You are not logged in.
For GW Johnson re #199
Thank you for continuing to assist with this stretching-the-envelope-to-the-breaking-point topic!
For SpaceNut re #200
Thank you for continuing to think about the alternatives that might be available for this venture ...
I'm picking up on a couple of points you've included in #200
First, regarding precision ... we can most definitely deal with that if we can get the strafing run concept to work. The vehicle able to deliver the payload onto a dime already exists on Earth. I'm going to make the assumption the same capabilities can be extended to the (of course) more challenging environment of Mars.
Second, regarding retro propulsion ... I find myself forced to stretch Void's original vision a bit as we go along, and observations from forum members come in.
Let us suppose (for this post anyway) that we can find a way to operate a Phobos-Phobos drop-off flight system.
There is a trade off between the velocity of the vehicle (needed to return to Phobos with available propellant) and the velocity of the payload, which needs to be as low as possible and at ** least ** under 500 mph to insure the safe arrival of non-perishable payloads.
Is there a sweet spot (I'm asking GW Johnson primarily, but all are welcome to contribute) in which a solid propellant decelerator might be able to slow a one ton payload to (for the sake of an ideal) zero velocity in all dimensions with respect to Mars?
In post 199, GW Johnson has opened the door for consideration of non-chemical propulsion systems.
Staying with all-chemical propulsion for a moment ...
If a payload is coming from Earth, then it could be packed for shipment with a solid rocket motor designed for the special circumstances of a Ballistic Delivery scenario. In practice, it appears (to me at least) that ** most ** if not all Mars landers have included solid rocket motors in the landing procedure.
Those have been designed to perform landing deceleration after parachute release, and the payloads in ** those ** cases must be delivered to Mars at zero relative velocity in all dimensions. However, the Ballistic Delivery payloads do not need to receive such gentle handling.
A shipment of Quaoar's Olive Oil might need a ** bit ** of TLC, but I doubt it needs "zero" relative velocity.
Let us suppose there is a sweet spot at which a chemical powered rocket plane (X37b equivalent) can deliver a retrorocket equipped payload package to a precise landing trajectory so that the solid rocket motor (with a bit of smarts) can decelerate the package to whatever the survivable velocity may be for that particular shipment.
GW Johnson has shown that conditions at the 500 mph minimum velocity suggested by Calliban is ** just barely ** doable.
Suppose we up the ante a bit ...
Is there a velocity for the space plane from Phobos that would allow it to return to Phobos using 245 ISP with a bit of margin for safety, while the retro rocket ahead of the payload handles the responsibility for landing precisely on the designated landing site at the minimum velocity needed by the given payload?
In other words ... the shipper is going to need to invest a bit more in the retro rocket for Olive Oil than will the shipper of highly refined steel.
(th)
Offline
TH:
Solid propellant rocket motors have a fixed thrust-time trace at any one soak-out temperature (which affects burn rate, hotter being faster). Hot-soaked motors have a taller, shorter thrust-time trace, cold-soaked motors have a lower, longer thrust-time trace. The area under the thrust-time trace is the motor total impulse. This, too, varies with soak-out temperature, but the percentage variation is lower than that of thrust level.
There is no control or adjustment of the delivered thrust-time trace. Once lit it burns to the end, just like a dynamite stick, except far slower than dynamite. The burn rate of dynamite is measured in km/s, which is why the explosion appears instantaneous. The burn rate of solid rocket propellants is measured in mm/s or at most a few cm/s.
To zeroeth order, the total impulse of your thrust-time trace changes the momentum of your object by the same amount (ignoring mass variation effects). That total impulse is equal to the average Isp of the motor, times the mass of propellant it contains. That's a rough-and-ready first estimate of the motor size you need.
GW
Last edited by GW Johnson (2021-04-12 15:46:53)
GW Johnson
McGregor, Texas
"There is nothing as expensive as a dead crew, especially one dead from a bad management decision"
Offline
The issue for solids is you get no second chance to light them and once on they run until the fuel is gone.
GW just checking on computer need, have been there last year and would extend help if wanted....
Offline
For GW Johnson re #202
Thanks for taking a first look at the proposition at hand!
SearchTerm:RoughAndReady estimate of solid rocket size for Ballistic Delivery retropropulsion from Phobos-Phobos drop off vehicle.
For SpaceNut ... You are certainly right to point out the risk of failure to light. That risk is faced by every mission planner who uses solid rocket motors, or pyrotechnic devices for that matter, to perform critical tasks. All I can do at this point is to note the high rate of success of the devices implemented by NASA/JPL for their on-Mars landings.
Another risk of equal weight is the risk of navigation error by the on-board electronics package. That package needs to have precise location information (probably by direct injection from the spacecraft until the moment of disconnect) as well as precise figures for velocity and vector.
On the other hand, the package is NOT delivering perishable goods (in the first iteration for sure) so even if the rocket fails or navigation is imprecise, the payload can be salvaged. The purpose of the rocket in this line of thinking is to reduce velocity (a) and (b) to insure accurate placement of the payload into the customer designated delivery zone on Mars.
At this point, it remains a speculation (on my part) that there exists a solution that allows the best of both concepts .... the orbit-dropoff-orbit vehicle, and the landing package for non-perishable supplies, to provide the most cost effective (and therefore competitive) solution possible.
The objective for this round is to find a solution that puts a metric ton of mass in a designated customer site at survivable velocity, while allowing the delivery vehicle to return safely to orbit.
(th)
Offline
There has been many a rocket made using the solid fuels to boost them off the earth to giving the final push of the probe but they have not been used to do retro propulsions as they are heavy, and burn to quickly with no control of the velocity of the thrust that exits the rockets nozzle.
So far the sample return for mars is slated to make use of a solid to get the samples on the way from mars surface.
Offline
For SpaceNut re #206 ...
I'm surprised by your report of solids NOT being used to slow vehicles landing on Mars, but assume you must have read reports more carefully than I have. The details of rocket type were probably printed in the various articles about the landings, but I must have overlooked them.
The nature of the rockets actually used is interesting, and it may turn out to be worth considering whatever rocket types have actually been used, since they've clearly been successful ( a few failures here or there notwithstanding).
