Seong-Ku Kim

Combustion Dynamics and Stability of a Fuel-Rich Gas Generator

The dynamic characteristics of fuel-rich combustion have been studied using an experimental combustor simulating a gas generator for a liquid rocket engine. The combustor burns liquid oxygen and fuel (Jet A-1) at a mixture ratio of about 0.32 and a chamber pressure ranging from 4.10 to 7.24 MPa, which covers subcritical to supercritical pressures of oxygen. For the investigation of combustion dynamics, pressure fluctuation measurements using piezoelectric sensors have been a major probe throughout the present study. Two different types of injector heads equipped with biliquid swirl coaxial injectors and either a short nozzle or a turbine manifold nozzle have been used in the study. Fuel-rich combustion of both injector heads with the short nozzle revealed pressure pulsation at frequencies of about 128 Hz, which attenuates along with an increase of a chamber pressure. Combustion tests with the turbine manifold nozzle conducted at chamber pressures lower than the oxygen critical pressure showed combustion instabilities at a frequency of 330 Hz, which has been identified as a longitudinal resonant mode by a linear acoustic analysis. The combustion instabilities seem to be induced by inherent pressure fluctuations from the biliquid swirl coaxial injector when the chamber pressure is below the oxygen critical pressure.

Effects of various baffle designs on acoustic characteristics in combustion chamber of liquid rocket engine

Effects of various baffle designs on acoustic characteristics in combustion chamber are numerically investigated by adopting linear acoustic analysis. A hub-blade configuration with five blades is selected as a candidate baffle and five variants of baffles with various specifications are designed depending on baffle height and hub position. As damping parameters, natural-frequency shift and damping factor are considered and the damping capacity of various baffle designs is evaluated. Increase in baffle height results in more damping capacity and the hub position affects appreciably the damping of the first radial resonant mode. Depending on baffle height, two close resonant modes could be overlapped and thereby the damping factor for one resonant mode is increased exceedingly. The present procedure based on acoustic analysis is expected to be a useful tool to predict acoustic field in combustion chamber and to design the passive control devices such as baffle and acoustic resonator.

A Pressure-Based Algorithm for Gaseous Hydrogen/Liquid Oxygen Jet Flame at Supercritical Pressure

Computational tools of turbulent combustion have practical applications for various fields including liquid rocket engines, but some numerical issues are still presented for solving supercritical combustion. In the present study, several of these numerical issues are studied and discussed. Turbulent flow and thermal fields of gaseous hydrogen/cryogenic liquid oxygen flame at supercritical pressure are simulated by a turbulence model. To realize real-fluid combustions, the modified Soave-Redlich-Kwong (SRK) equations of state (EOS) are implemented into the flamelet model with a look-up table as functions of mean and variance of mixture fraction, scalar dissipation rate, enthalpy, and pressure. For supercritical combustion flows, modified forms of the pressure implicit with splitting of operator (PISO) algorithm for solving the pressure-velocity linked equation are introduced. From a comparison of instantaneous temperature distributions for gaseous hydrogen/cryogenic liquid oxygen flame at supercritical pressure, the capability of each method based on the different solution sequence is examined and the effective sequence is explored. The results show that the updated mixture fraction reflected in the pressure correction loop is a critical factor for numerical stability. Also, the relative performance of six convection schemes for supercritical combustion is discussed.

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)

Development of Helmholtz Solver for Thermo-Acoustic Instability within Combustion Devices

In order to effectively predict thermo-acoustic instability within real combustors of rocket engines and gas turbines, in the present study, the Helmholtz equation in conjunction with the time lag hypothesis is discretized by the finite element method on three-dimensional hybrid unstructured mesh. Numerical nonlinearity caused by the combustion response term is linearized by an iterative method, and the large-scale eigenvalue problem is solved by the Arnoldi method available in the ARPACK. As a consequence, the final solution of complex valued eigenfrequency and acoustic pressure field can be interpreted as resonant frequency, growth rate, and modal shape for acoustic modes of interest. The predictive capabilities of the present method have been validated against two academic problems with complex impedance boundary and premixed flame, as well as an ambient acoustic test for liquid rocket combustion chamber with/without baffle.

