Micro-Space >> Ultralight Manned Spaceflight

[hop]

dtbaird - 27/12/2007  1:25 PM
By the way, the U.S. has never demonstrated skip entry either, although the guidance algorithm was included onboard the Apollo missions to be used in the event of a contingency.  The Russians tried it, and it is my understanding that they succeeded on 1 of 4 attempts.  Any information on this would be appreciated.

Zond 4 and 5 both had guidance failures which lead to ballistic reentry (4 was self destructed, and 5 recovered by ships pre-positioned at the expected landing zone for a ballistic re-entry) Both 6 and 7 seem to have flown the profile correctly, although 6 was a failure for other reasons.

Astronautic covers this: http://www.astronautix.com/craft/soyz7kl1.htm

I would expect that aerocapture is significantly more delicate than a skip re-entry. Zond could (and did) survive a straight ballistic descent. The g loads were a heavy, but crew would have stood a fair chance of surviving as well.

[kkattula2]

A manned aerocapture to High Mars Orbit , is going to have significantly more battery, computation & RCS margin than your typical automated probe. Not to mention a manual control option, and a non-solar secondary power source, (fuel cells?).

Depending on Earth return strategy, it may not even drop the heat shield. Mars return re-entry to Earth will be higher velocity, but much lower mass, so using the same shield for both may make sense.

A sensible strategy might be to plot an optimum aerocapture pass, but carry sufficient propellant (including lander's) to make up for the worst case estimates.  This could be graded into 3 or 4 levels:

1) Within full mission parameters

2) Lander operations limited (i.e. 1 lander instead of 2, or operating parameters/locations reduced)

3)  Lander operations cancelled.  (orbit only or maybe Phobos or Deimos exploration)

[dtbaird]

"A manned aerocapture to High Mars Orbit , is going to have significantly more battery, computation & RCS margin than your typical automated probe. Not to mention a manual control option, and a non-solar secondary power source, (fuel cells?). "

Battery margin will be more generous, and I would agree that a non-solar power source will likely be used, though fuel cell will not be the solution.  Let's assume that the spacecraft does not have to re-point upon exiting the Martian atmosphere.  Of course, this is complete BS because it still must properly point to raise periapsis and then point the antennas toward Earth, but I digress.  In this scenario, the spacecraft "only" has to enter the Martian atmosphere with a periapsis of ~40 km, properly fly the flight path while maintaining the integrity of the heat shield, exit the atmosphere, propulsively raise periapsis prior to re-entering the atmosphere (which would otherwise mean certain death), trim the orbit, safe the engines, and point the antenna toward Earth.  So, let's see what the added capability over an automated probe would do for the astronauts onboard.  Extra battery margin means nothing if solar arrays are not used.  Computational power for guidance would differ little from the probe. though a backup flight software system could enhance relaiability.  RCS margin may be greater to allow for more orbit trims and faster slews, but that, in and of itself, will not decrease the chances off a second (and deadly) re-entry.  Manual control will definitely be available, but humans would not be able to manually fly the drag profile, although I admit that the flexibility of human intervention could bee helpful outside of the atmopshere.  

"A sensible strategy might be to plot an optimum aerocapture pass, but carry sufficient propellant (including lander's) to make up for the worst case estimates."  
Estimating the errors incurred by flying the drag profile is no problem.  The problem is how to deal with an anomaly, and while in the atmosphere, major anomalies are unlikely to be survivable.  If you are not in the proper orbit, it is much more difficult and unlikely to be able to execute any kind of surface mission, though I admit that it would be an interesting trade to study.

"Depending on Earth return strategy, it may not even drop the heat shield. "
As for the heat shield, shuttle tiles could not be used, leaving us with only ablatives.  I have not heard of any ablative heat shield that we have that could withstand 2 entries at the speeds being contemplated, 1 at Earth & 1 at Mars with a huge thermal soak in between, (a dual heat pulse from a skip entry is hard enough).  In fact, I have not heard of a conventional ablative shield of the size required for a reasonable human mission that could even be fabricated.  Problems exist trying to build a 5m diameter PICA shield for MSL and CEV. Therefore, the question of being able to re-use the heat shield is moot as there would be major problems just building a heat shield that could be used once.  By the way, I have ignored the minor detail that it would have to be deployable.

