Would SSTO's Honestly be better than multistage rockets?

[Spaced Out]

So lets say you have two rockets. One is single stage, and the other is multistage. (These aren't accurate numbers, just an example) The single stage rocket's dry weight is 250 pounds, with another 750 as fuel. The first stage of the rocket weighs dry at 250 lbs, with 500 being fuel, and the second stage is 100 lbs with 150 being fuel.

The single stage rocket, once it has finished burning 500 lbs of fuel, has to use the rest of the fuel to lug push rocket which it can't just seperate from. The multistage rocket though, once it has finished burning 500 lbs of fuel, simply seperates, (and lets just say it is a Falcon 9 and had some extra fuel for landing) then lands to be reused. The second stage, unlike the single stage rocket, jettisons all of the extra weight and can use the rest of the fuel to push a lighter weight, and get more delta-v than the single stage rocket with the same amount of fuel, because the single stage rocket wasn't given the change to jettison extra weight.

Tell me what you think and if my argument makes sense. And here is one more thing. If SpaceX finds a way to reuse a second stage, then basically everything that matters has been recovered. They even found a way to recover the fairing. If the Falcon 9 becomes fully reusable, than why would an SSTO be any better? 

Edited August 20, 2017 by Spaced Out

[hms_warrior]

Well the basic idea behind SSTO is not only SSTO is it "reusable SSTO that you just need to refill".

In that case, a two stage rocket needs to be but together again after each launch. This means time and possible failure points. a "rSSTOyjntr" would just land, open it's cargo bay, get the next payload pushed in, tanked up and is good to go.

[MajorTomtom]

The problem with real life SSTO is that you have a single engine optimization throughout your entire flight, whereas a staged rocket can have multiple engines adapted to different altitudes, and therefore be more efficient.

If you recover your craft, whether it's SSTO or simply stages, you will have to go through a lot of refurbishing anyway to make it fly safely again (and convince people it is still safe). I don't know if recovering the second stage is really feasible/practical for Falcon 9. The first stage is in a "simple" parabolic flight whereas the second one is at orbital velocity or close to... It would have to survive reentry and therefore either be well shielded or spend a good amount of fuel to slow down enough.

[sevenperforce]

In theory, an SSTO has fewer engines, less maintenance, and fewer failure points.

In practice, it's just unattainable, and the theoretical advantages are miniscule.

[Bill Phil]

It depends on exhaust velocity. If your exhaust velocity is low (chemical engines), you need high mass ratios for your rocket. Only way to get around that is staging, without magic materials science. But if your exhaust velocity is high? In the 600 to 1000 second range (~6000 to ~10k m/s), then your mass ratios are smaller. This then increases payload mass fraction. More payload per launch, similar launch costs, (if you're lucky) cheaper access to space. 

A single stage is simpler, in theory. In practice, it doesn't work well with low exhaust velocities. As they say, the best system is the one that's not there. No interstage, no staging system, one set of engines for the propulsion system, and so on. But it only really works with higher exhaust velocities.

[magnemoe]

26 minutes ago, Bill Phil said:

It depends on exhaust velocity. If your exhaust velocity is low (chemical engines), you need high mass ratios for your rocket. Only way to get around that is staging, without magic materials science. But if your exhaust velocity is high? In the 600 to 1000 second range (~6000 to ~10k m/s), then your mass ratios are smaller. This then increases payload mass fraction. More payload per launch, similar launch costs, (if you're lucky) cheaper access to space. 

A single stage is simpler, in theory. In practice, it doesn't work well with low exhaust velocities. As they say, the best system is the one that's not there. No interstage, no staging system, one set of engines for the propulsion system, and so on. But it only really works with higher exhaust velocities.

This, and I would put the limit higher than 1000 seconds. At 1000 it would make sense for some uses, not bulk use and not outside LEO. An 2 stage with reuse of both stages will win in payload capacity unless ISP is many thousands.
Think Orion pulse nuclear is the only known design who would not benefit of staging because of the insane ISP and the heavy engine.

Skylon would have better performance as an dual stage setup. Even with second stage recovery and that is plausible as an SSTO, chemical rockets, forget it. 
 

[Kerbart]

The SSTO design has many, many advantages, as mentioned. Less parts, simpler, not throwing out the most expensive part (the engine), not lugging 2nd stage engines at take off, etc.

