• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2011.04a
• /
• pp.228-231
• /
• 2011

Burning rate of a propellent is an essential factor when designing a propulsion system. In order to come up with burning rate, first we need to design and build propellent grain to get neutral pressure curve. Then check the pressure with ground test and calculate the burning rate using burning rate equation. This burning rate is then compared to the burning rate of a propellent which was resulted from making a standardized specimen and combusting it using a strand burner. An accurate burning rate is calculated after comparing those two burning rates. For this study, compact propulsion system was designed, produced, tested and analyzed in order to get burning rates, an essential factor in propulsion system design, in an effective way.

• Journal of the Korean Society of Propulsion Engineers
• /
• v.15 no.4
• /
• pp.11-17
• /
• 2011

In order to see flame behavior in the annular reverse gas turbine combustor, sector combustion test was performed. Ignition test by using torch ignition system was carried out at various combustor inlet velocity and air fuel ratio. Also, flame blow out limit was measured by changing fuel flow rate with constant air mass flow rate. In test results, stable ignition is possible at air excess ratio of 6 and this limit is gradually increased with combustor inlet velocity. The minimum blow out limit is about 4 at 40 m/s of combustor inlet velocity. This blow out limit is also increased up to about 10 with increasing combustor inlet velocity. Test result shows that lean blow out limits are increased with air velocity. The highest blow out limit was found at the combustor inlet velocity of 65 m/s.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2010.11a
• /
• pp.153-159
• /
• 2010

In order to see the flame behavior in the annular reverse gas turbine combustor, sector combustion test was performed. Ignition test by using torch ignition system was carried out at the various combustor inlet velocity and air fuel ratio. Also, flame blow out limit was measured by changing fuel flow rate with constant air mass flow rate. In the test results, stable ignition is possible at air excess ratio of 6 and this limit is gradually increased with combustor inlet velocity. The minimum blow out limit is about 4 at 40 m/s of combustor inlet velocity. This blow out limit is also increased up to about 10 with increasing combustor inlet velocity. Test result shows that lean blow out limits are increased with air velocity. The highest blow out limit was found at the combustor inlet velocity of 65m/s.

• Fire Science and Engineering
• /
• v.29 no.6
• /
• pp.13-19
• /
• 2015

This study was performed to test the combustive properties of Pinus rigida specimens treated with phosphorus (P) and nitrogen (N) additives. Each Pinus rigida specimen was painted three times with 15 wt% P-N additive solutions at room temperature. After drying the treated specimens, the combustion properties were examined using a cone calorimeter (ISO 5660-1). The time to ignition (TTI) for the treated specimens was 90 to 148 s except for the specimen treated with PP/$4NH_4^+$, and the time to flameout (TF) was 556 to 633 s, which was longer than that of virgin plate. While the The specimens treated with P-N additives showed 12.5 to 43.4% higher mean heat release rate ($HRR_{mean}$) and 11.8 to 43.1% higher total heat release (THR) than virgin plate. The effective heat of combustion (EHC) was by 2.9 to 17.5% lower than that of virgin plate. It can thus be concluded that the combustion-retardation properties were partially improved compared to those of virgin plate.

• Journal of the Korean Society of Propulsion Engineers
• /
• v.26 no.4
• /
• pp.64-71
• /
• 2022

For the development of an improved upper-stage engine, research on a staged combustion cycle liquid rocket engine is in progress. A cold flow test, ignition test, and combustion test plans were established and performed to develop reignition combustion technology. In order to solve the problem of purge gas flowing into the fuel line, which may cause cavitation in the turbo pump during reignition, the test results of each stage were analyzed. Based on the analysis results, the purge gas inflow problem was solved by reducing the overlapping time between the operation of the bubble removal valve and the opening of the purge valve and the engine fuel valve. Based on this, the reignition combustion test was successfully performed.

• Proceedings of the Korean Nuclear Society Conference
• /
• 1998.05b
• /
• pp.285-290
• /
• 1998

KALIMER 고속로의 기본 핵연료인 U-l0Zr 핵연료봉의 노내 성능시험을 위해, 러시아의 BR-10 연구용 고속원자로에서 핵연료노내조사 Program이 1997년부터 수행되고 있다. 1 차년도에는 핵연료 시편의 설계 및 제조와 금속합금 핵연료의 균질도, 밀도, 열전도도 등의 노외 특성 시험이 수행되었다. U-l0Zr 핵연료심은 Arc 용해로 제조되었는데, 합금의 구성 원소들은 비교적 균일하게 분포되었다. 핵연료 시편은 2 개가 제작되었는데, BR-10 원자로에서 각각 연소도 1.08 % 및 2.15 %까지 연소된 후, 조사 후 검사가 수행될 것이다. 금속핵연료는 대개 낮은 연소도에서 급격한 변화틀 보이기 때문에, 본 핵연료 노내조사시험 Program의 결과는 금속핵연료봉의 성능해석 모델 개발에 활용될 수 있을 것이다.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2009.05a
• /
• pp.111-115
• /
• 2009

The method of test for observing the combustion instability and controling the instability actively by using secondary injection of fuel through flame stabilizer was studied by constructing model combustor of supersonic engine. The frequency of combustion instability was detected by measuring the pressure of combustor using pressure sensor and by optical sensing of flame intensity. Electro-magnetic valve was adopted as actuator for active control and the characteristics of modulated fuel was studied by measured pressure of valve outlet and scattering signal of spray at secondary fuel injection.

• Journal of the Korean Society of Propulsion Engineers
• /
• v.11 no.5
• /
• pp.1-13
• /
• 2007

Scramjet engine is one of the most promising propulsion systems for future transport. For the ground test with T4 shock tunnel, model scramjet engine is designed. Design flight Mach number is 7.6 and flight altitude is 30km. Engine intake is designed by Levenberg-Marquardt optimization method and Korkegi relation. Furthermore, cowl cut out region is installed by the rule of Kantrowitz limit. Inside the combustor, cavity type flame holder is installed. Cavity is designed by Rayleigh line relation and PSR model. Numerical analysis is performed for the design confirm.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2004.03a
• /
• pp.198-204
• /
• 2004

Using liquefied natural gas as a fuel, water, natural gas and liquefied natural gas-cooled firing tests were conducted. With the viewpoint of characteristic velocity, and specific impulse, the effect of OF mixture ratio and fuel inlet temperature into a combustion chamber were analyzed. OF mixture ratio and fuel inlet temperature into a combustion chamber have great influence on the performance. Characteristic velocity and theoretical specific impulse attain the maximum value at 0.72~0.75 and 0.75 of OF mixture ratio, respectively. Engine performance has a tendency to increase, proportional to fuel inlet temperature into a combustion chamber affected by the regenerative cooling.

• Journal of the Korean Society of Propulsion Engineers
• /
• v.24 no.6
• /
• pp.85-92
• /
• 2020

Ground hot-firing tests were conducted to analyze the combustion performance according to the characteristic lengths 1.37 m, 1.71 m, and 2.06 m of the combustion chamber in 200 N-class GCH4-LOx small rocket engine. Thrust, specific impulse, and characteristic velocity at the steady-state could be obtained as the key performance parameters of the rocket engine. The performance characteristics acquired through the test were compared and analyzed with the theoretical performance calculated from CEA analysis. Observation of the influence of characteristic length on the combustion performance indicates that an optimal characteristic length shall remain between 1.71 m and 2.06 m.