SLS Rocket + Orion + Apollo LEM

Impaler · 2014-12-26 19:49:28

RobertDyck wrote:

I have a radical idea: Mars
Since before Obama was elected, I said the Moon is "Been there, done that." President Obama himself repeated those words. But if you want to crunch the numbers, I can do that. And this demonstrates that Orion is optomized for one misison only; it is not "Multi-Purpose Crew Vehicle". And media announcements during the Orion test launch called it the ship that will take humans to Mars. Ha! Never going to happen. Although it's practically the same mass as Apollo, it's way to heavy for a capsule. We discussed a couple alternatives for Mars in other threads.

Mars is non-sequitur in this thread as were talking about moon and implicitly the shorter time-frame of a decade which rules out Mars.

I have no disagreement with the Orion not going to Mars, the only viable missions for it are cis-lunar so that's why people are trying to find a lander, I'm just saying if that lander is not more functional then the Apollo LEM then forget about it. We are not going to fund a billion dollar moon program that makes carbon-copies of the Apollo landings. Remember the original LEM could only land at the Lunar equator, all the excitement is centered on polar now, being able to land at any latitude was one of the selling points of Altair a decade ago and polar interest has only increased since then. Do you disagree with any of that?

RobertDyck · 2014-12-26 20:59:09

Not exactly equitorial.

To be picky, it was called "Lunar Module", acronym was "LM". Media said "LeM" to make it pronounceable.

The Apollo LM was designed for 5 decompress/recompress cycles, so limited exploration. And only carried 2 astronauts. But I don' t see any reason it couldn't land at a pole.

Last edited by RobertDyck (2014-12-26 21:33:49)

Impaler · 2014-12-26 21:39:08

Don't nit-pick, it makes you look like an ass, Apollo was confined to ~30 degrees above or below the equator (don't remember the exact amount off the top of my head) as part of the vehicle and mission SPEC, that was the amount of plane change they could do cause plane-change is expensive in terms of Delta-V. When I say equator any person should be able to understand I mean equatorial. For someone who crows about how great Apollo was I'm surprised your unaware of this limitation.

To get a polar landing you either need more Delta-V on the lander for two big plane changes or the whole capsule and service module, lander stack need to go into polar orbit of the moon on initial capture into lunar orbit and because we leave the Earth from an orbit that is closer to Equatorial then it it polar it would cost more Delta-V to go into a polar orbit. Orion service module is already in the works and I don't think it's going to be capable of that. Some kind of low-energy transfer orbit might manage to squeeze out the performance but it's by no means guaranteed.

My POINT still stands that you can't do the minimum mission with an old LEM (yes I'm going to keep calling it that).

Last edited by Impaler (2014-12-26 21:42:20)

RobertDyck · 2014-12-26 22:02:21

Yes, initial capture into lunar polar orbit. Do that and you can land at a pole. 1960s tech may have considered that difficult, but I don't see why today it would be any more difficult than equitorial. Lunar Prospector already did that, initially captured into polar orbit. Same delta-V as equitorial, same communications, same everything. Landing from equitorial orbit means the Moon is rotating in the same direction as your orbit, but landing as much off equator as Apollo 15 requires some navigation. Landing right at the pole should be easier. The only difficulties would be surface light and surface temperature.

RGClark · 2014-12-27 03:53:15

RobertDyck wrote:

Well, NASA has been working on a deployable fabric heat shield. Let them do it. I thought of using Nextel 440 because that's the fabric of DurAFRSI, an advanced thermal blanket developed by Ames Research Centre. But NASA is working on a carbon fibre parasol. Current focus is for Venus. If it works, then why not use it for aerocapture at Earth. That means return from either the Moon or Mars. And my intention is reusable.
Adaptive Deployable Entry and Placement Technology (ADEPT): Progress in Payload Separation Risk Mitigation for a Deployable Venus Heat Shield
Appears to be published in 2013, in a journal of American Institute of Aeronautics and Astronautics. This paper is hosted by the NASA technical report server.

Thanks for that link, but I wasn't able to find the weight of this deployable heat shield there. For this to be useful it will have to beat the ablative Apollo style heat shields. These were about 15% of the landed mass. Actually the deployables would have to beat the SpaceX PICA-X heat shields estimated at half the weight of the Apollo heat shields which would put them at about 8% of the landed mass.

