Young-Sung Ko

Modeling the Prediction of Helium Mass Requirement for Propellant Tank Pressurization

DOI: 10.2514/1.A32073 A numerical model that predicts the helium mass required for propellant tank pressurization during propellant outflow was developed. The model has the feature of including the effects of the internal hardware of the propellant tanksuchasbaffles,pressurantstoragetanks,andotherinstrumentation.Ituseda finitevolumemethodthatdivides the ullage and tank wall one-dimensionally along the propellant tank axis. A series of cryogenic propellant drainage testswerecarriedouttoverifythedevelopednumericalmodel.Therequiredamountofheliummasspredictedbythe model showed very good agreement with test data within an accuracy of 2:27% under the operating conditions. The developed model was applied to the pressurization system of Korea Sounding Rocket-III, and the results were compared withthe flight-test data.The comparisonresults showedthatthe developed modelwassatisfactory for the prediction of the required helium mass during flight. Additionally, a parametric study was performed to test the sensitivity of the developed model, and the results showed that the heat transfer coefficient between the ullage and the tank wall was the key factor in the accuracy of the model.

Acoustic Tests on Atmospheric Condition in a Liquid Rocket Engine Chamber

Acoustic characteristics of unbaffled and baffled combustion chamber are experimentally investigated under atmospheric condition to preliminarily determine baffle for mitigation of combustion instability. To investigate the effect of the baffle which has several configurations such as radial baffles and hub/blade baffle, resonant-frequency shift and damping factors of the chamber were analyzed and compared quantitatively with those of the unbaffled combustion chamber. From a view of acoustic characteristics, radial baffles with several configurations have not much difference in resonant-frequency shift and damping factor ratio with each other. On the other hand, hub and blade baffle is very effective to suppress the first tangential mode which was found to be the most harmful acoustic mode in KSR(Korean Sounding Rocket)-III engine. But more study on design parameters such as hub size and axial length should be done for complete optimization of hub and blade baffle. The present study based on linear acoustic analysis is expected to be a useful confirming tool to predict acoustic field and design a passive control devices such as baffle and acoustic cavity.

Construction and Validation Test of Turbopump Real-propellant Test Facility

ABSTRACT Liquid rocket engines of KSLV-II employ a turbopump feed system for propellants. A turbopump real-propellant test facility based on liquid oxygen and kerose ne has been constructed for the experimental verification of the turbopump performance using the real media of propellants(i.e., LOX/Kerosene). The verification tests of sub-systems were perfo rmed such as LOX/kerosene feed system and alcohol burner system. Finally, the performance of t he whole system was executed and verified through a sets of validation tests with the developmen t model of the KSLV-II turbopumps. It has been confirmed that the test facility satisfies the operating conditions and time of the turbopump at the design and off-design performance test using real-propellant.초 록 한국형발사체를 위한 액체로켓엔진은 터보펌프 추진제 공급방식의 시스템으로 이루어진다. 이 터보펌프의 실추진제를 사용하는 실험적 성능검증을 위해 액체산소와 케로신을 토대로 한 터보펌프 실매질 시험설비가 구축되어 왔다. 산화제/케로신 공급 시스템과 알코올버너 시스템과 같은 주요 서브시스템에 대한 검증시험이 이루어 졌고, 터보펌프 개발모델을 이용한 인증시험을 통해 전체 시험설비에 대한 성능검증이 이루어 졌다. 설계점 및 탈설계점을 포함한 터보펌프의 모든 운용조건과 운용시간에 대한 실매질 성능검증시험을 본 시험설비를 이용하여 수행할 수 있는 것으로 확인되었다.Key Words:Turbopump(터보펌프), Real-propellant Test Facility(실매질 시험설비), Liquid Oxygen(액체산소), Kerosene(케로신), Alcohol Burner(알코올버너)Received 22 December 2014 / Revised 12 June 2015 / Accepted 19 June 2015

