(page 18)

2.2.6 Other satellites (continuation)

Outline of NOZOMI's failure

NOZOMI was similar to Akatsuki in that propulsive substance was regulated and during OME firing both fuel tank pressure (P3) and oxidant tank pressure (P4) were supposed to maintain 1.4 MPa or thereabout.

However, when the mulfunctionhappened during OME phase P3 was maintained, but P4 kept falling. As a result propulsion and acceleration fell lower than anticipated. We concluded that CV-O caused valve closure (see following schematics and graph)

(Here, there is a small graph showing P3 and P4 values during the 600 seconds of OME burn. Blue dotted line is P4 and red solid line is P3, which is constant at about 1.4 MPa.

To the right of the graph is a schematic of the piping system layout with valves, tanks, and pressure meters etc. which reflects ; )

Design philosophy of NOZOMI propulsion system

Non reversibles valves (CV-O and CV-F) and other valves are placed in upstreams of the fuel and oxidant tanks so that accidental mixing of fuel and oxidant vapours cannot lead to an explosive pressure rise

(end of page 18)

(page 19)

(There are 2 schematics on this page, Nozomi's piping system and Akatsuki's piping system. They look simple enough, but I will have difficulties with the latter as you will see in my translation. Anybody who can help?,P)

2.2.6 Othew satellites (continuation)

Relationship between mulfunction of Akatsuki CV-F and Nozomi mulfunction

About the possibility of salt formation leading to valve closure on Nozomi it is indeed thought that salt was formed inside the piping system of Nozomi. However, salt forms only in the valve on the fuel side, so, salt formation on Nozomi does not have a direct influence on Nozomi mulfunction.

Improvements made with Akatsuki based on the lessons learnt from Nozomi mulfunction

• Simple redundancies in a series isabolished as much as possible as this may lead to flow channell closures. Put concretely, we placed stop valves in pararell on the oxidant side. Also, we placed, inside the fuel tank, a liquid/gas seperation membrane, thereby eliminating the need for stop valves.

• We selected a different type of stop valves

• (Here, I am at a loss, P. Translation is) The position of the stop valves was changed to the upstream side of non-reversible valve (of fuel?, or oxidant?, P) so that less effects of oxidant vapour are felt.

• Also, with Akatsuki the regulator valves were also made redundant (redundant layout was adopted is meant, P) so as to increase reliability compared with Nozomi

(end of page 19)

(since I managed to translate with • without replacing these with numbers here I should perhaps explain what I did on several occasions in the past translations invloving •. When there are •ed statements separated by somet other symbols and I try to copy and paste them they are regrouped despite the fact that they belonged in different places and I have to sort them out from the handwritten memo. In these cases I have often replaced • with numbers of my choice.

For example, if I am translating section 2, subsection 2.1, 2.2, 2.3 etc etc and encounter some •s I simply replace them with numbers without paying any attention at all to all previous section (and sub, subsub) numbers. Net result ise that these numbers appearing in any one of the pages that I translate have no direct link with the numbers on the contents list.

So, you encounter suspicious numberings please consult the original pdf. I have no difficulty if •ed statements are grouped together in one continuous block as is the case with this page, P)

P, just a note to say thanks for this tremendous effort to keep us informed....thanks!

Very impressed with the depth of the analysis as seen thus far. This is the serendipitous (i.e., unintended yet highly beneficial in other ways) product of the event: understanding malfunctions is key to not repeating them in future efforts. I hoped this would be so, and clearly it's happening.

I grow more optimistic that they will achieve Venus orbit in 2016!

I second that!
Thanks Pandaneko for all your translation work.
This is a huge effort on your part and makes fascinating reading for the rest of us.

could it be "orders of magnitude"? in which case it would mean hundred times larger

Ah..., what they (and I in translating) mean is something like;

2536.945 as opposed to 25.36945. Do you see what I mean? How do you describe this difference?

and anway, I should be able to cover two pages this evening right after this

P

(page 20)

2.2.7 Policy for designing propulsive systems with future missions

With future missions using 2 liquid engines we will have to take into account preventing the mixing and reaction of fuel and oxidant vapours inside the gas supply pipings


Measures to be taken to prevent mulfunctions due to fuel/oxidant vapour mixing (proposals)

• With respect to propellant vapour flow we will consider both leak and transmission effects

• With particular respect to oxidant vapour flow we will need to quantise the phenomenon making use of real liquid and real valves

• Taking into consideration the possibility of mulfunctions due to salt formation we will have to choose approapriate number, positions, characteristics, and piping layout of the valves to be used against the envisaged operational duration

(end of page 20)

(page 21)

2.3 Looking at the possibility of contaminants sitting in the gaps

Out of the six items shown in section 2.1 our investigation has shown what follows below with respect to the item relating to contamination

Contamination sources are itemised as follows and evaluation is added. Overall, we believe that mulfunctions due to contaminants sitting in the gaps are small in possibility

（１）Contaminants that existed there right from the time of manufacturing

By checking the cleanliness test records we do not think they led to CV-F closure

（２）Contaminants that sneaked into the system while being serviced on earth

This is not possible as liquids are accessed through filters and cannot lead to CV-F closure

（３） Contaminants produced at valving operation


Possible sources are the regulating valves in the upstream of CV-F, but their outlets are also filtered and cannot produce contaminants leading to CV-F closure.

However, we cannot rule out, by nature, the possibility of contaminants sitting in the gaps with absolute certainty as they can be created accidentally.

For these reasons we will be paying particular attention to the issue of filtering during the design stages of our future missions. We wil also be looking at more improved pipe line cleansing methods in order to reduce the possibility of grits sitting in the system.

(end of page 21)

(page 22)

2.4 Summary of the investigation into CV-F mulfunctions

We have looked at the 16 candidates filtered out as the causes of CV-F closure and what we were left with are salt formation and contaminant grits as the possibility. From our efforts in clarifying these two items we obtained following understandings.

• About the amount of oxidant vapour passing across the valves both in upstream and downstream directions we now know that the leak model where a standard gas moves across an equivalent orifice is utterly insufficient and that we will need to take into consideration the effects of transmission across the valve sealing material.

• When fuel and oxidant vapours mix within the gas supply piping system salts are produced and this may lead to valve closures/mulfunctions.

• With these understandings above we investigated propellant flow rates with other satellites and came up with measures to be taken with future missions with silimar propulsive systems.

