Ice. All these rockets use oxygen as the oxidizer component of their propellant. The Saturn 5 also used hydrogen in some of its engines (upper stages). They are stored in liquid state, which requires very low temperatures (below -183c for oxygen, below -253c for hydrogen). Despite insulation, some of the outside surfaces can get cold enough to condense and ...
You are missing how heat is distributed in exhaust.
Most of propellant ejected through the nozzle never makes contact with the nozzle surface or walls of the combustion chamber, and as result never has any chance to transfer its heat into them.
The exhaust gas primarily cools through adiabatic expansion - high pressure and high temperature both transformed ...
While in the atmosphere, a forward center of gravity contributes to stability. You want the center of gravity to be forward of the center of pressure for aerodynamic stability. A longer distance between the center of gravity and the gimbaling engines gives them more control authority as well.
The fuel used in the shuttle's Orbital Maneuvering System engines and used for the deorbit burn was not cryogenic; it was storable hypergolic fuel.
The cryogenic hydrogen and oxygen fuel burned by the main engines was used only during the ascent, which took only about 10 minutes.
As suggested by OrganicMarble in a comment, nitrogen is miscible with oxygen (you can thus make liquid air). According to NASA Technical Paper 2464, this is a major concern because using "enriched air" instead of pure oxygen as the oxidizer degrades the performance of the engine:
The transfer of liquid oxygen (LOX) from a storage vessel to a ...
Kennedy Space centre in Florida formerly had its own plant for the manufacture of cryogenic fuels and other liquids. This is now closed as noted earlier.
The procurement for cryogenic fuel for NASA is done via the office of procurement at KSC. It was determined that the US has only one supplier cryogenics that can supply the cape with the necessary 30 trucks ...
RP-1 isn't cryogenic actually. The subcooling for RP-1 is only to cool it to slightly below the freezing point of water, 20 F. At that temperature, no extreme cooling is required. The temperature difference is quite small compared to the larger LOX temperature difference, and the fuel is only there for 35 minutes. There likely isn't a need to cool the RP-1 ...
I wouldn't recommend it, and there's several reasons why.
First, even if we assumed you'd get chemically pure water out of the exhaust plume, in essence distilled water, it would lack any dissolved minerals and cause you to lose electrolytes. If that was all that was wrong with it, then small quantities wouldn't hurt you, but you'd have to substitute lost ...
Confusion abounds. Spaceflight 101 has this to say about the NK-33:
The NK-33 requires sub-cooled Oxygen with a temperature below its boiling point of -183 degrees Celsius to cool the turbopump bearings that would otherwise fail. Also, sub-cooled LOX has a higher density, close to that of Kerosene, reducing required tank volume and overall launch vehicle ...
...is the tank within a tank a sound engineering concept for rocket
I take this to mean that you are not talking about pressurant bottles or other small devices submerged in the propellant tanks. Instead you mean the large primary propellant tanks.
Then, No, this is not a good idea.
It doesn't need to be a double-walled, or "dewar-flask" tank....
NASA has done an extensive report on this, and in fact, cryogenic hydrogen tanks are considered to be one of the greatest technical achievements that NASA managed. Much of this is specific to the Centaur upper stage, but here's a few interesting quotes from the article:
Bossart led Mrazek out into the factory yard, where a Centaur tank
The biggest risk I foresee is nitric acid. A hot open flame in the atmosphere will cause N2 and O2 to react, creating a number of different molecules such as NO and NO2. These dissolve reasonably well in water and will form acids.
The existing answers do not accurately describe the procedure for the Space Shuttle system (or, I believe, for Apollo, but I am not 100% sure of that - see note at end of answer). The propellant tanks in the Space Shuttle's External Tank (ET) were never filled with nitrogen.
The initial condition for LH2 loading into the ET LH2 tank was with the tank and ...
The SSME is a staged-combustion rocket engine, which means that some small fraction of the propellant flow into the main combustion chamber is first diverted into a small pre-burner (two actually). These preburners combust (relatively) small amounts of fuel and oxidizer to produce hot exhaust gas which is expanded through a turbine, which is mechanically ...
Ariane 5 does not shed ice at liftoff. The first stage is covered in foam insulation that prevents ice buildup.
In this image, the insulation is the brown stuff. Later Ariane 5 versions switched to light-blue insulation tiles.
The Shuttle had insulation on its External Tank for the same purpose. For the Shuttle, it was critical not to have chunks of ice ...
