Yet another Mars architecture

RobS · 2015-06-29 15:25:56

I agree with you, Robert; this is all pork barrel. But it is sad that because of that pork barrel, we'd have to spend a billion to develop SEP and then spend billions extracting xenon from the atmosphere. I wouldn't be surprised if, by the time we can go to Mars, Falcon Heavy is flying regularly and reliably, it will "dejustify" SLS, and the outcry against the wasteful cost of SLS will force its cancellation. But right now, by all means, plan to use it, because it will be built!

RobertDyck · 2015-06-29 16:08:01

kbd512 wrote:

I don't claim that a repurposed ISS module isn't "enough space" for the astronauts. It's entirely dependent upon the attitude of our explorers and whatever goes on upstairs in the designers heads. Some at NASA don't think ISS modules provide enough habitation space for the astronauts.

Destiny science module is 8.53 metres (28.0 ft) long including end cones, and 4.27 metres (14.0 ft) diameter. Pressurized volume 106 cubic metres (3,700 cu ft). Without end cones it's 25.2 feet (7.68m) long. Each payload rack is 2m high x 1.01m wide (80" x 40"). Open space in the centre is faced by racks on both walls, ceiling, and floor. So the "floor" is 2m x 7.68m, or 15.36 square metres. Mars Direct Hab has 8.4m outside diameter, let's assume 8.0 metre inside diameter. That's 0.2 metre (7.87") thick walls. That provides 50.265 square metres of floor area. Just the upper floor alone is 3.27 times the floor area of Destiny. Someone want to quibble about size of science rack pivots, thickness of Mars Direct walls, or volume of equipment racks?

kbd512 · 2015-06-29 16:29:09

RobertDyck wrote:

kbd512 wrote:
Dragon V2, like Orion, isn't going to land anywhere but Earth in its present configuration.
I said Dragon. Gwynn Shotwell talked about using Dragon V2 to transport crew to ISS. I would prefer the current Dragon for Mars. Dragon Rider was a design that never flew; a full-size mockup was built, but no flight vehicle. It was the current Dragon (Cargo Resupply Ship), but with launch abort rockets on the side, an APAS hatch instead of CBM hatch, seat / controls / life support. I don't think any form of LAS is needed for a life boat. Thrusters built into Dragon CRS are enough for my proposal. Of course if you deliver crew to the Mars vehicle from Earth with the same Dragon that you use as life boat, then you need LAS. Just for launch from Earth. But that means I'm not arguing for Dragon V2. If SpaceX says the only version of Dragon is V2, and it's the same mass as Dragon Rider, then fine. Use that. But SpaceX is adding complexity, not me.

If your MTV proposal incorporates a lifeboat, then that lifeboat has to be rated for long duration space flight. Dragon was not designed for that. Would it be easy to redesign, if necessary, and certify it for that type of use? It should be, but nothing is ever that simple.

ISS modules are rated for long duration space flight. If a spare is required, then send two modules. The redundancy of two or more complete MTV's flown in formation would be the best solution, even if it costs more.

1 Node Module + 4 BEAM + SEP + chemical kick stages for TMI/TEI and 2 crew per MTV

1 Lab Module + 1 Node Module + SEP + chemical kick stages for TMI/TEI and 4 crew per MTV

Either of the two solutions above could be flown in formation with an active radiation shield module and multi-person MDV/MAV.

1 Skylab II Module (integrated SEP, active radiation shield, spares carried for repairs; no backup MTV) and 4 to 6 crew

Multi-person MDV/MAV prepositioned in LMO using a SEP tug.

If we do develop SEP tugs, then we can assemble/service/test the MTV's at ISS, preposition the MTV's at L1, use relatively small chemical kick stages for TMI/TEI, and use Dragon V2 for transits to/from L1.

The LEM lifeboat solution worked for the Apollo 13 crew because it was only a few days to/from the moon. That won't work for a Mars mission.

kbd512 · 2015-06-29 16:33:56

RobS wrote:

I agree with you, Robert; this is all pork barrel. But it is sad that because of that pork barrel, we'd have to spend a billion to develop SEP and then spend billions extracting xenon from the atmosphere. I wouldn't be surprised if, by the time we can go to Mars, Falcon Heavy is flying regularly and reliably, it will "dejustify" SLS, and the outcry against the wasteful cost of SLS will force its cancellation. But right now, by all means, plan to use it, because it will be built!