• We will continue to grapple with the issue of contaminants as they cannot be purged completely out of the system.

(end of page 22)

(from tommorrow on we will be moving into sections 3 (OME) and 4 (orbital re-insertion attempt, I think we have covered about half of the report by now, P)

"orders of magnitude" should be good

Thanks, it was just that the character they used was a little different from their usual expression. P

(page 23)


3. Influences suffered by OME as a result of CV-F closure

3.1 Influence candidates relating to OME mulfunction

(D1 to D5)

We estimated (section A.2a) through FTA that CV-F closure led to the mixing ratio of fuel and oxidant exceeding the design values and put OME into (one or more of, I think, P) following situations during VOI-1.
．
D-1) Thruster nozzle breaking due to excessive heat flow

D-2) Thruster nozzle breaking due to film cleaning jet direction anomally

D-3) Rear portion of throat burnout

D-4) Unstable burn

D-5) Injector thrust direction anomally

We conducted what follows, mainly through experiments, in order to narrow down on them.


1. Investigation into anomallies D1 and D5 (section 3.2) due to propellant thrust direction by carrying out injector thrust status confirmation tests

2 Investigation into anomallies D1, D3, and D5 (section 3.3) due to behaviours exceeding the design limits by carrying out burn tests

(end of page 23)

(page 24)

3.2 Results of injector thrust status confirmation tests


During VOI-1 phase the fuel supply pressure went down alarmingly compared with the design values. As a result, there is a possibility that this injector thrust status change led to burn anomally and/or insufficient cooling.

We therefore conducted investigation (experimental) tests particularly because we could not deny the possibility of film cooling (direction, I think, P) and injector thrust direction anomally under reduced fuel supply pressure (corresponding to D2 and D5).

(immediately after this there is one schematic (left) and one picture (right) and I find the schematic rather difficult to understand, but I will have a go at explaining the schematic. Picture is simple)

(I think this schematic is about the exit end of the injector and the inlet end is replaced schematically by arrows (red for oxidant and blue for fuel) coming in from above. The vertical broken lines below the yellow rectangular block are the burn chamber walls.

The character string right in the middle of the chamber reads: "core: burning of fuel and oxidant mixture".

Oxidant is injected into the direction of 5 o'clock and fuel 7 o'clock. Since some of the fuel arrow lines intersect with the dotted lines these are mean to be the film cooling, I think. Here, I am talking about the two lines going out of the chamber.

There are two more lines well inside the chamber, long lines and they are not pararell with the outgoing lines, indicating perhaps the result of momenta of oxidant and fuel colliding inside the injector inlet and producing resultant thrust directions. They call the angle between these lines and the vertical "fan angle", whatever that means. The lines not pararell with the outgoing lines are called "fan". It is not fun since I barely comprehend this schematic...

On the other hand the picture is OK. It is the photo of the injector status confirmation tests. So it says, anyway.

I have in above verbally(?) explained all of the characters with the schematic, but am keeping them here below as they are in the original schematic for easy reference in case anybody might be interested.)

燃料酸化剤酸化剤燃料
ｲﾝｼﾞｪｸﾀ
フィルムクーリング
(FC：燃焼器壁面の冷却)
コア：
燃料と酸化剤の混合による燃焼
燃焼器
壁面
ファン角
ファン噴射イメージ

Tests confirmed that during VOI-1 phase both film cleaning thrust direction angle and core propellant thrust direction angle (that angle after the propellant collision) were within the designed values and showed no significant anomally. Therefore, we concluded that the candidates D2 and D5 are innocent.

D2: thruster nozzle braek due to film cleaning thrust direction anomally
D5: Injector thrust direction anomally;

(end of page 24)

(page 25)

(I find this page awfully difficult to translate, P)

3.3 Burn tests recreating VOI-1

We conducted ground burn tests, by recreating the propellant (fuel and oxidant) supply system status to verify D-1，D-3，and D-4.

(immediately after this there is a graph and a photo (simple photo of ground burn tests), followed by Note;)(about the graph seperately down below translation, P)

Note: during VOI-1:
• Oxidant supply pressure was regulated to remain more or less constant
• Fuel supply pressure gradually decreased, due to the closure of CV-F, as burn time increased
→Oxidant/fuel mixing ratio (O/F)increased as burn time increased

(end of page 25)

(about the graph) (since I barely understand the graph itself let me translate those character strings inside and outside the graph itslef first so that our experts can make comments for the rest of us )

1. bottom horizontal scale is the fuel tank pressure in MPa and
2. vertical scale on left is the oxidant tank pressure in MPa
3. top horizontal scale is oxidant/fuel ratio

(about other lines and markings/signs/symbols just outside, to the right of the graph)

4. dark solid line is in-orbit burn during VOI-1
5. blue solid line is burn tests of VOI-1 (part 1)
6. red dotted line is burn tests of VOI-1 (part 2)
7. the square is the flight plan's operational range
8. the large blue dot is the condition(s) for the breaking of the burner during part 1 burn tests

Now, about the middle of the graph there is an arrow made up of dotted lines and the caption for this left pointing arrow reads:

"condition(s) move to left as burn time passes"

Also on the graph itself those lines slanted indicate propulsive power

My simple questions here are;

what do they mean by "burner"? Did it break during the tests?, what do they mean by "condition(s)?, etc etc. and in case I lose this translation I have been looking only at my sketch of the graph for the last half an hour and I am still uncertain as to exactly what they did... I may find some of the answer on the next page and I will move on. P

(page 26)

(before I move on I missed something out in page 25. There is a small dark sqaure on the graph and the caption for that is "at time of VOI-1 anomally" and within the square on the graph the righthand end of the blue solid line is indicated as "burn start condition(s) I am more and more confused...)

