requirements for 'NetSat' mission with respect to structural and thermal limitations
Bhardwaj Shastri
Space Engineering, master's level (120 credits) 2019
Luleå University of Technology
Department of Computer Science, Electrical and Space Engineering
Houston : Roger, Tranquility, we copy you on the ground.
You’ve got a bunch of guys about to turn blue.
We’re breathing again. Thanks a lot.
Tranquility base : Thank you."
With an advancement in space technology, nowadays it is feasible for private sector companies to perform their individual missions like Earth observation, technology demonstration, scientific experiment, etc. Zentrum Für Telematik e.V. (ZFT) is also a child whose dream is to climb space in search of new hopes for knowledge. Based on a successful foundation of UWE missions and with collaboration with University Würzburg, ZFT is looking forward for its upcoming mission
"NetSat". With a very less heritage missions on formation flying at nano-satellite scale, NetSat will be one of a kind of technology demonstration mission, whose primary objective will be to check in-orbit formation flying capabilities of four nano-satellites (3U+ CubeSat) using electric propulsion system. NetSat is expected to be launch in first quarter of year 2020.
In the scope of this master thesis work, the proposed design for NetSat was analyzed for mis- sion and system requirements with respect to structural and thermal limitations. In order to check whether the proposed design can withstand all structural loads during it’s mission life- time, multiple structural analysis simulations were run in Fusion 360. Also, to satisfy thermal requirements for the mission, this master thesis provides a design of passive thermal control system for NetSat. A thermal space environment was setup in Siemens NX simcenter 3D space system thermal solver and multiple simulations were run to analyzed the design with regards to thermal loads and constraints during it’s mission lifetime. Different load case scenarios for structural and thermal analysis were considered during the process which have been discussed.
Based on results, the design is qualified and expected to satisfy all mission and system require- ments with regards to structural and thermal limitations.
I would like to express my sincere gratitude to Zentrum Für Telematik e. V. (ZFT), Würzburg, Germany for providing me with an unimaginable opportunity to undertake my thesis work. The kind of work atmosphere and learnings I was offered here with helped me grow into an adaptive professional. I would like to thank Mr. Ilham Mammadov for patiently guiding and supervising me on my work throughout my journey. I would also like to thank Mr. Roland Haber for his valuable advice and suggestions.
I am gratefully indebted to Mr. Daniel Eck (CEO, ZFT) and Mr. Julian Scharnagl (Head of Space Department, ZFT) for always keeping their doors open and welcoming to help me re- solve the problems I ran into. My sincere gratitude to Prof. Dr. Klaus Schilling for being there and critically analysing my work and helping me learn newer things.
I would also like to express my sincere gratitude to Enpulsion GmbH for providing me data on propulsion module and allowing me to play with it for my thesis work.
I would like to take an opportunity to sincerely thank Dr. Victoria Barabash for supporting me in every which way possible during my thesis journey. I am also hugely grateful to Dr.
Johnny Ejemalm for providing his valuable inputs and examining my work. I would also like to thank Ms. Anette Snällfot-Brändström, and Ms. Maria Winneback for their continuous help to tackle administrative tasks.
Lastly and most importantly, I must express my very profound gratitude to my parents, my sister and my uncle for providing me with unfailing support and continuous encouragement throughout my years of study and through the process of researching and writing this thesis.
This milestone of accomplishment would not have been possible without any of the above mentioned people. Thank you.
Yours Sincerely, Bhardwaj Shastri
Abstract i
Acknowledgement ii
List of Figures v
List of Tables vii
1 Introduction 1
1.1 NetSat Mission . . . 2
1.2 Outline of the report . . . 4
2 State of the Art 5 2.1 CubeSat : A trend in nano-satellites . . . 5
2.2 Satellite formation flying . . . 8
2.3 Formation flying at nano-satellite scale . . . 9
2.3.1 CanX - 4 & 5 mission . . . 9
2.3.2 CANYVAL-X mission . . . 9
2.3.3 AeroCube-4 mission . . . 9
2.3.4 STARS mission . . . 10
2.4 NetSat heritage . . . 10
3 Design and Development of NetSat 13 3.1 CubeSat Design Specification . . . 13
3.2 CAD model . . . 14
3.3 Subsystems . . . 15
3.3.1 Structure . . . 16
3.3.2 Propulsion module . . . 17
3.3.3 AOCS module . . . 19
3.3.4 OBDH module . . . 19
3.3.5 EPS module . . . 20
3.3.6 Payload module . . . 20
3.4.2 Mass budget . . . 23
3.5 Analytical model . . . 24
3.5.1 Finite Element Method (FEM) . . . 25
3.5.2 Finite Element Mesh generation . . . 26
3.5.3 Mesh model . . . 27
4 Analysis with respect to structural requirements 30 4.1 Modal analysis . . . 30
4.1.1 Theoretical background . . . 31
4.1.2 Launcher requirements . . . 32
4.1.3 Natural frequency of NetSat . . . 32
4.1.4 Simulation results . . . 33
4.2 Quasi-static Launch analysis . . . 36
4.2.1 Launcher requirements . . . 37
4.2.2 Arrangement of NetSat (Z axis) parallel to flight axis . . . 37
4.2.3 Arrangement of NetSat (Z axis) perpendicular to flight axis . . . 39
4.3 Structural analysis conclusion . . . 41
5 Analysis with respect to thermal requirements 43 5.1 Principles for Thermal Control System design . . . 43
5.2 TCS problem definition for NetSat . . . 47
5.2.1 Thermal requirements . . . 47
5.2.2 Problem definition . . . 47
5.3 NetSat TCS design . . . 48
5.4 Simulation environment implementation . . . 50
5.4.1 Solver setup . . . 50
5.4.2 Thermal loads setup . . . 52
5.5 Thermal load scenarios . . . 52
5.5.1 Load case 1 . . . 53
5.5.2 Load case 2 . . . 54
5.5.3 Load case 3 . . . 56
5.5.4 Load case 4 . . . 57
5.6 Thermal analysis conclusion . . . 58
6 Conclusion 59
Bibliography 61
1.1 NetSat orbit and formation configuration . . . 3
2.1 Number of CubeSats launched per year and per end use . . . 6
2.2 Success rate of CubeSats as a function of time . . . 7
2.3 Overview of UWE-4 subsystems . . . 11
2.4 Integration of NanoFEEP thruster in UWE-4 support rail and cut-away view . . 12
3.1 3U+ CubeSat Design Specification Drawing . . . 14
3.2 CAD model of NetSat . . . 14
3.3 Anatomy of NetSat . . . 15
3.4 NetSat rail structure . . . 16
3.5 NetSat frame structure . . . 16
3.6 NetSat subsystem support structure . . . 17
3.7 IFM nano thruster . . . 18
3.8 Analytical model of NetSat (Cutaway view) . . . 24
3.9 Analytical model of NetSat . . . 25
3.10 Smooth surface approximated by finite element . . . 26
3.11 Mesh model in Fusion 360 . . . 28
3.12 Mesh model in NX simcenter 3D space system thermal . . . 29
4.1 Spring-mass system with 1 DOF . . . 31
4.2 Result of 1st mode of modal frequencies analysis . . . 33
4.3 Result of 5thmode of modal frequencies analysis . . . 34
4.4 Result of 6thmode of modal frequencies analysis . . . 35
4.5 Primary structure of NetSat . . . 36
4.6 Arrangement of NetSat (Z axis) parallel to flight axis . . . 37
4.7 Von Mises stress result for load case 1 . . . 38
4.8 Structural deformation result for load case 1 (parallel orientation) . . . 39
4.9 Arrangement of NetSat (Z axis) perpendicular to flight axis . . . 39
4.10 Von Mises stress result for load case 1 . . . 40
4.11 Structural deformation for load case 1 (perpendicular orientation) . . . 41
4.13 Comparison of displacement results for load case 1 between Al6061-T6 (left)
vs Al7075-T6 (right) . . . 42
5.1 Example of perfect contact heat transfer through conduction within body . . . . 44
5.2 Example of gap radiation heat transfer through radiation within body . . . 44
5.3 Thermal problem of NetSat . . . 47
5.4 CAD model of thermal interface . . . 48
5.5 Thermal interface prototype . . . 49
5.6 Monte Carlo settings . . . 51
5.7 Load case 1 - Side facing towards the Sun . . . 53
5.8 Load case 1 - Side facing towards the Earth . . . 53
5.9 Load case 1 - TCS connected to side facing radiative environment . . . 54
