Sensitivity study of the implementation of Air-Breathing Electric Propulsion systems as atmospheric drag compensation measures Earth Observation CubeSat missions at Very Low Earth Orbits

ESEIAAT Bachelor’s Thesis Sensitivity study of the implementation of Air-Breathing Electric Propulsion systems as atmospheric drag compensation measures Earth Observation CubeSat missions at Very Low Earth Orbits BACHELOR’S THESIS Degree: Bachelor’s Degree in Aerospace Technologies Engineering Delivery date: 30/09/2019 Student: Pau Nadal Vila Director: Silvia Rodríguez Donaire Co-Director: Miquel Sureda Anfres Bachelor’s Thesis 1 Abstract Placing a CubeSat in Very Low Earth Orbits can prove to be considerably beneficial for satellite imagery missions. To consistently benefit from its advantages though, any satellite in such orbits must compensate the drag force generated by the residual atmosphere or it will de-orbit in a matter of days. Any sort of thruster is therefore required. It can either be a conventional propulsion system with on-board stored fuel or an Air-Breathing Electric Propulsion system, which can collect its own fuel from the residual atmosphere. This project evaluates the performance that both engines would have in a possible Very Low Earth Orbit Earth Observation scenario and discloses whether the Air-Breathing Electric Propulsion system can be a viable candidate as CubeSat drag compensation system with regards to conventional thrusters. Keywords: Air-Breathing Electric Propulsion, CubeSat, Very Low Earth Orbit, Earth Observation, Satellite systems, Solar activity, Ion Engine, Field Emission Electric Propulsion, Drag compensation, Residual atmosphere, Small satellite 2 Acknowledgements To all coordinators of the DISCOVERER project: Silvia Rodríguez Donaire, Daniel García Almiñana and specially Miquel Sureda Anfres, for letting me take part in this project and for trying to help me even in quite difficult moments. To all my friends, either from outside or inside the university, for all the the amazing experiences lived during those last four years. To my family, for the unconditional support that they have given me during so many years of my life. 3 Honor Declaration I declare that, the work in this Degree Thesis is completely my own work, no part of this Degree Thesis is taken from other people’s work without giving them credit, all references have been clearly cited. I understand that an infringement of this declaration leaves me subject to the foreseen disciplinary actions by The Universitat Politència de Catalunya-BarcelonaTECH. Title of the thesis: Sensitivity study of the implementation of Air-Breathing Electric Propulsion systems as atmospheric drag compensation measures Earth Observation CubeSat missions at Very Low Earth Orbits Signed: Pau Nadal Vila September 2019 Bachelor’s Degree in Aerospace Technologies Engineering 4 Bachelor’s Thesis 5 Contents 1 Introduction 11 1.1 Aim............................................ 11 1.2 Scope........................................... 11 1.3 Requirements ...................................... 12 1.4 Justification ....................................... 12 2 Background and State of the Art 13 2.1 CubeSatstateoftheart................................. 13 2.2 Micro-propulsion systems state of the art . . . . . . . . . . . . . . . . . . . . . . 15 2.3 ABEP and micro-ABEP state of the art . . . . . . . . . . . . . . . . . . . . . . . . 16 2.3.1 ABEPdesignsreview.............................. 18 2.3.2 Air-Breathing Electrostatic Propulsion . . . . . . . . . . . . . . . . . . . . 18 2.3.3 Air-Breathing Plasma Propulsion . . . . . . . . . . . . . . . . . . . . . . . 19 2.4 ABEP lifetime expectancy . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 20 3 Earth Orbits and physical environment 21 3.1 Classification of Earth Orbits . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 21 3.1.1 HighEarthOrbits................................ 22 3.1.2 MediumEarthOrbits ............................. 22 3.1.3 LowEarthOrbits................................ 23 3.1.3.1 Atmosphere model . . . . . . . . . . . . . . . . . . . . . . . . . 28 4 Benefits and Challenges of orbiting in Low Earth Orbit 30 4.1 Benefitsandadvantages ................................ 30 4.2 Challenges and disadvantages . . . . . . . . . . . . . . . . . . . . . . . . . . . . 31 5 ABEP scenario proposal 32 5.1 Scenarioparameters .................................. 32 5.1.1 Orbitproposal.................................. 32 5.1.2 CubeSatconception .............................. 33 5.1.3 Conventional propulsion power plant decision . . . . . . . . . . . . . . . 34 5.1.4 ABEPdecision ................................. 35 5.1.5 Adopted assumptions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 36 6 CubeSat design using the IFM nano-thruster approach 38 6.1 Structuralarchitecture ................................. 38 6.2 Systemsandsubsystems................................ 39 6.2.1 Payloadsystem................................. 39 6.2.2 Attitude determination and control system . . . . . . . . . . . . . . . . . 40 6.2.2.1 Attitude sensors . . . . . . . . . . . . . . . . . . . . . . . . . . . 41 6.2.2.2 Attitude actuators . . . . . . . . . . . . . . . . . . . . . . . . . . 42 6.2.3 Communication system . . . . . . . . . . . . . . . . . . . . . . . . . . . . 43 6.2.4 Command and Data Handling system . . . . . . . . . . . . . . . . . . . . 45 6.2.5 Thermal control system . . . . . . . . . . . . . . . . . . . . . . . . . . . . 45 6 CONTENTS Bachelor’s Thesis 6.2.6 Propulsionsystem ............................... 45 6.2.7 Electric power system . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 47 6.3 Overall CubeSat system configuration . . . . . . . . . . . . . . . . . . . . . . . . 49 6.3.1 Totalmassbudget ............................... 49 6.3.2 Totalpowerbudget............................... 51 6.3.3 Totalelementbudget.............................. 54 6.3.4 CubeSat design conclusions . . . . . . . . . . . . . . . . . . . . . . . . . . 56 7 CubeSat design using ABEP approach 57 7.1 Structuralarchitecture ................................. 57 7.2 Systemsandsubsystems................................ 57 7.2.1 Propulsionsystem ............................... 58 7.2.2 ADCS ...................................... 59 7.2.3 Electric power system . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 59 7.3 Overall CubeSat system configuration . . . . . . . . . . . . . . . . . . . . . . . . 60 7.3.1 Totalmassbudget ............................... 60 7.3.2 Totalpowerbudget............................... 61 7.3.3 Totalelementbudget.............................. 63 7.3.4 CubeSat design conclusions . . . . . . . . . . . . . . . . . . . . . . . . . . 64 8 CubeSat propulsion system performance analysis 65 8.1 Dragforcedetermination................................ 65 8.2 Propulsion operating altitude range determination . . . . . . . . . . . . . . . . . 67 8.3 Propulsion system operating lifetime . . . . . . . . . . . . . . . . . . . . . . . . . 68 9 Environmental analysis 70 10 Conclusions 71 7 List of Figures 2.1 Small satellites classification [1] . . . . . . . . . . . . . . . . . . . . . . . . . . . . 13 2.2 Physical and power typical constraints for <10kg CubeSats [2] . . . . . . . . . . 13 2.3 Evolution of launched CubeSats over years[3] . . . . . . . . . . . . . . . . . . . 14 2.4 Decay time in orbit for various ballistic coefficients[4] . . . . . . . . . . . . . . . 15 2.5 Characteristics of current conventional micro-propulsion systems [5] . . . . . . 16 2.6 ABEPconcept[6] .................................... 16 2.7 Intake Setups: (a) Funnel concept, (b) Bypass concept[7] . . . . . . . . . . . . . . 18 2.8 ABEPclassification ................................... 19 3.1 OrbitEccentricity[8] .................................. 21 3.2 OrbitInclination[9] ................................... 22 3.3 Types of orbits according to altitude . . . . . . . . . . . . . . . . . . . . . . . . . 22 3.4 MolniyaOrbit[9] .................................... 23 3.5 Sun-synchronous circular orbits as function of altitude [10] . . . . . . . . . . . . 24 3.6 Solar activity over time [11] . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 25 3.7 Partial density of atmospheric species as a function of altitude[12] . . . . . . . . 25 3.8 Average density of atomic oxygen as a function of altitude for minimum,nominal and maximum solar conditions[12] . . . . . . . . . . . . . . . . . . . . . . . . . . 26 3.9 Comparison of different disturbing acceleration in LEO as function of altitude[4] 26 3.10 Drag force as a function of altitude for different cross-sectional areas and in mean solar activity conditions[7] . . . . . . . . . . . . . . . . . . . . . . . . . . . 27 3.11 Variations of semi major axis and eccentricity as function of true anomaly during maximum solar activity (a-b) and minimum solar activity (c-d) for PRIRODA satellite[13]........................................ 28 5.1 CubeSat orientation towards Earth . . . . . . . . . . . . . . . . . . . . . . . . . . 34 5.2 Schematic representation of FEEP operation [14] . . . . . . . . . . . . . . . . . . 35 5.3 Mock-up of the 6U AB-GIE thruster system [15] . . . . . . . . . . . . . . . . . . 35 5.4 Schematic representation of ABEP operation [15] . . . . . . . . . . . . . . . . . . 35 6.1 IFM nano-thruster relation between specific impulse, thrust and power consumption......................................... 47 6.2 Relative mass distribution for CubeSat design using an IFM engine . . . . . . . 51 6.3 Relative power consumption in imaging mode for the IFM CubeSat design . . 52 6.4 Relative power consumption in eclipse mode for the IFM CubeSat design . . . 53 6.5 Relative power consumption in downlink mode for the IFM CubeSat design . . 53 6.6 Relative power consumption in battery recharge mode for the IFM CubeSat design........................................... 54 6.7 Relative cost distribution for CubeSat design using an IFM engine . . . . . . . . 