Author Topic: SpaceX McGregor Testing Updates and Discussion THREAD 2  (Read 473154 times)

Offline Lars_J

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Re: SpaceX McGregor Testing Updates and Discussion THREAD 2
« Reply #40 on: 07/07/2013 02:24 am »
He said what he said. He called the first stage for the 1st v1.1 "F9-R". SpaceX can decide to call things whatever they want.

When did he say that? (Asked respectfully!  :))

https://twitter.com/elonmusk/status/341405518566395904
Quote
Elon Musk
‏@elonmusk
1st firing of Falcon 9-R advanced prototype rocket. Over 1M lbs thrust, enough to lift skyscraper pic.twitter.com/AUCsWTw77E
« Last Edit: 07/07/2013 02:25 am by Lars_J »

Offline ClaytonBirchenough

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Re: SpaceX McGregor Testing Updates and Discussion THREAD 2
« Reply #41 on: 07/07/2013 02:39 am »
https://twitter.com/elonmusk/status/341405518566395904
Quote
Elon Musk
‏@elonmusk
1st firing of Falcon 9-R advanced prototype rocket. Over 1M lbs thrust, enough to lift skyscraper pic.twitter.com/AUCsWTw77E

Yes, but that is sort of vague. The original question was:

Are we 100% that F9 v1.1 is going to fly the same 180s / Mach ~10 path to separation as v1.0?

While that looks like a v1.1 stage, it is stated as being the F9-R stage. They are essentially the same; the v1.1 stage is basically the F9-R stage minus the landing legs. Musk did not say the v1.1 first stage was the F-9R first stage. All in all, this still does not answer our question about first stage separation velocity!
« Last Edit: 07/07/2013 02:40 am by ClaytonBirchenough »
Clayton Birchenough

Offline Jason1701

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Re: SpaceX McGregor Testing Updates and Discussion THREAD 2
« Reply #42 on: 07/07/2013 06:42 am »
Yup - that's the reference I was basing my question on. So again, could we be looking at a 02:00.00 to ~02:20.00 first stage burn or is it too early to speculate?

F9-R is not Falcon v1.1. The statement Musk made may not apply to the v1.1 flights.

He said what he said. He called the first stage for the 1st v1.1 "F9-R". SpaceX can decide to call things whatever they want.
Falcon 9-R is an evolvable design so ever time the rocket flys, there may be some changes (upgrades) made.

Just to keep Jim happy. :)

Offline cambrianera

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Re: SpaceX McGregor Testing Updates and Discussion THREAD 2
« Reply #43 on: 07/07/2013 07:30 am »
About staging velocity of the vehicle we saw in testing at McGregor.

Second stage is double volume and double thrust than v1.0 with less increase in payload.
In another thread I did some speculations on this, numbers are consistent with a reduction of staging velocity.
http://forum.nasaspaceflight.com/index.php?topic=31514.msg1040840#msg1040840
Oh to be young again. . .

Offline docmordrid

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Re: SpaceX McGregor Testing Updates and Discussion THREAD 2
« Reply #44 on: 07/07/2013 07:48 am »
^^ Just quoting your summary -

Quote
Obviously this is approximate, but consistent with SpaceX intention to reduce staging velocity from 3000 to 2000 m/s.
« Last Edit: 07/07/2013 07:49 am by docmordrid »
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Offline malu5531

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Re: SpaceX McGregor Testing Updates and Discussion THREAD 2
« Reply #45 on: 07/07/2013 07:52 am »
Yup - that's the reference I was basing my question on. So again, could we be looking at a 02:00.00 to ~02:20.00 first stage burn or is it too early to speculate?

Then I calculate it will stage after 150s (2:30), with 20% residual fuel for a massive 5715 m/s residual delta-v and capable of lifting 10.5mT to LEO. (I did not anticipate the fly-back and landing delta-v requirements would be this big and suspect this is wrong).

Maybe my dry/wet assumption based on previous discussions here of 3.5% is too optimistic for the reusable version with legs and all? At 10%, with a bit lower staging velocity (still around mach 6), it would stage with 3140 residual delta-v after 140s (2:20) and lift 9.5mT to LEO.

Another error I have made could be the guesstimate of 236000 lbs  (wet, including payload) for the second stage from discussions on this forum, a larger second stage would be beneficial. (If anyone have better numbers, please let me know)

calculations:
http://tinyurl.com/Falcon9R-2

speculation:
It's been implied in the discussions on this forum, that SpaceX is "hiding" something. Since we expected the qualifying "full duration burn" of the stage to be 3 minutes (180 second) (the time used for 1.0 and each Merlin 1D) and we know no stage-testing burn achieved this, yet SpaceX consider the stage to be qualified. Perhaps, they are "hiding" something and the "gut feeling" here is right? - But what they are "hiding" (by lack of clarity) is not some conspiracy to rush qualification, but simply the true duration of the first stage burn, perhaps in order to keep their plans for reusability to themselves at this point. Knowing the duration would certainly help reduce the unknowns in this puzzle.

conclusion:
Yes, docmordrid, I believe we can look at a 2:00 - 2:30 first stage burn (probably closer to 2:30). However, much is still unclear.

