Catalog Catalog


"Orbit Maneuver Compensation of KAGUYA for its Safe and Accurate Lunar Transfer"
Kawakatsu Y., Terada H., Matsuoka M., et al.
Proceedings of the 26th International Symposium on Space Technology and Science, 2008-d-59, 2008.

Reported in this paper are the results of the orbit maneuver compensation in KAGUYA’s Lunar transfer. Because of the uncoupled allocation of the attitude control thrusters, extra velocity increment ( ) is induced whenever KAGUYA performs an orbit maneuver. Since the observed level of was unacceptable range from the point of maneuver accuracy requirement, it was compensated by means of deducting estimated from the orbit maneuver command. The estimation model was updated step-by-step during the Lunar transfer, which leaded to the evident improvement of the orbit maneuver accuracy and resulted in the omission of the last trajectory correction maneuver. The method of the compensation and its results are introduced in detail.

"KAGUYA (SELENE) Trajectory Reconfiguration Plans Prepared for Anomaly in Translunar Phase"
Kawakatsu Y.

18th AAS Space Flight Mechanics Meeting, AAS 08-106, 2008

In this paper, the reconfiguration of translunar trajectory in case of main engine anomaly is investigated. The objectives of the trajectory design are to reduce the excessive velocity at the Lunar encounter, as well as to reduce the total required to complete the sequence. 3-impulse Hohmann transfer based trajectory is adopted and possible trajectories are categorized under 2-body approximation. The solutions obtained are applied to more sophisticated models (3-body and 4-body approximation) and yields feasible trajectories.

宇宙技術、Vol. 6, pp. 87-96, 2007

Analyzed in this paper are Near-Earth Asteroids sample return mission opportunities in early 2010s. The mission sequences supposed are Keplerian orbits connected with implulsive velocity changes including planetary gravity assists. The sequences are constructed by “trajectory parts connecting method”, which is exploited by the author. The method enables to construct the possible sequences comprehensively under defined structure and given constraints. At the same time, it enables to assess the dynamical feasibility of the constructed sequences quantitatively by way of the total required velocity increment required to complete the sequences. Over 4000 Near-Earth Asteroids are taken into account as the candidates of the mission targets, and the mission sequences include not only the sample return from a single asteroid, but also the sample return from two asteroids. Lists of the asteroids (or their combinations) which have mission opportunities in early 2010s are provided and some example sequences are shown.

"Study on the Characteristics of Two-burn Translunar Trajectory"
Kawakatsu Y.
Transactions of the Japan Society for Aeronautical and Space Sciences Space Technology Japan, Vol. 5, pp. 9-15, 2007

In this paper, the translunar trajectory of the simple sequence is investigated. The trajectory discussed in this paper is a two-burn ballistic trajectory from the low earth parking orbit to the low lunar orbit. The problem is practical in that the geometric relation from the launch site to the moon is fully considered. The paper includes three subjects. First is the structure of the problem and the solution space. The problems are defined and the solutions are grouped by the properties of the lunar transfer sequence. Second are the characteristics of optimal solutions. The topics discussed are the transition of the required velocity increment based on the launch date, and the difference according to the property of the lunar transfer sequence. Third is the sensitivity analysis of a number of items to the deviation of the parameters from the optimal solution.

"Study on a Lunar Approach Strategy Tolerant of a Lunar Orbit Injection Failure"
Kawakatsu Y., Yamamoto M., Kawaguchi J
Transactions of the Japan Society for Aeronautical and Space Sciences Space Technology Japan, Vol. 5, pp. 1-7, 2007.

Discussed in this paper is a lunar approach strategy tolerant of a lunar orbit injection (LOI) failure. LOI is one of the most critical events for a lunar orbiting mission. If LOI is not performed, the spacecraft flies by the moon, and in the worst case, it escapes not only from the moon but also from the earth, which leads to mission failure. The proposed strategy is to design a trajectory so as to re-encounter the moon even when LOI is not performed. It provides an opportunity for mission recovery even in the case of an unexpected fly-by. A trajectory design procedure is introduced and an example of the designed trajectory is shown.

"Concept Study on Deep Space Orbit Transfer Vehicle"
Kawakatsu Y
Acta Astronautica, Vol. 61, pp. 1019-1028, 2007

In this paper, the concept of Orbit Transfer Vehicle for Deep Space Exploration (Deep Space OTV) is proposed, and its effectiveness and feasibility are discussed. Basic concept is the separation of two functions required for the deep space exploration, the transportation to the destination, and the exploration at the destination. Deep Space OTV is a spacecraft specialized for the transportation to the deep space destination. It is an expendable spacecraft propelled by solar electric propulsion. The payload of Deep Space OTV is Explorer, which is a spacecraft specialized for the exploration at the deep space destination. The effectiveness of the concept is discussed qualitatively, focused on the merits of the separations of two functions. The feasibility of Deep Space OTV is discussed based on the conceptual design of the spacecraft and its applicability to deep space missions. Several deep space missions are modeled and the payload capacity of Deep Space OTV is estimated. The missions include Asteroid rendezvous, Mars orbiter, Lunar lander, and so on.