5.10 Load case 2 - Two sides facing towards the Sun . . . 54
5.11 Load case 2 - Two sides facing towards the Earth . . . 55
5.12 Load case 2 - TCS connected to side facing towards the Sun . . . 55
5.13 Load case 3 - Side facing towards the Sun . . . 56
5.14 Load case 3 - Side facing towards the Earth . . . 56
5.15 Load case 3 - TCS connected to side facing towards the Sun . . . 57
5.16 Load case 4 - Side facing towards the Earth . . . 57
5.17 Load case 4 - TCS connected to side facing radiative environment . . . 58
2.1 Classification of satellites based on mass . . . 5
3.1 Performance details of IFM Nano thruster . . . 18
3.2 Mechanical and thermal properties comparison between Al6061 and Al7075 . . 21
3.3 Properties of solar cells . . . 22
3.4 Properties of FR-4 material . . . 22
3.5 List of materials used for different subsystems . . . 23
3.6 Iterative mass budget of NetSat . . . 23
3.7 Mesh parameters in Fusion 360 . . . 27
3.8 Mesh parameters in NX simcenter 3D space system thermal . . . 29
4.1 Natural frequency of NetSat in range of 5 - 2000 Hz . . . 32
4.2 Quasi-static launch requirements . . . 37
4.3 Load cases for parallel arrangement of NetSat with respect to flight axis . . . . 38
4.4 Load cases for perpendicular arrangement of NetSat with respect to flight axis . 40 5.1 Thermal requirements for NetSat . . . 47
5.2 List of available thermal control methods with their respective TRL and avail- ability . . . 48
5.3 Solver parameters for simulation setup . . . 50
5.4 Thermal loads for simulation setup . . . 52
AOCS Attitude and Orbit Control System.
AU Astronautical Unit.
CAD Computer-aided Design.
CDS CubeSat Design Specifications.
CMOS Complementary Metal-Oxide Semiconductor.
CoG Center of Gravity.
DAR Deviation Wavier Approval Request.
DOF Degree Of Freedom.
EPS Electrical Power System.
FAB Front Access Board.
FEA Finite Element Analysis.
FEEP Field Emission Electric Propulsion.
FEM Finite Element Method.
FRAM Ferro-electric Random Access Memory.
GPS Global Positioning System.
I2C Inter-Integrated Circuit.
IMU Inertial Measurement Unit.
JTAG Joint Test Action Group.
MoI Moment of Inertia.
NetSat Networked Nano-Satellite Distributed System Control.
OBC On-Board Computer.
OBDH On-Board Data Handling.
P-POD Poly Pico-satellite Orbital Deployer.
PCB Printed Circuit Board.
PPU Power Processing Unit.
PSLV Polar Satellite Launch Vehicle.
RAAN Right Ascension of the Ascending Node.
TCS Thermal Control System.
TI Texas Instruments.
TRL Technology Readiness Level.
UART Universal Asynchronous Receiver-Transmitter.
UHF Ultra High Frequency.
UNISEC University Space Engineering Consortium.
UWE University Würzburg Experimental.
ZFT Zentrum Für Telematik e.V..
Introduction
The in-orbit cooperation of several satellites in a formation will open new perspectives for a broad spectrum of applications and related theoretical background has been elaborated for years. Nowadays very small satellites offer the opportunity for realization at an affordable cost frame. Typical applications are Earth observation, navigation and communication, technology demonstration. For joint observations several satellites need to be pointed at the same target from different directions. Therefore position and attitude of satellites need to be measured with high accuracy in relation to the partner satellites. This information can then be used as basis for implementing cooperative control approaches.
In the past multi-satellite systems were mainly established in form of constellations (with each satellite individually controlled from ground) to increase coverage of Earth surface. Nowadays a paradigm change is evolving, aiming to use groups of very small satellites instead of single large satellite. The advantages are lower cost, better scanning ability and higher reliability. In a formation mission, satellites form a network to control closed loop in orbit position and attitude of the distributed satellite system. Thereby new perspectives for a broad spectrum of applica- tions open up, such as monitoring Earth surface or deep space, as well as measuring physical fields of the Earth.
In the past years many activities concerning research and development of small satellites with a size of a few tens of cm and a mass of only a few kg were carried out. This enables to de- crease launch cost, by mounting several of them on a single launcher. But the limitations of such pico/nano-satellites cause severe constraints for potential sensors and actuators as well.
1.1 NetSat Mission
With increase in more autonomous missions, spacecrafts require a high level of on-board self governing systems and inter-satellite communication. Such missions can be easily done with advantage of small satellite formations which can provide high spatial coverage and resolution at reasonable costs. Similarly, the Networked Nano-Satellite Distributed System Control (Net- Sat) mission under development at Zentrum für Telematik (ZFT) will demonstrate in-orbit the autonomous control of a formation flying of four nano-satellites (3U+ CubeSat) using electric propulsion and UHF inter-satellite links.[1]
The tentative launch for NetSat is in first quarter of 2020 and will be the first nano-satellites formation flying mission from Würzburg. The primary objective is to demonstrate technology that will enable future formation flying mission and thereby laying the foundation towards more application driven missions. NetSat will not carry any type of payload for Earth observation, but it will still test formation methodologies which will be require for future payload requirements and especially target pointing capabilities that make simultaneous Earth observations with mul- tiple satellites possible. These 3U+ CubeSats will test electric propulsion system that is IFM Nano Thruster from Enpulsion GmbH, which will enable precise orbit control for demonstrat- ing various formation within the available power budget.[1]
NetSat prepares contributions in areas of on-board self governing systems, distributed formation control, relative navigation, inter-satellite communication and protocols, as well as miniaturised attitude and orbit determination and control systems for future satellite formation flying mis- sions. NetSat heritage lies in previous missions in the area of pico-satellite technologies and distributed systems. Key technologies in the area of communications, attitude determination and control as well as robust on-board data handling and electric propulsion have already been demonstrated in-orbit in the scope of the University Würzburg Experimental (UWE) satellite program with UWE-1, UWE-2, UWE-3 and UWE-4.[1]
Orbit and formation configuration
The NetSat formation targets a near-polar sun-synchronous orbit with a mean altitude of 600 km.
Three satellites will be placed in nearly the same orbital plane in a Cartwheel type of configu- ration with the same eccentricity, and with the arguments of perigee separated by 120 deg and fourth satellite will be placed in a different plane with the same inclination but with an off- set in right ascension of the ascending node (RAAN) and a smaller eccentricity as shown in Figure 1.1.[2]
Figure 1.1: NetSat orbit and formation configuration[2]
The inter-satellite distances accounting for the radial and north-south gradients can be modified by adjusting the eccentricity of S1, S2 and S3. The inter-satellite distances accounting for the east-west gradient can be equally modified by adjusting the offset in RAAN between S4 and the plane defined by S1, S2 and S3.[2]
1.2 Outline of the report
For a successful mission, the structural design of the satellite plays a vital role in mission de- sign. It should be design in such a manner that it withstands the stress and other forces generated during the launch. CubeSat Design Specifications (CDS) provides a framework for a suitable design. However, excessive design would increase the mass, which will eventually increase the cost. Therefore, optimum design is the major necessity of CubeSat mission. The standard design of 3U+ CubeSat platform is used to integrate the propulsion module and to keep the mass distributed over entire satellite and the center of gravity (CoG) as per the CubeSat Design Specifications (CDS).