Bob Clark

RobertDyck · 2014-12-27 07:51:09

New document for ADEPT. This is a slide presentation in .PDF format. File with NASA's technical report server on November 17 of this year.
Deployment Testing of the ADEPT Ground Test Article
Caution: they optimized it for Venus. I'm saying this technology could be used for Earth or Mars, but may have to be modified. After all, this looks exactly like the deployable heat shield for the habitat in Mars Direct. Mars Direct was developed in 1989 & 1990, but this is actual physical development. Yet another required technology has been completed; can we go now?

slide 3 wrote:

Decreasing G-load (~10X Reduction)
Traditional Venus Entry:
· 200-300g's
· ~4000 W/cm^2
ADEPT Venus Entry
· 20-30 g's
· ~300 W/cm^2

slide 3 wrote:

Flight-system mass reduced by 25%

And another slide presentation, this one dated August 7.
Viability of 3 D Woven Carbon Cloth and Advanced Carbon-Carbon Ribs for Adaptive Deployable Entry Placement Technology (ADEPT) for Future NASA Missions
It includes the same slide, this time slide 5. But has more detail on thermal testing, and a reference for use on Mars. Slide 3 says "ADEPT can be scaled to deliver 40 MT payloads to the surface of Mars". Their timeline is slide 4; it concerns me:

- Human/Heavy Mass Mars Mission and Design Studies FY'11
- Full Scale Demonstrator FY'2014-16
- Ballistic Robotic Venus Lander (~2017)
- Ballistic Robotic Mars (1-5mT)(~2026)
- Lifting Concept Flight Demos (> FY'2026)
- Human Mars (~2035)

The good side is they're actually working on it. But from initial design study to first human flight, 24 years? How many administrations is that? How many chances for Congress to cancel it? From JFK's speech at Rice University to Apollo 11 was 8 years. Yes, JFK actually ordered NASA to start work a year before his speech. He ensured NASA could do it before making his big speech; smart man. That means 9 years from order to landing. Not 24 years for just one technology. Or I could be more harsh, include the time from Mars Direct. From 1990 to 2035 is 45 years. But to be fair, they didn't start real work until 2011. I suppose actually starting work is progress, but 24 years? Yes, I'm discounting some steps. Design study to Venus lander is only 6 years. But from that to Human Mars is another 18?

GW Johnson · 2014-12-27 10:31:06

To answer Bob Clark's question: yes, Nextel 312 was the woven fabric in my oddball insulation. The matrix was a pipe insulation potting compound called Cotronics 360(? not sure of number) M (M for "moldable" as I recall). The surface sealant was a cement called 901. Cotronics Corp is still around, and those products are still available today, if you ask specifically for them.

It looks like the AFRSI quilt is another ceramic-ceramic composite in a way. The "trick" to achieve low density is high porosity, but you need something else to provide some structural strength. It is the low-density batting sandwiched between the layers of 440 that gets you the low thermal conductivity. The fabric and the wire net plus foil supply the strength for it not to get torn away in the wind.

I've been looking at a reprise of my old low-density insulation scheme for another ramjet combustor that has to be reusable. This time around, I'm looking at yttria-stabilized fibrous zirconia products from Zircar. Haven't worked out the processing yet, but there's both furnace board and fabrics good to operating temperature limits of 4000 F.

These kinds of materials could serve as refractory heat shields on blunt objects, given some kind of redundant retention. The zirconia offers a lot higher temperature capability, which means with a black surface higher ballistic coefficients typical of capsules from LEO could be handled. The downside with zirconia is that the densities are higher than what I achieved with the alumino-silicates long ago: near to water.

I'm not sure, but I believe the low density that I got with the potting compound was an artifact of my processing and part shape. The steam wormholing its way out during cure left behind a myriad tiny paths. That probably won't happen with the zirconia materials.

The two applications are actually quite different. The combustor liner has zero re-radiation cooling, a hefty convective heating, and operates at very low conductivity to maintain the steep temperature gradient at very modest conduction heat throughflow, for a modest backside cooling requirement. The heat shield application is extraordinary convective heating, balanced by re-radiative cooling, with conduction heat throughflow essentially zeroed, for no backside cooling requirement at all.

The "ultra-high-temperature-ceramic" nose tip and leading edge materials work in a completely different way. They are very high density high conductivity materials of high strength and thermal shock resistance in a crystalline structure. Stagnation heating is conducted elsewhere through the material, producing a very large backside cooling requirement.

GW

Impaler · 2014-12-27 20:15:14

Ok, direct polar orbit entry for the whole mission stack at the moon, can we get some numbers on how much Delta-V that will cost and if Orion Service Module is capable of it. We would probably need to go for one of the high inclination frozen orbits too but I belive the highest inclination ones are around ~70 which should be adequate to do any desired polar landing.