In short, while aerocapture at Mars will be a necessary technology for a human Mars expedition, I strongly disagree with the assertion made by rpspeck that "It appears to me that aerocapture to a high, elliptical orbit is quite well proven."  Mr. Speck, with all due respect, aerocapture has not even been demonstrated at Earth, much less Mars.  It is a technology that is far from proven, though I truly wish that was not the case.  Being able to hit the proper aimpoint at the top of the atmosphere has been demonstrated.  In fact, for MER-A we hit the aimpoint to within 200 m without ANY optical cameras, and for MER-B we were well within 1 km.

[rpspeck]

Thank you for your informative response.  

I struggle with “It Hasn’t Been Done Before!” complaints since a small tech company stays in business by doing things which haven’t been done before, month after month.  It is called “progress” in technology.  I have been doing this for fifty years and have patents and licensed proprietary technology as proof of success.  Everything is of course more difficult to finish successfully then to describe as a suggestion, but the fact that I have developed and successfully flown 17 liquid fueled, bipropellant rockets and two successful finless, gyro referenced guided rockets using gimbaled thrust vectoring confirms an ability to complete projects with complex technology.

I am also perturbed when writers paint problems I have solved in those efforts as monumental obstacles.  Truthfully, reorienting a spacecraft toward the sun in the hours after an aerobraking pass is not very demanding compared to correcting attitude in the fraction of a second required for thrust vectored hovering!   I am well aware that the extreme cost of space “experiments” HAS PREVIOUSLY severely perverted the environment for technical progress in space.  The fact that this situation has changed has not yet sunk in.  Space experiments can be conducted at up to 95% lower cost that previously with a number of new flight options, but these have been “undersubscribed” by an industry accustomed to enormous financial barriers and glacial progress. (What other industry boasts of a leading product operational in 1965? (the Soyuz))

I appreciate any input which clearly identifies What, in my proposals, hasn’t been done before and Why it hasn’t been done.  Then I know what novel issues I am confronting.  In the aerobraking context, it is useful to point out that very low delta V passes have been used at Mars to allow the solar cells and antennas to remain deployed (and not burn off).  Given no benefit from rushing the orbit lowering process with a deeper plunge into the Martian atmosphere, simplifying the overall process by leaving the solar cells extended has a big payoff.  Once the option of aerobraking capture is rejected, there is little reason to accelerate the lowering process.  But that does not mean that faster braking is overly difficult or impossible.  

Recognizing a greater than 50% loss of payload mass with rocket braking into low Mars orbit (using no aerobraking before reentry) and its effect on mission cost, it is certainly useful to consider aerobraking in detail.  

I am surprised that achieving 900 meters per second aerobraking delta V is considered to be “Extremely Difficult” compared to the 11,000 meters per second accomplished by Apollo. At 5.7 km/sec total Mars intercept velocity, the specific energy involved is only 27% of the Apollo case, and the actual energy dissipated in 900 meter/sec aerocapture is only 29% of that total, or 8% of the specific energy absorbed by Apollo.  Since improved ablative materials have been developed since 1965 (Apollo design freeze), why is absorbing 1/12 the energy (per vehicle pound) a big issue?  Very low “lift” reentry body shapes of course produce low trajectory sensitivity to attitude.  

Regarding “deployment” of a 5 meter heat shield, I am staring at a chart showing 5.2 meter launch faring diameters:  skip the deployment and fly with a rigid heat shield. Since I am discussing unusually light weight Mars missions, that diameter is enough.

As indicated, I appreciate and welcome real information, but have a limited tolerance for inveterate “naysayers”.          

[hop]

rpspeck - 28/12/2007  4:51 PM
I struggle with “It Hasn’t Been Done Before!” complaints since a small tech company stays in business by doing things which haven’t been done before, month after month.

