Journey time to Mars...

louis · 2018-11-03 20:33:20

It's nothing to do with a You Tube video which makes me think perhaps you haven't actually looked at it.

Oldfart1939 wrote:

Louis-
I don't generally debate YouTube video presentations. I debate the absolute physics that apply. It's science, mathematics, and engineering. And you simply asked a question as the original poster. I answered to best of my ability, as did GW. Sorry it doesn't agree with wishful thinking and cheerleading.

Oldfart1939 · 2018-11-03 22:20:16

The charts bear little resemblance to what is actually achievable with safe landings/aerodynamic atmospheric entrances. The data was actually confusing, and the commentary later also tended to confirm my assertions.

louis · 2018-11-04 04:09:38

OK, so your conclusions lead you to think the chart is faulty, drawn up on the basis of erroneous assumptions. Thanks for that clarification.

Oldfart1939 wrote:

The charts bear little resemblance to what is actually achievable with safe landings/aerodynamic atmospheric entrances. The data was actually confusing, and the commentary later also tended to confirm my assertions.

SpaceNut · 2018-11-04 12:15:29

Part of the trouble is the next big thing link a page back is mixing rockets engines, electric versus nuclear and several types of transfer orbit paths to give you a false mix of numbers that make it seem like we can get to mars in a blink of the eyes.

Earth departure, destination arrival, Earth return, and Earth arrival dominates the overall mission V, even when launch is included. The rocket equation thus tells us that most of the launch mass, and therefore most of the cost.

https://ocw.mit.edu/courses/aeronautics … _Lec17.pdf

Transfer orbit types:
https://en.wikipedia.org/wiki/Hohmann_transfer_orbit

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

https://en.wikipedia.org/wiki/Bi-elliptic_transfer

https://space.stackexchange.com/questio … athematics

rocket engine types:

Vasimr plasma drive
http://www.adastrarocket.com/Andrew-SPESIF-2011.pdf
https://www.nextbigfuture.com/2018/03/a … times.html

We have a topic here and have debunked the claims

All trying to achieve the space x mission https://en.wikipedia.org/wiki/SpaceX_Ma … astructure

kbd512 · 2018-11-04 17:47:38

If you want to spend a lot of time and money shipping propellants, then an all-chemical mission architecture will assure that. If you want more payload and less propellant, then combine chemical with high-power SEP. It's that simple.

GW Johnson · 2018-11-05 09:48:52

Kbd512 is quite right that high-power electric propulsion would help increase payload fractions, but not all alone, especially manned items.

Chemical propulsion exists and is flying right now. Generating a new design from an existing technology base is not that big a deal, although certainly not trivial. You can have pretty much any chemical rocket engine burning any propellant combination that you want. You are NOT restricted to only pre-existing designs. Although those pre-existing designs (all of them back to the beginning) ARE your existing technology base.

Electric propulsion exists and is flying now, at rather small thrust levels. It gets used for satellite station-keeping, and a couple of interplanetary probes so far.

High-power electric propulsion is being worked-on, but is not yet flying! NASA and Ad Astra Rocket are the two largest outfits doing work on high-power electric propulsion. The NASA effort concerns me, because it has spent the best part of the last two decades, and a fair fraction of a trillion dollars, on a new giant launch rocket in one form or another, and is still years from flying. Ad Astra suffers mostly from inadequate funding, so the marketing hype is overstated to win more.

The fundamental problem with high power electric propulsion is the source of the power. This could be solar out to about Mars, but not much further. The other option is nuclear. The trouble is which grows much faster as you scale up, the watts or the kilograms? I think with solar we have finally reached the point where watts grow much faster than kilograms. Not so sure about nuclear.

Whatever that answer, you still need a vehicle acceleration capability nearer 0.01 gee than 0.0001 gee, in order not to waste all your high Isp advantages in (sun-centered) gravity losses that act to raise the effective delta-vee of your interplanetary mission. I'm talking about having to double the orbital mechanics delta-vee, to set the mass ratio requirement of a low acceleration vehicle here. Such is just far-outside-the-ballpark infeasible for LEO departure scenarios.

