On-Orbit Results and Lessons Learned from the ASTERIA Space Telescope Mission

The Arcsecond Space Telescope Enabling Research in Astrophysics (ASTERIA) was deployed from the International Space Station (ISS) on 20 November 2017, beginning a technology demonstration and opportunistic science mission to advance the state of the art in nanosatellite performance for astrophysical observations. The goal of ASTERIA is to achieve arcsecond-level line-of-sight pointing error and highly stable focal plane temperature control. These capabilities enable precision photometry—i.e. the careful measurement of stellar brightness over time—which in turn allows investigation of astrophysical phenomena such as transiting exoplanets. By the end of the 90-day prime mission, ASTERIA had achieved line-of-sight pointing stability of approximately 0.5 arcseconds root mean square (RMS) over 20-minute observations, pointing repeatability of 1 milliarcsecond RMS from one observation to the next, and focal plane temperature stability better than ±0.01 K over 20-minute observations. This paper presents an overview of the ASTERIA flight and ground system, summarizes the pre-delivery test campaign, and discusses the on-orbit performance obtained by the pointing and thermal control subsystems. We also describe the process for planning opportunistic science observations and present lessons learned from development and operations. Having successfully operated for over 200 days as of this writing, ASTERIA is currently in an extended mission to observe nearby bright stars for transiting exoplanets

Nanosatellite Store-and-Forward Communication Systems for Remote Data Collection Applications

Due to compact design, cost-effectiveness and shorter development time, a nanosatellite constellation is seen as a viable space-based data-relay asset to collect data from remote places that are rather impractical to be linked by terrestrial means. While nanosatellites have these advantages, they have more inherent technical limitations because of limited space for subsystems and payloads. Nanosatellite S&F communication systems are notably challenging in this respect due to requirements on antennas, transceivers, and signal processing. Although nanosatellites can be scaled up for better resources and capabilities, smaller platforms (i.e., ≤6U CubeSat) tend to be used for cost-effectiveness and lower risk. This thesis dealt with the problem of designing a nanosatellite S&F communication system for delay-tolerant remote data collection applications considering: (a) technical constraints in hardware, processing capabilities, energy budget and space in both the nanosatellite and ground sensor terminal (GST) sides; (b) physical communication layer characteristics and constraints such as limited available bandwidth, LEO channel Doppler, attenuation and fading/shadowing effects, low transmit power and data rate, and multi-user interference among asynchronously transmitting terminals. We designed, developed, and operated an amateur radio payload with S&F communication and APRS-DP capabilities, and performed a post-launch communication failure investigation. We also investigated suitability of E-SSA protocol for IoT/M2M terminals to nanosatellite communication by analyzing performance and energy efficiency metrics. The thesis comprises nine chapters. Chapter 1 describes the research background, problem, objectives, state of research, potential contributions of this thesis, and a gist of methodology detailed in later chapters. Chapter 2 and 3 provide an extensive literature review. Chapter 2 reviews the previous research works on using nanosatellites for S&F communication for remote data collection, and the previous nanosatellite S&F missions. Such research works and nanosatellite missions were undertaken primarily in the context of non-commercial/civil applications. Then, Chapter 2 surveys the recent commercial nanosatellite IoT/M2M players and examines their proposed systems in terms of satellite platform, constellation design, communication technology, targeted applications, requirements, and performance. Chapter 3 presents a literature review on communication system architecture, physical layer and random-access schemes, protocols, and technologies relevant to satellite IoT/M2M systems. In the context of IoT/M2M applications, the constraints in energy budget, transmit power and available bandwidth limit the system’s capacity in terms of amount of data that can be received and number of GSTs that can be supported. In both nanosatellite and GST sides, there are stringent limitations in hardware complexity, processing capabilities and energy budget. Addressing these challenges requires a simple, spectrally and energy efficient asynchronous random-access communication protocol. This research investigated using the enhanced spread spectrum Aloha (E-SSA) protocol for satellite IoT/M2M uplink (terminal to satellite) communication and analyzed its performance and suitability for the said application. Chapter 4 discusses the BIRDS-2 CubeSat S&F remote data collection system, payload design, development, tests, and integration with the BIRDS-2 CubeSats. Chapter 5 discusses the investigation on communication design issues of BIRDS-2 CubeSat S&F payload, tackling both the methodology and findings of investigation. It is noted that there are only a few satellites that have carried an APRS-DP payload but even some of these failed due to communication, power, or software issues. In BIRDS-2 Project, considering tight constraints in a 1U CubeSat equipped with other subsystems and payloads, we developed a S&F/APRS-DP payload and integrated it with each of the three 1U CubeSats of participating countries. After launching the CubeSats from the ISS, several amateur operators confirmed reception downlink beacon messages, but full two-way communication failed due to uplink communication failure. Thus, this research not only studied the design and development of a S&F/APRS-DP payload suitable for a CubeSat platform, but also systematically investigated the causes of communication failure by on-orbit observation results and ground-based tests. We found that uplink failure was caused by two design problems that were overlooked during development, namely, the poor antenna performance and increased payload receiver noise floor due to satellite-radiated EMI coupled to the antenna. Chapter 6 first describes the enhanced spread spectrum Aloha (E-SSA) based nanosatellite IoT/M2M communication model implemented in Matlab and derives the mathematical definitions of packet loss rate (PLR), throughput (THR) and energy efficiency (EE) metrics. Then, it tackles the formulated baseband signal processing algorithm for E-SSA, including packet detection, channel estimation, demodulation and decoding. Chapter 7 presents the simulation results and discussion for Chapter 6. Chapter 8 tackles the S&F nanosatellite constellation design for global coverage and presents the results and findings. Chapter 9 describes the laboratory setups for validating the E-SSA protocol and then presents the findings. Finally, Chapter 9 also gives the summary, conclusions, and recommendations. Simulation results showed that for E-SSA protocol with the formulated algorithm, THR, PLR and EE metrics are more sensitive to MAC load G, received power variation σLN and Eb/N0, due to imperfect detection and channel estimation. With loose power control (σLN=3dB), at Eb/N0=14 dB, the system can be operated up to a maximum load of 1.3 bps/Hz, achieving a maximum THR of 1.25 bps/Hz with PLR<0.03. Without power control (σLN=6dB,9dB), at Eb/N0=14 dB, maximum load is also 1.3 bps/Hz, but achievable THR is lower than ~1 bps/Hz and PLR values can be as high as ~0.23. Worse PLR results are attributed to misdetection of lower power packets and demodulation/decoding errors. Both are caused by the combined effects of MUI, channel estimation errors, imperfect interference cancellation residue power, and noise. The PLR and THR can be improved by operating with higher Eb/N0 at the expense of lower energy efficiency. Then, laboratory validation experiments using a SDR-based platform confirmed that with G=0.1, Eb/N0=14dB, σLN=6dB, the formulated algorithm for E-SSA protocol can still work even with inaccurate oscillator (±2 ppm) at GSTs, obtaining experimental PLR result of 0.0650 compared to simulation result of 0.0352. However, this requires lowering the detection thresholds and takes significantly longer processing time. For the S&F nanosatellite constellation design, it was found that to achieve the target percent coverage time (PCT) of more than 95% across all latitudes, a 9x10 Hybrid constellation or a 10x10 Walker Delta constellation would be required.九州工業大学博士学位論文 学位記番号：工博甲第506号 学位授与年月日：令和2年9月25日1: Introduction|2: Nanosatellite S&F Research, Missions and Applications|3: Satellite S&F Communication Systems and Protocols|4: BIRDS-2 CubeSat S&F Data Collection System, Payload Design and Development|5: Investigation on Communication Design Issues of BIRDS-2 CubeSat APRS-DP/S&F Payload, Results and Discussion|6: E-SSA-based Nanosatellite IoT/M2M Communication System Model and Signal Processing Algorithm|7: Simulation Results and Discussion for E-SSA-based Nanosatellite IoT/M2M Communication System|8: Nanosatellite Constellation for Global Coverage|9: Experimental Laboratory Validation for E-SSA Protocol, Research Summary, Conclusions and Recommendations九州工業大学令和2年

