• International Journal of Aeronautical and Space Sciences
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• v.15 no.2
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• pp.190-198
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• 2014

The pyramidal reaction wheel arrangement is one of the configurations that can be used in attitude control simulators for evaluation of attitude control performance in satellites. In this arrangement, the wheels are oriented in a pyramidal configuration with a tilt angle. In this paper, a study of pyramidal reaction wheel arrangement is carried out in order to find the optimum tilt angle that minimizes total power consumption of the system. The attitude control system is analyzed and the pyramidal configuration is implemented in numerical simulation. Optimization is carried out by using an iterative process and the optimum tilt angle that provides minimum system power consumption is obtained. Simulation results show that the system requires the least power by using optimum tilt angle in reaction wheels arrangement.

• Proceedings of the KIEE Conference
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• 1994.07a
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• pp.337-341
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• 1994

This paper describes an effort to develop a simulator of Active Power Filter (APF) by three dimentional(3-D) space current co-ordinate. System current is represented by 3-D vector composed of three current components - active, reactive and distorted. %THD (%Total Harmonics Distortion) can be converted to height-angle of system current vector and power factor can be defined on 3-D space current co-ordinate without loss of generality. Current of APF and power system can be analyzed by 3-D visualization of current vector trajectory. So, the computer simulation results show that the proposed method by 3-D space current co-ordinate make up for disadvantages of performance evaluation on time / frequency domain.

• Journal of information and communication convergence engineering
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• v.15 no.1
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• pp.21-27
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• 2017

In this research, we study the outage probability for distributed space-time coding-based cooperative (DSTC) systems with amplify-and-forward relaying over Rayleigh fading channels with a high temporal correlation where the direct link between the source and the destination is available. In particular, we derive the upper and lower bounds of the outage probability as well as their corresponding asymptotic expressions. In addition, using only the average channel powers for the source-to-relay and relay-to-destination links, we propose an efficient power allocation scheme between the source and the relay to minimize the asymptotic upper bound of the outage probability. Through a numerical investigation, we verify the analytical expressions as well as the effectiveness of the proposed efficient power allocation. The numerical results show that the lower and upper bounds tightly correspond to the exact outage probability, and the proposed efficient power allocation scheme provides an outage probability similar to that of the optimal power allocation scheme that minimizes the exact outage probability.

• International Journal of Aerospace System Engineering
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• v.3 no.2
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• pp.31-37
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• 2016

STEP Cube Lab (Cube Laboratory for Space Technology Experimental Project) is a 1U standardized pico-class satellite. Its main mission objective is an on-orbit verification of five fundamental core space technologies. For assuring a successful missions of the STEP Cube Lab with five payloads, electrical power subsystem (EPS) shall sufficiently provide an electrical power to payloads and bus systems of the satellite during an entire mission life. In this study, a design process of EPS system was introduced including power budget analysis considering a mission orbit and various mission modes of the satellite. In conclusion, adequate EPS hardware in compliance with design requirements were selected. The effectiveness and mission capability of EPS architecture design were confirmed through an energy balance analysis (EBA).

• Nuclear Engineering and Technology
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• v.53 no.2
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• pp.637-646
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• 2021

The free-piston Stirling engine (FPSE) has been widely used in aerospace owing to its advantages of high efficiency, high reliability, and self-starting ability. In this paper, a 20-kW FPSE is proposed by analyzing the requirements of space nuclear power reactor. A code was developed based on an improved simple analysis method to evaluate the performance of the proposed FPSE. The code is benchmarked with experimental data, and the maximum relative error of the output power is 17.1%. Numerical results show that the output power is 21 kW, which satisfies the design requirements. The results show that: a) reducing the pressure shell's thickness can improve the output power significantly; b) the system efficiency increases with the wire porosity, while the growth of system efficiency decreases when the porosity is higher than 80%, and system efficiency exhibits a linear relationship with the temperatures of the cold and hot sides; c) the system efficiency increases with the compression ratio; the compression ratio increases by 16.7% while the system efficiency increases by 42%. This study can provide valuable theoretical support for the design and analysis of FPSEs for space nuclear power reactors.