"Newspace" has been claiming that it can do things 10x faster and 10x cheaper than the NASA and big boys for years, making exactly the same sort of arguments. The failure to live up to these claims has been stunning it's totality, with only a couple of exceptions very specific areas. (To be I should point out the big boys have made similar claims with equally abysmal results.)

The few newspace companies which have actually gotten to the point of bending metal and flying things have almost universally found it far harder and more expensive than they expected.

Given this record, you shouldn't be surprised that your own claims are met with skepticism, especially when they don't seem to be well grounded in existing experience.

Truthfully, reorienting a spacecraft toward the sun in the hours after an aerobraking pass is not very demanding compared to correcting attitude in the fraction of a second required for thrust vectored hovering!

As pointed out earlier in the thread, for aerocapture you almost certainly need to control your trajectory precisely, in real time, while in the atmosphere. This is not impossible, but very few vehicles have made active control in this flight regime survival critical, and none have done it in mars' highly variable upper atmosphere.

To describe this "quite well proven" invites the sort of skepticism mentioned above.

This pdf has extensive discussion of a NASA/Boeing areocapture study:
http://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19930014062_1993014062.pdf

[rpspeck]

Thank you for an unusually informative response.  Obviously, it will take me some time to absorb the 674 pages of this document, but a quick survey shows that it contains information I have been looking for.

Relative to the “New Space” optimism (and remembering that projects from home building, to electronics, computer programs and spaceflight usually involve unexpected complexities and take longer than planned)  note that several low cost “CubeSats” have flown with Attitude Determination And Control systems.  At least one had operational “Plasma Thrusters”.  Progress in this under funded arena is pretty impressive.  I personally see no sign that such efforts have reached an insurmountable barrier and that hundreds of thousand of Dollars of funding (instead of tens of thousands) will not bring impressive performance increases.

Back to the present discussion:  Page 666 of this document summarizes the study. It indicates that for the desired mission (Fast Transfer Mars with High Lift Aerobrake) Mars aerocapture is strongly limited by poor knowledge of the Martian atmosphere and by the intercept accuracy that can be guaranteed. Under the reference conditions “some TPS advancement  needed” is listed.  This lesser difficulty had been clarified on document page 275 with the option of “ablator TPS use, lower approach C3, or modify the (lifting) aerobrake shape (along with other options).  The use of a separate aerobraking unit, entering ahead of the main vehicle, is discussed just as I have suggested.  

Even with 7.3 km/sec entry velocity (as opposed to the 5.7 km/sec I discuss) the main heating problem (indicated on page 239) is obviously the high level of radiative transfer expected in a CO2 atmosphere rather than in a N2 and O2 atmosphere.  But the steep heat transfer curve with C3 (evidenced by data on page 263) makes it obvious that the 115 W/cm^2 expected heating peak of shown on page 239 will drop below the existing 68 W/cm^2 tolerance with a modest drop in Mars intercept C3, and pages 164-165 show radically lower C3 mission options.

This information supports, rather than counters, my assertion that aerobraking at Mars is a reasonable (but not, I admit, “proven”) plan.  It should be quite reliable for low energy Mars missions.  The need for guidance during atmospheric flight is also apparent FOR THE HIGH LIFT AEROBRAKES discussed in this paper.  But consider the Russian Soyuz (essentially a sphere) for the ultimate in “low lift” aerobrakes, with no need for precision attitude control in atmospheric flight.  This does increase the accuracy of intercept prediction desired, but eliminates the atmospheric flight problem.

Thanks again for a wealth of information which I have added to my NASA research library.  Without this data base, no “New Space” adventures could succeed.  But when applied to a different “style” of missions this data continues to support the optimistic projections I have been discussing.                

[hop]

rpspeck - 29/12/2007  10:31 AM
Relative to the “New Space” optimism (and remembering that projects from home building, to electronics, computer programs and spaceflight usually involve unexpected complexities and take longer than planned)  note that several low cost “CubeSats” have flown with Attitude Determination And Control systems.

Various microsat and hamsat projects were exactly what I had in mind when I said "exceptions very specific areas". This is a lot of impressive work in this area, and it certainly gives hope.