I'd rather NASA not throw away all the development funding expended on SLS, wasteful as it was/is, but that's just me.

kbd512 · 2015-06-29 16:37:20

RobertDyck wrote:

kbd512 wrote:
I don't claim that a repurposed ISS module isn't "enough space" for the astronauts. It's entirely dependent upon the attitude of our explorers and whatever goes on upstairs in the designers heads. Some at NASA don't think ISS modules provide enough habitation space for the astronauts.
Destiny science module is 8.53 metres (28.0 ft) long including end cones, and 4.27 metres (14.0 ft) diameter. Pressurized volume 106 cubic metres (3,700 cu ft). Without end cones it's 25.2 feet (7.68m) long. Each payload rack is 2m high x 1.01m wide (80" x 40"). Open space in the centre is faced by racks on both walls, ceiling, and floor. So the "floor" is 2m x 7.68m, or 15.36 square metres. Mars Direct Hab has 8.4m outside diameter, let's assume 8.0 metre inside diameter. That's 0.2 metre (7.87") thick walls. That provides 50.265 square metres of floor area. Just the upper floor alone is 3.27 times the floor area of Destiny. Someone want to quibble about size of science rack pivots, thickness of Mars Direct walls, or volume of equipment racks?

That sounds great. Unfortunately, Mars Direct was and is affected by NIH syndrome, so that's not what NASA will develop. ISS hardware already exists. It doesn't require development. It's flying in space as I write this.

RobertDyck · 2015-06-29 21:46:21

kbd512 wrote:

If your MTV proposal incorporates a lifeboat, then that lifeboat has to be rated for long duration space flight. ... The LEM lifeboat solution worked for the Apollo 13 crew because it was only a few days to/from the moon. That won't work for a Mars mission.

That's not the purpose. My architecture carries a Dragon on for aerocapture. If aerocapture fails, then crew and Mars samples abort to Earth surface. That's it, the only purpose. If Dream Chaser is flying by then, Dragon can stay attached until the next mission. If not, then use that Dragon for return to Earth from ISS. Don't obsess about some other architecture.

The Apollo LEM (later called LM), was used for its life support. The contingency in my plan is to keep the Mars lander attached during transit from Earth to Mars. That way if a free return is necessary (like Apollo 13), then life support equipment, food, and supplies of the lander will still be attached and available.

kbd512 wrote:

ISS modules are rated for long duration space flight. If a spare is required, then send two modules. The redundancy of two or more complete MTV's flown in formation would be the best solution, even if it costs more.

Quote from the movie "Contact": Why build just one when you can build two for twice the price? Typical government thinking. Are you a bureaucrat?

Your mission plan includes several modules, and Skylab II that is the same diameter as Mars Direct, but several times the height. Much bigger. Bigger = more expensive. And RobS is challenging me for suggesting use of SLS at all.

kbd512 wrote:

That sounds great. Unfortunately, Mars Direct was and is affected by NIH syndrome, so that's not what NASA will develop. ISS hardware already exists. It doesn't require development. It's flying in space as I write this.

Yea. When Robert Zubrin got attention from NASA, they redesigned it. Mars Direct was estimated at $20 billion for the first mission, then $2 billion for each mission thereafter. In 1989 or 1990 dollars. NASA DRM (Semi-Direct) was estimated at $55 billion for 7 missions. Congress looked at the price increase with no hardware built. Already massive price creep, and just studies. So it was denied. Has anyone learned? The lessons are supposed to be: accept good ideas from wherever they come, and keep the cost down. Has NASA learned either?

kbd512 · 2015-06-29 23:27:08

RobertDyck wrote:

That's not the purpose. My architecture carries a Dragon on for aerocapture. If aerocapture fails, then crew and Mars samples abort to Earth surface. That's it, the only purpose. If Dream Chaser is flying by then, Dragon can stay attached until the next mission. If not, then use that Dragon for return to Earth from ISS. Don't obsess about some other architecture.

The only benefit to aerocapture is reduced propellant consumption. Impulsive transfers combined with spiraling in to LMO and L1 avoids the complexity of aerocapture altogether. If a reusable Falcon Heavy comes to fruition, making an all-chemical architecture affordable, why bother with the engineering and operational issues associated with aerocapture when propulsive captures are a realistic option? I know that NASA has the technical capability to successfully employ aerocapture, but such maneuvers are an avoidable mission risk.

RobertDyck wrote:

The Apollo LEM (later called LM), was used for its life support. The contingency in my plan is to keep the Mars lander attached during transit from Earth to Mars. That way if a free return is necessary (like Apollo 13), then life support equipment, food, and supplies of the lander will still be attached and available.

Shipping multi-component mission hardware sets to Mars as a single vehicle would require propulsion modules that even SLS can't orbit in its current form or using multiple propulsion modules. If each major mission component (MTV, MDV/MAV, surface habitat, kick stages) only required one launch for the component and one launch for the propulsion module, mission risks associated with launch coordination and multi-component assembly would be reduced. What major issue does staging the MDV/MAV in LMO create? Why is it necessary to assemble Dragon + MTV + MDV/MAV + propulsion? What's wrong with MDV/MAV + propulsion and MTV + propulsion?

RobertDyck wrote:

Quote from the movie "Contact": Why build just one when you can build two for twice the price? Typical government thinking. Are you a bureaucrat?

