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GW, that was also my thinking on aluminium use (still not the main reason to discard), but data provided by kbd512 suggest otherwise:
Here is the exact link:
http://www.dtic.mil/dtic/tr/fulltext/u2/429244.pdf
If I understand correctly tested aluminium alloy in low temperature seems to be doing quite well, The titanium suggestion is also from this paper's summary.
All in all I am still unable to prove that "LH2 fuel depot" is a wrong idea. (I have no reference for active cooling system design, but this is what probably kills the project in NASA analysis).
Oldfart1939, KISS solutions loose with more efficient ones in long run most of the time. My favourite example is DC electricity transfer (Edison) vs DC to AC conversion and AC transfer (Tesla). On the other hand first attempts with not KISS solutions fail miserably (like first reusable SpaceX rockets?)
Also - I agree that the LH2/LOX is not the best choice in this mission, but it makes NASA do some actually useful designs and research in the future.
Kdb512, does you post includes irony? It is hard to tell, since I am not a native speaker, but the sources provided suggest this. I was not aware that in low temperatures the stainless steel performs better than the aluminium - still the articles i found earlier suggest that the main reason for using stainless steel or other exotic materials to make long term LOX tanks is the corrosion. The information in last link point to titanium as best solution but I think that it is good that the proposed stainless steal is one of the most feasible materials.
Belter, update your information on thermal radiation.
GW, i think i made it very clear that I am talking about the tank designed with single purpose of long time propellant storage in space. Centaur is designed to reach orbit and do the burn and be done - the propellant does not need do be stored for very long so I think the boil-off issue is not addressed at all in this design, as long as the boil-off is low.
Still, I do not know on what assumptions you have based your opinion.
The insulation used in space is different that the one used in atmosphere. If you are filling the tank in space you can omit the insulation used on earth, so you are not thickening anything. The data i got on the subject suggest that very good insulation that is used in space weights 2 kg/m2 (only blankets, not walls - so you need to add the standard inner wall and very thin outer wall).
The bigger the tank the better ratio you have for useful propellant/heat leak since volume grows with ^3 and external surface with ^2. From NASA documents gathered i know that heat pump that can remove heat from LH2 tank has COP of 0.008 - but they do not specify mass of the installation, so i was wondering if you did the math with the machinery. Apparently, not.
Kbd512 - stainless steel is used not to insulate the tank, but to prevent corrosion issue.
GW, I do not agree at all. I will respond to your post by paragraph, since it mixes a lot of topics - most that do not apply to the given situation.
1. We are talking zero boil-off tank. The amount of energy removed form tank with active cooling system is equal to heat leak - pressure and temperature stays constant (and hopefully uniform). ZERO boil-off tank - no fuel loss EVER.
2. We are talking tank in space - convective does not apply. Rest addressed below.
3. This is clearly not the design for tank. OK- we have two walls and vacuum (provided by default in space), but material does not matter - the MLI blankets are placed in-between. Heat leak from support structures does not matter - in space the tank is weightless, so most of them can be disconnected after liftoff - a few can be made of material with very bad thermal conductivity and actively cooled on the hot side.
4. Does no apply. MLI in-between.
5. This is totally wrong. The surface on the sun side needs high reflectivity. On The dark side the high emissivity in IR is needed. All else surfaces need just low emissivity - there is no sunlight. Any material can be covered with proper layer to adjust for required properties, but actually only the sunny side needs it.
6. This rule applies to any tank is space - we have those already.
7. Does not apply. At all times the outer wall has equal pressure on both sides. (In atmosphere tank is empty - no insulation needed - no vacuum needed, and you vent it as you go up).
8. Material - does not apply. How big is the cryocoler? - no reference.
9. I cannot agree with this opinion - bad analysis based on wrong assumptions and no calculations.
Rest - does not apply.
I do not know how you actually imagine this tank, but clearly - a short description would be welcome, since is hard to connect the problems you specified with actual design. The designs i checked suggest construction as follows:
1 . Outer wall - aluminium with reflective layer on sunny side, emmisive layer on dark side.Layer Inside - does not care.