A detailed CAD model was created using Solidworks, later from the CAD model, the cen- ter of gravity (CoG) and the moment of inertia (MoI) were calculated. To validate the design, structural analysis was performed using Fusion 360. Proper loads distribution and constraints were applied as per the conditions during launch sequence. Also, a passive thermal control system (TCS) is designed based on the mission requirements. To analyze the thermal behaviour of designed TCS in space environment, software NX simcenter 3D space system thermal solver was used. A transient thermal analysis was performed to analyze the temperature distribution over the satellite and its subsystems. For analysis, orbital parameters, orientation of the satellite and radiation parameters were provided as input to simulate the orbital environment.
Chapter 2 contains the survey of the state-of-the-art for formation flying mission for nano satel- lites and trend of CubeSat.
Chapter 3 gives an overview of CDS and guidelines provided by CalPoly are discussed. Later in the chapter the overview on the design and development of NetSat platform including CAD model and its subsystems are discussed. It also discuss regarding the theory of Finite Element Method (FEM) and also details on implementation in the analytical model of NetSat in prepa- ration for further analysis is discussed. This chapter also provides information on the meshing process of the entire system.
Chapter 4, the environmental load conditions and the requirements for the structural analy- sis are discussed.
Chapter 5, design and analysis for TCS based on the orbital environment conditions and pa- rameters are discussed.
Lastly, chapter 6 summarize the obtained results and provide suggestions for future work.
State of the Art
2.1 CubeSat : A trend in nano-satellites
CubeSats have become a technology of center of attraction in the space sector. Nowadays Cube- Sats are being used for several space applications, such as education, Earth remote sensing and technology demonstration. Standard CubeSats are made up of 100 x 100 x 100 mm units, also called as 1U, designed to provide a cubic volume. 3U+ CubeSats are composed of three units of 1U stacked lengthwise along with an extension of cylindrical area of length 36 mm and diameter 72.5 mm. Table 2.1 shows different classification of satellites based on their mass. 1U CubeSat belong to the genre of pico-satellite, while 3U+ CubeSat belong to the class of nano-satellites.
On the other hand, 6U, 12U and 27U CubeSats belong to the group of micro-satellite.
Satellite Class Mass (kg) Large satellite >1000 Small satellite 500 - 1000 Mini satellite 100 - 500 Micro satellite 10 - 100 Nano satellite 1 - 10 Pico satellite <1
Table 2.1: Classification of satellites based on mass[3]
The initial purpose for the development of CubeSat was educational. The idea was to provide hands-on experience to students in space activities, allowing them to work on the entire cycle of a space project, from the initial concept until its operation in space. When CubeSats were launched in the early 2000s, there was a general perception that they were just toy satellites designed to fulfil the needs of student training or to meet some amateur demands.[4]
Figure 2.1: Number of CubeSats launched per year and per end use [4]
Later, it was understood that CubeSats could also be used for other applications, such as test- ing of technologies and science missions, like those related to astronomy and space weather.
From Figure 2.1, one can notice that after 2013, Earth remote sensing is the main application to which CubeSats have been devoted to. Technology demonstration also represents an important fraction of the total number of launched CubeSats in recent 5 years.
Moreover, one of the concern for CubeSat missions are their failure rate. CubeSats themselves cannot be blamed for all mission failures. Around 20% of all failures occurred either during launch or during the deployment phase [5]. When CubeSat fails during commissioning or dur- ing the early stages, it is said that “it died as an infant”. Infant mortality has been a big issue for CubeSats. Most of them failed as soon as they get into the space environment. [4]
Figure 2.2: Success rate of CubeSats as a function of time [4]
The evolution in CubeSat technology and the modernize techniques of new companies in man- ufacturing dedicated parts shows that CubeSats infant mortality rate is decreasing. Figure 2.2 shows the increase in success rate of the overall CubeSat missions as a function of time, cur- rently being about almost 75%.
The success rate was estimated by considering CubeSat missions as a Bernoulli experiment, so the success rate is considered as a parameter of a binomial distribution. It was considered a “success” if the CubeSat survived its early operational stages (deployed and commissioned, launch failures were excluded). Some factors have probably improved CubeSat success rate in the recent years, like the use of CubeSat international standards. As a consequence, CubeSats are becoming more reliable, as a strong foundation of flight heritage is being built up. [4]
2.2 Satellite formation flying
Small satellites are enabling multi satellite missions because of their small size and modular nature.Multiple small satellites can be flown instead of a much bigger and costlier conventional satellite for distributed sensing applications such as atmospheric sampling, distributed antennas, and synthetic apertures [6]. Missions with multiple small satellites can deliver a comparable mission capability than a monolithic satellite, but with significantly enhanced flexibility and robustness.
Multi satellite missions can be broadly divided into two categories, namely formation flying missions and constellation missions. The dynamic states of formation flying satellites are cou- pled through a common control law. In other words, in an formation flying mission, at least one satellite must track a desired state relative to another satellite, and its tracking control law must depend upon the status of another satellite. Multi satellite missions that do not satisfy the definition of formation flying missions are called constellation missions.[7]
For example, even though specific relative positions are actively maintained, GPS satellites constitute a constellation because their orbit corrections require only the individual satellite’s position and velocity. As per these definitions, satellites in formation flying missions must have active propulsion systems.[8]
Formation flying missions are further subdivided into two categories, namely formation fly- ing missions that involve rendezvous and docking and formation flying missions without dock- ing. The main challenges of formation flying missions occurs from dynamic couplings between satellites and the environment. If formation flying satellites are launched into LEO, they face environmental disturbances, such as atmospheric drag, solar pressure, and J 2 perturbations.
These disturbances can cause satellites to rapidly drift away from each other unless they main- tain their positions.[9]
Therefore, satellites need to counter these disturbances while maintaining their orbits and rel- ative distances and attitudes. If the desired positions are not in the same altitude, then the satellites have to expend additional control effort to synchronize their orbital periods and rel- ative distances. This challenge is exacerbated by the limited capabilities of the current sensor and actuator technologies for nano-satellites. Because of these challenges, formation flying missions is an active area for research and development.[10]
2.3 Formation flying at nano-satellite scale
Few formation flying missions at nano-satellites scale, consisting of only two satellites, have already been carried out by achieving most of their primary mission objectives. Few well- known formation flying missions are as follows :
2.3.1 CanX - 4 & 5 mission
The Canadian Advanced Nanospace Experiments 4 & 5 (CanX - 4 & 5) mission, led by the University of Toronto and primarily supported by Canadian Space Agency, is a dual-nano satel- lite mission that demonstrated satellite formation flying with submeter tracking error accuracy and low change in velocity requirements. Each nano-satellite was capable of attitude control accuracy of 1 deg using six coarse/fine sun sensors, a three-axis magnetometer, three rate gy- roscopes, three magnetorquer coils, and three orthogonally mounted reaction wheels. Satellites performed formation maneuvers using the Canadian Advanced Nano satellite Propulsion Sys- tem (CNAPS) that had a maximum thrust of 5 mN.
The CanX - 4 & 5 nano-satellites were successfully launched into sun-synchronous LEO on 30 June 2014. Using carrier-phase differential GPS techniques for extremely high precision relative navigation, two spacecrafts were reconfigured to perform projected-circular orbit for- mations in which one satellite appears to circle the other from a ground observer’s standpoint.