The only thing ready-to-fly "today" is chemical. The high power electric could be flying "soon", if the current actors get with the program and actually do it right. For NASA, that means proper management, not at all similar to how they managed Ares/SLS. For Ad Astra, that means far higher funding.

GW

Last edited by GW Johnson (2018-11-05 09:57:01)

SpaceNut · 2018-11-05 19:27:36

So going with the chart I posted the delta needed is 3.2 to 4.8 km/s before we would drop the stage and go to ion drive at that point as there is no sense to cart around dead weight to mars it could be expended. So if we did cart the eds stage to mars then we would leave it in orbit as a building block with the ion stages for later refueling and reuse.

That would leave 2.9 km/s to get to low mars orbit to be able to stage before landing.

That said lets play devils advicate on the BFR landing and tons of supplies changing the stages refueling numbers and size to correctly land and return to orbit. Now to resize the bfr lander section unit from GW's calculations, then add the ions stages and the right sized eds for the earth location of orbital start.

kbd512 · 2018-11-05 21:08:52

SpaceNut,

I'm going to start by making some assertions that I believe to be fundamentally true:

1. All first stage rocket boosters require moderate Isp propellants that produce a tremendous amount of thrust to minimize losses from gravity and atmospheric drag. Specific Impulse is important, but not all-important. Absolute reliability and good reusability are all-important for cost reduction. To that end, liquid propellants like LOX/RP1 or LOX/LCH4 are two of the best choices available with current technology. However, denser liquids like NTO/UDMH and solids like APCP or HTPB also provide the greatest Density Impulse and most thrust practical with current technology.

2. All second stage or upper / orbital stages require high Isp propellants that produce a lot of thrust with the best Specific Impulse we can reasonably manage. Again, absolute reliability and good reusability are all-important for cost reduction. To that end, LOX/LH2 is the best choice available with current technology. The accumulated body of knowledge associated with LOX/LH2 as a rocket propellant combination is greater than all other propellants combined. Entire branches of materials science and fields of study were created for the use of LOX/LH2.

3. To routinely travel to distant planetary bodies in an economical and therefore sustainable manner, Specific Impulse far above what chemical propellants can provide is required. Current working technology dictates the use of an inert gas (Argon, Krypton, Nitrogen, Xenon) that is easy to ionize for that purpose. Storable elements (Aluminum, Iodine, Lithium, Sodium) may prove useful as propellants in the future, but the propulsion technologies that use them are far from ready for prime time. The most advanced propulsion technologies available are hall thrusters or magneto plasma dynamic thrusters. Some thrusters use electrical energy alone, some use radio frequency energy, and a few like VASIMR and ELF actually use a combination thereof.

4. Shipping propellant to space detracts from useful payload, so there is an incentive to do as little of it as is technologically feasible. There's no reason to avoid shipping it entirely since it's required to go to the places we want to go to, but establishment of a space economy is fundamentally about shipping more useful payload and less propellant.

5. Whatever we come up with right now is primarily about establishment of a foothold on other planets, rather than a technology set that is optimal for mass colonization. I think Elon Musk was quite correct when he posited that even ITS was too small for effective colonization.

Therefore...

Ideally, BFS would be a fully reusable upper stage capable of putting a 100t+ payload in an elliptical lunar orbit. Whether it achieves that in a single shot or by using on-orbit refueling is only important in the sense that it needs to be achieved. An in-space propulsion stage that combines storable chemical propellants with SEP would take over from there. I favor the elliptical lunar orbits as staging points for Mars missions since those orbits aren't affected by the intense radiation trapped in the Van Allen belts, nor the ever-growing debris field surrounding Earth as more and more space junk accumulates.

