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what about conformal fuel tanks? what makes them so impossibel? is it forming the materials into complex shapes?

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 A SSTO formed from the first stage of a TSTO will carry less payload than the full TSTO. But it will be cheaper and most launches don't need the full TSTO launch capacity anyway.

 Key then is knowing how high that payload can be using altitude compensation for the SSTO. People dismiss the possibility because they don't realize how high the payload can be with alt. comp. Even though the calculation is no more difficult using the rocket equation than is the standard case without alt. comp., they don't do the calculation because they assume the answer is just about the same as the standard case.

 In fact the answer is not even in the same ballpark. In fact the answer is in the range of currently in use rockets that have billion dollar contracts.

 

  Bob Clark

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On 5/30/2016 at 9:39 PM, Exoscientist said:

 A SSTO formed from the first stage of a TSTO will carry less payload than the full TSTO. But it will be cheaper and most launches don't need the full TSTO launch capacity anyway.

 Key then is knowing how high that payload can be using altitude compensation for the SSTO. People dismiss the possibility because they don't realize how high the payload can be with alt. comp. Even though the calculation is no more difficult using the rocket equation than is the standard case without alt. comp., they don't do the calculation because they assume the answer is just about the same as the standard case.

 In fact the answer is not even in the same ballpark. In fact the answer is in the range of currently in use rockets that have billion dollar contracts.

 

  Bob Clark

ok, let's assume a tsto carries 28T to leo. pretty reasonable for the biggest satellites.

A altitude compensated ssto might carry around 5T, generously, which  is too small for most satellites. 

The performance gain from tsto is so great that the extra complexity of a staged vehicle is taken. It would also be more well-understood and less complex than air-augmenting a 1st stage.

On 5/29/2016 at 10:52 AM, Emperor of the Titan Squid said:

what about conformal fuel tanks? what makes them so impossibel? is it forming the materials into complex shapes?

not impossible, just difficult enough that no one bothers since they don't need to use it.

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On May 31, 2016 at 2:23 PM, fredinno said:

ok, let's assume a tsto carries 28T to leo. pretty reasonable for the biggest satellites.

A altitude compensated ssto might carry around 5T, generously, which  is too small for most satellites. 

The performance gain from tsto is so great that the extra complexity of a staged vehicle is taken. It would also be more well-understood and less complex than air-augmenting a 1st stage.

not impossible, just difficult enough that no one bothers since they don't need to use it.

  To estimate via the rocket equation the payload to LEO, you can just use the vacuum value of the Isp (rather than taking an average) since you can regard the loss of performance at sea level as just another loss such as gravity and air drag and add that onto the delta-v required to orbit. Then a value for the required delta-v to orbit is in the range of 30,000 ft/sec about 9,100 m/s:

http://yarchive.net/space/rocket/fuels/hydrogen_deltav.html

 The rocket equation for the rocket's delta-v is: ΔV = Isp * g0 * ln(Minitial/Mfinal) . 

 We need specs on the F9 first stage. In this lecture Elon Musk provides a key parameter, the propellant fraction, of the F9 v1.1 version:

 About 30 minutes in, he gave the propellant fraction of the Falcon 9 v1.1 first stage as around 96%, or perhaps 95.5%.The propellant load of the F9 v1.1 has been estimated as 385 metric tons. Then using the 95.5% propellant fraction number this would correspond to a gross mass of the first stage of 385/.955 = 403 metric tons. The dry mass is then .045*403 = 18 metric tons. Then this could loft 2 metric tons to orbit as a SSTO:

 ΔV = 311*9.81ln((403 + 2)/(18 + 2)) = 9,177 m/s.

 Now suppose we could get the first stage engines to have the same vacuum Isp of the Merlin Vacuum at 342 s by altitude compensation. Then, by this estimation method, we could get 9 metric tons to orbit:

ΔV = 342*9.81ln((403 + 9)/(18 + 9)) = 9,143 m/s

  However, both of these are approximations. To get a better estimate you need to take into account how the Isp changes with altitude. Anyone know how to do this in Kerbal?

  Bob Clark

 

 

Edited by Exoscientist
Typo

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3 hours ago, Exoscientist said:

  To estimate via the rocket equation the payload to LEO, you can just use the vacuum value of the Isp (rather than taking an average) since you can regard the loss of performance at sea level as just another loss such as gravity and air drag and add that onto the delta-v required to orbit. Then a value for the required delta-v to orbit is in the range of 30,000 ft/sec about 9,100 km/s:

http://yarchive.net/space/rocket/fuels/hydrogen_deltav.html

 The rocket equation for the rocket's delta-v is: ΔV = Isp * g0 * ln(Minitial/Mfinal) . 

