Implications of the coming era of commercial heavy launch: point-to-point transport for both cargo and passengers.

 Copyright 2025 Robert Clark

The new era of heavy launch.  
By Gary Oleson  
The Space Review  
July 24, 2023  
https://www.thespacereview.com/article/4626/1

 The author Gary Oleson discusses the implications of SpaceX achieving their goal of cutting the costs to orbit to the $100 per kilo range. His key point was costs to orbit in the $100 per kilo range will be transformative not just for spaceflight but because of what capabilities it will unlock, actually transformative for society as a whole.  

 For instance, arguments against space solar power note how expensive it is transporting large mass to orbit. But at $100/kg launch rates, gigawatt scale space solar plants could be launched for less than a billion dollars. This is notable because gigawatt scale nuclear power plants cost multiple billions of dollars. Space solar power plants would literally be cheaper than nuclear power plants.  

 Oleson makes other key points in his article. For instance:    

The Starship cost per kilogram is so low that it is likely to enable large-scale expansion of industries in space. For perspective, compare the cost of Starship launches to shipping with FedEx. If most of Starship’s huge capacity was used, costs to orbit that start around $200 per kilogram might trend toward $100 per kilogram and below. A recent price for shipping a 10-kilogram package from Washington, DC, to Sydney, Australia, was $69 per kilogram. The price for a 100-kilogram package was $122 per kilogram. It’s hard to imagine the impact of shipping to LEO for FedEx prices.

 Sending a package via orbit transpacific flight would not only take less than an hour compared to a full day via aircraft, it would actually be *cheaper*.  

 Note this also applies to passenger flights: anywhere in the world at less than an hour, compared to a full day travel time for the longer transpacific flights, and at lower cost for those longer transpacific flights.   

Oleson Concludes:   

What could you do with 150 metric tons in LEO for $10 million?
The new heavy launchers will relax mass, volume, and launch cost as constraints for many projects. Everyone who is concerned with future space projects should begin asking what will be possible. Given the time it will take to develop projects large enough to take advantage of the new capabilities, there could be huge first mover advantages. If you don’t seize the opportunity, your competitors or adversaries might. Space launch at FedEx prices will change the world.

 These are the implications of SpaceX succeeding at this goal. However, a surprising fact is SpaceX already has this *capability* now! They only need to implement it:  

SpaceX routine orbital passenger flights imminent.  
http://exoscientist.blogspot.com/2024/11/spacex-routine-orbital-passenger.html

 

  Bob Clark

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.

The SpaceX Raptor engine is still of unproven reliability.

 Copyright 2025 Robert Clark

 The explosion of the upper stage Starship during the IFT-7 test flight came as a surprise since SpaceX has promoted the idea the Starship is close to being operational to carry passengers.

 I had previously written in 2023 the Raptor engine was insufficiently reliable at least in regards to an engine intended for manned rockets:

SpaceX should withdraw its application for the Starship as an Artemis lunar lander, Page 2: The Raptor is an unreliable engine.

https://exoscientist.blogspot.com/2023/12/spacex-should-withdraw-its-application.html

 That conclusion has to be said still holds. Multiple lines of evidence lead to the conclusion SpaceX has not been completely forthright in regards to the Raptor reliability. 

 SpaceX has been disingenuous in regards to the Raptor reliability from the beginning of the Starship testing. In describing static fires of the Raptor, SpaceX referred to short 5 to 7 second burns as “full duration”. But in the industry the term “full duration” is understood to be short for “full mission duration.” It refers to static fires that last for the full length and full power level of an actual mission. They are meant to give confidence to the rocket company, and importantly also to potential customers, the engine can indeed fulfill the mission requirements needed during flight.

 Defenders of this use of the terminology have argued SpaceX is using it to mean “full planned duration”. But in the industry, if a rocket engine manufacturer wants to do a test for a shorter length they just call them tests of that shorter length. There is no logical reason for using a term well accepted in the industry with the meaning changed. The only reason that comes to mind is that SpaceX wanted to provide an unwarranted assessment to the Raptor reliability.

  The unreliability of the Raptor engine was seen in prior tests of the Starship landing procedures:

  As this video shows, the leaks and fires are seen quite commonly during restarts, though they do occur during the initial burns also. This points out another area where SpaceX has not been fully forthright about the Raptor reliability. For the SpaceX plan using multiple refuelings for their Moon and Mars flights it is absolutely essential the Raptor be reliable for 3-burns during a single flight, the initial burn, the boostback or reentry burn, and finally the landing burns.

 But astonishingly SpaceX has not done a single static test of the Raptor able to do all 3 burns for the full mission lengths, full mission wait times between burns, and full mission power levels.

