Miok Joh

A Trade-off Analysis between Combustion and Cooling Performance of a Liquid Rocket Combustor with Fuel Film Cooling Scheme

․ Seong-Ku Kim* ․ Hwan-Seok Choi*ABSTRACT Performance of a liquid rocket thrust chamber with regenerati ve cooling scheme has been numerically analyzed using in-house CFD code which can predict combustion/cooling performance and provide nozzle design parameters. This paper investigates trade -offs between combustion and cooling performance with varying amount of fuel directly injected into the chamber wall to form cooling films and mixture ratios for the peripheral injectors. Further efforts to verify/improve the simulation methodology including comparison with the firing test results a re planned to make it a reliable tool to optimize the film cooling and other major design parameters. 초 록 액체로켓 추력실의 성능 예측 및 초음속 노즐부 형상 설계에 활용 중인 in-house 해석 도구를 이용하여 재생냉각 연소기에 대한 성능/냉각 통합해석을 수행하였으며, 막냉각 유량 및 외곽 분사기열의 혼합비 변화에 따른 연소 성능과 냉각 성능 간 trade-off 경향을 고찰하였다. 향후 막냉각 및 주요 설계인자의 최적화 도구로 활용될 수 있도록 개발 연소기에 대한 시험 결과와의 비교 등을 통하여 수치해석 도구를 검증/개선해나갈 계획이다. Key Words: Liquid Rocket Combustor(액체로켓 연소기), Regenerative Cooling(재생냉각), Fuel Film Cooling(연료 막냉각), Combustion/Cooling Performance(연소/냉각 성능)

Effective Modeling of Conjugate Heat Transfer and Hydraulics for the Regenerative Cooling Design of Kerosene Rocket Engines

The present work is motivated to develop a unified framework to simulate multi-physical processes which are crucial for trade-off design of liquid rocket thrust chambers among propulsive performance, regenerative cooling, and pressure budget. In this paper, an effective modeling of conjugate heat transfer and hydraulics through the regenerative cooling passage has been performed to quantitatively evaluate detailed cooling designs, including spirally twisted channels and bidirectionally branched circuit, as well as to provide the wall heat flux to a compressible reacting flow solver in an interactively coupled manner. The kerosene fuel used as coolant is modeled by a three-component physical surrogate, and the fluid properties required for calculation of a Nusselt number correlation and empirical resistance coefficients are computed over the entire thermodynamic states from compressed liquid to supercritical fluid using the NIST SUPERTRAPP. The present method has been applied to an actual regeneratively cooled thrust chamber and validated against measurement of hot-firing tests in terms of temperature increase and pressure drop of the coolant through the cooling passages. Based on the numerical results, supplementary effects of peripheral fuel cooling injection and thermal barrier coating are addressed.

Flow Analysis of the Oxidizer Manifold for a Liquid Rocket Combustor using OpenFOAM

Flow in the oxidizer manifold of a liquid rocket combustor has been analysed using an open source CFD toolbox, OpenFOAM. The applicability of OpenFOAM to the problems with complex geometries involving porous media zones for simulating the pressure drop induced by the injectors has been evaluated by performing turbulent, incompressible steady-state flow analysis. The usefulness and applicable area of the OpenFOAM as a design evaluation and analysis tool will be confirmed and enlarged by further evaluation with various computational cases representing major physical phenomena in rocket combustion devices.

Effect of Thermal Barrier Coating and Film Cooling Condition on the Cooling Performance of Liquid-propellant Rocket Engine Combustor

The effect of ceramic thermal barrier coating thickness on the cooling performance of a liquid-propellant rocket engine combustor has been investigated through combustion/cooling performance analysis whose results verified against measured data from hot-firing tests. Also have been confirmed the effects of film cooling amount near the face plate on the coolant temperature and on the thermal barrier coating surface temperature. Some important points to be considered for designing cooling schemes for regeneratively cooled rocket engine combustor have been drawn and reviewed from present study and further verification of the analysis tool should be performed in the future.

Estimation of Thermodynamic/Transport Properties of Kerosene using a 3-Species Surrogate Mixture

Kerosene(Jet A-1), one of the propellants for each stage`s engine of the Korea Space Launch Vehicle-II(KSLV-II), functions as coolant at the same time as it flows inside the cooling jacket of the combustion chambers and is injected through the film cooling holes. A physical surrogate mixture model to reproduce the thermophysical characteristics of Jet A-1 has been selected and the thermodynamic/transport properties of the model fuel under high pressure including supercritical conditions have been estimated using SUPERTRAPP(NIST SRD4). Comparisons with the measured properties suggest that proposed database can be used to extract properties of Jet A-1 for conjugate heat transfer analysis of liquid propellant rocket engine thrust chambers. Predicted combustion/cooling performance of regeneratively cooled thrust chambers shall be validated through comparisons with upcoming firing test results.

Neural-Network Midcourse Guidance with Consideration of the Head-on Attack Condition

The objective of this study is to enhance neural-network guidance to consider the impact condition. Missile impact angle error, a measure of the degree to which the missile is not steering for a head-on attack, can have a significant influence on the final miss distance. Midcourse guidance using neural networks is employed to reduce the deviation angle from head-on effectively in the three-dimensional space. In addition, a coordinate transformation is introduced to simplify the three-dimensional guidance law and reduce training data for the neural network. Computational results show that the current neural-network guidance law with the coordinate transformation can be used to reduce the impact angle errors.

Real-time midcourse guidance with consideration of the impact condition

The objective of this study is to enhance neural-network guidance to consider the impact condition. The optimal impact condition in this study is defined as an head-on attack. Missile impact-angle error, which is a measure of the degree to which the missile is not steering for a head-on attack, can also have an influence on the final miss distance. Therefore midcourse guidance is used to navigate the missile, reducing the deviation angle from head on, given some constraints on the missile g performance. A coordinate transformation is introduced to simplify the three-dimensional guidance law and, consequently, to reduce training data. Computer simulation results show that the neural-network guidance law with the coordinate transformation reduces impact-angle errors effectively.