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.

Could Blue Origin offer it’s own rocket to the Moon?

 Copyright 2025 Robert Clark

 There is increasing concern within NASA that China could beat us back to the Moon. A big component of this concern derives from increasing delay in the development of the SpaceX Starship, tabbed to be the lunar lander for the Artemis lunar program.

 Surprisingly, it may turn out that Blue Origin’s New Glenn with some relatively small upgrades can operate as its own independent Moon rocket.

 One estimate of the Blue Origins first stage propellant mass has been in the range of 1,150 tons:

First Stage:
Fuel load: 1150 tonnes
How? BE-4 with 2,440kN of thrust, and an ISP ~310 should have a mass flow rate of ~803kg/s. We know from the payload users guide that the engines burn for 200 seconds. 200s x 7eng x ~803kg = 1,124,200kg. This number should have a pretty high fidelity, being off on ISP by 3 only changes the final number by ~10 tonnes. Subtract out some for a likely throttle down during Max-Q and during the end of the burn to limit Gs under 4 then add back in the landing fuel and you likely arrive at ~1,150 tonnes. 

Empty mass: ~100 tonnes
How? Falcon 9's first stage weighs about 27 tonnes vs. ~409 tonnes of fuel. Falcon 9 also has fuel that is 30% more dense, doesn't have large strakes, and the TWR for Merlin is probably about twice as high as BE-4.

https://forum.nasaspaceflight.com/index.php?topic=41146.msg2120895#msg2120895

 This seems pretty robust, based on propellant burn rates and published length of the first stage burn time as indicated there. However, it could be this doesn’t take into account the propellant that needs to be kept on reserve for the reentry and landing burns.

 On the other hand, this estimates it as smaller by a factor of 2.6 than that of the SpaceX Superheavy booster:

{Credit: Ken Kirtland}

https://x.com/kenkirtland17/status/1761481624548511916?s=61

Apparently taken from this graphic available in the New Glenn Users Guide:

 The SuperHeavy has a capacity of 3,600 tons, this would put the New Glenn propellant capacity at 1,380 tons.

 Estimate its’s dry mass as in that forum.NasaSpaceflight.com post as ca. 100 tons. For the hydrolox upper stage, based on it’s size estimate propellant load as ca. 200 tons, and assume a Centaur-like 10 to 1 mass ratio, giving it a dry mass of ca. 20 tons.

  Then estimate the payload to LEO using the rocket equation:

340*9.81Ln(1 + 1,380/(100 + 220 + 100)) + 465*9.81Ln(1 + 200/(20 +100)) = 9,300 m/s, sufficient for 100 tons to LEO. 

 But the point of the matter is 100 tons to LEO has long been seen as the needed payload capacity for a rocket to serve as a single launch manned rocket to the Moon.

 However, this will need greater thrust than the 1,750 tons cited by Blue Origin to lift off with adequate thrust/weight ratio(TWR). Indeed, the gross mass in this estimate is already above the 1,750 ton sea level thrust quoted by New Glenn.

 During the first test launch of New Glenn it was much commented on how slow was the acceleration at lift-off. Even at that lower estimated first stage propellant load of 1,150 tons the lift-off TWR would still be quite low. It seems likely this low TWR is the cause of the payload being initially only 25 tons rather than the previously announced 45 tons.

 Given the low liftoff TWR and the reduced payload, Blue Origin probably intended from the beginning to upgrade the New Glenn thrust.

This graphic shows a thrust level of 4.51 million pounds for a three stage New Glenn, or 2,050 tons compared to the current 1,750 tons, a 17% upgrade.

SpaceX has upgraded the thrust of the Raptor more than once at higher than this level of increase so this is a feasible upgrade. 