This mission currently sets the bar for state-of-the-art formation flying missions.[11]
2.3.2 CANYVAL-X mission
CANYVAL-X (CubeSat Astronomy by NASA and Yonsei using Virtual Telescope Alignment eXperiment) is a collaborative nano-satellite technology constellation of NASA, Korea’s Yonsei University and the KARI (Korea Aerospace Research Institute) with a primary objective of val- idating technologies that allow two spacecraft to fly in formation along an inertial line-of-sight (i.e., align two spacecraft to an inertial source). Demonstration of precision dual-spacecraft alignment achieving fine angular precision enables a variety of cutting-edge heliophysics and astrophysics missions. CANYVAL-X CubeSats were launched on 12 January 2018.[12]
2.3.3 AeroCube-4 mission
Three 1U CubeSats, called as AeroCube-4, were built by the Aerospace Corporation, which could control its attitude to 1 deg absolute accuracy using Earth and Sun sensors, a high-fidelity three-axis rate gyroscope and an IMU. These satellites could estimate its position with 20 m accuracy using a GPS receiver and control their position by varying their cross sectional area using extendable wings.
These satellites were launched into elliptical LEO on 13 September 2012. These satellites demonstrated formation flight by deliberately changing their drag profile and using different wing configurations, thereby re-configuring themselves over the course of several weeks.[13][14]
2.3.4 STARS mission
The Space Tethered Autonomous Robotic Satellite (STARS) mission, led by Kagawa University and Takamatsu National College of Technology, Japan, demonstrated undocking and docking of a daughter satellite with a mother satellite using a 10 m tether. The mother satellite first deployed the daughter by injecting an initial velocity,and then it retrieved it using the tether, and the daughter finally docked with the mother satellite. The mission was launched on 23 January 2009. The mother satellite determined its attitude using GPS, magnetometers, and gyroscopes and controlled it using magnetorquers. The daughter satellite determined its attitude with respect to the mother satellite using a camera and then controlled it using its own arm link motion under tether tension. Most of its objectives were achieved but faced instability problems in space.[15]
2.4 NetSat heritage
Starting with the UWE-1 launch in 2005, UWE pico-satellite program has seen so far four successful launches: UWE-1, UWE-2, UWE-3 and UWE-4. UWE-1, the first German pico- satellite, demonstrated in orbit the use of Internet protocols for space-ground communications, optimizing specific parameters for the space environment [16]. Following in 2009, UWE-2 ex- tended the internet in space experiments and emphasized technologies required for autonomous attitude and orbit determination [17].
In 2013, UWE-3 demonstrates autonomous real-time attitude determination and control. UWE- 3 introduced a significant redesign of the spacecraft bus, with focus on modularity, flexibil- ity and extensibility, constituting the reference architecture and design for the future program missions [18]. Within road map of formation flying CubeSats, UWE-4 was launched on 27 September 2018, with incorporation of propulsion system for the first time in UWE program.
Key satellite aspects to formation flying and technology transfer to products were handled by ZFT, while the University Würzburg focuses on basic research. Thus UWE-4 benefits from ZFT’s development related to OBC (On-Board Computer), as well as the attitude determination sensor suite and control system.[19].
The technical objective of the UWE-4 mission was to perform in-orbit demonstration and char- acterization of an electric propulsion system for 1U CubeSat. For this, the project cooperates with the TU Dresden that develops the Nano-FEEP propulsion system [20].
This system fits the CubeSat’s strict requirements in terms of size and power consumption, and has therefore been selected as the UWE-4 technical payload. The propulsion system consists of thruster heads that are being integrated into CubeSat rails and the PPU (Power Processing Unit).[21]
The primary technical mission objective was to activate thrusters and measure their thrust in different operating ranges. Secondary objectives include attitude control using four thrusters and eventually basic orbit control maneuvers [22]. The UWE-4 satellite is based on the archi- tecture introduced by UNISEC Europe which has been demonstrated first on UWE-3. It makes use of a backplane which interconnects all subsystems with a standardized interface and also interfaces the CubeSat’s panels.
This architecture supports rapid development, test, and integration of subsystems. It provides several redundant communication busses (I2C, UART) and power busses, dedicated synchro- nization signals, and complete debug access to each subsystem. The subsystem stack as shown in Figure 2.3 consists of OBC, AOCS, PPU, EPS ( Electrical Power Subsystem), UHF Commu- nication System, and FAB (Front Access Board).
Figure 2.3: Overview of UWE-4 subsystems [23]
The Nano-FEEP system, developed at TU Dresden, was selected for its compatibility with the 1U size and power restrictions. The system consists of thruster heads and two dedicated PPUs.
Thruster heads have been integrated into CubeSat rails as shown in Figure 2.4.
Figure 2.4: Integration of NanoFEEP thruster in UWE-4 support rail and cut-away view [24]
The thrust is generated through ionization and subsequent acceleration of small amounts of Gallium fuel. The fuel is stored in the thruster head (0.25 g each) and is heated to a temperature of about 50 °C at which the Gallium turns liquid and flows due to capillary forces along the porous needle to its tip. An electric voltage of up to 12 kV between the needle and the extractor cathode ejects the ions from the thruster by electrostatic force. The required voltage is generated from the unregulated battery voltage on one of the two PPUs which can provide up to 250 µA of current. Each PPU can interface and power two thruster heads and one neutralizer individually.
A single thruster can generate continuously a thrust level of up to 8 µN with peaks up to 20 µN and requires approximately 700 mW at 2 µN thrust.[24]
Design and Development of NetSat
3.1 CubeSat Design Specification
CubeSat Design Specification (CDS) was originally drafted by Californian Polytechnic State University in the year 1999. The document provides set of general as well as specific require- ments (Mechanical, Electrical and Testing) need to be followed during design life cycle of every CubeSat. CDS ensures the safety of the launch vehicle, CubeSat itself and other similar CubeSat launched during the mission. According to the CDS (CalPoly, 2014), 3U+ CubeSat’s mechanical requirements are stated below :
• The CubeSat shall use the coordinate system as defined in Figure 3.1. The -Z face of the CubeSat will be inserted first into P-POD. The CubeSat configuration and physical dimensions shall be per Figure 3.1.
• Exterior CubeSat components shall not contact the interior surface of P-POD, other than the designated CubeSat rails.
• Deployables shall be constrained by CubeSat. The P-POD rails and walls shall not to be used to constrain deployables.
• Rails shall have a minimum width of 8.5 mm. Rails shall not have a surface roughness greater than 1.6 µm. The edges of the rails shall be rounded to a radius of at least 1 mm.
• CubeSat shall not exceed 4.0 kg mass.
• Al7075 or Al6061 (aluminum alloy) shall be used for both the main CubeSat structure and rails. If other materials are used, the developer shall submit a DAR and adhere to the waiver process.
• The CubeSat rails and standoff, which contact P-POD rails and adjacent CubeSat stand- offs, shall be hard anodized aluminum to prevent any cold welding within P-POD.
Figure 3.1: 3U+ CubeSat Design Specification Drawing
3.2 CAD model
NetSat’s structure was designed with consideration to be more simple and efficient. With the intention to reduce the assembly time, the platform was designed to support all the subsys- tems through the backplane and FAB, which allows easy integration of all electronic modules through pin arrangement. The Figure 3.2 shows the CAD model of NetSat, indicating its outer components. Every side panel contains 7 solar panels along with GPS antenna and sun sensor.
The top plate contains 2 solar panels along with mechanism (designed according to CDS 3U+
CubeSat requirements) for deployment of four antennas (not shown in CAD model) while the base plate provides rigid support between rail structure and propulsion module.
