No Abstract, See Full Text

No Abstract, See Full Text

3-D compressible flow analysis was conducted by using mixing plane method for turbine system which is consisted of partial admission nozzle and rotor. Computational results are shown oblique shock wave in blade leading and trailing edge and also shown flow separation along suction surface of blade due to abrupt blade curvature. But computational results are well agree with 1-D calculation results and experimental data.

In this paper, we investigate the rotordynamic characteristics of turbopump system for 9.5ton thrust liquid rocket engine. A finite element method is used to analyze the vibratior characteristics of a rotor-bearing system. The turbopump rotating system is modeled by shaft with sixty elements, nine rigid disks, four ball bearings and four floating ring seals. The calculation results show that the margin of 1st critical speed is increased from 12% to 68% by use of elastic damping ring. In addition, the margin of the 2nd critical speed near the operating speed is increased from 30% to 63% by the stiffness and damping of floating ring seals.

소형항공기용 IO-540 왕복엔진을 대상으로 추진기관의 성능에 미치는 배압 특성을 조사하기 위해서 지상 성능시험을 수행하였다. 본 시험은 6가지 배기조건에 대해서 엔진의 배압과 흡기압력, 토크, 그리고 프로펠러의 정지추력을 측정하였다. 엔진의 배압은 대체로 머플러를 장착할 경우와 배기관 끝단이 상대적으로 직경이 작고 길이가 길수록 증가하였다. 배압 특성은 특히 엔진의 회전속도가 고속 영역으로 갈수록 프로펠러 추력에 변화를 나타냈으며 전반적으로 추진기관 성능에 미세한 영향을 미치는 것으로 나타났다.

A numerical parametric study is conducted to simulate shock-induced combustion with a variation in freestream conditions. The analysis is limited to inviscid flow and includes chmical nonequilibrium. A steady combustion front is established if the freestream Mach number is above the Chapman-Jouguet speed of the mixture. On the other, an unsteady reaction fi:ont is established if the freestream Mach number is below or at the Chapman-Jouguet speed of the mixture. The three cases have been simulated for Machs 4.18, 5.11, and 6.46 with a projectile diameter of 15 mm. Machs 4.18 and 5.11 shows an unsteady reaction front, whereas Mach 6.46 represents a steady reaction front. Thus Chapman-Jouguet speed is one of deciding factor for the instabilities to trigger. The instabilities of the chemical front with a variation of projectiles diameters will be investigated.

Flame propagation during the initial stages of ignition in a non-premixed swirl, having some of characteristics of the primary zone of an aero gas turbine combustor, has been investigated. Nd:YAG laser was adopted as the principal ignition source to allow arbitrary placing of the ignition site i subsequent flame development was monitored using a natural light high speed filming technique for many ignition site at two different swirl ratios and an overall equivalence ratio of 0.9. For ignition offset from the burner centreline, buoyancy force associated with radial pressure gradient produced a strong inward movement of the flame kernel. At the burner exit. flame kernels invariably developed into cylindrical form and a 'radial confinement /axia expansion' (RCAE) process was observed.

Recently, there have been many research efforts to improve the fuel regression rate in the hybrid rocket engines. In the present study, the ongoing research and development of the next general ion hybrid rocket engine are systematically reviewed. The detailed discussions have been made for the innovative combust ion technologies including the vortex hybrid rocket engines , cryogenic sol id propellant hybrid rocket engine, and the gas generator hybrid rocket engines.

Hybrid propulsion systems provide many advantages in terms of stable operation and safety. However, classical hybrid rocket motors have lower fuel regression rate and combustion efficiency compared to solid propellant rocket motor. The recent research efforts are focused on the improvement of volume limitation and regression rate in the hybrid rocket engine. The present study has numerically investigated the combustion processes in the hybrid rocket engine. The turbulent combustion is represented by the eddy breakup model and Hiroyasu and Nagle and Strickland-Constable model are used for soot formation and soot oxidation. Radiative heat transfer is modeled by finite volume method. To reduce the uncertainties for convective heat transfer near solid fuel surface having strong blowing effect, the Low Reynolds number k-$\varepsilon$ turbulent model is employed. Based on numerical results, the detailed discussion has been made for the turbulent combustion processes in the vortex hybrid rocket engine.

No abstract, See Full-text

In the present work, ultrasonic testing method has been developed to evaluate adhesive layers in rocket motor case assembly for the reliability of the rocket. The main objective of the present work was to find debonding between steel and rubber layers. The relationship between adhesion ratio and reflected ultrasonic amplitude was calculated by considering reflection coefficient at the interface between steel and rubber layers. It was found that the higher amplitude of ultrasound is reflected for the debonding area, and shown good agreements with experimental results. The ultrasonic C-scan images offers good implements for the determination of debonding area. The nondestructive testing results were compared with the micrography of destruective testing. As results, ultrasonic testing could be utilized for the evaluation of adhesive layer in the rocket motor case assembly.