(th)
Offline
The retrorockets used on Mercury were solid propellant.
The ullage motors used on Saturn-1 and Saturn-5 were solid propellant.
The old Scout satellite launcher was a 4-stage solid that put hundreds of satellites up for 30 years without any failures of the launcher.
All of the variants of the Minuteman ICBM are 3-stage solids, with a hydrazine-thrusted warhead "bus" to achieve extreme precision.
Nearly every single combat missile in the world is pushed by a solid rocket. A few are turbojet. A few are ramjet (and some of them feature a solid propellant fuel supply). But all have solid rocket boosters, even if airbreathers.
Just about all of these are very old and very successful technologies that employ solid rockets. You need to allow for the major effects on thrust level, and minor effects on total impulse, of propellant soak-out temperature, or else you need to control soak-out temperature.
These things are "wooden round" simple. It works when you hit the ignition signal. Period. All you need is redundant igniters, to achieve super high reliability. No prep, no verifications, Nothing. What they are checking before they launch a missile is the guidance/control/seeker electronics, not the rocket motor.
Does that give you some idea how reliable these motors are? Why they are so popular in weapons?
The counterexample is the shuttle SRB that destroyed Shuttle Challenger. But, consider: NASA insisted on the two O-ring seal design that caused that failure, when Thiokol who built them knew better. The same NASA that had never, ever built a solid motor themselves. They still haven't.
So, what REALLY caused that failure? Perhaps ignorant, arrogant mismanagement by the government customer? And maybe we've seen that behavior before? In many government agencies?
GW
Last edited by GW Johnson (2021-04-15 08:23:04)
GW Johnson
McGregor, Texas
"There is nothing as expensive as a dead crew, especially one dead from a bad management decision"
Offline
For GW Johnson re #207
SeachTerm:Solid rockets - review of history with assessment of reliability
SearchTerm:Rockets solid GW Johnson Post #207
Would you be willing to investigate the possibility there is a sweet spot for the Phobos-Phobos Ballistic Delivery scenario, using a carefully designed (of course) solid rocket motor with suitable guidance electronics to place a 1 ton payload on a dime at the surface of Mars with velocity under 500 mph/800 kph?
I am hoping this forum can attract persons with skills, knowledge and experience comparable and complementary to yours, but at the moment, you appear to be the ONLY member able to take a look at the question.
My working hypothesis/guess/assumption is that a well designed, well executed Phobos-Phobos-Strafing-Run maneuver will prove superior to any soft landing scenario, absent fusion power, which will surely happen at some point.
There may be a practical limit to the mass that might be delivered by this method. The goal I am suggesting for the first round to evaluate feasibility is 1 metric ton of non-perishable supply delivered to Mars in recoverable condition.
However, that 40 ton delivery objective of NASA (published earlier in this forum) remains "out there" as a tantalizing possibility.
First, we (forum members) need to establish if it is feasible to employ this delivery method at all, using 245 ISP propellant and Phobos-Phobos flight plan.
(th)
Offline
TH:
Using delta-vee as a surrogate for $, the cheapest route from Earth to Mars is a simple Hohmann transfer. At its apohelion, velocity is parallel to Mars's orbit, but a lot slower than Mars orbital velocity. The planet wants to run over your craft "from behind", so to speak. Once it gets close, the 3-body effect is to speed your craft up even more.
I use an energy-based approximation to estimate this speed-up effect, without resorting to software for a 3-body solution. Velocity at low orbit or entry interface is around 5.4 km/s with respect to Mars. This can be as high as 7.5 km/s, using the much faster 2-year "abort" ellipse. Hohmann averages 8.6 months 1-way. The 2-year abort transfer averages 4.3 months 1-way.
You had to make a departure burn from Earth. Period. Could be from the surface or from orbit. Smaller delta-vee is from orbit (about 3.3 or 3.4 km/s, or about 1-2 km/s faster for faster transfers). There is no avoiding that departure requirement, there is only the choice of how you are going to do it: with your vehicle or with a throwaway departure stage. This is a long-duration burn, so nuclear thermal could be used.
On the way, there are two course corrections to make. One is near midpoint, called the midcourse correction. The other is as you approach Mars, to hit the "window" for whatever you are choosing to do there. These are pretty much required as a budget, whether you actually use all of it or not. It totals near 0.5 km/s for Mars, as best I can estimate. These are short, sharp burns requiring great precision. Nuclear thermal is probably not a viable option.
Now that you are near Mars and about to smack it at high speeds (5.4 to 7.5 km/s) that are well over Mars escape (5 km/s), you MUST do "something". You have two (and only two) practical choices at this time in history, because time-to-impact is very limited: (1) aerobrake into a direct landing, or (2) conduct a burn to enter low Mars orbit.
The direct landing will always (ALWAYS!!!) have some sort of touchdown burn requirement, on the order of 1 km/s theoretically if you come out too low for chutes, for which you ought to budget 1.5 km/s or more.
The entry into low Mars orbit varies quite a bit with Mars's distance from the sun at your arrival, but is on the order of 2 km/s. You CANNOT avoid that burn, if you intend to reach low Mars orbit!
There is a third option, but it is not very practical until we have weather satellites around Mars that can gauge actual density at entry altitudes (40-140 km). That would be (3) aerocapture by repeated passes dipping into the Martian atmosphere, which is RobertDyck's favorite.
There is NO AVOIDING the requirement that the first drag pass reduce your velocity from beyond escape to something between escape and circular orbit speed. Fail, and you are lost in space, forever! There is simply no avoiding that requirement!
Bear in mind that entry altitude densities at Mars vary by an erratic factor of 2 from nominal. You have to know exactly what it really is upon your arrival, so that you know how deep to plunge on that first pass. That determines the arrival course correction you MUST make. There is no avoiding that issue!
Until we can have that knowledge of actual entry-altitude densities, this repeat-pass aerobrake option cannot be practical at Mars. Factor-2 variation in your pathwise drag integral on the first pass leads either to burnup/crash, or to bouncing off into deep space, at its extremes. The faster your speed at arrival, the worse this is.