 Additionally, Blue Origin in an employment announcement has mentioned an increase in the number of engines from 7 to 9:

 SpaceX has also shown increasing or decreasing number of engines on a stage is a relatively straightforward modification. Actually, given the competitiveness between New Glenn and SpaceX it may be that Blue Origin originally decided to go with 7 engines rather than 9 just so as not to be seen as copying SpaceX. It does seem mysterious why Blue Origin would field a rocket with such a low TWR from the beginning.

 So we’ll assume an upgrade of thrust level of the BE-4 by 17% to bring the 7 engine thrust to 2,050 tons and also an increase in the number of engines to 9, to bring the total thrust now to 2,635 tons. This results in a quite healthy liftoff TWR of 2,635/1,800 = 1.46.

 After their success in their first launch of reaching orbit with New Glenn, Blue Origin plans to top that with a launch of a lunar cargo lander in March with the Blue Moon Mk 1. This would be quite remarkable to advance so rapidly from an initial orbital launch to follow that in the next launch to a landing with a rather sizable 21 ton lander on the Moon. This larger than the Apollo lunar lander.

Note, a 100 ton LEO capacity of the upgraded New Glenn will allow a ca. 50 payload to trans-lunar injection(TLI). This is comparable to that of the Saturn V. 

 In a follow-up post we’ll show this New Glenn with an additional third stage and lander comparable to the Blue Moon Mk 1 can form a manned rocket to the Moon.

Half-size Ariane core stage for a reusable launcher.

Copyright 2018 Robert Clark

 Long-time space advocates will recall back in the late 90's there was a push for large numbers of communication satellites for the purposes of cell-phone communication. This led to the creation of several private launch companies then to serve what was expected to be hundreds to thousands of required launches.

 However, it turned out the great majority of cell-phone communications could be served by terrestrial cell towers. The large satellite constellation plans were then abandoned, and those private launch companies then collapsed.

 But now once again there are renewed plans for satellite megaconstellations containing hundreds to thousands of satellites, such as OneWeb or SpaceX's StarLink. This time it is primarily for high speed internet service. This time there is billion dollar backing for the projects and there have been preliminary launches to test the idea.

 It's very likely now that the projects will take place. For space advocates, an important result of the large numbers of launches required is that it provides a clear advantage for low cost reusable launchers.

 SpaceX always believed reusable launchers could be financially feasible. But other space launch providers were skeptical. They didn't think the number of launches under the current market would pay for reusability.

 But now with the advent of the new megaconstellation plans even previously skeptical Arianespace plans to transition to reusability. See for example the articles here:

https://satelliteobservation.net/2018/05/21/ariane-6-and-beyond/

https://satelliteobservation.net/2018/06/02/cnes-director-of-launchers-talks-reusable-rockets/

 One project Arianspace is planning is called Callisto. It is to be a small sized hydrolox test vehicle to test reusable, vertical landing boosters. It is to be analogous to the SpaceX Grasshopper tests.

 However, unlike the SpaceX Grasshopper that used the original Merlin engines and the same F9 propellant tanks, though perhaps only partially filled, the Callisto plan is to use an entirely newly designed and built stage.

 I see a problem with this. For the money spent on Callisto, it will not be an actual operational vehicle. This mirrors a problem with the X-33 test vehicle that was supposed to test the technologies for an operational SSTO vehicle. But for all the money spent on the X-33, it itself would not have been an operational vehicle.

 I believe this was a mistake. It would have been better if the X-33 itself was to be used as an operational vehicle. It could have been used as a reusable first stage booster to cut costs for a two-stage to orbit system, a la the SpaceX plan:

DARPA's Spaceplane: an X-33 version.

https://exoscientist.blogspot.com/2013/10/darpas-spaceplane-x-33-version.html

 Then my recommendation is not to repeat the mistake of the X-33 program by instead actually using operational stages to test reusability and vertical landing.

 This could be done with two existing Arianespace stages. The Ariane 5 core stage and the Ariane H10-3 cryogenic upper stage. In both cases you would use partially filled tanks, approx. half-filled so that the stage could lift-off on their single engines.

