• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.40 no.9
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• pp.752-759
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• 2012

This paper presents a mission orbit design method for spacecraft which use the sun-synchronous and ground repeat orbits. In this work, we have proposed a new design procedure, "Nonlinear simulation-based numerical optimization technique" using the commercial S/W's such as STK (Satellite Tool kit) and Matlab, which are widely adopted S/W's in the area of orbital mechanics and engineering analysis. Inclusion of all the perturbation effects on the spacecraft not only can more precisely satisfy the mission requirements for sun-synchronicity and repeated ground track, and also operational requirements such as minimum change in the S/C local time, maximization of the contact time with a specified ground station, etc. can be appropriately considered. Design examples for LEO sun-synchronous mission are presented to demonstrate the usefulness of the proposed method in this paper.

• International Journal of Aeronautical and Space Sciences
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• v.3 no.2
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• pp.88-94
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• 2002

This paper describes attitude determination algorithm for the low earth orbit(LEO) spacecraft using stellar inertial sensors. The cascaded gyro/star tracker extended Kalman filter is constructed to fuse two sensor data. And then the smoothing of the measurement are proposed for an unreasonable jump of star tracker. The smoothing algorithm for the rejection of star tracker error jumps is designed by scalar adaptive filter. The proposed algorithms operate to process the measurement of gyro/star tracker Kalman filter, therefore, it is comparatively simple to apply these methods to other integration systems. Simulations to gyro/star tracker integrated system show that the proposed method is effective.

• Bulletin of the Korean Space Science Society
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• 2003.10a
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• pp.47-47
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• 2003

Analytical formulae are presented to approximate the evolution of the semi major axis, the maneuver time, and the final mass fraction for low thrust orbital transfers with circular initial orbit, circular target orbit, and constant thrust directed either always along or always opposite the velocity vector. For comparison, the associated results for high-thrust transfers, i.e. the two-impulse Hohmann transfer, are summarized. All results are implemented in a computer code designed to analyze planar planetary and interplanetary space missions. This implementation yields fast and reasonably accurate approximations to trajectory performance boundaries. Consequently, the approach can provide trajectory analysis for each spacecraft configuration during the conceptual space mission design phase. As an example, a mission from Low-Earth Orbit (LEO) to Jupiter's moon Europa is analyzed.

• 제어로봇시스템학회:학술대회논문집
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• 1995.10a
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• pp.110-114
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• 1995

This paper presents the on-board attitude determination algorithm for LEO (Low Earth Orbit) three-axis stabilized spacecraft. Two advanced star trackers and a three-axis Inertial Reference Unit (IRU) are assumed to be attitude sensors. The gyro in the IRU provides a direct measurement of the attitude rates. However, the attitude estimation error increases with time due to the gyro drift and noise. An update filter with measurements of star trackers and/or sun sensor is designed to update these gyro drift bias and to compensate the attitude error. Kalman Filter is adapted for the on-board update filter algorithm. Simulation results will be presented to investigate the attitude pointing performance.

• Bulletin of the Korean Space Science Society
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• 2004.10b
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• pp.376-379
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• 2004

Operational orbit determination (OOD) depends on the capability of generating accurate prediction of spacecraft ephemeris in a short period. The predicted ephemeris is used in the operations such as instrument pointing and orbit maneuvers. In this study the orbit prediction problem consists of the estimating diverse arc length orbit using GPS navigation data, the predicted orbit for the next 48 hours, and the fitted 30-hour arc length orbits of double differenced GPS measurements for the predicted 48-hour period. For 24-hour orbit arc length, the predicted orbit difference from truth orbit was 205 meters due to the along-track error. The main error sources for the orbit prediction of the Low Earth Orbiter (LEO) satellite are solar pressure and atmosphere density.

