• Journal of Aerospace System Engineering
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• v.11 no.4
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• pp.48-53
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• 2017

Liquid rocket engines (LREs) are very complex systems that include combustion chambers, turbopumps, gas generators, ducts and tubes, valves and etc. Most components of the LREs require higher than or equal to level 6 IT (ISO Tolerance). The components along with pipe line and/or tubing must dispose not to interfere each other. In addition, effectiveness of maintenance and service after assembling should be considered when the allocation of the components are determined. Especially at the stage of the development, tolerance accumulations or unpredictable errors may result in misalignment and/or mismatches at interfaces of the parts. Namely, it is the engine assembling process that many inherent risks are realized and crises or incidents occur. Therefore, a rapid reaction system should be prepared. In this research, 7 tonf class liquid rocket assembling process was studied and actual building steps were introduced.

• Composites Research
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• v.36 no.2
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• pp.132-139
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• 2023

A linerless composite propellant tank was designed and manufactured by using the carbon fiber-reinforced composite materials which have superior strength-to-weight ratio in order to reduce weight of the tank. In this research, we designed a sub-scale composite propellant tank with a diameter of 800 mm to withstand an MEOP of 1.7 MPa. We manufactured the boss of the tank by using the same composite materials to reduce the thermal expansion difference between the boss and the secondary-bonded composite layers of the barrel in the cryogenic environment. We used the collapsible mandrel to manufacture the tank without any liner. The mandrel was made from epoxy-based composite tooling prepregs to reduce weight of the mandrel. We manufactured the test tanks by laying up the carbon fiber fabric prepregs manually on the mandrel and then applying the autoclave cure process. We performed a proof test, a helium tightness test, a repeated pressurization test, and a burst test in room temperature. The test results demonstrate that the proposed design and manufacture process satisfies all strength requirements as well as an anti-leakage requirement.

• Bulletin of the Korean Space Science Society
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• 2006.10a
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• pp.127-127
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• 2006

• International Journal of Aeronautical and Space Sciences
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• v.18 no.4
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• pp.807-815
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• 2017

For liquid rocket engine(LRE)-based space launch vehicles, longitudinal instability, often referred to as the pogo phenomenon in the literature is predicted. In the building block of system-level task, accurate dynamic modeling of a fluid-filled tank is an essential. This paper attempts to apply the virtual mass method that accounts for the interaction of the vehicle structure and the enclosed liquid oxygen to LOX-filled tanks. The virtual mass method is applied in a modal analysis considering the hydroelastic effect of the launch vehicle tank. This method involves an analysis of the fluid in the tank in the form of mass matrix. To verify the accuracy of this method, the experimental modal data of a small hemispherical tank is used. Finally, the virtual mass method is applied to a 1/8-scale space shuttle external tank. In addition, the LOX tank bottom pressure in the external tank model is estimated. The LOX tank bottom pressure is the factor required for the coupling of the LOX tank with the propulsion system. The small hemispherical tank analysis provides relatively accurate results, and the 1/8-scale space shuttle external tank provides reasonable results. The LOX tank bottom pressure is also similar to that in the numerical results of a previous analysis.

• Journal of Aerospace System Engineering
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• v.18 no.3
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• pp.70-75
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• 2024

The insulation performance of a common bulkhead propellant tank for small launch vehicles with variations in insulation thickness was analyzed. The common bulkhead propellant tank composed of a single part allows for lightweight design, as it eliminates the need for tank connections. However, problems such as propellant loss and ignition delay due to heat transfer caused by temperature differences between oxidizer and fuel may arise. Therefore, it is essential to verify the insulation performance of the common bulkhead structure that separates the oxidizer tank and fuel tank. In this study, transient heat transfer analysis was conducted for propellant tanks with insulation thicknesses of (50, 55, 60, 65, and 70) mm to analyze the insulation performance using boil-off mass. Subsequently, the boil-off mass of the oxidizer generated during the first-stage flight time of the propellant tank was determined. The results confirmed that increasing the insulation thickness reduces the boil-off mass, thereby improving the insulation performance.

• Journal of the Korean Society of Propulsion Engineers
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• v.24 no.1
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• pp.17-23
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• 2020

Upper stage of launch vehicle mainly injects a lunar explorer from low earth orbit to the moon at a distance of 380,000 km. In foreign lunar explorer, the upper stage is separated from the explorer after the explorer is injected into the earth-moon transfer trajectory, and the lunar explorer then uses on-board propellant to carry out mid-course correction maneuvers and lunar orbit insertion maneuvers. This study describes a newly presented small liquid upper stage. Using a small liquid upper stage with a wet mass of 2.9 tonnes, the lunar explorer not only can be injected earth-moon transfer trajectory but also can be performed lunar orbit insertion. This study provides acceptable mass range of the lunar explorer and the scope of acceptable mission range also describes based on the launch from Naro Space Center.

• Journal of the Korea Institute of Military Science and Technology
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• v.23 no.5
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• pp.459-465
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• 2020

In this paper, a compact antenna assembly for an amplitude comparison direction-finding(DF) method for a small flight vehicle is presented. Designed antenna assembly consists of four antennas and it is mounted on a radius of 1.45 λ_c where λ_c corresponds to the wavelength of the center frequency. To achieve compactness and robustness of the assembly, the elements are fed by end-launch feeding method and have modified aperture shapes of E- or H-sectoral horns. The feeding part consists of SMA connector, stepped impedance matching structure, and square waveguide of 0.6 λ_c × 0.6 λ_c. To achieve different main beam directions for every antenna which is required condition for amplitude comparison DF method, all apertures of the antennas are inclined and it makes the main beam direction of each antenna to top, bottom, left, and right with respect to the axis of the platform. To verify the validation of DF performance of the presented antenna assembly, amplitude comparison curves using measurement results are presented. The bandwidth of the antennas are above 3.2 % in Ku-band(VSWR ≤ 2:1).