RobertDyck · 2014-12-28 03:32:20

Here is the trajectory for Lunar Prospector in 1998. (Has it really been that long?)

Note it entered polar orbit directly, not any sort of plane change. That means entering Lunar polar orbit requires exactly the same delta-V as entering equatorial orbit. A human mission would leave Orion parked in the "12 Hour Initial Orbit". Lunar Prospector had to drop to a low mapping orbit, but Orion would stay in the higher orbit.

Here is a comparison of Apollo 11 vs proposed Constellation trajectories. Click image for high resolution.

Differences: Constellation launches Altair LM on Ares V, and Orion on Ares I. Dock before departure. One proposal for SLS was to use SLS block 2 instead of Ares V, and SLS block 1 instead of Ares I. That would allow Constellation to the Moon. But one of the two launch pads has been leased to SpaceX for Falcon Heavy, so how would they conduct two launches close enough to do that? The other change was to use the Altair LM for the Lunar Orbit Insertion burn. Apollo used the Service Module for that. But notice this chart is otherwise the same.

Only one more difference: "Apollo 11 Equatorial landing" vs "Constellation polar outpost landing". This is because delta-v is exactly the same for polar vs equatorial orbit insertion and departure. Launch from Lunar surface can use rotation of the Moon to assist with ascent, but the Moon rotates so slowly that this isn't much anyway.

This answers another question I had. Why is Altair so heavy? One reason is fuel of the descent stage is used for LOI burn. I suppose that's because they thought Ares V with SRBs was unsafe, while Ares I was safe. Now we have SLS, intended for human space flight. So don't be afraid of putting humans on SLS block 2. That allows a larger service module.

Apollo CSM was rated for spacecraft delta-v 2,800 m/s. Orion with ATV Service Module is rated for spacecraft delta-v 1,340 m/s. We need a better SM.

I don't have delta-v figures for Orion with "Orion 606 SM". I can quote mass figures.
ATV-based SM dry mass 4,430 kg, propellant 7,907 kg
Orion 606 SM dry mass 3,700 kg, propellant 8,300 kg
So the Orion 606 Service Module would be both lighter, and more propellant. That provides more delta-v. Not sure how much, but doesn't appear to be enough for both LOI and TEI burns. At least not with LM attached.

Note: Propellant on Apollo Service Module was UDMH & N2O4. Propellant for ATV-based SM and Orion 606 SM is MMH & N2O4. Orion was originally proposed to use liquid methane and LOX, but they switched to established technology. Main engines for Orion 606 SM were to be upgraded version of the Shuttle OME. Thrust of OME = 6000 pounds (2,721 kgf), proposed Orion 606 SM engine thrust 7500 pounds (3,402 kgf).
OME Isp = 316 seconds
Russian RD-160 (uses liquid methane & LOX) thrust 2,000 kgf, Isp = 381 seconds
Orion Propulsion Inc. Methane Engine, developed from 2001, tested 2005, thrust = 100 lbf (45.36 kgf), Isp not listed
Orion Propulsion Inc. was purchased by Dynetics in 2009.
May 19, 2014 Dynetics announced partnership with Aerojet/Rocketdyne.
methane engine is currently listed as completed, but the web page has nothing more than thrust.

Last edited by RobertDyck (2014-12-28 06:08:28)

SpaceNut · 2014-12-28 11:07:53

Thanks for the run down but while we are well on the way for the SLS and even Block 2 with the Orion capsule which you noted still needs a better SM; we still have very little work done with regards to Altair or any other version of lander for the moon which would include a reusable ETLV 2 new lunar lander. Most of the work seems to have stopped on the lunar truck, the inflateable lander and more which are needs for anything other than a sortie mission.

SpaceNut · 2014-12-28 15:07:07

The Government Accountability Office (GAO) is indicating that there are still [url=http://www.spacedaily.com/reports/Challenges_for_Orion_and_SLS_999.html]Challenges for Orion and SLS
[/url]

GAO Director of Acquisition and Sourcing Management Christina Chaplain testified on the progress of the Space Launch System (SLS) and Orion crew vehicle on Dec. 10 at the Space Subcommittee of the House Science, Space and Technology Committee which held a hearing on the progress of the nation's next generation deep space exploration vehicle and heavy lift rocket that is being developed for deep space human missions that will take astronauts to the Moon and Mars.

