Mission Analysis for Engineering Test satellite, Kitsat -3
Sungdong Park, Sungheon Kim, Dan Kenu Sung, Soon Dal Choi
Satellite Technology Research Center (Sa TReC)
Korea Advanced Institute of Science and Technology (KAIST)
373-1 Kusung-dung, Yusung-Ku, Taejon, 305-7-1
Abstract
The SaTReC has developed and been operating two micro-satellites, KISAT-1 and 2, and is now developing the third satellite system, KITSAT-3 is designed and operated from the basis of engineering test purposes. The KISAT-3 system is a small satellite with a mass of approximately 100 Kg, however, 3-axis stabilized. Each subsystem has been designed from similar concept to the KIST-1/2, however system architecture is so unique and modular that can be easily modified and expanded for future missions.
This paper presents how the KISAT-3 system has been designed at system level. It will also include critical mission analyses such as power budget, mass budget, thermal analysis, link budget, operational scenarios and attitude maneuvering under given constraints. The results of mission analyses will generate a baseline system for preliminary design.
1. Introducting Of Kitsat-3 System
The Satellite Technology Research Center (SaTReC) has launched two of micro-satellites, KITSAT-1 and KITSAT-2, in 1992 and 1993, respectively. The KITSAT-1 was the first Korean satellite and has been developed as a collaborative project with University of Surrey in UK. The KITSAT-2 has been completely newly developed and test in Korea by Korean engineers with major changes on payloads and corresponding bus systems.
The SaTReC is now aiming to design, develop, and operate the third KITSAT using experience and well trained engineers obtained from KITSAT-1/2. The KITSAT-3 system will be SaTReC's unique small satellite and will be used as bus platform for SaTReC's future space missions. The primary mission objectives are development and in-orbit test of newly designed small satellite system especially demonstrating new techniques such as 3-axis stabilized attitude control system, solar panel deployment mechanism, common bus networked architecture, high resolution and multi-spectral camera system, and payload data transmission system in high data rate.
The mission payloads on board KITSAT-3 are a remote sensing payload with 15m resolution and pushbroom type CCD camera in three spectral bands, a space science payload with a high energy particle telescope and a radiation monitor which can measure radiation environment around mission orbit and can also distinguish what types of particles and in which levels, and a data collection payload collecting data measured from buoys floating on ocean.
The KITSAT-3 is supposed to be launch ready at the end of 1996 and designed to be launched by Chinese Long March IV as a secondary launch. The weight of KITSAT-3 system shall be less than 100 Kg and the dimension of satellite shall be within 45 x 45 x 60 (cm).
2. Mission Analyses
In order to configure the satellite systems, it is necessary to allocate performance to the satellite elements and to specify corresponding subsystems. The process to allocate performance and to evaluate them is generally called mission analysis. It required iterations and provides overall system configuration for preliminary design. It begins with analyzing the requirements and allocating performance parameters by establishing budgets for power, mass, and propellant. For KITSAT-3 system, the propellant budget is not considered, as there is not propellant system. The pudgeting must be based upon the mission orbit and operational view point. In this paper, power budget, mass budge, and initial thermal analysis will be mainly presented.
2.1. Orbital analysis
Prior to analyze requirements and allocate performance parameters, the Mission orbit should be defined clearly and analyzed first. The mission orbit would be one of the most important constraints to designs satellite systems. Once satellite system has been designed to be operated in a mission orbit. It needs significant modifications to be operated in others. Different orbital characteristics give different solutions for available power, battery's charge and discharge characteristics, and thermal control which is the most significantly affected.
The mission orbit of the KITSAT-3 is approximately 800 km of altitude and sun-synchronous orbit. The descending node will be about 10.30 AM , therefore the sum incident angel toward orbit ,it is assumed that the satellite is positioned at the initial position and the orbit properties are not significantly changed during mission life .
Sun-synchronous gives a lot of benefit to satellite system designer. It will give nearly contact sun illumination to satellite therefore available power and depth of discharge is constants at every mission orbit .Thermal environment is also moderate to adopt passive thermal control . For the remote sensing payload as the areas taken by camera system are illuminated at well contact level, it helps to control gain and integration time .
At 800 km altitude, the node spacings between successive orbits are calculated as 25.22 deg for a day ,6.92 deg for 4 days , and 2.46 deg for 16 days . as the field of view of KITSAT-3 camera system is 5 deg, the kitsat-3 can achieve completed coverage and take images at any area at lest every 16 days .If the attitude control system is well qualified , the revisit period can be reduced to 4 days with small attitude adjustment using reaction wheels.
Through the detailed analysis on orbit property , it can be calculated that the mission control center at the SaTReC has the contact time of approximately 48 minutes over 5 deg and 17 minutes over 20 deg per day .
2.2.Power Budget
The power budget is estimated by adding the payloads power requirements to power estimates for the spacecraft system . If the satellite system has several different requirements to power estimates for the spacecraft system . It is necessary to estimate power requirement separately for each operation mode , power requirement, it is necessary to estimate power requirement separately for each operation mode , paying particular attention to peak power need for each subsystem . Once defined power requirements and their duty cycle the required power from energy source can be calculated . from this result, we can size the solar panel and the battery capacities to supply enough power for operations and recharging batteries until the end of mission life . During this process the degradation on power subsystem over the mission life should be taken into account . radiation damages to the solar panels and depth of discharge and number of charge and discharge cycles of the battery are main factors for degradation .
Table 1 Power requirements at subsystems
| Subsystem |
Estimated power (watt) |
Duty cycle(%) |
| Spacecraft |
| Electrical power system |
20.0 |
100 |
| Attitude control system |
17.2 |
†¹ |
| C&DH System |
2.5 |
100 |
| Transmitters/ Receivers |
10.0 |
100 |
| Structure/ Thermal |
0 |
|
| GPS Receiver /processor |
7.0 |
†² |
| Sub -Total |
56.7 |
|
| Payloads |
| Remote Sensing payload with Data Transmitter |
60.0 |
10 |
| Space Science Payload |
0.9/2.2 |
100/10, +¹ |
| Data collection payload |
2.0 |
50 |
| Sub-Total |
8.03 |
|
| Total |
64.73 |
|
Where and †¹, †²are variable to operation modes and +¹ corresponds to radiation monitoring mode and pitch angel rotation modes. Pitch angel rotation mode will be activated only after solar flare precaution for about 48 hours.
Remote sensing payload will be operated once a day basis , therefore 10%of duty cycle will be enough for estimation.
From the power budget estimation , the required power from solar panels and their area are roughly calculated as 100 Watt and 0.362m²,when GaAs solar cells (18 % of efficiency) are used. with the consideration of solar panels' degradation during 3 years mission life, which is about 20 % at 800 km's polar orbit and oblique Sun incident angle ,which is used about 22.5 deg, if the area or solar panel is over 0.56m², then it will be sufficient to supply enough power until the end of mission life.