A Design Of A Test Bed For Cubesat Attitude Determination And Control System

Tez (Yüksek Lisans) -- İstanbul Teknik Üniversitesi, Fen Bilimleri Enstitüsü, 2016Thesis (M.Sc.) -- İstanbul Technical University, Institute of Science and Technology, 2016Gelişen teknoloji ve keşfedilen yeni metodlar ile birlikte, daha küçük uyduları tasarlamak ve üretmek mümkün hale gelmiştir. Bu yeni küçük uydulara küp uydu denmektedir. Bir küp uydu 10cmX10cmX10cm boyutlarında ve 1.3 kg kütlesinde olarak tanımlanmıştır. İlk çıkış amaçları öğrencilere bire bir uydu geliştirme tecrübesi edinmeleri için yapılmıştır. 2000'ler yıllardan bu yana küp uydu geleştirilmesi ve de fırlatmaları sürekli artmaktadır. 2012'de fırlatılan uydu sayısı 25 iken, 2017 yılı için planlanan uydu sayısı 311dir. Şu anda geliştirilmekte olan bütün küp uyduların \%40'ı üniversiteler tarafından yapılmaktadır. Bu uydular çoğunlukla öğrenciler tarafından geliştirilmektedir. Bunun sonucu olarak uydulardaki risk artmaktadır. 2016 yılına kadar fırlatılan nano uyduların başarısızlık oranı \%33'tür. Sadece üniversiteler tarafından geliştirilen uydularda kısmi başarıların da tam başarı olarak kabul edilmesi durumunda tüm küçük uyduların başarı oranı \%40 civarında oluyor. 2010 yılına kadar yapılmış olan nanouyduların neredeyse yarısı 3U ve ondan küçük uydular olarak tasarlanmıştır. 2016 yılı verileri incelendiğinde fırlatılan ve planlananlar için bu oran \%70'lere ulaşmaktadır ve toplam sayı 1000'den fazladır. Uydu sayıları bu kadar artarken başarısızlık da artacaktır. Mevcut eğilim göz önüne alındığında uyduların fırlatılmadan önce daha çok test edilmesi gerektiği görülmüştür. Fırlatılan uyduların büyük çoğunluğu yörüngeye ulaşabilmekte fakat ardından kısa bir süre sonra işlevsiz hale gelmektedir. Buradan anlaşıldığı üzere termal-vakum testleri ve titreşim testleri uydunun sadece dayanıklılığını göstermektedir. Fakat uzun vade de uydunun herhangi bir yazılım ya da algoritma sonucunda başarısız hale gelip gelemeyeceği ya da farklı sorunlarda ne gibi sonuçların ortaya çıkacağı öngörülememektedir. Oranın azaltılması adına uyduların fonksiyonel testlerinin arttırılması gerekmektedir. Gerekli olan test ekipmanları uydulardaki alt sistemler temel alınarak belirlenebilir. Bir küp uydu başlıca elektrik düzenleme biriminden, pil biriminden, güneş panellerinden, yapıdan, yönelim belirleme ve kontrol sisteminden, uçuş bilgisayarından, modemden ve bilimsel yükten oluşmaktadır. Yörüngedeki bir uydunun durumu göz önüne alınırsa, fonksiyonel olduğu zaman boyunca çeşitli etkiler altında kalmakta ve görevler yerine getirmektedir. Sürekli olarak bir manyetik alan etkisinde kalmaktadır, değişen açılarla güneş ışığına maruz kalmaktadır, yıldızlara ve de dünyaya bakarak konum algılamakta, fotoğraf çekmekte, kendisini yönlendirmekte, enerji üretip dağıtımını yapmaktadır. Tüm bunları yeryüzünde yapabilmek için bütünleşik bir test sistemine ihtiyaç duyulmaktadır. Örnek olarak dünyanın manyetik modelini gerçeklemek için Helmholtz kafesi kullanılabilir. Bu kafes ile modellenen manyetik alan içerisinde uydunun manyetometresi denenebilir, toplanan verilere göre yönelim belirleme ve kontrol sisteminin yörünge tahmini ve kontrolü denenebilir. Değişen açılarla güneş ışığına maruz kalabilmesi için güneş ışığı benzetim düzeneği yapılabilir. Yapılan bu sistem ile güneş panellerinin üretimleri kontrol edilebilir, yönelim belirleme ve kontrol sisteminin alt birimleri ve algoritması kontrol edilebilir. Bir diğer sistem olarak dünya haritası ve de yıldız haritası ekranlarda oluşturularak uydu uzaydaymışçasına fotoğraf çekmesi sağlanabilir, yönelim belirleme ve kontrol sistemin yönelendirmesi, algoritması, tepki tekeri gibi sistemleri denenebilir. Tüm bu sistemlerin birleştirilebilmesi için ayrıca uydunun rahatça haraket edebileceği, düşük sürtünmeli bir düzenek gerekmektedir. Bu hava yatağı ile sağlanabilir. Bu sayede uydu üç eksende serbest olarak hareket edebilir. Bahsedilen bilgiler ışığında yönelim belirleme ve kontrol sisteminin uzaydaki bütün görev ve durumlarda aktif ya da dolaylı olarak rol oynadığı görülmektedir. Bu yüzden yönelim belirleme ve kontrol sistemi için bir düzeneğin geliştirilmesi diğer sistemler içinde rahat bir başlangıç oluşturacaktır. Yönelim belirleme ve kontrol sistemi küp