Solving Mars mission docking with Phobos

GW Johnson · 2020-05-17 19:05:41

In orbital mechanics, potential energy is defined to be zero at infinity for mathematical convenience. PE values are thus negative closer in, so that PE increases, as one recedes farther, as expected. It's just that the PE numbers happen to be negative.

Req is the equatorial radius and h the altitude above the surface. M is the planetary mass, and G is the universal gravitation constant. V is the speed of the body under consideration. Here are the formulas:

PE = -GM/(Req + h)

KE = 0.5 m V^2

For no drag or gravitational losses, the sum of these two items must be conserved (conservation of total mechanical energy, a nonrelativistic thing). Note that in the formula PE goes to zero as R = Req + h goes to infinity. Whatever is left over is KE, which must be a positive number.

For something whose velocity is zero at R = Req + h = infinity, the velocity at R = Req (h = 0) is (defined to be) the surface escape velocity Vesc. That case corresponds to PE + KE = exactly 0. Thus 0 = -GM/Req + 0.5 Vesc^2, and so Vesc = (2 GM/Req)^0.5, as per all of the texts. This is just a hair over 5 km/s for Mars.

What we are interested in, is the velocity at the entry interface at Mars, for something initially at rest at the distance of Phobos. If it is at rest at the distance of Phobos, the V = 0 and R = Req + h = 9410 km.

Thus the total mechanical energy out at that point is GM/9410 km + 0.5*0^2 (the KE term is zero). You must convert km to m to have consistent units, for use with the standard SI metric value of G = 6.67342E-11 Nm/kg^2, and the mass M listed for Mars = 6.42E23 kg.

I calculate ME = PE = -4.5528E6 Nm at the distance of Phobos, for zero velocity V. This mechanical energy is conserved until you hit atmosphere or apply thrust, all the way down to the atmospheric entry interface altitude. As PE decreases (further negative), KE increases (from zero).

The interface altitude at Mars is generally considered to be 140 km, with Req = 3386 km. The potential energy at 140 km altitude calculates as -12.1503E6 NM, for a KE = +7.5975E6 Nm. The interface velocity then calculates to be 3898 m/s = 3.898 km/s.

The entry angle associated with this is exactly 90 degrees below local horizontal, because this motion is strictly radial with respect to the center of Mars. One can then use the hypersonic constant ballistic coefficient analysis at some "realistic" ballistic coefficient, to estimate how deeply into the atmosphere this penetrates at Mars during hypersonic entry deceleration. That is a separate problem.

But for anything, anything at at all, launched from Phobos such that its velocity respect to Mars is zero, at the distance of Phobos, the velocity at Mars atmospheric interface (after falling ballistically) is 3.90 km/s. (It's just about 4 km/s at the surface.)

This is true no matter its shape or density. What happens to it during atmospheric entry is entirely dependent upon its shape and density.

For other situations such that velocity with respect to Mars is non-zero at the distance of phobos, the KE term out there is non-zero, and the mechanical energy is larger. So velocity at entry interface is larger. But such paths are transfer ellipses tangent to the orbit of Phobos.

The one with pretty near min delta-vee is the one also tangent to the surface of Mars (a surface-grazing orbit). For such, 2a = Req + Rphobos. Its apoapsis velocity is INHERENTLY smaller than Phobos's orbital velocity. Its periapsis velocity is INHERENTLY smaller than Mars's surface escape velocity.

GW

tahanson43206 · 2020-05-17 19:46:39

For GW Johnson re #51

Thank you for this detailed and comprehensive summary!

SearchTerm:OrbitalMechanicsMars
SearchTerm:OrbitalMechanicsPhobos

I've pulled a detail out of the post because I'm hoping I have the correct understanding of how you arrived at the number you gave:

But for anything, anything at at all, launched from Phobos such that its velocity respect to Mars is zero, at the distance of Phobos, the velocity at Mars atmospheric interface (after falling ballistically) is 3.90 km/s. (It's just about 4 km/s at the surface.)

The velocity of the object falling ballistically from the altitude of Phobos is less than the value I came up with, which was on the order of 6 km/s.

My ** guess ** is that the difference could be accounted for by your having used the gradient of the force of gravity (which is less at the altitude of Phobos), while I used the force of gravity (acceleration due to gravity to be more accurate) at the surface of Mars for the entire 3700 mile drop.

It is helpful to know that the ** actual ** velocity of a dropped object would be less than I had thought, but the number is ** still ** (very likely to be) greater than a balloon fabric envelope could be expected to survive upon encounter with the atmosphere of Mars, thin as it is.

(th)

GW Johnson · 2020-05-17 21:06:57

Your guess is correct. The lower effective average value of gravitational acceleration leads to a lower velocity. You have to figure this right, to get the right answer. I gave the equations in my previous post.

I took a good guess at a generic study of spheres dropped into Mars's atmosphere from rest, at the distance of Phobos. I did this versus diameters from 0.01 meter to 10 meter, by factors of 10.

I used a constant low density for the spheres of 100 kg/cu.m (sp.gr = 0.1), and a constant hypersonic CD of 1.5, based on the circular cross section area. The "nose radius" for heating rates is half the sphere diameter.

The square-cube scaling affects areas, volumes, and masses, and therefore ballistic coefficients. Those data are:
dia, m Ax, sq.m Vol, cu.m mass, kg beta, kg/cu.m
0.01 7.85E-5 5.24E-7 5.24E-5 0.4444
0.10 .007854 .000524 .05236 4.444
1.00 .785398 .523599 52.35988 44.44
10.0 78.53982 523.5988 52359.88 444.4

For all of those diameters, the entry interface velocity is 3.90 km/s, and the trajectory angle is 90 degrees below horizontal. I used my 1955-vintage warhead analysis spreadsheet calculator to get the hypersonic trajectory dynamics. Those results show:

dia,m h, km at max gee max gee
0.01 55 32.5
0.10 35 32.5
1.0 15 32.5
10. -5 32.5

dia, m h, km at max Q max Q/A, W/sq.cm
0.01 66 36.5
0.10 45 36.5
1.0 25 36.5
10. 5 36.5

dia, m h, km for M3 speed exit from hypersonics
0.01 44.5
0.10 24.7
1.0 4.9
10. -13.5

I was surprised to see the very same (unsurvivable) max gee values, and the same modest max heating rate values, for all the diameters I looked at.

Figure the gees on the mass, for the decelerating force. Then divide that by the cross section area, for the average wind pressure across the wetted surface. Peak wind pressure is about twice (or more) that value. For a balloon, the internal pressure must exceed that value, or the shape collapses.

The most alarming result was the negative altitudes at which exit-from-hypersonics occurred, in any "reasonable" size. This was for a 10m dia sphere, and it also showed up in the negative altitude for max deceleration gees at that size.

For shallower entry angles nearer 2 degrees below horizontal, much higher betas result in more manageable altitudes, even for very much higher velocities at entry interface. And, most tellingly, the peak deceleration gees are much, much lower. 32-33 gees is pretty much fatal to humans, especially those weakened by exposure to microgravity. At 45 gees, human bodies literally start coming apart, even if fully reclining.

GW

PS update 5-18-20: 36.6 W/sq.cm = ~ 116,000 BTU/hr-sq.ft (units I am more comfortable with). At emissivity ~ 0.5 and radiating to space at absolute zero, the equilibrium balloon skin temperature is just under 3000 F at the peak heating point. It's somewhat lower if emissivity is closer to 0.8, and substantially higher is emissivity is closer to 0.2. 3000 F is above the 2350 F solid phase-change embrittlement point for aluminosilicate materials, but below their 3350 F meltpoint. Such could only serve as single-use balloons, and only if the fabric porosity (around 30%) could be sealed in some unknown way to hold gas pressure.

