Polymath

Could Blue Origin develop a lander for Artemis III?

noreply@blogger.com (Robert Clark) — Sat, 15 Nov 2025 23:40:00 +0000

SpaceX is coming under increasing criticism for its delay in developing the Starship lunar lander:

U.S. Is Losing Race to Return to Moon, Critics Say, Pointing at SpaceX.
“The company’s Starship rocket, which has suffered a series of recent test explosions, is still years away from being ready for the mission, former NASA executives say. The SpaceX Starship rocket has exploded during three of its four recent tests, and its current version can carry only a fraction of its promised payload into orbit.”
By Eric Lipton
Reporting from Washington Sept. 20, 2025 https://www.nytimes.com/2025/09/20/us/politics/spacex-us-moon-race.html

Then other possibilities should be explored for a lander for Artemis. Blue Origin is planning on the multi-billion dollar Blue Moon Mk2 manned lander, but only for the later Artemis missions. However, due to the size of the Mk2, this plan would require a separate Cislunar Transporter to carry refueling propellant and 2 to 6 New Glenn launches for refueling the Mk2.

But instead of that, I discuss here using the smaller Blue Moon Mk1 lander for the purpose. Blue Origin already developed the Mk1 under its own dime, and so, one assumes, under much lower development costs.

I’ll argue here for using the Mk1 and which can be done in a single launch of the New Glenn in expendable format. But for this to happen the New Glenn’s capabilities will have to be extended to its reach originally announced payload.

Blue Origin had originally cited New Glenn’s capability as 45 tons to LEO under partial reusability. However, for its first test launch it was listed as only 25 tons. Looking at the surprisingly slow takeoff acceleration of that first test flight this must have been due to lower initial thrust of its BE-4 engines. It could not have been Blue Origin always intended for the takeoff thrust to be that low.

Then Blue Origin either needs to quickly upgrade the engines to reach their originally intended thrust value or increase the number of engines to 9. We have seen with SpaceX and the Starship, adding extra engines is relatively straight-forward and accomplished within months. Quite likely, the only reason why Blue Origin did not want to give New Glenn nine engines in the first place is because they did not want to be seen as copying the Falcon 9. But given the possible implications, getting a manned lander before SpaceX, and possibly even by the planned date for Artemis III, the option should be explored.

Then I’ll assume the New Glenn’s thrust can be raised in short order to have its payload in partial reusable format raised to the 45 tons originally cited. Then that would imply its expendable payload would be in the 60+ tons range.

But this means Blue Moon Mk1 could work as a lander to and from NRHO to the lunar surface, launched on a single expendable New Glenn. No refuelings would be required, unlike what’s required for the Mk2. You might need to use the Delta IV Heavy upper stage, currently used for the Interim Cryogenic Propulsion Stage (ICPS) on the SLS, or a Centaur V upper stage as a 3rd stage to do the TLI burn.

But it’s possible you wouldn’t need that extra stage at all. Considering New Glenn uses an efficient hydrolox upper stage, it may not need an extra stage to get the ca. 25 ton Mk1 to TLI at all.

See the attached image showing in the first row the delta-v’s needed to go one way from LEO to land on the lunar surface, 6.1 km/s. This is the plan for how Blue Moon Mk1 was intended to be used by Blue Origin as a one-way cargo lander. Now, look at the fourth row showing the delta-v’s making stops at NRHO. You see to first rendezvous at NRHO then go back and forth to the lunar surface totals 5.95 km/s. That is to say, the Blue Moon Mk1 has this capability in its current form carrying 3 ton payload.

Note though the highlighted part of the fourth row is not including the delta-v needed to get to TLI. That propulsive burn would have to be supplied by New Glenn or an additional 3rd stage I mentioned.

Then we need a manned capsule for this lander within the 3 ton payload capability of the Mk1. This is certainly doable since the Apollo LEM had a crew module at only 2 ton dry mass. Now, to get a rapid, low cost development, I would advise adapting the Cygnus capsule for the purpose, rather than developing one from scratch. It would only require adding life support and crew seats. It is my thesis using already existing and operational systems saves greatly on development costs in general.

Lightweight thermal protection for reentry of upper stages.

noreply@blogger.com (Robert Clark) — Sat, 19 Apr 2025 23:10:00 +0000

In the blog post “Reentry of orbital stages without thermal protection, Page 2”, http://exoscientist.blogspot.com/2025/04/reentry-of-orbital-stages-without.html , I discussed some possibilities of thermal protection for the SpaceX Starship. Chief among them was the possibility that lightweight wings added might allow the stainless-steel Starship to survive reentry without added thermal protection at all.

Other possible methods of thermal protection discussed there were a “parashield” of Dr. David Akin and inflatable conical shield experimented for the Cygnus capsule return.

The method used there to estimate the temperature reached was calculation of the ballistic coefficient,

β = (mass)/(drag coefficient*area). In a report by aerospace engineer Dr. David Akin, the estimated ballistic coefficient for the max temperature reached to be 800 C, so as not to need additional thermal protection, was ca. 20 kg/sq.m.

However, I calculated the ballistic coefficient for the Starship to be ca. 60 kg/sq.m. Note though this was using a much lower dry mass for the Starship than now obtains. The currently estimated dry mass of the reusable Starship is in the range 160+ tons. I believe this high mass for the reusable Starship is primary reason SpaceX is having difficulty getting effective TPS for it.

My opinion is that SpaceX should first get an expendable Starship and then proceed to reusability. This approach worked spectacularly well for the Falcon 9.

In this regard it is notable Elon Musk once estimated the dry mass of the expendable Starship as only 40 tons:

Then in the following I’ll use the 40 ton value for the Starship dry mass. In this case, there might be an example that would give us a reusable thermal shield for a vehicle the size of Starship. I’m thinking of the X-33/Venturestar.

The length in meters was 38.7m and width 39m. For the dry mass, the total gross weight was 2,186,000 lbs, propellant weight 1,929,000 lbs, and payload weight 45,000 lbs; giving a dry weight of 212,000 lbs, or 96,400 kg.

Using a hypersonic drag coefficient of 2, and considering the triangular planform requires multiplying by 1/2 the length*width to get the area, the ballistic coefficient calculates out to be 96,400/(2*1/2*38.7*39) = 64 kg/sq.m.

Remarkably close to the ballistic coefficient of the Starship at the 60,000 kg mass of the expendable’s dry mass + fairing mass.

But the added weight of the metallic shingle TPS of the X-33/Venturestar can’t be too high to allow the ballistic coefficient to remain close to this value.
The areal density of the metallic shingle TPS was about 10 kg/sq.m:

REUSABLE METALLIC THERMAL PROTECTION SYSTEMS DEVELOPMENT
Max L. Blosser*, Carl J. Martin*, Kamran Daryabeigi*, Carl C. Poteet **
*NASA Langley Research Center, Hampton, VA, USA
** JIAFS, The George Washington University, Hampton, VA, USA
https://ntrs.nasa.gov/api/citations/200 … 095922.pdf

The metallic tiles had better resistance to impact and rain than the ceramics at about the same weight.

Fig.3 Layered metallic sheeting separated by insulation.

Fig.21 Metallic TPS at same weight of ceramic tiles, ~10kg/sq.m.

At a 10 kg/sq.m. areal density, the added weight covering just the lower half of the Starship would be (1/2)*Pi*9*50*(10 kg/sq.m.) = 7,060 kg, proportionally small enough that the ballistic coefficient would still be ca. 60 kg/sq.m.

This would be advantageous in that you don’t need added wings and you don’t need an additional conical thermal shield.

BUT for this to work SpaceX would have to go back to the smaller, expendable mass of the Starship. SpaceX had tested the X-33 metallic shingles and concluded they were inadequate. But that was with temperatures developed with the higher 160+ ton Starship. With a lighter dry mass, much reduced temperatures result.

Thermal Protection for the Falcon 9 Upper Stage.

SpaceX had originally intended to make the Falcon 9 upper stage reusable as well as the first stage but decided it was too difficult and chose to only make the first stage reusable. They also engaged in attempts to recover the separated fairing half’s, but decided not to continue implementing this.

However, the metallic shingles of the X-33/VentureStar may provide a method to recover the upper stage and fairing.

This page gives the F9 upper stage as 12.6m long, 3.66m wide at a dry mass of ~4,000 kg, and the fairing as 13.1m long, 5.2m wide, at ~ 1,750 kg:

Falcon 9 FT (Falcon 9 v1.2)
https://web.archive.org/web/20230710234357/https://spaceflight101.com/spacerockets/falcon-9-ft/

The interstage has been estimated as weighing 1,000kg. Then using again a cylinder’s hypersonic drag coefficient of 2, the ballistic coefficient calculates out to be:

(4,000 + 1,750 + 1,000)/(2*(12.6*3.66 + 13.1*5.2)) = 29.5 kg/sq.m.

This is well less than the desired 60 kg/sq.m point for metallic shingle TPS. But we have to make sure the added weight of the TPS still allows the ballistic coefficient to stay below this point.

The weight of this added metallic shingle TPS would be (1/2)*Pi*(12.6*3.66 + 13.1*5.2)*10 kg/sq.m. = 1,800 kg. Adding this on, the ballistic coefficient would still only be 36 kg/sq.m.

Another possibility though arises from the low ballistic coefficient of 29.5 kg/sq.m from the bare upper stage+fairing without TPS. This is close enough to the 20 kg/sq.m ballistic coefficient point for a stainless steel spacecraft not needing TPS, that should be investigated for the F9 upper stage.

The tankage, fairing, and interstage would have to be replaced by stainless-steel. The tanks are aluminum-lithium. The specialty high-strength stainless-steel as used on Starship saves about 1/3rd the weight off aluminum-lithium tanks. But the fairing and interstage are composite. The stainless-steel alloys are about the same weight as the carbon-composites.

Doing some rough estimates it will be approx. at the 20 kg/sq.m point if the tanks and fairing are converted to stainless-steel but the interstage is jettisoned and just use a lightweight steel plate to block the engine from the high temperature air stream during reentry.

Reentry of orbital stages without thermal protection, Page 2.

noreply@blogger.com (Robert Clark) — Thu, 10 Apr 2025 15:03:00 +0000

SpaceX is having difficulty creating an effective thermal protection system. I think they should reconsider using wings for return. For instance using sufficiently lightweight wings, it may be possible no thermal protection would be needed at all.

This is the thesis put forward here. The idea behind it was an article that suggested with sufficiently low wing loading, weight per wing area, an orbital stage might need no thermal protection at all on reentry:

Wings in space.
by James C. McLane III
Monday, July 11, 2011
Wing loading (the vehicle’s weight divided by its wing surface area) is a prime parameter affecting flight. The antique aluminum Douglas DC-3 airliner had a big wing with a low loading of about 25 psf (pounds per square foot of wing surface). At the other end of the spectrum, the Space Shuttle orbiter has a high wing loading of about 120 psf. This loading, combined with an inefficient delta-shaped wing, makes the orbiter glide like a brick. A little Cessna 152 private plane features a wing loading of about 11 psf and modern gliders operate down around 7 psf. A space plane with huge lifting surfaces and a very low wing loading might not require any external thermal insulation at all. Building a space plane with a wing loading of, say, 10 psf should not be an impossible proposition. Perhaps some day it will be done.
http://www.thespacereview.com/article/1880/1

I discussed the possibility here:

Reentry of orbital stages without thermal protection? UPDATE: 7/1/2019
https://exoscientist.blogspot.com/2019/06/reentry-of-orbital-stages-without.html

(Note: I mistakenly used the half-surface area of a cylindrical stage in those calculations. This underestimates the wing loading, in psi units. So the stages discussed there appeared better than they actually were. In the calculations in this post, I’m using cross-sectional area.)

The proposer of this idea was the legendary spacecraft designer Maxime Faget, who was the chief designer of the Mercury capsule, the first U.S. manned space capsule. Then on that basis alone the possibility should be given serious consideration.

The parameter though used to measure the capability of a particular shape to slow down descent is not wing loading, weight divided by wing area, but the ballistic coefficient, (mass)/(drag coefficient*drag area), β = m/C_DA, given in metric units, where the drag area is by cross-section. This takes into account the fact different shapes are more effective in slowing down the spacecraft by including the coefficient of drag C_D as well as being more general than just looking at wings for the decelerator.

A couple of ways being investigated to get a lightweight decelerator are by using a inflatable and by using a foldable heat shield.

There are several variations of the inflatable heat shield idea, sometimes called a ‘ballute’. The most researched one is a conical inflatable heat shield. It’s being investigated for example as a heat shield to make the Cygnus cargo capsule reusable:

Here’s a research article on it:

HEART FLIGHT TEST OVERVIEW
9th INTERNATIONAL PLANETARY PROBE WORKSHOP 16-22 JUNE 2012, TOULOUSE
https://websites.isae-supaero.fr/IMG/pdf/137-heart-ippw-9_v04-tpsas.pdf

In this report, the mass used for their analysis is ca. 5,000 kg and the diameter of their conical decelerator is 8.3 meters. There is thermal protection applied but I gather less of it is needed since the conical aeroshell is just made of silicone rubber.

To get the low ballistic coefficient you want to minimize the dry mass of the upper stage or capsule being returned. This is a concept understood by spaceflight engineers: extra mass added to an upper stage subtracts directly from payload. So spaceflight engineers commonly try to minimize this dry mass.

I have discussed before I believe it is a mistake for SpaceX to want to go directly to a fully reusable upper stage. Elon Musk once estimated that an expendable Starship upper stage without fairing could be made at only 40 tons dry mass:

Elon Musk @ElonMusk
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.
https://x.com/elonmusk/status/1111798912141017089?s=61

Now, in their attempts at making it fully reusable the dry mass has ballooned to ca. 160 tons or more. That huge dry mass is the primary reason why they are having difficulty finding effective thermal shielding.

Then in the following I’ll assume SpaceX does go first for an expendable Starship upper stage at ca. 40 tons dry mass. Then when they do proceed to reusability of the upper stage, if it is just ballistic coefficient determining the effectiveness of the heat shield, then for a spacecraft or stage about 9 times heavier than the HEART shield for the Cygnus, say, at 40,000+ kg, then the area needs to be 9 times more, that is, a conical shell about 25 meters in diameter.

BUT, the key questions is how does the mass of the decelerator scale with the size of the reentry spacecraft? In this report the added mass of the inflatable shield is a small proportion of the spacecraft being returned, in the range of 25%. But that is for a returned spacecraft of ca. 5,000 kg dry mass, with decelerator mass of ca. 1,300 kg.The report doesn’t discuss how the mass of the decelerator scales with size. You could make an argument it should scale with the cube of the decelerator diameter. The reason is because of not just the area increasing but the shield thickness also increasing to maintain shield strength. Then for a cone shield of 3 times larger diameter the mass would be 3³ = 27 times heavier or 35,000 kg. That is quite larger percentage of the 40,000 kg stage dry mass. It is still much better than the 120+ ton added mass the Starship now has in the attempt to make it reusable.

On the hand, you could make an argument it should scale by the square of the diameter. The reason is you could use multiple copies of the smaller cone shields to cover the entire returning spacecraft. So it would be 3² = 9 times heavier or 9*1,300 = 11,700 kg being added to the dry mass. This would be a more palatable increase, if that is indeed the correct scaling.

This report though doesn’t give the maximum, i.e., stagnation temperature reached so it’s a little difficult to see if steel itself would be able to withstand the heating. It describes using layers of the Nextel thermal blankets so presumably this would also work for the Starship or other stages or capsules with the appropriate size conical shield for the reentry dry mass.

But we can make an estimate of what size wings for the Starship could get similar ballistic coefficient as the inflatable conical shell and therefore existing off-the-shelf Nextel thermal blankets would suffice for the thermal shield.

For the example considered in this report, the dry mass of the returning spacecraft is approx. 5,000 kg and the area on the inflatable is about 56 m². Then the ratio of mass to area is about 100 kg/m². But actually the ballistic coefficient also divides this by CD , the coefficient of drag. At hypersonic speeds the drag coefficient of a 55 degree half-angle cone is about 1.5, so the ballistic coefficient is of about 60 kg/m².

For the expendable Starship at 40 tons, adding on the fairing at ca. 20 tons brings the total mass to ca. 60 tons. At a diameter of 9 meters and length of 50 meters, the cross sectional area is 450 m². To calculate the ballistic coefficient also need the hypersonic drag coefficient. For a cylinder entering broadside that is about 2. Then the ballistic coefficient is 60,000/(2*450) = 66 kg/m².

This is close enough for the Starship itself without wings only using existing Nextel thermal blankets to survive reentry for reuse as long as SpaceX starts with the lightweight value of the expendable version.

Another method for lightweight thermal shielding via low ballistic coefficient would use a foldable heat shield. This is the approach investigated by Dr. David Akin, of the University of Maryland, he calls it a ‘parashield’. He described it as used for a lightweight manned space capsule here:

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.

Here’s a research article on it:

SpaceOps 2010 Conference
25-30 April 2010, Huntsville, Alabama AIAA 2010-1928
Applications of Ultra-Low Ballistic Coefficient Entry Vehicles to Existing and Future Space Missions
David L. Akin∗
Space Systems Laboratory, University of Maryland, College Park, MD 20742
https://spacecraft.ssl.umd.edu/publications/2010/SpaceOps2010ParaShieldx.pdf

The author here uses units of N/m² which is Pascals instead of kg/m², perhaps because he wants to use units of pressure and to make an analogy to aircraft design’s “wing loading” units of pounds per unit area. But it is easy to convert to the more common kg/m² by dividing by 9.81, i.e., approximately by 10.

Here he takes the desired ballistic coefficient as 200 Pa, about 20 kg/m². This article does give the max stagnation temperature so we can estimate the size of wings needed to reach that ballistic coefficient. Quite notably in this report the peak heating only reaches 800°C. The author notes this is within the temperature range of off-the-shelf Nextel blankets to withstand. But that’s what would be needed for a standard aluminum structure. Stainless steel has a melting point in the range of ca. 1,400°C. Then it maybe no additional thermal shielding would be needed at all as long as wings allow it to reach this low ballistic coefficient.

The ballistic coefficient calculated above is about 60 kg/m². Then we need 3 times higher cross-sectional area to bring that down to ca. 20 kg/m², or likely more if we take into account the added mass.

As before with the inflatable conical decelerator we need to know how the ‘parashield’ mass scales with size. This isn’t provided in the research report. It could be by the cube of the diameter or by the square.

I’ll make an estimate based on just the “wings” being a flat sheet of the needed size. The thickness of the Starship walls is about 4mm. But it’s been speculated in a weight optimized design they could be as low as 2mm thick. Then I’ll use stainless steel at 2mm thick.

The hypersonic drag coefficient of a flat sheet is similar to that of a cylinder at ca. 2. Take the added wing cross-section as 36m*50m, added on to the 9*50 = 450 m² cross-section of the cylindrical Starship. Then with a density of stainless-steel of 7,800 kg/m², the ballistic coefficient calculates out to be:

(60,000 + .002*36*50*7,800)/2*(450 + 36*50) = 19.6 kg/m², below the desired 20 kg/m² point.

According to the assumption the ballistic coefficient is the deciding factor, it could be horizontal wings, delta wing, or the parashields spherical section. Like in the parashield design, additional high-strength spars might need to added to withstand the applied dynamic pressure during the hypersonic/supersonic/subsonic regimes.