uydularda gün geçtikçe daha çok kullanılmaktadır. Bunun sebebi hem sistemlerin fiyatlarının azalması hem de görevlerin zorluklarının ve hassasiyetlerinin artmasından kaynaklanmaktadır. Yönelim belirleme ve kontrol sistemi uydunun yönlendirilmesinde, konumunun belirlenmesinde, uydunun denge halinde tutulmasında, fotoğraf çekerken sabitlenmesinde ve itki sistemleri için yönlendirme oluşturmak üzere kullanılmaktadır. Bir yönelim belirleme ve kontrol sisteminde birçok duyarga sistemi bulunmaktadır. Güneş ve ufuk senyörü uydunun hangi yöne doğru yöneldiğini, uydunun tutulma zamanlarının anlaşılmasında kullanılabilmektedir. Ebatlarından dolayı uydularda bu araçlardan birer adet olmaktadır. Odak noktası ayarlanmış kamera oldukları için güç ihtiyaçları vardır. Bunun dışında kaba güneş duyargaları da bulunabilir. Bu tarz duyargalar daha çok güneş hücreleri şeklinde olmaktadır. Uydunun her yüzeyinde en az bir tane konulmaktadır. Uydunun tam olarak hangi yüzeyinin güneşe baktığının anlaşılması ve de yörünge tahmin yazılımın hassasiyetinin arttırılmasında kullanmaktadır. Güneş hücresi şeklinde oldukları için kendi enerjilerini kendileri üretebilmektedirler ve bu da sistemi daha güvenli ve bağımsız yapmaktadır. Bir diğer alt sistem de ataletsel ölçüm birimidir. Elektronik, yazılımsal ve mekanik parçaların bir araya gelmesinden oluşmaktadır. İçerisinde manyetometre, ivmeölçer ve gyro bulunabilir. Hepsinin bir arada olmasından dolayı hassasiyetleri o kadar yüksek değildir. Manyetometre daha hassas ve daha düşük manyetik alanları da ölçebilir. Yıldız takip sistemleri de uydunun konumunu ve de yönelimini anlamak için kullanılmaktadır. Bu sayılan sistemler pasif sistemlerdir. Yönelim belirleme ve kontrol sisteminde aktif yani hareketli elemanlar da bulunmaktadır. Bunlar uydunun sabit tutulması ve de döndürülmesi için kullanılmaktadırlar. Bunlardan en basiti manyetik eğleyicilerdir. Dünya çevresindeki manyetik alanı kullanarak sahip olduğu sargılardan akım geçirmek suretiyle kuvvet oluşturarak uyduyu sabit tutabilir ya da tepki tekerlerinin doyuma girmesini engellemek için sönümlenmelerine yardımcı olmaktadırlar. Ardından sırasıyla tepki tekeri, momentum tekeri ve kontrol moment gyroskobu gelmektedir. Bu sistemler arasında sadece küçük farklılıklar vardır. Tepki tekeri genelde kapalı olup ihtiyaç halinde yüksek hızlara çıkarak uydunun istenilen bir doğrultuya yönlendirilmesini sağlamaktadır. Momentum tekerleri ise sürekli olarak dönmektedir. Bu yüksek hızlı dönme uydunun yörüngede ilerlemesi sırasında uydunun sabit tutulmasını sağlamaktadır. Uyduların yeryüzünde test edilebilmesi için tavsiye edilen başlangıç sistemi Helmholtz kafesidir. Çünkü diğer bahsedilen test sistemlerine kıyasla daha basit ve ucuzdur. Bu tez çalışmasında helmholtz kafesi tasarlanmıştır. İlk olarak dairesel kafes sonlu elemanlar yöntemi ile analiz edilmiştir. Planlanan uydu boyutları değerlendirildiğinde 3U ve daha küçük uyduların daha çok olduğu görülmektedir. Bu bilgi ışığında 30cmx30cmx30cm'lük bir hacim içresinde oluşturulacak manyetik alan birçok uydunun ihtiyacını karşılayabilecek nitelikte olmaktadır. Helmholtz kafesleri sahip oldukları sargıların yarıçapı kadar mesafe ile iki sargının yerleştirilmesinden oluşur ve bu aradaki mesafede homojen alan oluşur. Yapılan sonlu eleman analizleri sonucunda 30x30x30cm3 lük hacmin 30cm sargı arası mesafesi yerine 60cm sargılar arası mesafeye sahip bir sistemin içerisine konması manyetik alanın homojen dağılımında içeriye konulacak sistem için çok büyük farklılıklar oluşturmuştur. İkinci olarak da kafes şekillerine göre benzetim ve analizler yapıldı. Kare ve dairesel kafes yapıları incelendiğinde kare yapıların, dairesel yapılara göre daha homojen olduğu görüldü. Bütçe ve zaman yetersizliğinden dolayı sistemin çalışabilirliğinin kanıtlanması amacıyla küçük bir kafes tasarlanmasına karar verildi. Tasarlanan kafes 5cmx5cmx5cm lük bir hacimde homojen bir manyetik alan oluşturacak şekilde benzetimi ve analizi yapıldı.. Mevcut Helmhotz kafesinde manyetik alanın belirlenebilmesi için öncelikle uydunun bir konumunun olması ve o konumdaki manyetik alan bilgilerinin elde edilmiş olması gerekmektedir. Bunu gerçekleştirebilmek için Dünya manyetik modeli IGRF ve yörünge ilerletici SGP4 yazılımları kullanılarak küresel