From low Mars orbit, the deorbit burn is about 50 m/s = 0.05 km/s. You aerobrake, then depending upon how low you are, coming out of hypersonics, you may or may not be able to use a chute. If not, there the is very same touchdown burn requirement as for the direct landing: 1 km/s theoretical, 1.5+ km/s that should be budgeted.
So, for min-energy Hohmann departing from Earth orbit to a direct landing you have:
depart..........................3.4 km/s
mid/endcourse.............0.5 km/s
aerobrake.....................0
touchdown...................1.5 km/s
total..............................5.4 km/s (also a grand total)
And for min-energy Hohmann departing from Earth orbit to low Mars orbit you have:
depart...........................3.4 km/s
mid/endcourse..............0.5 km/s
orbit entry.....................2.0 km/s +/-
total..............................5.9 km/s +/-
plus what is required in order to land:
deorbit..........................0.05 km/s
aerobrake......................0
touchdown....................1.5 km/s
total...............................1.55 km/s
grand total to land via low Mars orbit:
grand total......................7.45 km/s +/-
If (and only if) the repeat-pass aerobraking into Mars orbit becomes feasible, then that looks about like the burn-into--orbit without the orbit entry burn, done as Hohmann transfer from Earth orbit:
depart............................3.4 km/s
mid/endcourse...............0.5 km/s
repeated aerobraking......0 (as a min)
total................................3.9 km/s (as a min)
plus what is required to land:
deorbit............................0.05 km/s
aerobrake........................0
touchdown......................1.5 km/s
total.................................1.55 km/s
grand total to land using repeat-pass aerobraking capture:
grand total........................5.45 km/s (as a min)
To do "ballistic delivery" to Mars, until and unless repeat-pass aerobraking capture becomes feasible with certainty of entry altitude densities, you have but two choices: (1) direct entry to a direct landing, or (2) entry into low Mars orbit followed by a landing from orbit.
For direct entry and landing, there is no need for a delivery spaceplane or rocket or anything, except that required to survive hypersonic entry and to touch down survivably. But, the landing location could literally be all over the planet.
If you go into low Mars orbit, you can use either the same craft, or a different one, to conduct a precision entry and landing, to a precisely-controllable location. The same would apply to repeat-pass aerobrake capture, once that becomes feasible.
There can be a lot of vehicle and propulsion choices about exactly how to go about doing either choice.
I didn't look at "ballistic capture" into a high, elongated orbit. Both the deorbit burn, and the entry interface speed, are higher (speed near Mars escape). You will expend more propellant to get a precision landing location from a high, elongated orbit like that. That consequence is unavoidable, too.
I didn't look at delivery to Phobos then delivery to Mars, either. That would have the same difficulties as the high, elongated "ballistic capture" scenario, plus some small burns to land on (and depart from) Phobos.
I think low Mars orbit makes the most sense, but you WILL pay a delta-vee price for a precise landing location. Low Mars circular orbit speed is close to 3.55 km/s. To be mass-ratio effective, you factor that up by about 2%. If you aerobrake, drop stores at high speed, then reaccelerate to orbit, your theoretical delta-vee back to orbit gets reduced by your speed at drop. This might help with descent of the cargo, but you incur a need to refuel the drop vehicle on-orbit about Mars ( which is NOT where the propellant is manufactured!!!), unless you intend to throw it away (if you do, it might as well be the lander).
GW
Last edited by GW Johnson (2021-04-15 20:43:48)
GW Johnson
McGregor, Texas
"There is nothing as expensive as a dead crew, especially one dead from a bad management decision"
Offline
For GW Johnson re #209
First, thank you for this comprehensive summary of the entire Go to Mars scenario
SearchTerm:Overview of Mars travel from Earth and various options when there Post #209 GW Johnson
Second, thank you for the narrow focus of looking at the scenario of a space plane (X35b is an example) able to drop from Phobos, decelerate using the (thin) atmosphere for braking, release a payload package with a solid rocket motor for precise delivery, and return to Phobos.
I think low Mars orbit makes the most sense, but you WILL pay a delta-vee price for a precise landing location. Low Mars circular orbit speed is close to 3.55 km/s. To be mass-ratio effective, you factor that up by about 2%. If you aerobrake, drop stores at high speed, then reaccelerate to orbit, your theoretical delta-vee back to orbit gets reduced by your speed at drop. This might help with descent of the cargo, but you incur a need to refuel the drop vehicle on-orbit about Mars ( which is NOT where the propellant is manufactured!!!), unless you intend to throw it away (if you do, it might as well be the lander).
Thank you (in particular) for the reminder that the strafing run delivery vehicle needs to refuel somewhere, in order for the scenario to make any sense.
The flight through the atmosphere is a reasonable place to look for the ingredients for the next flight, so collecting all that ** hot ** CO2 would seem like a reasonable aspiration. Solar power at Phobos can split the molecules if the vehicle can get them there.
Per your observations about the difficulties of lateral flight through the atmosphere, I would expect the delivery zone from Phobos to be within some (relatively) narrow band near the Phobos track. The width of the track might be extended over time, as engineers gain experience with the vehicle and its characteristics.
A business built around this concept might well create a class of vehicle that dips into the atmosphere of Mars to collect gas without worrying about making a delivery. In that case, the vehicle would be collecting gas not only for itself but for a stash that would be maintained at Phobos for actual deliveries.
The atmosphere collection flights can be run non-stop, regardless of whether there are any deliveries in the pipeline.
In fact (come to think of it) the package delivery vehicle already has enough of a challenge, so I expect the fuel collection task would NOT be one of its assignments.
As the package delivery scenario is shaping up in this early stage, I'm "seeing": [Executive Summary < ? >]
1) Preparation of the payload package at Earth, including payload, wrapping, deceleration rocket, electronics and whatever I've overlooked.