55 7.1 Relative mass distribution for CubeSat design using the AB-GIE . . . . . . . . . 61 7.2 Relative power consumption in imaging mode for the AB-GIE CubeSat design 62 7.3 Relative power consumption in eclipse mode for the AB-GIE CubeSat design . 62 7.4 Relative power consumption in downlink mode for the AB-GIE CubeSat design 63 8 2 Background and State of the Art Bachelor’s Thesis Figure 2.4: Decay time in orbit for various ballistic coefficients[4] 2.2 Micro-propulsion systems state of the art Satellites in Very Low Earth Orbits decay in just a few days after being placed in orbit and the Spacecraft’s Active Lifetime (SAL) is reduced drastically. This is the reason why satellites in such orbits required Corrective Propulsion Plants (CPP) as drag compensating measures to increase mission lifetime. Currently, many satellites in LEOs have got some sort of propulsion systems that activate when the satellites lose too much altitude. By expelling any kind of fuel, these conventional propulsion systems allow the satellites to recover altitude in order to return to the desirable orbit [5]. There are several types of micro-propulsion systems for small satellites [24][25]: •Cold gas thrusters (CGT). They basically produce thrust by expelling a gas (usually nitrogen, helium or butane) contained in a pressurized tank. It is the simplest propulsion system and requires really low electric power. However,its specific impulse and fuel efficiency is really low, so its fuel consumption is elevate. •Resistojet thrusters: with an operating principle similar to CGT, resistojets heat the propellant via an electrical resistance, at the cost of a really high power consumption. The fuel efficiency is improved. •Chemical thrusters: chemical engine take profit of the combustion of a mixture of propellants (usually consisting of fuel and oxidizer ) to provide thrust. •Electrostatic /electromagnetic thruster: Electric propulsion systems provide thrust by accelerating ionized propellant to high velocities. There are several kinds of electric propulsion systems, each one with a different method to ionize and accelerate particles. This thrusters have the highest specific impulse and therefore the lowest fuel consumption rate. However, they generally have low thrust-to-power ratio. 15 2 Background and State of the Art Bachelor’s Thesis Figure 2.5: Characteristics of current conventional micro-propulsion systems [5] 2.3 ABEP and micro-ABEP state of the art The problem is that the fuel used by these propulsion systems has to be carried inside the satellites. However, each extra kilogram of mass considerably increases the cost of launching the satellite. Therefore, small satellites do not carry large amounts of fuel, so, as soon as the stored propellant had run out, the propulsion system stops working and the satellites re-enter. With the considerable drag force effects in VLEO conditions, the fuel consumption by the CPPs can be huge and the amount of fuel that needs to be stored to sustain the satellite in orbit for a significant lifetime can be prohibitive. This is why Air-Breathing Electric Propulsion (ABEP) systems are being developed. The ABEP concept, also known as atmosphere-breathing electric propulsion or ram-electric propulsion, was created to overcome such problem, enabling longer mission lifetime and reducing propellant mass requirement. The main idea behind ABEP systems is to provide thrust by utilizing the residual atmosphere at VLEO as propellant so fuel is no longer carried on-board. Therefore, the satellite fuel consumption will drastically decrease and at the same time the lifetime of the satellite will noticeably increase, hence increasing the overall profit generated by a single satellite[26].Besides, eliminating the propellant mass also reduces the overall launch mass, significantly decreasing the launch costs of the spacecraft. Figure2.6 illustrates the basic ABEP operating principle: Figure 2.6: ABEP concept [6] Nonetheless, ABEP technologies are currently under research and development, and it is still unclear whether or not they will be a more feasible propulsion system than traditional ones used in small satellites. They are far from totally researched or tested. In fact, many of the 16 2 Background and State of the Art Bachelor’s Thesis ABEP projects have a low technological readiness level. Moreover, most of them are still in a preliminary phase of design. Therefore it is difficult to find fully assessed ABEP engines and many of their parameters. Some of them, which are exposed in this bachelor’s thesis, are just an estimation of the expected performance. This is the principal target of this bachelor’s thesis, to estimate if current ABEP systems could be a competitive means of propulsion systems used in satellites in EO missions at VLEO. In order to become feasible, ABEP technologies have to deal with several challenges. They need to be capable of efficiently using the peculiar atmospheric species as propellant. But also of operating with really low mass flow rates and preferably, with low available power. Properly ionizing the residual atmosphere into useful propellant and while avoiding atomic oxygen electrode erosion are crucial design factors of the overall system [7]. ABEP missions are only feasible in a specific range of altitudes. The lower the altitude the higher the density of the atmosphere, so propellant mass flow rate is higher. However, this implies that the drag force is greater, so more power is required to fully drag compensation. On the other hand, at higher altitudes the mass flow rate can be so low that the engine may not be capable of collecting enough propellant mass flow for adequate operation [27]. Over 400 km, the influence of the drag perturbations over the lifetime of the spacecraft are quite slight, so conventional CCPs are clearly preferable rather than ABEP. For lower orbits though, the atmospheric drag forces becomes the strongest perturbation force, severely limiting the lifetime of satellite missions[7]. For satellites orbiting between 250 and 400km, the atmospheric gases could theoretically replace the entirety of the on-board propellant. In this situation, the mission profile in terms of lifetime will remain the same, but the mass that would instead be allocated as propellant can now be used for extra commercial/scientific payload. Otherwise, an ABEP system could complement another sort of CPP, further increasing the satellite lifetime. For orbits in between 160 to 250 km, ABEP systems alone can totally compensate the drag perturbations and could maintain the spacecraft in orbit for an unlimited amount of time. For orbits <160 km, neither the current ABEP systems nor the being developed ones can provide enough thrust to totally compensate drag, though they can elongate the satellite lifetime for some weeks. However, according to the Jet Propulsion Laboratory (JPL), a minimum altitude has been established at 120km, due to heating effects.Actually, it is at 120km that it is considered that re-entry is [28].Moreover, the power that ABEP systems would require for sustained operation at such low altitudes would be prohibitive. This results agree with an ESA study conducted in 2007, where it was considered that air breathing electric propulsion systems are only truly competitive in the range of 120km to 250km [29]. For missions over 250km the amount of propellant required for conventional CCP is so low and the possible propellant flow rate for air-breathing systems is so small that ABEP can not compete with regular EP. For a full drag compensation, the mean thrust force needs to be at least equal to the mean drag force(see eq8.1 in section3.1.3). At orbit altitudes ranging from 120 km to 250km, the orbital speed in which particles ram the satellite is around 7.8 km/s. Theoretically, the exhaust ve- 17 2 Background and State of the Art Bachelor’s Thesis locity of the thruster ce, should be at least this value to counter the drag force. However, it has to be taken into account that only part of the incoming mass flow rate gathered in the intake will lately be used in the propulsion system. This collection efficiency,ηcis a critical design factor in any ABEP system and might drastically affect its feasibility as it will increase the minimum exhaust velocity . ηcdepends on the size and geometrical parameters and design of the intake, but also on its specular reflection of particles, thus decreasing as atomic oxygen degrades the collector. Generally, two different type of intake setups are being used currently: a funnel-type design and a bypass-type design. Each one has its particular advantages and disadvantages, but they are considered out of the scope of this bachelor’s thesis: Figure 2.7: Intake Setups: (a) Funnel concept, (b) Bypass concept[7] For an appropriate size of the intake and assuming compression factors of 100-200, collection efficiencies of ∼40% can be obtained[7]. 2.3.1 ABEP designs review In this section, some of the most relevant ABEP engines that have been designed or are undergoing a development phase will be reviewed. Historically, the concept of ABEP systems is not new. In fact, the idea of replacing stored space fuel with the residual atmosphere gases originated back at the 1960s[7]. However, it was not until 2003 that real ABEP systems started to be developed[15]. Over the last 16 years, several ABEP designs have been developed and some have even been used in space mission, usually for technological demonstration. Generally, ABEP systems can be classified according to the following scheme: 2.3.2 Air-Breathing Electrostatic Propulsion AB electrostatic propulsion devices provide thrust based in the ionization, acceleration and neutralization of the particles (mostly N2and O2) captured by the inlet. The most common AB electrostatic propulsion systems are the air-breathing ion engines (ABIE) and the Hall Effect Thrusters (HET).These two have been the most researched ABEP designs. 