3.3.1 Burn tests (part 1): where the burner was damaged

We conducted ground tests to check on D1, D3, and D5 by recreating the VOI-1 phase burn (and what follows are the results)

(There are two graphs and one picture immediately after above statement. I will come back to the graphs after translating the remaining statements as follows, P)

graph 1, picture, graph 2

• the burner was damaged when the rapid temp rise immediately after ignition came to and end and the temp was gradually increasing in quasi static manner reflecting changes in O/F ratio due to decreased supply of fuel

• after the burn tests we confirmed the starting point of the breakage on the broken surface area and we did not observe traces of any clear material deffects in the vicinity of this starting position, indicating that this breakage was not due to a simple manufacturing error (section C.2)

These tests showed us that an increased thermal pressure (burden) as a result of going out of the designed operational range of burn conditions will lead to burner breakage

(end of page 26)

(about the graphs)

the 1st graph is the history of propulsive power and burn pressure before and after burner breaking. Horizontal axis is time in seconds and the vertical scale is the burn chamber pressure Pg in MPa (or Pc, character too small!), red line is the burn chamber pressure and blue line is the power

The upward arrow and caption says "power after breakage"

the 2nd graph is the burner's outer wall temp against time in seconds. To the left of the vertical dotted lines is the transient temp rise region and to the right a quasi static temp rise region in response to changes in O/F

measurement range is between 3200 and 2000 degrees C and horizontal scale is time and the vertical is the burner outer wall temp in degrees C

On the graph itself there is a dot which is captioned as "breaking point"

The photo caption is "burner after ground burn test (thruster nozzle is broken) (to me nothing seems to be broken, perhaps the nozzle is gone compeletely?)

I am actually no longer sure if I am doing translation or simply adding my questions/doubts and comments. Anyway, up you go! P

Thanks again, mighty Pandacat!

One question I'd love to be able to ask: If they go with the oxidizer dump/ACS thruster orbital insertion schema, is there any way to dump the oxidizer so that it could aid in the required delta-V? It probably wouldn't produce very much thrust, but if it could be done safely (i.e., no backsplatter on the arrays or instruments) and from a favorable attitude, might save s small amount of ACS fuel.

I am feeling a lot happier now with the graph on page 25. This is because of my new discovery that there are 3 lines starting at the same position within the designed conditions, that is the oxidant and fuel pressure combinations enclosed within the square.

Last night, I could not see that there were these three lines. I only saw an entangled mess. Now I know what happened in flight. Burn started right in the middle of the square with 500N. Oxidant pressure remained constant at 1.4 MPa, but fuel pressure kept falling right down to 0.8 MPa and that is the start of VOI-1 anomally. AND, we do not know what happened to the engine.

So, they did (at least, I think) 2 burn tests on ground. During one test, that is the blue line test, the engine broke when fuel reached 1.25 MPa anfd the photo is showing that.

However, they did another test, which is red, and despite the fact that the fuel pressure reached 0.8 MPa beyond the breaking point of 1.25 MPa for the blue, the engine did not break.

So, with all other factors being equal (I think) we still do not know if the engine in orbit broke or not

So, translation will have to continue, to discover JAXA's views on this.... P

I am afraid I do not know. In fact, I am not technically qualified to answer such difficult questions. The minimum I understand from your question is the you want to reduce the total mass to make the burn more effective. However, our local newspaper here did say that JAXA will try to ignite OME in early September. Dumping the oxdizer, will it not make the burn impossible to start?

My understanding of the burn tests in September is that OME and ACS (or RCS) burns are independent. I suspect, though, that we may gain more insight as translation continues.

P

(page 27)

(There is a simple graph here. More about it after translation of main texts)
燃焼開始
時間の流れ
燃焼試験（その2）での燃焼器外壁最高温度の履歴Burn test (part 2): where the burner did not get damaged

We conducted a burn test using another burner which is different from that used for part 1 burn test.

Looking into the difference between two different burn tests

We simulated VOI-1 and our views are:

• Two different burners were subjcted to tests and with one of them the thruster nozzle was damaged

• Whether a burner will get damaged or not if it goes outside the designed parameter range by how much just depends on the burner itself

• We estimate that the difference comes from these individual burner characteristics

• For your information the flight burner was checked for its health by AT test within the designed parameter range and was offered for actual flight. With this second burn test we noted followings candidates were not reproduced even if the burner was subjected to conditions beyond O/F=1.13 when the mulfunction appeared

• Burner was not damaged (D1 and D2)
• Burn efficiency decrease was not noted and there was no throat rear burn (D3)
• Burn status was stable (temp distribution burn pressure) and there was no abnormal burn inside the burner (D4 and D5)

From these observations we estimated that D3 and D4 are innocent.

(end of Page 27)

The graph is the history of the maximum temp of the external walls of the burner, which is the vertical scale in deg C. Horizontal is O/F. There are three mountain like curves, green, red, and blue and I think these are three separate measurements. They all peak at O/F=1.13. Time lapse is intrinsic on the graph, from left to right.

All in all, we still d not know if the OME is OK or not!!!

Or rather (as your subsequent translation suggests) the combustion chamber or nozzle fails at a
fuel pressure that is not completely deterministic. It is easy to imagine a stochastic element to the
failure mechanism in a fluidic system like this where flow might stick or not (think about pouring tea
from a bad teapot).

My estimate from the angular acceleration in the Akatsuki telemetry was that the thrust was effectively
diverted by 45 degrees, likely by half of the nozzle breaking off. But it is easy to imagine the nozzle
in the ground test failing slightly differently (perhaps the whole nozzle coming off)

Thanks again for your great work in translating!

(page 28)

3.4 Summary of FTA verification about the influences suffered by OME

We can obtain following conclusions through injector thrust status confirmation test, burn test (part 1), and burn test (part 2)

(above statement is followed by a table of judgements. At this point in time of translation I table elements lined up immediately after this and I might as well translate them before saying where these elements will fall into in the table)

D1: Thruster nozzle break due to excessive thermal flux

D2: Film cleaning injection direction anomally leading to thruster nozzle break

D3: Throat rear burn

D4: unstable burn

D5: injector thrust direction anomally

(actually, these above are the items to be judged upon and naturally come into the first row, from left to right)
(I must now explain the table structure. There are 5 rows and 6 columns. First column contains captions as follows, followed by marks about which I will translate soon. Each mark indicates judgement in that particular position of the table)

R2C1: Thrust status confirmation test: N/A, x, N/A, N/A, x

R3C1: burn test (part 1): ○, ○, N/A , N/A, N/A

R4C1: burn test (part 2): × note1, × note 1, ×, ×, ×

R5C1: Overall verdict: ◎, －, －, －, －

×： did not occur
N/A： no judgement basis available
◎： very likely
○： small possibility of occurence
－： low possibility

note 1: there is a possibility that brekage did not occur due to individual chracteristics

We can conclude through burn tests (part 1 and part 2) that D1 and D2 are candidates. However, D2 occurence possibility is deemed low via gievn the result of visually confirmable thrust status confirmation test.

