Mars Colonial Transporter

SpaceNut · 2017-01-20 19:20:45

Found these links while searching for Dragon stuff.....

SpaceX Successfully Tests Carbon Fiber Tank for Mars Spaceship

The tank is 12 meters (~39 feet) wide, the largest such vessel ever produced. carbon fiber prepreg” – carbon fiber that’s pre-impregnated with a resin to make it tougher. “In theory, it should hold cryogenic propellant without leaking and without a sealing linker

MCT/BFR Realistic Analysis

Lots of details in this forum blog post.

Zubrin improves the Spacex Mars plan and points out reusable Falcon heavy can be a globespanning rocketplane with Boeing 737 capacity

elderflower · 2017-01-21 07:27:52

I wouldn't trust it for strong oxidisers without an impervious liner.

GW Johnson · 2017-01-21 15:19:48

I'm with Elderflower. Carbon touching LOX is a BAD idea.

GW

Oldfart1939 · 2017-01-21 16:45:19

Contact of LOX with carbon fiber has already been discussed in the SpaceX 1 September "anomaly." I'm very skeptical of this approach unless a PTFE-type polymer lines the tank.

SpaceNut · 2017-01-21 21:31:08

Reading the forums numbers a rocket having a first stage engine count of 60!!!! wow what the wrong way to go...

Making a two-stage rocket that is intended to deliver around 100 metric tons of cargo into LEO.. with the first stage being recovered in the same manner as the Falcon 9....

Of course the article talks about fuel type and what can be done...

The heavy version of the Falcon 9 is said to be able to get 53 ton to orbit and that will be with the triple barrel count of 27 engines so if we add 2 more cores to make a 5 barrel first stage then we are are just under 90 tons with that configuration and that's a huge rocket for launch.....

RobS · 2017-01-21 22:41:29

The current Heavy version doesn't even use transfer of fuel to the central core from the outer two stages, either, which the original Heavy was supposed to do. If that technology were added to it, even the Falcon Heavy could lift more; probably closer to 70 tonnes. It may be they dropped that feature because the core first stage would come in too "hot" to be recovered, though.

RobertDyck · 2017-01-22 05:33:20

Are you serious? Falcon Heavy was supposed to have side boosters slightly taller than the core, while the core was the same size as Falcon 9. Then Gwynne Shotwell obsessed over the number of models of rocket stage she would have to maintain in inventory. She had them change it so the core stage and side boosters were the same. The result was Falcon 9 "full thrust" aka Falcon 9 v1.2. Now fuel transfer is deleted? That would hamper Falcon Heavy so much that you can't call it Falcon Heavy any more. Falcon Heavy was supposed to lift 53 metric tonnes to LEO when it had the Falcon 9 v1.1 core stage and fuel transfer. Now that it has a v1.2 core, they're saying 54 tonnes. But without fuel transfer, it most likely won't be able to lift even that.

On the other hand, some of us talked about a modification using Raptor engines instead of Merlin. Using Merlin for side boosters means fuel transfer would not be possible. If they don't have fuel transfer now, that makes this modification fairly easy.

RobS · 2017-01-22 08:03:26

The Falcon Heavy is supposed to lift 54 tonnes to LEO WITHOUT fuel transfer. Consider that the Falcon 9 v1.2 can lift 22.8 tonnes to LEO if the first stage is expended (http://www.spacex.com/falcon9). Even without fuel transfer, they can cut back on the thrust of the central core as the weight of the vehicle diminishes and save some of its fuel for a later staging event. The Falcon was upgraded significantly when v1.2 came out; it was something like 60% more capable. This was necessary to make first stage return standard and not diminish its basic orbital capacity. With fuel transfer, the Falcon Heavy probably would approach the Space Launch System's first configuration in launch capacity. Maybe that's one reason they dropped fuel transfer; to avoid the political heat. I suspect they didn't anticipate the need for a vehicle able to launch more than 54 tonnes and there was also the issue of development time and cost (which at this point would delay development of the MCT). They also can't afford launch failures of experimental Falcons, because the entire Falcon family would get grounded.