 SpaceX has done a static test showing a quite large number of restarts in succession:

Adam Cuker  @AdamCuker

Guess that was only an appetizer. We end up having another test with 34 Raptor firings back-to back

https://x.com/AdamCuker/status/1849176567785967989

 This was offered as evidence of the Raptor able to do the needed burns for reusability. But actually it does the reverse. The Raptor will never have to do this number of burns in quick succession for a real flight. In contrast, the Raptor will have to do the cited 3-burns for both stages for every single flight. Why test an engine usage that will never happen in place of one that will always happen?

 The only apparent answer is SpaceX has no confidence in the Raptor to do the necessary burns for the needed burn times, wait times, and power levels.

 Several Raptors also exploded or otherwise failed on the first Superheavy/Starship test flight IFT-1. SpaceX has argued the Raptor reliability has improved with the Raptor 2. But a key failure shows the Raptor 2 still is lacking in reliability. Indeed this failure provides further support for the contention SpaceX has not been completely forthright on the Raptor reliability. This was the failure on IFT-4 during the booster landing burn.  

 

 During this booster landing burn a Raptor 2 actually exploded. SpaceX still has not “come clean” on this fact. By not acknowledging this explosive failure they are giving an inaccurate assessment of the Raptor reliability.

 There are other important implications of this failure however. SpaceX had previously told the FAA the Superheavy booster was expected after ocean touchdown to tip over and float. 

Starship/Super Heavy Vehicle Ocean Landings and Launch Pad Detonation Suppression System, p.5

https://www.faa.gov/sites/faa.gov/files/20230414_Starship_ReEvaluationEA.pdf

  It should have floated like the Falcon 9 did after a soft ocean landing:

 But in point of fact the Superheavy booster actually exploded. It appears likely the Raptor explosion during the landing burn compromised the vehicle integrity causing it to explode after ocean touchdown.

 But this has important consequences for the other booster landings over ocean and land. In IFT-6 the booster was waved off from the booster catch and did an ocean touchdown. Elon said it was likely that it would explode after ocean touchdown, which it did. But what about the SpaceX claim to the FAA that after soft ocean touchdown and tip over the booster would survive and float?

 Note that in this landing burn of the booster in IFT-6, flames were also seen shooting out the side of the booster. Even though there was no apparent engine explosion during this landing burn, as happened in IFT-4, that it also exploded after ocean touchdown suggests in this case also the booster was damaged. Then rather than the flames shooting up the side of the booster being an intentional venting it may be indicative of fires occurring in the engine bay as had been seen previously, thus compromising the vehicle integrity.

 In IFT-5 however, the booster was able to successfully complete the tower catch, despite the flames shooting up the side. In IFT-7 as well the tower catch was successful despite the flames also seen shooting up the side:

  We may hypothesize that it is the forces of the booster toppling over and impacting the ocean that cause the explosions that do not obtain during the tower catch, even though fires inside the engine bay occur in both scenarios.

 There is further evidence to suggest that fires occurring inside the engine bay are the underlying cause of the flames seen shooting up the sides of the booster during the landing burns.

 

 After the Starship explosion in IFT-7 Elon suggested the flames inside the Starship may have caused pressure build up that released flames, in this case small near the rear flap hinge:

 Elon suggested the position of the fire based on the position of this hinge:

Elon Musk @elonmusk

Preliminary indication is that we had an oxygen/fuel leak in the cavity above the ship engine firewall that was large enough to build pressure in excess of the vent capacity. 

Apart from obviously double-checking for leaks, we will add fire suppression to that volume and probably increase vent area. Nothing so far suggests pushing next launch past next month.

8:14 PM · Jan 16, 2025

 However, it should be noted because of their large size the Raptor Vacuum engines powerheads are also close to this area, so the leak could still have originated from them. You see in the image below the middle sea level engines’ powerheads are surrounded by fire shielding, and are below an apparent firewall. But the longer vacuum engines powerheads extend above this firewall.

 

 In that statement by Elon, it is also notable that Elons says the oxygen/fuel leak and resulting fire was in excess of the vent to handle. Applying that logic also to the Superheavy booster, the flames seen shooting out the side during the booster landing burns may have been due to propellant leaks and fires that were within the capacity of the larger vents on the booster to handle. 

 The second part of Elon’s statement also suggests this. Elon makes SpaceX sound sanguine about the leaks and flames within the rocket as long as they can be controlled.

 But whether they are controlled or not does not contradict the fact the flames seen shooting up the side of the booster during the landing burns are due to leaks and fires within the rocket.

