I wonder why they went with hydrazine RCS instead of xenon. SEP RCS+CMGs is now a well-proven combination. Not enough torque for a station this size? Or not enough for control during docking?
{snip}Is there anything about that image that makes you think it's a good idea or in any way useful? It's a hot mess in the middle of nowhere. Dump the circus act for a useful LLO and the electric power and propulsion requirements evaporate, as does funding requirements.
"Near rectilinear halo orbit" is a non-starter. [...] Dump the circus act for a useful LLO and the electric power and propulsion requirements evaporate, as does funding requirements.
"Near rectilinear halo orbit" is a non-starter. If you'd like to know more about the impossibility of station keeping, manned habitation and severe constraints on every mission plan it creates, NASA has plenty of papers - oddly pointing this out and concluding we should do it anyways. Much like the terrible idea of "Lagrange Gateways" that's been floated for years, now NASA pushes an even worse idea.Cislunar Near Rectilinear Halo Orbit for Human Space Exploration - Sept 2016https://ntrs.nasa.gov/search.jsp?R=20160003078Is there anything about that image that makes you think it's a good idea or in any way useful? It's a hot mess in the middle of nowhere. Dump the circus act for a useful LLO and the electric power and propulsion requirements evaporate, as does funding requirements.
Quote from: Propylox on 07/21/2017 03:09 am{snip}Is there anything about that image that makes you think it's a good idea or in any way useful? It's a hot mess in the middle of nowhere. Dump the circus act for a useful LLO and the electric power and propulsion requirements evaporate, as does funding requirements.1. Pretty picture.2. I thought low lunar orbits (LLO) were unstable, requiring between 0-400 m/s of station keeping.Putting the DSG in LLO would make lunar landing much easier.I suspect NASA wants to put the DSG in a high lunar orbit to reduce the delta-v needed by the large mass transfer vehicle to go to Mars. However if the DSG only has a 15 year life expectancy then the Mars Transfer Vehicle (MTV) will still be under construction at the end of the Gateway's life. DSG #2 may be the return point for the MTV.
Quote from: A_M_Swallow on 07/21/2017 04:18 amQuote from: Propylox on 07/21/2017 03:09 am{snip}Is there anything about that image that makes you think it's a good idea or in any way useful? It's a hot mess in the middle of nowhere. Dump the circus act for a useful LLO and the electric power and propulsion requirements evaporate, as does funding requirements.1. Pretty picture.2. I thought low lunar orbits (LLO) were unstable, requiring between 0-400 m/s of station keeping.Putting the DSG in LLO would make lunar landing much easier.I suspect NASA wants to put the DSG in a high lunar orbit to reduce the delta-v needed by the large mass transfer vehicle to go to Mars. However if the DSG only has a 15 year life expectancy then the Mars Transfer Vehicle (MTV) will still be under construction at the end of the Gateway's life. DSG #2 may be the return point for the MTV.Has nothing to do with MTV
The other item I noticed was max weight for this co-manifested payload showing there is likely going to this orbit with Orion on a SLS-1B only 7.5mt of excess capability. Which also means that the DSG itself will have to weigh no more than 7.5mt. This orbit was probably picked because for other orbits there just was not enough excess capability to have a useful co-manifested payload.
Quote from: oldAtlas_Eguy on 07/21/2017 04:57 pmThe other item I noticed was max weight for this co-manifested payload showing there is likely going to this orbit with Orion on a SLS-1B only 7.5mt of excess capability. Which also means that the DSG itself will have to weigh no more than 7.5mt. This orbit was probably picked because for other orbits there just was not enough excess capability to have a useful co-manifested payload.Yes mass restrictions seem to be popping up and causing development problems, see links below: http://russianspaceweb.com/imp-ppb.html#2017http://russianspaceweb.com/imp-lcub.html
NASA is pushing quite an aggressive schedule, delivery in 2021 for launch in 2022 for propulsion module. Habitation module to follow the following year.
Quote from: Khadgars on 07/21/2017 05:47 pmNASA is pushing quite an aggressive schedule, delivery in 2021 for launch in 2022 for propulsion module. Habitation module to follow the following year. Can this really be done by 2021? This looks like the "hard" part of the Gateway, with the most powerful SEP system ever, no? Four years to delivery, and no contracts yet? I could believe a docking node or hab module in 4 years, if there was a lot of money, but the propulsion module seems a lot more ambitious technologically.