koordinatlara bağlı olarak manyetik alanlar hesaplanmıştır. Ardından bu değerler kartezyen koordinatlara çevrilerek kafeste kullanıma uygun hale getirilmiştir. Kafesin etkin bir şekilde çalışabilmesi için güç kaynaklarının dikkatli bir şekilde hesaplanması gerekmektedir Çalışacak olan güç kaynaklarının tepki süreleri, kafes içerisinde kullanılacak olan manyetometreden yavaş olmalıdır. Aksi takdirde ölçüm bilgisi tam gelmeden, sistem değeri tekrar değişecek ve de düzenlenecek bu da sürekli olarak sistemin yanlış konumdaymış gibi düzenleme yapmasına sebep olacaktır. Ayrıca sistemin hassasiyeti güç kaynağının anahtarlama elemanın anahtarlama frekansına ve onu kontrol edecek olan mikro denetleyicinin kristal frekansına bağlıdır. İkisi arasındaki oran adım büyüklüğünü vermektedir. Yani adım büyüklüğü ne kadar küçük olursa, manyetik alan değişimi o kadar hassas bir şekilde gerçekleştirilebilir. Planlanan işler olarak ilk önce masa üstü modelinin güç kaynağının üretilmesi gelmektedir. Ardından yazılım ile bağlantısı sağlanarak sistem bir bütün olarak denenecektir. Çalışması durumunda yazılımda bir değişiklik yapmadan sadece kafes ve güç kaynağı ölçeklenerek daha büyük ve asıl sisteme geçiş yapılabilir. Bu tez çalışması sırasında bu tezin çok fazla konuyu bir arada barındırdığı görülmüştür. Aslında bu çalışmanın 3 farklı proje olarak yapılması daha kaliteli ve etkin bir sistemin çıkmasında etkili olacaktır. Manyetik alan ve yörünge ilerletici, kafes tasarımı ve de güç kaynağı tasarımı olmak üzere üçe bölünerek daha detaylı çalışmalar yapılabilir.Owing to the recent developments in miniaturization and integration technologies, CubeSat’s can handle more complex practical missions. Such missions require 3-axis control of the satellite along with a miniaturized 3-axis attitude determination and control systems (ADCS). With the QB0 project, the CubeSat’s developed in ITU-SSDTL have also started to use such ADCS systems. Whether they are developed or procured, an ADCS system usually requires a suitable test bed to test the behavior and the performance of it. An ADCS system used for LEO missions usually operates within the Earth’s magnetic environments. Therefore magnetic field sensors such as magnetometers are employed to measure the mediums’ magnetic field or the change in the magnetic field to determine the orientation and motion of a satellite. Magnetic actuators such as magneto torquers are used to align with the magnetic field of the earth or to damp the tumbling motion of a satellite resulting from unbalanced torque distribution on it. A major goal of the SSDTL is to have such a test bed available in the lab. In addition to related software, the major components are a Helmholtz cage and a suitable air bearing table. With this in mind, the purpose of the present thesis is to aid the development of such an ADCS test bed designing first a suitable Helmholtz Cage system. The sizing of the test system depends on requirements such as the maximum mass of the satellite to be tested, coordinates of center of mass and disturbance levels to be counteracted. In the present thesis, first a comprehensive nanosatellites literature survey is conducted. The success rates and mission failure reasons are also investigated. Based on this preliminary study it is observed that 3U CubeSat’s are generally adopted for most missions. Therefore a Helmholtz Cage that will house a 3U CubeSat is considered. A circular Helmholtz cage that will fit a 3U CubeSat is designed and analyzed. 1 axis and 3 axis Helmholtz cage cases are considered. Magnetic field lines present within the cage are demonstrated. Since the size of the Cage considered was just large enough to house 3U the magnetic field lines were not uniform enough. Therefore similar analysis is carried out for larger cage of double the size of the small one. Both are compared for uniformity of the magnetic fields. Then a small square cage is designed and analyzed. Again, 1 axis and 3 axis Helmholtz cage cases are considered. Magnetic field lines present within the cage are demonstrated. The large square cage was found to be the best choice for testing a 3U CubeSat. As a result, a square Helmholtz cage which can be used to test a 3U CubeSat at LEO from 250 km to 1500 km.Yüksek LisansM.Sc

