http://www.mext.go.jp/b_menu/shingi/uchuu/...ts/04061101.pdf

Above pdf file will be translated for aspiring students in aeronautics, control engineering etc. so that in future lay people like me will be able to enjoy planetary scenes and events without worrying about failures.

The overall title is "Looking into the causes of failure and trying to find the right measures to take for the future with respect to the 18th scientific satellite (PLANET-B ) not inserted into Mars orbit as planned" and it is dated 21 May 2004.

This file is very much detailed at 1.1 megabytes and the number of pages is about 40, I think. In addition, I will be translating 3 more files after this particular file. They will be;

1. ISAS file with views and comments on the failure
2. Another ISAS file, a newsletter written out in a series of 4 individual letters.
3. JAXA file, which is a press release and it is a very concise document with just sufficient details.

Re concise link making I tried a few times, but I simply failed and all the links will be fully pasted out as required.

Pandaneko

above for ease of reference

(preliminary, just a gist of SAC comments dated 26 May 2004 on the above report as follows)

We accepted above report. Most grateful to those who spent time on this report out of their own busy schedules.

This report talks about two major failures and the findings will be refelected in the future science satellites design philosophy particularly in the areas of;

1. design changes
2. ground tests
3. policy on imported parts
4. failure separation
5. software operation

We hope that JAXA will be making best use of this report for their routine inspection/checking procedures and R&D activities.

P

(I will upload the list of contents immediately after this)

above for ease of reference

(list of contents is as follows)

Preliminary･･･････････････････････････････････････････････････････････････････1

Ⅰ．Outline of Nozomi

１．Outline of the satellite････････････････････････････････････････････････････････････2

(1) Objectives of Nozomi･･･････････････････････････････････････････････････････2
(2) Outline of the satellite･････････････････････････････････････････････････････････････2
(3) Instruments on Nozomi and the outline of the knowledge obtained by the time orbit insertion was abandoned････2

２．History of development and its background････････････････････････････････････････････････2

(1) History of Nozomi development･････････････････････････････････････････････････3
(2) Science targets of Nozomi･････････････････････････････････････････････3
(3) Nozomi development philosophy･･･････････････････････････････････････････････4
(4) Nozomi design philosophy･･･････････････････････････････････････････････5
(5) Lighter Nozomi due to launch postponement････････････････････････････････6

３．Outline of the history leading up to abandonment of orbit insertion････････････････････6
(1) Occurrence of fuel system failure (20 October 1999)･･････････････････････6
(2) Occurrence of coms. and thermal control system faillure (25 April 2002)････････････････････7
(3) Operation for fault recovery･････････････････････････････････････････････････7
(4) Abandonment of orbit insertion (9 December 2003)･･････････････････････････7

Ⅱ．Looking into the causes of fuel supply system mulfunction

１．Circumstances of mulfunctioning･･････････････････････････････････････････････････････8

(1) Outline of propulsion system･･･････････････････････････････････････････････････････････8
(2) About the operation for Mars transfer orbit insertion･･･････････････････････････････････････8
(3) Grasping the telemetric data and analysis･･････････････････････････････････9
(4) Operation after telemetry and status of LV2････････････････････････････････････････････10

２．History of selecting LV2 valve for Nozomi･･････････････････････････････11

(1) History of LV2 valve selection･････････････････････････････････････････････････････････11
(2) Inspection contents of LV2･････････････････････････････････････････････････････････13

３．Estimated causes of the mulfunction･･････････････････････････････････････････････････････15

(1) Fault tree analysis (FTA)･････････････････････････････････････････････････････15

-ⅰ-

(2 Candidates for faults･････････････････････････････････････････････････････････15
(3) Estimated causes of the mulfunction･･･････････････････････････････････････････････････17

Ⅲ．About the mulfunctions in the coms and thermal control systems

１．Circumstance of these occurrences･･････････････････････････････････････････････････････18

(1) Outline of the power system･･･････････････････････････････････････････････････････････18
(2) About the operation between the times of contact loss prior to the mulfunction and mulfunction day･････････････････18
(3) Space environment on the day of mulfunction development･････････････････････････････････････････････････19
(4) "1 bit coms" and grasping of the probe status by autonomous function･････････････････････････19
(5) Grasping the status thanks to the operational recovery after fault development･････････････････････････････････20

２．Estimating the causes･･････････････････････････････････････････････････････22

(1) Fault tree analysis (FTA)･･･････････････････････････････････････････････････22
(2) Causes of the short circuiting･････････････････････････････････････････････････････23
(3) Estimating the causes of mulfunctions･･･････････････････････････････････････････････････････25

Ⅳ．For the future

１．For the future fuel supply systems････････････････････････････････････････27

(1) Measures to be taken in valve selection･････････････････････････････････････････････････27
(2) Measures to be taken in satellite operation･･･････････････････････････････････････････････････28
２． Measures to be taken in coms. and thermal control systems･･････････････････････････････････29
(1) Measures by seperating out the faults･････････････････････････････････････････････････････29
(2) Development of parts which will not develop latch-ups･････････････････････････････････････29

３．Reflecting into the design philosophy of future science satellites････････････････････････････････････････････30

(1) Design changes･･･････････････････････････････････････････････････････････････30
(2) Ground tests･･･････････････････････････････････････････････････････････････30
(3) Imported parts･････････････････････････････････････････････････････････････31
(4) Fault seperation･･･････････････････････････････････････････････････････････････31
(5) Software operation in contingencies･････････････････････････････････････････････････31
(6) Policy on future deep space missions･･･････････････････････････････････････32

Ⅴ．graphics and charts and tables･･･････････････････････････････････････････････････････････････34

‐ⅱ‐

Ⅵ． Glossary and abbreviations･･･････････････････････････････････････････････････････････78

（reference 1） Finding the causes of the failure of Nozomi and future measues to be taken･･････････････････82

（reference 2） Members of SAC (I will not be translating this reference, P)･･････････････････････････････････83

（reference 3） Schedule of the investigation meetings･････････････････････

(end of the contents list page)

P

Thanks for doing this. It is very interesting.

Phil Stooke

Yes indeed, THANK you, Pandaneko! In my opinion, what you're doing here is one of the most valuable things that multilingual UMSF people can do: provide translations of technical documentation, which of course is rarely affordable for individual projects or even national space agencies.

above for ease of reference

Page 1: preliminary (omitted)

Page 2:

１．Outline of the satellite

(1) Objectives set for Nozomi

Nozomi was conceived with its main objectives of looking into the direct interaction between the solar wind and Maritian atmosphere/ionosphere. In addition, Nozomi was conceived as the first planetary probe of this country trying to look into solid planets and serving as an engineering test satellite for future deep space missions.

(2) Outline of the satellite

Nozomi is a spin stabilised satellite with a high gain antenna fixed atop a pentagon shaped pillar boby. It is a small satellite, with its inertial mass of 540kg (of which 280kg is fuel) and the total height of 2.4m (from nozzle tip to end antenna) and diameter of 1.6m. It carried 15 different instruments (35kg inluding an extensible structure). Nozomi's ultimate shape (imagined) in its Mars circulating orbit is shown in the graph I-1-1.

(3) Outline of the knowledge obtained before insertion abandonment and its instruments「

Nozomi carried 15 different instruments such as an extreme ultraviolet imager, ultraviolet imager, ion energy spectrograph etc. Main findings using some of these include the world first image of Earth's plasma sphere and interstellar materials measurements. Altogether, 10 out of 15 instruments were actually operated.

In addition, Nozomi had 8 engineering objectives required for future deep space missions such as ultra high precision in orbit determination and autonomous control of the probe. These naturally form an important basis for our future deep space missions.

For your information, the outline of observational results, list of instruments, achieved engineering objectives, and the positions of each instrument on board are shwon in tables I-1-1, I-1-2, and schematics I-1 and I-2.



２．Background and history of its development (what a strange place for this to be!, P)

end of page 2

above for ease of reference

Page 3 (I believe)

(1) History of Nozomi development

It was decided in 1991 that we will start developping, with interplanetary missions in mind, the M-BV series of solid fuel rocktes. This marked a contrast with the earlier M-3S2 series of solid fuel rockets in that the launch capacity will increase from approx. 770kg to approx. 1800kg (ability for insertion into the lower earth orbit) and meant that planetary missions were suddenly within the grasp of scientists. Nozomi was thus conceived.

Since interplanetary missions require a large amount of launch energy it was decided to make use of planatery swing by method with the Nozomi mission. For your information the table I-2-1 shows the range of mission targets which became possible as the result of M-V rockets.

(2) Science targets of Nozomi

Nozomi's main aim was to look into the direct interaction between the solar wind and the upper atmosphere of a planet. About whether its target is to be Venus or Mars the then Institute of Space and Astronautical Sciences (ISAS) made an extensive investigation taking into the account the voices of scientific communities interested in planetary science.

Based on this it was finally decided that Nozomi's science target was to be Mars taking into the account the following points.



① There was very little observational result at that time.

② Earlier Viking lander (note 1) showed that Martian atmosphere extended to such a height that could not be fully explained by the pressure balancing of the atmosphere and the solar wind.



③ Earlier Phobos 2 probe's (note 2) observation suggested that an extremely large amount of oxygen ions flew into the interplanatary space which cannot be ignored in our reasoning of the evolution of Mars. For your information the table I-2-2 shows the status of missions to Mars by other contries at the time of Nozomi's planning.



(Note 1) : Viking Lander

This is a NASA Mars lander. Two of them landed on Mars in 19XX (I am afraid I do not have a year designation conversion table for this period ready for translation, P) and offered direct data on Martian atmosphere and ionosphere.

(Note 2) : Phobos 2

Former Soviet Union's Mars observer and stayed in orbit for 2 months from January 19XX (ditto, P) and discovered an extremely large amount of oxygen ions escaping Mars.

end of page 3

(in the earlier page I made a mistake, extensible should have been extendable, i.e., telescopic)

P

above for ease of reference


(3) Nozomi's design philosophy

Nozomi was designed to be launched in 1997 and its development and manufacturing spanned the four years starting in 1993. Its development was based on the following points.

① It should carry world first class instruments which can expect maximum returns.

② In order to maximise its scientific returns Nozomi should be designed to be proactively international.

③ Most reliable engineering technologies should be employed to secure its mission.

④ In manufacturing the probe only reliable and trustworthy parts and instruments should be procured.