If a mission requires true redundancy, the only way you achieve it is to have two or more of everything you need or enough spare parts to rebuild any mission critical component.

RobertDyck wrote:

Your mission plan includes several modules, and Skylab II that is the same diameter as Mars Direct, but several times the height. Much bigger. Bigger = more expensive. And RobS is challenging me for suggesting use of SLS at all.

I've proffered several ideas for potential mission hardware sets that are flexible and feasible. Good, better, and best are a matter of opinion. I'm not smitten with any particular idea or architecture. I just want NASA to decide on an architecture that's flexible and feasible and then work with it, perfect or not.

RobertDyck wrote:

Yea. When Robert Zubrin got attention from NASA, they redesigned it. Mars Direct was estimated at $20 billion for the first mission, then $2 billion for each mission thereafter. In 1989 or 1990 dollars. NASA DRM (Semi-Direct) was estimated at $55 billion for 7 missions. Congress looked at the price increase with no hardware built. Already massive price creep, and just studies. So it was denied. Has anyone learned? The lessons are supposed to be: accept good ideas from wherever they come, and keep the cost down. Has NASA learned either?

The two most costly aspects of both plans are more capable launch vehicles and a multi-person MDV/MAV. We're getting SLS, economy be damned. Remaining funding should have been directed towards MDV/MAV development, but was directed towards Orion development.

RobertDyck · 2015-06-30 06:16:04

You mentioned L1 again. You realize there's no advantage of any sort what so ever to putting any human mission at any Lunar Lagrange point. Total fuel to get there, then to Mars, is greater. Earth-Moon L1 is directly between the Earth and Moon, and getting there takes as much fuel as a Trans-Lunar trajectory. Lowest propellant trajectory is direct launch to Mars. Actually, lowest propellant is direct launch from Earth surface to Mars surface. If you have to do orbital assembly, do it in LEO. Assembly in LEO then launch to Mars is *MUCH* lower propellant than assembly in Earth-Moon L1 or L2.

You complain that aerocapture is complicated, but you want to use electric propulsion, that either requires gigantic solar arrays or a nuclear reactor. As soon as you say the word "nuclear", the cost skyrockets. And you get all the anti-nuclear activists. Aerocapture is the simpler technology.

Which electric propulsion are you looking at? Ion, TAL Hall (Russian equivalent to ion), MPD from Glenn Research Centre, or VASIMR?

kbd512 wrote:

Shipping multi-component mission hardware sets to Mars as a single vehicle would require propulsion modules that even SLS can't orbit in its current form or using multiple propulsion modules.

If we use your mission architecture it would. Your Skylab II would have greater mass than the entire stack of mine. It has the same diameter as a Mars Direct hab, but 11.15m (36.1ft) high. It's 4 times as high as a single story Mars Direct hab. And that doesn't include the Skylab II truss structure or arm. Plus you want ISS nodes and BEAM modules. How massive is your Battlestar?

RobS · 2015-06-30 06:35:11

There's nothing complicated about aerocapture at all. You don't need to know the precise density of the Martian atmosphere. It doesn't vary that much and you only need to get your velocity below Mars's escape velocity on the first pass. You can make additional passes to tweak your orbit, and that's well understood; that's what NASA does with all its Martian spacecraft now, using their solar panels as sources of wind resistance. Until someone actually uses aerocapture, it will remain speculative, but it shouldn't be any more difficult that supersonic retropropulsion that Space X has demonstrated. That hadn't been developed much, either.

kbd512 · 2015-06-30 07:40:52

RobertDyck wrote:

You mentioned L1 again. You realize there's no advantage of any sort what so ever to putting any human mission at any Lunar Lagrange point. Total fuel to get there, then to Mars, is greater. Earth-Moon L1 is directly between the Earth and Moon, and getting there takes as much fuel as a Trans-Lunar trajectory. Lowest propellant trajectory is direct launch to Mars. Actually, lowest propellant is direct launch from Earth surface to Mars surface. If you have to do orbital assembly, do it in LEO. Assembly in LEO then launch to Mars is *MUCH* lower propellant than assembly in Earth-Moon L1 or L2.

Stop insinuating that I want to do assembly at L1/L2. I want to assemble the MTV at ISS and have a SEP tug move the stack to L1. Nobody will be aboard, so there's no rush. I want to use SEP on the MTV to spiral in to LMO and to spiral in to L1 on the return home. I indicated that chemical kick stages should be used for the orbital transfers.

RobertDyck wrote:

You complain that aerocapture is complicated, but you want to use electric propulsion, that either requires gigantic solar arrays or a nuclear reactor. As soon as you say the word "nuclear", the cost skyrockets. And you get all the anti-nuclear activists. Aerocapture is the simpler technology.