Quote from: jgoldader on 07/21/2017 08:40 pmQuote from: Khadgars on 07/21/2017 05:47 pmNASA is pushing quite an aggressive schedule, delivery in 2021 for launch in 2022 for propulsion module. Habitation module to follow the following year. Can this really be done by 2021? This looks like the "hard" part of the Gateway, with the most powerful SEP system ever, no? Four years to delivery, and no contracts yet? I could believe a docking node or hab module in 4 years, if there was a lot of money, but the propulsion module seems a lot more ambitious technologically.Probably helps that its almost literally the latest design of ARM without any of the grabbing equipment and an extra docking port.
ARM had a docking port for Orion. PPE would just need the one. Right?
Quote from: Propylox on 07/21/2017 03:09 am"Near rectilinear halo orbit" is a non-starter. [...] Dump the circus act for a useful LLO and the electric power and propulsion requirements evaporate, as does funding requirements.Is part of it about wanting a big SEP project in a post-ARM world? Back-solving from the tech they want to fund?
What use is DSG at LLO when no one can visit it. Orion could make it to LLO but it would be oneway trip.
The problem is Orion cannot enter and then exit LLO, so no infrastructure using Orion will be built or moved there.
I thought low lunar orbits (LLO) were unstable, requiring between 0-400 m/s of station keeping.Putting the DSG in LLO would make lunar landing much easier.
7.5 mT is the weight of each module. The current plan is for 4 modules.https://www.nasaspaceflight.com/2017/04/nasa-goals-missions-sls-eyes-multi-step-marsIMHO A small spacestation containing 4 tiny modules to be lifted by 4 SLS and assembled by 4 Orions appears excessive.
... But this PPE design is useful for other uses than just the DSG.So it may have a life of it's own regardless of SLS/Orion or even DSG.
If the EC Lander is delayed then the Habitat could be 2 years earlier but would also require a human rated habitat to be completed/developed in 6 years from now. That is really pushing it since the habitat has yet to even be on contract, not even close to a PDR point (probably at least 3 years away 2020) , CDR a year latter (2021) then 3 years to build and certify the habitat (2024) then a year later launch (2025). So SLS flt #4 could be as late as 2025 if EC#2 (Lander) development is delayed into second half of 2020's.
Quote from: oldAtlas_Eguy on 07/22/2017 01:50 am... But this PPE design is useful for other uses than just the DSG.So it may have a life of it's own regardless of SLS/Orion or even DSG.I'd counter this PPE may have a thousand uses, but DSG ain't one.Manned habitation, docking attempts, or any sort of mission won't be occurring from that orbit. Meaning DSG will be somewhere else while PPE floats about. If/when DSG is planned for a useful orbit, the electric propulsion system this PPE sports won't be necessary or wanted. Meaning this PPE won't be part of DSG, though that's how NASA is selling it. You buying?
"Near rectilinear halo orbit" is a non-starter. If you'd like to know more about the impossibility of station keeping, manned habitation and severe constraints on every mission plan it creates, NASA has plenty of papers - oddly pointing this out and concluding we should do it anyways. Much like the terrible idea of "Lagrange Gateways" that's been floated for years, now NASA pushes an even worse idea.Cislunar Near Rectilinear Halo Orbit for Human Space Exploration - Sept 2016https://ntrs.nasa.gov/search.jsp?R=20160003078
Quote from: Propylox on 07/21/2017 03:09 am"Near rectilinear halo orbit" is a non-starter. If you'd like to know more about the impossibility of station keeping, manned habitation and severe constraints on every mission plan it creates, NASA has plenty of papers - oddly pointing this out and concluding we should do it anyways. Much like the terrible idea of "Lagrange Gateways" that's been floated for years, now NASA pushes an even worse idea.Cislunar Near Rectilinear Halo Orbit for Human Space Exploration - Sept 2016https://ntrs.nasa.gov/search.jsp?R=20160003078I was only able to read the abstract of that paper but from what I read it doesn't seem that the authors are declaring NRO orbits impossible or worthless. They are pointing out that NROs have never been used for human exploration and that different orbital models will be required. It's a challenge but not an impossible one. Attached is another paper that I ran across in NSF a while ago that is quite positive on the usage of NROs for human exploration.