2. MLI blankets in vented space.
3. Very few supports made of non-conductive material.
4. Inner wall - material suitable for given substance. Emissivity - does not care.
5. Active cooling system powered with solar panels.
The LH2 is not best option for this lander - agreed, still I am happy they are doing it with LH2/LOX engine. This may force NASA to design big, independent, actively cooled, permanent zero-boil tank that will be placed in some distant from the station, as a fuelling platform. A small dv needed to get to the tank assures that the tank - even with moving parts machinery - may be easy to maintain.
Such device would be of great value for future mining operations in solar system, since you can store in-situ produced propellant in orbit of practically every body. This technology is also useful for a long-range ships, which cannot refuel at destination.
GW - you have stated: "Dewar with a cryocooler is going to be heavy."
Do you have any reference design or did you do the calculations?
Louis, if it ain't gonna happen then the whole talk about the Mars colony is useless. There is no way to do more than flag planting and maybe a single small outpost for few scientists with BFR.
The reason is very simple. BFR needs fuel on Mars. The more stuff you want to move the more fuel you must produce to return. The bigger the refinery producing fuel, the more materials and more maintenance is required (more spare parts, more energy, more people - rotation required, more food, oxygen - production required = more spare parts, more materials etc). You may argue that ships may stay on Mars, but that is defeat to the point of BFR.
If you need to allocate all of the resources to produce fuel, colony on Mars ain't gonna happen. I mean - what is the point if it exist only to produce fuel? Even if you solve the unloading subject you cannot overcome the above problem - this is a design flaw and so you need to alter the design.
Belter, I had the same idea for one-way cargo transportation to Mars and I did some calculations and simple model.
Here are the calculations for reference:
http://spacetechs.ovh/myfiles/Calculati … 9_2018.ods
During this work I realised that there is a lot better way to deliver cargo to Mars surface (I did not update the calculations yet). It goes like this:
1. There is a lander with cargo in LEO.
2. You attach big tank with engines and control unit.
3. You do the burn into Hoffman.
4. When you reach the Mars you are on direct reentry trajectory (some small correction may be needed but that is not a problem) .
5. You detach the lander - the cargo stays on reentry trajectory and lands.
6. Meanwhile the tank-engine section do a small burn that lifts the periphrasis above the atmosphere - now you are on hyperbolic trajectory (ref. Mars).
7. You fly by the Mars - you are on elliptical Sun orbit that crosses Mars and Earth orbits - another burn is needed to adjust timing to hit Earth atmosphere and try aerobrake.
I think that this is the best setup for dead cargo delivery to Mars - no matter the engine type.
The biggest advantage of this approach is that there is a very few critical failure points to lose cargo. You can do a full maintenance and check in orbit - then you do a single (or staged) burn into Hoffman (first potential failure), detach the cargo (can this fail? - not likely i think) and finally reentry (second possible failure point). If this succeeds (and i cannot image a solution that does not have those failure points) the cargo is in place, so even if the very hard maneuver of aerobrake fails you loose only the empty ship and the colony is still successfully resupplied.
Sorry tahanson, the ice platform idea seems naive. Wont it melt from engine work? How would you protect it?
Seriously i cannot image any way to evaluate landing options. I was doing lately the calculations linked above and I do not know the subject very well. Are yours opinions based on any data? Do we have any ground parameters? Cross-section maybe? - (this is actually quite important) Also - keep in mind that if we are talking concrete and dynamic load - reinforcement with steel bars is also needed.
I think, given proper machines are available, the 1-way BFRs can be adapted to serve colony. I did not do any analysis on this, so I do not know if it is possible and feasible, but doing it i would start with this option:
Place BFR horizontally (remember - we are working in lower gravity and with empty tanks).
Remove engines - I do not see any use for this part.
Remove tanks from shell - use them as storage for water (fuel tank) and oxygen (LOX tank). They do not need to stand vertically.