18 2 Background and State of the Art Bachelor’s Thesis Figure 2.8: ABEP classification The AB Ion Engine (ABIE) was the first ever ABEP system to be developed. It was designed to handle very low input pressures.The key elements of any ABIE system are: First of all, an air intake in which the propellant mass flow rate is collected. Followed by an ionization channel in which those particles are ionized and temporary stored, if needed. Finished by accelerator grids that will accelerate the outcome flow to the desired exhaust velocity so that drag is compensated. It is worth mentioning that a neutralizer would be added after the accelerator grids, so that the outcome flow is neutral and there is not an aggressive chemical degradation. In fact, the only current example of ABEP system miniaturized for CubeSat applications is the ABIE thruster being developed by the University of Colorado,Boulder[15]. It is a low-weight, small-size ABEP system designed for 3U, 6U, 12U and 27U CubeSat Applications. As an alternative to AB ion thrusters, Hall Effect Thrusters (HET) have higher thrust density, the ratio between thrust and the exhaust area. Henceforth, this allows a reduction of satellite cross-sectional area, thus reducing the overall generated drag. However, preliminary studies show that both ABIE and HET systems have two major drawbacks. First of all, they require a high propellant flow rate that, depending on the collecting efficiency, might exceed the real possible mass intake for a certain altitude. Besides, for low mass flow rates, the engines will have a reduced thrust-to-power ratio. Furthermore, the flow of oxygen through the thruster can result in electrode erosion, which will severely limit the actual lifetime of the engine, becoming a substantial drawback[7]. 2.3.3 Air-Breathing Plasma Propulsion The AB plasma propulsion devices were developed in order to solve the electrode erosion constraint that AB electrostatic propulsion suffers from. The solution was the development of Inductive Plasma Thrusters (IPT), which are electrodeless concepts, based in inductively heated plasma flows coming from Inductively Heated Plasma Generators(IPG). Due to the electrodeless design and the heated plasma flow, IPTs can operate using chemically aggressive propellants such as atomic oxygen without a significant reduction of the engine lifetime. Currently, IPT systems are being researched in the Institute of Space Systems (IRS) of the University of Stuttgart. They are developing and characterizing an IPG6. 19 2 Background and State of the Art Bachelor’s Thesis Another example of plasma propulsion systems are the Air-Breathing Pulsed Plasma Thrusters (AB-PPT). They are still in a basic preliminary design phase, but it is expected that they will be able to efficiently operate with smaller mass intake and relatively low supply power: storage of the mass inflow will cut down the mass flow rate fluctuations, allowing further compression for a better thrust-to-power ratio, thus leading to a better performance. Furthermore, an inductively heated electrothermal plasma generator will allow the AB-PPT to handle hazardous gaseous mixture without negative side effects[7].One advantage of the AB-PPT over the IPT is that the discharge frequency of the AB-PPT permits the system to easily regulate the power level according to the requirements of each pertinent orbit, thus considerably increasing the thruster lifetime. So far,the project performance has not been tested in a proper test bench, so no experimental values exist. 2.4 ABEP lifetime expectancy As mentioned, the ABEP technologies are still undergoing research and development. Many of them are still in preliminary design. In fact, even test facilities for ABEP systems are still being developed, so they are still far away from being normalized. Therefore accurate experimental data regarding ABEP performance is unquestionably scarce. Moreover, if it is taken into account that the tests conducted have only been undertaken a small fraction of time, the only chance of estimating the ABEP lifetime is a well-educated guess. An ESA-CFD feasibility study report carried out in 2007 stated that, according to performance predictions, an average ABEP system could be used for missions with a spacecraft active lifetime from 3 to 8 years[29]. However, later experimental test with air-breathing electrostatic engines revealed that the thruster degradation due to interaction with reactive species such as oxygen had a more detrimental effect than expected, reducing the performance of the engine and severely shortening its total lifetime to ∼10000 hours, depending on the quality of the material. This is why AB plasma thrusters are being researched; in order to cope with the electrode erosion so to significantly increase the engine total lifetime. Nonetheless they are not ready well studied so its expected performance and lifetime are merely predictions[7]. 20 Chapter 3 Earth Orbits and physical environment 3.1 Classification of Earth Orbits In the Solar System, there are several kinds of orbits an object can follow, ranging from hyperbolic to elliptic. Nonetheless, most satellites in Earth’s close space follow an elliptic orbit. Those orbits can be described by three main parameters. First of all, a satellite in a elliptic orbit around the Earth will be at a distance from the Earth surface, or altitude. The closer to the Earth’s surface a satellite is, the higher it’s orbital speed. Therefore, satellites at low altitude will take less time to make a complete orbit around the Earth. Another parameter is the eccentricity of the orbit. Eccentricity indicates how deviated the shape of an orbit is in comparison to a circle. Orbits with no eccentricity become circular, orbits with an eccentricity in between 0 and 1 (none of them included) are considered elliptic, orbits with an eccentricity of 1 are parabolic and orbits with an eccentricity higher than 1 are hyperbolic: Figure 3.1: Orbit Eccentricity[8] The last parameter to basically identify an orbit is inclination. The inclination of an orbit refers to the angle between the orbit plane and the Earth’s equator plane, starting at 0 degrees when the orbit is directly above the equator: 21 3 Earth Orbits and physical environment Bachelor’s Thesis Figure 3.2: Orbit Inclination[9] According to the altitude of the orbit and considering them as nearly circular (eccentricity close to or 0), orbits can be classified in three essential types of Earth Satellite Orbits: High Earth Orbit, Medium Earth Orbit, and Low Earth Orbit. Each one with its particular advantages and disadvantages[9]: [9] Figure 3.3: Types of orbits according to altitude 3.1.1 High Earth Orbits Starting at 35780 km above the Earth’s surface, High Earth Orbits or HEO are the most distant satellite orbits. At exactly 35780 km, satellites take around 24 hours to orbit the Earth. It is a remarkable fact as this to say that a satellite’s orbital speed matches the Earth’s rotation. Therefore, a satellite orbiting at an altitude of 35780 km will remain in place over the same longitude, though it might drift north to south and vice versa. This very special orbit is called geosynchronous. Besides, if a circular geosynchronous orbit plane matches the equator plane (inclination and eccentricity 0) it will become a geostationary orbit. A satellite in geostationary orbit will not have relative movement to the ground, so it will always be over the same place on the Earth’s surface. This is of vital importance to weather monitoring satellites, as they will constantly cover the same wide area. It is also valuable for telecommunications (phone, internet, radio, television) satellites. 3.1.2 Medium Earth Orbits Ranging from 2000 km to 35780 km altitude, Medium Earth Orbits or MEO are particularly suitable for Global Navigation Satellite Systems (GNSS) and telecommunications satellites. It is in this type of orbits that can be found remarkable satellite constellations such as the U.S. Global Positioning System (GPS), the European Galileo satellites or the Russian GLONASS. There are two remarkable medium Earth orbits: •Semi-synchronous Orbit. It is a nearly circular orbit at 20200 km above the Earth’s surface. Satellites in this orbit take approximately 12 hours to complete an entire orbit. 22 3 Earth Orbits and physical environment Bachelor’s Thesis Therefore, satellites pass over the same spot on the equator twice a day, resulting in a highly predictable orbit. The GPS constellation orbits here. •Molniya Orbit. This orbit has an inclination of 63.4 degrees and an eccentricity of 0.72. This highly elliptical orbit is used by the Russian space agency, Roscosmos, to deploy the GLONASS satellite constellation as it is quite useful for observing high latitudes: Figure 3.4: Molniya Orbit[9] 3.1.3 Low Earth Orbits Starting at barely 120 km of altitude and extending up to 2000 km, Low Earth Orbits or LEO are the closest ones to the Earth’s surface. While this type of orbit has not been as exploited as much the other two (specially the geosynchronous case), it turned out to have its own advantages and disadvantages. In fact, it is only in the Low Earth Orbit or Very Low Earth Orbit (VLEO, from 120 km to 450 km)[4] circumstances that Air-Breathing Electric Propulsion systems are indeed viable. Henceforth, it is worth explaining a how the LEO environment is like. Amongst all the LEO orbits, it is worth highlighting the sun-synchronous orbit. These are orbits with an altitude in between 200 to 1680 km and a high inclination in which the sun lightning along the ground track remains almost constant over time, so the surface is always illuminated by the Sun at the same angle when viewed from the satellite. They are frequently used for Earth observation, solar study, weather forecasting and reconnaissance as there are not significant changes in shadows and lighting over time[10]: 23 3 Earth Orbits and physical environment Bachelor’s Thesis Figure 3.5: Sun-synchronous circular orbits as function of altitude [10] Low Earth Orbit has a peculiar difference from the either the MEOs or the HEOs, as there is still a presence of a residual atmosphere