Real world rockets` initial TWR?

[FennexFox]

I`m trying to find it out, but it`s very hard to.

How`s it going? For example, Initial TWR of Saturn V 1st, 2nd, 3rd stage, or of Delta IV?

I don`t think it`s classified ones but it`s really hard to find out...

Edited October 20, 2014 by FennexFox

[Master Tao]

Don't know of a list, but it's not too hard to calculate from the Wikipedia sidebars.

Saturn V:

• Stage 1: 2,300 T
  
• Stage 2: 480 T
  
• Stage 3: 121 T
  
• Payload to LEO: 118 T
  
• Thrust: 34,020 kN
  

So the launch TWR is: 34,020 / ((2,300 + 480 + 121 + 118) * 9.81) = 1.15

By the time you get to the second and third stages, acceleration due to gravity will have changed enough to affect the TWR, so you'll also need to figure out how high the rocket was at that stage.

IIRC, the Space Shuttle had an initial TWR of 1.1.

[r_rolo1]

Hum, have you tried to check the wiki pages of the rockets you want to know and make the TWR calcs yourself? The Saturn V one has all the needed intel there in the right side box ...

EDIT: Ninja'ed :/

[MarvinKitFox]

Actually, SaturnV is a very bad example, as its launching TWR is the lowest of ***any*** orbital rocket ever.

(although Ariane 6 is looking to beat that, it seems)

More typical is something like the Soyuz, at 1.62

or the space shuttle at about 1.54 - 1.57 (varied a bit, depending on which shuttle, and the cargo load.)

Proton launches at 1.53

Falcon 9.1 TWR at liftoff is a mere 1.19 with max cargo, usually 1.21 or so

Of course, one assumes from the context of your question that you are referring to ground-launched, LEO-reaching rockets.

Otherwise funny results pop up, such as the Nike Zeus rocket, at TWR 39

Yes, 39g's of pullaway at the pad, and increasing as it burns!

Edited October 19, 2014 by MarvinKitFox

[Guest]

Generally, real rockets use between 1.15 and 1.5. Soyuz isn't "more typical", since it's TWR is abnormally large. Soviet designs in general have more TWR than American, and newer rockets have less than older ones.

[FennexFox]

Well, yes calc it myself is a great way to find them out, but I'm From a country where doesn't use imperial measure so it gives my a lot of headache. Anyway, thank you all for reply, and I'll try it.

[MarvinKitFox]

Well, yes calc it myself is a great way to find them out, but I'm From a country where doesn't use imperial measure so it gives my a lot of headache. Anyway, thank you all for reply, and I'll try it.

Just where does Imperial Measure come into this, when you are working with Mass in kilogram + Thrust in Kilonewton?

[MarvinKitFox]

Generally, real rockets use between 1.15 and 1.5. Soyuz isn't "more typical", since it's TWR is abnormally large. Soviet designs in general have more TWR than American, and newer rockets have less than older ones.

Soyuz isn't "more typical"??

Considering that more Soyuz have been launched than any three other designs together, I would say it pretty much defines typical.

As both Soyuz and Proton falls outside your 1.15-1.5 range, and between them represent 60% of all launches done, ever, your "Generally" also falls short.

[magnemoe]

Now the next question is why do they use so low TWR? Performance should be better with higher TWR?

My guess is that its because tanks and fuel is cheaper than engines.

[Kulebron]

The reason is that at sea level, thrust is considerably lower, unlike KSP where it remains the same. But there was no reason to optimize the engines for the lowest 20km of atmosphere. Engineers sacrifices some gravity losses for the sake of better TWR and Isp at higher altitudes, where the most of the work is done.

Also, in Apollo missions, launch pad was destroyable.

Soviet rockets were launched from 50 degrees latitude and thus needed to accelerate 200-300 m/s more (this meant that they also had less centrifugal force working for them and had to do more work to raise apoapse).

IIRC, gravity loss IRL is about 1-1.5km/s among 9-9.5 km/s total dV (15%). In KSP gravity losses make about 1.2km/s among 4.4km/s (27%). So in KSP it's a bigger problem, that's why we deal with that. By the way, Better than starting manned mod tweaks engine performance so that atmo TWR is lower that vac TWR.

Edited October 20, 2014 by Kulebron

[MarvinKitFox]

Now the next question is why do they use so low TWR? Performance should be better with higher TWR?

My guess is that its because tanks and fuel is cheaper than engines.

