Trip to Mars on hypergolics

[lobe]

I was doing calculations for a hypergolics fuelled trip using tech from 1970-1980, and I came up with 40 tons to land, 292 tons to orbit, and 700 tons to inject to Mars. Is this correct?

Edited March 14, 2016 by lobe
Added in dates

[Steel]

You're going to have to be a little more specific. ISPs, fuel types, are you launching from earth or are you already in orbit?

[KSK]

700 tons seems a lot - how massive is your spacecraft?

[lobe]

I was assuming 40 tons to land on Mars for a three man crew (I added 15 tons for human resources and landing apparatus). The flight plan was leaving LEO into a Trans-Mars Injection, then orbital insertion, then landing and return to Low Mars Orbit (LMO). I was using 311 for lander Isp (two stage- landing and ascent) and 316 for transfer stages (one for Trans Mars Injection, second for capture). These are based on the Lunar Module engines and the RD-253 engines. 

 

Woops I added 15 to 25 tons for landing

Edited March 13, 2016 by lobe
MOAR INFO

[AngelLestat]

why you choose hypergolics fuel in the first place?

I dont see any trouble with methane + lox.
You have big volume in the tanks which reduce the surface and you are not close to earth shine (which does not matter either because earth albedo is not enough to boil methane or lox)
You dont even need active cooling or any kind of cooling,  just hide the tank from sunlight and other heat sources from the ship.
400 isp is much better than any hipergolic fuel, more when you need to save until the last kg in a mission to mars.

 

[kerbiloid]

28 minutes ago, AngelLestat said:

why you choose hypergolics fuel in the first place?

I dont see any trouble with methane + lox.
You have big volume in the tanks which reduce the surface and you are not close to earth shine (which does not matter either because earth albedo is not enough to boil methane or lox)
You dont even need active cooling or any kind of cooling,  just hide the tank from sunlight and other heat sources from the ship.
400 isp is much better than any hipergolic fuel, more when you need to save until the last kg in a mission to mars.

 

Buran was equipped with Syntin / LOx engines and its liquid oxygen was stored overcooled, in special tanks with stirring mechanism to keep it for 15-20 days without active cooling.
Mars mission needs at least 4 months just to get there, and 1.5 years - to return back. And many tons of fuel.
No problem to start with LH2/LO2 from the NEO, but looks very unclear how much cryofuel will be left near Mars.
(Looks like Methane is of no purpose here, as fuel cost doesn't matter.)

Edited March 13, 2016 by kerbiloid

[lobe]

I chose hypergolics (in this case Aerozine 50 and UDMH as fuels, nitrogen tetroxide for oxidizer) because they are not cryogenic, and won't have to worry about wasting any propellant due to boil off or equipment related to scavenging the boil off. As well, I was reading another thread concerning the benefits of a nuclear thermal rocket (using NERVA specs) and one post (thread was a month or two back) went through some calculations and came to the conclusion that if it was used, using hydrogen with a nuclear rocket would perform about the same as a regular chemical rocket due to added equipment for boil-off scavenging and the reactor itself. Or at least I thought that was the conclusion, if someone could find that thread that would be great.

The other reason was to look into what a 1970-80 era Mars mission might look like. Though I do say, the numbers look huge, but the mass suggests it would be about the size of a Proton Rocket.  

[AngelLestat]

4 hours ago, kerbiloid said:

Mars mission needs at least 4 months just to get there, and 1.5 years - to return back. And many tons of fuel.
No problem to start with LH2/LO2 from the NEO, but looks very unclear how much cryofuel will be left near Mars.
(Looks like Methane is of no purpose here, as fuel cost doesn't matter.)

First I said Methane.. no hydrogen..  -160c boiling point for methane  and -250c for hydrogen.
Space background temperature -270c
You can loose hydrogen due leaks due how small is its molecule with old tanks and longer times, but today we already had many options (as paint) to shield against h2. But h2 is not my option, is methane which does not leak and its boilling point is much higher keeping a good ISP.
You dont even need active cooling, and if you need it doesn´t add much mass neither consume energy (negligible). 

