Where are we now?

louis · 2018-09-02 14:47:19

It's what I got from the presentation. IIRC correctly there was more than one reference to Antarctic bases.

SpaceNut wrote:

Are there any links to the shift in business direction from colonization to exploration. As one is geared to make money while the other is about spending some.

RobertDyck · 2018-09-02 14:53:09

I editted my post above (#26) to include the table where I got the TMI figure for SLS. This is a table that SpaceNut posted in another thread where we discussed alternative boosters. The only configuration for SLS Block 2 uses 6 AR-1 engines per booster. The standard Block 1B is listed, which uses Exploration Upper Stage (EUS) with 4 RL10C-2 engines. As well as Block 1B with the original upper stage proposed for Block 2, which uses a single J-2X engine. However, the thing I find concerning is the row for Block 1, without any qualification. NASA lists Block 1 as capable of lifting 70 metric tonnes to LEO, but this table lists 87.2 metric tonnes. Why?

Block 1B with 4xRL10C-2 EUS is listed able to lift 105 mT to LEO, or 34.4 mT to TMI. Block 2B with 2xF-1B LRB can also lift 105 mT to LEO, but the point is it can lift 44.0 mT to TMI. Block 2B with 6xAR-1 LRB can also lift 105 mT to LEO, but it can lift 44.2 mT to TMI. All this makes sense. Block 2B with no other qualification means a core stage with 4 engines, and an upper stage with 4xRL10C-2 engines.

Block 2 without modification means a core stage with 5 engines, an upper stage with 1xJ-2X engine, and a pair of advanced SRBs. That could lift 130 mT to LEO. The listing in this table is modified, it lists 6xAR-1 LRB instead of SRB, and able to lift 206 mT to LEO.

All that makes sense. But why is Block 1 without any modification or qualification not what NASA lists? Although this is actual sourced data, I raise this as a caution.

Last edited by RobertDyck (2018-09-02 14:54:46)

louis · 2018-09-02 15:58:08

I got a slightly different impression.

Wooster's presentation I found very reassuring. They seem focussed on all the right issues and they appear to have junked a lot of the more utopian aspects of Musk's vision, at least for the time being.

I doubt Space X will be publicising the names of individual engineers working on specific Mars projects. I expect Space X have security advice. Let's be honest, Space X is now a target for hostile action by at least two governments around the world, and maybe more, who have an interest in frustrating their ambitions.

I have always commented on the premise that Space X were fully aware of the challenges of landing on unprepared ground on Mars. My view is that there were, are and will be techniques that can be used to mitigate risks, and possibly the robot-cargo mission can prepare ground for a landing by humans.

I think Wooster made it clear that you didn't necessarily need a glacier-style concentrated ice source, that ice in regolith could be good enough.

The "at least 100 tonnes" reference was to "useful payload". I don't know if there is a definition for "useful payload". It's clearly to be differentiated from a theoretical payload that fills the whole of the payload bay. Does it include packaging? I expect so.

I think 6 years from now is a reasonable timeline when we reference the Apollo Mission that went from rudimentary models to to circumnaviation of the Moon in 6 years, and a Moon landing one year later. That was basically a pre-computer mission. Now we have computer controlled everything, CAD and so on. We also have a huge bank of knowledge about rocketry and Space X specifically have built up their own stock of know-how about how to put together successful rockets. Obviously it's a very steep ascent they are on...they could fall off at any point. But in principle I don't think it's an impossible timeline.

