i MASTER OF SCIENCE IN AEROSPACE ENGINEERING
SPACE CURRICULUM
University Of Pisa
Preliminary Assessment of the Power
subsystem and the Communication
subsystem for D3SAT mission.
Anupam Parihar
Academic Year 2014-2015
iii
UNIVERSITY OF PISA
MASTER OF SCIENCE IN AEROSPACE ENGINEERING SPACE CURRICULUM
Preliminary Assessment of the Power
subsystem and the Communication
subsystem for D3SAT mission
Supervisors : Candidate:
Prof. Ing. S. Marcuccio
Ing. S. Gregucci
Anupam Parihar
3
“I listen and forget, I see and I remember, I make and I
understand.”(Ancient Chinese proverb)
“The true sign of intelligence is not knowledge but
imagination.” Albert Einstein
5
ACKNOWLEDGEMENTS
I sincerely thank my advisor Professor Salvo Marcuccio for his guidance, suggestion, and continuous support throughout my thesis. I greatly appreciate the opportunity given by him to work on this topic of D3SAT project.
My profound thanks to Engineer Stefan Gregucci at Sitael S.p.A for his valuable suggestion and help in carrying out this dissertation work. Lastly but not the least I express my sincere thanks to all my friends, colleagues and parents who have patiently extended all sort of help for accomplishing this undertaking.
Pisa, November, 2015
7
ABSTRACT
This thesis aims to design the power and communication subsystems of D3SAT (Drag make up and de-orbiting demonstration satellite) mission. This mission aims to demonstrate SITAEL's Hall-Effect thruster HT-100D in-orbit by mounting it on a microsatellite which has a mass less than 40 kilograms. This thesis deals with the analysis of orbital operations required for performance evaluation, namely drag compensation and end-of-life de-orbiting. .Starting with the communication subsystem, housekeeping data budgets are created and link budget analysis is performed by defining different RF transmission losses and exploiting UHF/VHF/S-band cases. Power subsystem discusses the power requirements of D3SAT in different modes of mission, several scenarios of drag compensation and de-orbiting are analyzed in order to minimize the solar array area, mass and de-orbiting time. An accurate thermal model of the solar cell has been developed to calculate the maximum power generation as function of the temperature in space. . Both of sub-systems are found to be compliant with the stringent volume, mass and power requirements on-board small satellites The HT-100D is found to efficiently satisfy the mission requirements with very low power requirements and low propellant consumption, thereby proving to be an ideal candidate to enhance the capabilities of small satellites.
9
TABLE OF CONTENTS
Thesis structure ... 1
Chapter
1 ... 3
D3SAT Mission ... 3
1.1 Small satellites: Classifications ... 3
1.2 Applications of small satellites ... 4
1.3 Predicted growth in satellite market ... 6
1.4 Flying at very low-Altitude and de-orbiting ... 9
1.5 D3SAT mission ... 11
Chapter
2 ... 14
State-Of-Art: Power, Communication and Propulsion
Subsystems ... 14
2.1 State of the Art ... 14
2.2 SITAEL’s SpA HT-100D: low power hall effect thruster ... 21
Chapter
3 ... 25
Communication Subsystem ... 25
3.1 Requirements and assumptions ... 25
3.2 D3SAT’s data acquisition system ... 26
3.3 D3SAT’s On-Board computer ... 27
3.4 OBC Architecture ... 28
3.5 House-keeping data budget ... 30
3.6 Selection of Hardware for communication subsystem .... 32
10
3.7.1 Link Budget Calculations ... 36
3.8 Calculation for the auxiliary memory space and transmission time ... 50
Chapter
4 ... 53
D3SAT’s Power Budget ... 53
4.1 Requirements and constraints ... 53
4.2 Power Consumption On-board D3Sat ... 54
4.3 Power Sub-system Modes ... 55
4.4 Power System Architecture of D3Sat ... 57
4.5 Selection of solar cells, battery for D3sat and Depth of Discharge (DOD) ... 58
Chapter
5 ... 62
Thermal Modelling of D3SAT’s solar array ... 62
5.1 Determination of Eclipse time for D3SAT ... 62
5.2 Solar Array allocation on D3SAT body frame ... 66
5.3 Degradation of D3Sat solar cells ... 70
5.4 Thermal Analysis of solar cell ... 71
5.5 Thermal model for D3SAT’s solar array ... 73
Chapter
6 ... 82
Power Sub-system simulation: ... 82
Determination of Solar Array Area ... 82
6.1 EPS Simulation ... 82
6.3 Power subsystem simulation for D3SAT’s End of life de-orbiting phase... 93
6.4 Summary of the Power Sub-system Simulation for the determination of Solar Array Area ... 104
11 6.5 Calculation of D3SAT’s semi-controlled de-orbiting time
... 105
Conclusion ... 111
Future Work ... 114
12
LIST OF FIGURES
CHAPTER 1
Figure 1.1: Classification of satellites [2] ... 4
Figure 1.2: Satellite launches forecast [5]Euroconsult 2014] ... 6
Figure: 1.3: Satellite Market distribution [8] ... 7
Figure 1.4: Global Satellite (1-100kg) launches 2015-2024[9] ... 8
Figure 1.5 : Forecasted SpaceX satcom constellation[8]Extreme tech] ... 9
Figure 1.12: NRLMSISE density model [15] ... 10
CHAPTER 2
Figure 2.1: High specific energy batteries (clyde space) and triangular advanced solar cells... 15Figure 2.2 : S-band patch antenna (Surrey tech); S-band transmitter(Cubesat shop) ... 16
Figure 2.3: Micro-cavity discharge thruster (CU aerospace) ... 17
Figure 2.4: Resistojet (SITAEL SpA) ... 18
Figure 2.5: Electrospray thruster [24] ... 18
Figure 2.6: Laser Plasma thruster (Photonic associates) ... 20
Figure 2.7: Hall Effect thruster (Goebel), HT-100D (Sitael SpA) 20 Figure 2.8: SITAEL SpA HT-100D[29] ... 22
Figure 2.9: HT-100D system architecture [29] ... 23
CHAPTER 3
Figure 3.1: temperature sensor [ intersil ] ... 26Figure 3.2: 12 bit Analog-digital converter (Texas inst.) ... 27
Figure 3.3: On-board computer (IMT) ... 28
Figure 3.4: D3SAT’s OBC star architecture [35] ... 30
Figure 3.5: Drag Compensation mode data budget ... 31
Figure 3.6: House-keeping data budget ... 32
Figure 3.7: a) S-band transmitter b) UHF/VHF transceiver (courtesy cubesat shop®) ... 34
Figure 3.8: D3SAT’s ground station antenna (cubesat shop ®)[44] ... 34
13
Figure 3.9: Transmission Losses [47] ... 38
Figure 3.10: Ionospheric Scintillation [49] ... 39
Figure 3.11: Gas absorption attenuation [50] ... 40
Figure 3.12: Attenuation (UHF-band) Vs Angle of elevation of D3SAT’s ground pass ... 41
Figure 3.13: Attenuation (S-band) Vs Angle of elevation of D3SAT’s ground pass ... 42
Figure 3.14: Antenna Misalignment [46] ... 42
Fig 3.15: D3SAT’s ground station and environment [52] ... 43