In addition, we find that the time history of the changes in power after D1 phenomenon during burn test (part 1) is similar to the history obtained during orbit insertion attempt (assumed extent of burner breakage and how it matches the data obtained during insertion attempt is shown in sections C.3 and C.4)

Taking all these into consideration we conclude that the damage experienced by OME is very likely to be the thruster nozzle breakage (D1) due to an excessive heat flux

Given above conclusion we see that we will need to reignite the broken burner for the expected orbital insertion and we will discuss issues relating to this point in the next section.

(end of page 28)

Ralph, thank you for above! and also, NEPREV, an answer to your earlier question is coming up in one of the sub-sections later on. That will be given in the context of OME being considered useless.

(page 29)

4. Considerations for Venus orbital re-insertion

It is thought, from the post abortion data of the propulsive system and through considerations conducted in section 2, that a minute amount of gas is still being supplied inside the propulsive system's pipings despite the fact that CV-F is closed.

What follows is based on this assumption with a view to arriving at a viable means of operation for re-insertion. (We also are aware that closure status of CV-F may be even now varying due to the possibility of continued mixing of fuel and oxidiser as mentioned in section 2)


・ Looking into a new orbit plan given Akatsuki's current status (section 4.1)

・Although Akatsuki's OME is damaged at the thruster nozzle section we are looking at the following two points (section 4.2)

①Orbit change using the borken OME and its re-firing
②Orbital re-insertion using RCS without using the damaged OME

・With both ① and ② above we look into approapriate operational methods for the spacecraft and the propulsion system (section 4.3)

・Current Akatsuki's orbit means 0.6 AU as the nearest approach to the sun and it is consequently subjected to harsh thermal environment beyond the designed range of conditions and a solution is sought (section 4.4)

We now plan, based on further ground tests and analyses, to conduct OME test burn in September this year (2011) and also carry out orbit change in November as Akatsuki nears the sun

(end of page 29)

(page 30)

4.1 Orbit plan for re-insertion into Venus orbit

In what follows we show our current plan for orbital changes for re-insertion

From now on we will be looking at;

・ΔV scheme based on the result of in-orbit test firing
・constraints on attitudes at the time of maneuver (such as thermal constraints)

in order to arrive at a practical operational plan

Our tentative plan is to go for re-insertion in Novemver 2015 by reducing the total ΔV required through orbital changes during the second and third nearest approaches to the sun

(after this there is a ΔV schedule (almost like a table) and a graph of orbits (which is too small for me and in any event captions seem all in English, so I am not translating) (what I will do here is to turn the schedule into a table of my own, 4 rows and 5 columns)

(1st row contains timings and they are as follows)

R1C2: April 2011, R2C2: Novemver 2011, R3C2: June 2012, R4C2: Novemver 2015

(1st column contains case studies and they are)

R3C1: Case A, R4C1: Case B

(2nd row contains descrption of the timings just above them and they are)

R2C2: 1st nearest approach to sun, R2C3: 2nd nearest, R2C4: 3rd nearest, R2C5: reunion with Venus

(then delta V numbers for the 2 cases as follows)

R3C3: 243 m/s, R3C5: 322 m/s
R4C4: 294 m/s, R4C5: 281 m/s

(243 m/s and 294 m/s are enclosed in botted line boxes and the character string pointing to 243 m/s reads "maneuver at time of nearest sun approach for case A)

(after the delta V schedule and the orbit plan in the right hand box, the caption of which reads "example of orbits for case A" there are 3 notes as follows)

(note 1) ΔV is after impulse conversion and result of many-body analysis
より．
]
（note 2） ΔV in the Venus reunion box is that for a 4 day circular orbit

（note 3） test firings not entered in the table

(end of page 30)

(page 31)

4.2 Trade-off in Venus orbit re-insertion operation

We aim at confirming the current status of OME by firing it in orbit

We are currently looking at following two cases, depending on OME firing tests, one where OME can be refired and the other where OME cannot be refired

(after this there are boxes containing planned actions and expected results in two streams and the starting box in the middle is as follows and after this node box I will follow two seperate streams in translation)

０．OME test firing in orbit

We will test fire OME several times and try to confirm the health of OME and attitude maitaining ability during OME burn
Timing: September 2011
ΔV：1～20 [m/s]

ＯＭＥ：Orbit Main Engine ? of 500 N class
ＲＣＳ： Attitude Control System, 4 x 23 N class engines

Plan 1： We can use both OME and RCS

１．Manuever at closest approach to the sun
Timing: Nov 2011 or June 2012
ΔV：230~300[m/s]

２．Venus orbit re-insertion mauever (4 day orbiting)
Timing：Nov 2015 (reunion with Venus)
ΔV：280~360[m/s]

３．Re-insertion into observation orbit (from 4 day orbiting to 30 hour orbiting)
Timing：after Nov 2015 during closest approach to Venus
ΔV：200[m/s]

With this plan we should be able to guide Akatsuki into the planned observation orbit


Plan 2： We cannot use OME and we will have to achieve ΔV using RCS only

1．Oxidiser jettisoning tests and jettisoning
Timing： before nearest sun approach manuever

2．Nearest sun approach manuever
Timing：Nov 2011 or June 2012
ΔV：230~300[m/s]

3．Venus orbit re-insertion manuever：
Timing：Nov 2015 (reunion with Venus)
ΔV：280~310[m/s]

With this plan 2 ΔV is too small and we will not be able to guide Akatsuki into the planned observation orbit

(end of page 31)

(While translating this page I realised that I may have made a mistake on the last page, page 30 about the description of the orbit they are talking about. I remember that was a 4 day "circular" orbit, but the same expression appears here and I see that there is no "circular" as such.

I do not even remember if the originally planned orbit was meant to be circular at all. It may have been similar to that with Bepi Colombo, mildly eccentric elliptical? Perhaps, not too important as long as it goes around?, P)

(page 32)

4.3 Looking at Venus orbit re-insertion operation

We are currently trying to obtain information for engine operation in order to achieve orbital re-insertion as outlined in sections 4.1 and 4.2

A) Investigating how to reduce shocks on ignition (section C.5)

We used a broken nozzled engine and conducted re-ignition tests and saw that the breakage will propagate and in some cases might not withstand the impacts. Therefore, we are currently looking at ways to reduce re-ignition impacts, such as adjusting the timings of oxidiser and fuel injection. We are also looking into the effects of OME destruction on re-ignition and how that might affect the spacecraft

Blowdown operation in OME firing (section C.6) (I am annoyed by this fellow at the start of this subsection. He should be read as bracket B!)