Max-Q

No use having a gonzo takeoff TWR, if you are having to throttle down drastically to keep your rocket intact by artificially lowering the max-q pressure.

Earth's soup-o-sphere is quite dense.

Earth's rocket bodies have an enormous Ludicrous Plaid mass fraction for fuel tanks. But this also means flimsy rocket structures, so aerodynamic buffeting is a real danger.

(in effect, it was the high shear wind it encountered, not the bad o-ring, that hammered the final nail in Challenger's coffin)

[Red Iron Crown]

You also get better payload fractions with initial TWR in the 1.2-1.3 range, if the real world is anything like KSP.

[cantab]

Yet another factor: in the real world drag on a streamlined rocket is much less than in stock KSP. So even with lower TWR, the rocket can get going faster sooner.

[FennexFox]

Just where does Imperial Measure come into this, when you are working with Mass in kilogram + Thrust in Kilonewton?

I found some information refers mass as pound, so I gave up immediately. I think I was impatient.

[Alshain]

Don't know of a list, but it's not too hard to calculate from the Wikipedia sidebars.

Saturn V:

• Stage 1: 2,300 T
  
• Stage 2: 480 T
  
• Stage 3: 121 T
  
• Payload to LEO: 118 T
  
• Thrust: 34,020 kN
  

So the launch TWR is: 34,020 / ((2,300 + 480 + 121 + 118) * 9.81) = 1.15

By the time you get to the second and third stages, acceleration due to gravity will have changed enough to affect the TWR, so you'll also need to figure out how high the rocket was at that stage.

IIRC, the Space Shuttle had an initial TWR of 1.1.

I just want to point out, all of this is payload dependent. The Apollo rocket that launched the Space Lab for example had barely enough TWR to get off the ground, fortunately it was only going to orbit and not the moon. The space shuttle had a TWR of anywhere from 1.1 to 1.25. STS-51-L being the absolute lowest TWR mission, but since that one never made it, I'm not sure if it would count.

Edited October 20, 2014 by Alshain

[LethalDose]

In my mind, the efficiency of low TWR at lift off is related to the mass ratio in the Tsiolkovsky equation, albeit indirectly. A low TWR at lift-off means your rocket is using , the bare minimum thrust needed to get off the ground in a stable manner. Assuming your rocket is well designed, using minimal required thrust should correspond with minimal engine mass, and as we all know that engine mass is basically "dead weight" in Tsiolkovsky's equation.

Hence, low TWR at lift-off is typically associated with increased dV efficiency in the lifting stage.

[Red Iron Crown]

Hence, low TWR at lift-off is typically associated with increased dV efficiency in the lifting stage.

Er, not exactly. Lower launch TWR typically means reduced dV efficiency but better fuel efficiency (within reasonable limits, of course).

[MarvinKitFox]

... The space shuttle had a TWR of anywhere from 1.1 to 1.25.

May I ask where you are getting this figure from?

Simple calc of Pad mass, and initial thrust, gives 1.51 for the heaviest ever load.

I suspect that this is *without* factoring that the SSME's normally ran at 109% throttle, and could be pushed to 114% if needed.

Simple calc of height-of-vehicle and time-elapsed from first motion gives me 5.3m/s2 of acceleration

And the graph of G-forces used for centrifuge simulation of the launch looks like this:

So, just *where* are you pulling your figure of 1.1 to 1.25 from?

[metaphor]

According to Wikipedia data (initial thrust / initial mass), the Saturn V had a TWR of 1.15, the Soyuz-FG has a TWR of 1.43, the Falcon 9 has a TWR of 1.18, the Atlas V 401 has a TWR of 1.25, and the Atlas V 551 has a TWR of 1.89.

[Frannington]

I suspect that this is *without* factoring that the SSME's normally ran at 109% throttle, and could be pushed to 114% if needed.

No SSME was ever run above 100% throttle for that particular engine.

100% throttle was designated as the maximum thrust that the original SSMEs could run at. Later ones had refinements that allowed them to deliver more thrust than the original engines, but they retained the 100% datum.

In other words, those later engines were running at 109-114% thrust of the engines that were originally fitted to Columbia, but they were at (or below) their own design specification.

[NathanKell]

LethalDose: you're forgetting gravity losses...low TWR means higher gravity losses.

