Quote from: philw1776 on 08/24/2015 01:07 pmMCT Mass 180 mTS2 Mass w/MCT 1025 mTS2 Mars 25mT Cargo 8.5 Km/sec Rocket EquationI don't think those three go together.A stage with 180mt dry mass, 8.5km/s delta v and 25mt payload has a wet mass of 1980mt.
MCT Mass 180 mTS2 Mass w/MCT 1025 mTS2 Mars 25mT Cargo 8.5 Km/sec Rocket Equation
So we're saying that the oft stated 200mT MCT landing on Mars is impossible?
I think we need to consider entry velocity when looking at any lander, it's the #1 driver of TPS, structural mass and retro-propulsion needs. Most lander heritage involves direct entry from interplanetary transit and very high entry velocity. The Viking lander was the exception and is probably the best comparison because it doesn't drop nearly as many parts along the way. It had a landed mass of 62% of entry mass compared with 10% for the Phoinex lander.
The lander would launch from Earth atop a 2 stage launch vehicle and would be loaded with ~50 mt of propellants providing ~900 m/s DeltaV to be used for emergency separation and propulsive landing in the even of an abort. This would put the total launch mass at 225 mt. And the lander could then be fully topped off via another launch of and transfer of 250 mt or propellant from a stretched 2nd stage tanker. Transit to Mars is then done via a hybrid propulsion system, first a large SEP vehicle would slow push the lander to high Earth orbit where crew would board via a dragon capsule.
Isn't PICA supposed to be significantly better than previous TPS materials though? Would that reduce the required dry mass?
The human lander concepts I referred to are all designed for either entry from Mars orbit or aerocapture into orbit + entry from Mars orbit. Granted, gs, heat peak rate and heat load are significantly higher in the aerocapture phase than in the entry phase, although its a lot better than direct entry."Parts dropped along the way" are part of the structural mass at entry, which I was refering to. I.e. the entry mass minus the fuel and payload.The best lander in terms of payload to structural entry mass I cound find is in the pdf attached on page 15, top left. A value of ~1.2. With SIAD, from orbit. It has a total entry mass of only 20t though. The higher the entry mass the worse usually.With HIAD it may get better, haven't seen the mass break down of such a lander yet. HIAD could be interesting for MCT.
Wait, you want to use SEP for transfering a 475t payload from LEO to HEO (e.g. LDHEO)? That would take about 8 years with a 300kw SEP.
Quote from: Vultur on 08/25/2015 01:26 amIsn't PICA supposed to be significantly better than previous TPS materials though? Would that reduce the required dry mass?PICA is the best ablative, but I intend for the lander to do repeated rapid cargo flights between mars surface and low orbit maybe as many as 100 such round trips before returning to Earth, no ablative can go through that many cycles so I reject it as an option.
As the Mars colony is built up the small SEP transit vehicle would be replaced with a large more powerful cargo hauler with extensive habitats aka the mother-ship. The lander would be fueled in LEO and used to rapidly deliver passengers (100 at a time) to the waiting mother-ship in high orbit and would then travel attached too it as the mother-ships speed is expected to be competitive with the earlier direct flight. At mars the lander is used to repeatedly ferry between surface and low orbit with containerized cargo being reloaded in orbit. At Earth the launch of chemical propellant and landers is almost eliminated in favor of launching naked cargo containers and SEP propellant to be loaded onto the mother-ship.
PICA is the best ablative, but I intend for the lander to do repeated rapid cargo flights between mars surface and low orbit maybe as many as 100 such round trips before returning to Earth
PICA is the best ablative, but I intend for the lander to do repeated rapid cargo flights between mars surface and low orbit maybe as many as 100 such round trips before returning to Earth, no ablative can go through that many cycles so I reject it as an option.
As opposed, I believe 50t could be the dry mass of a 820t fully loaded and fueled MCT, note that at most there would be 1.3km/s of ΔV or so a re-entry mass around 230t, and a landing mass approaching 150t.
