Sang Hun Kang

Cowl and cavity effects on mixing and combustion in scramjet engines

To investigate the supersonic combustion patterns in scramjet engines, a model scramjet engine was tested in the T4 free-piston shock tunnel. The test model had a rectangular intake, which compressed the freestream flow through a series of four shock waves upstream of the combustor entrance. A cavity flame holder was installed in the supersonic combustor to improve ignition. The freestream test condition was fixed at Mach 7.6, at an altitude of 31 km. This experimental study investigated the effects of varying fuel equivalence ratios, the influence of the cavity flame holder, and the effects of cowl shape. As a result, supersonic combustion was observed at equivalence ratios between 0.11 and 0.18. Measurements indicated that the engine thermally choked at a fuel equivalence ratio of 0.40. Furthermore, the cavity flame holder and the W-shaped cowl showed improved pressure distribution due to greater reaction intensity. With the aid of numerical analysis, the cavity and the W-shaped cowl are shown to be effective in fuel–air mixing.

Effects of Flameholder Configurations on Combustion in Scramjet Engines

The effects of flameholder configurations on combustion within a scramjet combustor are investigated through experiments, numerical simulation, andquasi-one-dimensional analysis. For the experimental test, a blowdownwind tunnel with a vitiated air heater is used. The test model is composed of an isolator, fuel injectors, and a diverging combustor. To investigate the effects of different flameholder configurations, three kinds of exchangeable fuelinjection plates are fabricated: no cavity, plain cavity, and zigzag cavity. The test results indicate that the zigzag cavity cases obtain the highest values of combustion-induced pressure among the three. The results of the numerical analysis show that the zigzag cavity generates a transverse-directional flow and enhances combustion. The zigzag cavity case also shows the highest combustion efficiency in the quasi-one-dimensional analysis. With a zigzag cavity installed, a scramjet engine combustor can be shortened, and its friction drag can be lowered.

Development of the Scramjet Engine Test Facility in Korea Aerospace Research Institute

A hypersonic air-breathing engine test facility was constructed at Korea Aerospace Research Institute for a hypersonic engine research. This facility, so called Scramjet engine test facility(SeTF) is a blow-down type, high enthalpy wind tunnel which has the pressurized air supply system, air heating system, free-jet type test chamber, supply system, facility control, and measurement system, and exhaust system. In this paper, Details of the specifications and configuration of the Scramjet engine test facility are described, and the performance data of the SeTF are also included.

Turbine Rotor-Pyrostarter Coupled Test for the Verification of Thermo-Structural Suitability of a Turbopump Turbine

ABSTRACT Turbine rotor-pyrostarter coupled test was performed for the verification of thermo-structural suitability of a turbopump turbine. Newly developed solid propellant and design concept were used in pyrostarter development. In case of turbine rotor, rotor config uration modification and post EDM machining process are adopted in rotor manufacturing respective ly for the thermal stress relief and the surface integrity improvement on the blade surfaces. In the tes t, combustion gas of pyrostarter was directly ejected from the nozzles and impinged on the stationar y turbine rotor specimen through the identically shaped flow passage of turbopump. Three kind of thermal load - design to extreme condition - test were performed and no damages were found on the turbine rotor specimens. 초 록 터보펌프 터빈로터의 열구조적 적합성 검증을 위한 터빈로터-파이로시동기 연계시험을 수행하였다. 새로운 추진제를 적용한 파이로시동기와 열응력 경감 설계 및 터빈 동익 표면 건전성 향상을 위한 후가공 공정이 적용된 터빈로터시편이 시험에 사용되었다. 시험은 75톤급 엔진시동을 위한 파이로 시동기의 연소가스를 터보펌프 터빈로터와 동일한 형상의 시편에 분사하는 방식으로 이루어졌다. 터빈에 가해지는 열 부하는 운용 설계점에서 극한 조건까지 세 종류로 구분하여 시험을 진행하였으며 모든 시험에서 터빈로터의 손상은 발견되지 않았다.Key Words: Turbopump(터보펌프), Turbine(터빈), Pyrostarter(파이로시동기), 연계시험(Coupled-test)Received 7 June 2013 / Revised 23 December 2013 / Accepted 2 J anuary 2014

Scramjet Engine Combustor Tests in a Supersonic Wind Tunnel with a Vitiated Air Heater

Effects of flame holder configurations on the combustion within the scramjet combustor are investigated by experiments and quasi-one-dimensional analysis. For the test, a blowdown windtunnel with vitiated air heater is used. The test model is composed of an isolator, fuel injectors and a diverging combustor. In order to investigate the effects of flame holder configurations, three kinds of fuel injection plates, “No cavity”, “Plain Cavity” and “zigzag cavity” are made for exchange. As a result, zigzag cavity showed the highest value of combustion induced pressure and estimated combustion efficiency among others. The zigzag cavity is expected to generate transverse directional pressure nonuniformity and enhance the fuel-air mixing due to irregularity of flame front. With application of zigzag cavity, scramjet engine combustor can be shortened and attain lower friction drag.