But the ‘wing’ is only intended to support the structure on return when the tanks are nearly empty. They might not be able to support it when fully fueled. So the Starship would have to insure to fly a non-lifting trajectory during ascent. It might further be required for the ‘wing’ to be folded away during ascent, only deploying on return.

That the wing only supports the nearly empty weight of the structure during return suggests it could be made even thinner. For instance the Centaur V upper stage at a gross weight of ca. 60 tons has stainless-steel tanks walls only 1mm thick:

Using now a smaller wing cross-section of 18*50, the ballistic coefficient would be:
(60,000 + .001*18*50*7,800)/2*(450 + 18*50) = 24.8 kg/m².

Given the leeway between the 800°C max temperature at 20 kg/m^2 and the 1,400°C melting point of steel this would likely still be sufficient for the stainless-steel structure not needing further thermal protection.

If the added wing could be made this thin then the added weight would only be .001*18*50*7,800 = 7,020 kg, compared to the 60,000 kg of the Starship dry mass+fairing weight.

The importance of such lightweight orbital decelerators has now increased importance now that the Air Force wants rocket cargo point-to-point delivery:

Air Force picks remote Pacific atoll as site for cargo rocket trials
By SETH ROBSON STARS AND STRIPES • March 4, 2025
https://www.stripes.com/theaters/asia_pacific/2025-03-04/cargo-rocket-pacific-johnston-atoll-air-force-17026030.html

The plan is for fully reusable launchers but note these decelerators could still be used to send the cargo back down to their delivery points even with expendable launchers or partially reusable launchers where the only the first stage is reused and the upper stage is expendable. This means they can be used to delivery cargo now with the Falcon 9, and soon with the Rocket Labs Neutron and Blue Origin’s New Glenn, also planned to be partially reusable.

Sierra Space has also already contracted with the Air Force to deliver supplies that were already preloaded and stored in space to distant locations:

Sierra Space Ghost: Revolutionizing Global Logistics
OCTOBER 3, 2024

https://www.sierraspace.com/press-releases/sierra-space-ghost-revolutionizing-global-logistics/

But the very same method can be used as the orbital decelerator for a rocket cargo point-to-point delivery system. Notably, the Sierra Space method is quite analogous to the Akin’s parashield method.

High Hypersonic Lift/Drag Ratio Used for Return From Orbit.

The above discussed methods are useful for just drag decelerators. But the discussion is incomplete for winged reentry because it does not include the effects of lift. For instance if wings with high lift/drag ratio at hypersonic speeds were used the descent rate would be decreased even further, thus further decreasing the heating rate, and thereby allowing a lighter reentry system. The hypersonic aerodynamics of the Space Shuttle have been described as falling “like a brick” with a quite low hypersonic L/D ratio of about 1, thus necessitating it’s heavy thermal protection. Then wings with high hypersonic L/D ratio could greatly improve on this.

This possibility is discussed here:

Clamshell wings for hypersonic reentry of rocket stages. UPDATED, May 4, 2023.

https://exoscientist.blogspot.com/2023/02/clamshell-wings-for-hypersonic-reentry.html

In this approach there would be no extra added weight for the decelerator at all for a returning rocket stage, just the fairing or propellant tank itself as the wings.

Robert Clark

Why SpaceX Needs a TRUE Chief Engineer.

noreply@blogger.com (Robert Clark) — Thu, 27 Mar 2025 21:30:00 +0000

The post below was written back in August last year. Back then any criticism of the approach Elon Musk was taking to develop the Superheavy/Starship was verbotten. Now with two successive failures of Starship upper stages for the very same reason, people are starting to ask tough questions about the development of the Starship:

SpaceX Needs A New Mini-Starship To Land Humans On The Moon And Mars.
By Kevin Holden Platt, Contributor. Kevin Holden Platt writes on space defense…
Mar 17, 2025 at 11:33pm EDT
…
“Our approach today has a very low probability to match the ‘before 2030’ milestone for landing humans on the Moon,” Daniel Dumbacher, who formerly served as Deputy Associate Administrator of NASA’s Human Exploration and Operations Mission Directorate, in charge of the Artemis lunar landings, testified at the hearing.

While he didn’t mention the fiery breakup of SpaceX’s Starship during its January flight demo, Dumbacher, now a professor in aeronautical engineering at Purdue University, said that the ship’s need to be refueled with super-cooled liquid oxygen and methane in low Earth orbit via multiple dockings with still-to-be-developed tankers - a complicated operation that has never been tested - before each flight to the Moon involves an assemblage of complex technologies that might not be perfected within the next five years.

“We might have to build a lander - we might have to scale down the current lander,” Dumbacher told the House, “so that we get to that 2030 landing.”

To avert potentially spiraling problems with testing the colossal Starships during the countdown to this new Moon quest, he said, “I’d get myself a simplified lander - so that I can get to the Moon - that does not require multiple launches.”
…
https://www.forbes.com/sites/kevinholdenplatt/2025/03/17/spacex-needs-a-new-mini-starship-to-land-humans-on-the-moon-and-mars/

__________________________________________

Why SpaceX Needs a True Chief Engineer.

Elon Musk has said that early on when there was still uncertainty if SpaceX would be viable they tried to get a Chief Engineer, but nobody good would come. Musk then designated himself as Chief Engineer.

Now they could get anyone they wanted as a Chief Engineer. My opinion, SpaceX should have as a priority to get a good Chief Engineer.

SpaceX seems to want it make a point of ignoring the lessons of Apollo. When they were building their first launch tower for the SuperHeavy/Starship they were told they needed a flame trench All large launchers use a flame trench. But SpaceX, and more specifically, Musk thought they could do without it:

For some reason, Musk became convinced early on that he did not want the launch tower to have:

*A flame-diverter or flame trench to redirect the blast from the booster’s engines

*A water deluge system to dump a massive amount of water around the launch tower during liftoff

The launch facilities at Kennedy have both of these. Even the launch pads used for the much smaller Falcon 9 have both a flame trench and a water deluge. They help to protect not just the launch pad, and the surrounding area, they also help to reduce the noise. Which sounds trivial, but that noise is energy. That’s what broke up the concrete under the Starship Stage Zero, not the fire. That’s what sent car-sized chunks flying in all directions.

A flame diverter and a water deluge would have greatly reduced, or even eliminated, the damage to the area around the pad. They would have prevented the blow back of debris that damaged Starship before it even left the ground. It might have headed off the whole cascade of events that resulted in that button being pressed 4 minutes into the flight.

We don’t have to guess about whose decision it was not to implement these systems, because Musk already said he decided to skip these systems over the recommendations of his engineers. Musk even had a preview of what was going to happen, as past test flights of the upper stage also resulted in significant spalling of concrete structures and damage to at least one of the ships. He just made them try different kinds of concrete.

The parts for a water deluge were actually on site, ready to install, but Musk decided to forego that installation—likely so he could enjoy the pun of launching his super-joint on 4/20. Which was something Musk had joked about doing months ago.[1]

We know now the result of that poor engineering choice with the launch pad being essentially destroyed, with concrete blocks being thrown hundreds of meters away.

Then, on top of that SpaceX made another inexplicable architecture choice from the beginning. The Apollo program and every other large launch system always constructed a full-up test stand to test all the engines of a stage at once. Instead of taking that approach, SpaceX chose to take the approach of the ill-fated Soviet N-1 rocket of only testing all the engines together during an actual flight. The N-1 became famous in the industry for its explosive failures. Instead of following the approach of the Apollo program of spectacular successes, they chose instead to follow the approach of the N-1 of spectacular failures.

This has important consequences. It might be hard to believe but we still do not know if all the Raptors together on the SuperHeavy can fire at full thrust for the full length of time of an actual mission.

The first test flight resulted in 8 of the 33 engines failing, some in actual explosions. A failure rate of 25%. Observers noted that all the concrete debris being thrown up could have damaged the engines. Elon in an interview however said there was no evidence the concrete being thrown up could have damaged the engines in a way to result in the failures.

SpaceX then was in a bad position. Either the poor engineering choice of not using a flame diverter caused the engine failures, or the engine themselves had such poor reliability that they failed all on their own.

SpaceX in defense of the Raptor reliability might argue there were no failures on the SuperHeavy engines during the ascent for test flights 2 and 3. However, taking note of the slow acceleration of the booster during ascent for these flights, there is very good reason to believe they achieved this reliability by throttling down the Raptors, perhaps to only 75%, [2].

This puts the lack of full-up, full thrust, full duration test burns in sharp relief. This last, of “full duration”, is another way SpaceX deviates from standard industry practice and even terminology. In preflight test burns of the SuperHeavy SpaceX runs these burns for perhaps 5 to 10 seconds, then refers to these burns as “full duration”. In the industry “full duration” is used to mean “full mission duration”. They are burns of the full flight length and full flight thrust and are meant to give confidence to the launch company and importantly to potential customers as well that the engines can run reliably on an actual flight.

Then because SpaceX has not done this with the SuperHeavy, SpaceX’s most important customer NASA has no idea if the Superheavy’s full complement of Raptors can be run reliably at full thrust and full duration of an actual flight. I’m asserting that SpaceX by calling their short burns of only a few second length “full duration” is obscuring this fact.

This again goes to my point of why SpaceX needs a true Chief Engineer. A Chief Engineer should have the persona of a scientist. A scientist, at least a good one, should be scrupulously forthright even on their own research. Instead of Elon Musk proudly boasting how great are the thrust levels of the Raptor engine and how much greater it will be in future versions, a true Chief Engineer would give a more realistic and accurate appraisal. A true Chief Engineer would say something of the nature of,

“The Raptor has reached very high maximum thrust and chamber pressure levels. However, we are working to improve its reliability when run at those levels. Right now when run at 75%, we are at the desired levels of reliability. We are confident though we can reach the needed reliability when run at 100% power and full mission duration.”

Instead of that, Elon Musk as Chief Engineer displays the persona of a PR maven.

NASA Must Know What the True Specifications Are.

My assertion has been SpaceX was intentionally throttling down the Raptors on the booster after the first test flight to improve reliability, [2]. This fact can be regarded as confirmed by the fact the SuperHeavy/Starship has drastically reduced payload capacity than first planned. Elon has stated the payload is only in the 40 to 50 ton range at full reusability when the initial estimate was 100 to 150 tons. If you run the numbers for calculating delta-v it is mathematically impossible for the SH/SS to have a payload that low with the dry mass and engine specifications SpaceX has publicly released. This means either the dry mass numbers or engine numbers or more likely both are significantly worse than what has been released by SpaceX.

NASA must know this. The kinds of calculations known in the industry as delta-v calculations are standard methods of estimating a launch vehicles payload capacity. NASA might be of the opinion that the specifications of a company’s launch vehicle are proprietary and should not be released by them publicly, but NASA has a $4 billion contract with this company to provide an essential component of a multi-billion dollar project being funded by the public, the lunar lander.

It’s a debatable point if NASA should require SpaceX to reveal the true numbers for a project funded by billions of tax-payers dollars. But I don’t think it unreasonable for the public to require NASA to answer the question of whether the low payload capabilities of the SH/SS are what they were expecting at this point in the program, since it is the public that is providing that funding.

Multiple Architecture Mistakes in the SuperHeavy/Starship.

It is irritating for those of us as observers of the space program for decades for SpaceX to simply dismiss the successes of the Apollo program, like Apollo was just some blip in the history of spaceflight. This dismissal is not just of Apollo but of standard industry knowledge. So SpaceX didn’t want to use a flame diverter, and didn’t, and still will not, construct a full-up, all engine test stand for full thrust, full mission duration tests.

But some of the mistakes SpaceX has made in the SH/SS development are just weird. It is standard industry knowledge that you have to have a physical mechanism for separating stages during staging. But SpaceX wanted to do it simply by flinging the upper stage away from the lower stage. Fortunately, they rectified that poor engineering decision with the hot-staging method now used, a method long known to be effective.

On the second test flight they vented oxygen on the second stage while the engines were still firing. Sometimes, you’ll vent propellant from the second stage once it has reached orbit to prevent the possibility of explosion, possibly harming other spacecraft already in orbit. But venting propellant while the engines are firing is a big no-no because of the possibility that itself will cause an explosion, as did indeed happen.

Pressurization gas taken directly from exhaust gases from the pre-burners??? The method used to pressurize the tanks is known as autogenous pressurization. This means the propellant itself is used to pressure the tanks rather than using separate gas such as helium for the purpose. This has been used successfully before with other engines such as the SSME’s. But the other times it was used heat exchangers were used to heat the propellant to pressurize the tanks. Placing exhaust gases directly into the tanks is a very poor engineering approach. Aside from the risk of causing an explosion, there is also the fact the exhaust products contain water and CO2. When these contact directly the cryogenic propellant directly they will tend to freeze. This is bad because the frozen ice or CO2 can then clog the engine inlet valves, as did indeed happen.

SpaceX Taking The Wrong Approach to Reusability.

SpaceX’s approach to reusability is ill-conceived. SpaceX was spectacularly successful in first getting an expendable Falcon 9 then gradually working towards a reusable booster. But instead of taking that successful approach, SpaceX wants the entire SH/SS reusable from the start.

SpaceX has the spectacular success of the Falcon 9’s reusability approach right in front of them but they are ignoring it. First, it is known reusability is most importantly done for first stages because they make up the greatest bulk of the cost, plus the fact in going much lower speed they are more easily made reusable and don’t need thermal protection.

But first just get the rocket to orbit successfully as an expendable. The F9 was able to do that and make a profit as an expendable, with SpaceX all the while working towards reusability.

It might be understandable that SpaceX wanted to take the approach they're taking with the SH/SS if their cost for the launches of the SH/SS were like those of the SLS of ca. $2 billion per launch, since they couldn't get customers at that high launch cost. But Elon has said it is only in the range of $100 million per launch.

Then, if you make the first launches expendable, with the much higher payload without the reusability systems, then SpaceX could be making profit on these launches all-the-while working towards reusability like with the F9.

Another reason why you want to get an expendable launcher first is that extra mass on the upper stage for reusability subtracts directly from payload. For that reason, dry mass on the upper stage is treated like gold. Or said another way extra mass added to an upper stage is like throwing away the equivalent mass in gold.

Elon has said an expendable Starship might have a dry mass of only 40 tons:

_________________________________________________

Elon Musk @elonmusk Mar 29, 2019
Replying to
@Erdayastronaut and @DiscoverMag
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.
_____________________________________________________

https://x.com/elonmusk/status/1111798912141017089?s=61

But the current reusable version has been cited as weighting 120 tons, tripling the dry mass. An extra 80 tons as added to the dry mass of an upper stage is huge in regards to reducing its capabilities.

To put this in perspective the dry mass of the Falcon 9 upper stage is about 4 tons. If this was tripled for reusability it would be 12 tons dry mass, i.e., an additional 8 tons. The LEO payload for reusable booster F9 missions would shrink from 15.8 tons to 7.8 tons. And quite importantly the payload for GEO missions would shrink to 0.

When you take into account the much lower expendable dry mass of the Starship and the fact you don’t wastefully leave propellant on the first stage unused on ascent, for return to launch site, the expendable SH/SS probably could get ca. 200 tons to LEO.

But remember those launches would still be $100 million per launch. That would be a cost of only $500 per kilo even as an expendable. That’s well less than even the reusable Falcon 9. At such low cost SpaceX would definitely be able to find customers just for the expendable SH/SS.

Again, SpaceX needs a true Chief Engineer. I simply cannot believe that the SpaceX leadership is aware of these numbers. A major impetus of Elon Musk in wanting to build the SH/SS is for manned interplanetary missions. A 200 ton payload capability is huge. With payload this high every SH/SS launch would have a capability of doing single launch missions to the Moon or to Mars, no multiple refueling flights, nor SLS required. With two launch towers and each tower accommodating a launch every month, this means every other week the public would be witnessing launches capable of doing manned launches to the Moon or Mars.

Keep in mind this is a capability we could have literally, like, tomorrow, if SpaceX went for the approach of first getting the expendable rocket, proven so spectacularly to work for the Falcon 9.

For those familiar with doing delta-v calculations this is based on the delta-v’s for doing missions to the Moon or Mars, [3]. But keep in mind this would require giving the SH/SS a 3rd stage/lander.

Again, SpaceX needs a true Chief Engineer. It’s basic spaceflight engineering that to do beyond LEO missions, requiring high delta-v’s, it’s most efficiently done by adding additional stages. But instead of doing that, SpaceX wants to take the approach of doing 10, 15 or perhaps even more refuelings for a two-stage launcher.

Proponents of this approach quote Elon’s dictum, “The best part is no part” in arguing against adding an additional stage. But they forget the second part of Elon’s saying. The full saying is, “The best part is no part. The best process is no process.” Rather than this additional, complicated process of multiple refuelings, it would be far cheaper, safer, and faster simply using that 3rd stage.

And even the complaint of having to construct an entire new 3rd stage is incorrect. It turns out a stage capable of filling the role already exists: the Falcon 9 upper stage. This is already a fully human-rated stage. A 200-ton to LEO capacity is so large it could launch the F9 upper stage to LEO fully-fueled and with additional extra capacity so that the F9 upper stage could perform all the additional burns needed for a round trip to the Moon and back.

For the crew capsule don’t use the overbloated in size and cost Orion capsule. Use the Dragon capsule. It was already designed to have a heat shield capable of withstanding the greater heating for return from the Moon compared to return from LEO. The only extra piece of equipment it would need is stronger communications equipment for communicating from the longer distance from the Moon compared to LEO.

And for the mission to Mars? Again, SpaceX needs a true Chief Engineer. Not only is SpaceX ignoring basic principles of spaceflight, it’s ignoring basic principles period.

For a first flight to Mars that might take years for a round trip, you don’t send 100 potential colonists who’ve never been in space before in accommodations akin to a cruise ship. You send a small crew of professional astronauts who have already experienced months long periods in space.

Imagine if for the first Apollo flights the architecture was to send 100 lunar colonists consisting of people who had never been in space before. Apollo would have never taken place under that design.

Then for a first flight to Mars, also add on a 3rd stage akin to the Falcon 9 upper stage, and a small habitat of ca. 10 tons size for a crew of 3 to 4 astronauts for the 6 to 8 month flight to Mars. For the return trip you follow the Mars Direct approach of generating the return propellant on Mars, and have air, food, water, and supplies for the stay on Mars and the return flight already in place on Mars beforehand.

Article In Nature Concludes Starship Missions To Mars Not Feasible.

The SpaceX plan is to send 100 colonists to Mars per mission. The authors of this paper conclude this is not feasible:

About feasibility of SpaceX's human exploration Mars mission scenario with Starship.

Scientific Reports

volume 14, Article number: 11804 (2024)

https://www.nature.com/articles/s41598-024-54012-0

A major issue the paper discusses is the power requirements for producing the propellant on Mars by ISRU. Robert Zubrin has also discussed this problem and estimated it would take 10 football fields worth of solar panels or a nuclear power plant sent to Mars to generate the needed power:

NRPLUS

SCIENCE & TECH

Elon Musk’s Plan to Settle Mars

By ROBERT ZUBRIN

February 22, 2020 4:30 AM

https://www.nationalreview.com/2020/02/mars-elon-musk-plan-to-settle-red-planet/

Note that simply using a small habitat for 3 - 4 astronauts and small third stage a la the Falcon 9 upper stage, makes the power requirements drastically lower and makes the mission doable in a single flight.