Literally this PPE is a "DEEP SPACE TUG". It has advanced docking adapters for in-space attachment of any "payload" and sufficient DV to push some significant sized items around. As I said before and probably is the reason NASA wants to build it in the first place is that this PPE has a multitude of uses outside of just the SLS/Orion/DSG program. Also with its fore and aft docking it can be stacked with multiple PPE to create a very large outer planetary DV delivery system. Think of mating this PPE to a dedicated Europa Lander. This then makes the EC Lander a simpler design such that the power and propulsion is designed leaving only the communication and experiment packages. Also if this vehicle was to also have a significant communication relay capability then the payload no longer need that either. This module then becomes if used to send stuff to Mars a orbital communication relay with its very high power solar arrays capable of multiple high data rate channels for multiple ground assets and links to Earth. 24KW is 2X the power used on HTS comm sats.Added:A question then is the intent to make this PPE have a diameter when stowed such that it could fit in a 5m fairing? If so then it definitely has a future for use in the planetary programs. Much less any HSF programs that eventually get approved regardless of SHLV that is used in that program.I like this vehicle. It shows some forethought into a "LEGO" in-space methodology. Many have thought that the customized each vehicle method has always been the wrong way to go for shortening the development time and development costs. A more "LEGO" approach where stuff is just docked together on the ground at launch or even in-space docking using smaller LV's gives a large set of options for programs to choose quicker and lest costly development paths. This will also help the DSG in that it offloads much of the design problems of the DSG into logical "LEGO" pieces that can be individually tested and improved/replaced if needed.
Quote from: Endeavour_01 on 07/23/2017 06:25 pmQuote from: Propylox on 07/21/2017 03:09 am"Near rectilinear halo orbit" is a non-starter. If you'd like to know more about the impossibility of station keeping, manned habitation and severe constraints on every mission plan it creates, NASA has plenty of papers - oddly pointing this out and concluding we should do it anyways. Much like the terrible idea of "Lagrange Gateways" that's been floated for years, now NASA pushes an even worse idea.Cislunar Near Rectilinear Halo Orbit for Human Space Exploration - Sept 2016https://ntrs.nasa.gov/search.jsp?R=20160003078I was only able to read the abstract of that paper but from what I read it doesn't seem that the authors are declaring NRO orbits impossible or worthless. They are pointing out that NROs have never been used for human exploration and that different orbital models will be required. It's a challenge but not an impossible one. Attached is another paper that I ran across in NSF a while ago that is quite positive on the usage of NROs for human exploration.The one distinct advantage of L2 is that launch windows to that destination are not overly complex. But NRO's require specific timing of the object in the NRO with the Earth's rotation which could be highly restrictive. But for SLS that is unlikly to launch more than once a year that is not really a concern. But for an active continuously manned DSG where commercial services are resupplying and possibly even delivering crews such orbits would represent significant launch scheduling conflicts and other possible lengthy delays when a window is missed.
We found that in order to minimize Orion propellent usage, the optimizer was adjusting the outboundtrip times to keep the arrival in and departure from the NRHO near the favorable regions ofthe NRHO for those maneuvers. In terms of Orion propellant used, the rendezvous missions wouldapproach the performance of the free-phase missions once per NRHO period. At these points nearthe phase match, there would typically be from 3 to 5 consecutive feasible rendezvous missionopportunities, with the best approaching the performance of the free-phase cases (see Figure 13).For the short stay missions examined, this means that there would be multiple sets of launch opportunitieseach month, with each set spanning 3 to 5 consecutive days. The results also indicatethat, at least broadly, the previous free-phase results can be used to gain insight into the generalperformance situation for fixed-phase trajectories.
I cannot see any mention of the fairing size in the RFI. So if an aerospace firm can design the PPE to fit into a 5m fairing they may get NASA to pay for development of their SEP tug.
Isn't it a rather *small* SEP tug at 7500 kg? Launching it together with an Orion puts some severe limitations on it.Also, isn't the whole point of a SEP tug to move payloads around? I'd expect a real tug to continuously carry payloads from LEO, perhaps synchronized with cargo flights. But it seems the plan is for this to sit in the same orbit for years.I'd rather describe this as a small space-station core module, providing power and station-keeping.
QuoteWe found that in order to minimize Orion propellent usage, the optimizer was adjusting the outboundtrip times to keep the arrival in and departure from the NRHO near the favorable regions ofthe NRHO for those maneuvers. In terms of Orion propellant used, the rendezvous missions wouldapproach the performance of the free-phase missions once per NRHO period.https://ntrs.nasa.gov/search.jsp?R=20170001352These are launch opportunities with Orion towing a 10 ton module to rendezvous with another object already in that orbit.