Figure 3.2: CAD model of NetSat
3.3 Subsystems
The Figure 3.3 shows the anatomy of NetSat. The subsystems included are :
• Structure
• Propulsion module
• AOCS module
• OBDH module
• EPS module
• Communication module
• Payload module
Figure 3.3: Anatomy of NetSat
3.3.1 Structure
The structure of CubeSat provides the framework for the satellite as a whole and incorporates the other subsystems into its design while optimizing the use of volume and mass. The struc- tural integrity of a CubeSat must be maintained throughout handling, launch, deployment, and mission life in order to provide support and protection for all subsystems. The structure must be able to passively dissipate heat generated by internal components as well as insulate against incoming radiation. The structural design must ensure the safety of CubeSat throughout three major mechanical environments: ground, launch, and orbital. Thus, keeping in mind these conditions, the rail of NetSat is designed which is shown in Figure 3.4.
Figure 3.4: NetSat rail structure
3U+ CubeSat primary structure requires 6 separate parts : 4 rails, 1 top plate, and 1 bottom plate.
The arrangement is shown in Figure 3.5. The solid side panels allow the internal components to be incorporated into the satellite. They get attach to the rail along the side and to the top and bottom plates. All parts are manufactured from aluminium alloy Al7075-T6 and fasteners are manufactured from brass material to increase the likelihood of mission success.
Figure 3.5: NetSat frame structure
All subsystem modules are attached to frame structure through spacers connection as shown in Figure 3.6. For subsystems which are aligned perpendicular to longitudinal axis (Z axis) of CubeSat are supported by spacers connecting the rails at both ends. While, for payload module which is aligned parallel to longitudinal axis of CubeSat is supported by spacers connecting the bottom end of propulsion module and mid-support of rails. EPS modules, which are also aligned in parallel, are supported by spacers connecting two mid-support provided on rails.
Figure 3.6: NetSat subsystem support structure
3.3.2 Propulsion module
The IFM Nano Thruster addresses the urgent need of a propulsion system for micro and nano- satellites : its wide range of thrust, the excellent throttlability, and a high specific impulse, allow to significantly increase the mission range of such satellites in low orbits. The high specific impulse (Isp) on the other hand allows for very high ∆V manoeuvres at high propellant mass utilization efficiency. The small volume and its light mass make the thruster suitable for all nano-satellites (especially CubeSat). The combination of high Isp with medium very well controllable thrust levels in a small and light package makes the IFM Nano Thruster a strong competitor for existing colloid, cold gas, or Hall effect thrusters.[25] The main performance details of IFM Nano Thruster are listed in Table 3.1.
Dynamic thrust range 10 µN to 0.4 mN
Nominal thrust 350 µN
Specific impulse 2000 to 6000 s
Propellant mass 230 g
Total impulse more than 5000 Ns Power at nominal thrust 40 W incl. neutralizer Outside dimensions 100 x 100 x 82.5 mm
Mass (dry/wet) 670 / 900 g
Total system power 8 - 40 W
Hot standby power 3.5 W
Command interface RS422/RS485
Temperature envelope
(Non-operational) −40 to 105 °C Temperature envelope
(Opertional) −20 to 40 °C
Table 3.1: Performance details of IFM Nano thruster [26]
Mechanically, the thruster comprises a tank filled with the metal propellant and rigidly con- nected to the crown emitter. The extractor anode, which is part of the top plate, is separated via an isolator from the electronics compartment and the tank. An optional housing encloses all parts but is not required for mechanical stability. Electrically, the thruster is commanded through the CubeSat bus via UART. The central command and control module is connected to the independent power supplies for the heater, the two cathodes (two separate supplies), the emitter, and the extractor. The model presentation of IFM nano thruster is shown in Figure 3.7.
Figure 3.7: IFM nano thruster [26]
3.3.3 AOCS module
NetSat’s AOCS inherits its basic setup from UWE-4’s AOCS. It is implemented as standard subsystem carrying a low-power microcontroller that fuses sensor data from magnetometers, sun-sensors, and gyroscopes, and computes attitude and orbit control outputs for the satellite’s magnetorquer and propulsion system.
The sensor suite is enhanced, but the original sensors have been kept as backup. The mi- crocontroller now has access to a primary highly integrated MEMS 9-axis IMU (each 3-axis magnetometer, gyroscope, and accelerometer) placed on the AOCS board itself. Each side panel contains a mangetic torquer. Each side panel also carries a redundant IMU and a high precision sun-sensor. All sensors’ data can be injected into the Kalman filter sequentially and independently, such that a coarse attitude determination is also available during eclipse.
The sun-sensors are based on an ultra-low power miniature CMOS camera with a field of view of more than 90° at 250 x 250 pixel resolution and nominal power consumption of only 4.2 mW.
The magnetic torquers are placed on each panel with a total magnetic moment of 0.1 Am2 per axis and are mainly used for angular rate control. NetSat also carries miniaturized 3 pairs of reaction wheel mounted in between AOCS and power board, intended for control maneuvers.
3.3.4 OBDH module
OBDH is the core of the satellite and has control over all the other subsystems. OBDH consists of two micro-controllers functions in a Master-Slave configuration. The ZFT’s OBC features two redundant low-power commercial microprocessors fully interconnected in order to repair and restore one another in case of radiation induced failures. Both processors are monitored by a watchdog cascade in order to detect faulty behavior. UWE-3 and UWE-4 have proven its reliability by seamless operation since launch in 2013.
The purpose is to monitor the overall health status of CubeSat and to allow communication from ground with all subsystems through the redundant communication system. It’s power consumption of <15 mW guarantees that it is always switched on, even in severe low-power conditions. Furthermore, the OBC carries its own latchup-protection with automatic power cycling capability and a backup power conditioning unit.
The new development focus for NetSat mission is on an enhanced implementation of the re- dundancy concept for increased robustness and on extending the system’s debug access to other subsystems. Therefore, the system now carries two independent high precision real-time clocks, employs several independent FRAM chips with a total storage capacity of 40 Mbit, a set of 4 Gbit NAND Flash memory chips, and a pair of microSD card slots.
Furthermore, it implements debug interfaces to all other subsystems’ microprocessors which includes 4-Wire JTAG and Spy-Bi-Wire (2-Wire JTAG) for other TI mixed signal microcon- trollers, and SWD (Serial Wire Debug) for support of ATMEL ARM processors. This debug access further extends the system’s capability for fail-safe in-orbit software updates, extending significantly the already successfully demonstrated capabilities of UWE-3 and UWE-4.
3.3.5 EPS module
The EPS of NetSat continues the distributed architecture from earlier UWEs mission, with the only change of increase in number of boards. NetSat will carry 3 EPS modules to satisfy the maximum power demand. The solar cells’ power is directly tracked on each panel and supplied to batteries. Each EPS board carries two redundant 2.6 Ah Li-ion batteries and several power conditioning units for the regulated 3.3 V and 5 V power busses. An extra board has been also added to EPS module which will supply 12 V to satisfy the increased power demand. Power dis- tribution is realized through the standardized subsystem interfaces which are controlled through OBC.
These batteries are monitored and protected against low voltage conditions and the power con- ditioning units passively distribute the actual power demand among them. The NetSat’s power busses can be generated by one of the two power paths (battery & power conditioning units) or both in parallel for power demanding applications. The standardized subsystem interfaces pro- vide power monitoring, latchup-, over-voltage and under-voltage protection for each subsystem individually. Hence, the power switching of the NetSat has been revised from it’s heritage and the changes made will further enhance the overall satellite efficiency.
3.3.6 Payload module
Since the primary objective for NetSat mission is to check in-orbit formation flying capabilities with proposed propulsion system, hence there won’t be any payload. But few missions have been proposed based on the foundation which will be laid by successful NetSat mission.