that has proven to be rather significant. The composition and density of this residual atmosphere is primarily driven by the Solar Extreme Ultraviolet (EUV) flux. The solar activity can be measured with the 10.7cm solar radio flux indicator (or F10.7), which indicates the output of the EUV radiation of 10.7 cm wavelength in units of 10−22W/m2∗Hz. The atmosphere also depends on the geomagnetic heating of Earth -designated by the APindexbut in a much lesser extent [13]. Table 3.1 shows the maximum, mean and minimum solar and geomagnetic activity of the last 11-year solar cycle. Maximum Average Minimum F10.7 250 140 65 AP45 15 0 Table 3.1: Maximum, average and minimum solar and geomagnetic activity levels of the 2005-2016 solar cycle Even though the solar activity cycle restarts once every 11 years, its F10.7 values never repeat themselves in the exact same way: The most frequent elements that constitute the residual atmosphere are dinitrogen N2and atomic oxygen AO, unlike the diatomic oxygen O2that makes up for the primary element for the atmosphere at sea level. Atomic oxygen is formed at the LEO environment when diatomic oxygen gets photo dissociated by short wavelength solar radiation (<243 nm)[12]. Figure 3.7 represents the average partial density of the most abundant species at LEO for mean solar activity: 24 4 Benefits and Challenges of orbiting in Low Earth Orbit Bachelor’s Thesis •Low revisit timeliness. Satellites in Low Earth Orbits have the highest orbital speeds. Hence, the periods of time in between flying over the same ground spots are shorter than satellites in higher orbits. •Reduced latency. The lower altitude of LEO satellites mean that the data trip between the satellite and the ground is reduced, thus significantly improving the latency. •Cheaper launchs. As a general rule, the highest the orbit required for the satellite, the more expensive launchs become. Therefore using VLEO can certainly reduce the investing money required to place a CubeSat in orbit which, actually, is one of the biggest expenditures of the satellite imagery industry. 4.2 Challenges and disadvantages On the other hand, orbiting at LEO has some important drawbacks compared to higher orbits, most of them due to the extreme LEO environment: •Strong aerodynamic forces and torques. The presence of the residual atmosphere, as function of the altitude and the solar activity, generates an important drag force that, generally, is considered as an unwanted effect that needs to be compensated using propulsion systems. Besides, the aerodynamic torques may also pose a challenge to attitude control systems. •Atomic oxygen erosion. As mentioned in the previous chapter, AO is one of most abundant atmospheric species at LEO, being a highly reactive element that can degrade the sensor surfaces and damage the satellite surfaces. •Short communication windows. The high orbital velocities also mean that the time apertures in which the satellite can transmit data to the ground station are considerably short. This is to say that the ability to downlink or uplink data to a same ground station is severely limited. Therefore satellites in LEO must have high data rates, leading to the use of higher bandwidth communications subsystem, increasing the cost. 31 Chapter 5 ABEP scenario proposal In order to successfully accomplish the aim of this bachelor’s thesis and therefore make an analysis of ABEP systems as nano-thrusters for EO CubeSat applications, an hypothetical scenario will be presented and upon it a feasibility study will be done. The scenario will be that of a short term EO commercial mission at VLEO, which is likely to happen in the near future. Such mission is going to be carried out by a nano-satellite following the CubeSat Design Specifications (CDS) [17]. In this case, a nano-propulsion system will be incorporated as a single primary propulsion plant for full drag compensation. A possible conceptual CubeSat design for EO commercial applications in VLEO will be presented, with two possible configurations. The first one will be a conventional propulsion system with an on-board stored fuel. The second time, an ABEP system will be set up. For both configurations, a summarised mass and power budget of the different CubeSat elements and systems, based on COTS products, will be conducted. Furthermore, a brief manufacturing cost budget for each layout will be done. Afterwards, the operating viability of each propulsion system regarding mass and power restrictions will be evaluated. Finally, considering the mass and power limitations of each propulsion system, a feasibility study for ABEP systems as nano-thrusters will be carried out. 5.1 Scenario parameters In this section, the scenario parameters and the decisions that were taken to realize this study -concerning the EO mission, CubeSat conception, the conventional propulsion power plant choice decision and the ABEP system choiceare explained. In addition, the assumptions adopted in the study are disclosed. 5.1.1 Orbit proposal Due to the fact that the study will be based on a commercial EO CubeSat, a sun-synchronous circular orbit has been picked, as it offers beneficial conditions to EO missions (see Section 3.1.3). The orbit altitude in will be varied among the VLEO range (160-450 km). However, for a preliminary orbit design an altitude of 250 km will be set as initial consideration. Under Newton’s law of universal gravitation, the orbital period of any circular orbit can be obtained by: T=s4π2(R+r)3 µ(5.1) 32 5 ABEP scenario proposal Bachelor’s Thesis Where T(s)is the orbital period, R(m)is the celestial body radius, r(m)is the orbit altitude and µ(3.986e14 m3/s2)is the standard gravitational parameter of the celestial body, calculated as the Earth body mass M(kg)multiplied by the universal gravitational constant G(6.67428e−11 m2/kg ∗s2): µ=G∗M(5.2) In the case of a circular Earth orbit at 250km of altitude, the resulting orbital period is of 89.36min, leading to 16.11 orbits/day. From the total orbit period, it is conservatively to assume that 40% of it will be in eclipse [33]. During those 35.74 min the solar arrays will generate no power, so the electric batteries will have to deliver it. As for total lifetime in orbit, a modest 1.5 year mission duration will be considered. Additionally, the orbital speed Vo(m/s)can also be calculated using the law of universal gravitation: Vo=rµ R+r(5.3) Where R is the Earth’s radius and r is the orbit altitude. Table 5.1 includes the orbit’s main parameters: Orbit type Sun-synchronous Earth orbit Altitude 250 km Eccentricity 0 (circular) Inclination* ∼95o Orbital speed 7759.02 m/s Orbits per day 16.11 Total orbital period 89.36 min Sun-light time 53.61 Eclipse time 35.74 min Table 5.1: Reference orbit for the CubeSat conceptual design For orbit density prediction as a function of altitude, the JB-2006 model will be used [11]. As explained in Section3.1.3, this model is specially useful to figure out the total density of the atmosphere over a certain altitude, which is needed to obtain the drag force (see Eq. 8.1) . 5.1.2 CubeSat conception The chosen spacecraft design has been that of a 6U (3Ux2U, dimensions ∼30x20x10cm, 1.33 kg per Unit) EO CubeSat carrying an optical camera as single payload, similar to many EO CubeSats that some satellite imaging companies are employing currently [34]. The Cube- Sat size has been selected as 6U because smaller configurations imposed too strict mass and power constraints, whereas bigger ones meant extra satellite complexity, see Figure 2.2. The optical payload was selected as its relatability with today’s imaging industry allowed for a simpler income forecast model [35]. A single payload was adopted as the lower the mass and power spent in the payload subsystem, the higher the mass and power is available for the propulsion subsystem. 33 5 ABEP scenario proposal Bachelor’s Thesis The CubeSat preliminary design will be done in compliance with the 13 Rev. CubeSat Design Specifications [17]. The already severe constraints of such an small spacecraft will need to be balance out so that it can hold a nano-propulsion system to perform in a possible VLEO commercial mission scenario. All of the CubeSat elements will come from a COTS origin when possible (see Section 2.1). The CubeSat orientation towards earth is represented in Figure ??: Figure 5.1: CubeSat orientation towards Earth Where the x-face will be the 2Ux1U face and the z-face will be the 2Ux3U face. This way, the cross-sectional area will be minimum, thus reducing the total drag force. 5.1.3 Conventional propulsion power plant decision Among all the the possible propulsion systems for CubeSats reviewed in Section 2.2, there is clearly just one option that might be suitable as a nano-propulsion thruster for an EO CubeSat mission: the electric thrusters, either electrostatic or electromagnetic [25]. Other propulsion systems such as cold/hot gas or liquid mono-propellant might provide higher thrust, but their low specific impulse will lead to a humongous fuel mass consumption, see Figure 2.5. There are two categories of electrostatic thrusters: ion engines and hall effect thrusters. Even though both have rather similar performances, ion engines usually have higher specific impulse, leading to a lower propellant mass flow rate. Therefore, as the critical design factor for the conventionally thrusted CubeSat is probably going to be the available mass for propellant, an ion engine will be adopted in this study. Besides, ion engine also offer a highly controllable very precise steady thrust, thus leading to lesser disturbances and easier, more accurate pointing [36][33]. This is outstandingly suitable for formation flying or even constellations [36]. However, most nano-ion engine systems are yet under development, so they are still being tested used in experimental technology demonstration missions [5]. The only nano-thruster with a high enough technological readiness level that is currently at miniaturized CubeSat scale and available for commercial missions is the IFM nano-thruster engine developed by FOTEC and manufactured by the Enpulsion company. It is a Field Emission Electric Propulsion (FEEP) system which generates thrust by accelerating ions via electric fields. The ions come from the propellant reservoir and are expelled at really high exhaust velocities: The IFM thruster uses indium as solid propellant. It’s incredibly high specific impulse allows for a really low fuel consumption. Its main drawback is the relatively low thrust-to-power ratio, which implies that the thruster requires a high amount of electric power supply and a complex Power Processing Unit (PPU) [37]. The characteristics of the IFM nano-thruster can be found at Table 6.14 in Section 6.2.6. 