But consider the Russian Soyuz (essentially a sphere) for the ultimate in “low lift” aerobrakes, with no need for precision attitude control in atmospheric flight.  This does increase the accuracy of intercept prediction desired, but eliminates the atmospheric flight problem.

Careful there. Nominal Soyuz re-entries (and Zond style skip re-entries) are lifting and require active control. Lives have been saved by the fact that passive ballistic re-entries are also survivable, but this is possible because the desired result in either case is hitting the planet. The fact that you endure a higher G loads and end up hundreds of miles off course is just an inconvenience. It's not clear to me that you can design an aerocapture profile that is purely passive.  Relatively small errors would lead either to lithocapture or failure to establish an orbit.

Even if a passive areocapture is possible, it would clearly require detailed current knowledge of the state of the atmosphere. It's also hard to see how you would sufficiently reduce the current uncertainties without some actual test flights, which unfortunately would each require a mars flight of their own.

[dtbaird]

"This information supports, rather than counters, my assertion that aerobraking at Mars is a reasonable (but not, I admit, “proven”) plan."

Well, at least you have come to your senses by admitting that it is not proven.  Of course, I also agree that aerocapture is reasonable.  It's shame that it has not been tried.
 

"It should be quite reliable for low energy Mars missions. The need for guidance during atmospheric flight is also apparent FOR THE HIGH LIFT AEROBRAKES discussed in this paper. But consider the Russian Soyuz (essentially a sphere) for the ultimate in “low lift” aerobrakes, with no need for precision attitude control in atmospheric flight. This does increase the accuracy of intercept prediction desired, but eliminates the atmospheric flight problem."

Huh???  Hope do jump to the conclusion that you don't need to worry about the "atmospheric flight problem?"  You have to fly a density profile to achieve the desired drag profile so that you can control your flight path.  There is no magic that allow you to escape this simple fact. You simply have to either have a drag flap or an offset center of mass and control the bank accordingly.  Predicting the density profile of the Martian atmosphere with the kind of precision required for the proposed "passive aerocapture" is totally impossible (not unlikely, but rather impossible).  We can't even predict the thermospheric density to that level of precision for aerobraking. We allow 100% margin in our density corridor to protect from sudden density increases that could damage the spacecraft via unacceptably high heating rates.  Aerocapture requires flight in the stratomesosphere, where there is virtually no in-situ density data to aid in any kind of density forecasting.


"Regarding “deployment” of a 5 meter heat shield, I am staring at a chart showing 5.2 meter launch faring diameters: skip the deployment and fly with a rigid heat shield. Since I am discussing unusually light weight Mars missions, that diameter is enough."

Please point me to a link to your proposed reference mission.  I skimmed through all pages of this thread and found no links of relevance.



"As indicated, I appreciate and welcome real information, but have a limited tolerance for inveterate “naysayers”. "

I, too, welcome REAL information, and to be honest, I am intrigued by any radical concept.  From what I can tell, your proposal is the epitome of radical, correct?  That being said, I have a limited tolerance for those who proclamations are really nothing more than a bunch of bombastic rhetoric.

[hop]

To further drive home dtbairds point (and if I'm not mistaken, he has significant relevant professional experience):
NASA-TM-101607 Preliminary investigation of parameter sensitivities for atmospheric entry and aerobraking at Mars

Initially Mars entry runs were made trying to achieve capture with a single fixed bank
angle. The problem was extremely sensitive to changes in initial flight path angle and bank angle
and no combination of the two angles was found that would result in capture. Changes of the
order of .0001° would result in either impact or skip-out. At this point, the multiple bank angle
approach was tried.

That's before you account for any global or local variation from the standard atmosphere model.

Searching the NTRs server for "mars areocapture" will find you many more relevant documents. For example NASA-TM-103544 Applications of low lift to drag ratio aerobrakes using angle of attack variation for control goes into significant details on many of the trade-offs involved.

The fact that you haven't even made a preliminary survey of the literature in this area suggests that you aren't in position to make a statements on the practicality of aerocapture for your mission. It also leads one to wonder how well founded many of your other claims might be.