Last edited by GW Johnson (2020-05-18 08:27:09)

tahanson43206 · 2020-05-18 09:06:25

For GW Johnson re Post #53

I had just called up NewMars to take a print of your earlier post, with details about Mars so I could run some calculations, when I was surprised (and delighted) to find you'd posted a serious study of the falling balloon problem.

In the parallel topic (Balloon landing) I've been attempting to find an optimum mix of technologies, and you've been over there helping as well!

Since we are mixing and matching subtopics within topics, I'd like to invite you to find the altitude i need to find for the bottom of the tether from Phobos, so that Void's "bubbles" can survive and deliver useful content to anxious customers on the surface.

Meanwhile, I'll print your earlier post and some of SpaceNut's links, to try to find how far Kevlar cable can stretch (I know, 2%) to provide a suitable dropoff point.

While there is absolutely NO guarantee this investigation is going to lead to anything other than bitter disappointment, it ** is ** exhilarating while the hunt lasts.

The rabbit in the bag, of course, would be to find a solution that is safe, cost-effective, reliable, and of sufficient validity to convince a skeptical funder.

The ** ultimate ** skeptical funder would be NASA. Success with NASA is probably too much to hope for, but perhaps one of the American billionaires would be willing to take a whack at it.

***
Please note the introduction of a book project idea. I would be hoping you might be interested in the PhD level component.

I have several authors in mind for the lower school levels.

Edit#1: I hope this next question interests you!

Can Kevlar rebound from that 2% stretch? If so, it could/would be a useful feature for the operator to use in real time to assess the functioning of the descent control system. If a descending vehicle is placing too much stress on the cable, then it may have to be allowed to descender faster than might be ideal, in order to protect the tether. In the case of a cargo vehicle, the risk would be transferred to the after-drop-off phase, where remedial measures might be taken, if they were provided for.

In the case of a passenger vehicle, I would expect that tolerances would not be pushed (except in an emergency such as a hospital trip), but the 2% measurement by sensors along the tether would be useful. Come to think of it, only one sensor (at the bottom) is really needed, but sensors along the length would be helpful as the tether deteriorates over time, so that a weak section can be coddled until a replacement section is available.

(th)

Last edited by tahanson43206 (2020-05-18 09:37:02)

GW Johnson · 2020-05-18 21:17:10

I looked at a surface-grazing transfer orbit from Phobos to Mars's surface. This requires escape from Phobos, then a delta-vee to get from Phobos's orbit speed about Mars to the apoapsis velocity of the transfer orbit.

escape Phobos 0.020 km/s
apoapsis delta-vee 0.582 km/s
entry velocity 4.204 km/s
entry angle 9.355 degree (below horizontal)

I looked at the same spread of density = 100 kg/cu.m spheres ranging from 0.01 to 10 m diamter. The nose radius for heating is half the sphere diameter. Hypersonic drag CD was assumed 1.5, as before.

I found a similar pattern of answers (constant peak gee and heating rate), but I also found this approach far more favorable with higher altitudes and lower gees and heating.

peak gees peak heating M3 point
D,m beta,kg/sq.m h,km gees h,km Q/A,W/sq.cm h,km
0.01 .4444 71 6.14 81 18.46 59.9
0.10 4.444 51 6.15 61 18.47 40
1.0 44.44 31 6.15 41 18.47 21.5
10. 444.4 11 6.15 21 18.47 0.3

Peak gee near 6 is much more survivable by humans, even those somewhat weakened by microgravity disease. Peak heating near 18.5 W/sq.cm = ~ 6700 BTU/hr-sq.ft. At 0.5 emissivity and the environment at absolute zero, this corresponds to a radiating material temperature near 1212 F, well with the reuse capability of aluminosilicate fabric materials, or even bare stainless steel.

The altitudes look good until you get close to 10 m size with its higher ballistic coefficient beta = 444.4 kg/sq.m. This traces to the nearly 10 degree down-angle at entry on this trajectory. This can be adjusted by raising the periapsis altitude a few km above the surface, instead of absolutely surface-grazing. There's a trade to be done, there.

This approach has a lower delta-vee requirement than the vertical-drop approach (under 1 km/s versus over 2 km/s), and a much easier and more survivable entry scenario. It gets into trouble with low altitudes somewhere between ballistic coefficients 444.4 and 44.44 kg/sq.m.

Entry pressures are also lower, leading to lower balloon pressures being required to resist them (this still begs the question of sealing a porous aluminosilicate fire curtain cloth to hold gas pressure at elevated temperatures).

Just musing about this, the possibility of having this balloon buoyantly float in Mars's atmosphere seems quite ridiculously remote! There will have to be some sort of landing apparatus. But with higher altitudes, chutes become feasible, to be combined with landing rockets.

Beware, the denser we make these things, the deeper they penetrate down to lower altitude during entry, because mass and ballistic coefficient increase for any given size. Nothing will change that.

What I have really shown is that more-conventional vehicle designs are feasible for the transfer orbit approach, but not the vertical drop approach. That costs too much delta-vee to kill Phobos's orbit speed, and it incurs far too stressful an entry because it dives straight down.

Shallow entry angle is simply essential. Nothing is going to change that, either.

GW

PS -- the ~2% elongation for Kevlar is mostly elastic.

Last edited by GW Johnson (2020-05-18 21:20:52)

tahanson43206 · 2020-05-19 06:24:54

For GW Johnson re Post#55

Thank you for the work you put into the study you reported! And thank you for confirming the "mostly elastic" property of the 2% elongation for Kevlar.
This tells me it would be reasonable to anticipate measuring stress on the material in real time using optical instrumentation (ie, lasers) to accurately measure elongation. That information could/would be fed into the control programs for the descent system in the balloon tether topic.

Your work on finding traditional de-orbit approaches for Void's gas-filled bubbles seems (to me at least) quite encouraging. I hope that folks with more education and experience than I can offer will begin to think about the approach you have outlined.

A distinct advantage of the traditional approach you are developing, with refinements for the unique features of Mars, is that the infrastructure required is MUCH less massive. I am arriving at the understanding that the tether approach can be done, but it would require a Normandy Invasion investment to pull it off. Several projects were carried out to fulfill the Normany Invasion. I'm thinking now of the British mobile dock systems, which were built in sections in Britain, towed over the channel and assembled for deployment on the shore and extending out to deeper water for cargo transfer.

A tether appears to be feasible, and it would (apparently) offer gentle dropoff of cargo and personnel above the atmosphere, but the investment would be massive. Per Google >> www.compoundchem.com, and pubchem.ncbi.nim.nih.gov, the formula of Kevlar is C14 H14 N2 O4.

It would be (should be) possible to assemble the components by harvesting asteroids and comets and the occasional moon, without bringing everything from Earth.

In contrast, as I understand your approach after a first reading, investment would be limited to individual vehicles, and therefore the return on that investment could be directly calculated.

If there is someone in the registered forum able to support GW Johnson in exploring this way of delivering Void's bubbles of gas to the surface of Mars, now is a good time to pitch in. If there is someone in the forum audience who can assist in refining this set of ideas, now is a good time to register and help out.