We found that in order to minimize Orion propellent usage, the optimizer was adjusting the outboundtrip times to keep the arrival in and departure from the NRHO near the favorable regions ofthe NRHO for those maneuvers. In terms of Orion propellant used, the rendezvous missions wouldapproach the performance of the free-phase missions once per NRHO period.
Quote from: A_M_Swallow on 07/24/2017 02:41 amI cannot see any mention of the fairing size in the RFI. So if an aerospace firm can design the PPE to fit into a 5m fairing they may get NASA to pay for development of their SEP tug.NASA is paying for the development of a SEP tug, that is what PPE is.
Correct. I have met salesmen and company directors who think because it is called a PPE that the machine cannot also be a SEP tug. They would expect the second machine to be designed from scratch.
Quote from: ncb1397 on 07/24/2017 05:43 pmQuoteWe found that in order to minimize Orion propellent usage, the optimizer was adjusting the outboundtrip times to keep the arrival in and departure from the NRHO near the favorable regions ofthe NRHO for those maneuvers. In terms of Orion propellant used, the rendezvous missions wouldapproach the performance of the free-phase missions once per NRHO period.https://ntrs.nasa.gov/search.jsp?R=20170001352These are launch opportunities with Orion towing a 10 ton module to rendezvous with another object already in that orbit. Fantastic. This implies there are a few opportunities each month to descend to the surface or to schedule a return from the surface. If there's a surface emergency, they'll just have to die waiting for orbital alignment. If an emergency arises while in orbit, can we safely say only half the orbit allows direct to Earth-return. The other half is also death? Great plan.
NRHO gives DSG days over one pole per orbit allowing for direct line of sight into that poles craters for hours if not days. Ideal for communicating with assets (rovers, landers) in those craters and maybe beaming power using laser. NRHO is just one of few orbits that DSG can use, between Orion missions it can shift to another.
1) But I'm having trouble evaluating how well this would work if used as an actual SEP tug between LEO and NRHO. Assuming 9000 ISP (snip)2) But ideally you would want a SEP tug capable of transferring payloads as large as you can place in LEO, right?
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.
--- continued on ion architecture ---Unfortunately there's only assumptions and hurdles with high-power applications for all three ion options mentioned. The PPE is a puny design, both in power levels and capabilities, that doesn't address any current and future needs or advance technology that is needed. And that orbit. - HET: Material science currently limits cathodes to around 30-50 amps without rapid degradation. It's why I suggested a variable voltage to use high amps/thrust deep in gravity wells and plane changes, then switch to high voltage/isp for spiraling. This preserves the engine, increases efficiency and transit time.Additionally a grid, or cluster of multiple HETs arc across each other, rapidly destroying one at a time. PPE uses broadly-spaced, low power HETs to avoid this while tests have used external magnetic containment poles to isolate each HET. Packing 20-30 HETS of 30-50kW together for a viable SEP tug seems highly problematic. - VASIMR: Weight and reliability are the major questions. This architecture requires active cooling - possibly regenerative, but that may not be enough. If not there's additional system weight and reliability questions.VASIMR's also never done long-duration testing like was planned aboard ISS to prove reliability. This may be a great design, but final operating parameters and design needs to be proven. - HiPEP: While ~9,ooos isp and ~40kW was shown, the efficiency study I linked was to test at up to 16kW for a 25kW mission, but never neared that. It focused on temperatures, coupling, efficiency and degradation around 1kW. I'd guess a final design wouldn't be near 9,ooos by increasing kg/s to keep it cool and reliable.
{snip}4) Adding some TDRSS comm system items to the PPE would increase its usefulness for many missions in the future. Such as a comm relay system for far side of the Moon surface missions. But this also adds those pesky large antennas.
This sounds like more of a 5 year development program than just a 3 year build program. In a three year build program as being requested (delivery date of 2021 and contract start sometime in FY2018), there is no time to develop technology.
With 2 IDSS in requirement, plus SEP... how that will be configured?Assuming SEP on the long-axle, and 1st IDSS on the other end, that will make 2nd IDSS mounted on the side?How that translated to center-of-mass for station keeping boost? SEP on gimbals?Attached image is from Orbital ATK's concept video, which have 2 IDSS on 2 end of long-axle. Small thrusters next to one of IDSS, which doesn't seem functional when that IDSS been used.Or we gonna see Orbital ATK MEV-style HET mounted on extended arm? Which I think is a smart design to somehow adjust the thrust alignment by moving extended arm.Titus
BL's copy of the RFI - #7: These berthing locations will support unpressurized logistics ... and #9b: The PPE will provide a minimum translation path for EVA.I'm assuming that's a tunnel for unpressurized cargo, possibly large enough to pass through or am I misinterpreting "translation path"?