• Journal of Aerospace System Engineering
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• v.16 no.5
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• pp.8-16
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• 2022

This paper presents the orbital control for positioning micro synthetic aperture radar (SAR) satellites for all-weather monitoring around the Korean Peninsula. In Small SAR technology experimental project (S-STEP) developed in Korea, multiple satellites are placed at equal intervals in multiple orbital planes to secure an average revisit period for the region around the Korean Peninsula. Satellites entering the same orbital plane use ion thrusters to control their orbits and the separation velocity from the launch vehicle to distribute them evenly across the orbit. For an orbital that places the satellites equally spaced in the same orbital plane, the shape of the satellite constellation is formed by adjusting the difference in drift rates between the satellites. This paper presents, different types of satellite constellations, and the results of satellite constellation placement according to launch strategies are presented. In addition, a method and limitations in shortening the duration of orbital deployment are presented.

• Current Industrial and Technological Trends in Aerospace
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• v.9 no.1
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• pp.158-164
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• 2011

Check valves developed for aerospace applications and commercially available for the applications are investigated. The examples include the ones for launch vehicles, SSME (Space Shuttle Main Engine) and GSE (Ground Support Equipment) purges developed by NASA, requiring high reliability, and the ones by KARI. Also the commercial ones for room and cryogenic temperatures by major valve US companies. Relations of design factors such as seal materials and spring rate to principal performances like operating temperature/pressure and cracking pressure are explained. Then potential operational problems such as chatter and contaminations are explained. Also, filters, fittings for end connections and cleanliness requirements for the applications are considered.

• The Korean Journal of Air & Space Law and Policy
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• v.5
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• pp.101-118
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• 1993

Generally there is no law and liability system which applies particulary to commercial space ventures. There are several international treaties and national statutes which deal with space ventures, but their impact on the liability of commercial space ventures has not been significant. Every state law in the United States will impose both tort and contract liability on those responsible for injuries or losses caused by defective products or by services performed negligently. As with the providers of other products and services, those who participate in commercial space ventures have exposure to liability in both tort and contract which is limited to the extent of the resulting damage The manufacturer of a small and cheap component which caused a satellite to fail to reach orbit or to operate nominally has the same exposure to liability as the provider of launch vehicle or the manufacturer of satellite into which the component was incorporaded. Considering the enormity of losses which may result from launch failure or satellite failure, those participated in commercial space ventures will do their best to limit their exposure to liability by contract to the extent permitted by law. In most states of the United States, contracts which limit or disclaim the liability are enforceable with respect to claims for losses or damage to property if they are drafted in compliance with the requirements of the applicable law. In California an attempt to disclaim the liability for one's own negligence will be enforceable only if the contract states explicitly that the parties intend to have the disclaimer apply to negligence claims. Most state laws of the United States will refuse to enforce contracts which attempt to disclaim the liability for gross negligence on public policy grounds. However, the public policy which favoured disclaiming the liability as to gross negligence for providers of launch services was pronounced by the United States Congress in the 1988 Amendments to the 1984 Commercial Space Launch Act. To extend the disclaimer of liability to remote purchasers, the contract of resale should state expressly that the disclaimer applies for the benefit of all contractors and subcontractors who participated in producing the product. This situation may occur when the purchaser of a satellite which has failed to reach orbit has not contracted directly with the provider of launch services. Contracts for launch services usually contain cross-waiver of liability clauses by which each participant in the launch agrees to be responsible for it's own loss and to waive any claims which it may have against other participants. The crosswaiver of liability clause may apply to the participants in the launch who are parties to the launch services agreement, but not apply to their subcontractors. The role of insurance in responding to many risks has been critical in assisting commercial space ventures grow. Today traditional property and liability insurance, such as pre-launch, launch and in-orbit insurance and third party liability insurance, have become mandatory parts of most space projects. The manufacture and pre-launch insurance covers direct physical loss or damage to the satellite, its apogee kick moter and including its related launch equipment from commencement of loading operations at the manufacture's plant until lift off. The launch and early orbit insurance covers the satellite for physical loss or damage from attachment of risk through to commissioning and for some period of initial operation between 180 days and 12 months after launch. The in-orbit insurance covers physical loss of or damage to the satellite occuring during or caused by an event during the policy period. The third party liability insurance covers the satellite owner' s liability exposure at the launch site and liability arising out of the launch and operation in orbit. In conclusion, the liability in commercial space ventures extends to any organization which participates in providing products and services used in the venture. Accordingly, it is essential for any organization participating in commercial space ventures to contractually disclaim its liability to the extent permitted by law. To achieve the effective disclaimers, it is necessary to determine the applicable law and to understand the requirements of the law which will govern the terms of the contract. A great deal of funds have been used in R&D for commercial space ventures to increase reliability, safety and success. However, the historical reliability of launches and success for commercial space ventures have proved to be slightly lower than we would have wished for. Space insurance has played an important role in reducing the high risks present in commercial space ventures.