(th)

GW Johnson · 2020-05-19 08:25:59

What we have been dancing about here is what the JPL guys have long termed the Mars EDL dilemma: that larger objects have higher ballistic coefficients, which makes them smack the surface before you can slow down. For direct entry trajectories, that's where the 1-ton lander size limit comes from. It's not a universal limitation, excepting for that entry scenario.

I followed up my own suggestion, and tried a periapsis at 70 km instead of 0 altitude. Entry speed was about the same, but entry angle reduced from 9.5 degrees to 8.3 degrees. That got my Mach 3 point up to 0.9 km from 0.3 km, at the 444.4 kg/sq.m beta. Positive, but not much improvement, so orbit details like that are not a sensitive variable to use.

But when I reduced the sphere to 7 m diameter for beta = 311.1 kg/sq.m (comparable to Apollo capsules), the Mach 3 point altitude went up to 4 km. Not surprising at all given other studies I have done, and what JPL has been publishing for years. Size and beta are VERY sensitive variables.

What you'll really want to do to get the endpoint altitude up is do a transfer orbit from Phobos to low Mars orbit, then a second small delta-vee into low Mars orbit. Deorbiting from there is a trivial delta-vee, a lower entry speed (about 3.7 km/s vs 4.2 km/s), and a much lower entry angle (nearer 2 degrees than 8). That'll make landing a whole lot more feasible at high beta.

No surprises there.

But the lure of decelerating to rest (2+ km/s delta vee) out near Phobos is clearly a false promise. You pick up to near 4 km/s entry speed when you fall to Mars, and you enter straight down, at unsurvivable gees and very high heating, and you smack the surface still moving straight down at hypersonic speed, when you try to do it this way.

What I have found applies to any and all vehicle designs.

I remain quite skeptical of delivering things to Mars inside balloons. You cannot deliver anything denser than a gas that way, and it has to be at a pressure higher than the wind pressure on the surface during entry. There are NO ready solutions for sealing or heat protection for a flimsy structure like that. And there are no solutions for the high-subsonic to supersonic impact with the surface, because NO structure capable of surviving entry forces could possibly be a buoyant structure in that near-vacuum of an atmosphere.

Sorry to burst your balloon, if you'll forgive my choice of words.

GW

PS -- this is easy to do with the textbook equations for orbital mechanics in one spreadsheet, and the 1955-vintage entry dynamics approximation for warheads in another spreadsheet. That's exactly how I did it. All this stuff is in the Justus & Braun tome on planetary EDL, except that they botched some of the converting of the old H. Julian Allen entry analysis correlations to metric units. I fixed that.

Last edited by GW Johnson (2020-05-19 08:32:08)

tahanson43206 · 2020-05-19 10:57:57

For GW Johnson re #57

No problem with "busting balloons" while they are just ideas << grin >>

Thanks for this further analysis!

I'll come back to study #57. Hopefully there are others in the forum readership who will do the same.

Edit#1 ... Still planning a slow reading, but I'm back anyway with a question about the results you showed ...

It seems (as I read your report) that you are anticipating that the vehicle will maintain the same face to the wind.

However, these are (theoretically at least) spheres, and they could be designed (with little fins) to rotate in the wind.

The first effect (and the most important one perhaps) is that the heating caused by the atmosphere would be spread over the entire surface over time.

The second effect (I'm not sure how useful it is) would be that heat could radiate to the sky while the heated surface rotates to the back side.

Third, I would expect some small lifting effect to occur, depending upon the rate of rotation and perhaps other factors I'm not aware of.

In other words, at the risk of further complicating your work, perhaps allowing for rotation of the balloon/bubble/ball/sphere can improve performance to some small degree. Perhaps a small adjustment like this could be enlisted to provide a minimal level of steering.

Edit#2: Some serious academic research has been done to investigate the effect of spinning on soccer balls. However, I decided not to post the link to the first study I found, because the results, while indicative of an effect of rotation on movement, were considered to suffer from difficulties with the experimental apparatus. The study included multiple types of soccer balls, with varying patterns of surface, from none to complex patterns.

Another field of study for rotating balls in air would be baseballs, which are observed to travel in paths influenced by spin as well as the nature of the surface of the sphere.

(th)

Last edited by tahanson43206 (2020-05-19 13:55:43)

SpaceNut · 2020-05-19 15:30:35

Current mars landing from orbit to surface is a 15:1 ratio of hardware, fuel and devices to slow the item down to have a payload that we can use.

So 15 tons of ship on orbit with payload of 1 ton on the surface is where we are for the technology.

We know from aero dynamics that the larger the diameter the heatshield is the more payload we can land.
The same holds true for upping the fuel mass for the engine run time is the other solution to slow a vehicle down for the larger payload.

Bottom line more mass for mars means more mass to launch out of earths gravity well which incurs costs for the mission.

GW Johnson · 2020-05-21 08:26:49

To get from Phobos to the surface of Mars one-way is not very hard technologically, if you use the transfer orbit and low Mars orbit flight path. You will need a minimal heat shield, even at Apollo capsule-class ballistic coefficients, and minimal but controllable rocket propulsion.

escape phobos...about 0.02 km/s
enter transfer.....about 0.6 km/s
enter low orbit...about 0.75 km/s
deorbit..............about 0.05 km/s

If you come out of hypersonics above 10-15 km, you can use a ringsail chute to decelerate down to high subsonic speed, and then land with rockets for a landing delta-vee allowance of perhaps 0.5 km/s.

If you are bigger and heavier, you will come out of hypersonics around 5 km. There is no time to deploy a chute, much less the time for one to decelerate you at all. Just go for the direct retropropulsive landing from Mach 3 class speed. That would be the higher landing delta-vee allowance of about 1.5 km/s.

Either way, we already know exactly how to do all of this stuff.

The two tiny delta vees would be best done with small thrusters. That's 0.07 km/s delta-vee required of your your attitude control system. Double or triple that to cover actual attitude control, for 0.14-0.21 km/s worth of thruster delta-vee.

The rest totals 1.35 km/s in space, plus no more than 1.5 km/s to land. That's 2.85 km/s total to get from Phobos to the surface of Mars one-way. Your peak gee loading is about 4 gees for entry from low Mars orbit. Your peak heating is low enough to use heat-sinked stainless steel (if you wish) as your heat shield. A thin paint layer of Avcoat on just about any metal substrate but aluminum, would work quite well.

The vehicle could be as simple as a heat shield and cargo deck, with a minimal metal aft aeroshell to protect the cargo stacked on the deck. Put your propellant tanks on that same deck. Fire your rockets through holes in the heat shield, from small sealed compartments around each rocket that stop throughflow. It's one-way, no landing legs, just plop down on the heat shield. Minimal.

GW

PS update:

For the main engine(s), delta-vee is 2.85 km/s. Assuming Hydrazine-NTO storables, with a very modest bell in vacuum, Isp ~ 300 sec, for which Vex ~ 2.94 km/s. The required mass ratio for that is 2.636, corresponding to a propellant fraction of 62.1%.

Assume the thrusters are just smaller engines using the same propellant supply at the same Isp. That delta-vee is 0.21 km/s, for a mass ratio of 1.074, for a propellant fraction of 0.069.

Add the two propellant fractions together, for .690. Now assume the vehicle inert fraction is 0.10 (10%), although it might actually be a bit lower. That leaves 21% for dead head cargo, as a worst case for all-retropropulsive landing (no chutes). Not too bad.