We assume that during VOI-1 CV-F closure led to O/F ratio going outside the designed range of parameters. Therefore, we are currently looking into ways of maintaining the right ratio despite CV-F closure and conducting experiments to blow down fuel and oxidiser tanks.

C) Investigating into ways of jettisoning oxidiser (section C.7)

If we find OME firing impossible, we will then be forced to use RCS to achieve orbital re-insertion. However, since RCS engines are single liquid engines and given their low performance we will need to reduce craft mass as much as possible and we wish to throw away our oxidiser in orbit and we are conducting experiments to that end.

(end of page 32)

Thanks again Panda!

Any orbit achieved only by the RCS would be a huge ellipse and a long long way from the original design orbit yes? Are there any indications of what kind of orbit the RCS alone might achieve?

P

If they get it in an orbit, can't we dream about aerobraking?

I sorta doubt it. IIRC, all the aerobraking Mars orbiters had articulating solar arrays partially for that purpose with beefed-up joints. They were mechanically designed to take the stress, and the articulation meant that they didn't have to consume as much ACS propellent during the maneuvers.

I'm ashamed to admit that I don't recall Akatsuki's array architecture, but it probably wasn't designed to deal with the stress of aerobraking...to say nothing of the additional, presumably significant drain on RCS propellent if it was even attempted.

But I would be very happy to be wrong about all the above. It's a good thought, Steve.

I am afraid, so far, there is no mention of the new orbit details. They may say something towards the end of this report. P

(page 33)

4.4 Responding to high temp environment in the vicinity of nearest sun approach

Akatsuki's failure meant that Akatsuki is right now cruising in an orbit which was not thought of initially and the possibility is real of onboard devices being affected by high temperatures through deterioration of surface control materials (see graphs below) in the vicinity of the nearest approach to the sun.

Maintaining device temp below allowed levels will lead to some attitude constraints and we ;

• maintained the allowed temp environment, during the sun approach on 17 April 2011 by pointing the Z- axis of the craft to the sun
• plan to carry out nearest sun approach manuever by taking into consideration the minimum time in which we can maintain best craft attitude for firing

(after this there are two diagrams and they are simple enough and I will describe them after main area translation)

Above diagram shows the time line of thermal environment when nearest sun approach manuever is not carried out. Nearest sun approach manuever is an operation by which furthest from the sun altitude is lowered without changing the nearest sun approach altitude. For this reason even the execution of nearest sun approach manuever will not change the high temp side of temp environment significantly

About the deterioration of the temp control materials we are right now conducting tests, taking into account future nearest sun approaches.

(end of main text translation of page 33)

1. Thermal environment diagram (time line) if Venus orbit insertion had been a success

Lef vertical axis is the solar ray intensity in W/m squared and horizontal is the time line. There are horizontal reference lines for the intensity, blue line and lowest is the thermal environment generally felt by earth bound satellites. Additionally, there are two higher red lines and the lower of these is the Venus environment and the highest nearest sun approach

Dates given are 1 May 2010, 2 May 2011, 2 May 2012, and 3 May 2013. Launch was on 21 May 2010 and after a transfer orbit the curve reaches 2621 W/mxm on 7 Dec 2010 during VOI-1 and should have stayed at 2649 W/mxm after that

2. Thermal environment diagram (time line) if no nearest sun approach manuever is carried out

Captioned dates are the same. The curve takes us to the highest temp line after transfer orbit, then into solar orbit with temp standing at 3655 W/mxm which was reached on 17 April 2011 and the curve tells us that there will be two more nearest approaches before May 2013.

(page 34)

I find this page awfully difficult and time consuming to translate and let me think how best. The page structure is simple enough as the whole page is a table of things done to date, to be done before Nov 2011 for eventual re-insertion attempt.

My gut feeling is that I will have to translate each table element starting with row/column number. While here let me describe the table structure.

There are only 4 rows and of these row 3 is very wide as it contains all sorts of information. There are 13 columns. The top row is the month numbers, starting with April (month of 4th) 2011 as R1C2 and after that May (month of 5th), June (month of 6th), July (month of 7th) etc etc.

I do not remember off hand what were in row 2. Row 3 is where they can use both OME and RCS and row 4 where they can only use RCS. So, this is the large picture of the table, simple, but elementwise it is a mess! I might be tempted to do this during the course of this evening, unsure... P

The original planned orbit was elliptical (as prior missions), but equatorial. It was cleverly designed such that the angular
rate at apoapsis is similar to that of the cloudtops (sort of temporarily 'nimbosynchronous'), which would be ideal
for making movies.

At this point, though, any orbit would be a great orbit. The payload is diverse enough to give good results on the
nightside (lightning/airglow camera and IR cameras) as well as dayside imaging.

(page 34)

4.5 Schedule for the next nearest sun aproach manuever

(after this there is the table and after the table is: )

Constraints on the timing of test firing in orbit

・since it doubles as orbital plane change manuever it will have to be carried out near the point of elevation
・we need to secure enough preparation period before nearest sun manuever
・we must ensure that in-orbit temp rise is moderate

(now I am going for the table and translate table elements if they are easy to identify on the table. Let me do this with the first few and I then will propose another scheme for translation with an example, P)

R2C2: passing of the nearest to the sun point (month of 4 in row 1)
R2C7: test firing in orbit (month of 9 in row 1)
R2C9: nearest to the sun manuever (case A, problem with this is that nowhere on this table do I find case B!, P)) (month of 11 in row 1)

R3C1: orbit insertion plan 1
R4C1: orbit insertion plan 2

(so far so good, but immediately after this and if I continue with row#column# scheme the first element to be translated will have to be designated like (and do not laugh!)

R3&4C2&3:

Test burn (VOI-1 reproduction and understanding)

Upon igniting the broken nozzle the burn chamber became a total loss (section 4.3a) and

we need approapriate measures with the next nearest sun approach

(these three lines are contained in the largest pentagon on the table with the sharpest angle pointing to the right and there are two arrows radiating out from this pentagon to other smaller pentagons. I needed multiple suffices to cover more than two months (mainly). So, my new scheme is to describe entries with respect to row 1's month numbers)

(I will have to start afresh on the next page, P)

This largest pentagon covers April and May of this year (2011) and is common to both plan 1 and plan 2.