BTOE sketches are fun and interesting, but there's usually very large devils lurking in the details.

[Lampyridae]

I'm no engineer myself, but I agree with what dtbaird and hop are saying. What you are proposing is just a sketch, and I would avoid getting too attached to any one technology. From this Mars mission as it stands, I'd expect to see a deranged astronaut coming back home with bones like twigs and a very high probability of getting terminal cancer.

Launching a human on a Dnepr, for example, would probably result in human jelly or at least a deafened astronaut. They talk about 140+ decibels in rocket fairings. Some novel padding would be required, and it drives weight up. Best to have performance margins coming out of your ears before you start.

As for new.space doing it better than old.space, Spaceship One was developed for $20M, whilst EADS Astrium want $1B for their suborbital rocketplane. Granted that one is a prototype and the other is a finished product, but I thought I'd point out that, besides the rocket engine, the whole SS1 operation was very much in Scaled Composites' field of expertise. They are VERY good at what they do. If I were you, I'd look at developing your strengths and playing to them.

Skintight pressure suits: I would seriously consider contacting Professor Dava Newman at MIT who's been developing the Biosuit.

Heat engines: http://unews.utah.edu/p/?r=053007-1 Seriously worth a look for temperature control and power generation. The Nazis had a similar idea with a sonic cannon based on the same principles.

Solar thermal propulsion: the only high Isp propulsion tech I can think of that can be done on the cheap. I don't know why it hasn't been used earlier. No examples have flown, so you'd have to start pretty much from scratch.

HTP/RP-1 propulsion: I'd have a look at the British experience with the stuff. There might be good data for you there. There's also fancy stuff now like HTP/quadricyclene.

Aerocapture: I would just suggest ditching the idea altogether.

[rpspeck]

UPDATE:

This thread follows Micro-Space efforts to develop and demonstrate affordable human “Deep Space” missions and their necessary systems hardware.  I focus on Mars expeditions as a next step beyond historic accomplishments, with the knowledge that pursuing that goal will both require and enable many people to participate in experimental space efforts in the vicinity of Earth and our Moon.    

Engineers who accept responsibility to lead ambitious projects must be people of  “great faith”, with confidence in their own abilities and in the resources they can tap to solve both the known and the inevitable unknown problems. (Faith is spelled R-I-S-K).  Adventurers must similarly be people of “great faith” since they will face similar problems and their lives are at risk.  Both groups benefit from faith in a supernatural “higher power” willing and able to assist them.  Radical Faith is required for trailblazers!  Many adventurers, in an appropriate context, will admit to times that their lives were saved by improbable, or downright impossible events.  People of faith will continue to lead humanity into new and fruitful fields: often with little or no understanding of the benefits which will flow from their audacious pioneering accomplishments.

The “Dream” of personal spaceflight is “Alive”, but it is comatose and on life support!  In the thousands of hours I have invested (since I discovered that reaching Mars by rocket is easier that reaching common GEO, communication satellite orbits) I have found that the difficulties and dangers of spaceflight have been systematically exaggerated, often massively!  The risks for a Mars expedition are modest compared to those faced by a team attempting to climb Annapurna, or even K2!  And the risks will plummet for those who follow the pioneers.  The minimum costs are modest considered as a world class sporting event, but the sponsorship potential is unusually great.  An adventurous soul has a better chance of succeeding as a Martian explorer than as a major league race car driver,  since the funding requirements are similar.  It is time: rise up and “Seize the Dream”(TM) .

I realized last year that I had taken a close look at each and every requirement for affordable human Mars missions without finding any “show stoppers” or factors which would discredit my risk or cost estimates.  That investigation included much review of published research, familiarization with the applicable theories and analytic techniques and the fabrication of  operational engineering prototypes of life support and other spacecraft systems.  The move into more detailed design continues as well as the inevitably more expensive move to space compatible, production prototypes.  The step beyond that will of course include the million dollar production of the actual systems to be used by space pioneers!  