Compared to the SSTO, the multi-stage design has only one advantage. But it's a truly magnificent one: it actually works.

Until someone manages to succesfully launch with an engine that's efficient from sea level to space (the Aerospike?) launches will likely be multi stage.

[Bill Phil]

1 hour ago, magnemoe said:

This, and I would put the limit higher than 1000 seconds. At 1000 it would make sense for some uses, not bulk use and not outside LEO. An 2 stage with reuse of both stages will win in payload capacity unless ISP is many thousands.
Think Orion pulse nuclear is the only known design who would not benefit of staging because of the insane ISP and the heavy engine.

Skylon would have better performance as an dual stage setup. Even with second stage recovery and that is plausible as an SSTO, chemical rockets, forget it. 
 

Most higher isp systems have heavier engines, making multi stage designs a bad idea...

Let's do some simple math:

Using the rocket equation, 1000 s of isp (9800 m/s) as our exhaust velocity, and 10k as our total delta-v, our mass ratio is barely much higher than e. Less than 3. Less than the A4/V-2 mass ratio. Of course, you need an engine with just as much isp on the ground, or averaging 1000s during the ascent. But with a mass ratio of e, and current high propellant mass fraction tank designs, even with a heavy engine setup we could get significant payload mass fractions. For a 500 Tonne rocket, empty mass would be about 180 tonnes. About 36 tonnes of that would likely be tankage (and that might be a stretch, considering that shuttle ETs were about 27 tonnes empty at their lightest). That gives us about 144 tonnes for engines, structure, and payload, guidance, and so on. 50 tonnes of payload shouldn't be impossible, seeing as an Atlas V can get more than 10 with inefficient chemical engines and something around 300 to 400 tonnes of rocket. 

Halving exhaust velocity squares the required mass ratio for a given Delta-v. Doubling the exhaust velocity reduces the required mass ratio to its square root. This is why just doubling chemical rocket's current isp would benefit rockets enormously.

[kerbiloid]

3 hours ago, Bill Phil said:

A single stage is simpler, in theory. In practice, it doesn't work well with low exhaust velocities. As they say, the best system is the one that's not there. No interstage, no staging system, one set of engines for the propulsion system, and so on. But it only really works with higher exhaust velocities.

It's more accurate to say "with higher exhaust / orbital velocity ratios".
Say, on Kerbin (ISP/orbital ~1) this works just fine.

P.S.
Multistage is a wheelchair until you get enough high ISP/orbit.
it has no advantages itself.

[Bill Phil]

20 minutes ago, kerbiloid said:

It's more accurate to say "with higher exhaust / orbital velocity ratios".
Say, on Kerbin (ISP/orbital ~1) this works just fine.

P.S.
Multistage is a wheelchair until you get enough high ISP/orbit.
it has no advantages itself.

Well, sure, on different planets you don't need as much Delta-v, but everyone lives on Earth at the moment.

[YNM]

4 hours ago, Bill Phil said:

4 hours ago, kerbiloid said:

It's more accurate to say "with higher exhaust / orbital velocity ratios".
Say, on Kerbin (ISP/orbital ~1) this works just fine.

Well, sure, on different planets you don't need as much Delta-v, but everyone lives on Earth at the moment.

He still says it correctly. ISP of 340s (~3300 m/s) on Kerbin, for example, is higher than orbital velocity, hence very efficient. On Earth you need ISP higher than 920s (~9000 m/s).

Edited August 21, 2017 by YNM

[Bill Phil]

28 minutes ago, YNM said:

He still says it correctly. ISP of 340s (~3300 m/s) on Kerbin, for example, is higher than orbital velocity, hence very efficient. On Earth you need ISP higher than 920s (~9000 m/s).

And Kerbin doesn't exist. 

I'm not saying that what Kerbiloid said is false, but it's not very applicable until we have the need to launch from other celestial bodies, which hasn't been done very often.

And besides, on Kerbin, 1000s of isp would still be extremely useful and would improve payload fractions enormously.

[p1t1o]

Today? In terms of launch-for-launch? No not really, especially since they are so difficult to build/exist on the edge of technological feasibility.

But fast forward to an arbitrarily far future. Is this a future where you see commuting to a job in space as a norm? Or at least people vacationing there for the same price as a skiing trip?