RobS · 2017-01-22 08:22:20

I should add that Musk has now announced a new "final" version of the Falcon 9 to replace v1.2, the "Block 5." It is supposed to go into production later this year. The main changes are higher thrust on the Merlin engines and improved landing legs. The thrust, however, will be 8,451 KN rather than 7,607 KN of the v1.2, which is another 11% increase. Presumably, it can put even more into LEO, and the Falcon Heavies flying with it will also have increased payload capacity. An 11% increase in 54 tonnes in 60 tonnes, for example. The information about the Block 5 is here: https://en.wikipedia.org/wiki/Falcon_9.

Oldfart1939 · 2017-01-22 13:02:57

From all my readings, the Falcon Heavy at ~ $98 Million per launch is a huge bargain when compared to the SLS. Just "wondering out loud," as to whether SpaceX will develop a new heavier-still version rocket, performance of which, exceeds the basic Falcon design by using CH4/LOX propellants and the new Raptor engines? This would seem to be a logical developmental pathway, from the business end of things, to proceed. I know there are some numbers floating around, but haven't taken the time to track them down, about the performance of the Raptor engines as compared to the Merlin D engines. I recall the Isp for the vacuum version being stated by Musk as "approaching 383 seconds." I'm just offering a rational/logical pathway forward towards a first Mars manned mission by 2024. I just don't see the massive Colonial Transporter being brought to fruition that soon, but an intermediate Super Heavy "Pseudo-Falcon" could be...

Last edited by Oldfart1939 (2017-01-22 17:05:50)

RobS · 2017-01-22 13:08:36

I have wondered that as well.

SpaceNut · 2017-01-22 21:43:41

Methane fuel means a heavier rocket for launching....

Oldfart1939 · 2017-01-23 00:25:08

SpaceNut wrote:

Methane fuel means a heavier rocket for launching....

On what basis is this comment made? The tankage for CH4 should be approximately similar in construction to the LOX tanks on a Falcon 9, and I'd assume a supercooled CH4 for densification. Methane has a lot higher Isp than RP-1, which should offset any weight disadvantages. Maybe GW Johnson can jump in here to clarify?

GW Johnson · 2017-01-23 09:50:50

Well, my ancient Pratt & Whitney vest pocket handbook lists 1000 psia c* as 5900 ft/sec for LOX/RP-1 and 6120 ft/sec for LOX/methane. Since Isp is CF c*/gc, that means your specific impulses are roughly 3.7% higher with LOX/methane, as long as your nozzles get designed to the same expansion conditions and nozzle kinetic energy efficiencies. The kind of cycle tap-off by which you run your propellant pumps makes a bigger difference than that.

For you metric guys, there are 3.2808333 feet in every meter, and the gc you use is 9.8067 instead of 32.174. It's the same CF regardless.

Same reference lists r = wox/wfuel as 2.55 for LOX/RP-1 and 3.15 for LOX/methane. Both use the same cryogenic-capable, venting-equipped tankage for the LOX, and it has to be a Dewar if more than multiple hours of boiloff lifetime are required. The kerosene is 28% of the propellant mass in that system, and can be a simple fuel tank, which is lighter. The methane is 24% of the propellant mass in that system, and requires the heavier cryo tank.

So, yeah, the methane system is slightly heavier because of the heavier cryo tankage for the fuel. But I doubt the real design difference is all that large, because the oxygen cryo tankage dominates the designs anyway, and both kinds of tanks must resist all the flight loads in addition to internal pressurization.

Methane has become popular because (1) you can make it, but not kerosene, anywhere you can find sources for CO2 and H2O, and (2) it may (or may not) offer some reusability advantages: kerosene tends to coke inside the regenerative cooling passages of the engine. I don't really know, but methane may not be much of a coking risk.

GW

PS -- that same reference lists c* at 1000 psia for NTO-plain hydrazine as 5860 ft/sec - almost identical to LOX-kerosene's 5900 ft/sec (NTO and 50-50 blend of hydrazine and UDMH is lower at 5560 ft/sec, so there are things to be lost getting the properties you want). Kerosene and methane/LOX are just not that much better! Room temperature storables that do not boil off for years and years. Hmmmm!!!