Its the last one you described. Per the renderings from NASA which match the last version of the ARM bus, the SEP thrusters are mounted on moveable arms around the IDSS port (for ARM, Orion was to dock at this port) but most of the gateway seems to be set up to be on the other end of the bus giving it a more traditional-looking configuration. I did a quick render here to show the arrangement in the aft compared with the (more exposed) CAD diagrams of the area from NASA. I do not know if the bus is pass-through (basically a tunnel, like Cygnus DS) or if the IDSS ports are just hard points for moving it around using any compatible vehicle- I'm leaning towards just hard points though simply because I cannot find any information about a tunnel in the design. There is some word of a small science airlock at the "front end" of the bus, but thats from much earlier international discussions and not any NASA documents.
I did a quick render here to show the arrangement in the aft compared with the (more exposed) CAD diagrams of the area from NASA.
Quote from: okan170 on 08/01/2017 04:56 pmI did a quick render here to show the arrangement in the aft compared with the (more exposed) CAD diagrams of the area from NASA.What is the source of that CAD picture?
I revisit the DSG pdf, and realize in the plan drawing, the only module connected with PPE is Hab (using rear port of Hab). Since Hab has 3 ports in total, front/rear/side, that leave 2 ports available to be used, not blocking station keeping thruster....
HEO Committee power point on Future Exploration Plans by Greg Williams has been posted, including several slides on the PPE https://www.nasa.gov/sites/default/files/atoms/files/nac_exploration_july_2017_4-2.pdf
Quote from: BrightLight on 08/01/2017 02:44 pmHEO Committee power point on Future Exploration Plans by Greg Williams has been posted, including several slides on the PPE https://www.nasa.gov/sites/default/files/atoms/files/nac_exploration_july_2017_4-2.pdf Is PPE (and ARM) still unfunded in NASA's 2018 budget, suggesting this RFI and assumptions about a DSG in NRHO are merely castoff program dreams? Is it better to wait until a new, actually viable plan is put forth to speculate on DSG and Lunar operations?
Quote from: Propylox on 08/05/2017 03:34 pmQuote from: BrightLight on 08/01/2017 02:44 pmHEO Committee power point on Future Exploration Plans by Greg Williams has been posted, including several slides on the PPE https://www.nasa.gov/sites/default/files/atoms/files/nac_exploration_july_2017_4-2.pdf Is PPE (and ARM) still unfunded in NASA's 2018 budget, suggesting this RFI and assumptions about a DSG in NRHO are merely castoff program dreams? Is it better to wait until a new, actually viable plan is put forth to speculate on DSG and Lunar operations?Not everything that NASA launches has specific funding. Space technology and Advanced Explorations Systems combined have a billion dollar per year budget with NASA having pretty wide latitude with how it is spent(besides the portion of that budget that Congress directs to specific items). It is how they worked on ARM with Congress against it.
Purely FWIW, does the 2026 configuration of DSG remind anyone else of Skylab? Maybe they can call it 'Moonlab'? (Yes, I know, I know, my Steven Baxter amazing people is showing again!)
Thanks for the document.There are a few items that requires mention. This PPE will be the primary communications device for the DSG as long as it is attached to the other DSG elements. It also provides the battery power during solar array eclipses. So the other DSG elements would have very minimal batteries and communications. Unfortunately the weights and other capabilities for orbit maneuvers and RCS are TBD. There is no mention of a air lock or capability of moving through the two docking ports by personnel. So this document does not require that these ports be anything more than a place holder for other elements that provide for power, communications, and prop transfer but nothing more than a structural attach point. No crew egress. That makes a small problem with the Orion in that it cannot dock at the PPE and still do EVA or crew swap. This makes the supply procedures more complex for the DSG. Requiring the supply VV to dock at a crew access port first to be unloaded then undock and move to the PPE to transfer propellant. This would have to be done unless the other DSG elements all have additional piping and valves to be able to transfer prop through these other elements from the VV to the PPE.
Quote from: oldAtlas_Eguy on 08/18/2017 06:50 pm...( Is it counter to politeness here to use ellipsis to show text has been removed? - tdperk)I suspect that the DSG will be refuelled by unmanned cargo vehicles. Since the PPE's fuel tank can only take 2000 kg of propellant the Commercial Resupply Service DSG (CRS-DSG) can be performed in several ways. Most cheaper than Orion on SLS.
