Elsevier

Acta Astronautica

Volume 52, Issues 2–6, January–March 2003, Pages 189-195
Acta Astronautica

Synthesis of an alternative flight trajectory for Mars explorer, Nozomi

https://doi.org/10.1016/S0094-5765(02)00156-XGet rights and content

Abstract

This paper deals with the orbital synthesis for the “Nozomi” spacecraft, which was launched in July 1998. The spacecraft trajectory was tactically designed to extract the gravity assist from the Sun, Earth and the Moon so as to be boosted up even beyond escape energy. This orbital maneuvers well demonstrated how this scheme works well. However, an incident occurred when the spacecraft was about to leave the Earth on 20 December 1998 owing to the pressurization problem of the reaction control system aboard. The correction capability was greatly lost and the original scenario was on the verge of collapse. Within two Weeks from then, the trajectory correction had to be performed if the Mars was targeted even though the fuel was not enough. An alternative flight was designed within the shortest available period and the correction was canceled and the spacecraft was directed to return to the Earth first. The altered flight utilizes double Earth swingbys to change its ellipse apsis direction completely to the opposite side so that the Mars Insertion delta-V can be drastically reduced. The paper describes how the orbital synthesis is devised. Trajectory correction strategy is also given for the “Nozomi” to fly back to the Earth in 2002. The contents, in part, were already reported in the IAF congress in Amsterdam in 1999.

Introduction

“Nozomi” spacecraft was launched by the Institute of Space and Astronautical Science (ISAS) in July 1998 from Kagoshima Space Center in the southern part of Japan. Nozomi is the Mars orbiter for the plasma physics observation carrying the international payload of about 30kg aiming at disclosing the solar interaction with the Martian intrinsic magnetic field. The spacecraft is illustrated in Fig. 1. The spacecraft that weighs about 540kg including the bi-propellant fuel of 285kg. After 5.5 month operations in phasing orbit, which included 2 lunar swingbys, the spacecraft left the earth in December 1998. At the time of 500N main engine firing to leave the earth gravity field, there was a problem with one of the latching valves resulting in insufficient, so corrective maneuvers followed on the next day. The total consumption of the propellant exceeded the originally planned one. Hence, the orbit plan was changed from the original one and a completely new orbit plan was developed which requires significantly smaller amount of the propellant. In exchange for it, the arrival time at Mars will be January 2004 as opposed to the originally planned time of November 1999. Another technical problem occurred with S-band power amplifier; however, the final orbit around Mars will not be affected and with 14 onboard science instruments in good shape, scientific rewards will be obtained as originally planned except part of the occultation observations. This paper focuses on the trajectory synthesis process immediately after the incident took place before the correction deadline. The paper here also deals with the correction strategy result minimizing the fuel amount with an admissible terminal dispersion at Earth swingby.

Originally it was planned to be launched in 1996; however, owing to the delayed launcher development, it was slipped to 1998. The window in 1996 was the best window among several years and the alternative trajectory sequence had to be introduced to compensate for the payload capability deficit. What ISAS devised for this purpose was to utilize the lunar and solar gravity assist to boost its speed up beyond the escape velocity from that corresponding to the trans-lunar trajectory. This pumping mechanism accelerates the spacecraft velocity by almost 120m/s that deserves 24kg of the dry spacecraft mass. In view of the total spacecraft mass, it is well understood how the use of this scheme is advantageous. The section presents the essence of the idea and how Nozomi has performed this sophisticated trajectory maneuvers. The orbital sequence taken by Nozomi was very unique in its boosting mechanism. The primary mechanism is, first of all, the lunar swingby to accelerate it. As obviously, a single lunar swingby is not adequate for increasing the orbital energy required for the journey to Mars. Just the repetition of the lunar swingbys does not help in pumping the energy up. The relative velocity to the moon needs to be higher so that another advantage can be extracted from the swingby. To this end, the spacecraft trajectory was designed to be thrown away once to the Earth gravity field boundary region where the solar gravity affects the trajectory around the Earth. The effect is so big as to make the orbit retrograde, resulting in the higher swingby velocity at the second encounter with the moon. The effect is, from dynamics point of view, the same contribution as that employed in our LUNAR-A mission where the solar gravity effect was utilized to lower the encounter velocity to the moon when it is captured around it [1], [2], [3], [9]. The idea was actually demonstrated in our Hiten mission ’91, in which the spacecraft took the advantage of the solar gravity assist having enabled it to orbit around the moon with less fuel [4], [5], [6]. Details of these ideas on how the trajectory can be synthesized are presented in [1], [7], [8], [10]. Three lunar swingbys strategy to escape were also shown, which may interest the reader more.