Last edited by GW Johnson (2020-05-21 11:34:42)

tahanson43206 · 2020-05-21 11:49:38

For GW Johnson re #60

Thank you for adding (what seems to me) substantial support for the Phobos docking maneuver.

What I am seeing/hearing is that if NASA were your customer, and you were to offer to land 40 MT on Mars with a high degree of confidence (which I am assuming you have << grin >>) then you would invite them to set up a landing package and ship it to Phobos using a Deep Space Only vehicle.

The total package would amount to (21% of X = 40 MT) so X = 190 MT and change.

Is this a strategy that would appeal to anyone?

In your vision (as I understand it) the vehicle would end up safe and sound on Mars, with fuel exhausted.

It might be designed to be useful there, or it might be designed to be refueled and returned to orbit.

Since Dr. Zubrin has spent some time on the the refuel-and-return option, I'm wondering if you would be willing to spend a few minutes thinking about how to design the lander for a long and fruitful life as part of the growing infrastructure on the surface.

(th)

Last edited by tahanson43206 (2020-05-21 11:50:39)

GW Johnson · 2020-05-21 18:40:31

Well, if you fly with zero payload, and a full load of propellant, there is mass ratio enough to reach low Mars orbit and a bit more, but not enough to return to Phobos. The design sizing is about 0.4-0.5 km/s short of that.

That's with hydrazine-NTO at 300 sec Isp. With LOX-LCH4, the Isp is higher (360+ sec), and I rather think it might make it all the way back to Phobos. The design mission is the return to Phobos, not the descent, under those circumstances. That is because the mass ratio-effective factored delta-vee is higher for the return.

GW

tahanson43206 · 2020-05-22 04:57:26

For GW Johnson re #62 in light of #60

In Post#62, if I am reading correctly, you have considered how the lander might return to Mars orbit, if it were refueled on the planet, from local resources.

From my interpretation, it might be more useful as a source of parts and material for the surface expedition.

Add the two propellant fractions together, for .690. Now assume the vehicle inert fraction is 0.10 (10%), although it might actually be a bit lower. That leaves 21% for dead head cargo, as a worst case for all-retropropulsive landing (no chutes). Not too bad.

In Post #62 (as I read it) you have thought about a return to Phobos, but it would appear that is not practical for the design of Post #60.

In Post #60 (again as I understood it) you were offering a vision of a way to land 40 MT (for NASA or any customer) on Mars, WITHOUT a return to Phobos.

Perhaps a useful way to configure the lander would be as a source of parts to be removed from the lander and employed for structures on Mars.

The tanks should be usable in manufacturing. It's hard for me to see a comparable applications for engines or some related piping, so those might be melted down for their valuable metals.

Can you confirm my understanding that the vehicle you have in mind could land NASA's 40 MT if 190 MT is allocated for the lander? That was based on the estimate of 21% for the payload to the surface.

NO vehicle (to this point) has been designed to return from the surface to anywhere, so leaving the vehicle on the surface would be "traditional".

The operational concept I've been hoping would evolve from this discussion is of a Deep Space vehicle able to deliver a 190 MT lander to Phobos, from which it could depart for a one-way trip to Mars. In this concept, the Deep Space vehicle would remain on station for a time and then resume the ellipse back to Earth.

The humans on board (assuming there are some) could interact with the lander and with the payload using teleoperation, in order to greatly accelerate the work plan for the payload, before resuming the flight.

I'm not sure how many days the vehicle can loiter at Phobos before returning, but I'm assuming there is a way to expend a bit more propellant on the way home to catch up with the Earth at the right time. After all, the 190 MT lander will no longer be present, and the propellant to dock with Phobos will be absent, so the vehicle will be lighter for the return burn.

That ** would ** be a long time to be cooped up in the Deep Space vehicle, but some EVA at Phobos could be a part of the mission.

Dropping off some equipment at Phobos could be (and from my point of view, should be) part of the mission plan. After all (in this scenario), the Deep Space vehicle will be returning two years later with another lander, another crew and more equipment to deploy at Phobos.

(th)

GW Johnson · 2020-05-22 16:16:40

Well, the issue under discussion seemed to be ways and means to deliver something one-way from Phobos to the surface of Mars. I clarified what class of trajectories was better, then did a bounding analysis to determine the proportions of a single stage craft to do exactly that, with relatively-minimal-performing storable propellants. I then checked its performance potential to ascend back to Phobos if unladen of payload, and found it wanting. So I reported that any two-way craft would be sized by the more-demanding ascent requirements, which is the usual case at Mars.

All that being said, I have to wonder about something very fundamental to the attractiveness of using Phobos at all, since trips between it and the surface are more demanding than trips between low Mars orbit and the surface. The usual assumption seems to be that volatiles can be mined from Phobos to be used for propellant or other supplies. I have to take serious issue with that assumption, and assumption it is (there is NO ground truth). I think Phobos, like all the other small asteroids, is devoid of volatiles. Those are proving to be dry rubble piles, of varying very-weak cohesion, but devoid of any ice content.

If that's the case, then what's the point of messing with Phobos? Why not just choose between direct entry or staging out of low Mars orbit? There are good arguments for both. As far as any orbit-to-orbit spacecraft is concerned, there is very little dfference between going to Phobos versus going to low Mars orbit. The delta-vee values are not very different. Not very different at all.

GW

Last edited by GW Johnson (2020-05-22 16:20:09)

SpaceNut · 2020-05-22 16:42:41

The vehicle from earth is sent via ion, with artificial gravity and with that the orbit choice to be in lmo or to land of phobos to be attached to the tether transfer cord only removes some dead mass and or is used to top off the tanks for mars landing once we start down the tether.
The energy to travel down the tether is via solar most likely for the climbing portion of the trip to mars.

tahanson43206 · 2020-05-22 18:18:01

For SpaceNut re #65 .... I understand that it is easy to merge topics when they contain similar content.

However, ** this ** topic (managed by GW Johnson) is NOT about a tether. Instead, what i am watching evolve over time, is a set of ideas (supported by calculations) that could lead to a proposal to NASA that NASA would actually consider. The tether topic is unlikely to receive such a favorable review.

For GW Johnson ... In an earlier post, (as I recall) you provided a figure of 1.8 Km/s as the optimum landing/docking DeltaV for a mission to Phobos.

My recollection is that a Low Mars Orbit will ** ALWAYS ** require a much greater DeltaV.

My understanding of the difference between the two is that Phobos will lend its momentum to slow the vehicle.

That is why I asked some time ago if there is any risk of a tall space vehicle tipping over due to Coriolis forces.

Beyond that, however, is the human side of the equation. A mission to Mars can certainly be organized to come and drop off a payload and then leave.

I don't think there have been any missions that fit that description, but there certainly could be in the future.

My reason for a focus on Phobos is that it provides a "There there!".

If you say you're going to Phobos, where a group of Nations have set up base camps, then it will instantly be understood what you're talking about.

If you're going to Mars Low Orbit to drop off a payload (such as NASA's target of 40 MT), then you are indistinguishable from any other mission planner.

If you're offering a trip to Phobos, in a Deep Space vehicle that has features to benefit travelers, AND you're offering to drop off a 40 MT payload while you're there, I'm pretty sure you have the game to yourself.

if you dress up your proposal with addons such as configuring your lander to convert into a tool shed, or something even more valuable, then (it seems to me) you are in a league by yourself.

***
Taking a slightly different tack ...

On multiple occasions, RobertDyck has reminded forum readers of the history of Newfoundland.

