Could Blue Origin offer its own rocket to the Moon, Page 2: low cost crewed lunar landers.

 Copyright 2025 Robert Clark

 In the last blog post, "Could Blue Origin offer it’s own rocket to the Moon?", I suggested that with technically feasible upgrades of the New Glenn booster engine, New Glenn might be transformed into a Saturn V-class, 100 tons to LEO, Moon rocket.

 An objection raised to the calculations I presented there was that the maximum New Glenn first stage tank size I was using did not include ullage space, i.e., the space left unfilled or filled with gas to account for boiloff. Three possible solutions: first, even with the commonly used estimate of ca. 1,150 tons propellant load it would require just a ca. 10% increase in tank size to get the prop. load in the 1,300 ton range. SpaceX has shown that additional tank rings have been swapped in and out of the Starship to get an additional propellant load increase of this size or more. 

 Second, an announcement from the Texas State Senate has indicated Blue Origin has been assigned a grant to increase the New Glenn prop.load by subcooling, i.e., densifying the propellant. Propellant subcooling typically results in an approx. 10% propellant load increase. 

 Third, New Glenn's, Moon lander uses hydrolox so it must make use of some zero-boil-off tech to not lose too much hydrogen over a mission lasting several days. This same tech might be able to be used on the New Glenn first stage to minimize the need for ullage.

 Therefore we'll work on the basis the New Glenn can be upgraded to get ca. 100 tons to LEO as expendable.

Getting a crewed lander.

 The space industry was pleasantly surprised by Blue Origin's New Glenn being able to reach orbit on its first launch. They were even more surprised by the announcement the next mission planned will take a cargo lander to the Moon as early as March, though more recently they've only said sometime in late Spring.

 The success of Blue Origin reaching orbit on the first launch with New Glenn and the rapidity at which they wish to progress to launching a lunar lander on the Moon shows the importance in having a top notch Chief Engineer such as David Limp making the technical decisions. If SpaceX had taken the route of hiring a true Chief Engineer, they would already be flying the Starship with paying customers at least in expendable mode. Moreover, they would recognize having a launcher as expendable with 250 ton capacity means they could do single launch missions to the Moon or Mars, no SLS, no multiple refueling flights required.

 As it is, SpaceX is in real danger of being lapped by Blue Origin in having a manned Moon rocket or even a Mars rocket.

  Blue Origin has stated their Blue Moon Mk1 cargo lander will have a 21,350 kg fueled mass, and payload of 3,000 kg payload to the Moon one-way.

Blue Moon Mk1 cargo lunar lander.

 Given the delta-v requirements for getting to the Moon we can make estimates of its propellant and dry mass values:

Delta-V budget.
Earth–Moon space.

https://en.wikipedia.org/wiki/Delta-v_budget#Earth%E2%80%93Moon_space%E2%80%94high_thrust

 Reports are the current version of the New Glenn has a payload to LEO of 25 tons. A 21,350 kg fueled mass of the Blue Moon Mk1 lander plus 3 tons cargo would be 24,350 kg, just under the payload capacity of the current New Glenn.

This though means Blue Moon has to provide the delta-v for trans-lunar injection(TLI) and insertion into lunar orbit as well as lunar landing. From the table the total of TLI and insertion into low lunar orbit and landing is 5.93 km/s, 5930 m/s.

 The engine on the lander is supposed to be the BE-7 hydrolox engine upgraded from the BE-3 used on the New Glenn's upper stage. We'll assume the BE-7 has about the same vacuum Isp of the BE-3, of 445 s. Then taking the propellant load of the Blue Moon as 18.35 tons and dry mass as 3 tons allows it to get 3 tons in cargo to the 5,930 m/s delta-v needed to go from LEO to the lunar surface, plus some margin:

445*9.81Ln(1 + 18.35/(3 +3)) = 6,110 m/s.

 The Blue Moon Mk1 is also already developed and paid for by Blue Origin on its own dime. And it is established fact at this point that spaceflight components, rockets or spacecraft, as developed by commercial space, and privately funded saves 90% off the previous governmentally financed approach that is paid for by governmental space agencies such as NASA. 

 A key fact not yet generally recognized is that we are already at the long desired point of having spaceflight being sufficiently low cost that it can be fully financed by commercial space and private funding only, no governmental financing required at all. BUT such low costs hold true only if it is privately funded.

 A majorly important example is the Mars Sample Return mission. There is much hand-wringing at NASA and among space science advocates about the $10 billion price tag estimated by NASA for MSL. But in point of fact this mission and all space science missions going forward can be paid for at 1/100th the costs estimated by NASA by following the commercial space approach. And in fact the costs as privately funded would be so low, such missions could even be mounted as privately financed at a profit. See discussion here:

Low Cost Commercial Mars Sample Return.
https://exoscientist.blogspot.com/2023/07/low-cost-commercial-mars-sample-return.html

 The argument for this is quite simple. SpaceX and now multiple other space startups have confirmed that development costs as privately funded are 1/10th the costs of governmental funded development costs. But then production costs of individual space components rockets or spacecraft are commonly 1/10th or less than their development costs. As a space company paying for a space project on your own dime, rather than paying the large development costs of a new component you would just naturally use ones that already exist, resulting in far smaller outlay on your end. Then taking into account 1/10th cheaper development cost overall as privately financed and 1/10th or lower cost using already existing components, rather than developing them from scratch, the result is 1/100th or less cost than the usual development costs estimated by NASA following the government financed approach.

 So we already have a lander in the Blue Moon Mk1. But could this serve as a crewed lander? Yes, it can because of a key fact being overlooked by NASA: Artemis is not Constellation's Apollo on steroids, It is in fact Apollo 2.0.

 Perhaps NASA didn't want to acknowledge this so that it would continue to get funding. Just saying Artemis is Apollo redone would not sound nearly as impressive or necessary. But it is important to understand this point. 

 The argument for this conclusion is quite elementary. The primary launcher of Constellation was the Ares V. It was intended to have a startling 188 tons to LEO payload capacity. But there was more to Constellation than that still. The crew were intended to be launched separately to LEO by the Ares I. This had the payload capacity to LEO of 25 tons. Then the Constellation plan with its two launchers could get ca. 210 tons to LEO. This is about twice that of Apollo, but more importantly its about twice as much as Artemis. So in point of fact in the key measure of payload mass to orbit Artemis is Apollo. It is far from Constellation was capable of.

 Once, this is understood then it is understood Artemis should not try to get a lander the size of the Altair lander of Constellation at 45 tons. It should try to get one comparable in size to Apollo. 

 Instead, NASA is seeking that Altair sized lander such as the crewed version of the Blue Origin lander, the Blue Moon Mk2 also at 45 tons, 

Blue Moon Mk2 crewed lunar lander.

or, worse seeking to get the 1,200 ton Starship HLS with multiple refuelings to fit in the Artemis architecture.

 Instead we'll show the Mk1 cargo lander can form the lunar lander for single launch crewed lunar mission format based on the New Glenn as launcher. 

