A kerosene-fueled X-33 as a single stage to orbit vehicle.
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From: Androcles on 6 May 2010 12:30 "Robert Clark" <rgregoryclark(a)yahoo.com> wrote in message news:b8ca9ed7-c432-4d9a-9ffc-685ad98ea60f(a)o8g2000yqo.googlegroups.com... On May 6, 7:29 am, "Androcles" <Headmas...(a)Hogwarts.physics_z> wrote: > "Robert Clark" <rgregorycl...(a)yahoo.com> wrote in message > > news:486970b5-ba74-434f-bcf9-0faf305a01f3(a)o8g2000yqo.googlegroups.com... > On May 4, 3:40 pm, "Androcles" <Headmas...(a)Hogwarts.physics_z> wrote: > > > > > "Robert Clark" <rgregorycl...(a)yahoo.com> wrote in message > > ... > > > What is the point of your argument? > > ==================================== > > Let's see, there are a few. > > 0) NASA, the Russian Federal Space Agency, the China National Space > > Administration and the European Space Agency I shall refer to as the > > agencies. > > 1) The agencies have already heard of kerosene and opted for hydrogen > > instead. > > 2) The agencies know they can combine oxygen and hydrogen to produce > > H2O. > > 3) The agencies know other exhaust gasses will be toxic, in particular > > carbon monoxide. > > 4) The agencies know they will have to lift heavy liquid oxygen (atomic > > mass > > 16), a point Clark ignores. > > 5) Clark thinks combining a mixture of carbon chains that typically > > contain > > between 6 and 16 carbon atoms (atomic mass 12) per molecule, aka > > kerosene, > > with oxygen, will yield more thrust than pure water because Clark has a > > Ph.D. in chemistry and the agencies are all idiots. > > > Now, do tell... what is the point of YOUR argument? > > I suggested a possible reason > ============================================== > You've not understood the question. I didn't ask for a reason, I asked > what the POINT of your ARGUMENT is (its purpose). > You can pontificate until you are blue in the face, but if you want to > argue then you'll answer the points I've made or I shall simply ignore > yours as you are ignoring mine, and that makes your argument futile. I'm arguing two points: one that SSTO is technically feasible, ============================================= I won't challenge that, the coal-fired steam locomotive is technically feasible; enthusiastic hobbyists still maintain them. However, diesel-electric replaced them. ============================================= and two that it is financially beneficial. ============================================= Now we have a problem, you failed to address the cost of producing and lifting the oxygen the other 190 miles. Your idea is dead on arrival. <handwaving snipped> You can pontificate until you are blue in the face, but if you want to argue then you'll answer the points I've made or I shall simply ignore yours as you are ignoring mine, and that makes your argument futile.
From: Robert Clark on 23 May 2010 07:06 I wanted to get some info on minimal trajectories to orbit. I found this page after a web search: Atmospheric Flight of a Launch Vehicle. http://astro.u-strasbg.fr/~koppen/launcher/launcher.html It has a JAVA applet for calculating the ascent to orbit for some known rockets. It requires you to tweak the ascent angles, but fills in much of the information for you. Problem is I haven't been able to get it to work right. For the given rockets, we know they were able to achieve orbit but every time I try it I get that the orbit was not achieved. Perhaps someone here can figure out how to work it. Bob Clark On Apr 28, 10:56 pm, Robert Clark <rgregorycl...(a)yahoo.com> wrote: > Another question asked in email and on different forums about this > is, if the Russians already had these high performancekerosene > engines that make SSTO possible why aren't they doing it? > I wondered about that too. I thought their new Angara rocket should > be SSTO capable since it will be using these high performancekerosene > engines. But then I checked the specifications of the Angara rocket on > the SpaceLaunchReport.com site: > > Space Launch Report: New Launchers - Angara.http://www.spacelaunchreport.com/angara.html > > I found that the Angara mass ratios are significantly worse than for > the SpaceX Falcon launchers. In fact, in general the Russian launchers > are not as well mass optimized as the American launchers. This > probably is a big part of the reason the Russians have had this great > drive to increase the performance of theirkeroseneengines - out of > necessity. > Then to get SSTO you use the best features of both the American and > Russian designs combined into one. > > Bob Clark > > On Mar 14, 9:24 pm, Robert Clark <rgregorycl...