Recently I met a chap who spends summers in Newfoundland, as an unexpected benefit of his interest in some aspect of the region which I have since forgotten.

I bring up this chap because his stories about the region matched up well with RobertDyck's reports.

Phobos has the potential to become a major shipping center like the Panama Canal on Earth today.

I've been working on the possibility of a tether, but Phobos is capable of becoming a major shipping hub without that.

Elsewhere in the forum there has been speculation about fishing CO2 from the atmosphere. That exploit seems unlikely to me for the same reason the tether is unlikely, but it is (most likely) possible.

What seems ** much ** more feasible is collecting valuable materials from passing comets and the occasional asteroid, and choosing favorable celestial conditions to arrange for shipment to Phobos.

The example of Newfoundland (it seems to me) is the success of private enterprise by individuals and small groups, motivated by the hope of return on hard labor and investment in wooden shipping and supplies to make it possible.

The trouble with Mars Low Orbit is there is no "there" there. With Phobos, there is the possibility of creating a Solar System Wide resource.

(th)

Last edited by tahanson43206 (2020-05-22 18:20:22)

SpaceNut · 2020-05-22 18:38:21

The point poorly made was there is little to no energy difference in fuels used for an orbit around Mars versus one that lands on phobos from the stand point of using ion drive and ag to get there.

That follow up landings from either stating point has little difference for fuels.

With that what you have said about leveraging the use of phobos is where the issue for mars changes as you noted as well as its about saving and replenishing resources.

These are similar related information for mars:

Mars landing from GW's website
to give information on mission to the mars area
mars with mass crew variables for mars
how the capsule shape and incoming speed make a difference for mars
post showing how that speed must burn off in the atmosphere
what we must do to refuel a ship that uses either location as a post

GW Johnson · 2020-05-23 08:18:20

Now I have never really understood what the tether thing is all about. I have heard of the "space elevator", which for Earth would require the invention of superstrong materials (what I have called in the past "manurium", "unobtainium", or "unbelievium").

However, there seems to be a different concept involving rotation which might be done with things we already know how to make. That one I do not understand yet. All I know is that you cannot push on a string. Or a chain, for that matter.

If instead you are looking at more conventional space vehicles, what I have done as a part of this posting is determine which class of trajectories might be the better choice for travel between Phobos and the surface of Mars. That turns out to be a routine transfer to low Mars orbit from minimal-escape from Phobos, driven as much (or more) by intolerable entry conditions as by delta-vee requirements.

All that begs the question of whether or not you would want to go to Phobos. I cannot answer that. So far, no one can.

For an orbit-to-orbit transport design, the total one-way or round-trip delta-vee requirements are a little less going to Phobos than to low Mars orbit, both starting from low Earth orbit. That is true, because of the lower orbit velocities further out from Mars, and taking advantage of Phobos's direction of movement and extremely-low gravity. But it is not a large effect.

The final leg of such a trip is travel between Phobos and Mars. Phobos costs a lot more delta-vee than low Mars orbit, especially the ascent. There is no way around that. To go to Phobos from the surface of Mars, you essentially have to escape from Mars. That's significant.

A solid surface with such low gravity as Phobos (escape velocity about 18 m/s) presents no advantage I can think of, for any purpose. That moon being most likely a captured tiny asteroid, it is unlikely in the extreme to have any buried ices from which to make propellant, drinking water, or oxygen to breathe. There might be oxygen atoms in the rock molecules, but it'll cost you way more than an arm and a leg's worth of energy to get it out. The thing is a dry rubble pile of loose rocks of varying size. What's point?

Now staging your mission out of Mars orbit instead of a direct landing has both advantages and disadvantages. Those have to be evaluated in terms of what you ultimately intend to do by going there. One huge advantage is that you do not have to make your transport ship meet the design requirements for landing and ascent. Similarly, your landing and ascent craft does not have to meet the design requirements for an extended interplanetary trip. Those are VERY different sets of design requirements. I think or at least hope that everybody here understands that point.

There is another advantage to basing out of low Mars orbit that pertains to exploration, if that is your purpose. It does NOT pertain to planting and maintaining bases or colonies. If (1) you have a 2-way-capable landing craft, and (2) you have more than one of them or can reuse them, then you can explore more than one landing site during the one trip to Mars. Considering how difficult it is (at this time in history) just to go to Mars at all, that is one HUGE advantage for exploration.

Note that I define exploration thusly: find out what all is there, and where exactly it is. And I mean that quite literally exactly as worded, word-for-word. Think about what that implies, then think about the explorations done 5 centuries ago. Which ones were the most successful? What exactly did they do?

Now there's a huge delta-vee cost for doing it that way. You incur that knowingly, up-front, if you do it that way. The arrival into Mars orbit is done with rocket thrust, not aerobraking. And it's multiple km/s. Direct entry trajectories kill most of that aerobraking straight in, at the cost of very low terminal altitudes for credible ballistic coefficients of large vehicles. Essentially on the surface at Mach 3. And EVERYBODY has a big delta-vee requirement to escape from Mars to come home.

I know some will argue that multi-pass aerobraking is the way around that arrival cost problem. But consider this: (1) we have never done that before with people, so it is not an off-the-shelf technology, and (2) Mars's atmosphere is factor-2+ variable in its density, even at high altitudes, which renders the results very uncertain at any given time. You want to risk lives on that? I do not.

If instead you are planting bases or colonies, supposedly you are interested only in one site because you are NOT exploring, and the cheapest way to get there is direct entry, pretty much the way we sent the probes, and the way Spacex wants to do it. You still have to pay the cost to come home. Like I said, everybody has to do that.

See now how your choice of where to base depends upon what you are fundamentally trying to accomplish? It's a question I don't often see considered, much less answered.

GW

Last edited by GW Johnson (2020-05-23 08:43:56)

tahanson43206 · 2020-05-23 09:12:16

For GW Johnson re #68

SpaceNut introduced the subject of tethers into this topic by accident. He corrected for that subsequently.

As I understand your theme in this topic (which SpaceNut created for you), it is to work with traditional chemical rocket concepts and various traditional maneuvers to achieve various objectives.

The objective I am hoping to entice you to concentrate on is NASA's stated objective of a landing system for 40 MT of cargo.

There is ** absolutely ** no need to return to orbit. That requirement is reserved for different missions.

What I'm hoping I can interest you in doing is designing a 190 MT lander that can safely and reliably deliver 40 MT to the surface of Mars.

The lander would be shipped to Mars on a Deep Space vehicle which would remain in orbit around Mars.

At this point, it is not important for meeting the mission objective to worry about the question of where in orbit around Mars the Deep Space vehicle would take up residence.

As an additional consideration ... You have mentioned a number of chemical fuels and oxidizers as you have been developing your thoughts in this topic.

I'd like to offer a suggestion that would be practical if you will accept as a working premise that the Deep Space vehicle can electrolize water while it is in transit from Earth LEO to Mars orbit. The vehicle would be loaded with clean water before departure from Earth LEO, and it would accumulate a stock of H2 and O2 during the trip. If it is considered necessary, the gases could be liquefied for loading into the landing vehicle.

Thus, in this scenario, you (the designer) would have a decent set of chemical components to work with.

As you are designing your lander, please consider how you would reach the multiple objectives of:

1) Safe reliable delivery of the NASA 40 MT payload to some reasonable point on the surface of Mars

As I understand your previous posts, the outlines of this objective are already present

2) Provide for re-use of as much of the lander as possible for construction projects on Mars (eg, tanks for storage or habitat, etc)

This objective is not necessarily one for which you might feel prepared. Enlisting others to help is a possibility.