Design and manufacture of 3DOF reaction wheels as actuators for attitude control of a 1U CubeSat

The main objective of this thesis is the design and additive manufacture of various assemblies of Reaction Wheels for the Attitude Control Subsystem of a 1U CubeSat. A CubeSat is a small satellite which dimensions are 10 × 10 × 10 cm and weights 1 kg. These nanosatellites make space exploration more accessible since a lower volume of materials is needed. The mathematical models to perform numerical simulations of the reaction wheels consist of the dynamics and kinematics equations used to describe the satellite, here a quaternion representation is used. Furthermore, the derivation of the distribution matrices of each reaction wheel configuration are presented, the considered configurations are: Orthogonal, NASA Standard, Pyramid and Tetrahedral. Thereafter, the control law used for the simulation model is introduced. It was chosen to apply the typical control law for spacecraft attitude, a PD controller. This is then expressed in the quaternion feedback control, including a counteract for the gyroscopic effect of the rotating body. The Reaction Wheels mechanical design and the simulation model were established. In the mechanical design, the motor selected is the Maxon 20 EC flat. From this motor, four reaction wheel concepts are created, each one of them has a version for the NASA Standard and the Pyramid configuration. These configurations are chosen because they offer redundancy and are suitable to fit in the CubeSat constraints. The fifth concept C5 is for another motor of interest due to its reduced size. Successively, the simulation model is introduced, with an explanation of each block with their inputs and outputs. Subsequently, the concepts were 3D printed and manually assembled. Thereafter, the simulation results showed that the chosen reaction wheels can efficiently control the attitude of the satellite even when one wheel of the redundant system fails. The outcomes for this thesis are that Concept 4 of the mechanical design is the most convenient since it takes advantage of the extra inertia of the external-rotor of the motor. Furthermore, in the numerical simulations the tetrahedral configuration performs the maneuver more effectively, however for a better layout distribution of the wheels, the pyramid configuration is the optimal with good performance results and the best space usability in the CubeSat

パルスプラズマスラスタを搭載したキューブサットのための軌道・姿勢制御特性の評価に関する研究

九州工業大学博士学位論文 学位記番号: 工博甲第476号 学位授与年月日：令和元年6月28日1. Introduction|2. Background|3. Aoba VELOX Missions|4. Satellite Dynamics|5. AOCS algorithm|6. Lunar Orbit lifetime analysis|7. AOCS testing platforms|8. Conclusions and future work|9. References九州工業大学令和元年