 We need specs on the F9 first stage. In this lecture Elon Musk provides a key parameter, the propellant fraction, of the F9 v1.1 version:

 About 30 minutes in, he gave the propellant fraction of the Falcon 9 v1.1 first stage as around 96%, or perhaps 95.5%.The propellant load of the F9 v1.1 has been estimated as 385 metric tons. Then using the 95.5% propellant fraction number this would correspond to a gross mass of the first stage of 385/.955 = 403 metric tons. The dry mass is then .045*403 = 18 metric tons. Then this could loft 2 metric tons to orbit as a SSTO:

 ΔV = 311*9.81ln((403 + 2)/(18 + 2)) = 9,177 m/s.

 Now suppose we could get the first stage engines to have the same vacuum Isp of Merlin Vacuum at 342 s by altitude compensation. Then, by this estimation method, we could get 9 metric tons to orbit:

ΔV = 342*9.81ln((403 + 9)/(18 + 9)) = 9,143 m/s

  However, both of these are approximations. To get a better estimate you need to take into account how the Isp changes with altitude. Anyone know how to do this in Kerbal?

  Bob Clark

 

 

No, using the vac. ISP gets way more Delta-V out of a rocket because the rocket equation is exponential, and most of the fuel is used in the atmosphere (ie why air-launch is useful). Also, you didn't add up landing equipment mass, heat shielding, reserve landing fuel, payload fairing mass (though that is ejected, it must be added for a good estimation)...

Add that up first, even just landing fuel.

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3 hours ago, fredinno said:

No, using the vac. ISP gets way more Delta-V out of a rocket because the rocket equation is exponential, and most of the fuel is used in the atmosphere (ie why air-launch is useful). Also, you didn't add up landing equipment mass, heat shielding, reserve landing fuel, payload fairing mass (though that is ejected, it must be added for a good estimation)...

Add that up first, even just landing fuel.

 

 What I mean to say is the reduction of the ISP of the rocket at sea level causes a reduction of the delta-v achieved by the rocket by a known amount. So you can add this reduction onto the delta-v you are requiring for orbit, just like you add ca. 1,000 m/s for gravity drag and add ca. 100 m/s for air drag onto the 7,800 m/s delta-v needed for just the orbital speed.

 Example, when it is commonly said that you need in the range of 9 km/s delta-v to reach orbit, does this mean the rocket is traveling at 9 km/s when it reaches orbit? No, but you can use this larger value than just the orbital speed in your rocket equation calculations because you are aware of how gravity and air drag detract from your actually achieved delta-v. In the same way, you can just use the vacuum ISP in your rocket equation calculations once you add back onto the required delta-v to orbit the loss due to sea level ISP reduction.

 Note this means you are using a higher required delta-v to orbit than if using an average value of the ISP. You could use some estimate of the average ISP of the flight based on the sea level and vacuum ISP and then you would use a lower value of the required delta-v to orbit. The result would be the same either way for the payload. 

 The calculation was for an expendable SSTO with altitude compensation to show the payload could be in the range of currently used multistage rockets in that case. To make it reusable you would lose payload but this also happens with the multistage rockets.

  Bob Clark

Edited by Exoscientist

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9 hours ago, Exoscientist said:

 

 What I mean to say is the reduction of the ISP of the rocket at sea level causes a reduction of the delta-v achieved by the rocket by a known amount. So you can add this reduction onto the delta-v you are requiring for orbit, just like you add ca. 1,000 m/s for gravity drag and add ca. 100 m/s for air drag onto the 7,800 m/s delta-v needed for just the orbital speed.

 Example, when it is commonly said that you need in the range of 9 km/s delta-v to reach orbit, does this mean the rocket is traveling at 9 km/s when it reaches orbit? No, but you can use this larger value than just the orbital speed in your rocket equation calculations because you are aware of how gravity and air drag detract from your actually achieved delta-v. In the same way, you can just use the vacuum ISP in your rocket equation calculations once you add back onto the required delta-v to orbit the loss due to sea level ISP reduction.

 Note this means you are using a higher required delta-v to orbit than if using an average value of the ISP. You could use some estimate of the average ISP of the flight based on the sea level and vacuum ISP and then you would use a lower value of the required delta-v to orbit. The result would be the same either way for the payload. 