It is more complicated than SEP because it's never been used, to my knowledge, in the manner in which you want to use it. It's also an unnecessary mission risk, which is my primary objection to using aerocapture.

RobertDyck wrote:

Which electric propulsion are you looking at? Ion, TAL Hall (Russian equivalent to ion), MPD from Glenn Research Centre, or VASIMR?

Regular hall thrusters. It's tech that already exists, has been used for decades, and is in active development.

RobertDyck wrote:

If we use your mission architecture it would. Your Skylab II would have greater mass than the entire stack of mine. It has the same diameter as a Mars Direct hab, but 11.15m (36.1ft) high. It's 4 times as high as a single story Mars Direct hab. And that doesn't include the Skylab II truss structure or arm. Plus you want ISS nodes and BEAM modules. How massive is your Battlestar?

You're fixated on Skylab II. It was a proposal that would work if we have SLS. You want to include a multi-person MDV/MAV in the hardware stack that you're shipping to Mars. How much will the MDV/MAV weigh fully fueled? NASA will not send a MDV/MAV to Mars with empty tanks for the first mission.

One MPLM + 4 BEAM would not weigh more than any fully fueled MDV/MAV and neither would Skylab II.

RobS · 2015-06-30 08:39:42

L1 has been proposed as a departure point in some Mars mission scenarios. It takes 300 meters per second to get into L1, after making the transfer burn (which is basically the minimum energy to get to the moon, which I think is 3.3 km/sec). Subsequently, to make your Trans-Mars injection burn, you need to return to low Earth orbit and do the burn there, to get the maximum advantage. That means altogether a stop in L1 adds 600 meters per second to your trip to Mars. It may be that you can get to the moon in a much smaller delta-v from L1 and use the moon to swing your spacecraft back to the Earth, so perhaps 400 m/second or 450 m/second will work instead. It's a delta-v penalty either way.

There are two reasons to use L1. One could use a high Earth orbit or a highly elliptical lunar orbit for the same purpose, by the way:

1. Use SEP to spiral your spacecraft most of the way out of the Earth's gravitational field. L1 is a convenient place to park while you wait for crew or assemble (dock) different modules together. If you use SEP to get the TMI kick stage to L1, you get it there with relatively little propellant.

2. Use propellant manufactured from lunar water for TMI. At the moment it is not clear this will ever be economic.

Terraformer · 2015-06-30 09:27:45

Huh? It doesn't add a delta-V penalty once you include the much reduced burn required for a craft coming in from L1 at near escape velocity, versus the one for a craft in low earth orbit...

RobS · 2015-06-30 10:11:37

The only delta-v penalty is the 300 meters/second it takes to "stop" at L1 and then to head back to LEO later. You can't just fire your vehicle from LEO to L1; it won't stay there without a circularization burn of 300 m/sec. To return to LEO you need to "decircularize" so that you are on a trajectory back to LEO and that takes 300 m/sec, unless you can use less than that to go to the moon and make a small burn when you pass the moon to head back to Earth.

RobertDyck · 2015-06-30 10:36:03

For my architecture, instead of a single Dragon spacecraft, you could use a pair of Soyuz descent modules. You wouldn't need the service or orbital modules, just the descent capsule. But Soyuz carries 3 cosmonauts, or one seat can be replaced with cargo. So 4 astronauts plus Mars samples would require 2 capsules. They mass 3 metric tonnes each, total 6 tonnes. Dragon with full propellant tanks masses 8 metric tonnes. Only saves 2 metric tonnes. And with the current political environment, I don't see a Mars mission using any Russian hardware.

Last edited by RobertDyck (2015-06-30 15:00:24)

RobertDyck · 2015-06-30 10:43:48

If you use electric propulsion, the best method is use chemical propulsion to raise your orbit just barely out of the atmosphere. Then raise the apogee while keeping the perigee where it is. Keep raising the apogee until you depart. Be aware that LEO is not completely out of the atmosphere, there is a little atmospheric drag. It's that wisp of atmosphere that keeps out Van Allen belts. But you need an electric propulsion engine that provides significantly more thrust than the drag in LEO. Ideal is to raise perigee above LEO, but then there's the trade-off of propellant to fight drag in LEO vs propellant to raise perigee. Bottom line: don't stop at L1. There's no point.

When using electric propulsion for Mars, you can also use the Moon for gravity assist. The Moon is small, it only provides 0.5 km/s, but every little bit helps.