Robert Clark

Dept. of Mathematics

Widener University

Chester, PA 19013

1.)https://www.dailykos.com/stories/2023/4/22/2165317/-A-Starship-Post-mortem-Why-the-giant-rocket-failed-and-why-it-s-Elon-Musk-s-fault

2.)Did SpaceX throttle down the booster engines on the IFT-2 test launch to prevent engine failures?

http://exoscientist.blogspot.com/2023/12/did-spacex-throttle-down-booster.html

3.)

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

noreply@blogger.com (Robert Clark) — Sun, 16 Mar 2025 22:38:00 +0000

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.

noreply@blogger.com (Robert Clark) — Thu, 20 Feb 2025 19:47:00 +0000

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?

noreply@blogger.com (Robert Clark) — Fri, 31 Jan 2025 20:04:00 +0000

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.

noreply@blogger.com (Robert Clark) — Tue, 28 Jan 2025 19:05:00 +0000

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.

noreply@blogger.com (Robert Clark) — Mon, 20 Jan 2025 20:13:00 +0000

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.

noreply@blogger.com (Robert Clark) — Wed, 27 Nov 2024 00:40:00 +0000

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

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:

Should the DoD be involved in returning us to the Moon?

noreply@blogger.com (Robert Clark) — Tue, 01 Oct 2024 14:11:00 +0000

Recent news is China expects to send manned missions before 2030. However, delays in the Starship HLS human landing system and the 2nd NASA mobile launcher suggest we might not be able to return to the Moon until 2029 or later.

Some space analysts such as Namrata Goswami have argued China has already surpassed us in the current race back to the Moon, [1]. Space analyst Jordan Wright makes the point that arguing that we already got there first anyway is like saying the U.S. should be called the United States of Norway because Leif Erickson got here first, [2].

The key point is China getting back to the Moon first gives them a significant advantage in establishing a base first there and developing resources, [3]. They could even claim the most useful locations for exploiting resources. When China makes a pronouncement of when they will reach the Moon it should be taken seriously. Unlike the U.S. and every other country’s space program even Russia’s, the Chinese space program is just an off shoot of their military. Then when they want to get to the Moon they’ll do it in the simplest way possible. They don’t have to apportion to each state their little portion of the federal spending to keep each Senator and Congressman happy so the program gets funded.

Remember at the beginning of the U.S. space program in the late 50’s when our rockets kept failing, while the Soviet Union kept succeeding, made famous in the book and movie the Right Stuff? We weren’t able to finally succeed until we gave it over to the military to manage.

Since the DoD now considers the Moon to be a strategic resource they should consider their own approach to getting to the Moon, not through the over bloated SLS. It is quite notable Robert Zubrin, founder of the Mars Society, observed a single launch of the Superheavy/Starship can get a manned mission to the Moon with no refueling flights, nor SLS required:

Mars Direct 2.0 - Dr. Robert Zubrin - IAC 2019.

What’s even more astonishing is each of those flight should only cost ~$100 million, as Elon Musk has said that is the cost of a SuperHeavy/Starship launch. I said ”should” because that’s what it would cost to SpaceX. But of course once the DoD is paying for it, SpaceX will find a way to charge them a billion dollars for it. But even that is just chicken feed to the annual DoD budget in the range of $850 billion a year.

It's quite notable as well Zubrin calculates that two launches of reusable Starship at a 100+ payload to LEO capacity could do manned missions to Mars. This means the expendable version at 200+ tons to LEO capacity[4] could so single launch manned missions to Mars. Remember both the possibilities of single launch manned Moon and Mars missions can be done now. It would require just stripping off the reusability systems and adding on an already existing small stage such as the Falcon 9 upper stage to serve as a 3rd stage/lander.

The low possible cost, based on the real cost to SpaceX, is astonishing in comparison to the projected multi-billion per launch cost of the Artemis architecture, with all of its /SLS/Orion/Starship HLS or Blue Moon landers/Boeing EUS/Advanced SRB's/Lunar Gateway space station/ components. The SLS is amortized at $2 billion per launch, the Orion capsule at $2 billion each, and a $4 billion contract to SpaceX for two launches of the Starship HLS amounting to $2 billion each, or a $3 billion cost for a single launch of the Blue Origin Blue Moon lander. That's already $6 to $7 billion per mission. Then the Boeing EUS/Advanced SRB's/Lunar Gateway are likely to add another $1 to $2 billion to the per mission launch cost. This is likely to be a total in the range of $8+ billion per mission for the Artemis missions.

I once asked ChatGpt what were the per launch costs of the Apollo program in current dollars. It’s response was:

____________________________________________________________________________

The Apollo program, which included a total of 17 missions (from Apollo 1 through Apollo 17), cost approximately $25.4 billion in 1973 dollars. To estimate the cost per flight in today’s dollars, we’ll follow these steps:

1. Adjust the total cost for inflation:

• The cumulative inflation rate from 1973 to 2024 is approximately 5.8 times. This means that $1 in 1973 is worth about $5.80 in 2024 dollars.

Total cost in 2024 dollars = $25.4 billion x 5.8 = $147.32 billion

2. Calculate the cost per flight:

• There were 17 Apollo missions, including the uncrewed and crewed missions.

Cost per flight = $147.32 billion/17 = $8.67 billion per flight

Summary:

The Apollo program cost approximately $147.32 billion in today’s dollars, which breaks down to about $8.67 billion per flight on average.
____________________________________________________________________________

And we already know Apollo was unsustainable.

In contrast the approach using the Superheavy/Starship launcher directly, no refueling flights nor SLS required, could potentially be as low as ~$100 million per flight. It is quite notable this is in the range of what NASA pays for crewed flights to the ISS. In other words the long-desired goal of having a sustainable human habitation on the Moon is doable now by following this far simpler, faster, and cheaper approach.

Unfortunately, large government programs, and Artemis certainly is one, have an inertia that make them difficult to budge. NASA is too attached to the Artemis program and SpaceX is too attached to its multi-refueling approach for flights to the Moon and Mars to change them. The only possibility is for the DoD to pay SpaceX for this alternative approach.

Not only would China getting first its footprints and a base on the Moon in the current lunar race offer the advantage of valuable resources, but it could in fact give it a devastating military advantage.

The concept of lunar mass drivers or railguns have long been written about as means of getting lunar resources back to Earth or into space, such as for orbiting propellant depots. But they also have a clear usage as a near indefensible weapon. China is planning on building such "mass drivers" on the Moon for delivering resources in the form of "spin launch", [5]. Publicly, their stated purpose is only for delivering resources back to Earth. But given that China's space program is just an off-shoot of their military, its military applications must be under consideration as well by the Chinese military.

Robert Clark

1.) Features | Security | East Asia
China Has a Head Start in the New Space Race
The “Age of Apollo” is over and the Chang-e era has dawned.
By Namrata Goswami
May 29, 2019
https://thediplomat.com/2019/05/china-has-a-head-start-in-the-new-space-race/

2.) China just took the lead over NASA and Artemis in the race back to the Moon!
https://youtu.be/u47T1PVBW4I?si=u609RHLABcoZ6P6Z

3.) China plans to build moon base at the lunar south pole by 2035.
News
By Andrew Jones published September 10, 2024
https://www.space.com/china-moon-base-south-pole-2035

4.)250 Tonnes to Orbit!?: SpaceX's New Expendable Starship Option.
https://www.youtube.com/watch?v=UutHG8Y2UuQ

5.) China will build a Super-Weapon on the Moon! Can NASA stop them?
https://youtu.be/eElDqTNe4oE?si=FC1TFw5TSubyUkCE

Low cost approach to winged, air-breathing and rocket SSTO's, Page 1.

noreply@blogger.com (Robert Clark) — Tue, 11 Jun 2024 03:59:00 +0000

Table of Contents.

1.)Radically Reduced Costs by using Private Financing and Already Existing Systems.

2.)B-58 Based Combined-cycle SSTO.

3.)Reducing Structural Weight.

4.)Reducing Weight of Ancillary Systems.

5.)Combined-cycle Engines.

6.)Could we get Mach 5+ for a kerosene-fueled air-breather vehicle?

7.)Estimated Performance for a B-58 Based Airbreathing/Rocket SSTO.

_______________________________________________________________

1.)Radically Reduced Costs by using Private Financing and Already Existing Systems.

I'll take a few approaches to estimate the technical feasibility of a combined jet/rocket SSTO. But first, note quite key to its financial feasibility is that costs can be 1/100^th that of what you get from traditional NASA estimates. First, it is just filtering in to the general mindset of those in the space industry how important it was that SpaceX demonstrated 90%(!) reduction in development costs by the private financing approach of commercial space. ESA for example had to be dragged kicking and screaming into the idea of using the private financing approach to new launchers and spacecraft. They are finally following that approach now, though.

Secondly, anyone familiar with the space program knows development cost for a new system developed entirely from scratch is far more than the individual cost of the launchers or spacecraft that result. So base your system around components already developed and in service. This can cut development costs again by an additional factor of 10, and commonly more.

These two factors together mean you can cut costs by more than 1/100^th what you would expect by traditional NASA cost estimates.

Then the technical plan is to base the airframe for the air-breathing portion of the flight on existing jet airframes. Then that airframe would carry an existing hydrolox rocket stage that would provide the rocket propelled portion of the flight. In this regard, you may be able to save additional dry mass by integrating the airframes fuselage and hydrolox core into one unit, rather than placing the hydrolox stage inside the existing fuselage.

To the end of reducing development costs and towards getting an operational prototype vehicle flying quickly, this next suggestion is key: use kerosene(jet) fuel for the first air-breathing portion of the flight.

First, it was long thought for SSTO’s that it had to be hydrogen fueled because it offered the highest ISP. Then later it was realized because of the heavy tankage mass for hydrolox rockets, that dense propellants actually were preferred for SSTO’s. Further analysis showed if you could have a tripropellant engine burning both kerosene and then hydrogen, you could improve performance further.

But an even more important practical consideration is that we don’t have an operational hydrogen-fueled turbojet engine operating now. For example, the jet engine company Rolls Royce is researching hydrogen fueled jet engines, but their CEO seemed to poor cold water on their own research program suggesting they won’t come into operation until the 2040’s. Remember now though, key to my approach is to use existing, operational components.

2.)B-58 Based Combined-cycle SSTO.

Now, what airframe to use? For jet aircraft, very different from rockets, the mass ratio between gross fully fueled weight and the empty weight is rather low, commonly only between 2 and 2.5. So, I was looking for aircraft that had the highest mass ratio. I was interested to see the legendary B-58 Hustler had a mass ratio of over 3. The B-58 is long retired now but it may be possible to buy one for renovation for operational use.

And its futuristic look would make it utterly badass for a SSTO-design as an actual rocketship that flies to space:

Specifications (B-58A)

General characteristics
Crew: Three
Length: 96 ft 10 in (29.51 m) ^[95]
Wingspan: 56 ft 9 in (17.30 m) ^[95]
Height: 29 ft 11 in (9.12 m)
Wing area: 1,542 sq ft (143.3 m²)
Aspect ratio: 2.09
Airfoil: root: NACA 0003.46; tip: NACA 0004.08^[96]
Empty weight: 55,560 lb (25,202 kg)
Gross weight: 67,871 lb (30,786 kg)
Max takeoff weight: 176,890 lb (80,236 kg)
Zero-lift drag coefficient: C_D0.0068
Frontal area: 10.49 sq ft (0.975 m²)
Powerplant: 4 × General Electric J79-GE-5A afterburning turbojet, 10,400 lbf (46 kN) thrust each dry, 15,000 lbf (67 kN) with afterburner
Performance
Maximum speed: 1,146 kn (1,319 mph, 2,122 km/h) at 40,000 ft (12,000 m)^[95]
Maximum speed: Mach 2.0
Cruise speed: 530 kn (610 mph, 980 km/h)
Range: 4,100 nmi (4,700 mi, 7,600 km)
Combat range: 1,740 nmi (2,000 mi, 3,220 km)
Service ceiling: 63,400 ft (19,300 m)
Rate of climb: 17,400 ft/min (88 m/s) at gross weight^[97]
Lift-to-drag: 11.3 (subsonic, "clean configuration")
Wing loading: 44 lb/sq ft (210 kg/m²)
Thrust/weight: 0.919

https://en.m.wikipedia.org/wiki/Convair_B-58_Hustler#Specifications_(B-58A)

It doesn't have to be the B-58 though. Other in use or recently in use jets have mass ratios in the range of 2.5 might be used such as the McDonald A-4 Skyhawk:

https://en.m.wikipedia.org/wiki/Douglas_A-4_Skyhawk#Specifications_(A4D-5_/_A-4E_Skyhawk)

Douglas A-4 Skyhawk - Wikipedia

en.m.wikipedia.org

Such retired aircraft probably could be purchased at relatively low cost as surplus.

Note, also that airframes based on fighters or bombers would get great attention to be supported by air forces around the world. But when they offer to finance the project, politely say, No. Remember key to keeping the development cost low is to use private financing. But you could use the interest the various air forces have in purchasing the vehicle as a selling point to get that private financing. (It would be quite amusing to see the Air Force officers jaws drop when they offer to pay for the development, and you respond, "Nah, we don't need that.")

3.)Reducing Structural Weight.

Now, the reason why we may not need the high mass ratio of the B-58 airframe is from a rather surprising fact about the B-58. The B-58 was quite innovative and still is for the weight saving measures it undertook. It turns out it's structural empty weight fraction, the fraction of its structural weight to its maximum takeoff weight, was only 13.8%. This corresponds to structural mass ratio of over 7.

The structural weight as the name implies includes the structural members only of the craft, such as wings, fuselage, tail, landing gear, engine nacelles. It excludes engines, avionics, wiring, APU's, hydraulics, instruments, seats, etc.

But after looking up references I found it is common for the structural empty weight to be significantly less than just the empty weight, more than can be accounted for by just removing the engine weight:

From:

Bonded Bomber — B-58.
Author(s): L. M. Smith and C. W. Rogers
Source: SAE Transactions , 1962, Vol. 70 (1962), pp. 477-486 Published by: SAE International
Stable URL: https://www.jstor.org/stable/44469505

You see in Table 2 the structural empty weight is much less than the known empty weights for all the aircraft in the table, in the range of half as much. The B-58 was designed in the late 50's. Then with modern materials we may be able to cut the structural weight again by an additional factor of two.

See this graphic of some modern high strength aluminum alloys:

There are some high strength steel alloys that also improve over standard aluminum in strength-to-weight ratio, by better than a 2 to 1 factor. Because of the heat induced by the high Mach flight we may want to use these or titanium over the high strength aluminum alloys. See discussion here:

DARPA's Spaceplane: an X-33 version, Page 2.
https://exoscientist.blogspot.com/2018/06/darpas-spaceplane-x-33-version-page-2.html

Other techniques that might be used to reduce structural weight is that of the "isotruss":

Isotruss Tower

280' tower installed in Spanish Fork, Utah

It is 4 times better in strength-to-weight in bending and 7.5 times better in buckling than standard aluminum grades. And compared to standard steel grades it is 10 times better both in bending and bucking on strength-to-weight basis. This would quite helpful in reducing the weight of wings.

The isotruss gets its great strength even beyond carbon-fiber composites from its geometry. However, like carbon composites the different strands have to be epoxied together. And like carbon composites also this is where it loses some of its strength. Then we may also try to form the isotruss from high-strength aluminum or steel alloys. These would be isotropic in strength so would not lose strength at the joints. See discussion here:

Horizontal landing for the BFR on Earth. UPDATED, 12/15/2018.
https://exoscientist.blogspot.com/2018/11/horizontal-landing-for-bfr-on-earth.html

Another possibility for reducing wing weight would be the idea proposed for the horizontal lift-off SSTO proposal from the 70's called Star-Raker. The technique the Star-Raker used for the wings was innovative. They used multiple cylindrical tanks in the wings such that weight of the wings as tanks was close to that of normal cylindrical tanks:

BLAST FROM THE PAST; a few good ideas may return to the light of day…

https://horizontalspace.wordpress.com/2015/08/10/blast-from-the-past-a-few-good-ideas-may-return-to-the-light-of-day/

As a familiar example of heavy tank weight, recall the X-33. The non-cylindrical shape of the tanks resulted in poor weight efficiency. That poor weight efficiency led to the perceived idea carbon fiber tanks were needed to save weight. The failure of those tanks led to the program cancellation. In effect, the poor weight efficiency of non-cylindrical tanks led to the program cancellation.

4.)Reducing Weight of Ancillary Systems.

The amount by which the structural weight is less than the empty weight is not accounted for by just subtracting the engine weight. In fact, the ancillary systems beside the structures and engines, such as avionics, wiring, APU's, hydraulics, instruments, seats, etc. weigh nearly as much as the engine weight. In this regard then its quite important to note than we may be able to cut the weight of these systems by significantly more than by just 50%.

See this table of weights of the various systems in the Boeing 737-200 to get an idea of the weights of the ancillary systems on aircraft:

From:

Comparison of Aircraft Conceptual Design Weight

Estimation Methods to the Flight Optimization System.

https://ntrs.nasa.gov/api/citations/20190000431/downloads/20190000431.pdf

The 737-200 was designed in the 60's. Quite likely the ancillary weights on the B-58 can also be reduced.

First, "Surface Controls", means "control surfaces", ailerons, elevators, rudder, flaps, etc. Like structural weight this can be reduced better than 50% with modern lightweight materials.

"Auxiliary Power" unit, APU, is used to start up the engines. For the Boeing 737-200 it can put out 45 KVA, 56 kW. For that 56 kW it weighs 800 kg. Now see the lists here for power-to-weight ratios of heat engines and electric engines:

https://en.m.wikipedia.org/wiki/Power-to-weight_ratio#Engines

You see several that are in the range of a few tens of kW, a la the 56 kW APU of the 737-200, have power-to weight ratios of 5 to 10 kW/kg. That would result in their weights being in the range of only 5 to 20 kg.

For example, Honeywell produces a 45 kVA electric generator at only 28.3 pounds, 13 kg.

For our unmanned launcher, the 400 kg for the "Instruments" would be removed.

For "Hydraulics", they can be replaced by electromechanical motors at about 60% the weight:

HYDRAULIC ACTUATOR REPLACEMENT USING ELECTROMECHANICAL TECHNOLOGY WEIGHT REDUCTION.
The primary reason the Aircraft Industry moved away from hydraulic actuation and into electro- mechanical actuation is the reduction in overall systems weight. In a case study completed for a small business jet. The weight of a full hydraulic system for landing gear actuation was compared to an electromechanical system. The weight savings for a 28VDC system was 15 pounds over the hydraulic system with tubing and actuators. For the electromechanical system the weight was 24 pounds as compared to 39 pounds for the hydraulic system. This represents approximately a 40% weight saving for the employment of electromechanical actuators.
https://www.jdtechsales.com/wp-content/uploads/2018/12/Hydraulic-Actuator-Replacement-using-Electromechancial-Technology_Whitepaper.pdf

For electrical systems, carbon nanotubes can cut weight by 80% over copper:

Can Carbon Nanotubes Replace Copper?
By John Sprovieri

But perhaps the most intriguing way to use CNT fibers is to spin them into conductive yarns that could someday replace copper wire in wiring harnesses and motor windings. The driving force behind this application is weight reduction.