⑤ Since the launching rocket is to be an M-V type the maximum weight of the probe should be kept within 530kg (later increased to 540kg thanks to the improvement of the rocket capability).

With respect to the points 1 and 2 above we are pleased to note that overseas groups were providing 4 instruments, ultrahigh stable resonator, image compression chips with world top class obervational capabilities.

In addition, all the data obtained by its mission was, ultimately, to be made available to all scientists across the world.

About the point 3 above, we adopted the dual liquid propulsion system because we concluded that aerocapture (note 3) and electric propulsion system (note 4) were still technically unreliable.

About the point 4 above, the valves that were right for the Nozomi specs were not produced by domestic manufacturers. This meant that we would have to use overseas parts with restrictions on the provision of technical indformation. For this reason, we decided that we should be sufficiently careful in order to ensure that they met our requirements in terms of reliability through quality assurance tests and related tests and inspections.

About the point 5 above, we reflected this weight limitation in our probe design (to be discussed later) and reduction of the weight of the probe was concretely put into action.


(Note 3) Aerocapture

This is a technique by which atmospheric pressure resistance is used to reduce the velocity of the probe for orbit insertion. This will allow a very large amount of reduction in fuel consumption. However, in the case of Mars the precision required for deviation from an optimum height is about 5km and is quite chaleenging. In addition, the target height will also vary according to the state of the Mars atmosphere at the time of velocity reduction and also the probe must be protected against heat.

（Note 4) Electrical propulsion

Artificially produced plasma is accelerated by high voltage and released into space for propulsion

End of page 4

(I am trying to be as accurate as possible in my translation. However, if anybody has any questions or require further clarification I will be very pleased to re-translate the bits in question. P)

I understand that the reason for the final failure of the Nozomi mission was as follows:

"In April 2002, on its way to Mars, NOZOMI had experienced a very strong solar energetic proton event associated with a strong solar flare. This caused a short circuit in one of the subsystems and a loss of telemetry signal, which made the Mars orbit insertion impossible."

http://www.spaceref.com/news/viewpr.html?pid=13182

I also understand that Spirit and Opportunity survived the same solar flare without suffering any problems. I have always presumed that the reason for this was either that Japan did not have access to the same space certified components that JPL had access to or that Japan did not have JPL's understanding of designing space hardware in a radiation tolerant way.

the reason why Spirit and Oppy survived the April 2002 solar flare so well was because they were still shielded by the Earth's magnetosphere...

above for quick reference

page 5

(What is immediately after this is the continuation from page 4. I am sticking to the original layout so as not to create confusion to those who may attempt their own translation)

It has higer efficiency compared with the weight of the fuel required. For your information this method of propulsion is one of the targets for testing with Hayabusa spacecraft which was launched in May 2003.


(4) Nozomi design philosophy


Nozomi design is based on the following considerations.

① We pay utmost attention to reducing the weight of the satellite.

② Since there are instruments which are very susceptible to surface (electrical) potential we should take an even more precaution in earthing them compared with usual measures taken for preventing accidnets due to charging and discharging.

③ We should have a certain criteria/standard for the electromagnetic noise leve so as not to affect instruments on board.


④ We should not use potting materials in order to prevent instrument deterioration through resulting contamination. In order to put into effect the point 1 above following improvements were made.


・With these science satellites we are expected to obtain world top class results and for this reason it will be desirable to keep as much ratio by weight of the instruments against the total weight of the satellite. Therefore, while retaining the reliablity standard comparable to that enjoyed by other earlier satelllites the results of the STRAIGHT project (note 5) were put into use with Nozomi. These included surface mounting of parts, batteries using nickel/hydrogen cells, semiconductor data recorder using large capacity memories etc.


・While the satellite was made as light as possible reliability assurance was of paramount importance and for this reason a redundant system was employed around the CPU relating to attitude and orbit control (AOCE). For your information the bus system design gave a higher priority to weight reduction rather than fault seperation ability given that the system used radiation tolerant parts and other parts for space use, all pointing to much lower possibility of mulfunctiioning.


・ Compared with the earlier central power distribution method a dispersed power distribution method was employed for the first time with a science satellite (17 lines).


・With the observation system a radical reduction in weight was pursued by taking into account the relative merits in reliability against on board weight and this included, for instance, a unified electronic control looking after more than one instrument.

(Note 5) STRAIGHT: Study on the Reduction of Advanced Instrument Weight

This is a project looking into the next generation probe technologies.

end of page 5

P

above for ease of reference

I realise I made some omissions about redundancies with onboard computers in the last page. The passage should have read;

the systems around CPU relating to data handling unit (DHU) and AOCE adopted "waiting" redundancy and the register for the common systems had a tripple redundancy incorporated in the system

("waiting" bit, I am unsure. I just simply translated direct from the original sentence, P)

page 6

From the viewpoint of point ② above all of the layers of the thermal blanckets (MLI) well over 100cm in length were earthed. In addition, conducting treatment to the cover glass of the solar cells and earthing were carried out. We also adopted approapriate design for keeping the instruments full capacity as outlined in ③ and ④ above.




(5) Further weight reduction due to launch postponement

Nozomi's launch date was sifted due to the delay in M-V development. It was decided in 1994 that the launch will be postponed to 1999. The most desirable launch timing was 1997 from the perspective of the satellite weight. The new launch timing of 1999 meant that the weight will increase by 30kg, coming from the fuel and this immediately meant that the dry weight of the satellite will have to de reduced by a further 20kg from the original design (10kg was to be covered by the increased rocket capacity).

Given all these we decided that an alteration to the then adopted shape of Nozomi and total reconsideration of the insturment layout was too risky at this stage and that the two years arising from the delay will have to be used to come up with further reduction in weight of individual components without changing the interface with other parts.

One such example includes the power supply to the heater control circuit (HCE) and data recorder (DR). Originally, the power to these was to be supplied from a dedicated source. However, they consumed a relatively small amount of power and the source of power was thus changed to the common systems power source (CI-PSU) and the number of power sources itself was also reduced from the original 17 to 15.

Individual weight reductions achieved during this period and their effects are listed on the table I-2-3.



３．Outline of the history of the failure of Mars orbit insertion of Nozomi

(1) Occurrence of mulfunction in the fuel supply system (20 December 1999)

Nozomi was launched on 4 July 1999 from the then ISAS Kagoshima space centre by an M-V 3 rocket. A mulfunction was detected in the fuel supply system during the escape from the earth gravity on 20 December 1999 and Nozomi failed to produce required propulsion. As a result, it became impossible to contemplate an insertion into Mars orbit during the middle of October 2000.

Therefore, it was decided to make use of 2 earth swingbys and reach Mars after a delay of 4 years some time in late December 2003 to early January 2004. This was a further change of plan.

For your information, the orbit plan after this change is shown with the schematic I-3-1.

end of page 6

above for ease of reference

page 7

(2 ） Occurrence of mulfunction in comms. and temp control systems (25 April 2002)

Nozomi had been flying without hitches atfer that. However, on 25 April 2002 a mulfunction developped in the comms. and temp control systems and it only became possible to receive beacon signals. At the same time it was no longer possible to control heaters. Because of this the fuel froze and it meant that we were unable to use the main and auxilliary engines.


(3) Operation designed to recover from these troubles

For above reasons we tried recovery operation from 15 May 2002, but we did not get anywhere. For your information, by end August of the same year the use of heat generated from onboard instruments and others led to the frozen fuel reaching the melting temp and it became possible to use the auxilliary engines at the start of September.

Therefore, we attempted the 1st earth swingby on 20 December 2002, the 2nd earth swingby on 19 June 2003 and these were both successful and we were able to put the probe into the Mars transfer orbit.

However, parts of the piping system which were meant to supply fuel to the main engine remained frozen and the main engine remained unusable. The temp change history from late July to end September measued at temp measurement points is shown on the graph I-3-2.

(4) Giving up of hope for Mars circulating orbit insertion (9 December 2003)

It would have been possible to place the probe into the Mars circulating orbit had the main engine recovered from the freeze before insertion into the transfer orbit. We continued recovery operation from 5 July 2003. However, above mentioned function did not come back before 9 December 2003 and the hope of Mars circulating orbit insertion was lost.

For your information, an orbit change to avoid the possibility of collision with Mars (approx. 1% possibility) was conducted during the night of 9 December using the auxilliary engines. Following this operation Nozomi passed the point above Mars surface at approx. 1,000km on 14 December 2003 and it is assumed that Nozomi finally escaped from Mars gravitational field by 16 December 2003.

end of page 7

P

above for ease of reference

page 8

Ⅱ．Trying to clarify the causes of mulfunction in the fuel supply system

１．How mulfunction occurred

(1) Outline of the propulsion system

Nozomi's propulsive system was meant for insertion into Mars transfer orbit before reaching Mars, orbital changes for insertion into Mars circulating orbit, and pointing the antenna towards the earth once in Martian orbit and consisted of the Orbit Maneuvre Engine (OME) with a 500N class dual liquid thruster and the Reaction Control System (RCS) with single liquid thrusters for atitude and orbit control.


Nozomi's propulsion system and main specs are listed with the schematic II-1-1 and the table II-1-1.

Nozomi's piping system was typical satellite propulsion system and its OME engine produced propulsion by burning the hydrazine from the fuel tank with NTO. In so doing the fuel and the oxidiser are pushed out by the helium gas into the engine. Both hydrazine and NTO are liquid and are highly reactive when mixed together with self-igniting capability and for that reason the engine does not have a special ignition mechanism as such.

(2) About the operation at the time of insertion into Mars transfer orbit

Nozomi was launched by an M-V 3 solid fuel rocket on 4 July 1999 and went through various orbital adjestments. During the unseen period (tracking station unable to see the satellite in line of sight) on 20 December of the same year an automatic control was executed to put the satellite into the Trans Mars Orbit (TMI).

It was during this sequence that the orbital change by the OME did not attain the required velocity increase. Time sequence of this event is shown on the schematic II-1-2.