Last edited by RobertDyck (2015-06-30 10:47:17)

RobS · 2015-06-30 10:56:46

Generally, electric propulsion needs to apply a low thrust over a long period of time, so you can't just raise the apogee; you have to raise your entire orbit. As a result, the delta-v to go to the moon with ion propulsion is about double it is for chemical propulsion. That would be true for a trip to Mars and a spiral down to low Mars orbit as well. That's one way that solar-thermal propulsion is superior. You use a mirror to heat up a carbon block massing hundreds of kilograms, then you run your hydrogen through it over a short period of time when you are at perigee. The result is a specific impulse of 800 or maybe more seconds; close to the Isp nuclear thermal. Thrust is limited to 50 to 100 pounds/250-500 newtons. That's a lot more than electric propulsion but a lot less than chemical. It takes weeks or months to escape earth; you'd want to use a chemical kick stage at the end to complete trans-Mars injection, during your last perigee. Zubrin notes its 80% or so as good as nuclear thermal. It'd be much less controversial to use and cheaper to develop, but no one has done it yet.

RobertDyck · 2015-06-30 11:00:10

kbd512 wrote:

RobertDyck wrote:
Which electric propulsion are you looking at? Ion, TAL Hall (Russian equivalent to ion), MPD from Glenn Research Centre, or VASIMR?
Regular hall thrusters. It's tech that already exists, has been used for decades, and is in active development.

NASA's work on Hall thrusters only produced 1800s Isp. Russians developed something called Thruster Anode Layer (TAL) Hall thrusters. In 2001, NASA hired engineers from Russia to teach engineers at Glenn Research Centre everything they knew. NASA discovered Russia had developed Hall thrusters much farther. NASA launched Deep Space One in 1998 to demonstrate an ion engine. The hope is it would be used after the technology demonstrator. But Russia had been using Hall thrusters for station keeping of military spy satellites since the late 1960s. Russian TAL Hall thrusters had as much Isp as the best ion engines that the guys at Glenn could make. And Russia has proposals on paper for high power Hall thrusters that would have even higher Isp. Actually, that's why the guys at Glenn were able to develop MPD so well. They used everything they had developed for ion engines, plus everything they learned from Russians about Hall thrusters, plus everything Princeton University had done with MPD. Combining all that together they produced an MPD engine that uses liquid hydrogen propellant and 8400s Isp. That happened to equal the best theoretical specific impulse of Russia's proposed, big, high power Hall thrusters. But the guys at Glenn actually built a prototype of their MPD engine, and tested it in the lab. It worked.

So why are you so fixated on Hall? Are you Russian?

kbd512 · 2015-06-30 18:19:27

RobertDyck wrote:

kbd512 wrote:
RobertDyck wrote:
Which electric propulsion are you looking at? Ion, TAL Hall (Russian equivalent to ion), MPD from Glenn Research Centre, or VASIMR?
Regular hall thrusters. It's tech that already exists, has been used for decades, and is in active development.
NASA's work on Hall thrusters only produced 1800s Isp. Russians developed something called Thruster Anode Layer (TAL) Hall thrusters. In 2001, NASA hired engineers from Russia to teach engineers at Glenn Research Centre everything they knew. NASA discovered Russia had developed Hall thrusters much farther. NASA launched Deep Space One in 1998 to demonstrate an ion engine. The hope is it would be used after the technology demonstrator. But Russia had been using Hall thrusters for station keeping of military spy satellites since the late 1960s. Russian TAL Hall thrusters had as much Isp as the best ion engines that the guys at Glenn could make. And Russia has proposals on paper for high power Hall thrusters that would have even higher Isp. Actually, that's why the guys at Glenn were able to develop MPD so well. They used everything they had developed for ion engines, plus everything they learned from Russians about Hall thrusters, plus everything Princeton University had done with MPD. Combining all that together they produced an MPD engine that uses liquid hydrogen propellant and 8400s Isp. That happened to equal the best theoretical specific impulse of Russia's proposed, big, high power Hall thrusters. But the guys at Glenn actually built a prototype of their MPD engine, and tested it in the lab. It worked.
So why are you so fixated on Hall? Are you Russian?

Rob,

If the MTV uses SEP to capture at LMO and L1, then the power and propellant mass required for propulsive capture are reduced. I want to use existing propulsion options for the mission, meaning storable chemical propellant kick stages and hall thrusters, to reduce development costs and risks. I'd prefer funding for CL-ECLSS and active radiation shielding over bleeding edge propulsion technology. By eliminating unnecessary risks, we're eliminating excuses for not doing the mission.

GW Johnson · 2015-07-03 08:53:32

NASA will point at any technological excuse not to go to Mars, some of which technologies are what we have been discussing here. Those are just excuses, we could have gone as long ago as 1995 with good prospects for getting the crew home alive.

The real reasons have to do with (1) excessive risk-aversity on NASA's part (3 dead crews and some near-calls, including problems with risk of drowning in water-cooled overly-complicated spacesuits), and (2) Congress mandating that all the money goes to favored contractors located in their districts, while completely ignoring whether those activities have anything to do with going to Mars. Until those two things change, NASA will not go to Mars. And we are stuck arguing angels-on-the-head-of-a-pin on these forums.

Unless somebody else besides NASA goes first.

