Journal of Automation and Control Engineering Vol., No., December A Concept Study of the All-Electric Satellite s Attitude and Orbit Control System in Orbit Raising Yoshinobu Sasaki Japan Aerospace Exploration Agency / Space Technology Directorate I, Tsukuba-city, Ibaraki, Japan Email: sasaki.yoshinobu@jaxa.jp Fig. compares the simulation result for the satellite mass ratio (drymass/propellant) for a conventional satellite and an electric satellite. It is assumed that the satellite mass (BOL) is kg, that the propellant used is for GTO to GEO transferring, and that out-of-plane control is in GEO. Abstract All-electric satellite systems and electric propulsion are becoming increasingly well-known technologies, because they help in reducing the propellant weight by a significant amount and enable the mounting of an increased mission payload. Unfortunately, the systems suffer from a much longer Geostationary transfer orbit (GTO) to Geostationary equatorial orbit (GEO) transfer time. This means there is more delay inimp service-in times. Since the transfer orbit is in the Van Allen radiation belt, it is advisable to shorten the transfer time. To resolve this problem, significant progress has already been made to optimize the orbit and attitude control. These solutions require complex orbits and attitude control; relatively little research has studied the attitude control method. In this paper, the concept of a simple attitude control method that achieves orbit-raising time optimization and has little design effect on the subsystem, except for on the electric propulsion subsystem, is reported. Total Propellant Satellite Mass (Wet) [kg] 6 Index Terms All-Electric Satellite, GTO, GEO, OrbitRaising, Attitude Control, Electric Propulsion Satellite Mass (Dry) Propellant Reduction I. INTRODUCTION Geostationary satellites have conventionally used a two-liquid propulsion system for transferring themselves from Geostationary Transfer Orbit (GTO) to Geostationary Equatorial Orbit (GEO). A two-liquid propulsion system enables satellites to transfer from GTO to GEO within a few days, but because of its low specific impulse (~ a few hundred seconds), significant amount of the satellite s mass is occupied with propellant. Additionally, some geostationary satellites have another propulsion system (for example, one-liquid or electric) that executes out-of-plane and attitude control, and reaction wheel s unloading. This design makes the satellite system more complicated. In contrast, all-electric satellites, utilizing a system that only uses electric propulsion, simplify the satellite system. Additionally, an electric satellite system significantly reduces the weight of required propellant and utilizes the high specific impulse (greater than a thousand seconds) of the electric propulsion system thus cause to increase the mission s payload. Conventional (Isp = sec) Figure. Propellant ratio comparison between a conventional satellite and an all-electric satellite The simulation result to calculate total propellant ( m p ) is simply derived from a equation (). m p m ( exp( v )) gi sp () m is satellite mass (Dry) [kg], v is necessary velocity from GTO to GEO [m/s], g is the gravitational constant, I sp is specific impulse of thruster. It can be seen that an electric satellite reduces the propellant weight by about kg as compared to a conventional satellite. However, it should be noted that an electric satellite s optimal transfer needs a greater change in velocity (approximately m/s) in comparison to conventional transfer (approximately m/s). Manuscript received August, ; revised December, Journal of Automation and Control Engineering doi:.88/joace...9 All Electric Satellite (Isp = sec) 9
Journal of Automation and Control Engineering Vol., No., December sin Gaua Geu e cos Gaua Geu e An all-electric satellite system possesses a significant advantage in comparison to a conventional satellite. Boeing s SP satellite was the first all-electric satellite launched [], and other aerospace agencies and companies are currently developing their all-electric satellite systems or electric propulsion subsystems [-]. However, an all-electric satellite system does have some problems, one significant issue being the higher GTO to GEO transfer time of several months. The higher the transfer time, the higher the time spent in the Van Allen radiation belt. Furthermore, there are increases in service-in time delays in the satellite operation. Therefore, it is vital that the transfer time is minimized. To resolve