1) Status at the time of TMI

According to the US JPL flash report issued around 12:00 (UTC) on 20 December 1999 there was a shortfall of velocity by some 100m/s against the required delta V of 423.22m/s

2) Operation during the visible period immediately after TMI

(part of what follows actually spills out onto page 9, but I am translating the whole paragraph for ease of reading)

The visible operation immdeiately after the TMI (around 17:00 on 20 December (UTC)) indicated that the onborad intergrated value was 327.1m/s as the velocity increase, close to JPL flash report. At the same time the pressure sensoer P4 indicated that the pressure of the oxidiser tank alone showed an abnormal value of 8.3kgf/cm (0.8MPa). (Schematic II-1-3)

end of page 8

P

above for ease of reference

page 9

(what follows follows directly from the last para on the last page)

For this reason a command was sent (18:20 UTC on the same day) to the oxidiser gas system latching valve LV2 to open it (schematic II-1-1) which is instaled as a non-reversible valve in the pusher gas piping system for the oxidiser tank. As a result we obtained following information and we concluded that the propulsion system returned to normal.

・ We obtained confirmation that the status of LV2 was open (schematic II-1-4).

・ Accelometer detected a vibration which seemed to have come from the functioning of the LV2 valve and the gas flow following it (schematic II-1-5)

・ It was confirmed that the oxidiser tank pressure recovered 15.1kgf/cm (1.48MPa) (schematic Ⅱ－１－ 2（1.48MPa）４)

(about this reference number, I am utterly unsure and we will have to wait until tommorrow if above is right. This is a result of my copy/paste and 1.48MPa sneaked into this reference and it ought ot be immdeiately after the kgf and I cut this out and put it in the right place as you can see here. However, if I cut out (1.48MPa) from the reference I will be left with a ridiculous ref number. It cannot be 24, and I suspect that it is actually 4, but there is no easy way for me to find out without losing what I have done by now)



(3) Grasping the events from the telemetry data

1) Telemetry data analysis

Analysis of the telemetry data relating to the propulsion system obtained during the period immediately after TMI indicated what follows.

a) Attitude change (at the time of the latching valve beeing open)


・Attitude change so that OME points in the velocity increase direction and a spinup (10 to 25 rpm) for securing attitude stability for when OME functions, both of these were activated by the RCS engines


・ A series of sequence until liquid and gas system latching valves (LV6, LV5, and LV2) indicator indicated open status were normally executed.


Functioning of OME

・The pressure of the oxidiser tank continued to fall without staying constant throughout the period in which OME burnt, from 14.6kgf/cm (1.43MPa) to 7.1kgf/cm (0.70MPa) at the time of OME stopage (schematic 2II-1-6).

(this makes me think above might have been 2II-1-4, P)


・ Acceleration also decreased from 0.97m/s to 0.76m/s (schematic II-1-7, 22) (again, unsure about this 22, P)


・ Integrated value of acceleration did not reach the required value of velocity increase. However, the duration of OME firing reached the maximum operational 397.5 seconds for safety and the firing was automatically terminated. It meant that there was a shortfall of 100m/s against the required value of 423.22m/s as the reached value was 327.1m/s.



c) Attitude change (at the time of latching valve close)

end of page 9

P

above for ease of reference

(I realise that I forgot to look at the earlier reference number. I will deal with it seperately.)

page 10

・After the termination of OME burn a close command was sent out following the sequence to liquid and gas system latching valves (LV2, LV5, and LV6) and the indicator showd "closed".

・After this a series of sequence including spin down and attitude change was excecuted without any problem.

2) Mulfunction situation

From these points above the mechanism of insufficient delta V is estimated to be as follows.


① Despite the fact that LV2's indicator is showing open status not enough pusher gas is being supplied to the oxidiser tank. From the relationship between the estimated value of empty tank volume change and and pressure change it is estimated that the supplied amount was 2% of the LV2 complete close case.


②Oxidiser tank pressure decreased.


③ Not enough oxidiser was supplied to OME, leading to an inefficient burn, resulting in OME propulsion drop.


④ Because of this unexpected propulsion drop OME firing was terminated as the maximum burn time was reached before attaining the planned delta V.


3) Where faults took place

Insufficient delta V during TMI was due to the decrease in the capacity to supply pusher gas to the oxidiser tank. Possible cause of this decrease is thought to be hitches relating to the check valve CV2 and the latching valve LV2 in the propulsion system. In this particular case it has been shown that this insufficiently opened passway occurred not with CV2 but with LV2 from the following reasons.

① The pressure value returned to normal as a command was sent out from the ground to open LV2.


② The output of the acceleration meter indicated that there was a shock as the LV2 valve opened and also there was a vibration as the gas flew as a result.


(4) Operation after this and the status of LV2


1) Evaluation of the status of the propulsion system

In order to make up for the insufficient delta V during TMI two additional corrective OME burns (TMI_C1, TMI_C2) were conducted and these were carried out normally. We also carried out an evaluation test on the OME propulsion and supply system in order to check on approapriateness of the corrective burns and subsequent health of the propulsion system and we are satisfied that propulsion system returned to normal.

a) Evaluation of OME propulsive power

We obtained, from the acceleration of OME burn and tank pressures data, OME propulsive power and its specific impulse as well as the amount of propellant flow. These characteristics were then sorted out according to the regulator pressure (P2). The result showed that during the subsystem burn tests (SFT) and during the early stage after the launch TMI_C1 and TMI_C2 both showed almost similar characteristics, meaning that there was no trouble with the OME burns. (Schematic II-1-8)

end of page 10

(actually, part of the very last paragraph spills into page 11, P)

above for ease of reference

for your information the reference for the (2) on page 9 is (schematic II-1-4)


page 11

Evaluation of OME supply system

We examined, in terms of OME functioning time, the pressure loss ratio of the gas and liquid systems due to OME burn during SFT and after the launch. All other indicators except that for the oxidiser system during TMI showed almost similar behaviours and we did not find any trace of characteristics change due to material deterioration and comcluded that the supply system was healthy. (schematics II-1-9 and II-1-10)

2) Operation of LV2

With Nozomi's propulsion system LV2 and CV2 prevent the reverse flow of vapour from mixing with hydrazine and leading to explosion. For an explosion to actually take place the vapour must condense upstream and a certain minimum amount must stay there.

However, it was confirmed that there was enough electrical power available from the track record of operation and that it was possible to maintain, due to the relaxation of the allowable tank temperature range, the temperature of the valve module A (VM-A) (schematic II-1-1) 10 degrees higher than that of the NTO tank.

Therefore, we concluded that even if LV2 is open we will be able to be free from above trouble with this newly acquired temp control leading to a dual safety mechanism with the use of CV2.

For this reason, the operation thereafter was such that both LV1 and LV2 were kept open all the time so as not to cause LV mulfunctioning.


２． History etc. of selecting LV2 for Nozomi

As shown above it became clear from the analysis of the telemetry data that the oxidiser gas system's latching valve LV2 developped a mulfunction. This LV2 in question had been selected and passed the verification test as shown below.

(1) History of LV2 selection

1) Method of valve selection

Valves for space use are little produced domestically. Therefore, in the case of Nozomi valves that met neccessary specs were procured from overseas. With this particular valve there was know-how related subtle technical information relating to its design and manufacturing. Therefore, details of its fine structure were not available and it was also prohibited to disassemble the valve in Japan.

LV2 was procured from a US manufacturer with excellent track records in space use. It was converted by the manufacturer, at the request of the then Institute of Space and Astronautical Sceinces (ISAS) so that it conformed to the structure of Nozomi by adding a status monitor (LVDT).

(from here on the remaining text spills into page 12)

LV2 is based on the combination of two components which have substantial flight records.　As mentioned above its delails are not available and what was done was to investigate the potential imaginable risks invloved upon selecting this particular valve. Detailed information of this investigation is shown in the table II-2-1.

end of page 11

P

above for ease of reference

page 12

2) Building in of LV2

We had debated about whether we should carry LV2 on Nozmi and our conclusion was that we should do so from the viewpoint of preventing a massive reverse flow of oxidiser vapour. LV2 is placed in the propulsive system at the downstream side of the regulator after the forking out of fuel and oxidiser pipes and between the check valve CV2 and the oxidiser tank (schematic II-1-1). Its function is to prevent the mixing of fuel and oxidiser at the gas side of the system.

With this particular propulsive system we are employing a dual redundancy system of accident prevention together with the CV2 placed in the upstream. This dual redundant system with LV2 and CV2 does not mean we did not have confidence in any of these valves. Rather, it was employed to increase reliability to an even higher stage against the reverse flow of vapour which might lead to a fatal accident, damaging the probe.

During the development stage of Nozomi in 1994 an accident occurred to a US Mars probe (Mars Observer) and it was lost. Its cause was estimated to be an accidental mixing of oxidiser and fuel vapours in the gas system pipeline. Reverse vapour flow is likely to lead to a total loss of a satellite and this point was also taken into consideration.


3) Adding LVDT

About whether we should install a valve status monitor LVDT with the LV2 our conclusion was that LVDT constituted an important source of information for the steady operation of LV2, reducing the pressure on the operating team and the possibility of human errors.

In the case of Nozomi, on the other hand, there was a factor of influence at the time of valve selection arising from the merger and acquisition of the US valve manufacturer in that it meant the addition of LVDT to a valve without a status monitor but with an ample proven record of operation.

Generally speaking the reliability evaluation of a design change to an operationally proven part has not been established just as in the case of newly developped parts. However, our judgement was that this was going to be based on an operationally proven valve and the design change was going to be introduced by a reliablbe US manufacturer and consequently the risks involved in this design change were small enough. For this reason we selected the LV2 with LVDT added to it.

As shown in the table II-2-1 the risk of valve mulfunction after the desing change was thought to be similar to that of operationally proven valves.

end of page 12

P

表Ⅱ－２－１に示すように、バルブ開閉ができなくなるリスクでは、改造品といえども、そのリスクは搭載実績品と同様であると判断されていた。

above for ease of reference

page 13


4) Structure of LV2

LV2 structural outline is shown by the schematic II-2-1. Inside the valve there are 2 electromagnets (one each on the right and left of the schematic) and they activate the inner piece (some kind of intermidiate piece, P).

いての独自の追加試験の検討
By letting an electrical current flow through the close side electromagnet (on the left) the inner piece is pushed in and the flow is stopped (state of the schematic) and by letting an electrical current flow through the open side electromagnet (on the right) the inner piece will leave the plug (or seperate from the plug, P) and in that state there occurs a small pressure difference between the up and down streams and this differential pressure pushes the plug and the flow will start.

Also, as shown on the schematic II-2-1, LVDT on the right will monitor the open/close status of the valve by detecting the position of the inner piece. Those original valves with track records of being used with satellites did not have this LVDT. Design change by adding this LVDT meant that the plug and the inner piece were seperate items.