PPS -- The "subway" chart was pretty neat. Thanks. For destinations with atmospheres, I noticed some pretty high delta-vees, reflecting gravity and drag losses factored in. Those would be specific to some unspecified design, so be careful. You cannot use those reversibly.

Last edited by GW Johnson (2017-01-23 10:05:02)

Tom Kalbfus · 2017-01-23 10:28:46

I notice its easier to get to Venus, Mars, Phobos, and Deimos than our own Moon. (Assuming the planets are lined up right) The Moon doesn't have an atmosphere to brake off of. It should be easier to drop off supplies on Mars if you do it ahead of time. Venus is easier to get to than Mars, but not easier to get from! So long as your mission to Venus is one-way, that works out fine.

Last edited by Tom Kalbfus (2017-01-23 10:30:28)

GW Johnson · 2017-01-23 11:34:43

Lessee, low Earth orbit equatorially is 8.0 km/sec. A slick launcher operated efficiently will have about 5% gravity and drag losses, for a net required 8.4 km/sec.

The factor for plane change to off-equatorial is roughly 2 sine (half the plane change angle) applied to orbit speed of 8 km/sec. This is an overkill at 3.2 km/sec for a 23 degree plane change, since you really change course early in the first stage burn, right off the pad at subsonic speed. Call it 8.5 or 8.6 km/sec as a good guess, not the 11.6 km/sec that you get by adding these two numbers.

So, why is LEO listed as 9.4 km/sec? That's a lot of drag and gravity loss!

GW

Last edited by GW Johnson (2017-01-23 11:36:29)

Oldfart1939 · 2017-01-23 11:38:12

My understanding of this problem may be somewhat different from the straight aerospace/rocket motor engineer, and as a chemist have a slightly different "take" on things. Use of RP-1 requires an extremely high LOX ratio to RP-1, greater than the stoichiometry would indicate in order to avoid the coking problem. Switching to CH4 should allow getting the ratio back closer to the "as written" chemical reaction, thus lowering the amount of surplus oxygen the vehicle needs carry. I've heard numbers bandied about by SpaceX regarding the actual "observable" Isp for the vacuum engine using CH4 as 382 seconds. On the other hand, reported numbers, again "observed" for RP-1 as being around 300 seconds.

Another advantage of CH4 is the extremely low Melting Point (not actually "freezing point") of -182.48 C, and a Boiling Point of -161.49 C. As a result, methane is extremely pure and free from contaminants which are usually present even in highly refined kerosene (RP-1). The density was also given, but the temperature at which listed was absent. d = 0.4, roughly half that of RP-1. As usual, a tradeoff game the constructors will decide...

Figures for CH4 are from the Handbook of Chemistry and Physics, 48th Ed.

Oldfart1939 · 2017-01-23 11:50:48

I believe the Red Dragon 2 that Musk is thinking of for his first Mars landing is fuelled by a Hydrazine/NTO couple. Not sure whether they are using UDMH or not, or possibly MMH?

Last edited by Oldfart1939 (2017-01-23 11:51:47)

GW Johnson · 2017-01-23 12:18:40

What's listed for v1 and v2 is NTO-MMH. My old Pratt handbook doesn't list data for NTO-MMH, but I've heard it's pretty similar to plain hydrazine performance-wise. There's a power-law dependence of c* with chamber pressure, exponent in the vicinity of 0.01. It's higher at higher chamber pressures. Using high chamber pressures (3000-5000 psia) and long expansion cones for higher vacuum performance is where the higher vacuum Isp numbers come from.

The data I was quoting above includes references to the oxidizer/fuel ratio used in the engine, which is a much weaker function of chamber pressure than c*. Supposedly these ratios were already "optimized" by for best Isp in real test data. LOX/RP-1 with r=2.55 is running rich on kerosene. Stoichiometric would be closer to 3.4. I haven't run a stoichiometry on methane, but I'd bet that r= 3.15 is rich on methane, not stoichiometric or excess-oxygen. I'd guess about 3.4 to 3.5 is stoichiometric with methane. Running fuel-rich gets them more regenerative coolant to work with.