Adapt the empty shell as habitat.
I also do not see any need to build the whole colony at once. For the first crews not connected habitats may suffice (as those will be mostly science missions), and as the time goes on and the needs grow a better shelters may be delivered.
"Follow the money" - i have no idea what you meant by that but that was a good advice. I searched internet for the grant number and here it is, the study in question:
https://www.ncbi.nlm.nih.gov/pmc/articles/PMC3903607/
It seems legit - i am not a biologist to dispute it and prove the claim on radiation effects is wrong. On the other hand I do not know how did NASA obtain data on REM values away from earth? I mean this value depends on the radiation type so you cannot just scale x-ray effects.
This does not mean we cannot fly to Mars, just adequate protection is needed. There seems to be a lot of support for protecting craft with water shield. I think it is a good idea, since you need to take the water anyway, but is the irradiated water suitable for drinking?
GW I think you are missing the point here.
Parking BFR in orbit is proposed as faster option for safe Mars mission.
Firstly you do not send crew on that flight. You pack just enough supplies for return trip and tank the rocket full, then you send it to Mars.
Secondly, since TPS can survive land, in this scenario you do not need a lot of aerobreak passes. You do a single heavy aerobrake (just a little weaker as the max to not crash with worst weather). If the weather is bad (a big density) you do not need anything more, but most probably you will need a small burn to adjust the orbit (or another very weak pass). Now, as you did not use the fuel to land and you did not fly with maximum payload (or part of the payload is just a tank with more fuel), you may end up having a very safe return option for the astronauts sent with next flight, as long as they have alternate way to go to LMO.
Imagine that the BFR with crew will damage the heat shield during a storm on Mars or we wont be able to provide enough fuel for return trip (inefficient ISRU, tank damaged, etc.). Now you cannot go back - you will burn on Earth or you wont make it. With BFR in orbit the crew may swap the ship and safely return home.
Now, at the first glance this plan looks good, but the numbers indicate that the amount of fuel left is not sufficient for return trip, so the SpaceX would need to send another rocket just to pump the fuel from both into one, which skyrockets the costs, and so this idea is useless. This is exactly why I have no idea why I did write the post...
BTW. Is there any progress on the GCR topic? The GW's reply on REM doses looks optimistic, but i think the researchers knew what they were doing. Maybe the radiation further away from Earth is somehow different and has larger impact on health? (like heavier particles are destroyed with very weak magnetic field so the ISS is safe but a spacecraft's crew would not last). The look at the study would clear the doubts but I am unable to find the link to that paper (maybe it is available only in US?).
Louis, I think they are going one step at a time - first design the rocket that can go to Mars, then go loud "we did it" and gather founds to design the shelters. The first crew will probably go to Mars just to plant the flag (from ignorant's perspective to be precise - crew will have their hands full reporting all problems, checking all devices and gathering a lot of useful data on the trip to Mars and on the Mars surface). Also, as I stated before, I think the automatic machinery (remotely controlled) will prepare and refuel the "goback" rocket prior to crew sending, so they will almost immediately go back. This also means that the "arrival rocket"'s tanks can be vented after touchdown for safety.
SpaceNut, why would you need two systems? I did not find any valid reason that is preventing the use of system working without gravity in 1G.
Reverse is more problematic. There is a problem of settling fluids on the bottom of the tanks in space, but solving that with small accelerations allows use of normal gravitational devices.
GW Johnson, I think you are wrong on that. You do not need to do a stop burn at Mars. BFR is designed with very good TPS and it can in theory perform aerocapture (the difference from reentry with landing seems to be only the reentry angle). There is around 1.6km/s of dv between crash and bouncing off into the void and the last visualisations show the BFR will be capable of some steering in atmosphere, but still doing this you will probably end up with not very precise long elliptical orbit. The problem is that the rocket will be stuck there since there are no means to refuel the craft.