2) Shipment of the package to Mars from LEO using slow-boat-to-Mars-least-cost-option
3) Arrival at Phobos - careful testing by personnel there - locking into payload delivery vehicle -
4a) Launch from Phobos (no vehicle fuel used for departure - Phobos provided impulse - exact nature to be determined)
4b) Delivery run .... Stay high and fast enough to be able to return to Phobos - drop payload with location data at optimum velocity and vector
5) Vehicle returns to Phobos with enough fuel left to dock
6) Package continues toward customer designated site using onboard navigation and external references as available
7) Depending upon customer investment, package arrives with some velocity that payload can tolerate, all the way to zero relative velocity
8) Customer retrieves delivery and separates all materials for use - nothing is wasted - nothing is lost
9) Vehicle refuels at Phobos for next delivery
Thanks again for the comprehensive overview of the entire Earth/Mars flight regime.
(th)
Offline
TH:
Seems like I put some numbers up about going to Phobos (and from there to Mars) on "exrocketman". I'd have to go look. Laptop is dead. I am trying to get the desktop files recovered; the source documents and spreadsheets are there. Until recovered, I have no access to them at all. Meanwhile I can correspond occasionally, using my wife's computer.
If I understand correctly, any sort of ship, even a 1-way throwaway vehicle, could bring the cargo to Phobos, in your plan. Once there, you load this cargo into some kind of reusable relatively-aerodynamic vehicle, and go from Phobos to Mars, and back to Phobos after dropping the cargo off. Then you refuel this thing on Phobos. Somehow.
Assuming refueling on Phobos is possible, what you seem to want to explore is how to aerobrake deep into Mars's atmosphere, drop off the cargo in flight, and then reaccelerate back out of the atmosphere, and return to Phobos. The main question here is at what flight speed to make the drop, and exactly how to get the cargo to survive the "landing".
The max speed value for the drop is about local Mach 3 (on Mars about 0.7 km/s = 700 m/s), when the aerobraking entry hypersonics end. For what amounts to high ballistic coefficient for a craft of nontrivial size, the altitude will be quite low: say, 5-10 km, and the trajectory bending steeply downward, in the vicinity of 45 degrees. You are a single handful of seconds from impacting the surface.
The max deceleration gee pulse and the max entry heating pulse will have only happened a single handful of seconds before. You just don't want to dump your cargo out, into an incandescent plasma slipstream. You have to wait for the heating pulse to end. And for the pulse of deceleration gees to end, too. They occur at slightly-different times.
The cargo item you drop will need some sort of attitude stabilization. I would suggest a small cluster of solid rocket motors on a pole sticking out the rear, oriented to decelerate the cargo as tractor retrorockets. If it needs a soft landing, use bigger motors, and very precisely control the speed and altitude at which you make the drop. If it can survive whacking the ground at significant speed, you can use smaller motors, and not care much about precision in your drop point.
As for the solid soakout temperature effect, your delivery vehicle cargo bay needs to be temperature-controlled (which implies an atmosphere). That way you know precisely what thrust level and total impulse your solid motors will have.
Supersonic store separation is an extremely serious issue, although the thinness of the Martian atmosphere helps alleviate that to one extent or another. The risk is for aerodynamic forces exceeding cargo package weight. A powerful cargo eject mechanism is required, and it cannot induce tip-off perturbations into the cargo package attitude. There is a reason (of potentially fatal consequences) that most military aircraft are restricted to max 485 knots indicated airspeed at sea level for store separation. It's a dynamic pressure limit, really.
If we have good solutions for all those pitfalls, then there is one VERY critical unaddressed pitfall: refueling the delivery vehicle on Phobos. There are many on these forums who seem convinced that Phobos contains volatiles buried within, from which propellant could be made. I STRONGLY disagree! I think it will prove dry as an old bone, nothing but a ball of sand, gravel, rocks, and boulders, with some space-weathered "crusty crust" carbonaceous crap at the surface.
Actually, a Phobos incapable of providing resources to make propellant is the very biggest, most fatal, objection to this plan! That is the issue that must be addressed FIRST, unless you plan on shipping propellants from Earth.
I can't do much more on this project until I have a new laptop, with appropriate software, and my recovered data is loaded into it. I don't have any insight yet into when that will be. -- GW
GW Johnson
McGregor, Texas
"There is nothing as expensive as a dead crew, especially one dead from a bad management decision"
Offline
GW I was offered help last year by several members when my communication were cut off for months until I got the newer computer going with its networking updated and would likewise put out that helping hand as I have stimulus money that can go with no strings to you if it would help in some way. I am very sure that others feel this way as well.
Thank you for all that you do to support this community...
Offline
For GW Johnson ...
Best wishes for a smooth recovery of your data!
I've learned this recently, so there might be something you could use (or a helper could use) ... there are adapter devices for small hard drives (such as may have been installed in your laptop). Unless the problem is the hard drive, which is a ** different ** problem altogether.
In any case, thank you for taking another look at the delivery challenge.
I particularly appreciate the tip about a specific velocity that US military aircraft are designed to fly at the moment ordnance is released. I'd not heard of that before.
SearchTerm:ordnance release velocity for US military aircraft GW Johnson Post #211
(th)
Offline
For SpaceNut ... by any chance, do we already have a topic for discussion of collecting atmosphere from Mars by a moving vehicle to make fuel and oxidizer.
I definitely remember the subject coming up, but I cannot remember where that happened. I ** think ** that Void may have been an one of his creative periods at the time.
GW Johnson has (in effect) issued a challenge for solution of the refueling at Phobos problem. I have posited capture of atmosphere by a vehicle designed for the purpose. Science Fiction writers have been describing such vehicles for decades, but they have the distinct luxury of not having to worry about the details.
In this case, we would need to be able to fly down into the upper level of Mars' atmosphere, collect enough CO2 to make fuel for that flight and at least one more, and return safely to Phobos.
The British SSTO design imagines collecting oxygen on the way to space, cooling it with hydrogen so it can be burned in an engine, and ... here I'm vague on details possibly ? using it as oxidizer for a rocket phase of the flight?
In any case, it is a model that implies capture of atmosphere during the flight of a vehicle, although cooling with hydrogen would be challenging for a vehicle based at Phobos.
(th)
Offline
The insitu co2 collection of a moving vehicle is a means to waste energy as the amount of improvement via the inlet is made up for the losses of the moving vehicle on the surface....
As for it being the rocket the amount collected is not going to be all that much as the thick part of mars atmospher is close to the planet where it needs to collect it. Its moving to fast through it to make use of what is there. The amount which could be liquified for use by the manufacturing of fuels on Phobos would be about all you could do with multiple trips....
We have a topic for co /co2 rockets where I am wondering if you start with liquid co and mix the sabre like reaction to cool co2 from the air would be sufficient? to make that style of rocket possible.
did a bump to a couple but there are more that could relate for mars
Offline
Skylon is the aircraft I was thinking of...
https://en.wikipedia.org/wiki/Skylon_(spacecraft)
Static testing of the engine precooler began in June 2011, marking the start of Phase 3 in the Skylon development programme,[25][46][47] In April 2012, Reaction announced that the first series of the precooler test programme had been successfully completed.[48] On 10 July 2012, Reaction announced that the second of three series of tests has been completed successfully, and the final series of tests would begin the following month after the testing facilities had been upgraded to allow testing of −150 °C (−238 °F) temperatures.[49][50] ESA's propulsion division audited the precooler tests during mid-2012 and found the results satisfactory.[9][51]
The Wikipedia article indicates that the concept is still under development, and that funding is still continuing.
I'm interested in the cooling aspect of the design, since (as you point out in #215) the vehicle would be moving swiftly to accumulate CO2 molecules and still be able to return to Phobos.
(th)
Offline
Skylon is planned to operate as a liquid air cycle airbreather from the surface up to about Mach 5 at about 25 or 30 km altitudes. Martian surface density is about like Earth at 40 km, and much thinner yet, further up. Entry altitudes on either planet are higher yet, with even thinner atmospheres. Compressing stuff that thin with an inlet is useless, because 4-5 times nothing is still nothing. Likewise, applying mechanical compression is still useless, because 10-20 times nothing is still nothing (pressure ratios typical of modern axial-flow gas turbine compressors).
Skylon's Sabre engines compress air and (I think separate-out the oxygen) and store it aboard, way down in the thicker portion of Earth's atmosphere under 25 km, and at speeds below Mach 5, even subsonic. That way, it need not lift the total load of oxygen off the surface. From the Mach 5 point, Sabre is nothing but a hydrogen-oxygen rocket, using the oxygen stored aboard. The cold of the hydrogen is integral to the air compression and liquefaction process. You can't do that with most other fuels.
Those kinds of physics are why I am exceedingly skeptical of plans to scoop up and contain anything so thin as the Martian upper atmosphere (or anything else at typical entry altitudes on any planet). Stuff that thin is still quite voluminous even if you compress its pressure by a factor of 100 to 1000 (something not technologically possible today with fast-operating equipment light enough to fit inside a vehicle). The compressed material just cannot fit inside a practical-size vehicle.
There are some very restrictive design constraints imposed by hypersonic flight, things the designers of Skylon have yet to address. Those engine nacelles at the wingtips will survive ascent, but that configuration will not survive entry/descent: the compression spike shock waves will cut the wings off! While hypersonic, even in the thin air, heating rates are magnified by a factor of at least 7, at only Mach 6-7, where a shock wave impinges directly upon an adjacent surface. Heating rates that high are just not survivable with the materials we have at this time in history.
I refer you to the late-1960's X-15 flight to Mach 6.7, testing both a ceramic coating and a scramjet test article on the ventral fin stub. The coating didn't work, primarily because its reflectivity interrupted re-radiation cooling (it was bright white in color). But the big issue was the spike shock from the scramjet article: it nearly cut the tail off the plane. Whose structure was largely Inconel-X. I posted photos of this result on "exrocketman" some years ago.
So I am also skeptical that Skylon as envisaged will ever fly back from orbit. Not without big, heavy, thick, replaceable ablative heat shield panels on its wings, and a lot of active cooling behind them. (There went the payload capability!)
GW
Last edited by GW Johnson (2021-04-17 09:09:12)
GW Johnson
McGregor, Texas
"There is nothing as expensive as a dead crew, especially one dead from a bad management decision"
Offline
Following up on the challenge from GW Johnson, I asked Google for help.
Google is a nascent AI, evolving before our eyes as we humans feed it requests, and as it is pushed and prodded by its developer/trainers.
The question I posted is: gas+capture+by+flying+scoop
The first response was for a company that calls itself the Flying Scoop ... it produces advertising media
I can't fault Google for that one ... it admits "missing gas and capture"
However, after that near miss, it found Dr. Robert Bussard's ramjet
Bussard ramjet - Wikipedia
en.wikipedia.org › wiki › Bussard_ramjet
The Bussard ramjet is a theoretical method of spacecraft propulsion proposed in 1960 by the ... The scoop field funnels interstellar gas into an "accelerator" (this could for example ... Although the prelaunched fuel for the ramjet negates one advantage of the Bussard design (collection of fuel as it ... The Star Flight Handbook.
The sixth item down appears to be a reasonably good match for what ** I ** have in mind, which Google cannot (of course) know ...
Low Earth Orbit Atmospheric Scoops - ToughSF
toughsf.blogspot.com › 2017/09 › low-earth-orbit-atmospheric-scoops
Sep 11, 2017 · The relative proportions of each gas this fourth step produces is determined by the atmospheric composition at the collection altitude. Number of ...
http://toughsf.blogspot.com/2017/09/low … coops.html
This turns out to be an article well worth a few moments of time (for me at least) ....
It contains some plausible looking math and reasonable/familiar looking statistics ...
It is focused upon orbital refueling, and considers a variety of ways of achieving that...
The Phobos-Phobos delivery scenario would be an example of orbital refueling ... so far so good.
Now we're (possibly) getting somewhere ....
Atmospheric Gas Scooping
This concept consists simply of running a gas scoop through the upper atmosphere and collecting the atmospheric gasses to be used as propellant. Some of the propellant is consumed by the scoop itself, the rest of made available for orbital refuelling of other craft.
The article provides a picture with this subheading ...
The PROFAC gas scoop operated at 120km altitude and was powered by a 10MW reactor (4) and had some rocket fuel on-board to burn with collected oxygen to produce thrust in case the reactor failed.
The authors spend several paragraphs laying the ground work for this:
We can quickly work out that propulsion systems with an exhaust velocity just above orbital velocity will have a very hard time collecting propellant. An ideal engine has an exhaust velocity several times the orbital velocity. However, as we will now see, getting enough power to these engines is problematic.
The proposed ISP for a Phobos-Phobos vehicle is 245. The velocity of Phobos in orbit is 2.138 km/s (per Google/Wikipedia)
I haven't learned how get from here to there in this situation, so can't evaluate how the 245 ISP vehicle exhaust velocity compares to 2.138 km/s.
The authors continue to the subject of drag to be experienced by an atmosphere scoop vehicle:
The power requirements of a gas scoop are determined by the thrust it needs to produce, which depends in turn on the drag it experiences. Aerodynamics play no role at orbital velocities, drag is only a function of cross-section and atmospheric density. This equation will give you an estimate within +/-10% of reality.
Drag: 0.5 * Orbital velocity^2 * Cross-section area * Gas density
This next quote is specifically for SpaceNut who posted a guess that molecules to be collected at altitude would be "not much" ...
Gas density is determined by altitude and noted in kg/m^3. If the scoop operates at a single altitude, we use a single value. If it changes altitude, we use an average value based on the time spent at each altitude. At very high altitudes, some sources will note a value in particles per cubic meter. This has to be converted into kg/m^3.
The article includes a chart showing "Particle counts for different elements..." (for Earth of course)
I note a detail that may be of interest to some forum readers ... The authors specify that the scoop should cover the entire frontal area of the vehicle, so that there are no losses to drag by non-productive surfaces.
So how much power do we need?
We first must estimate the drag generated and give the scoop ship enough power to produce sufficient thrust to counter the drag, plus a safety margin.
This is ** precisely ** what I'm looking for in the case of the Phobos-Phobos shipping business.
How this is going to turn out is unknown (to me for sure) but I have learned from GW Johnson not to be optimistic.
The authors provide development of the background needed to determine/predict engine power needed at various altitudes.
The engine power here is the effective power output going out of the nozzle. Rocket engines are not 100% efficient, and there are further losses in the systems that generate, transport and convert electricity going into the engine. These might double the actual power consumption.
There are many options for producing the required energy to power the engines, but only a few are practical or achievable with the technologies available in the near future.
Oh oh!
The two options remaining are solar panels and nuclear reactors.
Solar panels are considered but quickly discarded for this application. (Edit#1: I subsequently decided solar panels would work at Mars in diving scenario)
Nuclear reactors are very powerful and are unlikely to need much mass or volume to produce the output required. Experiments have been conducted and prototypes have been flown of space-rated nuclear reactors, but historical and political reasons have prevented their widespread use. We can send a 100kW reactor today into space massing only 512kg.
There ** is ** a scenario where solar panels would make sense for this application. If the vehicle is on a high elliptical orbit, then it can spend most of it's time out of the atmosphere recovering from the most recent scoop pass.
For that reason, I note the author's admonition to keep the solar panels behind the scoop. They would play no role in powering the vehicle during a scoop pass, and indeed the momentum of the vehicle would be bled off as it collects molecules and accelerates them to orbital speed.
However, the authors now continue on to the mechanism for gas collection using a scoop ...
Ah ha! Here comes the Skylon connection ...
The authors propose cooling the collected gases and saving them as liquids ...
In the case of CO2 the logical destination/state is solid, since the liquid state requires pressure.
The authors continue at great length and I recommend this article to all forum members who share an interest in the prospect of "mining" useful material from the atmospheres of various bodies.
However, since the article is about ** Earth ** they point out that they would collect Nitrogen as well as the more useful Oxygen.
They also point out the energy savings of having a kilogram in orbit vs on ground.
I deduce from this first reading that whatever the possibilities for this method to work at Mars may be, the mining of Oxygen and Nitrogen from the atmosphere of Earth for delivery to customers on Mars looks quite viable.
Oh ho! So my anticipation of the high elliptical orbit is recognized by the authors!
There is a design that allows a scoop ship to get away with much smaller engines and power sources. A 'diver' operates at an elliptical orbit, dipping into the atmosphere only at its lowest point. At that point, it rapidly scoops up gas and rams its way up and out with its momentum. For the rest of the orbit, it fires up its small engine to recover the lost momentum over a long period of time. Much smaller solar panels would be needed to feed propulsion and gas processing equipment. If the diver goes into the lower atmosphere with excess velocity, like an aerocapture maneuver, it could use a comparatively tiny scoop to collect a lot of gas, so overall the diver scoop will be much smaller and lightweight compared to other designs.
Well, there had to be a downside ...
The downside is that ramming through the lower atmosphere at higher than orbital velocity involves significant heating which lowers the ship's useful life. Collecting gasses quickly means that some intermediary store must be available before they are processed and liquefied: holding large volumes of hot gas would require voluminous tanks that add drag and weight to the scoop ship.
There is an (impressive to me for sure) series of interactions with the author after the article itself.
I'm coming away from that presentation and discussion with a sense that an atmosphere collection system for Mars is conceivable, although I'm not sure it is feasible. The diving mode makes the most sense to me, because photoelectric cells can collect power during the vast majority of the high elliptical orbit, and the momentum of that high orbit is consumed by collecting molecules during the atmosphere pass.
The collected gases have to be cooled, and in the case of Mars the logical end state is dry ice. Other gases collected during a pass might be of value.
The dry ice needs to be delivered (somehow) to processing facilities on Phobos.
I conclude that this article (by Matter Beam) shows a potential solution to the challenge of GW Johnson.
(th)
Offline
I suspect the article author used full theoretical stagnation pressure in the duct behind his scoop. You won't get that, not even in rarefied flow. The most likely scenario is a tube behaving like a pitot tube, with essentially zero through-flow compared to the swept-out area and density as a flow rate. There will be an enormous bow shock out front, with overheated plasma moving subsonically through the scoop into the duct. At 7800 m/s, the effective temperature of that plasma will be in the neighborhood of 7800 K. You would be lucky to recover even 1% of the theoretical stagnation pressure.
GW
GW Johnson
McGregor, Texas
"There is nothing as expensive as a dead crew, especially one dead from a bad management decision"
Offline
For GW Johnson re #219
Thank you for taking a look at the the imaginings of Matter Beam. In a scan of the feedback to the original article, I did not see any reference to the issue you have raised, so I'm taking it as a suggestion to be skeptical of the concept, since someone with your background must not have been in the group of commentators.
The ** only ** possible antidote to the heating you've predicted is cooling at a rate comparable to what Skylon is intended to achieve, and ** they ** are (reportedly) using liquid hydrogen to cool incoming air sufficiently to allow it to be fed into a subsonic combustion chamber.
The author (Matter Beam) ** does ** appear to be imagining a method of cooling, but that part of his paper may not be achievable in this universe.
What ** does ** seem true (in this case) is that the author has made a valiant effort to step beyond the well-trod science fiction scoop-ship trope.
Thanks again!
SearchTerm:Analysis of Scoop ship proposal by Matter beam by GW Johnson Post #219
SearchTerm:Scoop ship concept paper by matter Beam reviewed by GW Johnson
(th)
Offline
I'm not saying these things won't work. But an upper bound estimate being the real-world delivered performance is a very rare thing indeed (I have never seen such). You get a lot closer estimate if you go with a lower bound. From there, it may improve a bit a real life as we learn more about the technology details.
GW
GW Johnson
McGregor, Texas
"There is nothing as expensive as a dead crew, especially one dead from a bad management decision"
Offline
For GW Johnson re #221
Thank you for your continued interest in this topic! I have started an investigation of the nature of quantum level interactions between gas particles and solid (bound by cohesive electron bonds) matter. I've been too busy with other interests/duties to pursue it, but it is there waiting for someone to invest the time that will be needed.
The object of ** that ** topic is to try to understand (a) why elastic particles (molecules) bounce away from solid molecules instead of sticking to them, (b) to try to imagine a way to capture them. It ** is ** intuitively obvious that since electrons repel each other, electron shells of atoms would repel each other, unless an atom has enough momentum to overcome the objections of the electron shells and penetrate far enough to dislodge the target atom from the matrix of which it is apart, as happens when solid material (ie, heatshield) is exposed to gas molecules at sufficiently high speed.
The Skylon approach ** appears ** to be to cool the metal of the funnel collecting gases so the arriving molecules do not dislodge the metal and (b) to absorb energy from the elastic molecules as they bounce around frantically in the confines of the tunnel where they are being collected.
The ** ideal ** situation (possibly not available in the ** real ** Universe) is for the gas molecules to quietly accept capture by the arriving heat shield, to accept acceleration to the velocity of the heat shield without frantically trying to escape, and to contribute to the momentum of the heatshield vehicle by decreasing that momentum by the amount of the mass of the molecule.
To my knowledge there is ** no ** way to achieve that ideal, but I think the topic is worth exploring.
If the ideal could be achieved, then the molecules in the path of the vehicle would pile up in front of it in an accumulating stack until the vehicle releases them, at which point they would have a velocity lower than was the case when the descending vehicle first showed up.
(th)
Offline
Hi TH & Spacenut:
Data recovered from dead laptop. I still have to get another laptop. Not sure when that will occur. But once it does, I plan to look at the cargo to Phobos/vehicle to Mars and back to Phobos thing, dropping cargo at about M3/5 km, and using solid rockets to slow it for a survivable impact.
That leaves the Phobos refueling question unaddressed.
"Isp 245" will probably not look very good for this, and MMH-NTO is in the 300-320 range. If there is water inside Phobos, you could reasonably presume hydrogen-oxygen propulsion. I will probably look at that.
GW
Last edited by GW Johnson (2021-04-24 16:03:57)
GW Johnson
McGregor, Texas
"There is nothing as expensive as a dead crew, especially one dead from a bad management decision"
Offline
For GW Johnson re #223
Congratulations on successful recovery of your data from the laptop hard drive!
***
This next comes from someone who never owned an Apple product (except for 1 (one) digital audio device). The members of the tech groups of which I am a member seem to ** all ** be running on Apple lap and desk tops. They run Windows, Linux and Apple software simultaneously on these machines.
I have no idea how any of the Apple products compare to other company products, but assume they are more expensive.
***
Thanks for deciding to look at the Phobos-drop-Phobos scenario. I assume it is technically feasible but the trades may not allow it to be competitive with the soft landing folks. Naturally, I'm rooting for the sweet spot you (may) find to be competitive, but am (realistically) not holding my breath.
That unexpected figure-of-speech is a direct result of reading your reminder of the history of accidental decompression incidents.
***
The refueling on Phobos scenario would work best if it is possible to collect atmosphere (CO2) to make whatever fuel is going to be used for the drop off.
A supply of Hydrogen readily at hand would help, but other than shipping quantities of water from Earth there is no (presently known) source.
(th)
Offline
TH re post 222:
I don't know much about atomic theory, or the kinetic theory of gases, much less quantum stuff. We mechanical engineers don't use that sort of thing for our design analyses, because it is way too hard to manipulate for design purposes, and simply does not address the real-world situations we have to address.
For aerodynamical/fluid flow problems, we generally use the "continuum flow" math models, which take no account at all of individual atoms and molecules. Although, the ideal gas equation of state is a correlation of effective density in terms of pressure and temperature, in which pressure can be understood by the momentum reaction of gas molecules impacting a solid wall, and temperature can be understood as the average kinetic energy of the molecules.
But we don't care about the motions of those molecules. All that is imperceptibly small. We can measure temperature and pressure directly with simple devices, and deal with them directly in our continuum math models. A theory that gets us answers we can trust about the problems we have to solve, is useful. One that does not, is not useful, no matter how "accurate" it might be. That is perhaps the true distinguishing characteristic of mechanical engineers versus physicists and other scientists.
Most folks are familiar enough with the incompressible-flow math model of continuum flow mechanics to recognize it. That's where the classic textbook "streamtube variation" and "Bernoulli equation" stuff comes from. But, once speeds reach a point where density is no longer able to be modeled as a constant, you have to use the compressible fluid flow math model, a version of continuum flow mechanics that presumes variable density, and relates it to temperature and pressure, by the ideal-gas equation of state. The switchover to compressible flow analysis takes place in the range of half to 80% of the speed of sound, depending upon how accurate you want to be. It's a fuzzy boundary.
Ideal gas compressible flow (which looks very complicated to those not used to doing it), is a rather accurate model that serves quite well, right up to the point where the ideal gas equation of state no longer holds as an accurate model for density. Generally speaking, that point is in the Mach 6 to 8 range of flight speeds. Again, it's a fuzzy boundary. Depends upon just how accurate you have to be. A bit more precisely for the case of air flow, it is when the air starts ionizing significantly, and is no longer "air". That's at air thermodynamic ("static") temperatures around 5000 F. Again, a fuzzy boundary.
Above that limitation, you need to use empirical relations to determine density vs temperature and pressure for your specific mixture of gases, usually expressed in terms of enthalpy and pressure, and in the old days these were paper charts, entirely based on experimental data. Those charts gave you the better answers when real gas effects prevailed. Although, the ideal gas continuum fluid flow model is still a useful tool for a conceptual understanding of what is going on, despite the real gas errors. (That is in fact the case, even for re-entry analysis.)
The old entry rule-of-thumb, reflecting this disparity of real gas versus ideal gas, was that the effective thermodynamic temperature driving both convective and radiative heat transfer, if expressed in deg K, was numerically equal to the flow speed, expressed in meters/second. That's a crude approximation, wrong by about 10%. The difference between that and the ideal gas compressible-flow analysis estimate of stagnation temperature, was explained as kinetic energy no longer going solely into internal energy, but also partitioning into ionization. It's a rather good explanation.
There is an altitude above which the molecules are far enough apart that the "continuum flow" math model just inherently fails to be accurate, in that you begin to see disparities between what you would expect, and what you actually see, for pressures and temperatures. That is somewhere above the 50-60 km altitude at which meteors are most brightly burning, but lower than the nominal 140 km entry interface altitude.
But, it is still higher than the altitude where peak entry heating and peak deceleration gees occur. So the error using continuum flow math models all the way from the entry interface altitude is small, because the accelerations you get way up there are so very small. Twice nothing is still nothing as an error. Where the drag effects are largest, the continuum math model works just fine.
Which in turn is why my 1953-vintage crude entry spreadsheet model still works as remarkably well as it does. It's just not worth the effort to go to a free-molecule flow model above 50-60 km. You get just about the same answers regardless of which model you use, and kinetic theory of gases requires massive computer simulations, while continuum flow does not, not even real gas continuum flow. Continuum flow is just far simpler and easier to use. True for Earth. True for Mars. So far, true for anywhere else with a sensible atmosphere that we have been.
OK, that being said, using continuum flow, there is some point somewhere on the surface of the heat shield where the surface flow velocity behind the bow shock has decreased to exactly zero. That is the so-called "stagnation point". Pressure (and stream gas temperature) there, is the very highest that it can be. If symmetrical, this is in the center of the heat shield. If the vehicle is angled to generate "lift", it is located off-center. But, it is there, somewhere!
Away from the stagnation point, local flow is nonzero velocity, but still subsonic, usually to somewhere at (or near) the periphery of the heat shield. Pressure (and stream temperature) are still quite high, but decreasing as flow velocity increases. There is a peak in the boundary layer temperature distribution radially away from the surface, but it is not a large peak, in this subsonic flow region. The value of that peak is the "recovery temperature" in "continuum flow" mechanics and heat transfer, and its value is pretty much indistinguishable from what we term "stagnation temperature". It is the driving temperature for heat transfer to the surface of the heat shield, both convective and radiative.
Further away from the stagnation point, the flow goes supersonic again. This is usually near the heat shield periphery, for most practical heat shield designs. Pressures (and temperatures) of the flowing gas stream are lower there, although the stagnation and recovery temperatures are pretty much the same as they are at the stagnation point. So the driving temperatures for heat transfer are unchanged! These show as a large peak in the boundary layer temperature distribution. What is different is local pressure (and thus density), which is very much lower as the surface flow re-accelerates supersonic. That low pressure reduces the "heat transfer coefficient". And thus it reduces the heat transferred per unit area in this region, well below that seen near the stagnation point.
What usually also happens is flow separation near the heat shield periphery, just downstream of the sonic point. This is not predictable from fundamental theory, it is entirely empirical! It just happens. Every shape is unique. That leaves the aft surfaces of the body in a wake zone at very low velocities and modest pressures at or just below ambient, scrubbing the surface, with the high (and accelerating) main flow velocities not scrubbing the surface at all! Heat transfer coefficients (and heat transfer rates per unit area) are quite low here, despite the high stagnation and recovery temperatures. All this heat transfer stuff is entirely empirical. No theory predicts this "from scratch".
The separated wake more-or-less closes somewhere behind the body, with velocities (and pressures) outside the wake pretty much once again the same as freestream values ahead of the body. Actually, there is a sort of wake "core" that persists, in which velocities are lower, and temperatures higher, than freestream. That is the incandescent streak left behind by meteors and most re-entry vehicles. It can persist on time scales from a few seconds (tiny meteors) to a few minutes (the rather large space shuttle).
The stagnation point values of convective heat transfer per unit area pretty much correlate in terms of velocity^3, and both density^0.5 and radius^-0.5, for pretty much any entering shape. Below about 10 km/s velocity, you can pretty much ignore radiant heating from the plasma. Above about 10 km/s speed, radiant heat transfer quickly dominates, and correlates on velocity^6 power. It is by the body shape that local velocities and pressures and heat transfer rates per unit area away from the stagnation point get correlated empirically. Every shape ultimately has to be tested in a hypersonic wind tunnel to generate or confirm its correlations.
GW
Last edited by GW Johnson (2021-04-25 09:40:13)
GW Johnson
McGregor, Texas
"There is nothing as expensive as a dead crew, especially one dead from a bad management decision"
Offline