• Nuclear Engineering and Technology
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• v.56 no.5
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• pp.1584-1602
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• 2024

It is vital to output the required electrical power following various task requirements when the space reactor power supply is operating in orbit. The dynamic performance of the closed Brayton cycle thermoelectric conversion system is initially studied and analyzed. Based on this, a load tracking power regulation method is developed for the liquid metal cooled space reactor power system, which takes into account the inlet temperature of the lithium on the hot side of the intermediate heat exchanger, the filling quantity of helium and xenon, and the input amount of the heat pipe radiator module. After comparing several methods, a power regulation method with fast response speed and strong system stability is obtained. Under various changes in power output, the dynamic response characteristics of the ultra-small liquid metal lithium-cooled space reactor concept scheme are analyzed. The transient operation process of 70 % load power shows that core power variation is within 30 % and core coolant temperature can operate at the set safety temperature. The second loop's helium-xenon working fluid has a 65K temperature change range and a 25 % filling quantity. The lithium at the radiator loop outlet changes by less than ±7 K, and the system's main key parameters change as expected, indicating safety. The core system uses less power during 30 % load power transient operation. According to the response characteristics of various system parameters, under low power operation conditions, the lithium working fluid temperature of the radiator circuit and the high-temperature heat pipe operation temperature are limiting conditions for low-power operation, and multiple system parameters must be coordinated to ensure that the radiator system does not condense the lithium working fluid and the heat pipe.

• Journal of Power Electronics
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• v.6 no.1
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• pp.18-28
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• 2006

The Electric Power Subsystem (EPS) is one of the most critical systems on any satellite because nearly every subsystem requires power. This makes the choice of power systems the most important task facing satellite designers. The main purpose of the Satellite EPS is to provide continuous, regulated and conditioned power to all the satellite subsystems. It has to withstand radiation, thermal cycling and vacuums in hostile space environments, as well as subsystem degradation over time. The EPS power characteristics are determined by both the parameters of the system itself and by the satellite orbit. After satellite separation from the launch vehicle (LV) to its orbit, in almost all situations, the satellite subsystems (attitude determination and control, communication and onboard computer and data handling (OBC&DH)), take their needed power from a storage battery (SB) and solar arrays (SA) besides the consumed power in the EPS management device. At this point (separation point, detumbling mode), the satellite's angular motion is high and the orientation of the solar arrays, with respect to the Sun, will change in a non-uniform way, so the amount of power generated by the solar arrays will be affected. The objective of this research is to select satellite EPS component types, to estimate solar array illumination parameters and to determine the efficiency of solar arrays during both detumbling and normal operation modes.

• Bulletin of the Korean Space Science Society
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• 2004.10b
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• pp.231-234
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• 2004

Electrical grounding is defined as referencing an electrical circuit or a common reference plane for preventing shock hazards and for enhancing operability of the circuit and EMI control. In order to realize the best electrical grounding system of korean space launch vehicle, we should design the electrical grounding architecture of korean space launch vehicle of system-level at the earliest point in design procedure. To minimize the electrical grounding loop and the unnecessary electromagnetic interference or radiation among the electronic subsystems, we should establish the electrical grounding rules of the all electrical interfaces. The electrical interfaces among the electronic subsystems are generally classified into the electrical power and signal interfaces. Because of using the primary and secondary power system architecture in the korean space launch vehicle system such as the common space launch vehicle systems, we need to establish the electrical grounding rules between the primary and secondary power system. We also need to establish the electrical signal grounding interface rules among the electronic subsystems. In this paper, we will describe the grounding schemes of the common space launch vehicle system and propose a conceptual design for the electrical grounding architecture of korean space launch vehicle system.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.43 no.4
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• pp.318-325
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• 2015

Pico satellite has limited surface to install the solar cells due to its extremely limited size. Also, the sun incidence angle with respect to the solar panel continuously varies according to the attitude control strategy and its important parameter for the power generation. In this study, a concentrating photovoltaic system for pico satellite application has been proposed that can enhance the power generation efficiency in case of the unfavorable condition of the sun incidence angle with respect to the solar panel of the satellite using the fresnel lens. To prove the possibility of maximizing the power generation efficiency of the proposed concentrating power system, we have performed the power measurement test using a solar simulator and commercial fresnel lens. And on-orbit analysis of the power generation efficiency using the STK which is a commercial S/W has also been performed based on the test results.

• Current Industrial and Technological Trends in Aerospace
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• v.7 no.2
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• pp.33-39
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• 2009

On 1968 Dr. Peter Glaser introduced the concept of a large solar power satellite system in a high geosynchronous orbit for collection and conversion of solar energy into an electromagnetic microwave beam to transmit usable energy to rectennas on earth. With respect to it, U.S.A, Japan, E.U., etc. noted the Space Solar Power(SSP) as a future new energy resource, performed a substantial research and the concept design, and recently announced detailed plans for realizing SSP projects. While the new technology of SSP is developing, U.S.A. and Japan have a plan to provide the electric service by using SSP 2030. This paper presents the technology trend of advanced countries and the domestic strategies on the SSP development as a green energy and a new energy resource.