Consider an RG-58 coaxial cable. The weight of a standard copper construction is 38.8 grams per meter, says Stefanie E. Harvey, Ph.D., senior manager for corporate strategy at TE Connectivity. Replacing the copper braid with CNT tape would reduce the weight to 11.5 grams per meter. Replacing the center conductor with CNT yarn would further reduce the weight to 7.3 grams per meter, for a total weight savings of 80 percent.

Such a reduction equates to hundreds of pounds in an aircraft, says Harvey. For example, the F-35 fighter contains approximately 15 miles of cable. If the copper shielding on all that cable were replaced with CNT tape, the total weight of the cabling could be reduced by approximately 1,180 pounds, she says. If the shielding and the conductors were replaced with CNT materials, the total weight savings would be 1,975 pounds.
https://www.assemblymag.com/articles/93180-can-carbon-nanotubes-replace-copper

For the "Avionics", that would be practically nothing in weight now compared to its weight in the 50's and 60's. Recall, back then we were still using vacuum tubes and relays.

No "Furnishings and Equipment" for our unmanned, cargo vehicle.

No "Air Conditioning + Anti-Icing" for our unmanned, cargo vehicle.

Putting these weight savings together we can reduce the ancillary weights from 16,900 lbs of the 55,000 lbs empty weight, about 30%, down to only 1,300 lbs, only 2%.

5.)Combined-cycle Engines.

There is still the issue of the combined-cycle engines. Reaction Engines did testing of their precooler for their Skylon Sabre engine by sending the exhaust of a small jet engine into it to confirm it could deal with the high temperatures.

However, since we would be using a kerosene turbojet, I advise testing their precooler by placing it in front of an existing kerosene turbojet engine and confirming they could get the needed cooling in that case. For the cooling you could still use the hydrogen or perhaps kerosene in this case.

In this scenario, we would use oxygen-rich combustion for the jet engine Reaction Engines in their tests places in front their precooler, so that the exhaust exiting the precooler has enough oxygen to support combustion of a jet engine placed after the precooler.

Alternatively, instead of using a jet engine to send exhaust into the precooler we could use, for example, electrically heated air that is then accelerated by a de Laval nozzle, to have the needed speed and temperature to emulate air flowing into the precooler at ramjet speeds.

This would actually be a quite important test to do. It's long been known that the maximum speed a ramjet could be operational getting positive thrust is Mach 5 to 6. But there is little published work on this being actually achieved experimentally. The only thing I've found is almost anecdotal reports of it being achieved accidentally. For instance the research conducted by Onera in the 50's and 60's:

1946 to 1962: aeronautical research that is rapidly gaining momentum.
From the beginning, large teams (of 150 to 200 people) were assigned to the study of liquid propellant and solid propellant rocket engines, ramjets and turbomachinery. In particular, in 1951 in Hammaguir in Algeria, ONERA launched its Stataltex ramjet, which for a long time held the world record for speed and altitude for target devices of its class, reaching Mach 5 at an altitude of 38 km.
https://www.onera.fr/en/history/onera-70-years-1946-1962-aeronautical-research-that-is-rapidly-gaining-momentum

And this test series with the Martin Marietta ASALM ramjet powered missile in the 70's:

The Air Force Almost Got A Near Hypersonic Radar Plane Killing Cruise Missile Decades Ago.
The primary goal was to give bombers, such as the B-52, a means to destroy Soviet air defense sites and airborne early warning and control aircraft.
During one of the tests, the PTV test vehicle actually exceeded expectations, reaching a hypersonic speed of Mach 5.5 at an altitude of 40,000 feet. In at least one of the launches, an A-7 Corsair II combat jet was used as the launch platform, indicating the Air Force may have considered expanding the number of aircraft certified to carry the weapon. Mockups of the ASALM missiles were also shown mounted on the rotary launcher for the B-52.
https://www.twz.com/34036/the-air-force-almost-got-a-near-hypersonic-radar-plane-killing-cruise-missile-decades-ago

This last occurred reportedly when a fuel valve got stuck open.

Then if Reaction Engines with their Sabre engine could definitively show even on the test stand positive thrust from their precooled engine at Mach 5+ air inflow, that would be a significant advance for propulsion at the upper end of the the ramjet range.

Note though strictly speaking the Sabre is not a ramjet. The unique approach of the Sabre is that it first sends the Mach 5+ incoming air into their precooler and then sends the chilled air into the standard compressors of a turbojet.

The difficulty of the Sabre development is that Reaction Engines wants to use hydrogen fuel. But there aren't any hydrogen fueled turbojets available. And as mentioned the leading researcher on hydrogen-fueled turbojets Rolls Royce engines feels it may take another two decades to make them viable.

Then I advise Reaction engines try proving their concept with kerosene(jet) fuel. It may even be possible to do using existing jet engines.

So in regards to achieving Mach 5+ air-breathing propulsion we have a scenario where we have two possible methods: the precooler with turbojet approach and the ramjet approach. Because Reaction Engines precooler has already been proven to work for cooling high velocity, high temperature incoming air, there's a high probability it will work when attached to an an actual (kerosene-fueled) turbojet. Note though we don't have the capability, yet, for using the high Isp hydrogen for the fuel for this approach.

On the other hand for the ramjet approach we only have anecdotal evidence that it can work in the Mach 5+ speed range. But ramjets have been described as "flying stovepipes". They have no moving parts. It would seem they should be able to work up to the maximal speeds of Mach 5+, even using high Isp hydrogen fuel. Yet still we don't have definitive evidence they have been successfully accomplished at these high speeds.

I advise both methods be tried for our proposed airbreathing/rocket SSTO.

For the standard turbojet/ramjet approach, in keeping with the principle of using already developed components we could try adapting the famous J58 engine of the SR-71 to Mach 5+ flight. We would close off the intake's pathway to the compressors when we exceeded the Mach 3.2 operating speed of the J58, which likely can be extended to ca. Mach 3.5. We may need to replace the materials of the combustion chamber to deal with the higher temperatures of the ramjet combustion at Mach 5+ speeds. Another possibility is because we will be carrying hydrogen we can use that to cool the combustion chamber keeping the same combustion chamber materials.

Note also, that while getting a turbojet to operate on hydrogen fuel is a non-trivial technical problem, a ramjet is a much simpler operational engine. In that case, getting useful propulsion with hydrogen fuel is likely much simpler. So for our adapted J58 engine we may get significantly higher ISP by switching to hydrogen fuel for the ramjet portion of the flight.

That's for using the J58. We should investigate as well though combining separate operational turbojets and ramjets. Most turbojets use for example some proportion of bypass air that is routed away from the compressors. Then we could either use this air either in the original combustion chamber when the jet is adapted to ramjet operation or we could bolt instead onto the jet engine a new combustion chamber taken from an already in use ramjet engine.

As before we can use kerosene or hydrogen for the cooling. The possibility of using hydrogen for the cooling leaves open the possibility that we can even use a ramjet proven to work only at, say, Mach 3+, to work at Mach 5+ due to the increased cooling made possible using hydrogen.

6.)Could we get Mach 5+ for a kerosene-fueled air-breather vehicle?

I'll take a tentative conclusion of yes. I'll look at the SR-71. It's rated top speed was Mach 3.2. But on at least one mission it reached Mach 3.5.

Here are its specifications:

General characteristics

Crew: 2; Pilot and reconnaissance systems officer (RSO)
Length: 107 ft 5 in (32.74 m)
Wingspan: 55 ft 7 in (16.94 m)
Height: 18 ft 6 in (5.64 m)
Wheel track: 16 ft 8 in (5 m)
Wheelbase: 37 ft 10 in (12 m)
Wing area: 1,800 sq ft (170 m²)
Aspect ratio: 1.7
Empty weight: 67,500 lb (30,617 kg)
Gross weight: 152,000 lb (68,946 kg)
Max takeoff weight: 172,000 lb (78,018 kg)
Fuel capacity: 12,219.2 US gal (10,174.6 imp gal; 46,255 L) in 6 tank groups (9 tanks)
Powerplant: 2 × Pratt & Whitney J58 (JT11D-20J or JT11D-20K) afterburning turbojets, 25,000 lbf (110 kN) thrust each

JT11D-20J 32,500 lbf (144.57 kN) wet (fixed inlet guidevanes)

JT11D-20K 34,000 lbf (151.24 kN) wet (2-position inlet guidevanes)

Performance

Maximum speed: 1,910 kn (2,200 mph, 3,540 km/h) at 80,000 ft (24,000 m)
Maximum speed: Mach 3.3^{[N 8]}
Ferry range: 2,824 nmi (3,250 mi, 5,230 km)
Service ceiling: 85,000 ft (26,000 m)

Now use the fact the max velocity for an aircraft varies by the square-root of the thrust:

Are there formulas for estimating theoretical top speeds of aircraft?

Absolutely, and it's pretty much as simple as it gets. T=D, where T is the maximum thrust of your engine, and D is the drag that corresponds to the maximum velocity of the vehicle.

Now, both of those values are functions of a few other variables, and so getting the exact numbers, especially for drag, can be a bit tedious. However, the drag equation can be simplified to a function of velocity, which looks like this:

$D = \frac{1}{2} ρ V^{2} A C_{D}$

Where $ρ$ is the density of air at the flight condition (a constant), A is any reference area on the aircraft, usually wing area, and $C_{D}$ is an experimentally determined value, called the coefficient of drag, using that reference area. Wind tunnel testing is used to calculate the drag coefficient (and the corresponding lift coefficient), because it's a function of a whole lot of variables that is not easy (or even analytically possible) to solve.

So, all you do is replace D in the first equation with that mess, move all constants (everything but velocity) to one side, and out pops the following formula for the max speed of your aircraft:

$V = \sqrt{\frac{2 T}{ρ A C_{D}}}$

https://www.quora.com/Are-there-formulas-for-estimating-theoretical-top-speeds-of-aircraft

Now instead of the two J58 engines on the SR-71, imagine giving the SR-71 four of the F135 engines:

F135-PW-100

Data from Pratt & Whitney,^[4] Tinker Air Force Base,^[51] American Society of Mechanical Engineers^[52]

General characteristics

Type: Two-spool, axial flow, augmented turbofan
Length: 220 in (5,590 mm)
Diameter: 46 in (1,170 mm) max., 43 in (1,090 mm) at the fan inlet
Dry weight: 3,750 lb (1,700 kg)

Components

Compressor: 3-stage fan, 6-stage high-pressure compressor
Combustors: annular combustor
Turbine: 1-stage high-pressure turbine, 2-stage low-pressure turbine
Bypass ratio: 0.57:1

Performance

Maximum thrust:
- 28,000 lbf (125 kN) military thrust,
- 43,000 lbf (191 kN) with afterburner
Overall pressure ratio: 28:1
Turbine inlet temperature: 3,600 °F (1,980 °C; 2,260 K)
Thrust-to-weight ratio: 7.47:1 military thrust, 11.47:1 augmented

https://en.wikipedia.org/wiki/Pratt_%26_Whitney_F135#F135-PW-100

Two of the J58 engines have a 300 kN thrust in afterburner, while four of the F135 engines would have a thrust of 760 kN in afterburner, larger by a factor of 2.5. Since max speed varies by the square-root of thrust, the max speed would be larger by a factor of 1.6. From a max speed of Mach 3.5 to a max speed of Mach 5.6.

This would be just about the limit for ramjet and precooler/turbojet propulsion.

Note quite key about this estimated top speed is the assumption that the famous ramjet mode of the J58 engine above about Mach 2 can be extended to the Mach 5+ range. Quite likely it would require using additional cooling with either kerosene or hydrogen at these speeds. But more to the point I'm supposing replacing the J58 engine with the F135, due to its higher thrust at about the same weight.

But the F135 does not have a ramjet mode. However, if you look at the operation of the J58 in ramjet mode, it is not particularly complex additional structures that need to be added, mostly adjusting the position of the inlet spike and opening and closing additional air inlet doors:

Radian Aerospace Spaceplane.

noreply@blogger.com (Robert Clark) — Thu, 02 May 2024 16:26:00 +0000

I was interested to see this interview with Livingston Holder, co-founder of Radian Aerospace:

As well as the video lecture Mr. Holder did for Valispace:

I noted Radian is following the approach of winged horizontal launch for their near-SSTO. It’s a near-SSTO because it uses sled launch to reduce the propulsion requirements of the craft.

I wonder if Radian could do the trades instead for a vertical launch approach that would be a true SSTO. Former Air Force officer Major Mitchell Burnside Clapp who did stints both with the DC-X and X-33 programs did an analysis showing both hydrogen and kerosene fueled true SSTO’s are possible by vertical launch:

A LO2/Kerosene SSTO Rocket Design.

https://web.archive.org/web/20211001062714/https://erps.org/archives/321

Vertical launch has the advantage that the wings only have to support the dry mass of the craft during return, thus requiring much reduced wing weight. In horizontal launch on the other hand the wings have to support the entire weight of the fully-fueled craft, thus needing a much heavier wing weight.

As Mr. Holder described in the Valispace lecture though the horizontal launch approach has the advantage it can use a sled to provide some of the propulsion and also horizontal flight can reduce the gravity loss.

Still, those large heavy wings for horizontal launch induce a significant dry mass hit which is why I would like to see a trade analysis on both approaches.

In support of the feasibility of vertical launch where the wings only have to support dry mass on return, we can look at the Air Force’s X-37B spacecraft. This is certainly not a SSTO, but I’m referencing the short, stubby wings that enable winged return from orbit.

The small wings will correspond to a relatively small wing weight. Indeed, we could use the X-37B’s plan as a model, scaled-up, for what a vertical launch SSTO might look like. There are actually several aircraft/spacecraft with short stubby wings we could model the basic plan on.

Robert Clark

@RGregoryClark

Use short, stubby wings for lightweight. Use spiral path to return 1st stage to launch site.

1:13 PM · Mar 17, 2021

In Mr. Holder’s video discussion he mentioned that during the famous X-33 development and competition back in the 90’s to develop a SSTO, he decided with Boeing to pull out of the competition on financial viability reasons.

But two very important reasons suggest why that conclusion needs to be reconsidered, and why quite likely it is no longer valid.

First, an overwhelmingly important fact is SpaceX has shown space vehicles as privately financed can cut 90%(!) off the development cost compared to the usual government financed approach. And secondly, the large number of satellite launches required with several megaconstellation satellite systems soon coming online, requiring tens of thousands of satellites each, make reusable launchers imperative.

So the two factors of low development cost plus a definite market for the needed launches suggest a reusable SSTO may now be financially feasible.

Still, an operational SSTO has not been done before so there is still some financial and technical risk involved. Then I advise first doing a small scale vehicle that, most importantly, is based on a design of existing expendable rocket stages.

The importance of basing it on existing rocket stages is we can look at the most efficiently weight-optimized stages and then make estimates of the added weight needed for reusability systems.

Then there are two existing rocket stages I think might be useful in this regard: the kerosene fueled Falcon 9 upper stage and the hydrogen fueled Centaur upper stage. Again we want to use small stages at the beginning to limit the financial risk.

The F9 second stage has a quite impressive mass ratio:

Type	Falcon 9 FT Stage 2
Length	12.6m (Separated Length)
Diameter	3.66 m
Inert Mass	4,000 kg (est.)
Propellant Mass	107,500 kg (est.)
Fuel	Rocket Propellant 1
Oxidizer	Liquid Oxygen
LOX Mass	75,200 kg (est.)
RP-1 Mass	32,300 kg (est.)
LOX Tank	Monocoque
RP-1 Tank	Monocoque
Material	Aluminum-Lithium
Guidance	Inertial
Tank Pressurization	Heated Helium
Propulsion	1 x Merlin 1D Vac
Engine Type	Gas Generator
Propellant Feed	Turbopump
Thrust	934kN
Engine Dry Weight	~490kg
Burn Time	397 s
Specific Impulse	348s
Chamber Pressure	>9.7MPa (M1D Standard)
Expansion Ratio	165
Throttle Capability	Yes
Restart Capability	Yes
Ignition	TEA-TEB, Redundant
Pitch, Yaw Control	Gimbaled Engine
Roll Control	Reaction Control System
Shutdown	Commanded Shutdown
Reaction Control S.	Cold-Gas Nitrogen Thruster

https://spaceflight101.com/spacerockets/falcon-9-ft/

An est. 4 tons of dry mass and est. 107.5 tons propellant mass.

Its single Merlin could not launch it from the ground however, since, as an upper stage, the engine does not need thrust/weight ratio above 1.

Two solutions for that that also speak to the vertical vs. horizontal launch question: 1.) we can add a second Merlin, and 2.)we can add sufficient wings for ground launch.

For the 1st approach we would only use small, stubby wings a la the X-37B as we are using vertical launch and the wings only have to support the dry mass on return.

For the 2nd approach note for horizontal launch the engines do not have to have a T/W above 1 because aerodynamic lift is being used to loft the vehicle. But because the wings have to support the fully-fueled weight they need to be heavier, eventhough you don’t have the added weight of an additional engine.

The trades need to be done to see which approach is optimal. It may be both approaches would close to allow an operational SSTO.

As a first level estimate of wing weight, they are typically 5% to 10% of the lifted weight of an aircraft. But with composites we might be able to make it half to that at 2% to 5% of the lifted weight. Now for the case of adding a second Merlin to the F9 upper stage for vertical launch, that would be an additional 470 kg for the extra engine. Then if wing weight is as low as 2% of dry mass that would only be an additional 90 kg added to the dry mass.

On the other hand, if you have to use wings to loft the entire fully-fueled weight that’s 2% to 5% added on to the 108 ton gross weight. That’s 2 to 5 tons added onto the vehicle dry mass in wings alone.

That makes it difficult for this approach to close without the addition of sled launch. Sled launch though adds more to the capital cost of development.

By the way, Holder also discusses in his interview that the engines would use altitude compensation to achieve high vacuum Isp from engines that can launch from ground. There are many ways of accomplishing this. One way that would involve the least technical risk because it is a technique that has been in use already for decades is to use extendible nozzles.

Robert Clark

@RGregoryClark

1st image shows how the extendable nozzle on the RL10 looks; 2nd shows it can even work for multiple engine stages; 3rd image shows how an inflatable nozzle would look.

GIF

5:16 PM · Oct 16, 2022

Note that while the suggestion is of basing the SSTO core on the Falcon 9 upper stage it is not necessary to use the actual stage or even the Merlin engine. The tech involved in developing the Falcon 9 was not particularly advanced. Indeed the performance of the Merlin is no more advanced than the kerolox engines already available in the 60's. Then probably we could use just surplus engines lying around such as the RS-27 engines. Quite key is the increased performance needed will come from the altitude compensation nozzles applied to them, able to increase both thrust at sea level and Isp at vacuum.

That’s the discussion for the Falcon 9 upper stage case. For the hydrogen-fueled Centaur case there is again the question of adding additional engines vs. large wings for horizontal launch. For the Centaur to use vertical launch though using the RL10 engine an important consideration is how much the thrust is reduced for the sea level version, to ca. 6 ton force. Then you would likely need 4 to 5 engines instead of the one or two used on the Centaur. This would add in the range of 1 ton to the original 2 ton dry mass just in added engine weight.

