• Journal of Space Technology and Applications
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• v.4 no.1
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• pp.27-47
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• 2024

In the New space era, satellites are being developed to perform on-orbit service (OOS) missions. Various missions for orbital service include failure repair, refueling, towing, component replacement, and space construction, and in order to do so, a robot arm payload must be mounted. Unlike conventional satellite payloads, the robot arm payload is not move in a fixed state, but is a payload that must move continuously to perform the mission. It is also characterized by the need to perform the mission while being directly exposed to outer space, rather than existing inside the structure of the satellite. Due to the characteristics of these payloads, thermal design and interpretation that can be operated smoothly in an extreme space thermal environment is essential, but there are not many papers on thermal design and interpretation of the robot arm. This paper introduces and summarizes cases of thermal design and interpretation of robot arm payloads developed so far, and finally, it intends to suggest directions for thermal design and interpretation of robot arm payloads to be developed in the future.

• Proceedings of the KSRS Conference
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• 2005.10a
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• pp.493-496
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• 2005

The PMU (Payload Management Unit) in MSC (Multi-Spectral Camera) is the main subsystem for the management, control and power supply of the MSC payload operation. The PMU shall handle the communication with the BUS (Spacecraft) OBC (On Board Computer) for the command, the telemetry and the communications with the various MSC units. The PMU will perform that distributes power to the various MSC units, collects the telemetry reports from MSC units, performs thermal control of the EOS (Electro-Optical Subsystem), performs the NUC (Non-Uniformity Correction) function of the raw imagery data, and rearranges the pixel data and output it to the DCSU (Data Compression and Storage Unit). The BUS provides high voltage to the MSC. The PMU is connected to primary and redundant BUS power and distributes the high unregulated primary voltages for all MSC sub-units. The PSM (Power Supply Module) is an assembly in the PMU implements the interface between several channels on the input. The bus switches are used to prevent a single point system failure. Such a failure could need the PSS (Power Supply System) requirement to combine the two PSM boards' bus outputs in a wired-OR configuration. In such a configuration if one of the boards' output gets shorted to ground then the entire bus could fail thereby causing the entire MSC to fail. To prevent such a short from pulling down the system, the switch could be opened and disconnect the short from the bus. This switch operation is controlled by the BUS.

• Advances in aircraft and spacecraft science
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• v.4 no.5
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• pp.503-514
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• 2017

The attitude control of an aircraft is usually fulfilled by means of thrusters at high altitudes. Therefore, the possibility of using also aerodynamic surfaces would produce the advantage of reducing the amount of fuel for the thrusters to be loaded on board. For this purpose, Zuppardi already considered some aerodynamic problems linked to the use of a wing flap in a previous paper. A NACA 0010 airfoil with a trailing edge flap of 35% of the chord, in the range of angle of attack 0-40 deg and flap deflections up to 30 deg was investigated. Computer tests were carried out in hypersonic, rarefied flow by a direct simulation Monte Carlo code at the altitudes of 65 and 85 km of Earth Atmosphere. The present work continues this subject, considering the same airfoil and free stream conditions but two flap extensions of 45% and 25% of the chord and two flap deflections of 15 and 30 deg. The main purpose is to compare the influence of the flap dimension with that of the flap deflection. The present analysis is carried out in terms of: 1) percentage variation of the global aerodynamic coefficients with respect to the no-flap configuration, 2) increment of pressure and heat flux on the airfoil lower surface due to the Shock Wave-Shock Wave Interaction (SWSWI) with respect to the same quantities with no SWSWI or in no-flap configuration, 3) flap hinge moment. Issues 2) and 3) are important for the design of the mechanical and thermal protection system and of the flap actuator, respectively. Under the above mentioned test and geometrical conditions, the flap deflection is aerodynamically more effective than the flap extension, because it involves higher variation of the aerodynamic coefficients. However, tests verify that a smaller deflection angle involves the advantage of a smaller increment of pressure and heat flux on the airfoil lower surface, due to SWSWI, as well as a smaller hinge moment.

• Journal of Astronomy and Space Sciences
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• v.25 no.4
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• pp.471-480
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• 2008

COMS Image Data Acquisition and Control System (IDACS) plays a key role in real time ground processing of Meteorological and Ocean observation data. Beyond processing, it serves processed image data and additional data to end users through the spacecraft in the internationally recommended format. The IDACS will be installed at three location (MSC, KOSC, and SOC) and automatically operated 24h/365days. After the IDACS subsystem tests and inter -subsystem interface tests had been completed in the first half of 2008, the acceptance test which was a comprehensive test performed as an integrated form to verify function performance and operational requirements. This paper introduces test objective, preparation, and major result of the COMS IDACS acceptance test.

• The Journal of Korean Institute of Electromagnetic Engineering and Science
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• v.25 no.3
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• pp.311-318
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• 2014

Electromagnetic compatibility(EMC) performance of the first Korea space launch vehicle(KSLV-I) should be ensured and verified in order to guarantee the normal operation among the spacecraft, ground facilities which are installed in the space center, and other wireless communication networks. For the purpose of the EMC performance verification, pertinent EMC test specifications, methods, and procedures for both the subsystems and the system should be established in consideration of operational properties and electromagnetic environmental effects. And it is required to maintain and control the EMC properties consistently in accordance with the determined specifications up to the program closing phase. In this paper, sequential management work conducted during the overall development process of the KSLV-I is explained, and not only the phased EMC test plan for each model of the KSLV-I and its subsystem but also test method, specification, and results of the verification tests are presented. And also, multipaction analysis results are presented.

