• Journal of the Korean Society of Propulsion Engineers
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• v.10 no.3
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• pp.73-81
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• 2006

In order to evaluate the thrust of a propulsion system, generally the captive thrust stand is used Based on the applied propulsion system, the various captive ways between thrust stand and the propulsion system are considered. In this paper, the effect of motor deformation generated during firing in horizontal thrust stand on the measured thrust is theoretically derived from classical beam theory. New flexure performance index is defined and estimated on the basis of the thrust measurement error. Its result is good agreement with numerical result of ABAQUS. This study showed that measurement reliability and safety of test can be highly upgraded, in case of two flexure-type captive thrust stand.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 1998.04a
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• pp.22-22
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• 1998

기존의 공력 조타에 의한 비행 자세 제어 방법은 속도의 2승에 비례하는 제어력을 발생하지만, TVC(Thrust Vector Control)를 이용하면 추력 방향을 변경하여 제어력을 얻음으로써 방향 제어에 보다 월등한 성능을 발휘하는 것으로 알려져 있다. 후자의 방법으로는 저속도 경우와 공기가 희박한 고 고도에서도 충분한 제어력을 얻을 수 있다. 보다 효율적인 제어력을 얻기 위해서는 TVC 방법이 우수하지만 그 성능에 대해서는 충분한 자료가 없는 것이 현재의 상태이다. 제트 베인 방식의 TVC는 베인이 직접 고온 고속의 가스 흐름 내에서 작용하기 때문에 편향추력 발생 측면에서 아주 우수한 방식이며 추력 편향각, 추력 손실 등의 유체역학적인 특성은 제트 베인의 형상, 위치 등으로 결정된다.

• Proceeding of EDISON Challenge
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• 2016.03a
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• pp.537-540
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• 2016

유체역학적 추력 방향 제어(Fluidic Thrust Vector Control) 방법 중 하나인 동축류 제어 유동 분사를 이용한 추력 방향 제어(Co-flow Thrust Vector Control)의 작동 특성에 대해서 연구하였다. 이 제어 방법은 점성 유동이 벽면에 부착되어 흐르는 코안다 효과(Coanda Effect)를 이용하여 주 유동을 편향 시키는 방법으로서 그 편향각은 이러한 제어 유동 노즐 출구의 플랩 형상에 영향을 받는 것으로 알려져 있다. 이 연구에서는 출구 플랩 형상을 여러 가지로 바꾸어 가며 주 유동의 전압력 300kPa일 때 제어 유동의 편향각이 포화되는 제어유동의 전압력을 측정하였다. 그 결과 쐐기형 플랩의 각도가 증가할수록 포화 영역에서의 편향각은 증가하며 그 각은 플랩의 각도와 일치한다. 그러나 각도가 증가할수록 제어 유동이 플랩의 벽면을 지나면서 팽창파에 의해 가속되어 충격파을 발생시키게 되고 이 충격파는 주 유동에게까지 전파되어 주 유동 제트의 속도를 감소시킨다.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2010.11a
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• pp.55-62
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• 2010

In addition to the previous study for development of a 1,200 N-class bipropellant rocket engine with concentrated hydrogen peroxide, the effect of characteristic length and thrust measurement were experimentally evaluated. Tests with characteristic lengths of 0.95, 1.07, and 1.20 m were performed and $C^*$ and Isp efficiencies were increased as increasing characteristic length. The maximum $C^*$ and Isp efficiencies were 98.4% and 93.1% respectively. Based on the evaluation of the designed engine, the optimized characteristic length was proposed in using the engine adapted decomposed hydrogen peroxide and the engine performance at vacuum-level was evaluated using thrust and Isp efficiency at the designed equivalence ratio. As a result, 218.4 s at sea-level, 253.3 s at vacuum-level, and vacuum thrust of 1035.3 N can be estimated.

• Journal of the Korean Society of Propulsion Engineers
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• v.15 no.3
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• pp.1-8
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• 2011

In addition to the previous study for development of a 1,200 N-class bipropellant rocket engine with concentrated hydrogen peroxide, the effect of characteristic length and thrust measurement were experimentally evaluated. Tests with characteristic lengths of 0.95, 1.07, and 1.20 m were performed and $C^*$ and Isp efficiencies were increased as increasing characteristic length. The maximum $C^*$ and Isp efficiencies were 98.4% and 93.1% respectively. Based on the evaluation of the designed engine, the optimized characteristic length was proposed in using the engine adapted decomposed hydrogen peroxide and the engine performance at vacuum-level was evaluated using thrust and Isp efficiency at the designed equivalence ratio. As a result, 218.4 s at sea-level, 253.3 s at vacuum-level, and vacuum thrust of 1035.3 N can be estimated.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2010.05a
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• pp.66-70
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• 2010