• Advances in aircraft and spacecraft science
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• v.11 no.1
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• pp.23-31
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• 2024

This paper addresses the enhancement of long-endurance solar-powered aircraft through the integration of a rechargeable battery and Maximum Power Point Tracking (MPPT) controller. Traditional long-endurance aircraft often rely on non-renewable energy sourcessuch as batteries orjetfuel, contributing to carbon emissions. The proposed system aims to mitigate these environmental impacts by harnessing solar energy and efficiently managing its storage and utilization. The MPPT controller optimizes the power output of photovoltaic cells, enabling simultaneous charging and discharging of the battery for propulsion and servo control. A prototype is presented to illustrate the practical implementation and functionality of the proposed design, marking a promising step towards more sustainable and enduring solar-powered flight.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.37 no.7
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• pp.669-677
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• 2009

In general, the pitch momentum biased system that induces inherently nutational motion in roll/yaw plane, has been adapted for geosynchronous communications satellites. This paper discusses the method of roll attitude control using yaw axis momentum management method for a low earth orbit(LEO) satellite which is a pitch momentum biased system equipped with only two reaction wheels. The robustness of wheel momentum management method with PI-controller is investigated comparing with wheel torque control method. The transfer function of roll/yaw axis momentum management system that is useful for attitude controller design is derived. The disturbance effect of roll/yaw axis momentum management system for attitude control is investigated to identify design parameters such as magnitude of momentum bias and to get the insight for controller design. As an example, the PID controller design result of momentum management system for roll/yaw axis control is provided and the simulation results are presented to provide further physical insight into the momentum management system.

• Journal of Astronomy and Space Sciences
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• v.15 no.2
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• pp.459-473
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• 1998

As the satellite system becomes more complex, the probability of unpredictable failures may be increased due to design inadequacy, experience deficiency, lack of problem recognition. Poor quality control, improper testing, and workmanship fault. Consequently, these problems can lead to the reduction or end of the satellite mission lifetime. This article addresses general satellite failure modes and factors influencing satellite mission life. The mission life factors of LEO sun-synchronous KOMPSAT spacecraft are investigated, in which its mission life is predicted based on these factors. Since the end of mission due to random failures is not predictable, the predictable mission life factors such as power budget, propellant budget, battery charging/discharging cycle, radiation effects payload reliability, single point failure, and redundancy are primarily investigated.

• Proceedings of the KSRS Conference
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• v.2
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• pp.791-793
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• 2006

In the satellite camera, the incoming light source is converted to electronic analog signals by the electronic component for example CCD (Charge Coupled Device) detectors. The analog signals are amplified, biased and converted into digital signals (pixel data stream) in the video processor (A/Ds). The outputs of the A/Ds are digitally multiplexed and driven out using differential line drivers (two pairs of wires) for cross strap requirement. The MSC (Multi-Spectral Camera) in the KOMPSAT-2 which is a LEO spacecraft will be used to generate observation imagery data in two main channels. The MSC is to obtain data for high-resolution images by converting incoming light from the earth into digital stream of pixel data. The video data outputs are then MUXd, converted to 8 bit bytes, serialized and transmitted to the NUC (Non-Uniformity Correction) module by the Hotlink data transmitter. In this paper, the video data streams, the video data format, and the image data processing routine for satellite camera are described in terms of satellite camera control hardware. The advanced satellite with very high resolution requires faster and more complex image data chain than this algorithm. So, the effective change of the used image data chain and the fast video data transmission method are discussed in this paper

• 제어로봇시스템학회:학술대회논문집
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• 1999.10a
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• pp.37-40
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• 1999

ETRI has successfully completed and delivered to KARI the KOMPSAT Mission Control System. This system was designed to work in the conventional TT&C modulation scheme with the pre-assigned frequencies. As a way to accelerate in catching up with future TT&C technology evolutions, a preliminary study needs to be carried out to prepare for the development of a spread spectrum applicable to TT&C. A brief study was carried out to review some points to be considered in designing and implementing spread spectrum schemes to the ground TT&C system intended for a LEO spacecraft. Also a simulation and link design revisit was performed to see the operational and technical benefits with the KOMPSAT TT&C parameters. An experiment concept is proposed to test as many functions at a time once the prototype is developed. In this configuration, a ground-model TT&C transponder is connected via LAN to the ETRI-developed KOMPSAT S/W simulator and linked to the KOMPSAT TM/TC processing s/w via spread spectrum signals through a GEO satellite bent-pipe link. A satellite data relay link simulation could be carried out in this configuration.