 For the operational versions, you would make the tanks themselves half-size, instead of half-filling a full-sized tank, to save dry mass, at least for the Ariane 5 core. For the Ariane H10-3 for the upper stage use it might be able to carry its full propellant load dependent on the propellant load on the Ariane 5 core to be able to lift off on its single Vulcain engine.

 Another advantage of this approach is that it would finally provide Europe with an independent manned spaceflight capability.

 There is a key problem that would need to be solved. Discussion on on a space forum was that the Vulcain II is not throttleable. The HM7-B engine used on the Ariane H10-3 upper stage is also not throttleable. Then both engines would need to be upgraded to be throttleable. As support for the idea this should be feasible, it should be noted the original versions of the SpaceX Merlin engines prior to the Merlin 1D were not throttleable. SpaceX has also shown with its "hoverslam" approach to vertical landing, it would not have to have a high degree of throttleability. Probably the degree of throttleability common to liquid fuel engines in the range of 60% would be sufficient.

See discussion here:

A half-size Ariane for manned spaceflight.

https://exoscientist.blogspot.com/2014/11/a-half-size-ariane-for-manned.html

 Bob Clark

UPDATE: 6/10/2018

 In the discussion above I forgot a key point. The most important factor in
regards to cost is not the development cost.
The key cost factor is what they would charge per flight for a reusable
launcher. Robert Zubrin made this point insightfully in one of his books. He
recounts that he made the argument for reusable launchers in his former job
with one of the big launch companies.

 He argued that they could cut the cost of launch by an order of magnitude.
The company execs responded: why would we do that? Their view was their
revenue would then be slashed by a factor of ten. They were assuming the
market would still be the same but they would be getting one-tenth the
revenue.

 So the OldSpace companies were acting quite rationally in a business sense
in discounting reusability. They were saying the market was not enough to
make it advantageous to them.

 But if there were a large market then they would make more money making more launchers at the lower price. That is, the price would be reduced by a factor of ten but the number of launches would be increased by more than a factor of ten.

 Also, the importance of the large market and lowered prices for satellite
launches extends beyond that of just the satellite market. By making
launches at such reduced prices, that increases the possible market for
passenger flights to space. So the impending megaconstellation launches may
also bring to fruition the long desired routine passenger flights to space.

About the launch abort system for the New Shepard capsule.

Copyright 2017 Robert Clark

 Blue Origin has revealed the format of its suborbital tourism capsule for the New Shepard suborbital launcher:

Take a Peek Inside Blue Origin’s New Shepard Crew Capsule.
Published: 29 Mar , 2017
by Nancy Atkinson

http://www.universetoday.com/134783/take-peek-inside-blue-origins-new-shepard-crew-capsule/

   The cylinder in the middle is the launch abort motor. It is only supposed to fire in case of an emergency to pull the capsule away from the rocket launcher.

 Normally, it would not even fire. Still its presence inside the passenger cabin is rather disconcerting. Moreover, it is a solid rocket motor. For solid motors, the combustion chamber is the entire rocket, so if a failure, i.e., a breech does occur it can happen anywhere along the motors length.

 A Blue Origins video animation from 2015 shows the solid rocket escape motor with handholds at about the 2:25 point:

 Be careful to mind your head while floating though!

   The reason Blue Origin decided to put the abort motor inside the cabin likely was for reasons of positioning of the center of gravity(CG) with respect to the center of pressure(CP). A well known rocket stability rule of thumb is the center of pressure should be below the center of gravity. 

The trunk and fins helped that for the SpaceX launch abort test by bringing the CP rearward:

 But compare this to the Blue Origin abort test:

  Notice that the capsule is gyrating while the rocket motor is firing. This would be very unpleasant for the passengers since they would be subjected to high g's while being thrown right and left, albeit while strapped in.