Also I doubt there exist possibility to land BFR on the exact spot. Mass is big and the atmosphere density on Mars may be too diverse to calculate proper reentry parameters to stop at right moment. Even if this is manageable I think just for safety the landing sites will be far apart (Like wing failure may direct rocket straight at colony).
Louis, the BFR is designed for take-off from earth (so it must be parked on earth vertically) so we can safely assume that all internal components will also work well in lower gravity (when parked on Mars vertically). Also i think the layout will be designed for vertical use (floor, walls, stairs) since in space the orientation does not matter.
For rocks... Well, the structure must withstand it either way - you need to fly the ship back.
Also, I personally think the people will fly to Mars when there will be a fully tanked, parked, well pictured from space and ready to go back BFR. All task needed will be completed by automatic machinery beforehand.
The shallow articles like this one are exactly why I attempted to design my own version of "spaceship". I trust that you will correct me in this subject if I am wrong, since my understanding is still limited.
Indeed, the GCR is problematic, but we can solve the problem quite simply by limiting exposure via crew rotation on space stations or reasonable space travel time to Mars.
Currently the sun flares are bigger obstacle. Single short travel to Mars and back might be unlucky and get struck by it.The chances are low, so this is the risk we are willing to take. For the space stations with the rotational crews we are 100% percent sure that at some point we have dead people.
Shame no one knows a guy in NASA to get some tips on their solution to this problem with their new plan for the Moon space station.
EDIT: Ok, so i found better article on the subject (closer to source):
https://gumc.georgetown.edu/news/Animal … Astronauts
Still I am unable to find the original study to evaluate data. Any of you Americans had a better luck? The problem seems serious, since the author claims an "equivalent of a months-long period" radiation was delivered and this is exactly the time of BFR travel one way.
My personal opinion on the state of different subject related to the space exploration:
1. Lifting cargo into orbit - with current technology we can actually do it quite "simply" - some efficiency boost would be welcome (like BFR - more payload, less costs) but I do not think like there is a strong need to pursue better designs in this field.
2. Space travel and payload transportation - we are capable of constructing "spacecrafts" that can carry all necessary resources to achieve all other tasks in this list. It is very expensive, and very troublesome but the work is progressing (BFR). Still i think that a lot more effort should be directed into designing and creating a better and more efficient solutions. Currently, this subject needs a strong advertising, promotion and consequently investors. Also better society awareness of the benefits of the space exploration could bring more funds from the governments (like USA - military/NASA funding balance)
3. Space stations (anywhere) - critical problem: sun flares - not solved.
4. Mars colony - in theory - technology allows but solutions not yet developed. No agreement on "how we should do it" - underground?
problems: habitats - not solved, Mars-Earth transportation (and long term colony resupply) - in progress (BFR), mining insitu resources - not solved, creating new constructions from available materials - not solved. There is a lot more problems here, but since there are many proposed approaches, each one creates new challenges. No point on dwelling here until there is no agreement on the pursued solution.
5. Mun base - now this is interesting. A very low interest in pursuing this goal is somehow difficult to understand. Challenges similar to the ones in Mars Colony, but no fame from this achievement. Still much easier and cheaper, provides a lot of engeeniering experience, may be supplied from earth - and a magnitudes better for space turism than the Mars. Quick arrival and return, plus lower gravity (a lot of fun). Transmitting soccer match (or any sport event) played in those conditions would surly be a something to look forward in TV. Health issues due to low gravity may require workers rotation but it would be much easier to develop the solution to the whole "gravity" problem in test ground in place.
Thanks for the link. I am moving my question with numbers to that topic (somehow i missed it when searching earlier).
As for the electric propulsion, I cannot find any up to date article, so does anyone know if they even completed this 100 hour constant work test?
Also even if they did, to generate 400N of thrust one would need a 200 engines and 20MW reactor (one engine seem to generate 2N of thrust with 100kW power). The reactor design seems to be pursued, so this part may be doable. Still there is problem of heat dissipation - my best guess is that just the radiators array would weight around 50t. (btw. why people write "metric tons" and not just "t"? ).
