• 한국군사과학기술학회지
• /
• 제23권1호
• /
• pp.18-27
• /
• 2020

In order to estimate the launch point of a high-speed vehicle, predicting the various characteristics of the vehicle's movement, such as drag and thrust, must be preceded by the estimation. To predict the various parameters regarding the vehicle's characteristics, we build the IMM filter specialized in predicting the parameters of the post-launch phase based on flight dynamics. Then we estimate the launch point of the high-speed vehicle using Inverse Dynamics. In addition, we assume the arbitrary error level of the radar for accuracy of the prediction. We organize multiple-dimensioned IMM structures, and figure out the optimal value of parameters by comparing the various IMM structures. After deriving the optimal value of parameters, we verify the launch point estimation error under certain error level.

• 제어로봇시스템학회논문지
• /
• 제20권4호
• /
• pp.414-421
• /
• 2014

This paper presents a method of ballistic missile launch point estimation using phase division least squares. The proposed algorithm employs smoothing to enhance estimation accuracy and generates functions of time for total velocity, flight path angle and heading angle, allowing extrapolation to estimate the launch point. Performance of the proposed algorithm is tested in conjunction with the extended Kalman filter and the Kalman filter.

• 아태비즈니스연구
• /
• 제12권1호
• /
• pp.151-163
• /
• 2021

Purpose - The break-even point refers to the point where total profit and total cost coincide, and from this point on, the entrepreneur's decision-making takes a different route. Strategic decisions can be made for more efficient operation and eventually for more likelihood for growth and sustainability if a startup figures out when it recoups the investment and switches to a net profit. Design/methodology/approach - 748 creativity and skill-based sole proprietors in manufacturing industry were examined to demonstrate the effect of the entrepreneur's entrepreneurial experience and education level, the business launch preparation time, or the self-financing on the achievement of break-even point. Findings - While the business launch preparation time lowered the likelihood of reaching a break-even point, self-financing increased the likelihood. As a result of further analysis by subdividing into subgroups according to skill level, only the business launch preparation time was statistically significant in the highly skilled industries. In the low skilled industries, in addition to the business launch preparation time, the CEO's education level and the self-financing were statistically significant. Research implications or Originality - The longer the business launch preparation time, the higher the start-up cost, which increases the burden of initial cost recovery, and the agile response to market changes is thereby delayed, resulting in the business idea losing its appeal. Self-financing not only provides stability and strong motivation for the business operation but also promotes careful spending which contributes to the achievement of break-even point. In particular, it is found that practical experience is more useful than theoretical knowledge in low skilled industries. Due to the limitation of secondary data based on the recollection, the time required to reach a break-even point, percentage of financing sources, etc. may include cognitive errors. In addition, variables are not included that explain the characteristics of creativity and skill-based sole proprietorship, so it is necessary to exercise caution with the actual application.

• Journal of Astronomy and Space Sciences
• /
• 제38권2호
• /
• pp.145-155
• /
• 2021

In this work, preliminary launch opportunities from NARO Space Center to the Sun-Earth Lagrange point are analyzed. Among five different Sun-Earth Lagrange points, L1 and L2 points are selected as suitable candidates for, respectively, solar and astrophysics missions. With high fidelity dynamics models, the L1 and L2 point targeting problem is formulated regarding the location of NARO Space Center and relevant Target Interface Point (TIP) for each different launch date is derived including launch injection energy per unit mass (C3), Right ascension of the injection orbit Apoapsis Vector (RAV) and Declination of the injection orbit Apoapsis Vector (DAV). Potential launch periods to achieve L1 and L2 transfer trajectory are also investigated regarding coasting characteristics from NARO Space Center. The magnitude of the Lagrange Orbit Insertion (LOI) burn, as well as the Orbit Maintenance (OM) maneuver to maintain more than one year of mission orbit around the Lagrange points, is also derived as an example. Even the current work has been made under many assumptions as there are no specific mission goals currently defined yet, so results from the current work could be a good starting point to extend diversities of future Korean deep-space missions.

