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Trajectory Design for the JAXA Moon Nano-Lander OMOTENASHI

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Trajectory Design for the JAXA Moon Nano-Lander OMOTENASHI SSC17-III-07 Trajectory Design for the JAXA Moon Nano-Lander OMOTENASHI Javier Hernando-Ayuso, Yusuke Ozawa The University of Tokyo 7-3-1 Hongo, Bunkyo-ku, Tokyo 113-8656, Japan; +81-42-336-24309 [email protected] Shota Takahashi The Graduate University for Advanced Studies 3-1-1 Yoshinodai, Chuo-ku, Sagamihara, Kanagawa 252-5210, Japan; +81-42-336-23042 [email protected] Stefano Campagnola∗ Jet Propulsion Laboratory 4800 Oak Grove Drive, La Canada˜ Flintridge, CA 91011, USA [email protected] Toshinori Ikenaga Tsukuba Space Center, Japan Aerospace Exploration Agency 2-1-1 Sengen, Tsukuba-shi, Ibaraki 305-8505, Japan [email protected] Tomohiro Yamaguchi, Tatsuaki Hashimoto Institute of Space and Astronautical Science, Japan Aerospace Exploration Agency 3-1-1 Yoshinodai, Chuo-ku, Sagamihara, Kanagawa 252-5210, Japan [email protected] Chit Hong Yam ispace Inc. 3-1-6 Azabudai, Minato-ku, 106-0041 Tokyo, Japan [email protected] Bruno V. Sarli Catholic University of America 620 Michigan Ave NE, Washington, DC 20064 USA; +1 301 286 0353 [email protected] ABSTRACT OMOTENASHI (Outstanding MOon exploration TEchnologies demonstrated by Nano Semi-Hard Impactor) is a JAXA 6U cubesat that aims to perform a semi-hard landing at the Moon surface after being deployed into a lunar fly-by orbit by the American Space Launch System, Exploration Mission-1. In this paper, we present the analysis and design of OMOTENASHI trajectory, divided in an Earth-Moon transfer using a cold gas thruster and a landing phase using a solid rocket motor. Strong constrains exist between the two phases, making the mission design a very challenging task. The flight path angle at Moon arrival must be shallow in order to minimize the effect of delay of the deceleration maneuver. This, together with the execution error of the cold gas maneuver, demands a correction maneuver to compensate for these errors. Requirements on the ground station tracking are also deduced from this analysis, and it was found that the use of DDOR is an enabling technology for a safe lunar landing. Under the current subsystems design, we found that the most critical factors in the landing success rate are the maneuver orientation, thrust duration and total delta-v errors. Results suggest accuracy requirements to the landing devices, solid rocket motor and attitude accuracy, as well as to the transfer phase trajectory design. ∗The work of Stefano Campagnola was carried out as an Interna- tional Top Young Fellow in ISAS/JAXA, Japan Hernando-Ayuso 1 31st Annual AIA/USU Conference on Small Satellites INTRODUCTION nition to achieve the required deceleration. Finally, the Small satellites are being considered for missions of surface probe will separate from the retromotor module increasing complexity and interest. They offer a re- at burnout to reduce the load on the energy absorption duced cost and development time, which allows to re- mechanisms. spond to technological and science demands in a shorter timescale. Their use in Low Earth Orbit has already been Figure 1 shows the current state of the design of the proven, and there is an growing interest on applying the spacecraft for different parts of the mission. On the concept to interplanetary missions. This was already the top, Fig. 1a shows the orbiting configuration of OMOTE- case of PROCYON, the first interplanetary small satel- NASHI, featuring solar arrays in the +Y face. The solid lite, developed and launched by The University of Tokyo rocket motor, including its sealing lid, is also visible. Be- and JAXA in 2014 as a secondary payload of Hayabusa2 fore DV2, OMOTENASHI will deploy its airbag as can mission. 1 be seen in Fig. 1b. The orbiting module is ejected after the solid rocket motor ignition, being the configuration One type of mission that can greatly benefit from the ad- during the deceleration maneuver as shown in Fig. 1c. A vantages of small satellites is Moon exploration. The use detailed view of the internal parts of OMOTENASHI is of cubesats detaching from Moon-orbiting spacecraft has presented in Fig. 2. Figure 2a shows the Reaction Con- been proposed in the past . 2 However, if a piggyback op- trol System (RCS), attitude control module, communica- portunity in a mission that features a lunar flyby is avail- tion devices, rocket motor and surface probe. Looking able, the mission scenario can be considerably simplified. from a different angle, Fig. 2b shows the battery module, the laser diode (LD) used to ignite the motor, and the de- This opportunity will arise in the first launch of Amer- vices in charge of inflation of the airbag: N2 gas tank and ican Space Launch System (SLS), called Exploration shape memory allow (SMA) opener. Mission-1 (EM-1). After launch in 2019, thirteen 6U cubesats will be injected into a lunar flyby orbit. 