GW

Last edited by GW Johnson (2015-07-03 08:54:05)

kbd512 · 2015-07-03 19:22:49

GW Johnson wrote:

NASA will point at any technological excuse not to go to Mars, some of which technologies are what we have been discussing here. Those are just excuses, we could have gone as long ago as 1995 with good prospects for getting the crew home alive.

I think if a reliable CL-ECLSS, active radiation shielding, and artificial gravity are developed into reliable mission hardware, there are no excuses left for not going.

GW Johnson wrote:

The real reasons have to do with (1) excessive risk-aversity on NASA's part (3 dead crews and some near-calls, including problems with risk of drowning in water-cooled overly-complicated spacesuits), and (2) Congress mandating that all the money goes to favored contractors located in their districts, while completely ignoring whether those activities have anything to do with going to Mars. Until those two things change, NASA will not go to Mars. And we are stuck arguing angels-on-the-head-of-a-pin on these forums.
Unless somebody else besides NASA goes first.
GW

NASA isn't risk averse, its management is common sense deficient. None of shuttle crews would've died had NASA management exercised a little common sense. When $2B worth of flight hardware is on the line, suspend all disbelief and have faith that bad things will happen if you ignore obvious problems.

The space suits currently in use are simply not well designed. NASA has had decades to replace them and thus far has not. The MCP suits originally developed in the 1960's were the design that NASA would have selected if anyone in management with a modicum of common sense was making technology decisions based on how well a technology works. It doesn't matter what the astronaut office "thinks" about a given space suit design, it only matters how well it protects the astronauts from space.

As far as a manned Mars mission is concerned, NASA could have all the development funding required for the task if SLS and Orion are defunded. I don't blame NASA for funding them, but I blame NASA for approving impractical designs and continuing to develop designs that don't support current mission hardware requirements.

SLS could have been a comparatively inexpensive development proposition if NASA elected to use existing hardware rather than completely redesigning every critical component in the rocket. I think most, if not all, of the upgrades have a lot merit, but not for the amount of money expended.

RobertDyck · 2015-07-04 01:14:57

kbd512 wrote:

If the MTV uses SEP to capture at LMO and L1, then the power and propellant mass required for propulsive capture are reduced. I want to use existing propulsion options for the mission, meaning storable chemical propellant kick stages and hall thrusters, to reduce development costs and risks. I'd prefer funding for CL-ECLSS and active radiation shielding over bleeding edge propulsion technology. By eliminating unnecessary risks, we're eliminating excuses for not doing the mission.

You keep failing to take into account all propellant. It doesn't matter how much you save by only calculating part way there. Astronauts go from Earth to Mars, and back. You have to include all propellant, including Earth to L1, or whatever way point you use. Your solution does NOT save propellant.

And I said before, but obviously have to repeat: SEP means this...

Notice 700 metre wide spacecraft, but only a single ISS module for accommodation. The ITV I described is 3 times the size. And you are the one calling for bleeding edge propulsion. You are introducing unnecessary risk. I talked about chemical propulsion: J-2X or RL10.

GW Johnson · 2015-07-04 09:36:39

I have no problems with EP, whether it is solar or nuke-powered. That kind of stuff is now working routinely at small scale for satellite station-keeping, and for a couple of the deep space probes. The trouble arises when you try to use it as the only main propulsion for a manned vehicle.

It does not scale up very well, because we haven't yet developed EP engines that big, and the power supplies get impossibly large and heavy. This is partly driven by the need to fly faster with men: longer missions have heavier life support and invite more problems statistically. You can spiral-out unmanned from Earth and send the men fast out to meet it, of course, but now you have two vehicles. There ain't no such thing as a free lunch.

Your best bet, as I have argued before, is chemical propulsion for getaway and capture maneuvers (I think we have not enough experience yet with aerocapture to bet lives on it remote from LEO where rescue is possible). Use EP, (solar or nuke as desired) to speed up the long transits. That kind of EP need not be such a large engine, so the scale-up developments are not so demanding, and the system weights more tolerable.

Trade added EP weight against smaller life support weights for the shorter trip. It might be favorable. Who knows? We have yet to fly something like that.

BTW, I get irritated when I see spacecraft designs based on large space trusses. That's just inert weight buildup, which is to be avoided at all costs. That structural role should be fulfiulled by something you have to have along with you anyway. There's a role for truss structures, to be sure, but not as primary spacecraft structure, not as a major component. Not at this time in history. Not with rocket propulsion.