 An earlier statement by Elon also suggests SpaceX had accepted the leaks, and resulting fires, by the Raptors and just sought to contain them:

Elon Musk @elonmusk

We could build a lot more, but the next version of Raptor is really the one to scale up production. We begin testing it in McGregor within a week or so. 

Regenerative cooling and secondary flow paths have been made integral to the whole engine, thus no heat shield is required. Nothing quite like this has ever been done before.

Taking away the engine heat shields also removes the need for 10+ tons of fire suppression behind the engine heat shield, as any gas leaks simply enter the already super-heated plasma surrounding the engines, rendering the leaks irrelevant.

Raptor 3 also has higher thrust and Isp.

9:38 AM · Jun 23, 2024 · 404.1K Views

https://x.com/elonmusk/status/1804871620114214978

 

 SpaceX managed to convince the FAA not to require mishap investigations when an engine didn’t start or or fire the expected length of time, as long as the public was not endangered. This was a mistake because it allowed even IFT-4 not to require a mishap investigation. But this meant SpaceX didn’t have to admit an Raptor exploded during the booster landing burn on this flight. This had the effect of giving a misleading understanding of the reliability of the Raptor.

 However, with Starship exploding in IFT-7 with a chance conceivably of the public having been endangered, it must be noted the possibility it was an engine explosion can not be ruled out, along with the possibility it was a plumbing explosion. Then the Raptors tendency to leak and catch fire must be given more serious review. 

 For these reasons, the FAA should require SpaceX to release any and all videos of the engine bays of both stages while the engines are firing, most specifically during restarts.

Currently existing fire equipment could solve the California wildfires within days.

 Copyright 2025 Robert Clark

 Several fire fighting equipment manufactures have mobile, high power water pumps:

 Such pumps can send water kilometer distances: 

23 Jul 2018

New firefighting water cannon said to produce up to 81000 liters/minute.

“Extremely long hose lines.

With some pressure stabilization along the way, the pumps should have the capacity to enable crews to reach fires up to 3 kilometers / 1,7 miles away from the water source.”

https://www.ctif.org/news/new-firefighting-water-cannon-said-produce-81000-litersminute

 Then the idea be would be to transport the high power mobile pumps to a water source, ocean, lake, pond, river, stream, and then use multiple fire hoses connected end-to-end to transport the water to the site of the fire.

 Placing multiple fire trucks along the path from the water source to the fire you can extend the distance even longer than the 3 km mentioned in that article. Fire trucks typically carry fire hoses of lengths totaling 2,000 feet or more and their onboard pumps can further extend the distance which the high water pressure can be presented.

 Note that many people have questioned why the nearby ocean water is not being used to fight the wildfires. An issue is that saltwater can be damaging to forest land. That is why I want to consider also other, freshwater sources. In looking up other wildfires I noticed it is common that water sources do exist within kilometers of wildfires. For the Pacific Palisades wildfire the Topanga Creek and the Santa Ynez Reservoir lie within kilometers of the wildfire, as can be confirmed by Google Maps.

 Unfortunately, in the last few days it’s been reported that the Santa Ynez Reservoir has been out of operation and empty for nearly a year. This has raised quite a bit of consternation since it should have had quite a large amount of water and could have helped resupply the empty fire hydrants in the region. 

 Still, there is the Topanga Creek. But as its name implies it is rather shallow. A large, high volume pump may not be able to draw from it. However, actually it may be using smaller pumps would be of necessity anyway. 

 The Topanga Creek and the Palisades wildfire are in forested areas.  They would not be accessible to fire trucks or large trucks to transport high volume pumps. You would need smaller pumps for the purpose. For transporting them and the needed fire hoses, there are ATV’s specialized for firefighting:

https://www.kimtekresearch.com/category/utv-vehicle-guide/

 Using several of these could accomplish the same thing as one large, high volume pump:

 You might also be able use small “Bobcat”-style bulldozers for clearing a paths for the ATV’s and fire hoses.

 Because of the regulations and red-tape that hinder new government projects taking place in a timely fashion I would advise the companies that offer these mobile pumps to offer to implement the proposal gratis, just asking to get the go ahead. If the project succeeds they would get world-wide acclaim and contract offers to implement it to fight other wildfires.

 

SpaceX routine orbital passenger flights imminent.

 Copyright 2024 Robert Clark

 An approximate $100 per kilo cost has been taken as a cost of space access that will open up the space frontier. For then instead of a price for a private citizen to go to space instead of millions of dollars, it could be priced at a few tens of thousands of dollars. This is in the range of the price of a first-class roundtrip ticket from Los Angeles to Australia. SpaceX now has the capability to offer such a launcher at such a low per kilo rate.