34 5 ABEP scenario proposal Bachelor’s Thesis Figure 5.2: Schematic representation of FEEP operation [14] 5.1.4 ABEP decision As stated in Section 2.3, ABEP are still in early phases of development. Even more so, miniaturised ABEP systems capable to fit inside CubeSats are in really low stages of technological readiness. Nonetheless, the University of Colorado has developed a conceptual design of an ABEP thruster which would be able to work inside the small CubeSat frame [15]: Figure 5.3: Mock-up of the 6U AB-GIE thruster system [15] It is an Air-Breathing nano Gridded-Ion Engine (AB-GIE) which collects some of the species left in the residual atmosphere (mainly N2and O2), ionizes them and accelerates them so that they exhaust at high velocities. It’s inner workings are quite similar to the IFM thruster, but in this case there is not need for a fuel storage tank which makes for a great advantage over traditional propulsion units: Figure 5.4: Schematic representation of ABEP operation [15] Even though this design is still undergoing simulation, it is expected that the engine could provide relatively high thrust compared with other nano-propulsion systems. However, this ABEP system currently presents some major drawbacks. First of all, the inlet and ionization efficiencies are quite low, so the overall system itself has a low thrust-to-power ratio which means that it requires a lot of electric power consumption. Secondly, collecting atmospheric species at VLEO does unfortunately imply gathering AO. As mentioned in Section 3.1.3, such element is highly corrosive so a continued operation regime of the ABIE at VLEO would lead to a rapid deterioration of the propulsion system, highly reducing its operating lifespan. Although the ABIE is still in early phases of development and testing, it’s expected perfor- 35 5 ABEP scenario proposal Bachelor’s Thesis mance characteristics can be found at Table 7.1 (see Section 7). 5.1.5 Adopted assumptions It has to be taken into account that the purpose of this bachelor’s thesis is not to provide an in-depth CubeSat design or to determine a meticulous mass and power budget of a CubeSat. Neither it is to implement an detailed value chain of CubeSat platforms nor to establish a rigorous income model for CubeSat satellites. Furthermore by the time of this project was conducted, many nano thruster technologies are still undergoing a phase of development and testing, specially concerning ABEP. Therefore, some assumptions were taken to simplify this study in order to keep it inside its scope. Furthermore, some values in this study may turn out not to be completely accurate. Moreover, as ABEP systems are more researched over time, some performance parameters are likely to significantly change. The following statements and hypothesis have been embraced in this study, selected for several criteria: 1. As it will be a short-term EO mission, the variations of the solar cycle over time will not be considered. 2. Atmosphere changes due to orbit inclination will be neglected. 3. Atmosphere changes in the day-night cycle will not considered. 4. Earth oblateness and shadowing effects will not be considered. 5. AO insulation techniques will not be studied. 6. Degradation due to AO will only be accounted in the degradation of solar arrays. It will not be taken into account CubeSat structural degradation, changes in the CubeSat optical or thermal surfaces due to AO degradation. Nor changes in the thruster performance or efficiency. 7. No mechanical or structural analysis of the CubeSat will be done. 8. No satellite heat transfer analysis or thermal budget will be done. 9. No bus or link design will be done. 10. Satellite stability due to torques will not be analyzed. 11. Volume constraints will not studied. CubeSat internal volume element distribution not analyzed. The internal arrangement of components will not be disclosed. 12. It will be considered that the power-to-thrust ratio of the ABIE thruster will remain constant, independently of the altitude and solar activity. 13. The transitory interval in between launcher deployment and the stationary establishment will not be considered 14. The dimensionless drag coefficient CDwill be assumed as a canonical 2.2, used in the literature [7]. 15. No redundancy design will be done in neither of the CubeSat configurations.. 16. In order to minimize drag, the solar arrays surface normal vector will be perpendicular to the incident flux. Doing so implies that the solar radiation does not reach the solar array at an optimum angle [33]. In some specific cases, this drawback could drastically reduce the total supplied power generated by the photo-voltaic cells. Therefore, 36 5 ABEP scenario proposal Bachelor’s Thesis many commercial satellites have auto-adjustable solar arrays that keep rearrange for the appropriate incidence angle. However, doing such thing in VLEO would imply a significant increase in drag as the cross-sectional area would increase too. It this study, the effects of solar incidence angle in the solar arrays will not be considered. 17. At VLEO range, the total density of the atmospheric species is so low that the stream of particles is not longer continuous, but rather each particle is seen individually. Such state is know as free molecular flow. Nonetheless, drag calculation in free molecular flow requires highly complex models. Therefore, for the simplicity of this study it will be considered that the particle flow impacting into the satellite will be continuous and homogeneous. Besides, it will be supposed that such flow does only come in the ram direction. 18. It will be assumed that the cross-sectional area will always remain the same. 37 Chapter 6 CubeSat design using the IFM nano-thruster approach In this chapter, a conceptual design and sizing will be done for the 6U CubeSat carrying the IFM nano-thruster on behalf of a conventional propulsion system (see 5.1.3). As with every satellite, the design of this CubeSat can be decomposed in 2 key segments: the structural architecture and the satellite systems and subsystems. In this particular case, that of a CubeSat employing a conventional propulsion system, the design will be primarily driven by the strictly limited available propellant mass.Besides, the small electric power supply in nanosatellites will also be a restricting factor (see Figure 2.2). The preliminary description, sizing and arrangement of this aforementioned key elements is presented below: 6.1 Structural architecture The structure will be the primary chassis of the CubeSat. It has to physically support all systems and subsystems while being able to resist to the mechanical loads that the satellite might experience during its lifetime, for example during the launch phase. It might as well provide thermal and radiation shielding for sensitive components on the inside. It usually consists of a hollow frame with interior braces and brackets serving as mounting points for the different components inside the CubeSat. The frames are generally made from aluminium as it is a low weight, low cost material with high specific strength. Additionally to the core frame, once the components are mounted inside the CubeSat, external metallic or fiberglass panels cover the satellite to shield the internal components and to provide mounting points for the solar array. A 6U CubeSat structure, which follows the CubeSat Design Specifications, has been found online: 38 6 CubeSat design using the IFM nano-thruster approach Bachelor’s Thesis Element: 6U CubeSat structure Parameter Value Total mass 1100 g Inside envelope 960x960x849 mm Outside envelope 1000x2263x3405 mm Thermal range -40 to +80 oC Manufacturer ISIS Price 7850e Table 6.1: 6U CubeSat structure parameters Both mass and price include the frame, the interior braces and brackets and the external metallic panels. 6.2 Systems and subsystems Now that the structural architecture has been found, its time for the several systems and subsystems. Generally, there are 7 different satellite systems: •Payload system. •Attitude Determination and Control System (ADCS). •Communication system. •Command and Data Handling System. •Propulsion system (not common in CubeSats). •Thermal control system. •Electrical power system. 