GW Johnson wrote:

What I saw in Wooster's talk regarding BFR/BFS: relative to the 2017 presentation, BFS has now 7 engines instead of 6. Payload was 150 tons, is now "greater than 100 tons". It would appear they have been running design analyses, and have found some needed changes.
Another thing I saw: recognition of the danger of rough-field landings in a tall vehicle. Used words like "rocks" and "pits". Talked about initial "challenging landings" (meaning high risk of overturn or crash). Talked about "prepared landing fields" for subsequent flights. I wonder who they will buy the electric-powered, rechargeable road grader and bulldozer from?
Another thing I saw: "focused on transportation" and "opportunities for others" to make this possible. I believe this refers to makers of a real propellant factory and real earth-moving equipment suitable for Mars, not just somebody's probes to confirm the buried glaciers they know they need (which also says they know that processing damp regolith for water is not a practical solution).
6 years from now (2024) is an awfully short timeline to do engineering development and harsh-condition field testing of the necessary equipment and machinery to make a manned trip to Mars in the BFR/BFS system feasible.
One last thing I noticed: he gave no names for anybody actually working on this stuff they will need.
GW

SpaceNut · 2018-09-02 17:39:36

Well what would be the tonnage for an Unmanned mission versus a manned mission for payload for the BFR archetecture?
We have run some numbers for power (solar/battery vs nuclear), Life support (iss style, greenhouse support), and of course food all of which eat into a payload...but these all hinge on time on the surface, crew count and what we will be doing.

GW Johnson · 2018-09-02 20:27:59

Just to let you all know, I posted 5 closely-related articles today (Sunday 9-2-18) on "exrocketman". These deal with fair comparisons of launchers for landing payloads directly on Mars. "Everybody lies" about throw weight capabilities. You need to know how to get past the hype and put some realism in it. I've got that for you.

One article is a cross-comparison of the 7 launcher configurations I examined. The other 4 articles examine in detail (1) the SLS (all 3 Block configurations), (2) Falcon-9 and -Heavy, including Falcon-Heavy with Red Dragon, (3) BFR/BFS as presented in 2017, and (4) Atlas-V 551 and Delta-IV Heavy. That leaves Ariane, Proton, Soyuz, and the other new concepts like Vulcan and New Glenn "for others to do". I just gave you a very fair way to do that, with examples.

Sources of data and specific methods are explicitly given in the 4 articles that characterize the rockets. The SLS article uses data I got from an on-line release of the NASA Mission Planner's Guide for SLS. None of these are estimated from the rocket equation and jigger factors. All use published figures for thrown mass onto TMI. Where data is assumed or uncertain, I explicitly point that out.

I only considered direct entry from a Hohmann min energy trajctory, with a retropropulsive landing from low altitude and 0.7 km/s speeds. That last was because all landed items exceeded 2 tons. Retropropulsion is flying operationally now (although only one company so far), while extendible or inflatable heat shields are only flying very experimentally. That's why I only considered retropropulsion.

The fair method I used takes the advertised thrown mass to TMI, deducts 10% for the custom payload adapter, and uses the remainder as the direct lander carrying the delivered useful payload. Its ignition mass is the same as its entry mass for ballistic coefficient. I presumed storable propellants, and the typical mass fraction for a one-way, one-stage, one-shot retropropulsive lander is 50%. In the case of BFR/BFS, the payload adapter and lander mass fractions do not apply. For Falcon-Heavy with Red Dragon, there is no payload adapter, but I used a 2 ton max payload in Red Dragon in order to have a tad more than the theoretical minimum for the touchdown delta-vee.

For those who don't already know, the "exrocketman" site is my blog at http://exrocketman.blogspot.com. There is a navigation tool on the left that works by date and title. Click on the year, then the month, then the title. Right now, these 5 articles are "up top", all dated 9-2-18, so you can just scroll down, unless you want to see the earlier reference articles. Enjoy.

GW

louis · 2018-09-03 06:28:56

Thanks for that work GW. Very helpful.

Seems to me that BFR/BFS stands out as quite different to all the others. It requires a lot of launches, to allow for LEO refuelling, but achieves a high tonnage to surface and at a much, much lower cost. Is that a fair summary? Did you assume one reusable BFR/BFS doing the 7 launches, or more than one...?