...( Is it counter to politeness here to use ellipsis to show text has been removed? - tdperk)
Quote from: A_M_Swallow on 08/19/2017 09:36 pmQuote from: oldAtlas_Eguy on 08/18/2017 06:50 pm...( Is it counter to politeness here to use ellipsis to show text has been removed? - tdperk)I suspect that the DSG will be refuelled by unmanned cargo vehicles. Since the PPE's fuel tank can only take 2000 kg of propellant the Commercial Resupply Service DSG (CRS-DSG) can be performed in several ways. Most cheaper than Orion on SLS.If NASA for some reason wants to continue avoiding necessary technology development, then instead of mastering bulk liquid fuel transfer in weightlessness, a propulsion module which docks to a DSG or other structure to maneuver it and undocks to permit a fresh one to attach should be developed. That propulsion module should include or be compatible with a return capability for refurbishment.
Quote from: tdperk on 08/20/2017 03:25 pmQuote from: A_M_Swallow on 08/19/2017 09:36 pmQuote from: oldAtlas_Eguy on 08/18/2017 06:50 pm...( Is it counter to politeness here to use ellipsis to show text has been removed? - tdperk)I suspect that the DSG will be refuelled by unmanned cargo vehicles. Since the PPE's fuel tank can only take 2000 kg of propellant the Commercial Resupply Service DSG (CRS-DSG) can be performed in several ways. Most cheaper than Orion on SLS.If NASA for some reason wants to continue avoiding necessary technology development, then instead of mastering bulk liquid fuel transfer in weightlessness, a propulsion module which docks to a DSG or other structure to maneuver it and undocks to permit a fresh one to attach should be developed. That propulsion module should include or be compatible with a return capability for refurbishment.Too complex, and easier to just ship 1-2mt of prop up with a regular commercial re-supply vehicle.
After the technology of bulk weightless liquid transfer is established it would be more complex, maybe.And I have had people tell me the reason SpaceX cannot possibly succeed with the ITS/BFT is that no one knows how to transfer liquids in bulk--so that's a showstopper. That we cannot count on that task being handled.
Quote from: tdperk on 08/21/2017 12:29 amAfter the technology of bulk weightless liquid transfer is established it would be more complex, maybe.And I have had people tell me the reason SpaceX cannot possibly succeed with the ITS/BFT is that no one knows how to transfer liquids in bulk--so that's a showstopper. That we cannot count on that task being handled. The technology is well established. They already do bulk propellant transfer for ISS.
Orion does not have enough propulsive capability to enter and return from LLO.
Quote from: JacobLutz7 on 01/20/2018 06:23 pmOrion does not have enough propulsive capability to enter and return from LLO.I actually requested more information on the Orion delta-v front in another thread, the discussion on which starts here.
This is from memory havn't confirmed it, Orion is about 1800m/s.
The attached NASAfacts sheet from 2011 indicates a delta-V of 4920 ft/s, i.e., 1500 m/s.The mass to orbit is quoted as 50,231 lbm, while the SM's propellant load is 17,433 lbm, giving a mass ratio of 1.5315, assuming negligible propellant residuals. The delta-V of 1500 m/s then in turn implies an effective exhaust velocity of 3518.9 m/s, i.e., a specific impulse of 359 s, which seems unlikely for storable propellants.I think what's missing in this analysis is that some propellant is burned on the way to orbit, since SLS places Orion only into a transfer orbit with a very low perigee.
You're confusing the SLS Core stage with iCPS/EUS. The former stages slightly suborbital, the latter reaches a circular parking orbit first and then performs TLI, and then completes a disposal burn afterwards
Looks like in the budget proposal from a couple weeks back the Deep Space Gateway has been renamed the "Lunar Orbital Platform", the propulsion module now has funding attached to it ($504 million next year, $2.7 billion over five years), and the targeted launch date is actually being moved up to 2022. It is also now is planned for launch on a commercial vehicle instead of on the EM-2 SLS flight.
Quote from: Toast on 02/26/2018 11:27 pmLooks like in the budget proposal from a couple weeks back the Deep Space Gateway has been renamed the "Lunar Orbital Platform", the propulsion module now has funding attached to it ($504 million next year, $2.7 billion over five years), and the targeted launch date is actually being moved up to 2022. It is also now is planned for launch on a commercial vehicle instead of on the EM-2 SLS flight.That's a steep 'initial' price tag for an SEP tug.What part of the technology is so expensive?