GW

kbd512 · 2015-07-04 21:58:59

RobertDyck wrote:

kbd512 wrote:
If the MTV uses SEP to capture at LMO and L1, then the power and propellant mass required for propulsive capture are reduced. I want to use existing propulsion options for the mission, meaning storable chemical propellant kick stages and hall thrusters, to reduce development costs and risks. I'd prefer funding for CL-ECLSS and active radiation shielding over bleeding edge propulsion technology. By eliminating unnecessary risks, we're eliminating excuses for not doing the mission.
You keep failing to take into account all propellant. It doesn't matter how much you save by only calculating part way there. Astronauts go from Earth to Mars, and back. You have to include all propellant, including Earth to L1, or whatever way point you use. Your solution does NOT save propellant.
And I said before, but obviously have to repeat: SEP means this…
Notice 700 metre wide spacecraft, but only a single ISS module for accommodation. The ITV I described is 3 times the size. And you are the one calling for bleeding edge propulsion. You are introducing unnecessary risk. I talked about chemical propulsion: J-2X or RL10.

Rob,

The manned Mars mission hardware that the Russians proposed uses SEP propulsion, exclusively, and mounts the MDV/MAV to the MTV. Naturally, it's totally unrealistic, even if a 200t capable launch vehicle was available. That is most definitely not what I proposed.

Let's review what I proposed, in detail, and then go from there:

1. Use Falcon Heavy for launch services if there is not a law that forces NASA to use SLS.

* If SLS is required by law, then develop Skylab II into a MTV and incorporate an artificial gravity module into the MTV design

2. Develop a 100kW class SEP tug to move the unmanned MTV from ISS to L1 and back.

* The same hardware developed for the SEP tug will be incorporated into the MTV design to reduce development cost and increase mass delivered to Mars.

* The SEP hardware incorporated into the MTV has two tasks to perform:

- Insert the MTV into LMO

- Insert the MTV into L1

* The SEP tug has four tasks to perform (to be performed by multiple tugs, if that isn't completely obvious):

- Move the assembled MTV (habitat + mini-mag + kick stage) from ISS to L1 and back to ISS for refurbishment between missions

- Move the chemical kick stage to Mars for TEI

- Move the M113/MTVL mobile surface exploration rovers to Mars

- Move the MDV/MAV to Mars

3. Develop a chemical kick stage that uses storable propellants for TMI and TEI. The TEI stage will be transferred to Mars using a SEP tug.

* The TMI/TEI kick stages use storable chemical propellants because this technology already works and has worked for decades.

* The chemical kick stages have two, maybe three, tasks to perform:

- TMI for the MTV from L1

- TEI for the MTV from LMO

- Optionally, TMI for the MDV/MAV from L1

4. Perform all MTV assembly and refurbishment at ISS.

* Each MTV would require a minimum of 3 Falcon Heavy flights:

- Flight #1: SEP tug

- Flight #2: habitat and mini-mag

- Flight #3: chemical kick stage

5. Develop CL-ECLSS to improve consumables mass requirements

* NASA intends to use their improved ECLSS aboard ISS and there's no reason not to incorporate it into the MTV

6. Develop active radiation shielding to help limit the crew's radiation exposure

* Apart from SPE's radiation exposure is a non-issue, but NASA made it an issue, so the stupidity must continue.

7. Use Dragon v2 to transfer the Mars mission crews to and from the MTV after the MTV has been positioned at L1 using a SEP tug.

* If it's not clear, Falcon Heavy launches Dragon v2 directly to L1 where the MTV will be waiting for the crew to board. After the crew has boarded and determined that the MTV is ready for TMI, the SEP tug that transferred the MTV to L1 detaches. The empty Dragon detaches and returns to ISS so that a departing ISS crew can use it to return to Earth.

* Once the MTV returns to L1 after a Mars mission has been completed, another Falcon Heavy launches another Dragon v2 to return the crew and samples to ISS from L1. A refueled SEP tug transfers the empty MTV back to ISS for refurbishment.

8. Use Falcon Heavy to launch M113/MTVL rovers to Mars using SEP tugs. This means the surface exploration habitats require single Falcon Heavy launches.

* A SEP tug transfers the MTVL and EDL hardware to Mars.

* The unmanned MTVL's aerocapture and then land using ADEPT + parachute + retrorockets.

* 4 MTVL's will be landed before the MTV arrives.

9. The MDV/MAV can be one of two designs.

* The first design, which everyone else seems to prefer, is a massive vehicle that lands the entire crew using ADEPT and retrorockets for soft landing.

* The second design, which I favor, lands each crew member individually in an unpressurized micro capsule employing HIAD for entry and a large ringsail parachute forced open using an inflatable collar for soft landing.

* In either case, I would send the MDV/MAV ahead of the crew using a chemical kick stage and SEP tug.

* In either case, a minimum of two Falcon Heavy flights are required.

* Using the multi-person MDV/MAV design, the first flight launches the MDV/MAV and the second flight launches the chemical kick stage and SEP tug.

* Using the single person MDV/MAV design, the first flight launches 4 MDV micro capsules, 4 MAV micro capsules, and the SEP tug. The second flight launches the chemical kick stage required to send the MDV's and MAV's to Mars. The MDV's stay attached to the SEP tug in LMO and the 4 MAV's land using ADEPT for entry and retrorockets for soft landing. The MTVL's will then move the individual MAV's away from each other to a pre-designated launch area.