The NetSat-SPG (NetSat-StereoPhotoGrammetry) mission will be an attempt to use a satel- lite formation consisting of four nano-satellites to obtain 3D pictures of the Earth surface in the visible and thermal-IR range for further processing by stereo-photogrammetric methods.
Stereo-photogrammetry measures the spatial position, shape and size of objects on a stereo pair of photographs. The final goal is the definition of the position and size of any objects on the surface of the Earth using a mix of four pictures of the same section of the Earth’s surface, obtained from four different perspectives [1].
The NetSat-4G (Global GeomaGnetic Gradiometry (4G) ) mission is another proposed con- cept, which will consists of four nano-satellites flying in a Cartwheel-Helix formation at low altitude and carrying vector magnetometers. The use of four satellites makes possible the real- isation of a full gradiometry mission, by simultaneously measuring the geomagnetic gradients in all three directions: east-west, north-south and radial.[1]
3.4 Physical properties
Frame structure According to CDS requirement, aluminium 6061 or 7071 alloy must be used for manufacturing of CubeSat frame structure. Typical mechanical and thermal properties of Al6061-T6 and Al7075-T6 are listed in Table 3.2.
Aluminium type Mechanical and Thermal properties Unit
Al 6061-T6 Al 7075-T6
Density g/cm3 2.7 2.81
Elastic modulus GPa 69 70
Elongation at break % 10 8
Fatigue strength MPa 96 160
Poison’s ratio 0.33 0.32
Shear strength MPa 210 330
Ultimate tensile strength MPa 310 560
Yield tensile strength MPa 270 480
Melting point °C 580 480
Specific heat capacity J/kg.K 900 870
Thermal conductivity W/m.K 170 130
Table 3.2: Mechanical and thermal properties comparison between Al6061 and Al7075
Multi-junction cells Multi-junction cells consist of multiple thin films, each essentially a solar cell grown on top of another, typically using MOVPE (metalorganic vapour phase epitaxy).
Each layer has a different band gap energy to allow it to absorb electromagnetic radiation over a different portion of the spectrum. Example of multi-junction cells used for space applications are GaAs, Ge, GaInP. Typical properties of multi-junction cells are listed in Table 3.3.
Base material GaInP / GaAs / Ge on Ge substrate
AR - coating TiOx / Al2O3
Dimensions 40.15 x 80.15 mm ± 0.1mm
Cell area 30.18 cm2
Average weight ≤ 116 mg/cm2
Thickness 280 + 25 µm
Cover glass CMX 100
Cover glass thickness 100 µm
Table 3.3: Properties of solar cells
FR-4 FR-4 (Flame Retardant) is a NEMA (National Electrical Manufacturers Association) grade designation for glass-reinforced epoxy laminate material. FR-4 is a composite material composed of woven fiberglass cloth with an epoxy resin binder that is flame resistant. FR-4 is commonly used material for printed circuit boards (PCBs). A thin layer of copper foil is laminated to one or both sides of FR-4 glass epoxy panel, usually referred to as copperclad laminates. FR-4 does not specify any specific material, but instead a grade of material. Typical physical and electrical properties of FR-4 are listed in Table 3.4.
Material FR-4
Density 1.850 g/cm3
Thermal conductivity 0.29 W/m.K Tensile strength 415 MPa
Young’s modulus 24 GPa
Poison’s ratio 0.136
Temperature index 140°C Relative permittivity 4.4
Table 3.4: Properties of FR-4 material
3.4.1 Materials
The list of materials used for different subsystems of NetSat are listed in Table 3.5. The same list of materials are applied to respective components during structural and thermal analysis.
Subsystem Components Material
Rail Al7075-T6
Subsystem support Brass Structure
Screws Brass
Thruster Al7075-T6
Propulsion module
PPU FR-4
Thermal control module Thermal interface Copper
Solar cells GaInP/GaAs/Ge Solar panels
Base plate Al7075-T6
Battery case Al7075-T6 EPS module
Battery Lithium
AOCS module Reaction wheels Al7075-T6
All PCBs FR-4
Table 3.5: List of materials used for different subsystems
3.4.2 Mass budget
When applied to a space mission, budgeting refers to more than just money. Due to the limi- tations of launch vehicle’s requirements and design constraints, project capacities are generally managed in a budget format. In order to understand the structure mass complexity, the mass budget was broken down by subsystem which is listed in Table 3.6.
Subsystem Mass (g)
Structure 1353.585
Propulsion module 900.267
EPS module 656.850
AOCS module 213.450
Communication module 81.880
OBDH module 57.600
Total mass (g) 3263.632
Table 3.6: Iterative mass budget of NetSat
3.5 Analytical model
A simple analytical model is the easiest model to work with while doing finite element analysis.
While some application may not allow for a simple model, efforts to focus on required analysis solution and simplifying the detailed CAD model into analytical is rewarded with accuracy and time. If CAD model contains small regions or small geometries, this could cause to have pro- portionally small elements in that area during meshing and cause higher stresses than what is actual. This will cause up unnecessary computation power and time. So getting rid of that area or converting them into a single element is a key to create analytical model.
The process of setting up the analytical model can have an impact on number of elements required in the analysis. We could increase the number of elements, but they would not result in dramatic improvement in the results, this is referred to as convergence. Convergence is when that an increase in elements does not result in more accurate results for the additional compu- tation time. The analytical model of NetSat, derived from its detailed CAD model is shown in Figure 3.8 and Figure 3.9.
Figure 3.8: Analytical model of NetSat (Cutaway view)
Figure 3.9: Analytical model of NetSat
3.5.1 Finite Element Method (FEM)
FEM is a numerical technique that provides approximate solutions to the governing equations of a complicated system through a discretisation process. Before an analysis is carried out, the entire system has to be divided into a number of individual subsystems or components whose behaviour is readily understood. The basic units of the discretised subsystems are called finite elements, which should neither overlap nor have gaps between each other. The finite elements used for a domain need not be of the same type, and the properties could also vary. Figure 3.10 shows how a smooth curved surface, as defined by function φ, is modelled by elements of various types.
Figure 3.10: Smooth surface approximated by finite element [27]
When three-node triangular elements are used, the φ surface is approximated by flat triangular facets, whereas the four-node and eight-node quadratic elements are able to represent warped and curved surfaces and can thus better approximate the actual function. Obviously, the ap- proximation can also be improved by using more elements instead of increasing the order of the interpolation polynomial. This sketch illustrates the basic idea of the finite element method:
moreover, approximation of a smooth function by means of simple polynomials, each of which is defined over a small region (element) and represented in terms of the values of the function at the element nodes.[27]
3.5.2 Finite Element Mesh generation
A finite element mesh is a partition of a given domain into sub-domains, which are called ele- ments, such that every point of the domain is found in one of the elements. The entire domain has to be covered by the elements without overlapping, and the conditions of compatibility be- tween finite elements on the boundary have to be satisfied as well. Two-dimensional domains can be discretised into triangular, quadrilateral or a mixture of triangular and quadrilateral el- ements. Over three-dimensional domains, tetrahedral and hexahedral elements can be used;
however, in some situations, wedges or pentahedral elements and pyramid elements could also be employed.[27]
As the topology of curved surfaces locally resembles that of a planar domain, similar to a two-dimensional problem, triangular and quadrilateral elements can be generated on surfaces.