Architecture 1: this will be analogous to the Early Lunar Access proposal of NASA, a proposed follow-on to Apollo.

https://web.archive.org/web/20081106190735/https://nss.org/settlement/moon/ELA.html

 The salient feature of this proposal is it used a single crew capsule for the full round trip from Earth orbit, all the way to the lunar surface, and back to Earth, thus no separate lunar module, i.e., no lunar orbit rendezvous(LOR).

 You see from the table of delta-v's the delta-v needed from the lunar surface back to Earth is 2.74 km/s, 2,740 m/s. This would not put you in Earth orbit though but on a ballistic return trajectory to reenter Earth's atmosphere, a la the Apollo command module. 

 The total round-trip delta-v would be 2.74 km/s + 5.93 km/s = 8.67 km/s, 8,670 m/s.

 The extra delta-v could be provided by the Delta IV Heavy's upper stage, now being used for the interim upper stage of the SLS. This stage would be put atop the New Glenn as a 3rd stage performing the role of a "Earth Departure Stage" for the push to translunar injection. Carrying the Mk1 with a 3 ton crew module it could get:

465*9.81Ln(1 + 27.2/(3.5 + 24.35)) = 3,110 km/s, sufficient for translunar injection(TLI) of the 24.35 ton total mass of the Mk1 lander and crew module.

 This 3rd stage plus the Mk1 and crew module would have a total mass of 30.7 + 24.35 = 55.05 tons. The cited 45 ton payload capacity of the New Glenn to LEO was a for a partially reusable version, with the booster landing downrange. Then for an expendable use it should get ca. 60 tons to LEO, sufficient for the purpose. 

 However, the key question is of a crew capsule that would be analogous to the Apollo Command capsule or the Orion capsule or the Dragon capsule but only at ca. 3 tons dry mass. This is only half the dry mass of the Apollo Command capsule but required to play a similar role.

 A research report of Prof. David Akin of the University of Maryland aerospace department suggests this is indeed possible:

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.

 Below is page 3 from this report:

 Since the Cygnus cargo capsule of Orbital Sciences, now a division of Northrop Grumman, of comparable size to the Phoenix proposal, already exits I suggest basing it on the Cygnus just given life support and heat shield. Remember our dictum is, "Use existing resources to save on costs if available."

 The proposed heat shield for the Phoenix was a "parashield", a combined parachute and heat shield:

 And a proposed heat shield of the Cygnus to make it reusable was an inflatable:

  These may indeed work. But to get to an operational system minimizing development work and cost I advise simply making the Cygnus tapered like most manned capsules and using a traditional heat shield beneath it:

 For both the Soyuz and Dragon, they have relatively small taper angle so you would lose a relatively small size in capsule interior volume by giving the Cygnus a similar side taper.

 Quite notable is with this option you can get a crewed Moon mission with only a single launch of a 60 ton to LEO launcher. Then both the New Glenn as expendable or the Falcon Heavy as expendable could do it in a single launch.

 Robert Zubrin had proposed a Moon mission architecture using the Falcon Heavy with his "Moon Direct" proposal but it would require two launches of the Falcon Heavy to do it. This alternative approach could do it in a single launch provided it is indeed possible to produce an Apollo Command module analogue of dry mass only 3 tons.

Architecture 2: an Apollo sized capsule.

 The Apollo architecture that had the Apollo Command Module to carry the astronauts for the in space portion of the trip from LEO to lunar orbit with a separate smaller capsule for the lander, had an advantage in providing backup capability. This was quite fortunate during the Apollo 13 mission when the Apollo LEM had to sustain the crew for a part of the time on the way back to Earth.

 There is still the question of whether you can make the Apollo Command Module analogue only at 3 tons dry mass. So here we'll do the calculations for an analogous architecture to that of Apollo with a main crew capsule for the in-space portion of the flight and a smaller, separate crew module for the lander.

 I estimated above the Blue Moon Mk1 lunar lander has about a 6 to 1 propellant load to dry mass ratio, at 18.35 tons prop load to 3 tons dry mass. But the Mk1 was designed to do all the propulsion from LEO, to translunar injection(TLI), to low lunar orbit insertion, to lunar landing, with a 3 ton cargo. If the only thing required is to go from low lunar orbit to the lunar surface and back with a 3 ton crew module then a much smaller lander can be used. 

 I'll assume you can a smaller lander at 1/3rd the Mk1 size with a 6 ton prop load while maintaining the 6 to 1 prop mass to dry mass ratio, so 1 ton dry mass. First, from the Earth-Moon delta-v table, the delta-v one way from low lunar orbit to the lunar surface is 1,870 m/s. Then the round-trip delta-v is 3,740. Note now, the smaller lunar lander can provide a delta-v of:

 445*9.81Ln(1 + 6/(1 + 3)) = 4,000 m/s, sufficient for the round-trip from lunar orbit to the surface and back to lunar orbit.

 Now we need a propulsive stage to do the burn to insert the 6 ton main crew capsule and 10 ton lander into low lunar orbit, and to do the burn to bring the main capsule back to Earth, a la the Apollo architecture. For this we'll use a stage half-size to the Mk1 at 9 ton prop load and 1.5 ton dry mass.

 The burn to escape low lunar orbit is commonly estimated as 800 m/s to 900 m/s, same as that for the burn to enter into low lunar orbit. Then 2 tons of propellant is required to be left over as reserve for the return of the primary capsule to Earth, the lander being jettisoned a la the Apollo architecture:

445*9.81Ln(1 + 2/(1.5 + 6)) = 1,030 m/s.

 Then 7 tons of propellant out of 9, with the 2 tons left in reserve for the return, is sufficient to put the 6 ton primary capsule and the 10 ton lander into low lunar orbit:

445*9.81Ln(1 + 7/(1.5 + 6 +10 +2)) = 1,340 m/s.

 The rather large margin of 1,340 m/s over the maximum 900 m/s needed to insert into low lunar orbit suggests we might be able to do with a somewhat smaller stage for this purpose, perhaps 7 tons instead of 9 tons prop load.

 Now the total mass that needs to be sent to TLI is 9 + 1.5 + 6 + 10 = 26.5 tons. We'll use again the upper stage of the Delta IV Heavy to do the TLI burn:

465*9.81Ln(1 +27.2/(3.5 + 26.5)) = 2,940 m/s. 

 This is slightly less than the value commonly given for TLI in the range of 3,000 m/s to 3,100 m/s. But the propulsive stage that's used to insert into lunar orbit had so much margin that it could be used to provide the slight extra push to make TLI.

 Or as I mentioned that propulsive stage for the lunar orbit insertion, essentially reprising the role of the Apollo's Service Module, had so much margin we could make it smaller to ca. 7 tons prop load. Then the TLI total mass would be the same as the Architecture 1 case. And the Delta IV Heavy's upper stage could get the total mass to TLI on its own. 

 It's still quite notable that doing it either way we still could launch the full system to orbit on a 60 ton to LEO launcher.

Flights to the Moon at costs similar to costs of flights to the ISS. 

