• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.31 no.3
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• pp.72-78
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• 2003

In this paper, the feasibility study and the parametric design of hybrid motor with HTPB/LOX were conducted for micro air launch system. Design results were compared with 1st stage of Pegasus XL for verification of hybrid motor. Results showed that hybrid motor replace solid booster if Isp of hybrid motor reaches 330sec. In addition, mission analysis was established for micro air launch system, and parametric design was conducted with design variables: number of port, initial oxidizer flux, and chamber pressure. And the region of Isp was identified by parametric study which satisfied design constraints and mission analysis.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2002.04a
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• pp.77-77
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• 2002

초소형 공중발사체 설계 시 하이브리드 모터의 적용가능성에 대한 연구를 실시하였다. HTPB/LOX를 추진제로 하여 마차바퀴형 연료 그레인, 산화제 탱크 가압방식을 사용하였고, 성능특성을 계산하기 위하여 하이브리드 연료의 연소율이 일정하다고 가정 하였다. 본 연구에 사용된 임무는 중량 3.5kg의 나노위성을 근지점 고도 200km, 원지점 고도 1,500km의 타원궤도로 진입시키는 것을 목적으로 하는 로켓의 1단 부분에 관한 것으로 1단의 발사속도는 M=1.3, 발사고도는 12km, 연소종료 고도는 40km이다. 1단에 대한 페이로드 중량은 127.5kg이고, 속도증가분($\Delta$V)은 3,330m/s이다. 모선은 F-4E를 사용하였고 모선의 특성상 발사체의 총 중량이 1,000kg이하로 제한되고 길이와 직경이 5m${\times}$5m로 제한되나 1단에 대한 길이의 제한조건은 현재까지 명확히 정립되지 않은 상태이다. 설계과정에서의 변수는 연료 그레인 포트 개수, 초기 산화제 플럭스, 연소실 압력을 사용했고, 설계 제한조건은 추진제 중량, 평균 비추력, 평균 추력, 연소시간, 1단 길이, 직경, 연소시간이고, 이들의 범위는 모선의 특성과 초소형 공중발사체의 임무특성에 맞게 설정하였다.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2009.11a
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• pp.507-510
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• 2009

The small size of hybrid rocket using PE-$LN_2O$ was designed, constructed and launched for a development basic technology of Hybrid rocket vehicle. The hybrid engine ignition system was designed with valve system using external actuator and confirmed working without any fault. To design fuel grain an internal ballistics design was carried out, and to estimate rockets flight path an external ballistics analysis was carried out. So the rocket was designed and constructed, and the launch test proves that hybrid rocket's design was suitable. The hybrid rocket(weight : 9kg, diameter : 110 mm, height : 1.7 m) was launched successfully. But parachute was deployed on mid-flight and the mission could not finish its purposed flight. Some of problems were found in this activity but next launch vehicle will be improved.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2010.11a
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• pp.820-824
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• 2010

The purpose of this study is to develop basic technology of hybrid rocket vehicle by constructed and launched. This small hybrid rocket using HDPE/LN2O and Aluminium steel for its body (weight : 12.5 kg, diameter : 114 mm, height : 1.8 m) was designed. The fuel grain and injector were designed for 50 kgf thrust and burning time 2.5 sec. This rocket was loaded the data acquisition device for obtaining data of pressure and velocity during its flying and equiped an automatic ejector system using spring/motor and timer to collect the rocket more safely after launching. It was launched successfully, but found some problem that the rocket's weight was heavier than expected and the thrust was not enough to reach the designed altitude and analyzed its flight path way.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2005.11a
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• pp.491-494
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• 2005

This paper describes the fundamental research of hybrid-propellant rocket system focusing the flight. The performance of the hybrid-propellant rocket system (using the background of lab-scale tests) for a small-scale launch vehicle could have been validated by static thrust tests and flight tests. Based on this system, it was found that hybrid rocket propulsion system was available in an application to a launch vehicle system.

• 한국항공운항학회:학술대회논문집
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• 2015.11a
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• pp.13-19
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• 2015

본 연구의 목적은 자체 설계 제작한 하이브리드 발사체를 이용하여 대기 중의 미세먼지를 측정하는데 있다. 대기 중의 미세먼지를 측정하기 위한 방법으로는 중량농도법을 이용하였으며, 미세먼지 측정을 위한 포집장치를 자체 제작하였다. 엔진은 5port의 HDPE(High Density Polyethylene)를 연료 그레인으로 사용하였고 $LN_2O$(Liquefied Nitrous Oxide)를 산화제로 사용하였다. 압력, 가속도, GPS 등을 수집하며, 하이브리드 발사체의 동체는 FRP(Fiberglass Reinforced Plastics)로 제작한다. 전체 비행시간은 총 95초로 예상되며. 연소시간은 3초, 고도는 800 m 이다.

• The Journal of the Acoustical Society of Korea
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• v.39 no.4
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• pp.351-363
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• 2020

Launch vehicles are subject to airborne acoustic loads during atmospheric flight and these effects become pronounced especially in transonic region. As the vibration due to the acoustic loads can cause malfunction of payloads, it is essential to predict and reduce the acoustic loads. In this study, a complete process has been developed for predicting airborne vibro-acoustic environment inside the payload pairing and subsequent noise reduction procedure employing acoustic blankets and Helmholtz resonators. Acoustic loads were predicted by Reynolds-Averaged Navier-Stokes (RANS) analysis and a semi-empirical model for pressure fluctuation inside turbulent boundary layer. Coupled vibro-acoustic analysis was performed using VA One SEA's Finite Element Statistical Energy Analysis (FE-SEA) hybrid module and ANSYS APDL. The process has been applied to a hammerhead launch vehicle to evaluate the effect of acoustic load reduction and accordingly to verify the effectiveness of the process. The presently developed process enables to obtain quick analysis result with reasonable accuracy and thus is expected to be useful in the initial design phase of a launch vehicle.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.32 no.4
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• pp.53-60
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• 2004

Optimal design of the hybrid motor has been performed for the first stage of nanosat air launch vehicle using F-4E Phantom as mother plane. Selected design variables are number of ports, the initial oxidizer flux, the combustion chamber pressure, and the nozzle expansion ratio. GBM(Gradient Based Method) and GA(Genetic Algorithm) are simultaneously used to compare the versatility of each algorithm for optimal design in this problem. Also, two objective functions of motor weight, and length are treated separatedly in the optimization to study how the objective function can affect the optimal design. The design results show that the optimal design can be successfully achieved either using GBM or GA regardless of the choice of the objective function; motor weight or length. And nanosat air launch vehicle which has total mass of 704.74kg, and length of first stage 3.74m is designed.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• v.y2005m4
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• pp.79-82
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• 2005

The hybrid rocket has been known for over 50 years. It is safe and cheap but wasn't widely used for the deficit of low regression rate. However, the hybrid rocket propulsion system will replace a lot of fields of missiles, rockets and propulsion systems of launch vehicles with new development of paraffin based solid fuel composition

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2007.11a
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• pp.248-251
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• 2007

The blowdown oxidizer feeding system is effective in the respect of higher reliability by the small number of parts and the absence of additional pressurization tanks, but it also has the unfavorable disadvantage such as thrust variation during the operation. Thus, in order to understand the these performance characteristics inherent in the Hybrid Rocket Motor (HRM) with blowdown oxidizer feeding system, this study proposed the integrated mathematical model to describe physical phenomena in the following parts: the oxidizer tank, combustion chamber, fuel grain, nozzle and injector.