CubeSat Interface Standard Draft and Project Update

CubeSats are often built with emphasis on low-cost and fast-delivery. The low-cost can be achieved by extensive use of non-space-qualified commercial-off-the-shelf parts and units. The fast-delivery is, however, often difficult to achieve when the interface of different units, such as printed circuit board (PCB), do not match each other. The incompatibility can cause significant delay in the satellite project, leading to the loss of business opportunity or academic/technology competition. There is also increasing trend that a CubeSat platform that contains all the satellite bus functionalities by a single vendor is combined with a mission payload. If there is a common standard on the interface between the CubeSat platform and the mission payload, it will broaden the choice for the those who want to do a space mission but not want to build a satellite to select the platform depending on their needs. The standard will make it easier for CubeSat vendors to enter the market of CubeSat platforms. In 2019, a new project to standardize the CubeSat interface started with participations of worldwide CubeSat developers and vendors. After going through a series of meetings, now a standard draft is ready to be submitted to ISO (International Standard Organization) for registration as a New Work Item. The draft standard document describes internal and external interface of CubeSat. The document is made of eight chapters. In the fifth chapter, the document lists the requirements for unit (i.e.component) to unit interface. After providing general requirements, it deals with the PC-104 style and the backplane style. The document then describes the requirements for mission payload to platform (i.e.bus) interface. In the sixth chapter, the document lists the datasheet requirements for CubeSat units. The purpose is to ease the production selection process, satellite integration and testing. In the seventh chapter, the document lists the datasheet requirement for CubeSat platform, which basically can serve as the interface control document. In the eight chapter, the document lists the requirements for the external electrical interface, i.e.umbilical

Lean Satellite Concept

In 2014, International Academy of Astronautics (IAA) initiated Study Group (SG) 4.18, Definition and Requirements of Small Satellites Seeking Low-Cost and Fast-Delivery . Its objectives of the study group are to examine the definitions of small satellites, identify the requirements every satellite should follow and then reflect some of the findings to an ISO standard draft ISO/WD/20991, Space systems - Requirements for Small Spacecraft that is being developed recently at ISO/TC20/SC14. The purpose of the present paper is to present the latest findings in the SG activity especially in terms of small satellite definition. During the SG meeting in 2014, a round-table discussion was held to discuss the terminology to describe small satellites. The majority of the opinions were that neither mass nor size is suitable for defining small satellites. Rather, philosophy of design, manufacturing, mission, program management, etc., should be used for the definition. The round-table discussion came to the conclusion that the term Lean Satellite is more suitable than Small Satellite . Historically, the word of lean originated from Toyota Production System (TPS). There are few things in common between satellites and automobiles. It is very difficult to apply lean concepts as they are to satellites. But some concept of lean is necessary for satellites. New types of customers are emerging today who want more value from satellites through lower unit prices and faster system delivery. Currently, mega-constellations consisting of hundreds or thousands of satellites are being proposed. Traditional satellite development philosophy cannot be applied to mega-constellations because the total cost would be prohibitively high. Small satellites and mega-constellations can benefit from the application of the lean satellite concept, although it must be modified to accommodate the differences between satellites and automobiles. Space systems engineering has put emphasis on delivering a perfectly working system. On the other hand, lean concept has put emphasis on delivering a high-quality product with the minimum cost and shortest time. Developing the lean satellite concept can be an interesting new subject for space systems engineering. As a part of the SG activity, 16 questions were identified as good measures to scale the characteristics as a lean satellite. The 16 questions are made of 9 categories with different weighting. They are (1) total cost, (2) delivery time, (3) simplicity, (4) risk taking, (5) risk mitigation, (6) reliability requirement, (7) mission duration, (8) launch, and (9) waste minimization. Some categories are further divided to multiple questions. Each question has its weight. Each answer has its score. By adding the points of all the 16 questions, the total point is between 0 and 100. The questionnaire made of the 16 questions was distributed in SG and answers by more than 41 satellites from all over the world were collected. At the conference, the lean satellite concept will be presented more in detail along with the analysis of answers made by the 41 satellites

