Compressed gas balloon rocket for Mars launch

SpaceNut · 2020-06-11 19:01:30

Here are a few more links for the mass shift of inflateables.

https://ntrs.nasa.gov/archive/nasa/casi … 014351.pdf
Estimating Mass of Inflatable Aerodynamic ... - NASA

https://ntrs.nasa.gov/archive/nasa/casi … 031869.pdf
Aerocapture Inflatable Decelerator for Planetary Entry - NASA

This is what parts have been through testing concepts that have similar dynamics

GW Johnson · 2020-06-12 15:36:55

I took a look at the balloon rocket idea, restricting it to oncoming wind pressures as the max possible pressurant level inside the balloon. Even with no losses, you must be moving Mach 1 to barely choke the engine nozzle, and that's with no expansion cone at all. With real flow losses, this is closer to Mach 1.1, just to get a throat choke. So I don't think there is much potential in this idea.

Why? because there is no wind pressure at all upon the forward surface of the balloon when sitting still, and it takes (in CO2) a 1.85 chamber-to-ambient pressure ratio to even choke a sonic-only nozzle throat. There is very little thrust coefficient CF at 1.85 pressure ratio with 1:1 "expansion". So there is very little thrust, primarily due to the low Pc, and secondarily due to the low CF.

You cannot get an internal pressure in the balloon equal to the pitot pressure, unless the balloon is contained within a hard tube with the open end facing forward. An exposed balloon will see pitot pressure at its stagnation point, decreasing around its sides to something very near atmospheric ambient along the lateral sides. Whatever that distribution integrates to, that average is your internal pressure. (And you lose a tad of that in flow losses getting into the chamber.)

Pitot pressure means local stagnation (or total) pressure, Which is also reduced by the normal shock pressure ratio once you are supersonic. THAT is what a pitot tube facing directly into the wind will read. And the dynamic pressure that you need for drag is NOT the difference between pitot and static, the way it is for well-subsonic flight. Those things are calculated with the equations of the compressible flow model. I know quite well how to do that. I did it, and that is too much detail to describe here.

What you have to do is sharply pressurize the balloon with your gaseous propellants, so that it will work statically as a pressure-fed engine. Otherwise, this concept will never work for launch on Mars or anywhere else. And you need a high pressure indeed to get decent thrust out of your engine. Megabars, not millibars.

Your thrust potential at pitot pressure is quite low compared to drag, even in that near-vacuum of an atmosphere. At 1.85 sonic-only (around Mach 1.1) on Mars, your surface ambient pressure is 6 mb, and so your full-pitot-pressurant chamber pressure is about 11 mb. CF will not be much different from 1. F = Pc At CF = (0.011 for the milli)(100,000 N/m2 for the bar) (At in m2)(1.0 for CF) ~ 100 N thrust per square meter of throat area. N, not KN!

Drag on the balloon is CD q Aref, where CD will be near .3 subsonic, near 0.8 transonic, and dropping back toward 0.4 low supersonic. Aref is the balloon frontal cross section area. At Mach 1.1, q = 0.5 gamma P M^2 ~ .65 6mb 1.21 = ~ 5 mb. So drag D = ~ .8 (0.005 for the milli) (100,000 N/m2 for the bar) (Aref in sq.m) = ~ 4000 N per square meter of frontal cross section area.

4000 N/sq.m of balloon cross section is a whale of a lot bigger than 100 N/sq.m of engine throat area. And THAT is why I said what I said!

Sorry to be a balloon-buster, but that's just the physics of high-speed aerodynamics and of rocket nozzles. It was a fun idea to investigate, but it will lead us nowhere.

GW

Last edited by GW Johnson (2020-06-12 15:37:30)

SpaceNut · 2020-06-12 19:15:05

So lets go learn about pressure feed engines in the mass of payload on the top of the balloon once balloon is filled will cause internal balloon pressure to rise with the fabric bag being the restrictor to expansion of the balloon. The cable winch system to increase force pressure for the tank would add to the drag force for feeding fuel at higher pressure for accelerating the rocket....

So posting links for all as I learn about liquids versus a gas at which its a volume to psi equation where as the liquid is density to flow.
https://en.wikipedia.org/wiki/Pressure-fed_engine

The pressure-fed engine is a class of rocket engine designs. A separate gas supply, usually helium, pressurizes the propellant tanks to force fuel and oxidizer to the combustion chamber. To maintain adequate flow, the tank pressures must exceed the combustion chamber pressure. To maintain adequate flow, the tank pressures must exceed the combustion chamber pressure.