To reduce discretisation (numerical) error in a finite element analysis, the quality of the finite element meshes has to be optimised such that the element size is in compliance with the speci- fied nodal spacing and that the shape of the elements ought to be as equilateral as possible. For conforming meshes, the boundary nodes of the finite element mesh have to lie on the boundary surface of the given domain, and for constrained meshes, apart from the geometrical require- ments of a conforming mesh, additional topological requirements such as specified edges and faces have to be present in the mesh as well. Furthermore, a higher-order curvilinear can also be employed to fit domains with curved boundaries to reduce discretisation error.[27]
3.5.3 Mesh model
Since two different types of software were used : Fusion 360 for structural analysis and NX space system thermal solver for thermal analysis, hence two mesh models from analytical model were created in respective softwares. The brief overview of both mesh models are explained below.
Fusion 360 2-D (quad) elements were used to mesh analytical model in Fusion 360. This is because the 2-D meshes need a thickness definition to accurately represent a body. The four- noded elements are also best suited for rectangular surfaces, which are in abundance in the CubeSat. The mesh paramaters used are listed in Table 3.7 and Figure 3.11 shows the mesh generated for NetSat and its subsystems.
Finite element mesh summary
Solids -
Scale mesh size per part No
Average element size (absolute value) 10 mm
Element order Parabolic
Create curved mesh elements Yes
Max. turn angle on curves (deg.) 60
Max. adjacent mesh size ratio 1.5
Max. aspect ratio 10
Minimum element size (% of average size) 20
Number of nodes 725271
Number of elements 410370
Table 3.7: Mesh parameters in Fusion 360
Figure 3.11: Mesh model in Fusion 360
NX space system thermal solver The combination of Quad4 elements and Tri3 elements were used to understand more efficiently the temperature differences during different iteration.
The mesh parameters used during the simulation are listed in Table 3.8 and Figure 3.12 shows the mesh model generated in NX simcenter 3D space system thermal solver.
Finite element mesh summary
Meshing method Subdivision
Average element size (absolute value) 3.68 mm
Create curved mesh elements Yes
Max. turn angle on curves (deg) 50 Min. element size (% of average size) 5 Total number of elements in the part 125011 Total number of nodes in the part 123991
Number of Quad4 elements 122867
Number of Tri3 elements 2144
Table 3.8: Mesh parameters in NX simcenter 3D space system thermal
Figure 3.12: Mesh model in NX simcenter 3D space system thermal
Analysis with respect to structural requirements
4.1 Modal analysis
All things vibrate at some particular frequencies. Prolonged vibration is detrimental to struc- tures. It is also frequently unavoidable. Unchecked vibration can lead to eventual structure failure. Vibration is all about frequencies. Frequency is defined as so many cycles in a given time period. One cycle per second is equivalent to 1 Hertz.
Structures exhibit multiple natural frequencies of vibration when excited by an imposed force, acceleration, or displacement. The way that the structure moves for a particular natural fre- quency is referred to as the mode shape. Mode shapes might involve bending, twisting, elon- gation and contraction, or a combination of these effects. There are several types of modal frequencies: Rigid body mode, fundamental mode and harmonic mode. The following factors influence the natural frequencies and mode shapes:
• The shape of the structure.
• The mass of the structure and how that mass is distributed.
• The way the structure is constrained.
• The stiffness of the material and structure.
• The tensile or compressive loads applied to the structure.
The major purpose of modal frequencies analysis is to design the structure so that resonance does not occur due to any potential driving frequency or design must withstand the momentary resonance produced.
4.1.1 Theoretical background
The mathematical process behind determining natural frequencies and the corresponding mode shapes through FEA is complex. Let’s consider a very simple linear spring-mass system with one degree of freedom (DOF) as shown in Figure 4.1, where m is mass and k is stiffness of spring. The mass is only free to move in a single direction.
Figure 4.1: Spring-mass system with 1 DOF
Now, Hooke’s law defines the force F required to displace a spring of stiffness k at a distance X:
F = k.X (4.1)
Newton’s second law of motion, specifies the relationship between force F, mass m, and accel- eration a for our 1 DOF system:
F = m.a (4.2)
From equation 4.1 and equation 4.2, we get the relationship between mass, acceleration, spring stiffness and displacement,
m.a = k.X (4.3)
Now, the angular frequency of oscillation ω, in radians/second, is given by
ω = (k/m)1/2 (4.4)
There are 2π radians per vibration cycle. Therefore, the natural frequency of oscillation (fn), in Hertz is given by
fn= (k/m)1/2.2π (4.5)
In an actual finite element model, we have 3D motion and many DOF due to all of the nodes and elements in the mesh. In a continuous system, there is an infinite number of vibrating modes. However, in a finite element model, there is a finite number of DOF, and therefore a finite number of vibration modes. Determining the vibration frequencies and mode shapes for these complex 3D systems involves matrix operations, eigenvalues, and eigenvectors.[28]
4.1.2 Launcher requirements
The selection of launch vehicle is one of the important steps to determine the launch scenario.
Quasi-static launch loads and natural frequencies of rocket will determine the CubeSat launch requirements [29]. The potential launcher planned for NetSat mission is Polar Satellite Launch Vehicle (PSLV) from India. Predicting of loads that comes from the launch vehicle is one of the hardest steps of designing a spacecraft. Because of the complexity and high variety of mis- sion environments little inaccuracies in the finite element models are capable of causing large errors.[30]
During its launch, a satellite is subject to various external loads resulting from steady-state booster acceleration, vibro-acoustic noise, air turbulence, gusts, propulsion system engine vi- brations, booster ignition and burnout, stage separations, vehicle maneuvers, propellant slosh, payload fairing separation and ejection. These sources’ characteristic feature is being random and independent [31]. According to PSLV requirements, the CubeSat mounted shall not have resonance frequencies below 100 Hz. Also, natural frequencies in the range of 5 - 2000 Hz of CubeSat mounted shall be analysed.[32]
4.1.3 Natural frequency of NetSat
According to launcher requirements, natural frequencies in range 5 - 2000 Hz were calculated through analysis. The corresponding modes obtained are shown in Table 4.1.
Mode
Natural frequency
(Hz)
Mode
Natural frequency
(Hz)
Mode
Natural frequency
(Hz)
Mode
Natural frequency
(Hz)
1 227.3 11 591 21 943.9 31 1658
2 280.8 12 636.6 22 999.2 32 1722
3 306.3 13 678.3 23 1021 33 1773
4 320 14 731.8 24 1087 34 1815
5 398 15 751.4 25 1133 35 1847
6 406 16 787.9 26 1259 36 1864
7 427.9 17 856.6 27 1358 37 1942
8 461.3 18 901.4 28 1443 38 1946
9 559.1 19 924.3 29 1489 39 1954
10 584.6 20 932.2 30 1557 40 1986
Table 4.1: Natural frequency of NetSat in range of 5 - 2000 Hz
4.1.4 Simulation results
The displacement results are displayed as a color contour plot. Modal displacement results are only meaningful in that they demonstrate the deformed shape of the structure for each vibration mode. The absolute magnitudes of the modal displacements are not indicators of how much the structure actually deflects when vibrating. The modal displacements are not scaled to any specific excitation. They are normalized, which means that the maximum displacement is al- ways one length unit, and the minimum displacement is always zero. The range is fixed from zero to one length unit for every model and every vibration mode. The legend shows the natural frequency and mode number associated with the currently displayed deformed shape.[28]
Figure 4.2 shows the simulation result for 1st (lowest) mode of natural frequencies analysis for NetSat. According to the result, the lowest mode is approximately 227.3 Hz and since this value is higher than the launcher requirement (100 Hz), it is acceptable.
Figure 4.2: Result of 1st mode of modal frequencies analysis
Figure 4.3 shows the natural frequency corresponding to 5th mode. At approximately 398 Hz, structure starts to vibrate with minimum displacement.
Figure 4.3: Result of 5thmode of modal frequencies analysis
Figure 4.4 shows the natural frequency corresponding to 6th mode. At approximately 406 Hz, structure starts to vibrate with maximum displacement.
