Author Topic: Deep Space Gateway Power/Propulsion RFI  (Read 42027 times)

Online envy887

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Re: Deep Space Gateway Power/Propulsion RFI
« Reply #120 on: 03/16/2018 04:22 pm »
So is an RFP expected for the PPE? Or does that exist somewhere and I missed it?

The latest front page article mentions that the budget proposal indicates that the RFP will ask for the company proposing the PPE design will also commercially source a launch for it.

https://www.nasaspaceflight.com/2018/03/cislunar-station-new-name-presidents-budget/

Offline speedevil

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Re: Deep Space Gateway Power/Propulsion RFI
« Reply #121 on: 03/16/2018 04:35 pm »
So is an RFP expected for the PPE? Or does that exist somewhere and I missed it?

The latest front page article mentions that the budget proposal indicates that the RFP will ask for the company proposing the PPE design will also commercially source a launch for it.

https://www.nasaspaceflight.com/2018/03/cislunar-station-new-name-presidents-budget/

Quote
“The targeted release of the draft solicitation will be in the April 2018 timeframe with final proposals anticipated to be due in the late July 2018 timeframe"

Offline Proponent

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Re: Deep Space Gateway Power/Propulsion RFI
« Reply #122 on: 03/18/2018 02:35 pm »
But what is the point with LLO really? It adds 1000 m/s of requirements on the Orion(vs NRHO), but subtracts 1000 m/s of requirements from the lander (vs staging at NRHO). The lander can be smaller and lighter compared to Orion that has to re-enter and support crew for longer durations and thus it could very well take less fuel to move the lander an extra 1 km/s vs move Orion an extra 1 km/s

I have not thought through all of the implications of the whole range of possible lunar parking orbits, but off hand it seems to me the LLO probably offers more frequent return-to-Earth opportunities.  Certainly for near-equatorial landing sites, LLO gives an abort option every 2 hours.  For near-polar sites, if you have the delta-V for big plane changes (as was to be the case in Constellation), then you also have frequent abort options.

You could say that a rendezvous point at L1 or L2 gives you anytime abort.  But the catch is the getting from the lunar surface to the L-point is going to take about 3 days all by itself.  At some point, humanity is going to have to cut the cord and not worry about being able to beat a speedy return to Earth, but in the near term, taking a week or so to get home might seem risky.

Offline Proponent

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Re: Deep Space Gateway Power/Propulsion RFI
« Reply #123 on: 03/18/2018 03:01 pm »
While speculation is fun, I prefer to know what I'm talking about.

Always a good idea!

Quote
I am looking for the thread on ideal Earth orbit transfer points for SEP tugs and search results have failed so far.  If no such thread exists, can someone help me identify the homework that needs to be done to justify thread creation?  I'm really curious just how much TLI(or any other destination) Dv can be transferred from Earth LVs.  100 km 86° LLO appears to be the best compromise to use for proper exploration of the Lunar poles while allowing access to other points of interest.  Fixing the Earth orbit departure point(Please let it be my hypothesized EML1-synchronous elliptical parking orbit with a perigee between 6000-10,000 km.) is required so that I can move beyond speculation about overall architecture mass budgets.

I don't understand what you mean by transferring TLI delta-V.  But if we're talking about using electric propulsion to get things to the moon, the most efficient thing to do (i.e., the thing that would make most use of electric propulsion and the least use of chemical) would be to shift from chemical to electric in LEO.  The delta-V needed for a constant-low-thrust transfer between two circular orbits is simply the difference in the circular velocities of the two orbits.  You can think of escape as being a circular orbit at infinity, i.e., one with a circular velocity of zero.  TLI is a little bit short of escape, but the difference is not large.

The inclination of the lunar orbit has little impact on the delta-V needed.  The moon's radius being about 1738 km, the difference between aiming for lunar equatorial orbit and lunar polar orbit at an altitude of 100 km is only about (1738 km + 100 km)/(384,400 km) = 0.00452 radians = 0.259 degrees.