Quote from: Impaler on 08/25/2015 04:00 amPICA is the best ablative, but I intend for the lander to do repeated rapid cargo flights between mars surface and low orbit maybe as many as 100 such round trips before returning to EarthAh, I was assuming one landing, one launch direct to Earth. MCT is supposed to be "land the whole thing".So only one Earth launch, one Mars entry, one Mars launch, and one Earth entry between servicings (on Earth).I think PICA-X will be used for MCT.
It is supposed to get the mission done as efficiently as possible, I'm making the case that this hybrid architecture is 1) Actually achievable without using pixie-dust and 2) more efficient to boot. Compared to that any 'supposed to be' is moot.
So I propose a vehicle designed to act as a ferry between the surface and orbit, able to land 100 mt and return to orbit with 25 mt PLUS enough propellant to make another landing on Mars. At 75 mt dry mass for the vehicle the landing requires 40 mt of propellants. The launch mass is 400 mt of which 260 is assent (4.1 km/s) propellant, 25 cargo, 40 return propellant (.8 km/s) and 75 the vehicle dry mass. Total propellant load is just 300 mt meaning propellant production on Mars can be a fraction of that needed for Direct return. The structural mass fraction at launch is a very conservative 18.75% and payload is 46% of entry mass (including propellant) both very reasonable figures.The lander would launch from Earth atop a 2 stage launch vehicle and would be loaded with ~50 mt of propellants providing ~900 m/s DeltaV to be used for emergency separation and propulsive landing in the even of an abort. This would put the total launch mass at 225 mt. And the lander could then be fully topped off via another launch of and transfer of 250 mt or propellant from a stretched 2nd stage tanker. Transit to Mars is then done via a hybrid propulsion system, first a large SEP vehicle would slow push the lander to high Earth orbit where crew would board via a dragon capsule. Then the SEP would separate and the lander would make a lunar-Earth slingshot burn of ~1 km/s for fast transit to Mars and use propulsive capture and perhaps some airobraking to reach low mars orbit leaving just enough propellant for landing. On the surface the cargo is unloaded and a 25 mt return cabin is loaded in it's place. The SEP system would make a slow transit to mars arriving well after the lander and wait in low orbit for rendezvous after which it would bring the lander back to high Earth orbit, crew would disembark again via Dragon capsule and then the unoccupied combined vehicle would do a down-spiral to low Earth orbit where it would be refueled again and use it's propulsive capacity to reduce entry velocity to the 3.6 km/s velocity it can tolerate followed by landing on Earth for reloading and reintegration, SEP remains in orbit ready to be reused.Total BFR launches would be 6, one for the lander, 2 tankers of chemical propellant, 1 for SEP hardware, 2 estimated for SEP propellant. In addition 2 trans-lunar Dragon capsule launches likely on Falcon Heavy. Each subsequent launch would need 5 launches when SEP is reused. The cycle would be one round trip every 2 synods. If philw1776 can run the numbers of the size of the BFR needed to do a 225 mt launch that would be most informative to see how it compares.
Quote from: Impaler on 08/25/2015 06:36 amIt is supposed to get the mission done as efficiently as possible, I'm making the case that this hybrid architecture is 1) Actually achievable without using pixie-dust and 2) more efficient to boot. Compared to that any 'supposed to be' is moot.To be clear on this. You are arguing that you know better than SpaceX and Elon Musk?
Perhaps you should write it up as a paper.
Quote from: guckyfan on 08/25/2015 06:49 amTo be clear on this. You are arguing that you know better than SpaceX and Elon Musk?Since the amount of information known about MCT outside of SpaceX can fit on a postage stamp, written with a crayon, and that SpaceX plans change considerable over time, how do we know this isn't the plan? No point in saying "MUSK SAID THIS", because aforesaid plans change.
To be clear on this. You are arguing that you know better than SpaceX and Elon Musk?