Astrowatch.net: What are the main challenges that NASA's human space exploration programs have to face now?
Christina Chaplain: We reported on a funding gap for SLS but by delaying the launch date to November 2018 (and if the budget passes as-is), NASA will have some relief from that gap. The delay will also alleviate schedule pressure for the core stage. So right now, the challenges for SLS are the types of inherent engineering and technical challenges that face any program of its nature. Orion is facing a funding risk as well and it is facing challenges with the heat shield and parachute system. But perhaps more challenging is the fact that there is considerable development left for human support systems for Orion. We have not reviewed the ground system in detail. In the long run, funding and affordability issues are likely to challenge all the programs.
Astrowatch.net: What is the most probable date of the first SLS launch?
Chaplain: NASA has a 70 percent confidence level in the November 2018 date but Orion has not set its official date yet and even November 2018 may be a challenge.
Astrowatch.net: William Gerstenmaier, NASA's Associate Administrator for Human Exploration and Operations said at the hearing, that even more funding won't keep the 2017 date for SLS, is it true? What are your thoughts on this?
Chaplain: I would agree. At this point the schedule for the core stage is very compressed. There is very little room to address problems to meet the December 2017 date. Much of the work is serial so money can't really speed things up too much at this point.
Astrowatch.net: What technical or other problems need to be resolved to allow Orion's first manned mission launch?
Chaplain: I mentioned the Orion challenges earlier. One key thing is that the European Space Agency will be a partner in developing human support systems. It may be challenged to deliver on time.
Astrowatch.net: GAO urges NASA to specify its plans beyond the Exploration Mission-2 (EM-2) in 2021, what has your Office already learned from the agency about this stage of human space exploration?
Chaplain: At this point there are no official missions beyond EM-2 and there are no cost estimates or even ranges depending on possible missions. The uncertainty puts NASA at risk of making decisions today that might not make sense tomorrow. While there seems to be consensus that Mars is the long-term destination, decisions on how and when to get there can impact decisions today as well as costs in the long term.
Astrowatch.net: Do you think the costs of SLS, Orion and Ground Systems Development and Operations (GSDO) Program will increase?
Chaplain: If NASA sets Orion cost estimates through EM-2 at 70 percent, then there will be a reasonable chance that costs will not increase significantly before the second flight. But complex programs like this tend to encounter problems that they did not anticipate and are expensive to resolve. Establishing cost estimates at 70 percent confidence is a good practice and should help prevent such growth.
Astrowatch.net: Is the $18 billion for NASA in 'CRomnibus' spending bill enough to push the program forward and avoid other delays?
Chaplain: The Omnibus should help keep the programs on track at least for the next year.

Impaler · 2014-12-29 02:56:25

Rob: I'm still skeptical that initial polar insertion to the moon is exactly the same Delta-V as equatorial though your making a good case that it is not significantly more, the fact that it has been done before is very good, do we have the numbers on the trajectory of Lunar prospector, what was it's time of flight? Can the trajectory simply be reused?

The under-powered Orion service module (effectively underpowered when pushing such a fat capsule) seems to be emerging as a big barrier. The key change of having the lander perform LOI rather then the service module seems to have been a very significant and in my opinion bad move. This means the Delta-V budget for Orion and it's service module alone is likely to be completely used up for the Earth return burn, so the vehicle can't even do an Apollo 8 mission equivalent of orbiting and returning from the moon. I think this may be why 'distant retrograde' orbits are mentioned as destinations for the ARM mission asteroid, cause if it was closer to the moon Orion wouldn't reach it with enough propellent to return to Earth. When they made Orion SM this size it made the failure to develop the lander even more egregious because that lander was in fact a stage for getting humans back to Earth.

Worse it may rule out a reusable lander because even the Orion flying alone (and wasting half the SLS capacity) and making lunar orbital rendezvous with a fully fueled and disposable lander would then be stranded around the moon. Only some kind of long-lifespan cryo-stage which dose LOI (and then ideally fuels a reusable lander) would be needed and that's yet another asset which dose not exist yet.

Maybe the ESA service module can be stretched to give the necessary Delta-V, but that would certainly rule out any kind of Lander riding along with it on SLS. If that were done you would see Orion launch on another rocket like Falcon Heavy (which would likely be cheaper then the original Ares I roman-candle death-trap anyways) and then you have the whole SLS for the lander.