GW Johnson wrote:

Just to let you all know, I posted 5 closely-related articles today (Sunday 9-2-18) on "exrocketman". These deal with fair comparisons of launchers for landing payloads directly on Mars. "Everybody lies" about throw weight capabilities. You need to know how to get past the hype and put some realism in it. I've got that for you.
One article is a cross-comparison of the 7 launcher configurations I examined. The other 4 articles examine in detail (1) the SLS (all 3 Block configurations), (2) Falcon-9 and -Heavy, including Falcon-Heavy with Red Dragon, (3) BFR/BFS as presented in 2017, and (4) Atlas-V 551 and Delta-IV Heavy. That leaves Ariane, Proton, Soyuz, and the other new concepts like Vulcan and New Glenn "for others to do". I just gave you a very fair way to do that, with examples.
Sources of data and specific methods are explicitly given in the 4 articles that characterize the rockets. The SLS article uses data I got from an on-line release of the NASA Mission Planner's Guide for SLS. None of these are estimated from the rocket equation and jigger factors. All use published figures for thrown mass onto TMI. Where data is assumed or uncertain, I explicitly point that out.
I only considered direct entry from a Hohmann min energy trajctory, with a retropropulsive landing from low altitude and 0.7 km/s speeds. That last was because all landed items exceeded 2 tons. Retropropulsion is flying operationally now (although only one company so far), while extendible or inflatable heat shields are only flying very experimentally. That's why I only considered retropropulsion.
The fair method I used takes the advertised thrown mass to TMI, deducts 10% for the custom payload adapter, and uses the remainder as the direct lander carrying the delivered useful payload. Its ignition mass is the same as its entry mass for ballistic coefficient. I presumed storable propellants, and the typical mass fraction for a one-way, one-stage, one-shot retropropulsive lander is 50%. In the case of BFR/BFS, the payload adapter and lander mass fractions do not apply. For Falcon-Heavy with Red Dragon, there is no payload adapter, but I used a 2 ton max payload in Red Dragon in order to have a tad more than the theoretical minimum for the touchdown delta-vee.
For those who don't already know, the "exrocketman" site is my blog at http://exrocketman.blogspot.com. There is a navigation tool on the left that works by date and title. Click on the year, then the month, then the title. Right now, these 5 articles are "up top", all dated 9-2-18, so you can just scroll down, unless you want to see the earlier reference articles. Enjoy.
GW

Oldfart1939 · 2018-09-03 10:27:43

It seems that the BFS/BFR system is still in a state of flux. The engineering maxim is coming into play: "No concept survives contact with reality." According to GW's website, and the talk at the Mars Society meeting last week, BFS has undergone a payload reduction of some 30%. Perhaps the bold concept presented by Musk in 2017 is proving more difficult to actualize than even he considered? IMHO, if the progress on BFS/BFR slows significantly, I wonder whether a revised Falcon Heavy/ Red Dragon style mission could not be resurrected? I've argued here on this site for a parallel set of missions, one in which much of the support infrastructure is thrown to Mars. i.e. Mars GPS system, nuclear reactor, Sabatier plant, Moxie plant. Each landing could incorporate a radar transponder, making each subsequent mission incorporate better landing accuracy. The way things are currently structured at SpaceX, it's putting all the expensive eggs into one basket (conceptually). I do not regard my proposal as having two mutually exclusive systems, but ones which are interlocking and complimentary. This would allow a continuous advance instead of order of magnitude giant steps.

SpaceNut · 2018-09-03 11:27:11

The Space x BFR launcher is just that as he does not make any of the other things that make a mission for mars possible.

Well 7 launches at $100-200 million for 7 cargo of 100- 150 Mt each is $700 -1400 million at min of 700Mt to 1050 Mt is going to be a range at for 100Mt of $1 million for Mt to 2 Million and for the other end of the scale of 150 Mt of 670 thousand to 1.2 million but the question is can the second stage be recovered as that will make the costs rise if they are not?

Now if the tonnage is less the range of launch to payload will also rise in the same proportion....

louis · 2018-09-03 11:45:49

I really don't see Space X setting off down another path to Mars now. The BFR-BFS project puts the "expensive eggs" in six different craft.