If you do, then this world will need vastly increased capacity for launching vehicles. An SSTO might not be better than multistage today, but what if you want to launch 10,000 flights per day, globally?

Then I think the savings and turnaround time, might add up.

Maybe.

Edited August 21, 2017 by p1t1o

[kerbiloid]

24 minutes ago, Bill Phil said:

And Kerbin doesn't exist. 

???

25 minutes ago, Bill Phil said:

until we have the need to launch from other celestial bodies

ITS, so on...
Say, Martian orbital speed is just 3.6 km/s.

[Bill Phil]

1 minute ago, kerbiloid said:

???

ITS, so on...
Say, Martian orbital speed is just 3.6 km/s.

Higher specific impulse, no matter what body you're launching from, leads to higher payload fractions and lower mass ratios. In the context of SSTO vehicles, launching from Earth, we need higher isp values. We're not discussing Mars SSTO, or Lunar SSTO, or any other solar system body, not at this moment. Discussion on those isn't really what we're talking about here. We can't get to them until we leave Earth. One problem at a time.

ITS is, at best, decades away. At worst, the cause of hundreds, or even thousands, of preventable deaths. 

[YNM]

2 hours ago, Bill Phil said:

And Kerbin doesn't exist. 

I'm not saying that what Kerbiloid said is false, but it's not very applicable until we have the need to launch from other celestial bodies, which hasn't been done very often.

And besides, on Kerbin, 1000s of isp would still be extremely useful and would improve payload fractions enormously.

Well yes. It is.

So, no SSTOs then for the next century.

[wumpus]

2 hours ago, YNM said:

Well yes. It is.

So, no SSTOs then for the next century.

There was a small startup called Escape Dynamics which made an effort to develop a SSTO with an Isp of ~800s.  The "engine" would consist of ground based lasers (or masers) that would heat up the reaction mass tank (full of hydrogen) and launch that way.

The power output from the lasers doesn't really seem to be something Escape Dynamics seriously attempted to provide, but it is an ongoing military research program by the US Navy.  While it is only concerned with power during pulses (which is a great way to destroy missiles), it is also the most likely means of providing the power needed to launch a heat-powered rocket such as Escape Dynamics proposed.

No idea if 800s is enough for SSTO to be viable, although reducing dry weight enough to change a log by 80% is heroic.  Perhaps once the lasers get it going supersonic, it could use Ramjets/SCRAMJETS to assist to mach 3-7.  In any event, you need an unbelievable number of flights to justify such a contraption over reassembling conventional rockets (and don't expect to be saving much fuel, if any).  I doubt that missing any "magic Isp number" was what did in Escape Dynamics, lack of available lasers was much more likely.

Another possibility (based on a test vehicle that has already flown) would be based on the X-43.  This would greatly resemble any effort using the SABRE, although it might extend the staging velocity a bit.  I should be noted that the paper that the NASA investigators published didn't include any measurements of thrust/acceleration.  Acceleration was clearly positive (although not labeled) at mach 7 (faster than SABRE) and roughly similar to the negative acceleration due to drag (so presumably pretty good).  Acceleration was only "most likely positive, statistically speaking" at mach 10 and clearly not useful for accelerating a spacecraft (although more research might allow sustained mach 10 flight).  I've mentioned another paper before (likely by students digging through NASA data) which included Isp measurements, but couldn't find it today.  Isp is projected to be 2900s at mach 6, measured Isp was not included (although I'm pretty sure it was over 2000 for mach 7).

https://hapb-www.larc.nasa.gov/Public/Documents/AIAA-2006-1-317.pdf

Note that I think Spacex can a put commercial payload into orbit for less than the projected "government accounting" price listed for the proposed scramjet rocket, with the possible exception of the final SSTO (no speed listed for switch to supplying oxidizer).  I have no doubt that Spacex can break the $1000/lb mark first (note that any marketing saying as much is quoting prices for "with recovery of first stage" and payload mass "without recovery").  Don't be too surprised if Blue Origin beats Spacex to it (if and only if you have a truly massive payload).

[Nothalogh]

On 8/20/2017 at 8:51 PM, Kerbart said:

the multi-stage design has only one advantage. But it's a truly magnificent one: it actually works.