RGClark · 2014-12-30 05:29:06

RobertDyck wrote:

New document for ADEPT. This is a slide presentation in .PDF format. File with NASA's technical report server on November 17 of this year.
Deployment Testing of the ADEPT Ground Test Article
Caution: they optimized it for Venus. I'm saying this technology could be used for Earth or Mars, but may have to be modified. After all, this looks exactly like the deployable heat shield for the habitat in Mars Direct. Mars Direct was developed in 1989 & 1990, but this is actual physical development. Yet another required technology has been completed; can we go now?
...

The 25% reduction in "flight mass" might be for Mars landing requirements. I don't think it's relating to Earth landing since an Apollo ablative heat shield was only 15% additional mass so if you reduced the "flight mass" by 25% you would be reducing the actual capsule mass.

This is similar to the "parashield" concept investigated by Prof. David Akin at the Univ. of Maryland's aerospace dept.:

Phoenix: A Low-Cost Commercial Approach to the Crew Exploration Vehicle.

Abstract: Since the announcement of President Bush’s Vision for Space Exploration (VSE) in early 2004, the architecture of Project Constellation has been selected. The system will be centered around the Crew Exploration Vehicle (CEV), which has been dubbed by NASA administrator Michael Griffin as “Apollo on steroids”. The CEV is to be launched on a new launch vehicle, derived from existing shuttle technology. The development of this new
spacecraft and launch vehicle is a very costly proposition. An alternate approach is proposed in this study. The Phoenix is a smaller spacecraft designed specifically to be launched on the Falcon 5 vehicle under development by SpaceX. Because the SpaceX vehicle will cost only a fraction of today’s launch costs, the Phoenix is estimated to cost less than half of the price of the CEV. This reusable three person capsule utilizes an innovative re-entry concept, which allows for a cylindrical spacecraft with greater interior volume. This extremely cost-effective spacecraft is an attractive option for fulfilling VSE requirements.

http://nia-cms.nianet.org/RASCAL/2010-W … SC-AL-2006.

From the report:
Screen+Shot+2013-04-25+at+1.03.42+PM.png

From the specifications listed it looks like the parashield would weigh about 12% of the dry capsule mass, i.e., without propellant.

Bob Clark

SpaceNut · 2014-12-31 23:10:30

At times I think Nasa is read our posts...

Several recent articles in space publications have called into question NASA's ability to assemble a reasonable space exploration program plan.

Revisiting a 2009 Space Exploration Architecture Proposal

In summary, NASA is apparently building the Space Launch System (SLS) which supports an exploration architecture that is quite similar to that used for the Apollo Lunar Program, i.e., launch the entire mission equipment and crew on one launch vehicle for travel beyond low earth orbit.

We should not just repeat the past of Apollo but build to make the next steps possible....

Use the ISS as an assembly, integration and test facility for a modular reusable solar system human exploration transfer stage (HETS). In the early years, a rather simple HETS can be used for flybys of near-Earth points of interest such as asteroids and the moon.
In the later years, the HETS might be evolved into a Mars vehicle. Upon return to Earth, the HETS could rendezvous with a space tug at high altitude, enabling the tug to tow the HETS back to the ISS.

Some would say ditch the ISS as well and give all the contracts to Space x and others that would build cheaper launching vehicles and hope that they can lift the mass that we want to do orbital assembly...

RobertDyck · 2015-01-01 08:25:20

I have advocated using ISS as a construction shack for orbital assembly. And for long duration testing of life support. The lift support system on ISS is almost ready for Mars, but not quite. And I want to see the life support system tested in space for the full duration of a Mars mission, before sending humans to Mars. That's best done on ISS. So no, I would not agree with ditching ISS.

But yes, NASA does read this forum. I have spoken with people at NASA who are concerned about what is said here. And NASA appreciates our efforts to lobby congress.

RobertDyck · 2015-01-01 08:47:08

From the article that SpaceNut linked...

Let's use the Space Shuttle to ferry astronauts and valuable cargo to the ISS and use expendable launch vehicles to transport consumables and low-value cargo to low orbit.

Um, what? That was the plan, but the Shuttle has been decommissioned. The article is dated December 30, 2014, so just Tuesday. Why does he mention Shuttle?

GW Johnson · 2015-01-01 13:37:26

The article was re-edited by "staff writers" from a 5-year old document, just to have something to publish. Those "staff writers" actually know very little about their subject, or they would have realized there is no shuttle anymore.

Credentials and experience actually do matter, after all. Out on the internet in general, I find that 99+% of what is posted is really ignorant BS. I'm happy to say that percentage is far lower on the forums.

