nuclear thermal rocket do better than chemical rocket

[nec208]

<p>&nbsp;There is more energy production capacity from a nuclear fission reactor than from chemical reactions.&nbsp; What I said was that this is not the reason nuclear thermal rockets give you better efficiency (Isp) than chemical rockets.&nbsp; The reason it gives you better efficiency is because you can use a low molecular weight propellant (hydrogen).</p><p>------------------------------------------------------------------------------------</p><p>How could there be more energy capacity in nuclear fission than chemical.They use chemical it just they don't burn it.It is heated and splitting of hydrogen atoms.</p><p>&nbsp;</p><p><br /><br />&nbsp;</p> <div class="Discussion_UserSignature"> </div>

 

[KosmicHero]

Replying to:<BR/><DIV CLASS='Discussion_PostQuote'>&nbsp;There is more energy production capacity from a nuclear fission reactor than from chemical reactions.&nbsp; What I said was that this is not the reason nuclear thermal rockets give you better efficiency (Isp) than chemical rockets.&nbsp; The reason it gives you better efficiency is because you can use a low molecular weight propellant (hydrogen).------------------------------------------------------------------------------------How could there be more energy capacity in nuclear fission than chemical.They use chemical it just they don't burn it.It is heated and splitting of hydrogen atoms.&nbsp;&nbsp; <br />Posted by nec208</DIV><br /><br />I dont know what you're trying to say.&nbsp; Much more heat (energy) can be produced from a nuclear reactor than from hydrogen and oxygen combusting.&nbsp; <div class="Discussion_UserSignature"> kosmichero.wordpress.com </div>

 

[nec208]

<p><BR/>Replying to:<BR/><DIV CLASS='Discussion_PostQuote'>I dont know what you're trying to say.&nbsp; Much more heat (energy) can be produced from a nuclear reactor than from hydrogen and oxygen combusting.&nbsp; <br />Posted by KosmicHero</DIV></p><p>Yes because when heated&nbsp;hydrogen and splitting atoms you get more energy than burning hydrogen and oxygen .</p><p>And you don't need a&nbsp;oxygen tank.</p><p>&nbsp;</p><p><br /><br />&nbsp;</p> <div class="Discussion_UserSignature"> </div>

 

[neilsox]

<p>We need to reach an altitude of about 1000 kilometers before we energize a nuclear rocket engine otherwise we put radioactive issotopes in Earth's atmosphere which will shorten the life expectancy of humans and other criters. Less than half of the humans consider even a slightly&nbsp;reduced life expectancy an acceptable sacrifice to get us to Mars or elsewhere in our solar system. If the chemical rockets to lift us though Earth atmosphere fail, we may get the reduced life expectancy even though we did not energize the nuclear rocket.</p><p>Clearly we need a constuction facility in a 1000 plus kilometer orbit to build the nuclear rockets, and we need to mine (and refine)&nbsp;the uranium on asteroids instead of on Earth. With present technology the cost would be several (assuming higher risk and working smarter than we have)&nbsp;times what we have spent on the ISS = international space station, and 2020 is an optimistic date for the first test firing of nuclear rockets off planet. Let's go for it.&nbsp;&nbsp; Neil</p>

 

[annodomini2]

<p><BR/>Replying to:<BR/><DIV CLASS='Discussion_PostQuote'>Those are excellent questions and I would be happy to explain.&nbsp; For a rocket that is acclerating a fluid through converging-diverging (C-D) nozzle, the equation for the exit velocity of the fluid is given by (I hope this works)&nbsp;&nbsp;This is the simplified equation if you assume that the pressure is expanded to ambient which is usually what nozzles are designed to do so we'll ignore that part.&nbsp; It doesnt really impact this explaination anyway.&nbsp; So in this equation the gamma is the ratio of specific heats which is affected slightly but the type of propellant that you use but its pretty small (1.6 for hydrogen, 1.4 for air).&nbsp; The R is the gas constant and the M is the molecular mass of the propellant (2 for hydrogen (actually since it will disassociate at these high temperatures it will be closer to 1, and water is 18 which is the primary product of combusting hyrdogen and oxygen which is the highest performing liquid combination).The last term is the 'stagnation' or 'chamber' temperature.&nbsp; This is how hot we can get the combustion (or in the case of the NTR the reactor).&nbsp; You are right when you say that a higher chamber temperature will give you higher Isp (specific impulse), and an NTR can be run to extremely high stagnation temperatures.&nbsp; Unfortunately the reactor (made from some very exotic materials usually tungsten/nickel alloys) cannot handle the temps the reactor can put out.&nbsp; A liquid rocket can produce around 2500 K to 3000 K.&nbsp; NTRs are run upto maybe 3500 K but after that material limitations start to become the constraint.One more thing, the connection between specific impulse and exhaust velocity is from the ideal rocket equation and the definition of Isp.&nbsp; Ue = Isp*g.&nbsp; There is a lot of confusion here, but the g you use is the g at sea level no matter where you are.&nbsp; This is because this is converting velocity and 'weight' and weight as it is thought about here is weight at sealevel. &nbsp;Let me know if this suffices or if you have any more questions.&nbsp; If the equation doesnt work I'll just make a cruder one.&nbsp; <br /> Posted by KosmicHero</DIV></p><p>&nbsp;</p><p>Just to see if I understand this correctly, you want a lower molecular weight to reduce the overall mass of the vehicle?</p><p>As the reactor is producing the same amount of energy the total energy in the equation remains constant and you end up with a higher nozzle velocity resulting in a higher ISP?</p> <div class="Discussion_UserSignature"> </div>