For forcing a liquid fuel into the engine is heated helium to cause the flow rate on earth so if the engine is the same engine for earth then why so low a chamber pressure for mars if we are supplying it with the same source pressures. Sure more cubic feet of fuel flows in gaseous form for liquid to burn.

https://courses.lumenlearning.com/physi … ropulsion/

https://mrbremer.wordpress.com/2014/11/ … le-design/

tahanson43206 · 2020-06-13 06:34:46

For GW Johnson re #27

First, thank you for taking a look at Calliban's idea! I am thinking of making a sketch to show what I think your vision might have been. I refreshed my understanding of Monsieur Pitot's invention, because (as I understand your presentation) you used that as a model for the hypothetical Mars launcher.

For SpaceNut re #28 .... thank YOU for your interest in the learning opportunity Calliban has provided, and which GW Johnson has just enhanced with a significant dose of "real world" analysis! I am encouraged by your willingness to keep pushing the envelope (pun intended!). I think your cable crank idea is interesting, but there may be an alternative that is worth considering.

In another post, I suggested using an ordinary rotary turbine to pull oxygen gas from the balloon to feed into the reaction chamber. There will already be a rotary turbine (or conceivably a piston pump) moving kerosene from the fuel tank to the engine.

What I'm hoping you will do is to to investigate to see how much space must be allocated for the gaseous oxygen to permit a successful flight.

For GW Johnson ... Please develop your ideas for the lander component of your Earth/Phobos/Mars comprehensive transportation system. In particular, I'd like to know (on behalf of those who are trying to understand the potential of Calliban's idea) the amount of kerosene you would need, and the amount of liquid oxygen you would need, to deliver the lander and its cargo back to Phobos.

Given the quantity of oxygen you would need, SpaceNut (and perhaps others) can compute various combinations of volume and pressure that would be needed to accommodate the oxygen in gaseous form.

I am NOT expecting to succeed with this exploration, but the effort will leave behind (on this forum) a trail of work that can be studied in future times by others who may be curious about the question. In this way, the result would be similar to what I have seen on Quora, with the unique flavor that the NewMars forum has established in its 20+ years.

Finally, as a reminder, I had proposed (earlier in this topic) that we consider a central shaft/spine/core for the vehicle, which would NOT employ the traditional outer shell made of metal used by every rocketeer on Earth since the days of Goddard and perhaps even before.

If the rocket is designed with a central shaft to bear the loads imposed by the oxygen and nosecone, then the volume around the shaft can be allocated to fabric cylindrical storage for gaseous oxygen.

I don't have a good feeling for the volume of oxygen that would be required to achieve orbit.

Google came up with a citation form sciencing.com that considers the volume of oxygen that is obtained from evaporation of 70 liters of liquid oxygen.

I'll have to come back to this later today, when I am on a modern computer.

Edit#1: In GW Johnson's Phobos/landing topic, I have offered the suggestion that a vendor competing for the opportunity to refuel GW's lander will need to make sufficient oxygen and fuel ahead of the planned flight, ** and ** to schedule liquefaction to finish just as Gw's flight plan calls for it.

While gas is being produced, it will be in gaseous form, and it must be stored in containers while it awaits liquefaction, where it will be stored in different containers with characteristics unique to the liquefied state.

The work done in **this** topic an be productive for the purposes of the vendor on the surface, regardless of whether the balloon fed rocket itself is practical.

In other topics in this forum, work is underway to deliver cargo and passengers safely, reliably, and cost-effectively to the surface of Mars, using vehicles that are designed to offset part of the pull of gravity of Mars by holding some of the atmosphere at bay using hydrogen filled volume.

It has been proposed that the hydrogen enclosure material can find application on Mars after a landing.

One of the applications for which such an enclosure can be put to immediate use is storage of gas for fuel production facilities.

In addition, in another topic recently started by Void, there is discussion of various designs for greenhouse enclosures.

Among the designs under discussion is the simple fabric cylinder, and by happy coincidence, that is exactly the form that a gas filled enclosure for a Slow Mars Lander would take. Thus, in addition to sales to vendors of supplies for rocket launches, the lander company can provide ready-to-go enclosures for greenhouse builders.

(th)

Last edited by tahanson43206 (2020-06-13 07:27:23)

SpaceNut · 2020-06-13 16:43:10

From documents
LOX/CH4 propellant at mixture ratio = 3.45 delivering Isp = 360 seconds
Propellant – Retro-Propulsion 6,821 kg

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

Ascent Mass 94,242 kg dry plus fuel for launch

Volume of oxygen Gas at 70°F (21°C) and 1 atm 38,633 m3 but for mars temperatures we will be lots less in volume

volume of methane 26,803 m3 and since the oxygen and methane are similar it as well will be less in volume as well.
The volume ratio of gaseous to liquid methane is approx. 1:600

The greater the pressure when cooled down will mean the volume of the tanks will be smaller.

tahanson43206 · 2020-06-13 18:36:51

For SpaceNut re #30 and continuing development of this topic ...