An electrical power system for CubeSats

The advent of CubeSats has provided a platform for relatively low-budget programmes to realise space missions. In South Africa, Stellenbosch University and the Cape Peninsula University of Technology have impressive space programmes and have been involved in numerous successful satellite launches. A number of CubeSat projects are currently in progress and commercial-grade Attitude Determination and Control Systems (ADCS), and communications modules, are being developed by the respective universities. The development of a CubeSat-compatible Electrical Power System remains absent, and would be beneficial to future satellite activity here in South Africa. In this thesis, some fundamental aspects of electronic design for space applications is looked at, including but not limited to radiation effects on MOSFET devices; this poses one of the greatest challenges to space-based power systems. To this extent, the different radiation-induced effects and their implications are looked at, and mitigation strategies are discussed. A review of current commercial modules is performed and their design and performance evaluated. A few shortcomings of current systems are noted and corresponding design changes are suggested; in some instances these changes add complexity, but they are shown to introduce appreciable system reliability. A single Li-Ion cell configuration is proposed that uses a 3.7 V nominal bus voltage. Individual battery charge regulation introduces minor inefficiencies, but allows isolation of cells from the pack in the case of cell failure or degradation. A further advantage is the possibility for multiple energy storage media on the same power bus, allowing for EPS-related technology demonstrations, with an assurance of minimum system capabilities. The design of each subsystem is discussed and its respective failure modes identified. A limited number of single points of failure are noted and the mitigation strategies taken are discussed. An initial hardware prototype is developed that is used to test and characterise system performance. Although a few minor modifications are needed, the overall system is shown to function as designed and the concepts used are proven

Design, Development and Testing of a Low-cost Sub-Joule μPPT for a PocketQube

Small satellites are unmanned spacecraft with small size and mass weighing less than 500kg. A small satellite called the CubeSat was created by two university professors to help students understand satellite design. The idea of small satellites caught on and they became popular due to their low cost, quick development time and easy deployment. The inexpensive nature of small satellites has helped lower the entry barrier to space and led to a movement called the “democratisation of space”. The popularity of small satellites has also caught the eye of private companies that recognise the potential of commercialising small satellite technologies. Nowadays, small satellites are being considered for more complex and challenging space missions. However for a small satellite to reach its full potential, it needs to be equipped with a proper propulsion system. Governments, space agencies, companies and universities around the world have started to research new innovative miniaturised space propulsion technologies. Nowadays, there are many newly developed miniaturised propulsion technologies available. The new propulsion systems are either sold by the companies and universities at a very high price, or research and development is closely guarded due to the potential commercial value of the propulsion system. Companies and universities have primarily focused on researching and developing top-of-the-line micro-propulsion devices to win lucrative research funds. This has resulted in a lack of research into cheap reliable micropropulsion as there have been no incentives for companies and universities to develop this area. As a result, fund-limited students and individuals have been left behind, defeating the purpose of small satellites. This dissertation focuses on designing and developing a low-cost sub-joule micro-PPT propulsion system for a PocketQube satellite. The first section covers the literature review, which looks at the different space propulsion technologies currently available. The next section covers the micro-PPT propulsion system’s mechanical and electrical design and development process. After the development process, the performance of the prototype is tested using various input parameters, as well as in vacuum conditions and over its lifetime. The test results show that the optimal performance is obtained with an input voltage supply of 5V at a pulse frequency of 0.5Hz, which achieves a minimal impulse bit v of 0.698μNs and thrust range of 0.349~1.071μN. In comparison to the STRaND-1 3U CubeSat’s PPT, performance data show that the developed μPPT propulsion system is a competitive propulsion solution, as it achieves more thrust with similar minimal impulse bit, using only one third of the power consumption. The μPPT propulsion system is able to produce 1980 shots so far, which is far lower relatively than other established PPTs due to the limitations resulting from capacitor failure.Thesis (MA) -- Faculty of Engineering, the Built Environment, and Technology, 202