11 minutes ago, lobe said:

I chose hypergolics (in this case Aerozine 50 and UDMH as fuels, nitrogen tetroxide for oxidizer) because they are not cryogenic, and won't have to worry about wasting any propellant due to boil off or equipment related to scavenging the boil off. As well, I was reading another thread concerning the benefits of a nuclear thermal rocket (using NERVA specs) and one post (thread was a month or two back) went through some calculations and came to the conclusion that if it was used, using hydrogen with a nuclear rocket would perform about the same as a regular chemical rocket due to added equipment for boil-off scavenging and the reactor itself. Or at least I thought that was the conclusion, if someone could find that thread that would be great.

The other reason was to look into what a 1970-80 era Mars mission might look like. Though I do say, the numbers look huge, but the mass suggests it would be about the size of a Proton Rocket.  

You dont lose proppelent, also.. Elon Musk MCT will use methane.  
Can you tell me how much you think it weight a active cooling mechanism?  I can understand that you might have issues with hydrogen with -250c, but methane??  or liquid ox?   No sure what is the problem, I can assure you that even in the hydrogen, no matter how high is your active cooling mass estimation, never would be higher than the deltav lost using hipergolics.

(you are using 1970 papers and tech as guide?   we are in 2016!  )

[lobe]

I don't think delta-v is the problem here, the mass and mass fraction certainly are. As I stated before, I just want to see what a three-man mars mission in the years 1970-1980 would have looked like. Yes NERVA was around and almost flight ready but I wanted to see what hypergolics could do and appreciate the size of this spacecraft.

[B787_300]

okay  a couple of things. without baseline masses, isps, launch dates (in widow or not) this is a REALLY hard question to answer.  

Also all of these numbers are Really sensity to other factors.  I (and my senior design group) have a SSTO MAV that can carry 3 kg of payload to LMO and it only weighs 500 kgs, but we increase payoad the mass goes WAY up

[Streetwind]

Hypergolics is a competitive option, yes. Without having to manage the temperature of your propellant much, your tankage is going to have less dry mass. Their much higher density than typical high-Isp propellants (which by definition must be light and fluffy) also plays to that same strength. This can actually weigh up the Isp difference, because your spacecraft has a higher fuel mass fraction.

A lot of amateurs often make the mistake of going "highest Isp is best", while forgetting that the rocket equation is a multiplication of Isp and fuel mass fraction. If you drop the latter too much in relentlessly chasing the former, your dV is going to end up less, not more. And having to maintain the temperature of cryogenic propellants is unfortunately a great way to reduce your fuel mass fraction.

Because of this, I'm not surprised that you got a valid result for doing the math with hypergolics. I'm not saying it's necessarily the best option, as determining that can get arbitrarily complex (factoring in such things as the payload of your launch vehicle, which dictates how the parts of your spacecraft must be segmented, which in turn can play to the strengths of certain propellant choices over others etc). But it's a reasonably competitive option.

[fredinno]

 

2 hours ago, AngelLestat said:

First I said Methane.. no hydrogen..  -160c boiling point for methane  and -250c for hydrogen.
Space background temperature -270c
You can loose hydrogen due leaks due how small is its molecule with old tanks and longer times, but today we already had many options (as paint) to shield against h2. But h2 is not my option, is methane which does not leak and its boilling point is much higher keeping a good ISP.
You dont even need active cooling, and if you need it doesn´t add much mass neither consume energy (negligible). 

You dont lose proppelent, also.. Elon Musk MCT will use methane.  
Can you tell me how much you think it weight a active cooling mechanism?  I can understand that you might have issues with hydrogen with -250c, but methane??  or liquid ox?   No sure what is the problem, I can assure you that even in the hydrogen, no matter how high is your active cooling mass estimation, never would be higher than the deltav lost using hipergolics.

(you are using 1970 papers and tech as guide?   we are in 2016!  )

Yeah, and he wanted to see what a 1970 mission would look like. Elon is only using Methane on Mars as you can make it via ISRU- which adds risk, and was not in any major proposals until Mars Direct in the '90s.

And the mission mass still seems to be pretty big. What are the masses for the HAB, mission trajectory, etc?

[Streetwind]

14 minutes ago, fredinno said:

And the mission mass still seems to be pretty big. What are the masses for the HAB, mission trajectory, etc?