This arrangement of having a seperate structure had been proposed by the US valve manufacturer in an answer to the concern that the addition of LVDT meant a longer inner piece and a resulting small positional erorr might influence the open/close operation of the valve.

(2) Contents of LV2 verification

1) Verification method for LV2

Starting from valves used for scientific satellites those valves used for space use are almost all imported from overseas. Consequently, verification of the selected valves are carried out by those overseas manufacturers and the Japanese procurers. In the case of Nozomi too, this verification was carried out by both US manufacturer and in Japan as shown below.

a)LV2 verification method

① Contents of verification requested by the Japanese side to the US valve manufacturer

・ Verification of design validity

・ Verification of manufacturing validity concerning the shipped flight items

・ Flight track records

Flight tracking records include, if possible and approapriate, concrete mission names and usage. We regarded the usage as important because if the past programmes' mission duration and environment had been different it might be difficult to count it as a proven track record.

② Contents of verification conducted by the Japanese side

・ Confirmation of the justification of the verification methods for each of the items which relate to the verification of the validity of design and manufacturing

(up until here, translation from page 13, but the circle 2 continues into page 14 as follows, P)

- 14 -

・ Carrying out a consolidated system test in the state in which the items were actually built into the satellite

・ Investigating into the possibility of our own additional tests with those items whose flight tracking records could not be confirmed and where details of the verification methods were not available.

end of page 13

P

above for ease of reference

page 14

History of LV2 tests

Confirmation tests of LV2 were carried out in two major parts. One is the design quality confirmation test where another test piece different from the flight piece was fabricated and offered to testing. Second is the confirmation test for flight model manufacturing approapriateness, which is conducted on the flight model. Their outlines are shown below and the ground test history of LV2 is carried on the table II-2-2.


① Design quality confirmation test

This is the test whereby design validity is confirmed. In addition to the test conducted by the US valve manafacturer another test was carried out in Japan.

The valve manufacturer conducted their own test using two test pieces, LV2 and a similarly designed LV1 together with HLV. In Japan we used a spare part, common to both LV1 and LV2, and conducted a confirmation test on the adaptability to the oxidiser (NTO) environment.

We had been given a report from the manufacturer that the valve in question had resistance to NTO and ours was carried out seperately to confirm this report using the flight piece of the valve. The number of valve actions during this test is shown on the table II-2-3.

② Confirmation test for the manufacturing validity of the flight model

This was the test conducted on the flight model. Confirmation was sought with the part built into the satellite system for its health. Below is the number of delivered pieces.

・LV1、LV2 : one each
・LV1、LV2 common piece: two
・HLV: one and HLV spare part: one

2) Action history of LV2

The number of counted valving actions during TMI with the valve in question was 42nd since the day of the delivery and 6th from the day of launch. It was confirmed that all valving actions prior to that had been normal. LV2 action history is shown on the table II-2-4.

end of page 14

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above for ease of reference

page 15

３．Estimating the causes of mulfunction

(1) Fault Tree Analysis (FTA)

We conducted an FTA with the following events at tree top in order to estimate where LV2 mulfunction occurred.

・ A command had been sent from the ground for opening of LV2 and LVDT did show that LV2 was open. However, in reality the opening was minimal and the flow rate was very small. The result of this FTA is shown on the table II-2-5. We at JAXA did exchange information with the US valve manufacturer after the accident. However, as mentioned earlier, the information relating to the valve structure etc was limited in availability.

We summarised the possible mulfunction candidates based on the FTA analysis as shown under (2) below.


(2) Possible mulfunction candidates

1) Bad sliding motion of the plug

One strtucural problem with LV2 is that the plug and the inner piece are separated physically. In addition, there is a possibility of the material surface of the sliding component inside the valve suffered from a fletching wear due to the valve motion and this led to the issue of valve material's compatibility with the oxidiser.

It is thought possible that these two factors might have led to bad sliding motion of the valve.


(note 6) : fletching wear

surface damage arising from repeated slidings at the sliding area


a) Valve structure


We did conduct a repeat evaluation on the past flight records and had confirmation that valves with this particular material had flown many times in the past. However, these valves had a monostructure of the plug and the inner piece and the valve was forced to activate with an electromagnet.


With respect to the valve that we used its (valve) open mode was similar to that of so called check valves in that it opens passively when there is a diffential pressure before and after the plug. The driving force arising from this differential pressure was smaller compared with the force originally ensured by an electromagnet and the valve could have suffered if the plug's sliding force resistance (note 7) had increased.


（note 7） : sliding force resistance (please note that this sentence astrides pages 15 and 16)

The resistance against the sliding motion involving two objects in contact. If this force becomes large

- 16 -

then objects become harder to slide against each other, requiring a larger force for sliding.

end of page 15

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above for ease of reference

page 16

oxidiser resistance of the material

Material combination of the sliding parts is austenite (spelling unsure, P) stainless steel at the sliding area of the plug, and ferrite staeinless steel for the valve body. Ferrite stainless is magnetic by nature and is used extensively with electromagnetically activated valves. On the other hand it is slightly less resistant against NTO compared with non magnetic austenite stainles steel.

It is thought possible that fletching wears would have led to corrosion of the valve body area in question after being filled with NTO, itself further leading to an increase in the sliding friction.

2) Gritting of the plug into the opening area

It is thought possible that the plug which had been pressed against the opening area for a long time might have gritted into the opening, making the plug less mobile. It is also thought that from the status of the lowering of the oxidiser tank pressure a small amount of helium gas was still being supplied to the oxidiser tank and that in the event of plug gritting this might have led to , with a high possibility, the complete closure of the sealing area.

With respect to LV2, the duration in which it was kept closed immediately prior to TMI operation was not especially long compared with similar operational duration with other launches. For that reason it is thought that LV2 did not suffere from the same trouble. For your information the history of LV2 use in operation is shown on the table II-2-6.


3) Plug mis-allignment

The valve in question had been subjected to Quality Test (QT) and undergone more than 1,000 times of open/close operation. All this was monitored using LVDT monitor and also pressurised heilum gas flow confirmed the valving actions. Also, quantitative flow volume tests had been put into action before and after this open/close test and it was confirmed that everything was normal as designed.

For these reasons we consider that there was no design problem with the clearance between the valve body and the plug.

In order for this particular mulfunction to take place we could suspect, as its cause, bad manufacturing of the valve. However, the valve in question operated normally 42 times before mulfunction including the period immediately after launch and there was no sign of bad manufacturing.

Furthermore, the valve continued to function normally even after the occurrence of the mulfunction and we believe that the valve is not a culprit.


4) Glitching by foreign pieces

The diamterwise clearance between the plug and the valve supporting body is such that if foreign pieces larger than the gap migrate into

栓とバルブ本体の直径クリアランスは、より大きい異物がバルブ内に入り込むと噛

- 17 -

the area inside the valve they may cause gritting there. However, the pre-launch gas filling was done using a filter whose mesh is smaller than the clearance dimension and it is thought unlikely that there were foreign pieces of that size present inside the piping system and that foreign pieces were the culprit.

Also, generally speaking, once gritting takes place it is rare for the system to resume normal operation immediately. However, in this case we are looking at the function of LV2 returned to normal after the hitch and consequently it is thought highly unlikely that this phenomenon happened to LV2.

end of page 16

above for ease of reference

page 17

5) Temporary immobility (or solidification) by material growth (or some kind of crystal formation?)

If there had been residual water inside the propulsion system it could have been possible that the reaction of fuel and oxidiser led to the formation of ammonium nitrate. However, this possibility has been discarded as extremely low for the following reasons.


・ dryness at the time of propellant filling was very good (condensation temperature being minus 55 degrees C or below)


・ Had there been a mixing with fuel the check valve upstream would have caused mulfunction


・ It would have shown up in a short span of time (a few days)


(3) Result of estimating for the causes of mulfunction

If we are to summarise what we have been talking about so far, we think that the causes of LV2 mulfunction are due to, as shown in 1) of (2), the fact that the valve was susceptive to increased sliding friction of the plug given that the valving was powered by the differential pressure by the seperate inner piece and the plug

and also the fact that the sliding surface caused a fletching wear leading to corrosion by the oxidiser environment, further leading to an inceased sliding friction.

That is to say that the multiplier effect by these two factors was the most likely cause of the failure.

end of page 17

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above for ease of reference

page 18

Ⅲ．About the mulfunctions of comms. and thermal control systems

１. Environment in which mulfunctions took place

(1) Outline of the power source system

Unlike all other scientific satellites which had adopted a centralised power source Nozomi was the first satellite to use a decentralaised power source. Within the Nozomi power system there were 15 power sources and one of them was the Common Instrument Power Supply Unit (CI-PSU) which was used to supply power to a multiple number of commonly used systems (equivalent to power source 2 in fig. III-1-1).

This CI-PSU receives power from solar cells and batteries placed on the primary side and supply power to 10 secondary side subsystems such as the telemetry command interface (TCI) for the X-band transmitter (TMX) and the heat control circuit (HCE).

For your information, the structure of CI-PSU is shown by the schematic III-1-2 and the table III-1-1.

(2) About the operation from the time of losing signals until the day of mulfunction taking place

The mulfunction was first detected during the operation on 25 April 2002 and the location of Nozomi at that time is shown by the schematic III-1-3. Also, the operational sequence from a day earlier (24 April 2002) prior to losing signals until the day of detecting the mulfunction is shown on the schematic III-1-4. This operational sequence tells us the following story.


① At 18:05 on 25 April 2002 (UTC) signals came in in the beacon mode despite the expectation that the transmission mode had changed to the telemetry mode (note 8). We then sent a command from the ground so that the mode changed to the telemetry mode, but this was not successful. From all this it is obvious that the status remaining was such that we could not switch the mode.

（note 8） : telemetry mode

mode in which telemetry information from the satellite is carried on the satellite wave

（note 9） : beacon mode

satellite is emitting waves, but these waves are not carrying telemetry information.


② We had anticipated an attitude change (about one degree) at 09:00 on 25 April (UTC). However, from the reception level of the satellite waves on 25 April it is estimated that this attitude change did not take place (fig.19-III-1-5). It is also estimated that the command sent from the ground for the attitude change did not succeed.

From all above it is estimated that the attitude change remained impossible.