 Then for these reasons I suggest giving the New Shepard a trunk with fins as has the SpaceX Dragon capsule.

 This could be done by instead of having the ring structure at the top of the New Shepard stay attached to the New Shepard, let it act as the trunk for the capsule:

 Then you would move the solid rocket abort motor down into this structure, so it is no longer inside the passenger compartment.

 However, this ring structure does have a function as far as the landing of the New Shepard rocket; it holds the fins and the speed brakes used during the landing:

 So how could we maintain those functions if that ring structure is instead attached to the capsule? Two possible approaches you could duplicate it so the New Shepard has its own as does the capsule. 

 Or another possibility would be to have the ring structure only detach along with the capsule only during an abort scenario. For the normal launch, with no abort, the ring structure would stay attached to the New Shepard rocket, carrying also inside the abort motor, while the capsule detaches for the normal flight to suborbital space.

 But if there is a need for an abort, the solid rocket abort motor would fire carrying the ring structure and the capsule away from the New Shepard. In this scenario where there would need to be an abort presumably there would be a failure of the New Shepard anyway and you would not expect to recover it.

  Bob Clark

New Shepard as a booster for an orbital launcher.

Copyright 2016 Robert Clark

 Blue Origin scored another first by successfully relaunching their vertical landing New Shepard suborbital rocket:

 In the blog post "Triple Cored New Shepard as an orbital vehicle", I suggested using three cores of the New Shepard rocket with a small upper stage could form an orbital launcher. However Jonathan Goff on his blog page SelenianBoondocks raised the possibility a single New Shepard could serve as the first stage booster of an orbital rocket:

Random Thoughts: New Shepard for Pop-Up TSTO NanoSat Launch.
http://selenianboondocks.com/2016/01/random-thoughts-new-shepard-for-pop-up-tsto-nanosat-launch/

  I think it should be doable using a similar small cryogenic upper stage as for the triple-cored case. The stage I suggested there was the cryogenic upper stage of the Ariane 4, the Ariane H10-3, or one developed by Blue Origin similar to it. It had a dry mass of 1,240 kg and a propellant mass of 11,860 kg. The Isp was 445 s with a vacuum thrust of 64.8 kN. However, simply using a nozzle extension as on the RL-10B-2 can give it likewise an Isp of 462 s and vacuum thrust of 110 kN. So we'll use these values.

 To make the estimate of the payload we need the vacuum values for the Isp and thrust of the BE-3 engine. In the "Triple Cored New Shepard as an orbital vehicle" blog post I estimated these to be 360 s and 568.8 kN respectively.

 However, to loft the vehicle with the additional weight of the upper stage we'll need to increase the BE-3 thrust slightly. This should doable. For instance the SSME’s could operate at 109% of their originally rated thrust, and the Merlin 1D had a 15% thrust upgrade. So say the BE-3 vacuum thrust is increased 9% to 620 kN, keeping the same Isp.

 Now use Dr. John Schilling's payload estimator program. For the "Restartable upper stage" option check "No", otherwise the payload will be reduced. Select Cape Canaveral as the launch site and enter 28.5 for the launch inclination in degrees to match the latitude of the launch site. Then the calculator gives the result:

Mission Performance:

Launch Vehicle:	User-Defined Launch Vehicle
Launch Site:	Cape Canaveral / KSC
Destination Orbit:	185 x 185 km, 28 deg
Estimated Payload:	1690 kg
95% Confidence Interval:	1298 - 2153 kg

"Payload" refers to complete payload system weight, including any necessary payload attachment fittings or multiple payload adapters

This is an estimate based on the best publicly-available engineering and performance data, and should not be used for detailed mission planning. Operational constraints may reduce performance or preclude this mission.

 Altitude Compensation to Increase Payload.
  As I discussed in the "Triple Cored New Shepard as an orbital vehicle" blog post, altitude compensation provides a simple, low cost method of improving payload.  For instance by attaching a nozzle extension the vacuum Isp of the BE-3 can be increased to the 462 s range of the RL-10B-2 engine. The vacuum thrust will then be increased proportionally to (462/360)*620 = 796 kN.

