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The "standard" Block 2 (about to become Block 3) Starship does its deorbit burn and its flip-and-landing burn on the propellant in the header tanks, so SpaceX says. They also said that those were transferred to the main tanks in the Block 2 design to make the landing burn. (Note that there is really nothing "standard" about Starship until all the experimental flight testing is successfully done.)
That transfer from headers to main tanks may or may not still be true in Block 3, as I saw in some illustrations plumbing direct to the engines from the header tanks. Those header tanks hold something like 20-25 tons at most, not anything like 100 tons! The ship explodes upon toppling over in the Indian Ocean, because not all of that 20-25 tons was used to make the landing. There are still fuel and oxidizer aboard when it falls over and breaks open.
I do not know of any variants yet being seriously considered except the HLS and some sort of propellant tanker. And that will likely be true until after the next round of Block 3 flight tests are done! Then it gets bigger yet again in the Block 4 design, which must again be proven in test before it can be considered as a viable prototype for any sort of real mission work. Excepting maybe HLS and the tanker, any other variants must wait until everything through Block 4 has been tested and found adequate.
As for the ballistic coefficient of a Block 2 or Block 3 Starship, the hypersonic drag coefficient of a round cylinder dead broadside to the wind is about 1.20 based on the cylinder blockage area, coming from my old Hoerner "drag bible". The drag coefficient of a flat plate normal to the wind in hypersonic flow is about 1.84, based on its blockage area. Same data source. I took a good guess for the effects of the pointed nose, and for the relative blockage areas of body and flaps, and determined a CD = 1.22 on a blockage area of just about 462 sq.m normal to the wind.
Starship does NOT enter normal to the wind, it enters at a nominal angle of attack of 60 degrees, although that varies some about that nominal angle. Thus the normal blockage area is not actually normal (90 deg) to the wind, it is about 30 deg off. That's no big impact on CD, but the effective blockage area is 462 sq.m times a cosine factor of 0.866 for that angled entry.
I used Bob's numbers of 120, 160, and 40 metric tons for the inert mass of Starship, even though I thought (and still think) the 40 ton figure is ridiculously unrealistic. To that I added 20 tons landing propellant, and 100 tons payload at Mars, and I added the same 20 tons of propellant and 0 tons of payload at Earth. That puts mass-at-entry 240, 280, and 160 metric tons at Mars, and 140, 180, and 60 metric tons at Earth.
I did entries at Mars at 7.5 km/s off a fast trajectory, and 2 deg below horizontal. I did entries at Earth at 7.9 km/s out of low circular LEO and 2 degrees below horizontal.
That's where my numbers came from! They should be pretty good, as good as the masses at entry are. The whole thing turned into a sensitivity study with ballistic coefficient the independent variable. The low inerts are ballistic coefficient roughly half those of the heavier values.
GW
"Building things in orbit is still a very future wish list item."
But that is EXACTLY how the ISS was built over 30 years ago! Docking together stuff sent to orbit. The only requirement to do that was the arm on the shuttle to hold things where you wanted them to go, instead of using thruster thrusts. That docking thing is the primary reason there is an arm on the ISS today.
ISS is in the wrong high-inclination orbit to go anywhere else without huge plane-change dV's. But a station in a low inclination orbit does not suffer that problem. Simply equip it with multiple arms, and you can assemble lots of things by docking, by nuts and bolts, by clamps, perhaps even by welding. Just hold the items together in the right position with the arms. If any of this is manual work, the EVA astronauts "stand" attached to other arms. Demonstrated decades ago on the shuttle servicing Hubble and others.
Add another truss section out there with propellant tanks attached to it, and plumbing and power routed inside the truss, and you have a refueling facility as well. Use my rotating-vanes tanks, and transferring cryogenics becomes as easy as handling storables. And there is no reason you cannot also stockpile the gases used by electric propulsion as propellants.
All of which makes space tugs possible for sending and receiving things using high-elliptical capture orbits. It takes at the very least a refueling capability to do that, and even SpaceX is talking about a Starship variant that stays on orbit as a refueling facility, to be kept filled with propellants brought by other tankers, propellants stored on orbit for other craft to use.
The same tug stages that can go onto, and back off of, an elliptic orbit, based from low circular LEO, can also go get depleted satellites for their refilling at the station, and then putting them back where they were before. Even geosynch is less challenging than an elliptic capture orbit that goes more than halfway to the moon. Some of this might be done remotely, some of it might be manned. So what?
We not only could do this today, we could have done most of this at least 30 years ago! My rotating vane tank would make cryogenics transfer a lot easier (and faster) than by using ullage thrust, which disturbs your orbit. That did not exist 30 years ago, but the arms and ullage thrust did!
And a "tug" is just any convenient stage of adequate dV and payload capability, plus a guidance and control rig, and maybe a place for a crew capsule on manned missions, sent up there and kept filled for use on many different missions.