Figure 4.4: Result of 6thmode of modal frequencies analysis
4.2 Quasi-static Launch analysis
Every event generates structural loads in the life of a CubeSat starting from launch till deployed into orbit. Even though launch causes the highest value loads for structures; any other event can be critical and significant for some parts of the structure, such pre-launch preparations, payload separation, etc. During the launch phase, the CubeSat has to endure the acceleration loading governing from the launching process. In order to determine the stress level on the CubeSat due to these acceleration loads, quasi static launch analysis are performed.
Basically a CubeSat structure can be classified into two types : primary and secondary. The primary structure consists of those elements which react to the overall structure deformation like bending, axial, shear, and torsional loads. The primary structures of NetSat comprise of the CubeSat rails, spacers, midplane PCB and top plate of propulsion module. The secondary structure comprises of those elements which do not appreciably contribute to overall stiffness.
In NetSat, the secondary structures include the side panels, base plate and panel screws.
Quasi-static launch analysis includes all the loads acting on the primary structure of CubeSat.
The two important parameters of the analysis are Structural deformation (amount of change in elements of the structure) and Von-Mises Stress (stress tensor of a material at a given time).
Figure 4.5 shows the primary structure of NetSat.
Figure 4.5: Primary structure of NetSat
4.2.1 Launcher requirements
According to PSLV’s requirements, the loads applied during launch are shown in Table 4.2 [32].
Also, the requirement from P-POD has to be fulfilled. The structural deformation of CubeSat shall not exceed 100 microns.[33]
Quasi-static Launch requirements
Axis Load
Flight axis 11 g
Lateral axis (both) 6 g
Table 4.2: Quasi-static launch requirements
One of the important notes is the orientation of the deployment system in the launch vehicle.
The orientation of CubeSat with respect to the flight axis is generally not known during early design stages. Therefore to verify the robustness of the structure, extreme load conditions on all axis (X, Y, and Z) are considered. The loading conditions of NetSat are analysed under two arrangements : Arrangement of NetSat (Z axis) parallel to flight axis and Arrangement of NetSat (Z axis) perpendicular to flight axis.
4.2.2 Arrangement of NetSat (Z axis) parallel to flight axis
Figure 4.6 shows the arrangement of NetSat (Z axis) with respect to flight axis.
Figure 4.6: Arrangement of NetSat (Z axis) parallel to flight axis
Considering parallel orientation of NetSat, different load cases were analysed, which are shown in Table 4.3. From the table, one can notice that structure develops a maximum stress of 27.23 MPa in load case 4, which is well below the yield stress for Aluminum 7075-T6. Simu- lation results for load case 1 are shown in Figure 4.7 and Figure 4.8.
Load case Axis Loads
Maximum Displacement
(microns)
Maximum Von Mises stress (MPa)
-Z 11 g
1 -X, -Y 6 g 14.28 27.17
-Z 11 g
2 +X, +Y 6 g 14.25 26.38
-Z 11 g
3 +X, -Y 6 g 14.24 26.67
-Z 11 g
4 -X, +Y 6 g 14.31 27. 23
Table 4.3: Load cases for parallel arrangement of NetSat with respect to flight axis
Figure 4.7 shows Von Mises stress of maximum 27.17 MPa generated at bottom of rails where it is constrained.
Figure 4.7: Von Mises stress result for load case 1
Figure 4.8 shows the maximum displacement of 14.28 microns on upper section of NetSat, which is well within mentioned requirements.
Figure 4.8: Structural deformation result for load case 1 (parallel orientation)
4.2.3 Arrangement of NetSat (Z axis) perpendicular to flight axis
Figure 4.9 shows the arrangement of NetSat (Z axis) with respect to flight axis.
Figure 4.9: Arrangement of NetSat (Z axis) perpendicular to flight axis
Considering perpendicular orientation of NetSat, different load cases were also analysed, which are shown in Table 4.4. From the table, one can notice that structure develops a maximum stress of 14.25 MPa in load case 2, which is well below the yield stress for Aluminium 7075-T6.
Simulation results for load case 1 are shown in Figure 4.10 and Figure 4.11.
Load case Axis Loads
Maximum Displacement
(microns)
Maximum Von Mises stress (MPa)
-Y 11 g
1 -X, -Z 6 g 1.248 13.61
-Y 11 g
2 +X, +Z 6 g 1.637 14.25
-Y 11 g
3 +X, -Z 6 g 1.459 13.86
-Y 11 g
4 -X, +Z 6 g 1.403 14.02
Table 4.4: Load cases for perpendicular arrangement of NetSat with respect to flight axis
Figure 4.10 shows Von Mises stress of maximum 13.61 MPa generated on upper rails.
Figure 4.10: Von Mises stress result for load case 1
Figure 4.11 shows the maximum displacement of 1.248 microns, which is also well within mentioned requirements.
Figure 4.11: Structural deformation for load case 1 (perpendicular orientation)
4.3 Structural analysis conclusion
The primary structure proved to be capable of withstanding all launch scenarios and extreme loading conditions as per requirements. Hence, there is no need of providing extra "x" cross connections between rails. Also, the lowest natural frequency is above the launcher resonance frequency requirement, still it is suggested to provide extra support for FAB module in order to increase the lowest natural frequency with good margin from required launcher resonance frequency.
All the above mentioned simulations were done by using Al7075-T6 material as potential can- didate. Also, the same case scenarios were run considering Al6061-T6 as comparison option.
One of the comparison between Al6061-T6 and Al7075-T6 is discussed. Taking into account parallel orientation of NetSat with respect to flight axis and considering load case 1, the simu- lation results are compared, which are shown in Figure 4.12 and Figure 4.13.
Figure 4.12 shows comparison of Von Mises stress results. Figure on left indicates results for Al6061-T6 while figure on right indicates results for Al7075-T6. Maximum stress generated by Al6061-T6 structure is 26.06 MPa while by Al7075-T6 is 27.17 MPa. Similarly, Figure 4.13 shows comparison of structural deformation results. Maximum displacement shown by Al6061- T6 structure is 14.7 microns while by Al7075-T6 is 14.28 microns. Henceforth, the choice of material will be totally depend on factors of availability and manufacturing cost.
Figure 4.12: Comparison of Von Mises stress results for load case 1 between Al6061-T6 (left) vs Al7075-T6 (right)
Figure 4.13: Comparison of displacement results for load case 1 between Al6061-T6 (left) vs Al7075-T6 (right)
Analysis with respect to thermal requirements
5.1 Principles for Thermal Control System design
Most of heat transfer takes place through conduction or radiation (within the components or with space environment). There is also a small amount of convection when propulsion module (propellant) or heat pipes (fluid in heat pipes) are used. The space environment is vacuum at a temperature ∼3 K. When there is direct contact between two surfaces, heat transfer through conduction takes place, while radiation does not require any surface contacts. The direction of the conductive heat flow is in the direction of the temperature differential. The most commonly used steady and transient conduction equations in one dimensional flow are
Q = k.A.˙ δT
δx; (5.1)
δT
δx = K.δ2T
δx2 (5.2)
where, ˙Q is rate of heat transfer, k is thermal conductivity of the medium, A is Area of cross- section of the heat flow, T is Temperature, x is linear dimension, δTδx is gradient of temperature;
K is thermal diffusivity.[34]
The radiation flow depends on the difference of the fourth powers of the respective temper- atures of the surfaces. The net radiative heat transfer law is given by the Stefan-Boltzmann equation,
Q = .σ.A.(T˙ 14 − T24) (5.3)
where, is emissivity; σ is Stefan-Boltzmann constant, A is Area of emitting surface, T1 is Temperature of emitting surface, T2is Temperature of surroundings.[34]