Development Test of Alcohol Burner for Turbopump Real-propellant Test Facility

ABSTRACT A turbopump real-propellant test facility(TPTF) is to verify the performance of a turbopump unit(TPU) based on liquid oxygen and kerosene. One of the most important sub-facilities is a hot-gas generation system which makes the driving force of the TPU with an alcohol burner. The alcohol burner generates the required flow rates and temperature at the facility using high pressure air and ethanol. In the study, the verification tests of the alcohol burner which was manufactured entirely with domestic technology were performed and fabrication technique and operation skill for the burner could be obtained ahead of the construction of the facility. Two burners will be operated simultaneously for the real-propellant test of 75tf class turbopump and satisfy the power requirement from the turbine of the TPU.초 록 터보펌프 실매질 시험설비는 액체산소와 케로신을 토대로 터보펌프 조립체의 성능을 검증하는 시험설비이다. 이 시험설비 중에서 가장 핵심적인 부분은 터보펌프의 구동력을 생성시키는 고온공기 생성시스템이다. 고압공기와 에탄올을 사용하는 알코올버너에 의해 시험시설에서 요구되는 유량과 온도가 생성된다. 본 연구에서는 국내기술로 제작한 알코올버너의 개발시험을 수행하였으며, 터보펌프 실매질 시험설비 구축에 앞서 버너의 제작기술과 운용경험을 확보할 수 있었다. 75톤급 터보펌프 실매질 시험에서는 이 알코올버너가 2기 동시 운용되어 터빈이 요구하는 구동력을 충족시킬 것으로 보인다.Key Words: Turbopump(터보펌프), Real-propellant Test Facility(실매질 시험설비), Alcohol Burner(알코올버너), High Pressure Air(고압공기), Ethanol(에탄올)

A Study on Quantification of Damping Efficiency of Acoustic Cavities by Absorption Coefficient

A Helmholtz resonator as a stabilization device to control high-frequency combustion instabilities in liquid rocket engine is adopted and its damping capacity is verified by linear acoustic analysis and atmospheric acoustic tests. To compare the results of acoustic attenuation effect in accordance with uni-resonator`s geometry, quantitative analyses were made in the cases of various orifice diameters and lengths. Next, in the experiments to compare the results of acoustic attenuation effect by a difference in the number of resonators, damping capacity of harmful resonant frequency was improved by the increase of the number of resonators. On the other hand, attenuation efficiency of the frequency tended rather to lower due to over damping from the point of view of absorption coefficient. As the result, tuning the suitable geometry for the resonator to the resonant frequency is required for the control using the resonator. Also, the design of resonator`s geometry and the choice of its number are important to put up the optimal efficiency in consideration of restriction of its volume.

Effects of CO2 dilution on combustion instabilities in dual premixed flames

There has been a rapid increase in the demand for biogas applications in recent years, and dry low NOx and dry low emission gas turbine combustors are promising platforms for such applications. Combustion instability is the most important drawback in dry low NOx gas turbine combustors and has, therefore, attracted considerable research interest lately. As a fundamental study towards the use of biogas in dry low NOx and dry low emission gas turbine combustors, this article investigates the influence of CO2 in surrogate biogas on combustion instability. Tests were conducted using a dry low NOx type, a dual lean premixed gas turbine combustor. For a dual flame with dual swirl, the pilot fuel mass fraction affects the flame structure, and the flame structure, in turn, determines the temperature distribution in the combustion chamber and the combustion instability. The effects of the pilot fuel mass fraction, which is an important parameter of the combustor, and the CO2 dilution rate, which is a major contributor of biogas combustion, on the combustion characteristics and instability are investigated through dynamic pressure signal and phase-resolved OH* images. Combustion instability decreases for higher CO2 dilution rates, whose effects depend on the pilot fuel mass fraction. The instability reaches its maximum at a pilot fuel mass fraction of 0.3. Tests confirm that combustion instability diminishes with CO2 dilution, as it reduces the perturbation in the heat emission, and the flame speed decreases resulting in a greater flame surface or volume. Further, investigation of the Rayleigh Index, which represents the coupling strength of the heat release fluctuation and the natural frequency, shows that CO2 dilution weakens the coupling strength, resulting in less combustion instability.