The rocket gods may be cruel, but they are honest

[magnemoe]

On 21.8.2017 at 4:59 PM, wumpus said:

There was a small startup called Escape Dynamics which made an effort to develop a SSTO with an Isp of ~800s.  The "engine" would consist of ground based lasers (or masers) that would heat up the reaction mass tank (full of hydrogen) and launch that way.

The power output from the lasers doesn't really seem to be something Escape Dynamics seriously attempted to provide, but it is an ongoing military research program by the US Navy.  While it is only concerned with power during pulses (which is a great way to destroy missiles), it is also the most likely means of providing the power needed to launch a heat-powered rocket such as Escape Dynamics proposed.

No idea if 800s is enough for SSTO to be viable, although reducing dry weight enough to change a log by 80% is heroic.  Perhaps once the lasers get it going supersonic, it could use Ramjets/SCRAMJETS to assist to mach 3-7.  In any event, you need an unbelievable number of flights to justify such a contraption over reassembling conventional rockets (and don't expect to be saving much fuel, if any).  I doubt that missing any "magic Isp number" was what did in Escape Dynamics, lack of available lasers was much more likely.

Another possibility (based on a test vehicle that has already flown) would be based on the X-43.  This would greatly resemble any effort using the SABRE, although it might extend the staging velocity a bit.  I should be noted that the paper that the NASA investigators published didn't include any measurements of thrust/acceleration.  Acceleration was clearly positive (although not labeled) at mach 7 (faster than SABRE) and roughly similar to the negative acceleration due to drag (so presumably pretty good).  Acceleration was only "most likely positive, statistically speaking" at mach 10 and clearly not useful for accelerating a spacecraft (although more research might allow sustained mach 10 flight).  I've mentioned another paper before (likely by students digging through NASA data) which included Isp measurements, but couldn't find it today.  Isp is projected to be 2900s at mach 6, measured Isp was not included (although I'm pretty sure it was over 2000 for mach 7).

https://hapb-www.larc.nasa.gov/Public/Documents/AIAA-2006-1-317.pdf

Note that I think Spacex can a put commercial payload into orbit for less than the projected "government accounting" price listed for the proposed scramjet rocket, with the possible exception of the final SSTO (no speed listed for switch to supplying oxidizer).  I have no doubt that Spacex can break the $1000/lb mark first (note that any marketing saying as much is quoting prices for "with recovery of first stage" and payload mass "without recovery").  Don't be too surprised if Blue Origin beats Spacex to it (if and only if you have a truly massive payload).

Think escape dynamic folded, main issue would be laser / maser costs and complexity with all the stations along the flight path. 
They planned do use air until they got high and fast reducing the reaction mass need and get higher TWR at the start.
 

[wumpus]

1 hour ago, magnemoe said:

Think escape dynamic folded, main issue would be laser / maser costs and complexity with all the stations along the flight path. 
They planned do use air until they got high and fast reducing the reaction mass need and get higher TWR at the start.

Air would make a lot of sense and easily hit any "magic Isp number" (although I think it also means that you wouldn't get 800 above the atmosphere, there were a few catches in delivering that number too far away from the laser.  But getting mach 6 roughly drops the "magic Isp" number in half (and I think Isp bottomed out around ~600, in theory).

I also don't think laser "complexity" was so much the issue as that lasers with the needed power simply don't exist (or only deliver said power for milliseconds).  Somebody might try this once the Navy has fielded tested laser weapons, but I suspect that such small startups can't develop powerful continuous lasers out of burst ones.

[magnemoe]

12 hours ago, wumpus said:

Air would make a lot of sense and easily hit any "magic Isp number" (although I think it also means that you wouldn't get 800 above the atmosphere, there were a few catches in delivering that number too far away from the laser.  But getting mach 6 roughly drops the "magic Isp" number in half (and I think Isp bottomed out around ~600, in theory).

I also don't think laser "complexity" was so much the issue as that lasers with the needed power simply don't exist (or only deliver said power for milliseconds).  Somebody might try this once the Navy has fielded tested laser weapons, but I suspect that such small startups can't develop powerful continuous lasers out of burst ones.

You would switch to internal hydrogen at some speed and attitude. 
Agree on the laser issue, this require high power lasers so wait until we have them, less an issue than laser weapons as system don't need to be mobile 
Takeoff would be another issue, not sure how to handle it. Air drop would be simple but have size constrains. 