(in what follows I will refer to structural shapes, from left to right and from top to bottom)

(there are 3 pentagons astriding months of June and July)

1st pentagon reads "discuss if continued OME burn will be possible" and the exaplanation just below it reads "find out if OME can be fired by blow down, (section 4.3b)"

2nd pentagon reads "find ways to reduce ignition impacts" and the explanation just below it reads "prevent the total loss of engine, (section 4.3a)"

3rd pentagon reads "find ways to jettison oxidiser" and the explanation just below it reads "reduce total mass by as much as possible, (section 4.3c)"

(now, move to the right)

(the next pentagon covering months of July to October says "test burn (continuous)"

the box in September (month of 9) says "in-orbit test firing"

there is a judgement box in September and it reads "OME can or cannot be used?" and the line going below it reads "preparation for orbit re-insertion" and the other line goint to the right from this branching box reads "preparation for the planned orbit re-insertion"

There is another pentagon in the month of September and it reads "in-orbit oxidiser jettison tests"

There is another pentagon in October and it reads "oxidiser jettison"

(moving again to the right)

November box reads "OME manuever in orbit"

Another box in Novemver just below it says "RCS manuever in orbit"

There are notes: Solid lines mean "either done, accomplished, or scheduled" and dotted lines mean "only if required"

(end of page 34)

Don't forget Magellan! It was done at Venus first.

True! Mea culpa.

But still: The ship was designed to do it, and presumably Akatsuki was not.

Magellan wasn't designed for it either

(page 35)

5. Summary of this 3rd report

1. We have looked into the causes of CV-F closure and gained following conclusions.


• We need to take into consideration the transmission of the oxidiser across the valve seal
• CV-F can be closed by salt formation as a result of chemical reaction of fuel and oxidiser

2. By looking into the effects OME suffered from CV-F closure we have obtained following conclusions for re-imsertion.

• Thruster nozzle is very likely to have been damaged during VOI-1
• We need to conduct ground tests and experiments to see if re-ignition of the damaged thruster nozzled engine is possible and test firing in orbit, and depending on the outcome we will have to carry out preparations for re-insertion.

３．We must come up with a viable operational plan for orbit changes.

• OME test firings ---> early September 2011
• Orbit change at nearest approach to sun, depending on the outcome of test firings ----> November 2011

(end of page 35)

(I am feeling a little uneasy because I remember saying at the start that this report was a 48 page report. I do not remember now where this number came from, but I could not have said it without a firm basis...

Anyway, this JAXA 3rd report came to an abrupt end (to me) at page 35 and what is left is a set of appendices and the last page is 61, yes, 61, I just had a look. What follows is the contents list. I will carry on translating as these seem more detailed and we should be able to finish the rest well before OME test firings.)

Appendices

A. FTA results presented at 2nd Investigation Meeting

A.1 Looking into the causes of Akatsuki mulfunction
A.2 Akatsuki FTA

B.Looking into the causes of CVF-closure

B.1 Discussing cause candidates relating to CV-F design and manufacturing
B.2 Discussing cause candidates relating to dynamic behaviour of the valve
B.3 Discussing cause candidates relating to excessive insertion of the valve
B.4 Discussing cause candidates relating to wear and tear
B.5 Evaluation of the speed of propellant flow
B.6 Estimation of the propellant (fuel) flow amount inside Akatsuki's propulsive sysytem
B.7 Investigation of the past mulfunctions relating to oxidiser flow
B.8 Examples of gas supply piping system with 2 liquid propulsion system on board for prolonged flight in space

C. Effects suffered by OME

C.1 Current understanding of the status of OME through analyses
C.2 Breakage surface observation of the damaged burner
C.3 History of acceleration and angular velocity during the latter half of VOI-1
C.4 Propulsion characteristics of the damaged burner
C.5 Trying to find ways to reduce re-ignition impacts
C.6 Discussing the possibility of continued firing of OME
C.7 Discussing ways of jettisoning oxidiser

(end of appendices page)

(My translation will start from B. Here, suffice to say that the very end of A talks about doing another FTA for P3 pressure drop as an additional analysis and finding that its cause was also judged to be CV-F closure. P)

(page 43, start of appendix section (I seem to have bad keyboard contact...)

B.1 Cause candidates for CV-F relating to design and manufacturing

We have had talks with manufacturers regarding design, manufacturing and materials used including factory visits and obtained following information



E‐1) Material incompatibility used for the seals

We checked manufacturing and inspection records and confirmed that the seals had been manufactured with propellant compatible materials and according to design specifications

E‐4) Sliding area mulfanction due to incompatible materials (&, P)
E‐10) Bad manufacturing of the sliding parts

From the records (manufacturing and inspection) such as material verification records and surface treatment records we confirmed propellant compatibility of the materials used

E‐6) Bad design/manufacturing of the clearance of the sliding portion (&, P)
E‐7) Bad alignment of the valve and its moving section

Records inspection confirmed that valves had been manufactured according to specifications


E‐5) Clearance changes due to inapproapriate fixing method

We investigated the valve body deformation due to inapproapriate torque and this confirmed that the deformation was sufficiently small.

From all above we confirmed CV-F design/manufacturing information and are satisfied that above candidates are sufficiently innocent.

(end of page 43)

(page 44)

B.2 Cause candidates relating to the dynamic behaviour of the valve

１．Possibility of the valve sliding vibrationally due to the resonance of the regulator and the piping system

E‐12) Sliding products gritting due to an unexpected number of sliding motions

We conducted dynamic characteristics simulation and tests assuming fuel tank pressurisation. As a result, we confirmed that speedy (?, quickly reached?) vibrations between regulators themselves, between regulator and CV-F, and vibration of components did not occur.

２．Possibility that CV-F itself reached resonance over a certain working range, with mechanical parts breaking, falling away, and leading to gritting

E‐12) Sliding products gritting due to an unexpected number of sliding motions
E‐14) Mechanical parts breaking, dropping (or falling) off, and gritting due to an unexpected number of valve workings
E‐15) Falling off of coil springs

We conducted experiments and analyses where we changed the pressure both up and down streams of CV-F, covering the whole range of pressure status obtained from the flight history. As a result, we confirmed that CV-F closure, chatterings, frettings and other unexpected vibrations were not detected.