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.

Fuel-Side Cold-Flow Test and Pressure Drop Analysis on Technology Demonstration Model of 75 ton-class Regeneratively-Cooled Combustion Chamber

․ Jong-Gyu Kim** ․ Byoungjik Lim* ․ Munki Kim** ․ Donghyuk Kang** ․ Seong-Ku Kim* ․ Hwan-Seok Choi*ABSTRACT Fuel-side cold-flow tests were performed on the technology demonstration model of a 75 ton-class liquid rocket engine combustion chamber for the first stage of the Korea space launch vehicle Ⅱ. Pressure drop in the cooling channels of the combustion chamber was measured by changing fuel mass flow rate through a pressure regulating system. Pressure d rop in each segment of the chamber could be obtained and a lot of pressure drop was caused by high flow velocity in the nozzle throat segment. The accuracy of a hydraulic analysis method for calculating a pressure loss in cooling channels could be verified by applying it to the cold-flow test conditions. 초 록 한국형발사체 1단에 사용될 75톤급 액체로켓엔진 연소기의 기술검증시제를 설계, 제작하여 연료 수류시험을 수행하였다. 가압압력을 조절하여 연료 유량을 변경함으로써 주어진 유량에서 발생하는 연소기 재생냉각 채널에서의 압력 손실을 측정하였다. 연소실 각 부에서의 압력 손실을 측정할 수 있었으며, 상당량의 압력 손실이 유속이 강한 연소실 노즐목부에서 발생함을 확인하였다. 주어진 연료 수류시험 조건에서 수력학 해석을 통하여 수력학 해석 방법의 정확도를 검증할 수 있었다. Key Words: Liquid Rocket Engine(액체로켓엔진), Regeneratively-Cooled Combustion Chamber( 재생냉각 연소기), Fuel-Side Cold-Flow Test(연료 수류시험), Pressure Drop Analysis(차압 해석)

Eulerian Particle Flamelet Modeling for Combustion Processes of Bluff-Body Stabilized Methanol-Air Turbulent Nonpremixed Flames

The present study is focused on the development of the RIF (Representative Interactive Flamelet) model which can overcome the shortcomings of conventional approach based on the steady flamelet library. Due to the ability for interactively describing the transient behaviors of local flame structures with CFD solver, the RIF model can effectively account for the detailed mechanisms of NOx formation including thermal NO path, prompt and nitrous NOx formation, and reburning process by hydrocarbon radical without any ad-hoc procedure. The flamelet time of RIFs within a stationary turbulent flame may be thought to be Lagrangian flight time. In context with the RIF approach, this study adopts the Eulerian Particle Flamelet Model (EPFM) with mutiple flamelets which can realistically account for the spatial inhomogeneity of scalar dissipation rate. In order to systematically evaluate the capability of Eulerian particle flamelet model to predict the precise flame structure and NO formation in the multi-dimensional elliptic flames, two methanol bluffbody flames with two different injection velocities are chosen as the validation cases. Numerical results suggest that the present EPFM model has the predicative capability to realistically capture the essential features of flame structure and NOx formation in the bluff-body stabilized flames.

Numerical modeling of combustion processes and pollutant formations in direct-injection diesel engines

The Representative Interactive Flamelet (RIF) concept has been applied to numerically simulate the combustion processes and pollutant formation in the direct injection diesel engine. Due to the ability for interactively describing the transient behaviors of local flame structures with CFD solver, the RIF concept has the capabilities to predict the auto-ignition and subsequent flame propagation in the diesel engine combustion chamber as well as to effectively account for the detailed mechanisms of soot formation, NOX formation including thermal NO path, prompt and nitrous NOX formation, and reburning process. Special emphasis is given to the turbulent combustion model which properly accounts for vaporization effects on the mixture fraction fluctuations and the pdf model. The results of numerical modeling using the RIF concept are compared with experimental data and with numerical results of the commonly applied procedure which the low-temperature and high-temperature oxidation processes are represented by the Shell ignition model and the eddy dissipation model, respectively. Numerical results indicate that the RIF approach including the vaporization effect on turbulent spray combustion process successfully predicts the ignition delay time and location as well as the pollutant formation.

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.