[wumpus]

For takeoff, I'd assume that such a spacecraft could be launched off a rail.  This is a "build *lots* of infrastructure to make *many* launches cheaper" strategy.  I can't see the rail approaching the cost of the lasers/power supplies for same.

[Iskierka]

On 20/08/2017 at 7:56 PM, Spaced Out said:

So lets say you have two rockets. One is single stage, and the other is multistage. (These aren't accurate numbers, just an example) The single stage rocket's dry weight is 250 pounds, with another 750 as fuel. The first stage of the rocket weighs dry at 250 lbs, with 500 being fuel, and the second stage is 100 lbs with 150 being fuel.

The single stage rocket, once it has finished burning 500 lbs of fuel, has to use the rest of the fuel to lug push rocket which it can't just seperate from. The multistage rocket though, once it has finished burning 500 lbs of fuel, simply seperates, (and lets just say it is a Falcon 9 and had some extra fuel for landing) then lands to be reused. The second stage, unlike the single stage rocket, jettisons all of the extra weight and can use the rest of the fuel to push a lighter weight, and get more delta-v than the single stage rocket with the same amount of fuel, because the single stage rocket wasn't given the change to jettison extra weight.

The situation presented is naive and doesn't represent the actual situation you would have with an SSTO, thus gives a bad comparison.

Why give them equal weight? The constraint on launch is cost. You can always throw more engines on it. Given the complexity of staging, by removing a stage, you immediately save money that can increase the thrust, and therefore just add more fuel (which has negligible cost so we can effectively throw as much in as we need). You also don't have to pay for the second stage's engine, allowing a way bigger rocket for the same cost, and a much larger relative amount of fuel to rocket+payload.

Given a vertical launch SSTO, yes, this would give a bad payload fraction - but payload fraction makes no difference. Are you getting the payload up there? For the same or less cost? Then the rocket is superior to the two-stage it's potentially replacing, as it's definitely more reliable for having less failure points.

Vertical launch on Earth doesn't quite work for SSTO due to fuel performance, as LH2/LOx didn't live up to Tsiolkovsky's predictions, and ozone/fluorine/FOOF oxidizers turned out non-viable on practical sides. It is technically possible, but the margin is too small for you to convert the staging costs into fuel well enough to make a good return. It does, however, only need a small improvement to be quite viable. The failed suggestions above were in the range 500-570 seconds, and would largely be capable of reasonable SSTO, as IRL has excellent ability to create mass ratios as compared to KSP. Only small increases are needed.

Given that we can't use those fuels, though, you need to consider the other approach. Up to 450 seconds at 9500-10000 m/s doesn't work, 570 at 9500-10000 does. Okay, what about 450 seconds at 6500-7000? Just quick back of the envelope, 10000/570 is nearly 18, 7000/450 is about 15.5, so if you could reduce the orbit dV requirement, then you can get back into comfortable SSTO territory, with a little bit of margin to play with to attain that dV reduction. We need to take off 3000 m/s - 1500 m/s is often taken from gravity losses quite early just getting up to speed, so if we can start at some altitude and non-zero speed we get that down to 1500 remaining to take off.

1500 is just over mach 5 - can we get our non-zero speed to be 1500 m/s? Turns out, possibly, with fancy enough engines! This is why IRL SSTOs are generally thought of as spaceplanes, as they would need to fly up to speed & altitude on lift and airbreathing power. But, given how little dV reduction is actually required compared to the total, less than a third, in order to put SSTO in a comfortable area, there are potentially engines that can airbreathe that fast, that can work on current technology rather than finicky scramjets. Potential payload fraction isn't fantastic, although it is still quite good, but that doesn't matter if you can use just one vehicle for the entire job, that can then do the next job with minimal maintenance.

 

Really, payload fraction is a bad thing to focus on at all. There's always the potential to improve it by adding stages, even if the mass ratio required is only 1.5 - but do you really want to pay for two stages to make it 1.3 overall, when you could just stick with one stage at 1.5?

 

As an additional note, it's really not that hard to get an engine that's efficient across the whole range. There are even flown examples of bell nozzle engines of both hydrolox and kerolox that pretty much equal or beat suggestions for altitude compensation like aerospikes across the whole range, at high TWR too.