Oldfart1939 wrote:

It seems that the BFS/BFR system is still in a state of flux. The engineering maxim is coming into play: "No concept survives contact with reality." According to GW's website, and the talk at the Mars Society meeting last week, BFS has undergone a payload reduction of some 30%. Perhaps the bold concept presented by Musk in 2017 is proving more difficult to actualize than even he considered? IMHO, if the progress on BFS/BFR slows significantly, I wonder whether a revised Falcon Heavy/ Red Dragon style mission could not be resurrected? I've argued here on this site for a parallel set of missions, one in which much of the support infrastructure is thrown to Mars. i.e. Mars GPS system, nuclear reactor, Sabatier plant, Moxie plant. Each landing could incorporate a radar transponder, making each subsequent mission incorporate better landing accuracy. The way things are currently structured at SpaceX, it's putting all the expensive eggs into one basket (conceptually). I do not regard my proposal as having two mutually exclusive systems, but ones which are interlocking and complimentary. This would allow a continuous advance instead of order of magnitude giant steps.

EdwardHeisler · 2018-09-03 12:08:57

Hi all,

This past week, we have started our official channel on Reddit (called a subreddit) and it has already added over 500 subscribers since we started it on Thursday.

Please join us there and follow the latest on the Exploration of Mars and the growth of the Mars Society. We will be posting all of our convention videos there so they are easy to cross-post on other subreddits, and also to comment on.

https://www.reddit.com/r/MarsSociety/

Don't forget to add some "Flair" to your Reddit user in the form of "Mars Society Member", "Chapter Member", or even "Founding Member" if you are one of those (like me) who joined the organization in our founding year 1998.

Thanks,
-James

James Burk | Webmaster & IT Director | The Mars Society

SpaceNut · 2018-09-03 12:12:48

All with a variant of what is the second stage adapted for its purpose.

Here is where we put the description Kbd512's Space BFR variants

louis · 2018-09-03 17:49:27

An interesting video from the prolific Martian Wolf:

https://www.youtube.com/watch?v=JFmqPoA_CNo

I hadn't realised that Melas Chasma in Valles Mariensis won a vote on the 45 locations presented to NASA gaining 9 votes from 90 plus people voting at the conference.

At 06:50 Wolf mentions that this proposal had a novel way of solving the water problem by heating up polyhydride sulphates to produce water, which would deliver 50% water from the raw material mass. That's an interesting figure. Not sure if that's been discussed in detail.

SpaceNut · 2018-09-03 20:20:41

A little investigation into the chemical properties:

https://en.wikipedia.org/wiki/Polyhydride
high hydrogen stoichiometry FeH5, LiH6, and LiH7 with lots of other combinations.

https://en.wikipedia.org/wiki/Sulfate
Not so many types of SO4

I do know that hydrides were identified by the Mars Orbiter as the Lyman-Alpha Photometer (LAP) – a photometer that measures the relative abundance of deuterium and hydrogen from Lyman-alpha emissions in the upper atmosphere. Measuring the deuterium/hydrogen ratio will allow an estimation of the amount of water loss to outer space.

So while we might not find the super hydrides and loaded sulfides we will be able to do this heating process even with the lower levels of each to produce water.

louis · 2018-09-04 03:10:45

Thanks for that Spacenut, very helpful.

Of course the problem with planning a Mars Mission, especially Mission One,is that you have to work to the "worst case" scenario, rather than what you hope the data is telling you - since you have no fall back. In this case the data might suggest you can get 50% water by mass of regolith but, you are probably going to have to plan for 10%.

SpaceNut wrote:

A little investigation into the chemical properties:
https://en.wikipedia.org/wiki/Polyhydride
high hydrogen stoichiometry FeH5, LiH6, and LiH7 with lots of other combinations.
https://en.wikipedia.org/wiki/Sulfate
Not so many types of SO4
I do know that hydrides were identified by the Mars Orbiter as the Lyman-Alpha Photometer (LAP) – a photometer that measures the relative abundance of deuterium and hydrogen from Lyman-alpha emissions in the upper atmosphere. Measuring the deuterium/hydrogen ratio will allow an estimation of the amount of water loss to outer space.
So while we might not find the super hydrides and loaded sulfides we will be able to do this heating process even with the lower levels of each to produce water.