• The Bulletin of The Korean Astronomical Society
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• v.35 no.1
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• pp.30.2-30.2
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• 2010

Numerous operational anomalies and satellite failures have been reported since the beginnings of the "space age". Space weather effects on modern spacecraft systems have been emphasized more and more as increasing their complexity and capability. Energetic particles potentially can destroy and degrade electronic components in satellites. We analyzed the geostationary (GEO) satellite anomalies during 1997-2009 to search possible influences of space weather on the satellite anomalies like power problem, control processor problem, attitude control problem, etc. For this we use particle data from GOES and LANL satellites to investigate space weather effects on the GEO satellites' anomalies depending on Kp index, local time, seasonal variation, and high-energy electron contribution. As results, we obtained following results: (1) there is a good correlation between geomagnetic index(Kp) and anomaly occurrences of the GEO satellite; (2) especially during the solar minimum, occurrence of the satellite anomalies are related to electron flux increase due to high speed solar wind; (3) satellite anomalies occurred more preferentially in the midnight and dawn sector than noon and dusk sector; (4) and the anomalies occurred twice more in Spring and Fall than Summer and Winter; (5) the electron with the lowest energy channel (50-75keV) has the highest correlation (cc=0.758) with the anomalies. High association between the anomalies and the low energy electrons could be understand by the facts that electron fluxes in the spring and fall are stronger than those in the summer and winter, and low-energy electron flux is more concentrated in the dawn sector where the GEO satellite anomalies occurred more frequently than high-energy electron flux. While we could not identify what cause such local time dependences, our results shows that low-energy electrons (~100keV) could be main source of the satellite anomaly, which should be carefully taken into account of operating satellites.

• Advances in aircraft and spacecraft science
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• v.7 no.6
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• pp.553-570
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• 2020

To address the problems caused by the strong coupling of an airbreathing hypersonic vehicle's airframe and propulsion to the integrated control system design, an integrated airframe-propulsion model is established, and the coupling characteristics between the aircraft and engine are analyzed. First, the airframe-propulsion integration model is established based on the typical nonlinear longitudinal dynamical model of an air-breathing hypersonic vehicle and the one-dimensional dual-mode scramjet model. Thrust, moment, angle of attack, altitude, and velocity are used as transfer variables between the aircraft model and the engine model. The one-dimensional scramjet model can accurately reflect the working state of the engine and provide data to support the coupling analysis. Second, owing to the static instability of the aircraft model, the linear quadratic regulator (LQR) controller of the aircraft is designed to ensure attitude stability and height tracking. Finally, the coupling relationship between the aircraft and the engine is revealed through simulation examples. The interaction between vehicle attitude and engine working condition is analyzed, and the influence of vehicle attitude on engine safety is considered. When the engine is in a critical working state, the attitude change of the aircraft will not affect the engine safety without considering coupling, whereas when coupling is considered, the attitude change of the aircraft may cause the engine unstart, which demonstrates the significance of considering coupling characteristics.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.43 no.10
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• pp.868-874
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• 2015

Thrust variation is an essential parameter in a space mission such as landing on an atmosphereless planet or docking a spacecraft. In order to achieve the thrust variation control, using throttleable injector is a representative and general method. A dual-manifold injector, one of throttleable injectors, was used to control mass flow rate. Five kinds of injectors were designed and investigated in order to compare the spray characteristics of the dual-manifold injector with various tangential entries. Spray angles and patterns were measured to determine external flow characteristics and film thicknesses were measured in order to investigate the internal flow patterns.

• Aerospace Engineering and Technology
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• v.1 no.2
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• pp.57-65
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• 2002

The Electrical Power System (EPS) shall supply required power to maintain spacecraft and payload during the mission. The EPS sizing are based on space environment, satellite mission and lifetime, and allocated budgets. The type of the primary and secondary power is determined according to satellite design-level and allocated subsystem budgets. The design of EPS has closely related to system and others' subsystems design. To supply the sufficient power to the satellite, the implementation of the larger power source and energy storage is impossible actually. And there will be some problems of the attitude control of the satellite, the handling power capability of the electronic boxes, and launch vehicle selection caused by EPS oversizing. Also, the thermal control is not easy in the space by extra power. And the maintenance of the satellite within the specific orbit from orbit-drag is a big design burden of the thruster. So the various technologies have been developed to optimize the EPS sizing and to operate the power system efficiently.

• Aerospace Engineering and Technology
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• v.4 no.2
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• pp.107-116
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• 2005

As a unique payload of Komsat-2, MSC, comprising EOS(Electro-Optical Sub-system), PMU(Payload Management Unit) and PDTS(Payload Data Transmission Sub-system), is supposed to take pictures of one panchromatic and 4 multi-spectral image between wavelength 450mm~900mm, and is being under final Satellite I&T. It will perform the earth remote sensing with applications such as acquisition of high resolution images, surveillance of large scale disasters and its countermeasure, survey of natural resources, etc.. Under the hostile influence of the extreme space environmental conditions due to deep space and direct solar flux, the thermal design is especially of major importance in designing a payload. There are tight temperature range restrictions for electro-optical elements while on the other hand there are low power consumption requirements due to the limited energy source on the spacecraft. This paper describes details of thermal control system for MSC.