This is a really good question and the answer is probably not 100% known, even by SpaceX at this moment.
No doubt they will have some active cooling to minimize boil off.
Structurally there are tricks they can play. For example the landing fuel is stored in a smaller tank, which is submerged in the main tank. Thus the surface area to warm up over is ...
This is complicated, but here is the gist:
Thrust is generated by flowing high pressure gas into a lower pressure environment. This flow is supersonic, so what goes on downstream of the throat (top of the the nozzle) cannot be sensed by what is going on in the combustion chamber (sound is just a pressure wave). The nozzle is supersonic, the combustion ...
For the Apollo 11 launch, at S-IC ignition the frost weights were:
S-IC: 1400 lb/635 kg
S-II: 450 lb/204 kg
S-IVB: 300 lb/136 kg
But by holddown arm release:
S-IC: 650 lb/295 kg
S-II: 450 lb/204 kg
S-IVB: 200 lb/91 kg
So about half the ice was shed (due to vibrations) before the stack left the ground.
The Saturn V Skylab flight evaluation report has ...
It was filled through the "T-0 umbilicals" (referring to the time of disconnect).
LO2 through an umbilical on the Orbiter boattail, LH2 through one in the midbody.
This schematic shows the plumbing from the umbilicals to the Centaur through the CISS (Centaur Integrated Support System).
There will likely be significant differences in the required tankage, if nothing else.
The paper Lunar Lander Conceptual Design shows a comparison between landers with similar payload requirements and different engine systems. Note the different in tankage and propellant weights for the two options.
At Kourou, there's a production plant for liquid hydrogen and oxygen. The plant is run by Air Liquide, a French company that specializes in industrial gases.
There's also a casting facility for solid rocket stages.
The melting point of hydrogen isn't much lower than its boiling point (6K), so the temperature isn't necessarily that much of an obstacle. However, using solid fuel requires either melting it, or burning it in place. Trying to burn solid hydrogen would likely result in the whole thing flash-boiling from the radiated heat, so that's a bit of a non-starter. ...
An additional reason that the forward position is advantageous for the LOX tank is the pressure head created by the distance from the tank to the engines. The turbopumps that drive the engines need sufficient inlet pressure to prevent cavitation. Moving the LOX tank forward allows the weight of the LOX to contribute to that pressure, which reduces the ...
Different stages use different means; Centaur uses(used) a thermodynamic vent mixer system for its LH2 tank.
This device served to keep the bulk propellant well mixed, and ensured that only gas was vented.
Typically they drain the cryogenics back, if they can.
Now it depends on the cryogenic. Oxygen can be drained and reused. Hydrogen in liquid form is very hard to keep cool enough, and they will drain it, but they will lose lots of it to venting.
Liquid Nitrogen that they use for pressurization they may just let vent, since the quantities are pretty ...
There is no problem running Kerosene and Oxygen on the same turbopump shaft, at any temperature. provided both are liquid, the density variations are not sufficient to make any practical difference to the feasibility of a turbopump.
From the OP, the densities of Oxygen and Kerosene are 1.18 and 0.8 g/cm3 , a ratio of 1.475. The pressure produced by a single ...
Short answer, they let some boil to keep the rest cool and also 'waste' some to cool the plumbing and tanks before loading.
The bulk cryogenics are made by compression, not cooling and from that point on normally the liquid is kept liquid by evaporation during shipping and storage.
This can become a complex engineering problem for large rockets to handle ...
I think the most important factor in avoiding boiloff is Not Using Hydrogen.
The atmospheric boiling points of the chosen propellants are as follows:
Oxygen 90K (-183C, -287F)
Methane 111K (-161C, -258F)
Compare Hydrogen 20K (-253C, -423F)
Space is a place of temperature extremes: roasting in the sun, but pretty cold in the shade. https://en.wikipedia....
Unsurprisingly, it worked exactly like it did in shuttle.
To assure uniform flow, the capillary restrictors are coiled around a
warm water-glycol line to increase the oxygen temperature.
The aforementioned oxygen supply capillary restrictors are wound
around the line routed to the space radiators and relief valves. The
other line is ...
A lander with storable propellants needs to keep them at close to room temperature, for a minimum of several days. A hydrolox system will take up much more volume due to the low density of LH2, and that big LH2 tank has to be kept at around 20 K.
You are going to need a major structural redesign just to deal with the greater volume of the liquid hydrogen ...