An all-in-one solution that goes from the surface of the Earth to the surface of Mars and back again is not a worthwhile pursuit in the immediate future as a function of the unknown and unanswerable technological challenges associated with doing that. Immediate reuse of expensive reusable upper stages is far more important than sending a handful of such vehicles to Mars for several years and perhaps retrieving them several years later after two reentry events, provided that all systems continue to function as intended during that time.

BFS is still the best concept available for realistic sub-orbital transport and heavy cargo delivery to orbit and CIS-Lunar space. However, mandating 1,100t of propellant production on Mars to come back to Earth is a step to far. After we establish that humans can live on Mars indefinitely, we can see how well that concept works by experimenting with it. Until then, storable chemicals or modest propellant production for modest vehicles only intended to attain orbit and return to Earth aboard an interplanetary transport vehicle is required.

If we break this problem into smaller problems that are solvable with current and near-term technology, we can go to Mars in the next decade. If not, then ten years later we'll still be spinning our wheels, hoping for development of technology that still doesn't exist.

Oldfart1939 · 2018-11-06 00:53:27

kdb512-

I cannot really disagree with most of your assertions stated above. But what is really needed is somewhat of an eclectic approach to the "grand problem." That's something I've attempted to do over the past few years in which I've been a participant on this forum by suggesting a "modular" system. So what is needed, as you state in your final paragraph , is breaking the problem into a series of smaller ones.
(1) Get to LEO, which has been done by Soyuz, and the Space Shuttle to date. Hopefully we'll see the SpaceX Dragon 2 achieve that in the next 6 months to a year. Ditto, the Boeing Starliner capsule.
(2) Trans-Mars departure, for which we have only the promise of BFS sometime in the future. My interim solution is a combination of an expanded Dragon 2 with a dedicated module powered either by you choice of liq H2/LOX or a storable combination of UDMH/NTO or UDMH/LOX. I propose using either BFR or a Falcon Heavy to orbit into LEO a fully fueled but dockable module, capable of docking with the modified Dragon, which would have an interim landing module interposed. But that brings us to (3).
(3) Landing on Mars. Since there would be some residence time on the Red Planet, and also a long flight along a Hohmann trajectory of possibly 150 to 180 days, a storable propellant system makes most sense. NTO/UDMH, but if some boil off is permissible, then the high Id fuel could still be UDMH but with LOX as the oxidizer.
(4) Mars departure and Earth return. My system would be capable of doing a moxie Oxygen concentration/liquefaction, so the fuel could be incorporated in the Mars lander with provision of resupply of O2.

This is sort of the bare bones of my proposed modular approach. I could actually write a lot more and in mathematical detail, but this is enough to get the idea. Requires no Sabatier reactor, but only Moxie. Oxygen is the much larger mass required for Mars departure than say CH4 or UDMH.

Added in edit, as a P.S. I've leaned heavily on Falcon Heavy, as it is a vehicle which HA FLOWN, and for which there is already a manifest of 4 launches scheduled.

Last edited by Oldfart1939 (2018-11-06 01:02:30)

kbd512 · 2018-11-06 05:10:13

Oldfart1939,

My remarks are as follows:

1. Dragon 2 and Starliner figure heavily into this proposed mission architecture. I also think LOX/LCH4 is the best way to move forward with the boosters, along with virtually every major American space launch services provider not engaged in government contracts specifying LOX/LH2 or LOX/RP1 and solids. NASA now baselines LCH4, rather than LH2, as the fuel of choice for their landers. It's the cheapest oxidizer and fuel combination presently obtainable that does the least amount of damage to the engines. The new engines developed to use LCH4 are nearing the end of their development cycles and all have adequate government funding backing them.

2. I would like to see a Falcon Heavy booster with an ACES upper stage variant. SpaceX is already providing the boosters and spacecraft. ULA needs to do something useful for our human space flight program by providing their IVF (Integrated Vehicle Fluids) equipped upper stages to permit brief on-orbit storage of LOX/LH2. BFS will ultimately require some sort of IVF system to avoid the well-known mass and complexity penalties associated with GHe COPV's for propellant pressurization. ULA already has their IVF technology at a very advanced state of development, so it's the natural technology provider for that subsystem.