GW

SpaceNut · 2015-01-01 13:43:25

I think is was a metaphoric slip meaning that the space shuttle's successor in the SLS was the new ferry to the iss....but ya does lead one in the wrong way....
Nice to hear that those in Nasa are reading the forum, that does explain alot of what we posted in the past with regards to the path taken via Nasa on the SLS.

RGClark · 2015-01-03 11:25:04

The Early Lunar Access plan would follow this model in not needing a hugely expensive super heavy lift launcher, but use smaller launchers:

Lunar Base Studies in the 1990s
1993: Early Lunar Access (ELA)
by Marcus Lindroos
http://www.nss.org/settlement/moon/ELA.html

As this stems from the early 90's it uses discontinued launchers such as the shuttle and Titan IV. But currently existing launchers such as the Delta IV Heavy, the Proton, Ariane 5 would work, as well as the upcoming Falcon Heavy. These would not need to be man-rated. Any manned launcher available could be used to send the astronauts to the ISS. And they would board the lunar elements in space which were sent up to rendezvous and be combined at the ISS.

This would cost a fraction of the Apollo or Constellation plans since all the launchers and the Earth departure stage are already existing, or will be fully paid for by the developer such as the Falcon Heavy, and the only extra element that would need to be developed is the lunar lander stage.
And according to Dave Masten a reusable lunar lander could be developed for a few 10's of millions of dollars, not the $10 billion estimated for the Altair:

The future of NASA’s commercial partnerships
by Jeff Foust
Monday, May 19, 2014
http://www.thespacereview.com/article/2515/1

Bob Clark

Last edited by RGClark (2015-01-04 07:09:28)

RobertDyck · 2015-01-03 13:31:35

To save fuel, the LEV makes a direct landing rather than enter an intermediate lunar parking orbit as Apollo did. The vehicle retains sufficient propellant to perform a later ascent burn to return the crew to Earth.
...
The crew capsule would be derived from the Apollo Command Module that last flew in 1975. It retains the external size and shape of the original Apollo CM design to take advantage of the existing aero- and thermodynamic databases developed during that program. The interior has however been scaled down since it only supports a crew of two instead of three, and the capsule is lighter since its design is based on modern materials, lightweight electronics and construction methods.

This assumes the capsule is landed on the Moon. That means landing the capsule hull capable of Earth atmospheric entry, heat shield, parachutes, and floatation air bags. That sounds heavy. Apollo capsule (command module) without the vague upgrades massed 5,560 kg. This study counts on reducing dry weight to 3,688 kg. Parking CSM in lunar orbit means you don't have to land/lift all that. The service module used for TEI also does not have to land/lift from lunar surface. But then you have two crew vehicles (capsule and lunar module), as well as propellant for LOI. Apollo was originally designed to land directly on the Moon, which is why the service module engine was so large. But they found it was way too large for Saturn V. The first design had a single service module for landing and launch, but later changed to two stage: descent and ascent stages. That was still too heavy. Splitting to into lunar module and "mother ship" reduced total launch mass. They kept the previously designed capsule, and shrunk the service module, and added a new lunar module. The resulting stack was lower total mass than the lift capacity of Saturn V. They were pleasantly surprised to find they had extra lift mass available.

So I'm sceptical about returning to a design that lands the capsule. They use LH2/LOX instead of Aerozine 50 / N2O4. Apollo LM engines had Isp 290 s. Assuming LH2 doesn't boil off, the Shuttle's engines had Isp 453 s in vacuum, and dedicated engines designed for vacuum only increased that. RS-44 engines mentioned have Isp 481 s. That helps, but the reason Apollo used storable propellants was to avoid boil-off. Both the Apollo CM and SM used UDMH/N2O4, with Isp 314 s. Still, would a single stage design work?

Orion capsule is 8,913 kg, so it won't work. Dragon has a dry mass of 4,200 kg, and that's the cargo resupply version so doesn't include life support. And Congress would want to include Orion in any Moon plan, to justify what they spent. I don't have mass figures for the CST-100 capsule, the only thing I found said the total mass of capsule + service module + propellant is 10 tonnes.

Note: Orion capsule + ATV-based SM + propellant, not including launch abort system or fairing, is 21.25 metric tonnes. CST-100 is 10 tonnes. Dragon is 8 tonnes. And Dragon has side mounted launch escape rockets, and CST-100 is designed to use its service module for launch escape, so their escape systems are included.