"Initial Phase AOCS Operation of Infrared Astronomy Satellite 'AKARI'"
Kawakatsu Y., Hashimoto T., Bando N., et al.
Advances in the Astronautical Sciences, Vol. 128, pp. 755-774, 2007

Reported in this paper is the initial phase operation of the attitude and orbit control system (AOCS) of Japanese satellite “AKARI”. AKARI is the first Japanese satellite dedicated to the infrared astronomy. AKARI was successfully launched by M-V rocket from Kagoshima Space Center on February 22, 2006. Just after the launch, AKARI faced to a serious problem. It was found that something interferes with the fields of view of two (out of two) sun sensors. The two sensors were out of use in the subsequent AOCS operation. Fortunately, by using two star trackers and gyros, the attitude accuracy required for the scientific observation can be achieved without the sun sensors. However, as other many satellites, the sun sensors were to play important roles in AKARI’s AOCS operation. Firstly, they were to be used for sun acquisition at the very initial phase. The anomaly prevented us from carrying out the pre-planned sequence. However, the crisis was overcome with the appropriate ground support operation. Secondly, for the observation in far infrared wavelength, the telescope and the scientific instruments of AKARI are stored in the cryostat and cooled by liquid Helium. To prevent the sun light inflow to the telescope, the attitude of AKARI against the sun is strictly constrained after the scientific observation starts. The sun sensors were to be used to watch the sun direction to keep the appropriate attitude against the sun even when AKARI falls into the safe mode. To save the cryostat from the sun light in the absence of the sun sensors, AOCS architecture was reconfigured so that the equivalent function is achieved by the remaining sensors. The experiences in this recovery operation are mainly reported in this paper. In addition, unexpected continuous orbit rising was observed in the operation. The cause of the phenomenon and the investigation process are also reported.

"Mission Analysis of the Sample Return from Primitive Type Near Earth Asteroid"
Kawakatsu Y., Abe M., Kawaguchi J.
Advances in the Astronautical Sciences, Vol. 127, pp. 2119-2131, 2007.

Reported in this paper is the result of the mission analysis of the asteroid explorer mission. Following the results of HAYABUSA, the Japanese asteroid explorer, JAXA has started the study of the next asteroid exploration mission. The mission now under study gives priority on “early” achievement of the sample return from an asteroid with primitive composition. Therefore, the design of the spacecraft follows that of HAYABUSA basically as it is, and the spacecraft is planned to be launched in early 2010s. The objective of the mission analysis is to design a mission sequence, which has launch window in early 2010s, which is feasible by a HAYABUSA-type spacecraft, and whose target asteroid complies with the science objective. The result includes the selection of the target asteroid, the design of nominal mission sequence, and some back up plans.

“Mission Analysis of Multiple Near Earth Asteroids Exploration by Miniature Asteroid Interceptors”
Kawakatsu Y., Mori O., Tsuda Y., et al.
Advances in the Astronautical Sciences, Vol. 124, pp. 1773-1787, 2006.

Discussed in this paper are the results of the mission analysis of near Earth asteroid flyby missions using miniature Asteroid Interceptors. The Interceptor is an autonomous self-contained interplanetary probe with 10kg mass which is now under development in ISAS/JAXA. It has the capability of navigating itself autonomously to flyby the target asteroid using optical navigation system. The image of the asteroid taken by the camera onboard at the closest approach is the main science output of the mission. Firstly discussed is the mission by a single Interceptor, which enables the minimum size interplanetary mission. The interceptor is launched as a piggy back mission on a geostationary mission, separated on a geostationary transfer orbit (GTO), kicked by a solid rocket motor, and injected into an orbit suitable for encountering the asteroid. It is shown that the utilization of the Earth synchronous orbit and the Earth swing-by drastically increase the number of the possible target asteroids, which enables the selection of more scientifically interesting target for a given opportunity. The second mission concept discussed is the multiple asteroids exploration with a single launch. A straightforward application of the single Interceptor mission, that is, the mission by several independent Interceptors is shown firstly, and an option to overcome the difficulty in performing critical operation of multiple spacecrafts simultaneously is also discussed. The list of the target asteroid candidates, detailed mission sequence and maneuver parameters are shown for the assumed example mission.

“System Level Investigation Example in SELENE Spacecraft Design”
Kawakatsu Y., Sasaki S., Takizawa Y.
Proceedings of the 24th International Symposium on Space Technology and Science, pp. 1092-1102, 2004.