If long duration cryogen storage and orbital refueling concepts are proven to be viable, then we can design missions around them. The propellant tanks in our workhorse vehicles (Falcon 9, Falcon 9 Heavy, Atlas V, Vulcan, New Glenn) need to be switched to use these new carbon composites to first demonstrate that the lighter materials are truly suitable for hard use. Once again, some sub-scale validation testing is required and ULA is closest to being ready to demonstrate it using IVF.

There's no point in designing BFR/BFS to use composite primary structures if the performance of those structures is not up to par in the expected thermal and radiation environments. The workhorses are of sufficient size to adequately characterize composite performance. It's not the case that no experience is available, rather that insufficient operational experience is available from any technology provider to base any human mission on it.

3. The landers must use NTO/UDMH (the Russians would use this combination) or NTO/Aerozine-50 (most likely, given American experience) for the first missions. Nothing else is ready for prime time until we know where the water is, what varieties of nasties are mixed in with it, and whether or not it's mostly garden variety H2O or heavy water. There's also no telling how well MOXIE will work in a dust-laden atmosphere and it requires a lot of electrical power. The first ISPP experiments can be LOX and Argon production. The LOX is dual use for life support and as an oxidizer. For Aerozine-50, the specific impulse and density impulse between LOX and NTO is similar enough that a vehicle can be modified to use either, as necessary. I don't see a practical application of RFNA as an oxidizer here, given the lower performance, lack of known Nitrogen sources on Mars, and the more complicated processes required to produce it.

Argon is the only truly cost-effective inert gas for electric propulsion and it also happens to be available on Mars and Venus. Here on Earth, Argon costs about $5,000 per metric ton. The grade of LCH4 suitable for rockets would cost about $1,000 per metric ton. LOX is pretty inexpensive, but we'd still have to add a bit more to that cost to actually use the LCH4. The Argon is comparatively expensive, but you need a lot less of it.

4. For Earth return, I'd favor the use of electric propulsion. That avoids the requirement to manufacture tons of propellant or store propellants in orbit for a couple of years. It drastically reduces the Mars surface electrical power and Mars ascent stage propulsion requirements as well. Achieving orbit is all that's required. Instead of returning directly to Earth, the astronauts would be headed back to the moon. I'm mixing some of the old chemical propulsion technology with some of the new electric propulsion technology, but I think the benefits of electric propulsion outweigh the requirement to obtain anything from Mars just to come home.

5. I'd agree that Falcon Heavy must factor greatly into any realistic near-term technology development and testing effort. I'm hoping that New Glenn and Vulcan make rapid progress to provide like-kind alternatives if Falcon Heavy becomes unavailable for any reason. As far as I'm concerned, both SLS and BFS are just paper rockets until they've flown at least one mission and I don't care how earnest or advanced the development efforts are.

Oldfart1939 · 2018-11-06 10:18:36

kdb512-

There are many requirements before I would send crewed flights using this strategy, and having a constellation of communication/GPS satellites in orbit is primary concern. I would also make this a 3 lander operation, with only one vehicle for Earth return. One would carry a huge supply of food and supplies, the other would be a nuclear reactor to power the moxie system and provide energy for the base camp. I've made an attempt to maximally utilize vehicles we currently have flying (Falcon 9 Block 5, Falcon Heavy, but with Block 5 upgrades). I envision the Falcon Mars lander as having the cargo trunk be incorporated into a pressurized segment of the spacecraft for crew quarters and supplies. The interim landing stage would carry the engine(s) for the lander and landing legs would be attached. The trunk of this stage would be left behind with the legs after Mars is reached and during departure therefrom.

louis · 2018-11-06 16:01:24

This confirms that Space X is claiming 80 to 150 days and an analysis of a Non-Hohmann Transfer possibility by a contributor (Hop David) suggests that 140 days (just over 4 months) - presumably an average - would be possible.