Design, Development and Testing of a Low-cost Sub-Joule μPPT for a PocketQube

Small satellites are unmanned spacecraft with small size and mass weighing less than 500kg. A small satellite called the CubeSat was created by two university professors to help students understand satellite design. The idea of small satellites caught on and they became popular due to their low cost, quick development time and easy deployment. The inexpensive nature of small satellites has helped lower the entry barrier to space and led to a movement called the “democratisation of space”. The popularity of small satellites has also caught the eye of private companies that recognise the potential of commercialising small satellite technologies. Nowadays, small satellites are being considered for more complex and challenging space missions. However for a small satellite to reach its full potential, it needs to be equipped with a proper propulsion system. Governments, space agencies, companies and universities around the world have started to research new innovative miniaturised space propulsion technologies. Nowadays, there are many newly developed miniaturised propulsion technologies available. The new propulsion systems are either sold by the companies and universities at a very high price, or research and development is closely guarded due to the potential commercial value of the propulsion system. Companies and universities have primarily focused on researching and developing top-of-the-line micro-propulsion devices to win lucrative research funds. This has resulted in a lack of research into cheap reliable micropropulsion as there have been no incentives for companies and universities to develop this area. As a result, fund-limited students and individuals have been left behind, defeating the purpose of small satellites. This dissertation focuses on designing and developing a low-cost sub-joule micro-PPT propulsion system for a PocketQube satellite. The first section covers the literature review, which looks at the different space propulsion technologies currently available. The next section covers the micro-PPT propulsion system’s mechanical and electrical design and development process. After the development process, the performance of the prototype is tested using various input parameters, as well as in vacuum conditions and over its lifetime. The test results show that the optimal performance is obtained with an input voltage supply of 5V at a pulse frequency of 0.5Hz, which achieves a minimal impulse bit v of 0.698μNs and thrust range of 0.349~1.071μN. In comparison to the STRaND-1 3U CubeSat’s PPT, performance data show that the developed μPPT propulsion system is a competitive propulsion solution, as it achieves more thrust with similar minimal impulse bit, using only one third of the power consumption. The μPPT propulsion system is able to produce 1980 shots so far, which is far lower relatively than other established PPTs due to the limitations resulting from capacitor failure.Thesis (MA) -- Faculty of Engineering, the Built Environment, and Technology, 202

Design and verification of Guidance, Navigation and Control systems for space applications

In the last decades, systems have strongly increased their complexity in terms of number of functions that can be performed and quantity of relationships between functions and hardware as well as interactions of elements and disciplines concurring to the definition of the system. The growing complexity remarks the importance of defining methods and tools that improve the design, verification and validation of the system process: effectiveness and costs reduction without loss of confidence in the final product are the objectives that have to be pursued. Within the System Engineering context, the modern Model and Simulation based approach seems to be a promising strategy to meet the goals, because it reduces the wasted resources with respect to the traditional methods, saving money and tedious works. Model Based System Engineering (MBSE) starts from the idea that it is possible at any moment to verify, through simulation sessions and according to the phase of the life cycle, the feasibility, the capabilities and the performances of the system. Simulation is used during the engineering process and can be classified from fully numerical (i.e. all the equipment and conditions are reproduced as virtual model) to fully integrated hardware simulation (where the system is represented by real hardware and software modules in their operational environment). Within this range of simulations, a few important stages can be defined: algorithm in the loop (AIL), software in the loop (SIL), controller in the loop (CIL), hardware in the loop (HIL), and hybrid configurations among those. The research activity, in which this thesis is inserted, aims at defining and validating an iterative methodology (based on Model and Simulation approach) in support of engineering teams and devoted to improve the effectiveness of the design and verification of a space system with particular interest in Guidance Navigation and Control (GNC) subsystem. The choice of focusing on GNC derives from the common interest and background of the groups involved in this research program (ASSET at Politecnico di Torino and AvioSpace, an EADS company). Moreover, GNC system is sufficiently complex (demanding both specialist knowledge and system engineer skills) and vital for whatever spacecraft and, last but not least the verification of its behavior is difficult on ground because strong limitations on dynamics and environment reproduction arise. Considering that the verification should be performed along the entire product life cycle, a tool and a facility, a simulator, independent from the complexity level of the test and the stage of the project, is needed. This thesis deals with the design of the simulator, called StarSim, which is the real heart of the proposed methodology. It has been entirely designed and developed from the requirements definition to the software implementation and hardware construction, up to the assembly, integration and verification of the first simulator release. In addition, the development of this technology met the modern standards on software development and project management. StarSim is a unique and self-contained platform: this feature allows to mitigate the risk of incompatibility, misunderstandings and loss of information that may arise using different software, simulation tools and facilities along the various phases. Modularity, flexibility, speed, connectivity, real time operation, fidelity with real world, ease of data management, effectiveness and congruence of the outputs with respect to the inputs are the sought-after features in the StarSim design. For every iteration of the methodology, StarSim guarantees the possibility to verify the behavior of the system under test thanks to the permanent availability of virtual models, that substitute all those elements not yet available and all the non-reproducible dynamics and environmental conditions. StarSim provides a furnished and user friendly database of models and interfaces that cover different levels of detail and fidelity, and supports the updating of the database allowing the user to create custom models (following few, simple rules). Progressively, pieces of the on board software and hardware can be introduced without stopping the process of design and verification, avoiding delays and loss of resources. StarSim has been used for the first time with the CubeSats belonging to the e-st@r program. It is an educational project carried out by students and researchers of the “CubeSat Team Polito” in which StarSim has been mainly used for the payload development, an Active Attitude Determination and Control System, but StarSim’s capabilities have also been updated to evaluate functionalities, operations and performances of the entire satellite. AIL, SIL, CIL, HIL simulations have been performed along all the phases of the project, successfully verifying a great number of functional and operational requirements. In particular, attitude determination algorithms, control laws, modes of operation have been selected and verified; software has been developed step by step and the bugs-free executable files have been loaded on the micro-controller. All the interfaces and protocols as well as data and commands handling have been verified. Actuators, logic and electrical circuits have been designed, built and tested and sensors calibration has been performed. Problems such as real time and synchronization have been solved and a complete hardware in the loop simulation test campaign both for A-ADCS standalone and for the entire satellite has been performed, verifying the satisfaction of a great number of CubeSat functional and operational requirements. The case study represents the first validation of the methodology with the first release of StarSim. It has been proven that the methodology is effective in demonstrating that improving the design and verification activities is a key point to increase the confidence level in the success of a space mission