Figure 3.16 Galaxy noise influence in noise temperature [47] ... 44
Figure 3.17: Cold sky temperature VS angle of elevation ... 45
Figure 3.18: Ground passes of D3SAT over Pisa (1 month simulation) ... 46
Figure 3.19: BPSK modulation [54] ... 47
Figure 3.20: Link margin, elevation angle VS passes (S-BAND) . 48 Figure 3.21: Link margin, elevation angle VS passes (UHF-BAND) ... 48
Figure 3.22: D3SAT’s uplink Budget ... 49
Figure 3.23: Dwell time VS Passes ... 50
CHAPTER 4 Figure 4.1: PCDU (ThalesAlenia space) ... 54
Figure 4.3: Azurspace Solar Cell [62] & SAFT Li-ion Secondary battery [63]... 59
Figure 4.4: Battery Cycle Vs DOD [53] ... 60
CHAPTER 5 Figure 5.1: Orbit Beta Angle [65] ... 62
Figure 5.2: RAAN and orbit angle beta (in this figure RAAN > angle beta) [53][65] ... 64
Figure 5.3 a) Orbit angle beta variation throughout year (starting on 21st March) ... 64
Figure 5.3b: Variation in eclipse time throughout the year starting on 21st March. (21st March) ... 65
Figure 5.3 c) Orbit angle beta variation throughout year (starting on 21st March) ... 65
14 Figure 5.3b: Variation in eclipse time throughout the year starting on 21st March. (21st March) ... 66 Figure 5.4: Solar Unit vector in spacecraft body frame [56] ... 67 Figure 5.5a: Solar Array allocation on D3SAT for Orbital injection parameter RAAN=0 degrees ... 68 Figure 5.5b: Solar Array allocation on D3SAT for Orbital injection parameter RAAN= 90degrees ... 68 Figure 5.5: Solar Unit vector in spacecraft body frame (Top view) [56] ... 69 Figure 5.6: Effect of temperature on the I-V characteristics [68] .. 72 Figure 5.7: Maximum Power (single cell) VS temperature [62] ... 73 Figure 5.8: Simplified Solar Array thermal model [69] ... 74 Figure 5.9: Example of Deorbit scenario of D3SAT for thermal modelling example... 75 Figure 5.10: angle phi and lambda [56] ... 76 Figure 5.11: D3SAT’s Solar Array assembly [69] ... 77 Figure 5.12: temperature profile of D3SAT solar array (shade represents the earth shadow on the orbit) ... 79
CHAPTER 6
Figure 6.1: 1st part of determining the temperature profile in-orbit ... 83
Figure 6.2: 2nd part of EPS flow chart determining the power generation and total solar array area. ... 84 Figure 6.3: 3rd part of simulation showing the battery charging .... 85 Figure 6.4 : In-orbit power modes of D3SAT (firing during daylight) ... 86
Figure 6.5 : Power generation (in blue) and Power Requirement profile (in red) VS time. (shade-eclipse period) ... 87 Figure 6.6: Battery load VS time ... 87 Figure 6.7: Battery charge VS time (6 orbits) for daylight HT100D firing ... 88 Figure 6.8 : In-orbit power modes of D3SAT (firing during eclipse ) ... 89
Figure 6.9: Power generation (in blue) and Power Requirement profile (in red) VS time... 89 Figure 6.10: Battery Load VS time ... 90
15
Figure 6.11: Battery Charge VS time (6 orbits) ... 90
Figure 6.12: Power generation (in blue) and Power Requirement profile (in red) VS time... 91
Figure 6.13: Battery Load VS time ... 92
Figure 6.14: Battery Charge VS time (6 orbits) ... 92
Figure 6.15: Deorbit orbit profile of D3SAT ... 95
Figure 6.16: Power generation (in blue) and Power Requirement profile (in red) VS time... 95
Figure 6.17: Battery Load VS time ... 96
Figure 6.18: Battery Charge VS time (5 orbits) ... 96
Figure 6.19: Battery Load VS time ... 97
Figure 6.20: Battery Charge VS time (5 orbits) ... 98
Figure 6.21: Power generation (in blue) and Load (in red) VS time( 3 orbits) ... 99
Figure 6.22: Battery Charge VS time (2 orbits) ... 99
Figure 6.23: Power generation (in blue) and Load (in red) VS time( 3 orbits) ... 100
Figure 6.25: Battery Charge VS time (6 orbits) ... 101
Figure 6.26: Power generation (in blue) and Load (in red) VS time( 6 orbits) ... 102
Figure 6.27: Battery Charge VS time (6 orbits) ... 103
Figure 6.28 : NRLMSISE-00 density model ... 107
Figure 6.29: semi-controlled orbit decay of D3SAT at EOL ... 109
Figure 6.30: Orbit lowering of D3SAT at End-of-Life from 350 to 250 km ... 110
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LIST OF TABLES
Table 1.1: Orbital Injection Parameter ... 12
TABLE 2.1 HT-100D Specifications [30] : ... 22
Table 3.1: On-board computer specification ... 28
Table 3.2: Technical specification of S-band transmitter [40] ... 33
Table 3.3: Technical specification of UHF/VHF transceiver [41] 33 Table 3.4: technical specification of S-band patch antenna [42] ... 33
Table 3.5: Omni directional antenna specification UHF (GOMspace®)[43] ... 34
Table 3.6: Ground-Station Specifications [44] ... 35
Table 4.1 Power consumption by different subsystem: ... 55
Table 4.2: Drag Compensation Mode (Peak Power requirement) . 56 Table 4.3: Housekeeping Mode (Min Power Requirement) ... 56
Table 4.4: Summary for the selection of components. ... 60
TABLE 5.3: Thermal loads on solar array from fig 5.9 [69]: ... 75
Table 5.4: Specific Capacity of solar array module ... 77
Table 6.1: Solar Array Area and Power Mass budget for different cases: ... 104
1
Thesis structure
This thesis aims to design the power and communication subsystems of D3SAT (Drag make up and de-orbiting demonstration satellite) mission. This mission aims to demonstrate SITAEL's Hall-Effect thruster HT-100D in-orbit by mounting it on a microsatellite which has a mass less than 40 kilograms.
The thesis is structured into several chapters
Chapter 1
This chapter starts by classifying the small satellites into different groups and their application depending on the user requirements. In the second part, there is statistical data shown about how the market growth is forecasted to take place in the small satellite industries. In the third part, the advantages of flying in the very low earth orbit is described in terms of profit in the revenue, example of earth observation imagery is given, this part also describes the consequences of flying in the very low earth orbit and as a countermeasure D3SAT mission is proposed with mission description is given at the end of the chapter.
Chapter 2
This chapter describes the recent state-of-art technologies in power, communication and propulsion subsystems. In the propulsion subsystem several miniaturized propulsion working principles are demonstrated and in the end part of the chapter the D3SAT’s payload Sitael’s HT-100D Hall Effect thruster and its system architecture is explained quantatively.
