Exoscientist

Think that’s an April Fools joke. Bob Clark

Avi Loeb has received a $1.5 million private grant to conduct his search for the interstellar meteorite fragments beneath the ocean’s surface: https://avi-loeb.medium.com/a-gift-from-a-silver-star-af2993e0169a Bob Clark

I agree with you number of launches is the more important parameter. Likely prices will decrease and as you said if it decreases by half, that means 2,000 Falcon 9 equivalent launches. But that means an even larger market for other launch companies to get into and even greater necessity for them to get into reusables. Robert Clark

Global Space Launch Services Market is projected to reach at a market value of US$ 47.6 Billion by 2030: Visiongain Research Inc October 05, 2021 09:33 ET | Source: Visiongain Ltd https://www.globenewswire.com/news-release/2021/10/05/2308874/0/en/Global-Space-Launch-Services-Market-is-projected-to-reach-at-a-market-value-of-US-47-6-Billion-by-2030-Visiongain-Research-Inc.html Several independent market research surveys have put the space launch market at about the ~$40 billion range by ~2030. That amounts to about 1,000 reusable Falcon 9’s. That’s a large market for other launch companies to also take part in. BUT they have to use reusables to stay competitive with SpaceX. Otherwise, they’ll also go the way of ULA having to ask for buyers to avoid bankruptcy. In that regard ULA and Arianspace should upgrade their coming launchers, Vulcan Centaur for ULA and Ariane 6 for Arianespace, to add an additional engine to their first stages. As it is now, neither can lift off without using solid side boosters. But the space shuttle showed solid side boosters do not save costs on reuse, while SpaceX showed all-liquid rockets do. Robert Clark

I had estimated that the Vulcan Centaur given three BE-4's could get ~27 tons to LEO. But that was using a weight-optimized Vulcan dry mass estimate. However, I tried the SilverbirdAstronautics.com estimator on the Vulcan with the usual two BE-4's in the two side booster configuration and it gave results well above what was given on the ULA page on the Vulcan Centaur. So I think my dry mass estimate was too optimistic, i.e., too low. What I used was 25,000 kg for the Vulcan booster dry mass given three engines. The basis for that estimate is the weight-optimized Falcon 9 booster, using aluminum-lithium for the tanks, gets about a ~25 to 1 mass ratio. Then since methalox is at about 80% density of kerolox it should get about 20 to 1 mass ratio. This would give the Vulcan booster a dry mass of 25,000 kg. But that dry mass results in a badly overestimated V/C with two side boosters payload compared to that stated by ULA. So here's a another stab at a dry mass estimate. Make a comparison to the Atlas V booster mass ratio. This is hardly weight-optimized using just standard aluminum for the tanks as does the Vulcan stage. The Atlas V booster has a mass ratio of only 15 to 1. Then the Vulcan with methalox at only 80% the density of keralox might get ~12 to 1 mass ratio. So the Vulcan dry mass might be as high as 45,000 kg. This results in a much closer Silverbirdastronautics.com payload estimate to the ULA numbers. Bob Clark

Not like I know him. I asked via his twitter account: Bob Clark

“PD”? I use the numbers provided by the companies when provided. But some key numbers often are not provided such as stage dry masses, and often propellant masses. I asked Tory Bruno once about this and he said it was for competitive advantage. Bob Clark

Not lying. Just puzzling. Why not just give give it 3 BE-4’s so can launch fully loaded without side boosters? With side boosters you then get even higher payload. I used the payload estimator of Silverbirdastronautics.com to estimate a LEO payload of ~27 tons using 3 engines on the first stage w/o side boosters. This would beat the ~22 ton payload of the Falcon 9. BTW, it could then also do SSTO without the Centaur V upper stage. ;-) Robert Clark

Can Vulcan Centaur launch on just two BE-4 engines without side boosters? This Blue Origin page says Vulcan Centaur can get 10.8 tons payload to LEO on its two BE-4 engines without using side boosters: But we can calculate the mass of the Vulcan booster to show its beyond that of the thrust of the two BE-4 engines. See here: According to the heights given along the side of the image on the right, the bottom of the booster tank is at about 20 feet high, and its top at about 110 feet, for a length of 90 feet, 27.4 meters. The diameter is 5.4 meters, for a radius of 2.7 meters. Then the volume is π*2.72*27.4 = 627.5 m3 . The density of methalox or methanolox at the mixture ratio of liquid O2 to CH4 of 3.6 is about 800 kg/m3. Then the mass of the propellant is 502,000 kg. But the sea level thrust of the BE-4 engine is 550,000 pounds, so two is 1,100,000 pounds, i.e., 500 metric tons force. So two BE-4’s could not even loft the booster stage, let alone the upper stage and payload. Robert Clark

Yes. I copied this from another site with a different font. I’ll correct it. Robert Clark

Thanks. I took this image for the February Superheavy static test to be size of the exclusion zone for an actual launch: This is only an area 3 minutes, 15 seconds of latitude wide, that's 3.7 miles, 6 km. But this means the radius from the launch site is 2 miles, 3 km. But here is the image from the FAA report: https://www.faa.gov/space/stakeholder_engagement/spacex_starship/media/Draft_PEA_for_SpaceX_Starship_Super_Heavy_at_Boca_Chica.pdf This shows the exclusion radius for the public about 4 miles, 6.4 km. Perhaps the smaller map shown in the "Marine Safety Information Bulletin" was only for the SuperHeavy static test in February. I understand that only had 1/3rd propellant load. The larger area shown in the FAA map might be sufficient for the public. Still, one should be wary of the amount of damage even kilometers away as demonstrated by the Halifax and Texas City disasters, both in the ~3 kiloton range: https://en.wikipedia.org/wiki/Halifax_Explosion#Explosion https://sometimes-interesting.com/texas-city-disaster-deadliest-industrial-accident-in-u-s-history/