It's apparently not final yet, which is why SpaceX keeps delaying to talk about it. Elon Musk only wants to go public when he can be sure that they won't suddenly decide to redo the whole thing differently at some point.

Though there's that unconfirmed rumor that the BFR will be able to put 236 metric tons into LEO (without specifying what orbit exactly). And the broad-strokes goal for MCT is to put 100 metric tons down onto the surface of Mars. Those two figures help constrain things a bit, and some maths with the Raptor's projected vacuum Isp of ~380s might let you make some predictions.

Edited March 13, 2016 by Streetwind

[lobe]

Well, this might get more involved than I thought. This is only for one way, because I assume that since hypergolics are used a return stage can be launch separately. For the delta-v, I used the image below to simplify things. The lander is in two stages based roughly on the Apollo Lunar Module, the dry mass for the ascent stage is 2150 kg (wet 7480) and the mass without fuel in the descent stage is 14614 kg (50845 for it completely fuelled). Now, we need a orbital insertion stage, assuming a 10 ton dry mass for the tank and engines, I got 121613 kg. Finally, for Earth departure and Trans-Mars Injection, assuming a 15 ton dry mass, the total mass comes to 388952. I realize this is quite a bit smaller than my numbers from this morning, that is because I took out the burn to LEO which I accidentally included. I also realize this is much smaller than what it needs to be. Looking into this further, assuming a three man crew on a 670 day mission (http://www.brighthub.com/science/space/articles/6612.aspx ), I figure that our crew would use 12243 kg of food for the trip, which we can round to about 13 tons for safety's sake. For the Hab I would use Skylab ( https://en.wikipedia.org/wiki/Skylab ) at 68 tons, so our orbiting crew module is now 81 tons. For the lander, most of us here know my use of a heavily modified Lunar Module is incredibly dangerous, but if a hab has already been landed there it might not be so bad. If anybody has a better landing system I would accept it to make this a little less ham fisted.

 

For delta-v

For engines and fuel

RD-253

https://en.wikipedia.org/wiki/RD-253

Lunar Module ascent and descent engines

https://en.wikipedia.org/wiki/Apollo_Lunar_Module

Edited March 14, 2016 by lobe
Math error

[fredinno]

23 minutes ago, lobe said:

For the Hab I would use Skylab ( https://en.wikipedia.org/wiki/Skylab ) at 68 tons, so our orbiting crew module is now 91 tons. For the lander, most of us here know my use of a heavily modified Lunar Module is incredibly dangerous, but if a hab has already been landed there it might not be so bad. If anybody has a better landing system I would accept it to make this a little less ham fisted.

A LEM would be destroyed during ascent- the LEM is only designed for Vacuum. I would use a Apollo CM derived lander, despite the extra mass, as it is much more equipped. And you're going to need a HAB on the surface if you want to stay there for more than 14 days (the max. duration of the Apollo CM).

Skylab is also way too big, maybe just use the O2 tank of the S-IVB instead, plus the Skylab airlock module. You might need a storage module depending on how much stuff you need for the journey.

24 minutes ago, lobe said:

Looking into this further, assuming a three man crew on a 670 day mission (http://www.brighthub.com/science/space/articles/6612.aspx )

I would prefer 6 people, since even 4 is considered by NASA to be a little too much on the low end. That would also justify needing the H2 tank of the S-IVB for a mars mission.

25 minutes ago, lobe said:

For engines and fuel

RD-253

https://en.wikipedia.org/wiki/RD-253

No, Russian engines will not work in the '70s.

[AngelLestat]

2 hours ago, lobe said:

I just want to see what a three-man mars mission in the years 1970-1980 would have looked like. Yes NERVA was around and almost flight ready but I wanted to see what hypergolics could do and appreciate the size of this spacecraft.

And you mention that now? No sure from what topic this discussion is coming, but that was no mention in the OP.

1 hour ago, Streetwind said:

A lot of amateurs often make the mistake of going "highest Isp is best", while forgetting that the rocket equation is a multiplication of Isp and fuel mass fraction. If you drop the latter too much in relentlessly chasing the former, your dV is going to end up less, not more. And having to maintain the temperature of cryogenic propellants is unfortunately a great way to reduce your fuel mass fraction.