The lower thrust does not only donates longer travel time. Some maneuvers must be completed within time frame (like hyperbolic to eclipse orbit change near Mars), so to weak engine is not an option.
I do not think that electric propulsion is of any use in cargo transportation. I cannot give a specific figure, but last time i checked the amount of electrical power needed to generate enough thrust and dv to push heavy cargo (250t+ for payload and ship) to Mars cannot be provided in any feasible way.
My choice is LH2/LOX as it gives the best Isp and I am not a specialist to evaluate the cons of this propellant, although I am currently trying to do so, which brings me to the topic.
Although many NASA publications I found suggest the technology needed for long time LH2 storage seems to be almost ready, I find myself unable to scale provided systems to suit the needs for Mars mission. My first though on this subject was: "Is that not simple? Why not to put a big refrigerator into ship and we are done". Now, as I do not consider NASA scientist to be fools, my understanding is that there must be a serious problem with this approach, yet the thing must be very obvious, since I cannot find any good source that can explain it. I hope someone in this forum will enlighten me.
I did some calculations to comprehend the problem, below the numbers in case i did mistake and so my perspective is flawed.
Model - external shell (const temperature calculated from heat equilibrium - Th=262.15K ), insulation (2xMLI blankets, 22 layers each - E=0.001 ), LH2 tank(Tc=18K). Stefan-Boltzman cost: sigma=5.67E-08
LH2 tank: volume: 464.218m3, surface area: A=353.320m2
Heat Leak: sigma*E*(Th^4 - Tc^4)*A = 94.62W
To remove this amount of heat from tank using Helium as working medium would require a heat pump capable of generating dT=282 with COP=0,03.
That means that we need a little more than 3kW electric power (this much can be provided with 2x18m2 solar panels). If i am correct the 8m2 radiator working at 300K would remove 3,3kW of heat generated by system summed with heat from tank.
The heat pump proposed is 100% theory. The subject of efficiency of cryogenic heat pumps capable of working in space does not seem to be popular on the internet.
GW:
Well, this computer model is mostly the equations from your blog encoded with c++ instead of spreedsheet for better presentation. I written it myself so this may be considered more by hand calculations that proper FEM.
About the nose radius... I do not think that this value can be taken directly from the above pdf. Here at the page 22 of this pdf:
https://tfaws.nasa.gov/TFAWS12/Proceedi … Course.pdf
The heat load seem to differ a lot with just a minor shape change.
The thing about the speed at atmospheric entry is very worrying. Is there no way to reduce the relative Mars-object speed at entry to 5000 m/s? Above this speed the aerocapture is no go for the heavy loads with my current BC, so some burn will be required to reduce the speed. Also the whole analysis as for now gives this conclusion: The minimal heat protection for any valid design with heavy loads must use the multi-use heat shield equivalent to 4,2mm PICA layer (this seems to be not reusable protection).
For now the assumption is that covering the whole front area of spacecraft with reinforced carbon-carbon is sufficient heat protection to deliver heavy loads to Mars (ballistic coefficient below 1000 kg/m2 - speed 5300 m/s, periphrasis 20-22 km).
GW Jonhnos:
You are right about the atmospheric model. I found out it gives a false values for high altitudes some time ago, so I corrected it with some reference data found on internet. Now it gives similar values to the ones posted on your blog. It is not exactly the solution from your site, but the difference is very small and i think drag calculations for objects in my simulation are quite good.
Here is the sheet for my model:
http://spacetechs.ovh/myfiles/MarsAtm-SimpleModel.ods
Now for the data provided, I asked for the values specified exactly in this pdf:
http://www.ssdl.gatech.edu/sites/defaul … 7-0164.pdf
I know how the ballistic coefficient alters the trajectory of the object and how to change the relations so they would suit my needs. The problem is - when I apply those values to the object, the drag force and so the trajectory are good with the data in that pdf, but to get the same heat graph ploted using: q=k*sqrt( ro(alt) / Rn ) * v3 I need to input Rn = 8m (!).