 Then the Schilling calculator gives the result:

Mission Performance:

Launch Vehicle:	User-Defined Launch Vehicle
Launch Site:	Cape Canaveral / KSC
Destination Orbit:	185 x 185 km, 28 deg
Estimated Payload:	2324 kg
95% Confidence Interval:	1841 - 2895 kg

"Payload" refers to complete payload system weight, including any necessary payload attachment fittings or multiple payload adapters

This is an estimate based on the best publicly-available engineering and performance data, and should not be used for detailed mission planning. Operational constraints may reduce performance or preclude this mission.



  Bob Clark

UPDATE, Feb. 28, 2016:

 This considered an Ariane hydrolox upper stage as the upper stage for this New Shepard launcher. This would be problematical since it would be a direct competitor to Arianespace's Vega rocket at a much lower cost than the Vega's $35 million.

 Blue Origin very likely could develop a hydrolox upper stage that would be cheaper than the Ariane one. But that would take time and significant development cost. Instead of that, Blue Origin could produce a New Shepard derived launcher for cubesats at minimal extra development cost since the required small upper stages already exist.

 Existing upper stages that could work would be the large Star solid rocket upper stages such as the Star 63F:

Star 63F:
http://www.astronautix.com/engines/star63f.htm

 Using this for the upper stage, Schillings launch performance 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:	293 kg
95% Confidence Interval:	174 - 443 kg

 This is in the range being considered for the cubesat launchers that NASA has already awarded million dollar contracts to:

Firefly, Rocket Lab and Virgin Galactic Win CubeSat Launch Contracts from NASA.
By Caleb Henry | October 15, 2015 | Feature, Government, Launch, North America, Regional, Satellite TODAY News Feed
http://www.satellitetoday.com/launch/2015/10/15/firefly-rocket-lab-and-virgin-galactic-win-cubesat-launch-contracts-from-nasa/

 Considering the quoted prices there, this New Shepard based launcher very likely could beat these prices especially using the reusable New Shepard.

 And since the upper stage already exists, it very likely would also beat to launch these other systems still in development.

 About the quick route to operational status of this orbital rocket, I think it is significant that Blue Origin was able to beat SpaceX on a relaunch of its returned booster. The argument has been made that New Shepard is not an orbital launcher. But if Blue Origin developed this orbital launcher from New Shepard then they would be able to beat SpaceX at reusing a booster for a true orbital launcher as well.

 My opinion is SpaceX will have difficulty with getting their booster to land in reliable fashion as long as it does not have hovering ability. And because the New Shepard does have hovering ability it will be more reliable as a reusable booster.

 BTW, as Blue Origin develops its large high performance dense propellant engines, it will have the same problem as SpaceX it getting its booster to be able to hover, resulting in the same problem of reduced reliability on landing. For this reason I think Blue Origin should investigate methods of giving its large planned boosters hovering ability such as discussed here:

Hovering capability for the reusable Falcon 9, page 3: hovering ability can increase the payload of a RLV.
http://exoscientist.blogspot.com/2015/12/hovering-capability-for-reusable-falcon.html

 Surprisingly, it turns out that hovering ability when properly implemented can actually improve the the payload for a reusable rocket.

Triple Cored New Shepard as an orbital vehicle.

Copyright 2016 Robert Clark

 Blue Origin made a significant achievement in successfully landing their New Shepard rocket after a suborbital spaceflight:

 As their next development Blue Origin intends to make a several million pound thrust rocket capable of sending 25 metric tons to LEO. This would be a very large and expensive development for their first orbital rocket, comparable in size to the largest orbital rockets available now, larger for example than the Falcon 9.