For hydrolox, the RS-25 SSME gets near-optimal performance at all altitudes, and has a TWR of 68.5 - one of (or the?) highest TWR of any hydrolox engine ever made. Is expensive to operate, but the RS-25 arguably came at a bad time before much computer aid could be used in the design; hydrolox is a very gentle fuel, as shown by the consistent reliability and refire-ability of the RL10, and there's really no reason its performance shouldn't be possible to match on an engine that is much easier to maintain with modern design.

For kerolox, there's the NK-33 and -43, and the whole RD-171, -180, -191, -193 family, which have a surface-level impulse that beats the vacuum impulse of many or most kerolox engines, for some of the highest TWR ever observed on any type of rocket. The only two engines that can beat the NK-33 for TWR don't come close to it in impulse.

Regardless of fuel choice, a high-performance engine in both TWR and Isp, across the whole range, is entirely possible. This is the basic requirement of an SSTO - the second is simply to pull the dV requirement down to where the fuels available to choose from can do it.

Edited August 26, 2017 by Iskierka

[sevenperforce]

17 minutes ago, Iskierka said:

The situation presented is naive and doesn't represent the actual situation you would have with an SSTO, thus gives a bad comparison.

Why give them equal weight? The constraint on launch is cost. You can always throw more engines on it. Given the complexity of staging, by removing a stage, you immediately save money that can increase the thrust, and therefore just add more fuel (which has negligible cost so we can effectively throw as much in as we need). You also don't have to pay for the second stage's engine, allowing a way bigger rocket for the same cost, and a much larger relative amount of fuel to rocket+payload.

Given a vertical launch SSTO, yes, this would give a bad payload fraction - but payload fraction makes no difference. Are you getting the payload up there? For the same or less cost? Then the rocket is superior to the two-stage it's potentially replacing, as it's definitely more reliable for having less failure points.

Vertical launch on Earth doesn't quite work for SSTO due to fuel performance, as LH2/LOx didn't live up to Tsiolkovsky's predictions, and ozone/fluorine/FOOF oxidizers turned out non-viable on practical sides. It is technically possible, but the margin is too small for you to convert the staging costs into fuel well enough to make a good return. It does, however, only need a small improvement to be quite viable. The failed suggestions above were in the range 500-570 seconds, and would largely be capable of reasonable SSTO, as IRL has excellent ability to create mass ratios as compared to KSP. Only small increases are needed.

Given that we can't use those fuels, though, you need to consider the other approach. Up to 450 seconds at 9500-10000 m/s doesn't work, 570 at 9500-10000 does. Okay, what about 450 seconds at 6500-7000? Just quick back of the envelope, 10000/570 is nearly 18, 7000/450 is about 15.5, so if you could reduce the orbit dV requirement, then you can get back into comfortable SSTO territory, with a little bit of margin to play with to attain that dV reduction. We need to take off 3000 m/s - 1500 m/s is often taken from gravity losses quite early just getting up to speed, so if we can start at some altitude and non-zero speed we get that down to 1500 remaining to take off.

1500 is just over mach 5 - can we get our non-zero speed to be 1500 m/s? Turns out, possibly, with fancy enough engines! This is why IRL SSTOs are generally thought of as spaceplanes, as they would need to fly up to speed & altitude on lift and airbreathing power. But, given how little dV reduction is actually required compared to the total, less than a third, in order to put SSTO in a comfortable area, there are potentially engines that can airbreathe that fast, that can work on current technology rather than finicky scramjets. Potential payload fraction isn't fantastic, although it is still quite good, but that doesn't matter if you can use just one vehicle for the entire job, that can then do the next job with minimal maintenance.

 

Really, payload fraction is a bad thing to focus on at all. There's always the potential to improve it by adding stages, even if the mass ratio required is only 1.5 - but do you really want to pay for two stages to make it 1.3 overall, when you could just stick with one stage at 1.5?

A better number to look at is the ratio of payload to vehicle dry mass.

[Iskierka]

Better, as it correlates more with how much vehicle you have, but still imperfect. Totally hypothetically, no example vehicles to give, would you rather have a 20-tonne big dumb booster that can single-stage 20 tonnes for $500,000 with 400 tonnes of fuel, or a fancy plane that's 50 tonnes and 200 tonnes of fuel, but can single-stage that same 20 tonnes for $100,000? Choice should be obvious. Bring it back to cost, always - dry mass is just sometimes a useful guesstimate of cost.

Edited August 26, 2017 by Iskierka