Update, after some searching the forum on the June 19 test
According to this witness, the design qualification burn was 2.5 minutes (2:30);
« Last Edit: 07/07/2013 08:16 am by malu5531 »

Offline ClaytonBirchenough

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Re: SpaceX McGregor Testing Updates and Discussion THREAD 2
« Reply #46 on: 07/07/2013 10:28 am »
Another error I have made could be the guesstimate of 236000 lbs  (wet, including payload) for the second stage from discussions on this forum, a larger second stage would be beneficial. (If anyone have better numbers, please let me know)

Wouldn't a mass of 236000 lbs. imply a thrust to weight ration <1 for the second stage?
Clayton Birchenough

Offline Joffan


Wouldn't a mass of 236000 lbs. imply a thrust to weight ratio <1 for the second stage?

The most important second stage task is to get the lateral speed up to achieve orbit. T/W only determines how quickly this will happen. The major push against gravity to get above the atmosphere has been done by the first stage.

Disclaimer: I am not a rocket scientist.
« Last Edit: 07/07/2013 11:18 am by Joffan »
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Offline malu5531

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Re: SpaceX McGregor Testing Updates and Discussion THREAD 2
« Reply #48 on: 07/07/2013 11:27 am »
Wouldn't a mass of 236000 lbs. imply a thrust to weight ration <1 for the second stage?

Edit: After running the numbers I've found a better answer, considering any stage 2 mass, as Joffan suggested.

I simulated LEO performance for stage 2 mass between 100´000 lbs - 400´000 lbs (graph attached) using the corresponding required delta-v for LEO, which depends on time-to-orbit, which depends on stage configurations. For LEO I use 200x200, 26 deg inclination from Boca Chica.

Depending on how you optimize stage 2 mass, for reusability or not, there will be two different optimal configurations; 240klbs for non-reusable and 330klbs for reusable.

Stage 2 MassScenarioLEO (kg)Staging (mach)Staging (s)Time to LEO (s)
240 klbsuse once16191~9184595.4
240 klbsreuse12905~7.6168594.9
330 klbsuse once15793~7.2164760
330 klbsreuse13312~6.2149.63757

conclusion
If my calculations are correct, and SpaceX prioritise design for reusability, the mass of stage 2 should be ~330'000 lbs, with a stage 1 burn time of ~150s (2:30), delivering ~13mT to LEO.

If we assume SpaceX will do no "use once" launches, all launches will be capable of delivering at least 13 mT to LEO and have spare capacity to iteratively progress towards full reusability.

Given this, it's reasonable to assume SpaceX design requirement for stage 1 full duration burn was 150s

guesstimates:
Dry/wet ratio s1: 3.5% when non-reusable (no legs), 5% when reusable (with legs, etc)
Dry/wet ratio s2: 3.55% (from cambrianera's estimate)
Residual fuel: 1% when non-reusable, 8% when reusable (2800 m/s residual delta-v)
Total mass of stack: 0.8 times thrust at liftoff (1.323 Mlbf @ SL)

calculation and full table:
http://tinyurl.com/Falcon9R-3
« Last Edit: 07/07/2013 02:10 pm by malu5531 »

Offline sanman

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Re: SpaceX McGregor Testing Updates and Discussion THREAD 2
« Reply #49 on: 07/07/2013 09:01 pm »
Keep in mind the GH is a test bed with very over designed legs for off nominal landing during testing, not a production rocket. Presumably the insulation is to keep the metal in those legs from losing temper or otherwise taking damage but needing to replace that insulation after every *TEST* is not a big deal. It's not like having a bit of extra flame in the vicinity of the Merlin exhaust is likely to hurt anything.

Could some of the heat or exhaust from the engine be used to power the leg actuators? Would that be a useful approach to take?

Offline Kabloona

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Re: SpaceX McGregor Testing Updates and Discussion THREAD 2
« Reply #50 on: 07/07/2013 09:14 pm »
Wouldn't a mass of 236000 lbs. imply a thrust to weight ration <1 for the second stage?

Edit: After running the numbers I've found a better answer, considering any stage 2 mass, as Joffan suggested.

I simulated LEO performance for stage 2 mass between 100´000 lbs - 400´000 lbs (graph attached) using the corresponding required delta-v for LEO, which depends on time-to-orbit, which depends on stage configurations. For LEO I use 200x200, 26 deg inclination from Boca Chica.