 I suggest an intermediate development for their first orbital rocket. Running the numbers, their New Shepard suborbital rocket could be used to make an orbital rocket using three cores with a smaller upper stage, a la the Delta IV Heavy.

 It would have a payload to LEO in the range of 3,000 kg, about the size of the Arianespace Vega rocket. The Vega costs in the range of $35 million. Considering the small size of the New Shepard, even at three cores, Blue Origin should be able to beat this price.

 Moreover, this version would have the capability to be reusable. SpaceX is planning to make the three cores of the Falcon Heavy reusable by returning the two side cores to the launch site and recovering the central core by a barge landing out at sea. Quite likely this would work for a 3-cored New Shepard launcher as well.

Specifications of the New Shepard BE-3 engine.

 Here's a formula for calculating the sea level thrust from the vacuum thrust and back pressure:

F = q × Ve + (Pe - Pa) × Ae
where F = Thrust
q = Propellant mass flow rate
Ve = Velocity of exhaust gases
Pe = Pressure at nozzle exit
Pa = Ambient pressure
Ae = Area of nozzle exit
http://www.braeunig.us/space/sup1.htm

 Estimating the nozzle exit diameter as 1 meter, the exit plane area would be: π*0.5^2 = .7854. Then the back pressure to be subtracted off would be 101,000Pa*.7854 = 79,325 N. 
Blue Origin has given the sea level thrust as 110,000 lb, 110,000*4.45 = 489,500 N. So the vacuum thrust is 489,500N + 79,325N = 568,825 N. 

 We also need to calculate the Isp. One other piece of information will allow us to calculate this. This Blue Origin page gives the horsepower of the BE-3 as over 1,000,000 hp:

https://www.blueorigin.com/technology

 The power of a jet or rocket engine is (1/2)*(thrust)*(exhaust velocity). The 1,000,000 hp at sea level is 1,000,000*746 = 746,000,000 watts. Then using the formula the exhaust velocity at sea level is 3,048 m/s, and the Isp is 310 s.

 Since (thrust) = (exhaust velocity)*(propellant flow rate), we also get the propellant flow rate as 489,500/3,048 = 160.6 kg/s. Now we can get the exhaust velocity and Isp at vacuum. From the 568,825 N vacuum thrust, we get the vacuum exhaust velocity as 568,825 N/160.6 = 3,540 m/s, and the vacuum Isp as 360 s.

  It is interesting that the diameter and sea level and vacuum Isp's are close to those of the RL-10A5,  the sea level version of the RL-10 used on the DC-X:

http://www.astronautix.com/engines/rl10a5.htm

Size Specifications for the New Shepard.
 The Blue Origin environmental impact statement:

Final Supplemental Environmental Assessment for the Blue Origin West Texas Launch Site.
February 2014
https://www.faa.gov/about/office_org/headquarters_offices/ast/media/Blue_Origin_Supplemental_EA_and_FONSI.pdf

on p. 4 lists the max dry mass as 30,000 pounds (13,600 kg) and max propellant load as 60,000 pounds (27,300 kg). This corresponds to estimates made of the New Shepard gross mass based on its dimensions.

 We need also a small upper stage. The cryogenic upper stage of the Ariane 4 will suit the purpose, the Ariane H10-3. It had a dry mass of 1,240 kg and a propellant mass of 11,860 kg. The Isp was 445 s with a vacuum thrust of 64.8 kN. However, simply using a nozzle extension as on the RL-10B-2 can give it likewise an Isp of 462 s and vacuum thrust of 110 kN. So we'll use these values.

 Use now Dr. John Schilling's Launch Performance Calculator to estimate the payload. We'll also use cross-feed fueling to increase the payload. Cross-feed fueling is not an unknown technology having been used on jet aircraft such as the Concorde for decades and also on the Space Shuttle's OMS engines.