Quote from: AncientU on 02/27/2018 10:55 amQuote from: Toast on 02/26/2018 11:27 pmLooks like in the budget proposal from a couple weeks back the Deep Space Gateway has been renamed the "Lunar Orbital Platform", the propulsion module now has funding attached to it ($504 million next year, $2.7 billion over five years), and the targeted launch date is actually being moved up to 2022. It is also now is planned for launch on a commercial vehicle instead of on the EM-2 SLS flight.That's a steep 'initial' price tag for an SEP tug.What part of the technology is so expensive?No part. It's just that it will be done "NASA-style", much like SLS and Orion.Remember: gravy train...
The SLS upper stage is hydrolox, isn't it impossible for it to last 3 days and relight for orbit insertion? I know ULA's ACES claims to be able to do it but that's only after they develop IVF.I also remember reading that NRHO was picked partly because of limited Orion delta-v, is that true?On a related note, isn't it also very difficult to bring a SEP craft to lunar orbit Moon using only its own power? Switching the PPE to a commercial launch brings many new issues.
On a related note, isn't it also very difficult to bring a SEP craft to lunar orbit Moon using only its own power? Switching the PPE to a commercial launch brings many new issues.
Quote from: DreamyPickle on 02/27/2018 07:25 pmOn a related note, isn't it also very difficult to bring a SEP craft to lunar orbit Moon using only its own power? Switching the PPE to a commercial launch brings many new issues.It's certainly slower than using chemical propulsion as far as lunar-orbit insertion, but there's no rush. SMART-1 went as far as GTO on chemical propulsion and from there went electric.
Forty-two minutes after launch SMART-1 was placed into a geostationary transfer orbit, 742 x 36 016 km, inclined at 7° to the Equator.
The last firing of the EP before lunar capture was a thrust arc around the third lunar resonance, and ended on 14 October 2004. Apart from a 4 hour correction burn on 25 October, the EP remained inactive until lunar capture. Up to 26 October, and the 289th engine pulse, the SMART-1 EP system had cumulated a total on time of nearly 3650 hours, consumed about 59 kg of xenon and imparted to the spacecraft a velocity increment of approximately 2735 ms-1 (9850 kmh-1).
Much like the terrible idea of "Lagrange Gateways" that's been floated for years, now NASA pushes an even worse idea.
Why are "Lagrange Gateways" a terrible idea?
Quote from: Propylox on 07/21/2017 03:09 am Much like the terrible idea of "Lagrange Gateways" that's been floated for years, now NASA pushes an even worse idea.Why are "Lagrange Gateways" a terrible idea?
Quote from: JacobLutz7 on 01/20/2018 06:23 pmOrion does not have enough propulsive capability to enter and return from LLO. Yes, that's true. It does have enough delta-V to leave LLO though. An EUS designed to last the three day journey to the Moon could do LLO insertion with Orion (as well as an LM on a separate mission).
Orion doesn't have the delta-V to return from lower lunar orbits, ...
And SLS doesn't have anywhere near the launch rate to support a traditional crewed Mars mission architecture like DRM 5.0, so NASA will have to rely on an unproven, high-power electric transit stage (the Deep Space Transport) to get to Mars with fewer launches.
b) we've already been to the moon!
Of course, another rocket would need to place Orion in LLO for this to work.
You can do a Mars mission from LEO using only three SLS Block II launches (first launch carries Orion, Hab and Lander/MAV in re-entry fairing, second and third launches refuel SLS upper stage for TMI).
NASA's gateway plan requires a lot more launches than that.
Quote from: UltraViolet9 on 03/03/2018 07:46 pmOrion doesn't have the delta-V to return from lower lunar orbits, ...Yes it does. Orion has a delta-V of 1.2 km/s. That's sufficient for TEI from LLO. Of course, another rocket would need to place Orion in LLO for this to work.QuoteAnd SLS doesn't have anywhere near the launch rate to support a traditional crewed Mars mission architecture like DRM 5.0, so NASA will have to rely on an unproven, high-power electric transit stage (the Deep Space Transport) to get to Mars with fewer launches.Have you actually compared the number of SLS launches for both architectures? You can do a Mars mission from LEO using only three SLS Block II launches (first launch carries Orion, Hab and Lander/MAV in re-entry fairing, second and third launches refuel SLS upper stage for TMI). NASA's gateway plan requires a lot more launches than that.Quoteb) we've already been to the moon!We've also been to LEO, but we still keep going there. That's because its worthwhile to do so. Same with the Moon.