We've been given a challenge to prepare a proposal to enter into competition with more established groups. (See quotes from NASA/Kempton)

I have a level of confidence that members of ** this ** forum collectively are able to compete with the best the planet has to offer.

This forum is ** not ** limited in any way that I can see. Admission policy is so lax as to be amazing. If a hacker of minimal talent can gain entrance to this forum, a highly educated, experienced person with the skills needed for a particular part of the project can surely gain admission.

I have been inviting highly qualified persons to help with undertakings on this forum, and so far no one has taken up the invitation.

That could change at any time, if the forum proves able to initiate and sustain activities of a caliber that would match the standing of the folks who are available.

(th)

Last edited by tahanson43206 (2020-05-23 09:16:58)

GW Johnson · 2020-05-23 12:27:03

OK, here's a nozzle for a chamber that throttles between 200 and 1000 psia, sized to have 20:1 area ratio, as fired in the near-vacuum of Mars surface pressure (0.087 psia). This was figured for a nominal gas specific heat ratio of 1.20, and a conical exit cone with a half-angle of 15 degrees.

If unseparated, the vacuum thrust coefficient at any chamber pressure is 1.791264. If tested on Earth, it has an expected separation backpressure of 17 .1 psia at the max 1000 psia chamber ptressure. At the min throttled-down chamber pressure, the separation backpressure is 3.43 psia (meaning it would require testing with an altitude diffuser system on Earth, if throttled down).

Now for the engine that uses that nozzle. Figured with the chamber c* for LOX-LCH4 and no hot gas bleedoff to run turbopumps, and sized for 100,000 lb vacuum thrust, the Isp figured at Mars is 341 sec at full power, 336 sec throttled to 20% thrust. For a required L* of 90 inches and an Ac/At = 5, such an engine would have a chamber 18.85 inches ID, and 18.00 inches long, a throat 8.43 inches ID, and a nozzle cone with 37.72 inches exit dia and about 54.65 inches throat-to-exit.

You can sacrifice a tad of Isp to run turbopumps, or you can add inert mass fraction to have pressure-fed engines. Take your pick.

Now for a one-way vehicle from Phobos to the surface of Mars, no possibility of return or reuse-as-a-vehicle. The point is to carry just about 40 metric tons (44 US tons) payload. Here are the delta-vee requirements:
escape Phobos-----0.02 km/s
enter transfer----0.6 km/s
enter low orbit---0.75 km/s
deorbit burn------0.05 km/s
retropr.landing---1.05 km/s (all-rocket, no chutes)
total-------------2.47 km/s

Assume we spend more time throttled than at full thrust, so use the fully-throttled Isp as a lower bound. Isp = 336 sec, for an exhaust velocity of 3.295 km/s. The mass ratio required is exp(dV/Vex) = 2.111, for a propellant fraction = 1 - 1/MR = 0.5275.

Typical modern rocket stages are near 5% inert, but we need a cargo deck and some sort of heat shield, plus some sort of aeroshell to protect that cargo during entry. So, assume 10% instead.

That's 0.6275 for inert plus propellant, leaving 0.3725 for payload. If that payload is 40 metric tons, here is the rough-sized weight statement for the vehicle, in metric tons:
payload----40.00
inerts-----10.74
burnout----50.74 (Mars weight, pounds, 42,515)
propellant-56.64
ignition---107.38 (Mars weight, pounds, 89,974)
max thrust = 100,000 lb
min thrust = 20,000 lb

If I do that for the lower Isp and Vex of a storable propellant, the vehicle is bigger for the same payload. So what?

Either way, I do not see what is so damned hard about designing to land a big payload on Mars, once you realize the solution to the "EDL dilemma" there is retropropulsive landing! There's just no time to deploy and slow down with chutes on big objects. They come out of hypersonics too low for that. So fire up the thrust instead.

You simply have to do something different, when what you did before with small craft no longer works with big ones! What is so hard to understand about that?

There is no real "EDL dilemma", if you quit thinking inside the "we have to use chutes" box.

GW

PS -- Look at full thrust to Earth weight for somewhere around half the propellant gone. That would be sort of typical for max retropropulsive deceleration after the hypersonics are over. That's somewhere near 75 metric tons, or about 165,000 pounds. One 100,000 lb thrust engine will decelerate this bird at around 2/3 gee, and that's just not enough, not by a long shot. We need about 4 gees.

We really need somewhere around 6 times that design thrust to get 4 gees retro deceleration capability, near the same as peak 4 gees of aerobrake deceleration. We'd really need something like 5 or 6 or 7 of these engines at full thrust to get the retro thrust we need for final 4-gee deceleration near the surface. Which ambiguity takes care of engine-out.

At actual touchdown, only two or three engines need be running. If two, they'd both be near 21% thrust, right at minimum. If three, we'd need 14,000 lb thrust throttle-down capability. If one, it would be near 42% throttle. That takes care of engine-out at touchdown.

What that says is 6 or 7 of these engines is OK, but they need more than 5:1 turndown ratio. Which can be done, these days. Probably something closer to 8:1 would do the job OK. There I went and specified an engine design. 100,000 lb thrust max per engine, 8:1 turndown ratio, 7 engines. Nozzle expansion limited to 20:1, to control the size of the blast craters these things dig.

So yeah, I really do worry about the very practical stuff that can kill you, when I look at rough-sizing a Mars lander design.

Otherwise, put the cluster in the center of the heat shield/cargo deck, and mount them in a sealed compartment. Let them fire through a hole (or individual holes) in the heat shield. The sealed engine compartment stops throughflow of entry plasma through that hole (or those holes). You can always add a gas pressurization-balance system for the engine compartment, but I rather doubt that you really need one.

You need no engine port covers to open or to discard. Which is also thinking way outside prior boxes! But it WILL work! There is no better insulator than a gas column. You cannot overheat one of those. All else can be overheated.

Somebody ask me why you only want one O-ring seal on a pressurized system, especially if it is a dirty-gas solid. Or why you NEVER, EVER seal the insulation joint leading to one of those O-ring seals. I dare you! Ask me! NASA not understanding those issues is exactly why one of two Shuttle crews died. And they STILL do NOT understand!

If you do this Mars lander design instead with storable NTO-hydrazine (any of them) at around 300 sec Isp, you get a higher propellant weight and a higher inert weight for otherwise the same payload weight. Big deal. Same answer otherwise.

If you keep the exit diameters relatively small relative to the vehicle diameter, then the blast craters the engines want to dig also stay small. Small enough relative to the heat shield diameter you intend to plop down upon. We have no landing legs to stick or sink into the soft sand (landing straight on the heat shield). If rocks up to a meter or two in size get in the way, so what? As long as your vehicle length is less than your heat shield diameter.

So my question is twofold. (1) how can you lose doing it this way? and (2) why has this not already been done?

The answer to the first is open for debate. The answer to the second is too many people still thinking inside too many traditional boxes, when engineering physics does NOT demand that!

Now, NASA ought to understand supersonic retropropulsion better than they let on. They tested its effects upon the Mercury capsule in a wind tunnel back about 1959-1961 (1961 is the date on the NASA report cited as a reference in my Sighard Hoerner "drag bible" from 1965). Retro thrust plumes reduce basic capsule drag a little bit. That was NASA's own data.

