• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2010.11a
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• pp.170-176
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• 2010

A numerical analysis was peformed for the supersonic aircraft with variable pitch thrust vector nozzle. Based on the requirement of the mixed turbofan engine of the supersonic aircraft, two dimensional thrust vector nozzle with variable pitch angle was designed. To investigate the effect of the thrust vectoring nozzle, the numerical analysis was conducted by using Fluent under the several pitch deflection angle.

• Journal of the Korean Society of Propulsion Engineers
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• v.26 no.2
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• pp.1-11
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• 2022

The heat transfer coefficient on the wall, which is used as a boundary condition in the thermal analysis of general contract-divergent supersonic nozzles, affects the thermal analysis accuracy of the entire nozzle. Accordingly, many methods of deriving a heat transfer coefficient have been proposed. In this study, the accuracy of each method was compared. For this purpose, the heat transfer coefficients were calculated through theoretical-based analogy methods, semi-empirical equations, and CFD simulations for the previously performed heat transfer experiment with an isothermal wall and compared with the experimental results. The results show that the Prandtl-Taylor analogy methods and the CFD results with the k-ω SST turbulence model were in good agreement with the experimental results. Furthermore, the Modified Bartz empirical formula showed an overall over-prediction tendency.

• Journal of the Korean Society of Propulsion Engineers
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• v.4 no.1
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• pp.46-52
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• 2000

A computational analysis has been made on fluid flow in a regenerative cooling Passage for a reduced size liquid rocket engine to predict pressure drop and heat transfer rate in it. The contraction/expansion of the cross sectional area of the passage turn out to increases both the pressure loss and the heat transfer rate of the duct. The changes of the cross sectional area near the nozzle throat are effective to protect the throat which suffers from severe thermal load. Also given is the qualitative characteristics of the performance of the regenerative cooling system due to the variation of coolant flow rate.

• Journal of the Korean Society of Propulsion Engineers
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• v.4 no.4
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• pp.70-77
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• 2000

Supersonic Dual Bell Nozzle (SDBN) is an altitude-adaptive propulsion nozzle achieved only by a nozzle wall inflection. In order to investigate the altitude adaptive capability and the effectiveness of this nozzle concept, the present study addresses a computational work of the flow through SDBN. Several types of the SDBNs are tested for a wide range of the pressure ratio which covers from an over-expended flow to a fully under-expended flow at the exit of the SDBN. Axisymmetric, compressible, Wavier-Stokes equations are numerically solved using a fully implicit finite volume differencing scheme. The present computational results reveal that the base nozzle length affects the shock wave system occurring inside SDBN. For a quit wide range of the pressure ratio the flow separation occurs at the nozzle inflection point. It is found that the maximum thrust coefficient is obtainable for the correct expansion state at the exit of SDBN.

• Journal of the Korean Society of Propulsion Engineers
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• v.5 no.2
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• pp.51-58
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• 2001

Supersonic, axisymmetric, jets issuing from several kinds of dual, coaxial, nozzles were experimentally investigated. Four different kinds of coaxial, dual nozzles were employed to characterize the major. features of the supersonic, coaxial, dual jets. Two convergent-divergent supersonic nozzles with different impinging angle on the jet axis of were designed to have the Mach number 2.0 and used to compare the coaxial jet flows with those discharging from two sonic nozzles. The primary pressure ratio was changed in the range from 4.0 to 10.0 and the assistant jet ratio from 1.0 to 4.0. The results obtained show that the assistant jets from the annular. nozzle affect the coaxial jet flows and an increase of both the primary jet pressure ratio and assistant jet pressure ratio lead to a longer supersonic length of the dual, coaxial jet.

• Journal of Advanced Marine Engineering and Technology
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• v.24 no.6
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• pp.89-95
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• 2000

To obtain an exact flow loss in piping systems is very important in the face of efficiency anticipation and work control of plant. The object of this study is to get the flow loss through the experiment for sudden expansion and contraction part of circular pipe nozzle. The experiment in this study is performed after getting the flow loss factor for sudden expansion and contraction through preliminary experiments. It is confirmed that the results of this study agreed with the approximated equation of Ikeda and Matsuo. It is proved that flow loss factor ${\zeta}_3$for sudden expansion and contraction part of circular pipe is dependent on $L/D_1$in these experimental conditions.

• Journal of the Korean Society of Propulsion Engineers
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• v.16 no.2
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• pp.42-50
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• 2012

The goal of heat transfer studies is the accurate prediction of temperature and heat flux distribution on material boundaries. To this purpose, general-purpose computational fluid dynamics(CFD) code is used : FLUENT. Mass fluxes and pressure ratio are calculated for two types of nozzle. The comparative studies reveal that the computational results are in agreement with the experimental data. Also, heat transfer coefficients from FLUENT for one type of nozzle are very similar and agree well with the experimental data in the diverging part of the nozzle, but the calculated results are large in the converging part. The heat transfer coefficients from Bartz equation are over-predicted. We can consider various reasons for these differences, i.e., laminarization by the highly accelerated flow in the nozzle, turbulent flow model and grid generation.

• 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.18 no.4
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• pp.1-8
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• 2014

Overexpanded flow of an axisymmetric thruster nozzle is numerically simulated to investigate effects of nozzle pressure ratio (NPR) on the shock structure and thrust performance. The Reynolds-averaged Navier-Stokes equations with k-${\omega}$ SST turbulence model are solved utilizing FLUENT solver. As the NPR is raised, thrust performance monotonically increases with the shock structure and flow-separation point being pushed toward the nozzle exit. It is also discussed that the flow structure at nozzle-exit plane which is immediately affected by a position of nozzle-interior shocks and expansion waves, has strong influence upon the thrust performance of thruster nozzle.

• Journal of the Korean Society of Propulsion Engineers
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• v.17 no.5
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• pp.1-9
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• 2013

Variable area nozzle, where both throat and exit area vary, is required for optimal expansion and optimal nozzle shape upon operation of after-burner. Steady-state and transient analyses are carried out for each condition with and without afterburner operation and as a function of the location of the nozzle flap. Effects of that nozzle displacement on flow and thrust characteristics are analyzed from numerical results. With variable area nozzle adopted, the combustion field is variable in time, leading to periodically variable thrust. For off-design conditions, flow separation shows up due to over expansion at the flap tips and shock wave does in the nozzle due to under expansion. The undesirable phenomena can be solved by control of variable area nozzle.