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above for ease of reference

page 19

(3) Space environment at the time of mulfunction being detected

A solar flare of class X occurred on the western area of the Sun's surface on 21 April 2002 with a magnitude of X 1.5 (X-ray flux of 0.15mW/m in the vicinity of the earth). Nozomi received a direct hit of high energy particles associated with this flare on 22 April. As the strength of high energy particles it was the largest since the launch of Nozomi and it is estimated that it was one magnitude higher in energy than anything that Nozomi had experienced earlier. This magnitude is estimated to be such that you normally expect this magnitude in interplanetary space only once or twice during any one solar period (11 years).

For your information the temporal variation of solar proton monitoring by the device on board Nozomi is shown on the graph III-1-6. Also, the high energy particles strength variation experienced by Nozomi since 2000 is shown on the graph III-1-7.


(4) "One bit comms." and grasping of the phenomenon by the autonomous function etc.

It became possible to obtain the data neccessary for the operation of Nozomi (table III-1-2) by establishing a measurement method using Nozomi's autonomous function (note 10). For instance, if you wish to measure the temperature of the inside of the satellite, the beacon is turned either on or off, depending on whether the temp. is above or below the temp. set by the command sent from the ground. (1 bit comms, hereafter) (grapgh III-1-8)

As a result, it became clear that CI-PSU remained in the state of being off. Also, since the possibility of the primary circuit going wrong is extremely low it is estimated that the protection circuit against excessive current flow resisted the CI-PSU from becoming on.

Under these circumstances we expect following phenomena.


① Since TCI is in the off mode a command cannot be issued to TMX. For this reason, it remained impossible to switch between beacon and telemetry modes.


② Since the heat control circuit (HCE) is in the off mode all the heaters inside the satellite controlled by this device are also in the off mode. For this reason, the temp. of the inside of the satellite decreases rapidly to that temp. which is determined by the heat generated by the satellite, sun angle, and the distance to the sun.

Consequently, as shown by the data obtained (table III-1-2) the inner temp. of the satellite decreased below the freezing point of the fuel and attitude control became impossible.

(note 10） : autonomous function

the function which issues a particular set of commands which have been programmed in advance depending on the inside status of the satellite

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(and in this instance, we obtained information by making the satellite execute the command which will turn off the power of the power amplifier of the X-band transmitter).

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With above file I seem to be translating some entries which must be rather obvious to UMSF members. Please put up with this as these reports are meant for the Space Activities Commision (SAC) whose members include usually one civilian in order to reflect the opinions of the lay community. One such member as I recall from a few years ago was a newspaper cartoonist.

P

above for ease of reference

page 20

(5) Grasping events by recovery operation after the mulfunction

Grasping of satellite status, information gathering for the causes of mulfunction, and satellite operation after the mulfunction was carried out as follows.

1) Confirmation of the health of data handling unit (DHU) and command decorder (CMD)

A command was sent from the ground on 28 April 2002 and we succeeded in powering on and off of the X-band transmitter power amplifier (XPA)(XPA is normally ON). From this we concluded that DHU and CMD were operable normally.

2) Grasping probe status by "1 bit comms." using the autonomous function

During early part of May 2002 we managed to obtain probe status through "1 bit comms." using the autonomous function. As will be shown in (4) we discovered that the common power supply (CI-PSU) was in the state of being OFF and also it could not be made to switch into ON mode and that the satellite temp. was far below the freezing point of the fuel on board.


3 Loss of the beacon waves due to the recovery operation of the short circuited portion


This was the operation on 15 May 2002. By sending continuous ON commands to CI-PSU (about 100 times) we tried to burn out the faulty line, but we ended up in losing the beacon waves during this process. We carried out ground tests in order to find out why beacon waves were lost and we estimated that the beacon was made into OFF state because when TCI is started up it also meant that TMX's ON and OFF commands are simultaneously issued leading to erroneous action of TMX relay circuit.

It was also felt possible that the wrong command issued by TCI (such as TMX ON/OFF) in response to the ON command to CI-PSU might lead to TMX going back to ON mode again. Thus, we carried out beacon recovery operation by sending out a series of single ON commands.

4) Recovery of beacon waves

Beacon waves were recovered on 15 July 2002 after sending to CI-PSU a series of single ON commands (about 7,500 times). This confirmed our estimate shown in (3) above. It also became clear that the satellite power (primary power supplied to CI-PSU) was healthy and that CI-PSU were able to supply secondary power at least for a short period of time.

For this reason, it was thought likely that CI-PSU itself was healthy and that the power OFF state was caused by


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the protection circuit against excessive current arising from the short circuiting problem within the secondary devices connected to CI-PSU.

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page 21

5) Defreezing of the fuel tank and RCS thrusters

By the end of August 2002 thanks to the self generation of heat by on-board devices, distance to the sun, and improvements in the satellite attitude part of the propulsion system had been defrozen and in the upper part of September of the same year it became possible to control attitude and try small scale orbital control.

It has been confirmed since then that maintaining of the correct attitude helped to keep the fuel tank, attitude and orbital control thrusters from freezing. Owing to this recovery of attitude and orbital control thrusters we managed to succeed in the 1st Earth swingby operation on 20 December 2002, and also in the 2nd Earth swingby on 19 June 2003.

However, it was also confirmd that recovery of heater function was absolutely vital in defreezing the main thruster (required for orbital insertion around Mars) which was always on the shade side.


6) Recovery operation (operation for recovery of , P) CI-PSU and heater control function

Given that for orbital insertion it was vital to have CI-PSU and heating function back to normal we checked through the gournd tests that a continuous and rapidly-issued series of commands for CI-PSU recovery will not lead to action anomally by CI-PSU.

Based on this we started on 5 July 2003 to issue continuous ON-commands for CI-PSU so as to burn out the troublesome short circuited line on the secondary side. During this operatioh we managed to lose beacon waves. This recovery operation continued until 9 December 2003 without success. We therefore gave up the hope for orbital insertion.

7) Items confirmed through this recovery operation

Following the recovery operation explained as from 1) to 6) above we were able to confirm:

①DHUand CMD were operational normally.

②CI-PSU cannot be made to be ON.

③We cannot switch between telemetry and beacon modes.

④Loss of heating function led to fuel freezing and attitude control could not be achieved.

⑤Primary power supply voltage to CI-PSU was normal.

⑥CI-PSU can provide secondary voltage for a short time only.

From these reasons it was thought that CI-PSU was likely to be healthy and that with a high probability the short circuiting on the secondary device side led to the activation of the over-current protection mechanism, which in turn led to the OFF state of the power supply.

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thanks again for the translations, panda! I knew that they had regained trajectory control but I didn't know that the main engine remained unusable

above for ease of reference

page 22

２．Estimated causes of mulfunction

(1) Fault Tree Analysis (FTA)

1) We conducted a fault tree analysis (FTA) with the fact at the top of the tree that we received no waves despite being in the telemetry mode (TLM_ON). The result of this particular FTA is shown on the schematic (or graph) III-2-1.


① As for the possibility of data processing unit mulfunctioning we discarded it because command(s?) were issued normally and we were able to obtain data.


② As for the possibility of TMX mulfunctioning we discarded it because it fails to explain why we could not change attitudes and also because there was no abnormal current consumption by TMX.


③The fact that TCI power was OFF, that is to say we could not keep CI-PSU to be ON can explain without contradiction why we could not change attitudes. It also explains why we could not swtich between beacon and telemetry modes.


2) We conducted another FTA, following one of the the results of above FTA, that is to say ③of 1) above, with the fact at the top of the tree that we could not keep CI-PSU power to be ON all the time. The result of this particular FTA is shown on the schematic (or graph) III-2-2.

There are three causes we could think of as follows.


① There was an anomally in the function whereby CI-PSU inside the satellite power source (PCU) is turned ON/OFF and this resulted in power not being available to CI-PSU.


② There was a mulfunction inside the primary side of CI-PSU and this meant that although the power was being supplied to CI-PSU the overcurrent protection mechanism worked and CI-PSU could not be brought to be ON.


③ There was a short circuiting inside the secondary side of CI-PSU and although the function of CI-PSU was normal the fact that there was an excessive current flowing downstream meant that protection mechanism activated within a few milliseconds, bringing the CI-PSU to be OFF.


3) Discussions about the result of above FTAs mentioned in section 2)

These are as follows.



As for ① of 2) above this was thought to be unthinkable because all other functions inside PCU were

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normal.

(this particular paragraph continues into page 24 and is rather lengthy for it being translated as part of page 23. So, I stop here and will continue immediately after this as page 24)

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I am confused about this myself. I will look at this tommorrow and try and do correction.

Pandaneko

above for ease of reference

Paolo, I am so glad my translation is being of help with your work! What was strange about yesterday's translation, I had a look at its original. My conclusion is the number 23 towards the end should have been 22, as simply I was translating the bulk of page 22. Having said that I am still feeling very uneasy because I did not type this number 23 myself. I was simply overwriting the copy pasted from the original. Anyway,

page 23

As for ② of 2) it might be thought of arising from the FET (field effect transistor, I think, P) switch failure for DC/AC conversion or IC failure for controling the pulse widths in the primary system (PWM). However, these two possibilities can be discarded because voltages are present on the secondary side through the transformer inside CI-PSU.

As for ③ of 2) components (except those imported from overseas) which could have caused short circuiting on their own singly include 24 ceramic condensors, 7 resisters, and unprotected 43 ICs. It is possible that any one of these could have caused short circuiting mulfunction, or alternatively, short circuiting inside the imported components could be assumed to have caused the mulfunction we have been looking at.

From all these above based on FTA we may summarise the failure causes whose responsibility cannot be ruled out completely as shown in the following (2

(2) Failure causes of short circuiting mulfunction

1) Influence of high energy particles arising from solar flare

a) Deteriolation effects by total dosage

Deteriolation by total dosage is often seen as the cause of solar cell deteriolation by the cumulative effects of high energy particles and appears as an increase in power consumption.

It is without doubt that NOZOMI encountered a very rare and massive groups of high energy particles. However, as of the peak flux on 22 April 2002 plus or minus a few days there was no significant power consumption increase which suggested above mentioned deteriolation (see schematic III-2-3).

In fact, the cumulative dose estimated from the lower portion of schematic III-1-7 suggests that it was similar to NOZOMI's design value (10krad equivalent assuming 1mm Al thickness) was within the tolerance limit. For this reason we may discard, as unlikely, the possibility of the total dosage leading to deteriolation which caused the mulfunction.


Failure by a single event upset

For this to be the cause following two conditions must be satisfied at the same time to explain the mulfunction of this time

① Unexpected switching over took place by a single event upset (note 11)


② Devices which could be switched on by above ① had already caused short circuting, or caused short circuiting following it

Of these, one possibility with NOZOMI is the INS-SA which is used only at the time of launch. Since it is used only at launch time components after the relays are all meant for commercial uses.