Chapter 3
The first part of this chapter describes the requirement which D3SAT’s communication subsystem should fulfill for which housekeeping data budgets are made. On-board-computer and its architecture is designed and in the end part link budget analysis is done with the help of STK
2 ® to determine the link margin as function of elevation angle during D3SAT ground pass
Chapter 4 to Chapter 6
These three chapters are concentrated on designing the power subsystem of D3SAT.Power budgets are made for different modes of D3SAT mission and Depth-of-discharge of secondary batteries are calculated. Thermal model of solar array in-orbit with the help of numerical integration is developed in chapter 5 which later is implemented in the power sub-system simulation. With the help of power sub-system simulation, the solar array area is determined for different scenarios of HT-100D firing during the drag-compensation phase and de-orbiting phase. At the end of chapter 6 the time taken to de-orbit D3SAT at the end-of life is determined
3
Chapter
1
D3SAT Mission
1.1 Small satellites: Classifications
In 1957, the Soviet Union launched the first artificial satellite, Sputnik 1. The weight of that satellite was over 80 kilograms. After Sputnik 1, satellites have grown heavier, larger and more complicated till the last decade. But with each kilogram costing about 10,000$ [1] to put into orbit. Through technical minimization, micro-satellite are currently of increasing interest. Their possibilities as well as their scope of missions is growing steadily. Today, payloads with a mass of just few kilograms are able to perform measurements that would have been unthinkable a few years ago. The major advantage is the fast and cheap development of micro-satellites, which makes them a suitable platform for technology evaluation. Hence they provide the ideal opportunity to test new systems in space within a short timeframe and low budget.
4 Figure 1.1: Classification of satellites [2]
1.2 Applications of small satellites
One of the more useful ways to explain ongoing interest in designing, building, launching, and operating small satellites is to break the market down by categories of users.
i) Military Applications: The outbreak of hostilities or an emergency situation prompted by a terrorist attack can occur with little or no warning. The military has adapted to such needs by having small and dedicated satellites that can be launched with little advanced warning. This has motivated the satellite component suppliers to develop components (i.e., antennas, power supply, processors, stabilization systems, and thrusters) that could be quickly assembled and launched on short-term notice. These innovations have helped others be able to order small satellites with much quicker delivery schedules. Military relies heavily on the satellite communication, for instance several military operations are now relying extensively on the Iridium and Globalstar small satellite constellation for mobile communications services [3].
ii) Small satellites for scientific and educational applications: Perhaps the predominant application of small satellites is for educational
5 projects and scientific experimentation. On the other end of spectrum, there can be quite small but sophisticated experimental satellites designed by the world’s leading universities or space agencies. New Millennium Space Technology 5 project .This project consists of three micro-satellites (each 25 kg in mass) that have been measuring Earth’s magnetic field [3]
iii) Small satellites for amateur radios, emergency, disaster relief: The other important application of small satellites can be to support emergency services, disaster relief, or medical or health services in very rural and remote areas where conventional communications or other services are not present. Livesat operated a two-satellite low Earth orbit messaging service to provide medical information as data relay on demand service. [3]
iv) Start-up programs in various countries: Many countries that are just beginning a space program – or embarking on scientific measurement programs where a spacecraft is the optimal approach to take – typically embark on a small satellite program. [3]
v) Commercial constellations: For some missions only one spacecraft required to accomplish a mission but for some mission applications more than one satellite has to be used. Some of the examples are Communication relays, Earth Observation, Earth Science, Global Navigation Satellite systems etc. When talking about constellation the focus is on mission utility rather than spacecraft utility for example in the first DMC (Disaster Monitoring constellation) one of the key requirements from the user needs was the DMC was to have a revisit period of 24 hours on every location on the globe in order to detect natural disasters, the requirement flow down produced constellation consisting of 5 satellites of mass 68kg placed in 686 km sun-synchronous orbit. [4]
vi) Technology demonstration: Technology demonstration
will continue to use small satellites for the foreseeable future.
In addition to validating improvements for existing
technologies, small satellite technology will be used to
6
demonstrate new technologies such as interferometry, new
propulsion systems, solar sails, inflatable structures and
constellation flying for various structures. Number of
technology demonstration satellite to increase in direct
relation to the increasing number of space-faring countries,
and accelerating rate of technological developments. [5]
1.3 Predicted growth in satellite market
It is estimated that between 2014 and 2023, a total of 1.155 satellites
(Fig 1(a)) will be manufactured and orbited, worth a collective of 248 billion dollars. The report [6] attributes 25 percent of this projected business to commercial space companies, with the majority, 75 percent, coming from government programs.As seen in the report [6], it is predicted that the demand of Earth Observation will increase in the coming decade due to following:
7 1)High ground resolution and frequent revisit are key requirements for defense sector.[7]
2)Natural Resource Monitoring will be a priority for developing economies.[7]
3)Drop in oil prices is expecting to increase the demand of resource monitoring.[7]
4)Location based Services and users are expected to diversify.[7] 5)Disaster management efforts co-ordinated by International organisations will increase which will demand responsive high resolution data.[7]
6)Public sector users and Civil sector users will grow in Maritime operations.[7]
7) Research community and public users will increase in Environment monitoring sector which also demands specific instrumentation due to strong variation in datas.[7]
8) Europe will be good pool of market in EO due to increasing border tensions in the east.[7]
Figure: 1.3: Satellite Market distribution [8]
In the report produced by [8], it is forecasted that in the next 5 years till 2019 the Europe will increase its share in satellite market by 26% .
8 Figure 1.4: Global Satellite (1-100kg) launches 2015-2024[9]
NSR’s Nano and Microsatellite Markets, 2nd Edition found that Earth observation (EO) has emerged as key driver for this industry’s growth, with large constellations being planned and deployed by startups such as Planet Labs, Spire, Satellogic, PlanetiQ, and Black Sky Global, to name a few. Earth Observation, a market dominated by defense & intelligence needs, scores above other small satellite applications due to a situation of ‘data poverty’ in industry verticals such as agriculture, disaster management, forestry and wildlife, and financial services. In its research, NSR found that 40% of the nano and microsatellites planned to be launched by the end of year 2024 will be for EO applications, which will contribute a staggering 58% of the total manufacturing market (~ $400 million).[9]
Application of small satellites in providing satellite internet service also forcasted to boom in fututre. Space X announced to deploy a network of 4000 small satellites[10] in Low earth orbit that would beam wireless Internet access to Earth from space. These SpaceX satellites would be set up in a low-Earth orbit, allowing them to reach more areas and hand off connections to improve reliability and speed. Similar plans have been announced by Google and microsoft motivating satellite component manufacturers to boost the development and research for optimizing and miniaturizing subsystem components in order to maximize their share in commercial satellite market missions.[10]
9 Figure 1.5 : Forecasted SpaceX satcom constellation[8]Extreme tech]
Statistical data presented in this chapter shows a rapid growth in small satellite market (application in Earth observation and communication) in coming decade which is expected to set new standards in the satellite industry in terms of innovation in manufacturing, operations.
1.4 Flying at very low-Altitude and de-orbiting
Nowdays, the driving parameter for the survival of a company in EO market is dictated by the revenue collection (users civil,government or defence) ,majority of revenue collection depends on the imagery price and imagery price depends on the resolution of the image , the image will earn a good price if is more detailed i.e having a large number of bits (having low resolution)[11] . The resolution depends on the altitude and diameter of the aperture of the payload . By looking at the relationship of Rayleigh diffraction limit[12] :
𝑋′ = 2.44ℎ𝐷𝜆 (1.1)
Where X’ is the image resolution, h is the orbit altitude, λ is the wavelength of the electromagnetic wave that the sensor is capturing, and D is the diameter of the aperture. This relation is for nadir pointing sensor, but the relation between resolution and altitude still hold. It can be seen that at very low orbit altitude we can employ a small payload (small diameter of the aperture) and get an improved resolution
10 (low value of X’ ) image just like the large payload (large diameter aperture) operating at an high altitude.More savings in mission costs are achieved due to the impact of the payload size on the satellite and on the launch cost .[13]
The main problem of flying at low altitude is the loss of orbit energy due to the atmospheric drag which is projected in the direction against the spacecraft motion along the orbital trajectory. Force induced by atmospheric drag is given by[14]:
𝐹𝐷𝑅𝐴𝐺 = 12𝜌 𝑉∞2 (𝐶
𝐷 𝑆) (1.2) Where 𝜌 is the variable density of the atmosphere (fig1.12, x-axis is the height and on y-axis density is plotted) 𝑉∞ is the orbital velocity 𝐶𝐷 𝑆 are the product of the the drag coefficient and frontal cross sectional area of the spacecraft. In these parameters the most variable parameter is the atmospheric density which changes with altitude and the time of the year. In fig 1.12 it can be seen that the atmospheric density is quite high for very lower altitudes.
Figure 1.12: NRLMSISE density model [15]
The atmospheric drag can lead to dicomissioning a satellite in very less time of its mission and might not attain large profits, but this
11 problem of atmospheric drag can be very well compensated by using miniaturised space propulsion engines which are available in the market. With the help of State-of-Art propulsion, small satellites can compensate for drag by using electric propulsion and chemical propulsion techniques [18]. One of the state of art propulsion Hall Effect thruster HT-100D developed by Sitael S.p.A is a space qualified Hall Effect thruster (described in chapter 2) to perform in-orbit its drag make-up capabilities and end-of-life disposal. The Hall effect thrusters high specific impulse can provide the microsatellite with several years of mission duration at very low altitude ensuring the existence and competetivness of company in the EO market with very high quality imagery to sell and if there is propellant saved on board, the spacecraft can also deorbit itself clearing its orbit and preventing the formation od space debris since, the increase of the spatial debris in outer space leads to tremendous potential safety hazard [15] to the satellites which are in working condition. There are 22000 satellites and other traceable objects around earth and in [16] it shows that the cumulative distribution of collision generated debris in low earth orbit peaks around the orbit inclination angles at sun-synchronized orbits. In near future the concentration of satellites in this orbit would increase. Therefore the development of the space debris mitigation and removal measures are being taken seriously nowadays. Among several techniques: chemical propulsion, electric propulsion, drag augmentation techniques and electrodynamic tethers are used at the end of life of mission to perform a re-entry maneuver [17]. Hall-Effect thruster which comes in the family of electric propulsion could be used to deorbit the spacecraft at the end of its life
1.5 D3SAT mission
The proposal of D3SAT mission is to successfully demonstrate the drag make-up capabilities of HT-100D in Very low earth orbit by the integration of the thruster on a microsatellite platform whose dimensions are constrained to 40cm X 40cm X 40cm and mass below 40kg and since the specific impulse of HT-100D is very high [18] , the propellant left on-board D3SAT could be use to deomnstrate a successful end of life disposal of D3SAT.
12 At the time of writing of this work the orbital injection parameters of D3Sat were not fixed because the injection altitude of the spacecraft will be strictly dependent on the selection of launcher and its apogee altitude. The D3SAT team of ADCS and mission analysis worked on determining the probable altitude from several launchers and it was estimated that there would be an opportunity to launch D3SAT in a 350km and 750km circular Sun-synchronous orbit. The value of the orbital parameter RAAN will depend on the launcher chosen and is unknown at the time of thesis writing.
Table 1.1: D3SAT’s Orbital Injection Parameter
Altitude [19] 350 Km, 750Km Eccentricity [19] 0 (Circular)
Inclination [19] Sun-Synchronous 96.8 degrees Sun-Synchronous 98 degrees
RAAN Unknown
14
Chapter
2
State-Of-Art: Power, Communication and
Propulsion Subsystems
2.1 State of the Art
In recent years small spacecraft have become more attractive due to lower development costs and shorter lead times. There is a natural trade-off to be made between spacecraft size and functionality, but advances in both miniaturization and integration technologies have diminished the scope of that trade-off. Commercial off the shelf components (COTS) and consumer electronics are commonly used to build small spacecraft at the lower end of the cost range.[18] Since the work in this thesis mainly deals with Power subsystem, Communication subsystems and Propulsion subsystem, therefore State of the art of these subsystems will be discussed in brief .
2.1.1 Power generation and Energy storage
Triple juction Solar cells are being employed nowdays reaching upto the efficiency of 29% (38% in laboratory) [18]. Advancement in the coverglass material has ensured protective covering of solar cells
15 in order to resist light-reflection, darkening and ultraviolet radiation damage and increase overall End of Life effeciency of the solar cell which is the most important design parameter in power generation subsystem. Triangular Advanced Solar Cells (TASC) are being manufactured which has the advantage of odd form factors removing the necessity to cut individual solar cells. Increasing the number of junctions generally offers the potential to reach even higher efficiencies, but material quality and the choice of bandgap energies turn out to be even more importance than the number of junctions. Several four-junction solar cell architectures with optimum bandgap combination are found for lattice-mismatched III–V semiconductors as high bandgap materials predominantly possess smaller lattice constant than low bandgap materials. Direct wafer bonding offers a new opportunity to combine such mismatched materials through a permanent, electrically conductive and optically transparent interface[19]. Wafer bonded multijunction solar cells are in research reaching maximum effeciencies upto 34% for space applications[20]. Lithium- ion batteries due to their high specific energy densities (200W/kg) are used as secondry batteries and the latest lithium polymer batteries in flat form are like the ones which are used in modern mobile phones. In some cases acceptance testing (accoustic pressure,vibrations,radiation testings) are performed on individual COTS battery cells and then they are assembled in battery packs according to mission requirements.[18]
Figure 2.1: High specific energy batteries (clyde space) and triangular advanced solar cells
16
2.1.2 Communication Subsystem
The general purpose of any communication device onboard spacecraft is to maximize the data rate and minimize hardare constraints, power consumption and complexity. Nowdays UHF/VHF transmitters are reliable and are low cost (few thousand dollars) solution for missions requiring nominal amount of data transfer. These systems are typically used in LEO with omnidirectional antennas[18], and therefore do not require a high level of pointing accuracy.S-band transmitters are a popular communication system being used on recent small satellite launches.These transmitters provide a high data rate upto 10Mbps and has very small size to fit in small satellites. Microstrip antenna have also gain popularity in recent small satellite missions as they minimize mass and size requirement and still maintaining a good signal strength output.
Figure 2.2 : S-band patch antenna (Surrey tech); S-band transmitter(Cubesat shop)
2.1.3 Propulsion Subsystems
Small satellites are becoming increasingly popular tools for Earth-imaging, communications, and other applications. But they have major control issues: Once in space, they can’t accurately point cameras or change orbit, and they usually crash and burn within a few months.[21] Small spacecraft propulsion is rapidly growing technology field. The SOA in this field consists of various propulsion technologies satisfying the size constraints of cubesat and microsat. The future of propulsion technology is diverse with both chemical and electric propulsion
17 options on track to mature within the next few years.Some of them would be described in the proceeding sections.
i) Micro-cavity discharge thruster : The MCD thruster (see Figure 1.8) is an extremely small-scale electro thermal thruster which utilizes a capacitively coupled dielectric barrier plasma discharge as the primary heating mechanism. The fundamental structure of the MCD thruster is two aluminum foils coated in a nano-porous aluminum dioxide dielectric layer, where the nano-porous Al2O3 layer has a honeycomb-like composition. A micro-cavity of approximately 100-300 μm diameter is drilled through each foil. The foils are aligned with concentric cavities, creating a single cavity between the two foils. Applying an AC waveform to the foils causes the configuration to act as a capacitor. The micro-cavity between the two foils concentrates the fringe electric field effects of the capacitor, resulting in an electric field. [22] A high pressure feed source supplies a propellant gas to the thruster, and as it enters the cavity the electric field initiates a Paschen-type breakdown of the gas into a plasma, which in turn increases the bulk temperature of the propellant. The high-temperature gas is then exhausted through a moderate Reynolds number nozzle producing thrust levels in the range of 1-5 mN.
Figure 2.3: Micro-cavity discharge thruster (CU aerospace)
ii) Resistojet propulsion system: Resistojet propulsion systems pass a propellant (typically gas) through an expansion slot with heated walls.
18 As the gas molecules become more heated, their energy increases, and they pass through the expansion slot with greater speed, producing the thrust and specific impulse necessary to power satellites [23] .
Figure 2.4: Resistojet (SITAEL SpA)
iii) Electrospray thruster: The principle of electrospray thrusters is straightforward: a high voltage is applied between a conductive liquid in a capillary and an annular extractor electrode. Above a critical voltage, the electrostatic forces deform the liquid into a cone and a spray of charged droplets or ions is accelerated towards the extractor electrode [24] Micro machined Propulsion systems for very small satellites].
Figure 2.5: Electrospray thruster [24]
iv) Cold gas system: The simplest propulsion system available to small spacecraft vents a cold, pressurized gas through a nozzle. The specific impulse of a cold nitrogen gas system is less than 75 sec and thrust levels are less than 5 N. The system does not have a pump and is
19 referred to as a blow down system, where the pressure of the system decreases with time. It is possible to have a high-pressure tank with a regulator to vent the gas at a lower pressure for a longer amount of time, but the total impulse delivered is the same since it is a function of the pressure force over time. [25]
v) Monopropellant and bipropellant: Chemical propulsion systems use a chemical reaction to produce a high-pressure, high-temperature gas that accelerates out of a nozzle. Chemical propellant can be liquid, solid or a hybrid of both. Liquid propellants can be a monopropellant passed through a catalyst. A more conventional bipropellant is a mix of oxidizer and fuel. A solid rocket motor contains both an oxidizer and a fuel that are molded into various grain patterns. The benefits of monopropellants and solid systems include relatively low-complexity/high-thrust output, low power requirements, and high reliability. Liquid and hybrid systems can be stopped and re-started, and in some cases throttled, whereas solid motors can only be used once. The highest thrust and highest specific impulse systems are bipropellant but they are more complex, not miniaturized, and are not meant for low thrust applications during the writing of this work no bipropellant propulsion were considered suitable for small satellite application[18]. v) Laser plasma thruster: Its operation is based on laser-controlled ignition of an energetic material in a proprietary fuel tape (Figure 2). The fuel cannot detonate in an uncontrolled way. Because the firing rate and tape speed can be controlled over a factor of 100 in the current design, thrust can be varied from 0.1 to 10 mN, because of the contribution of the chemical energy in the tape coating. The input power is given in terms of micro-newton thrust per watt. [26].
20
Figure 2.6: Laser Plasma thruster (Photonic associates)
vi) Hall Effect thruster: propellant (Xenon) is used and principle relies on electrons source emitted from hollow cathode placed outside thruster. These electrons follow the opposite way of the electric field created between the anode through xenon is injected and the cathode. A coaxial channel placed between the two electrodes is fed with xenon. Inside this channel is realized an ionization of the neutral gas, then ions go on the direction of the hollow cathode and are exhausted outside the thruster providing propulsion. A radial magnetic field is applied across the channel to contain electrons and reduce their velocity increasing number of collisions and ionizations [27].
21
Conclusion
The future of propulsion technology is diverse with both chemical and electric propulsion options on track to mature within the next five years [18]. Electric propulsion engines are more efficient then chemical engines, in the sense that they require much less propellant to produce the same overall effect i.e. a particular in spacecraft velocity. The propellant is ejected up to twenty times faster than from classical thrusters and therefore the same propelling force is obtained with twenty times less propellant [28] which means saving the propellant on board which could be utilized for other purposes of station-keeping, attitude controls for prolonged mission duration or disposal of spacecraft at the end of mission. Since the aim of the thesis is technology demonstration of HT-100D, a variant of state of art hall effect thruster manufactured at SITAEL S.pA , the next section will describe the HT-100D propulsion system.
Mounting low-power SITAEL’s HT-100D (Hall Effect Thruster) on the microsatellite platform and implementing a suitable drag compensation strategy while taking into consideration power and mass constraint can yield a very good EO performance of the microsatellite at a very low altitude and also extending the mission duration.
2.2 SITAEL’s SpA HT-100D: low power hall effect thruster
Sitael’s HT 100 Hall Effect Thruster (HET) is the smallest and lowest power-consuming HET ever developed in Europe [29]. HT 100D has been conceived for application on mini- and micro-satellites with limited onboard available power and volume. Its most relevant design feature is the use of permanent magnets instead of electromagnets for the generation of the required magnetic field. This design choice resulted in an extremely compact device that has an overall weight of 436g. Its design has been performed in order to optimize the thruster performance and to increase the lifetime. In addition, the thermal design has been improved using a passive cooling system able to reduce the thruster working temperature for a fixed power level or to increase the operating power range. Furthermore22 HT100D is equipped with an internal stainless steel stiffness structure able to sustain all loads, even launch loads. [30]
Figure 2.8: SITAEL SpA HT-100D[29]
TABLE 2.1 HT-100D Specifications [30] : Performance parameter Thrust (mN) 6-18 Power (W) 120-400 Specific impulse [s] 1000-1600 Thrust efficiency 40 % Lifetime [hr] 2000 Mass of HT-100D [kg] 0.436 Mass of Cathode 0.2 Mass of PPU 0.7
PPU power input 28 V; 2.5 A
2.2.1 HT-100D propulsion system architecture
As shown in Figure 1.13, HT-100D propulsion system is composed of the thruster assembly (HET and Cathode), the power processing unit (PPU), and the Propellant Management assembly.
23
a) Thruster Assembly: The thruster assembly, via electric and magnetic
circuit elements that energize or manipulate the propellant flow, imparts kinetic energy to the propellant reactive mass for thrust generation. An overview of its functioning is presented in the state-of-art propulsion section. The thruster assembly also consists of a neutral cathode to maintain overall spacecraft charge balance [31].
Figure 2.9: HT-100D system architecture [29]
b) Power Processing Unit: The PPU does the following [31]:
i) Converts incoming source power to the voltages and currents required for thruster operations.
ii) Receives and relays thruster operational commands from the host spacecraft.
iii) Provides operational telemetry to the host spacecraft.
iv) Protects the power electronics from thruster-induced electromagnetic interference (EMI)
c) Propellant feed system: This system consists of the propellant tank,
pressure regulators for maintaining proper tank and line pressures, and flow controllers for delivering the required propellant mass flow rates to the thruster. It consists of
i)latch valve (LV): Shown in Fig 2.9, is in the feed system to provide a way to fully isolate the reservoir from the remainder of the feed system,
24 especially during the launch phase of the mission when the secondary must be totally inactive and benign[32]
ii) Flow Control Valves (FCV): Its purpose in the system is to provide accurate, fast response feedback control of the propellant flowrate. The PFCV responds quickly to changes in the applied voltage, making it possible to quickly vary the flowrate and also to control the flowrate for changing conditions within the system. [32]
25
Chapter
3
Communication Subsystem
3.1 Requirements and assumptions
D3SAT mission is a technology demonstration mission during which the demonstration by D3SAT’s payload HT-100D will be verified through the reception and transmission of commands and data. The data mainly consists of several readings from the sensors installed on different components of different subsystems whose data were first stored inside the auxiliary mass memory during the drag compensation mode, house-keeping mode and then transmitted to the ground during a ground pass.
Due to the stringent size and mass of the primary bus structure mentioned in section 1.5 of chapter 1, the component selection for communication subsystem is strictly based on their size, mass and device simplicity. In this primary analysis it can be assumed that the data generation on D3SAT would be quite less considering only housekeeping data generated for the transmission to the ground station, therefore 2 cases for downlink has been done one with UHF-band and another with S-band. It is required that the ground station
26 communication assembly has to have a simple architecture, being mobile, easily maneuverable and less expensive components.
3.2 D3SAT’s data acquisition system
3.2.1 Elements of D3SAT’s data acquisition system
1) Sensors: Device that is in charge of interpreting a physical property in analog form as an electrical signal such as a voltage or current. [34] The type of sensor varies depending on the observable data on D3SAT. For example: during the mission it is very important to monitor the temperature of the skin panel of D3SAT during the operation of HT-100D so thermistors would be employed on skin panels and other components like batteries and on communication subsystem, current and voltage sensors on solar array, the PPU of HT-100D is mounted with mass flow sensor which would provide telemetry of mass flow rate inside the thruster. A photo camera would be installed close to the HT-100D to take image of the thruster while operation.
Figure 3.1: temperature sensor [ intersil ]
2) Signal conditioning circuitry: It is used to modify and improve the electrical signal by performing operations like linearization, amplification or filtering. These steps are required to make sensor output suitable for the analog-to-digital converter. [34]
3) Analog-to-digital converter: It is responsible for converting the electrical analogue signal into digital data with numeric value that represents the quantity’s amplitude.[34]
27 Figure 3.2: 12 bit Analog-digital converter (Texas inst.)
4) Computer software: Here the main thing is that the software must be compatible with the operating system of both the computer and the whole system. It will be in charge of logging data for a further analysis and processing.
3.3 D3SAT’s On-Board computer
The OBC is a computer or a system of computers that processes various information transmitted to the satellite or from other on-board subsystems. The purpose of main OBC is to provide a platform for CDHS that interfaces with other subsystems of satellite and controls their operations. The other standard subsystems of a satellite are EPS, Communication Subsystem (COM), Attitude Determination and Control System (ADCS) and the payloads.
The OBC for space application have to provide following characteristics [35]:
1) Mechanical robustness to withstand launcher induced loads with respect to sinusoidal loads and release shocks at Karman points. 2) Should withstand thermal variation on-board the spacecraft. 3) Radiation robustness against high energetic particle.
4) Low power consumption.
For D3SAT the mainly required OBC elements are [35]: 1) Processor.
2) Mass memory for housekeeping data.
3) Data buses and bus controllers, debug and service interface, power supplies, and clock.
28 After market survey On-board computer from IMT is selected shown in fig 2.3 satisfying the requirements mentioned in the above paragraph.
Table 3.1: On-board computer specification
Power consumption 7 watts
Mass 0.7 kg
Size 130mm X 130mm X 100mm
Figure 3.3: On-board computer (IMT)
3.4 OBC Architecture
The architecture of a computing system (whether centralized or distributed) refers to the physical and logical framework used to interconnect components. It determines the pathways for inter-subsystem data transfer and may have a large bearing on both wiring harness size and modularity [36]. Three different architectures can be used to connect D3SAT’s subsystems to the OBC: Centralized architecture, Ring architecture and bus architecture.
1) Centralized architecture: A Centralized Architecture has Point-to-Point interfaces between OBC and other satellite subsystems [36]. It is also referred as Star Architecture.
Advantages [36] [37]:
1) Failure of a component or line does not cause system loss. 2) Individual interfaces are possible for secure data.
29 Disadvantages:
1) A lot of wiring is needed.
2) Adding a new node requires both hardware and software changes in the central node.
2) Ring Architecture: Often called ring architecture, it establishes a way to arbitrate information flow control as the data are passed in a circular pattern.[36][37]
Advantages:
1) Low cabling needed and can be distributed throughout the satellite structure.
2) Components can be tested independently.
3) Adding new nodes will have limited impact on OBC. Disadvantages:
1) It is less reliable, since each node is in-line and thus required to achieve transmission to the next node (e.g. short circuit can cause loss of system).
3) Bus Architecture: This architecture utilizes a common data bus with all subsystems sharing the bus. It uses standard protocols and communication schemes for all nodes. [36][37]
Advantages:
1) The system realization is simple.
2) Data transmissions are deterministic which reduces test and troubleshooting time while increases reliability.
Disadvantages:
1) All components must be developed with a specific interface - physically as well as electrically.
2) Failure (e.g. short-circuit) of a component or line may cause loss of system.
Selection of the OBC architecture
Out of these 3 architectures, the bus architecture is most flexible with less harness involved, it is flexible in the sense that it can add any extra component or subsystem [36] but according to the requirement for D3SAT the main priority would be to see the least risks involving feature of maintaining connection of the OBC with other components
30 of subsystem in case of a component failure, therefore star architecture is preferred for D3SAT shown in fig 3.4.
Figure 3.4: D3SAT’s OBC star architecture [35]
3.5 House-keeping data budget
In this preliminary analysis to calculate the amount of data acquired, data acquisition is divided into 2 budgets , the first budget in fig 3.5 corresponds to the data acquired rate during the drag compensation mode and fig 3.6 corresponds to the data acquired rate during the house-keeping mode. These budgets gives us the information that what is the rate at which the data is getting transmitted to the OBC where it would be processed or stored in the auxiliary memory space so that this data would be transmitted to the ground station during ground pass through the communication subsystem. Reference cases [38], [39] were studied for the preparation of budget shown in Fig 3.5 and Fig 3.6. 12 bits of word size is selected because the On-board-computer [IMT] uses 12 bits of word size to store the data.
31 Figure 3.5: Drag Compensation mode data budget
DRAG COMPENSATION mode
SENSOR TELEMETRY TYPE QUANTITY SAMPLING RATE(hz) ACQUIRED DATA RATE (bps) POWER SUBSYSTEM
solar array voltage voltage 3 0.016 0.576 solar array current current 3 0.016 0.576 PCDU voltage voltage 1 0.016 0.192 battery charge sensor voltage 2 0.016 0.384 temperature sensors voltage 7 0.016 1.344 THERMAL (SKIN PANEL)
temperature sensors voltage 6 0.016 1.152
COMMUNICATIONS
GPS power level sensor voltage 1 0.016 0.192 GPS temperature sensor voltage 1 0.016 0.192 Antenna Temp voltage 2 0.016 0.384 Transmitter power voltage 1 0.016 0.192
CD & H
CPU power level sensor voltage 1 0.016 0.192 CPU temperature sensor voltage 1 0.016 0.192
PROPULSION
Mass flow rate sensor voltage 1 1 12 Voltage sensor voltage 1 1 12 current sensor current 1 1 12 temperature sensor voltage 1 1 12
ADCS
Magnetometer XYZ positions 3 0.016 0.576 Gyro XYZ rotations 3 0.016 0.576 Temperature sensors voltage 2 0.016 0.384
32 Figure 3.6: House-keeping data budget
3.6 Selection of Hardware for communication subsystem
Hardware selection consists of selecting suitable transmitter, receiver, uplink and downlink antenna based on their size, complexity and power consumption. For the calculation of link budget it is assumed that 2 cases of downlink will be analyzed, first case of S-band downlink and VHF uplink (S-band/VHF) and the second case is of UHF downlink and VHF uplink (UHF/VHF). VHF uplink is selected for D3SAT because it is assumed that the commands to be transmitted from the ground station would not correspond to large amount of data.After the market survey for S-band/VHF case, S-band transmitter and VHF receiver by cubesat shop is selected due to its portable size which
HOUSE-KEEPING mode
SENSOR TELEMETRY TYPE QUANTITY SAMPLING RATE(hz) ACQUIRED DATA RATE (bps)WORD SIZE (bits) POWER SUBSYSTEM
solar array voltage voltage 3 0.016 0.576 12 solar array current current 3 0.016 0.576 12 PCDU voltage voltage 1 0.016 0.192 12 battery charge sensor voltage 2 0.016 0.384 12 temperature sensors voltage 7 0.016 1.344 12
THERMAL (SKIN PANEL)
temperature sensors voltage 6 0.016 1.152 12
COMMUNICATIONS
GPS power level sensor voltage 1 0.016 0.192 12 GPS temperature sensor voltage 1 0.016 0.192 12 Antenna Temp voltage 2 0.016 0.384 12 Transmitter power voltage 1 0.016 0.192 12
CD & H
CPU power level sensor voltage 1 0.016 0.192 12 CPU temperature sensor voltage 1 0.016 0.192 12
PROPULSION
Mass flow rate sensor voltage 1 0 12 Voltage sensor voltage 1 0 12 current sensor current 1 0 12 temperature sensor voltage 1 0 12
ADCS
Magnetometer XYZ positions 3 0.016 0.576 12 Gyro XYZ rotations 3 0.016 0.576 12 Temperature sensors voltage 2 0.016 0.384 12
33 would easily fit inside D3SAT. For the second case UHF/VHF transceiver by cubesat shop ® is selected. Fig 3.7.
Table 3.2: Technical specification of S-band transmitter [40]
Table 3.3: Technical specification of UHF/VHF transceiver [41]
For the selection of antenna, S-band patch antenna and monopole
antennas selected from cubesat shop®, their specifications are given in Table 3.4 and Table 3.5.
Table 3.4: technical specification of S-band patch antenna [42]
Operating frequency 2200 MHz – 2300 MHz Gain 6db Beamwidth 85 degrees Polarsation RHCP Area 50 mm x 50 mm x 3.2 mm Mass 62 g Power 4 watts Modulation BPSK Frequency 2,1-2,3 GHz Data speed Max 100kbps Size 90 x 96 x 15 mm Mass 75 g Power <4w(transmitter); 0.5w (receiver) Modulation BPSK(downlink) FSK(uplink) Frequency 420-450MHz(downlink) 140-150MHz(uplink) Data speed @ 9.6 kbps (downlink)
@ 1.2 kbps (uplink) Size 96 x 90 x 15 mm
34
Figure 3.7: a) S-band transmitter b) UHF/VHF transceiver (courtesy cubesat shop®)
Table 3.5: Omni directional antenna specification UHF (GOMspace®)[43]
Operating frequency 400-480 MHz
Gain 1.5db
Polarsation RHCP
Mass 30 g
35
Table 3.6: Ground-Station Specifications [44]
Antennas S-band dish antenna (downlink, optional) Yagi Antennas (UHF(downlink),VHF(uplink)) Steerable assembly
Gain 29.2 dB (S-band antenna) 16.3 dB (UHF antenna) 13.2 dB (VHF antenna) Beam-width 5.1 degrees (S-band antenna)
30 degrees (UHF antenna) 52 degrees (VHF antenna) Overall noise
Figure
0.9 dB (S-Band) 2 dB (UHF band)
equipped with one dish antenna to support S-band downlink, it has a gain of 29.2 dB, a beam width of 5.1 degrees and 2 Yagi antennas for UHF and VHF communication with gains of 16.3dB, 13.2 dB and beam widths of 30 degree and 52 degree respectively. The overall noise figure of the system is 0.9 dB for S-band, 2dB for UHF band the whole assembly is steerable. [44]
3.7 Link Budget
Link budget is a method to evaluate the received power and the noise power in a radio link. A radio link consists of three basic elements [45]: 1) Effective transmitting power: transmitter power – (cable + connector) losses + antenna gain.
2) Transmission losses
3) Effective receiving sensibility: antenna gain – cable loss – receiver sensitivity.
For a proper radio link performance, the transmitting power + propagation loss + receiving sensitivity must be greater than 0. The remaining quantity out of them is the link margin of the radio link. A good radio link has 6 dB -10 dB [40].
When working with radio signals, the levels of power or intensity can be very large or very small number and logarithms are best suited
36 to express them[46] , specifically decibel (dB) is used which is a logarithmic unit used to express the ratio of two values of a power or intensity. In short, the logarithmic scale nature of the decibel means that a very large range of ratios can be represented by a convenient number, in a similar manner to scientific notation.
3.7.1 Link Budget Calculations
STEP-1: Determine EIRP
Effected isotropic radiated power: It is the output power when a signal is concentrated into a smaller area by the antenna. An isotropic radiator radiates power equally in all direction, however a perfect isotropic radiator is only theoretical as even the simplest antenna concentrate the signal in one direction [47]. EIRP is given by the product of the transmitter power with the gain of transmitting antenna in logarithmic notation it can be written as
𝐸𝐼𝑅𝑃 = 𝑃𝑇+ 𝐺𝑇+ 𝐿𝑇 (3.1) The unit of 𝑃𝑇 is dBW and 𝐺𝑇 is dBi, 𝐿𝑇 in dB is the transmitter to antenna line loss.
STEP-2: Determine Free Space loss
Power flux density: This parameter refers to received power density on a surface located at a distance r from the transmitting antenna. Expressed in W/𝑚2[47], [48].
