5^{\circ}$) flow for the perfect gas, the equilibrium and the nonequilibrium chemically reacting hypersonic flow. The effective gamma(${ \tilde{\gamma}}$), enthalpy to internal energy ratio was used to couple chemistry with the fluid mechanics for equilibrium chemically reacting air. The influences of the various altitude(30km, 50km) at Mach number(15.0, 20.0) were investigated. The equilibrium shock position was located farthest downstream when compared with those of perfect gas in a quasi one-dimensional nozzle. The equilibrium shock thickness over the blunt body region was much thinner than that of perfect gas shock.

• International Journal of Aeronautical and Space Sciences
• /
• v.1 no.1
• /
• pp.21-28
• /
• 2000

The flowfield of hypersonic shock-shock interaction has been simulated using a two-dimensional Navier-Stokes code based on AUSMPW+ scheme. AUSMPW+ scheme is a new hybrid flux splitting scheme, which is improved by introducing pressure-based weight functions to eliminate the typical drawbacks of AUSM-type schemes, such as non-monotone pressure solutions. To study the real gas effects, three different gas models are taken into account in the present paper: perfect gas, equilibrium flow and non equilibrium flow. It has been investigated how each gas model influences on the peak surface loading, such as wall pressure and wall heat transfer, and unsteady structure of flowfield in the region of shock-shock interaction. With the results, the value of peak pressure is not sensitive to the real gas effects nor to the wall catalyticity. However, the value of peak heat transfer rates is affected by the real gas effects and the wall catalyticity. Also, the structure of the flowfield changes drastically in the presence of real gas effects.

• Journal of the Korean Society for Aeronautical & Space Sciences
• /
• v.36 no.4
• /
• pp.328-336
• /
• 2008

The shock tunnel as a hypersonic ground test facility was designed, constructed and its performance test was conducted to reproduce the high speed flow which the hypersonic propulsion system is encountered. The design points were understood and the conceptual design was completed using the quasi one dimensional operation analysis code. After that, the specific performance and compartment design were completed using CFD simulation as the part analysis. The facility was then constructed according to those design results and the performance test was conducted for various operation conditions. In this paper, we suggested the compartment design method using CFD analysis, construction process and various performance test results in detail.