However, this particular device had been turned off after launch and it is confirmed that it stayed that way until the day when we lost signals on 24 April 2002. The possibility of this particular relay device being induced to be turned ON, as estimated from the solar proton minitor's count number (graph III-1-6), is 1,000 times higher on 22 April compared with the signal loss date of 24 April.

Furthermore, all this cannot explain the mulfunction of this time unless a short cuircuiting had already taken place in the system after the relay by the time the relay was turned ON, or alternatively a short circuiting did take place within a few hours of being turned ON.


- 24 -

(See, 24 above, it is happening again! this should be 23. It has been there all the time until I noticed it now, I think)

(note 11) single event upset (this ref is on page 24)

Bit flipping by passgae of high enery particles through ICs

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above for ease of reference

page 24

c) Mulfunction by latch-up

About the possibility of ICs latching-up (note 12) by high energy particles those ICs used on board except one of them (59MeV/(mg/cm)) had line (or linear, P) energy investment (this must be wrong and it is most certainly "tolerance" or something starting with T, because abbreviation is, P) (LET) of more than 75MeV/(mg/cm ).

Even if we assume more than 59MeV/(mg/cm ) the possibility of latching-up by galactic cosmic rays is thought to be once in a thousand years or so.

Furthermore, the solar flare of 21 April was such that protons should not have reached the sort of energy which might have caused the latch-ups. Even if we turn our attention to heavier particles such as iron they only appear during the initial stage of the event and it is unthinkable for them to cause failre after a few days of the solar flaring event.

(note 12) : Latch-up

Short circuiting inside ICs

2) Destruction/electromagnetic interferance by discharge

It is possible that discharging may have caused a current flow through the satellite body and onboard instruments/devices may have been affected through latch-ups etc. About this discharging there are a few possibilities as follows.


a) Charging and discharging by high energy particles

Discharging could be caused by accumulation of a large amount of electric charge inside conductive layers which are not earthed. However, in the case of NOZOMI, because of observation requirements an absolute care had been taken regarding charging and discharging issues and as shown per below there are few places for charge accumulation and the possibility is considered to be extremely low.


・All external surfaces of the satellite

These surfaces are earthed to the satellite frame structure with 1MΩ or below and only one exception are a few patches on the rear side of the solar cell panels.


・ Thermal blanckets (MLI)

All layers with areas larger than 100cm (sic, P) are earthed. Also, the surface material is "Black C(or K, P) apton" (carbon coating) and no cracks are possible and for this reason earth lines being cut is unimaginable.


・Inside of the satellite

- 25 - (here again, this should read 24 and when I checked this before translation with the original it was 24... I will no longer be yapping about this as it should not have been any of your concern, P)

All conductors are earthed down to the level of shielding layers for electrical appliances and unearthed circuit ground pattern does not exist as far as the design is concerend. For this reason it is unimaginable for charge accumulation to take place which may lead to discharging.

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above for ease of reference

page 25

Destruction/EM interference by high voltage instruments discharging

In the case of Nozomi all high voltage devices were thoroughly designed as shown per below so that they do not affect the working of other instruments and it is not thinkable that the secondary side of CI-PSU were affected.


① Dust counter (MDC)（±200V）

The high voltage part of the high energy particle counter (EIS)(3kV) is shrouded by the secondary ground of the instrument and even if discharges happen the design is such that no current will directly flow into the satellite structure.


② Ultraviolet imaging system (UVS)(3kV)

The sensor area which receives the high voltage is contained in a vacuum filled glass case. The part exposed to the external area is only the connecting part of the high tension cables and even those are designed so as not to affect the satellite structure.

(3) Estimating the causes of mulfunction

In the discussion concerning (2) above it is true that Nozomi did encounter a very rare magnitude of solar flare which produced groups of high energy partilces. However, there is no data available to verify the relationship between the mulfunction and the the solar flare. Furthermore, the possibility of destruction/EM interference by discharging is thought to be extremely low given the degree of thorough preparation at design stages.

For these reasons the only remaining cause of mulfunction is an accidental component failure. However, here again, the candidate component groups which could lead to single point of failure were rated at quality assurance level as class S equivalent (space use components) and from the viewpoint of failure rate the possibility of failure isthought to be low.

Table III-2-1 shows the accidental failure rates of main class S components.

Those components which may cause mulfunction are those that exist in the secondary side of the circuit and failure candidates and their properties are as follows.


①Ceramic condensors (２４ of them)

Even the largest accidental failure rate per one component is something like once in 90,000 years and this rate is regarded as extremely low as the single failure candidate. For your information, in the case of Nozomi, all the ceramic condensors were selected on the basis of anti-deteriolation and anti-high tension.

② Resisters （７ of them）

Here, the largest accidental failure rate is something like once in 20,000 years and furthermore, the possibility of line breakage leading to mulfunction is more likely. Thus, we do not believe that they were the causes of single failure event.

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For your information, in the case of Nozomi, all resister were selected on the basis of anti-power deteriolation.

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page 26

③ Unprotected ICs (43 of them)

Probability of accidental failure rate is something like once in 10,000 years and this probability is considered to be extremely low. For your information, it is becoming increasingly difficult with latest ICs to insert protective registers given their large current amplitude and allowable voltage width, leading to a larger number of candidate (suspect, I think. P) ICs.


Also, we can think of, as candidates, the status monitor (LVDT) which is an imported unit for monitoring open/close status of the valve, ultra highly stable resonator (USO), internal short circuting of pressure sensors. With these, we find it difficult to make evaluation for the cause of mulfunction as although they have good track records to have been on board many satellites we simply do not know the circuit inside these black box units.

As you have seen above we have been able to whittle down to a few units which might have caused the mulfunction. However, it is difficult to pin down on specific single units for cause clarification.

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above for ease of reference

page 27


Ⅳ．For the future

１．For the future of fuel supply system

(1) Measures to be taken in selecting the valves

From the mulfunction of this time we can point out two major issues relating to the selection of valves for Nozomi. These are the fact that we made a design change to the valve which had an ample track record of going on board in space and also the verification methodology of that valve. We show measures to be taken in relation to these two issues as follows.

1) About design changes

The valve we are talking about as developping a mulfunction is LV2. Originally, this valve made by a certain US manufacturer with ample flight records was changed at the request of the then Institute of Space and Astronautical Sciences (ISAS) to include a status monitor (LVDT) for monitoring of valve opening and closing.

This LVDT was added to the existing valve to ensure reliable operation of LV2. It meant that the valve with enough track records but lacking the monitoring function was changed to include LVDT as a post desgin alteration.

As for this change of design, since it was made by a specialist US manufacturer based on existing design with track records we thought, at the time of decision making, that the risk involved was small enough, but we now think that our study at that time was not satisfactory.

At this time, given the structure of Nozomi, we had to introduce a design alteration to LV2, but we think that inherent risks involved in desgin altered componets must be treated adequately.


2) About the verification methodology

Most of space use valves are imported from overseas and in some cases detailed structural indormation and availability of technical information is limited. It is vital to establish high reliability with these products. With Nozomi we did conduct LV2 verification and evaluation as shown below, but we now think that it was not enough.

With Nozomi verification tests were conducted as shown on tables II-2-2 and II-2-3. We also requested the US manufacturer for similar verification and in Japan we conducted an independent study on the NTO vapour arising from the oxidiser. We conducted an even harsher test of sealing NTO liquid inside a valve for keeping and for action tests, thereby verifying the durability of the valve against NTO environment and confirmation of its health.

This particular test was conducted so as to verify that LV2 had durability against NTO and lasted only for two months equivalent length. In that sense it was, strictly speaking, not an accelerating test approapriate for Nozomi's operational lifetime (one year had been assumed)


- 27 -

The number of valving actions during the accerleratin test is important from the viewpoint of verifying the influence of that number affecting the condition of the sliding part of the valve. In our case this time it was less than 10 cycles of opening and closing that were tested in the oxidiser environment and we cannot deny the possibility of it being inadequate for taking into consideration the possibility of fletching wears etc.

From all these reasons mentioned as above we should have paid a lot more attention to the issue of conditions in which verification is sought. Valves in particular have wide areas of concern relating to electricals, materials, fluid dynamics etc and we should be closely working with specialists in these areas both inside and outside our organisation in order to check if proposed verification methods are adequate for our purpose.

With imported valves we may think of, as a means to further improve on the adequacy evaluation technique, an early detection of valve deteriolation from the changes appearing in the current wave profile imprinted (or imposed?, P) on the valve. Clearly, we should be spending a lot more on strengthening our verification stance.


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above for ease of reference

page 28


(2) Measures taken in our satellite operation

Given Nozomi's launch window the timing of OME firing for transfer orbital insertion (TMI) had to be within the invisible time zone in our orbit plan design. We outline the measures taken in operating this satellite as follows.

In order to combat invisible operation during TMI we thought of requesting the support of NASA's Deep Space Network (DSN) or ESA so that they may allow the use of of their ground stations. However, our final decision was that we will not be requesting their support for the following reasons.


・Nozomi was equippred with automatic OME firing function.

・ This automatic OME firing function was also going to be used in Mars orbital insertion and a thorough ground verification had been made.

・It had been planned so that an orbital firing test of this automatic function was going to be made during visible operational period (in fact, this automatic OME firing was tested about 4 months before TMI during ΔV5).

・ Nozomi's TMI timing coincided with the period in which Nozomi was also invisible to DSN and ESA ground stations.

However, we cannot deny the fact that had we been able to send an immdeiate response command to the event that happend in real time monitoring in visible operation through telemetry we may have been able to carry out the originally planned task. This does suggest that securing operational visibility is very important. One such measure that can be taken in this respect is to request the support of overseas ground stations. We can think of a few things as follows for achieving this.


・ Make preparations through international cooperation so that quick response commands can be sent out by DSN and ESA ground stations.


・We establish our own overseas ground stations (for example, one such station in South America)

In the case of 20th scientific satellite (Hayabusa) launched in 2003, in this regard, we did request such support from DSN for emergency measures.


- 28 -

However, we must point out that even in the case of visible operation we still have this issue of time lag in deep space operation and that for this reason we will still have to rely on autonomous operation despite the risks inherent in this kind of satellite operation.