But we're using it's kissing cousin in DSG/LOP-G as a point solution band-aid to keep a deficient exploration architecture from falling apart.
Quote from: UltraViolet9 on 03/03/2018 07:46 pmOrion doesn't have the delta-V to return from lower lunar orbits, ...Yes it does. Orion has a delta-V of 1.2 km/s. That's sufficient for TEI from LLO. Of course, another rocket would need to place Orion in LLO for this to work.{snip}
Quote from: Steven Pietrobon on 03/04/2018 04:59 amQuote from: UltraViolet9 on 03/03/2018 07:46 pmOrion doesn't have the delta-V to return from lower lunar orbits, ...Yes it does. Orion has a delta-V of 1.2 km/s. That's sufficient for TEI from LLO. Of course, another rocket would need to place Orion in LLO for this to work.{snip}A depot in LLO could refuel the Orion's service module.
Quote from: A_M_Swallow on 03/04/2018 04:21 pmQuote from: Steven Pietrobon on 03/04/2018 04:59 amQuote from: UltraViolet9 on 03/03/2018 07:46 pmOrion doesn't have the delta-V to return from lower lunar orbits, ...Yes it does. Orion has a delta-V of 1.2 km/s. That's sufficient for TEI from LLO. Of course, another rocket would need to place Orion in LLO for this to work.{snip}A depot in LLO could refuel the Orion's service module.In theory yes, but adds another potential point of failure. Having crew stuck in LLO because of refuelling failure is not good outcome.
Round trip for lander is 5.5km/s for NRO compared to 3.7km/s for LLO so extra 1.8km/s. If DSG is already in place better to design lander to stage from.Best to wait a few years and see what comes out of robotic exploration. With ISRU lunar refuelling lander only needs to be capable of 2.7km/s. Can be LH LOX as boil off is not an issue over a day. The extra development costs of 5.5km/s lander compared to 2.7km/s would go long way to help pay for small ISRU plant.
Would that EDS also brake the group of spacecraft into Martian orbit, and return them later to Earth? Or would there have to be a Earth Return Stage(s), waiting there for them to send the Orion and Habitat back?
Quote from: TrevorMonty on 03/04/2018 05:43 pmQuote from: A_M_Swallow on 03/04/2018 04:21 pmQuote from: Steven Pietrobon on 03/04/2018 04:59 amQuote from: UltraViolet9 on 03/03/2018 07:46 pmOrion doesn't have the delta-V to return from lower lunar orbits, ...Yes it does. Orion has a delta-V of 1.2 km/s. That's sufficient for TEI from LLO. Of course, another rocket would need to place Orion in LLO for this to work.{snip}A depot in LLO could refuel the Orion's service module.In theory yes, but adds another potential point of failure. Having crew stuck in LLO because of refuelling failure is not good outcome.You could always refuel in NRHO. Orion would use about 3300 kg of propellant to get into NRHO, be topped off to go to LLO and back (~1400-1500 m/s). It might be at a slight ~100 m/s deficit compared to a benchmark 100 km altitude circular orbit. Then it would just have to be fueled with about 2500 kg of propellant for the trip back to earth. Alternatively, it could be directed back to earth(TEI from LLO is about the same as LLO to NRHO(~750-800 m/s).Or you could launch something else in the USA with a docking port to attach to the Orion. A lot of things would work to make up the modest 500 m/s deficit that Orion has to enter LLO and return to earth. A second SM with a docking port kit, a lunar descent stage equivalent in performence to th LEM descent module or even a ~8 mT chemical propulsion comsat with about 4200 kg of propellant(that it would need for GEO circularization - ~1.6 km/s, stationkeeping at 50 m/s per year anyway). The Block 1B is so overpowered for putting Orion alone into LLO that all kinds of not necessarily efficient schemes could work. 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
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/
“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"
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
While speculation is fun, I prefer to know what I'm talking about.
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.
Quote from: Joseph Peterson on 03/08/2018 05:15 amWhile speculation is fun, I prefer to know what I'm talking about.Always a good idea!QuoteI 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.
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
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.
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.
...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...
TLI + 450 m/s > to NRHO (uncertain)
TLI + 829 m/s > NRHO (uncertain)
Due to the shutdown, the contract start date has moved to May 31.Also, I found these NAC HEO slides from December which indicate more than one partner may be selected. If that had been mentioned before, then I missed it.