 I said Artemis is really Apollo redone based on its payload size. It is not Constellation. It is not "Apollo on Steroids". Does it have any value then? I am arguing the goal of getting sustainable lunar habitation is important and doable now. It probably can't be done by Artemis though in a sustainable fashion considering that both the Orion capsule and SLS already each, separately cost $2 billion per flight. When you add on the over-large proposed landers the SpaceX HLS or the New Glenn MK2 each costing ca. $2 billion per flight, and the the Boeing EUS, advanced composite casing SRB's, and lunar Gateway, the total per flight would be in the range of $8 billion to $10 billion per flight.

 It is now becoming increasing likely that Artemis will be cancelled. The only question now is will it be cancelled before Artemis II or will Artemis II be allowed to fly and then the program would be cancelled.

 However, the most important fact is sustainable lunar habitation can be done following the commercial space approach making use of already existing space assets. As I mentioned the combined effect of both these factors can cut the costs of such missions by a factor of 1/100. For example both the Falcon Heavy and the New Glenn cost in the range of ca. $100 million. The small size of the additional in-space stages probably can be done for less than $100 million under the commercial space approach.

 And the crew capsules? An unexpected calculation suggests they can be done together for less than $100 million. For instance back in 2009, Orbital Science contracted Thales Alenia  to construct the Cygnus capsule for 180 million euros for 9 capsules, about 20 million euros each.

 A further contract Thales Alenia made with Axiom Space illustrates how low cost such modules can be while illuminating also how much more expensive space systems are when government funded compared to being privately funded. A contract Thales Alenia made to Axiom Space for two space station modules was only $110 million for two:

THALES ALENIA SPACE TO PROVIDE THE FIRST TWO PRESSURIZED MODULES FOR AXIOM SPACE STATION
14 JUL 2021
Rome 15 July, 2021 – Thales Alenia Space, Joint Venture between Thales (67%) and Leonardo (33%), and Axiom Space of Houston, Texas (USA), have signed the final contract for the development of  two key pressurized elements of Axiom Space Station - the world’s first commercial space station. Scheduled for launch in 2024 and 2025 respectively, the two elements will originally be docked to the International Space Station (ISS), marking the birth of the new Axiom Station segment. The value of the contract is 110 Million Euro.

https://www.thalesgroup.com/en/worldwide/space/press_release/thales-alenia-space-provide-first-two-pressurized-modules-axiom-space

 The individual modules have about 75 cubic meters pressurized space for four crew members, and already have life support systems.

 Now compare that to the HALO module Northrop Grumman contracted with NASA to produce at a cost of $935 million:

Northrop charges on lunar Gateway module program reach $100 million.

by Jeff Foust

January 25, 2024

Northrop received a $935 million fixed-price contract from NASA in July 2021 to build the module, which is based on the company’s Cygnus cargo spacecraft. HALO will provide initial living accommodations on the Gateway and includes several docking ports for visiting Orion spacecraft and lunar landers as well as additional modules provided by international partners. It will launch together with the Maxar-built Power and Propulsion Element (PPE) on a Falcon Heavy.

https://spacenews.com/northrop-charges-on-lunar-gateway-module-program-reach-100-million/

Based on the "Super" 4-Segment version of the Cygnus, it might have a volume of ca. 33.5 cubic meters:

 The Axiom Space AxH1 habitation modules at 70 cubic meters have double the space of the HALO modules but, as privately financed, cost less than 1/10th as much as government financed HALO modules.

 The needed crew module would be well cut down in size from the 70 cubic meters of the Axiom space station habitation module, with a comparable reduction in cost. Addition of a heat shield would cost a fraction of the total cost of the crew module itself.

 Then the crew modules for the main capsule or of the lander module might cost in the range of a few 10's of millions of dollars.

A contingency plan for a fast return of the U.S. to space.

Copyright 2014 Robert Clark

Why NASA and Congress Spent Four Hours Shouting At Each Other About Russia.
April 8, 2014 // 04:17 PM EST

http://motherboard.vice.com/read/why-nasa-and-congress-spent-four-hours-shouting-at-each-other-about-russia

 The congressmen kept asking for a short-term contingency plan to return America to space in case of seriously deteriorating U.S/Russia relations and Bolden kept responding with the three-year plan to have commercial crew flying. But there is a shorter term plan. BOTH SpaceX and Boeing have said they could be flying crew by next year with funding. So if the congressmen want a shorter term contingency plan, provide that required extra funding.

 At the Humans 2 Mars 2014 conference I asked Bolden about such a contingency plan. It's about at the 15 minute mark in this video:

 He responded that SpaceX has not been selected yet as the crew launch provider. OK, then also fund Boeing so they can also return crew to the ISS by 2015.

 There has been talk in Congress of only having one crew launch provider. I strongly disagree with that plan. We all saw what can happen when you only have one launch provider and that one goes down, as happened with the shuttle. SpaceX is furthest along so they should be one of the providers. But on the other hand the Boeing capsule would be carried on the Atlas V which has had a remarkable string of successful launches, which SpaceX is nowhere near to matching yet.

 Russian Deputy Prime Minister Dmitry Rogozin mocked the U.S. space sanctions against Russia saying NASA would need to get a trampoline to get its astronauts to the ISS. This led Elon Musk to state through his twitter account that SpaceX would be revealing its man-rated Dragon 2 at the end of May:

 Now, if SpaceX is flying their own crews to LEO in 2015 and there is still a strained relationship between the U.S. and Russia then, then it would be extremely embarrassing for NASA to still be paying Russia to ferry NASA astronauts to the ISS when SpaceX will already be flying American crews to LEO.

 A solution would be for NASA to at least draw up a contingency plan including cost estimates of how much extra funding it would take to also take NASA astronauts to the ISS. Then the onus would be on Congress to decide if they want to provide NASA with the extra funding to do so.

   Bob Clark

"Golden Spike" Circumlunar Fights, Page 2.

Copyright 2014 Robert Clark

 In the blog post "Golden Spike" Circumlunar Flights, I suggested that the Falcon 9 1.1/Dragon could do a circumlunar mission with a half-sized Centaur, of ca. 10 mT propellant load, to do the translunar injection.

 Interestingly it might be possible to do without even needing the extra in-space stage. Elon Musk has said through his Twitter account that the 13 mT payload capability was actually a reduction of the F9 V1.1's true capability due to reusability considerations. Gwynne Shotwell confirmed this on a TheSpaceShow interview on Friday, Mar. 21 at about the 9 minute mark. She said the quoted payload on their web site for the F9 v1.1 is about 30% less than that of a one-use version.

 This would put the expendable version in the range of the 16.6 metric tons to LEO given on NASA's site:

NASA Launch Services Program's
Launch Vehicle Performance Web Site.

http://elvperf.ksc.nasa.gov/elvMap/staticPages/nls1.htm

 The point is this would be just about at the payload capability to do translunar with the Dragon using just its onboard Draco, or upgraded SuperDraco, thrusters. On the "NASA Launch Services Program" site, click the link for the Performance Query Tool and select the Falcon 9 and "elliptical" orbit option. Enter in 36000 km for the altitude corresponding to geosynchronous transfer orbit (GTO) and 28.5 degrees for the orbital inclination corresponding to launch from Cape Canaveral. Then the calculator gives the payload to GTO as 5745 kg.