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page 29


２．About the measures to be taken in future with comms. and thermal control systems

With the mulfunction in comms. and thermal control systems it is difficult to pinpoint single causes of the mulfunction given that as mentioned before it is difficult to be specific about the causality with solar flares and also that thorough measures had been taken with respect to the possibility of discharging in the design of Nozomi.

For this reason the remaining possible mulfunction candidate is the accidental failure of components and measures to be taken in this regard can be as follows. For your information other measures to be taken with respect to the cause candidates which are extremely unlikely to be true given the thorough preparations that had gone into Nozomi's flight are seperately listed on the table IV-2-1.

(1) Responding by seperating out failures

With these interplanetory missions such as one attempted by Nozomi we can gain precious engineering and physical knowledge from the mere fact that the mission existed at all in outer space. For this very reason we should endevour to make sure that we produce a design which will allow continuation of the mission even if some of the on-board devices develop mulfunctions.

With this in mind we then must reduce the number of mulfunction possibilities and even in the event of mulfunction our design should be able to localise its effects and prevent its ripple effect eating into the more superior systems by placing more emphasis on seperating out failure causes.

In this regard in the case of Nozomi we must admit that not enough care had been taken to prevent an initial mulfunction of the component(s) which occurred on the secondary side of CI-PSU system from spreading into the more superior CI-PSU. For your information the method and characteristics of failure separation and some examples are listed on the tables IV-2-2 and IV-2-3.

(2) Development of components which will not cause latch-ups

With latest components it is becoming increasingly difficult to insert protective registers since their current amplitudes are large and allowable voltage bands are narrow. Also, even with those components which have small possibility of latching-up we are still talking about probability of mulfunction and there is no way we can say that they will not fail.



For this reason developments are underway for semiconductor devices such as "silicon on insulator" (SOI).

- 29 -

We are aware that these devices are being developped for consumer uses. However, we should also try and carry out further research so that these may be used as space flight components.

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page 30

３．Reflecting into the design philosophy of scientific satellites

About the two accidents that befell on Nozomi and the causes of mulfunction and measures to be taken for future we have been talking about fuel supply and comms./thermal control systems respectively. Scientific satellites come in different shapes and internal structures depending on mission requirements.

It is therefore important that we should reflect the measures we have discussed as a result of Nozomi's failure and particularly those ones that are commonly applicable to future scientific satellites into their design philosophy and development.

What follows are those items which we think should be incorporated into the design philosophy of future scientific satellites.


(1) Alterations to exisitng design

Whenever we try to make changes in the desing of the components to go on board we should remind ourselves that this will carry the same degree of risks as in designing them from scratch even if these changes are to be made on those that have enough track records.

For this reason it should become our design philosophy to examine the risks involved in changes made to proven designs and if we had to we should be taking every possible caution in determining the neccesity of desin alteration and possible repurcussions/verification methods etc. by calling for specialist advice from a wide range of desciplines.


(2) Ground tests

Naturally, with components and instruments to go on board any scientific satellites it is imperative that they are highly reliable with enough proven records. However, this is not always easy as mission contents and development times are all different and we may not always be able to fulfill these requirements.

Therefore, it is important that pre-flight ground tests should amply verify and evaluate the reliability/functions/capacities of those components going on board. Furthere more, we should make sure that an exccessive loading is not placed on the pieces to be tested and that the test contents are sufficiently approapriate and effective (including the influence of operational environment) for the purpose of verification leading to reliable data by inviting professional advice from specialists both within and outside our organization.

In the event of ground tests not leading to convincing reliability we should take a renewed look into the desing steps of to-be-on board components and insturments with a view to completely returning to design board.

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page 31


(3) Dealing with imports

There are cases where we have to procure imported items for space use with scientific satellites such as their components and devices because they are not produced in this country. With these imported items even if information on verification tests and flight records against our specs can be obtained from manufacturers it may not always mean that detailed internal structural information is available.

In these cases of limited availability of information on internal structure of devices and their electrical interfaces we must ensure that not only verification tests from user's point of view are carried out but also we must give utmost priority to the issue of seperating out failure causes.

Furthermore, we must, in the event of import troubles, make sure to cooperate with the manufacturers to clarify the status with a view to coming up with approapriate measures for rectifying the situation not just by oursleves but also with the cooperation of all organisations involved.

(4 ) Seperating out failure causes

Space systems cannot exclude the possibility of mulfunctions completely. Nor can we expect to be able to carry out repairt works in orbit. For these reasons we mus ensure that partial mulfunction by onboard devices will not lead to a total mission loss. We lay out points to note in trying to seperate out failure causes as follows.


① About the degree of seriousness of mulfunction in hand we must be able to evaluate the seriousness of possible repercussions to the system as a whole by making use of evaluation methods such as Failure Mode Effects Analysis (?, P)(FMEA) and Failure Mode influence fatality analysis (FMECA) so that the trouble in hand will not spread to other important systems by giving preferential priority to the failure cause seperation.

② Based on the priority judged by ① above we must make selective (given seriousness, repercussions and importance of the troubles) judgement on the possibility of mulfunction seperation and containment such as adoption of redundancy, seperation of power sources, building in of protective resisters etc.

③ With those items whose functional loss will not lead to serious faitality or those which are used only at the time of launch etc., that is to say, those items whose mulfunctions will not spread into secondary fatalities and therefore do not need our utmost attention we must select as much as possible the least troublesome failure cause seperation measures (such as the use of protective registers and installation of switches etc.).

(5) Trouble shooting by software

In the case of Nozomi our operation continued even after the accident in 2002. In fact, our continued operation of Nozomi lasted for 5 years from the launch in July 1999 to December 2003. This was made possible by several factors such as an improvement made on autonomous function and re-writing of data handling unit (DHU) software etc.

- 31 -

Being able to re-write the software after launch from the ground is an extremely effective measure to deal with various troublesome situations as we can pliably deal with different events by making the on-board devices carry out different functions. Naturally, similar capability has been adopted for the scientfic satellites which are in the pipeline and it is thought that changes in hardware functions by software re-writing will be popular from now on.

On the other hand we should note that use of unverified software is very dangerous. Therefore, it is imperative that we conduct sufficient ground tests with each of the software functions in order to establish reliability. Also, in the event that on-board software re-writing is deemed paramout we must first thoroughly check the safety of re-writing with a flight model and its electrically equivalent functional devices. Actual re-writing will have to be made based on the result of these ground tests.

end of page 31

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above for ease of reference

page 32

(6) Measures for deep space exploratory mission


In addition to above measures we may add a few other measures as follows, which may turn out to be effective with future Nozomi like deep space missions.


・Development of low bit rate comunication function for emergency cases

・Development of a system which is further improved on the system with new functions as follows used for the 20th scientific satellite (Hayabusa) for verification in deep space

Hayabusa was launched in 2003 as a deep space mission after Nozomi (18th scientific satellite). Hayabusa has following functions in addition to those available to Nozomi.


・Report packeting function

Function whereby results of autonomous actions are sent out each time by dedicated packets. If labour saving is wanted in operation then operation can be achieved only by report packeting.


・Choice of sampling rate at reproducing HK telemetry

Function whereby coarse transmission rate is used for coarse data scanning to be followed up later with only the required portion of data by higher sampling rate.

・System timer function

Function whereby commands are executed after a specified amount of time lapse, timing function for general purposes.

The measures intended for future as described by this current report as a result of Nozomi's failure are issues specific to all scientific satellites designed for deep space mission.

- 32 -

However, these are also issues commonly relating to space use devices and system development. Therefore, it is the wish of our organisation that these measures will be fully utilised for future space activities with increased reliability.

end of page 32

(I have not looked at the portion immediately after this, but I suspect that this might be the end of this report proper, only to be followed by supporting materials (appendices), which I will also translate, P)

above for ease of reference

page 34 (all pages before page 34 except those already translated are somehow missing from this report, P)

fig. 1-1-1 Nozomi final shape upon extending everything

page 35

Table Ⅰ－１－１ List of instruments on board (1/2)

Scientific observations achieved during cruising period

Instrument name:: Result
[collaborators ]

ＭＩＣ：Mars imaging camera:: First view of the other side of the Moon by a Japanese camera etc.
[Kobe Univ,ISAS/JAXA etc. and CNRS]

ＵＶＳ：Ultra violet imaging camera:: observation of interstellar winds outside solar system etc.
[Tohoku Univ, National Inst. of Polar Research etc. ]

ＸＵＶ：Extreme ultraviolet imaging camera:: 1st imaging of earth plasma sphere etc.
[Nagoya Univ, Boston Univ, etc.]

ＭＤＣ：Dusts conter:: detection of interstellar dusts etc.
[Munich Inst. of Tech, Tokyo Univ, ISAS/JAXA･LFM･MPIK･STMS, ESA etc.]

ＥＩＳ：High energy particle counter:: observation of solar flares etc.
[Tamagawa Univ, ISAS/JAXA, MPIA etc.]

ＥＳＡ：Electron energy analyser:: observation of lunar wake (unsure, P) etc.
[Kyoto Univ, ISAS/JAXA etc.]

ＩＳＡ：Ion energy analyser:: observation of interstellar winds etc.
[ISAS/JAXA etc.]

ＩＭＩ：Ion mass analyser:: long term monitoring of solar winds etc.
[IRF･Rikkyo Univ, ISAS/JAXA, etc.]

ＭＧＦ：Magnetic fields measuerer:: long term monitoring of soalr winds etc.
[ISAS/JAXA･Nagoya Univ, NASA/GSFC etc.]

ＲＳ：scientific observation of EM waves:: observation of solar corona structure etc.
[ISAS/JAXA]

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above for ease of reference

page 36


Table Ⅰ－１－１ List of Nozomi insturments （２／２）

Instrument:: result expected to have been obtained after insertion into Mars circular orbit
[participants]

ＰＷＳ： Plasma oscillation sounder:: radar probing of ionosphere etc.
[Tohoku Univ. ISAS/JAXA etc.]

ＬＦＡ：Low frequency wave observer:: observation of perturbation in ionosphere etc.
[Kyoto Univ RASC, Toyama Prefectural Univ., ISAS/JAXA etc.]

ＰＥＴ： Electron probe temperature:: First imaging of plasma regio0n of Earth etc.
[ISAS/JAXA, Michigan Univ, MPIA, a Korean inst. etc.]