Depending on how you optimize stage 2 mass, for reusability or not, there will be two different optimal configurations; 240klbs for non-reusable and 330klbs for reusable.

Stage 2 MassScenarioLEO (kg)Staging (mach)Staging (s)Time to LEO (s)
240 klbsuse once16191~9184595.4
240 klbsreuse12905~7.6168594.9
330 klbsuse once15793~7.2164760
330 klbsreuse13312~6.2149.63757

conclusion
If my calculations are correct, and SpaceX prioritise design for reusability, the mass of stage 2 should be ~330'000 lbs, with a stage 1 burn time of ~150s (2:30), delivering ~13mT to LEO.

If we assume SpaceX will do no "use once" launches, all launches will be capable of delivering at least 13 mT to LEO and have spare capacity to iteratively progress towards full reusability.

Given this, it's reasonable to assume SpaceX design requirement for stage 1 full duration burn was 150s

guesstimates:
Dry/wet ratio s1: 3.5% when non-reusable (no legs), 5% when reusable (with legs, etc)
Dry/wet ratio s2: 3.55% (from cambrianera's estimate)
Residual fuel: 1% when non-reusable, 8% when reusable (2800 m/s residual delta-v)
Total mass of stack: 0.8 times thrust at liftoff (1.323 Mlbf @ SL)

calculation and full table:
http://tinyurl.com/Falcon9R-3

Elon says the payload penalty for reusable F9R vs. expendable will be about 40%. Your graph is showing less than 20% penalty. (13klbs vs. 16klbs.)

http://www.popularmechanics.com/science/space/rockets/elon-musk-on-spacexs-reusable-rocket-plans-6653023

Edit: does your graph assume  an expendable S2 for both curves, but a reusable S1 for one curve and an expendable S1 for the other?

That would explain the discrepancy, since I assume Elon's comparison was between expending *both* stages and reusing *both* stages.
« Last Edit: 07/07/2013 10:21 pm by Kabloona »

Offline Kabloona

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Re: SpaceX McGregor Testing Updates and Discussion THREAD 2
« Reply #51 on: 07/07/2013 09:19 pm »
Keep in mind the GH is a test bed with very over designed legs for off nominal landing during testing, not a production rocket. Presumably the insulation is to keep the metal in those legs from losing temper or otherwise taking damage but needing to replace that insulation after every *TEST* is not a big deal. It's not like having a bit of extra flame in the vicinity of the Merlin exhaust is likely to hurt anything.

Could some of the heat or exhaust from the engine be used to power the leg actuators? Would that be a useful approach to take?

Unnecessary complication. I would expect a simple mechanical solution, eg springs.

Offline bubbagret

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Re: SpaceX McGregor Testing Updates and Discussion THREAD 2
« Reply #52 on: 07/07/2013 09:21 pm »
Keep in mind the GH is a test bed with very over designed legs for off nominal landing during testing, not a production rocket. Presumably the insulation is to keep the metal in those legs from losing temper or otherwise taking damage but needing to replace that insulation after every *TEST* is not a big deal. It's not like having a bit of extra flame in the vicinity of the Merlin exhaust is likely to hurt anything.

Could some of the heat or exhaust from the engine be used to power the leg actuators? Would that be a useful approach to take?

Unnecessary complication. I would expect a simple mechanical solution, eg springs.

Aka: http://en.wikipedia.org/wiki/Rube_Goldberg_machine
...or for the other side of the pond: http://en.wikipedia.org/wiki/Heath_Robinson

Offline docmordrid

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Re: SpaceX McGregor Testing Updates and Discussion THREAD 2
« Reply #53 on: 07/07/2013 09:24 pm »
If the 40% v 20% disparity is from a mis-guesstimate of upper stage mass &  performance will that necessarily change the estimated lower state burn time of 180s v 150s (the question I posed upthread)?
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Offline Oli

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Re: SpaceX McGregor Testing Updates and Discussion THREAD 2
« Reply #54 on: 07/07/2013 10:29 pm »
^

Its lbs, not kg? But 150t is still far too heavy.

Offline Kabloona

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Re: SpaceX McGregor Testing Updates and Discussion THREAD 2
« Reply #55 on: 07/07/2013 11:28 pm »
^

Its lbs, not kg? But 150t is still far too heavy.

For comparison, spacelaunchreport estimates S2 total mass at around 78 tonnes.

Offline Zed_Noir

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Re: SpaceX McGregor Testing Updates and Discussion THREAD 2
« Reply #56 on: 07/08/2013 01:12 am »
Keep in mind the GH is a test bed with very over designed legs for off nominal landing during testing, not a production rocket. Presumably the insulation is to keep the metal in those legs from losing temper or otherwise taking damage but needing to replace that insulation after every *TEST* is not a big deal. It's not like having a bit of extra flame in the vicinity of the Merlin exhaust is likely to hurt anything.

