Sangyeop Han

Numerical simulation of supersonic jet instability modes of plane and notched rectangular nozzles

The effect of exit geometry of a rectangular nozzle on the instability modes and mixing characteristics of under-expanded supersonic jets is computationally investigated. The unsteady three-dimensional viscous simulation is based on the Proteus code developed at NASA Glenn Research Center. A shock adaptive grid generator was developed to enhance the shock simulation. The nozzle aspect ratio for both plane and notched rectangular nozzles in this study is 5.0 and the fully expanded jet Mach number is 1.526. For the case of a plane rectangular nozzle, the ‘flapping’ oscillations, which are extensively observed in schlieren flow visualization and reported in acoustic measurements, are also captured in the presented computations. The flapping frequency in experimental measurements (7400 Hz) is closely predicted in the presented computations (7500 Hz). The symmetrical instability mode is also observed as viewed from the nozzle’s small and large dimensions at twice the flapping frequency. For the notched rectangular nozzle, the flapping oscillations ceased and instead a spanwise oscillation mode was observed as viewed from the nozzle’s large dimension. The instantaneous mass flowrate at nine jet widths downstream of the nozzle exit showed an increase of 8.5 per cent in mass flowrate in the jet emerging from the notched as compared to the plane rectangular nozzle.

Passive Control of Supersonic Rectangular Jets through Boundary Layer Swirl

Abstract Mixing characteristics of under-expanded supersonic jets emerging from plane and notched rectangular nozzles are computationally studied using nozzle exit boundary layer swirl as a mean of passive flow control. The coupling of the rectangular jet instability modes, such as flapping, and the swirl is investigated. A three-dimensional unsteady Reynolds-Averaged Navier-Stokes (RANS) code with shock adaptive grids is utilized. For plane rectangular nozzle with boundary layer swirl, the flapping and spanwise oscillations are captured in the jet’s small and large dimensions at twice the frequencies of the nozzles without swirl. A symmetrical oscillatory mode is also observed in the jet with double the frequency of spanwise oscillation mode. For the notched rectangular nozzle with boundary layer swirl, the flapping oscillation in the small jet dimension and the spanwise oscillation in the large jet dimension are observed at the same frequency as those without boundary layer swirl. The mass flow rates in jets at 11 and 8 nozzle heights downstream of the nozzles increased by nearly 25% and 41% for the plane and notched rectangular nozzles respectively, due to swirl. The axial gross thrust penalty due to induced swirl was 5.1% for the plane and 4.9% for the notched rectangular nozzle.

Flow Visualization and Calculation at the Outlet of Propellant Tank Pressurizing Gas Injector

Propellant tank pressurizing gas injector is used in the pressurization system of liquid propellant rocket to reduce incoming gas velocity and distribute the gas in the tank. Temperature distribution in the propellant tank ullage is varied according to the gas injector shape, and it has influence on the required pressurant gas and thermal phenomena in the tank. In this paper, diffuser type gas injector was studied to make the ullage have stratified temperature distribution. Injected gas flow at the outlet of prototype diffuser was visulized using particle image velocimetry method and it was compared with the results of calculation. Calculation was well agreed with measurement and was used as an inlet condition of propellant tank ullage calculation.

The Past and Future Perspectives of Hydrogen Peroxide as Rocket Propellants

In the field of rocket propulsion system hydrogen peroxide has been used as mono-propellant and as the oxidizer of bi-propellants. At the beginning, hydrogen peroxide was used as mono-propellant for thrusters, but later it had been replaced by hydrazine, which has better specific impulse and storability. On the other hand, to drive turbo-pumps, hydrogen peroxide is still being utilized. As the oxidizer of bi-propellants it was used until 1970`s and from 1990`s hydrogen peroxide once again got back to developer`s interest, because one of the recent development purposes of rocket propulsion system is low-cost and ecologically-clean. Until now the storability of hydrogen peroxide has been remarkably improved. The combination of Kerosene/ also shows similar accelerating performance to Kerosene/ combination because of higher propellant density and higher O/F ratio, even though the propulsion performance is not as good as the combination of Kerosene/. Moreover, its combustion products are much cleaner than Kerosene/ combination.

Effect of boundary layer swirl on supersonic jet instabilities and thrust

This paper reports the effects of nozzle exit boundary layer swirl on the instability modes of underexpanded supersonic jets emerging from plane rectangular nozzles. The effects of boundary layer swirl at the nozzle exit on thrust and mixing of supersonic rectangular jets are also considered. The previous study was performed with a 30° boundary layer swirl (S=0.41) in a plane rectangular nozzle exit. At this study, a 45° boundary layer swirl (S=1.0) is applied in a plane rectangular nozzle exit. A three-dimensional unsteady compressible Reynolds-Averaged Navier-Stokes code with Baldwin-Lomax and Chien’sk-ε two-equation turbulence models was used for numerical simulation. A shock adaptive grid system was applied to enhance shock resolution. The nozzle aspect ratio used in this study was 5.0, and the fully-expanded jet Mach number was 1.526. The “flapping” and “pumping” oscillations were observed in the jet’s small dimension at frequencies of about 3,900Hz and 7,800Hz, respectively. In the jefs large dimension, “spanwise” oscillations at the same frequency as the small dimension’s “flapping“ oscillations were captured. As reported before with a 30° nozzle exit boundary layer swirl, the induction of 45° swirl to the nozzle exit boundary layer also strongly enhances jet mixing with the reduction of thrust by 10%.