 The calculation was for an expendable SSTO with altitude compensation to show the payload could be in the range of currently used multistage rockets in that case. To make it reusable you would lose payload but this also happens with the multistage rockets.

  Bob Clark

But a reusable SSTO would fall down into the negatives, or into the smallsat class. Considering how much larger the F9 is, I highly doubt it'd compete vs even an expendable methane TSTO in cost.

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Posted (edited)

In his presentation last year on the Interplanetary Transport System (ITS), at about the 54 minute mark Musk discusses that the second stage in its tanker form or in its spaceship form will be able to reach orbit when used as a single stage. He states though the tanker will not be able to land, presumably because of insufficient reserve fuel. Then it could be an expendable SSTO.

 

 

 Here's a simulation of the ITS upper stage tanker as an SSTO:

ITS Tanker SSTO - YouTube.

It suggests it can get a total mass of 190 metric tons to LEO as an expendable. Since the dry mass is 90 metric tons, this means a 100 metric ton payload to orbit.

 Musk has said this ITS tanker upper stage will be available for testing by 2020, then we may have a viable high payload SSTO then. In fact I consider the next big advance in space access will be SSTO's.

   Bob Clark

Edited by Exoscientist

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If it could get 90 tons to LEO as an expendable, then it could probably get at least 20 tons and carry enough propellant to land.

I haven't looked into the details, but it sounds fishy to me. 

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On 3/15/2017 at 10:17 AM, Nibb31 said:

If it could get 90 tons to LEO as an expendable, then it could probably get at least 20 tons and carry enough propellant to land.

I haven't looked into the details, but it sounds fishy to me. 

  Perhaps one of the Kerbal heads on the forum could use the specs on the ITS upper stage and do their own simulation.

 

  Bob Clark

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I think the reason it's not reusable is probably the issues getting something that big to reenter safely/be controllable in atmosphere going the wrong way.

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8 minutes ago, Steel said:

I think the reason it's not reusable is probably the issues getting something that big to reenter safely/be controllable in atmosphere going the wrong way.

Not sure what you mean. The ITS Tanker is already going to need to reenter from orbit controllably.

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1 hour ago, sevenperforce said:

Not sure what you mean. The ITS Tanker is already going to need to reenter from orbit controllably.

It is?! Well I completely missed that part!

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41 minutes ago, Steel said:

It is?! Well I completely missed that part!

The entire ITS is supposed to be reusable. It's pretty much the whole point.

That simulation doesn't make much sense at all. If the ITS can put 90t into LEO as an SSTO, then there really is no use for the booster at all. You could do the refuel runs without it, you would just need to do more of them, but it would save a huge effort.

90t to LEO as an SSTO is more than the SLS Block I, which is multi-stage. It really doesn't make any sense at all.

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Posted (edited)

4 hours ago, Exoscientist said:

Perhaps one of the Kerbal heads on the forum could use the specs on the ITS upper stage and do their own simulation.

The ITS booster needs 3.3 km/s of dV for its boostback and return, which requires m0/m1 of 2.7; an orbital vehicle would need less due to not needing a boostback burn. I know the Falcon 9 goes transonic off drag alone, so the dV for the landing wouldn't be more than 500 m/s. 

1 hour ago, Nibb31 said:

That simulation doesn't make much sense at all. If the ITS can put 90t into LEO as an SSTO, then there really is no use for the booster at all. You could do the refuel runs without it, you would just need to do more of them, but it would save a huge effort.

90t to LEO as an SSTO is more than the SLS Block I, which is multi-stage. It really doesn't make any sense at all.

The ITS tanker carries 2,880 tonnes of propellant and has a projected dry mass of 90 tonnes. But since only three of the engines can be used at SL, with a combined thrust of 9,150 kN, achieving a TWR of 1.1 limits launch mass to 848 tonnes, which requires that the tanker launch with its tanks 74% empty. An SSTO launch would burn nearly straight up in order to ignite its higher-ISP engines as soon as possible; I'm guessing it will be able to do so at around 1 km/s but will incur 1 km/s of gravity and aerodynamic drag in order to get there. At this point, fuel reserves have dropped to 350 tonnes. Adding another 0.2 km/s of gravity drag means we still need 7.0 km/s, but those 350 tonnes of fuel can only get us 5.8 km/s of dV even with the higher-ISP Vacuum Raptors.

So I don't see how the ITS tanker can achieve SSTO at all.