Could some of the heat or exhaust from the engine be used to power the leg actuators? Would that be a useful approach to take?

Unnecessary complication. I would expect a simple mechanical solution, eg springs.

IIRC the F9-R uses high pressure Helium to actuated the landing gears and probably use subsequently as  shock absorbing fluid.

Offline justineet

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Re: SpaceX McGregor Testing Updates and Discussion THREAD 2
« Reply #57 on: 07/08/2013 01:18 am »



"SpaceX's Merlin 1D engine has achieved a full mission duration firing  and multiple restarts at target thrust and specific impulse (Isp). The  engine firing was for 185 seconds with 147,000 pounds of thrust, the  full duration and power required for a Falcon 9 rocket launch. The tests  took place at SpaceX's rocket development facility in McGregor, Texas.  The Merlin 1D builds on the proven technology of the Merlin engines used  on the first three flights of Falcon 9, including the recent historic  mission to the International Space Station."  --- Space X(explanation on 1D Melrlin Engine Testing...Youtube)




As you can see from Space X Merlin 1D engine test explanation, the "full duration" of stage one for Falcon 1.1 is 3 minutes. So far Space X has only reached about two minutes testing with all 9 engines of Falcon 1.1 together.  It seems to me quite a leap of faith Space X would go on with production and launch w/o full duration 3-minute by concluding the cause for the aborts is test stand environment related not hardware issues with the stages. It's useful to note  Merlin 1D might qualify perfectly in stand alone full duration tests again and again. But when configured with other 9 engines together the additional thermal and kinetic energy produced from the 50% more powerful thrust may be surpassing the coping margins of some of the engine components. Or, if not the engine components, other parts of the stages!

For reasons stated above I find very hard to think Space X would go on with launch w/o doing a full duration 3-minute test. If full duration of F1.1 first stage is truly 3 minutes as mentioned by Space X, it would be a gigantic leap of faith for Space X to conclude that the rocket would operate properly past 2-minutes when that has not been proven on the ground!!!

Offline beancounter

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Re: SpaceX McGregor Testing Updates and Discussion THREAD 2
« Reply #58 on: 07/08/2013 01:51 am »



"SpaceX's Merlin 1D engine has achieved a full mission duration firing  and multiple restarts at target thrust and specific impulse (Isp). The  engine firing was for 185 seconds with 147,000 pounds of thrust, the  full duration and power required for a Falcon 9 rocket launch. The tests  took place at SpaceX's rocket development facility in McGregor, Texas.  The Merlin 1D builds on the proven technology of the Merlin engines used  on the first three flights of Falcon 9, including the recent historic  mission to the International Space Station."  --- Space X(explanation on 1D Melrlin Engine Testing...Youtube)




As you can see from Space X Merlin 1D engine test explanation, the "full duration" of stage one for Falcon 1.1 is 3 minutes. So far Space X has only reached about two minutes testing with all 9 engines of Falcon 1.1 together.  It seems to me quite a leap of faith Space X would go on with production and launch w/o full duration 3-minute by concluding the cause for the aborts is test stand environment related not hardware issues with the stages. It's useful to note  Merlin 1D might qualify perfectly in stand alone full duration tests again and again. But when configured with other 9 engines together the additional thermal and kinetic energy produced from the 50% more powerful thrust may be surpassing the coping margins of some of the engine components. Or, if not the engine components, other parts of the stages!

For reasons stated above I find very hard to think Space X would go on with launch w/o doing a full duration 3-minute test. If full duration of F1.1 first stage is truly 3 minutes as mentioned by Space X, it would be a gigantic leap of faith for Space X to conclude that the rocket would operate properly past 2-minutes when that has not been proven on the ground!!!

Well it's not like SpaceX doesn't have any flight experience.  If they believe they've characterised their engine performance sufficiently well and believe that it's not flight issues that are preventing full duration burns, then it's not surprising that they move to the next stage.
It's their engines and vehicle after all and their history speaks for itself.

If that is the case, we'll find out just how well they do know their business on the first F9v1.1 launch.  Sure looking forward to it.

I also believe that SpaceX is not too worried about schedule.  They seem to simply take the long view and talk to their customers.

Cheers,
Beancounter from DownUnder

Offline Joffan


For reasons stated above I find very hard to think Space X would go on with launch w/o doing a full duration 3-minute test. If full duration of F1.1 first stage is truly 3 minutes as mentioned by Space X, it would be a gigantic leap of faith for Space X to conclude that the rocket would operate properly past 2-minutes when that has not been proven on the ground!!!

This was achieved some time ago. See http://www.nasaspaceflight.com/2013/06/reducing-risk-ground-testing-recipe-spacex-success/, noting the paragraph highlighted with "UPDATE".
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