Offline Joseph Peterson

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Re: Deep Space Gateway Power/Propulsion RFI
« Reply #124 on: 03/21/2018 06:04 am »
While speculation is fun, I prefer to know what I'm talking about.

Always a good idea!

Quote
I am looking for the thread on ideal Earth orbit transfer points for SEP tugs and search results have failed so far.  If no such thread exists, can someone help me identify the homework that needs to be done to justify thread creation?  I'm really curious just how much TLI(or any other destination) Dv can be transferred from Earth LVs.  100 km 86° LLO appears to be the best compromise to use for proper exploration of the Lunar poles while allowing access to other points of interest.  Fixing the Earth orbit departure point(Please let it be my hypothesized EML1-synchronous elliptical parking orbit with a perigee between 6000-10,000 km.) is required so that I can move beyond speculation about overall architecture mass budgets.

I don't understand what you mean by transferring TLI delta-V.  But if we're talking about using electric propulsion to get things to the moon, the most efficient thing to do (i.e., the thing that would make most use of electric propulsion and the least use of chemical) would be to shift from chemical to electric in LEO.  The delta-V needed for a constant-low-thrust transfer between two circular orbits is simply the difference in the circular velocities of the two orbits.  You can think of escape as being a circular orbit at infinity, i.e., one with a circular velocity of zero.  TLI is a little bit short of escape, but the difference is not large.

The inclination of the lunar orbit has little impact on the delta-V needed.  The moon's radius being about 1738 km, the difference between aiming for lunar equatorial orbit and lunar polar orbit at an altitude of 100 km is only about (1738 km + 100 km)/(384,400 km) = 0.00452 radians = 0.259 degrees.

Thanks for the reply.  Hopefully I can rephrase my question in a more intelligible manner.

SEP provides a ~90% reduction in required propellant mass.  I want to argue in favor of is a version of chemical limbo, or 'how low can you go.'  SEP's primary limitation in Earth orbit is the inner Van Allen belts, or a perigee of 6.000-10,000 km.  There should exist an elliptical parking orbit with a perigee of ~10,000 km that allows for 2nd stage disposal/recovery, and, maximizes total payload to Lunar orbit, while, allowing support of dozens of Lunar landings per Earth year.  What I want to know is if anyone has already done the math to find out where we can expect ideal cargo staging points are?

If not, I'll be asking for help checking my methodology.
I hate feeling like I'm playing chess when everyone else is playing checkers.  Reality proving my feelings are correct is even worse.

Online Ronsmytheiii

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Re: Deep Space Gateway Power/Propulsion RFI
« Reply #125 on: 08/08/2018 06:38 pm »
Seems applicable:

Quote
Space Systems/Loral, L.L.C., (SSL) in Palo Alto, California, $2 million
  Proposal: In-Space Xenon Transfer for Satellite, Servicer and Exploration Vehicle Replenishment and Life Extension

https://www.nasa.gov/press-release/nasa-announces-new-partnerships-to-develop-space-exploration-technologies

Offline DreamyPickle

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Re: Deep Space Gateway Power/Propulsion RFI
« Reply #126 on: 08/14/2018 05:35 pm »
Shouldn't this thread be moved to the moon section with the rest of LOP-G? It's especially silly to have this in the SLS section when there's talk of moving it to a commercial launch.

Offline libra

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Re: Deep Space Gateway Power/Propulsion RFI
« Reply #127 on: 12/08/2018 11:50 am »
Quote
Also launch from Earth is implied NRHO having a few opportunities each month vs L2 having an opportunity every day. So for regular opperations NRHO imposes mission planning/scheduling restrictions.

This is a very interesting point. So NRHO is a little "compromised" when compared to Farquhar EML-2 Halo orbit ?

What are the other drawbacks of NRHO when compared to EML-2 / EML-1 halo orbits ?

Quote
I understand that NHRO is easy to use to get to other Lunar orbits. It is practically a transfer orbit between HLO and LLO with very small DV to change orbits. But as a more permanent orbit location it has many disadvantages.