The performance of 75 ton liquid rocket engine combustion chamber for a space launch vehicle was predicted through firing tests at low pressure. In low pressure tests of 75 ton LRE combustor chamber, the combustion characteristic velocity of 1750 m/sec and the specific impulse of 240 sec were obtained which are higher than the low pressure performance of 30ton combustion chamber. The combustion characteristic velocity of 1770 m/sec and the specific impulse of 278 sec at design point for 75 ton LRE combustion chamber were predicted by using the low/high pressure performance correlation of 30ton LRE combustion chamber.

• Journal of the Korean Society of Propulsion Engineers
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• v.3 no.1
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• pp.1-8
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• 1999

In the present paper an attempt has been made to simulate the secondary injection-primary flow interaction in the conical rocket nozzle and to derive the performance of secondary injection thrust vector control(SITVC) system. Complex three-dimensional flowfield induced by the secondary injection is numerically analyzed by solving unsteady three-dimensional Euler equation with Beam and Warming's implicit approximate factorization method. Emphasized in the present study is the effect of secondary injection such as secondary mass flow rates and the momentum of secondary/primary nozzle flow mass rates upon the gross system performance parameters such as thrust ratio, specific impulse ratio and deflection angle. The results obtained in terms of system performance parameters show that lower secondary mass flow rate is advantageous for to reduce secondary specific impulse loss. It is further found that the nozzle with secondary jet injected downstream and interacting with fast primary flow is preferable for efficient and stable SITVC over the wide range of use with the penalty of side specific impulse loss.

• Journal of the KSME
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• v.56 no.9
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• pp.38-43
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• 2016

초임계 난류유동이 활용되고 있는 가장 대표적인 시스템은 액체로켓연소기이다. 액체로켓엔진은 목적하는 임무에 따라 추력이 결정되고 엔진의 추력이 결정된다. 이러한 추력을 결정하는 핵심부품이 분사기(injector)이다. 실용 인공위성을 발사하기 위한 액체로켓엔진의 연소기는 수백 개의 분사기를 통해 연료와 산화제가 혼합되는 구조를 가지고 있다. 따라서 로켓연소기의 연소특성 및 성능은 분사기의 혼합특성에 좌우된다. 그러므로 단일 분사기의 연료/산화제 혼합특성에 대한 많은 실험과 해석연구가 진행되고 있다. 그런데 초임계압력에서는 액체의 표면장력이 사라지게 되어 독특한 혼합특성을 가지고 있기 때문에 성능을 높여주기 위하여 초임계 압력을 선택할 경우 분사기에 대한 연구가 선행되어야 한다. 이 글에서는 이러한 초임계 작동압력에서 분사기에 대한 연구들을 소개하고자 한다.

• Journal of the Korean Society of Propulsion Engineers
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• v.20 no.3
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• pp.54-62
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• 2016

Micro plasma thruster (${\mu}PT$) was studied experimentally with a dual-stage micro-hollow cathode discharge (MHCD) plasma. Electrostatic-like acceleration exhibiting more directional and elongated exhaust plume was achieved by a dual layer MHCD at the total input power less than 10 W with argon flow rate of 40 sccm. V-I characteristic indicated that there was an optimal regime for dual-stage operation where the acceleration voltage across the second stage remained constant. Estimated exhaust plume length showed a similar trend to the analytic estimate of exhaust velocity which scales with an acceleration voltage. ${\mu}PT$ with multiple holes exhibited similar performance with single-hole thruster indicating that higher power loading is possible owing to decreased power through each hole. Boltzmann plot of atomic argon spectral lines showed average electron excitation temperature of about 2.6 eV (~30,170 K) in the exhaust plume.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.32 no.9
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• pp.122-129
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• 2004

An accurate thrust measurement is one of the critical paths to the successful test and evaluation program of small liquid propulsion engines. This study describes the design factors for the development of thrust measurement system (TMS) as well as manufacturing practice of TMS hardware. We investigate characteristics of the TMS and its performance through hot-firing test of small liquid engine in a vacuum test cell which is capable of simulating 100,000 ft of altitude or higher. For performance test of TMS, we measure thrusts by changing propellant injection pressure at steady state firing mode as well as at pulse firing mode. Measured eigen frequency of the TMS is 67 Hz. Linearity test of the TMS shows good performance with less than 0.5% of linearity error.