• Journal of Astronomy and Space Sciences
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• v.35 no.3
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• pp.195-200
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• 2018

The present paper describes the design of a Solid State Telescope (SST) on board the Korea Astronomy and Space Science Institute satellite-1 (KASISat-1) consisting of four [TBD] nanosatellites. The SST will measure these radiation belt electrons from a low-Earth polar orbit satellite to study mechanisms related to the spatial resolution of electron precipitation, such as electron microbursts, and those related to the measurement of energy dispersion with a high temporal resolution in the sub-auroral regions. We performed a simulation to determine the sensor design of the SST using GEometry ANd Tracking 4 (GEANT4) simulations and the Bethe formula. The simulation was performed in the range of 100 ~ 400 keV considering that the electron, which is to be detected in the space environment. The SST is based on a silicon barrier detector and consists of two telescopes mounted on a satellite to observe the electrons moving along the geomagnetic field (pitch angle $0^{\circ}$) and the quasi-trapped electrons (pitch angle $90^{\circ}$) during observations. We determined the telescope design of the SST in view of previous measurements and the geometrical factor in the cylindrical geometry of Sullivan (1971). With a high spectral resolution of 16 channels over the 100 keV ~ 400 keV energy range, together with the pitch angle information, the designed SST will answer questions regarding the occurrence of microbursts and the interaction with energetic particles. The KASISat-1 is expected to be launched in the latter half of 2020.

• Journal of Korean Society of Industrial and Systems Engineering
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• v.30 no.3
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• pp.165-173
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• 2007

A space launch vehicle system represents a typical example of large-scale multi-disciplinary systems, consisting of subsystems such as mechanical structure, electronics, control, telecommunication, propulsion, material engineering etc. A lot of cost is required to develop the launch vehicle system. A precise planning of R&D cost is very essential to make a success of the launch vehicle development program. Especially in the early development phase of a new space launch vehicle system, cost estimation techniques and analogy from past similar development data are very useful methods to estimate a development cost of the launch vehicle system. Now Korea Aerospace Research Institute is in charge of the KSLV-I (Korea Space Launch Vehicle-I) Program that is a part of Korea National Space program. KSLV-I Program is a national undertaking to develop launch capabilities to deliver science satellites of a 100kg-class into a low earth orbit. It is hereafter, going to plan to develop a new korean space launch vehicle. In this paper, first the development costs of well-known launch vehicles in the world are presented to provide a reference to make a development plan of a new launch vehicle. Second this paper introduces the present status of cost estimation applications at NASA. Finally this paper presents the results from application of a TRANSCOST, a parametric cost model, to derive a cost estimate of a new launch vehicle development, as an example.

• Journal of Astronomy and Space Sciences
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• v.26 no.1
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• pp.31-46
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• 2009

An Unscented Kalman Filter (UKF) for estimation of the attitude and rate of a spacecraft using only magnetometer vector measurement is developed. The attitude dynamics used in the estimation is the nonlinear Euler's rotational equation which is augmented with the quaternion kinematics to construct a process model. The filter is designed for small satellite in low Earth orbit, so the disturbance torques include gravity-gradient torque, magnetic disturbance torque, and aerodynamic drag torque. The magnetometer measurements are simulated based on time-varying position of the spacecraft. The filter has been tested not only in the standby mode but also in the detumbling mode. Two types of actuators have been modeled and applied in the simulation. The PD controller is used for the two types of actuators (reaction wheels and thrusters) to detumble the spacecraft. The estimation error converged to within 5 deg for attitude and 0.1 deg/s for rate respectively when the two types of actuators were used. A joint state parameter estimation has been tested and the effect of the process noise covariance on the parameter estimation has been indicated. Also, Monte-Carlo simulations have been performed to test the capability of the filter to converge with the initial conditions sampled from a uniform distribution. Finally, the UKF performance has been compared to that of the EKF and it demonstrates that UKF slightly outperforms EKF. The developed algorithm can be applied to any type of small satellites that are actuated by magnetic torquers, reaction wheels or thrusters with a capability of magnetometer vector measurements for attitude and rate estimation.