6.2.1 Payload system Even though the main target of this bachelor’s thesis is the study the nano-propulsion systems in VLEO, the actual purpose of every CubeSat commercial missions is to generate value out of a hosted payload. This is why a single but yet competitive high quality payload has been included in this preliminary design. Nonetheless, its optical properties and performance will not be analyzed in depth. The chosen payload has been a high resolution nadir-looking Chameleon Imager, configured as a high frame rate RGB Bayer-pattern panchromatic (PAN) camera: Element: High framerate RGB Bayer-pattern camera Parameter Value Total mass 1350 g Dimensions 200x94x94 mm Power consumption (Readout mode/ imaging mode) 2.5 W / 3.5 W Price 116000e Manufacturer SAC Image resolution (Mp) 3.6 Spatial resolution (at 500km) 9.6 m Swath (at 500km) 32 km Table 6.2: Chameleon Imager panchromatic camera parameters 39 6 CubeSat design using the IFM nano-thruster approach Bachelor’s Thesis The last 3 parameters of Table 6.2 are optical properties of the payload itself. On one hand, the image resolution is the number of pixels captured in a single digital image, generally expressed in Megapixels (Mp). The amount of Megapixels per image and bit/pixel ratio dictates the size of the file. As this camera has a JPEG2000 frame data (with a typical compression ratio of 10:1) and a 8 bit/pixel ratio, each photograph of 3.2 MP will have a size of 0.32MB [38]. On the other hand, the spatial resolution or, more appropriately, the Ground Sample Distance (GSD) points out the ground distance equivalent to the distance of the center of two pixels in the image. This parameter is of vital importance as the it is dictates the selling prices of the outcome satellite images. Table6.3 illustrates current commercial resolution types and their corresponding selling prices: Very high resolution High resolution Medium resolution Low resolution GSD (m) 0.3-1 1-5 5-10 >10 Custom requested imagery price (e/km2)13.322 7.105 1.261 Free* Archive imagery price (e/km2)7.105 3.552 0.888 Free* Table 6.3: Current imaging resolution parameters Custom requested images refer to customer ad hoc specifically demanded images, while archive images refer to images that the satellite has already taken and that have been stored in its archive. The asterisk for low resolution imaging prices mean that any customer can get the image for free as long as self-accessed. Regarding Table6.3, it can be seen that the Chameleon Imager operating at 500 km would be included into the medium resolution range. Nonetheless, taking into account Eq.4.1 from Section4.1, it can be remarked that when the same payload operates at around ∼260km its GSD goes down to 5m, which would be considered as High resolution, thus substantially increasing the selling imaging prices (specially for requested images). Finally, the swath measures the portion of the Earth’s surface that the camera payload is able to capture as the the satellite orbit around the Earth. It can be calculated by: Swath(km) = GSD(m)∗Imageresolution(Mp)(6.1) Again, considering that the image resolution does not change, the swath of the Chameleon Imager operating at ∼250km would be of 16 km. Additionally, the payload has an integrated mass storage of 160 GB, so no extra payload storage will be needed. A big storage capability will allow the CubeSat to keep making images without overwriting them in the hypothetical case that the satellite could not establish communication with ground station. However, having such data storage would be pointless if the communication system is not able to transmit (to the ground station) the stored data at a sufficient downlink rate. 6.2.2 Attitude determination and control system The purpose of the attitude determination and control system is double. Firstly, it has to measure the position of the satellite’s orientation with respect the center of mass (attitude sensors). Then, in case the attitude has been disturbed, the ADCS has to rapidly correct it so that it is accurately maintained (attitude actuators). An appropriate attitude maintenance is essential for optimal communications, solar panel orientation and payload performance. Furthermore, in the case of satellites in VLEO it is also important to control the thrust direction and the satellite orientation towards the incoming air mass flow in order to minimize the 40 6 CubeSat design using the IFM nano-thruster approach Bachelor’s Thesis Figure 6.1: IFM nano-thruster relation between specific impulse, thrust and power consumption propellant tank, the feed systems (which brings the propellant from the tank to the thruster) and the required control hardware. The Enpulsion company sells the entire engine, with all three elements already assembled together. 6.2.7 Electric power system Being one of the most important systems, the electrical power system has to supply enough power for the entire satellite platform through all its lifetime. The system consists of a power source, energy storage, Power Distribution Unit (PDU) and other power regulation elements. The most common power source of CubeSats are photo-voltaic solar cells while the energy storage is usually done by batteries. In this study, the electric power system will be designed using a maximum power consumption approach. This is to say that the system has to be characterised according to the End of Life (EOL) maximum power requirements. Therefore, to properly size the electric power system, the rest of the satellite systems and subsystems have to be set into maximum power consumption, even if it is not needed to. For example, at the specified orbit of 250km, the IFM engine will not operate at maximum thrust or power conditions, but it will be assumed so. This criteria will allow for the characterisation of the maximum capabilities that a 6U CubeSat can offer currently. In the case of an EO CubeSat, it can be deduced that the power requirements will change according to four different operational modes: imaging, eclipse, downlink and battery recharge: In imaging, the payload will be at imaging mode and the S-Band transmitter and antenna (transmitting subsystem) will be off. In eclipse, the payload will be at off mode and the transmitting subsystem will be off too, but the solar arrays will not be active so the required power will be supplied from the batteries. In downlink, the payload will be off but the S-Band transmitter and antenna will be on, thus considerably increasing power consumption. Finally, in battery recharge, the payload will be in read-out mode and the transmitting subsystem off again. The ADCS, the UHF transceiver and UHF antenna (receiving subsystem), the propulsion system and the command and data handling system will be powered at all time. The power consumed by the PDU can be neglected. Table 6.15 shows the summarised power budget consumption for every system in each operational mode (for systems with variable power configurations, the maximum power setups have been selected). The power needed to recharge the batteries is yet To Be Determined (TBD): 47 6 CubeSat design using the IFM nano-thruster approach Bachelor’s Thesis System / Mode Imaging (W) Eclipse (W) Downlink (W) Battery Recharge (W) Payload system 3.5 0 2.5 0 ADCS 2.8 2.8 2.8 2.8 Communications 3.6 3.6 12.6 3.6 Command and data handling 0.2 0.2 0.2 0.2 Propulsion system 40 40 40 40 Electric power system ∼0∼0∼0 *TBD Total power consumption 50.1 46.6 58.1 46.6 Table 6.15: CubeSat summarised power budget for the several operation modes (without battery power consumption determined) and the IFM-design case In eclipse mode, the batteries will have to provide 46.6 W. Considering that eclipse will last for 35.74 min, the total energy that must be delivered to the CubeSat during eclipse will be 27.75 Wh. A suitable commercial battery for this study could be: Element: Pegasus Class BA01/D Parameter Value Total mass 180 g Total dimensions 89x95x14 mm Energy storage 44.4 Wh Nominal supply power 3.7 V Nominal supply current 12000 mAh Number of cells 8 Price* 5800e Manufacturer Exa Table 6.16: Battery parameters As it can be concluded, the most demanding mode will be battery recharge, due to the high electric power demands of the CubeSat and all the conservative assumptions. Therefore, the battery recharge mode will be used for the overall electric power system sizing and in the later results as maximum power consumption mode. Then, assuming that imaging time is 1 min and downlink will last for 5 min, the remaining sun-light time available for battery charging is 47.61 min (see Table 5.1). Then, assuming a 90% recharge efficiency, the power consumption required to fully recharge the batteries in 47.61 min is 38.86 W. Those 38.86 W are added in the battery charge power consumption for a total of 85.46 W. Finally, adding a 15% safety margin leads to a total maximum power consumption of 98.28 W. However this is the EOL required power, so to properly size the solar arrays the Beginning Of Life (BOL) power should be found. Current solar arrays of satellites at about 500 km degrade at a 3% rate per year [16]. To be conservative, a 10% yearly degradation due to atomic oxygen will be assumed at 250km. Then, taking into account that the CubeSat EO mission described in this study will last for around a year, it can be concluded that the minimum BOL power generation should be 109.19 W, a tremendously high amount for 6U CubeSat standards (see Figure 2.2. The only commercially available solar array for 6U CubeSats capable of such power generation is manufactured by dhv technology. It is a deployable Triple Junction GaAS solar array formed by 2 wings of 3 solar panels per wing: 48 6 CubeSat design using the IFM nano-thruster approach Bachelor’s Thesis Element: BCT 6U Solar Panel Parameter Value Total mass 300 g Total dimensions 360x189.5x1.6 mm Solar array power 20 W Bus voltage 7-23 VDC BOL efficiency 29.5% / 1267 W/m2 Price* 8600e Manufacturer dhv technologies Table 6.17: Single BCT 6U solar panel parameters Element: BCT 6U-H Triple Wing Solar Array Parameter Value Total mass <1800 g Total dimensions 360x189.5x4.8 mm each wing Solar array power 120 W Bus voltage 7- 23 VDC BOL efficiency 29.5% / 1267 W/m2 Price* 51600e Manufacturer dhv technologies Number of panels 6 (2x3) Table 6.18: BCT 6U-H Triple Wing Solar Array parameters *The price of the total solar array was not available online, but similar commercial 6U deployable solar panels can be bought for 8600e. The solar array will provide extra 8.33 W, so there is even more power supply margin. 6.3 Overall CubeSat system configuration Now that the entire CubeSat systems have been explained, sized and selected, a review of its characteristics as a whole platform will be done. Again, it will be part of a conceptual design. Mechanical, structural, thermal, optical analysis and testing should be done in order to optimize the overall CubeSat performance. 