• Aerospace Engineering and Technology
• /
• v.10 no.1
• /
• pp.1-12
• /
• 2011

The error of onboard propellants mass which is mostly occupied in total mass of launch vehicle and The error of residual affect the performance of launch vehicle seriously. In other words, the errors directly cause the error of total impulse. Therefore, optimization of performance of launch vehicle can be achieved by the minimization of the residual. For minimizing the residuals, the active control for completely depleting the propellants and the calculation method using probability for minimizing the residuals have been researched. In this paper, the added fuel was calculated for minimizing the residual and the minimized residual was predicted by the presented method.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2006.11a
• /
• pp.203-206
• /
• 2006

We have studied, for maximizing the total impulse of liquid propulsion system, Propellant Utilization System (PUS) to minimize outage of propellant. Propellant outage is mainly influenced by propellant mixture ratio during flight and real quantity of loaded propellant. If one employs cryogenic propellant, the variation of propellant density due to the temperature change has major effect on outage control. Feedback control of propellant level of each tank during flight could deplete both tanks simultaneously. To introduce this system, however, the mixture ratio control system of rocket engine is necessary.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2005.11a
• /
• pp.191-195
• /
• 2005

Propellant remains as outage at engine shutdown contributes no useful impulse to the rocket and produces an unwanted increase in burnout weight. Minimization of outage, is therfore is a basic consideration in attaining the maximum performance capability of my bipropellant liquid rocket. This paper present the calculation procedures of outage and optimum loading propellant mass. And some control methods and measurement techniques for outage are presented.

• Journal of the Korean Society for Aeronautical & Space Sciences
• /
• v.43 no.1
• /
• pp.88-95
• /
• 2015

This paper considers the effects of propellant mixture ratio and loading errors on the performance of a liquid rocket. Propellant residuals generated by error sources are analyzed for a launch vehicle model whose first stage consists of a cluster rocket of four 75-tonf class engines using a statistical Monte-Carlo approach and then the optimal fuel biases minimizing residuals are computed. The results are validated through comparison with analytic method using approximate formula, which have been applied for other space launch vehicles.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2000.11a
• /
• pp.8-8
• /
• 2000

하이드라진 단기액체엔진을 장착하고 궤도에서 임무를 수행하고 있는 다목적실용 위성 추진시스템 궤도비행 초기운용 자료에 근거하여 추진제 소모율을 산정 한다. 추진시스템은 위성의 궤도각과 비행고도 조정을 위한 속도증분($\Delta$V) 및 자세제어를 위한 추력을 발생시킨다. 단기액체 추진시스템에서 추진제 소모량은 추력기 밸브의 개폐시간에 비례하고 추력 생성 효율은 추진제의 연소기 유입압력에 종속한다. 일정질량의 가압 기체 압력에 의해 연료를 공급하는 추진시스템에서 잔류 추진제 량의 감소는 연소기 유입압력의 감소를 유발하고 추진기관의 효율을 저하시키는 요인으로 작용하여 임무말기로 진행함에 따라 동일한 운동량 생성에 보다 많은 연료소모가 이루어진다.(중략)

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2017.05a
• /
• pp.968-971
• /
• 2017

In this study, manufacturing the Single Base propellant using a Nitroguanidine(NGD A%, NGD-B%, 2A=B). The Factors affecting the combustion rate are Moisture and Volatiles, Residual Solvents, Dimension. These Factors were analyzed and compared. Also, NGD-A% propellants and NGD-B% propellants were compared with firing test Test. As a result, NGD-A% was confirmed to have a higher speed at a similar pressure than NGD-B% propellants.

• Journal of the Korean Society of Propulsion Engineers
• /
• v.16 no.3
• /
• pp.82-89
• /
• 2012

The temperature of cryogenic propellant in the propellant tank 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.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2011.11a
• /
• pp.709-716
• /
• 2011

The temperature of cryogenic propellant in the propellant tank 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.

• Composites Research
• /
• v.19 no.4
• /
• pp.31-38
• /
• 2006

In this study, effects of curing temperature and autofrettage pressure on a Type 3 cryogenic propellant tank, which is composed of composite hoop/helical layers and a metal liner, were investigated by thermo elastic analysis and composite/aluminum ring specimen tests. Temperature field of a Type 3 tank was obtained from solving the heat transfer problem and, in turn, was used as nodal temperature boundary conditions during the elastic analyses for curing temperature and autofrettage pressure effects. As a result, it was shown that the higher curing temperature was, the more residual compressive stress and tensile stress were induced in composites and metal liner, respectively. On the contrary, autofrettage pressure brought the reduction of these residual thermal stresses caused by cryogenic environments to the tank structure. This tradeoff for curing temperature and autofrettage pressure must be considered in the design and manufacturing stages for a Type 3 cryogenic tank.

• Bulletin of the Korean Space Science Society
• /
• 2003.10a
• /
• pp.69-69
• /
• 2003

우주발사체의 비행 목표는 위성의 궤도 투입이다. 이를 위해서는 발사체에서 요구되는 추력값과 총추력을 보장하는 추진기관이 개발되어야 한다. 엔진은 엔진 자체의 작동 안정성을 위해서 유량제어를 필요로 하지만, 이뿐만이 아니라, 발사체의 비행임무 수행을 위해서도 추진제가 모두 소진되는 시스템(TDS:Tank Depletion System) 개념이 도입되어야 하며, 이는 유량 제어를 통해서 실현된다. 본 연구에서는 우주발사체의 비행임무 수행에 필요한 즉, 총추력 오차 범위, 추력 오차 범위, 추진제 탑재량 및 잔류량 오차범위 관점에서 필요한 추진기관에 요구되는 성능을 검토하였고, 이를 위해 TDS 개념의 도입과 더불어 이를 구현할 수 있는 유량제어 개념을 제시하였다.