３．Possibility that the resonance due to transient response to the rising tank pressure in orbit might have affected mechanical parts

E‐14) Mechanical parts breaking, falling off, and gritting due to an unexpected number of valve workings
E‐15) Falling off of coil springs

Based on the telemetry data of the rapid tank pressure rise in orbit we conducted experiments controling the upstream pressure of CV-F, and for the down stream we conducted experiments using same piping diameter, same length, and same flow (?) volume.

As a result, we confirmed that the spread of cracking pressure and reseat pressure was contained within 0.002MPa and not leading to the closure.

(There is a simple picture of the valve test)

Top left caption on the picture reads "Valve working tests"

An arrow points down to the picture from "CV-F (spare CV-F)"
An arrow points up to the picture from "Acceleration sensor: Data obtained : acceleration during valve working (vibration monitor)and valve capability trend after load imprinting"

From these analyses and tests of the dynamic behaviour of the valve we are satisfied that above candidates are sufficiently innocent.

(end of page 44)

(page 45)

B.3 Cause candidates relating to excessive valve insertion

E‐3) Excessive valve insertion due to long term reverse pressure imprint

We conducted a 21 day test simulating the 35 day long in-orbit reverse pressure state by applying the pressure 1.5 times that in orbit

The cracking pressure (that pressure which changes CV-F status from close to open) and the re-seat pressure (that pressure which changes CV-F status from open to close) before and after the test are shown in the following graph. (Note: see section 1.2 for CV-F valving actions)

From the test we confirmed that there was no excessive valve insertion (valve moving beyond the regular close position into further close state with too much seal pressing) and we are satisfied that above candidate is sufficiently innocent.

(end of the main text of this page 45)

(the graph is simple with the vertical showing the % change after AT. Horizontal is the quantum time line and there are 9 of them from left to right. They are;)

1. AT (whatever this is, P)
2. Pre-assembly inspection
3. before valve action test
4. after valve action test
5. before dynamic characteristics simulation test

6. after 1st in-orbit tank pressure rise simulation
7. after 2nd in-orbit tank pressure rise simulation
8. after 3rd in-orbit tank pressure rise simulation
9. after 21 days of reverse pressure imprint

(end of page 45)

(page 46)

B.4 Cause candidates relating to wear and tear

Most of the ground tests for checking the health of the valve were conducted in helium gas environment. However, in real situation the environment is the mixture of helium and propellant vapours. We therefore set out to see how this difference might have led to increased friction and wears.

E‐8) wears and surface roughing due to sliding motion
E‐9) surface corrosion due to incompatible materials used for sliding mechanism

We conducted fuel environment friction tests. These tests use a pin-on-disc setup and we observe static/dynamic friction coefficients, amount of wear, wear particles by subjecting the test pieces (which have been immersed either in helium or propellant vapour environment) to sliding motion.

As a result we confirmed that there was no worsening by the propellant vapours. Rather, we observed that we obtain markedly larger results from the tests under helium gas environment. (surprise! helium should be less sticky!!!, P)

E‐11) Propellant vapour environment producing products which grit

Measuring the disc wear amount

The depth of the grooves left on the test pieces in the wake of sliding motion under helium environment was 5 to 10 microns. In contrast to this the maximum depth unde propellant environment was 4 microns and fuel environment did not contribute to the worsening.

Products under propellant fuel vapour environment

Upon sliding tests we dis see a minute amount of metal-metal friction induced metal particles, but we did not find any chemically produced products reacting with fuel or others

(there is a diagram and captions are;)

Fuel environment friction test
Pin
Disc

＜data obtained＞
Friction coefficient, wear amount, wear particles

As a result of these tests we did not observe any phenomena where fuel environment gave adverse effects to friction issues and we are satisfied that above candidates are sufficiently innocent.

(end of page 46)

(page 47)

B.5 Evaluating the propellant flow speed

We are now going to show two different evaluation methods, 1) method employed at the time of design and 2) method introduced as the transmission model based on our new insight into the mulfunctions

1) Evaluation at design time: by leak model

1-1) From the helium leak speed we assumed a viscous flow through an equivalent orifice and estimated the equivalent orifice diameter. In helium leak measurements we only have one component pressurising the downstream of CV-F and the upstream is vaccum, leading to adoption of viscous flow hypothesis.


FORMULA for viscous flow: (I cannot type it out. Please refer to the original page, P) where

Q(Flow,mass)： Flow mass speed ( mg/s)
ρ： density at the average pressure of up and down streams (g/m3)
μ：viscocity coeeficient （Pa・s）
d, L： bore diam. and length (mm)
ΔP： differential pressure (MPa)

1-2) We evaluated the propellant flow speed assuming difusion (or dispersion?, P) in a ficticious orifice due to the differential pressure within the two component system (helium + propellant).

In estimating the propellant flow speed we have a configuration where against a constant one atmospheric pressure (of helium and propellants) we have a saturated propellant vapour pressure downstream of CV-F and no propellant concentration upstream.

We therefore evaluated the propellant leak speed as due to propellant diffusion by the in-orifice differential pressure.

FORMULA for diffusion: (I cannot type it out. Please refer to the original page, P) where

Q(Leak,mass)： leak speed (mg/s)
Dgass： mutual dispersion constant of gas （m2/s）
RT： gas constant （8 3 J/mol K）×temp.（K）
ΔP： differential pressure（＝saturated vapour pressure） (MPa)

1-3) Evaluation result (for Q(Leak,mass))

(this is a simple table and I will narrate its contents as follows, P)

For fuel Q(Leak,mass) is 2x10-10 mg/s @0.0014MPa

For oxidiser Q(Leak,mass) is 2x10-8 mg/s @0.1MPa

(-10 and -8, to the power, naturally!, P)

(See section 2.2.1, item A on the table "Comparison of measured values and model values of propellant leak speed")

In all above we assumed all of the flow was due to leakage, ignoring transmission

(end of page 47)

One addition to page 45's graph:

I simply forgot to translate those captions on the graph. The red line is the re-seat pressure and the blue line the cracking pressure. Apologies!