What you often see with NASA (and the other players) is what I term "institutional forgetfulness", which is why knowledge obtained by older guys fails to get passed on to younger guys, on-the-job, in ANY institution. Not just NASA. All of them. It's just that this was actually written down, yet still forgotten.

There, does that satisfy what you might want to propose to NASA? I hope so, I need to go outside and start shredding again. I may be retired, but I live on a working farm. There's more I have to do than just sit here and figure things to answer questions.

GW

Last edited by GW Johnson (2020-05-23 15:34:41)

SpaceNut · 2020-05-23 14:50:09

How do the space x falcon engines compare for the firing versus incoming atmosphere which from what from what I remember we can not get an engine to fire up due to incoming pressure. So we must already have them running long before we get deep into the atmosphere in which mars requires a heat shield but does the plume from the running engines do something similar....

As for game changing technology for NASA/Kempton and of the forum.. NASA is the one with space x that killed the retro landing proof of concept for landing larger mass on mars surface as they did not like legs or engine openings that would be made to make landing possible for what was called the Red Dragon mission.

As to the forum we are all in the same boat for hacking, pirating of information; as all can be read unless you shut that off and enforced a logged in credential for being able to even read its content.

Here is the website for the big challenge http://bigidea.nianet.org/judges/

SpaceNut · 2020-05-23 15:11:14

What makes going to mars for humans is found in 6 challenges

The Big Six Challenges That NASA Has To Overcome For Mars Mission

The First Challenge – The Orbit
You can’t just point and shoot when it comes to space. Rocket scientists have to rely on a maneuver known as Hohmann transfer orbit for sending a vehicle from a small circular orbit to a bigger orbit. The transfer orbit needs to be precisely timed so that when the spacecraft leaves Earth, it reaches the destination orbit at the same time that Mars reaches the same position. Earth and Mars only make it to the right orbital alignment for a Hohmann transfer to take place every 26 months for allowing a six-month journey that is survivable.
Going back to the moon helps because it is relatively quite easy for a vehicle that is Mars-bound to make a ‘stop by’ Gateway at its very high orbit at a lower cost in terms of propellant and velocity change.

This where with what we have we are looking at reusability and size of vehicle plus building blocks to allow for a low cost to e incurred for going to Mars to stay.

The Second Challenge – The Long Away Mission
The mission is just too long. Astronauts will have to remain on the Red Planet or orbit it for months at a time. A single Mars mission, from the start till the end, will take more than two years. That is actually quite a long time period to be away from the home planet while breathing recycled air. There will be a forty-minute lag in sending messages to Earth and receiving responses back.
NASA is hoping to acquire more knowledge about how astronauts psychologically deal with spending extended time on the moon. Scientists will also be gauging the effects of microgravity and space radiation on astronauts.

The duration for stay is why we are looking at cost in step one as longer stays require more mass and that is what will cost man when going to Mars.

The Third Challenge – Sending Supplies
In order to deliver big payloads, larger rockets will be required. A Mars lander and the trans-Martian cruise vehicle will have to be built using the payloads that are delivered into the orbit by NASA’s Space Launch System (SLS) rocket and Orion spacecraft over the course of many missions.
NASA will be developing this deep-space capability by relying on the SLS and Orion vehicles for constructing the Gateway lunar orbiting platform and habitats that are based on the moon.

We look at prestaging of supplies for the mission design but some things can only last so long and then we need more. Limiting how often and for what we need is part of the cost lowing concepts.

The Fourth Challenge – Sending More Supplies
Resupplying at Mars is not an easy option. In fact, Mars explorers will have to wait for two years, most likely, and this implies that enough food and fuel must either be brought from Earth or created on Mars.
The lunar base can become the proving ground for technologies that will be used for producing the propellant fuel from water ice that is mined from under the moon’s surface.

Sending what we need for insitu use and for growth on mars is a must and will need to be started within the first steps or we will not be able to stay.

The Fifth Challenge – A Better Space Station
A more versatile space station will be needed. While NASA along with its dozens of aerospace contractors carry on with the work on SLS, a new space-station sized cruiser known as Deep-Space Transport will have to be created to allow for the shuttling of crews between Mars and Earth.
The key systems for DST, including radiation shielding and the closed-loop-life-support will undergo flight testing on Gateway.

Going direct is a space area question area for protection and safety needs with the mass that must be assembled to go to mars but its also going to cost you to build a station to travel in.

The Sixth Challenge – Next-Gen Landers
Making it to Mars means that there will be a plethora of expected and unforeseen problems. The DST will only be able to make it to the orbit of Mars, thus giving rise to the need of a big lander that will be capable of performing a soft touchdown on Mars. This lander will have to be built in space, and Lockheed Martin has already proposed a big space plane that will rely on SLS rocket for launching from Earth.
NASA is hopeful that new moon missions will help it to speed up the mission to Mars by expediting the development of the technologies that are required to get there.

This is the current topics interests as to be able to make use of insitu, we must have started the process with that being the end goal and not just a short one time stay.

GW Johnson · 2020-05-23 15:43:49

When you first come out of hypersonics, the wind pressure still isn't very high (edit update 5-23-20: relative to engine chamber pressure). I see no reason why an engine cannot be ignited at min flow rate, and quickly throttled up. Why would it care what dynamic pressure might be infiltrating backwards up the nozzle? The chamber and injector plate see a local static pressure that is the entry pitot pressure, not anything truly dynamic. (edit update 5-23-20: that is freestream total pressure knocked down by the normal shock total pressure ratio at flight Mach.)

Note that I added a "PS update" to my post 70 above, addressing trajectory kinematics and how many engines to have. This includes engine-out considerations all the way to touchdown. Just how realistic do you want that answer?

As for NASA versus Red Dragon with retropropulsive landing (edit update 5-23-20: and landing legs through the heat shield), the issue is more about "not invented here" than it is about anything technical. Period. End of issue.

The Spacex landings of their Falcon cores PROVES THAT beyond any shadow of a doubt.

Not to mention the one-and-only USAF test flight of a Gemini with a hole sawn in the heat shield. This was the unmanned reflown-test of the Gemini-3 capsule given to USAF by NASA. This was the one and only flight of the USAF "Manned Orbiting Laboratory (MOL)" program back in 1968, before that program was cancelled in 1969. That very capsule is in a museum on public display, AFTER being flown through re-entry with a hatch in its heat shield. You stop throughflow, you survive entry just fine. Period! End of issue!

How much evidence do you want for "not invented here" being an exceedingly-serious problem for NASA? Or Boeing? Or Lockheed-Martin? Or ULA (which literally is both Boeing and Lockheed-Martin)? And don't think for one minute that the expertise from outfits like McDonnell, Douglas, Northrop, Grumman, Rocketdyne, or anybody else still survive within Boeing or Lockheed-Martin! Because it so very definitely does NOT! In acquisitions, what always happens is keeping the upper management team, and firing everybody else. Most of those upper managers are truly technically incompetent. I saw this for too many decades from the inside. It is (unfortunately) too true!

And just WTF do we need another space station for, if we want to go to Mars (or the moon)? This is a whole lot more about businesses and government lab sites within congressional districts, and nothing to do with the technicalities of going outside Earth orbit. Same is true of the SLS/Orion system that NASA itself cannot afford to use, for doing its Gateway lunar space station boondoggle. And useless boondoggle it is.