• Journal of Astronomy and Space Sciences
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• v.34 no.4
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• pp.303-314
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• 2017

With increased human activity in space, the risk of re-entry and collision between space objects is constantly increasing. Hence, the need for space situational awareness (SSA) programs has been acknowledged by many experienced space agencies. Optical and radar sensors, which enable the surveillance and tracking of space objects, are the most important technical components of SSA systems. In particular, combinations of radar systems and optical sensor networks play an outstanding role in SSA programs. At present, Korea operates the optical wide field patrol network (OWL-Net), the only optical system for tracking space objects. However, due to their dependence on weather conditions and observation time, it is not reasonable to use optical systems alone for SSA initiatives, as they have limited operational availability. Therefore, the strategies for developing radar systems should be considered for an efficient SSA system using currently available technology. The purpose of this paper is to analyze the performance of a radar system in detecting and tracking space objects. With the radar system investigated, the minimum sensitivity is defined as detection of a $1-m^2$ radar cross section (RCS) at an altitude of 2,000 km, with operating frequencies in the L, S, C, X or Ku-band. The results of power budget analysis showed that the maximum detection range of 2,000 km, which includes the low earth orbit (LEO) environment, can be achieved with a transmission power of 900 kW, transmit and receive antenna gains of 40 dB and 43 dB, respectively, a pulse width of 2 ms, and a signal processing gain of 13.3 dB, at a frequency of 1.3 GHz. We defined the key parameters of the radar following a performance analysis of the system. This research can thus provide guidelines for the conceptual design of radar systems for national SSA initiatives.

• The Transactions of the Korean Institute of Power Electronics
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• v.11 no.6
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• pp.535-542
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• 2006

A large signal stability analysis of the solar array regulator system is performed to facilitate the design and analysis of a Low-Earth-Orbit satellite power system. The effective load characteristics of every controllable method in the solar array system are classified to analyze the large signal stability. Then, using the state plane analysis technique, the stability of various equilibrium points is analyzed. A nonlinear transformation algorithm, which changes the effective load characteristic of the solar array regulator as constant resistive load, is also proposed for the large signal stability. The proposed resistive current mode control system can control the solar array output for purposes such as peak power tracking control and battery charging control. For the verification of the proposed large signal analysis and resistive current mode control, a solar array regulator system consisting of two 100W parallel module buck converters has been built and tested using a real 200W solar array.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2008.03a
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• pp.330-339
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• 2008

Ground-based interceptors(GBI) comprise a major element of the strategic defense against hostile targets like Intercontinental Ballistic Missiles(ICBM) and reentry vehicles(RV) dispersed from them. An optimum design of the subsystems is required to increase the performance and reliability of these GBI. Propulsion subsystem design and optimization is the motivation for this effort. This paper describes an effort in which an entire GBI missile system, including a multi-stage solid rocket booster, is considered simultaneously in a Genetic Algorithm(GA) performance optimization process. Single goal, constrained optimization is performed. For specified payload and miss distance, time of flight, the most important component in the optimization process is the booster, for its takeoff weight, time of flight, or a combination of the two. The GBI is assumed to be a multistage missile that uses target location data provided by two ground based RF radar sensors and two low earth orbit(LEO) IR sensors. 3Dimensional model is developed for a multistage target with a boost phase acceleration profile that depends on total mass, propellant mass and the specific impulse in the gravity field. The monostatic radar cross section (RCS) data of a three stage ICBM is used. For preliminary design, GBI is assumed to have a fixed initial position from the target launch point and zero launch delay. GBI carries the Kill Vehicle(KV) to an optimal position in space to allow it to complete the intercept. The objective is to design and optimize the propulsion system for the GBI that will fulfill mission requirements and objectives. The KV weight and volume requirements are specified in the problem definition before the optimization is computed. We have considered only continuous design variables, while considering discrete variables as input. Though the number of stages should also be one of the design variables, however, in this paper it is fixed as three. The elite solution from GA is passed on to(Sequential Quadratic Programming) SQP as near optimal guess. The SQP then performs local convergence to identify the minimum mass of the GBI. The performance of the three staged GBI is validated using a ballistic missile intercept scenario modeled in Matlab/SIMULINK.