Estimation of Heat Transfer Coefficient at the Upper Layer of Cryogenic Propellant

ABSTRACT The temperature of cryogenic propellant in the propellant tan k increases during flight due to heat input from surroundings. The propellant which temperature rises up over the required condition of turbo-pump remains as unusable propellant at the end of flight. In this paper the estimation method of the heat transfer coefficient at the upper layer of cryogenic propellant was presented. The heat transfer mode at the propellant upper layer was considered as conduction. Temperature distributions near propellant surface obtained from heat transfer coefficient were compared with test data to show the possibility of this method.초 록 추진제탱크 내의 극저온 추진제는 발사체의 비행 과정동안 주변으로부터 에너지를 흡수하여 온도가 상승한다. 비행 종료 시점에 있어 터보펌프 입구 요구조건 이상으로 온도가 상승된 추진제는 사용할 수 없는 잔류추진제로 남게 된다. 본 논문에서는 극저온 추진제 상층부의 온도변화를 살펴보기 위하여 추진제 표면 근처에서의 열전달계수를 구해보고자 하였다. 추진제 상층부의 열전달을 전도로 단순화하여 열전달계수를 예측하는 방법을 제시하였다. 이를 통해 얻어진 추진제 상층부의 온도를 시험데이터와 비교하여 열전달계수 예측 방법의 적용 가능성을 확인하였다. Key Words: Cryogenic Propellant(극저온 추진제), Heat Transfer Coefficient(열전달계수), Propellant Tank(추진제탱크), Propellant Surface(추진제 표면)

Experimental study on the Supersonic Jet Noise and Its Prediction

this paper the acoustic signature from a supersonic nozzle is measured and compared to the result of a program developed for a gas turbine noise prediction. In order to measure the jet Mach Number, the pressure and temperature at the settling chamber was measured along with pressures from a pitot-tube placed near the exit. The results are also compared to the ones obtained with a shadow graph technique. Jet noise produced by an imperfectly expanded jet contains shock associated noise, which consist of broadband noise and screech tone noise. For subsonic condition, the directivity is dominant to the downstream direction due to turbulence mixing noise. For supersonic conditions, however, the directivity is dominant toward upstream direction due to shock associated noise. The comparison with a jet exhaust noise prediction code shows good agreement at supersonic conditions but needs to be improved at subsonic speeds.

Characteristics of Dynamic Pressures in Liquid Rocket Thrust Chambers

The objective of this paper is to characterize dynamic pressure traces measured at various experimental conditions of liquid rocket thrust chambers. Stability rating test and self-excited combustion instability results for subscale and full-scale thrust chambers show various aspects of dynamic behavior of the pressure field depending on combustion system hardware and operating conditions. External perturbations facilitated by a shock wave generated from an explosion of solid material are able to alter the stable combusting flow to an unstable one when it has lack of dynamic stability margin. The direct comparison of pressure traces for self-excited and perturbed high frequency instabilities suggests that coupling mechanisms behind these two cases are totally different from each other. Naturally occurring combustion instability seems to be first coupled with a low frequency wave that affects flame more vigorously with the smaller phase difference between heat release rate and acoustics than a high frequency wave. However, for the artificial perturbation, the combusting flow triggered by a shock wave becomes located at physical conditions irrelevant to those before the perturbation.

Performance Test of A Reverse-Annular Type Combustor (TS2) for APU

Development of a small gas-turbine combustor for 100㎾ class APU(Auxiliary Power Unit) has been performed. This combustor is a reverse-annular type and has a tangential swiller in the liner head to improve the fuel/air mixing and flame stability. Three main and three pilot fuel injectors of the simplex pressure-swirl type are used. The performance target at the design condition includes a turbine inlet temperature of l170k, a combustion efficiency of 99%, a pattern factor of 30%, and an engine durability of 3000 hours. Under developing the combustor, we conducted the performance test of our first prototype(TS1) with some variants. As a result of the test, the performance targets of the combustor are satisfied except that the pattern factor is about 4% higher than the target value. Therefore, the second prototype(TS2) was redesigned and the performance test was conducted with the critical focus on the pattern factor and the exit mean temperature. We adopted TS2 four variants to check the improvement of the pattern factor. As a result, the pattern factors of several variants were satisfied with the performance target. Finally, the TS2A variant was chosen as a final combustor fur our APU model.