 As shown here the delta-v to GTO is 2,500 m/s:

 

 Then translunar injection (TLI) at 3,100 m/s would only require an additional 600 m/s delta-v. The Dragon has a dry mass of 4,200 kg and a propellant mass of 1,290 kg. SpaceX has not released the Isp of the hypergolic thrusters on the Dragon, but they typically are in the 320 s range in vacuum. Then it could carry 1,800 kg payload to the 600 m/s needed to reach TLI:

320*9.81ln(1 + 1290/(4200 + 1800)) = 610 m/s.

  Actually that 1,800 kg payload would put the total mass beyond the 5745 kg capacity to GTO. Smaller payload say in the 250 kg range would be doable.

 Such missions would be important to do since at a perhaps $120 million total launch price for the Falcon and Dragon it would show lunar missions are possible without requiring huge launchers such as the Saturn V, Ares V or SLS.

   Bob Clark

UPDATE, April 1, 2014:

 On TheSpaceport.us forum, DocM informed me via PM that in an environmental impact statement SpaceX gave the propellant load for the Dragon as 1,388 kg. This would raise the max payload to reach 600 m/s delta-v to 2300 kg. Again though this would put the total mass outside the range that could be lofted to GTO. Likely you would still have to limit the payload to ca. 250 kg or so.

  

 

 

The Coming SSTO's: Page 2.

Copyright 2013 Robert Clark

 In the blog post, The Coming SSTO's I calculated some delta v's that suggested we already have the capability to do SSTO's with significant payload. However, here I'll provide some more accurate estimates by using Dr. John Schilling's Launch Vehicle Performance Calculator page. I'll go back to the Atlas rocket SLV-3 Atlas / Agena B. The specifications are given here:

SLV-3 Atlas / Agena B.
Family: Atlas. Country: USA. Status: Hardware. Department of
Defence Designation: SLV-3.
Standardized Atlas booster with Agena B upper stage.
Specifications
Payload: 600 kg. to a: 19,500 x 103,000 km orbit at 77.5 deg
inclination trajectory.
Stage Number: 0. 1 x Atlas MA-3 Gross Mass: 3,174 kg. Empty Mass:
3,174 kg. Thrust (vac): 167,740 kgf. Isp: 290 sec. Burn time: 120 sec.
Isp(sl): 256 sec. Diameter: 4.9 m. Span: 4.9 m. Length: 0.0 m.
Propellants: Lox/Kerosene No Engines: 2. LR-89-5
Stage Number: 1. 1 x Atlas Agena SLV-3 Gross Mass: 117,026 kg.
Empty Mass: 2,326 kg. Thrust (vac): 39,400 kgf. Isp: 316 sec. Burn
time: 265 sec. Isp(sl): 220 sec. Diameter: 3.1 m. Span: 4.9 m. Length:
20.7 m. Propellants: Lox/Kerosene No Engines: 1. LR-105-5
Stage Number: 2. 1 x Agena B Gross Mass: 7,167 kg. Empty Mass: 867
kg. Thrust (vac): 7,257 kgf. Isp: 285 sec. Burn time: 240 sec. Isp(sl): 0
sec. Diameter: 1.5 m. Span: 1.5 m. Length: 7.1 m. Propellants: Nitric
acid/UDMH No Engines: 1. Bell 8081
http://www.friends-partners.org/partners/mwade/lvs/slvgenab.htm

 We see stage 1 called the sustainer stage has nearly a 50 to 1 mass ratio. However, the Atlas had an unusual "stage and a half" structure where engines needed to lift off from the pad were jettisoned later on in the flight, leaving only a smaller, lower thrust engine behind. This engine which is the one used in stage 1, did not have enough thrust to lift off from the pad. So as in The Coming SSTO's post,  I'll replace it with the NK-33 engine which has now flown successfully on the Orbital Sciences Antares. 

 The propellant load remains 114,700 kg as in the original Atlas but the dry mass increases to 3,086 because of the heavier engine. The vacuum Isp is 331 s for the NK-33, and the vacuum thrust is 1,638 kN. Now input these numbers into Schilling's calculator. Select "No" for the "Restartable Upper Stage?" option and Cape Canaveral for the launch site. For the orbital inclination choose 28.5 degrees to match the latitude of Cape Canaveral. Then the Calculator gives these results:

====================================================

Mission Performance:

Launch Vehicle:	User-Defined Launch Vehicle
Launch Site:	Cape Canaveral / KSC
Destination Orbit:	185 x 185 km, 28 deg
Estimated Payload:	4113 kg
95% Confidence Interval:	2860 - 5625 kg

====================================================

 This value of 4,113 kg is remarkable in being close to that of the payload capability of the full Antares at 5,000 kg, a rocket of twice the gross mass, using two stages and two of the NK-33 engines on the first stage.

 Based on this, this SSTO version could be significantly cheaper than the current Antares. Plus in being only liquid fueled, it could be used as a manned launcher. Note that Orbital already has the Cygnus capsule which with the addition of a heat shield and life support could be a manned capsule.

 The mass ratio of 50 to 1 for the original Atlas is so high it would be interesting to calculate the payload capacity if we used instead the lower Isp Merlin 1D engine. By the SpaceX page, nine Merlin 1D's have total vacuum thrust of 6,672 kN. So one is 741 kN. We will need two to lift off, at 1,482 kN vacuum thrust. The two Merlin 1D's together weigh about 330 kg less than the NK-33 case, so subtract that much from the dry mass of the NK-33 case. However, the Isp is also reduced to 311 s Isp for the Merlin:

 Then Schilling's calculator gives:

====================================================

Mission Performance:

Launch Vehicle:	User-Defined Launch Vehicle
Launch Site:	Cape Canaveral / KSC
Destination Orbit:	185 x 185 km, 28 deg
Estimated Payload:	3025 kg
95% Confidence Interval:	1952 - 4331 kg

====================================================

 It is the quite high mass ratio that leads to these rather high payload capabilities.

 SpaceX might not be inclined to support such an experiment, as they are deeply invested in keeping the Falcon 9 first stage and Merlin 1D engines. However, Orbital Sciences farms out its construction of the Antares first stage to a company in the Ukraine.  So they may be inclined to try a new stage that would at the same time prove to be a revolutionary step of creating an operational SSTO.                                                                                            

   Bob Clark

On the lasting importance of the SpaceX accomplishment, Page 4: how the Ariane 6 can beat both SpaceX and the Russians.