Let's see. Mass numbers in Table 1 from the article doesn't add up. If you add up the components for LEV, it adds up to 20,040 kg, but the table lists total mass of 20,140 kg. And total for the Capsule is 3,688 kg, LEV is 20140 kg, and payload adapter is 6,000 kg, but total Shuttle payload is 25,723 kg. I don't even know how they calculated that total Shuttle payload. But let's take the 20,040 kg figure for the LEV, and add the total capsule weight, and assume Isp for RS-44 engines = 481 s. A delta-V calculator says this gives 5634.09 m/s. Tried to find delta-V to directly land on the Moon; Google returned a result from another forum posted by RGClark. He linked a Wikipedia article here. That has a table, LEO-Ken to Moon is 5.93 km/s. "LEO-Ken" means Low Earth Orbit with an inclination equal to the Kennedy Space Center. But the article talks about an upgraded Centaur G stage for TLI. Ok. The Wikipedia table lists return from the Moon to LEO with aerobraking as 2.74 km/s. From lunar surface to Earth-Moon Lagrangian point 5 is 2.58 km/s, which is the gravitational mid-point, so presumably direct entry would involve falling to Earth. Hmm. These numbers aren't clear.

Last edited by RobertDyck (2015-01-04 15:36:26)

SpaceNut · 2015-01-03 19:53:28

Thanks for the links and ya I have had trouble with the Nasa numbers on other items as well....

Its interesting that it would seem we need more work on Nasa's plans to go back to the moon when we already had a working model to follow....

GW Johnson · 2015-01-04 11:02:04

I agree, landing the big, heavy Earth reentry capsule on the moon as part of a lander design is a stupid idea, based on what we learned from Apollo. That land-the-whole-ship idea was where NASA was with the Apollo design in 1963. It took two Saturn-5 launches per moon mission, plus cryogenic refueling from one 3rd stage to the other in Earth orbit, to make such a cluster landing possible back then.

The breakthrough idea came from outside NASA, and was resisted very strongly due to "not invented here" attitudes already in place at NASA that early. That idea was "lunar orbit rendezvous", which translates to "take only what you need to the surface, leave the rest in orbit about the moon". That got them down to one Saturn-5 per mission. By 1965 they had broken down and accepted lunar orbit rendezvous as the only way forward.

Strangely enough, using that plus LEO rendezvous and docking assembly could have gotten a moon mission down to 3 or 4 Saturn-1 launches, but they didn't need to do that , they had the Saturn-5 available.

BTW, the Saturns were not originally designed to be moon rockets. Von Braun was working for the Army before there was a NASA. Saturn-1 was a very large ICBM, and Saturn-5 started as concept for a "troopship rocket" to put 100+ men on the ground in Russia from the US in an ICBM's flight time. Von Braun knew these rocket could be, or could lead to, space launch rockets, just like a decade earlier with the V-2 for the Wehrmacht. Building ballistic missiles is how he got to play with space travel concepts, such as the "Nova" rocket we never built.

GW

SpaceNut · 2015-01-04 18:16:23

So the not invented here by Nasa needs to go and we need to run new numbers to help Nasa fix its Orion and lunar lander for a "lunar orbit rendezvous", if we want to go at all... Or we need to give Nasa a new 2 rocket approach that creates a base on the moon in one shot.

RobertDyck · 2015-01-04 21:29:57

I did a bit more reading. "Aerozine 50" used by the Apollo LM was a 50:50 mixture of UDMH with hydrazine. UDMH is hydrazine with 2 methyl groups added for stability. Modern MMH is hydrazine with just one methyl group. Raw hydrazine is hypergolic, which means it'll explode if you hit the tank hard enough. UDMH and MMH are stabilized, they don't do that. Furthermore, hydrazine freezes at +2°C; while UDMH and MMH remain liquid much colder. It turns out the 50:50 mixture of hydrazine with UDMH provides stability and lower freezing temperature of UDMH. Although Aerozine 50 has lower Isp (290s instead of 314s), it's more dense. That allows tanks for the lunar lander to be smaller. Looks like the 1960 engineers did total system optimization: Aerozine requires more propellant mass, but lower tank mass. Total of tank + propellant is what you have to land and lift. Good for them!

Today the same principle applies, but we would use more modern and higher performing propellants. Don't use liquid hydrogen because low density requires ridiculously large tanks. And LH2 boils off too damn fast. Instead use liquid methane. According to "The Case for Mars", LCH4/LOX has Isp = 380s. But in the 1990s, the Russians actually built several engines to do that. RD-160 has Isp = 381s. And two American contractors were hired to developed LCH4/LOX engines for Orion. They both only completed RCS thrusters, but both websites claim they could easily scale up for larger engines. Ok, do so. LCH4/LOX engines for main engines and RCS thrusters of the service module for Orion. And use the same fuel mixture for the lunar module. Could you use the same engines?