All of the ventures I have considered are lightweight, minimum cost, “pragmatic” expeditions.  I am demonstrating techniques which make personal spaceflight radically less expensive short term, and will eventually become industry standards.  (As an analogy: a typical American can afford a vacation in Europe – unless he insists on flying his Motor Home to Europe with him for the trip!) Most of the cost of any “pragmatic” deep space expedition will be for commercial “freight service” to Earth orbit.  A “pragmatic” plan is one which does not use the mission as a pretext to argue for development of a new launch system or radical technology, but makes do with commercial offerings.  These apply alpine and SCUBA philosophies to both technology and safety.  They avoid the use of massive, metallic spacecraft hulls (which have proven, SPECTACULARLY, to provide no protection against real hazards)  and as a result may be safer than traditional efforts, as well as radically less expensive.  I will review the results of this research soon.  

We continue to offer transferable options – for specific expedition objectives – to individuals as either an investment vehicle or a way to use early funding to lock in their right to pilot a historic space mission!  

I address the technical requirement of one more “engineering detail” in my following post.  This considers “Aerobraking”, a well understood (but “unproven”) technique which can reduce the mass – and cost – of a Mars expedition by a factor of 2 to 4.  NASA research papers brought to my attention evaluate this in sufficient detail to allow me and others to plan both small demonstration  spacecraft and operational human systems.  I needed, and greatly appreciate, this particular material being brought to my attention since it allows me to move forward with  detailed design.  I certainly see no reason to reject the conclusion of the authors of these three reports which were unanimously positive!  All believed that these procedures were practical and should be employed.  A few uncertainties requiring follow on study, or the demonstration spacecraft mentioned above, were spelled out in each report.

I have complained about inveterate “nay sayers” because people who expect to fail in an endeavor almost always have their expectations fulfilled.  Overconfident braggarts, on the other hand, are merely annoying and may actually stumble on something useful.

One resource I have in these efforts is YOU!  Lacking a large staff to search out and sort through publications and discussions which highlight both real and perceived problems to be faced, I have the hundreds (actually thousands) of individuals who read this material to assist me!

While the accomplishment of a lightweight human Mars mission may not seem overly significant in the grand scheme of things, the same was certainly true of the short flights at Kitty Hawk, 104 years ago.  Yet pioneering aviation efforts eventually transformed our planet and its interconnected economies.  A breakthrough in affordable spaceflight will begin a related process.  

Each of you who contributes to this work – technically or financially (in large or small dollar amounts) – advances this process.  You can take pride in joining others to pioneer this new frontier!

Eventually, the resources of a larger portion of God's creation will be tapped.  We will no longer need to do mining in our “living room”! The resources of the inner Solar System can support a thousand years of growth for a prosperous human community as Earth's “limits to growth” are sidestepped!

Richard P. Speck,  Micro-Space,  1/23/2008,  Denver, Colorado 80204

[Sid454]

Jim - 7/4/2007  2:04 PM

rpspeck - 7/4/2007  2:51 PM

Space Lizard - 8/4/2007  3:32 AM

rpspeck - 6/4/2007  9:29 PM
Anyone who thinks a MMU can’t be coupled with a reentry heat shield will need A WHOLE LOT OF PROOF!

So you want to rendezvous a space station with 8 hr of life support autonomy and almost no ?v capability?

I plan to orbit with 48 hour life support supply (1 liter of LOX + 4 kg LiOH), equipment to reenter at will and Delta V capability (the mass for which also scales linearly with spacecraft mass) as necessary depending on how sloppy the launch vehicle is.  

The Dnepr quotes +/- 4.0 km altitude, +/- 0.04 degree inclination (= 5 km lateral).  With 6 kg (3% 0f 200 kg vehicle mass) of modest performance fuel  I can correct far more than this error: +/- 70 km altitude,  +/- 0.3 degrees inclination.  Using that 70 km altitude correction, I can correct orbital phasing by 1.5% = 1.35 minutes per orbit.  If the Russians can launch this vehicle within one minute of schedule, orbital rendezvous won’t take long.  (I will also be using GPS so that orbital parameters and relative position will be known in minutes to very high accuracy.)

Richard P. Speck,  Micro-Space, Inc.