Really?  I thought that the biggest mistake that someone can make in rocket science is sent the high density low isp fuel to orbit instead the other way around.
mmm let's see what the manual says:
https://en.wikipedia.org/wiki/Multistage_rocket#Optimal
1. Initial stages should have lower Isp, and later/final stages should have higher Isp.
2. The stages with the lower Isp should contribute more ΔV.
3. The next stage is always a smaller size than the previous stage.
4. Similar stages should provide similar ΔV.

The rocket rise that payload from earth dont you?  Or is mining in space..  even if is the second, methane would be still a better option. 
Methane does not need active cooling!  I repeat.. just passive and it has a lot more density than hydrogen.
 

53 minutes ago, fredinno said:

Yeah, and he wanted to see what a 1970 mission would look like.

he should mention this before.

Quote

Elon is only using Methane on Mars as you can make it via ISRU- which adds risk, and was not in any major proposals until Mars Direct in the '90s.

And the mission mass still seems to be pretty big. What are the masses for the HAB, mission trajectory, etc?

Add risk? you know that the isru happens before the manned trip leaves earth, this also reduce the huge mass you need to brake on mars, which this is highest risk and constraint on a mars mission.
Mass seems to be pretty big?  can you elaborate?
---------------------------------------------------------------------
But well, now that we are in 1970..  mmm...  no sure..  everything seems hard in that time, the first probe landings in mars was way beyond that, and you want to land 40 tons?
You can not use aerobrake with that so the amount of fuel rise a lot, which is no good for a low isp fuel.

Edited March 14, 2016 by AngelLestat

[lobe]

Just now, AngelLestat said:

And you mention that now? No sure from what topic this discussion is coming, but that was no mention in the OP.

 I was a little rushed in the morning. Sorry for the confusion.

8 minutes ago, AngelLestat said:

Methane does not need active cooling!  the third time I repeat this..  just passive and it has a lot more density than hydrogen.

2 minutes ago, AngelLestat said:

You can not use aerobrake with that so the amount of fuel rise a lot, which is no good for a low isp fuel.

Methane doesn't, but liquid oxygen still boils off. I did calculate for a powered landing. A 40 or 50 ton lander is not small, and as I have state before that is unrealistically small for a manned landing of any duration.

8 minutes ago, fredinno said:

No, Russian engines will not work in the '70s.

I was willing to ignore the political situation and focus purely on the technical. However the Service Propulsion System has almost the same impulse as the RD-253, at 314 s.

[fredinno]

On 3/14/2016 at 6:10 PM, AngelLestat said:

And you mention that now? No sure from what topic this discussion is coming, but that was no mention in the OP.

Really?  I thought that the biggest mistake that someone can make in rocket science is sent the high density low isp fuel to orbit instead the other way around.
mmm let's see what the manual says:
https://en.wikipedia.org/wiki/Multistage_rocket#Optimal
1. Initial stages should have lower Isp, and later/final stages should have higher Isp.
2. The stages with the lower Isp should contribute more ΔV.
3. The next stage is always a smaller size than the previous stage.
4. Similar stages should provide similar ΔV.

The rocket rise that payload from earth dont you?  Or is mining in space..  even if is the second, methane would be still a better option. 
Methane does not need active cooling!  I repeat.. just passive and it has a lot more density than hydrogen.
 

he should mention this before.

Add risk? you know that the isru happens before the manned trip leaves earth, this also reduce the huge mass you need to brake on mars, which this is highest risk and constraint on a mars mission.
Mass seems to be pretty big?  can you elaborate?
---------------------------------------------------------------------
But well, now that we are in 1970..  mmm...  no sure..  everything seems hard in that time, the first probe landings in mars was way beyond that, and you want to land 40 tons?
You can not use aerobrake with that so the amount of fuel rise a lot, which is no good for a low isp fuel.

ISRU was considered risky, and still is, and would need numerous robotic tests before humans would be allowed to depend on it to get off Mars. That's what NASA is doing right now on Mars 2020. ISRU wasn't seriously considered until Zubrin demonstrated it was possible to NASA. Thus a 70s proposal is unlikely to use it.