Edit: Sorry! I hasted the post and divided the value by 1000, not 10000 when (W/m2 to W/cm2). I run the program again - it turns out to get the value in rage of 100W/cm2 the Rn is 0.12m - the screenshots not updated - the graph just scales).
I uploaded the screenshot for reference. The first one is the pdf example. The second is an objects with m =350t, A=64m2, Rn=8m. The last are both runs in single graph. Object data are shown in the bottom right window, the graphs as in the pdf are on the bottom. To get the values on graphs there are 2 grayed out boxes under each graph - first is the max value, second the min (sorry - the software is still in development).
For other important data: the Apogee and Perigee are given after the aerobrake! (Perigee before the atmosphere entry is in rage 21km-23km but since the objects slows downs before and after the lowest point it changes).
First:
http://spacetechs.ovh/myfiles/Screensho … -30-31.png
Second:
http://spacetechs.ovh/myfiles/Screensho … -37-28.png
Last:
http://spacetechs.ovh/myfiles/Screensho … -39-08.png
You may have noticed that for the heavy object the v0 = 5000m/s, not 6000 like in the pdf.
Such a heavy object with v=6000m/s cannot obtain enough drag to lock the trajectory into ellipse to start the aerobrakes phase. Still, the v0=5000m/s at 150km is the hyperbolic trajectory, so i think that this is a realistic assumption on Mars approach (still i did not check it, so this may be wrong - how close to the parabola can we get when approaching Mars?)
So maybe someone has an e-mail to NASA and can send them a message with questions regarding the radiation problem solution?
I mean, if they announced that they are going to do this, although the solution may still be a theory, they are going to pursue the design. If they succeed, existing radiation protection system for non-earth construction would probably alter all mission plans for any space expansion, since all would include it.
Bleter, You design is interesting but are you sure that there exist right now a 3D printing technology that can do what you require? If i understand correctly you are proposing a 3D printing in space. I mean even best big scale 3D construction done on earth in atmosphere with oxygen (which i quite important when we are talking about some 3D printing materials) are not very precise. There is also this gravity issue and pressure issue. Also you do not cover at all the problems of heat equilibrium and space radiation.
GW Johnson, SpaceNut - thank you for the help provided.
I corrected the error in the energy equation used for estimating power needed for aerobrake. I know this is not a 1:1 kinetic to craft thermal conversion but it gives a good perspective for comparing light or heavy spacecraft aerobrake vs heat protection in landing or aerocapture.
GW J. I dismissed your proposition not just for the off charts values that it gave (it was probably for my wrong unit conversion), but mostly for i was unable to locate the stagnation point in MRO craft. Now, as i did more reading on the subject and examined the data provided by SpaceNut i incorporated this equation in my software written to do the calculations. The numbers obtained from equations seem to be within the same magnitude that those given in SpaceNut's pdfs (but still i cannot apply the results for MRO).
I know that my constant reference to the MRO may be irritating but this spacecraft did the maneuver without any apparent heat shield. Now you may argue that is nothing special for an unmanned probe but look at the size of this thing (the picture on the bottom where the people are stating next to panel):
https://en.wikipedia.org/wiki/Mars_Reco … ce_Orbiter
Currently, as the software i written plots the graphs quite similar in shape to the ones given in SpaceNuts examples, i am trying to calibrate the model so it will give proper output values for this article Figure 6:
http://www.ssdl.gatech.edu/sites/defaul … 7-0164.pdf
The problem is the data in the publication do not agree with outputs from the equations. Can i inquiry how one should calculate the ballistic coefficient beta? I mean exactly step by step solution with numbers from the article since any given mass, area or drag coefficient cannot produce published results. (beta = 66.4 kg/m2, 302kg/m2).
I also did some reading on Mars atmosphere model. Currently for density calculations i used this simple solution:
https://www.grc.nasa.gov/WWW/K-12/airpl … mosmrm.gif
The results are ok, but i am aware that this is for the very rough estimates.