 

[DrRocket]

<p><BR/>Replying to:<BR/><DIV CLASS='Discussion_PostQuote'>&nbsp;Nuclear Thermal Rockets do NOT derive their added efficiency over chemical rockets from the increase in energy density of nuclear material or from significant increases in heat transfer to the working fluid (i.e. propellant).&nbsp; The advantages from NTRs are almost entirely derived from the independence in working fluid.&nbsp; &nbsp;When you&nbsp;operate a chemical rocket you require a fuel and oxidizer but when you&nbsp;operate an NTR the source of your energy (heat) is not chemically dependent on the fuel/oxidizer you choose.&nbsp; Because of this, you want to choose the lowest molecular weight propellant you can.&nbsp; This is hydrogen.&nbsp; Hydrogen is the only propellant that will make an NTR&nbsp;competitive with a chemical rocket.&nbsp; <br />Posted by KosmicHero</DIV></p><p>Absolutely correct.&nbsp; Isp varies about like sqrt(temp/mol wt).&nbsp; You want the lowest molecular weight and the highest temperature that you can get.&nbsp; However, another consideration is the inert weight that is necessary to contain and operate the rocket, since delta v = Isp ln(initial mass/final mass) (if you use Isp in seconds rather than in consistent units you need a gc factor).&nbsp; The problem with hydrogen, even liquid hydrogen, is that the density is low so you get a lot of inert weight in the packaging.&nbsp; Because of the density problem ammonia is also sometimes considered as a working fluid.</p> <div class="Discussion_UserSignature"> </div>

 

[DrRocket]

<p><BR/>Replying to:<BR/><DIV CLASS='Discussion_PostQuote'>Those are excellent questions and I would be happy to explain.&nbsp; For a rocket that is acclerating a fluid through converging-diverging (C-D) nozzle, the equation for the exit velocity of the fluid is given by (I hope this works)&nbsp;&nbsp;This is the simplified equation if you assume that the pressure is expanded to ambient which is usually what nozzles are designed to do so we'll ignore that part.....&nbsp; <br />Posted by KosmicHero</DIV></p><p>Your equation did not appear, but that is Ok.&nbsp; Your explanation is about for the idealized one-dimensional flow case.&nbsp; There are several versions of Isp, similar but not quite identical, used.&nbsp; In understanding the physics the easiest to understand is simply the velocity of the exhaust gasses.&nbsp; But when data is measured in a static test, what is determined is the thrust-time profile and the propellant weight expended.&nbsp; That requires a correction in meassured thrust for the effect of ambient pressure, and&nbsp;average Isp is then just&nbsp;the total impulse divided by the propellant mass expended.&nbsp; You can carry this further and by analysis determine the rate of propellant expulsion as a function of time and calculate an instantaneous Isp.&nbsp;Isp depends on the expansion ratio of the nozzle, and because of that sometimes one calculates Isp based on a hypothetical firing at sea level and expansion to sea level pressure.&nbsp; Moreover these calculations usually assume a chamber pressure of 1000 psi --- This is called "Isp zero-one thousand". One can also consider Isp in a vacuum with the gasses fully expanded --- vacuum Isp.&nbsp; Then there is delivered Isp, which is what you actually get.&nbsp; That is based on static test data with some analysis thrown in and includes effects like, two-dimensional effects in the nozzle, two-phase flow effects for solid propellant combustion, etc.&nbsp; The advertised accuracy for the code that is used to predict this for solid propellants is 1/2 sec., but it does not always fulfill its advertised performance.&nbsp; I have seen it miss by 3 seconds, and that is a big deal in a high performance application.&nbsp; So, don't take the theoretical calculation overly seriously, but do recognize that Isp generally varfies like the square root of temperature divided by molecular weight.&nbsp; </p> <div class="Discussion_UserSignature"> </div>