Copyright 2013 Robert Clark

Europe Urged To Halt Work on ‘Dead End' Ariane 6 Design.
By Peter B. de Selding | May. 30, 2013

The academy is urging the agencies to stop work on the Ariane 6 they approved in November with a view to beginning full development in 2014. The academy-favored rocket would use liquid propulsion instead of solid, and would face four more years of preparatory work before moving to full development in 2018.
In the meantime, the academy says, Europe should focus on an upgraded heavy-lift Ariane 5 that would fly for a decade before both it and the Europeanized version of Russia’s medium-lift Soyuz rocket are replaced by the all-liquid Ariane 6 in 2027. This rocket, called Ariane 5 ME, has been in design for several years. Continued work on it was approved, alongside Ariane 6, at the November meeting of European Space Agency (ESA) governments.

http://www.spacenews.com/article/launch-report/35546europe-urged-to-halt-work-on-%E2%80%98dead-end-ariane-6-design

 The Academy should also emphasize another key advantage of the liquid-fueled version of the Ariane 6 that it could be used for a manned launch vehicle.
 Note that Russia is raising their prices to $73 million per seat or $220 million for three. This is greater than the launch cost of the full 20 metric ton class Ariane 5. The smaller Ariane 6 would certainly be cheaper than that. By producing this liquid fueled Ariane 6, Europe could also get their own manned space flights and more cheaply than by paying the Russians.
 Both Russia and China have their own manned spaceflight programs, as will the U.S. in the near, short time frame. And even India and Japan are planning their own manned spaceflight programs. The Japan case is quite notable in that their plan is to use twin cryogenic engines of similar characteristics to the Vulcain II.
 The European Union has been the highest economic power or a close second to the U.S. in the world over the last few years. It should be regarded as unacceptable by European space advocates, private, governmental, and industry, that there has been no plan to give Europe a manned space program as with these other space agencies.
 Such a manned-capable launcher could be done more quickly and cheaply by using a commercial space approach. The Falcon 9 and the Antares only took 4 years and a few hundred million dollars in development cost that had to be paid by NASA.
 I also estimate the cost per launch of a single stage version could be done for half the $127 million cost given by the Academy in that report for their version of the Ariane 6, vastly undercutting the Russians:

On the lasting importance of the SpaceX accomplishment, Page 3: towards European human spaceflight.
http://exoscientist.blogspot.com/2013/05/on-lasting-importance-of-spacex.html

 Here's an argument for producing the Ariane 6 at a faster time frame than just 2027. The Ariane 6 is supposed to be one-half to one-third as expensive as the Ariane 5. The Ariane 5 is already being used to deliver cargo to the ISS but using the very expensive to develop and produce ATV. In fact ESA doesn't want to produce any more ATV's after the last one to launch in 2014.
 But if you have this less expensive launcher in the Ariane 6 then you have a much less expensive route to sending cargo to the ISS. But then you need a pressurized capsule to transport it. Why spend the expense of developing a new small pressurized capsule when you already have one in the European developed Cygnus? (By the way this raises an interesting economic question I'll discuss at the end.)
 SpaceX is charging NASA $133 million to transport a maximum of 6,000 kg to the ISS. Note this is well above the launch cost of the Falcon 9 alone. The large extra cost is due to the use of the expensive Dragon capsule. The Ariane 6 would have comparable payload capacity as the Falcon 9 but using a 2,000 kg lighter capsule in the Cygnus. Then it could be at or above the cargo capability of the Falcon 9 to the ISS. And from the estimated launch cost of the Ariane 6 and the low cost of the Cygnus compared to the Dragon their price could be at or below that of the Falcon 9/Dragon. How's that for wanting to be competitive with SpaceX?
 Now, the Academy wants ESA to make a liquid-fueled version of the Ariane 6 instead of the planned solid-fueled one. Imagine you have that and it is being used to send cargo via the Cygnus capsule to the ISS. It's not much of leap at all that if you add life support and a heat shield to the Cygnus then you would have a European vehicle capable of sending astronauts to the ISS as well. And you could do it at a price to undercut the Russians.
 I want to argue again here for the commercial space approach for accomplishing this. The 2027 time frame for such a liquid fueled Ariane 6 is following the usual glacial pace of government financed space programs. This would be near the end of the ISS (expected) extended life time. However, both SpaceX and Orbital Sciences by following the commercial space approach were able to develop their launchers in 4 years. Commercial space is both cheaper and faster than government space.
 To do the cost sharing of commercial space though the industry partners, or their investors, would have to be convinced it could be profitable. Note that SpaceX has gotten a $1.6 billion contract from NASA for delivering cargo to the ISS. The $127 million per launch cost estimated by the Academy is coming from the large, billion dollar, development costs under the usual governmental financing approach that would need to be recouped. Commercial space has proven though that both total development cost and the portion paid by the government are a fraction of those of the usual governmental financing. Then getting a similar billion dollar ISS supply contract as SpaceX and with a development cost that, literally, might only be a few hundred million dollars, would result in such a contract being highly profitable.

 About that economic question I mentioned above, Orbital Sciences paid for the development of the Cygnus to the Italian Space Agency(ISA). But certainly the ISA would not want to turn over the full rights to the Cygnus to a foreign company. It's quite likely ISA retains ownership of the Cygnus. This becomes interesting in regards to the price they would charge for the Cygnus compared to the price Orbital Sciences would charge.
 Because Orbital paid for the development of the Cygnus they would want to recoup that cost in the price they charge. But the ISA does not have to recover that cost. This means they could charge much less. But then why would anyone pay for the higher cost from Orbital when they could get it cheaper from the ISA?
 A puzzling question. It may be Orbital retains the rights to sell the Cygnus to NASA or even for all American launches.


    Bob Clark

On the lasting importance of the SpaceX accomplishment, Page 3: towards European human spaceflight.

Copyright 2013 Robert Clark 

 

European Human Spaceflight

The EU released a report critical of the ESA's policy on new launchers:

The EU Seems to Really Dislike ESA’s New Launch Vehicle Policy.
Doug Messier
on March 17, 2013, at 5:57 am
www.parabolicarc.com/2013/03/17/the-eu-seems-to-really-dislike-esas-launch-vehicle-policy/

  The report is rather opaque about what changes the EU wants in space policy as opposed to what the ESA is proposing. One thing I noted is that it wants the ESA to keep up with technological advances the other space programs in the world are embarking on.

 This possibly might relate to the proposal of the Ariane 6 to use all solids on the lower stages. This is going backwards, not forwards in technology. A forwards suggestion for the Ariane 6 would have been the option that uses liquid fuel for a core stage simply by adding a second Vulcain to the Ariane 5 core stage.
 Note this would have high commonality with the current Ariane 5 which the ESA also wants to save on costs rather than having to design entire new solid lower stages. But the most important advantage of this is a key technological advance it would provide to keep up with the other space faring nations.

 Russian and China have manned orbital launchers, and the U.S. will again also in the short near term. India is even planning on manned launchers. But the ESA has no plans on producing a manned launcher. Space advocates in Europe should regard this as unacceptable. But the key point is by using the multi-Vulcain option for the Ariane 6 this would provide Europe with a manned spaceflight capability.