One responder to that twitter thread claimed he heard that booster is to be scrapped. Robert Clark

Updated discussion of the topic: SuperHeavy+Starship have the thermal energy of the Hiroshima bomb. UPDATED, 3/8/2023. https://exoscientist.blogspot.com/2023/03/superheavystarship-have-thermal-energy.html Key points: 1.)While the explosive force of the SuperHeavy/Starship (SH/ST) is not likely to reach that of its full thermal content of 13.3 kilotons of TNT, comparable to the Hiroshima bomb, it is still likely to be in the range of 3 to 5 kilotons of TNT. 2.)The Halifax and Texas City disasters of comparable explosive force suggests damage can extend kilometers away. 3.)The hazard or exclusion zones of only 2 miles, 3 km, for SH/ST is likely inadequate based on the Halifax and Texas City disasters. 4.)SpaceX ignored FAA warnings not to launch SN8 due to weather conditions exacerbating the effects of a possible blast wave from an explosion. 5.)The Starship SN11 explosion in midair may have been a BLEVE, which introduces an additional detonation mode for cryogenic fuels. 6.)At least one Raptor leaked methane and caught fire on multiple test flights of the Starship. 7.)Since the SuperHeavy static test lasted little more than 5 seconds, a strong possibility exists that multiple engines will fail during a full burn of an actual flight. Recommendations. 1.)It should be revealed to the public the SH/ST has the thermal energy content of the Hiroshima bomb. 2.)Experts on launch vehicle explosions and fuel-air detonations should present a report to the public explaining what the likely explosive force would be if the vehicle exploded. 3.)SpaceX should not be granted a launch license for the SH/ST until SpaceX constructs a separate engine test stand sufficient to test all 33 Superheavy engines at the same time time, at full power, and at full flight duration, and for such tests to complete successfully for multiple tests.

There is speculation on the new Starship nosecone: SInce ship 26 will be expendable maybe the new nose cone is without header tanks in the nose. Robert Clark

The difference is speed. It took 10 yeas for New Horizons to each Pluto. The high Isp of the ion engines would allow them to get to one of the outer planets or dwarf planets more quickly. The high Isp also means they could slow down to actually go into obit. Bob Clark

And let’s not forget: Anyone ever do a Kerbal sim to the outer planets of such a mission? Robert Clark

I only saw these calculation for the two stage. Robert Clark

Try the Silverbird calculator for the SSTO payload. Note that for an actual SSTO, you won’t use the full nose cone/payload section of the full two-stage . Estimate the SSTO payload with a fairing 1/10th the size of the current one. Robert Clark What’s the scuttlebutt on the NasaSpaceflight.com forum for the purpose of Ship 26? I’m persona non grata on that site for my SSTO speculations. Robert Clark

Thanks for that. I didn’t know you could do static fires there, which would need hold down clamps for example. I’ll be interested to find out if they’ll be installing 6 engines on it, as that is the current plan for an operational Starship. If yes, then it can do a launch fully fueled. Robert Clark

Interesting article: Much web discussion is going on on space forums about the Starship version Ship 26. This surprised everyone in being a completely expendable format. It has no top or bottom flaps, heat shield, or legs. Since it is to be expendable it likely also has no ballast tanks. The most frequent speculation is its a test vehicle for orbital refueling. But it has no visible external connections for linking up to another Starship. The key clue is it’s moved to the suborbital launch pad. This means it can launch without the SuperHeavy booster. With 6 Raptor 2 engines it can launch fully fueled unlike the previous Starship test flights meant just to test landing. The key question: what is the dry mass of this expendable version without flaps, legs, heat shield, or ballast tanks? If you know that you can calculate how much payload it can lift to orbit in a single stage. Elon said the expendable version with only 3 engines might mass only 40 tons: https://twitter.com/elonmusk/status/1111798912141017089?s=61&t=A46qVnS2GH4VVA-pQSUlkg Add another 5 tons for 3 more engines and this version might mass 45 tons. However, the increased thrust may require strengthening of the tanks which would increase the dry mass. On the other hand, this version would have to support far less payload atop it than the max 250 tons of the full two-stage so would need reduced tank strengthening. The argument can be made that just being moved to the suborbital launch pad does not mean it is going to be launched. It might be just used for pressure testing for example. However, the “Angry Astronaut” did a video from Boca Chica showing the Raptor work station being moved towards Ship 26: https://www.youtube.com/live/MmUwHVji9b4 He says that’s only done if you are installing engines on the Starship. You don’t do that if you are only doing pressure testing. He notes though that it could be putting engines either on S26 or S25. Probably we’ll know by the end of today which ship is having engines installed. Robert Clark

The Europa Clipper mission will be an orbiter mission to Europa to be launched on the Falcon Heavy in the 2024 time frame. But an actual lander mission to Europa could also be lauched in the same time frame on a Falcon Heavy. Having both orbiter and lander missions at the same time would be as revolutionary for planetary space science as were the Viking missions to Mars. Can we adapt the Antarctica IceCube Neutrino Observatory drilling technique to explore the subice oceans of Europa? https://exoscientist.blogspot.com/2023/01/can-we-adapt-antarctica-icecube.html Bob Clark

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

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

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.

If you do the calculation the amount of propellant used for landing is far less than 30 tons. Robert Clark