• Title/Summary/Keyword: Chemical Propulsion

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Environmental test campaign of a 6U CubeSat Test Platform equipped with an ambipolar plasma thruster

  • Stesina, Fabrizio;Corpino, Sabrina;Borras, Eduard Bosch;Amo, Jose Gonzalez Del;Pavarin, Daniele;Bellomo, Nicolas;Trezzolani, Fabio
    • Advances in aircraft and spacecraft science
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    • v.9 no.3
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    • pp.195-215
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    • 2022
  • The increasing interest in CubeSat platforms ant their capability of enlarging the frontier of possible missions impose technology improvements. Miniaturized electrical propulsion (EP) systems enable new mission for multi-unit CubeSats (6U+). While electric propulsion systems have achieved important level of knowledge at equipment level, the investigation of the mutual impact between EP system and CubeSat technology at system level can provide a decisive improvement for both the technologies. The interaction between CubeSat and EP system should be assessed in terms of electromagnetic emissions (both radiated and conducted), thermal gradients, high electrical power management, surface chemical deposition, and quick and reliable data exchanges. This paper shows how a versatile CubeSat Test Platform (CTP), together with standardized procedures and specialized facilities enable the acquisition fundamental and unprecedented information. Measurements can be taken both by specific ground support equipment placed inside the vacuum facility and by dedicated sensors and subsystems installed on the CTP, providing a completely new set of data never obtained before. CTP is constituted of a 6U primary structure hosting the EP system, representative CubeSat avionics and batteries. For the first test campaign, CTP hosts the ambipolar plasma propulsion system, called Regulus and developed by T4I. After the integration and the functional test in laboratory environment, CTP + Regulus performed a Test campaign in relevant environment in the vacuum chamber at CISAS, University of Padua. This paper is focused on the test campaign description and the main results achieved at different power levels for different duration of the firings.

Analysis on the Internal Flow of the Hydraulic Dual Chambers Applying Various Orifice (다양한 오리피스를 이용한 연결형 공압 챔버 내부 유동 해석)

  • Cho, Kihong;Park, Jungho;Kim, Euiyong
    • Journal of the Korean Society of Propulsion Engineers
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    • v.18 no.1
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    • pp.58-64
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    • 2014
  • Hydraulic dual chamber, as the simulator for a dual pulse rocket motor, was tested by a high pressure device with various orifice-hole size being applied. Pressure difference occurs between 1st chamber and 2nd chamber depending on area ratio of the orifice to nozzle throat. Studying a design configuration of the orifice is essential to the motor development because pressure difference severely affects the rocket motor performance. It is noticed in this study that energy dissipation is caused by the vortex flow originating from the orifice as the 2nd chamber is operated. The flow field is simulated by a commercial computational fluid dynamics program, ANSYS FLUENT V14.5.

Analysis of Three Dimensional Liquid Ramjet Engine with Spray and Combustion (액체 램제트 엔진의 3차원 분무 및 연소 반응 해석)

  • 오대환;임상규;손창현;이충원
    • Journal of the Korean Society of Propulsion Engineers
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    • v.3 no.2
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    • pp.18-24
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    • 1999
  • Liquid ramjet combustor is closely connected with complex phenomena due to a series of processes such as intake air, spray, mixing, and combustion. The present numerical experiments were peformed to investigate these flow characteristics for two and three dimensional liquid ramjet combustor. Grid system was made with three domains: intake region where air is supplied and fuel is injected, combustor and nozzle region, and exit atmosphere region. The numerical results showed that two and three dimensional flow patterns in recirculation region of combustor were significantly different each other and spray model was necessary to predict correctly the chemical reaction flow characteristics. Numerically examined for two different location of fuel injector, one is located on the bottom position of curved intake and the other is located on the top position. We found that bottom position of fuel injector is better than top position because fuel influx to the recirculation region which is need to sustain chemical reaction is more than the latter.

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Internal Flow Characteristics of Simulated Dual Pulse Rocket Motor by Using the Hot Gas and Cold Gas (Hot Gas와 Cold Gas를 이용한 모사 이중펄스 로켓 추진기관의 내부 유동 특성)

  • Cho, Kihong;Park, Jungho;Kim, Euiyong
    • Journal of the Korean Society of Propulsion Engineers
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    • v.19 no.2
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    • pp.1-8
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    • 2015
  • Dual pulse rocket motor is a variant of solid rocket motor with two propellant grain separated by a pulse separation device. The major performance of such a rocket motor is influenced by the change in the hole area of pulse separation device to nozzle throat area ratio. In this study, we performed flow analysis to investigate the internal flow characteristics according to the pulse separation device hole area to nozzle throat area ratio change. Gases used flow analysis were used combustion gas of HTPB/AP composite propellant and nitrogen gas. Flow analysis results of the dual pulse rocket motor were validated by comparison with experimental results of pneumatics. Commercial CFD code ANSYS FLUENT 14.5 is used in this study to simulate flow analysis.