SpaceNut · 2018-09-04 18:21:37

We know the science, the require identification equipment and what it will take to process what we want from insitu materials but what we do not have defined is the quantity, quality of specific insitu items that we can gain from mars. They only 1 that we have confirmed for quality and quantity is in mars atmosphere.

louis · 2018-09-04 19:46:47

Indeed - which is one of the reasons I feel we shouldn't dismiss the atmospheric option...I am not sure how it synchs into Mars Missions because in the N Hemisphere it (water vapour) is only available in quantity in the summer months (ie maybe something like 8 or more earth months). So the issue presumably is how that synchs in with your Mars landing. I recall a NASA study which produces 3 kgs of water per sol from an 800 kg rig. Maybe you'd be looking at 2000 kgs per sol from a 600 tonne rig on that basis. But I am sure economies of scale would apply so the rig would mass in at far less than 400 tonnes. If you could get it down to say 150 tonnes, that might be a viable option.

SpaceNut wrote:

We know the science, the require identification equipment and what it will take to process what we want from insitu materials but what we do not have defined is the quantity, quality of specific insitu items that we can gain from mars. They only 1 that we have confirmed for quality and quantity is in mars atmosphere.

SpaceNut · 2018-09-04 20:50:11

An important part of any system that takes in a very dusty mars air at times....
Martian Atmospheric Dust Mitigation for ISRU Intakes via Electrostatic Precipitation

Mars Water In-Situ Resource Utilization (ISRU) Planning (M-WIP) Study
I see this one has 4x kilowatt power plants in one of the pages....

Of course this what we need to even get fuel creation from the atmosphere and that is because we have landed and are ready to go home in time.
Mars Ascent Vehicle Design Considerations

elderflower · 2018-09-05 04:32:01

If we were to park storable return fuel in LMO, we would only need to get off the planet and rendezvous with it. CO/LOX could do that, and be a useable fuel for surface equipment, despite its relatively poor Isp. We would need a new ascent engine, though.
Alternatively we could bring a storable fuel to the surface and generate LOX there, discarding surplus CO. The majority of rocket propellant, by mass, is the oxidiser. Neither of these involves extracting large quantities of water from any place on Mars. We would have to clean up atmospheric gases and split CO2 into O2 and CO, so we will still need a large power source.
Argon is not so easy to separate from LOX as nitrogen and CO, so the LOX product from this process may contain some of it, depending on the gas separation technology employed. This will slightly reduce the available Isp.

elderflower · 2018-09-05 04:39:44

Isp using Mars hydrogen (whether it is in the form of LH2, CH4, NH3 or any other compound) will be reduced relative to earth hydrogen due to the increased abundance of Deuterium on Mars. Molecular weight of Mars hydrogen is about 1.3. Not 1.0! I would expect electrolysis to favour production of the lighter isotope, but I've no idea to what degree, if it does.

Last edited by elderflower (2018-09-05 04:41:12)

spacetechsforum · 2018-09-05 05:44:48

As for Elderflower solutions, bringing any amount of fuel to Mars orbit or even worse as you need to land it - to surface, scales costs of the missions in very bad way. The choices are no return or make fuel on site for the landed cargo to make the way home. This is why the return trips should be reduced to minimum.

Last edited by spacetechsforum (2018-09-05 05:45:21)

louis · 2018-09-05 06:39:23

This isn't the case. Have a look at GW's analysis, which shows that the BFR is much the most cost effective approach and the BFR takes a a lot of fuel to Mars orbit.

http://exrocketman.blogspot.com/2018/09 … eness.html

I think you are over-focused on fuel usage as opposed to overall cost per tonne. If you want anything more than a flags and fooprints mission or a very slow build up of a colony, you need something like the BFR which can deliver large tonnages to the surface at a low cost per tonne.

spacetechsforum wrote:

As for Elderflower solutions, bringing any amount of fuel to Mars orbit or even worse as you need to land it - to surface, scales costs of the missions in very bad way. The choices are no return or make fuel on site for the landed cargo to make the way home. This is why the return trips should be reduced to minimum.