6.3.1 Total mass budget Once all the different systems and subsystems of the CubeSat are assembled together the total mass budget accounting for the mass of all the systems and subsystems of the CubeSat can be made: 49 6 CubeSat design using the IFM nano-thruster approach Bachelor’s Thesis Element Mass (g) 6U Structure 1100 Payload system 1350 Chameleon Imager (+storage) 1350 ADCS 479 Star Tracker 250 Sun sensor 6.5 Reaction Wheel 200 Magnetorquer (x3) 22.5 Communication system 343 S-Band transmitter 64 S-Band antenna 100 UHF Transceiver (Rx) 94 UHF antenna 85 Command and data handling 70 Computer 70 Propulsion system (dry) 670 IFM Nano Thruster 670 Electric power system 1980 BCT 6U-H Triple Wing Solar Array 1800 Pegasus Class BA01/D 180 Dry mass 5992 Propellant 250 Wet mass 6242 15% Margin 899 Total mass 6891 Table 6.19: Mass budget for the CubeSat design using an IFM engine It can be seen that the entire CubeSat mass is 6.89 kg, below the maximum allowed total mass for 6U CubeSat satellites, according to CDS (see Table 2.2). However, it has to be taken into account that this result is due to the fact that the propellant mass is already fixed at 250g by the provider of the IFM thruster. Such low fuel amount is going to be the limiting factor of the mission lifetime in orbit. Nonetheless, an improved version of the IFM CubeSat design could be done employing the spare mass for extra fuel storage, greatly increasing the time in orbit. The spare mass could also be used, totally or partially, to add extra payload instruments, increasing the potential income generated by the satellite. 50 6 CubeSat design using the IFM nano-thruster approach Bachelor’s Thesis Figure 6.2: Relative mass distribution for CubeSat design using an IFM engine From the relative mass distribution in this CubeSat configuration it can be surprisingly noticed that the heavier subsystem is not the propulsion one, but rather the electric power system. This is due to low thrust-to-power ratio of the IFM engine, which requires a massive amount of solar cells to supply enough power in order to operate. This fact is so remarkable that in this setup the electric power system mass (1.98kg) doubles the total propulsion system mass (920 kg). 6.3.2 Total power budget Once the power supply required for each CubeSat system has been acquainted, the power budget can be presented, either in an absolute way or in a relative way with regards to the different systems. It has to be reminded that the CubeSat has 4 possible operating modes: 51 6 CubeSat design using the IFM nano-thruster approach Bachelor’s Thesis System / Mode Imaging (W) Eclipse (W) Downlink (W) Battery Recharge (W) Payload system 3.5 0 2.5 0 ADCS 2.8 2.8 2.8 2.8 Star Tracker 1 1 1 1 Sun sensor 0.12 0.12 0.12 0.12 Reaction Wheel 0.18 0.18 0.18 0.18 Magnetorquer (x3) 1.5 1.5 1.5 1.5 Communication system 3.58 3.58 12.58 3.58 S-Band transmitter 0 0 5 0 S-Band antenna 0 0 4 0 UHF Transceiver (Rx) 0.08 0.08 0.08 0.08 UHF antenna 3.5 3.5 3.5 3.5 Command and data handling 0.2 0.2 0.2 0.2 Propulsion system 40 40 40 40 Electric power system 0 0 0 38.86 Power consumption 50.08 46.58 58.08 85.44 15 % Margin 7.512 6.987 8.712 12.816 Total power consumption 57.59 53.57 66.79 98.26 Table 6.20: Power budget for the CubeSat design using an IFM engine Figure 6.3: Relative power consumption in imaging mode for the IFM CubeSat design 52 6 CubeSat design using the IFM nano-thruster approach Bachelor’s Thesis Figure 6.4: Relative power consumption in eclipse mode for the IFM CubeSat design Figure 6.5: Relative power consumption in downlink mode for the IFM CubeSat design 53 6 CubeSat design using the IFM nano-thruster approach Bachelor’s Thesis Figure 6.6: Relative power consumption in battery recharge mode for the IFM CubeSat design From the total power budget it can be noticed how impacting the propulsion system power consumption is, as it drains the major part of the power supply available (around 80%) in imaging and eclipse operation modes. However, it is the battery recharge mode that is the most restricting one in terms of power consumption, as it practically doubles the power consumption in regards of the other three operation modes. It has to be noted that even in the battery recharge mode, the power supply will provide an surplus of 21,74 W. This spare amount of power supply could be used to decrease the battery recharge time. This power budget proves that the overall high power consumption will become a real challenge in the design of electric thrusted CubeSats in any kind of mission in the VLEO range. 6.3.3 Total element budget If all the different element prices are accounted together, the total element cost of a single CubeSat using this IFM engine configuration can be obtained. This resulting element cost indicates how much expensive buying all the different parts of the CubeSat would be. It is not to be confused with the total manufacture unit cost, as this should take into account the assembly costs, the determination of which is out of the scope of this project. 54 6 CubeSat design using the IFM nano-thruster approach Bachelor’s Thesis Element Price (e) 6U Structure 7850 Payload system 116000 Chameleon Imager (+storage) 116000 ADCS 42560 Star Tracker 30000 Sun sensor 3660 Reaction Wheel 6500 Magnetorquer (x3) 2400 Communication system 17500 S-Band transmitter 8500 S-Band antenna 2500 UHF Transceiver (Rx) 3500 UHF antenna 3000 Command and data handling 24500 Computer+software 24500 Propulsion system (wet) 40000 IFM Nano Thruster + Propellant 40000 Electric power system 57400 BCT 6U-H Triple Wing Solar Array 51600 Pegasus Class BA01/D 5800 Net cost 305810 15 % Margin 45871.5 Total mass 351681.5 Table 6.21: Element budget for the CubeSat design using an IFM engine Figure 6.7: Relative cost distribution for CubeSat design using an IFM engine Unlike the mass and power budgets, the element manufacture budget is shaped by the payload system. This is due to the high resolution camera that composes the system. As beforehand stated, such a quality camera setup has been chosen in order to provide commercially competitive imagery. However, the impact in the CubeSat element budget, although useful 55 6 CubeSat design using the IFM nano-thruster approach Bachelor’s Thesis in regard to possible high quality payload setups for EO missions at VLEO, should not be given much importance as it is not a critical element of the CubeSat critical design. It is much more relevant noticing that, concerning the other CubeSat systems, the costs of the electric power system plus the costs of the propulsion system make up for a total of almost one third of the entire CubeSat element manufacture cost, up to almost 100,000 euros. Such high increase in the manufacturing costs drifts apart from the CubeSat standard key elements (simple and cheap components, see Section 2.1). This indicates that perhaps the CubeSat platform is not curently suitable for VLEO environment, as the propulsion system required for continuous orbit maintenance stretches to the limit the already mince nano-satellite capabilities. This is to say that nowadays the power supply required for VLEO operation is so high that such region is more convenient for bigger more capable satellites. 6.3.4 CubeSat design conclusions Here concludes the conceptual design and sizing of a CubeSat carrying a miniature IFM engine. It has been proved that a CubeSat configuration carrying a conventional propulsion system could be constructed. Despite the fact that initially seemed that it’s design was going to be restricted by the propulsion system available mass, it turned out that the most critical system design is going to be the electrical power unit, either by mass limitations or by maximum power available. The significantly high power consumption for the proper operation of the propulsion systems as drag compensations systems will pose a challenge in the design of any thrusted CubeSat operating at VLEO. As stated in Chapter 5, this was just a preliminary sketch, proposed under several assumptions. It is by no means a definitive functional design as a more meticulous technical analysis and further testing ought to be done. 56 7 CubeSat design using ABEP approach Bachelor’s Thesis Figure 7.4: Relative power consumption in downlink mode for the AB-GIE CubeSat design Figure 7.5: Relative power consumption in battery charge mode for the AB-GIE CubeSat design 7.3.3 Total element budget For the third time, the AB-GIE budget parameter is still unknown. In this case though, the maximum possible price that the AB-GIE could have can not be properly estimated. However, both the AB-GIE system costs plus the electrical power unit costs will probably lead to a considerable increase in the manufacturing costs of the complete CubeSat, just like in the IFM CubeSat design. 63 7 CubeSat design using ABEP approach Bachelor’s Thesis Element Price (e) 6U Structure 7850 Payload system 116000 Chameleon Imager (+storage) 116000 ADCS 49060 Star Tracker 30000 Sun sensor 3660 Reaction Wheel(x2) 13000 Magnetorquer (x3) 2400 Communication system 17500 S-Band transmitter 8500 S-Band antenna 2500 UHF Transceiver (Rx) 3500 UHF antenna 3000 Command and data handling 24500 Computer+software 24500 Electric power system 57400 BCT 6U-H Triple Wing Solar Array 51600 Pegasus Class BA01/D 5800 Net cost 272310 15 % Margin 40846.5 Total element manufacture cost 313156.5 Table 7.5: Element manufacture budget for the CubeSat design using an AB-GIE engine 7.3.4 CubeSat design conclusions Here concludes the conceptual design and sizing of a CubeSat carrying a miniature AB-GIE system. It has been proved that a CubeSat configuration carrying an ABEP could be constructed. Again, the most critical design factor has been the total power consumption. However, it has been found out that the AB-GIE does not have much maximum available mass. Therefore, depending on the total mass that the AB-GIE does finally get, there could be a mass problem in this configuration. Overall, both IFM and ABEP CubeSat designs have become quite similar. This is due to the fact that both configurations were selected for the exact same mission in the exact same conditions. Furthermore, both propulsion systems have quite resembling characteristics. 