this problem, it is necessary to optimize satellite orbit and attitude control, into which substantial research has already been conducted [ 6],[9]. For example, the Sequential quadratic programming (SQP) method [] and Particle swarm optimization method [8] have been proposed to optimize the transfer time. These solutions require complicated orbit and attitude control, but few studies have demonstrated an acceptable method of attitude control. In this paper, the concept of a simple attitude control method that is based on the orbit optimization method [] is reported. It achieves both orbit-raising-time minimization and minimal design effects to the subsystem (except for to the electric propulsion subsystem). () Acceleration Orbit β α Figure. (in-orbit-plane) and (out-of-orbit-plane) definitions x GEO - GTO - - II. CONCEPT STUDY OF ATTITUDE AND THE ORBIT CONTROL SYSTEM - - x The out-of-orbit-plane angle, yaw angle expressed as equation (). is argument of perigee, f is true anomaly. A. Precondition of Satellite Mass, System, and Initial Orbit In this study, the initial orbit is GTO, which is defined as the initial semi-major axis: a =,6 km, initial Gi ( / ) cos( f ) eccentricity: e =.6, and initial inclination: i =. degrees. The satellite mass and electric propulsion subsystem parameters are defined as, satellite mass (BOL): m = kg, propulsion type: Hall thruster, Then, () Ga is fixed to, since these three gains are relative values. Next, the orbit transfer time is divided by and an initial value of Ge and Gi is assumed. Following this process, optimization is executed. More detailed information about this optimization method is available []. specific impulse of thruster: I sp = seconds, thruster force: F = 8 mn, and two thruster heads. The s eclipse is not considered in this study. C. Simulation Result of Optimal Transfer Orbit and Acceleration Direction The conditions and method execute simulation. Fig. shows orbit trajectory and Figs. 6 show the satellite acceleration direction in the first, middle, and end stages. According to the simulation result, acceleration direction (in-orbit-plane) is close to the satellite s direction of movement at the first stage in order to increase semi-major axis. In contrast to the first stage, acceleration direction (inorbit-plane) is fixed in inertial space at the end stage in order to decrease eccentricity. B. Method of Transferring Orbit Optimization The transferring orbit-time optimization method is based on the SQP method []. The method followed is summarized here. Initially, the unit vector u a and ue are calculated, enabling the maximization of the rate of variability of each orbital element, semi-major axis a and eccentricity e. The in-orbit-plane angle (pitch angle as shown in Fig. ) is determined by normalizing the sum of first gain Ga multiplied by u a and second gain Ge multiplied by ue as equation (). Journal of Automation and Control Engineering - Figure. Orbit trajectory from GTO to GEO (green line)
Journal of Automation and Control Engineering Vol., No., December x Time =.8 [h], I = 9 [-] Acc Vector dra dt - - - - - - x Figure. Satellite acceleration vector at the first stage (light blue line) x Figure. Apogee radius (Ra) and Perigee radius (Rp) changing in orbit-raising Time = 6. [h], I = 8 [-] D. Concept of Attitude Control Method in Orbit-raising There are several methods that can control satellite attitude in orbit-raising. In this study, two simple controls were investigated. (a) Active attitude control method The first solution for controlling satellite attitude involves taking two restrictions (satellite solar array direction to sun and acceleration direction) into consideration. To be specific, this necessitates fixing one axis to the thrust vector in the acceleration direction and fixing the other axis to the solar array paddle axis so that it will be confronted by the direction of the sun. This low allows the satellite to maximize the solar power it receives, and there is no requirement to change the design of the electric propulsion subsystem. However, this method has several problems which must be taken into consideration: i. This attitude law occurs at two singular points in one circulation. To avoid this, it is necessary for the coasting attitude to change mode near the singular point. ii. Satellite attitude is always moving in orbit-raising, so the subsystem will need vision or a specific direction (for instance, a star tracker, sun sensor, geostationary GPS receiver, or feeder link antenna), and