Talk:Rocket propellant/Archive 1

Archive 1

Circular Reference

The "Solid rocket fuel" links to an article which links right back here to this one.

Ertyseidel (talk) 17:33, 23 March 2009 (UTC)

Performance analysis questionable

The current article contains several discussions relating to the effects on rocket performance of such things as mixture ratios, propellant density, propellant fractions, combustion temperature, molecular weight, etc. These are legitimate subjects to address but I am bothered that amateur analysis is being used to draw conclusions based on single design factors that do not determine system performance to the extent being asserted. The tendency to take statements out of context and develop them to arrive at overstated or questionable conclusions is resulting in a lot of equally questionable text being introduced into the article. I encourage the persons doing this research to spend more time on the details and comparing their conclusions with materials on the same subject from other sources.

These performance discussions contained at least half a dozen erroneous conclusions that stood out as I read them. Example: Both solid propellant stages and hydrogen-fueled stages were stated to inherently suffer from low propellant fractions (high hardware weight fractions) compared to stages using other liquid propellants. These obvious conclusions were based on the facts that 1)solids require a high-pressure vessel which surrounds all of the propellant (as opposed to pump fed liquid stages which can store their propellant at low pressure) and 2)hydrogen requires extra large tanks because of its low density. But in actual fact, both solid stages and some hydrogen-fueled stages have propellant mass fractions as good as any pump-fed liquid propellant upper stages using LOX/kersoene or storable liquid propellants. How can that be? Solid propellants have much higher density than liquid propellants and require no pumps or injectors. Some high performance Star space motors have propellant fractions as high as 94.6% while Castor first stage solid motors exceed 92.2%. And the hydrogen-fueled Centaur upper stage of Atlas V has a propellant fraction of 91.6% using lightweight, pressure-stabilized balloon tanks. LOX/kerosene and storable liquid propellant upper stages do not have higher mass ratios. The assertion of low relative mass ratios for solids and hydrogen-fueled stages are not givens as stated. Solids may still have lower performance due to lower specific impulse but not necessarily for the reason given in the article. Rather than interpreting the implications of assumptions, there might be more to be gained from studying real systems. "All else being equal" does not really apply between alternate systems using different propellants. Magneticlifeform (talk) 10:13, 26 June 2011 (UTC)'

In general comparisons need to be like-for-like using similar technology as far as possible, and also comparing stages that are operated in similar situations.

The Centaur system for example uses balloon tanks, but those techniques can be applied to dense propellants as well, and give better significantly mass ratios than you can get with hydrogen. -Rememberway (talk) 12:31, 27 June 2011 (UTC)

In addition you seem to confuse atmospheric stages, which generally require high pressure combustion, with upper stages, which can be run at low pressure and often significantly lower thrust and accelerations. That makes a very significant difference with respect mass fractions that can be achieved. -Rememberway (talk) 12:31, 27 June 2011 (UTC)

• • • Your points are well taken. The use of so-called balloon tanks did in fact originate with the kerosene fueled MX-774 test vehicle and was then applied to Convair's kerosene-fueled Atlas ICBM. The balloon-tanked Centaur was then developed as an upper stage for Atlas, yielding the Atlas Centaur launch vehicle. To my knowledge the current Centaur is the only balloon-tanked stage in use, with the Atlas 5 first stage now having conventional aluminum isogrid tanks. Kerosene or storable liquid upper stages could use balloon tanks but do not. The thin-walled pressure stabilized tanks are difficult and expensive to manufacture. The need to always keep them pressurized also requires special handling. The cost is only justified if it is really needed. When hydrogen fuel is used, the weight saving is quite significant. I am fully aware of the different requirements for sea level operation of first stage engines as opposed to vacuum operation of lower thrust upper stages. I have designed both. It is some of the conclusions drawn from the like-to-like comparisons that bothered me, or rather the assumption that like-to-like actually applies to hardware systems that are rarely of the "all else being equal variety". The balloon tanks were merely a case in point since they allowed Centaur to have a mass ratio comparable to (actually existing, as opposed to hypothetical) stages using higher density propellants. At any rate, we seem to be on the same page in most respects. Best wishes and keep on writing.Magneticlifeform (talk) 21:10, 27 June 2011 (UTC)'

Isp

Can somebody explain what is the "Isp"?

There is no linking to other articles, it just appeared out of nowhere. Thanks — Preceding unsigned comment added by Sedimin (talk • contribs) 10:33, 14 February 2007 (UTC)

See specific impulse (commonly abbreviated I_sp) Sdsds 04:54, 17 February 2007 (UTC)

Specific impulse is the average exhaust velocity divided by the acceleration of gravity, taken to be 9.80665 meters/sec/sec (approximately 32.174 ft/sec/sec). From Newton's Second Law, the thrust of a rocket engine equals the mass flow rate times the average exhaust velocity. Hence, average exhaust velocity is a measure of rocket engine performance. Typical average exhaust velocities range from about 800 m/sec for black powder rockets to 4565 m/sec for the highest performing oxygen/hydrogen rocket engines. (Electric thrusters have much higher exhaust velocities.) In English speaking countries it is customary to describe flow rates in pounds/second, but pounds are units of force, not of mass. The unit of mass is the slug, which actually weighs 32.174 pounds. To convert pounds/second to slugs/second, one must divide by the acceleration of gravity, 32.174 ft/sec/sec. That is how the custom arose of describing rocket engine performance in terms of "specific impulse" which is the average exhaust velocity divided by "g" (9.80665 m/s/s or 32.174 ft/s/s). To convert specific impulse back to average exhaust velocity, simply multiply by the acceleration of gravity. Magneticlifeform (talk) 23:09, 11 October 2010 (UTC)Magneticlifeform

• Sorry but the above explanation is not even interesting as history (dividing by g to get convert imperial unit to mass). It's worse than wrong. First, I don't see anything good in introducing arcane stuff like slugs, when you're right in the middle of explaining something else. Moreover, I think the explanation above is physically wrong, inasmuch as it tries to "explain" the need to divide by g to get units of force (thrust) from "mass flow" * "exhaust velocity". The latter is m/t * v = mv/t, but that does indeed give units of thrust = force. HOWEVER, specific impulse is not thrust and it never has force units. So, all this is more complicated, and bad physics AS WELL.