Testing of Micro/Nano Satellites and their On-orbit Performance

Demand of an environment test facility suitable for micro/nano satellites has increased as the number of new satellite developers such as universities and small business companies has increased. In Japan, development of micro/nano satellites (10cm to 50cm) is very active since Japan Aerospace Exploration Agency (JAXA) started the piggyback program via H2A rocket in 2009. Micro/nano satellite systems and components can be developed at a small laboratory. For testing, however, expensive and special test machines such as thermal vacuum chamber and high power shaker are necessary. To reduce the burden on the micro/nano satellite developers, Center of Nanosatellite Testing (CeNT) at Kyushu Institute of Technology (KIT) was established as a comprehensive test facility for micro/nano satellite up to 50 cm in 2010. The purpose of CeNT is to provide low-cost and easy-to-use testing service for universities and small business companies. CeNT has a vibration machine (~ 33 kN), shock machines (~10000G), small and large thermal vacuum chambers (0.3 m and 1.7 m) and thermal cycle chambers. Offering the one-stop service for environment testing ensures the traceability of verification processes necessary for anomaly investigation. Since 2010, 11 satellite projects used the facility. 2 satellites were already launched in 2012 and 9 satellites will be launched in 2013. This paper describes the testing results and their on-orbit performance of the following 3 satellites. 1. Horyu-2 (Kyushu Institute of Technology, launched in 2012) Horyu-2 was launched on May 18, 2012. The satellite was developed by KIT students. The satellite’s size and weight are 30 cm cube and 7 kg. During the project, STM, EM and FM were developed from 2010 to 2012. The main mission of Horyu-2 is high voltage photovoltaic power generation. Vibration, shock and thermal vacuum were carried out for each model. JAXA set the strict test conditions for the vibration test and the shock test to avoid accidents such as dropping parts, accidental switch-on and unexpected RF emission during the launch and the fairing separation that could cause the critical damage to the main satellites. The satellite was launched without any anomaly. The temperature variation during the nominal operation matched with the prediction. We kept a detail record of the cost associated with each test. The effectiveness of each test will be reviewed. 2. FITSAT (Fukuoka Institute of Technology, launched in 2012) FITSAT was launched on October, 2012. The satellite was one of the first nano satellites released from Japan Experiment Module (JEM) on International Space Station (ISS). FITSAT was launched by HTV (H2 Transfer vehicle). It was packed within the rocket body, very different conditions from typical piggy-back satellites. The missions of FITSAT were high- speed data communication by 5.8 GHz amateur band and high power LED illumination. At CeNT, vibration and thermal vacuum test was carried out for EM and FM model. FITSAT carried a lithium-ion battery to drive the high power LED. Thermal vacuum test was necessary to ensure the safety of the battery because FITSAT was stored on ISS (International Space Station) until the satellite was released. The satellite missions were successful. 3. QSAT-EOS (Federation of universities and companies in Kyushu, launch in 2013) QSAT-EOS (Kyushu Satellite for Earth Observation System Demonstration) was developed by a consortium of universities and companies in Kyushu. The satellite is 50cm cube and 50kg weight. It will be launched in 2013 via a Dnepr rocket. QSAT-EOS was the first 50 kg satellite tested in CeNT. For QSAT-EOS, thermal vacuum test was conducted for STM and FM. In this paper, we will compare thermal vacuum test results and the on-orbit data. At the conference, the lessons learned from these environmental tests will be presented

Computer simulations on the initiation and morphological difference of Japan winter and summer sprites

[1] Two‐dimensional (axisymmetric) computer simulations (electromagnetic code) have been performed to study the initiation and morphological difference (summer, carrots, and winter, columns) of sprites for simulating Japan summer and winter sprites. By changing the physical parameters of a parent lightning, we have found the following findings by looking at the spatial‐temporal distribution of reduced electric field, etc. (1) There are three important factors (the height where the charge is removed (ds), charge transfer (Q), and lightning current risetime (τ)) in the initiation of sprites and their morphological differences. (2) For the initiation of sprites, the charge transfer (Q) should exceed a certain value for possible charge heights (giving us a threshold of charge moment change (Qds) of the order of ∼120–200 C·km), with a combination of small risetime of lightning current waveform. (3) Further, the height for positive charge is much higher than that for negative charge in a typical lightning configuration, which is the essential factor in determining the morphological difference of sprites in summer and winter. Positive charges for summer are located at a much higher altitude than those in winter in Japan, which might result in carrot‐type in summer and columnar‐type in winter even for the same positive polarity. (4) A combinational effect of (Ids) and (Qds) is important for having sprites. Finally, the present computer results on the initiation of sprites for Japanese lightning have been compared extensively to the well‐documented properties of summer continental sprites and future subjects to study have been suggested