∅𝑚 = 𝐺𝑇𝑃𝑇 4𝜋𝑅2
(3.2) The first step in the calculations for free space loss (FSL) is to determine the losses in clear-sky conditions. These are the losses that remain constant with time. FSL derive from the spreading of signal in space. The received power is given by:
37 𝑃𝑅 = ∅𝑚 𝐴𝑒𝑓𝑓𝑒𝑐𝑡𝑖𝑣𝑒
(3.3) Where 𝐴𝑒𝑓𝑓𝑒𝑐𝑡𝑖𝑣𝑒 is the effective aperture of the receiving antenna. It is possible to write:
∅𝑚 = 𝐸𝐼𝑅𝑃/4𝜋𝑅2
(3.4) The effective aperture of the antenna is provided by:
𝐴𝑒𝑓𝑓𝑒𝑐𝑡𝑖𝑣𝑒 = 𝜆2𝐺 𝑅/4𝜋
(3.5) So, the received power may also be calculated by:
𝑃𝑅 = (𝐸𝐼𝑅𝑃 4𝜋𝑅2) ( 𝜆2𝐺 𝑅 4𝜋 ) 𝑃𝑅 = 𝐸𝐼𝑅𝑃 𝐺𝑅 (𝜆/4𝜋𝑅)2 (3.6) In dB can be written as 𝑃𝑅 = 𝐸𝐼𝑅𝑃 + 𝐺𝑅 − 20 log(4𝜋𝑅/λ) (3.7) The last term in the above equation represents the free space loss. We know
𝜆 = 𝑐 𝑓
Where c is the speed of light and f is the frequency represented in MHz then free space loss can be given by expression:
𝐹𝑆𝐿 = 20 𝐿𝑜𝑔 (4𝜋𝑅𝑓 𝑐 )
(3.8) In these equation R is considered distance between the spacecraft and the ground station receiver, eq (3.8) can be considered a basic equation during the calculation of link budget and is the ideal case when additional losses are not present.
38
STEP 3: Determine Additional losses
Figure 3.9: Transmission Losses [47]
In this preliminary analysis we are not going to focus on local losses, a safety margin will be taken at the end of link budget to account local losses. In the previous step free space loss was already defined. Scintillation Effects: Irregularly structured ionospheric regions can cause diffraction and scattering of trans-ionospheric radio signals. When received at an antenna, these signals present random temporal fluctuations in both amplitude and phase. This is known as ionospheric scintillation. Ionospheric scintillation may cause problems such as signal power fading etc., and degrade the quality of satellite navigation systems. The ionosphere can deviate from the expected behavior. This is the case when the ionosphere includes irregularities in which the electron density differs significantly from the “ambient” plasma. These irregularities can cause diffraction effects, i.e. scintillations, on the signals passing through them. The formation, evolution and dynamics of such irregularities are ruled by the interplay between the
39 geomagnetic field, the Interplanetary Magnetic Field (IMF) and the solar wind (that is the emission of energetic particles coming from the Sun) [49].
Figure 3.10: Ionospheric Scintillation [49]
In [48] mentioned that the frequency band between 30 MHz and 3GHz are mostly to suffer by scintillation effect. The ground station of D3SAT is located in Pisa having latitude of 43 degrees and from fig 3.10 we see that the latitude of 43 degrees is untouched by ionospheric irregularities for L-band, it is not exactly D3SAT frequency but it is close to it. Therefore, ionospheric effects are neglected in the case of D3SAT.
Attenuation: Troposphere is composed by a miscellany of molecules of different compounds, such as hail, raindrops or other atmospheric gases. Radio waves that pass by troposphere will suffer their effects and will be scattered, depolarized, absorbed and therefore attenuated [47]. There will be small percentage of radio wave that will not reach the receiver antenna at the ground station because when the radio wave crosses troposphere by coming into contact with miscellany of molecules, their radio frequency energy will convert to thermal energy attenuating the signal and will also be scattered in different direction. The main particles in atmosphere like hail, clouds, ice fog poses a problem of attenuation for the frequencies greater than 10GHz,
40 D3SAT’s radio frequency is quite less than this therefore this attenuation will be neglected [47].
Rain attenuation: This attenuation is also neglected because rain attenuation is a punctual event which could take place during one transmission (ground pass), during the next ground pass rain attenuation might not be present and D3SAT can transmit its signals. [50]
Gas attenuation: For frequencies less than 80GHz only water vapor and
Figure 3.11: Gas absorption attenuation [50]
Oxygen attenuates the frequencies and for frequencies less than 3 GHz only oxygen is responsible for attenuation. In figure 3.11 Y-axis represents the zenith attenuation per km i.e. attenuation per thickness of the atmosphere, when the spacecraft is just overhead the ground station [47]. So, to calculate the total attenuation experienced by the radio link of the spacecraft during the ground pass would be a function of the thickness of atmosphere (taken as 20km from [47]) and function of elevation angle. The atmospheric absorption is given by [47]:
41 (3.9) In eq(3.9), 𝐿𝑎𝑏𝑠|=90°(𝑑𝐵 𝑘𝑚)⁄ is the value taken from fig 2.11 for the
corresponding frequency of D3SAT’s com link and 𝜀 is the angle of elevation, there is no clue about the angle of elevation and in particular the range of angle of elevations will be different for different ground pass therefore to know the attenuation as a function of angle of elevation, 2 plots in fig 3.12 and fig 3.13 are produced whose values will be interpolated for their respective angle of elevation during D3SAT’s ground pass.
Figure 3.12: Attenuation (UHF-band) Vs Angle of elevation of D3SAT’s ground pass
42 Figure 3.13: Attenuation (S-band) Vs Angle of elevation of D3SAT’s
ground pass
Pointing loss: Pointing losses are the result of the misalignment of angles of transmission and reception antenna due to a not lined up bore sight between the earth ground station and, such that the received signal is outside the peak of the antenna beam at reception [51].
Figure 3.14: Antenna Misalignment [46]
The loss due to antenna misalignment at reception and transmission is given by [51]:
𝐿𝑇 = 12(𝜃𝑇⁄𝜃3𝑑𝐵)2 (3.10a) 𝐿𝑅 = 12(𝜃𝑅⁄𝜃3𝑑𝐵)2 (3.10b)
43 For D3SAT’s UHF-downlink case the pointing loss can be assumed to be negligible [11] with use of monopole antennas due to their radiation pattern. And for S-band downlink can be calculate by eq (3.10) where 𝜃3𝑑𝐵the antenna’s half power beamwidth is available in [51].
STEP 4: Calculate system noise temperature
The reception quality of the satellite receiving system is commonly defined through a Receiving System Figure of Merit as 𝐺 𝑇⁄ where 𝑇𝑆 𝑆 is the system temperature at the ground station given by:
𝑇𝑆 = 𝑇𝐴+ 𝑇𝐶𝑂𝑀𝑃 (3.11) Where 𝐺 is receiving antenna gain, 𝑇𝑆 is receiving system noise temperature, 𝑇𝐴 is antenna noise temperature and𝑇𝐶𝑂𝑀𝑃 is composite noise temperature of the receiving system, including lines and equipment. The Figure of Merit 𝐺 𝑇⁄ expresses the impact of external 𝑆 and internal noise factors [52]. 𝑇𝐴 may be known if the total attenuation due to rain and gas absorption (A), the temperature of the rain medium (Tm) and the temperature of the cold sky (TC) are also known. Then, the following expression may be applied:
𝑇𝐴 = 𝑇𝑀 (1 − 10−𝐴 10⁄ ) + 𝑇𝐶10−𝐴 10⁄ (3.12)
44 For cloudy condition the value of 𝑇𝑀 considered as 280K [47]. Sky noise is a combination of galactic and atmospheric effects, according as both these factors influence the quality of the signal in the reception. Galactic effects decrease with the increase of frequency (Fig 3.15). They are due to the addition of the cosmic background radiation and the noise temperature of radio stars, galaxies and nebulae. The value cold sky temperature in case of D3SAT’s UHF downlink can be taken as 100k from fig 3.16 and in the case of S-band downlink the value of sky temperature would be interpolated fig 3.17 depending on the range of angle of elevation of D3SAT during ground pass by using the data from the fig 3.17.