• 한국시뮬레이션학회:학술대회논문집
• /
• 한국시뮬레이션학회 2005년도 춘계학술대회 논문집
• /
• pp.106-111
• /
• 2005

LCC(Launch Control Center) in NARO Space Center perform a data monitoring and control through the interface to the external system of launch vehicle. Launch Control function needs a high reliability and processing speed. Hence, LCC's remote control system configure a real time system. An important role of the Simulation system is discovering a risk element and minimize it When developing a launch control system. Also, secure a development technique to solve the risks. Launch Vehicle simulator is composed of various component at characteristic of the Launch Vehicle. To be like this each function component the developer will be able to develop easily in order, it using the LabVIEW which is a Graphical Program and it programs, The LabVIEW GOOP(Graphical Object-orinted Programming) which supports an Object-orinted programming it uses with the Component it develops will have a strong point which reusability and a unit test, maintenance, size of program and individual developments.

• 항공우주시스템공학회지
• /
• 제1권3호
• /
• pp.37-42
• /
• 2007

The flight safety system for the satellite launch vehicles is required in order to minimize the risk due to launch vehicle failure. For prompt and reliable decision of flight termination, the flight safety system usually uses multiple sensors to estimate launch vehicle's flight trajectory. In that case, multiple types of observed tracking data makes it difficult to identify the flight termination condition. Therefore, a fusion tracking filter handling the multiple tracking data is necessary for the flight safety system. This research developed a simulation software for generating multiple types of launch vehicle tracking data, and then processed the data with fusion filters.

• Journal of Astronomy and Space Sciences
• /
• 제16권2호
• /
• pp.199-208
• /
• 1999

In this paper, KOMPSAT satellite launch and deployment operations are discussed. The U.S. Taurus launch vehicle delivers KOMPSAT satellite into the mission orbit directly. Launch and deployment operations is monitored and controlled by several international ground stations including Korean Ground Station (KGS). After separation from launch vehicle, KOMPSAT spacecraft deploys solar array by on-board autonomous stored commands without ground inter-vention and stabilizes the satellite such that solar arrays point to the sun. Autonomous ground communication is designed for KOMPSAT for the early orbit ground contact. KOMPSAT space-craft has capability of handing contingency situation by on-board fault management design to retry deployment sequence.

• Journal of information and communication convergence engineering
• /
• 제7권3호
• /
• pp.372-376
• /
• 2009

In this paper, new and simple optical transmission link with fixed dispersion management (DM) scheme, i.e., pre(post) compensation and residual dispersion per span (RDPS) are fixed to net residual dispersion (NRD) = 0 ps/nm, and optical phase conjugator (OPC) having optimal position depending on launch power in WDM transmission system is proposed. Also, effective launch power range of WDM channels resulting 1 dB eye opening penalty (EOP) is induced as a function of OPC position. First, it is confirmed that, for applying DM into WDM transmission link fixed pre(post)compensation and RDPS, which are independence on exact system parameters except launch power, sufficiently are used in WDM links, but OPC with optimal position is needed for effective compensating impairments of WDM channels. And, it is confirmed that effective launch power is broader in case of RDPS = 100 ps/nm than in RDPS = 50 ps/nm. But, it is shown that the best OPC position offset is -0.6 km from a point of view of power window, which is defined as difference between maximum and minimum effective launch power.

• International Journal of Aeronautical and Space Sciences
• /
• 제11권2호
• /
• pp.118-125
• /
• 2010

This paper describes a mid-course guidance strategy based on the earliest intercept geometry (EIG) guidance. An analytical solution and performance validation will be addressed for generalized mid-course guidance problems in area air-defence in order to improve reachability and performance. The EIG is generated for a wide range of possible manoeuvres of the challenging missile based on the guidance algorithm using differential geometry concepts. The main idea is that a mid-course guidance law can defend the area as long as it assures that the depending area and objects are always within the defended area defined by EIG. The velocity of Intercept Point in EIG is analytically derived to control the Intercept Geometry and the defended area. The proposed method can be applied in deciding a missile launch window and launch point for the launch phase.

• 제어로봇시스템학회:학술대회논문집
• /
• 제어로봇시스템학회 1989년도 한국자동제어학술회의논문집; Seoul, Korea; 27-28 Oct. 1989
• /
• pp.59-64
• /
• 1989

In this paper, the launch trajectory of the Japan scientific satellite M-3H-3 from launch to orbit injection is investigated. For the terminal conditions at a guidance target point, a guidance and control system is used. An open-loop and a closed-loop guidance schemes are used simultaneously. For the closed-loop guidance scheme, the velocity polynomial algorithm represented by the velocity difference between the target point and present velocity is used. A PD control system is used for activating gimbal type engines. The simulation result shows that all the terminal position and velocity conditions are satisfied and the trajectory for the M-3H-3 scientific satellite is reasonable.