Google computed the cube root of the volume you gave for Oxygen as = 33.8054051 meters. (just under 112 feet)

Google computed the cube root of the methane volume you gave as = 29.9268588611 meters (just under 100 feet)

For comparison, the volume of the Hindenburg was:

Length: 245 m / 803.8 feet. Diameter: 41.2 m / 135.1 feet. Gas capacity: 200,000 cubic meters / 7,062,000 cubic feet.
Hindenburg Statistics | Airships.net
www.airships.net › hindenburg › size-speed

The figures you quoted for the gaseous propellant Mars launch vehicle ore "only" 38,663 m3 and 26,803 md for a total of 65466.

I am wondering if the volume you quoted for oxygen would be much greater than the amount shown for Earth standard sealevel pressure.

Balloons swell on Earth as they rise from the surface, so I would expect the same behavior on Mars, which starts with the equivalent of a significant elevation on Earth.

All in all, this discussion has certainly confirmed the advantage of liquefaction of propellants << grin >>

I'm going to throw out a guess that if you want to achieve the 38,633 m3 volume of oxygen on Mars, then the balloon is going to have to be able to withstand the equivalent of Earth sea level pressure.

101.325 kilopascals
Standard sea-level pressure, by definition, equals 760 mm (29.92 inches) of mercury, 14.70 pounds per square inch, 1,013.25 × 103 dynes per square centimetre, 1,013.25 millibars, one standard atmosphere, or 101.325 kilopascals.May 27, 2020
atmospheric pressure | Definition & Variation | Britannica
www.britannica.com › ... › Earth Sciences

So what fabric can hold a standard atmosphere of pressure?

I decided to ask Google about parachutes:

Parachutes exhibit a rather narrow range of pressures between points in and around the canopy, even during inflation (around . 25 PSI), but absolute pressure sensors have a very wide range (typically 0-15 PSI).
The Use of Pressure Sensors in Parachutes
www.pcprg.com › psexcerp

So it would appear that a parachute material could withstand sea level pressure with some safety margin.

That material is (most likely) not impermeable, so gas would leak through without a lining.

***
Air bags are a technology on Earth that have received some study. I found a wide variety of citations but the consensus seemed to be that in operation, air bags used for collision protection experience about 5 PSI.

I'm inclined to let this topic come to a relaxed conclusion.

We can think of it as all the gas leaking out into the atmosphere of Mars.

Edit#1: Here is an example of a bladder manufactured for the US military:

The manufacturer advertises its products for liquids and "low" pressure gases.

(th)

Last edited by tahanson43206 (2020-06-13 19:06:52)

SpaceNut · 2020-06-13 20:24:25

The Bigelow inflatable on orbit as in genesis and the beam unit attached to the ISS all have 1 atm for internal pressure and that said they save 2/3 of the mass for a similar metal structure. That calculates into need for less fuel to achieve orbit.

The external bag is ,meant to contain the pressure of the filling of the balloon with Kevlar weaving.

The difference is how much does the volumes shrink when cooled but not all the way to cryogenic temperatures but we do know that space x does have composite tanks that can do that and they do have less mass as well.

Now to go learn some more...

tahanson43206 · 2020-06-14 06:37:42

For SpaceNut re #32

It appears we are egging each other on to pursue Calliban's will-o-the-wisp despite his having dropped it like a hot potato.

I am NOT unmindful of GW Johnson having carefully studied the proposition and found it wanting.

However, since (to my delight) you are forging ahead with this idea, I would like to point out that (a) your observation about the Bigelow design is helpful, because it provides a 1 atmosphere baseline and (b) because it reminds us of what the wall of a hypothetical Mars native un-refrigerated launch vehicle might look like.

We already have dimensions, thanks to your previous work a few posts back, and the example of the Hindenberg.

We can be reasonably confident that the vehicle we are discussing will be smaller than the Hindenburg, although at this point, I am not sure how ** much ** smaller it would be.

The central core can be postulated with some confidence, since the walls of the Bigelow design would not be load bearing.

It would be advantageous to be able to apply pressure greater than 1 atm to the tanks, to further reduce volume.

It has been proposed earlier in this topic to use turbine pumps to pull fuel (gaseous or liquid) from their tanks, and to use a suitable variation of that idea to pull oxygen from the oxygen tanks.

I'd like to offer an enhancement to the central core concept.

The central core can include feed pipes for the components of the propellant supply that are located along the spine, and the storage tanks (fabric) can be positioned (stacked) along the spine to keep the cross section of the vehicle as small as possible.

While drag in the rarefied atmosphere of Mars is much less than on Earth at the sea level launch site, it is a good idea to keep the cross section of the vehicle as small as possible. This means the volume of compressed gas will be arranged along the spine for as great a distance as is required.