The TRUE challenges for going to Mars are (1) artificial gravity, (2) solar flare radiation shielding, (3) having well over 300 m^3 volume per person available inside the transit vehicle in order to stay sane during many months confinement, (4) a way to land multiple-dozen-ton objects on Mars (whether from low Mars orbit or from direct entry trajectory makes little difference), (5) some way to get your crew home without killing them, especially since budget-cutting will likely stop the sending of supplies from Earth, and (6) some sort of decision about why we send people all the way there in the first place (is it "exploration", or it is planting colonies or bases)?

NASA has absolutely nothing to do with planting colonies or bases. And never will have anything to do with such! Nor will any other government space agency, anywhere around the world. Better to face THAT, up front!

If it is to be "exploration", then the first thing to do is to DEFINE that word! And note that what we did on the moon during Apollo was NOT exploration! It was engineering flight test. (edit update 5-23-20: in point of fact, the real purpose of Mercury, Gemini, and Apollo combined, was engineering flight test to pull off a political upstaging of Russia.)

GW

Last edited by GW Johnson (2020-05-24 10:42:34)

GW Johnson · 2020-06-07 13:36:33

If you look at the Mars mission articles I posted over at "exrocketman", you can see the evolution of the orbital transport with landers idea into something that could be done with all-existing technology, and for about 10% of cost proposed by the "big boys" to NASA. You can also see that the purpose/focus of the whole mission design is for exploration as I defined it, not the planting of bases or colonies. Although I always left something behind as the core of a possible base.

What I got to was a recoverable and reusable orbital transport, fueled by storable propellants. Only some of the emptied tanks were jettisoned. The Earth departure stage was LOX-LH2, jettisoned, although it could be resized for recovery. Crew radiation protection, centrifugal artificial gravity at a tolerable spin rate, and a huge available volume per person, were all available in my rough-sized designs.

It carried enough propellants to reach Mars and enter orbit there, and refill with propellants sent ahead from Earth for the return home. The return trip included enough propellants to capture into Earth orbit for recovery, refurbishment, and reuse on another mission. I included a free-return capsule as a "way out" for the crew in the event the transit vehicle propulsion failed to do the Earth arrival job. That capsule was the crew ferry home from Earth orbit in the nominal case.

This thing was intended to base out of low Mars orbit, but there is not a reason in the world it could not base from Phobos.

The landers were single-stage reusable chemical rocket vehicles, sized by the sum of EDL and ascent requirements. The ascent requirements were far larger, and I met them by brute force, plus some ascent payload reduction reflecting the use of supplies while on the surface. I did this with the same storable propellants and engine sizings that I used for the transit vehicle. They are the surface habitat for crews while on the surface. And they are the "Mars ascent vehicles", all in the one hardware design.

You send multiple landers to cover rescue capability for a crew stranded on the surface by a lander failure, and to cover the possibility that one quits for no damned reason other than Murphy's Law. My mission plan for a crew of 6 has 3 below for a month at most, with 3 in orbit doing science up there while watching over the surface crew. The orbit transport spins for 1 gee artificial gravity while in orbit.

Landers and the lander propellant supply, and the transport's Earth return propellant supply, all get sent ahead unmanned to Mars orbit using electric propulsion. Radiation crossing the Van Allen belts is no risk to unmanned assets. You don't send the men until all this crap is docked, or at least flying in formation, in low Mars orbit. You send enough to make 6 to 20-something landings. I usually planned on 12 landings at up to 12 different sites. Every person in the crew gets to go to the surface. No one has to go all the way to Mars and never set foot there.

This whole thing is sent from Earth, including all the life support supplies. I assume not one blessed thing in the way of local propellant manufacture, local water and oxygen production, or the growing of any food. You try all those things, yes, but the mission assumes they all fail. So any successes with any of that stuff are just "gravy".

THAT is how you design-in a "way out" at every phase of the mission. And, it is EXACTLY HOW you get human ground truth for local resources from up to 12 different sites around Mars, all in the one human trip there! Is that not the VERY ESSENCE of exploration? When exploration is defined to be answering the question "what all is there, and where exactly is it?".

There is no reason at all why this mission could not be based from Phobos instead of low Mars orbit. The landers get bigger because the delta vee requirements are about 40-50% bigger. Not much else needs to change.

Tahanson43206 has talked about the advantages of having a "there" there, where you base your activities. I'm not so sure about the emotional value of having a physical object to point at, but having something big enough to see and to land on, does make good sense to me. My mission plan is all about rendezvous and docking or flying in formation. Landing on Phobos makes that easier to do, actually.

So if we want to put together a mission proposal to submit the way Tahanson43206 has suggested, then my input to that activity is let's modify my Mars mission plan to land on Phobos instead of rendezvousing in low Mars orbit. The one I am talking about is "Mars Mission Outline 2019", dated 14 September 2019, located on http://exrocketman.blogspot.com.

GW

Last edited by GW Johnson (2020-06-07 14:07:46)

tahanson43206 · 2020-06-07 18:02:02

For GW Johson re #74

SearchTerm:ExplorationMission

For SpaceNut ... do you have any members or former members in your contact list, who could help to write up GW's work here?

For all ... please think about folks you know who could contribute to development of a proposal to be submitted to NASA.

The theme of designing an expedition so that nothing depends upon in situ resource utilization (for safe return of the crew) should appeal to many who consider the proposal. The potential bonus of successful exploitation of local resources will surely add to the interest value of the concept.

RobertDyck ... Any chance GW's suggestions here might be of interest to your audience at the next Mars conference?

(th)

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#51 2020-05-17 19:05:41

Re: Solving Mars mission docking with Phobos

#52 2020-05-17 19:46:39

Re: Solving Mars mission docking with Phobos

#53 2020-05-17 21:06:57

Re: Solving Mars mission docking with Phobos

#54 2020-05-18 09:06:25

Re: Solving Mars mission docking with Phobos

#55 2020-05-18 21:17:10

Re: Solving Mars mission docking with Phobos

#56 2020-05-19 06:24:54

Re: Solving Mars mission docking with Phobos

#57 2020-05-19 08:25:59

Re: Solving Mars mission docking with Phobos

#58 2020-05-19 10:57:57

Re: Solving Mars mission docking with Phobos

#59 2020-05-19 15:30:35

Re: Solving Mars mission docking with Phobos

#60 2020-05-21 08:26:49

Re: Solving Mars mission docking with Phobos

#61 2020-05-21 11:49:38

Re: Solving Mars mission docking with Phobos

#62 2020-05-21 18:40:31

Re: Solving Mars mission docking with Phobos

#63 2020-05-22 04:57:26

Re: Solving Mars mission docking with Phobos

#64 2020-05-22 16:16:40

Re: Solving Mars mission docking with Phobos

#65 2020-05-22 16:42:41

Re: Solving Mars mission docking with Phobos

#66 2020-05-22 18:18:01

Re: Solving Mars mission docking with Phobos

#67 2020-05-22 18:38:21

Re: Solving Mars mission docking with Phobos

#68 2020-05-23 08:18:20

Re: Solving Mars mission docking with Phobos

#69 2020-05-23 09:12:16

Re: Solving Mars mission docking with Phobos

#70 2020-05-23 12:27:03

Re: Solving Mars mission docking with Phobos

#71 2020-05-23 14:50:09

Re: Solving Mars mission docking with Phobos

#72 2020-05-23 15:11:14

Re: Solving Mars mission docking with Phobos

#73 2020-05-23 15:43:49

Re: Solving Mars mission docking with Phobos

#74 2020-06-07 13:36:33

Re: Solving Mars mission docking with Phobos

#75 2020-06-07 18:02:02

Re: Solving Mars mission docking with Phobos

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