High Thrust to Weight Ratio (TWR) rockets means get into fairly quickly and constantly experiencing G-Force, are there any downsides to high TWR rockets other than engineering? Seems like there are no problems accelerating 50 m / s^2 and get into orbit in 3 minutes.
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1$\begingroup$ Do you mean problems that come from having high TWR, or tradeoffs that must be made to achieve that? $\endgroup$– Nathan TuggyFeb 5, 2017 at 3:40
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1$\begingroup$ @NathanTuggy I mean problems regarding space flight, difficulty of engineering is excluded. $\endgroup$– RazeFeb 5, 2017 at 3:46
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1$\begingroup$ @Raze it would be helpful for you to improve your question a bit by expanding the explanation of your question. Instead of answering in a comment, can you edit your question directly and explain there? $\endgroup$– uhohFeb 5, 2017 at 8:28
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2$\begingroup$ I personally would not mind having TWR explained in the question ... $\endgroup$– user10509Feb 5, 2017 at 13:50
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5 Answers
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An extremely high TWR at launch makes for a rocket that will be traveling very fast in relatively dense air. This means more energy is spent fighting aerodynamic drag, and leads to greater friction heating and aerodynamic stress on the rocket.
Note also that an initial TWR of 5:1 will increase rapidly as fuel is consumed, and implies a peak acceleration on the order of 20g or more. This is way too much for manned flights and a big issue for unmanned - most space cargos are rated for something around 6g.
Finally, there's no real advantage to getting into orbit quickly, except in the possible case of military applications. It takes months or years to prepare for a launch and it takes 92 minutes to go around once in LEO; what are you going to do with the 10 minutes saved in the middle?
In practice, the ideal TWR is usually the lowest TWR that gets the launcher safely clear of the tower in a reasonable amount of time.
Thought experiment: Consider a 75-ton launcher with a TWR off the pad of 2:1 (i.e. the engines produce 150 tons of thrust at sea level). If you were to add 25 tons of "drop tanks" - say 1 ton of plumbing and separation equipment, 2 tons of tank structure, and 22 tons of fuel - you'd raise the liftoff mass to 100 tons, and you'd have a liftoff TWR of 1.5:1. If the drop tanks feed propellant into the main tanks at the same rate the engines consume them, then at the point when you empty the tanks and drop them, you effectively have the exact same 75-ton rocket you started with (okay, with a very small penalty in plumbing mass), with full tanks, but you've added some altitude and vertical velocity - something like 7m/s for each second it took to empty the drop tanks. In practice, the mass of empty tankage is so small that real launchers don't even bother with drop tanks - they simply design for a very low initial TWR.
Falcon 9's liftoff TWR is around 1.2. Saturn V was around 1.16. Shuttle was on the high side at about 1.5.
answered Feb 5, 2017 at 15:40
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2$\begingroup$ There is an advantage to getting to orbit fast once you are out of the atmosphere. Until you reach orbit you are suffering gravity losses. If you point your rocket straight up for 1 minute you spent 9.8 * 60 = 588 m/s of fuel for no gain. That's why rockets tip over as soon as they can--gravity loss is 9.8m/s * (1 - your fraction of orbital speed) so you want to build horizontal speed as soon as it won't cost you too much drag. During the 8 minutes the rocket roars at least 17% of it's energy is lost to drag and gravity. $\endgroup$ Feb 6, 2017 at 4:11
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$\begingroup$ Gravity losses obviously need to be accounted for, but my point is that there's no inherent advantage in getting to orbit in a shorter time. $\endgroup$ Feb 6, 2017 at 5:12
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1$\begingroup$ @RussellBorogove: The inherent advantage is lower delta-V. This is at odds with other factors, so given compromise must be found, but a rocket of initial TWR of 1.1 will be a very, very lousy rocket; at certain point adding more fuel means adding more tank dry mass, and due to gravity losses you're worse off than with less fuel (resulting in higher TWR) as the lousy initial delta-V gain does not offset lost acceleration at the end of the burn when you haul all the extra tank mass. $\endgroup$– SF.Feb 6, 2017 at 13:17
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1$\begingroup$ Reductio ad absurdum, your launch time TWR is less than 1, and the rocket sits on the launchpad blasting at full thrust futively until it burns up enough fuel to be able to lift itself off. A rocket of TWR a notch below 1 will perform better than one with twice as much fuel (and TWR of order 0.5), because the latter will have a higher dry mass, and their actual launch mass is the same. This holds true for values a good bit above 1.0. $\endgroup$– SF.Feb 6, 2017 at 13:21
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$\begingroup$ I'm not making myself clear. OP's post as written seems to presume that a 3-minutes ascent is in and of itself a good thing. I mean that given two rocket designs both of which can bring a given payload to orbit, there is no particular operational advantage to doing to in 3 minutes rather than 15. $\endgroup$ Feb 6, 2017 at 15:26
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The main disadvantage of high TWR is that engines are expensive but fuel and tankage is relatively cheap. This makes it more economical to use smaller engines and lower TWR, even though more fuel will be burned counteracting gravity the fuel is of negligible cost.