There not a reason in the world why this could not be done, except government and corporate "not invented here" prejudices. With my spinning-vanes tank approach for cryogenics, ALL the critical enabling technologies now exist to make this very practical!
GW
From AIAA’s Daily Launch email newsletter for 2-23-2026:
NEW YORK TIMES
Problem With Artemis Rocket Is Likely to Delay NASA’s Moon Mission
A day after NASA officials optimistically said they were on track to launch astronauts around the moon early next month, a problem with the rocket’s upper stage will require rolling it off the launchpad for repairs.
My take on it:
“Upper stage” would refer to the interim upper stage that was the Delta-IV upper stage. That is something Boeing builds. The pattern really is complete.
GW
That's the latest thing I was looking at. It would be assembled from components shipped up to LEO. I might use 1-way/1-shot solids for the dV onto the Hohmann trajectory, and have NTO-MMH do the course corrections and "fine-tune" the speed getting onto that trajectory.
If the departure stage is to be recovered, I would use LOX-LH2 for departure, and NTO-MMH for course corrections and for capture back into an extended ellipse about the Earth. The transfer trajectory would be an ellipse past Mars with exactly a 2 year period. One would need a tug from LEO to go get that transfer stage and bring it back to LEO.
A space station with assembly facilities and refueling facilities would be perfect for this. It needs to be a low-inclination orbit for this, to reduce the dV's going anywhere beyond Earth orbit. At 53 deg, the ISS orbit is totally wrong for that purpose.
GW
There you go, you heard it from me and you heard it from Oldfart1939. We both quoted the same news article, I think.
GW
From what I have read, we have not been told whether the helium problems were in the core rocket or the spacecraft and its service module.
The Europeans built the service module. Lockheed-Martin built the capsule. Boeing built the SLS core stage, and until it was retired, the upper stage of Delta-IV which is the second stage of SLS. The SRB's do not use helium at all. 2 of the 4 things that use helium are built by Boeing, and the helium problem was also one of two near-fatal difficulties for the Dreamliner, also built by Boeing.
See the pattern here? The odds favor the helium problem being either in the 1st or second stages of SLS.
GW
SpaceX has long said Starship enters at about 60 degree angle-of-attack, not dead broadside at 90 degree angle of attack. This does 2 things: (1) it generates a lift vector (albeit at low L/D ~ 0.1 or less) for trajectory control by controlling roll angle around the vector of the oncoming wind, and (2) it limits the massflow throughput of hot plasma into the open-ended engine compartment (and it accomplishes that at zero inert mass contribution for a closure across the engine bay outlet).
That angle of attack varies some, for trajectory control. Sometimes it’s under 60 degrees, sometimes over. That puts the stagnation point somewhere on the bottom of the nose, in the whitened zone of the Flight 10 ship that was stained orange. That point moves around on the bottom of the nose of the vehicle as angle of attack changes, so the whitened zone is a large as it was. Supposedly, the white stain was from the ablative products of whatever the ablative backup layer was, underlying the tile.
Once the hypersonics are over and the heating and deceleration has eased, the vehicle's trajectory is rapidly bending over nearer vertical, and the vehicle controls angle of attack to dead broadside, 90 degrees to the oncoming wind, for the "belly-flop maneuver" we have seen since the very first suborbital flight tests. There is no longer a threat of super-hot plasma intruding into the engine bay (air is OK, and even at Mach 3, it is not all that hot), obtaining max drag is more important. The vehicle is very near its terminal velocity, which is still supersonic at high altitude in the thin air, slowing to low subsonic down in the thick air near the surface.
I used a CD = 1.2 for a cylinder broadside to the oncoming wind, from Hoerner at Mach 6-10-ish, which is hypersonic. Hoerner says a flat plate dead broadside is CD = 1.84, so the round cross section shape of Starship is definitely a drag reducer compared to a flat plate. But I did 1.2 on the projected body area plus 1.84 on the projected flap areas, for an overall CD = 1.22 on the total projected area. For ballistic coefficient, I multiplied the area in the denominator by the sine of 60 degrees, since it is NOT dead broadside during entry.
Hoerner's "drag bible" is full of real test data for a multitude of things. His English does show a German accent in his writing, though. Some folks have difficulty with that. But for fast things since the X-15 that were designed before computers, Hoerner was THE source. And that includes the SR-71, the X-20, and the early concepts for the space shuttle, as well as the Mercury, Gemini, and Apollo capsules.
The stagnation zone heating correlations I put into the spreadsheet were the convective model from the 1950's from H. Julian Allen's work, which equation closely resembles the Sutton-Graves thing. But I added a radiation heating model from the 1960's to it, that for Earth entries shows little effect below 10 km/s speeds, but dominates above it. That is precisely what Apollo saw! The same model shows a more dominant plasma radiation heating effect for entries at Mars, even at lower speeds, because the density lapse rate with altitude is different, so that peak heating and peak deceleration are taking place at different densities than at Earth.
Further, the convective model shows higher heating with smaller nose radius, while the radiational heating is higher as nose radius increases! This neatly explains why Earth entry heat shields are very blunt, while Mars entry heat shields have been conical with a small tip radius on the tip of the cone. And I just told you why that has been true!
All in all, I think I have just about the right models in the entry spreadsheet file. The Justus and Braun entry, descent, and landing "bible" document certainly seems to think so.
As for surface temperature, estimating that has a multiplicity of options with a different model for each of them. Ceramic tiles that do not ablate cool by thermal re-radiation and by conduction into the interior. The cooling must balance the incident convective plus plasma radiation "steady state", because the peak heating transient is closer to a minute long than it is to a second.
What I did was just zero the conduction inward, and let all the cooling be re-radiation. That slightly over-estimates equilibrium surface temperature. But the choice of higher infrared emissivity is a stronger influence upon equilibrium temperature than the deletion of conduction inward. That's why most heat shield tiles on windward surfaces are black.
GW
Spacenut:
I think they fixed most of that in the last Block 2 flight, the one after the flight where the picture in post 2243 came from. The explosive burn-through into the engine bay seems unique to that one flight. Although, whether that was caused by a deliberately-missing tile or not, has never been made clear, as far as I can tell.
They also added something (not clear what) to the flap-body joint where the burn-throughs had been happening, and I noticed that there was no flap sheet metal damage visible during that last ascent. The previous pictured flight showed some kind of aft edge sheet metal damage during the ascent right after liftoff. So they changed something there, too. It is just entirely unclear what that something was.
Bob:
There's a lot more going on with inert mass than just scale and material. Most of it is not publicized or made clear. But by the time Block 2 flew, they found they had to add a lot of stringers and frames inside the propellant tanks than they had used on Block 1. They also changed the shape of the aft dome to which the engines are mounted, making it more conical, if memory serves. That also requires additional framing not to bend-toward-round when the tank is pressurized, and to avoid crippling-collapse with the higher engine thrusts (and they are higher!).
Both of you:
Something else real-world to consider: the ballistic coefficient would be lower by a factor of about 87% if reentry were made dead broadside, exposing the largest possible blockage area to the oncoming stream. But they cannot do that! Dead broadside, there is too much reentry plasma (thousands of K effective) getting into the engine bay. They fly reentry at about 60 degree angle of attack as the compromise that limits plasma intrusion into the engine bay, while at the same time presenting the largest possible blockage area to the hypersonic flow. You'll notice that when the hypersonics are over and the hot plasma danger is no more, the belly-flop maneuver really is flown just about dead-broadside to the relative wind, for the biggest-possible drag area. It shows as a near-horizontal axis while flying almost straight down.
GW
From CBS News website 2-20-2026 (my take on it appended at the end):
NASA's new chief rebukes Boeing, space agency over problem-plagued Starliner mission that left astronauts stuck in space for months
By
William Harwood
Updated on: February 19, 2026 / 8:10 PM EST / CBS News
An independent review of the first — and so far, only — piloted flight of Boeing's troubled Starliner spacecraft concluded that the test represented a potentially life-threatening "Type A" mishap resulting from multiple technical problems and management miscues, NASA officials said Thursday. The findings prompted NASA's new chief to make openly critical comments about his own agency and Boeing.
"This was a really challenging event and...we almost did have a really terrible day," said Amit Kshatriya, NASA associate administrator. "We failed them."
He was referring to now-retired astronauts Barry "Butch" Wilmore and Sunita Williams, who were launched in June 2024 expecting to spend eight to 10 days in space. They ended up remaining in orbit for 286 days, hitching a ride home aboard a SpaceX Crew Dragon capsule in March 2025 after NASA ruled out landing aboard the Starliner.
NASA Administrator Jared Isaacman, who took the reigns of the agency in December, said NASA will continue working with Boeing to make the Starliner a viable crew transport vehicle, adding that "sustained crew and cargo access to low Earth orbit will remain essential, and America benefits from competition and redundancy."
"But to be clear, NASA will not fly another crew on Starliner until technical causes are understood and corrected, the propulsion system is fully qualified and appropriate investigation recommendations are implemented," he said.
He made the comments as the agency was releasing the results of a months-long independent investigation of the Starliner mission. The panel's report cited a long list of management failures and technical issues that were not fully understood at the time, but were still considered acceptable for flight.
The panel concluded the problems experienced during the mission were representative of a "Type A mishap," meaning an unexpected event that could have resulted in death or permanent disability, damage to government property exceeding $2 million and the loss of a spacecraft or launch vehicle.
Isaacman said the eventual cost of the Starliner's woes exceeded the $2 million threshold "a hundred fold."
"Starliner has design and engineering deficiencies that must be corrected," he said. "But the most troubling failure revealed by this investigation is not hardware. It's decision-making and leadership that, if left unchecked, could create a culture incompatible with human space flight."
Isaacman said the investigation revealed pressure within NASA to ensure the success of the agency's Commercial Crew Program, which is based on having two independent astronaut ferry ships. That advocacy "exceeded reasonable bounds and placed the mission the crew and America's space program at risk."
"This created a culture of mistrust that can never happen again and there will be leadership accountability," Isaacman said.
The report quoted unnamed personnel saying things like, "There was yelling in meetings. It was emotionally charged and unproductive."
Another said, "If you weren't aligned with the desired outcome, your input was filtered out or dismissed."
Yet another told the panel, "I stopped speaking up because I knew I would be dismissed."
Equally troubling, according to one NASA worker quoted in the report, "NASA wasn't blaming Boeing, but everybody else was. [...] You know, it's our program. We're responsible too. Nobody said that. And nobody within NASA [or outside of NASA] has been held accountable. Nobody. We're 11 months after it happened, and there's been no accountability at all, from any organization."
Isaacman promised that "lessons will be appropriately learned across the agency and there will be accountability."
In the wake of the space shuttle's retirement in 2011, NASA awarded multi-billion-dollar contracts to Boeing and SpaceX in 2014 to build independent ferry ships to carry astronauts to and from the space station. SpaceX, awarded an initial $2.6 billion contract, has now launched 13 piloted Crew Dragon flights for NASA and seven purely commercial missions.
In contrast, Boeing, awarded an initial $4.2 billion contract, ran into multiple problems during an unpiloted Starliner test flight in 2019 that eventually required a second crew-less test flight before Wilmore and Williams were finally launched on June 5, 2024, on what has been the ship's lone crewed test flight.
The trip to space atop a United Launch Alliance Atlas 5 rocket went smoothly and the crew successfully docked with the International Space Station the next day. But the capsule experienced multiple helium propulsion system leaks along the way and several maneuvering jets did not produce the expected thrust.
"During the rendezvous and proximity operations, propulsion anomalies cascaded into multiple thruster failures and a temporary loss of six-degree-of-freedom control," Isaacman said Thursday. "The controllers and the crew performed with extraordinary professionalism ... and docking was achieved.
"It is worth restating what should be obvious," he said. "At that moment, had different decisions been made, had thrusters not been recovered or had docking been unsuccessful, the outcome of this mission could have been very different."
Williams and Wilmore downplayed the malfunctions during the flight, which was originally expected to last about eight days. But NASA and Boeing ended up extending their stay in orbit, carrying out weeks of tests and analysis to determine whether the Starliner could be trusted to safely bring its crew back to Earth.
By August 2024, Boeing managers were convinced engineers understood the problems and the crew could safely come home in the Starliner. But NASA managers ruled that option out. Instead, they decided to keep the astronauts aboard the station until early 2025 when they could hitch a ride back to Earth aboard a SpaceX Crew Dragon ferry ship.
To make that possible, a Crew Dragon was launched in September 2024 with just two astronauts aboard instead of four as originally planned. That freed up two seats for Wilmore and Williams after the SpaceX crew completed their six-month stay in space.
The Starliner, meanwhile, successfully made an uncrewed return to Earth in September 2024 even though, the investigation report revealed, additional propulsion problems left the craft with no available backup options had another failure occurred.
The mission, "while ultimately successful in preserving crew safety, revealed critical vulnerabilities in the Starliner's propulsion system, NASA's oversight model and the broader culture of commercial human spaceflight," the investigation team concluded.
The panel issued 61 formal recommendations "across technical, organizational, and cultural domains to address these issues before the next crewed Starliner mission."
"The report underscores that technical excellence, transparent communication, and clear roles and responsibilities are not just best practices, they are essential to the success of any future commercial spaceflight missions," the team said. "The lessons from CFT must be institutionalized to ensure that safety is never compromised in pursuit of schedule or cost."
For its part, Boeing said in a statement the company had made "substantial progress" on corrective actions "and driven significant cultural changes across the team that directly align with the findings in the report."
"NASA's report will reinforce our ongoing efforts to strengthen our work...in support of mission and crew safety, which is and must always be our highest priority. We're working closely with NASA to ensure readiness for future Starliner missions and remain committed to NASA's vision for two commercial crew providers."
------
My take on it:
This report understates, but confirms, my contentions about NASA management culture valuing money and schedule (which is also money) above crew lives. It is quite apparent that NASA management never learned the lessons of the two lost shuttle crews. Or if any of them did, they no longer work there.
This report also confirms something not common knowledge up to now: that Butch and Sunita’s Starliner experienced even more propulsion problems during its unmanned return from the ISS, leaving no backups.
It would appear that Jared Isaacman might do some “housecleaning” among NASA management. He at least talks like it, in this report. We will see if he really does. However, it does concern me about him that, so far, he shows no sign of concern over a flawed heat shield flying crewed on Artemis-2, and the same flawed design being built for Artemis-3.
It would also appear, based on this report, that Jared Isaacman wants Boeing to properly “fix” Starliner, so that he has a second capsule to use, going to ISS. If it were me, I’d plow that money into Dreamchaser instead. But that is just me.
I would very strongly recommend to NASA that they fly Dragon with all 7 seats installed, even when only flying crews of 4 to the ISS. With the extra 3 seats, up to 3 astronauts in trouble at a time, could hitch a ride home with any full-size Dragon crew sent to ISS, that is coming home. Having that capability is simply far more important than the weight of the 3 seats, or any egress time issues associated with them. But, that importance assessment is true only if you value lives over money!
GW
Bob:
Check your email. I was able to rerun the entry study for circular LEO at Earth. I sent you a pdf file.
GW
Bob:
I think the 40 ton inert figure for Starship is unrealistic in the extreme. Myself, I never heard him say anything under 80 tons. But as far as I know, Block 1 was in the vicinity of 120 tons. And it has grown since then, about 6 meters longer.
I thought you were asking about Mars entry, which is why I looked at that. It's easy enough to run the spreadsheet here at Earth. I can use most of the same inputs, just the Earth atmosphere model. And Earth entry from low circular orbit would hit the atmosphere at about 7.9 km/s. One variation could be deleting the payload weight, on the assumption it was delivered on-orbit. That still leaving landing propellant aboard, though.
I am not familiar with the X-33 metallic shingle thing. But I do know that one of the last two Block 2 Starship flights had at least some "metallic tiles" which were apparently iron-based or iron-containing. These apparently experienced high rates of oxidation, resulting in the "rust-color" staining seen on the vehicle.
GW
Harold:
I haven't yet had a chance to go back and look at what I did trying to land 40 tons of payload 1-way into Mars. If memory serves, and it may not, I did not like my final results. The harshest constraint seemed to be getting enough thrust to fit the space in the vehicle, and second-harshest was protecting those engines during entry.
Riding out in the open behind a heat shield exposes you to the same effective temperatures the heat shield sees, just at an order-of-magnitude-or-so less scrubbing velocity on surfaces, and pressures near atmospheric ambient at extreme altitudes. The heat shields see more scrubbing at much higher pressures, which is why those heating rates are so much higher.
GW
Bob:
Check your email. I ran the entry study you suggested, using the entry spreadsheet available for free download right off the links here on the forums for the "orbits+" course materials. I sent you a pdf document of what I did. I looked at 120 mt, 160 mt, and 40 mt inert dry masses, but I added 100 mt payload and 20 mt landing propellant to those masses. I also used a block 3 length of 56 m.
That's about a factor-2 range of ballistic coefficient, and all 3 showed peak heating at about 35-40 km, and 6 to 6.3 km/s speeds. Using the entry old rule-of-thumb that is about 10% accurate, effective temperature deg K is numerically equal to speed in m/s. It falls between about 6000 and 6300 K, for all 3 configurations.
The amount of plasma radiation heating varied by about a factor of 2 across that range of ballistic coefficients. The altitude at end of hypersonics was around 10-16 km, highest at the lowest ballistic coefficient. Not much else varied at all, not even the speed and altitude for peak heating.
I did find that plasma radiation stagnation heating dominates by far, quite unlike at Earth. I did NOT do the heat transfer balance trying to determine where the tile surfaces might equilibriate. But since the driving temperatures are about the same, the equilibrium temperatures might not be that much different, despite the crudely factor-2 difference in heating rates.
GW
Rob:
I have no way to contact Jared Isaacman directly. All I can do is write a paper letter to him at NASA headquarters. He will not see it, some staffer will. And THAT is the problem. No one but that staffer will see it, and he/she will shit-can it. I had the very same problem writing to his predecessor with the heat shield "fix" almost 2 years ago now.
Rob & Harold:
What I eventually did was send the heat shield stuff to a former colleague whom I found out was working at NASA. He took it to the thermal protection group in Houston, and verified to me that they got it, and that they pretty much agreed with it!
My friend has since retired, so that doorway in, is now closed! After that, I never heard another peep out the heat shield group in Houston, and never from anybody else inside NASA, either.
Those are the verified facts. What follows is informed speculation.
That heat protection group in Houston took it to the Artemis managers, who promptly shit-canned the notion, and ignored the NASA engineers. For them, it was either the Apollo hand-gunning that they did on Orion EFT-1, or else do exactly what they did on Artemis-1. Exclusively! And they saved tons on money doing it the Artemis-1 way. My way to "fix" this fell way outside that either/or thinking. It was NEVER considered at all!
Does that sound hauntingly familiar?
As in the NASA (and Thiokol) engineers who argued against Challenger launching so cold? And the NASA engineers who wanted an inspection of Columbia's wing, either from an EVA, or from ground photography, or both? Both sets of engineers were ignored by management, who arrogantly (and as it turned out ignorantly) assumed that "they knew better". Both times! Many years apart!
Which is EXACTLY why I SIMPLY MUST CONCLUDE that NASA management culture learned absolutely NOTHING from the loss of 2 shuttle crews! Despite the billion-$ inquests and almost-2-year-delays after each one!
GW
I just followed a link on today's AIAA Daily Launch to a Space.com article on the Orion heat shield for Artemis. This was a reporter summarizing the latest press release from NASA, explaining why they do not think Artemis-2's heat shield will show spalling. They think deleting the skip and second heat pulse will prevent the spalling and cratering.
Most of us who actually did insulation work with ablatives DO NOT think so! NASA ran some tests and went to the same kinds of analyses to "prove" they were eliminating risks, the same stuff that failed to predict what happened to Artemis-1's heat shield.
The most disturbing thing about the press release is that they intend to build the Artemis-3 heat shield the very same way as Artemis-1 and Artemis-2, despite that landing being at least 2 years away. That mission cannot fly until (1) there is a reliable lander ready to fly crewed, and (2) their Gateway space station is in place around the moon (since SLS/Orion block 1 cannot get into and out of low lunar orbit).
They have 2 years to figure out a way to put the hex into the bonded Avcoat tiles. I gave them the way to do that over a year ago, and I know for a fact their heat shield group in Houston got it! There is NO EXCUSE not to do it, or else revert to the hand-gunned Avcoat into hex cells that worked on Orion EFT-1 and all of Apollo.
This is top level managers not listening to engineers within and outside the agency, valuing money above lives. The very same flaw that killed 2 shuttle crews!
GW
I never said it didn't.
The same criteria and soil bearing strength numbers apply to both the moon and Mars. Only the local weights are different, and only by approximately a factor of 2.
What I see routinely presented to the public in these press releases is not just wrong, but wrong by orders of magnitude.
GW
I've since learned the GEM 63XL solids do not have thrust vectoring. At least that is not the problem!
From the pictures I have seen, it is impossible to really tell whether the burn through and leaking plume occurred at the nozzle-case joint, the nozzle throat region, or somewhere in-between. It did not appear to be somewhere down on the supersonic expansion bell.
There have been 4 Vulcan flights. Two have had booster anomalies attributed to nozzle problems. The other one supposedly "lost its nozzle" entirely, whatever that really means.
All I can say is that somebody somewhere in the conglomerate that is Northrup-Grumman, specifically its solid motor outfit (which is one of the old large motor contractors from when I worked in that business decades ago), is not paying enough attention to adequate thermal ablative protection in that joint or in that nozzle throat approach design.
With aluminized propellants and AP oxidizer, you are looking at chamber temperatures approaching 6000 F, with a high solids content (mostly aluminum oxide particulates and droplets, in the exiting stream. This thermal environment is way worse than anything the liquid boys have ever seen, even when they tried hydrogen with fluorine at Santa Susanna back in the 50's and early 60's. It's worse in some ways than entry heat shields. Even today.
And a lot of that knowledge was engineering art that was lost in the mass layoffs of the huge industry drawdown about 1993-1996, after the Soviet Union fell apart.
I do have to wonder, though, if NASA insisted on yet another idiotically-dangerous multiple-O-ring "seal" at the nozzle to case joint.
GW
They still have not thought rough-field landings through, because they have never, ever made one! There are few alive today who ever really did. The young crowd today seems blissfully unaware of what it takes to make a rough-field landing on soft ground, with a rocket vehicle making a powered landing. The exception seems to be Firefly. Their design, which worked on the moon, appears to be inspired by what worked decades ago with Surveyor, Apollo, Viking on Mars, etc.
Lunar regolith might be a tad stronger than Martian regolith because the lunar particles are sharp and the Martian ones are not. But the difference is of order factor 2, not orders of magnitude. Both regoliths resemble nothing so much as sand dune sand here on Earth. The presence of scattered rocks within that do not touch provides no reinforcement whatsoever. Earthly sand dune sand is listed in most Earthly foundation design references as having an allowable bearing pressure of 1 to 2 US tons per square foot = 0.1 to 0.2 MPa. The failure pressure is factor 2 to at most 2.5 above that allowable. Those are VERY LOW values to deal with!
When landing, your vehicle is lightweight, unless it is to take off again. Whatever that local-gravity weight is, there are dynamics of touchdown, and off-angle effects leading to one pad touching first. Both require factoring up the static weight by a factor of 2, for a transient landing pressure under the first pad of 4 times what the static value would be. None of those dynamics affect takeoff, the pads see only the static weight, until the vehicle lifts off. But if you are refilling propellant locally, as is often proposed now, that weight is the full max takeoff weight. Which may be some 5 to 7 times the near-empty weight at touchdowndown.
I see nothing in the HLS illustration to show any recognition of those pad bearing pressure issues. I also see nothing in the illustration to suggest any cognizance of the risk of coming down on uneven ground, which is actually depicted nearby, in the illustration, and not as bad as some actually seen during Apollo! The cg is about halfway up the vehicle, with the weight vector hanging from it. If that vector points to a location on the surface outside the polygon defined by the landing pads, the vehicle WILL fall over (and explode)! And that is if there is zero horizontal speed at the moment of touchdown! If there is horizontal speed, the lead pad will dig-in and "trip" the vehicle, even if the weight vector falls well inside the polygon.
I see no cognizance in any of the press release illustration from SpaceX (or anybody else) about these issues and risks. The same ones that twice tripped-up Intuitive Machines on the moon, and the Japanese lander, too.
The criteria to avoid this have been known since the early 1960's. (1) Min dimension of the pad polygon must be greater than the cg height. (2) Make sure the transient bearing pressure under the pads is closer to the allowable bearing pressure than the failure pressure on landing, and similar for the static bearing pressure, for takeoff. (3) Tip the pads on spring mounts inward a little, so that the lead pad won't dig in so easily if you touch down with some horizontal speed. (4) Make sure that you can hover just above the surface, for a significant time, so that whatever is controlling the vehicle during the landing can "see and avoid" uneven high-slope ground, big boulders, and cavities or other holes.
And, also since the early 1960's, if you do not have test data for the target regolith, presume it is like Earthly sand dune sand, and use the lower range of values.
Only parts of this were ever formally written down. This was mostly the engineering art that was known among those who actually did those landings decades ago. Most of them are dead now. Apparently very little of this art was passed-on, and what of it was passed-on, was lost over the time since.
GW
Spacenut:
Did you see the paragraph in the linked report that described how they fixed the leaky seals? By going to a material that better resists NTO?
Q: If the better material was known in the first place, then why did Boeing use something different? Ans -- cheaper.
What does that tell you about the thrusters? Think "cheaper" might have something to do with those troubles, too?
Certainly did for the B-737MAX debacle. And there's a couple of other airplanes Boeing is in deep kimchee with troubles. One is the 777-X, the other is the new USAF tanker. All built by the same Boeing. As is the SLS core stage.
GW
Should not be a surprise. Money talks.
His SpaceX and Tesla baseline incomes come from government contracting. He is being paid to land people on the moon, if he can. He IS NOT being paid by anybody to go to Mars. If he defaults on his moon landing commitments, he will lose far too much credibility to be a major space contractor anymore, to NASA or DOD. That is "default" by not doing what he promised to attempt. "Failure" by not being able to do it, is a different issue.
GW
Whether the Starship really will be cheaper to launch than a Falcon-9 or Falcon-Heavy is still debatable. That answer CANNOT be known until experimental flight testing and design changes are done. Be aware that SpaceX is talking about an even larger Block 4, while the somewhat-larger Block 3 still has yet to fly at all! Block 2 pretty much worked mostly right in the last two experimental flight tests.
The change from "block to block" is somewhat closer to a new design than just a minor variation on an old one. Always has been, always will be. That's just the inherent nature of vehicle developments and the experimental flight tests to "prove them out".
The cost question also depends upon how big your crews are. If 3-7, a Falcon-9 or Falcon-Heavy and Dragon makes very good sense at ~ $90 M/(3 to 7) for $12-30M per seat. That may look like a high seat price, but if all you are flying is a small crew, that's a small portion of your overall mission cost.
If you are flying a big crew 20+, then you would need more than one Dragon, and something big like Starship makes a lot more sense to shoot them all up there in one launch. Per seat price? Nobody can know yet!
Early exploration does NOT need big crews! That comes later when you do some sort of experimental base where you can prove out and verify that you really can start to "live off the land". That capability will NOT be the case in the first explorations, and it will still NOT be the case, when you first set up that experimental base. That is inherent.
Let's just say I am extremely skeptical of the "$10 M launch cost" some have been bandying about for Starship. I rather think $100 M is a better wild guess! But it is ONLY a wild guess! NOBODY has anything any better than a wild guess! Not Musk, not ANYBODY! $100M / (20+) is under $5M per seat. Still a small portion of any real mission cost.
GW
from the AIAA email newsletter "Daily Launch" for Thurs 2-12-2026:
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Spaceflight Now
Vulcan suffers solid rocket booster problem during USSF-87 launch
ULA said an issue affected one of the four solid rocket boosters that helped propel its Vulcan rocket into space Thursday on a mission for the United States Space Force. Despite the problem the rocket, making only its fourth flight, continued on its planned trajectory, the company said. The rocket thundered away from pad 41 at Cape Canaveral Space Force Station at 4:22 a.m. EST but less than 30 seconds into the flight, there appeared to be a burn through of one of the nozzles on a Northrop Grumman-built graphite epoxy motor (GEM) 63XL solid rocket boosters (SRBs).
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My take: This is not the first such issue with this solid's nozzle. Somebody is not paying enough attention to the highly-erosive ablative insulation environment approaching and through the throat of a solid rocket nozzle. Whether out of ignorance or being too cheap is not knowable from this. If you complicate the design by adding thrust vector capability, you raise the odds of such burn-through or nozzle-loss failures, I do know that!
Reinforced rubbers do not hold up in the high shear flow. It takes hard silica phenolic plus a graphite throat insert, and that does NOT deflect to vector the nozzle bell! You have to deflect to vector out near the full case diameter where the scrubbing fluid shear, to get the rubber to hold up. Even then, you must do that deflection while still holding the full compressive thrust load aligned where you want it. Vectoring a solid rocket is a bad thing to attempt.
GW
Here's variation on what KBD512 suggested. Put together a reusable electric-drive transfer vehicle that is 2-way capable and reusable. Do this in LEO where humans can build the thing from docked modules like the ISS was. Load it there, too. Spiral it out unmanned to an orbit outside the outer Van Allen belt, say 60,000 or 70,000 km out. That's where you send the crew to board it, and to recover them coming home.
But, you do not need a Starship to do that!
Falcon-Heavy with a Dragon has plenty of dV to take a crew of 4 to 7 (there are up to 7 seats in the capsule) across the belts quickly to board the ship there. It's not rated for usefulness beyond about 6 months, but I bet it could be. That's your emergency bailout capsule coming home. Falcon-Heavy also has the dV to fly up empty and bring a returning crew of 4 to 7 home. Once the reusable orbit-to-orbit craft is built, all you have to do is resupply it. It might be just about as easy to spiral in to LEO uncrewed for that, since the dV to LEO is only about 7.8 km/s + loss coverage, while the dV to 60-70,000 km circular is near 10.5 km/s + loss coverage + circularization.
Starship might be more appropriate for larger crews, but remember, its first and best role is as a transport to LEO, without any refueling tanker flights at all. Why use a big vehicle when a smaller one will serve? Just build some sort of tug based in LEO, which would require a lot fewer tanker flights to keep supplied if it is smaller than a Starship, and let it push a small hab about the size of Skylab at the most, for transporting crews quickly across the Van Allen belts. Just keep the tug and hab in LEO, based there.
The orbit-to-orbit interplanetary (or lunar) transport, ought to have spin gravity. Crews will be out there a long time, especially spiralling in and out at the moon or Mars. It will need a solar flare storm shelter, too. The rest of that stuff (the ordinary rockets and capsules) is short-term zero-gee stuff.
If we are going to send Dragons (or similar) out beyond the Van Allen belts, we will need better forecasting for solar weather, lest an unanticipated event kill a crew needlessly. And we will need a station in LEO at which to assemble things and to fill them up with mission propellants. It would seem likely one station could perform both functions.
All that being said, I am wondering why we would even consider sending tall, narrow Starships with landing legs and pads that are inherently too small, to the moon or Mars, before big, flat, hard-surfaced paved landing pads have been constructed. That's the fatal shortfall in Musk's vision. And don't kid yourself, it is a fatal shortfall.
GW
Same is true returning home, too! You have to cross the Van Allen belts fast, then either decelerate into orbit or do a free return. You cannot kill your crew trying to spiral-in slowly with electric.
GW
Void:
You cannot use electric propulsion to take humans from Earth orbit to Mars. Electric propulsion is extremely low thrust/low acceleration. It takes a month or two, maybe three, to spiral out through the Van Allen radiation belts. That long an exposure is a fatal dose for any crew.
You need to cross the Van Allen belts in only a day or so, something known since just before Apollo. That takes high-thrust chemical or nuclear rockets to get significant vehicle acceleration. You need around 0.1 to 0.5-ish vehicle gee capability to make that happen. Electric is typically under 1/10,000 gee.
The departure dV from Earth orbit onto the transfer trajectectory to Mars (or the moon, or anywhere else out there) is the biggest dV by far! If you have to use chemical (or nuclear) for that, then why bother with electric?
My point: there is a whale of a lot more to worry about than just Isp, when looking at travel to Mars, or the moon, or anywhere else. Focusing on only one aspect is a guaranteed trip down the wrong path.
GW