The reason for low TWR in real life is not because it's more efficient; it's less efficient. The reason is because engines cost money, and propellant is cheap. It's *far* cheaper to load down a rocket with twenty extra tons of propellant (necessary for, say, 150m/s extra gravity losses) than to increase engine power to avoid those gravity losses.

It's also worth noting that, absent Falcon 9, most modern launchers are optimized for GTO missions, which means the upper stage needs very little thrust at all (consider the abysmal TWR of the Single-Engine Centaur).

There are three linked, but nonetheless separate, variables here, when talking about taking a payload to a specific orbit (all three change greatly when orbit changes):

1. How much delta V it takes to reach orbit - a function of gravity and drag losses, and the orbit.

2. What the final payload mass fraction is - a function of engine (and other components like structure) efficiency, number of stages, and (1).

3. How much the LV costs.

Optimizing for (1) will hurt (3) the most, and may lead to a lower (2). Optimizing for (2) will also hurt (3), and may well lead to a larger (1). Finally, optimizing for (3) will throw (1) and (2) out the window--it's far cheaper to make a 9.5-9.7km/sec ascent with high gravity losses than to have more engine and lower gravity losses.

Example: the optimum delta V to orbit can be achieved by a rocket that is perfectly aerodynamic and has infinite (and infinitely variable) thrust. That obviously will not carry much payload, and be very expensive.

Now, as for how to play the game:

If you're shooting for LEO (and not MEO, let alone GTO/GEO), then you will need a high-thrust upper stage, or accept a shedload of gravity and steering losses from lobbing high and then lowering your apogee once in orbit (to have the time to get into orbit on low thrust).

That can be done by something that starts at 1.15 and reaches 9.0 with a small second stage; that can be done by starting at 1.5, ending at 3, then going 1.0 to 3.0 on the second stage; that can be done by many 1.8-2.4 asparagus stages; whatever floats your boat (or lifts your payload). The things you need to watch for are:

1. Is the liftoff TWR, and not just that but the rate at which TWR climbs, sufficiently high that you will get out of the atmosphere, while low enough to ensure controllability and thermal survivability?

2. Is the total burn time short enough to allow you to circularize before falling back into the atmosphere? (Note that you may need to loft your second stage well above your desired apogee and circularize after apogee, then burn to lower apogee to desired level, if your total burn time is too long)

Edited October 23, 2014 by NathanKell

[The_Rocketeer]

No SSME was ever run above 100% throttle for that particular engine.

100% throttle was designated as the maximum thrust that the original SSMEs could run at. Later ones had refinements that allowed them to deliver more thrust than the original engines, but they retained the 100% datum.

In other words, those later engines were running at 109-114% thrust of the engines that were originally fitted to Columbia, but they were at (or below) their own design specification.

Not that I dispute this, but what is your source for this? My understanding was that the engines were rated for a given thrust (called 100%), which could be exceeded (up to 109%) for limited durations. I'd be interested to read whatever material you are referencing.

[Frannington]

Not that I dispute this, but what is your source for this? My understanding was that the engines were rated for a given thrust (called 100%), which could be exceeded (up to 109%) for limited durations. I'd be interested to read whatever material you are referencing.

Well it actually turns out that I'm not entirely correct. I was right about the 100% datum, but not about why it could be exceeded on later motors. I actually heard this verbally from a technician at Kennedy, so had to go looking for an actual source. As it happens, Wikipedia and its sources disagree with her, which is a bit worrying.

http://en.wikipedia.org/wiki/Space_Shuttle_main_engine

Specifically, about 2/3 down the page...

Specifying power levels over 100% may seem nonsensical, but there was a logic behind it. The 100% level does not mean the maximum physical power level attainable, rather it was a specification decided on during engine developmentâ€â€the expected rated power level. When later studies indicated the engine could operate safely at levels above 100%, these higher levels became standard. Maintaining the original relationship of power level to physical thrust helps reduce confusion, as it created an unvarying fixed relationship so that test data (or operational data from past or future missions) can be easily compared. If the power level was increased, and that new value was said to be 100%, then all previous data and documentation would either require changing, or cross-checking against what physical thrust corresponded to 100% power level on that date.

Their reference is a PDF which I won't link here, but you can retrieve it if you like.

I apologise for misleading you.

[The_Rocketeer]

Well, looks like I wasn't quite right either. Jolly good :-)

[Alshain]

May I ask where you are getting this figure from?

I was recalling from memory, which apparently isn't that reliable. You are correct.