3 In this paper we perform a detailed analysis of a semi- JAXA will seize this opportunity with OMOTENASHI hard lunar lander like OMOTENASHI trajectory, includ- (Outstanding MOon exploration TEchnologies demon- ing the Earth-Moon transfer (DV1, TCM) and the land- strated by NAnoSemi-Hard Impactor). OMOTENASHI ing phase (DV2). We propose a design methodology for mission also seeks to study the radiation environment be- DV1 by analyzing the set of feasible solutions that ar- yond Low Earth Orbit in order to support human space rive at the Moon with a small FPA. Results of sensitivity exploration. 4 analysis under OD and maneuver execution errors sug- gest that a TCM must be considered. Finally, we design However, OMOTENASHI is a challenging mission. One the landing phase imposing zero vertical velocity and a of the main challenges comes from trajectory, which specified height over the Moon surface at burnout. We must be robust to execution and navigation errors. As we identify critical errors in the system, which can be seen present in this paper, a robust trajectory must have a small as requirements for the spacecraft to achieve a safe land- flight path angle (FPA) at Moon arrival. In particular, ing. we found that it must satisfy −7 deg ≤ FPA ≤ 0 deg in TRAJECTORY OVERVIEW order to be error-robust. To this end, the design of the In order to design a trajectory that leads to a safe semi- different arcs of the trajectory cannot be performed inde- hard landing on the surface of the moon, the trajectory is pendently, as they are strongly coupled. divided in two arcs: the transfer and the landing phase. After detaching from SLS, OMOTENASHI must per- During the transfer phase, OMOTENASHI must modify form two deterministic maneuvers that will make this its orbit from the Moon fly-by injection orbit to a Moon cubesat the first one to perform a semi-hard landing on intersection orbit. 6 Additionally, health check-ups and the Moon. A first maneuver, DV1, will inject OMOTE- orbit determination are key aspects of this phase. OD is a NASHI into a Moon-impacting orbit. After perform- critical resource during the first hours of operation, as the ing midcourse trajectory correction maneuvers (TCM) as 13 delivered cubesats have the same need of accurately needed, a solid rocket motor will be ignited shortly be- assessing the orbit they are flying, and the time before the fore the expected Lunar surface collision at a speed of ap- Moon fly-by/arrival is limited. During the trajectory de- proximately 2:5 km/s. After the deceleration maneuver sign process, this was identified as one of the key aspects (DV2), OMOTENASHI will experience a free-fall from that the small satellite community should address in the a low height (close to 100 m) and arrive at the Moon sur- near future. face with a speed of around 20 m/s. 4, 5 In order to reduce the mass budget, OMOTENASHI is composed of an or- The landing phase starts minutes before arriving at the biting module, a retromotor module and a surface probe. Moon surface, and the main event is the deceleration of The orbiting module must be ejected at rocket motor ig- the spacecraft by the use of a solid rocket motor. 7 An- Hernando-Ayuso 2 31st Annual AIA/USU Conference on Small Satellites the mass to be decelerated. In the absence navigation and maneuver execution er- rors, any Moon-intersecting trajectory would lead to a successful landing, provided that the surface probe can absorb the residual kinetic energy after braking. How- ever, the uncertainty on the actual trajectory introduces very strong constraints between the two phases. From the point of view of a safe landing, a trajectory with a very shallow FPA is preferred to minimize the effect of timing errors on the vertical displacement of the space- craft. A high position error on the vertical direction may (a) OMOTENASHI in its orbiting configuration lead to a premature landing during the solid rocket motor burn, jeopardizing the mission. On the other hand, a very shallow FPA might cause missing the Moon in the pres- ence of errors. To reduce the fly-by probability, a TCM may be introduced if necessary. A TCM must be care- fully planned in order not to hinder the orbit accuracy during the landing, as it reduces the time to perform OD before the landing phase. The current analysis and design were conducted with the initial conditions provided by Marshall Space Flight Center 8 and shown in Table 1. The position and veloc- ity components are expressed in a Moon-centered refer- ence frame whose axes are parallel to the J2000 Ecliptic frame. We considered the Sun, Earth and Moon Grav- ity as point masses and an impulsive DV1 maneuver.
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