If Elon plans on swapping three of the Vacuum Raptors out for SL raptors in an SSTO demonstrator, then launch thrust would jump to 18,300 kN and the ITS tanker could launch with 1,605 tonnes of propellant. Using the same numbers, the tanker would reach orbit with approximately 37 tonnes of residuals, which gives it 1.1 km/s of SL dV for a landing.

Edited by sevenperforce
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1 hour ago, Nibb31 said:

The entire ITS is supposed to be reusable. It's pretty much the whole point.

That simulation doesn't make much sense at all. If the ITS can put 90t into LEO as an SSTO, then there really is no use for the booster at all. You could do the refuel runs without it, you would just need to do more of them, but it would save a huge effort.

90t to LEO as an SSTO is more than the SLS Block I, which is multi-stage. It really doesn't make any sense at all.

 That's as an expendable SSTO. In the comments to that video the author estimates only 30 metric tons for payload as a reusable when you take into account the propellant that would need to be retained for the landing. If it is only 30 metric tons as a reusable that would be quite a huge number of flights needed to fully refuel a stage in orbit.

 

 BTW, about the SLS Block 1, many people believe the payload will be higher than the quoted 70 metric tons, perhaps in the range of 90 metric tons.

  Bob Clark

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40 minutes ago, sevenperforce said:

The ITS booster needs 3.3 km/s of dV for its boostback and return, which requires m0/m1 of 2.7; an orbital vehicle would need less due to not needing a boostback burn. I know the Falcon 9 goes transonic off drag alone, so the dV for the landing wouldn't be more than 500 m/s. 

The ITS tanker carries 2,880 tonnes of propellant and has a projected dry mass of 90 tonnes. But since only three of the engines can be used at SL, with a combined thrust of 9,150 kN, achieving a TWR of 1.1 limits launch mass to 848 tonnes, which requires that the tanker launch with its tanks 74% empty. An SSTO launch would burn nearly straight up in order to ignite its higher-ISP engines as soon as possible; I'm guessing it will be able to do so at around 1 km/s but will incur 1 km/s of gravity and aerodynamic drag in order to get there. At this point, fuel reserves have dropped to 350 tonnes. Adding another 0.2 km/s of gravity drag means we still need 7.0 km/s, but those 350 tonnes of fuel can only get us 5.8 km/s of dV even with the higher-ISP Vacuum Raptors.

So I don't see how the ITS tanker can achieve SSTO at all.

 

 I'm fairly sure the author was taking it as having 9 sea level raptors for the engines. Perhaps that's why Elon didn't think his version, with only 3 sea level Raptors, could be reusable.

 

  Bob Clark

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Posted (edited)

36 minutes ago, Exoscientist said:

 I'm fairly sure the author was taking it as having 9 sea level raptors for the engines. Perhaps that's why Elon didn't think his version, with only 3 sea level Raptors, could be reusable.

Well, that also changes the math. Nine SL Raptors can develop 2,798 tonnes-force, enough to lift 2,543 tonnes off the pad at a 1.1 T/W ratio. That's 2,453 tonnes of propellant, meaning the tanks would be about 85% full. That first 2 km/s (to get to 1 km/s) will cost 1,163 tonnes of fuel. But since there are no higher-efficiency, higher-thrust engines to ignite, this will kick the additional gravity drag up to 0.5 km/s, meaning we need 7.3 km/s more. The vacuum specific impulse of the SL Raptors is 361 m/s, which means orbit is reached with 85 tonnes of residuals. Allowing about 20 tonnes of fuel for deorbit and landing means 55 tonnes (of fuel) can be delivered in a single SSTO launch, as opposed to 380 tonnes in a single TSTO launch. I guess they'd rather do five TSTO launches than 40+ SSTO launches.

Especially since any decrease in performance or increase in the amount of fuel required for the landing could instantly double the number of SSTO launches required.

Edited by sevenperforce

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I know its too hard to dock even with automatic systems. But what about 5 full stack launches and then 2 reusable launches which refuel the 5 depleted on orbit to deorbit too?

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14 minutes ago, Toonu said:

I know its too hard to dock even with automatic systems. But what about 5 full stack launches and then 2 reusable launches which refuel the 5 depleted on orbit to deorbit too?

Each tanker in a full-stack TSTO launch launches with 2880 tonnes of propellant. It burns about 2480 tonnes to get from staging (2.4 km/s) to orbit, transfers 380 tonnes of propellant to the target vessel, and then uses about 20 tonnes of residuals to deorbit and land. There are no tankers left depleted in orbit.