GW Johnson · 2018-09-05 09:26:46

Louis:

I've said multiple times that BFS cannot stop in, or deliver anything to, Mars orbit. So your statement just above about what I claimed is simply incorrect. Payload delivery is ONLY to the surface.

When it arrives at Mars, it still has on board propellants that are only sufficient to provide about 1 km/s delta-vee, just barely enough to cover the final retropropulsive touchdown. The delta-vee required to enter low Mars orbit, instead of landing directly from the interplanetary trajectory, falls in the 1.6 to 2 km/s range. That's just a no go! The same is true of not being able to stop in Earth orbit on its return from Mars, and by a far larger margin (delta-vee in the 4 km/s class).

And I explicitly (!!!) pointed that out, in the related document that details where the comparison numbers come from. That is reference 3 in the comparison article you provided the link to, and is dated 9-2-18 and titled "Future Spacex Rockets", on my "exrocketman" site. I make that statement in the first paragraph of the subsection titled "Missions to Mars".

The 9-2-18 articles also explicitly acknowledge the differences between the configurations I analyzed from the 2017 presentation still posted on Spacex's website, and what Mr. Wooster said at the recent Mars Society 2018 convention. The two most obvious were the change to "over 100 tons" payload in 2018 versus "150 tons" payload in 2017, and "7 engines" in the BFS second stage in 2018, versus "6 engines" in the 2017 presentation.

Clearly, this still-only-on-paper design is still in flux, and any evaluations by any of us are approximate and subject to changes they make in the design. My numbers (also explicitlity stated) are based on the 2017 presentation.

As for how much development has been done, yes, the Raptor engines have had a lot of testing, and that continues. One big tank of the wrong diameter has been made and pressure-tested to destruction. And one big composite wrapping mandrel has been installed in California. Everything else is paper. Paper only!

And I would point out that the entry and landing sequence is different for the BFS second stage from its BFR first stage, which is like the Falcon cores. The second stage BFS entry sequence will require all-new test demonstrations. It does not go tail-first for retropropulsion until after the hypersonics are over. The first stage and Falcon cores go tail first while still exoatmospheric before hypersonic entry begins. That is an entirely different item, no matter how you look at it.

The BFS will face the risk of dead-broadside airloads that always crush and break up flight vehicles, unless this is done very carefully on a flight dynamic transient resembling an airshow stunt (tail slide). Plus, it's not the same air loads on Earth as on Mars. It's like this has to be designed and tested twice, at two wildly-different conditions.

By the way, my full reverse engineering of the 2017 presentation reveal is explicitly documented on the "exrocketman" site as an article dated 4-17-18 and titled "Reverse-Engineering the 2017 Version of the Spacex BFR". That is listed as ref 1 in the "Future Spacex Rockets" article. The April 2018 reverse-engineering article is quite long, and goes through the performance estimates in great and realistic detail. The updates I made to it take on such things as landing stability, and other similar issues.

GW

Last edited by GW Johnson (2018-09-05 09:45:48)

spacetechsforum · 2018-09-05 09:52:40

Louis, BFR does not bring any fuel to any place. It has only necessary load to land on Mars and is unable to proceed if stopped in orbit. I did look at those calculations very closely and they do clearly show that BFR is not the most efficient solution when the cargo is not live. Furthermore BFR is so very risk that i do not know if i would board this ship even for free.

Firstly they must land immediately when they arrive at Mars so if there is a dust storm or BFR is in state of emergency the crew is dead. Next the landing... They need to do quite dangerous reentry procedure then fire the engines in atmospheric flight(!) and stabilize very heavy load that needs to hit the specific flat surface. I am assuming that the whole in situ operation on the Mars surface will be conducted before first human mission so the resources to get back will be available but still i do not think that after refilling the rocket crew will be able to do the maintenance, so the way back will look exactly the same except the rocket will be vary.
Now, the flight is hard but can be done, yet we can expect the same situation as with the first spacex rockets, when first attempts failed.