The Analysis on the Effects of Hygrothermal Aging to THPP Using DSC and XPS (DSC와 XPS를 통한 수분노화가 THPP 점화제에 미치는 영향 분석)

  • Oh, Juyoung;Kim, Yoocheon;Yoh, Jai-ick
    • Journal of the Korean Society of Propulsion Engineers
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    • v.23 no.1
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    • pp.79-92
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    • 2019
  • Titanium hydride potassium perchlorate (THPP) is one of the commonly utilized pyrotechnic materials in aerospace industries. The current study elucidates the effects of hygrothermal aging on the combustion of THPP experimentally. First, applying the Differential Scanning Calorimetry (DSC) and isocoversional method, both the delay of reaction start and decrease in maximum reaction rate were observed. The kinetics parameters tended to fluctuate depending the thermal reaction or intermediate product formation of THPP. Also, the oxidants decomposition and fuel oxidation phenomenon were discovered by X-ray photoelectron spectroscopy (XPS). The experimental heat from DSC data were verified as reasonable by comparing with the theoretical heat obtained utilizing both THPP formulation from XPS and NASA Chemical Equilibrium with Applications (CEA). Both data had identical variation trend, which expecially had the highest heat value at 10 weeks aged sample.

Unsteady Ignition in the Pulse Combustor with Counter Jet Flows (대향분출류가 있는 맥동연소기의 비정상 점화현상)

  • 이창진
    • Journal of the Korean Society of Propulsion Engineers
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    • v.1 no.1
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    • pp.64-72
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    • 1997
  • An analytical study has been performed to investigate the unsteady ignition characteristics of pulse combustion. In many combustion applications, strain rate of the flow can significantly affect the combustion features; ignition, extinction, and reignition. In the pulse combustion, two jets (hot combustion gases and fresh mixtures) coming from the opposite side of the combustor will collide in the combustor forming a stagnation region where the chemical reaction is suppressed by the strain rate until this becomes below the critical value. In this research, the method of large activation energy asymptotic is adopted with one step irreversible kinetics to examine the ignition response to the periodic variation of the strain rate of flow. The results show the variation of the maximum value of strain rate can determine whether the ignition or extinction occur.

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Thermochemical Performance Analysis of Hydrazine Arc Thruster (하이드라진 아크 추력기의 열화학적 성능해석)

  • Shin Jae-Ryul;Oh Se-Jong;Choi Jeong-Yeol
    • Proceedings of the Korean Society of Propulsion Engineers Conference
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    • v.y2005m4
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    • pp.35-38
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    • 2005
  • The computational fluid dynamic analysis has been conducted for the thermo-chemical flow field in an arcjet thruster with mono-propellant hydrazine ($N_{2}H_4$) as a working fluid. Coupled Reynolds Averaged Navier-Stokes (RANS) equations and Maxwell equations were used to account for the Ohm heating and Lorentz forces. Hydrazine chemistry and thermal radiation were also incorporated to the fluid dynamic equations by assuming infinitely-fast reactions and optically thick media. In addition to the thermo-physical understandings of the flow field inside the arcjet thruster, results shows that performance indices are improved by amount of $20\%$ in thrust and $200\%$ in specific impulse with the 0.6kW are heating.

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A Study on Ablation Behavior of Graphite Nozzle using Liquid Rocket Engine (액체로켓엔진을 이용한 Graphite 노즐의 삭마 거동 연구)

  • Cho Nam Choon;Park Hee Ho;Keum Young Tag
    • Proceedings of the Korean Society of Propulsion Engineers Conference
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    • v.y2005m4
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    • pp.119-122
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    • 2005
  • Ablation phenomena is very complicated because it includes momentum, energy and mass transfer, chemical reactions as well as phase change. In this paper, ablation at the rocket nozzle throat is modeled as unsteady one dimensional axi-symmetric with proper boundary conditions and field equation is solved numerically. Analytical results are compared with measured ablation data from firing experimental liquid rocket engine. Test variables are combustion pressure and mixture ratio. for low combustion pressure and low mixture ratio, the discrepancy between analysis and experiments are large but for the normal rocket operation range, two results show a simliar trend with maximum discrepancy of $100\%$.

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Ignition Characteristics of JP-7 as a Scramjet Fuel (스크램제트 연료로써 JP-7 의 점화 특성)

  • Choi Jeong-Yeol;Puri Puneesh;Ma Fuhua;Yang Vigor
    • Proceedings of the Korean Society of Propulsion Engineers Conference
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    • v.y2005m4
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    • pp.377-380
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    • 2005
  • A study of ignition characteristics of cracked JP-7 fuel with both oxygen and air has been conducted over a wide range of pressures (1-20 atm), temperatures (1200-2000 K), and equivalence ratios (0.5-1.5). Correlations of ignition delay, of the form, $\tau=Aexp(E/RT)[F]^{a}[O_2]^{b}$ are established using the Chemkin-II package and least square analysis. The effect of $C_3$ hydrocarbons in cracked JP-7 fuel is examined by comparing the ignition delays for two different cracked compositions. A comparison for ignition delay is also made with the experimental results obtained by injecting liquid JP-7 fuel in air using a shock tube apparatus.

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Effects of Aspect Ratio of a Fuel Injection Nozzle into a Supersonic Air Stream on Combustion Characteristics (초음속 공기유동으로의 연료 분사노즐 종횡비 변화에 대한 연소특성 연구)

  • 김경무;백승욱;김윤곤
    • Journal of the Korean Society of Propulsion Engineers
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    • v.8 no.1
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    • pp.44-53
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    • 2004
  • This paper is to investigate the combustion characteristics with a three dimensional chemical reacting flow on the aspect ratio of an exit configuration of the slit type nozzle for the fuel injection and to device the methods of combustion/mixing enhancement. The results show that both inside inflow and slit side vertices should be considered from a viewpoint of the mixing. The combustion efficiency becomes the smallest at aspect ratio, where the aspect ratio is less and more than unity, respectively. The total pressure loss becomes the largest at aspect ratio of unity due to the high penetration. All results imply that a streamwise very long slit is desirable with respect to the combustion and the pressure loss.