  Dimensionally, specific impulse (Isp) is presented in two different basic dimensional forms: as momentum/mass = velocity, AND as momentum/weight = mv/mg = v/g = time. This is independent of whether SI or Imperial units are used. The "explanation" of needing to divide by a weight, works to explain the "time unit" specific impulse (Isp in sec), but not for the "velocity unit" specific impulse (Isp in ft/sec or m/sec or whatever velocity units you like). SBHarris 18:31, 25 January 2011 (UTC)

• • Dr. Harris, the term "specific impulse" originated as follows. Rocket engine thrust is measured in pounds. During engine testing, engineers measure the total impulse produced by integrating thrust (in pounds) over time (seconds) to obtain the total impulse produced in pound-seconds. The quantity of interest to the engineer is the "specific impulse" or the total impulse that is generated divided by the weight of propellant consumed to generate that impulse. Hence specific impulse has the units of pound-seconds of impulse per pound of propellant. (pound-seconds/pound)=seconds. So, propellant specific impulse is given in seconds, which, as you pointed out, is not a unit of thrust (i.e. force).

• • Thrust, from Newton's 2nd Law, is equal to the mass flowrate times the average exhaust velocity (a vector quantity ignoring off-setting, off-axis velocity components within that flow). Hence thrust divided by mass flowrate is equal to the average exhaust velocity. Average exhaust velocity is the true measure of the performance of a propellant in terms of total impulse divide by the mass (not weight) of propellant consumed. That's much simpler and does not involve the injection and removal of gravity into the equations. But those of us using English units of measure where "pound" is used as both force and mass are not spared the annoyance. We are stuck with "specific impulse" instead of exhaust velocity as a measure of performance. The conversion factor between the two is g, the standard acceleration of gravity. Yes, it is confusing and detracts from the simplicity of thrust=(mass flowrate x exhaust velocity). I wish "specific impulse" would go away, but since we all must deal with it, it is useful to know that specific impulse is just (exhaust velocity)/g with the odd unit of seconds. It's not bad physics but rather a messed up system of units.Magneticlifeform (talk) 01:36, 27 June 2011 (UTC)

Yes, see the excellent article on specific impulse. "Specific" means, "per amount of something" and in this case, the something is fuel. But amount can be in mass or weight. Specific impulse is not always given in seconds (what you get with weight), but sometimes is indeed given in units of velocity (what you get if you use fuel mass), and the latter is the specific impulse (given in seconds) multiplied by g. In that case, it is indeed the "effective exhaust velocity" or what you call the "average exhaust velocity." It's easy to see which units of "specific impulse" are being used, by looking at the units. If specific impulse is given as a velocity, then you know automatically that it's the "effective exhaust velocity" type of specific impulse. That's a perfectly good type, and you WILL see it. If given in seconds, it's the other type. The conversion factor between the two is g.

Again, it is NOT true that specific impulse is always given in seconds. Sometimes it is expressed as a velocity, and then it's not the end of the world. You simply know that you need to divide by g, if you want the Isp in seconds (if you like that better). SBHarris 02:22, 27 June 2011 (UTC)

• • • • The "specific impulse" article you recommend is indeed excellent and I also recommend it to anyone interested in understanding the subject. Curiously the article does not support your own statements, so I recommend you read it again yourself.Magneticlifeform (talk) 23:38, 20 February 2012 (UTC)

(separate comment:) Under the liquid fuels section: The first LOX should have "liquid oxygen" written after it in parentheses (I'm pretty sure that's what it means- you might want to double check). —Preceding unsigned comment added by 97.117.33.102 (talk) 21:58, 26 July 2010 (UTC)

Monopropellant is not Hybrid

The "Delchev" motor reference should be removed as it is a one component Monopropellant. The MnO2 catalyst to ignite the mixture is not a solid fuel component but an igniter for non-hypergolic fuel system, similar to a glow plug for Nitromethane-Methanol or Methanol-Ammonium Nitrate solutions. A reference to [Messerschmitt_Me_163] would be nice because the high energy of decomposition of Hydrogen Peroxide is the primary propellant, not attitude control (see [list_of_stoffs]). Also see [water_injection] as a method of increasing usable energy obtained from engine systems (e.g. used on takeoff by early model jet aircraft).

In hydrogen-oxygen liquid fuel rockets, theoretical impulse is not reached because the dissociation temp of water is ~2000 degrees and the combustion temp inside the engine is over 2500 degrees, a major part of the Hydrogen burning externally. Shjacks45 (talk) 14:51, 17 January 2012 (UTC)

SpicyItalianMeatball uh, no, the article is terrible and I can fix it. It will take some time because it is so bad. I know physics and propulsion and orbital mechanics and also have a friend at NASA who thinks I should trash the whole thing and start over. Actually...that would be easier. We are going to get this thing right. Bwtranch (talk) 06:12, 14 October 2012 (UTC)