 Another source of friction with the EU is that ESA is constrained to apportion work according to members financial participation, while the EU is under no such constraints:

CNES Design Team Sets ‘Triple-seven’ Goal for Ariane 6.
By Peter B. de Selding | Jan. 2, 2013
https://spacenews.com/33019cnes-design-team-sets-triple-seven-goal-for-ariane-6/

From the article:

...after months of hard selling that saw them pitted against much of France’s industry, CNES officials last year convinced Fioraso that Ariane 6 — less expensive and less powerful than Ariane 5, and carrying just one satellite at a time to orbit — is the way of the future.
The design of the rocket — two solid-fueled lower stages and a cryogenic upper stage, plus solid-fueled strap-on boosters — was frozen Nov. 21 during a meeting of ESA government ministers.ESA Launcher Director Antonio Fabrizi said this design, and no other, is what ministers approved.

and:

Ariane 6 has been conceived from the start as a “next-generation” rocket that in many ways looks like a throwback — more of a less-expensive Lockheed Martin Atlas 5, or a Proton launched from the equator. Ariane 5 can do more things for more customers.
But if it meets its design goal, Ariane 6 will reach a financial equilibrium that has eluded Ariane 5. CNES officials say economic criteria account for 43 percent of the design decisions made for the rocket, with technical criteria accounting for just 30 percent.
The remaining 27 percent of the design choices are being made on the basis of Europe’s existing industrial capacity.
French industry is responsible for around 50 percent of the construction of Ariane 5. Eymard said the agency assumes France will carry about the same load for Ariane 6.
Beyond the French contribution, all bets are off. CNES has penciled in Germany at 25 percent, and Italy at 10-15 percent. The Italian share should be relatively easy to secure because Italy already is heavily involved in production, with Snecma of France, of the solid-fueled strap-on boosters used on the Ariane 5 rocket. Italy is also the lead investor in the new Vega small-satellite launcher, which made its inaugural flight in early 2012.
Because of the all-but-guaranteed work share of Italian industry in the Ariane 6 solid-fueled stages, the Italian government is not likely to resist taking its 10-15 percent stake despite its public-debt crisis.
Ensuring German industry sufficient work will not be as straightforward, European government and industry officials said.

 This article shows the difficulty the ESA will have in developing innovative launch solutions. The biggest factor in deciding which launcher to develop is how much work it can provide to the ESA, member countries. This supersedes even lowered costs.

The ESA could develop a low cost launcher that would be comparable in cost to the SpaceX Falcon 9, AND moreover would give Europe an independent manned launch capability simply by adding a second Vulcain to the Ariane 5 core. Ironically though, this option is not chosen because it would be TOO low cost: it would be simple, quick - and not provide enough work to the ESA member countries.

The only way Europe is going to get low cost space access, it now appears, is if it is done under the commercial space approach. As proven by SpaceX this can cut 90% (!) off the development costs when privately financed. And in fact it should be even easier and cheaper than the SpaceX case since the components already exist in the Ariane 5 core, built in France, and Vulcain II engines, built in Germany. Even the capsule for the manned launchers is largely already designed in the Orbital Sciences, Cygnus capsule, which is actually built in Italy. You would just need to supply life support and heat shield to the capsule already designed to be pressurized.

 The only thing needed are entrepreneurs in Europe like Elon Musk in the U.S. with the insight to carry it out. In the blog posts "On the lasting importance of the SpaceX accomplishment" and "On the lasting importance of the SpaceX accomplishment, Page 2" I discussed the fact that space development costs were cut dramatically by SpaceX by private financing.

 NASA has found with its commercial crew program that it can develop manned launchers in general at lower costs by opting for a more commercial approach to their development. In fact NASA's commercial space program was presaged by the Air Force's Evolved Expendable Launch Vehicle (EELV) program. The Air Force only had to pay $500 million out of a $3.5 billion development cost for the Delta IV and $500 million out of a $2 billion development cost for the Atlas V. For the Delta IV, that's a 86% (!) savings in development cost.

 NASA also has saved in development cost on Orbital Science's Antares launcher. It only had to pay $288 million out of a development cost of $472 million for a 5 metric ton class launcher. 

 Then the suggestion to the EU is to institute a similar program for European manned launchers. Politically the ESA appears to be set on the all-solid Ariane 6. But what the EU could do is put out a request to European industry for commercially developed man-rated launchers that would be largely privately funded aside for perhaps some seed money, a la SpaceX. To sweeten the pot, the EU could state that as part of their policy they will use these European launchers for their manned flights as long as they are comparable in price to say what they are paying the Russians for their launchers.

 The Russians are charging $63 million per seat for flights on the Soyuz, so for three crew in the range of $190 million. This is almost the cost of a full Ariane 5 launch, a vehicle capable of 20 metric tons (mT) to LEO.

 A vehicle capable of carrying a manned capsule could be done at a 5 mT payload capability, a quarter the size of the Ariane 5. SpaceX spent $300 million developing the Falcon 9, capable of 10 mT to LEO. Then a vehicle half the size, that was also largely privately funded as was the Falcon 9, might cost ca. $150 million.

 Considering the payload for our twin-Vulcain Ariane likely will be above 5 mT though, we might instead estimate the development cost as $200 million based on how much JAXA spent to add a second cryogenic engine to the H-IIA core.

 Also, I've been informed by people who aware of CNES studies on a multi-Vulcain Ariane that the estimated price for the two-Vulcain Ariane 5 core would be only 50 million euros, about $60 million(!) So for only a ca. $200 million development cost and a $60 million launch cost the ESA could have manned spaceflight ability.

 Another source of income for such a launcher with the Cygnus capsule would be deliveries to the ISS. SpaceX is charging NASA about $133 million for ca. 6,000 kg delivery of cargo using the Falcon 9. Part of this inflated cost above the $54 million cost of the Falcon 9 is the use of the expensive Dragon capsule. The Cygnus is a smaller capsule with a much smaller development cost, so would be much cheaper than the Dragon. Using a ca. 8,000 kg payload for the launcher and ca. 2,000 kg mass for the Cygnus, this launcher could match the 6,000 kg delivery capacity of the Falcon 9 at a much reduced price.

 European Moon Flights

 According to NASA administrator Charles Bolden, NASA will not be returning us to the Moon but may engage in partnerships with other space agencies or private entities who could. Then it's interesting the ESA has the required lightweight in-space stages and lightweight capsule in the Cygnus to accomplish this at low cost.

Another key fact is that NASA has shown with SpaceX and now with Orbital Sciences that development costs can be cut drastically (by 80 to 90% !) by following a commercial approach. Then this could be a project NASA could encourage, at low cost to NASA, by partnering with ESA and private entities like Golden Spike, Planetary Resources, Inc., etc, while at the same time satisfying the critics who want us to return to the Moon.


   Bob Clark

"Golden Spike" circumlunar flights.