GPS doesn't help find the station.  

The ISS would not allow that type of fast approach.

To catch ISS you also need a radar/lidar to know how fast you are closing with it.

Strapping a heat shield to an MMU will only give you 9 to 10 hours of life support so you'd have to use a three orbit skylab type rendezvous.

And no you can't just add O2 tanks it doesn't work that way for one you have to also drink  space suits when operating properly have a very dry atmosphere the lithium hydroxide also is a desiccant and there's the need go to the bathroom.

Even on normal 7 hour EVAs astronauts bring a drink bag that is mounted inside the suit.

 Also being in the pressure suit for 48 hours could cause pressure sores and possible lung damage.

Now a realistic concept might be to use something like a modern Gemini but made from composites or a scaled up discoverer capsule.

It should be possible to get a small one place craft down to lets say 1500lbs which means a falcon 1 could be used.

[Sid454]

On propulsion for a manned mission to mars we already have a winner in VASIMR it can reduce the trip betwen earth and mars to 90 days using a 500ton space craft and 300KW of electrical power.
If you are crazy enough one could use a small salyut sized vehicle and a 6 mega watt reactor and earth to mars transfer times can be low as 30days.
http://www.adastrarocket.com/vasimr.html
But you still have to stay on mars for 6 to 10 months to wait for earth to be in the right position unless you want to do a close and I mean close as in inside mercury's orbit solar flyby.
Yes it performance is high enough to do that it's up to 10 to 100x more efficient then a hydrogen rocket or 5K to 50K impulse seconds.
http://www.adastrarocket.com/Advantages.html
But even a close solar flyby is several orders of power less risky then a 600 day long chemical only mars mission .

[meiza]

And Sid, VASIMR needs a big power source that masses a lot. Nuclear power is hard in space because you need a temperature difference and you need big radiators for that. So it becomes heavy.
Vanilla once commented that advertising VASIMR is like advertising wheels that can go huge speeds - the only problem is that there isn't a car.

So just saying VASIMR is great misses most of the point. Currently it is useless.

[Sid454]

The way I see if you don't have high ISP high performance propulsion and don't have the motivation to do the research on it you have no business at all being serious on trying to go to mars.

Advanced propulsion is a do the laundry list type item since it's such an enabling technology it literally in this case means the difference between life and death.

Because as it stands now a crew will likely come home in very poor condition and quiet insane or possibly die.
Also they will likely get cancer later in life since the spacecraft would have lacked shielding which for a very long mission can double the mass of the hab section of the spacecraft.

They might not even  be able to come home after nearly 2 years in space and have to live out the rest of their lives on the moon or something.

BTW the same power source also is being considered for unmanned missions be cause more power means a more powerful transmitter and higher data rates.
The mass wouldn't be that high either the entire JIMO probe which had 30KW was able to fit in an EELV.

The power ,propulsion and propellant of a vasimr mission would only weigh 250tons thats two ares Vs which is a lot less then the 800 ton all chemical 1989 mars reference mission.
The hab or living quarters 25 to 50tons "use a BA330" and the lander maybe 30 to 80tons.
A longer mars stay means a bigger surface hab.

This even compares well to Zurbin's in situ resource missions which nasa criticized for having living quarters that were too small.

Also means shorter missions times which can mean cheaper as one cost often not cited and often missed by many amateurs is the cost of ground support which can often cost almost as much as the rest of the mission.


http://en.wikipedia.org/wiki/Jupiter_Icy_Moons_Orbiter

The lightest realistic chemical mission one that provided it's crew with artificial gravity so their bones don't turn to chalk weighs over 400tons just for the earth mars transfer vehicle.


ISS does no propulsion and it's power requirement just for science and life support is over 100KW so 300KW of nuclear power really isn't that huge.

When you put it in perspective  300KW isn't that much power it's equivalent to the out put of a corvette's engine.

Missions like JIMO are actually more important to future space exploration then studies like the mars 500 and all that closed loop research.

The only other real alternative to VASIMR is nuclear thermo such as NASA"s BNTR study 3x chemical's ISP and still could provide the crew with 30KW.