This pdf in section 4:
https://ntrs.nasa.gov/archive/nasa/casi … 003184.pdf
claims that the MARS-GRAM may be obtained free of charge, yet i cannot find any existing copy.
As for the ethical issues, I am trying to evaluate an unmanned(!) cargo delivery system for heavy loads to Mars orbit or surface. NO HUMANS ON BOARD.
If BFR can land on Mars and withstand the reentry heat then aerobrake should not be a problem for a similar design.
The wildly variable Mars atmosphere is not an issue. The BFR must be designed to land in even most hostile condition two times (once on earth). Compare the energy difference between the landing and aerocapture and then divide the rest by 450. This will not give you the exact value but look at the difference.
The trajectory is problematic, but was done before. Is not like you need to do exactly 450 orbits. Just do the almost perfect correction on the first and last ones. You would need to do a very big mistake to crash the ship aiming at perigee that reduces dv by 3.6m/s, and you can afford to do a more than necessary initial brake since your heat shield based on BFR design allows it.
SpaceNut,
I am not thinking about landing. Just aerobraking, then orbit stabilization.
GW Johnson,
I do not think that this method will work. Look at the data from MRO:
https://ntrs.nasa.gov/archive/nasa/casi … 005139.pdf
The temperatures to withstand are much lower than the ones in your solution.
My current thinking:
Spacecraft to park in LMO needs to change velocity by 1600m/s.
In 5 months there is about 450 orbital passes, so in single pass dv=3.556m/s
Aerobrake is kinetic to thermal energy transformation. Kinetic Energy in single pass is E=1/2*m*v2.
Spacecraft mass is equal to 300t, so E = 1896.3kJ.
The temperature can be calculated as dT = Q/(m*cp). The ship is covered by 0.001m thick aluminium plates. Total area covered equals 1700m2. Aluminium density is 2700kg/m3, so total mass equals 4.59t. Cp for aluminium = 910J/kg*K.
If load is equal on the whole ship then: dT = 0.46K
Ship frontal area (nose + solar panels as airbrakes) = 135m2. Cover mass on this part = 364.5kg.
dT for the frontal area: 5.71K.
Too good to be true. Any suggestions?
GW Johnson, you are mostly right. The thing is that i have all those values and i can calculate heat but what i need is temperature. I am no expert in this field so my assumption my be wrong, yet currently i think that the calculation should look like this:
During the space flight the ship is in some sort of static thermal state. Then on the atmospheric entry the ship starts generating heat - the lower the altitude the more heat is generated. At this time generated heat rises the temperature of the shield. After the perigee heat generation starts to fall but temperature still goes up (if the radiation is not strong enough yet). This last until the ship leaves the atmosphere- then the accumulated heat radiates into space. Finally the ship is back to the original state.
I assume that there are two design constraints:
Firstly, the temperature. After perigee, when the shield radiation exceeds the heat generated by atmosphere the temperature starts to fall and so this is the max that the material needs to withstand.
Secondly, the heat leak. While the ship is not in it original state, the heat passes through shield. Total energy that goes in cannot rise the internal temperature by more than several degrees, so the shields needs to have a high heat conductivity resistance. This part is quite tricky. The heat that the ship absorbs is (atm generated heat) - (radiated heat). The radiated heat is tied to shield surface temperature and high resistance makes the shield hot faster (as the heat cannot go in). Also if the external material has a big heat capacity and low conductivity there is a lot of time before the internal surface of the shield starts radiating the heat and as the whole pass last somewhere about 10 minutes the unsteady heat analysis for heat transfer may be of use.
Can you calculate this by hand?
RobertDyck, it is very bad that you did not complete that project. I would like to see a small private community initiative in space as this would surly rise the interest in space industry.
As for the subject, i am looking for something like this:
https://www.cs.odu.edu/~mln/ltrs-pdfs/N … tm4674.pdf
https://ntrs.nasa.gov/archive/nasa/casi … 009520.pdf
Does anyone posses the copy of this LAURA program or know of a alternative similar software?
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.