Design Point Operating Characteristics of an Oxidizer Rich Preburner

ABSTRACT It was designed and tested at the design point that an oxidiz er rich preburner for a staged combustion liquid rocket engine propelled by kerosene and LOx. The oxidizer rich preburner was designed as some of LOx injected from the mixing head was burne d with kerosene and the rest of LOx injected from injection holes in the regenerative cooling c hamber was vaporized by combustion gas. The preburner is operated at OF ratio of 60 and combustion pressure of 20 ㎫. The Preburner has a honey-comb type mixing head with simplex swirl injectors, a turbulence ring improving combustion stability and uniformity of product gas temperature distribution, and a nozzle simulating the duct. With the combustion test results at the design point, the oxidi zer rich preburner showed high combustion stability and uniformity of product gas temperature distribution.초 록 케로신과 액체산소를 추진제로 하는 다단연소방식 액체엔진용 산화제 과잉 예연소기를 설계하여 설계점에서 연소시험을 수행하였다. 설계된 산화제 과잉 예연소기는 산화제 일부와 연료를 혼합헤드를 통해 연소실에 공급하여 연소시키고 나머지 산화제를 연소실 재생냉각채널을 거쳐 연소실 중앙의 분사공을 통해 연소실로 주입하여 기화시키는 형태로 최종적으로 연소압 20 ㎫, 혼합비 60에서 작동한다. 혼합헤드에는 단일 와류형 분사기를 벌집형태로 배열하였으며 가스 온도 균일성 향상과 연소 안정성 향상을 위한 혼합링과 터빈까지의 배관을 고려한 노즐을 장착하였다. 설계점 연소시험에서 산화제 과잉 예연소기는 높은 연소 안정성과 생성가스의 균일한 온도분포를 보였다.Key Words: Liquid Rocket Engine(액체로켓엔진), Staged Combustion(다단연소), Preburner(예연소기), Oxidizer Rich(산화제 과잉), Combustion Test(연소시험)Received 2 December 2012 / Revised 11 July 2013 / Accepted 17 J uly 2013

Research about Thermoacoustic Resonance Ignition

ABSTRACT The unique phenomenon that jet flow kinetic energy is converted to thermal energy through thermoacoustic resonance can be applied for the multiple ignition of liquid rocket engines. The present article includes the basic principle and theory behind the phenomenon as well as major outstanding, previous research works. The thermoacoustic phenomenon is affec ted by underexpanded jet flow characteristics from a nozzle, geometries of a nozzle and a resonance tube, and chemical composition of jet flow. The paper concludes with discussion what should be considered as crucial issues for the future research on the development of a multiple ignition syste m of liquid rocket engines.초 록 초음속 유동의 운동에너지가 열에너지로 바뀌는 열음향 공진 현상을 활용하여 액체로켓엔진의 다중 점화를 위한 점화기 구현에 적용할 수 있다. 본 논문은 이와 같은 열음향 공진 현상의 기본 원리와 이론, 현재까지 진행된 주요 연구결과를 본문에 수록하였다. 열음향 공진 현상에 의한 열발생 특성은 초음속 노즐을 통해 분출되는 유동의 특성과 노즐과 공진관의 형상 그리고 에너지를 전달하는 기체의 종류에 의해 영향 받는다. 이와 같은 열음향 공진 현상을 적용한 액체로켓엔진을 위한 다중 점화기를 구성하는데 있어서 향후 연구 개발이 필요한 부분에 대한 논의를 진행하였다.Key Words:Thermoacoustic(열음향), Resonance(공진), Heat Generation(열발생), Ignition(점화), Liquid Rocket Engine(액체로켓엔진)Received 3 December 2015 / Revised 8 January 2016 / Accepted 14 January 2016

Study on Pressure Fluctuations Observed in Combustion of Oxygen-Rich Preburners

․ Sang Hun Kang** ․ Soo-Yong Lee***ABSTRACT The paper includes the analytic results of pressure fluctuati on data from the combustion of an oxidizer-rich preburner applicable to high-performance, closed- cycle liquid rocket engine systems. The combustion experiments went through two different steps of a ch amber pressure during single run. Self-excited pressure fluctuations with a frequency of 78 Hz were observed only at the relatively low chamber pressure condition. These pressure fluctuations are reg arded as a bulk mode. The intensity of pressure fluctuations by a root-mean-square value is 13.3% norm alized by the chamber static pressure and no pressure excitation was observed at the design pressure condition. The bulk mode has an identical phase across the inside of the chamber and reveals th e similar characteristics to the Helmholtz resonator.초 록 고성능 폐회로 액체로켓엔진 시스템에 적용되는 산화제 과잉 예연소기 내에서 발생하는 연소 압력 섭동 데이터에 대한 분석 결과를 수록하였다. 연소실험은 두 단계의 압력 값을 순차적으로 형성하였는데 상대적으로 저압 조건에서만 78 Hz의 특성 주파수를 갖는 자발적인 압력 섭동이 발생하였다. 이와 같은 연소 불안정은 bulk mode에 해당하는 것으로 판단된다. 압력 섭동의 세기는 rms 기준으로 연소실 정압 대비 13.3%에 해당하며, 설계압 구간(19.3 MPa)에서는 매우 안정적인 모습을 보였다. 이와 같은 bulk mode는 연소실 내부에서 특성 압력파의 위상차가 없이 전체 공간에서 압력이 증가와 감소를 반복하는 것으로 Helmholtz 공진기와 비슷한 특성을 보이는 것으로 판단된다.Key Words:Pressure Fluctuations(압력 섭동), Oxygen-Rich(산화제 과잉), Preburner(예연소기), Combustion Dynamics(동적 연소특성), LOx(액체산소), Kerosene(케로신), Bulk Mode(벌크 모드)