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Nozomi in perspective, Revisiting the causes of failure
pandaneko
post Nov 16 2011, 09:53 AM
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QUOTE (pandaneko @ Nov 16 2011, 06:49 PM) *
above for ease of reference

page 22

2.Estimated causes of mulfunction


- 23 -

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)

end of page 23

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

Pandaneko
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pandaneko
post Nov 17 2011, 10:27 AM
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QUOTE (pandaneko @ Oct 23 2011, 06:12 PM) *


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.


cool.gif 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

end of page 23

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pandaneko
post Nov 18 2011, 09:25 AM
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QUOTE (pandaneko @ Oct 23 2011, 06:12 PM) *


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.

end of page 24

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pandaneko
post Nov 19 2011, 09:46 AM
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QUOTE (pandaneko @ Oct 23 2011, 06:12 PM) *


above for ease of reference

page 25

cool.gif 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.

- 25 -

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

end of page 25

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pandaneko
post Nov 20 2011, 09:22 AM
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QUOTE (pandaneko @ Oct 23 2011, 06:12 PM) *


above for ease of reference

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.

end of page 26

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pandaneko
post Nov 21 2011, 09:35 AM
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QUOTE (pandaneko @ Oct 23 2011, 06:12 PM) *


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.


end of page 27

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pandaneko
post Nov 22 2011, 09:12 AM
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QUOTE (pandaneko @ Oct 23 2011, 06:12 PM) *


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.

end of page 28

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


above for ease of reference

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.

end of page 29

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pandaneko
post Nov 24 2011, 09:39 AM
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QUOTE (pandaneko @ Oct 23 2011, 06:12 PM) *


above for ease of reference

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.

end of page 30

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pandaneko
post Nov 25 2011, 09:41 AM
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QUOTE (pandaneko @ Oct 23 2011, 06:12 PM) *


above for ease of reference

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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pandaneko
post Nov 26 2011, 09:42 AM
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QUOTE (pandaneko @ Oct 23 2011, 06:12 PM) *


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


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]

end of page 35

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pandaneko
post Nov 28 2011, 08:44 AM
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QUOTE (pandaneko @ Oct 23 2011, 06:12 PM) *


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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pandaneko
post Nov 29 2011, 09:05 AM
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QUOTE (pandaneko @ Oct 23 2011, 06:12 PM) *


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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pandaneko
post Nov 30 2011, 09:04 AM
Post #45


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QUOTE (pandaneko @ Oct 23 2011, 06:12 PM) *


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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