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1 hour ago, sevenperforce said:

Each tanker in a full-stack TSTO launch launches with 2880 tonnes of propellant. It burns about 2480 tonnes to get from staging (2.4 km/s) to orbit, transfers 380 tonnes of propellant to the target vessel, and then uses about 20 tonnes of residuals to deorbit and land. There are no tankers left depleted in orbit.

 

 The amount I've seen for the propellant is 2,500 metric tons. Where are you getting 2,880 mT?

 

  Bob Clark

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14 minutes ago, Exoscientist said:

The amount I've seen for the propellant is 2,500 metric tons. Where are you getting 2,880 mT?

From here: http://spaceflight101.com/spx/its-spaceship/

"2,500 metric tons of propellant are carried by the tanker for use during its own mission plus 380 t of propellant upmass that can be transferred to the Spaceship."

That may be an error, of course. 

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Posted (edited)

1 hour ago, sevenperforce said:

From here: http://spaceflight101.com/spx/its-spaceship/

"2,500 metric tons of propellant are carried by the tanker for use during its own mission plus 380 t of propellant upmass that can be transferred to the Spaceship."

That may be an error, of course. 

 

 Thanks for that. Based on this image I was assuming the 380 mT was just for payload:

GsyREf7.png

 

  But for the tanker version, if it is carried inside the tanks with the regular propellant then presumably it could also be used for propulsion. Imagine the payload if we could squeeze another engine in there so we could launch this as full propellant load!

 Another possibility to increase the thrust might be a thrust-augmented nozzle(TAN). This acts like an afterburner for the rocket engine where additional fuel is injected into the nozzle:

Thrust Augmented Nozzles.

Posted on November 12, 2007 by Jonathan Goff
http://selenianboondocks.com/2007/11/thrust-augmented-nozzles/

  Bob Clark

Edited by Exoscientist
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Depending on how early the Vacuum Raptors can be ignited without dangerous levels of overexpansion, the optimized approach is either nine SL Raptors or seven SL raptors and two balanced Vacuum Raptors.

Even so, the best it can do as an SSTO is probably around 55 tonnes of fuel to LEO. But that's just much, much less than what it can do on the top of the ITS booster.

Another option would be to give it a parallel booster system, either with COTS solids, FH-style asparagus staging, or a launch skirt as with the Saturn-1D concept.

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On 3/20/2017 at 8:31 AM, Exoscientist said:

 Perhaps one of the Kerbal heads on the forum could use the specs on the ITS upper stage and do their own simulation.

I decided to go ahead and crunch the numbers for a small SSTO launcher using a production version of the 1000 kN dev Raptor and a basic air-augmentation ejector shroud. The math actually comes out pretty nicely.

The Raptor is supposed to have a better TWR than the uprated Merlin, which boasts 180:1, so let's set the mass of a production dev Raptor at 550 kg. Rule of thumb on an ejector shroud is that it will be 3-5 times the mass of the engine; I'll guess at 2 tonnes. With the ejector shroud giving a static thrust increase of 15%, the pad thrust will be 117.3 tonnes. Set GLOW at 100 tonnes, slightly less than the mass of the Falcon 9FT expendable second stage.

Base specific impulse for the SL Raptor is 334 seconds; underexpansion pressure will cause it to climb to 361 seconds as altitude increases. At the same time, the ejector shroud will boost the effective specific impulse, starting at 15% at zero airspeed and climbing to 50% at Mach 2. Starting around Mach 4.4, ram drag due to the increasing airspeed will start to sap the efficiency boost; the boost will drop to zero around 3.4 km/s. However, the ejector shroud will still be able to increase the expansion ratio slightly...probably to around 375 seconds. This is still less than the specific impulse of the Vacuum Raptor.

Working these numbers iteratively gives 11.2 tonnes of payload+vehicle+residuals in LEO. Reserving 500 dV at SL for landing leaves 9.59 tonnes. Fuel consumption is 90.41 tonnes. The ITS Tanker has a structure+tankage ratio of 97.4% including TPS and auxiliary thrusters; adjusting for some square-cube losses, I'll place the ratio here at 96%, for a tankage+airframe+TPS mass of 3.77 tonnes. This provides a vehicle dry mass of 6.32 tonnes, for a total payload of 3.27 tonnes.

3.27% is a fantastic payload fraction to begin with. Even better for a fully reusable vehicle smaller than the Falcon 9 second stage.

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