 To emulate cross-feed fueling with the Schilling calculator for two side boosters, enter in 2/3rds of the actual propellant load into the propellant field for the side boosters. And for the central core enter in (1 + 2/3) times the propellant load in the field for the first stage. (See  discussion here for explanation of how the Schilling calculator emulates cross-feed fueling.)

 So in the dry mass fields for the side boosters and first stage enter 13,600 kg. And in the propellant field for the side boosters enter 18,200 kg and 45,500 kg for the first stage. For the second stage enter 11,860 kg for the propellant and 1,240 kg for the dry mass.

 In the thrust fields and Isp fields enter in the vacuum values. So for the side boosters and first stage enter 568.8 for the thrust in kilonewtons and 110 for the second stage. In the Isp fields enter 360 for the side boosters and first stage Isp in seconds and 462 for the second stage. 

 For the "Restartable upper stage" option check "No", otherwise the payload will be reduced. Select Cape Canaveral as the launch site and enter 28.5 for the launch inclination in degrees to match the latitude of the launch site.

 The 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:	3420 kg
95% Confidence Interval:	2766 - 4205 kg

"Payload" refers to complete payload system weight, including any necessary payload attachment fittings or multiple payload adapters

This is an estimate based on the best publicly-available engineering and performance data, and should not be used for detailed mission planning. Operational constraints may reduce performance or preclude this mission.


  This would be using an Arianespace upper stage. But this would be a competitor to their Vega launcher so that is problematical. Blue Origin could use instead the Rl-10B2 engine and their own constructed upper stage. The RL-10 though is a rather expensive engine. Another possibility is the 25,000 lb thrust hydrolox engine being developed by XCOR.

Altitude Compensation Increases Payload Even for Multistage Vehicles.
 It is unfortunate that SSTO's have (incorrectly) been deemed unviable. Since altitude compensation has only been thought of in terms of improving the payload of SSTO's, little research has gone into such methods, with SSTO's not being considered worthwhile.

 However, in point of fact altitude compensation improves the payload even for multistage rockets. As with the RL-10B-2 we can get a vacuum Isp of 462 s on the New Shepard hydrolox engine simply by the addition of a nozzle extension. Other methods of accomplishing it are discussed in the blog post "Altitude compensation attachments for standard rocket engines, and applications."

 Increasing the Isp will also increase the thrust proportionally. So at a 462 s Isp for the BE-3, the thrust becomes 568.8*(462/360) = 730 kN. Entering these values into the thrust and Isp fields for the side boosters and first stage gives the result:

Mission Performance:

Launch Vehicle:	User-Defined Launch Vehicle
Launch Site:	Cape Canaveral / KSC
Destination Orbit:	185 x 185 km, 28 deg
Estimated Payload:	5302 kg
95% Confidence Interval:	4359 - 6438 kg

"Payload" refers to complete payload system weight, including any necessary payload attachment fittings or multiple payload adapters

This is an estimate based on the best publicly-available engineering and performance data, and should not be used for detailed mission planning. Operational constraints may reduce performance or preclude this mission.


 This now is a serious payload capability. Note for example NASA awarded Orbital Sciences with a billion dollar contract to deliver payload to the ISS with their Antares rocket with a 5,000 kg payload to LEO capacity.


 Bob Clark

UPDATE, February, 3, 2016:

 Jonathan Goff on his SelenianBoondocks.com blog raised the possibility that a single New Shepard could serve as a booster for an orbital rocket. I confirmed it could at the 1 to 2 metric ton payload range by using the same type of hydrolox upper stage as discussed above in the triple-cored case:

New Shepard as a booster for an orbital launcher.
http://exoscientist.blogspot.com/2016/01/new-shepard-as-booster-for-orbital.html

 It could also serve as a booster for a smaller launcher by using instead one of the Star solid rocket upper stages, giving a few hundred kilos payload. This would have the advantage that little extra development would be required.

 Plus, it may allow Blue Origin to beat SpaceX at reusing a booster for an orbital launcher.