In addition, an engine capable of putting out more thrust generally needs to be larger and heavier - but that engine is not itself useful payload. Accelerating a heavy engine halfway to orbital velocity will reduce available delta-V compared with using a lighter engine which is still able to provide an acceptable TWR.
Solid Fuel, which is naturally conducive to a high TWR, also tends to be more expensive than liquid fuel. While solid fuels have their uses, it is cheaper if the bulk of the required orbital velocity comes from liquid fuels which gets back to using smaller engines at the lowest TWR that is still practical.
Higher thrust will also inflict more damage to the launchpad and surrounding infrastructure. I don't know whether that counts as a mere engineering problem, but it would certainly require a more robust launchpad or more repairs after a launch.
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$\begingroup$ I think any answer that suggests rockets and engines are designed to save money on propellant has gone a bit off-course. $\endgroup$– uhohFeb 5, 2017 at 17:39
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$\begingroup$ @uhoh hmmm doesn't my answer imply exactly the opposite? Maybe it's not as clear as I intended: Gravity drag can be counteracted quite effectively by having more thrust which means larger, more expensive engines, on the other hand it can be counteracted just by carrying and burning more fuel, which is cheaper. I did say "fuel" where I meant "deltaV" though. $\endgroup$ Feb 6, 2017 at 12:59
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$\begingroup$ If we call the 'more fuel' $\Delta \text{fuel}$ and the extension of the length of the rocket and of each tank to accommodate the $\Delta \text{fuel}$ as $\Delta \text{tank}$ and the fortification of the structure of the rocket to handle the extra weight $\Delta \text{rocket}$, I think the $\Delta \text{cost}$ of the $\Delta \text{fuel}$ is trivial in comparison to the $\Delta \text{cost}$ of the $\Delta \text{tank}$ plus $\Delta \text{fuel}$ - which is not so negligible. $\endgroup$– uhohFeb 6, 2017 at 13:24
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$\begingroup$ @uhoh most likely. But I was thinking, that the cost of propellant is only negligible if the propellants being used are of negligible cost - I believe the cost of filling a SRB would not be trivial compared with the cost of the shell - a quick back of an envelope calculation indicates that each SLS SRB's propellant would have cost ~$1 million in raw aluminium alone - which was still cheap in the context of a SLS launch but that would have to be expensive enough to matter for private sector and I'm sure this would have to at least factor into the choice to not use SRB's much. $\endgroup$ Feb 6, 2017 at 13:42
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1$\begingroup$ @uhoh not necessarily, the OP asked if there are downsides to high TWR rockets. Some fuels are better for high TWR than others, the best high thrust fuels are also (much) more expensive. That is a downside. Is it a downside that factors into rocket design decisions? Maybe, maybe not. Still a downside. (I agree it would be an interesting question though) $\endgroup$ Feb 6, 2017 at 14:16
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For manned launches, 5G is very uncomfortable, so manned rockets are often designed with a lower TWR and an ascent profile where engines are shut down to keep acceleration within limits.
This is less of an issue for unmanned rockets, but at some point G-loads will impose structural demands on the spacecraft that will make it heavier than it needs to be.
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As a spacecraft uses propellant, it loses weight which means the thrust-to-weight ratio of the spacecraft continues to increase. In many cases half or more of the weight could be propellant, meaning that it can increase substantially.
This make it more difficult to do carefully controlled thrust maneuvers, such as small delta-v corrections. These are usually executed by time, and the smaller the duration, the larger the possible error, depending on the level of control and stability and repeatability of ignition.
So a potential disadvantage would be increased difficulty with small delta-v maneuvers, or the necessity of developing, testing, and calibrating additional throttling capability, or the addition of additional small thrusters that might not have been necessary with a lower TWR engine.
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Because if you have a high TWR you might as well put more relatively cheap fuel in and get more deltav. Spacex does exactly that with the falcon 9, they have high TWR engines with a long fuel tank and they even super chill their fuel to get even more in. A high TWR rocket would encounter less gravity drag but the amount of delta v you save wouldn't be nearly as much as more fuel would add. There are other reasons but this is the main one.
