Solving Mars mission docking with Phobos

SpaceNut · 2020-06-07 18:13:25

This is the http://exrocketman.blogspot.com/2019/09 … -2019.html that GW indicates.

as indicate Murphy's law requires planning to accomplish regardless of what hardware used and assembly required there in for use to get men to mars and back.

AS GW has indicated we could go today with the combinations of what we already have once assembled in space for the trip. Our many topics have shown that we could if it were not for the funding issue.

That said we need the list of all parts and pieces with current values to assemble the task of man going to mars as that's what it will take to convince those that read a report structured with those parts with the numbers to back it up.

GW Johnson · 2020-06-08 08:17:21

What I like about landing on Phobos is relief from stringent docking requirements or long-term formation-flying requirements.

How about prepositioning the return propellant tanks, the lander operation tanks, and the landers themselves as items landed on Phobos? The orbit transport with the crew then goes to Phobos, picks up the landers and the landing propellant supply, and moves to low Mars orbit. From there, the mission runs just like I had it planned as based from low Mars orbit.

At end of mission, land the landers robotically at the best site or sites for a base. That leaves a usable habitation to serve as the core of a future base or colony mission. Then take the transport plus any unused propellant back to Phobos.

The transport leaves empty tanks on Phobos (to eliminate lost materials as space junk), adds the return propellant tanks, and departs from Phobos for Earth.

This way, the landers need not be redesigned for the more demanding ascents all the way to Phobos. All that is needed is more dumb propellant tanks to take care of moving (some not all of the) stuff between Phobos and low Mars orbit.

The one-stage reusable lander design is the critical design item here, not the orbital transport. The transport is nothing but a big propelled space station type of device. With places all over its exterior to bolt on and connect propellant tanks. Those tanks are part of the radiation shielding.

GW

Last edited by GW Johnson (2020-06-08 08:20:45)

tahanson43206 · 2020-06-08 09:09:02

For GW Johnson re #77 and topic

I'm interested in helping you develop this concept into a full NASA proposal.

To my knowledge we have no one in the active membership who has experience preparing such a proposal, but we ** do ** have encouragement from a NASA researcher with the required experience.

For new readers of this forum, please see: http://newmars.com/forums/viewtopic.php … 21#p168221

What we ** can ** do is to study existing proposal submissions, and seek out persons with the needed experience to provide coaching.

I see no reason to let this situation sit untended.

***
New item for GW Johnson ... please take a quick look at Calliban's new topic suggesting use of un-liquefied oxidizer for a home build Mars launcher.

While Calliban himself pulled back from the sheer boldness of his original idea, there is a possibility that if you combine it with SpaceNut's contribution of a bellows design for the gas container, and the (to me quite new) concept of a central collapsible spine for the vehicle, there might be the kernel of a very-low-cost vehicle design unique to the Mars situation.

One refinement I'm picking up from Calliban's hesitation is the idea of using kerosene (or similar liquid-at-room-temperature) propellant for the non-oxidizer component of the system.

For new readers of the forum, the critical element of the Calliban inspired design is use of atmospheric pressure on the nose cone to push oxidizer into the reaction equipment, by collapsing the forward section of the vehicle along its collapsible spine, compressing the bellows gas container.

(th)

GW Johnson · 2020-06-09 08:01:35

Tahanson43206:

With regard to the balloon rocket idea, I have not run any numbers. My hunch, or impression, going into this is that you will have very low rocket chamber pressure indeed, and therefore very low thrust, because thrust is Pc At CF, where Pc is very low, and CF depends upon pressure ratio to ambient, not Pc alone.

You cannot use the usual rocket experiences for c*, to get Isp = CF c* /gc, because combusted temperature and gas properties vary with combustion pressure, significantly. This shows up as correlations for c* (both theoretical and experimental) that vary as Pc^m where m is a number on the order of .01 to .05. I have never tried to extend these correlations below about 2 atmospheres. But the effect at 2 atm is quite the reduction, from 1000-psia values typically reported..

The pressure you can achieve upon the gas balloon is achieved only on the windward surfaces, and is some fraction of the difference between the (shocked if supersonic) pitot pressure and ambient static. That pressure must be larger than your chamber pressure, or the gas will not move from the balloon into the chamber. You'll be lucky indeed to achieve a compression pressure/ambient static ratio exceeding about 2 to at most maybe 3. You're probably talking about no more than 18 mb chamber pressure at the surface at supersonic speeds, and less at altitude. But I'd have to run some numbers to know.

The drag of this thing is a CD q Aref, where Aref is usually the frontal cross-section area. CD is a number somewhere between 0.2 and 2, varying with speed and shape. q is the dynamic pressure, which is 0.5 gamma P M^2, where for CO2 gamma will be about 1.33, and P is the ambient static pressure. M isMach number, for a sound speed on the order of 240-250 m/s on Mars. Although it varies a lot with the square root of the absolute temperature, which varies wildly on Mars.

All in all, my best hunch is that the thrust potential will always prove to be less than the drag. But as I said, I have not run any numbers.

GW

Last edited by GW Johnson (2020-06-09 08:04:36)

tahanson43206 · 2020-06-09 08:37:20

For GW Johnson re #79 and Calliban's gaseous oxidizer topic

My guess is that Calliban may regret having offered the balloon propellants idea at all, since he very quickly realized its impracticality.

However, SpaceNut picked up on it (as is his role as site evangelist) so off we go to the races, or at least, to the first hurdle.

The first hurdle may have been overcome, because with Calliban's caveat, the vision seems to have changed so that the fuel is liquid.

We are then left with the question of the oxidizer, and specifically whether it is practical to consider storing the oxidizer as a gas.

It is that question which (I think) your post addresses.

I am going into this with a ** very ** low expectation of a successful outcome, because the considerations you have (very kindly) laid out are significant and (to me at least) daunting.

I'll lead into whatever this thought process is going to be, by reminding the readers of the forum who may chance upon this topic unawares, that on Earth, the use of a gaseous oxidizer with a liquid fuel is a fairly well established art.

However, as GW Johnson has reminded us, there exist NO examples of pressure fed vehicles on Earth that feed oxidizer into a combustion chamber.

There ** are ** examples of pressure fed vehicles, but those are NOT using combustion. Instead, the stored energy of the compressed gas is fed to mechanical systems that translate the potential energy of the bouncing gas molecules into piston movement which feeds into rotary motion.

Those examples use extremely ** high ** pressure in the "propellant" tank and that is NOT the situation that Calliban's topic is about.

However, on Earth, gaseous oxidizer is fed into combustion chambers under pressure sufficient to meet the needs of the engines.

The mechanism by which gaseous oxidizer is prepared for combustion is primarily mechanical (to the best of my knowledge (*)).

A long established mechanism is the use of a piston to achieve needed pressure levels. In more recent times, rotary compression equipment has proven capable of achieving the needed pressures required by the engine combustion process.

Thus, I am led to inquire (the word propose seems overly ambitious) if Calliban's vision might be realized by adding a turbine pump to the mix.

I think that a means of applying external pressure to the oxidizer bladder remains a useful capability, and the sliding nose cone ring is a possible mechanism.

Assuming for a moment that a turbine pump can supply oxidizer to the engine (it would be more accurately described as a jet engine) at a rate sufficient to achieve orbit from a ground launch at Mars, we are now left with the question of whether it is (energy) cost effective to depart so dramatically from the Earth-way-of-doing-things.

I'll copy this post and drop it into Calliban's topic, for the sake of continuity there.

(*) Caveat to statement about gas fed into engines ... It is a stretch, but fuel cells do in fact accept gaseous oxidizer and fuel. However, that example seems inapplicable to Calliban's topic.

(th)

Last edited by tahanson43206 (2020-06-09 08:40:10)

SpaceNut · 2020-06-09 17:05:50

I have given a bit more thought for the inflatable tank rocket in that topic so thank you for the follow up GW.

The main issue if fuel temperature and other such aspects to making it work for mars as well as other attributes.

So start with an engine pod and payload mass with the numbers to trade for metal tanks for a given payload to bring to orbit.

Bounding these will allow for the tanks to turn flexible and inflatable as we would then know the mass to target for those to work.

GW Johnson · 2020-06-12 15:51:43

I looked at the thrust and drag numbers. That stuff is in the balloon rocket topic in the interplanetary transportation group. I didn't like what I found. The main problem is millibars instead of megabars for chamber pressure.

GW

tahanson43206 · 2020-06-13 07:16:01

For GW Johnson re #82 and work on gaseous oxygen idea in Calliban's topic ...

Thank you for your work on this idea. I have offered some feedback in Calliban's topic.

However, I am ** here ** to offer the suggestion that your lander will need to be refueled on Mars, so the folks on duty there (or perhaps the robots guided by teleoperators on Phobos) will need to make fuel and oxidizer for your vehicle.

It would be helpful to know how much fuel and oxidizer you would be buying (from a supplier on the surface) for a return flight to Phobos.

I am interested in kerosene, although (at this point) I have no idea if it can be made from methane on Mars. I **know** that methane can be made on Mars, because there has been consensus on that point from Zubrin and many others for a number of years.

Kerosene has the distinct advantage of being a storable liquid that does not require pressure vessels on Earth. It just occurred to me (somewhat belatedly) that kerosene may require pressurization on Mars. In that case, the question would be whether there might be other advantages of working with a liquid rather than a gas for launch. It may turn out that there are no advantages of going to the extra effort of making kerosene on Mars, and that methane is the optimum fuel for that environment.

However, MORE to the point in thinking about your purchase of supplies from vendors on Mars, the amount of liquid oxygen you will need should be knowable, because you will have designed your vehicle with cargo capability and performance in mind, and chosen the materials and their mass for the structure.

From ** that ** (quantity of LOX) your vendor will know how much oxygen to make from CO2, in anticipation of your vehicle's arrival.

And ** that ** leads to the point I am trying to make ... Regardless of how the investigation of a balloon enclosure for oxidizer for a rocket turns out, the vendor on Mars will want (I feel confident) to hold oxygen extracted from the atmosphere in gaseous form until the time has come to start liquefaction. Energy costs for maintaining a quantity of LOX will add up, so the vendor will want to try to schedule liquefaction as closely as possible to your flight schedule.

In the mean time, the quantity of volume for storage of gaseous oxygen should be knowable, and (by coincidence) that volume should be quite close to the volume that would have been needed for the balloon rocket if it had been practical.

(th)

elderflower · 2020-06-13 07:23:06

I would suggest the use of benzene as fuel. It uses a lot less of your precious hydrogen than methane and has about the same ISp, and you don't have to refrigerate it. The main drawback is that it is a severe carcinogen.

tahanson43206 · 2020-06-13 07:55:46

For elderflower re #84 ...

This is GW's topic so I'll just note the point you made, about increasing the ratio of Carbon (which is abundant) to Hydrogen, which is not.

I came back to follow up on my intuition that kerosene would boil if not contained in a pressure vessel on Mars.

Google quickly came up with useful data to confirm the intuition, although the situation is not as bad as I had feared.

Liquids - Vapor Pressure
Fluid Name Vapor Pressure (kPa)
Heptane 6
Hexane 17.6
Isopropyl alcohol (rubbing alcohol) 4.4
Kerosene 0.7
51 more rows
Vapor Pressure common Liquids - Engineering ToolBox

And, for Mars itself:

Compared to Earth, the atmosphere of Mars is quite rarefied. Atmospheric pressure on the surface today ranges from a low of 30 Pa (0.030 kPa) on Olympus Mons to over 1,155 Pa (1.155 kPa) in Hellas Planitia, with a mean pressure at the surface level of 600 Pa (0.60 kPa).
Martian Atmosphere – Planetary Sciences, Inc.

So a refueling station in Hellas Plantia would have a modest containment requirement for Kerosene

***
For elderflower ....

If a vendor were to consider manufacture of fuels for GW's lander, there would be several factors to think about (that I know of).

No doubt there are others.

1) The energy required to extract Carbon from the atmosphere (varies by elevation?)
2) The considerations for storage .... Is it better to move from CO2 to the target compound directly, and avoid intermediate storage?
3) Long term storage of fuel while awaiting customers
4) Transportation issues if fuel manufacturing facility is some distance from landing and takeoff locations.

If you were advising a venture fund on how to supply GW's landers, what would you suggest.

Your suggestion of benzine is interesting.

Google came up with this comparison:

https://www.itrcweb.org/PetroleumVI-Gui … roleum.htm

The article at the citation mentions the volatility of Aliphatics as compared to Aromatics including Benzene.

If you were advising a vendor, or GW for that matter, how would you lean when considering the risks and benefits.

(th)

GW Johnson · 2020-06-13 08:55:43

I'm a bit unsure why we are talking about refueling "GW's landers". What I have done (so far) is point at an orbit-based design-sizing study that uses an orbit-to-orbit transport design and an orbit-based two-way/one-stage/reusable "landing boat" lander design. That was in my 2019 version of the Mars mission study posted over at "exrocketman". Then I looked at the effect of one-way and two-way travel from Phobos to the surface upon that sort of a lander design. And I did not like what I saw. The delta-vee requirements upon the lander were too high for any sort of practical one-stage/two-way design.

Then I came up with the (as yet unevaluated) idea of basing the orbit mission upon Phobos instead, and leaving the Earth return assets there. Then bringing the transport and the lander assets from there to low Mars orbit. This would increase the delta-vee requirements on the transport, but not the landers. After mission conclusion, the transport returns to Phobos, refuels, and departs for Earth.

But in all of these, I looked at hydrazine-NTO storables brought from Earth to power all the needs. Anything else with higher Isp just looks that much better. So, think of all this as a sort of bounding analysis that shows the minimum performance, not the maximum.

With storables, you can do something in your tankage designs that you simply cannot do with cryogenics (including liquid oxygen). You can use an elastomeric expulsion bladder inside your tank shell. All that is needed is low gas pressure outside the bladder to make its contained volume exactly that of the liquid propellant still inside the tank. That eliminates the zero-gee ullage problem entirely.

You cannot do that with cryogenics. Not even LOX. There are no elastomers fit for service at such temperatures. There might (or might not) be such materials in future centuries, but not now. You much put small thrust on the vehicle to settle the propellant globules at one end of the tank so you can suck them up for the engine pumps. THAT IS the ullage problem.

That small thrust has three known sources at this time: small solid-propellant rocket cartridges (ullage motors), thrusters powered by bladder-expelled storable propellants, and cold-gas thrusters. Those are the only known solutions to the ullage problem at this time in history. And those solutions have been known to work for over half a century now.

For an exploration-type mission, the kind of orbit-based mission I have been talking about makes the most sense in terms of most sites explored for the enormous effort of making the one trip. You literally do not know the up-close-and-personal ground truth about the local resources, or it would not be exploration! Period! Ethically, you cannot bet lives on what you are uncertain about, so you MUST send everything needed to keep the crew alive and get them home FROM EARTH! You are ethically compelled to do that.

Once you know what is really there at some site, and what it is going to require in order to use that resource (or resources), then (and only then) you can count on them for life support and propellants, etc. That's when you are doing bases or colonies. That's a DIFFERENT mission ballgame! The cheapest transport for that is direct entry from the interplanetary trajectory, because the delta-vee to land is 90+% done with aerobraking, not rocket thrust. Once you are in that situation, it makes the most sense to make the return propellant right there on Mars.

Why? (1) Because you know you can, with the resources you know are there, you know in what condition they are, and you know they are there in what quantities. (2) Unless you can make enough propellant fast enough, then the vehicle designs cannot have practical payload fractions or single-stage reusability. It really is that simple.

What Musk's Spacex Starship/Superheavy does is provide a base/colony planting vehicle. His design is a one-way suicide trip unless he can actually make massive quantities of propellant fast enough to support return trips in a timely fashion. And if the ships sent there do NOT return, they are by definition NOT reusable, and therefore expensive! Best to face up to that from the outset. Then you can deal with it.

Now, you can misuse the Musk concept to do the initial exploration missions. (And that is what he wants to do.) It does limit the exploration you can do. You only get to land on, and attempt to exploit the resources, at one site. (The orbit-based approach can visit several in the one trip.) The first few Musk ships sent there are one-way. (The orbit-based mode offers the possibility of using the landers multiple times, even if the resources prove unexploitable.) And unless you stoop to sending suicide missions, in Musk's approach, the initial explorations must be done robotically. Yet men can cope with so much more unexpected stuff than robots, which severely limits what you do with an unmanned Starship on Mars.

Nobody is really going to build my orbit-based exploration mission, not with Musk's Starship on the horizon. I recognize that. The only practical thing I see to do is land an appropriate robot lander at the intended site or sites. The critical resource is water for electrolysis. That enables breathing oxygen, drinking water, and LOX production. Wringing moisture out of damp dirt is NOT the way to do this. Processing tens to hundreds of tons of regolith to get ounces to pounds of water is just NOT the way to manufacture 1200 tons of propellant (80% of which is LOX and 20% is LCH4) in only 2 years! Period! You might as well face that!

You need buried glacial ice you can mine, preferably by drilled-well techniques to "get the mostest the fastest", by steam extraction. While you might tolerate some salt in it, you probably cannot tolerate perchlorates in it. And you CANNOT know things like that until you drill and try! You need a robot lander that can do EXACTLY those tasks. And I have NEVER seen one like that proposed by anybody. Plus, there is no longer any possibility of sending a Red Dragon there to do that. That leaves Musk's Starship. No one else has anything to offer in the way of a concept.

If Musk has really thought this through, he could send such a robot driller and processor in an unmanned Starship to Mars, along with a subscale Sabatier propellant production plant to try out making methane. Where he will get these things, no one knows! How he gets them as quickly as he wants them, nobody can say! But he needs to send one-way unmanned Starships equipped with exactly those types of robots to candidate landing sites, until he finds the site where recoverable water exists and demonstrates processibility.

That's a lot of one-way/nonreusable Starships! But that's just the cost you (knowingly!! ?? !!) incur when you jump straight to the base/colony-planting vehicle, skipping the initial exploration vehicle. Which is fundamentally quite the different vehicle.

But ONLY THEN is it ethical to send men and supplies to set up the base/colony/whatever at that best site.

However, I will say this, too. It is far more likely that Musk will send crews to Mars in the mid 2030's than it is that NASA ever will, and by the 2040's at the earliest. Probably well after 2050, now that the Gateway boondoggle is going to dominate their budget for many years.

GW

Last edited by GW Johnson (2020-06-13 09:48:14)

tahanson43206 · 2020-06-13 11:59:40

For GW Johnson re #86

Thank you for your detailed clarification of the exploration mission ... I admit to having overlooked this point:

But in all of these, I looked at hydrazine-NTO storables brought from Earth to power all the needs. Anything else with higher Isp just looks that much better. So, think of all this as a sort of bounding analysis that shows the minimum performance, not the maximum.

That specific point was probably clearly made in the original posts, but I admit to having missed it, or not let it register.

Google came up with this:

around $100/lb
Typical cost is around $100/lb. The real cost of hydrazine is probably in the protective measures we take to deal with it.Jun 20, 2017
Hydrazine - Toxic for humans, but satellites love it. | Iridium

That quote was from www.iridium.com

I was curious to see what's involved in making hydrazine, so Google came up with this:

Showing results for manufacture of hydrazine-INTO
Search instead for manufacture of hydrazine-NTO
Search Results
Featured snippet from the web
The Raschig process, the original commercial production process for hydrazine, involves oxidation of ammonia to chloramine with sodium hypochlorite, then further reaction of the chloramine with excess ammonia and sodium hydroxide to produce an aqueous solution of hydrazine with sodium chloride as a by-product.
4. PRODUCTION, IMPORT, USE, AND DISPOSAL 4.1 ...

Is this correct for NTO: ?

Did you mean: define not nitrous tetroxide
Search Results
Web results
Dinitrogen tetroxide - Wikipedia
en.wikipedia.org › wiki › Dinitrogen_tetroxide
N verify (what is ☑ Y ☒ N ?) Infobox references. Dinitrogen tetroxide, commonly referred to as nitrogen tetroxide, and sometimes, usually ... In the first step, the ammonia is oxidized into nitric oxide: 4 NH3 + 5 O2 → 4 NO + 6 H2O. Most of the ...
Production · Use as a rocket propellant · Chemical reactions

Can you clarify the shipping conditions for these materials, in an expedition heading to Mars?

It would appear that a bladder would be used in the lander. I assume larger ones would do the job in the expedition vehicle(s).

Dinitrogen tetroxide - Wikipedia
en.wikipedia.org › wiki › Dinitrogen_tetroxide
Dinitrogen tetroxide, commonly referred to as nitrogen tetroxide, and sometimes, usually ... In the first step, the ammonia is oxidized into nitric oxide: ... most important rocket propellants because it can be stored as a liquid at room temperature.
Production · Use as a rocket propellant · Chemical reactions
Nitrous oxide - Wikipedia
en.wikipedia.org › wiki › Nitrous_oxide
Its high density and low storage pressure (when maintained at low temperature) enable it to be highly competitive with stored high-pressure gas systems. In a 1914 ...

Given that hydrazine and NTO would be your preferred energy storage materials for the lander(s), can I assume you'd employ the same material for the expedition vehicle(s)?

I would think that policy would simplify design, procurement and operation.

The risks seem significant (to me at least) but they are (have been) taken routinely ....

The descent propulsion system (DPS - pronounced 'dips') or lunar module descent engine (LMDE) is a variable-throttle hypergolic rocket engine invented by Gerard W. Elverum Jr. [1] [2] [3] and developed by Space Technology Laboratories (TRW) for use in the Apollo Lunar Module descent stage. It used Aerozine 50 fuel and dinitrogen tetroxide (N
2O
4) oxidizer. This engine used a pintle injector, a design also used later in the SpaceX Merlin engine.

It seems to me that a space faring nation ** other ** than the United States might be a better bet for your expedition proposal.

India has achieved a notable success in orbiting Mars (I believe) on its first attempt:

Mars Orbiter Mission (MOM), the maiden interplanetary mission of ISRO, launched on November 5, 2013 by PSLV-C25 got inserted into Martian orbit on September 24, 2014 in its first attempt. MOM completes 1000 Earth days in its orbit, today (June 19, 2017) well beyond its designed mission life of six months.

Mars Mission: India creates history as Mangalyaan successfully enters Mars orbit in first attempt
India has become the first nation in the world to have entered the Mars orbit in the first attempt. ISRO's MOM is also cheapest such mission till now.
Read more at:
https://economictimes.indiatimes.com/ne … aign=cppst

I'd like to try to encourage you to build out your design for an expedition. There might be a nation that would be interested in adapting it to their culture and capabilities.

Regarding the design considerations that (as I understand your lander return plan) require a two phase process to deliver the vehicle safely to Phobos ...

Is it feasible to plan for orbital refueling for the lander?

The alternative (as I understand your plan) is to place the expedition vehicle (or one of them) in low Mars orbit, so the lander can reach it.

An alternative to ** that ** plan is to simply meet the lander with a tender as soon as it reaches low Mars orbit, and either refuel it or attach to it and boost it up to Phobos.

That would split your lander concept into two vehicles. The orbit-only tender could perform other missions of exploration between lander recovery missions.

Edit#1: The tender could also serve as an emergency recovery vehicle, in case the lander is able to make orbit with passengers but unable to continue on for some reason.

(th)

Last edited by tahanson43206 (2020-06-13 12:02:26)

GW Johnson · 2020-06-13 12:26:51

On hydrazine-NTO systems:

There's more than one hydrazine. There's hydrazine N2H4, unsymmetrical dimethyl hydrazine UDMH, mono-methyl hydrazine MMH, and Aerozine-50, which have all been used. Aerozine-50 is a 50-50% blend of UDMH and straight hydrazine. The ULA-related outfits like Aerozine-50, while Spacex and a few others prefer MMH. 6 of one, half a dozen of the other, really. The differences show up mostly in vapor pressure, BP, and FP. And very modest effects upon Isp. The recommended treatment for a spill is massive water dilution, just like with ammonia.

NTO is just N2O4. It has vapor pressure and BP, FP in more-or-less the same range as the hydrazines. It is more poisonous to handle than any of the hydrazines, being about as dangerous as military nerve gases, albeit not the same mechanism. You wear a full p-suit and self-contained breathing gear when you are around it. Handling any of the hydrazines is about like handling anhydrous ammonia. Farmers do that. If you are suited-up for NTO, you are more-than-adequately protected against hydrazine.

The other proven storable oxidizer is IRFNA (inhibited red fuming nitric acid). That stuff is typically 83% nitric acid HNO3, 2% water H2O, and 15% NTO N2O4. It's more corrosive than it is toxic, but it is toxic. Isp performance is lower than NTO straight, for the same fuel at otherwise the same pressure and expansion. There are other known oxidizer liquids that are not cryogenic, but they are not proven by long experience except for high-test hydrogen peroxide H2O2. And that simply is NOT long-term stable. I also put RP-1 kerosene in the list as the other storable fuel.

item FP, F BP, F sp.gr.
N2H4 34.7 236.3 1.004
MMH -62.3 189.5 .874
UDMH -71.0 146 .786
Aero50 18.8 158 .898
NTO 11.8 70.1 1.433
IRFNA -63.4 150 1.555
RP-1 -76 390** .810 **at the 10% evaporated point on the distillation curve

For sea level designs at 1000 psia, the Isp performance potential of these follows. The r parameter is the oxidizer/fuel ratio by mass. r/(r+1) is your fraction of oxidizer out of total propellant. 1/(r+1) is your fuel fraction. The reference density for specific gravity is 1 g/cc = 1 kg/L = 1000 kg/cu.m = 62.43 lbm/cu.ft = 0.0361 lbm/cu.in. Divide mass by density to get the volume required.
fuel oxidizer r Isp, s
N2H4 NTO 1.33 292
UDMH NTO 2.6 287
MMH NTO 2.17 288
Aero50 NTO 2.0 287
UDMH IRFNA 3.1 272
RP-1 IRFNA 5.0 263

With NTO, all of the hydrazines are hypergolic ignition-upon-contact. With IRFNA, both the hydrazine and the RP-1 are hypergolic ignition-upon-contact. I didn't show RP-1-H2O2 because the peroxide is just NOT long-term stable at rocket-grade concentrations. However, that combination can be hypergolic. It's not always fully reliable as hypergolic, but such designs can be made reliably hypergolic.

The only other storable liquid fuel is ethyl alcohol as was used in the V-2 in the 1940's, and the Redstone/Jupiter family in the 1950-60's, long ago. Most folks are uninterested in that one anymore. No one has used it since then. It boils at 173 F, and its freezepoint is well below -110 F.

Almost none of these propellants would be feasible to make on Mars in the scenario of a base or small settlement there. Either LOX-LH2 or LOX-LCH4 are the propellant combinations actually feasible to make on Mars for decades to centuries yet. Bear in mind that handling LCH4 is in most ways just about like handling LH2, except the tank sizes are somewhat smaller. At BP, the sp.gr of LCH4 is 0.424, while that of LH2 is .071. LOX is 1.143. At 1000 psia for sea level designs, the r for LOX-LCH4 is 3.15, and LOX-LH2 is 4.0. Isps list as LOX-LCH4 310 s vs LOX-LH2 388.

There's buried ice on Mars, if you can find it, extract it, clean it up, and electrolyze it. All big if's, but nothing insuperable. There probably is no ice on Phobos, that likely being a dry rubble-pile asteroid. Given decent water, that makes manufacture of hydrogen and oxygen feasible. Mars has a largely-CO2 atmosphere. It's too thin for conventional compression to be effective, but there are ways to purify it and compress it. That makes methane manufacture feasible, given the water, too.

If you are making propellants on Mars, then LOX-LH2 or LOX-LCH4 are the best choices. And with high performance, too. If you are operating a vehicle from the surface to Phobos, then you have surface-orbit delta-vee, a delta-vee to the transfer ellipse, and a delta-vee to circularize adjacent to Phobos. The landing delta-vee at Phobos is trivial. That sum is in the same ballpark as escape from Mars (about 5 km/s ballpark).

Returning to the surface, there is a trivial escape from Phobos, a delta-vee into the transfer ellipse, a delta-vee to circularize in low Mars orbit, a trivial de-orbit burn, and the final landing burn. Aerobraking kills most of the low Mars orbit velocity, so you can afford a lot more payload during descent. That total delta-vee is about twice the difference between Mars escape and low Mars orbit (2x 1.5 km/s = 3 km/s ballpark).

You'll be refilling only on the surface. Payload up to Phobos is restricted below max, payload down to Mars is not restricted. Your Isp levels are high, so payload fractions will be high enough to be attractive in a one-stage vehicle. 8 km/s is quite doable single stage with LOX-LH2 or LOX-LCH4, especially if the ascent payload is light.

Now, my orbit-orbit mission used NTO-MMH, and it used the same engine designs in the landers as for the orbit transport, except for the LOX-LH2 departure stage. My design constraints were thus quite different for that scenario than this one we are discussing here. As everyone should be aware, optimization is generally all about constraints.

Those were storable propellants, lower performance. The vehicle was sized to make the descent at full payload, counting on aerobraking so the fairly small landing burn dominated, and then make the ascent with only a slightly-reduced payload, where the ascent delta-vee dominates. Refilling was on-orbit from supplies sent from Earth.

GW

Last edited by GW Johnson (2020-06-13 13:18:14)

tahanson43206 · 2020-06-14 07:28:49

For GW Johnson re #88

You have already provided a search term that is likely to be helpful.

I'll formalize it here: SearchTerm:hydrazine-NTO

We have a Hohmann window coming up in July of this year. NASA is planning a launch for sure, and it is possible other nations are planning to catch this one as well.

China set to launch Mars probe and rover mission in July. China's space program will launch a Mars mission in July, according to its current plans. This will include deploying an orbital probe to study the red planet, and a robotic, remotely-controlled rover for surface exploration.May 25, 2020
techcrunch.com › 2020/05/25 › china-set-to-launch-ma.
China set to launch Mars probe and rover mission in July ...

The next window will occur in August of 2022 (per http://clowder.net/hop/railroad/EMa.htm)

Your vision of an expedition able to insure a safe return for the crew ought to be of interest to multiple nations.

I'm hoping to encourage you to think about how we (forum members) might be able to help you achieve some visibility.

An article in some space oriented publication, or even something as generic as Popular Mechanics.

I just confirmed that Popular Mechanics is still in print. They are down to 6 issues a year.

Popular Mechanics, published by Hearst, currently publishes 6 times annually. Your first issue mails in 8-10 weeks. Cover price is 8.99 a copy. Wood Magazine, published by Meredith, currently publishes 7 times annually.
Magazines.com www.magazines.com › popular-mechanics-magazine

Popular Science is still in existence, and it too has cut back to 6 issues per year:

In January 2016, Popular Science switched to bi-monthly publication after 144 years of monthly publication. In April 2016 it was announced that editor-in-chief Cliff Ransom would be leaving the magazine.
Popular Science - Wikipedia

I'd like to offer a suggestion to consider ...

Designing a mission with BOTH safety assurance (ie, return not dependent upon insitu resources) which your hypergolic design provides, AND the liberal inclusion of various techniques to draw upon energy and material supplies likely to be available ought to increase the interest value of the proposal for a variety of audiences.

In addition, developing a plan that imagines a global cooperative venture ought to be of interest to the generation coming along, who appear to be interested in setting the squabbles of the past aside.

National entities that should be interested in a joint expedition in 2022 would (should) include:

The US
Russia
China
India
European Union (France)
Japan
Israel

For completeness, Iran and North Korea are listed by www.spaceanswers.com as having space flight capability.

A comprehensive mission plan would/should include cooperative and mutually supportive agreements between and among participating nations.

The lead nation ** should ** be the United States, if it can (somehow) get its act together in the next few months.

(th)

Last edited by tahanson43206 (2020-06-14 07:29:29)

GW Johnson · 2020-06-14 10:00:21

Well, if one allows a nominal 5 years to develop and prove-out a lander vehicle and a transport vehicle, or for Spacex to prove-out its vehicle in both roles, then we humans are just not ready to go until 2025. That makes the 2026 opposition the "target" for going. To me, shorter timelines are just not at all credible. They could be longer, but I am trying to be optimistic here.

Given the unrest associated with the pandemic-and-depression and the surprising world-wide participation in the racism protests, 2022 is just "right out". No one is going to spend massively to go to Mars with all this going on. Not even if the mission is $50B instead of "old space's" $500B. But much of this could be over in about 5 years. Climate change will return, however, as the polar icecaps destabilize further, and sharp sea level rise begins.

I think the climate emergency will be the main pre-empting factor that might stop us going to Mars that soon. We have screwed around too long, and are past the tipping point for some degree of icecap destabilization. Both Greenland and West Antarctica are beginning to destabilize seriously. Assuming 50% loss of each, that's about 6-7 meters of sea level rise, on a time scale of 2-4 decades. Yes, I said meters! NOT inches. It'll cost many, many $B's to cope with that. (BTW, East Antarctica's potential rise is 60 meters, all by itself; but fortunately, its stability is not yet in question.)

But being optimistic about this, a mission might be put together for late 2026. If Spacex succeeds with its Starship/Superheavy as an LEO freighter, that is the likely dual-role vehicle we are looking at for the mission. If not, we are screwed, because SLS/Orion/Gateway isn't even good for the moon, much less Mars. There is nothing else. Going from a concept idea (like mine or anybody else's) to proved hardware just takes longer than 5 years. Sorry, but it's true.

Musk will more likely get international and private funding, than US government funding, to go. If he ends up actually ready to go at all. The jury is still out on his vehicle, which has yet to fly.

I don't think NASA will get its act back together anytime this century. Too swamped in pork-barrel political boondoggles to succeed at anything really big. And going to Mars is BIG.

As for international participants, it's a long way between able to fling a satellite into orbit, and being able to fling men into orbit as anything but a suicide mission. The Russians with their giant R-7 rocket (now mistakenly referred to as "Soyuz", when that is properly the name of the capsule it flings) were the exception to the rule back in the late 1950's. I think your field narrows to us, the Russians, the Chinese, and (stretching it a bit) Europe, when you consider who has the technology mature enough to fly men into space.

That's not to say the others shouldn't participate, because they should. Although I have grave doubts about allowing North Korea and Iran.

As for the US leading, I doubt that the US government as embodied by NASA can lead this thing. A US corporate entity like Spacex might. There are a few, like Blue Origin and Virgin, with Spacex in the best position at present (already just barely flying men into orbit).

GW

Last edited by GW Johnson (2020-06-14 10:17:36)

tahanson43206 · 2020-06-15 13:50:41

For GW Johnson re #90

Thank you for your detailed and considered evaluation of the potential for a multi-nation expedition to Mars in 2026.

Your prediction the human race won't be ready for an expedition including astronauts until 2026 seems reasonable. However, it seems likely (to me at least) that results of NASA and China's venture this year should encourage (a few?) other nations to see if they can do an unmanned flight.

I'm logged in on this occasion to follow up on your post #51 in this topic:

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.

I've been trying to puzzle out how to work with the figures SpaceNut came up with for a Mars launcher, which he published in the Balloon Rocket topic.

Before going to that, I'd like to report that the Mars Reconnaissance Orbiter is reported to be in a 112 minute orbit, which I compute as (about) 3,400 meters per second, which coincides with the (reported) target velocity for Surveyor, way back when.

The MRO orbit is reported to be 250 km to 340 km. The difference in velocity between your computation for 140 km of altitude and the 250-340 orbit of the MRO seems likely to be accounted for by the difference in altitude of (about) 100 km.

Your estimate for a minimal orbit for a vehicle launching from Mars was 3898 meters per second.

I'm working my way to the question ...

SpaceNut came up with numbers (from documents) which gave:

Isp = 360 seconds (using LOX/CH4 at a mixture ratio of 3.45)
Vehicle mass: 160617 kg
Dry mass: 94242 kg
Fuel/Oxidizer 66388 kg

The engine system for the vehicle SpaceNut quoted needs to produce enough thrust to put the vehicle into orbit at a velocity of 3898 m/s (or close to that)

The little calculator program I have asks for:

Mass of vehicle: Check
Mass of fuel: Check
Exhaust speed: ?
Burn rate: ?

An online tool provided by guantumg.net/rocketeq.html offers this equation:
dv = 3.724 * 360 * Math.log(160617 / 94242) and that yields a result of 714

The 3.724 is the gravitational field of Mars, computed as 38% of that of Earth ... your value was slightly different but close

In any case, I am not seeing the velocity change needed for orbit around Mars.

Can you spot the missing elements here?

Edit#1: for the Balloon Rocket topic, I'm trying to determine how much pressure (and for how long) has to be applied to the gases (oxygen and methane) to deliver the performance that SpaceNut's quoted rocket appears to achieve.

Edit#2: You have already studied that idea and confirmed that it cannot be made to work. I understand I am beating a dead horse.

(th)

Last edited by tahanson43206 (2020-06-15 14:02:02)

GW Johnson · 2020-06-15 16:40:37

Tahanson43206:

Your dV equation should read as gc Isp LN (Mign/Mbo), where the gc Isp product is an approximation for the effective exhaust velocity Vex ~= gc Isp. You don't use Mars gee, you use Earth gee (9.80667 m/s^2). It is nothing but a consistent-units conversion constant. This comes from a time long before SI, when not only the US was using lbm and lbf, but the pioneering Germans at Peenemunde were using kgm and kgf (something still seen today on torque wrenches, which read in lbf-ft and in kgf-meters. You NEVER see N-m!!).

The units of Isp are not, and have never been, consistent. For Isp to be measured in seconds, you divide thrust in kgf by kgm/sec flow rate as if a kgf and a kgm are the same (they are not). If you divide N of thrust by kgm/sec flow rate, you must also divide that result by earth gc = 9.80667 to get the right number for Isp in seconds.

In US units, Isp is lbf of thrust divided by lbm/sec flowrate, as if a lbm and a lbf are the same, which they are not. You get the right number, unlike metric, without also dividing by 32.174 ft/sec^2, because of the definition of what a lbm is. It's NOT a consistent unit at all. N and kgm are. If we are using lbf, the consistent mass unit is a slug. For lbf thrust / slug/sec flowrate, you also have to divide by 32.174 ft/sec^2 to get seconds for Isp. THAT is the analog to N thrust and kg/sec flowrate in consistent SI units.

What this points to is the traditional, but inconsistent, definition of Isp. It's not really measured in seconds, it's actually measured in lbf-sec/lbm = kgf-sec/kgm. You can thank Von Braun and his associates and mentors for that. But if you improperly divide-out lbf with lbm or kgf with kgm, you get seconds. Most folks today no longer even know what a kgf really is. You can thank SI training for that.

Gas balloon rocket:

I don't think the gaseous balloon rocket idea will work for a chemical rocket engine, because you cannot practically get into a chamber pressure range that will result in a decent frontal thrust density. But I think it would work fine for any of the electric propulsion schemes (I think I already said that in one of those other replies). Their chamber pressures are pretty much zero. Which very neatly explains the extreme low frontal thrust density that seems so inherent in the electric schemes, restricting them to in-space use only. And the low thrust/weight.

Published Isp Data:

Most of the Isp data you see quoted in the various handbooks and other sources for comparing propellants are computed for two expansion conditions. One is perfect expansion to standard sea level atmospheric pressure as the ambient backpressure, and from exactly 1000 psia chamber pressure = 68.05 standard atmospheres = 6894.7 KPa = 6.8947 MPa = 68.947 bars. That results in a rather modest nozzle area expansion ratio, which they never put in the tables. The nozzle kinetic energy efficiency is always 100% in such tables, when any practical nozzle is usually quite near 98.3%.

The vacuum data are done to a fixed expansion ratio, usually now 100, and a definite chamber pressure, often only 100 psia in the older references. That would be 6.805 std atmospheres = 689.47 KPa = 0.68947 MPa = 6.8947 bars. The ambient backpressure is zero. The same 100% nozzle kinetic energy efficiency is used. You have to look close at the fine print to see what that expansion ratio and chamber pressure were, for the data in the table.

What I like to do is use the chamber c* data, if it is reported, curve fit it versus pressure using the c* = K Pc^m model between the two pressures, and just go do the real rocket nozzle ballistics using that, at the chamber pressures I prefer. Nozzle kinetic energy efficiency is for all practical purposes merely a streamline divergence factor off of axial. It applies to the mdot Vex term, but NOT to the (Pe - Pamb)*Aex term. C* does not depend upon nozzle expansion, only on chamber conditions. It is sensitive to mixture ratio, which sometimes throws me a monkey wrench trying to generate c* = K Pc^m.

The real kicker here is the "engine cycle", which is really exactly how the turbopumps are driven. Many cycles tap off chamber hot gas to drive the pumps, then dump it overboard. That lost gas plus what goes through the nozzle is the total propellant flow drawn from the tanks. To get a realistic Isp for that cycle, you must divide your nozzle thrust by that total flow rate drawn from storage. If you don't do that, you have a garbage-in/garbage-out problem with the rocket equation. In some of the old cycles, that dumped flow can be 5-10% of the total. Supposedly it is pretty near 0% for the Raptor. It is not 0% for the Merlin, but I don't have an accurate figure.

GW

Last edited by GW Johnson (2020-06-15 17:28:30)

tahanson43206 · 2020-06-15 17:54:46

For GW Johnson re #92 ... this is just a quick note of thanks for correcting the value of G ... and for the historical considerations that led to its use.

I'll study the rest of your interesting looking post in the next day or so, but the detail about allocating a portion of fuel to the compressor function is particularly intriguing.

Edit#1 ... returning after a break

The ** real ** positive outcome that I am seeing (after rereading Post#92) is that you seem to be giving some level of endorsement to the proposition of storing gas for use by an ion engine in gaseous form. Without going to the historical (and current) examples to be certain, i am fairly sure that is how gases shipped with NASA designed ion engines are supplied with throw mass.

The following is offered with apology for not being able (at this point) to figure out the pressure at the feed pumps myself.

For GW Johnson ....

This set of calculations is intended to explore conditions that apply to a launch from the surface of Mars.

The point of departure is a single data point .... An article about NASA's study of a personnel return vehicle that would launch from the surface of Mars reported that the burn would take 7 minutes.

No other details about the vehicle were provided, except for this:

The expedition ship was assumed to be in a highly elliptical orbit, which would dip toward the planet periodically.

The mission of the crew return vehicle was designed so the crew would be able to meet the expedition vessel after a flight of 43 hours.

It was (apparently) thought that the crew could tolerate confinement in space suit-like garments for 43 hours.

The purpose of ** this ** post is to try to identify all the conditions that would have been taken into account by the NASA mission planners who came up with the 7 minute burn time.

I'm planning to add input from a document found and reported by SpaceNut.

In that document, SpaceNut found specifications for amounts of fuel (CH4) and oxygen, and dry weight of a vehicle.

SpaceNut did not identify the source, but I'm going with the assumption the source was authentic and creditable.

Quoting from SpaceNut's post:

LOX/CH4 propellant at mixture ration = 3.45 delivering Isp = 360 seconds

Mars Ascent to 100 km x 250 km: Propellant: 66,388 kg
Fuel Oxidizer
14,849 kg 51,229 kg

Ascent Mass 94,242 dry mass plus fuel

This would put the vehicle dry mass at 27,854 kg (structure, cargo and passengers)

The tools to be used for this post include Physics 101 (this program has been described in detail elsewhere in the forum)

Activity #1: Find Acceleration:

Given V1 = 0
Given V2 = 3898 m/s (orbital velocity from GW Johnson estimate for 140 km altitude)
Given time = 7*60 seconds (420 seconds)
Acceleration computed is: 9.281 m/s^2

The acceleration shown would have been achieved despite the pull of Mars' gravity on the vehicle.

Edit after re-reading the post ... the acceleration shown is well outside the range that humans would be able to tolerate.
The fact that I am using data from two different vehicle designs may account for the discrepancy.

Per Google: 3.711 m/s^2

General Thrust Equation - NASAwww.grc.nasa.gov › WWW › rocket › thrsteq

In the web page shown above, m(dot) is given as = mass flow rate = mass/time

I'm assuming that in the case of the rocket SpaceNut found, if the burn time is 7 minutes, then the mass flow rate would be;

66,388 kg / 420 seconds, or 158.0667 kg per second

That's as far as I take the analysis at this point.

For you (if willing) .... is there enough information there to know other characteristics of the launch?

The rocket will have a thrust that produces the required acceleration.

The mass against which the thrust is operating will be decreasing at the rate of 158~ kg per second

The rocket engine will be designed (presumably) to operate as efficiently as possible, over the range of use (on Mars to space)

The pressure with which the fuel and oxidizer are supplied by mechanical devices must be greater than the pressure generated by combustion in the engine.

From this it would be known what pressure would be required in a storage container to deliver the fuel / oxidizer at the same flow rate, if the fuel and oxidizer were gaseous instead of liquid.

It seems reasonable to suppose that the pressure at the inlet to the combustion will decrease over time, if the thrust produced by the engine to achieve the needed acceleration is reduced as the mass to be accelerated decreases.

Whatever that pressure is I understand it is substantially greater than any fabric storage container could withstand. The purpose of this inquiry is to (hopefully) put the balloon launch from Mars topic to a well deserved rest.

If I could tempt you with another inquiry ... would you be willing to post a mission plan, based upon your earlier design for an expedition with multiple lander flights, that would show all the factors that would go into planning the flight?

If you ** are ** willing to take something like that on, I'd like to suggest a new topic dedicated to it.

it is possible you have already done that in the exrocketman blog.

(th)

Last edited by tahanson43206 (2020-06-16 12:47:36)

GW Johnson · 2020-06-17 12:14:31

OK, here are the "exrocketman" articles that relate to Mars landers other than a Spacex Starship or Red Dragon, and not including Mars mission design studies going back to about 2010. By date and title; on the site, click on the year, then on the month, then on the title, using the navigation tool on the left.

8-23-18 Back-of-the-Envelope Rocket Propulsion Analysis
8-6-18 Exploring Mars Lander Configurations
8-31-13 Reusable Chemical Mars Landing Boats Are Feasible
3-18-13 Low-Density Non-Ablative Ceramic Heat Shields
1-21-13 BOE Entry Analysis of Apollo Returning From the Moon
1-21-13 BOE Entry Model User’s Guide
12-31-12 Mars Landing Options
9-3-12 Using the Chemical Mars Lander Design at Mercury
8-28-12 Manned Chemical Lander Revisit
8-19-12 Ballute Drag Data
8-19-12 Blunt Capsule Drag Data
8-12-12 Chemical Mars Lander Designs “Rough-Out”
8-12-12 Direct-Entry Addition to Mars Entry Sensitivity Study
8-10-12 Big Mars Lander Entry Sensitivity Study
8-5-12 Ballistic Entry from Low Mars Orbit
7-25-12 Rough Correlation of Entry Ballistic Coefficient vs. Size for “Typical” Mars Landers
7-14-12 “Back of the Envelope” Entry Model
6-30-12 Atmosphere Models for Earth, Mars, and Titan (this is excerpts from the Justus & Braun EDL book)
6-24-12 Mars Atmosphere Model (Glenn RC) (superseded by the Justus & Braun book, useful only at very low altitudes)
6-3-12 Deceleration by Drag Devices (and more) on Mars

What these articles show is the evolution of both my lander concepts and my design analysis methods. There is a similar long history list for my Mars mission articles as well, starting with the one I presented at the 2011 Mars Society convention. There is also yet another long history of my reverse-engineering of Spacex rockets and capsules, including what has become Starship/Superheavy.

The landers are pretty much intended to travel unrefueled from low Mars orbit to the surface, and back to low Mars orbit. They get refueled on-orbit from supplies sent from Earth. But, if you can do that, you can also fly from the surface to low Mars orbit, and back to the surface, unrefueled. The refueling would be done on the surface from supplies manufactured on Mars.

If you can do the one, you can do the other. I showed it possible to do with hydrazine/NTO storables. A Mars surface-based design fueled with LOX-LCH4 would have even better performance potential.

All of these take the form of a Gemini capsule scaled way up: a cone of modest tumble-home angle with a central core that sticks out a ways. Landing gear are at the periphery of the heat shield. The engines fire through holes in the heat shield, and are housed inside otherwise-sealed compartments, which stops the throughflow through the holes in the heat shield.

Ballistic coefficient is 300+ kg/sq.m, which precludes chutes. You fire up the engines as you decelerate through Mach 3, and make a simple retropropulsive landing. Those retro plumes do reduce capsule drag coefficient slightly. That was cited in my Hoerner drag bible from NASA-funded wind tunnel tests in 1961. NASA has forgotten-to-death all the innovative work like that which it chose not to use back then. Most current NASA vehicle designers were born after that time.

And all the newer guys hated Hoerner, because he predated computers completely, and used wind tunnel data, not software predictions, for his drag recommendations in his book. Myself, I think his self-published book is a real jewel. Hoerner was an immigrant to the US after WW2. He was the aerodynamicist for the Messerschmidt 109 fighter plane project in Germany in the 1930's.

GW

Last edited by GW Johnson (2020-06-17 12:30:22)

tahanson43206 · 2020-06-17 15:16:48

For GW Johnson re #94

Thank you for the time you invested in preparing the summary in Post 94.

Beyond that though, is the awesome investment in the cited articles.

Thank you for the reminder of Dr. Hoerner, and the Messerschmidt 109.

https://airandspace.si.edu/collection-o … 9600327000

I took a quick look at the history recorded in the site above.

For Dr. Hoerner, I found this:

Did you mean: hoerner aerospace research
Search Results
Web results
Sighard F. Hoerner - Wikipedia
en.wikipedia.org › wiki › Sighard_F._Hoerner
Dr. Sighard F. Hoerner was an important figure in the aerodynamics field and is known worldwide for his two compendiums of aerodynamic knowledge, Fluid-Dynamic Drag and Fluid-Dynamic Lift. He is also notable for his design work on the pioneering STOL aircraft, the ... Dr. Hoerner studied mechanical engineering at the Technical University of ...
Career · Contributions

Hopefully I got the right gent, although the text in the Wikipedia article matches up pretty well with your notes.

The article describes his difficulty finding a publisher. Ultimately his persistence led to publication and even US Navy funding!

In your post, I would like to highlight this:

Ballistic coefficient is 300+ kg/sq.m, which precludes chutes.

It seems to me that planning to land (safely and reliably) without chutes or other paraphernalia is the key to winning a favorable review by NASA, for their stated goal of placing a forty metric ton payload on Mars.

A detailed plan for an expedition to Mars, designed from the outset (as you have indicated you would prefer to do) for safety of the crew and their safe return, might be of interest to a modern publisher in these times of ambitious statements by Elon Musk, and forward looking statements by India and China to name just two active players in this field.

For SpaceNut .... please think about how (or if) you would be interested in encouraging development based upon the foundation GW Johnson provides.

There is an opportunity to attract younger members of the active aerospace community, with the caveat that the undertaking needs to appear safe for them to risk their reputations.

I have so far been unsuccessful in persuading anyone to take the chance.

SearchTerm:ExRocketmanSummary

Pasting from summary in Post #94 above:

8-23-18 Back-of-the-Envelope Rocket Propulsion Analysis

8-6-18 Exploring Mars Lander Configurations

8-31-13 Reusable Chemical Mars Landing Boats Are Feasible

3-18-13 Low-Density Non-Ablative Ceramic Heat Shields

1-21-13 BOE Entry Analysis of Apollo Returning From the Moon

1-21-13 BOE Entry Model User’s Guide

12-31-12 Mars Landing Options

9-3-12 Using the Chemical Mars Lander Design at Mercury

8-28-12 Manned Chemical Lander Revisit

8-19-12 Ballute Drag Data

8-19-12 Blunt Capsule Drag Data

8-12-12 Chemical Mars Lander Designs “Rough-Out”

8-12-12 Direct-Entry Addition to Mars Entry Sensitivity Study

8-10-12 Big Mars Lander Entry Sensitivity Study

8-5-12 Ballistic Entry from Low Mars Orbit

7-25-12 Rough Correlation of Entry Ballistic Coefficient vs. Size for “Typical” Mars Landers

7-14-12 “Back of the Envelope” Entry Model

6-30-12 Atmosphere Models for Earth, Mars, and Titan (this is excerpts from the Justus & Braun EDL book)

6-24-12 Mars Atmosphere Model (Glenn RC) (superseded by the Justus & Braun book, useful only at very low altitudes)

6-3-12 Deceleration by Drag Devices (and more) on Mars >> Includes Soviet tank landing system

Image from June 2020 update:

(th)

Last edited by tahanson43206 (2020-06-18 06:39:47)

GW Johnson · 2020-06-17 15:53:47

I have had Hoerner's two books in my library for many, many years now, and I have used them many times. Along with many, many others.

GW

tahanson43206 · 2020-07-08 06:18:06

This is a follow up to Post #37 .... I am still concerned about the forces at work as a vehicle approaches Phobos ...

The most efficient docking at Phobos would appear to involve navigating so that Phobos is receding while Mars itself is advancing on the vehicle at the peak of the Hohmann ellipse. As I tried to interpret your chart (as shown in Post #35) it ** appears ** that the docking vehicle would need to accelerate 1.88 km/s to match orbit with Phobos while Phobos is itself receding from the vehicle at 2.138 km/s.
I have another concern, and (hopefully) it will prove of interest.
If the vehicle planning to dock at Phobos arrives at the optimum time, when Phobos is receding at its fastest rate with respect to Mars, then the vehicle will become part of the mass of Phobos. However, the vehicle itself will (presumably) still retain the momentum which (had it not been disturbed) would have carried it past Mars. Since Phobos is on a circular orbit, it is constantly accelerating toward Mars.
I'm assuming Coriolis effect would be present, although I don't have a sense of how strong it would be. Is there any risk of the landing vehicle tipping over, if it is long and thin, like one of Mr. Musk's Starships?
The gravity of Phobos itself is given as 0.0057 m/s² (by Wikipedia).
Thanks again for your patience!

It follows from the discussion above that Phobos will be responsible for accelerating the spacecraft by 2.138 km/s. That acceleration will occur as the moon swings around Mars and carries the spacecraft with it. It would seem clear by inspection, that if the spacecraft is not anchored securely it will fly off Phobos at the rate of 2.138 km/s less the gravity of Phobos, which will impose a minuscule acceleration to the spacecraft.

A reasonable guess at a suitable "anchoring" method is to place the spacecraft on the surface of Phobos at a point which will be directly "under" the spacecraft as the moon rotates in its travel around Mars. This guess leads to the prediction that the optimum landing location would be at a point on the surface which will rotate under the spacecraft while the acceleration due to momentum of the moon occurs.

SearchTerm:DockingPhobos

From another post above we have:

Google gives 3.711 m/s^2 as the acceleration of gravity at the surface of Mars.
https://www.universetoday.com/14859/gravity-on-mars/
The surface gravity of Mars can therefore be expressed mathematically as: 0.107/0.532², from which we get the value of 0.376. Based on the Earth's own surface gravity, this works out to an acceleration of 3.711 meters per second squared.Dec 16, 2016

Planned sketch and problem statement here >>> http://newmars.com/forums/viewtopic.php … 84#p169784

Edit#1: It should be possible for a person with the necessary knowledge, experience and computing hardware and software to predict with great accuracy exactly how much "g force/acceleration" due to the momentum of Phobos will be imparted to the spacecraft as it "rests" on the surface.

The amount of acceleration will increase as the moon curves around Mars, and then decrease to the minuscule gravity of Phobos.

The time over which this acceleration will occur would be no greater than half the period of the orbit of Phobos, which is given by Google (wikipedia) as 7 hours and 39 minutes.

Edit#2: Since Stickney Crater always points toward Mars, it seems reasonable (to me at this point) to suggest navigating the spacecraft to the interior of Stickney crater while Phobos is receding most rapidly from the spacecraft in its orbit around Mars. As Phobos continues its orbit, it will move toward the spacecraft at a low rate of acceleration until the spacecraft is firmly resting upon the surface of the interior of the crater. Acceleration on the spacecraft will continue until all the momentum needed has been imparted to the spacecraft. At that point, the spacecraft will be at rest with respect to Phobos.

(th)

Last edited by tahanson43206 (2020-07-08 10:51:52)

tahanson43206 · 2020-07-08 10:51:14

Image to go with #97

Removed version with landscape orientation.

For SpaceNut ... is there a way to change the orientation of this image?
Answer: it turns out that www.imgur.com provides a "rotate and select" option!

OKwmSUf.jpg?1

(th)

Last edited by tahanson43206 (2020-07-08 11:01:10)

GW Johnson · 2020-07-09 12:08:27

Tahanson43206:

I don't understand the concerns over Coriolis forces (should be vanishingly small), or where the velocity numbers you quoted are coming from. I do understand and share your concern about something tall and slender toppling over on rough ground after landing on Phobos.

My old (and probably slightly-wrong) data on Phobos say its surface gravity (highly variable) averages 1.67 cm/s^2 = 0.0167 m/s^2, and its mean escape velocity is just under 19 m/s. The semi-major axis (average radius) of its orbit is 9408 km (center-to-center). I presume it is tidally locked to Mars, so that it rotates in one Mars day (24 hr 37m).

Wikipedia matches my old dynamics textbooks defining the (apparent) Coriolis acceleration as 2 x spin rate x velocity, which in SI units would be spin rate in radians/s, and velocity in m/s, so that acceleration comes out in m/s^2. The real formulas deal in vector cross products, which have the values I show only when the velocity vector and spin rate vector are perpendicular. For angles between them not 90 degrees, you multiply the terms I show by the sine of the angle between them, which maxes-out at sin 90 = 1.000.

The spin rate would be Phobos's rotation rate = 1 full turn in 24 hr 37 m, or 2 pi radians in 88,620 s = 7.09E-5 rad/s. The velocity would be AT MOST Phobos's orbital speed of 673.9 m/s, typically near 10 m/s across the surface during landing or takeoff when close to the surface, and zero while sitting on the surface.

The corresponding Coriolis acceleration would be 0.048 m/s^2, and this would be identically zero if speed were zero with respect to Phobos. For a more typical 10 m/s during landing, the Coriolis acceleration would be 7E-4 m/s^2, which is quite small, even compared to the very low average surface gravity acceleration of Phobos (4.2%).

I think I had a 5% of local gravity acceleration figured as the surface slope limit for the 2018 and 2019 versions of Starship. I rather doubt Spacex has yet looked at sloped ground in the landing leg designs for Starship, any more than they have looked at surface bearing pressures on soft sand-like regolith surfaces. Which they so very obviously have not.

Topple-over is a severe risk, even at zero velocity (for zero Coriolis acceleration), just because the surface is quite rough and irregular! Fix that, and I think any Coriolis effects will take care of themselves.

Mars orbits the sun in a posigrade direction at a speed that is somewhere between its apohelion and perihelion values of 21.964 and 26.490 km/s, respectively, measured with respect to (wrt) the sun.

Phobos orbits Mars in a posigrade direction at an average center-to-center distance of 9408 km, and an average speed value of 0.6739 km/s, measured wrt Mars. What that means is the velocity of Phobos wrt sun is the velocity of Mars wrt sun plus the velocity of Phobos wrt Mars, just standard vector mechanics. Where these are parallel vectors, you may add magnitudes, elsewhere you must do real vector addition.

For Mars at its perihelion, the velocity of Phobos wrt sun is 26.49 km/s + 0.6739 km/s = 27.164 km/s in the posigrade direction, when Phobos is exactly on the opposite side of Mars from the sun. And, when Phobos is exactly lined up on the same side of Mars as the sun, its velocity wrt sun is 26.49 km/s - 0.6739 km/s = 25.816 km/s posigrade, for Mars at its perihelion.

The numbers are smaller when Mars is at its apohelion, because Mars's speed about the sun is significantly slower there. With Phobos on the backside, its velocity wrt sun is 21.964 + 0.6739 = 22.638 km/s posigrade. With Phobos on the solar side, its velocity is 21.964 - 0.6739 = 21.290 km/s posigrade.

Now ONLY for a Hohmann min-energy transfer ellipse, at its apohelion, the trajectory path is posigrade and tangent to Mars's orbit, so that trajectory and Mars velocity vectors are parallel, and their magnitudes add arithmetically. Faster trajectories have crossing points no longer parallel to Mars's orbit, for which true vector addition must be used.

This Hohmann apohelion point has maximum orbit length and minimum velocity when both planets are at their apohelion distances, and maximum velocity at minimum orbit length when both planets are at their perihelion distances. Thus the max transfer apohelion velocity is 24.9 km/s, which should be associated with the larger Mars perihelion velocity, and the minimum is 20.1 km/s, which should be associated with the smaller Mars apohelion velocity. These are wrt sun.

Phobos's surface escape velocity is a bit variable, but 19 m/s = 0.019 km/s would be typical.

So, for a spaceship trying to reach Phobos from a Hohmann transfer ellipse, we can estimate the relative velocity wrt Phobos (far from Phobos) as ship speed wrt sun - Phobos speed wrt sun = ship speed wrt Phobos. A negative sign on that result indicates retrograde direction of the relative velocity vector, which simply reflects that faster Phobos (and Mars) is running over the slower ship, from behind.

Using the min and max values given above thus bounds the problem: data for all other orbital positions will lie in between these values, no matter what. If you can handle the upper bound, you can handle anything.

Higher-speed Mars perihelion values, far from Phobos: ship velocity wrt Phobos = 24.9 km/s - 25.18 km/s = -0.28 km/s for Phobos on the sunward side, and 24.9 - 27.164 = -2.26 for Phobos on the backside.

Lower-speed Mars apohelion values, far from Phobos: ship velocity wrt Phobos = 20.1 km/s - 21.290 km/s = -1.19 km/s for Phobos on the sunward side, and 20.1 - 22.638 = -2.54 km/s for Phobos on the backside.

You clearly get the lower velocities wrt Phobos (far from Phobos) when you approach it on the sunward side. We shall consider only that case from this point. That puts the ship velocity wrt Phobos far from Phobos within these limits:
-0.28 to -1.19 km/s.

Those can be considered kinetic energies far from Phobos, to which the escape speed kinetic energy adds, to produce the kinetic energy (and thus the speed) very close to Phobos: Vfar^2 + Vesc^2 = Vnear^2. Those near-Phobos numbers are very little different, because Phobos's escape velocity is so small. Vnear wrt Phobos is 0.281 to 1.19 km/s. That range is the kinematic delta-vee requirement to make a direct landing on Phobos.

Phobos has no atmosphere, so there can be no drag loss. It has very little gravity (0.0017 gee), so there can be an almost trivial gravity loss (factor = 1.000085). Thus the mass ratio-effective delta-vee requirement range is essentially that same range 0.281 to 1.190 km/s.

Yes, going to Phobos has a lower orbit-to-orbit transport ship delta-vee requirement than going to low Mars orbit. But as I have determined elsewhere, the cost of going to Mars's surface from Phobos is substantially higher than from low Mars orbit (in point of fact very near Mars escape).

GW

Last edited by GW Johnson (2020-07-09 12:11:07)

SpaceNut · 2020-07-09 17:07:10

Coriolis Effect Formula The following formula can be used to determine the force acting on an object in motion due to the Coriolis effect.

F = 2 * m * v * w * sin (a) F is the force (newtons)

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

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

But I think this is the force you might be thinking of
https://en.wikipedia.org/wiki/Centripetal_force

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#76 2020-06-07 18:13:25

Re: Solving Mars mission docking with Phobos

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Re: Solving Mars mission docking with Phobos

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Re: Solving Mars mission docking with Phobos

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Re: Solving Mars mission docking with Phobos

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Re: Solving Mars mission docking with Phobos

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Re: Solving Mars mission docking with Phobos

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Re: Solving Mars mission docking with Phobos

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Re: Solving Mars mission docking with Phobos

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Re: Solving Mars mission docking with Phobos

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Re: Solving Mars mission docking with Phobos

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Re: Solving Mars mission docking with Phobos

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Re: Solving Mars mission docking with Phobos

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Re: Solving Mars mission docking with Phobos

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Re: Solving Mars mission docking with Phobos

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Re: Solving Mars mission docking with Phobos

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Re: Solving Mars mission docking with Phobos

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Re: Solving Mars mission docking with Phobos

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Re: Solving Mars mission docking with Phobos

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Re: Solving Mars mission docking with Phobos

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Re: Solving Mars mission docking with Phobos

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Re: Solving Mars mission docking with Phobos

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Re: Solving Mars mission docking with Phobos

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Re: Solving Mars mission docking with Phobos

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Re: Solving Mars mission docking with Phobos

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