Design of Space Launch Vehicle Solenoid Valve for Cryogenic Environment

Solenoid valves for space launch vehicles require the strict limitations on the size, weight and current consumption comparing to industrial solenoid valves. The preliminary design of a cryogenic and high pressure solenoid valve for propellant tank pressurization which can ensure the operation of solenoid valve under such strict limitation conditions was preformed. The Copper and Constantan materials in coil design have used to prevent the excessive rise of the current at cryogenic state. The measured current of solenoid valve at cryogenic temperature satisfies a design requirement.

Study on Leakage Measurements of Oxygen and Helium Using Standard Gas Flow Rates in a Orifice Flow

In this study, correlation equations were arranged about mass flow rates of oxygen and volume flow rates of helium using a mouthpiece method. The mouthpiece method can reduce examination cost by using similar empirical formula. Instead of liquid oxygen, in the mouthpiece method, gas helium can be measured in order to determine the leakage amount of liquid oxygen conveniently. Experiment was conducted and compared to understand leakage amount relation between the helium and the oxygen for prototype item under a room and a cryogenic temperature conditions. The leakage volume flow rate [A.ml/s] of the helium was 174 times higher than mass flow rate [g/s] of the oxygen leakage at liquid state. The derived correlation equations were verified using data from the National Institute of Standards and Technology (NIST).

Analysis of Dynamic Characteristics and Performance of Solenoid Valve for Pressurization Propellant Tank

․ Byunghun Kim* ․ Sangyeop Han*ABSTRACT A 2-way solenoid valve regulates to maintain the pressure of ullage volume of propellant tanks when the command is given by control system for the liquid-prop ellant feeding system of space launch vehicle. The simulation model of solenoid valve for pres surization is designed with AMESim to verify the designs and evaluate the dynamic characteristics and pneumatic behaviors of valve. To improve the accuracy of the model, numerical flow analysis by u sing FLUNET code. The simulation results of their operating durations of valve by AMESim analysi s are matched up with the results of experiments and validate valve model. Using the model, we analy ze performance of valve; opening/closing pressure, operating time on various design factors of basic valve and control valve; geometrical size of valve seat, ratio of basic valve and sealin g area.초 록 우주발사체 추진기관 공급계에서 2-way 솔레노이드밸브는 제어시스템의 명령에 의해 추진제 탱크를 가압하여 탱크내의 압력을 조절한다. 가압용 솔레노이드밸브의 제작에 앞서 설계검증 및 기본적인 작동특성을 분석하기 위해 AMESim상용코드를 이용하여 해석모델을 수립하였다. FLUENT 상용코드를 이용하여 내부유동해석을 수행하여해 해석의 정확도를 높이고, 모델을 검증하기 위해 동특성 해석을 통해 입구압력에 따른 작동시간을 시험결과와 비교하였고 잘 일치함을 확인하였다. 또한 해석 모델을 이용하여 컨트롤밸브와 기본밸브의 설계변수에 대한 밸브의 동특성 해석을 수행하였다. 해석을 통해 기본밸브의 시트형상과 직경비에 따른 밸브의 작동시간, 작동성능, 개폐압력을 예상하였다.Key Words: AMESim(아메심 상용코드), Dynamic Characteristics(동특성), Solenoid Valve(솔레노이드밸브), Pneumatic Valve(공압밸브), Propellant Tank(추진제탱크)

Optimum Performance Analysis of KSR-III LRE

To understand the each performance parameter correlation of flight type liquid-propellant rocket engine for KSR-III(Korea Sounding Rocket-III), the analysis of engine stand-alone combustion test results was carried out. Considering the variation of ablative material combustion chamber caused by erosion, linear regression analysis that ignores oxidizer/fuel ratio effect and two-variable 2nd-order polynomial regression analysis that considers oxidizer/fuel ratio change were performed. It can be described that linear regression analysis is simple and very practical method, and can predict the performance within 1% error inside analyzed region. And two-variable 2nd-order polynomial regression analysis can predict with very high accuracy inside region and shows that KSR-III engine`s optimum oxidizer/fuel ratio for thrust(or specific impulse) is 2.22 and that for combustion chamber pressure(or characteristic velocity) is 2.17.

Determination of Cyclogram for Liquid-Propellant Rocket Engine

A vertical test stand based on launcher propulsion system was constructed and several tests for the determination of cyclogram were carried out. To make an accurate estimation, static and dynamic pressures were measured and analyzed. Especially, static pressure measurements using fast response sensors without extension tubes were used to determine operation sequence more evidently. The standard operation times of final valves were determined in cold flow tests with an engine head, and fire formation time in combustion chamber was checked in an ignition test with an ignitor only. On the basis of these tests, ignition sequence was established and combustion test cyclogram was finally determined. According to combustion test, test results were well matched with the determined cyclogram within 0.05 sec.