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This topic is a branch from the one started by Calliban in the Interplanetary transportation Index level topic.
That one may be found as: Compressed gas balloon rocket for Mars launch
GW Johnson has looked closely at the Mars situation, and found it highly unlikely an un-refrigerated fuel/oxidizer vehicle could work on Mars.
However, the Moon is quite a different environment.
There is no pesky atmosphere to impose drag, and the gravitational field is 1/6th of Earth's, or 1/2 of that of Mars (approximately)
The theme of the original topic was the idea that balloons/fabric containers could hold fuel and oxidizer for a launch vehicle.
While the concept may not be practical for Mars, if should be quite practical for the Moon.
It has been pointed out that relying on pressure of gas containers to feed a rocket engine is likely to be unproductive, so mechanical pumps would appear to be required to provide the needed pressure of reactants into the engine(s).
I'm hoping this new topic will generate some interest, and perhaps even a realistic, practical vision for an achievable launch system.
Reminder ... the vehicle needs to be able to deliver a payload to orbit (or escape) and then return safely to the surface of the Moon.
The energy savings to be achieved is elimination of refrigeration of fuel and oxidizer gases, and perhaps some expenses related to operating cryogenic equipment.
(th)
Last edited by tahanson43206 (2020-06-14 07:44:24)
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Essentially what you’re asking for is a pressure-fed rocket to launch from the Moon’s surface. Because the orbital speed for Moon is so much smaller than for Earth, about 1.8 km/s for the Moon compared to about 7.8 km/s for Earth, this isn’t particularly hard.
Instead of balloons though use metallic or carbon fiber tanks commonly used for rocket propellant tanks with the propellant held under pressure.
You may have been conflating this question with another question: suppose you had a gas, could be even non-flammable such as air, or nitrogen, that is simply compressed under high pressure. Could this be used for a rocket that could reach lunar orbit?
My guess is probably yes. The Isp would be low though so it wouldn’t be very efficient. See table here: https://en.m.wikipedia.org/wiki/Cold_ga … ropellants
But the real problem is the low density of the gas, as indicated in the table on that wiki page, would result in a low propellant mass ratio.
A possible answer to that is to store the “propellant” as a liquid, then allow the expansion of the propellant into a gas to provide the thrust, such as here:
Liquid Nitrogen Rocket with Steve Spangler on DIY Sci S01E06.
https://www.youtube.com/watch?v=yFNLq7qnmMQ
Bob Clark
Last edited by RGClark (2020-06-15 08:05:07)
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I see that the main issue for the engines is flow rate for the liquid versus gaseous. As a liquid the pressure is very low to get a given volume to the engine while the pressure for a gas must be high to obtain a volume in a give time unit.
For mars I was using the movement to cause a pressure increase with a cable assisting collapsing mechanism to keep pressure going.
Sure heating will produce pressure and giving the fuel shade prior to filling makes the fuel condense. So if the fuel is compressed when cool the fuel as dense gas as it can be before going to a liquid state and that makes the fabric tank which will have considerable less mass a bigger buy for a larger payload.
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For RGClark re #2
Thanks for giving this new topic a boost! (pun not intended ... it just sort of happened)
And ** thanks ** for the link to the Steve Spangler video. The physics of that is intriguing, and I expect I'll be looking for explanations of how the two liquids contributed to the vigorous reaction. The water contributed heat energy of course, and it added mass to the effluent.
As far as who is looking for what is concerned, it would probably make sense for the forum reader to go back to Calliban's original topic, to see what he had in mind.
What I have in mind is very specific, and (to the best of my knowledge) it is not conflated with anything else.
The purpose of this topic is to establish as rigorously as possible, the practicality of feeding gaseous fuel and oxidizer into a lunar launcher, which would carry out the 'normal' chemical reaction that would have occurred if the components had been liquefied.
The suggestion of making the compression tank out of metallic or carbon fiber is interesting and (it seems to me) well worth pursuing.
However, ultimately, the practicality of the idea comes down to energy expense ...
It is going to cost X amount of energy (I'm assuming solar in this discussion) to make the fuel and oxidizer from lunar water.
It will cost Y amount of energy to liquefy those components for use in a traditional liquid fuel rocket.
It will cost Z(a) amount of energy to make the vehicle for the gaseous propulsion system.
It will cost Z(b) amount of energy to make the vehicle for the liquid propulsion system.
One of those two systems is going to deliver the customer's cargo to lunar orbit more cheaply than the other.
At this point, I have no idea which of the two systems is going to win that competition.
For further clarification ... GW Johnson has given the matter some study (in the original topic) and concluded that a pressure fed design is not practical for Mars.
I am making the assumption that a pressure fed design is not practical for the Moon as well.
Thus, this topic assumes mechanical assistance to deliver fuel and oxidizer at the pressures needed for efficient combustion in the same rocket engine that would be used for the liquid version.
Mechanical assistance has been in use for the purpose of driving jet engine combustion for a number of decades now. That experience should be transferable to the lunar launch with gaseous components situation.
Energy will be required to drive the mechanical compression systems (whatever they are). With advances in electrical power storage, it might prove most efficient to operate the pumps (whatever they are) using electrical power.
(th)
Last edited by tahanson43206 (2020-06-14 10:19:16)
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The problem is less about liquid versus gas storage, and more about low chamber pressure in the rocket engine for a low-pressure-fed system. F = Pc At CF, where CF is a dimensionless number somewhere between 1 and 2, lower for lower Pc. If your storage feed pressure is millibars, your chamber pressure is millibars. If bars, Pc is bars. But you DO NOT get KN of thrust per sq.m of At from a practical small At until Pc is Megabars. The bigger the At, the bigger and heavier the engine hardware. You want something on the order of 0.01 to 0.1 sq.m worth of At.
So if you do a pressure-fed system, your storage pressure must be megabars! Period! Or else you MUST accept very large and heavy rocket engines to get very little thrust. Even off the moon you are wanting to lift tons, which means you must have at least 10's of KN of thrust. You cannot put megabars inside an inflatable. That requires a hard shell tank. There are no materials to support a lightweight megabar-level inflatable.
What that says is "go ahead and use pumps of some kind" for launch purposes. They do not have to be turbopumps. But you do have to power them some way. Just bear in mind that is it far, far, far easier to pressurize liquids than it is to compress gases, by at least an order of magnitude.
And THAT is why the very best rockets (high thrust and high Isp) have been fueled with liquids for almost a century since Robert Goddard. Goddard used pressure-fed; nearly everything since has been pumped. (Solids get the high thrust, but not the high Isp. Useful during first stage boost. Not so efficient later in the ascent, although both Scout and the Minuteman family used solids to orbit or near-orbit, for other reasons of practicality.)
What your gaseous propellant/pressure-fed notion is far better suited to, is delivering the reaction mass to any of the electric propulsion schemes. Those do not need high chamber pressure, in point of fact, they use almost no density in their chambers at all. Unfortunately, all of those are unsuited to launch, as all of us already well know. They generate milli-N thrust from tens of kg (up to tons) of equipment, precisely because the stream densities are virtually vacuum.
So my skepticism has good reasons.
GW
Last edited by GW Johnson (2020-06-14 10:45:54)
GW Johnson
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"There is nothing as expensive as a dead crew, especially one dead from a bad management decision"
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The shuttle tank LOX Operation pressure: 34.7–36.7 psi (239–253 kPa) (absolute) and the LH2 tank Operation pressure: 32–34 psi (220–230 kPa) (absolute).
Each propellant tank has a vent and relief valve at its forward end. This dual-function valve can be opened by ground support equipment for the vent function during prelaunch and can open during flight when the ullage (empty space) pressure of the liquid hydrogen tank reaches 38 psi (260 kPa) or the ullage pressure of the liquid oxygen tank reaches 25 psi (170 kPa).
RS-25 burns cryogenic liquid hydrogen and liquid oxygen propellants, with each engine producing 1,859 kN (418,000 lbf) of thrust at liftoff. The engine produces a specific impulse (Isp) of 452 seconds (4.43 km/s) in a vacuum, or 366 seconds (3.59 km/s) at sea level, has a mass of approximately 3.5 tonnes (7,700 pounds), and is capable of throttling between 67% and 109% of its rated power level in one-percent increments. Components of the RS-25 operate at temperatures ranging from −253 to 3,300 °C (−400 to 6,000 °F).
The low-pressure oxidizer turbopump (LPOTP) is an axial-flow pump which operates at approximately 5,150 rpm driven by a six-stage turbine powered by high-pressure liquid oxygen from the high-pressure oxidizer turbopump (HPOTP). It boosts the liquid oxygen's pressure from 0.7 to 2.9 MPa (100 to 420 psi).
Which means flow rate of the liquid from the tank is at low pressure and the pumps are boosting pressure to spray it into the chamber.
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For SpaceNut re #6
Thank you for your detailed report on the operation of the Space Shuttle main engines, and ** particularly ** the contribution of the LPOTP, which itself is (I gather) driven by the HPOTP.
GW Johnson has advised that the pumps that would be required to achieve the same performance (ie, output of 2.9 MPa) with gaseous oxygen as input would have to operate at 10 times the performance of the pumps which were moving liquid into the RS-25.
I was curious to know what jet engines do with gaseous oxidizer, so asked Google, which came up with this source:
https://cs.stanford.edu/people/eroberts … s/how.html
The image above shows how a jet engine would be situated in a modern military aircraft. In the basic jet engine, air enters the front intake and is compressed (we will see how later). Then the air is forced into combustion chambers where fuel is sprayed into it, and the mixture of air and fuel is ignited. Gases that form expand rapidly and are exhausted through the rear of the combustion chambers. These gases exert equal force in all directions, providing forward thrust as they escape to the rear. As the gases leave the engine, they pass through a fan-like set of blades (turbine), which rotates a shaft called the turbine shaft. This shaft, in turn, rotates the compressor, thereby bringing in a fresh supply of air through the intake. Below is an animation of an isolated jet engine, which illustrates the process of air inflow, compression, combustion, air outflow and shaft rotation just described.
In his post recently in this topic, GW Johnson pointed out that a pump (of whatever kind) that would pull gaseous oxygen from a tank would have to be powered. The quote above shows that on Earth, in an aircraft, power to drive the pump is obtained from the combustion of fuel with the oxidizer.
It would not surprise me to learn that a jet engine cannot compare in thrust to the performance of a rocket engine, so it will not surprise me to learn that a turbojet engine could not provide the thrust that would be needed to launch a vehicle from the Moon.
Never-the-less, the path we are on, as initiated by Calliban in a moment of inspiration, would appear to require that we examine the potential of a turbojet engine to drive the acceleration of a lunar vehicle designed to hold un-refrigerated gaseous fuel and oxidizer.
In the case of the Moon, the candidate chemicals would seem most likely to be Hydrogen and Oxygen, derived from (hypothetical though likely) deposits of water on the Moon.
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The ideal gas law says that if you collect one liter of oxygen at three, seven or 10 times standard atmospheric pressure, you'll get three, seven or 10 times as many oxygen atoms in that liter as you would at standard pressure.
Daytime on one side of the moon lasts about 13 and a half days, followed by 13 and a half nights of darkness. Over the course of a full lunar day and night, the temperature on the Moon can vary wildly, from around +200 to -200 degrees Celsius (+392 to -328 degrees Fahrenheit), so it’s natural to wonder how lunar astronauts survived this huge temperature variation.
https://en.wikipedia.org/wiki/Liquid_oxygen
At this latter temperature, LOX vaporizes into the gaseous state. Gaseous oxygen will turn into liquid at atmospheric pressure by cooling to a temperature below -297 F. By increasing the pressure, gaseous oxygen can be liquified at higher temperatures, up to its critical temperature, - 182 F ( -119 C).
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.
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To understand the limitations of the turbine engine technology, you need to understand a couple of things about gas turbine engines.
There are three critical items:
(1) Higher compression pressure is higher frontal thrust density for the engine, all else being equal. Higher frontal thrust density is also associated with higher engine thrust/weight.
(2) About the largest turbine-driven compressor pressure ratio I ever heard of is 18:1. At about 1.1:1 per axial stage, that's something like 13-14 stages in series, or more, because it is unlikely you can achieve the same 1.1:1 ratio in each stage. Centrifugal is limited to 3-4:1 per stage, but if you use more than 1 stage, the machinery gets very large and heavy, sharply reducing frontal thrust density and decreasing thrust/weight.
(3) Turbine survival absolutely requires the oncoming hot gas temperature not to exceed something in the 2000 F class. To do that when stoichiometric flame temperatures are about 4000 F with air at 14.5:1 by mass, means that you drastically lean the mixture down to around 30:1. If you are using oxygen instead of air, the stoich ratio is nearer 3.4 with temperature near 6000 F. That means you run around 10 or 12:1 to reduce the temperature to acceptable levels. There are severe problems getting it to burn that far from stoichiometry. It gets even worse if you go fuel-rich instead of oxygen-rich.
Now, for 30 psia (2 bar) tank pressures, a 20:1 pressure rise puts you near 600 psia. That's about 41 bar. Modern rocket engines operate in the 1000-4000 psia chamber pressure range, which is 69-276 bar. Your frontal thrust density achievable is more-or-less proportional to your chamber pressure. It also affects thrust/weight, but not linearly.
To reach, say, 3000 psia chamber pressure, your gaseous tank pressure needs to be in the 150 psia (10 bar) class. An elastomer container capable of holding that much pressure looks more like a road tire than a balloon. And weighs just as much.
GW
Last edited by GW Johnson (2020-06-14 17:59:57)
GW Johnson
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"There is nothing as expensive as a dead crew, especially one dead from a bad management decision"
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For GW Johnson re #9
Calliban's venture is leading to some remarkable learning opportunities!
SearchTerm:TurbineEngine
SearchTerm:JetEngine
GW Johnson's summary: http://newmars.com/forums/viewtopic.php … 73#p169073
Key ideas for the Lunar Launcher application:
A turbine engine which derives power for the compressor by passing combustion gases over the turbine is at risk of damage to the power take-off turbine due to the elevated temperature of the exhaust gases.
A related issue (which has not been discussed in this topic until now) is the advantage that liquid fuel offers the designer of a rocket engine for cooling.
A hypothetical gaseous input would be much less efficient at collecting heat energy from the bell of the reaction engine than is liquid used as a coolant.
The use of electricity to power the compressor turbine remains on the table. However, the power to supply such a turbine would require mass to be invested in batteries and control electronics.
Rocket Lab uses electric pumps:
The pumps (one for the fuel and one for the oxidizer) in electric-pump feed engines are driven by an electric motor. The Rutherford engine uses dual brushless DC electric motors and a lithium polymer battery. It is claimed that this improves efficiency from the 50% of a typical gas-generator cycle to 95%.
Cycle: Electric pump-fed engine
Pumps: 2 electric pumpsRocket Lab Rutherford - Wikipedia
en.wikipedia.org › wiki › Rocket_Lab_Rutherford
However, it is immediately to be noted that GW Johnson has already addressed the difference between pumping a liquid and attempting to scavenge a gas in a non-liquid state. He estimated the difference as an order of magnitude, which would imply (to me at least) that the pumps and electric supply would have to be 10 times (at least) more robust than the ones in service with Rocket Lab.
The gaseous state rocket idea appears to be gasping for oxidizer at this point.
(th)
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