But for the wings for the horizontal launch, if it is 2% added weight that would be 2% of the 23 ton gross mass or 460 kg. But if the added wing weight is at 5% that would be 1,150 kg added weight. This is about the same as the added engine weight case.

So it is possible here horizontal launch is more competitive to the vertical launch case.

Robert Clark

A route to aircraft-like reusability for rocket engines.

noreply@blogger.com (Robert Clark) — Thu, 18 Apr 2024 17:04:00 +0000

A general fact about aircraft jet engines may offer a route to achieve aircraft-like reusability for rockets:

https://www.airliners.net/forum/viewtopic.php?t=739359

The key question: does this fact about jet engines also hold for rocket engines?

If so, increasing a turbopump rocket engine power just 10% to 15% cuts engine life in half. And conversely, decreasing it by 10% to 15% doubles engine life. And if so, would this still work if we repeated the concept multiple times? If we reduced the thrust by .9^5 = .60, i.e., to 60%, which most turbopump engines can manage, then we could increase the lifetime by a factor of 2^5 = 32 times? Then a Merlin engine with a lifetime of, say, 30 reuses by running it only 60% power could have its lifetime extended to 1,000 reuses?

This would be in the range of number of reuses of the type of jet engines used on long haul flights. Rocket engines with that number of reuses probably also would allow “gas and go” operation. That is, no major refurbishment needed in between flights, as with jet engines.

Is reduced temperature the key?

In examining this question of rocket engine longevity versus jet engine longevity I once hypothesized it had to do with the high temperatures rocket engines operated at, typically ca. 3,000 °C, whereas jet engines might only operate at ca. 1,200 °C to 1,500 °C.

It might be thought it would be the high pressures of rocket engines but that can’t be the primary reason since automobile diesel engines can operate at hundreds of bars of pressure, above even that of rocket engines for many hours of service:

ToughSF

@ToughSf

The pump pressures in rockets are impressive, but let's not forget that the injection system in modern diesel engines operate at 2500 bar. They are also fast enough to accomplish up to eight separate fuel injections with each cylinder cycle.
Bosch CRS3-25.
https://youtu.be/T7o2hvoJE-Q

0:20

8:00 AM · Sep 12, 2020

However, temperature can not be the only factor involved either. The hydrogen turbopump on the Space Shuttle Main Engine only operates at 1,310 °F, 710 °C, and turbopumps for rocket engines are a primary limiting factor in rocket engine longevity.

Still, the SSME hydrogen turbopump is still a type of turbomachinery and the longevity extension mentioned above known to hold with turbomachinery in general will still apply to it: derating its power level by a small relatively small percentage will result in a quite large increase in longevity.

Note an intriguing possibility also exists. The increase in longevity is actually in regards to turbomachinery specific power, that is power per weight. Note for example diesel engines are quite ruggedly built. So rather than derating the engine to, say, 60% power we could ruggedize it, i.e., using thicker, heavier components, increasing weight by ~50% to achieve the same reduction of specific power to 60% while keeping the same power level.

The reason you might want to do it that way is you may want to maintain the same power, i.e, thrust level. Lowering the thrust level would increase gravity losses, thus reducing payload, especially for a first stage engine, where most gravity loss is incurred. On the other hand at least for first stages increasing the mass of the stage only results in a fraction of that increase in the first stage mass being loss in payload, typically only 5% to 10% of that increase being lost from payload capability.

On the other hand rebuilding the turbopumps and possibly the combustion chamber using heavier components would be an expensive rebuild. The other approach of just derating its operational power level simply means running the existing engine at the lower power.

Fast ramp up to full power as cause of shortened longevity.

There is another interesting fact in that passage I cited on jet engine longevity:

Once a turbine engine has been shut down, usually it needs to cool down before restarting, depending on power levels prior to shut down. Cooling down can be done at ground idle power setting. Turbine engines generally don't like to be shut down straight from take-off power. They also require warming up before slamming to take-off power.

Take a look at this video of the SSME startup sequence:

STS-134 - The final launch of Endeavour - Full Launch in HD.

At about the 9:52 point the video shows the thrust level of the SSME goes from 0% thrust at T minus 7 seconds to 100% thrust at T minus 2 seconds, i.e., in only 5 seconds.

For jet engines, it is known such rapid changes are detrimental to engine longevity. It is quite likely the same is true for rocket engines. But for turbopump rocket engines you need almost immediate ramp up to the safe 60% to 100% throttle range.

But pressure-fed engines are usually more forgiven in this regard. They have a much wider range of throttleability levels. In that case you could have more gradual ramp up to full power, full thrust, or whatever you set the operational thrust level now to be.

Then it would be helpful if the turbopump engine could be operated in a pressure-fed mode, both in start-up and cool-down. This is known to be possible for some gas-generator engines which have an "idle mode". This is when the turbopumps are bypassed and the engine runs just on the pressure on the propellant in the tanks, i.e., a pressure-fed mode.

It's not known if staged combustion engines such as the SSME, RD-180 or Raptor can be run in idle mode. Given the potential of having such high performance engines have reuse lifetimes measuring in the thousands, I think it is a possibility that should be explored. A couple of possible ways need to be explored as possible ways of accomplishing it: one, the propellant could be allowed to flow through the turbopumps to the engine but the turbos being locked so they don't operate. And two, different piping being installed so that the turbos could be bypassed entirely if desired and the propellant go directly to the engines.

Another possibility is that there are turbopump engines that can run at highly reduced throttle level, to below 10% throttle for example. One such is the Common Extensible Cryogenic Engine(CECE), variant of the famous RL10 engine. These engines operate on a lower power engine cycle however known as the expander-cycle. Rather than using preburners to run their turbopumps as do gas-generator and staged-combustion engines, they run their turbos on the heat produced from the hydrogen run through the engine cooling channels. It may be possible to get higher power gas-generator and staged-combustion engines to run in a low power mode by adding additional piping to bypass the preburners, sending the hydrogen from the cooling channel directly to the turbopumps.

Temperature variations as cause of longevity reduction.
I'm informed by someone in the industry of another cause reducing rocket engine longevity: temperature variations across the combustion chamber causing differential thermal expansion.

This is primarily because of the high temperature in the combustion chamber at thousands of degrees compared to the low temperatures in the cooling channels at hundreds of degrees. If this is a problem of steady-state operation, i.e., not an issue of temperature ramp up speed, then it would seem to be the only solution would be operating the engine at lower temperature.

I mentioned that typically rocket engines operate ca. 3,000 °C, whereas jet engines might only operate at ca. 1,200 °C to 1,500 °C. This can not be the only issue with rocket engine longevity as we saw as the relatively low temperatures turbopumps operate at shows. Still, it could be an issue with combustion chambers.

There is a way to run rocket engines at the lower temperatures for most of the time of the flight. For rocket engines at near vacuum which is most of the flight, to get high performance, high Isp, what's really important is high expansion ratio. You don't really need high chamber pressure and the associated high temperature then.

Then if you had an adaptive nozzle, i.e., altitude compensation, that could extend to high expansion ratio at high altitude, near vacuum, you could get high Isp even using relatively low pressure and low combustion temperature in the near vacuum regime. This though would result in reduced thrust.

At low altitude soon after takeoff though you do need the high thrust. You might then run the engines at high pressure, high temperature then. The engine lifetime would still be extended longer than before in this being a relatively short portion of the engine firing time.

Alternatively, there might be a way to run the engines even at low altitude at low pressure, low temperature. It's related to the concept of altitude compensation. It's a method called the "thrust augmented nozzle(TAN)". It's quite analogous to afterburners for jet engines.

In TAN, additional propellant is injected into the nozzle and combusted there. This allows relatively low chamber pressures to operate in the engine while at the same time having a high expansion ratio even at sea level, and providing higher thrust.

Is SpaceX using the reduced throttle effect on longevity to improve Raptor reliability?I have argued that multiple lines of evidence suggest SpaceX throttled down the Raptor on the second and third test flights to improve reliability:

Did SpaceX throttle down the booster engines on the IFT-2 test launch to prevent engine failures?

https://exoscientist.blogspot.com/2023/12/did-spacex-throttle-down-booster.html

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

SpaceX should withdraw its application for the Starship as an Artemis lunar lander, Page 3: Starship has radically reduced capability than promised.

https://exoscientist.blogspot.com/2024/04/spacex-should-withdraw-its-application.html

Elon Musk gave the current version of the Superheavy/Starship launcher in his recent update to the Starship program as only having a payload capability of 40 to 50 tons to orbit. The problem is going by the specifications SpaceX has given for the dry mass and propellant mass of both stages and for the Raptor thrust and Isp values, the vehicle should easily be able to reach 100+ tons to orbit even as reusable. Then either the dry mass values or Raptor values or more likely both must be significantly worse than the cited values.

If so, then SpaceX is missing a golden opportunity to develop aircraft-like reuse capability for the Raptor and for the Merlin. It would require acknowledging however they are running the Raptor at reduced power to improve reliability, and in effect, longevity.

Bob Clark

SpaceX should withdraw its application for the Starship as an Artemis lunar lander, Page 3: Starship has radically reduced capability than promised.

noreply@blogger.com (Robert Clark) — Thu, 11 Apr 2024 16:34:00 +0000

Elon Musk presented an update on the plans for the Starship post the third Starship test flight:

Elon Musk Starship Presentation: IFT-4 Master Plan, Starship V2 & V3, Raptor V3, Mars, IFT-3 & More.

About 31 minutes in Elon suggests the current version V1 would be capable of 40 to 50 tons to orbit. This is bad because SpaceX sold NASA on the idea the Starship HLS could serve as an Artemis lander based on 150 tons to orbit reusable and “10ish” refueling flights. If the capability is max 50 tons, then it would take “30ish” refueling flights.

If they intend to use version V2 then this is bad because it would require further qualification flights for the larger version and more importantly further qualification of the more powerful Raptor 3 engine needed.

This last is doubly bad because I’d be willing to bet dollars to donuts that they never informed NASA that the current version couldn’t do it and further development would be required for the larger version.

SpaceX needs a true Chief Engineer. Elon once said that early on when there were still doubts about its viability, they tried to recruit a Chief Engineer for SpaceX but no one good was willing to come. So Elon designated himself Chief Engineer. It is not a role Elon is well suited for. A good Chief Engineer should be scrupulously forthright. He would not refer to the little 5 or 10 second static burns SpaceX does for the SuperHeavy or Starship as "full duration".

A true Chief Engineer would be aware that "full duration" in the industry is short for "full mission duration". These static fires in the industry are conducted at the full length and the full thrust of an actual flight and are meant to give confidence to potential customers that the engines can perform as expected for the promised capabilities of the launchers.

However, SpaceX in using the term "full duration" for these little few-second burns, doesn't even tell the public, or its major customer NASA for which they have a billion-dollar contract, if these little burns are even conducted at full thrust.

This has had majorly negative consequences. The FAA had great concerns in the Raptor reliability after the first test flight. In the "corrective actions" they required of SpaceX prior to a second Starship test flight, at the top of the list was correcting the tendency of the Raptor of leaking fuel and catching on fire while in flight.

I have argued multiple, independent lines of evidence suggest SpaceX intentionally reduced the throttle of the Raptors on the booster on the second test flight, IFT-2, to improve reliability of the engines:

Did SpaceX throttle down the booster engines on the IFT-2 test launch to prevent engine failures?https://exoscientist.blogspot.com/2023/12/did-spacex-throttle-down-booster.html

Running an engine at reduced throttle reduces the pressure levels within the engine, high pressure being a major cause of engine fuel leaks. The Starship upper stage though was run at near full throttle on IFT-2, perhaps because performance would be reduced too much if it also was run at reduced throttle.

The result was the booster engines worked fine, at least on ascent, while the Starship exploded on ascent on IFT-2. SpaceX has said the Starship RUD was due to an intentional LOX dump they performed to keep that flight as suborbital. However, many knowledgeable observers doubted the LOX dump alone would have caused a RUD. They argue due to the tendency of the Raptor to leak fuel, it's more likely that plus the LOX dump caused the RUD.

For the third test flight, IFT-3, after reviewing both propellant burn rates and the acceleration profile of the flight, I'm suggesting SpaceX learned their lesson from the second test flight, and this time both stages were run at reduced throttle on this flight. And this time both stages were able to complete the ascent stage of the flight successfully.

However, this does reduce the payload capability of the launcher. Elon has acknowledged this radically reduced payload capability in his recent update. But it needs to be explained by SpaceX why the payload is so greatly reduced. If it is because the Raptor needs to be run at reduced thrust in order to be reliable then that is an extremely important thing to acknowledge, and to inform NASA on it, because the thrust levels of a rocket go into assessing what its actual capabilities are.

There is another very important issue about Raptor reliability. Multiple times a Raptor has undergone a RUD doing a relight during prior testing of the Starship planned landing procedure. And on this last Superheavy/Starship test flight as well Raptors underwent a RUD during the booster landing procedure. The boostback back burn appeared to have occurred successfully. But there was venting gas after the bostback back burn suggesting there may have been a fuel leak here as well.

Note for a successful reuse of the Starship and booster, successful relights have to occur both for boostback burns and landing burns. Then in none of the prior Starship landing tests nor of the Superheavy/Starship flight tests have any flights shown successful Raptor relights without leaking fuel and catching fire, and often undergoing a RUD.

SpaceX has called one test of the Starship landing test, SN15, successful because it managed to land without exploding. But it is important to note even in that test a Raptor leaked fuel and caught fire prior to landing. It's just on that test SpaceX managed to extinguish the fire before the ship exploded:

Note that in the SpaceX plans for a reusable Starship it absolutely can not work if the Raptor can not be made to relight reliably. SpaceX in not publicly providing full mission duration, full thrust testing information on the Raptors have not shown this also for relights of the Raptor.

That is why it is so important for a launch company to publicly provide details on full mission duration, full thrust level static engine testing.

SpaceX needs a true Chief Engineer to provide such details in a forthright manner.

Robert Clark

SpaceX should explore a weight-optimized, expendable Starship upper stage.

noreply@blogger.com (Robert Clark) — Sun, 03 Mar 2024 17:56:00 +0000

To me it’s just stunning SpaceX is ignoring that an expendable Starship could be done for 40 ton dry mass, choosing instead the current 120 tons for the reusable version:

Elon Musk

@elonmusk

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

114

1.4K

https://twitter.com/elonmusk/status/1111798912141017089

Keep in mind that every kilo of extra mass in an upper stage subtracts directly from the payload possible. Then that 80 tons difference in the dry mass between the reusable and expendable versions is a huge difference.

Now, note because of size, that, just like with the Falcon 9, the 1st stage is 2/3rd of the cost. So for ~$90 million total for the SuperHeavy/StarShip, the SuperHeavy is $60 million of that. But as the Falcon 9 shows it is much easier to get reusable 1st stage. So assume with reuse of SuperHeavy, its cost, is now, say, $5 million per launch. Now it’s a $35 million total cost for the partially reusable SuperHeavy/StarShip. BUT now because of the radically reduced upper stage dry mass, we have ca. 300 tons payload this version!(Assume SuperHeavy lands down range if you wish to maintain the high payload.) But this is about the same cost per kilo as fully reusable 100 to 150 ton payload fully reusable version at $10 million per flight cost.

Then the question is how realistic is it the Starship could have 40 ton dry mass as an expendable? I think it is quite realistic.

Consider the original Atlas rocket first used to send John Glenn to orbit:

SLV-3 Atlas / Agena B.

Family: Atlas. Country: USA. Status: Hardware. Department of Defence Designation: SLV-3.

Standardized Atlas booster with Agena B upper stage.

Specifications

Payload: 600 kg. to a: 19,500 x 103,000 km orbit at 77.5 deg

inclination trajectory.

Stage Number: 0. 1 x Atlas MA-3 Gross Mass: 3,174 kg. Empty Mass:

3,174 kg. Thrust (vac): 167,740 kgf. Isp: 290 sec. Burn time: 120 sec.

Isp(sl): 256 sec. Diameter: 4.9 m. Span: 4.9 m. Length: 0.0 m.

Propellants: Lox/Kerosene No Engines: 2. LR-89-5

Stage Number: 1. 1 x Atlas Agena SLV-3 Gross Mass: 117,026 kg.

Empty Mass: 2,326 kg. Thrust (vac): 39,400 kgf. Isp: 316 sec. Burn time: 265 sec. Isp(sl): 220 sec. Diameter: 3.1 m. Span: 4.9 m. Length:

20.7 m. Propellants: Lox/Kerosene No Engines: 1. LR-105-5

Stage Number: 2. 1 x Agena B Gross Mass: 7,167 kg. Empty Mass: 867 kg. Thrust (vac): 7,257 kgf. Isp: 285 sec. Burn time: 240 sec. Isp(sl): 0 sec. Diameter: 1.5 m. Span: 1.5 m. Length: 7.1 m. Propellants: Nitric

acid/UDMH No Engines: 1. Bell 8081

http://www.friends-partners.org/partners/mwade/lvs/slvgenab.htm

The Atlas had an unusual design however. It dropped its main lift-off engine at altitude and continued on with what was called the “sustainer” engine. This engine due to much of the propellant mass being burned off had much lower thrust, and so much reduced required engine weight. Then looking at the specifications of this stage, note it had nearly a 50 to 1 mass ratio(!)

The comparison of this sustainer stage to the 3-engine Starship upper stage is appropriate since an upper stage typically doesn’t need to have the thrust of a stage needing to lift off from the ground. Weight growth of the Starship now at 120 tons dry mass required adding 3 additional engines, to now have 6 engines.

However, a key reason why the Atlas was able to achieve such a high mass ratio was that it used what was called “balloon-tank” design. This was a design that used pressurization to maintain its structure even on the ground. It would actually collapse under its own weight when not pressurized.

However, methanolox is at about 80% of the density of kerolox. So a corresponding methanolox version would be at 40 to 1 mass-ratio, better than the 30 to 1 mass ratio Elon suggested. But its not likely SpaceX would want to deal with the operational difficulties of having a stage be continually pressurized even when on the ground, unfueled, especially for a stage intended to have high launch rates.

So I’ll look at another stage, the S-II hydrolox 2nd stage of the Saturn V rocket. The Saturn V launcher of the Apollo program was remarkable in the lightweight features of its upper stages, the S-II and the S-IVB. This page gives a list of the fueled weights and empty weights of the Saturn V stages:

Ground Ignition Weights

http://history.nasa.gov/SP-4029/Apollo_18-19_Ground_Ignition_Weights.htm

The later versions of Apollo had improved weight optimization. We'll use the specifications for Apollo 14. The "Ground Ignition Weights" page gives the Apollo 14 S-II dry weight as 78,120 lbs., 35,510 kg, and gross weight as 1,075,887 lbs., 489,040 kg, for a propellant mass of 997,767 lbs., 453,530 kg, resulting in a mass ratio of 13.77 to 1.

Now, methanolox is 2.5 times greater density than hydrolox. Then the corresponding mass ratio for methanolox would be at 33 to 1. This comparison is particularly apt because the mass in the same size tanks would be approx. at the 1,200 propellant mass of the Starship.