If you want to run a few numbers, you could experiment with a calculator for the volume of oxygen and methane you have provided a few posts back.

Try a diameter of 10 meters for the vehicle, and see what height you come up with for the tanks.

I would presume it reasonable to add 10 meters for the nose section (navigation and passengers) and 10 meters for the aft section (propulsion and cargo).

I expect you'll come up with a length shorter than the Hindenburg, but the Hindenburg had a much greater girth, so my expectation could be wrong.

Something to keep in mind as we go forward is that the mass of the fabric enclosures, and the mass of the spine to support them, are being added to the total mass of the vehicle, so in the end, we may discover there is no solution.

This discussion takes into account that GW Johnson (directly) and Calliban (indirectly) have both indicated the proposition of a gaseous state for a Mars launcher is hopeless.

I am curious to see if the concept is even ** possible ** on Mars. It would certainly be more likely to work on the Moon (of Earth), but even there the dimensions of the vehicle would be impressive.

(th)

Last edited by tahanson43206 (2020-06-14 07:32:23)

SpaceNut · 2020-06-14 09:52:26

There is a new branch of the topic Compressed gas balloon rocket for lunar launch by RGclark

where by the drag component of launch has been removed that mars or earth would see.

a link from that topic:
https://en.m.wikipedia.org/wiki/Cold_ga … ropellants

One might use as the telescoping pole in the center a hydraulic system to make the payload and upper mass be stable until we are ready to light the engines at which time the fluid is slowly vented into a tank to save it for later. The mass of the payload and upper shell is the force that would be used to pressurize the tanks until its in motion.

tahanson43206 · 2020-06-14 10:34:27

For SpaceNut re #34

Thanks for noting the new branch from Calliban's topic here, with the interesting contribution by RGClark.

Earlier in this topic, GW Johnson published analysis which appears to show that a gaseous fuel/oxidizer system would fail to deliver mass to the rocket engine with sufficient force to succeed on mars. It seems to me (as I read your post) that you are not in agreement with GW Johnson's analysis?

Perhaps you ** are ** in agreement, and are arguing for a giant piston to drive the gas into the engine(s).

An alternative is to employ a mechanical pump (turbine or piston) to pull gas out of the tanks and force it into the engine at the pressures needed.

Your giant piston would (could?) have the same effect.

Mechanical devices to deliver quantities of gas into combustion chambers at the required pressure have been working on Earth for a while.

If we could (somehow) pull Calliban back into the discussion, I would be interested to know if any existing turbine or piston pumps could solve the delivery problem that GW Johnson has identified.

For SpaceNut ... if you feel up to it, please consider developing a conceptual model of a Mars native non-refrigerated launch vehicle, with a diameter of 10 meters, able to hold the quantities of oxygen and methane you identified as necessary in an earlier post.

I would be quite interested to know how tall such a vehicle would be. It might be as tall as a Saturn, for example, or perhaps even taller. It would be ** much ** less massive, of course.

In the Lunar topic, RGClark suggested looking at metallic or carbon fiber walls for the tanks. While I still think the central spine is a good idea, to avoid placing a load on the thin walls of the tanks, the fiber walls suggested by RGClark might be able to sustain their own weight, which would take that burden off the spine.

(th)

Last edited by tahanson43206 (2020-06-14 10:35:21)

SpaceNut · 2020-06-14 17:06:34

https://www.engineeringtoolbox.com/humi … d_677.html

p V = n Ru T
p = absolute pressure (N/m2, lb/ft2)
V = volume (m3, ft3)
n = is the number of moles of the gas present
Ru = universal gas constant (J/mol oK, lbf ft/(lb mol oR) = 8.3145 J/mol K = 0.08206 L atm/mol K = 62.37 torr /mol K
T = absolute temperature (oK, oR)

A summer day on Mars may get up to 70 degrees F (20 degrees C) near the equator so we will need to shade the fuel during the day and or make use of ground temperatures to moderate it. Night the temperature can plummet to about minus 100 degrees F (minus 73 C) for air temperatures.

Of course day and night the ground being way different. So fuels would pass from day through the ground storage to make it need less energy so to use less for making the fuels and oxidizer cold.oxygen cannot be liquified above a temperature of -119 degrees Celsius (-182 degrees Fahrenheit), no matter how much you compress it. At its critical temperature, it takes a pressure of 49.2 atmospheres to liquify oxygen. As you lower the temperature below critical, you need less pressure to liquify oxygen.

tahanson43206 · 2020-06-14 19:45:18

For SpaceNut re #36

Thanks for your report on considerations for liquefaction of oxygen, particularly as it applies to Mars.

In the parallel Lunar compressed gas rocket topic recently opened as a branch from this one, GW Johnson has provided helpful guidance regarding the use of a jet engine to compress gaseous fuel or oxidizer by passing combustion products over a power-take-off turbine. That guidance leads to the conclusion that the temperatures likely to be created will exceed the melting point of the turbine components.