Robert Zubrin has said in an interview that Elon Musk informed him he believes he can build the Starship, i.e., the upper stage of the Superheavy/Starship launcher, for $10 million:

SpaceWatch.Global on X: "“We don’t go to Mars to desert the Earth. We go to Mars to expand the capacity of the human race, to create new branches of human civilization.” - Dr. Robert Zubrin In this insightful Space Café Podcast episode, Dr. @robert_zubrin dives into the real challenges of building https://t.co/A41iFzClpY" / X

 But Zubrin notes what SpaceX is aiming for is reusability. Say a new Starship might cost $10 million to build with a purchase price of $20 million to the customer. Then allowing conservatively 10 reuses SpaceX might only charge $2 million per use. However, he does not mention it here but he is implying the use of this as a launcher independent of the first stage Superheavy. Then you would need a smaller upper stage, a mini-Starship. 

 Zubrin has discussed use of a mini-Starship, but in the context of a 3rd stage for the Superheavy/Starship. Presumably here though, he is suggesting a smaller launch system consisting of the Starship now as a first stage and a mini-Starship as an upper stage.

 An upper stage is commonly 1/4th to 1/3rd the size of the previous stage. So call the cost of the mini-Starship new of, say, $3 million. However, in regards to reusability, SpaceX has found that difficult to implement for an upper stage, particularly in regards to the thermal protection system. Afterthe last test flight IFT-6 for example, Elon Musk has suggested they might have to change to a completely different kind of TPS than the ceramic tiles now used. In contrast though, SpaceX has ampy demonstated with the Falcon 9 booster the first stage is much easier to reuse.  So I'll estimate a cost here of a partially reusable Starship/mini-Starship as $5 million.

 We'll calculate here that this smaller Starship/mini-Starship launcher will still be a a Saturn V-class expendable launcher at 100+ ton payload capacity to LEO. It comes from this Elon Musk estimate of the dry mass of the Starship as an expendable:

Elon Musk

@elonmusk

·

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Probably no fairing either & just 3 Raptor Vacuum engines. Mass ratio of ~30 (1200 tons full, 40 tons empty) with Isp of 380. Then drop a few dozen modified Starlink satellites from empty engine bays with ~1600 Isp, MR 2. Spread out, see what’s there. Not impossible.

9:14 PM · Mar 29, 2019

 But that is for an upper stage use where it did not have enough engines for liftoff from ground. Assume for 1st stage use it needs 9 engines. Increase the dry mass now to 50 tons for the greater engine mass. 

 For the mini-Starship, an upper stage commonly is 1/3rd to 1/4th the size of the lower stage, so call it 420 tons propellant mass. As an upper stage it doesn’t need high engine thrust so assume same mass ratio of ~30 to 1 as for Elon’s expendable Starship estimate, giving it a dry mass of 14 tons.

Take Starship exhaust velocity as ground launched as comparable to that of the Superheavy, 3,500 m/s. And take the upper stage’s vacuum exhaust velocity as 3,800 m/s. Then we could get ~120 tons to LEO: 

3,500Ln(1 + 1,200/(50 + 434 + 120)) + 3,800Ln(1 + 420/(14 + 120)) = 9,200 m/s, sufficient for launch to LEO.

 This a price of only $5 million launch cost for a Saturn V-class launcher as partially reusable. If we make the landing downrange we lose only ~20% off the expendable payload judging by the Falcon 9 example, so still ~100 tons to LEO as partially reusable. That amounts to radical reduction in launch cost down to only ~$50 per kilo. This means that rather than a price to orbit for a passenger being millions of dollars as it is now, it could be in the tens of thousands of dollars range. As noted by Zubrin in that SpaceWatch.Global interview this is in the range of a first-class round trip tocket from Los Angeles to Australia.

 Those price estimates though are based on a $10 million cost of the Starship. But that undoubtedly is for high production rates. 

 So we'll use an estimate based on the current production cost for Superheavy/Starship at about $90 million, with about 30%, $27 million, for Starship:

STARSHIP COST ANALYSIS
OVERVIEW

Note: This is Payload's current estimate and not based on access to any internal Space data or proprietary information.

Current Estimated Starship & Booster Full Stack Cost (S in thousands)

39 Raptor Engines                                   39,000

Labor.                                                       35,000

Structure, plumbing, tiles, parts               13,000

Avionics                                                     3,000

Total                                                         90,000

*Payload costs estimates are based on a post-R&D 1-2 year forward-looking model. This is an educated best estimate and not based on Space internal data. Further cost reductions are expected in the long-run. $90M cost: Payload estimates it costs $90M to manufacture a fully integrated Starship based on a post-R&D/test production phase near-term model. The go-forward cost does not factor in the near $5B SpaceX has spent on R&D to date.