So Starship could reach ca. 30 to 1 mass ratio when using the weight optimizing methods used during the Apollo program.

But if the price per kilo of this partially reusable version would be at about what the current version is what is the advantage? One advantage is as mentioned is you would not have the difficulty of making the upper stage reusable, no problematical heat shield tiles.

There is another advantage not as concrete, but in my mind just as important if not more so. In my opinion the approach SpaceX is taking with the SuperHeavy/Starship is ill-conceived. It is based on the idea the SuperHeavy/Starship should be the be-all-end-all for ALL of spaceflight.

But if you look at transport methods throughout history even going back to the horse-drawn era transports always came in different sizes. A comparison to the air traffic is most instructive. It turns our the largest air transports the jumbo-jet size aircraft actually make up a tiny percentage of air traffic. The great bulk of air traffic is carried by smaller aircraft.

And even looking at SpaceX’s own Falcon Heavy demonstrates this. The per kilo cost is less than that of the Falcon 9. But the number of Falcon Heavy flights is tiny compared to the number of Falcon 9 flights.

The fixation on the reusable Starship as the be-all-end-all for all spaceflight also leads to the poorly-conceived notion that a Mars or Moon mission must be carried out by multiple refuelings of the reusable Starship. The number of refueling flights for the Artemis lunar missions might be 8 to 16 flights.

But it is a basic principle of orbital mechanics that high delta-v missions such as to the Moon or Mars are more efficiently carried out by using additional stages. Simply by giving the SuperHeavy/Starship an additional 3rd stage, flights to both the Moon and to Mars could be carried out in a single launch.

An expendable Starship would mean it being regarded as just another stage. And a 3rd stage could be set atop it as needed, such as for high delta-v missions.

As another illustration of the fact this approach to the SuperHeavy/Starship is ill-conceived, the payload of the SH/ST to GEO is nearly zero because that Starship dry mass is so high. This is the most lucrative satellite market, but a single SH/ST launch could not service that market. In order to just launch satellites to GEO the SH/ST would have to do multiple refuelings just to launch a satellite to GEO, just like when it had to launch manned interplanetary missions. This is an odd state of affairs for a rocket simply to launch satellites to GEO.

Or of course it could utilize a 3rd stage. But if you are going to use a third stage then, why not just use it also for the manned interplanetary missions that would allow you to do such missions in a single flight?

Robert Clark

Altitude compensation is more efficient than staged-combustion engines.

noreply@blogger.com (Robert Clark) — Thu, 29 Feb 2024 21:43:00 +0000

Staged combustion engines such as the Russian RD-180 and American SSME are regarded as the utmost in efficiency because they achieve high vacuum Isp while being able to achieve high thrust at sea level. They achieve this by operating at high chamber combustion pressure.

SpaceX is developing the Raptor engine also as a staged combustion engine. However, a surprising fact is a medium performance, mid-level pressure and cheaper engine such as the Ariane 5’s Vulcain engine or Delta IV’s RS-68 can get higher performance than a staged combustion engine by using altitude compensation.

See the graphic of the Isp of the Vulcain engine with an altitude compensating nozzle:

The vacuum Isp of the SSME is 452.3 seconds (4,436 m/s), and the sea level, 366 seconds (3,590 m/s). You see from the graphic with altitude compensation the Vulcain sea level Isp would be ca. 3,850 m/s. And already at ca. 20,000 m, its altitude compensating Isp would match that of the SSME. An thereafter the the Isp would exceed the maximum vacuum Isp of the SSME. Indeed its Isp could reach 4,850 m/s, and above

That the Isp with adaptive nozzles can be this high is supported by calculations for hydrogen/oxygen engines at ultra large expansion ratios. This report concludes at a 600 to 1 expansion ratio we can get ca. 480 s vacuum Isp:

ORBITAL TRANSFER VEHICLE (OTV) ENGINE STUDY, PHASE A - EXTENSION
CONTRACT NO. NAS8-32996

https://ntrs.nasa.gov/api/citations/19800015867/downloads/19800015867.pdf

The method to get the altitude compensation does not have to be the aerospike nozzle. Better actually would be to add an altitude compensation nozzle extension to an existing engine such as the Vulcain or RS-68. Redesigning such an engine to use an annular combustion chamber for an aerospike nozzle would be expensive. Far cheaper would be to use an altitude compensating nozzle attachment to the already existing engine.

Such nozzle extensions already have been in existence for decades on upper stage engines, such as the extendable nozzles on for example the RL-10B2 engine.

But in actually the increase in efficiency would be higher for a first stage engine. For instance the vacuum Isp for the Vulcain or RS-68 could be increased from 432s or 412s to 480+ s and above.

The nozzle extension is just a well-known, simple way to accomplish it but there may be simpler or more lightweight methods of accomplishing it.

Aerospike in 3D exhaust injection. UPDATED, 1/10/2023: Extension to single nozzles.
https://exoscientist.blogspot.com/2023/01/aerospike-in-3d-exhaust-injection.html

SSME based SSTO’s. UPDATED, 6/28/2021 - Extension to the Delta IV Heavy.
https://exoscientist.blogspot.com/2021/06/ssme-based-sstos.html

ESA's Callisto reusability testbed as an *operational* TSTO and SSTO. UPDATE, 7/1/2019.
https://exoscientist.blogspot.com/2019/05/esas-callisto-reusability-testbed-as.html

Altitude compensation attachments for standard rocket engines, and applications, Page 6: space shuttle tiles and other ceramics for nozzles. UPDATED: 3/6/2018
https://exoscientist.blogspot.com/2017/12/altitude-compensation-attachments-for.html

Altitude compensation attachments for standard rocket engines, and applications, Page 5: metal foil expandable nozzles.
https://exoscientist.blogspot.com/2017/08/altitude-compensation-attachments-for.html

Altitude compensation attachments for standard rocket engines, and applications, Page 4: the double aerospike.
https://exoscientist.blogspot.com/2016/10/altitude-compensation-attachments-for.html

Altitude compensation attachments for standard rocket engines, and applications, Page 3: stretchable metal nozzles.
https://exoscientist.blogspot.com/2016/06/altitude-compensation-attachments-for.html

Altitude compensation attachments for standard rocket engines, and applications, Page 2: impulse pressurization methods.
https://exoscientist.blogspot.com/2016/01/altitude-compensation-attachments-for.html

Altitude compensation attachments for standard rocket engines, and applications.
https://exoscientist.blogspot.com/2014/10/altitude-compensation-attachments-for.html

The Coming SSTO's.
https://exoscientist.blogspot.com/2012/05/coming-sstos.html

Altitude Compensation Improves Payload for All Launchers.
https://exoscientist.blogspot.com/2016/01/altitude-compensation-improves-payload.html

Robert Clark

Could meteor impacts be the cause of the coronal heating problem?

noreply@blogger.com (Robert Clark) — Thu, 22 Feb 2024 18:21:00 +0000

A puzzle in solar science that has existed for 150 years is the corona heating problem:

Why is the sun’s corona 200 times hotter than its surface?
The paradox has astronomers scratching their heads over magnetic waves, nanoflares, and the now-debunked element coronium.
BY BRILEY LEWIS | PUBLISHED APR 12, 2023 6:00 AM EDT
https://www.popsci.com/science/how-hot-is-the-suns-surface-corona/

The Sun's surface is at about 10,000 F, 5,500 C. But the solar corona reaches millions of degrees. How is it possible to get so much hotter hundreds of thousands kilometers away from the Suns surface?

Noted solar astronomer Eugene Parker for whom the Parker Solar probe was named suggested it was due to nanoflares small flares emanating from the solar surface much smaller than the usual solar flares:

ScienceCasts: The Mystery of Nanoflares.

But what causes the nanoflares? Could it be asteroidal impacts? The argument could be made they are too small to cause any visible reaction on the Sun. But the question is of the local impact. The Sun’s escape velocity at its surface is 600 km/s. That is a tremendous amount of energy for a body impacting it at that speed. When material is thrown up after the impact the high temperature could be maintained far above the surface.

Nanoflares and coronal heating.

Micro-flare observed on 4 September 2016 with NASA SDO/AIA and the Swedish 1-m Solar Telescope.

The image shows a micro-flare observed on 4 September 2016. Magnetic reconnection in the corona as sketched in the cartoon in the lower left produces a hot loop of more than 7 million degrees. This hot loop is visible as the bright area in the green background image taken with the Solar Dynamics Observatory (AIA 94 Å). The active region with bright magnetic loops is shown in more detail in the yellow inset, corresponding to plasma of less than 1 million degrees (AIA 171 Å). The reconnection event in the corona produces fast electrons that hit the lower atmosphere with high energy. The impact region is very small and is shown at high resolution in the image taken with the Swedish 1-m Solar Telescope on La Palma. With the European Solar Telescope, we will be able to study the magnetic environment of the impact region in even finer detail.

https://est-east.eu/?option=com_content&view=article&id=920&Itemid=622&lang=en

For instance Jupiter’s escape velocity is 60 km/s and we saw the tremendous resulting impact from comet Shoemaker-Levy when it impacted Jupiter.

Jupiter in infrared, Shoemaker-Levy 9 collision (left) and Io (right) by Max Planck Institute for Astronomy

But the major, key reason for suspecting it is this: there is a type of nuclear fusion called impact fusion. It arises when bodies are made to collide at hundreds of kilometers per second relative impact speed.

Proceedings of the
Impact Fusion Workshop ~ National Security and Resources Study Center
LOS Alamos Scientific Laboratory Los Alamos, New Mexico
LOSALAMOS SCIENTI
LABORATORY
PostOfficeBox 1663 LosAlamos,New Mexico87545
July 10—12, 1979

https://sgp.fas.org/othergov/doe/lanl/lib-www/la-pubs/00203259.pdf

There are private fusion research concerns now investigating this to bring about controlled nuclear fusion.

Recent observations of nanoflares have observed million degree temperatures locally around the nanoflares origin point on the Sun’s surface, while the surrounding area is at the normal 5,500 C temperature.

So why don’t we see the asteroids during imaging of the nanoflares? It could be their small size. The Sun is so bright it completely washes out the asteroids that may be only a few kilometers across.

Recent observations and theoretical modeling suggest the million degree temperatures seen in the vicinity of the nanoflare origin point on the Sun’s surface should be able to be communicated to the corona-sphere thousands of kilometers above the Suns surface:

This May Be the First Complete Observation of a Nanoflare.
Heating the corona.
So far, these bright loops appeared to be tiny flares – but did their heat actually reach the corona?Bahauddin looked to NASA’s Solar Dynamics Observatory, which carries telescopes tuned to see the extremely hot plasma only found in the corona. Bahauddin located the regions right above the brightenings shortly after they appeared. “And there it was, just a 20-second delay,” Bahauddin said. “We saw the brightening, and then we suddenly saw the corona got super-heated to multi-million degree temperatures,” Bahauddin said. “SDO gave us this important information: Yes, this is indeed increasing the temperature, transferring energy to the corona.” Bahauddin documented 10 instances of bright loops with similar effects on the corona. Still, he hesitates to call them nanoflares. “Nobody actually knows because nobody has seen it before,” Bahauddin said. “It’s an educated guess, let’s say.”From the perspective of the theory that says nanoflares heat the corona, the only thing left to do is to show that these brightenings occur often enough, all over the Sun, to account for the corona’s extreme heat. That’s still work in progress. But observing these tiny bursts as they heat solar atmosphere is a compelling start.
https://www.nasa.gov/solar-system/this-may-be-the-first-complete-observation-of-a-nanoflare/

Additionally I was startled see to what would be the kinetic energy of an asteroid impacting the Sun at the 600 km/s escape velocity. Asteroids have been estimated to have densities in the range of 2,000 kg/m3 to 5,000 kg/m3 . The iron-nickel asteroids would have the higher density. This is important because they could also maintain their cohesiveness as they impacted the Sun.

Searches of a population of asteroids inside the orbit of Mercury, called vulcanoids, have been unsuccessful. This is because you have to look at the bright solar disk to detect them. But such searches put a size limit of 6 km wide on them. So assume the asteroid has size, say, 5 km across, with density, say, 4,000 kg/m3 . At that density the mass would be 4,000 kg/m^3 * (6,000 m)^3 = 8.64 * 10^14 kg. Now suppose this impacted the Sun at 600 km/s. Then the kinetic energy of that pact would be:

(1/2) * 8.64 * 10^14 * (600,000 m)^2 = 1.55*10^26 Joules. That is a tremendous amount of energy! To put in perspective the energy the Sun puts out each second is 3.86 * 10^26 watts. So if the asteroid deposited that energy in, say, 1 second, it would be a significant percentage of the total energy the Sun puts out in a second!

However, asteroids of kilometers size impacting the Sun must be quite rare, judging from this graph of asteroid impacts to Earth by size:

https://www2.tulane.edu/~sanelson/Natural_Disasters/impacts.htm

One meter and below must be more common. If the meteor impacting the Sun was 1 meter wide, then the kinetic energy would be (1/2)*4,000*(1)^3*(600000)^2 = 7.2*10^14 joules, nearly a quadrillion joules of energy.

About the likelihood of asteroids impacting the Sun in accordance to the change needed in their established orbital velocity, if they started further out in the Solar System, much less velocity change (delta-v) would be needed to direct them to impact the Sun.

Key confirmation required is to confirm the existence of these small solar impactors. Observations in the visual light spectrum have not succeeded. This is the solar irradiance spectrum showing the range of intensity’s according to wavelength:

You see it is vanishingly small at extreme ultraviolet wavelengths and at radio wavelengths around 10,000 nm, 10 microns, and above. The problem with observations at the extreme ultraviolet is that not large enough telescopes have been launched to observe them at less than 6 km diameters (the extreme UV is absorbed by the Earth’s atmosphere.)

Then the suggestion is to use large radio telescopes at the micron and above wavelengths to detect the close in asteroids.

One radio telescope that might manage it is the ALMA radio telescope array:

ALMA Demonstrates Highest Resolution Yet

November 15, 2023 | Science

The Band-to-band (B2B) method demonstrated this time to achieve the highest resolution with ALMA. In the B2B method, atmospheric fluctuations are compensated for by observing a nearby calibrator in low frequency radio waves, while the target is observed with high frequency radio waves. The top right inset image shows the ALMA image of R Leporis that achieved the highest resolution of 5 milli-arcsec. Submillimeter-wave emissions from the stellar surface are shown in orange and hydrogen cyanide maser emissions at 891 GHz are shown in blue. The top left inset image shows a previous observation of the same star using a different array configuration with less distance between the antennas and without the B2B method, resulting in a resolution of 75 milli-arcsec. The previous resolution is too coarse to specify the positions of each of the two emission components. (Credit: ALMA (ESO/NAOJ/NRAO), Y. Asaki et al.) Download image (1.3MB)

https://www.nao.ac.jp/en/news/science/2023/20231115-alma.html

At a max resolution of 5 milli-arc it should be able to detect kilometer wide asteroids at the distance of the Sun. The detection sensitivity should also be improved for iron-nickel meteorites for radio astronomy.

Note the importance of this is that if it is confirmed then we know impact fusion does indeed work.

Bob Clark

Alternative explanations for the CMB, universe expansion, and dark matter.

noreply@blogger.com (Robert Clark) — Thu, 08 Feb 2024 21:02:00 +0000

James Webb was promised to provide revolutionary results in cosmology and has not disappointed. Several observations have shown well-developed galaxies that stem from the earliest time after the Big Bang, which current theories suggest should not be possible.

JWST finds very early galaxies that should not exist.

The observations have led some scientists to question the accuracy of the current models for the beginning of the universe. Further, the cosmological microwave background(CMB) had been regarded as strong confirmation of the Big Bang theory for the origin of the universe. But there is a discrepancy between the rate of expansion of the universe based on the CMB and measurements of galactic motion.

This discrepancy has existed for several years now, but it was hoped with better instruments the discrepancy would be found due to measurement error. Instead, the JWST has provided further evidence the discrepancy is real. Then either the CMB estimate or the interpretation of the redshift measurements or both are wrong.

Here I'll discuss the possibility there is a problem with the interpretation of the origin of the CMB. It seems to me there should be some contribution to the CMB due to highly red shifted infrared and optical radiation from galaxies at high red shifts, but I never see this mentioned. The CMB is only described by relic radiation of the intense heat at the beginning of time that gradually cooled as the universe expanded.

An argument can be made that the CMB is seen in all directions but there are blank areas in some part of the sky. This does not support the idea of the CMB deriving from redshifted light from primordial galaxies.

But the Hubble Deep Field showed abundant galaxies in areas previous thought to be devoid of galaxies. It was a revolutionary advance in our knowledge of the extent of the universe. Hubble deep field images integration times ranged from 10 days to 23 days. But the JWST deep field image only went for a day:

STARTS WITH A BANG — APRIL 17, 2023
JWST surpasses, enhances Hubble’s deepest image ever
With infrared capabilities and image sharpness far beyond Hubble's limits, JWST looked at Hubble's deepest field, revealing so much more.
https://bigthink.com/starts-with-a-bang/jwst-surpasses-hubbles-deepest-image/

In view of the startling find of fully formed galaxies going back to near the time of the Big Bang by the JWST, such long integration times as for the Hubble must also be done for the JWST.

The estimate of the number of galaxies in the universe from the Hubble deep field was 170 billion. But numerical simulations put it at perhaps one hundred times more at 6 to 20 trillion galaxies:

STARTS WITH A BANG — JUNE 22, 2022
There are more galaxies in the Universe than even Carl Sagan ever imagined
Forget billions and billions. When it comes to the number of galaxies in the Universe, both theorists' and observers' estimates are too low.
https://bigthink.com/starts-with-a-bang/galaxies-in-universe/

By doing the longer integration times JWST may be able to confirm this larger number of galaxies. Such a large number of galaxies going back to near the time of the Big Bang may allow the CMB to be equally well explained by highly redshifted light, infrared and optical, from these earliest galaxies.

However, there may be another even greater contributor to the observed CMB. A little known fact is that for most of the galaxies in the universe they are receding from us faster than the speed of light(!)

97% of Galaxies Are Moving Faster Than Light, HOW IS THIS POSSIBLE?

This is explained as not being in conflict with relativity by the virtue of the fact that space itself is expanding. It is not the case that objects are moving through space as these superluminal speeds.

Nevertheless, this raises an interesting possibility. If it is the case that these galaxies are moving away from us at these apparent superluminal speeds, would we observe a luminal "boom" from these galaxies when they appear to cross the light-speed barrier relative to us?

The luminal boom is a concept that is analogous to the sonic boom for sound waves. This is actually seen for some subatomic particles traveling though matter, where the speed of light is reduced below that of the vacuum speed. In cases where the particles exceed that materials light speed, a phenomenon known as Cerenkov radiation is observed. Note that the particles are still not traveling faster than the vacuum speed of light, only the speed of light in the material. So relativity is still upheld.

An analogous phenomenon is seen in cosmic ultra high energy gamma ray bursts, GRB's:

Faster-Than-Light Speeds Could Be Why Gamma-Ray Bursts Seem to Go Backwards in Time.
SPACE
30 September 2019
By MICHELLE STARR

https://www.sciencealert.com/faster-than-light-speed-in-jets-that-produce-gamma-ray-bursts

This blog post contained the discussion of an alternative explanation for the CMB. Follow up posts will discuss alternative explanations for universe expansion and dark matter.