Edit#1: Air (on Earth) contains (about) 79% Nitrogen, which does not participate in combustion (directly) but which DOES help to absorb thermal energy from combustion of fuel and oxygen, and thus (may be presumed to help) keep temperatures of the exhaust gases below what they would otherwise be.

Powering compressors with electricity is possible, but the working example (Rocket Lab) uses electrically powered pumps with liquid propellants. GW Johnson (earlier in this topic) advised that trying to compress unliquefied gas will be about an order of magnitude more difficult.

(th)

Last edited by tahanson43206 (2020-06-14 19:57:10)

elderflower · 2020-06-15 04:13:36

There are plenty of compressors around. These include vacuum pumps and jet engine compressors. To compress a large quantity of gaseous product from low pressure, they need to be very large and therefore quite heavy, due to the low density of gas entering the machine,

tahanson43206 · 2020-06-15 06:12:06

For elderflower re #38

Thank you for your contribution to this topic.

It was on life support, and about to receive last rites.

Grasping at this last opportunity to pull out of a spin toward the graveyard of unsalvageable topics, I'll toss out this question:

Is it possible to design a very large, ultra lightweight compressor that can realistically be driven by electric power long enough to complete liftoff from the surface of Mars? There are two use scenarios in play at the moment. The lunar topic is different because the gravitational pull is less but the likely fuel (hydrogen) is more challenging than methane would be.

In a recent post in another topic, GW Johnson reported the inert mass of SpaceX Falcon first stage booster as 5% of takeoff mass, which includes the second stage with fuel and payload.

Using that figure as a benchmark, in order for this (somewhat speculative) topic to survive, it will be necessary (or at least helpful) to try to design the vehicle so that the total inert mass is 5% of the total of which oxygen and methane in gaseous form are 95%.

It is probably not possible to design such a vehicle and stay within the 5% figure, but that is a good place to start.

Inert components (that I am aware of) include:

1) Core/spine (might be made of carbon)
2) Gas enclosures (light and strong)
3) Nose section with passenger compartment, navigation equipment, some cargo
4) Engine section with: compressors, batteries, rocket engines and plumbing

Another element of a gaseous fuel and oxidizer design is the problem of cooling the rocket engine.

That function is efficiently and effectively done by circulating fluid through channels in the rocket in vehicles using liquid fuel and oxidizer.

Google came up with a number of citations relating to the various approaches to rocket engine cooling that have been considered or put into practice.

Search Results
Web results
liquid rocket engine fluid-cooled combustion chambers - NTRSntrs.nasa.gov › archive › nasa › casi.ntrs.nasa.gov
PDF
Flow-Control Thermal Protection. Flow-Control ... or nozzle. Several methods exist for cooling the walls so that the temperature is maintained ... are circulated as coolants ... cooling- A porous inner wall is cooled by forced flow of coolant ... rocket engines used on tile Saturn. V vehicle, the ('enlaur stage, and the Titan and.
by C CHAMBERS - Related articles

and ...

Review on film cooling of liquid rocket engines - ScienceDirectwww.sciencedirect.com › science › article › pii
Liquid film cooling with fuel or oxidizer as the coolant can be employed in the ... a layer of coolant fluid between the surface to be protected and the hot gas stream. ... The overall dump cooled nozzle extension used on Vulcain-1 has been replaced ... Film cooling of rocket engines will play a major role in the development of ...
by SR Shine - 2018 - Cited by 7 - Related articles

For SpaceNut ... this topic is still lacking a figure for size of the (hypothetical) gas-only rocket for the Mars case.

An approach to arrive at that figure would be to plan a vehicle diameter of 10 meters, filled with the quantities of oxygen and methane you provided in an earlier post in this topic at one (Earth) atmosphere. Temperature is so variable on Mars that some benchmark is needed, so I'll offer 25 degrees Celcius as a figure to use for the calculation.

I'm looking for a length/height in a range between 100 meters and 200 meters.

The height/length is needed in order to be able to estimate the mass that can be allocated to structural components.

This topic has been speculative from the outset, so arriving at a solution which is both practical and doable seems highly unlikely.

Edit#1 ... in the Lunar branch of this topic, RGClark tossed out a reminder that a chemical reaction might provide energy for compression.

In the example given in that post, RGClark shows a link to a demonstration of physics. Liquid nitrogen was poured into a 2/3's full plastic bottle of water. The bottle was then inverted, and the physical interaction of the warm water and the cold nitrogen produced a vigorous reaction. A very lightweight compressor for the Mars gas-only rocket might be imagined, but it would require walls of the compression chamber to be strong enough to withstand the pressures needed for efficient operation of the rocket engine.

The pressure-fed engine is a class of rocket engine designs. A separate gas supply, usually helium, pressurizes the propellant tanks to force fuel and oxidizer to the combustion chamber. To maintain adequate flow, the tank pressures must exceed the combustion chamber pressure.
Pressure-fed engine - Wikipedia

It would be a remarkable feat of chemical engineering to design a pressure delivery system capable of gradual increase of pressure over an extended period of time needed for acceleration of the vehicle to orbital velocity. It may not be achievable by humans for thousands of years.

Edit #2: gunpowder is designed to burn slowly instead of rapidly, in order to create sustained but not overwhelming pressure behind a bullet to be accelerated along the bore of a weapon. Perhaps there exists (or might be created) a slow burning reaction capable of driving gaseous fuel and oxidizer toward the engine intake ports.

(th)

Last edited by tahanson43206 (2020-06-15 06:41:10)

SpaceNut · 2020-06-15 17:01:29

The liquid cryogenic fuels are auto expanding due to tank insulation and external heating such that the boiloff continues to pressurize a tank as its draw down and must be vented externally once it exceeds the release values. The use the engine heat exchanger to also heat the fuel to keep it pressurized.

While we will not be fully liquid but we will still have a natural heating for the fuel since the tanks are not going to be all that great for isolation thermally acting to expand causing the pressure to rise inside the tanks we are attempting to work through.
.
Even if the tanks are made the same for the moon as they are for mars we will still have different heating rates for the same fuel.

Something to think about is a small cryogenic fuels amount used for the pressurization rather than the total amount which would normally be used in a tank similar to how helium is used for kerosene use. Just add the heat source and watch the pressure build...

Maybe in the end there will be little to no mass saving and only just energy but its worth looking at since will in time building ships on mars.

The tank will work, it would seem for on orbit use for ion just fine.

tahanson43206 · 2020-06-15 17:24:50

for SpaceNut re #40

Your post here reminds me of the suggestion by RGClark, to think about the reaction of liquid Nitrogen interacting with water at room temperature in a YouTube demonstration he showed us.

While the demonstration was dramatic, the underlying physics may be applicable to the present discussion.

In recent days, I recall one of us posting a note to the effect that there is a 1:600 ratio between the volume of liquefied gas and unliquefied gas.

As I understood the physics demonstration RGClark showed us, the water in the plastic bottle provided thermal energy to the liquid Nitrogen, thus encouraging the Nitrogen to return (rapidly in the example) to a gaseous state, which led to significant expansion and thus to thrust as material was driven out the nozzle.

In the case of this topic, and the one in the Lunar topic, the issue at hand (as made clear by GW Johnson earlier in this topic) is that any gas that might be considered for this application would need to be driven into the engine(s) at a pressure greater than the combustion chamber pressure where ignition and combustion are under way.

In a ** real ** rocket, fed by liquid fuel, a mechanical compressor delivers the fuel and oxidizer to the combustion chamber at the required pressure.

There are examples of pressure fed liquid fueled rockets. It is my understanding that helium is commonly chosen as the gas to provide pressure under those circumstances.

In the case of the hypothetical gaseous supply rocket we are discussing here, and picking up on your post #40, I can imagine that liquid oxygen could be introduced into a tank full of gaseous oxygen, with the result that heat would transfer from the gaseous oxygen and thus liberate the molecules of liquid oxygen to become part of the gaseous community.

The effect of this physical change would be to increase pressure in the chamber.

There are numerous unknowns to be addressed in this scenario.

An overall concern is how to design a fuel/oxidizer storage tank able to sustain the high pressures needed to force gas into the combustion chamber.

I am under the impression that a fabric tank does not currently exist with the required strength.

In addition, there would be computation needed (a candidate would be CFD software (Computational Fluid Dynamics)) to determine how much liquid oxygen should be fed into a container of gaseous oxygen to produce the needed pressure for the duration of a flight from the surface of Mars to orbit.

Edit#1: As a reminder to any forum reader who may be seeing this discussion for the first time ... the original proposal at the beginning of this topic was offered to launch discussion and not with any serious expectation of success. It has subsequently been shown by GW Johnson that the basic proposition is highly unlikely to work in practice. The present inquiry continues in an effort to tie up as many loose ends as possible.

The most recent development in the sequence is the introduction (by SpaceNut) of the possibility of creating pressure in a container of non-liquefied gas, by introducing a measured quantity of liquefied gas, for the purpose of generating pressure through physical rather than chemical transformation.

(th)

Last edited by tahanson43206 (2020-06-15 17:33:10)

SpaceNut · 2020-06-15 19:27:29

That was post 30 for the gas at room temperature fully expanded for 1 atm.

Still researching testing for the inflatables and materials used .

https://en.wikipedia.org/wiki/Inflatable_space_habitat
https://en.wikipedia.org/wiki/Bigelow_E … ity_Module

https://ntrs.nasa.gov/archive/nasa/casi … 017032.pdf
Creep Burst Testing of a Woven Inflatable Module

Four webbing variants have been studied in two strengths and two materials: 6,000 lbf and 12,500 lbf rated Kevlar and Vectran.

Vectran (a liquid crystalline polymer) is a high-performance multifilament yarn spun from liquid crystal polymer. The fiber is five times stronger than steel and offers a unique combination of outstanding properties.

Page 8 gives the stress equations for the pressure vessel structure.

tahanson43206 · 2020-06-15 21:10:21

This topic is generally about launching a spacecraft from Mars.

It is specifically about trying to use unliquefied fuel and oxidizer, but that seems unlikely to be feasible. Never-the-less, because the pursuit of the idea to the last possible gasp seems compelling, the topic keeps going.

This article showed up when I was looking for examples of rocket design for departure from Mars. The article is from National Geographic. It was published in 2015 so there may well have been some changes in thinking since then. Still, the article is written for a general audience, with enough technical detail to hopefully interest the NewMars forum readers.

https://www.nationalgeographic.com/news … urn-space/

A detail is the prediction that thrust from the launch vehicle would last for 7 minutes. This is NOT MAVEN. The 2015 article is about return of a crew.

Another detail is that methane is planned to be shipped with the vehicle, while oxygen is to be made on Mars. In 2015, the risks of not finding hydrogen were considered too great. This is an interesting variation on Dr. Zubrin's original idea which (as I recall it) was to ship Hydrogen to Mars to make methane.

(th)

Last edited by tahanson43206 (2020-06-15 21:12:06)

GW Johnson · 2020-06-16 12:40:43

Tahanson43206:

Be careful about weight statements. The inert mass and payload mass are two separate categories. I sort everything into those two plus propellant. The sum is ignition mass Wig = Winert + Wpayload + Wpropellant, and Wig - Wpropellant is the dry-tank burnout mass, which is also Winert + Wpayload. For a first stage booster, "payload" is the loaded second stage and payload it carries. The booster inert is a separate item, as is its propellant load.

Once you think you have a representative Isp to play with, then gc Isp = Vex for the rocket equation. Use 9.80667 = gc for metric units, 32.174 = gc for US customary. You need to knock down the handbook values a tad to model nozzle efficiencies. And, if the source lists theoretical potential Isp from a thermochem code, you need to knock it down further for c* efficiency, and for turbopump bleed-and-dump effects.

There are kinematic delta-vees for just about anything. You must factor them up as mass ratio-effective delta-vees before you put them into the rocket equation. Out in vacuum and microgravity, those factors are just 1. But for first stage boost here on Earth, I use 5% each for gravity and drag losses (for clean vertical-launch designs), resulting in a factor of 1.1 upon the delta-vee required of the first stage. The second stage is essentially flying in vacuum, and perpendicular to the gravity vector, so a factor of 1 is OK for that. For rocket landings, I look at the velocity to be killed, and factor that up by 1.5 to cover hover and divert usages.

Once you do all that, you are avoiding the garbage-in/garbage-out problem that happens with any equation, not just the rocket equation. And if you think that's complicated, you ought to see what it takes to do accurate ballistics of solid propellant systems, where all the things the textbooks treat as constants are really variables.

GW

Last edited by GW Johnson (2020-06-16 12:42:17)

tahanson43206 · 2020-06-16 20:53:54

For GW Johnson re #44

Thank you for the caution about interpreting figures from reports, and specifically mass figures for vehicles.

***
SpaceNut brought back into view a topic about reusable landers ... I was surprised (but not surprised) to find your presentation in 2015 on the subject was almost identical to your recent posts about that subject. The change that I see in your more recent thinking is the addition of the large "mother ship" from which the landers depart and to which they return.

http://newmars.com/forums/viewtopic.php … 85#p127185

If you are at all interested in running a serious discussion (something like a graduate seminar but with some slack for out-of-practice participants) we might be able to persuade SpaceNut to support it.

The NewMars forum provides a place for creative thinking, but it has shown tolerance for the occasional realistic sequence of posts.

If you and SpaceNut like the idea, it could be organized in Interplanetary transportation at the Index level, with a title that indicates the serious nature of the work. I am hoping forum members and readers are keeping in mind the opportunity to write a proposal for submission to NASA. Work of that quality and substance would be a step up from the day to day meanderings of the forum, but it is (in my estimation) within the realm of possibility.

Recently one of the regular contributors ... it might have been Louis ... or it might have been Void ... described a scenario in which SpaceX might toss a version of Starship into LEO in a single stage burn with no fuel left. Such a vehicle might be designed for merging into a large expedition cluster of components.

The engines might be designed for deep space (with extended bells). They would be less efficient at sea level than would be ideal, but they would only be making the trip from sea level once.

When configured in the expedition cluster, they could be mounted to swivel for various purposes, such as all pointing one way for boost from LEO, or pointing away from a central shaft for rotation to provide a minimal artificial gravity during flight.

Your frequent reminders about radiation protection in case of a solar burst suggest planning for configuration of the expedition cluster to put as much mass as possible between the oncoming particles and the crew. Some thought would seem required, to accommodate that need while rotation is in progress.

You've already discussed the landers in some detail, but more detail is needed for a proposal.

(th)

SpaceNut · 2020-06-16 21:17:40

Starship cannot be used as a single stage to orbit and when it is launched with the first stage and fully loaded starship is totally empty all but for the fuels once on orbit for the landing fuels for the return back to earth surface landing.

tahanson43206 · 2020-06-17 06:26:34

For SpaceNut re #46

Can a Starship reach orbit by itself if it does not lift a second stage? Apparently there is some speculation that it can.

If the vehicle can reach orbit by itself without a second stage, then it ** is ** the payload! The vehicle can be joined with others to make an expedition cluster.

There is no need for fuel to land on Earth, because in this scenario, it will never return to Earth.

It would (in that case) become part of a much larger vehicle.

The engines can be designed for deep space use from the outset.

This means they will be somewhat less efficient during liftoff from Earth than their cousins designed for atmosphere, but they will only be used once in atmosphere. After that, they would be used to boost from LEO for a flight to Mars, for matching orbit with Mars, and for reducing dV to return to Earth.

In the Earth-Mars-Earth cycle, these vehicles could serve for many years, with modest maintenance.

Edit#1: It appears that the idea that a Starship could reach LEO without a Super Heavy may have come from Elon Musk himself. Here is a discussion that reports on the idea:

An interesting aspect of the Starship is that it will be able to fly without the Super Heavy. That will allow it to return from other planets and moons to the Earth. It will also be capable of single stage to orbit launch from the Earth without any useful payload. The capability is intended to be used for testing of entering planets atmospheres from super orbital velocities like Mars/Moon transfer velocities. According to Musk there are certain heating parameters that scale to the eighth power with regard to speed.

https://www.metaculus.com/questions/168 … fore-2021/

(th)

Last edited by tahanson43206 (2020-06-17 07:31:10)

tahanson43206 · 2020-06-17 07:40:24

SpaceNut, Calliban's topic about the possibility of making a practical rocket for launch from Mars, using non-liquefied fuel and oxidizer may have run its course.

Meanwhile, GW Johnson has been developing ideas for an expedition to Mars for at least the past five years.

There might be some interest in a new topic for the Interplanetary transportation Index level, if you can persuade GW Johnson to take it up.

GW Johnson has outlined a concept for an Earth-Mars-Earth expedition that would perform all operations without depending upon collecting resources from Mars, while at the same time allowing for collection and use of insitu resources as part of the expedition mission.

As I understand the concept, there would be a large expedition cluster of equipment that would be built in LEO and then used for repeated missions to Mars.

The landers would be designed to make multiple flights from the expedition complex to Mars and back, with astronaut/explorers/scientists on board.

A topic designed ** just ** for this concept could become a repository of knowledge and practical advice for expedition planners.

** That ** is why I am interested in the possibility that Elon Musk can place Starships in orbit without needing a booster.

A creative team of NewMars forum members may be able to imagine what an expedition cluster would look like, if it were made from empty Starships.

(th)

GW Johnson · 2020-06-17 12:51:38

I just dunno about putting Starship into low Earth orbit without a Superheavy booster. I think that is an unrealistic notion. My numbers say the mass ratio to do that is just not there, even at zero payload. And the thrust ain't there to lift off at all, unless you install all 6 engines as sea level Raptors, and even then you are still too short to lift off on thrust, relative to fully fueled but zero payload weight. I went through all that doing my 2019 and 2020 reverse-engineering estimates.

GW

tahanson43206 · 2020-06-17 15:26:58

For GW Johnson re #49

Thank you for your review of the notion of putting a Starship into orbit as its own "payload".

The traditional way of building an expedition vehicle is to replicate the building of the ISS, with as many launches as are needed to put the components in place.

With the SpaceX progress with re-usability of first stage boosters, it seems (to me at least) quite likely the components can be shipped to LEO with only modest second stage propulsion, and that can be down with vacuum qualified engines, which can ** then ** be put into long term service moving the expedition vehicle where it needs to go.

SpaceX has been throwing away perfectly good Merlin engines for quite a while now.

I'm reminded of the policy to throw away perfectly good aluminum tanks for the Space Shuttle. There are quite a few of them littering the floor of the Atlantic, because NASA decided to resist the many appeals for a long term use.

(th)

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