~70% of costs accrue to Super Heavy and ~30% to Starship upper stage.

Future Starship (upper stage) cost reductions: As Starfactory comes online and Raptor production is refined, Space aims to reduce costs even further. A focus on Starship's upper stage: When SpaceX achieves full reusability, production of Starship second stage vehicles will be an order of magnitude higher than booster production.

• The company plans to eventually build multiple second stage Starships per week and reduce

Raptor engine's production cost to $250K a pop. If successful, the long-term cost to mass produce second-stage Starships could drop to $10M to $15M a vehicle. However, for purposes of this report, we will analyze costs as they are today.

Raptor 2 engines ($39M) Payload estimates each Raptor 2 engine costs ~$1M to build. The 39 engines-which include three additional upper-stage engines that will be added in the future-are by far the biggest Starship cost, adding $39M to total cost. SIM per Raptor 2 engine is half as expensive as its $2M+ Raptor 1 predecessor. 20 SpaceX hopes to eventually bring the cost per engine down to ~$250K.

Payload Research
18. Elon Musk on X 19. Space 20.Elon Musk on X 21. Elon Musk on X

https://docsend.com/view/fi9wuazzeex57iig

 Say, for a mini-Starship upper stage its cost would be a 3rd of the $27 million current production cost of the Starship as new, so $9 million. So the full vehicle production cost at $36 million as new. So a Saturn V-class launcher capable of 100+ tons to LEO at ca. $36 million price new. Note this is about half that of the price of the Falcon 9 but at 5 times the payload capability. This is a cut in price per kilo by a factor of 10 down to $300 per kilo from $3,000 per kilo.

 Note again though with SpaceX amply demonstrating practicality of first stage reusability we can do better than this still. Say the Starship now as first stage could be reused 10 times cutting its cost to, say, $2.7 million per launch, for the total partial reuse cost of $11.7 million per launch. But this is the price to SpaceX. Double this for a price to the customer of ~$23 million as partially reusable. As before landing downrange for the booster would still allow ~100 ton payload to LEO, for a price per kilo of $230 per kilo.

 This may still allow passenger tickets to orbit at say the hundreds of thouasands range depending on how many passengers could be carried in a passenger cabin. Note this is price range charged by Virgin Galactic and Blue Origin on New Shepard just going to suborbital space. Most importantly, large numbers of launches carrying private passengers to orbit of wealthier customers at the price point doable now will increase the production numbers for the Starship thus enabling the lower price point to be reached. 

Launch costs for manned Moon or Mars flights at only ~$20 million per launch.
 This is a Saturn V-class vehicle capable of single launch Mars or Moon missions we could launch now. No thermal tile problems, or needing to master orbital refueling, or stretching tanks, or increasing Raptor thrust. It literally could have been launched on the last few test launches and can literally be launched on the next test launch, providing a proof-of-principle for manned flights to the Moon or Mars. Note this is less than the cost now for sending astronauts to the ISS.

 I argue that this is better than the currently planned SpaceX/NASA approach. For instance the fully reusable Superheavy/Starship V2 will have payload at 100+ tons, and still need all of orbit capable TPS, orbital refueling, tank stretch and upgraded raptors. And it will need all of these to get a ca. $10 million reusable launch cost. But the Starship/mini-Starship will reach the same payload capability, at only be 1/3rd the size of the Superheavy/Starship, without the difficult technical advances.

The current approach to the Moon is not just worse, it's multiply times worse. SpaceX wants a multiple refueling, multiple launch approach of the full Superheavy/Starship for lunar missions. This will be in the range of 18 total flights with refuelings, the orbital depot, and the Starship HLS itself.

 In contrast the Starship/mini-Starship can do it in a single launch. And the current plan needing ~18 launches will actually be 18*4 = 72 times bigger than using a single Starship. Taking into account the size also of the mini-Starship, the current multiple Superheavy/Starship lunar plan will be about 50 times the size of just the single Starship/mini-Starship.

 Using an existing Falcon 9 upper stage or Centaur V as the 3rd stage/lander this is a capability we have now to do single launch Moon or Mars missions. 

 Keep in mind the ~$20 million price might be the customer price for the launch-to-LEO vehicle of the Moon or Mars flight. But SpaceX's own cost would be ~$10 million per launch. This is a capability SpaceX has now. 

 A comparison of the relatively sizes of the two approaches for getting to the Moon:

 Compared to: