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Exoscientist

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  1. Your mentioning of air brakes brings up another advantage of the rotatable rear flaps: in the fully rotated position they would create a great deal of drag to further slow down the descent. See the image displayed on the video start up screen here: That would be like a parachute in regards to slowing the spacecraft down. It might be SpaceX wants to give the Starship the ability to also land payload, but that amount of reserve propellant shouldn’t be used for all launches when it is not needed. That is unnecessarily subtracting from its normal payload. By the way, the “dry mass” Elon has been quoting for the reusable Starship also seems excessive. The latest is 120 tons(!) That’s nearly 3 times the bare dry mass of the expendable version, i.e., no reusability systems, of only 45 tons. That seems an excessive weight added for reusability. Bob Clark
  2. By the way the two SSTO projects compared by Burnside would be a good Kerbal design project to see if they could really get that much payload to orbit. Bob Clark
  3. Dense propellant first stages already get well better than 10% structural fraction, i.e., 90% propellant fraction. For a long time it was felt a SSTO had to use a light fuel such as hydrogen because it had the highest Isp, ca. 450 s. But more careful analysis showed actually dense propellants would be better for a SSTO because a big component of the dry mass of a rocket is tankage and dense propellants, such as kerosene or methane, can carry more fuel for the same size tanks.( More on this below) As an example of a first stage with very high propellant fraction, or said the other way, low structural fraction, take a look at the Falcon 9 first stage: Type Falcon 9 FT Stage 1 Length 42.6 m (47m w/ Interstage) Diameter 3.66 m Inert Mass ~22,200 kg (est.) Propellant Mass 411,000 kg (According to FAA) Fuel Rocket Propellant 1 Oxidizer Liquid Oxygen LOX Mass 287,430 kg RP-1 Mass 123,570 kg LOX Volume 234,700 l RP-1 Volume 143,900 l LOX Tank Monocoque RP-1 Tank Stringer & Ring Frame Material Aluminum-Lithium Interstage Length 4.5 m (est.) Guidance From 2nd Stage Tank Pressurization Heated Helium Propulsion 9 x Merlin 1D Engine Arrangement Octaweb https://spaceflight101.com/spacerockets/falcon-9-ft/ This is a structural fraction of 22,200/(22,200 + 411,000) = .05, 5%, so a propellant fraction of 95%. However, the sea level Merlins don't have a very good Isp at vacuum ~311 s, so as an SSTO would get minimal payload to orbit , if any. Here's an analysis that shows a dense propellant SSTO can carry more payload to orbit than hydrogen fueled: From: [email protected] (burnside) Newsgroups: sci.space.policy Subject: A LO2/kerosene SSTO rocket design, w/o AOL Date: 2 Feb 1997 15:11:33 GMT A LO2/Kerosene SSTO Rocket Design (long) Mitchell Burnside Clapp Pioneer Rocketplane (view with a fixed pitch font such as courier or monaco) Abstract The NASA Access to Space LO2/hydrogen single stage to orbit rocket was examined, and the configuration reaccomplished with LO2/kerosene as the propellants. Four major changes were made in assumptions. First, the aerodynamic configuration was changed from a wing with winglets to a swept wing with vertical tail. The delta-V for ascent was as a result recalculated, yielding a lower value due to different values for drag and gravity losses. The engines were changed to LO2/kerosene burning NK-33 engines, which have a much lower Isp than SSME-type engines used in the access to space study, but also have a much higher thrust-to-weight ratio. The orbital maneuvering system on the Access to Space Vehicle was replaced with a pump-fed system based on the D-58 engine used for that purpose now on Proton stage 4 and Buran. Finally, the wing of the vehicle was allowed to be wet with fuel, which is a reasonable practice with kerosene but more controversial with oxygen or hydrogen. Additionally, in order to reduce the technology development needed, the unit weights of the tankage were allowed to increase by 17 percent. After the design was closed and all the weights recalculated, the empty weight of the LO2/kerosene vehicle was 35.6% lighter than its hydrogen fuelled counterpart. Introduction NASA completed a study in 1993 called Access to Space, the purpose of which was to consider what sort of vehicle should be operated to meet civil space needs in the future. The study had three teams to evaluate three different broad categories of options. The Option 3 team eventually settled on a configuration called the SSTO/R. This vehicle was a LO2/hydrogen vertical takeoff horizontal landing rocket. The mission of the Access to Space vehicle was to place a 25,000 pound payload in a 220 n.mi. orbit inclined at 51.6 degrees. The vehicle had a gross liftoff weight of about 2.35 million pounds. The thrust at liftoff was 2.95 million pounds, for a takeoff thrust to weight ratio of 1.2. The empty weight of the vehicle was 222,582 pounds, and the propellant mass fraction (defined here as [GLOW-empty]/GLOW) was 90.5%. Main power for this vehicle was provided by seven SSME derivative engines, with the nozzle expansion ratio reduced to 50. This resulted in an Isp reduction from 454 to 447.3 seconds. Each engine weighed 6,790 lbs, for an engine sea level thrust to weight ratio of 62. Aerodynamically the vehicle was fairly squat, with a fineness ratio (length:diameter) of 5. The overall length of the vehicle was 173 feet and its diameter was 34.6 feet. It had a single main wing (dry of all propellants) of about 4,200 square feet total area, augmented by winglets for directional control at reentry. The landing wing loading was about 60 lb/ft2. The oxygen tank was in the nose section. The payload was mounted transversely between the oxygen and hydrogen tanks, and was 15 feet in diameter and 30 feet long. This design exercise was among the most thorough ever conducted of a single stage to orbit LO2/LH2 VTHL rocket. It was probably the single greatest factor in convincing the space agency that single stage to orbit flight was feasible and practical, to borrow from the title of Ivan Bekey's paper of the same name. A LO2/kerosene alternative A number of people have been asserting for some time that higher propellant mass fractions available from dense propellants may make single stage to orbit possible with those propellants also. The historical examples of the extraordinary mass fractions of the Titan II first stage, the Atlas, and the Saturn first stage are all persuasive. Further, denser propellants lead to higher engine thrust to weight ratios, for perfectly understandable hydraulic reasons. It has not usually been observed that higher density also leads to significant reductions in required delta-v. There are two major reasons that this is so. First, the reduction in volume leads to a smaller frontal area and lower drag losses. The second, and more significant, reason is that the gravity losses are also reduced. This is because the mass of the vehicle declines more rapidly from its initial value. The gravity losses are proportional to the mass of the vehicle at any given time, and hence the vehicle reaches its limit acceleration speed faster. NASA itself has implicitly recognized this effect. When the Access to Space Option 3 team examined tripropellant vehicles, the delta-v to orbit derived from their work was 29,127 ft/sec, for precisely the reasons described in the previous paragraph. This compares to a delta-v of 30,146 ft/s for the hydrogen-only baseline, as reported in a briefing by David Anderson of NASA MSFC dated 6 October 1993. To be clear, these delta-v numbers include the back pressure losses, so that no "trajectory averaged Isp" number is used. They did not, however, report any results for kerosene-only configurations. To come to a more thorough understanding of the issues involved in SSTO design, I have used the same methodology as the Access to Space team to develop compatible numbers for a LO2/kerosene SSTO. There are four major changes in basic assumption between the two approaches, which I will identify and justify here: 1: The ascent delta-v for the LO2/kerosene vehicle is 29,100 ft/sec, rather than 29,970 ft/sec. The reason for this is argued above, but I ran POST to verify this value, just to be sure. The target orbit is the same: 220 n.mi. circular at 51.6 degrees inclination. The detailed weights I have for the NASA vehicle are based on a delta-v of 29,970 ft/sec rather than the 30,146 ft/sec reported in Anderson's work, but I prefer to use the values more favourable to the hydrogen case to be conservative. The optimum value of thrust to weight ratio turns out to be slightly less than the hydrogen vehicle: 1.15 instead of 1.20. 2: The aerodynamic configuration is that of Boeing's RASV. Without arguing whether this is optimal, the fineness ratio of 8.27 and large wing lead to a much more airplane-like layout, better glide and crossrange performance, and reduced risk. The single vertical tail is simpler and safer than winglets as well. Extensive analysis has justified the reentry characterisitics of this aircraft. The wing is assumed to be wet with the kerosene fuel, as is common on most aircraft. The fuel is also present in the wing carry-through box. The payload is carried over the wing box, and the oxidizer tank is over the wing. This avoids the need for an intertank, which in the NASA Access to Space design is nearly 6,600 pounds. 3. The main propulsion system is the NK-33. The engine has a sea level thrust of 339,416 lbs, a weight of 2,725 lbs with gimbal, and a vacuum Isp of 331 seconds. Furthermore, it requires a kerosene inlet pressure of only 2 psi absolute, which dramatically reduces the pressure required in the wing tank. It also operates with a LO2 pressure at the inlet of only 32 psi. The comparable values for the SSME are about 50 psi for both propellants. This will have a substantial effect on the pressurization system weight. 4. The OMS weight is based on the D-58 engine. This engine is used for the Buran OMS system and the Proton stage 4. As heavy as it is the Isp is an impressive 354 seconds. NASA's vehicle used a pressure fed OMS, which is a sensible design choice if you're stuck with hydrogen and you wish to minimize the number of fluids aboard the vehicle. But because both oxygen and kerosene are space-storable, there is no reason to burden the design with a heavy pressure fed system. Using the same methodology for calculating masses, and accepting the subsystems masses as given in the Access to Space vehicle, a redesign with oxygen and kerosene was accomplished. The results appear in Table 1. Table 1: Access to Space vehicle and LO2/kerosene alternative Name O2/H2 LO2/RP Wing 11,465 11,893 lb Tail 1,577 1,636 lb Body 64,748 33,741 lb Fuel tank 30,668 - lb Oxygen tank 13,273 17,271 lb Basic Structure 14,610 10,274 lb Secondary Structure 6,197 6,197 lb Thermal Protection 31,098 21,238 lb Undercarriage, aux. sys 7,548 5,097 lb Propulsion, Main 63,634 36,426 lb Propulsion, RCS 3,627 1,234 lb Propulsion, OMS 2,280 823 lb Prime Power 2,339 2,339 lb Power conversion & dist. 5,830 5,830 lb Control Surface Actuation 1,549 1,549 lb Avionics 1,314 1,314 lb Environmental Control 2,457 2,457 lb Margin 23,116 16,105 lb Empty Weight 222,582 141,682 lb Payload 25,000 25,000 lb Residual Fluids 2,264 1,911 lb OMS and RCS 1,614 1,261 lb Subsystems 650 650 lb Reserves 7,215 8,895 lb Ascent 5,699 7,587 lb OMS 679 541 lb RCS 837 767 lb Inflight losses 13,254 17,445 lb Ascent Residuals 10,984 15,175 lb Fuel Cell Reactants 1,612 1,612 lb Evaporator water supply 658 658 lb Propellant, main 2,054,612 3,034,972 lb Fuel 293,604 843,048 lb Oxygen 1,761,008 2,191,924 lb Propellant, RCS 2,814 2,556 lb Orbital 2,051 1,756 lb Entry 763 800 lb Propellant, OMS 19,357 15,452 lb GLOW 2,347,098 3,246,156 lb Inserted Weight 292,486 211,185 lb Pre-OMS weight 271,482 186,152 lb Pre-entry Weight 252,125 170,700 lb Landed Weight 251,362 169,900 lb Empty weight 222,582 141,682 lb Sea Level Thrust 2,816,518 3,733,080 lb Percent margin 11.6% 12.8% Assumed Isp(vac) 447.3 331.0 s Ascent Delta-V 29,970 29,100 ft/s OMS delta-V 1,065 987 ft/s RCS delta-V 108 107 ft/s Deorbit Delta-V 44 53 ft/s Reserves 0.28% 0.25% lb/lb Residuals 0.53% 0.50% lb/lb Wing Parameter 4.56% 7.00% lb/lb TPS parameter 12.37% 12.50% lb/lb Undercarriage parameter 3.00% 3.00% lb/lb Wing Reference Area 4,189 5,528 ft2 Density of fuel 4.4 50.5 lb/ft3 Density of oxygen 71.2 71.2 lb/ft3 Volume of fuel 66,276 16,694 ft3 Volume of oxygen 24,733 30,785 ft3 Fuel tank parameter 0.42 - lb/ft3 Oxygen tank parameter 0.48 0.56 lb/ft3 Some discussion of the results and justification is in order. The wing is about 40 percent heavier as a percentage of landed weight than for the hydrogen fueled baseline. When considered as a tank, it is about 60 percent heavier for the volume of fuel it encloses. Its weight per exposed area is about the same and the wing loading is half at landing. No benefit is taken explicitly for the lack of a requirement for kerosene tank cryogenic insulation. The tail is assumed to have the same proportion of wing weight for both cases. This is conservative for the kerosene wehicle because its single vertical tail is structurally more efficient. The body of the kerosene vehicle has three components. The oxidizer tank has an increased unit weight of about 17 percent. This is done in order to avoid the need for aluminum-lithium, which was assumed in the Access to Space vehicle. The basic structure group is unchanged, except that the intertank is deleted and the thrust structure is increased in proportion to the change in thrust level. The secondary structure group is mostly payload support related, and was not changed. The thermal protection group is in both cases about 12.5% of the entry weight. This works out to 1.107 lbs/ft2 of wetted area for the kerosene vehicle, which is common to many SSTO designs. The undercarriage group is 3% of landed weight for both vehicles. There is no benefit taken for reductions in gear loads for the kerosene vehicle due to lower landing speed and lower glide angle at landing. The main propulsion group includes engines, base mounted heat shield, and pressurization/feed weights. The engines are far lighter for their thrust than SSME derivatives. The pressurization weights are reduced in proportion to the pressurized volume for the kerosene vehicle. No benefit is taken for reduced tank pressure. Here is as good a place as any to point out the erroneous assertion that increased hydrostatic pressure is going to lead to increased tankage weights. There is no requirement for a particular ullage pressure except for the need to keep the propellants liquid. It is the pressure at the base of the fluid column rather than the top of the column that is of engineering interest. The column of fluid exerts a hydrostatic load on the base of the tank, but this load does not typically exceed the much more adverse requirement for engine inlet pressurization. For the kerosene vehicle, the hydrostatic load at the base of the oxygen tank is 49 psi, which is compatible with the pressures normally seen in oxygen tanks for rocket use. The load declines after launch because the weight goes down faster than the acceleration goes up. The bottom line here is that dense propellants may require you to alter a tank's pressurization schedule, but not to overdesign the entire tank. Structures are sized by loads and tankage for rockets is sized principally by volume, and if the vehicle is small, by minimum gauge considerations. This is not completely true for wet wings, however, as discussed previously. In this particular example, there is no need for high pressure in the wing tank either, because of the low inlet pressure required by the NK-33. The OMS group is the only other major change, as discussed above. The reliable D-58 engine has been performing space starts for decades and will serve well here. The acceleration available from the OMS is about 0.12 g, which is standard. All the other weights are pushed straight across for the most part. A brief inspection suggests that this is very conservative. Control surface actuation requirements are certainly less, electrical power requirements less, much better fuel cells available than the phosporic acid type assumed here, and reduced need for environmental control. Nonetheless, rather than dispute any of these values it is easier simply to accept them. The margin is applied to all weight items at 15% execpt for the engine group at 7.5%. The justification for this is that the main and OMS engine weights are known to high accuracy. The vehicle has an overall length of 1955 inches, and a diameter of 236.4 inches. The wing has a leading edge sweep of 55.5 degrees and a trailing edge sweep of -4.5 degrees. Its reference area is 5,632 square feet, of which 3,992 square feet is exposed. The wing encloses 16,694 ft3 of fuel, with a further 5% ullage. The carry-through is also wet with fuel. The wing span is 1293 inches, and the taper ratio is 0.13. The payload bay has a maximum width and height of 15 feet. It sits on top of the wing carry through box. The thrust structure from the engines passes through and around the payload bay to the forward LO2 tank. The payload bay is 30 feet in length. It has a pair of doors, the aft edge of which is just forward of the vertical tail leading edge. The engine section encloses 11 NK-33 engines, with a 4 - 3 - 4 layout. The engines are each 12.5 feet long, and additional structure and subsystems take up another 6.5 feet. The oxygen tank comprises the forward fuselage, which encloses 30,785 ft3 of oxygen, with a further 5% ullage. The length of the tank is about 100 feet. The ventral surface of the tank is moderately flattened as it moves aft, to fair smoothly with the wing lower surface. This flattening reduces its length by about 5% with respect to a strictly cylindrical layout. The aft edge of the oxygen tank is about even with the forward payload bay bulkhead. A compartment of about 13.9 feet provides room for some subsystems and a potential cockpit in future versions. Conclusion The methods of the NASA Access to Space study were used to design a single stage to orbit vehicle using existing LO2/kerosene engines. An inspection of the final results shows that the vehicle weighs about 36.5% less than its hydrogen counterpart, with reductions in required technology level and off the shelf engines. The center of mass of the vehicle is about 61% of body length rather than 68% for the Access to Space vehicle, which should improve control during reentry. The landing safety is considerably improved by lower landing speed and better glide ratio. Structural margins are greater overall. The vehicle designed here appears to be superior in every respect: smaller, lighter, lower required technology, improved safety, and almost certainly lower development and operations cost.
  4. If you do the calculation the amount of propellant used for landing is far less than 30 tons. Robert Clark
  5. I’m suggesting they should do more than just fold up against the sides to varying degrees. They should also rotate forward and backwards. This would provide better control of the pitching moment. The Chinese 5th-generation fighter J-20 is another example with extreme rotatable control surfaces: https://www.reddit.com/r/aviation/comments/ynb0am/chinese_j20_fighter/?utm_source=share&utm_medium=ios_app&utm_name=iossmf The 30 ton propellant kept on reserve for landing is far more than needed for the landing burn. The landing burn might require only ~250 m/s delta V when you consider the low terminal velocity of ~90 m/s plus gravity drag: This would require a small amount of propellant to be burned considering the Raptor has 330s sea level Isp. And Elon has acknowledged the 30 tons ballast in the nose is to help stability. Robert Clark
  6. The whole root of the problem is that Boeing proposed this over expensive upper stage that NASA balked at paying for. But there really was no need for it to be that expensive: Why does the Boeing Exploration Upper Stage(EUS) cost so much? https://exoscientist.blogspot.com/2022/11/why-does-boeing-exploration-upper.html Robert Clark
  7. Actually, cislunar space includes all the space with the vicinity of the Earth and Moon. Robert Clark
  8. It looked like the Air Force buzzed the Starbase with their most advanced fighter: Actually, they were just practicing for a nearby air demonstration. But this is rather ironic because I advise rather than using 30 tons of ballast to maintain CG ahead of CP position for stability, thus subtracting that amount from payload, use computerized control to maintain stability as commonly used on fighter jets. This is taken to an extreme level with the F-22 with its large control surfaces at the rear of the plane: The flaps on the Starship could serve the same function. Instead of just letting them fold upwards to various degrees against the sides of the rocket, allow them also to rotate forwards and backwards as done with the F-22. Robert Clark
  9. This article suggests nuclear thermal propulsion could get to the Moon in hours: The US military is getting serious about nuclear thermal propulsion “Activity in cislunar space is expected to increase considerably in the coming years.” ERIC BERGER - 6/15/2020, 8:18 AM “With the DRACO program, the US Defense Department could potentially move large satellites quickly around cislunar space. For example, moving a 4-ton satellite from point A to point B might take about six months with solar electric propulsion, whereas it could be done in a few hours with nuclear thermal propulsion.” https://arstechnica.com/science/2020/06/the-us-military-is-getting-serious-about-nuclear-thermal-propulsion/ Robert Clark
  10. I was surprised when reading this: DARPA moving forward with development of nuclear powered spacecraft. by Sandra Erwin — May 4, 2022 https://spacenews.com/darpa-moving-forward-with-development-of-nuclear-powered-spacecraft/ The article discusses that DARPA is funding nuclear powered propulsion to cislunar space. This is space in the vicinity of the Moon. The only reason why you would want it nuclear powered is you want to get there rapidly, in a matter of hours instead of days. What military purpose could there be for getting to the Moon in hours? Robert Clark
  11. An upper stage for the SLS could be made by combining two Centaur V’s. This might allow a single launch Artemis lunar landing architecture, no SpaceX Starship launches required: Possibilities for a single launch architecture of the Artemis missions. http://exoscientist.blogspot.com/2022/10/possibilities-for-single-launch.html However, I used the Silverbirdastronautics.com payload estimator that has rather large error bars. I’d like to see a Kerbal Real Solar System mod to get a better estimate. Robert Clark
  12. The X-33 tiles were well tested, as well were as the Starship tiles: 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/20040095922/downloads/20040095922.pdf The metallic tiles had better resistance to impact and rain than the ceramics at about the same weight, and would not require water proofing. They also would have closer thermal expansion properties to the steel Starship. Fig.3 Layered metallic sheeting separated by insulation. Fig.21 Metallic TPS at same weight of ceramic tiles, ~10kg/m^2
  13. Problems with the heat tiles discussed here: Perhaps they should try the metallic heat tiles developed for the X-33? About same weight and thermal protection as ceramics , but were screwed on with bolts, and have superior impact and rain resistance. Robert Clark
  14. Actually, it was. And SpaceX went so far as detailing how long they expected it would take to make these total 16 launches per mission: SpaceX CEO Elon Musk details orbital refueling plans for Starship Moon lander. By Eric Ralph Posted on August 12, 2021 First, SpaceX will launch a custom variant of Starship that was redacted in the GAO decision document but confirmed by NASA to be a propellant storage (or depot) ship last year. Second, after the depot Starship is in a stable orbit, SpaceX’s NASA HLS proposal reportedly states that the company would begin a series of 14 tanker launches spread over almost six months – each of which would dock with the depot and gradually fill its tanks. … In response to GAO revealing that SpaceX proposed as many as 16 launches – including 14 refuelings – spaced ~12 days apart for every Starship Moon lander mission, Musk says that a need for “16 flights is extremely unlikely.” Instead, assuming each Starship tanker is able to deliver a full 150 tons of payload (propellant) into orbit after a few years of design maturation, Musk believes that it’s unlikely to take more than eight tanker launches to refuel the depot ship – or a total of ten launches including the depot and lander. https://www.teslarati.com/spacex-elon-musk-starship-orbital-refueling-details/ Blue Origin didn’t raise this objection arbitrarily about the number of launches. It’s because it was in the actual proposal SpaceX made to NASA, and on which SpaceX based their charge to NASA for their mission plan. If they now claim the number of refueling launches will only be 4, then they should amend the amount they are charging NASA for their plan. Robert Clark
  15. Actually, the mass of the Starship as an expendable is the entire point of the matter. From that you see you can get quite high payload as an expendable rocket. In fact, it’s in the same percentage of gross mass range of other currently in use expendable rockets. About the SpaceX lunar plan, it is an extremely important thing to know if NASA, and the U.S. tax payers, are getting jobbed if the original ~$3 billion price was based on ~16 flights per mission when it will only take ~4 launches per mission. Robert Clark
  16. The point I’m making is that in reading the proposal to NASA the ~16 launches per mission were built into the contract. Suppose instead that in the original contract they said it would only take ~4 launches per mission using the stripped down 40 ton mass of the lander. Then that contract price should have been less. Robert Clark
  17. Starship will use 30 tons of propellant as ballast in the nose on landing to help maintain stability: Note this will only be used on landing so would not be included for accounting of the mass as an expendable stage. The nose cone and nose cone barrel weighs ~17 tons. SpaceX will also save about ~7 tons off the tank mass on shaving down the tank thickness from 4mm to 3 mm. That's 54 tons off the 120 tons often cited for the Starship "dry mass", bringing it down to 66 tons. The attempt here is to estimate the dry mass as an expendable stage. But there's still the mass of the TPS, landing legs, and flaps. For the landing legs we can estimate that as following the Falcon 9 booster model of ~10% of the stage dry mass, so ~12 tons for the legs on Starship, when SpaceX is going by 120 tons as the dry mass. So we're now down to ~54 tons as an expendable dry mass. And the TPS? On the NasaSpaceflight.com forum that has been estimated as from 5 to 10 tons. So the expendable dry mass might be down to 44 tons. Then there is still the mass of the flaps that needs to be subtracted off for the expendable dry mass. Quite conceivable then that the expendable dry mass might be less than 40 tons. This is important to know because this is the form of the stage that would be used as the lunar lander, as SpaceX is intending it to be expendable in their lunar plan. Note also the payload section would also be removed in this configuration, with the lunar crew module being directed attached to the tank section via an adapter. Elon acknowledge this configuration would be much lighter and would therefore take fewer refuelings: If the Starship expendable without payload section or reusability systems would only mass ~40 tons then that needs to be acknowledged by SpaceX if that means the launches that needs to be paid for the U.S. taxpayers could be cut from ~16 to ~4. See discussion here: https://exoscientist.blogspot.com/2022/09/the-nature-of-true-dry-mass-of-starship.html Robert Clark
  18. With the hurricane threatening to force another rollback, leaving you with only one rollback left before the $2 billion vehicle has to be scrapped for parts, this becomes of increasing importance to find a solution. That NASA is dithering on the decision makes it clear the limit on rollback is a serious consideration. Even this fan solution if it works as a stop gap measure might be acceptable. One possible problem though is it might work too well. It might obscure the fact that a serious leak is occurring. Robert Clark
  19. Thanks for that. Surprising scientifically that the same material can be magnetic or non-magnetic depending on how you work it. Robert Clark
  20. This NASAspaceflight video also speculates SpaceX might be considering an expendable Starship version, with no heat shield or flaps: Robert Clark
  21. Or could some very strong neodymium magnets be included in the tiles so they are kept on magnetically against the steel airframe? Robert Clark
  22. Saw this discussed on the NASASpaceflight forum: ____________________________________________________________ Don't know if you know about it but this site has some useful articles https://www.nasaspaceflight.com/2022/09/starship-next-phase-of-testing/ https://www.nasaspaceflight.com/2022/08/booster-7-additional-tests/ ____________________________________________________________ https://forum.nasaspaceflight.com/index.php?topic=50748.msg2409134#msg2409134 With every test firing of the Starship some tiles pop off. There is some speculation Starship may be launched on the first test flight without the tiles, obviously in an expendable mode. Here’s a picture of the pins that hold on the tiles: It seems you could get stronger type pins than that. For example you could use spring-loaded wing nuts: Apparently SpaceX does not want to glue the tiles on like what happened with the shuttle because it takes too much maintenance time for replacement and refurbishment. But with the wing nuts you can adjust the strength to be removable but strong enough to hold on during flight. By the way, according to the discussion in that thread on NasaSpaceflight.com SpaceX is looking to improve on the tiles they are using. I advise considering mathematician and engineers GW Johnson’s ultra lightweight, reusable TPS material: Robert Clark
  23. The current plan is to not reuse the lander because it would take too much propellant weight to return it. The Starship without payload section could get twice the payload as an expendable SSTO as the Falcon 9 as an expendable. Here's a more representative image of the Starship as a lunar lander without a crew capsule(Orion): Robert Clark
  24. Note that SpaceX does not need the passenger section, or thermal protection, or upper and lower flaps for the lunar lander version of the Starship. In fact it probably can even do without a fairing, with the Orion capsule attached above the tank with an adapter for its smaller diameter. SpaceX wants to argue it might only take 4 refuelings for the lunar mission. Then they should acknowledge that this form of the Starship with no passenger section, i.e., with the upper rings above the tank removed, might only mass 50 to 60 tons. But that would mean this format of the Starship could do 40 to 50 tons payload to LEO as an expendable SSTO. Then any reasonable estimate for reusability systems would still allow significant payload as a reusable SSTO. Robert Clark
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