Apollo LM was two stage: descent and ascent. One reason was Isp = 290s. But today with Isp = 380s, can we build a single stage lunar module? That would allow reuse. Design the Orion SM to be large enough for LOI burn with the LM attached, and enough propellant for TEI burn. If you left the LM in lunar orbit, a second mission could carry a cargo module instead. The cargo module would carry propellant to refuel the LM, plus additional equipment such as lunar rovers, surface science equipment, construction equipment, etc. Transfer that to the LM for landing. That means the LM would be designed to land with substantially more cargo mass than the first mission, provided propellant tanks were full. All that equipment would be unloaded before ascent. Each mission launched by a single SLS block 2. Does that mean, for the first mission, LM propellant tanks would not be completely full? And of course, the LM would carry all 4 astronauts.

On orbit transfer of LCH4 and LOX would avoid toxic residue issues. Even if traces are tracked by spacesuits into the capsule, neither of these are toxic. Oxygen is obviously good for the spacecraft. And methane in trace quantities are produced by the human body; they're called farts. The activated charcoal filter in the spacecraft is already designed to deal with that. Methane is the primary gas in natural gas, my house is heated by natural gas, so that's obviously safe for household use. That isn't liquefied, but this looks like a safe propellant. Liquefied Natural Gas (LNG) is transported by train or ship all the time, so this technology is established. What issues would we have keeping it liquid for a lunar mission? Or in space propellant transfer?

Last edited by RobertDyck (2015-08-05 23:33:25)

Impaler · 2015-01-04 22:21:29

I tend to agree that for me moon is "permanent research base or don't bother", that base dose not need to be OCCUPIED permanently, intermittent occupation is acceptable and may be necessary, but it would ideally be seeing occupancy at least 25% of the time, say a 3 month stay once per year would be reasonable and yet still compelling. Seasonal availability of solar power may dictate a years 'moon season', the Russians can be convinced to come along by calling it a 'lunar-dacha'. Gradually longer mission stays can evolve as confidence grows. Crew of 4 seems reasonable and in line with potential lander capabilities.

To do this we would really need to move two an architecture in which a reusable lander AND the habitat segments are put on a heavy lift vehicle, and crew and supplies are sent on something much smaller, basically similar to ISS in which different 'build' and 'supply' vessels are used. Orion with just it's Service Module (maybe with slightly more propellent) is a fine lunar-taxi craft and should not be put onto SLS as the remaining mass budget can't be used effectively, get a smaller cheaper launcher under it by opening it up for a competitive bid like we did ISS. Then use that same launcher to send propellent and cargo (in a minimalistic Cygnus like tincan that can be switched-out with the habitat module on the lander) to the reusable lander in lunar orbit, that will be the whole 'not mixing crew and cargo' Augustine principle that SLS is ignoring. The established base now needs a much more realistic 4 launches per year of medium class (2 Propellent, 1 Orion, 1 Cargo), which will be hugely safer and more reliable then 1 Super-heavy launch per year.

The bigger launch for the lander and habitats is still not going to need to be the SLS if it is done intelligently, both the lander AND habitats have PLENTY of LEO loiter time available to them, they are going to live in space for decades. No reason to not do 2 launches and LEO rendezvous with a push stage that will send them to the moon and for the habitats do some of the de-orbit burn and topping off of the engines that will do the touch down, lander brakes itself into lunar orbit and needs first refueling to make landing. Falcon Heavy is the obvious launcher to do both missions, but again if you release this as a competitive bid it's open to everyone. At 4 launches a year of this Heavy class you put 2 element in place per year, total elements would be in the range of 6-8, possibly 2 landers being used for redundancy (and allowing them to be rotated out as they reach end of life after ~10 years/landings). Again the sweet spot of 4 launches a year is aimed for and the duration of just 3-4 years gets the base up and operating in a more reasonable time-frame the ISS took to do.

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Re: SLS Rocket + Orion + Apollo LEM

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Re: SLS Rocket + Orion + Apollo LEM

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Re: SLS Rocket + Orion + Apollo LEM

#40 2015-01-01 08:25:20

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Re: SLS Rocket + Orion + Apollo LEM

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Re: SLS Rocket + Orion + Apollo LEM

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Re: SLS Rocket + Orion + Apollo LEM

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Re: SLS Rocket + Orion + Apollo LEM

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