Guess what? Boil off is still a problem, for methane. http://www.eng-tips.com/viewthread.cfm?qid=312901

A 70s landing can also not aerobrake until they send enough probes to understand Mars' atmosphere sufficiently to do so.

[AngelLestat]

On 3/14/2016 at 5:25 PM, lobe said:

Methane doesn't, but liquid oxygen still boils off. I did calculate for a powered landing. A 40 or 50 ton lander is not small, and as I have state before that is unrealistically small for a manned landing of any duration.

Only 20 degree of diference with methane and lox boils at 70 degree higher than hydrogen boiling point.
Is still in the passive cooling range.
About the mass, I dont care if is realistic or not for a manned mission..  just pointing that we need a lot of fuel and Supersonic Retropropulsion which was first achieved by spacex some years ago at lower reentry speeds.
AeroBraking does not work, because the volume -surface -  mass relation, you increase just a bit the surface of your heatshield capsule design (S2) and the volume increase (V3), the volume in these cases is directly proportional to your mass, this mean that every time you increase your mass, the heatshield surface does no increase in proportion to allow an aerobrake.
Even drag chutes are out of the question at those speeds, so the only option is to use your whole tank+capsule in vertical to increase the surface drag (with supersonic retropropulsion), but it would be hard to keep that inclination all your way down, by hard I mean almost impossible with 1970 tech. 
A manned mission to venus (havoc mission style) would be much easier for that time, you have a thick atmosphere to brake and no land to hit. 

 

On 3/14/2016 at 5:32 PM, fredinno said:

ISRU was considered risky, and still is, and would need numerous robotic tests before humans would be allowed to depend on it to get off Mars. That's what NASA is doing right now on Mars 2020. ISRU wasn't seriously considered until Zubrin demonstrated it was possible to NASA. Thus a 70s proposal is unlikely to use it.

The ISRU part of Zubrin was no the detail of his mission considered risky, it was the time, the crew and the low payload and simplicity of the mission, and those who said it was risky does not respond to any quality review panel, they would said that no matter the type of mission, because there was no real interest from NASA and the government to go mars in that moment.

Also you said:

Quote

Elon is only using Methane on Mars as you can make it via ISRU- which adds risk, and was not in any major proposals until Mars Direct in the '90s.

Can you show me other proposals from NASA for a manned mars mission using hypergolic fuels? 

Quote

Guess what? Boil off is still a problem, for methane. http://www.eng-tips.com/viewthread.cfm?qid=312901

hahaha, lets start pointing that this source has a doubtful quality
And you did not even understood what they are talking about.
He shows a leak rate, but you dont know nothing about the tank dimensions or the environment (in earth, in orbit facing the sun, or in any other place)
He did not even clarify what was its unit of time.
Now, lets see if you still understand the main point here, if your tank is under the boiling point temperature, it does no boil off..  

[lobe]

I looked into this a little bit more, and came up with some still quite amazing numbers. To start off the mission profile is 670 days, now with 6 crew as per fredinno's recommendation and has a Skylab sized habitation module. Food, water, and air for this crew comes to 24.5 tons. The applicable delta-v budgets are 3.6 km/s for Trans-Mars Injection, 2.7 km/s for Martian orbital insertion, and 3.8 km/s for descent/ascent, as per the delta-v map I posted earlier in this thread. I will assume that the specific impulse for all engines are based on some variation of the Service Propulsion System engine, which is 314 seconds.

The lander is still two stage. The ascent stage I based off of the dry mass of the Apollo CSM, at about 12 tons, adding another 1.2 tons for supplies of a 3 crew for 30 days on the ground, coming to 13.2 tons. Fuelled mass is 45.4 tons. The descent stage takes that number and adds 10 tons, when this is fuelled there is a total lander weight of 190.4 tons. Yes, AngelLestat, this does and my previous calculations include the the total 7.6 km/s it takes to land and depart Mars, because aerobraking is for wimps. Also, this is tech pretty much from Apollo that I am using, and the end date for the tech I can use is up to 1980, the launch doesn't need to be in that time frame. By 1980 we already landed 2 probes on Mars (Viking 1 and 2, 1976), so I assume that the development program they used could be scaled up. 