Copyright 2012 Robert Clark

"Golden Spike Co" has released a paper describing their return to the Moon plans:

An Architecture for Lunar Return Using Existing Assets.
by James R. French et. al.
http://goldenspikecompany.com/wp-content/uploads/2012/02/French-et-al.-Architecture-Paper-in-AIAA-Journal-of-Spacecraft-and-Rockets.pdf

 It gives several different architectures and types of missions. But on page 8 it gives the payload capability of the Falcon 9, presumably the new version Falcon 9 v1.1, as 16,700 kg. However, on the SpaceX site it's given as 13,100 kg:

http://www.spacex.com/falcon9.php#launch_and_placement

 Interestingly at the 16.7 mT number you can do a manned circumlunar mission on a single Falcon 9 + Dragon, even including a launch abort system(LAS), by using a half-size Centaur as the in-space stage. But at the 13.1 mt number it becomes more problematical .
 Such a mission would be very important to accomplish. Recall the Apollo 8 mission was a manned lunar flyby that served as the prelude to the Apollo 11 mission. It is regarded then as being a part of the costly Apollo program, requiring the expensive Saturn V launcher.
 The skepticism among many about the Golden Spike plan or other commercial lunar plans is the idea it would require large, highly expensive Saturn V class launchers. However, if the manned flyby could be done by a single launch by what is still just a medium size launcher in the Falcon 9 v1.1 it would show that by going small and following a low cost, commercial approach, that a low cost return to the Moon is feasible.
 The Falcon 9 v1.1 will cost in the $60 million range, and we might estimate the half-size Centaur to be in the $15  million range. So the launch cost for such a mission might be in the $75 million range.
 As I discussed before in regards to using the first test flights of the Falcon Heavy for unmanned BEO test flights, the test flights of the Falcon 9 v1.1 could serve for unmanned test flights for this lunar flyby. Since SpaceX needs to do such tests anyway most of the cost of the Falcon 9 and Dragon capsule would be borne by SpaceX. Then you could have Golden Spike only paying ca. $15 million for the half-size Centaur.
 There would be some development cost of course beyond that for this half-size Centaur. For one thing you would have to make the cryogenic propulsion undergo less boiloff for 1 to 2 week missions. ULA has done studies on this so should be doable but still it has to be carried out in practice. An advantage of this would be that this half-size Centaur is about the size you need for the lander. So the lander could be derived from this, and the development cost for the two stages could be reduced.

CALCULATIONS
 The Golden Spike landing plan specifies using two Falcon Heavy's even though it uses a Dragon sized capsule. This is more than 100 mT to LEO. This is puzzling since the advantage of using a lightweight capsule is that it should require smaller amounts to be launched to orbit, known as IMLEO, initial mass to low Earth orbit. For instance the Early Lunar Access plan only requires 52 mT to orbit using a small two-man capsule. However, I believe the Golden Spike paper by French et. al. explains where the discrepancy arises.
 On page 13 is given a table of some masses for different possible propulsive stages. The mass for the Dragon with trunk and crew and supplies is 8,853 kg, well above the given dry mass of the Dragon capsule at 4,200 kg. The trunk section is less than 1,000 kg and the propellant for the Dragon is at 1,290 mT. The mass for crew and supplies in the Golden Spike paper is given as 300 kg. Evidently then the extra mass to get to a 8,853 kg mass is coming from the launch abort system (LAS).
 In any case a 8,853 kg mass would be at the mass of the Orion capsule and we would lose any advantage of a lightweight architecture. Then I suggest an alternative to the SpaceX LAS that has the LAS permanently integrated into the capsule.
 We could use again a tower type LAS that would be jettisoned prior to reaching orbit. To estimate its mass we might make a comparison to the Orion LAS. The Orion LAS is at 6,000 kg. The Dragon is at half the mass of the Orion capsule. Then we can estimate the mass for a tower type LAS for the Dragon as 3,000 kg.
 This is also a high additional mass. However, typically a tower LAS is jettisoned soon after first stage separation. So we can estimate how much this will subtract from the payload to orbit by making a comparison to how much the payload is reduced for an increased dry mass to the first stage. A rule of thumb is that every kilo of mass added to the first stage dry mass subtracts off 1/10 of a kilo from the payload. So we can estimate 300 kg being subtracted off the payload.
 Now we'll estimate the size of a cryogenic stage needed to take the Dragon to a circumlunar mission. In the table on page 13 in the Golden Spike paper is given a cryogenic, LH2/LOX, stage at 1,196 kg dry mass and 7,534 kg propellant mass. This is a mass ratio of 7.3 to 1. Notably this is less than that of current Centaur upper stages at about 10 to 1. This is because mass ratio improves as you scale up your rocket stages.
 This rocket stage would be sufficient to carry the Dragon's 4,200 kg dry mass plus the 300 kg for crew and supplies using RL-10 engines. The delta-v for trans lunar injection(TLI) is 3,150 m/s. Using a 451 s Isp for the RL-10 engines we get a delta-v of:

451*9.81ln(1 + 7534/(1196 + 4500)) = 3,700 m/s.

 But because of the loss of payload capacity due to the LAS from  SpaceX's cited payload to LEO of the Falcon 9 v1.1 of 13.1 mT, this would be slightly more mass than can be carried to LEO. So we'll use a slightly smaller stage. Let the propellant mass be 7,000 kg. Keeping the same 7.3 mass ratio, this corresponds to a dry mass of 1,100. Then the delta-v will be:

451*9.81ln(1 + 7000/(1100+ 4500)) = 3,600  m/s, still sufficient for the TLI.


  Bob Clark

SLS for Return to the Moon by the 50th Anniversary of Apollo 11.

Copyright 2012 Robert Clark

Credit: NASA's Space Exploration Vehicle (SEV).

Very interesting report about using NASA's proposed Space Exploration Vehicle for cislunar space exploration:

Lunar Surface Access from Earth-Moon L2.
A novel lander design and study of alternative solutions.
1 October 2012 | Washington, DC

http://www.sei.aero/eng/papers/uploads/archive/SEV-L2-Lander-Presentation_1Oct2012.pdf

