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Nozomi in perspective, Revisiting the causes of failure
pandaneko
post Oct 23 2011, 09:12 AM
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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
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pandaneko
post Oct 24 2011, 08:53 AM
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QUOTE (pandaneko @ Oct 23 2011, 06:12 PM) *


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)

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pandaneko
post Oct 24 2011, 09:37 AM
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QUOTE (pandaneko @ Oct 23 2011, 06:12 PM) *


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
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Phil Stooke
post Oct 24 2011, 11:04 AM
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Thanks for doing this. It is very interesting.

Phil Stooke



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nprev
post Oct 25 2011, 01:51 AM
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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.


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pandaneko
post Oct 25 2011, 08:58 AM
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QUOTE (pandaneko @ Oct 23 2011, 06:12 PM) *


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
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pandaneko
post Oct 26 2011, 09:26 AM
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QUOTE (pandaneko @ Oct 23 2011, 06:12 PM) *


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

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pandaneko
post Oct 27 2011, 09:09 AM
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QUOTE (pandaneko @ Oct 23 2011, 06:12 PM) *


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)

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PaulM
post Oct 27 2011, 11:52 AM
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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.
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Paolo
post Oct 27 2011, 12:09 PM
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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... wink.gif


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pandaneko
post Oct 28 2011, 09:43 AM
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QUOTE (pandaneko @ Oct 23 2011, 06:12 PM) *


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
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pandaneko
post Oct 29 2011, 09:39 AM
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QUOTE (pandaneko @ Oct 23 2011, 06:12 PM) *


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.

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pandaneko
post Oct 30 2011, 09:03 AM
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QUOTE (pandaneko @ Oct 23 2011, 06:12 PM) *


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.

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pandaneko
post Oct 31 2011, 09:38 AM
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QUOTE (pandaneko @ Oct 23 2011, 06:12 PM) *


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)

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pandaneko
post Nov 1 2011, 09:19 AM
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QUOTE (pandaneko @ Oct 23 2011, 06:12 PM) *


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.


cool.gif 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

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