10. A quick review of the launch costs alone indicates why this plan has some chance of success.

* Each MTV requires 5 Falcon Heavy launches (one launch is required during the first launch opportunity to transfer the TEI kick stage to Mars)

* Each MDV/MAV requires 2 Falcon Heavy launches

* Each MTVL requires 1 Falcon Heavy launch

* Assuming a crew of 4 and a 500 day surface stay, 7 (edit: 4 for MTVL's, 2 for MDV/MAV, 1 for TEI kick stage) Falcon Heavy launches are required during the first launch opportunity and then 5 (edit: 3 for MTV, 2 for crew transfer) Falcon Heavy launches are required at the next launch opportunity. The total launch costs for one surface exploration mission amount to the cost of one SLS flight. For the initial orbital mission, the launch costs are half the cost of one SLS flight.

* Apart from technology development, the most significant operational cost for a manned Mars mission is launch cost.

- 12 Falcon Heavy (~$1.2B): 636t to LEO

- 1 SLS, assuming development of a suitable upper stage (~$1B to ~$1.5B): 95t to LEO; without 5 RS-25's more than ~110t to LEO is pure fantasy

* By separating the surface stay hardware from the transit hardware, we can assemble and ship manageable payloads to Mars. There's absolutely no need to assemble and transfer an ISS sized spaceship to Mars.

Last edited by kbd512 (2015-07-04 22:06:42)

kbd512 · 2015-07-04 22:49:46

Using my mission architecture, adding more surface stay hardware does not affect the mass of the MTV at all and has minimal impact on the cost of the mission. The MTV can rendezvous with the MDV/MAV in LMO. This is simply routine for crew and cargo transfer to ISS. The MTV is just a miniature space station. The only difference between ISS and the MTV is that crew and cargo are transferred to/from Earth instead of to/from Mars.

If someone at NASA can't live without including a stationary surface habitat module as part of the mission, two additional Falcon Heavy launches would deliver it to Mars. My take on Mars surface operations is that mobility is life. The MTVL's provide the mobility required for real surface exploration and contingencies, whereas a stationary habitat module and unpressurized rover does not. A stationary habitat module, ISS MPLM or inflatable, would be nice to have, but adds nothing to a surface exploration mission and is a single point of failure unless two or more are landed in close proximity to each other.

RobS · 2015-07-20 21:11:21

This is a fascinating 100-page report that was just announced by the National Space Society today:

http://www.nss.org/docs/EvolvableLunarArchitecture.pdf

You all should take a look. It proposes, among other things:

1. A public/private partnership to go to the moon, then Mars. It estimates such a partnership will be able to accomplish their goals for 1/8 to 1/10th the cost that NASA could accomplish them.

2. They propose using Falcon and Falcon Heavy, though they note the new rocket proposed by Boeing could be used instead, but their report was too far advanced to add it in detail.

3. They use the SLS for only three launches and I don't think they use Orion at all.

4. They propose to use stretched Falcon second stages and refuel them in LEO using Falcon 9 Rs (reusables).

5. They propose to use the Dragon with two "trunks" to serve as the command/service module. The second trunk would contain 10 tonnes of additional hypergolic fuel to provide lunar orbit insertion and trans-Earth injection. The resulting vehicle would be launched by a Falcon Heavy, I think with a second stage refueled in LEO.

6. They propose a lunar module using Superdracos and hypergolic propellant. It would be capable of landing a 7-tonne ascent stage PLUS 7 tonnes of cargo, or 14 tonnes of cargo on the lunar surface. Altogether the lunar module would mass 30 tonnes. It would be orbited by a Falcon Heavy, I think, and use a refueled second stage for TLI. I suspect the dragon would provide some of the architecture.

7. The lunar transport system would start with landing crews at the equator and robots at the poles, then would start landing crew at the poles (the delta-v is 200 m/second more than at the equator), where they would set up a 4-person base using a Bigelow inflatable. Their purpose would be to maintain equipment to produce LH2/LOX propellant from polar water.

8. The lunar architecture would transition to a hydrogen/oxygen reusable vehicle for transport between the lunar surface and space.

9. Hydrogen and oxygen manufactured at the lunar poles would be used to fuel a Martian vehicle at L2. They say this would cut the cost of Martian transport.

10. The moon project would eventually fall under the responsibility of a public/private institution, the "Lunar Authority" modeled on the multinational institution that builds and expands CERN (the European atom smasher) or the Port Authority of New York and New Jersey. The later raises all its money commercially to expand its ports, harbors, bridges, and tunnels.

Quite interesting set of suggestions. Not much is offered about the Martian architecture, except a lot of the lunar architecture would be designed for Martian use as well, especially life support and ISRU.

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