64 Chapter 8 CubeSat propulsion system performance analysis Now that the conceptual designs of the 2 different CubeSat configurations have been presented and compared, its time to compare the performance that each engine would have in VLEO conditions, always taking into account the assumptions in Section 5.1.5. 8.1 Drag force determination First of all, the actual drag force received by the satellite shall be obtained. Taking into account the assumptions made in Section 5.1.5, the drag force can be calculated as shown in Eq. 8.1: FD=1 2AρCDV2(8.1) First of all, the total density of the residual atmosphere species must be found. To do so, it can be simulated using the JB-2006 atmospheric model from the European Corporation for Space Standardization [11]. As mentioned in Section 3.1.3, this is the most accurate model regarding the total density of the residual atmosphere. Figure 8.1: total atmospheric density as a function of altitude and for different levels of solar activity As it can be observed, at really low orbits the total density is not heavily influenced by the 65 8 CubeSat propulsion system performance analysis Bachelor’s Thesis solar activity, but as the altitude increases the it has a higher influence over the total density. This means that in the lowest orbits the propulsion systems’ performance will not depend much on the solar activity but it will progressively do when the altitude increases. Another parameter that has to be calculated is the relative velocity between the CubeSat and the atmospheric particles. Again, for simplifying purposes it will be considered that the thermal velocity of the atmospheric species can be neglected and that such speed will be equivalent as just the spacecraft velocity, which can be calculated as follows: Vo=rµ R+r(8.2) The meaning of each term is explained in Section 5.1.1. Using Eq.8.2 we can plot the satellite’s velocity over altitude: Figure 8.2: CubeSat’s velocity a function of altitude Then, the cross-sectional area has already been found of 0.02m2in Section 5.1.5. Finally, the current literature stipulates that the drag coefficient for conventionally shaped satellites is around CD=2.2 [7]. Therefore, the drag force that the 6U CubeSat (in either of both configurations) is going to experience when orbiting at VLEO can be seen at Figure 8.3: 66 8 CubeSat propulsion system performance analysis Bachelor’s Thesis Figure 8.3: Drag force as a function of altitude and solar activity The magnitude of this drag force is the exact amount of thrust that each propulsion system has to provide in order for full drag compensation. If more thrust is provided, the orbit of the CubeSat will rise. In this case, the amount of drag experienced will decrease so the orbit will rise even more and faster. It is not a particularly dangerous situation: if the original orbit was to be recovered, the CubeSat could simply generate less thrust and progressively descend till the desired orbit. This is a particularly interesting strategy if the propulsion system is to operate using a Duty Cycle. On the other hand though, if the CubeSat generates less thrust than the required in Figure 8.3, the orbit will rapidly decay. Such event can be quite serious if unwanted, as there will be an orbit in which drag force will get bigger than the maximum available thrust so the satellite will de-orbit inevitably. Nonetheless,the same event can be used to effortlessly de-orbit the CubeSat once its active lifetime has expired. 8.2 Propulsion operating altitude range determination However, both IFM nano-thruster and AB-GIE propulsion systems will not be able to provide full drag compensation in all of the VLEO range. In the case of the IFM thruster there will be a minimum altitude below which drag force will always be greater than the maximum available thrust provided by the engine. Concerning the AB-GIE system, there will be both a minimum altitude but also a maximum altitude. Again, the minimum altitude will be determined by the maximum available thrust. Nonetheless, the maximum available altitude or ceiling will depend on the minimum possible propellant required for a proper engine operation, see Section 2.3. Such maximum ceiling is expected to be at 600 km for 6U CubeSats[15]: 67 8 CubeSat propulsion system performance analysis Bachelor’s Thesis Figure 8.4: Maximum available thrust for each propulsion system and its correlation to the drag force as a function of altitude and solar activity From Figure 8.4, the minimum operating altitudes can be attained by finding the intersection points between the maximum available thrust for each engine and drag force as a function of altitude: Low Solar Activity Medium Solar Activity High Solar Activity Minimum altitude (km) 160.1 168.3 172.5 Maximum altitude (km) 600 600 600 Table 8.1: Minimum and maximum operating altitude of the AB-GIE system Low Solar Activity Medium Solar Activity High Solar Activity Minimum altitude (km) 194.6 203.2 215.9 Maximum altitude (km) - - - Table 8.2: Minimum and maximum operating altitude of the IFM system 8.3 Propulsion system operating lifetime The most important parameter of both engines performance is the period of time that they will actively operate or, in other words, the amount of time that they will be able to maintain the desired orbit, thus heavily deciding the overall mission lifetime. On one had, as mentioned in Section 2.3, the operating restraining factor for the AB-GIE is the system electrode corrosion due to AO, which severely reduces its total lifespan. Such lifespan is expected to be round 10000 hours, or 1 year and 51 days. On the other hand the operating lifetime of the IFM propulsion system can either be limited by the IFM own lifespan (17000h) or by running out of fuel. In order to determine so, the 68 8 CubeSat propulsion system performance analysis Bachelor’s Thesis following Equation can be used [25]: mp=24 ∗365 ∗t∗fD Isp ∗g0 (8.3) Where MPis the total fuel mass, t is the operating lifespan in hours, fDis the drag force compensated by the engine, Isp is the specific impulse and g0is 9.81m/s2. It should be noted that the specific impulse depends both on the IFM power consumption and on the amount of thrust provided (see Figure 6.1). The determination of the actual operating lifespan can be done assuming for 3 possible scenarios: best possible scenario, nominal scenario and worst possible scenario: Case Isp(s) Thrust(mN) Supply power (W) Operating lifespan (h) Best 6000 0.3 40 5599.32 Nominal 4500 0.35 40 3599.56 Worst 2000 0.25 20 2239.73 Table 8.3: IFM operating lifespan for the 3 different scenarios As it can be seen, in every scenario the IFM operating lifespan will be determined when the engine runs out of fuel. Moreover, even in the best possible scenario, the IFM maximum operating lifespan is still lower than the AB-GIE lifespan. It must be taken into account that these computes are made considering that the maximum fuel mass is 250g. Nonetheless, in the IFM CubeSat design there was a spare mass of 1.11kg that could be almost entirely used to store extra fuel. Making the computes again with the extra fuel mass the following updated IFM operating lifespan can be obtained for the different scenarios: Case Isp(s) Thrust(mN) Supply power (W) Operating lifespan (h) Best 6000 0.3 40 30236.30 Nominal 4500 0.35 40 19437.62 Worst 2000 0.25 20 12094.52 Table 8.4: IFM operating lifespan for the 3 different scenarios and taking advantage of the spare mass Now, it can be observed that in all three different scenarios the IFM operation lifespan has become greater than the AB-GIE one. However, it has to be said that for the nominal and best case scenario the maximum lifespan will actually be 17000h, as it will be limited by the own IFM engine lifespan, and not the fuel mass available. 69 Chapter 9 Environmental analysis The development of this bachelor’s thesis has been entirely theoretical and analytical. All of the obtained data has been acquired via electronic research or via simulation using software. None of the propulsion systems or the several CubeSat configurations have been actually recreated nor tested physically. Therefore, there is no need for an analysis of the environmental repercussions of this project. Nonetheless, it is worth noting that the development of the ABEP technologies are environmentally friendly as they do not consume any kind of fuel but instead only require of electrical power supply, which can be gathered using just solar cells. Furthermore, they could even replace some traditional satellite propulsion systems which use toxic fuels such as cesium or hidrazine [25]. 70 Chapter 10 Conclusions After the development of both CubeSat design configuration, its subsequent comparison and result analysis, this project has disclosed some interesting points. It has proved that the critical design factor for thrusted CubeSat EO missions at VLEO is not the maximum available total mass -as it could have seem in the beginningbut rather the maximum available supply power. This is due to the relatively low thrust-to-power ratio that many propulsion systems have, including ABEP technologies. This leads to the overwhelming challenge to stretch the already restrictive CubeSat power supply to a borderline limit (The most powerful solar arrays were required for a proper operation of the satellite systems). Such drawback is balanced out by the significantly high specific impulse that electric propulsion systems have nowadays, which leads to a low fuel consumption. As well, the project has shown that both the two presented CubeSat configurations, both ABGIE technologies and conventional electric propulsion systems are capable of providing full drag compensation for a determined VLEO altitude range. The minimum altitude was lower for AB-GIE systems thus allowing for lower orbits but it had a maximum altitude. Instead, the IFM engine had a higher minimum altitude but it does not have a ceiling, so the CubeSat can go as high as desired. 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