As discussed earlier it is the fact that it takes less DV from Earth to reach a NHRO than L2 is the probably the main reason it is being picked because of SLS/Orion shortfalls when carrying a co-payload. Also NASA has yet to figure out exact how the DSG will ultimately be used.

Use also specifies the orbit. By picking NHRO initially the usage determination can wait until the DSG is actually orbiting around the Moon. A delayed Mars program means that Lunar surface becomes a higher priority and with LLO being more desirable, although the same could be said for L2 but that depends on the lander hardware designs used. An accelerated Mars program would make L2 a desirable orbit.

From the old Kirk Sorensen thread (an alternative lunar architecture)
- EML-2 "the Farquhar way" took 3.3 km/s but 8 days.
- EML-1 was faster (4 days) but took 3.8 km/s.

So acess to a NRHO would be around 3.6 km/s ?

...and we still don't know what Orion SM delta-v is: 1200, 1500 or 1800 m/s ? what is sure is that a) it is inferior to Apollo CSM 2500 m/s and b) if lower than 2000 m/s, it can't enter / exit LLO...

what is sure is that
- it all starts from a 3.15 km/s TLI before entering any cislunar orbit (libration, retrograd or LLO)
- then if Orion SM provides only 1200 m/s, this need to be split into two equal halves - to enter and to exit the orbit (think Apollo TEI)
- hence, 3.15 + 0.6 = 3.75 km/s

So the question is, where can Orion go with 3.75 km/s ? 
- Barely to L1 or L2 on low energy, long transit time trajectories.
- Not to LLO, for sure, since it would need 1 km/s IN, and 1 km/s OUT (Apollo TEI, once again)
- Hence I suppose that NRHO must be somewhere these two, around 3.6 km/s...
« Last Edit: 12/08/2018 12:21 pm by libra »

Offline Steven Pietrobon

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Re: Deep Space Gateway Power/Propulsion RFI
« Reply #128 on: 12/09/2018 08:41 am »
...and we still don't know what Orion SM delta-v is: 1200, 1500 or 1800 m/s ? what is sure is that a) it is inferior to Apollo CSM 2500 m/s and b) if lower than 2000 m/s, it can't enter / exit LLO...

We do know. Its 1.2 km/s. See Table 6 in paper below.

S. S. Pietrobon, "Fly me to the Moon on an SLS Block II," Int. Astronautical Congress, Adelaide, Australia, IAC-17-D2.8-A5.4, Sep 2017.
http://www.sworld.com.au/steven/pub/IAC17pap.pdf
Akin's Laws of Spacecraft Design #1:  Engineering is done with numbers.  Analysis without numbers is only an opinion.

Offline libra

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Re: Deep Space Gateway Power/Propulsion RFI
« Reply #129 on: 12/09/2018 03:46 pm »
Thank you Steven. This weekend I've downloaded a whole bunch of NRHO / DRO recent documents. I'll try to extract some numbers from them. Ideally, I'd like to have a detailed view of all the insertion delta-v into those different orbits - L1 Farquhar, L2 Farquhar, NRHO, DRO, LLO, L1 direct, L2 direct... the difference is delta-v are sometimes very small. The whole gang is stuck between 3.5 km/s and 4.2 km/s, from LEO - one way, of course.
« Last Edit: 12/09/2018 03:48 pm by libra »

Offline libra

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Re: Deep Space Gateway Power/Propulsion RFI
« Reply #130 on: 12/11/2018 04:17 pm »
My short list so far. If somebody has a valid number for NRHO, that person is welcome. Some numbers come from that old thread by Kirk Sorensen, an alternative lunar architecture...