Development and Qualification of an FPGA-Based Multi-Processor System-on-Chip On-Board Computer for LEO Satellites

九州工業大学博士学位論文 学位記番号:工博甲第374号　学位授与年月日:平成26年9月26日Chapter 1: Introduction||Chapter 2: Background and Literature Review||Chapter 3: Multi-Processor System-on-Chip On-Borad Computer Design||Chapter 4: Space and Time Redundancy Trade-offs||Chapter 5: Radiation and Fault Injection Testing||Chapter 6: Thermal Vacuum Testing||Chapter 7: Results and Discussion||Chapter 8: Conclusion and Future Perspectives||ReferencesDeveloping small satellites for scientific and commercial purposes is emerging rapidly in the last decade. The future is still expected to carry more challenging services and designs to fulfill the growing needs for space based services. Nevertheless, there exists a big challenge in developing cost effective and highly efficient small satellites yet with accepted reliability and power consumption that is adequate to the mission capabilities. This challenge mandates the use of the recent developments in digital design techniques and technologies to strike the required balance between the four basic parameters: 1) Cost, 2) Performance, 3) Reliability and 4) Power consumption. This balance becomes even more stringent and harder to reach when the satellite mass reduces significantly. Mass reduction puts strict constraints on the power system in terms of the solar panels and the batteries. That fact creates the need to miniaturize the design of the subsystems as much as possible which can be viewed as the fifth parameter in the design balance dilemma. At Kyuhsu Institute of Technology-Japan we are investigating the use of SRAMbased Field Programmable Gate Arrays (FPGA) in building: 1) High performance, 2)Low cost, 3) Moderate power consumption and 4) Highly reliable Muti-Processor System-on-Chip (MPSoC) On-Board Computers (OBC) for future space missions and applications. This research tries to investigate how commercial grade SRAMbased FPGAs would perform in space and how to mitigate them against the space environment. Our methodology to answer that question depended on following formal design procedure for the OBC according to the space environment requirements then qualifying the design through extensive testing. We developed the MPSoC OBC with 4 complete embedded processor systems. The Inter Processor Communication (IPC) takes place through hardware First-In-First-Out (FIFO) mailboxes. One processor acts as the system master controller which monitors the operation and controls the reset and restore of the system in case of faults and the other three processors form Triple Modular Redundancy (TMR) fault tolerance architecture with each other. We used Dynamic Partial Reconfiguration (DPR) in scrubbing the configuration memory frames and correcting the faults that might exist. The system is implemented using a Virtex-5 LX50 commercial grade FPGA from Xilinx. The research also qualifies the design in the ground-simulated space environment conditions. We tested the implemented MPSoC OBC in Thermal Vacuum Chambers (TVC) at the Center of Nano-Satellite Testing (CeNT) at Kyushu Institute of Technology. Also we irradiated the design with proton accelerated beam at 65 MeV with fluxes of 10e06 and 3e06 particle/cm2/sec at the Takasaki Advanced Radiation Research Institute (TARRI). The TVC test results showed that the FPGA design exceeded the limits of normal operation for the commercial grade package at about 105 C°. Therefore, we mitigated the package using: 1) heat sink, 2) dynamic temperature management through operating frequency reduction from 100 MHz to 50 MHz and 3) reconfiguration to reduce the number of working processors to 2 instead of 4 by replacing the spaceredundancy TMR with time-redundancy TMR during the sunlight section of the orbit. The mitigation proved to be efficient and it even reduced the temperature from 105 C° to about 66 C° when the heat sink, frequency reduction, and reconfiguration techniques were used together. The radiation and the fault injection tests showed that mitigating the FPGA configuration frames through scrubbing are efficient when Single Bit Upsets (SBU) are recorded. Multiple Bit Upsets (MBU) are not well mitigated using the scrubbing with Single Error Correction Double Error Detection (SECDED) technique and the FPGA needs to be totally reset and reloaded when MBUs are detected in its configuration frames. However, as MBUs occurrence in space is very seldom and rare compared to SBUs, we consider that SECDED scrubbing is very efficient in decreasing the soft error rate and increasing the reliability of having error-free bitstreams. The reliability was proven to be at 0.9999 when the scrubbing rate was continuous at a period of 7.1 msec between complete scans of the FPGA bitstream. In the proton radiation tests we managed to develop a new technique to estimate the static cross section using internal scrubbing only without using external monitoring, control and scrubbing device. Fault injection was used to estimate the dynamic cross section in a cost effective alternative for estimating it through radiation test. The research proved through detailed testing that the 65 nm commercial grade SRAM-based FPGA can be used in future space missions. The MPSoC OBC design achieved an adequate balance between the performance, power, mass, and reliability requirements. Extensive testing and applying carefully crafted mitigation techniques were the key points to verify and validate the MPSoC OBC design. In-orbit validation through a scientific demonstration mission would be the next step for the future research