Skylab weighed about 68 tons, adding 24.5 to that gives us 92.5 tons. This is added to the lander, and will assume that the tank Skylab is attached to is 25 tons. This brings the unfuelled mass to 307.9 tons, fuelled 740.4 tons. Since we are now moving something with the fuelled mass over the Proton rocket (693.8 tons) the tank and the amount of engines are going to be pretty massive, so I assumed another 50 tons for this departure stage. This now makes the rocket 2,546 tons fuelled. To put this in perspective, this is about 3 and 2/3 Proton rockets, or about 85% of a Saturn V. It would take 18(actually 18.1, but you aren't going to launch 0.1 of a rocket) Saturn V launches to complete this single spacecraft.

Edited March 14, 2016 by lobe
Gnarly math error

[fredinno]

1 hour ago, AngelLestat said:

Only 20 degree of diference with methane and lox boils at 70 degree higher than hydrogen boiling point.
Is still in the passive cooling range.
About the mass, I dont care if is realistic or not for a manned mission..  just pointing that we need a lot of fuel and Supersonic Retropropulsion which was first achieved by spacex some years ago at lower reentry speeds.
AeroBraking does not work, because the volume -surface -  mass relation, you increase just a bit the surface of your heatshield capsule design (S2) and the volume increase (V3), the volume in these cases is directly proportional to your mass, this mean that every time you increase your mass, the heatshield surface does no increase in proportion to allow an aerobrake.
Even drag chutes are out of the question at those speeds, so the only option is to use your whole tank+capsule in vertical to increase the surface drag (with supersonic retropropulsion), but it would be hard to keep that inclination all your way down, by hard I mean almost impossible with 1970 tech. 
A manned mission to venus (havoc mission style) would be much easier for that time, you have a thick atmosphere to brake and no land to hit. 

 

The ISRU part of Zubrin was no the detail of his mission considered risky, it was the time, the crew and the low payload and simplicity of the mission, and those who said it was risky does not respond to any quality review panel, they would said that no matter the type of mission, because there was no real interest from NASA and the government to go mars in that moment.

Also you said:

Can you show me other proposals from NASA for a manned mars mission using hypergolic fuels? 

hahaha, lets start pointing that this source has a doubtful quality
And you did not even understood what they are talking about.
He shows a leak rate, but you dont know nothing about the tank dimensions or the environment (in earth, in orbit facing the sun, or in any other place)
He did not even clarify what was its unit of time.
Now, lets see if you still understand the main point here, if your tank is under the boiling point temperature, it does no boil off..  

When you run out of arguments is time to start to talk in CAPS?
Repeat what?  that link that shows the lack of boiling understanding?
 

Aerobraking is possible, but you'd need to do it very slowly (only good for short duration surface missions) and need to develop large ballutes.

Yes, ISRU was considered risky at that time- NASA would not consider their proposal until Zubrin experimentally proved ISRU worked.

http://www.permanent.com/space-transportation-propellants.html

LOx and CH4 (which has a slightly higher boiling point) do boil off at Mars distances, unless you put a low of insulation on, which I wanted to avoid for the increased amount of mass. I know you don't have any boil off if you keep the lander cool, but then, that means a more complex mission and lander to make sure it stays cool.

https://books.google.ca/books?id=abIKvmDXh_kC&pg=PA121&lpg=PA121&dq=hypergolic+mars+lander&source=bl&ots=557GifVghg&sig=fXHR09y4zk9Ogikb-QkltsbTAT8&hl=en&sa=X&ved=0ahUKEwit_LH1qr_LAhWJKGMKHZ26A8AQ6AEIJjAF#v=onepage&q=hypergolic%20mars%20lander&f=false

It's possible to do mars landers sans ISRU, but using methane, Von Braun proposed it, but I went with the safest option (probably the best option on something as essential as landing and getting off Mars)

37 minutes ago, lobe said:

I looked into this a little bit more, and came up with some still quite amazing numbers. To start off the mission profile is 670 days, now with 6 crew as per fredinno's recommendation and has a Skylab sized habitation module. Food, water, and air for this crew comes to 24.5 tons. The applicable delta-v budgets are 3.6 km/s for Trans-Mars Injection, 2.7 km/s for Martian orbital insertion, and 3.8 km/s for descent/ascent, as per the delta-v map I posted earlier in this thread. I will assume that the specific impulse for all engines are based on some variation of the Service Propulsion System engine, which is 314 seconds.