In this paper discussed is the system level investigation in the spacecraft design process. The term "system level” used in this paper involves the meanings “inter-subsystems” or “over-subsystems” or sometimes “inter-systems”. System level investigation is performed in various aspects of the design process. Those are, the allocation of the function and performance at the initial step of the design or various kinds of system analysis performed during the design. Methodology or procedures are established to some extent as to the investigations mentioned above. However, as space mission is sophisticated and spacecraft technology is advanced, the problems of new type arise that cannot be coped with in the way of the investigations established so far. What we want to discuss in this paper are the two examples of the system level investigations we faced in the spacecraft design of SELENE (SELenological and ENgineering Explorer). The first is the mission analysis of LISM, one of the main scientific missions of SELENE. Analysis as to the requirements of observation coverage is discussed in detail. The second is the onboard operation of the reaction wheel unloading.

“Application of Phasing Orbit on SELENE Translunar Trajectory”
Kawakatsu Y., TakizawaY., Kaneko Y., et al.
Proceedings of the 22nd International Symposium on Space Technology and Science, pp. 1570-1575, 2000.

SELENE is planned to be injected directly into the translunar trajectory. In this case, it takes only five days from launch to moon arrival, which leads to simple straightforward translunar trajectory sequence. On the other hand, it has a little weight demerit to cope with the worst translunar conditions in the launch window. However, by applying the idea of phasing orbit, better translunar condition can be achieved while holding adequate launch window. The translunar trajectory can be fixed for acceptable width of launch window by setting the buffer (phasing orbit) to adjust the duration from launch to translunar injection. Preliminary analysis using two-body model is done to estimate the effects of the phasing orbit. The weight merit of applying phasing orbit turns out to be more than 100kg at the lunar orbit injection considering additional DV of 50m/s prepared for multibody disturbance.

川勝康弘, 岩本祥広, 金子豊, 他
第16回誘導制御シンポジウム, 平成11年

n this paper, the lunar landing technology of SELENE and the next lunar explorer is discussed. Technical topics mainly related to guidance and control aspects are listed. The topics are classified into five subjects, those are navigation, guidance, control, instruments including sensors and thrusters, and system. In each subject, problems, difficulties, their solution in SELENE, and the prospect in their future trends are described. The landing technology now planned for the next lunar explorer is also introduced.

“Trajectory Design of SELENE Lunar Orbiting and Landing”
Kawakatsu Y., Kaneko Y., Takizawa Y.
Advances in the Astronautical Sciences, Vol. 100, pp. 269-280, 1998

This paper focuses on the three topics related to the trajectory design of SELENE. The first is the orbit maneuver of the orbiter. The altitude of the orbiter is 100km and the orbit is strongly perturbed by high order term of the gravity potential. In order to satisfy the mission requirements, ten maneuvers are scheduled during one year mission. The second topic is the orbit design of relay satellite. The relay satellite has no propulsion system and has no orbit maneuver capability. The orbit of the relay satellite is perturbed mainly by earth's gravity and the shape of the orbit changes through the one year mission. The initial orbit is selected carefully to meet the mission requirements through the mission considering the effect of perturbation. The third topic is the trajectory design of the landing mission. The navigation error in the landing phase is expected to be large value. Main reason of the error is orbit determination error and long duration of inertial navigation. The landing trajectory is designed to permit this navigation error and assure the safe landing.

“Feasibility Study on Enhancement of OTV Capability by Tether for Orbital Transfer Operations”
Kawakatsu Y., Watanabe T., Nakasuka S.
Transactions of the Japan Society for Aeronautical and Space Sciences, Vol. 38, No.122, pp370-382, 1996.

The orbital transfer operation utilizing the OTV (orbital transfer vehicle) equipped with a tether is proposed. There have been many ideas of the transportation using the tether since 1960’s. Most of them simply apply the gravity gradient effect to change the orbit of the payload attached at the tip of the tether. However, it is clear that a larger orbital change can be achieved by utilizing a librating or rotating tether. In this paper, the basic model of the mission using the rotating tether is introduced. The methods to rotate the tether are explained and their effectivenesses are verified by numerical simulations. At the same time, the fuel consumption for typical orbital transfer mission is studied. The results indicate that this type of OTV can reduce the fuel consumption by 45% maximum compared with conventional nontethered OTV system.

“Efficient Design of Launch Vehicle Trajectory Utilizing Design Knowledge”,
Kawakatsu Y.
45th International Astronautical Congress, ST-94-562, 1994.

A method of designing the practical 3D-launch trajectory is proposed. In the initial phase of the trajectory design, two capabilities are required for the design method. The first is the ability to cope with the practical problem, especially the constraints and the requirements. The second is the applicability to the management operations. The management operations involve dissolution of the conflicts between the constraints and the requirements, or the suggestion about the enhancement in ability by the relaxation of the constraints. The basic design methodologies employed in our method are 1) using the simulation to take any kind of constraints into account, 2) searching the trajectories from the sets of the important parameters’ combinations, 3) the reduction of the search space by the application of design knowledge. The effectiveness of this methodology is verified by its application to an example problem. The example is the launch trajectory design of Mos-1b satellite by H-1 rocket. Additionally, three techniques utilizing design knowledge to reduce the design time are introduced and their effect is verified.