Robert Clark

Towards Every European Country's Own Crewed Spaceflight, Page 2: saved costs and time using already developed, operational engines.

noreply@blogger.com (Robert Clark) — Thu, 25 Jan 2024 22:07:00 +0000

Vulcain-based launchers.

ESA head Josef Aschbacher made the remarkable statement that the Ariane 6 can not be guaranteed to be the launcher of choice in the European launch market:

“We are worried,” says European rocket chief at prospect of launch competition
On the continent, Ariane 6 may be the last launcher with a monopoly.
PEGGY HOLLINGER AND SYLVIA PFEIFER, FT - 1/9/2024, 9:18 AM
https://arstechnica.com/space/2024/01/we-are-worried-says-european-rocket-chief-at-prospect-of-launch-competition/

In the blog post, "Towards Every European Country's Own Crewed Spaceflight", I suggested any European country could build their own manned spaceflight capable launcher by buying an Ariane 5 or 6, disposing of the side boosters, and adding 1 or 2 additional Vulcain engines to the core. ArianeSpace might raise a squawk however since it would be using their tech to build a direct competitor to the Ariane 6 and at a cheaper price in not using the large, expensive side boosters.

Another approach might be to design their own launcher designed around the Vulcain engine. The Vulcain engine developer Snecma, now Safran Aircraft Engines, is independent of ArianeSpace so likely the Vulcain could also be purchased from Safran. Purchasing an already developed and operational engine would save on costs since engine development is typically the biggest development cost for a new launcher. See for example this breakdown on the costs of the Ariane 5:

Development budget

Again, Ariane 5, from 'Europäische Tragerraketen, band 2', Bernd Leitenberger:

Studies and tests 125
solid boosters 355
H120 first stage 270
HM60 (Vulcain) engine and test stands 738
other elements of the first stage and boosters 95
upper stage and VEB 200
ground support in Europe 80
Buildings and other structures in Kourou (launch pad) 450
Test flights 185
Total 2498
ESA and CNES management 102

https://space.stackexchange.com/questions/17777/what-is-the-rough-breakdown-of-rocket-costs

For our scenario we would not be using solid rockets. So that rather large cost would be saved. Note also in our scenario using the already developed and fully operational Vulcain, the engine development costs and test stand costs would also be saved. For the Ariane 5, the ESA also built entire new launch facilities in Kourou, Guyana in equatorial Africa. For this new launcher we'll assume it will use the already constructed launch facilities at Kourou, or the country where the new launcher is being developed would construct an independent launch facility for their nascent space industry.

To sure, we'll assume this new launcher would be developed using the commercial space approach spearheaded by SpaceX. SpaceX demonstrated development costs could be cut by a factor of 10 following this approach:

Falcon 9.

1Development history

1.1Conception and funding

In 2011, SpaceX estimated that Falcon 9 v1.0 development costs were on the order of US$300 million.^[39] NASA estimated development costs of US$3.6 billion had a traditional cost-plus contract approach been used.^[40] A 2011 NASA report "estimated that it would have cost the agency about US$4 billion to develop a rocket like the Falcon 9 booster based upon NASA's traditional contracting processes" while "a more commercial development" approach might have allowed the agency to pay only US$1.7 billion".^[41]

https://en.wikipedia.org/wiki/Falcon_9#Conception_and_funding

Now, several companies world-wide have also shown that following the commercial space approach of using private financing can cut development costs by a factor of 10.

The total development cost of the Ariane 5 was $2.5 billion in 1990's dollars. Now take into account the costs that wouldn't need to be included, solid booster development, engine development, and launch facilities. This reduces the development cost to $955 million in 1990's dollars. Now consider by following the commercial space approach this could be cut by a factor of 10 to ca. $95 million, or about $200 million in 2024 dollars. Quite remarkable also in particular is the development of the core stage without engines could be done for only about $54 million.

So an approx. 10 ton payload capacity all-liquid launcher could be developed for approx. $200 million, by using already developed and operational engines. This launcher would have the advantages, by not using solid rocket boosters, of being capable of reusability and being made manned flight capable.

Quite surprising also is how quickly such a manned-flight capable launcher might be developed. ArianeSpace could develop it the most quickly, probably in less than a year. All it would have to do is acknowledge large solid side boosters are not price competitive. As I discussed previously, JAXA showed with its H-II rocket, an additional engine can be added to a core stage for less than $200 million. And SpaceX showed with its Raptor engine that additional Raptors can be added to a core stage on a time scale of just months, not years, even if a new thrust structure is required to accommodate the new engines.

But even for those countries making the new launcher from scratch quite surprisingly it could also be done quite rapidly, assuming it used an already developed and operational engine. A fact not generally appreciated is how rapidly SpaceX was able to develop the Falcon 9 rocket by using the already developed and operational Merlin engine. After the first successful flight of the Falcon 1 in 2008, SpaceX built and successfully launched the Falcon 9 in only two years in 2010. Note because the Falcon 9 had a larger diameter and used 9 engines instead of just one, SpaceX had to use completely different tooling in constructing the Falcon 9.

Then following the SpaceX example, and the SpaceX commercial space approach, a company could build and launch a 10-ton payload capable launcher in only 2 years by using already developed and operational engines.

Methane-fueled Prometheus-based launchers.

ESA has received much criticism in not keeping up with SpaceX on reusability. The Ariane 6 in fact won't be reusable and it is now acknowledged it won't be competitive to the SpaceX Falcon 9 in price, necessitating hundred million dollar subsidies yearly to stay afloat.

Recognizing the need for reusability in future launchers, ESA has begun the development of the methane-fueled, reusable Prometheus engine. And through its subsidiary Maiaspace, ArianeSpace is developing an all-liquid reusable launcher using the Prometheus engine for launch:

ArianeGroup to Increase MaiaSpace Investment to €125M

https://europeanspaceflight.com/arianegroup-to-increase-maiaspace-investment-to-e125m/

The MaiaSpace launcher will be capable of about 1,500 kg payload to LEO as an expendable rocket, using three Prometheus engines at ca. 100-ton thrust capability. It is expected to make its first launch in 2025.

It is illuminating to make a comparison to the early development of SpaceX. The Falcon 1 had an approx. 600 kg to LEO capability using a single ~100-ton thrust Merlin engine. It had its first successful launch in 2008. Remarkably just 2 years later in 2010, SpaceX had the 9 Merlin-engine Falcon 9 rocket make a successful launch at a ca. 10-ton payload to LEO capacity.

Then following the SpaceX example, MaiaSpace using 9 Prometheus engines could have a 10-ton to LEO capable launcher available in 2027. This could be man-rated to be manned flight capable.

Then going by the SpaceX example of the $300 million development cost of Falcon 9, and considering engine development cost makes up the bulk of launcher development cost, any European country using an already developed and operational Prometheus engine could have a 10-ton to LEO capable launcher at less than $150 million development cost following the commercial space approach.

And again following the SpaceX example such a launcher could be built and launched within 2 years.

Manned Space Capsules.

ESA has announced opening a competition among European companies for cargo capsules to deliver supplies to the ISS, with manned capsules to follow in development:

ESA to start commercial cargo program

Jeff Foust

November 6, 2023

https://spacenews.com/esa-to-start-commercial-cargo-program/

SpaceX and Orbital Sciences, now a subsidiary of Northrup Grumman, with their Dragon and Cygnus cargo capsules, showed space capsules like launchers also could be developed at costs 1/10 that of the usual government-financed ones following the private financing approach of commercial space.

Then I advise the European companies entering the competition follow the commercial space approach in developing their space capsules. They could accept seed funding from ESA to get started, but the bulk of the development costs should come from private funding. Note that winning these seed dollars from the ESA could be used as a selling point in acquiring the private funding.

According to the SpaceNews article the cargo capsules are expected to be ready by 2027 or 2028. It is notable that this is around the same time MaiaSpace might be able to have a 10-ton to LEO capable launcher ready.

Because of this I advise the cargo capsules and manned capsules be developed concurrently. It is my thesis that manned capsules can also be developed at costs in the few hundred million dollars cost range by following the commercial space approach as found with cargo space capsules.

It is notable in this regard that when SpaceX accepted NASA funding for the development of the manned version of the Dragon capsule, costs ballooned to the billion dollar range. I'm arguing the costs were that high because NASA was paying for it.

Robert Clark

Possibilities for a single launch architecture of the Artemis missions, Page 4: lightweight landers from NRHO to the lunar surface.

noreply@blogger.com (Robert Clark) — Tue, 23 Jan 2024 16:42:00 +0000

Congress is becoming increasingly concerned that with the continuing delays of the Artemis missions that China may beat the U.S. back to the Moon:

US must beat China back to the moon, Congress tells NASA.
By Mike Wall
'It's no secret that China has a goal to surpass the United States by 2045 as global leaders in space. We can't allow this to happen.'
https://www.space.com/us-win-moon-race-china-congress-artemis-hearing

I had previously proposed correcting an error in the design of Orion's service module that instead of making it larger than Apollo's service module because of Orion's twice larger size, instead made it 1/3rd smaller:

Possibilities for a single launch architecture of the Artemis missions, Page 2: using the Boeing Exploration Upper Stage.

https://exoscientist.blogspot.com/2023/08/possibilities-for-single-launch.html

The proposal was to give an additional approx. 10 tons propellant to the service module. This would allow the Orion capsule/service module stack plus an Apollo-size lander to be carried all the way to low lunar orbit, not just to NRHO(near-rectilinear halo orbit).

This though because of the higher payload may require use of the higher thrust J-2X engine on the Boeing EUS(upper stage) rather than the 4 RL-10 engines now planned on the SLS Block 1B. It's higher thrust would result in a greater payload to LEO and TLI, perhaps to ca. 120 tons to LEO rather than the 105 tons planned to LEO.

This approach requires additional propellant tanks be added to the service module and a change in the EUS upper stage engine to the J-2X. As I discussed in that blog post, it may also require an additional Centaur V sized third stage be added atop the Boeing EUS. This is dependent on what is the TLI(trans lunar injection) payload for the Boeing EUS using the J-2X engine. It may be it can perform the needed TLI payload without an additional Centaur V 3rd stage.

In any case, I'll propose here an alternative approach to a single launch Artemis architecture without increasing the service module propellant load. This again will use a light-weight Apollo-sized lander with all the components of Orion capsule/Service Module/lunar lander all carried on that one single SLS launch. Because of the lower propellant load on the service module though I'll also send it to NRHO instead of to low lunar orbit.

Note the NRHO was chosen by NASA as the orbital location because it has a lower delta-v requirement to get there than going to low lunar orbit. Here’s the the delta-v requirements:

The second group of delta-v’s shows the delta-v to NRHO as 0.45 km/s and the delta-v to and from the lunar surface from NRHO as 2.75 km/s, or 5.5 km/s round trip.

I’ve seen various numbers for the Orion and service module dry mass and propellant mass. I’ll use 16.5 total dry mass for the Orion+service module together, and 9 tons of service module propellant mass, but only 8.6 tons of this as usable propellant because of residuals.

Then we'll use 6 tons of Service module propellant to get the Orion/Service Module/lunar lander to NRHO after being placed on TLI trajectory by the EUS, for the 16.5 ton Orion/Service Module dry mass, and 15 tons gross mass Apollo-sized lander with 2.6 tons left over for the return trip.

We'll need every bit of performance to accomplish the mission within these constraints. So we'll assume we can get a 324 s Isp out of the storable propellant engines on the service module. This is higher than specified for the Orion service modules engines but is doable because of the storable propellant Aestus engine on the Ariane 5 EPS storable propellant upper stage which gets this vacuum Isp. We'll assume we can get this increased Isp by using a larger expansion ratio nozzle or even by swapping out the engine on the service module to use the Aestus engine. Then we get:

324*9.81Ln(1 + 6/(16.5 + 15 + 2.6 + 0.4)) = 510 m/s, or 0.51 km/s, sufficient for placing the stack in the NRHO orbit, where the 0.4 in the equation is for the unburnt residuals.

Then with the 2.6 tons usable propellant left over for the return trip, after the lander is jettisoned, we get:

324*9.81Ln(1 + 2.6/(16.5 + 0.4)) = 450 m/s, 0.45 km/s, sufficient for the Orion return.

To increase performance even more we may want to switch even to the RS-72 engine. This is a turbopump-fed storable propellant engine with a vacuum Isp of 340s. It achieves this by using a higher chamber pressure of 60 bar and higher nozzle expansion ratio of 300 to 1 than the Aestus engine. A turbopump engine also has lower residuals, typically less than 1%. A disadvantage is that pressure-fed engines are simpler with fewer moving parts, and so higher reliability, important for an engine to place the spacecraft in orbit and for leaving orbit.

Now for the ca. 15 ton gross mass lander, because of the higher delta- v needed from NRHO we’ll use hydrolox rather than storable propellant stage. The Ariane 4 H10 hydrolox upper stage had a 11.8 ton propellant mass and 1.2 ton dry mass. We’ll use a 2 ton dry mass of the crew module:

ORBITAL PROPOSES FUTURE DEEP SPACE APPLICATIONS FOR CYGNUS.
SPACEFLIGHT INSIDER
MAY 1ST, 2014
Orbital’s proposal, outlined in this PDF, involves docking a Cygnus spacecraft with Orion to serve as a habitation and logistics module on longer flights. For these missions, the re-purposed Cygnus would be called the Exploration Augmentation Module (EAM). With its current life support systems used to transport pressurized cargo and experiments to the ISS, Cygnus is stated as being already suitable for the long term support of a crew. While berthed to Orion, Cygnus could support a crew of four for up to 60 days. Cygnus also has the capability of storing food, water, oxygen, and waste and features its own power and propulsion systems. The EAM would utilize the enhanced configuration Cygnus, which will begin flying larger cargoes to the ISS beginning with CRS-4 in 2015. An even larger version is also being proposed, featuring a 4-segment pressurized cargo module.

https://www.spaceflightinsider.com/missions/commercial/orbital-proposes-future-deep-space-applications-cygnus/

Note though the phrasing here is ambiguous. The Cygnus capsule as used as a cargo transport to the ISS contains air, as it would have to for the astronauts at the ISS opening it to retrieve the cargo, but not life support systems. I'm inclined to believe for the usage cited in this article it would be taking life support from the Orion capsule. Then the calculations need to be made for how much mass it would take for life support, thermal management, consumables for an independent crew module.

Now for the delta-v calculation for our hydrolox lander, we'll assume we can match the max 465 s Isp of the RL-10 engine by giving the Ariane 4 upper stage engine a nozzle extension as used on the RL-10, then we get:

465*9.81Ln(1 + 11.8/(1.2 + 2)) = 7,000 m/s, 7 km/s. This is quite a bit higher than the 5.5 km/s needed for the round trip from NRHO to the lunar surface and back again. But it uses hydrolox propellant so needs extra mass for low-boiloff tech.

Low boiloff-tech and long duration hydrolox stages are an important enabling technology. ULA engineers and ULA CEO Tory Bruno have written about this extensively in regards to for example the proposed ACES derivative of the Centaur upper stage. Because of the prior research on low-boiloff tech, an operational version to be fielded in a short time frame to be used on the Artemis missions likely can be done.

This shows a single launch mission is doable if going to NRHO, but it is not my preferred plan. A complete orbit around the Moon at NRHO altitude takes about a week, and for the Orion capsule being at NRHO and not low lunar orbit, the lander's crew would have to remain on the Moon about a week before they could return to the Orion in the NRHO orbit. The landers crew module would have to be larger with heavier life support and consumables in this scenario.

If instead the Orion was at low lunar orbit it takes two hours to complete an orbit and the lunar lander could launch every two hours to rendezvous with the Orion.

Since the Orion's service module being given an insufficient propellant load is such an obvious design mistake, the preferred route to take would be to correct that error, thereby allowing the missions to take place from low lunar orbit instead of from NRHO.

Robert Clark

Towards a manned Indian launcher: an all-liquid LVM3.

noreply@blogger.com (Robert Clark) — Tue, 16 Jan 2024 16:12:00 +0000

In the blog post, "A liquid-fueled Indian manned launcher. UPDATED", I suggested the launcher based on the liquid-fueled LVM3 core stage, but replacing the 2 solid side boosters by 4 of the liquid-fueled strap on boosters used on the earlier design, the GSLV Mk. II. Here I'll suggest instead a version using the LVM3 core but getting the added thrust needed for lift-off by adding a 3rd Vikas engine.

GSLV Mk. III Specifications

I have argued using such large SRB's are not price competitive:

Towards a revolutionary advance in spaceflight: an all-liquid Ariane 6.

It is very likely the same is true for the GSLV Mk III. Then we'll replace the two SRB boosters by an additional core engine.

The GSLV Mk. III core stage has specifications listed as:

Core Stage


Type	L-110
Length	21.26m
Diameter	4.0m
Fuel	Unsymmetrical Dimethylhydrazine
Oxidizer	Nitrogen Tetroxide
Inert Mass	10,600kg
Propellant Mass	115,000kg
Launch Mass	125,600kg
Propellant Tanks	Aluminum Alloy
Fuel	UH25 - 75% UDMH, 25% Diazane
Oxidizer	Nitrogen Tetroxide
Propulsion	2 Vikas 2
Thrust (SL)	677kN
Thrust (Vac)	766kN
Specific Impulse	293 sec
Engine Dry Weight	900kg
Engine Length	2.87m
Engine Diameter	0.99m
Chamber Pressure	58.5bar
Mixture Ratio	1.7 (Ox/Fuel)
Turbopump Speed	10,000rpm
Flow Rate	275kg/s
Area Ratio	13.88
Attitude Control	Engine Gimbaling
Ignition	T+110s
Burn Time	200s
Stage Separation	Active/Passive Collets

The Vikas 2 engine provides a thrust of 677 kN at sea level, 69 tons-force. The two on the core would be enough just to loft the core only. But we need enough thrust to liftoff a second stage and payload also. So we'll give the core a third Vikas engine.

The weight of the Vikas is 900 kg. Then the dry mass of the stage with an additional Vikas will be 11,500 kg.

The cryogenic upper stage has specifications listed as:

Cryogenic Upper Stage


Type	C-25 Cryogenic Upper Stage
Length	13.32m
Diameter	4.0m
Fuel	Liquid Hydrogen
Oxidizer	Liquid Oxygen
Inert Mass	~4,000kg
Propellant Mass	25,000kg
Launch Mass	~29,000kg
Propellant Tanks	Aluminum Alloy
Propulsion	CE-20
Engine Type	Gas Generator
Thrust - Vacuum	200kN
Operational Range	180-220kN
Specific Impulse Vac	443s
Engine Mass	588kg
Chamber Pressure	60bar
Mixture Ratio	5.05
Area Ratio	100
Thrust to Weight	34.7
Burn Time	580s
Guidance	Inertial Platform, Closed Loop
Attitude Control	2 Vernier Engines
Restart Capability	RCS for Coast Phase

Now plug in the data for the Silverbirdastronautics.com payload estimator:

Where we assume by just using a nozzle extension the Isp can be raised from 443s to the 465s max Isp of the RL10 engine.

The resulting payload to LEO is:

This is half the 10 ton payload of the current version of the LVM3 with the large solid side boosters. However, it has the advantage of not using the problematical solid side boosters with their safety concerns for manned flights.