The lander is still two stage. The ascent stage I based off of the dry mass of the Apollo CSM, at about 12 tons, adding another 1.2 tons for supplies of a 3 crew for 30 days on the ground, coming to 13.2 tons. Fuelled mass is 45.4 tons. The descent stage takes that number and adds 10 tons, when this is fuelled there is a total lander weight of 190.4 tons. Yes, AngelLestat, this does and my previous calculations include the the total 7.6 km/s it takes to land and depart Mars, because aerobraking is for wimps. Also, this is tech pretty much from Apollo that I am using, and the end date for the tech I can use is up to 1980, the launch doesn't need to be in that time frame. By 1980 we already landed 2 probes on Mars (Viking 1 and 2, 1976), so I assume that the development program they used could be scaled up. 

Skylab weighed about 68 tons, adding 24.5 to that gives us 92.5 tons. This is added to the lander, and will assume that the tank Skylab is attached to is 25 tons. This brings the unfuelled mass to 307.9 tons, fuelled 740.4 tons. Since we are now moving something with the fuelled mass over the Proton rocket (693.8 tons) the tank and the amount of engines are going to be pretty massive, so I assumed another 50 tons for this departure stage. This now makes the rocket 2,546 tons fuelled. To put this in perspective, this is about 3 and 2/3 Proton rockets, or about 85% of a Saturn V. It would take 9 (actually 8.2, but you aren't going to launch 0.2 of a rocket) Saturn V launches to complete this single spacecraft.

Sounds about right. Which is why you need to use Aerobraking and ISRU, plus robotic missions to prove it works. Otherwise, it's profibitively expensive. 

This may also be a good idea if you have to have a lot of launches regardless: http://www.astronautix.com/lvs/satv25sb.htm

[lobe]

38 minutes ago, fredinno said:

This may also be a good idea if you have to have a lot of launches regardless: http://www.astronautix.com/lvs/satv25sb.htm

Thank you for posting this, it helped me find an error. I said 9 launchs, that was because I read 310000 kg instead of lbs, so this made it twice as practical. In reality, it takes 18 Saturn-V launches, or 12 Saturn-V25(S)B launches. I was wondering why the SuperSaturn was weaker than the Saturn.

As for aerobraking, I omitted any need of that with my propellant calculations. If Mars where completely devoid of atmosphere, my spacecraft would still be able to complete its mission.

Edited March 14, 2016 by lobe

[DBowman]

1 hour ago, lobe said:

To start off the mission profile is 670 days, now with 6 crew as per fredinno's recommendation and has a Skylab sized habitation module. Food, water, and air for this crew comes to 24.5 tons.

I think you may be able to go lighter on the consumables. I used NASA 'Baseline Assumptions and Values' doc and derived a 1,400 kg for 500 man days. Your mission is 8.04 of those, so 11.256 tons + a couple hundred kilos fixed costs - so call it 11.5 tons. I did a few things to keep it minimal, NASA was talking about 70s versions of all these techniques for extended flights:

• 85% water recovery from urine using forward osmosis. For Apollo extensions they had planed long flights using evaporative recovery of water from urine - close enough...
• recovery of metabolically created water & sweat/respiration vapor recovery
• CO2 removal via ISS style CDRA (CO2 Removal Assembly)
• No hygiene water - wet wipes, drink your toothpaste, and use any excess of the metabolically created water

[lobe]

14 minutes ago, DBowman said:

I used NASA 'Baseline Assumptions and Values' doc

Do you have the link for this? I would like to look at it, I got my numbers from a blog of all places http://www.brighthub.com/science/space/articles/6612.aspx

 

[DBowman]

2 minutes ago, lobe said:

Do you have the link for this

a post of mine the doc is linked in that post and also a bunch of other docs you might find interesting - the Apollo 70s Venus flyby especially, they have mass estimates for a not Skylab Hab (cannot recall the numbers though).