Programmable CubeSat Interface Board to Reduce Costs and Delivery Time

A standardized interface for different CubeSat missions is one of the keys to reduce costs and delivery time. A backplane interface approach, proposed by the University of Wuerzburg in Germany as UWE-3, was implemented in three CubeSat projects at the Kyushu Institute of Technology (Kyutech) in Japan to shorten the development and assembly times. The backplane approach also helped to reduce the risk of workmanship errors associated with the harness. The proposed standard interface board, however, needed changes in every CubeSat project to comply with the mission requirements. To obtain more flexibility especially for data connections, this work introduces a novel idea of a software-configurable bus interface with the backplane board. A Complex Programmable Logic Device (CPLD) was used instead of the hardware routing so that we can reconfigure the bus interface by reprogramming the CPLD. The concept was validated by a functional test with a breadboard module. A radiation test verified that the selected CPLD has enough strength to survive total ionization dozes of more than 2 years in low Earth orbit. A new backplane board with CPLD have been integrated with Engineering Model and Flight Model of the fourth CubeSat project at Kyutech, BIRDS-3 project, and system level verification was conducted. The flight model is now ready for delivery to JAXA in February 2019 for a planned launch to International Space Station in April 2019. The initial on-orbit data will be obtained by the time of the conference in August 2019 and will be presented to the audience

Classification of Countries Worldwide according to Satellite Activity Level

Emerging countries worldwide can benefit technologically, economically, and socially from domestic space-related activities. However, limited resources and lack of know-how prevent many non-space faring nations from initiating space projects, much less building sustainable space or satellite programs. New opportunities exist to overcome these barriers. The number of small satellites launched in the last three years has increased by an order of magnitude, and the market of satellites less than 50kg is projected to grow from $700M USD in 2014 to ∼$2B USD by 2019. This paper gives an overview of space-related activities in various emerging countries and categorizes countries in nine regions worldwide according to level of satellite activity.The 30th International Symposium on Space Technology and Science (30th ISTS), July 4-10, 2015, Kobe, Hyog

New star identification algorithm using labelling technique

A new star identification algorithm is proposed for the attitude determination of a star sensor in the lost-in-space case, where prior attitude information is not available. The algorithm is based on a labelling technique, which uses label values to represent each group of stars. Using label values, multiple stars are simultaneously identified without repetition of search work. This labelling algorithm allows for a fast identification speed with efficiency, and provides the capability of more reliable identification by redundant confirmation. The proposed algorithm was verified by simulation study under various conditions

OPTIMIZATION OF A SUN VECTOR DETERMINATION FOR PINHOLE TYPE SUN SENSOR

The sun vector is commonly used for defining a satellite attitude and many types of sensors exist for its determination. An attitude determination system is designed for each satellite project based on missions’ requirements. A fine pinhole sun sensor type was chosen and designed for HORYU-IV nanosatellite of Kyushu Institute of Technology. This sensor has a round-shaped hole and uses commercial off-the-shelf silicon photodiode, which consists of four small sensitive elements arranged close to each other. This type of sensors commonly uses look-up tables for providing high accuracy, which requires a large amount of data to be saved. Linear and polynomial methods for sun vector determination were considered instead of look-up tables to avoid having a large amount of data to be saved. Moreover, the influence of dead spaces between photodiodes on sensor accuracy was also investigated. Six real sun sensors and their theoretical models with different configurations were designed for investigating the difference between various calculating methods. The comparison of accuracies between proposed methods for real sun sensor models leads to two main findings: 1) on average, a polynomial method decreases error level of determined angle by 70% when compared with linear method; 2) accounting for gaps between photodiodes further decreases the average error of the angle determined by 15 % for polynomial, and by 6% for linear method when compared with methods not accounting for gaps