Another promising technology is gas core reactor rockets which can offer high thrust and high ISP.

The best non nuclear alternative  might be solar electric using vasimr,ion or plasma thrusters.
 The array will have to be over sized about 2x to 4x the size of the arrays on ISS as the sun is only half as intense at the distance mars is from the sun actually about 43% earth intensity.
It only even become possible due to very recent advances in solar cell technology.

Both the solar and nuclear electric mother ships also have one more trick they can do for a mars mission they can beam down power for use on the surface of mars.

[Sid454]

Oh I forgot the vehicle needed to get off the surface of mars but since it doesn't have to go very far just get into orbit and dock with the mother ship it doesn't have to be very big.
Something as small as Dragon or Orion can be used it can even be a modified off the shelf version of either vehicle plus a launch vehicle large enough to get it into mars orbit this could be landed ahead of time with the main habitat and ISRU used to make the fuel for it.

[meiza]

JIMO was cancelled, one of the reasons being the advanced reactor. And it was a heavy craft with mostly radiators and a 1500 kg science payload. Of course the delta vee was huge too.

ISS has the advantage of being near earth. There is less solar power at Mars distance. Though ISS is in shade a significant portion of time.
You can use hall thrusters for low ISP initial acceleration and ion thrusters for high ISP later if you want and get a lot of similar benefits as VASIMR, using tech that has flown for over a decade already.

[Sid454]

Yah canceling JIMO was one of the stupidest things ever as we need to learn advanced propulsion or the moon will become a prison just as LEO has been a prison and constellation will have failed in it's missions of opening up the solar system.

You are too much of a defeatist on this subject as there were no big engineering issues with jimo.
The reactor was no engineering show stopper and the navy was going to help design it.

The real reason it got cut was because Orion started to over run it's budget and pretty much ate up the funding for advanced unmanned missions.

Getting to mars  we really only have three options the first is  NEP or nuclear electric propulsion .
The second is solar electric which actually will be heavier then NEP plus 40% slower as I stated the sun light at mars is only 43% as intense but solar will still work but still with in reach using HLLV from the lunar program.

The last option and the lowest risk plus the cheapest of them all Nuclear thermo such as the old nerva engines or even gas core engines that use UF6 as a fuel a mordern version of this BNTR also adds a low power mode to the reactors which allows them to generate 30KW each this can power mag bubble shields and or ion/vasimr engines.


BTW a 1000 day mission will not work you would be sending that crew to a very horrible and lonely death the real issue here is we need faster travel times more payload and the option to abort to earth at any point in the mission.

Also it's just stupid as can be as even chemical if you got the HLLV  can perform said mission in 500 days tops.

Also on hall thrusters they are low thrust like ion a better combination would be nuclear thermo and hall/ion since guess what they are functionally the same thing.

[meiza]

I've heard otherwise, how the VASIMR reactor technology was a low temperature dead end, to paraphrase. When you put a naval reactor into space, it doesn't work very well because of the fundamentally different environment where the only way to get rid of heat is radiating. There are threads about this from years ago.

Gas core could be nice but it's very hard. Not as hard as fusion though.

There are many problems with the "immediate" nuclear thermal rockets like NERVA, mainly that they most often spew radioactive stuff and can explode when something goes wrong, spreading the highly active core around.

People haven't proposed anything exact that wouldn't be as radioactive and could still be developed in a short amount of time AND would have high ISP and thrust. Clongton has come the closest with a pebble bed reactor IMO.

[hop]

It seems to me that anything involving the N word is pretty much out of the question for Mr. Specs concept of cheap, simple spaceflight developed by a small private organization on a shoestring.

Many of us find his proposals extremely dubious (and he has completely failed to address the problems pointed out earlier WRT aerocapture), but the idea of such an organization developing new nuclear propulsion systems is completely ridiculous.

Just the paperwork to obtain the materials, never mind develop, test and fly such a thing is beyond the reach of anything short of a major defense contractor or government program. Even getting RTGs or isotope heaters (which otherwise have the potential to work well in a minimal complexity mars mission) would be a major challenge.