TLI = 3150 m/s from LEO
TLI + 0 m/s > Zond, a lunar flyby.
Alternatively...
TLI + 1 m/s > to EML-1 or EML-2 (but it takes three months and a half... 103 days!)
TLI + 220 m/s >  to DRO
TLI + 330 m/s > to either EML-1 or EML-2 – powered lunar swingby, the Farquhar way
TLI + 450 m/s > to NRHO (uncertain)
TLI + 535 m/s > retrograde to EML-1, 13 days
TLI +  710 m/s > direct to EML-1
TLI +  829 m/s > NRHO (uncertain)
TLI + 900 m/s > LLO (Apollo 8, here we go)
TLI + 1100 m/s > direct to EML-2

There are some surprises and also some logic. For example, all four paths to EML-1 trade longer transit times against smaller delta-v (103 days, 21 days, 13 days, 4 days...)
Direct to EML-1 is far better than direct to EML-2, which is pretty bad, and actually worse than LLO. So all hail Lagrange and the three bodies problem. And all hail the late Robert Farquhar, the astrodynamics wizzard.

DRO is one of the best probably because it amounts to a capture orbit for the Moon - that is, the spacecraft is only in a lose, very high orbit. Only a little faster or farther, it would not be captured by the Moon and thus would make a Zond trajectory - a flyby. The limits of the Moon sphere of influence seems to be near EML-1 and EML-2, and DRO is not far away - 70 000 km above Earth satellite.
A NRHO is essentially a variant of a L2 halo orbit, except it is somewhat tweaked to get closer from the Moon and notably, to cover the South pole - Shackleton, Aitken, and those ice deposites to be turned into ISRU... someday.
As mentionned earlier in the thread, the issue with NRHO is that it lose a bit of L1 / L2 halos absolute flexibility - any time access from any place on the Moon (and on Earth, it seems).
« Last Edit: 12/11/2018 04:18 pm by libra »

Offline Steven Pietrobon

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Re: Deep Space Gateway Power/Propulsion RFI
« Reply #131 on: 12/12/2018 07:06 am »
TLI + 450 m/s > to NRHO (uncertain)

This is correct. See

https://forum.nasaspaceflight.com/index.php?topic=46645.msg1885927#msg1885927

Quote
TLI +  829 m/s > NRHO (uncertain)

This is not correct.
Akin's Laws of Spacecraft Design #1:  Engineering is done with numbers.  Analysis without numbers is only an opinion.

Offline libra

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Re: Deep Space Gateway Power/Propulsion RFI
« Reply #132 on: 12/12/2018 05:15 pm »
Thank you.

Updated list

TLI = 3150 m/s from LEO
TLI + 0 m/s > Zond, a lunar flyby.
Alternatively...
TLI + 1 m/s > to EML-1 or EML-2 (but it takes three months and a half... 103 days!)
TLI + 220 m/s >  to DRO
TLI + 330 m/s > to either EML-1 or EML-2 – powered lunar swingby, the Farquhar way
TLI + 450 m/s > to NRHO (no longer uncertain)  ;D
TLI + 535 m/s > retrograde to EML-1, 13 days
TLI +  710 m/s > direct to EML-1
TLI + 900 m/s > LLO (Apollo 8, here we go)
TLI + 1100 m/s > direct to EML-2

Makes a lot of sense. Being a variant of L2 / L1 & halo, NRHO logically fall right between them, delta-v wise.

TLI delta-v is incompressible, and so is descent / ascent to the lunar surface, a minimum of 2.5 km/s... delta-v is a harsh mistress.
« Last Edit: 12/12/2018 05:22 pm by libra »

Offline meberbs

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Re: Deep Space Gateway Power/Propulsion RFI
« Reply #133 on: 12/12/2018 08:04 pm »
The RFP for the PPE was released at the beginning of September, with proposals due in mid-November, and award expected in March 2019.

I have not seen any specific discussion on this here though. Has there been discussion somewhere that I missed? Are there any confirmations on which companies submitted proposals?

I assume the top candidates are the ones that had won NextStep-2 Appendix C awards:
•Boeing of Pasadena, Texas 
•Lockheed Martin of Denver, Colorado
•Orbital ATK of Dulles, Virginia
•Sierra Nevada Corporation’s Space Systems of Louisville, Colorado
•Space Systems/Loral in Palo Alto, California

For that matter, I wasn't able to find much discussion of the NextStep-2 contracts either, but I think there was a thread on that which I lost due to difficulties searching on this forum.

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