https://space.stackexchange.com/questio … mi-delta-v

kbd512 · 2018-11-06 16:38:28

Louis,

Possible vs practical are two entirely different questions. It's possible to make a four banger crank out 2,000hp with enough boost, but when you do that the engine has a lifespan measured in seconds. All wishful thinking runs into math and physics, whereupon reality quickly sets in. Getting there as fast as possible is not a priority. Getting there with adequate backups to ensure survival is of paramount importance.

louis · 2018-11-06 17:00:37

Did you read Hop David's analysis?

kbd512 wrote:

Louis,
Possible vs practical are two entirely different questions. It's possible to make a four banger crank out 2,000hp with enough boost, but when you do that the engine has a lifespan measured in seconds. All wishful thinking runs into math and physics, whereupon reality quickly sets in. Getting there as fast as possible is not a priority. Getting there with adequate backups to ensure survival is of paramount importance.

kbd512 · 2018-11-06 17:12:13

Louis,

Yes, I did. The arrival velocity is sufficient to impart mild brain damage to the crew from the shaking they'll receive from the buffeting during reentry. Problem solved. Humans can indeed get to Mars in 3 months... Just not without some mental impairment. Perhaps a minor detail, but I'm pointing it out to you. Ares I didn't work out for that reason (the vibration problem, but caused by the SRM in that case, rather than extreme reentry velocity and the wildly varying pressure gradients in the Martian atmosphere). That's why NASA dropped it. The vibration was too great. The astronauts couldn't read the control panel inches in front of their face, much less manually control the spacecraft if required to do so. The vehicle could handle it without a problem. The humans couldn't. Why is that so hard to understand?

Once again, possible vs practical are two entirely different things.

Terraformer · 2018-11-06 17:20:54

On the other hand, departing from EML1 and using the Oberth effect leaves you with a lot more propellent to use at the Mars end of the trip. It can cut off ~2km/s, if I remember correctly.

kbd512 · 2018-11-06 17:53:11

A 10km/s reentry velocity would work. An 11km/s reentry velocity is pushing what the crew can physically take in terms of vibration, not G-force. A 12km/s reentry velocity or greater would require some sort of vibration isolation system to prevent the crew from being physically shaken to death. They tried that during the Ares I program, with limited success, before the idea was finally abandoned entirely when it became abundantly clear that Orion would be too overweight for the booster to lift to the required orbit. Orion later suffered from cracking of the pressure vessel during EFT-1 from those aerodynamic forces that refuse to be ignored, which lead Lockheed-Martin to completely redesign it. Shaking infants does more damage because they're more delicate, but shaking an adult can still injure or kill them. Shocking, but true nonetheless.

Why injure or kill your crew to try to make some point about trip time that was lost before it started, for lack of understanding of all the physical forces involved?

What magic is there in a 4 month vs 6 month transit if you're going to spend the rest of your life on Mars?

Alternatively, we could face the fact that the mechanics of doing what Louis wants to do may not be in the best interest of the crew, whether we're talking about the cost of shipping propellant, abnormal wear and tear on BFS (composites have brittle failure modes), or the lives of the crew members. Thankfully, whether he understands the limitations of human physiology or not is irrelevant.

Oldfart1939 · 2018-11-07 00:39:54

Louis-
You will note that the Ares I made only a single flight, not to be repeated. The vibration that developed by that very long solid booster was excessive. In addition, the Orion capsule was already overweight and could not have achieved orbit if built as designed. I had several communications with the NASA engineer at Ames-Moffit Field who made the recommendation that it be cancelled. The possibility of TBI (Traumatic Brain Injury) was too extreme for ultra conservative and very risk adverse NASA.

louis · 2018-11-07 07:02:30

Re the Ares 1 this is what I found on Wikipedia:

"In January 2008, NASA Watch revealed that the first stage solid rocket of the Ares I could have created high vibrations during the first few minutes of ascent. The vibrations would have been caused by thrust oscillations inside the first stage. NASA officials had identified the potential problem at the Ares I system design review in late October 2007, stating in a press release that it wanted to solve it by March 2008. NASA admitted that this problem was very severe, rating it four out of five on a risk scale, but the agency was very confident in solving it. The mitigation approach developed by the Ares engineering team included active and passive vibration damping, adding an active tuned-mass absorber and a passive "compliance structure" – essentially a spring-loaded ring that would have detuned the Ares I stack. NASA also pointed out that, since this would have been a new launch system, like the Apollo or Space Shuttle systems, it was normal for such problems to arise during the development stage. According to NASA, analysis of the data and telemetry from the Ares I-X flight showed that vibrations from thrust oscillation were within the normal range for a Space Shuttle flight."

My translation: the vibration issue was not an insoluble problem. Also, the BFR is not Ares 1, on that we can agree.

My reading of the Hop David non-Hohmann analysis (note - it's where he states Here - some people might have missed it) is that the incoming velocity is 9.8 Kms per second at which point there's a retro burn. Musk showed in his presentation how they expect to reduce speed - there's a graph he displays.

I think we ought to start from the presumption that Space X are well aware of the risk of TBI from vibration and are eliminating the risk.

Last edited by louis (2018-11-07 07:02:52)

kbd512 · 2018-11-07 07:19:22

Oldfart1939,

Ironically, Lockheed-Martin's engineers put Orion on an extreme diet following EFT-1, creating a design substantially lighter capsule (2t+) and stronger at the same time.

NASA came up with seats that could isolate the crew members from the extreme vibration levels that the SRM produced, but there were still problems with manual control of the spacecraft. The "floating" seats would actually prevent injury to the crew, but then they're left with the inability to even hit the right switch on a control panel. That problem was also solvable by mounting the displays and flight controls on the seats, but now we're talking about a complete redesign of the displays and flight controls and an associated rewiring job that would've produced an inability to remove the seats for maintenance without disconnection of the flight controls. The seats were intended to be stowed / folded in orbit to give the crew more room in orbit for activities. I believe that same design feature is present in the Dragon 2 and Starliner designs. The interior volume of the capsules is a lot less than that of the Space Shuttle, thus the requirement.

There was also a problem with the vibrations potentially breaking electrical connections and damaging the heat shield. Those were the more intractable problems.

What should be the takeaway lesson here?

In the same way that exposing the spacecraft and crew / passengers to excessive acceleration (G-force) is bad, so is excessive vibration. Even if the spacecraft is so solidly built that it can take the abuse without issue, the primary goal of all this design and test work is to ensure that the crew is no worse for wear after the ordeal. If the Challenger and Columbia crews escaped unharmed, nobody would really care about the fact that we lost a couple of expensive machines. That happens every day in aviation. Newer and better machines are built when that happens. Sadly, we can never replace the experience and training lost when we lose a crew.

louis · 2018-11-07 09:09:24

But the Hop David analysis says the Mars approach velocity would be 9.8 kms per second, not 11 or 12.

Clearly vibration is an issue. And clearly Space X are already well aware of it:

https://www.reddit.com/r/spacex/comment … customers/

kbd512 wrote:

Oldfart1939,
Ironically, Lockheed-Martin's engineers put Orion on an extreme diet following EFT-1, creating a design substantially lighter capsule (2t+) and stronger at the same time.
NASA came up with seats that could isolate the crew members from the extreme vibration levels that the SRM produced, but there were still problems with manual control of the spacecraft. The "floating" seats would actually prevent injury to the crew, but then they're left with the inability to even hit the right switch on a control panel. That problem was also solvable by mounting the displays and flight controls on the seats, but now we're talking about a complete redesign of the displays and flight controls and an associated rewiring job that would've produced an inability to remove the seats for maintenance without disconnection of the flight controls. The seats were intended to be stowed / folded in orbit to give the crew more room in orbit for activities. I believe that same design feature is present in the Dragon 2 and Starliner designs. The interior volume of the capsules is a lot less than that of the Space Shuttle, thus the requirement.
There was also a problem with the vibrations potentially breaking electrical connections and damaging the heat shield. Those were the more intractable problems.
What should be the takeaway lesson here?
In the same way that exposing the spacecraft and crew / passengers to excessive acceleration (G-force) is bad, so is excessive vibration. Even if the spacecraft is so solidly built that it can take the abuse without issue, the primary goal of all this design and test work is to ensure that the crew is no worse for wear after the ordeal. If the Challenger and Columbia crews escaped unharmed, nobody would really care about the fact that we lost a couple of expensive machines. That happens every day in aviation. Newer and better machines are built when that happens. Sadly, we can never replace the experience and training lost when we lose a crew.

Oldfart1939 · 2018-11-07 09:57:48

I'm sure that if excessive vibration becomes an issue, it will be dealt with before anyone has his or her brains rattled beyond repair. I was simply commenting that NASA cancelled Ares I for a multitude of reason, but primarily due to the Obama administration's refusal to throw more money in that direction. My friend at NASA indicated that it was a flawed concept, however.

RGClark · 2018-11-07 11:33:07

Oldfart1939 wrote:

Louis-
The Tsilkovsky rocket equation is really tyrannical since it involves an exponential function. I just number-crunched the payload fractions of 3 different delta V scenarios: delta V = 7000meters/sec; payload of 5% (assuming 10% of total vehicle weight is structural); delta V of 8000 meters/sec; payload = 1.8% (same assumption regarding vehicle structure); delta V of 8500 meters/sec; payload = 0.2 %. These are the limits of the CH4/LOX system with an Isp of 383 sec. No matter what you would LIKE, the limits of the physics do not ALLOW the sort of transit times Musk tosses around without some serious modification to the mission architecture.
As GW has stated above, the entry velocities into Mars atmosphere MUST be considered, otherwise the possibility of skipping off into interplanetary space looms for those riding the spacecraft. The numbers in the Zubrin table that I have referenced are for FREE RETURN to Earth, should the vehicle miss the planet due to mistakes in trajectory refinement.
I don't care what Musk SAYS at this juncture, but only what the Tsilkovsky equation tells us what MUST happen and the constraints involved.

I seem to remember Musk talking of smaller missions, i.e., crew and cargo. Perhaps in those cases he would get faster travel times.
Also the structural weight is actually smaller than 10% according to this image of the BFR:

Presentation Slides – Elon Musk Update on SpaceX’s Interplanetary Transport System.
September 29, 2017

https://spaceflight101.com/iac-2017-spacex-slides/

It's closer to 4% there. The 10% value is closer to the total of structure weight plus payload weight.

Bob Clark

Last edited by RGClark (2018-11-07 11:34:53)

kbd512 · 2018-11-07 13:52:53

Louis,

To begin with, Dr. Zubrin is still correct when he says 6 months is the proper transit duration. Next, I'll presume you understand that you're not actually traveling in a straight line when traveling along an interplanetary transfer orbit. However, straight line velocity approaching the asymptote of the hyperbola (of a non-Hohmann transfer orbit) can be expressed using the term "vinf" since it's so close to actual straight line velocity that it's an acceptably accurate substitute. A Hyperbola is effectively what you're traveling along when you approach the target of your transfer orbit and the target is a planet of substantial size, like Mars. The straight line velocity that BFS must reduce to zero to actually land on Mars, whether by firing the engines and/or aerobraking into the Martian atmosphere upon arrival at Mars, can be roughly derived and termed "vinf". This is how fast you're going at the interface between your orbit and the upper atmosphere of Mars. If BFS doesn't kill enough velocity to land, then it becomes a man-made crater, gets ejected into interplanetary space if remaining velocity exceeds escape velocity, or enters into an orbit around Mars that it can't get out of without using propellant it doesn't have.