its design is thus affected. iii. This attitude low is dependent on the sun s direction, so the satellite s attitude would change depending on its launch date (and what season it was in). It makes more complicate to design a common GEO satellite. (b) Simple attitude control method The second solution is a simple attitude control method focusing on its characteristic acceleration direction in orbit-raising (as already described in section C). Acc Vector - - - - - - x Figure. Satellite acceleration vector at the middle stage (light blue line) x Time =. [h], I = [-] Acc Vector - - - - - - x Figure 6. Satellite acceleration vector at the end stage (light blue line) Journal of Automation and Control Engineering
beta(out of plane) [deg] beta(out of plane) [deg] beta(out of plane) [deg] beta(out of plane) [deg] Journal of Automation and Control Engineering Vol., No., December /6 /6 /6 - - - - - - - - - Figure 8. Alpha and beta in -pointing attitude low (from to ) - - /6 /6 /6 - - - Figure. Alpha and beta in inertial fixed pointing attitude low (from to ) /6 /6 /6 6/6 - - - - - - /6 /6 /6 6/6 - - - - - - - - Figure 9. Alpha and beta in -pointing attitude low (from to 6) This method s attitude direction is divided to two. Fig.. shows apogee radius (Ra) and perigee radius (Rp) changing in orbit-raising. To be specific, an -pointing attitude in the first half of orbit-raising time (dra/dt >=) and an inertial fixed attitude in the second half time (dra/dt < ). Fig. 8-9 shows pitch angle and yaw angle changing trajectory in the -pointing attitude in orbitraising time divided by six. Fig. - also shows angles in the inertial fixed attitude. As can be seen from Figs. 8-, if the satellite attitude is set to the -pointing attitude in the first half and to the inertial fixed attitude in the second half, then the direction of acceleration is within a certain range. Pitch angle is within to + degrees. This inplane angle change is dealt with by satellite attitude. The solar paddle moving axis is the same as, so this method manages to ensure both solar power and optimum orbit-raising. Figure. Alpha and beta in inertial fixed pointing attitude low (from to 6) The Yaw angle is within ~ + degrees. This out-of-plane angle change is dealt with by the electric propulsion subsystem (described in section E). This method provides the following three advantages in comparison to the method described in section D(a): i. This attitude control law doesn t occur at a singular point, so the satellite doesn t have to move quickly or coast. ii. This is a simple attitude law and therefore there are minimal effects on the design (except for the changes in the electric propulsion subsystem). iii. Satellite design is standardized because the law doesn t get affected by the launch date (and what season it was in). However, this method has two problems which must be taken into consideration: Journal of Automation and Control Engineering
Journal of Automation and Control Engineering Vol., No., December i. The electric propulsion subsystem requires a significant number of design changes. ii. On occasion, the solar power charge is reduced (approximately 8% at an inclination of. degrees). F E. Concept design of the electric propulsion subsystem In this section, two concepts of the electric propulsion subsystem which are used to resolve the problems encountered in D(b) are reported. (a) Fixing thruster heads on the satellite Fig. shows the electric propulsion thruster heads position on a satellite. Thruster heads (F) has two thruster heads, and generates acceleration at a direction of + degrees from the orbit plane. Additionally, F has two thruster heads which have degrees acceleration vectors. It is assumed that each thruster head force has normalized force as one. Fig. shows the acceleration direction when the satellite needs two normalized forces. Although most of the actual electric propulsion system only has specific force points. Therefore, two operating points (., ) are assumed in Fig.. Several angle ranges (from to from to ) have little difference from, but the other angle range is near, so it is acceptable. As previously described, if the electric propulsion subsystem has three force points (,., ), it is possible to conduct orbit control without controlling the out-ofplane attitude control. However, there are a lot of force-changing points in one orbit circulation, so evaluation of the lifetime of the electric propulsion system is required. (b) Dynamic gimbal to control thrust vector In order to avoid