GW Johnson · 2018-09-05 10:01:55

A follow-up: I noticed that the 9-2-18 article "SLS Capabilities on Mars" that I posted on "exrocketman" has drawn more readership than the rest. Just to be clear, while SLS Block 2 can fling 39-44 tons onto a trajectory to Mars, a whole lot less tonnage can actually be landed.

First you must deduct a custom payload adapter from the throw weight. This thing is a significant percentage, because it must support the rest of the payload against launch gees and vibration, which are not trivial, by any means! My guess of 10% of the thrown weight for this adapter is as good as anybody else's.

The least costly outcome is the direct-entry landing, for which only the touchdown propulsion is required. The delta-vee to stop in low Mars orbit will cost you much more mass ratio than any conceivable direct lander. Based on realistic inerts for a retropropulsive single stage one-shot lander, and storable propellant Isp, I determined a realistic payload fraction of 50% of the lander weight at ignition, which is also its weight at entry. This is way beyond the mass limit for being able to use chutes, so it must be retropropulsive from end of hypersonics at 0.7 km/s, only about 5 km up, and angled somewhere in the vicinity of 45 degrees downward. It'll require 3-4 gee deceleration to keep from crashing.

So, that's 50% of only 90% (in effect 45%) of the SLS throw weight that you can consider useful payload to the surface of Mars. THAT is where my 17.5-19.8 metric tons of useful landed payload came from.

That same 45% of theoretical Mars throw weight applies to ANY vehicle that sends a payload to mars inside a payload shroud. All such MUST deduct the adapter and the entry/touchdown lander from throw weight. Don't worry about the shroud weight itself, it is not included in the throw weight anyway.

The problem is far worse for delivering payload to low Mars orbit, because the propulsion delta-vee is around twice as high. It affects mass ratio exponentially.

GW

Last edited by GW Johnson (2018-09-05 10:16:15)

Oldfart1939 · 2018-09-05 10:54:13

spacetechsforum wrote:

Firstly they must land immediately when they arrive at Mars so if there is a dust storm or BFR is in state of emergency the crew is dead. Next the landing... They need to do quite dangerous reentry procedure then fire the engines in atmospheric flight(!) and stabilize very heavy load that needs to hit the specific flat surface. I am assuming that the whole in situ operation on the Mars surface will be conducted before first human mission so the resources to get back will be available but still i do not think that after refilling the rocket crew will be able to do the maintenance, so the way back will look exactly the same except the rocket will be vary.
Now, the flight is hard but can be done, yet we can expect the same situation as with the first spacex rockets, when first attempts failed.

Please see my post #57, this thread.

There is an almost mandatory requirement for some preliminary landings in order to site several RF landing beacons, such as radar transponders, capable of triangulating the desired touchdown site. Then it matters not having zero-zero visibility conditions. The final descent and landing can be accurately accomplished by radar. This is how the big airliners are landed at major airports. Virtually ALL are using GPS-WAAS interfaced autolanding systems. Even in broad daylight and VFR conditions.

My fundamental argument is we need as many as 6 or 7 Falcon Heavy missions to establish the necessary support for BFS/BFR landings. Need to have several Mars GPS satellites in orbit. Minimum of 3, prefer to see 6. Have landed and in place several radar transponders, preferably incorporated as secondary payloads to food and construction equipment as well as a self-contained Nuke power unit. The downside to my proposal is it takes time and money to do these things, but the upside is the Falcon Heavy is already/still is in production. If the booster recovery of the central FH core can routinely be accomplished, and side boosters can be reused, this cost could be reasonable. It amounts to expendable second stages and Dragon-style landers plus payloads.

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Re: Where are we now?

#73 2018-09-05 09:52:40

Re: Where are we now?

#74 2018-09-05 10:01:55

Re: Where are we now?

#75 2018-09-05 10:54:13

Re: Where are we now?

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