Here's a link to Hop David's blog explaining the concept:

What the heck is Vinf?

Further explanation of the concept from Hop David:

Do transfer orbits toward the central star necessarily result in a higher velocity on arrival due to the star's gravity?

From the linked article above:

"Vinf is a hyperbolic orbit's speed at infinity. For practical purposes a Hohmann transfer becomes hyperbolic when entering a planet's sphere of influence and Vinf refers to the difference in speed between heliocentric orbits."

Could you please tell us what number you're looking at on the "NonHohmannEarthToMars.xlsx" spreadsheet?

Are you looking at cell M25?

If so, that's the dV increment to achieve TEI and has nothing to do with his calculated straight line arrival velocity at Mars. I guess Hop David thought that SpaceX might want their BFS back after it lands on Mars.

Bob,

Since you or someone with your screen name appears to have read Hop David's blog post from 21 March 2013, could you explain this concept to Louis in a way he understands?

Maybe I'm just terrible at math today, but could you tell me how you derived a 4% structural mass fraction from the total mass?

1,100t (propellant) + 85t (vehicle) = 1,185t (total mass with no payload)

85 / 1185 = .07172 (structural mass fraction of total mass with no payload)

1,100t (propellant) + 150t (payload) + 85t (vehicle) = 1,335t (total mass with max payload)

85 / 1,335 = .06367 (structural mass fraction of total mass with max payload)

I believe .06367 (6.367%) is as low as it will ever be, given the structural mass and maximum propellant mass figures provided by SpaceX. I've no clue how they arrived at 85t without building their vehicle, so it must be an estimate of some kind. If you short load propellant or take away payload, then the structural mass fraction only goes up, not down. The only way that structural mass fraction can go down is if you can cram more propellant mass into the tanks without increasing structural mass or, of course, if structural mass goes down.

SpaceNut · 2018-11-07 18:11:57

Louis the delta map I put in post #29 gets all of the engine burns for where you start as you progress to the destination.

Using the equation for the time of the engine firing you can work out payload and fuel loading to make it happen.

New Mars Forums

Announcement

#26 2018-11-03 20:33:20

Re: Journey time to Mars...

#27 2018-11-03 22:20:16

Re: Journey time to Mars...

#28 2018-11-04 04:09:38

Re: Journey time to Mars...

#29 2018-11-04 12:15:29

Re: Journey time to Mars...

#30 2018-11-04 17:47:38

Re: Journey time to Mars...

#31 2018-11-05 09:48:52

Re: Journey time to Mars...

#32 2018-11-05 19:27:36

Re: Journey time to Mars...

#33 2018-11-05 21:08:52

Re: Journey time to Mars...

#34 2018-11-06 00:53:27

Re: Journey time to Mars...

#35 2018-11-06 05:10:13

Re: Journey time to Mars...

#36 2018-11-06 10:18:36

Re: Journey time to Mars...

#37 2018-11-06 16:01:24

Re: Journey time to Mars...

#38 2018-11-06 16:38:28

Re: Journey time to Mars...

#39 2018-11-06 17:00:37

Re: Journey time to Mars...

#40 2018-11-06 17:12:13

Re: Journey time to Mars...

#41 2018-11-06 17:20:54

Re: Journey time to Mars...

#42 2018-11-06 17:53:11

Re: Journey time to Mars...

#43 2018-11-07 00:39:54

Re: Journey time to Mars...

#44 2018-11-07 07:02:30

Re: Journey time to Mars...

#45 2018-11-07 07:19:22

Re: Journey time to Mars...

#46 2018-11-07 09:09:24

Re: Journey time to Mars...

#47 2018-11-07 09:57:48

Re: Journey time to Mars...

#48 2018-11-07 11:33:07

Re: Journey time to Mars...

#49 2018-11-07 13:52:53

Re: Journey time to Mars...

#50 2018-11-07 18:11:57

Re: Journey time to Mars...

Board footer