 The report proposes using the lightweight SEV, at only a 3 mT empty weight, and all cryogenic propulsion as a shuttle between the L2 space station NASA has recently discussed and the lunar surface. However it could also be used as the crew capsule between LEO and the Moon's surface.
 The architecture discussed is very interesting in that the SEV would be used as the single crew module to carry the crew all the way from the L2 station to the lunar surface and back again, i.e., no separate lander crew module. There would also only be a single propulsive stage to carry the SEV from low lunar orbit to the lunar surface and back to lunar orbit, i.e., no separate lunar descent and ascent stages.
 This has similarities to the architecture for the Early Lunar Access(ELA)[1] proposal of the early 90's. This also used all cryogenic space stages to save weight, only 52 mT required to LEO. ELA also saved weight and cost by using a single crew capsule for the entire flight from LEO to the lunar surface and back again. It also used a single propulsive stage for lunar descent and ascent. But instead of linking up with a stage waiting in lunar orbit for the return, the ELA proposal was to have this single lander stage return all the way back to LEO.
 An alternative architecture discussed on page 23 in this report on using the SEV for cislunar travel does not use the method of first stopping in lunar orbit, then having a separate lunar lander stage. Instead it uses the "direct descent" method of descending directly to the lunar surface. This landing method is analogous to that used in the ELA proposal to save propellant. Interestingly the SEV report on page 23 gives the delta-V for the direct descent method as 2,610 m/s. This compares to the 760 m/s + 2,150 m/s = 2,950 m/s for the method that first stops in lunar orbit, then descends to the surface as indicated in the image above. So according to this report a savings of 300 m/s in delta-V for the trip from L2 to the Moon is possible using direct descent, a significant savings.
 I had wondered if it was possible to save delta-V and propellant in this blog post 'Delta-V for "direct descent" to the lunar surface?'[2]. The SEV report suggests it may be possible to save in the range of 300 m/s by the direct descent method.
 The only technical complaint raised against the feasibility of the ELA proposal back in the 90's was the suggestion of getting a 2-man crew capsule at only a 3 mT empty weight. So the fact the SEV is expected to have this low an empty weight is important, since it suggests the possibility with just the 70 mT first version of the SLS of a manned lunar lander mission using currently existing cryogenic stages.
 Actually the 70 mT payload of the SLS is so much better than the 52 mT needed for ELA that likely we could even use a heavier hypergolic stage for the lunar ascent stage. During the early planning of the Apollo program when the possibility an engine might not ignite was regarded as a definite possibility, it was decided to use hypergolics, which ignite on contact, for the lunar ascent stage. At this point though the cryogenic RL10 engines have had decades of use and are regarded as highly reliable.
 Still for these first versions of these new lunar landers we might still want the certainty of using hypergolics for the ascent stage. I suggest using the engine and propellant tanks of the shuttle orbiter OMS pods for the purpose. This would be quite appropriate actually since the OMS pod engines were derived from the Apollo lunar lander engines. By the Astronautix page on the OMS pods[3], they are each about 10 mT propellant mass and 1.8 mT dry mass. Then using its 316s Isp, one of them would suffice for the ca. 2,740 m/s delta-V to go from lunar surface to LEO even with a 4 mT crewed and supplied mass for the SEV with plenty of margin: 316*9.81ln(1 + 10/(1.8 + 4)) =  3,100 m/s.
 The first version of the SLS, called Block 1, is expected to launch by 2017. I would expect a test lunar lander mission, especially if using all cryogenic in-space propulsion, to be done first before a crewed mission is sent. But certainly by 2019, the 50th anniversary of Apollo 11, a crewed mission could be sent. This is in contrast to a post-2030 proposed time frame for a crewed lunar landing using the full 130 mT version of the SLS when it first becomes available.
 There is the cost issue of mounting a manned lander mission. Oddly, the high cost of the SLS might be helpful in this regard. The cryogenic Centaur-like upper stages are already available at a cost in the range of $30 million [4], so the modifications there would be comparatively low cost, compared to the already high cost of the SLS. As for the development cost of the SEV, I suggest use of NASA's commercial crew program's financing procedures. SpaceX was able to develop the Dragon as largely privately financed for reportedly $300 million. And Boeing is paying much of the cost of the development of the CST-100 capsule. It is highly dubious they would be spending a billion dollars of their own money for its development. Then likely its total development cost is in the few hundred million dollar range. Therefore it is likely the development cost of the smaller SEV under commercial crew procedures would also be in the few hundred million dollars range, again comparatively low cost compared to the SLS.
 As I discussed in the blog post "SpaceX Dragon spacecraft for low cost trips to the Moon", SpaceX will also be able to mount a manned lunar landing mission using the 53 mT Falcon Heavy by following, it turns out, the ELA architecture. This will be much cheaper than using the SLS launcher, perhaps only in the few hundred million dollars range cost. But you would have to get private financing for that, since NASA would not fund it as it would undercut NASA's own program.
 In contrast, NASA using the SLS in such an early time frame for a manned return to the Moon would provide further support for continuing the SLS funding. No longer would the SLS be referred to as "a rocket to nowhere".


  Bob Clark

Update, Sept. 28, 2013:

 Finally, NASA has acknowledged that the Block 1, first version of the SLS to launch in 2017 will have a 90+ mT payload capacity not the 70 mT always stated by NASA:

SLS Dual Use Upper Stage (DUUS).
http://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20130013953_2013013757.pdf

 This is important since it means we will have the capability to do manned lunar landing missions by the 2017 first launch of the SLS:

SLS for Return to the Moon by the 50th Anniversary of Apollo 11, page 5: A 90+ metric ton first launch of the SLS.
http://exoscientist.blogspot.com/2013/09/sls-for-return-to-moon-by-50th.html

REFERENCES.

1.)Lunar Base Studies in the 1990s. 
1993:  Early Lunar Access (ELA). 
by Marcus Lindroos 
http://www.nss.org/settlement/moon/ELA.html 
(Note a typo on this page: the payload adapter mass should 
be 2,000 kg instead of 6,000 kg.) 

2.)Delta-V for "direct descent" to the lunar surface?
SATURDAY, SEPTEMBER 15, 2012
http://exoscientist.blogspot.com/2012/09/delta-v-for-direct-descent-to-lunar.html 

3.)Encyclopedia Astronautica.
Shuttle Orbiter OMS.
http://www.astronautix.com/stages/shueroms.htm

4.)Encyclopedia Astronautica.
Centaur IIA.
http://www.astronautix.com/craft/cenuriia.htm

SpaceX Dragon spacecraft for low cost trips to the Moon, page 2: Comparison to 'Early Lunar Access'.

 Copyright 2012 Robert Clark

 Early Lunar Access lander stage.

 The Early Lunar Access [1],[2], proposal of General Dynamics came as quite a surprise to those in the industry when it was first proposed in the early 90's. It suggested manned lunar missions at half the mass needed to LEO and at 1/10th the cost of the Apollo missions.

 It was based on using existing launchers with small upgrades to keep costs low. The only part of it that was technically doubtful at the time was that you could get the lightweight 2-man capsule they were proposing at only a ca. 3.7 mT crewed mass.

 Based on such a small sized capsule, they were able to get a manned mission to the Moon at only about 52 mT required to LEO using all cryogenic space stages. However, the 7-man Dragon capsule at a ca. 4mT dry mass suggests this is indeed feasible.

 It is also interesting the architecture they were proposing for low costs was similar to what I suggested for the SpaceX Dragon via the Falcon Heavy launcher. It would use a single capsule to take the crew all the way from LEO to the Moon's surface and back again, i.e.,no separate lunar crew module. Also it would use as I suggested a single lander stage to take the crew capsule from low lunar orbit to the Moon's surface and then all the way back to LEO, rather than linking up with a return stage waiting in lunar orbit for the return.

 This gives further confidence in the feasibility of the lunar lander plan using the Dragon with Centaur-style stages launched on the 53 mT Falcon Heavy.

  Bob Clark

1.)Encyclopedia Astronautica
Early Lunar Access.

https://web.archive.org/web/20160315095322/http://www.astronautix.com/craft/earccess.htm

2.)Lunar Base Studies in the 1990s.
1993:  Early Lunar Access (ELA).
by Marcus Lindroos

https://web.archive.org/web/20170204225428/http://www.nss.org/settlement/moon/ELA.html

(a typo on this page: the payload adapter mass should

be 2,000 kg instead of 6,000.)