• Journal of the Korean Society of Propulsion Engineers
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• v.24 no.3
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• pp.18-30
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• 2020

This study deals an estimation method of thrust measurement uncertainty in solid rocket motors. Guidelines of the force measurement uncertainty estimation have been provided by ISO, domestic and international organizations. However, all of them are described by focusing on the force calibration machines and force transducers with a conceptually-driven way. Thus the guidelines cannot be directly applicable to uncertainty estimation of calibration equation and its linear approximation, which are critical error sources in the thrust measurement. In this paper, the equations taking into account effects of both error sources are derived based on fundamental concepts of measurement uncertainty. These are applied to the real thrust measurement system where a relatively simple estimation method for the thrust measurement uncertainty is proposed.

• Bulletin of the Korean Space Science Society
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• 2009.10a
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• pp.45.3-46
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• 2009

홀 추력기는 비교적 간단한 구조와 작은 크기 및 높은 연료효율로 미래 소형위성의 핵심기술로 주목 받고 있다. 이 연구실에서는 2010년 발사예정인 과학기술위성 3호에 탑재할 소형위성용 저 전력 홀 추력기를 연구 개발하였다. 성능에 가장 큰 영향을 미치는 자기장 구조는 FEMM전산코드를 이용한 해석을 통해 설계되었으며, 제작된 프로토타입의 실험을 통해 자기장의 세기 및 모양, 양극전압 및 기체유량에 따른 성능 특성을 관찰하였다. 또한 Faraday Probe와 Retarding Potential Analyzer (RPA), 랑뮈어 탐침 등을 이용해 이온빔의 분사각도 및 전류밀도, 이온에너지 분포, 플라즈마 전위 등을 측정하고 관찰된 특성을 물리적으로 분석하였다. 이러한 최적화 과정을 통해 설계된 비행모델의 시험 결과 양극전력 200 W, 제논 연료유량 0.85 mg/s 을 통해 11.2 mN 추력, 1350 s 비추력, 37% 추력효율을 획득하여 개발목표를 상회하는 만족할만한 결과를 얻었다.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2009.11a
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• pp.298-302
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• 2009

Low power Hall thruster is under development as one of the core technologies for STSAT-3. The Hall thruster has several advantages such as its simple structure, high thrust density and specific impulse etc. Development target values deduced by analyzing requirements are consumed electrical power, thrust, thrust efficiency, and specific impulse of < 300 W, > 10 mN, ~ 35%, and > 1000 s, respectively. In order to achieve the target specifications, two prototype Hall thrusters were developed and compared. To date, thrust and efficiency are 11 mN and 37% under the total power of 290 W with 0.97 mg/s Xe propellent supply.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2017.05a
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• pp.310-319
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• 2017

This paper introduces the method how to measure the thrust in impulse facility. In a Facility having such a short duration time of steady flow, there's no time to reach a steady state of the forces acting on model so that the test model vibrates until the end of the flow. The forces exerted on an engine exist with vibration so that the usual force balance can not be used. SWFB(Stress Wave Force Balance) technique is utilized in a shock tunnel to get the thrust. As an example, a model force balance has been calculated its strain against impulse force by using FEM(Finite Element Method). A transfer function between the impulse force and strain has been obtained by the way of de-convolution.

• Journal of the Korean Society of Propulsion Engineers
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• v.8 no.2
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• pp.54-61
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• 2004

Miniaturization of subsystems including propulsion systems is recent trends in spacecraft technology. Small space vehicle propulsion is not only a technological challenge of a scaling system down, but also a combination of fundamental flow/combustion constraints. In this paper, physical constraints of micronozzle for cold gas micro-thruster are reviewed and discussed. Method to measure small thrust are also described.

• Aerospace Engineering and Technology
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• v.9 no.2
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• pp.8-13
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• 2010

An effective control of the axial thrust of a turbopump is one of the critical issues for obtaining its operational stability. Axial thrusts of the fuel pump for the 75-ton class rocket engine under development were measured with water as a test propellant at a room temperature. According to the test results, the axial thrust of the fuel pump seemed to satisfy the axial force condition of its bearing. Also, the thrust was increased as a whole when the flowrate of the pump was decreased. Furthermore it was found that the thrust and the leakage flowate were modified when the gaps between the floating ring seals and the impeller were changed.

• 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.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• v.y2005m4
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• pp.39-43
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• 2005

In this paper, we describe the contents of the achievement of basic experiments for manufacturing and evaluating a micro-rocket using solid propellant, as the device of the micro thrust generator which can be used for attitude control of satellites. We try to analyze performance of the rocket through visualization of the motor jet, and adopt various ignition methods to obtain confidence in ignition problem. For this purpose, we develop and test two different types of thrust measurement system.

• Journal of the Korea Academia-Industrial cooperation Society
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• v.20 no.10
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• pp.100-106
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• 2019

Ground and high altitude simulated combustion experiments were conducted using a liquid rocket engine with hydrogen peroxide and kerosene as the propellant. A ground and high altitude simulated combustion test facility was constructed by installing a high altitude model diffuser and TMS (Thrust Measuring System) on a vertical combustion test bench. The thrust characteristics according to altitude were investigated using the combustion test equipment. The diffuser was designed on a 1:4.8 scale to verify the characteristics of the high diffusing diffuser and starting pressure. The cold flow tests were conducted using nitrogen gas, and the performance characteristics and starting characteristics of the scale down diffuser were verified. A diffuser and TMS were installed on the vertical combustion test bench, and the thrust correction equations for the system resistance were derived. The thrust correction equations were derived from the step test and vacuum step test before the actual hot firing test. Nozzles with an operating altitude of 10km were designed. Hot firing tests were conducted to analyze the thrust characteristics according to the operating altitude changes. The actual thrust was calculated using each correction equation with the thrust value measured by the TMS.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• v.y2005m4
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• pp.381-385
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• 2005

Recently, spacecraft technology trends can be expressed three words, i.e. 'faster, cheaper and smaller'. Among these systems, micro propulsion device is an essential component. Also micro nozzle is the most important part in the micro propulsion device. In case of cold-gas thruster, micro nozzle converts the stored energy in a pressurized gas into kinetic energy through expansion ratio. In this paper we report characteristics of micro nozzle with throat expansion ratio and ratio of specific heats. We measure thrusts using strain gauge based thrust measurement system. We can estimate the micro nozzle performance through experiments.