P

(page 48)

(There are two tables here, one of them is simple and I will type it all out, but the second table will be translated after the main texts, P)

B.5 Evaluation of the propellant flow speed (continuation)

2) Evaluation at time of re-evaluation based on transmission model

(I now realise that my translation of "transmission" may be wrong, because the suffix to the formula below has "Per" and it probably means perforation...?, but I will carry on with "transmission", P)

2-1) Actual measurements and literature survey of transmission coefficient of the sealing material (polymeric)

(Here below is the table to be translated later on)


Transmission coefficient of gaseous molecules in polymeric materials is proportional to the product of degree of solvance (am afraid I do not know the right word, P) (itself proportional to the partial pressure of the gas in contact) and the dispersion coefficient.

Of these, dispersion coefficient has positive proportional relationship with molecular weight, but degree of solvance depends critically on molecular type (such as polarity, or polarisation). It is for this reason that helium has a larger transmission coefficient than fuel.

Also, the absolute value of the transmission coefficient is critically dependent on the crystaline structure of the polymeric material (itself dependent on production process) and the measured values may differ from the values given in literatures.

(here below, transmission coefficient, Per, is defined as; )


Per = S times Dsolid

Per： transmission coefficient (m2/sMPa)
S： degree of solvancy (1/MPa)
Dsolid： dispersion coefficient inside solids (m2/s)

2-2) Evaluation of transmission speed through the valve

Based on the transmission formula we evaluated the propellant flow speed by estimating the sealing material's geometric parameters (A/t) using helium flow speed

(here again, am unable to type out the forlura, P)

transmission formula

Q(Per, mass) ： transmission speed (mg/s)
A, t : sealing material's contact area with gas (mxm, m)
ΔP： differential partial pressure (= saturated vapour pressure) (MPa)
P： avareage partial pressure （＝(1/2) saturated vapour pressure） （MPa）

2-3) result of evaluation (of Q(Per, mass))

This is a table, but I will spell it out as;

Q(Per, mass) = 1x10-10 mg/s @ 0.0014MPa for fuel
Q(Per, mass) = 3x10-5 mg/s @ 0.1MPa for oxidiser

Note that item (B, bee) of the table in section 2.2.1 "Comparison of propellant flow speeds, model values and measures values" is based on the assumption that all the flows both up and down streams of the valve originate from transmission through the sealing material.

(end of main texts of page 48)

(here below the table mentioned above) (it is simple and carries the values (measured and literature) of Per, S, and D solid for oxidiser vapour and fuel vapour and helium)

(In this matrix the top entry in the 1st column is oxidiser vapour, next down is fuel vapour and at the bottom is helium)

(1st character string from the left in the 1st row is "measued value" and the next string to the right of it is "literature")

(the rest are all alphanumeric and please refer to the table on page 48)

Source: Polymer Handbook, 3rd ed., J. Brandrup and E.H. Immergut, John Wiley & Sons, 1989

(end of page 48)

(page 49)

B.6 Estimating the amount of propellant flow in Akatsuki propulsive system (fuel)

(hereafter is a schematic of propellant supply system)

(Leftmost box: fuel tank where diaphgram is located
section to its right: section D, between fuel tank and CV-F just before the step in the middle
small box just to the right of this step is a valve, then further to its right is CV-O, then finally oxidiserl tank, the box at the right end
between oxidiser tank and CV-O is section A
between CV-O and the simple valve is section B
between the simple valve and CV-F is section C to which helium from above flows)

Using the fuel flow speed across the valve and diaphgram which we measued we esimated the amount of propellant vapour crossing in the gas supply piping system. We found that the amount of fuel vapour moving across CV-F is almost zero and most stayed within section D (downstream of CV-F)

Note: Volume of D is the initial volume of empty space and does not include the increase as a result of fuel consumption

(hereafter there are 6 graphs in 2 rows and 3 columns)

(1st row from thre left are sections D, C, and B in this order and the same order goes to the second row as well)
(horizontal is the time line for 6000 hours, and the vertical for the 1st row is the amount of existing fuel in mg and the vertical for the 2nd row is fuel's partial pressure in MPa)

(red solid line is the measued value and red dotted is the pre-flight analysis)


Note: Measued values and pre-flight anaysis values are almost in agreenment in all sections

(end of page 49)

(page 50)

B.7 Looking at past examples of mulfunctions relating to oxidiser movement

Given Akatsuki's mulfunctions we conducted a survey of mulfunctions (minor mulfunctions included) during ground tests and in flight of overseas probes. We noted that there was no example of lost mission arising from salt formation. However, we also noted that there were a few cases of crystal formation of salts involving MMH (mono methyl hydrazine) as fuel.

 Example of severe mulfunction leading to loss of probe - reflected during design stage (direct translation, meaning unknown, P) -

• Mars Observer （launched in September 1992, approached Mars in August 1993, fuel : MMH and oxidiser ：NTO※)

It is estimated that during its flight the oxidiser condensed and liquefied at a cold portion of the gas supply system flew to the fuel side when the pyro valve was opened and led to explosive reaction

 Examples of mulfunction in orbit (mission was completed)

• Viking‐1 (August 1975 launch, Mars arrival June 1976, fuel : MMH oxidiser ：NTO※)

Regulator valve internal leak was observed during flight. Salt formation was assumed to be its cause.

• Intelsat‐603 (Launch 1991 fuel ：MMH oxidiser ：NTO※)

During its 1sy manuever an internal leak of regulator valve occurred. During 2nd and 3rd manuevers fuel side CV-O (or something similar, P) closure was observed. Salt formation was estimated to be its cause due to its long mission period.

 Example during ground tests

• Marienr‐9 (May 1971 launch, arriving at Mars in November, fuel ：MMH oxidiser ：NTO※)

CV closed during ground test burnt. Crystal formation of iron nitrate was confirmed upon dis-assembly of the system. Its cause was put to swelling of TFE.

※NTO is an abbreviation of N2O4. In actual use NO is often added to reduce possibility of metal corrosion. In the case of Akatsuki 3% of NO was added to N2H4 (forming MON-3).

(end of page 50)

"Examples of mulfunction in orbit (mission was completed)"

Interesting! I didn't know this.

in fact, it was the problem with the pressure regulator on Viking which prompted controllers to delay pressurizing the tanks on Mars Observer to just before orbit insertion, and the system (borrowed from telecom and meteo sats) was not designed for that...