Red Dragon is a variant of V2 with the seats, chutes, and life support deleted, and cargo/payload racks and equipment installed. And probably a slightly-thinner heat shield. Mars entry even from an interplanetary trajectory is far less demanding than that of a from-Mars free-return to Earth, for which the v2 heat shield (listed as 8 cm thick) was designed.

The old data sheets show a larger propellant tank set in v2 (and presumably Red). Just over 1200 kg in v1, just over 1800 kg in v2.

GW

Last edited by GW Johnson (2017-01-23 12:28:31)

RobS · 2017-01-23 13:02:13

All the engines run fuel rich, I understand, because (1) it avoids oxidation of the engine parts, and (2) if you run fuel-rich, you get water and carbon monoxide, and that has a lower molecular weight than water and carbon dioxide. As a result, you get a higher ISP even though there is incomplete combustion. I think somewhere in The Case for Mars, Zubrin says the stoichiometric ratio of oxygen to methane is 4:1 but the ideal mix for maximum ISP is more like 3.5:1.

Oldfart1939 · 2017-01-23 16:50:52

The stoichiometric ratio of O2 and CH4 is a ratio of 64 to 16 in mass. CH4 + 2 O2 ----> CO2 + 2 H2O. If less oxygen is used there will be carbon monoxide formed, CO. There is also as a consequence, some deposition of carbon, or coking. I haven't seen the fuel/oxidizer ratio for either RP-1 or CH4 in the Merlin 1D+ engine versus the new Raptor.

louis · 2017-01-23 18:53:21

It's a great map but why do these sorts of maps never show how much fuel you need to carry to LEO for each "stop"? - that would be much more meaningful for the layperson.

SpaceNut wrote:

Methane fuel means a heavier rocket for launching....
https://i.imgur.com/AAGJvD1.png

SpaceNut · 2017-01-23 22:22:15

It is due to not knowing the payload being pushed to the destination, fuel type to engine count and time wanted to arrive at a destination for a burn time....

SpaceNut · 2017-02-05 09:22:41

More details to size of the vehicle from new topic

SpaceNut wrote:

From Article section just on Mars response:
Real Mars and Spacex Plans
The current Mars plan is:
1. Send Dragon scouting missions, initially just to make sure we know how to land without adding a crater and then to figure out the best way to get water for the CH4/O2 Sabatier Reaction.
2. Heart of Gold spaceship flies to Mars loaded only with equipment to build the propellant plant.
3. First crewed mission with equipment to build rudimentary base and complete the propellant plant.
4. Try to double the number of flights with each Earth-Mars orbital rendezvous, which is every 26 months, until the city can grow by itself.
Item 1
Test landing system on Red Dragon Mission, but is it scalable for a human mission when we can not launch it as a whole system from Earths gravity well. That said we will doing subassembly on orbit to go to mars.
Explore for water requires drilling or carving out pieces of mars icefield with procesing of said collected waters depending on content value and contamination level contain complicates a direct means to providing water for electrolysis for the Sabatier reactor to make fuel with.
Create Mars Fuel, relies on finding water that is useable. No useable water means shipping in water or seed hydrogen to start the process for making fuel on mars using the atmosphere. Which brings me to the methods to capture mars atmosphere in great enough quantity to make the desired levels of fuels for ascent back to orbit or all the way directly back depending on what we land and its mars archetecture used.
Item 2
This Heart of Gold spaceship is the Mav ISPP testing phase for mounting a human mission to follow. It is based on item 1 success and scaled up for what we expect to amke fuel for in the later sized mission to follow.
Item 3
This first crew is sized based on out come of not only item 1 but item 2 scaling for this mission archetecture. It does not tell us if its archetecture is battlestar galatica, small chuck assembly in orbit, mars direct, mars semi direct.... its the least informative. It also relies on the duration of the mission as well.
Item 4
Is the flag and foot step defense ti item 3 calling on funding to stay going long after starting mission to mars for the purpose of colonizing and not just for science. This also relies on the MTV or Interstellar Transport developement that he is planning for a using at some point in time for mars missions.
Not listed item goal 1
The Interplanetary Transport system can launch 550 tons to low earth orbit which is nearly four times as much as the Saturn V. It would be over four times as powerful as the SLS in the final version of the SLS.
Have the second stage go only out to the distance of the moon and return to enable 5 payloads to be sent instead of one.
Leave the 100 person capsule on Mars and only have a small cabin return to earth.
This is a ship which has roughly 4 to 5 time the current first stage and a half of the Ares V which nasa is designing to fly soon.
Another not listed goal 2
use the refueling in orbit and other optimizations to enable a Falcon Heavy to deliver 40 tons to Mars instead of 12 for exploration missions in 2018, 2020 etc.
SpaceX ITS would consist of a very large two-stage fully-reusable launch system, powered by methane/oxygen chemical bipropellant.
Since the second-stage-plus-spaceship will have used its fuel in getting to orbit, it would need to refuel in orbit, filling up with about 1,950 tons of propellant (which means that each launch carrying passengers would require four additional launches to deliver the necessary propellant). Once filled up, the spaceship can head to Mars.
Other than Russia transfer of fuels to the ISS from Progress the US has not done so yet let alone methane Lox.
Last item 3 goal
The duration of the journey would of course depend on where Earth and Mars are in their orbits; the shortest one-way trip would be around 80 days, according to Musk’s presentation, and the longest would be around 150 days. (Musk stated that he thinks the architecture could be improved to reduce the trip to 60 or even 30 days.)
How does this happen when a mars orbital alignment cycle is 2 yrs 7 weeks for earth launch to mars....All things equal then 6 - 8 months earth leads mars for launch get there is 6 - 8 months to return as well as mars leads earth upon returning. That leave surface stay time of 1 yr 7 weeks or down to 8 months 7 weeks and anything in between as possible.

SpaceNut · 2017-02-05 09:29:44

There are issues with the fuel quantity reload and mass of the stage to orbit as payload or am I really off.....
See unlisted goal quotes.

New Mars Forums

Announcement

#26 2017-01-20 19:20:45

Re: Mars Colonial Transporter

#27 2017-01-21 07:27:52

Re: Mars Colonial Transporter

#28 2017-01-21 15:19:48

Re: Mars Colonial Transporter

#29 2017-01-21 16:45:19

Re: Mars Colonial Transporter

#30 2017-01-21 21:31:08

Re: Mars Colonial Transporter

#31 2017-01-21 22:41:29

Re: Mars Colonial Transporter

#32 2017-01-22 05:33:20

Re: Mars Colonial Transporter

#33 2017-01-22 08:03:26

Re: Mars Colonial Transporter

#34 2017-01-22 08:22:20

Re: Mars Colonial Transporter

#35 2017-01-22 13:02:57

Re: Mars Colonial Transporter

#36 2017-01-22 13:08:36

Re: Mars Colonial Transporter

#37 2017-01-22 21:43:41

Re: Mars Colonial Transporter

#38 2017-01-23 00:25:08

Re: Mars Colonial Transporter

#39 2017-01-23 09:50:50

Re: Mars Colonial Transporter

#40 2017-01-23 10:28:46

Re: Mars Colonial Transporter

#41 2017-01-23 11:34:43

Re: Mars Colonial Transporter

#42 2017-01-23 11:38:12

Re: Mars Colonial Transporter

#43 2017-01-23 11:50:48

Re: Mars Colonial Transporter

#44 2017-01-23 12:18:40

Re: Mars Colonial Transporter

#45 2017-01-23 13:02:13

Re: Mars Colonial Transporter

#46 2017-01-23 16:50:52

Re: Mars Colonial Transporter

#47 2017-01-23 18:53:21

Re: Mars Colonial Transporter

#48 2017-01-23 22:22:15

Re: Mars Colonial Transporter

#49 2017-02-05 09:22:41

Re: Mars Colonial Transporter

#50 2017-02-05 09:29:44

Re: Mars Colonial Transporter

Board footer