reduction in lifetime of the electric propulsion system, it is preferable to keep the thruster power constant. Fig. shows a dynamic gimbal to control the thruster vector. This design enables thrusting to continue at orbitraising, although this new type of development system obliges us to evaluate many issues, for instance, lifetime evaluation. F Figure. Out-of-plane thrust vector image: degrees (left), degrees (middle), + degrees (right).8..6. F F.. F+F Vector propulsion efficiency η [-] Normalized Thrust Force [-]. η - - - - Beta [deg] Figure. Total thrust force / angle from orbit-in-plane / vector propulsion efficiency (thrust force changing: unlimited (~)).8..6. F F F+F η.. Vector propulsion efficiency η [-] Normalized Thrust Force [-]. - - - - Beta [deg] III. CONCLUSIONS In this study, we reported the concept of a simple attitude control method that achieves orbit-raising time optimization and has little effect on the design of the subsystem except for the electric propulsion subsystem. This study covers part of a necessary issue that concerns realizing advances in all-electric satellites, so it is important to continue with this research, including the viewpoints of a variety of experts. Figure. Total thrust force / angle from orbit-in-plane / vector propulsion efficiency (thrust force changing:,., ) F ACKNOWLEDGMENTS I would like to thank Mr. Kentaro Nishi and Satoru Ozawa (JAXA) for their useful discussions, comments, and support for further work. F Figure. Dynamic gimbal to control thrust vector (an example) Journal of Automation and Control Engineering
Journal of Automation and Control Engineering Vol., No., December REFERENCES [] S. A. Feuerborn, D. A. Neary, and J. M. Perkins, Finding a way: Boeing s all electric propulsion satellite, 9th AIAA/ ASME/ SAE/ASEE Joint Propulsion Conference, San Jose, CA, AIAA 6,. [] J. J. Delgado, et al, Space systems Loral electric propulsion subsystem: Years of on-orbit operation, in Proc. Joint Conference of th ISTS, th IEPC and 6 th NSAT,. [] E. Rezugina and A. Demaire, All EP platform mission design challenges and subsystem design opportunities, in Proc. rd International Electric Propulsion Conference, Washington, USA, IEPC--9,. [] O. B. Duchemin, et al., Ariane -ME and Electric Propulsion: GEO Insertion Options, th AIAA/ASME/SAE/ASEE Joint Propulsion Conference & Exhibit, [] A. Dutta, et al, Minimum-fuel electric orbit-raising of telecommunication satellites subject to time and radiation damage constraints, American Control Conference (ACC),. [6] E. Rezugina and A. Demaire, All EP platform mission design challenges and subsystem design opportunities, in Proc. rd International Electric Propulsion Conference, Washington, USA, IEPC--9,. [] M. Utashima, Optimization of orbital transfer to geostationary earth orbit by solar electric propulsion, NASDA-TMR-, (in Japanese) [8] K. Kitamura, et al., Minimum time orbit raising of geostationary spacecraft by optimizing feedback gain of steering-law, ISTS- -d-, [9] K. F. Graham and A. V. Rao, Minimum-time trajectory optimization of low-thrust earth-orbit transfers with eclipsing, J. Spacecraft and Rockets, vol., pp. 89-, 6. [] A. Ruggiero and P. Pergola, Low-thrust maneuvers for the efficient correction of orbital elements, in Proc. nd International Electric Propulsion Conference, Wiesbaden, Germany, IEPC--,. [] B. A. Conway and C. A. Kluever, Spacecraft trajectory optimization, Cambridge University Press, New York,, pp. -. [] J. E. Pollard, et al., Simplified analysis of low-thrust orbital maneuvers, Aerospace Report, No. TR-(86)- [] N. S. Gopinath, K. N. Srinivasamuthy, Optimal low thrust orbit transfer from GTO to geosynhronous orbit and stationkeeping using electric propulsion system, in Proc. th International Astronautical Congress of the International Astronautical Federation, the International Academy of Astronautics, and the International Institute of Space Law, [] A. Dutta, et al., On the design of power and propulsion subsystems of all-electric telecommunication satellites, in Proc. SPACE Conferences, nd AIAA International Communications Satellite Systems Conference, Yoshinobu Sasaki was born in Nagasaki, Japan, in 98. He received B.E. and M.E. degrees in mechanical engineering from the University of Kyushu, Fukuoka, Japan, in, 9, respectively. In 9, he joined Japan Aerospace Exploration Agency (JAXA), Tsukuba, Japan, where he has been studying systems and components of guidance, navigation and control on satellite as engineer. Journal of Automation and Control Engineering