The all-liquid version is also likely to be significantly cheaper than the one with solid side boosters as large solid boosters are not price competitive to just using an additional liquid fueled engine.

It is notable a 5 ton class launcher is sufficient to launch a crewed capsule to orbit since the Gemini capsule had a toal mass of 3,800 kg:

GEMINI SPECIFICATIONS

First flight: 8-Apr-1964; first manned flight 23-Mar-1965 (Gemini 3)
Last flight: 11-Nov-1966 (Gemini 12)
Number of flights: 13 total; 10 manned
Principal uses: manned earth orbit rendezvous, docking, EVA tests
Unit cost: $13.00 million
Crew size: 2
Overall length: 5.7 m
Maximum diameter: 3.05 m
Habitable volume: 2.55 m3
Launch mass: 3,851 kg
Propellant mass: 455 kg total
RCS total impulse: 1,168 kNs
Primary engine thrust: 710 N
Main engine propellant: NTO/MMH
Total spacecraft delta v: 323 m/s
Power: fuel cells/batteries; 155.0 kWh total
https://www.braeunig.us/space/specs/gemini.htm

The payload to LEO also can be increased by weight optimizing the first stage. The stage is similar to the first stage of the Titan II that launched the Gemini capsule to space, except the Titan II's first stage dry mass was 6,000 kg less. Reducing the first stage dry mass input in the SilverbirdAstronautics.com payload estimator by 6,000 kg increases the payload to ca. 6,000 kg.

Robert Clark

Towards advancing the SpaceX Starship to operational flight: SpaceX should lower the Raptor chamber pressure and thrust level.

noreply@blogger.com (Robert Clark) — Mon, 08 Jan 2024 16:49:00 +0000

In the blog post, "Did SpaceX throttle down the booster engines on the IFT-2 test launch to prevent engine failures?", I suggested that SpaceX was able to get the SuperHeavy booster to complete its portion of the ascent in the last test flight by throttling down the thrust on the Raptor engines to less than 75%, and that the Starship upper stage was not able to because its Raptors were run at ~90%.

Reducing the throttle level reduces the chamber pressure also, from 300 bar to ~225 bar, allowing the Raptor to fire without leaks.

But if SpaceX lowered the thrust level on the booster to prevent engine failures, why did they not also do this on the Starship upper stage? From the propellant level indicators SpaceX provided on the launch video page, we can estimate the remaining propellant on the Starship just prior to FTS as under 100 tons out of a max 1,200 tons.

(The ship is not visible in this image because of its extreme distance just before it exploded.)

Then we can estimate if the engines were run at less than 75%, the remaining propellant would be in the range of 300 tons. This likely would have been too much SpaceX to reach its goal of getting the Starship to just under orbital velocity:

SpaceX Starship megarocket launches on 2nd-ever test flight, explodes in 'rapid unscheduled disassembly' (video) News.
By Josh Dinner published November 18, 2023
The spacecraft was never expected to reach full orbit around Earth, instead flying on a suborbital trajectory to splash down in the Pacific Ocean off the coast of Hawaii. "We're not targeting orbit today; we're targeting almost orbit," said Siva Bharadvaj, a SpaceX operations engineer, adding that the goal was to "get to a thrust profile similar to what we would need for orbit, but also energy level that the ship would need to dissipate for reentry."
https://www.space.com/spacex-starship-second-test-flight-launch-explodes

So what would be the payload possible if we ran both stages at ~75% thrust? A rough estimate would be at approx. 100 tons as a fully reusable launcher, instead of the 150 tons now. A problem with that is the estimated number of refuelings for the Starship HLS used as an lunar lander was perhaps 16. But if the max payload was only 100 tons, then the number of refuelings would rise to 24.

The idea of using so many refuelings for a lunar landing mission has been controversial. Then we'll look at other approaches. Robert Zubrin has noted it can be done in a single launch by giving the SuperHeavy/Starship a small 3rd stage, a mini Starship.

Elon Musk once suggested an expendable version of the Starship could have a 30 to 1 mass ratio:

__________________________________________________________________

Elon Musk @elonmusk

Mar 29, 2019

Replying to

@Erdayastronaut and @DiscoverMag

We'll explore this capability. We'll first look at a case where both the SuperHeavy and Starship are expendable stages. A stage generally gets better mass ratio scaled up, so to begin with we'll start off with an estimated mass ratio for the SuperHeavy as about 30 to 1 with the same propellant load as it is now of 3,400 tons but a dry mass of only 115 tons.

But that 30 to 1 mass ratio of the Starship is with just 3 engines since an upper stage does not need enough thrust to lift off from the ground. The SuperHeavy scaled up then would only have 9 engines, but this would not be enough for liftoff. We'll take the number of engines on this version of the SuperHeavy to be again 33. Then we need an additional 24 engines.

The thrust/weight ratio of the Raptor is about 140 to 1. Assuming we're reducing the operational thrust level of the Raptor to 75%, its thrust would be 0.75*230 tons-force = 172.5 tons-force. At a 140 to 1 T/W ratio its mass would be 1.2 tons. So the additional 24 engines needed would add 24*1.2 = 28.8 tons to the dry mass: 115 + 28.8 = 143.8 tons. Round it up to 150 tons with the additional plumbing and larger thrust structure required.

However, with the lower sea level thrust of our 75%-throttle Raptors, the 2-stage SH/SS would have low T/W ratio. We'll also look at a case of adding small 3rd stage of size ~300 tons. Adding a ~300 ton 3rd stage reduces the T/W to barely above 1. Then it would be advisable to add 2 more Raptors to the booster, bringing it to 35. There should be sufficient room beneath the booster for an additional 2 Raptors.

The total thrust of the SuperHeavy at sea level is now 35*172.5 = 5,950 tons-force, 58,400 kiloNewtons(kN). We'll use the SilverbirdAstronautics.com payload estimator. This takes the vacuum Isp and vacuum thrust as inputs even for first stage engines. For the lower chamber pressure, ~ 225 bar corresponding to the reduced thrust, we would also have lower vacuum Isp. Call it about that of the Raptor 1 so ~ 350s. Then the vacuum thrust for this version of the Raptor would be (350/327)*172.5 = 184.6 tons-force. For the 35 SuperHeavy engines that's 35*184.6 = 6,460 tons-force, 63,400 kN vacuum thrust. For the dry mass for the 35 engine stage we'll raise it to 152.4 tons.

For the Starship upper stage, the Raptor Vacuum will still have a vacuum Isp of 380s. However, we are keeping 3 sea level Raptors and 3 vacuum Raptors. Then the average Isp for the upper stage will be 370s. The vacuum thrust calculates out to be 11,500 kN.

With the 3 additional Raptor engines being vacuum engines the added weight will be higher than that of just 3 additional sea level engines. We'll take the dry mass of the new upper stage to be 45,000 kg.

Then the input page to the SilverbirdAstronautics.com page looks like:

And the payload results are:

The expendable payload value of 264 tons is more than the expendable payload of the current SH/SS of 250 tons even-though the new version would be at 75% thrust level. This surprising result must be due the greatly reduced dry mass of both stages.

But even more surprising is the payload possible to translunar injection(TLI). In the SilverbirdAstronautics.com calculator this option is indicated by selecting "Escape Trajectory" for "Destination". In the "Hyperbolic, C3" field enter, -1.0. This number indicates how far beyond escape velocity the flight needs. In this case, it's negative since you don't quite have to get to escape velocity to reach the Moon.

Then the estimated payload is:

The payload would be ~75 tons sent to rendevous with the Moon.This is far beyond what Apollo at ~43 tons or SLS even in its later Block 2 version at ~46 tons to TLI could do as a single launch architecture. This would allow wide latitude in how you would design the in-space stages and lander to reach the surface of the Moon.

A Three-Stage SuperHeavy/Starship/mini-Starship.

A basic principle of spaceflight is high delta-v missions, such as missions to the Moon or Mars, can be done more efficiently with more stages. Robert Zubrin has proposed a "mini-Starship" as a 3rd stage for the SuperHeavy/Starship. He trenchantly observed, "The Starship is a reusable Saturn V. It is not a LEM". Addition of a 3-rd stage would allow even higher payload to the Moon. Say, the mass ratio of ~30 to 1 can be retained for the 3rd stage at ca. 300 propellant load and ca. 10 tons dry mass, using a single Raptor vacuum of 380s Isp and 1,900 kN vacuum thrust. This gives an input page result:

Then the payload to TLI would be ~105 tons:

This is regarding the SH/SS/mini-SS as a launcher analogous to the 3-stage Saturn V for the Apollo missions. This would require additional in-space stage(s) for doing the landing. However, the payload to TLI is so high this mini-Starship could itself serve as the lander for a round-trip mission to the Moon's surface with a capsule or habitat of ~15 tons mass. This is a mass 50% higher than the mass of the Orion capsule that could be carried not just to a lunar orbit planned for Orion, but all the way to the lunar surface and back to Earth again.

Flights to Mars.

Addition of a third stage also allows more efficient flights to Mars in a single launch. This graphic gives the C3 needed for flights to the planets:

For Mars it is about 14 km2 /s2. The TMI payload to Mars is then:

A payload of 86 tons sent towards Mars for trans-Mars injection(TMI) as a single flight architecture.

These are for expendable versions of SuperHeavy/Starship. Elon Musk has estimated the cost of the SH/SS as in the range $100 million to $200 million. Even as expendable this is a greatly reduced cost than the SLS cost of $4 billion per flight.

Estimates of Reusable Payloads.

The reduction of cost of a flight to the Moon and Mars to only a couple of hundred million dollars rather than multi-billions makes the expendable versions worthwhile. But SpaceX is committed to reusability.

Elon has said full reusability would lose 40 to 50% off payload. And even the reusability only for the first stage loses a quite significant 30% from the payload. These are from using powered boost back approach to landing. The propellant that must be kept on reserve, unused during the ascent for the boost back and landing is in the range of 7%. For the 3,400 ton prop load SuperHeavy this ~240 tons. This is doubly disadvantageous for payload in that not only can this prop not be used during flight to orbit but it also acts like additional dead-weight that must be carried in flight.

Elon Musk has said powered, vertical landing is preferred over horizontal, winged landing because it can be used on airless or low atmospheric worlds such as the Moon and Mars. However, SpaceX needs to do a trade study to see which method results in the least payload lost. Wings typically take up about 5 to 10% of the gross weight of an aircraft. For a vertical launch rocket using non-lifting trajectory on ascent to orbit, this 5 to 10% would only have to be of the dry weight of the craft since aerodynamic lift would be used only during return when the stage is nearly empty.

So for the SuperHeavy the added weight of the wings would only have to be 7.5 to 15 tons compared to the 240 tons needed for the vertical landing method. Actually, it likely can be even smaller than the 5%. For one thing rather than using large wings such as used on the Shuttle. It could use short, stubby wings instead. The large wings of the shuttle were due to competing requirements by NASA and the Air Force for where the Shuttle should be able to land.

Examples of how they could look instead might be those of the X-37 and of the Skylon:

For the Skylon by using carbon-fiber the wing weight was only 2% of the landed weight. This is 2% of the full gross weight because it used a horizontal liftoff. But since the Starship will be using a vertical liftoff and non-lifting trajectory, the wings only have to support the weight of the vehicle on return, so that 2% only has to be calculated on just the dry weight. The specialty high strength stainless steel used on the Starship might be preferred for the wings as well since it would have about the same strength but have reduced thermal shielding requirements.

The landing gear weight can be taken as only 3%:

https://yarchive.net/space/launchers/landing_gear_weight.html

This likely can be reduced by half to ~1.5% by carbon-fiber or the specialty steel used on the Starship.

Finally, the thermal protection such as SpaceX’s PICA-X might only add on additional 8% of the dry weight.

So these extra systems required for reusability will only add a proportionally small amount to the dry mass, and so subtract only a proportionally small amount from the payload.

Bob Clark

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

noreply@blogger.com (Robert Clark) — Thu, 28 Dec 2023 20:50:00 +0000

I had earlier argued that SpaceX should withdraw the Starship as a lunar lander. The primary basis for this was for safety of the surrounding population in case of an explosion on launch, SpaceX should withdraw its application for the Starship as an Artemis lunar lander.

However, an additional reason why the Starship should not be used for a lunar lander is for safety of the crew. In the blog post, Did SpaceX throttle down the booster engines on the IFT-2 test launch to prevent engine failures?, I noted two separate methods of calculation suggest the SuperHeavy booster was throttled down to <75%. I also suggested the Starship upper stage was fired at ~90%. Given this difference in thrust power levels, I suggested the booster completed its portion of the ascent because it was throttled down and the upper stage did not because it was at close to full thrust.

Even though the booster engines successfully fired during the ascent, the booster exploded during the attempted return. One explanation offered was the engines were damaged by fuel slosh during flip of the booster. However, it should be noted the Starship during tests of the landing procedure, that at least one Raptor always leaked fuel and caught fire.

Note even in the last two shown here, SN10 and SN15, there were engine fires on landing. For SN10 the engine fire led to the vehicle exploding a few minutes after landing. For SN15 the fire was extinguished before it caused an explosion. SN15 was called a “successful” landing test because it did not explode. But that a Raptor still caught fire during this test gives further evidence the Raptor is still not a reliable engine.

And SN11 experienced a catastrophic explosion after a fuel leak and engine fire:

Since relighting the Raptors in flight always resulted in an engine fire, that is the most likely explanation for the IFT-2 booster explosion as well.

SpaceX Misleadingly Characterizes Raptor's Qualification for Flight.

SpaceX has been using the term "full duration" for their Raptor static fire tests when they might only last 5 seconds. In the rest of the industry other than SpaceX, a full duration static test means firing for the full duration of an actual launch.

Rocket Factory Augsburg

@rfa_space

280 seconds of glorious hot fire! 
We are incredibly proud to be the 1st private company in #Europe () to hot fire a staged-combustion upper stage for its full duration. This qualifies our upper stage and Helix engine for flight 
Enjoy the video and read more in our press release  https://bit.ly/3WJY2G4

https://x.com/rfa_space/status/1664683388928655374?s=20

And for the four SSME's on the SLS core stage:

SpaceX calling their 5 second long test fires "full duration" misleadingly gives the impression that is sufficient to qualify the engines for full mission flight time.

No estimates for Raptor engine reliability publicly provided.

For engines for a craft intended to carry astronauts and for which billions of dollars of public funds are earmarked there should be provided some indication about the safety and reliability of such engines. For instance this report provides estimates of the reliability of the different components of the SLS:

SLS-RPT-077
VERSION: 1
National Aeronautics and Space Administration
RELEASE DATE: MARCH 8, 2013
SPACE LAUNCH SYSTEM PROGRAM (SLSP)
RELIABILITY ALLOCATION REPORT

https://foia.msfc.nasa.gov/sites/foia.msfc.nasa.gov/files/FOIA%20Docs/42/SLS-RPT-077_SLSP-Reliability-Allocation-Report.pdf

But no such estimates for the Raptor have been provided. That so many engines have consistently failed in actual flights suggest they have quite low reliability.

In the scenario of the Merlin engines used for crewed flight, over 80 missions of the Falcon 9 were successfully flown before the first crewed flight. That means over 800 successful firings of the Merlins during that time. And added on after that the many launches since then, over one thousand successful firings of the Merlins have been made.

Robert Clark

Did SpaceX throttle down the booster engines on the IFT-2 test launch to prevent engine failures?

noreply@blogger.com (Robert Clark) — Fri, 15 Dec 2023 16:16:00 +0000

 Given the Raptors repeated history of leaking fuel and catching fire I was surprised the booster was able to complete its portion of the ascent with no engine failures.

 Hypothesis: the booster flew without engine failures because it throttled down to < 75%. The Starship had engine failures because it ran at ~90%, like the booster did on the first test flight with its multiple engine failures.

Throttle Down Calculated by Propellant vs. Time Graph.

Two separate observers, u/jobo555 and @space_josiah found fairly constant propellant flow rate, and therefore throttle, before where the booster begins to prepare for stage separation. Rocket thrust is given by (thrust) = (exhaust speed)*(propellant flow rate). So can get degree of throttle by propellant flow rate.

 The graphs give the percentage of propellant remaining vs time. From this we can calculate the percentage change rate as the slope. For the booster it’s about 0.5%/s, 0.005/s as a decimal. Then given the total propellant load of 3,400 tons, in absolute term that propellant flow rate is 17 tons per second.

 But the full thrust propellant flow rate  for each Raptor v2 can be calculated as:

props flow rate = thrust/exhaust speed = 230,000*9.81/(327*9.81) = 700 kg/s.  Then for all 33 engines on the booster that’s 33*700 kg/s = 23,100 kg/s, 23.1 tons/s. Then the throttle down for the booster amounted to: 17/23.1 = .736, less than 75%.

 For the Starship, from the first image below, in its second graph we see from 4 minutes to 8 minutes, 240 seconds, the propellant level dropped from  ~80% to ~5%, for a percentage rate drop of 75/240, 0.313%/s. Then the absolute flow rate for a 1,200 ton prop load is 3.756 tons per second. But for the 6 engines the flow rate at full thrust would be 6*700 = 4,200 kg/s, 4.2 tons/s. Then the throttle is .894, ~90%.

 Note that throttling down to 75% also correspondingly drops the combustion chamber pressure from 300 bar to about 225 bar, allowing the Raptor to operate without leaks. 

 But this reduced thrust would also mean the SuperHeavy/Starship could carry less payload. I estimate a drop in payload to ca. 100 tons reusable. In such a scenario, the 16 refueling launches needed for a Starship HLS would be increased to 24 launches.

Throttle Down Calculated by Acceleration Graph.
 A completely separate argument allows us to conclude the thrust was throttled down to less than 75%. This observer @meithan42 looked at the velocity and altitude data and derived the acceleration data.

  On the acceleration graph I marked where the horizontal acceleration visually appears about 10 m/s2. The vertical acceleration there visually appears as about 6 m/s2. Visually this occurs at about the 90 second point. 
 Note that gravity subtracts ~10 m/s2 from the vertical acceleration, the actual vertical acceleration produced by the engines thrust is about 16 m/s2. Then the actual acceleration generated by the engines thrust is SQRT(102  + 162) = 18.87 m/s2 .  

 But now lets calculate the actual acceleration that should be produced by the engines assuming they were running at full throttle at the 90 second point. The thrust is, (thrust) = (exhaust speed) * (flow rate). Since we are near vacuum the Isp will be 363 s and the exhaust speed 363*9.81 = 3,560 m/s. Then the thrust at full throttle with a total prop flow rate of 23,000 kg/s, should be thrust = 3,560*23,000 = 81,880,000 N. 

 We'll take the total mass of the rocket as 4,850,000 considering the tanks are filled slightly less than 100%. If the engines are at full throttle then the mass after 90 seconds is 4,850,000 - 90*23,000 = 2,780,000, and the actual acceleration generated would be 81,880,000/2,780,000 = 29.45 m/s2. This is well beyond amount observed. 

 In contrast, if we take the throttled down propellant flow rate as 17,000 kg/s, then we calculate the actual acceleration as:

363*9.81*17,000/(4,850,000 - 90*17,000) = 18.23 m/s2 ,a value much closer to what is actually observed.

   Robert Clark

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