The first swingby took place on 24 September and the second one on 18 December 1998. When Nozomi finished its second lunar swingby, the orbital energy had already exceeded well beyond the escape level and the last perigee passage is therefore Earth swingby with delta-V, in other word, the Earth powered swingby. The swingby is a relatively tight swingby in which the lowest altitude was 1000km above the surface. In order not only to add the velocity but also to bend the escape direction in compliance with the flight to the Mars, the escape (not really for escape) maneuver of 420m/s was attempted, applied to on 20 December 1998. This is the trans-Mars Insertion (TMI) maneuver. The plan view of the launch to escape is shown in Fig. 2. All the sequences of the flight were perfect, however, an incident occurred in the propulsion system onboard and the associated fuel consumption to correct the trajectory was so large that the original timeline had to be given up. It occurred in the midst of Pacific ocean followed by the flight over the northern America and European continent (see Fig. 3). The pressure history is shown in Fig. 4 that simply indicates the oxidizer tank pressure gradually decreased while the fuel tank pressure was maintained properly. The resulted thrust was not enough. The spacecraft was not visible and was not controlled from ISAS, Japan, for almost half a day. As mentioned above, when ISAS had AOS, the energy sensitivity had reduced greatly and this lead to the extra fuel consumption. The troubled latching valve makes the thrust dwindled and the burn did not give the exact speed to the spacecraft toward Mars. ISAS anyhow performed the compensation maneuver to make it reach the Mars in October 1999, even though the fuel shortage was anticipated.

After the compensation maneuver, ISAS set about devising the orbital sequence that may be substituted with. During two weeks, there were found calculated four ideas to it. They are listed in Table 1 below. If the maneuver had to be attempted even though the fuel is never sufficient for the science [11], the correction was needed to be done as early as possible.

The delta-V requested grew up as Fig. 5 shows and the deadline for the new orbit synthesis was very early January. The period left for us was less than two weeks. ISAS orbit synthesis team made a strenuous effort to devise the trajectory and had come up with the new sequence below. After the TMI ended, the remaining fuel available was estimated 1060n/s including that for the attitude and trim maneuvers. Idea 3 may have been thought as barely satisfying the fuel budget. But it was turned down by taking potential extra compensation maneuvers into account. As clearly understood by Table 1, proposal 4 was decided to be taken. It did not assume any powered swing by that must be performed accurately and may request a little tough specification for the RCS aboard not in perfect condition. The fourth idea only postulates the bi-propellant burns at the TCM scheduled at the end of February and at the MOI. Therefore, it is a safer way. Besides, the fuel amount required for it is surprisingly low and good enough for the whole science missions unchanged. Idea 1 trajectory is shown in Fig. 6.

Section snippets

Current status of “Nozomi” spacecraft

As mentioned, the flight sequence was amended and an alternative strategy was developed to relax the fuel consumption, leaving the sufficient amount left onboard.

Fig. 7, Fig. 8 show the new trajectory developed in both planview and birdview. Fig. 9 also schematically shows the heliocentric path back to Earth. The remaining flight period is approximately 2 years and the trajectory correction schedule needs to be fixed currently.

Associated with this alternative scenario, the angles properties and

TCM plan in near future

The paper presents the orbit correction example. Suppose, here, first that the first correction is done on 15 June 1999 and that the terminal dispersion specification is 11.9km. Table 3, Table 4 list the TCM plan when the total number of TCM is four including the initial correction and it show that the total delta-V will be around 8m/s.

Fig. 13 shows the delta-V history corresponding to Table 5 and it also draws the terminal dispersion error. The first delta-V occurs when the delta-V profile

Concluding remarks

The PLANET-B spacecraft of the Institute of Space and Astronautical Science (ISAS) of Japan was successfully launched on 4 July 1998 and is currently on its journey to Mars. It was renamed “Nozomi” (Hope) after the launch. It has so far performed two lunar swingbys together with the solar gravity assist to accelerate its speed to save the fuel onboard. This escape scheme was first taken by the spacecraft and exhibits the new path to the interplanetary flight widely available for the smaller

References (11)

  • K. Uesugi, Space odyssey of an angel—summary of the HITEN's three years mission, AAS/GSFC International Symposium on...
  • K. Uesugi, J. Kawaguchi, et al., Follow-on mission description of HITEN, Proceedings of the 18th ISTS, Kagoshima, 17–23...
  • K. Uesugi, J. Kawaguchi, et al., GEOTAIL launch window expansion and trajectory correction strategies: analysis and...
  • H. Mizutani, et al., Lunar interior exploration by lunar penetrator mission, 41st IAF, IAF-90-039, Dresden, Germany,...
  • E.A. Belbruno, J.K. Miller, Sun-perturbed Earth-to-Moon transfers with ballistic capture, AIAA Journal of Guidance and...
There are more references available in the full text version of this article.

Cited by (11)

  • An Earth-Mars microsatellite mission leveraging low-energy capture and low-thrust propulsion

    2022, Acta Astronautica
    Citation Excerpt :

    The gravity assist maneuver has been extensively applied to reduce the total ΔV required by interplanetary transfers [7–11] and it was proposed to transfer the small satellite Nozomi from the Earth to Mars [12]. Nozomi's trajectory included a double lunar swing-by, providing the ΔV required to escape from the sphere of influence (SOI) of the Earth, and a powered swing-by at the perigee, redirecting the spacecraft toward Mars [13]. Because of an incident that occurred during the last swing-by, Nozomi could not conclude its transfer to Mars, but it defined a milestone in the design of interplanetary transfers.

  • Earth-Mars microsatellite mission using ballistic capture and low-thrust propulsion

    2021, Proceedings of the International Astronautical Congress, IAC
View all citing articles on Scopus
View full text