• KSII Transactions on Internet and Information Systems (TIIS)
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• v.13 no.3
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• pp.1260-1283
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• 2019

Low earth orbit (LEO) satellite broadband network is a crucial part of the space information network. LEO satellite constellation design is a top-level design, which plays a decisive role in the overall performance of the LEO satellite network. However, the existing works on constellation design mainly focus on the coverage criterion and rarely take network performance into the design process. In this article, we develop a unified framework for constellation optimization design in LEO satellite broadband networks. Several design criteria including network performance and coverage capability are combined into the design process. Firstly, the quality of service (QoS) metrics is presented to evaluate the performance of the LEO satellite broadband network. Also, we propose a network stability model for the rapid change of the satellite network topology. Besides, a mathematical model of constellation optimization design is formulated by considering the network cost-efficiency and stability. Then, an optimization algorithm based on non-dominated sorting genetic algorithm-II (NSGA-II) is provided for the problem of constellation design. Finally, the proposed method is further evaluated through numerical simulations. Simulation results validate the proposed method and show that it is an efficient and effective approach for solving the problem of constellation design in LEO satellite broadband networks.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.49 no.2
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• pp.155-166
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• 2021

In this paper, design study of a small two-stage launch vehicle is undertaken for the dedicated launch of the Compact Advanced Satellite 500 (CAS500)-class satellite into the Low Earth Orbit (LEO) by modifying the second and third stages of the Korean Space Launch Vehicle II (KSLV-II). Since the KSLV-II has three stages, velocity increment is newly distributed for the two-stage small launch vehicle. For this end, the staging design is carried out for the design parameters such as stage mass ratios, structural coefficients and engine options for each stage followed by trajectory analysis. Investigation of the results provides the combination of design parameters for the small launch vehicle for the dedicated launch of 500 kg-class satellite into LEO.

• Journal of Space Technology and Applications
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• v.2 no.1
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• pp.1-12
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• 2022

In this paper, we described the development of a drag augmentation device for nanosatellite. Recently, space industry has entered the New Space era, and barriers to entry into Low Earth Orbit (LEO) for artificial objects such as small rockets and nanosatellite mega constellations have been significantly lowered. As a result, the number of space debris is increasing exponentially, and it is approaching as a major threat to satellite currently in operation as well as satellites to be launched in near future. To prevent this, international organizations like Inter-Agency Space Debris Coordination Committee (IADC) have been proposed space debris mitigation guidelines. The Korea Aerospace Research Institute (KARI) conducted KARI Rendezvous & Docking demonstration SATellite (KARDSAT) project, the first nanosatellites for rendezvous and docking technology demonstration in Korea, and we also developed drag augmentation device for KARDSAT Target nanosatellite that complied with the international guideline of post-mission disposal.

• Journal of Space Technology and Applications
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• v.2 no.2
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• pp.121-136
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• 2022

Automatic Identification System used in ship communication is required for marine control way, including monitoring of vessel operation in coastal and exchanging of information for safety navigation between them. But, it uses a very high frequency band of approximately 160 MHz, and at the same time, due to the curvature of Earth, there is a limit to the communication distance. Several demonstrations were made successfully over satellite, but not much work has been done yet through cube-satellite which has low-orbit at 500 km altitude. Here, we demonstrate a reception test of AIS (automatic identification system) receiver for a cube-satellites using software-defined radio (SDR). We collected AIS data from ship at port of Busan, Korea, using R8202T2 SDR and established to transmit them using Adam-Pluto and Matlab Simulink. The process of weakening the signal strength to a satellite was constructed using attenuator. Through above process, we demonstrated whether AIS data was successfully received from the AIS payload.