ＮＭＳ： Neutral particle mass analyser:: detection of interstellar dusts etc.
[NASA/GSFC, Michigan Univ, Arizona Univ, Univ. of Hawaii,ISAS/JAXA etc]

ＴＰＡ： Thermal plasma analyser:: Observation of constituents of ionosphere etc.
[Univ. of Calgary, ISAS/JAXA･NRC･CSA, Univ. of Victoria, Univ. of Western Pntario, Univ. of Alberta, etc.]

end of page 36

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above for ease of reference

I had a quick look at the rest of this report and am satisfied that it does relate directly to failure causes. Therefore, I will continue with my translation, but there are a very few pages (I suspect 2 or 3such pages) which are irrelevant. One such is as follows as page 37. I am translating this page for completeness sake.

page 37

Table Ⅰ―１－２ List of engineering objectives achieved with Nozomi

Engineering objective:: Outline (as follows)

mission analysis:: coming up with an optimum mission scenario by trading-off, given limited resources and time, all those engineering options available

orbit planning:: ability to design orbits peculiar to planetary mission such as swingby techniques with the Moon and the Earth

high precision orbit determination:: by waves from the ground obtain velocity and line of vision distance data to be fed into precision dynamical modelling in order to determine the deep space probe's orbit with high accuracy


autonomous operation:: AI technique for letting the onboard computer make judgements

ultra long distance communication:: communication equipment and operational knowhow for the long distance (max. 4 times 10 to the power of 8 km) communication

weight reduction of the onboard instruments:: reducing the weight of all onboard devices including propulsion, solar batteries, antenna, batteries, electronics given that deep space probing requires considerably more energy at launch

ground support software:: ground software with AI capability given that safe operation of the deep space probe requires operation under complex constraints including a long term cruising phase


end of page 37

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above for ease of reference

page 38

fig I-1-2

instrument location on board

page 39

Table Ⅰ―２－１ Targets which became possible as a result of M-V launcher

Research targets

１） internal structure of planets

Observation of earthqua： Lunar penetrator
→Lunar-A (1997→2004) (development of penetrator)

２）pristine astonomical bodies

sample return from near earth asteroid
→MUSES-C「Hayabusa」(launch May 2003)

３） planetory environemtn

Venus/Mars：PLANET-B（「Nozomi」1996→1998）（Ｍ－Ｖ development）

Table Ⅰ―２－２ other Mars probing missions in plan at the time of Nozomi concept

Launch year : satellite name: planned by

1988: Phobos1 & Phobos2: former Soviet Union
1992: Mars Observer: US
1996: Mars96: former Soviet Union

Table Ⅰ―２－３ Weight reduction and some examples

target: result

adoption of nickel/hydrogen batteries: ２ｋｇ

CFRP treatment of GHe tank: ６ｋｇ

development of high gain light weight antenna: ４ｋｇ

adoption of new connectors for wiring within common use devices: ３ｋｇ

Weight reduction of S band receiver: １．５ｋｇ

adoption of dispersed power source (comparison with centralised system）: ３．６ｋｇ

end of page 39

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above for ease of reference

page 40

fig. I-3-1

New orbital plan for Nozomi

(here, I will have to explain, rather than translate as follows)

red: Mars
green: Earth
purple: Nozomi

(on this Earth-Mars system main character strings clockwise are)

1. 1st earth swingby: Dec 2002
2. leaving Earth gravitational field: Mars Dec 1998
3. insertion into Mars orbit: end 2003 to early 2004
4. 2nd earth swingby: June 2003

(on this page there is another Earth-Mars system in a square. The Monn is in yellows and Mars is depicted in red, with a short caption near Mars which says "into Mars orbit: Oct 1999)

end of page 40

page 41

fig I-3-2 Temp changes at temp measurement point(s?) by 1 bit communication

(Here, I have no idea as to what WANT is, but anyway)

Vertical axis is WANT temp and horizontal is the dates, from 23 July to 1 October. On the upper lef a box contains 5 character strings from top to bottom. The last one at bottom says "fitting of WANT-1".

Similarly, there are 5 character strings all pointing to the slanted line on the graph and the last one at bottom of this group of 5 character strings says "fitting of WANT-1"

The cpation of the arrow pointing vertically upward to the slanted line says "melting point under hydrazine tank pressure (2.2 deg C)"

There is another character string qualifying the same slanted line and this says "temp. increase rate per day of 0.32 deg C"

end of page 41

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above for ease of reference

page 42

(main feature of this page is a figure and a table)

Fig. II-1-1 Nozomi propulsion system outline

Most of this figure is in English. There are 5 oblong squares in dark blue. 3 of them to the left and 2 of them to the right.

Left ones from top to bottom are:

Pusher gas tank He
Fuel tank Hydrazine
Auxilliary engine RCS (mono propellant)

Right side one from top to bottm are:

NTO tank
Main engine (dual propellant) and the caption for this says "To be used for insertion into Mars transfer orbit"

In addition, there are 5 squares in light blue: They indicate the status of the valves. In this diagram all of these 5 squares carry "OPEN".

In further addition there are one red oblong square and the characters in there say: LV2: CLOSE->OPEN->CLOSE and one pale blue oblong square and the characters in there say: LV5,6: CLOSE->OPEN->CLOSE

There are two circles in this figure and one near the top and the caption for that says "Valve which developped mulfunction" and the other one near the bottom is OME.

There is a note also in the figure and it says "Valve status is all during TMI"

The qualifiers in the figure are as follows.

HLV: high pressure gas system latching valve
LV1-2, LVm: low pressure gas system latching valve
LV3-6: liquid system latching valve
RG: regulator
CV1-2: check valve
F1-3: filter
P1-4: pressure sensor

FDV1-5: inlet/outlet valve
TP1-5: test port
A&T1-4, R1&2: RCS thruster
OME: OME
FTNK-A&B: hydrazine tank
OTNK-A&B: NTO tank
GTNK: He tank

Table II-1-1 Main specs for propulsion system

Here, I will translate column by column, from left to right and from top to bottom. Sometimes, I will add row numbers so that we know where exactly we are down the columns.

Leftmost column and from top to bottom:

pressure
thruster specific impulse
thruster specific impulse
torque (9 and 10 are row numbers)
gas container or gas reservoir?
fuel tank (x2)
oxidiser tank (x2)
onborad propellant amount (20 and 21 Ditto)
effective propellant amount
propulsion system DRY mass
propulsion system WET mass

Now, second column from top to bottom:

initial pressure in the primary pressure system
initial pressure in the secondary pressure system
axial (x4)
radial (x2)
tangential (x4)
OME (x1)
RCS
OME

3rd column is as follows:

precession
spin
inner volume
MDP
destruction safety rate
inner volume
MDP
destruction safety rate
fuel (hydrazine)
oxidiser (NTO)
for RCS
for OME

the next column is nominal specs and the contents are from top to bottom (and I omit those in English):

2.3N per base
2.3N per base
2.3N per base

27.8 litter

98 litter per tank

40 litter per tank

and the last column is meant for notes and the contents are:

(=24.5 MPa)
(=1.37 MPa)

at time of pulse is given
lower or lowest limit of somekind of difference
R=1m at time of two in action
R=1m at time of two in action
CFRP/titanium alloy sphere
tatanium alloy tear drop shaped (=2.01 MPa)
titanium alloy tear drop shaped

end of page 42

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above for ease of reference

page 42

fig. II-1-2 Status at the time of failure in December 1999 (time is UTC)

(Here, most of the time line is in English. I simply suppllement those with captions in Japanese)

07:17 Start of attitude change (sun angle 60 degrees)

07:33 attitude control completed (sun angle 120 degrees)

07:41 SPIN-Up (->25 RPM) time taken is 1 minute

08:06:20 to 08:13:04 OME fired

08:32 SPIN-Down (->10 RPM) time taken is 1 minute

08:40 Start of attitude change (sun angle 120 degrees)

09:00 Attitude control completed (sun angle 43 degrees)

(character string inside this large box says):

all monitored values and status were normal except that the monitored value of NTO upstream pressure started declining as the firing started. (confirmed by reproduced TLM)

(the 1st red arrow pointing at 12:00 says): News flash from JPL that there was an insufficiency (shortfall of delta V) of about 100 m/s against the planned value of 423.22 m/s

(2nd red arrow just past 14:00 says): similar insufficiency confirmed by onboard TLM integration

end of page 42

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above for ease of reference

page 44

fig. II-1-3 Telemetry data during visible period after insertion into Mars transfer orbit (TMI)

(here, the only thing that needs translation is the caption inside a Mexican hat like area and it says):

oxidiser tank pressure P4 is low

end of page 44

page 45

fig. II-1-4 Nozomi status at time of LV2 open command (part 1)

(here, there are 9 squares in yellow and they are, from top to bottom):

P1: pressure of gas storage
P2: regulator exit pressure
P3: fuel tank pressure
P4: oxidiser tank pressure

injector temp.
OME firing duration (or time?)
delta V
LV2 status
LV2 open

(There are 3 blocks in pink and they are):

(referring to P2): transient decrease in regulator pressure due to rapid supply of He gas to oxidiser tank which was judged to be lower than normal

(referring to P4): pressure rise as LV2 is opened -> He gas is being supplied

(last small square): pressure drop is judged to be due to imperfect opening of LV2

(there are 2 more small squares in pale blue and they are):

(top square): Open
(bottom square): Close

end of page 45

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above for ease of reference

page 46

fig. II-1-5 Nozomi status at time of LV2 open command (part 2)

(caption on the left says): Shock due to LV2 opening

(caption on the right says): Vibration due to He gas flow

end of page 46

page 47

fig. II-1-6 Status of OME firing at time of insertion into Mars transfer orbit (TMI)

(here, there are 11 pale yellows areas with captions and I will go from top to bottom, then bottm to right, in this order)

P1: pressure of storage tank
P2: pressure at regulator exit
P3: fuel tank pressure
P4: oxidiser tank pressure

injector temp.
OME firing time (or duration/)
delta V
status of LV2
LV2 open (this is the bottom and I will move right from here)

OME firng start
OME firing end/LV2 close

(now lines in red)

(dotted lines in red next to P4 says): originally planned regulator level
(red solid line next to delta V says): target delta V (423.33m/s)

(there are cations in 3 pink areas. they are, from top to bottom)

Not regulated like P2, P3 -> insufficient supply of He gas

insufficient amount of delta V

status changes in response to OPEN/CLOSE command

(what remains are 2 pale blue areas and they are from top to bottom)

OPEN (top)
CLOSE (bottom)

end of page 47

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