(a)yahoo.com> wrote: > > > The SpaceLaunchReport.com site operated by Ed Kyle provides the > > specifications of some launch vehicles. Here's the page for the Falcon > > 1: > > > Space Launch Report: SpaceX Falcon Data Sheet.http://www.spacelaunchreport.com/falcon.html > > > Quite interesting is that the total mass and dry mass values for the > > Falcon 1 first stage with Merlin 1C engine give a mass ratio of about > > 20 to 1. This is notable because a 20 to 1 mass ratio is the value > > usually given for akerosene-fueled vehicle to be SSTO. However, this > > is for the engine having high vacuum Isp ca. 350 s. The Merlin 1C with > > a vacuum Isp of 304 s probably wouldn't work. > > However, there are some high performance Russiankeroseneengines that > > could work. Some possibilities: > > > Engine Model: RD-120M.http://www.astronautix.com/engines/rd120.htm#RD-120M > > > RD-0124.http://www.astronautix.com/engines/rd0124.htm > > > Engine Model: RD-0234-HC.http://www.astronautix.com/engines/rd0234.htm > > > However, I don't know if this third one was actually built, being a > > modification of another engine that burned aerozine. > > > Some other possibilities can be found on the Astronautix site: > > > Lox/Kerosene.http://www.astronautix.com/props/loxosene.htm > > > And on this list of Russian rocket engines: > > > Russian/Ukrainian space-rocket and missile liquid-propellant engines.http://www.b14643.de/Spacerockets_1/Diverse/Russian%20engines/engines... > > > The problem is the engine has to have good Isp as well as a good T/W > > ratio for this SSTO application. There are some engines listed that > > even have a vacuum Isp above 360 s. However, these generally are the > > small engines used for example as reaction control thrusters in orbit > > and usually have poor T/W ratios. > > For the required delta-V I'll use the fact that a dense propellant > > vehicle may only require a delta-V of 8,900 m/s, compared to a > > hydrogen-fueled vehicle which may require in the range of 9,100 to > > 9,200 m/s. The reason for this is explained here: > > > Hydrogen delta-V.http://yarchive.net/space/rocket/fuels/hydrogen_deltav..html > > > Then when you add on the fact that launching near the equator gives > > you 462 m/s for free from the Earth's rotation, we can take the > > required delta-V that has to be supplied by thekerosene-fueled > > vehicle as 8,500 m/s. > > I'll focus on the RD-0124 because of its high Isp, 359 s vacuum and > > 331 s sea level. On the "Russian/Ukrainian space-rocket and missile > > liquid-propellant engines" page its sea level thrust is given as > > 253,200 N, 25,840 kgf. However, the Falcon 1 first stage weighs 28,553 > > kg. So we'll need two of them. Each weighs 480 kg, so two would be 960 > > kg. This is 300 kg more than the single Merlin 1C. So the dry mass of > > the Falcon 1 first stage is raised to 1,751 kg. There is a RD-0124M > > listed on the Astronautix page that only weighs 360 kg, but its sea > > level Isp and thrust are not given, so we'll use the RD-0124 until > > further info on the RD-0124M is available. > > Taking the midpoint value of the Isp as 345 s we get a delta-V of > > 345*9.8ln(1 + 27102/1751) = 9,474 m/s (!) Note also the achieved delta- > > V would actually be higher than this because the trajectory averaged > > Isp is closer to the vacuum value since the rocket spends most of the > > time at altitude. > > This calculation did not include the nose cone fairing weight of 136 > > kg. However, the dry mass for the first stage probably includes the > > interstage weight, which is not listed, since this remains behind with > > the first stage when the second stage fires. Note then that the > > interstage would be removed for the SSTO application. From looking at > > the images of the Falcon 1, the size of the cylindrical interstage in > > comparison to the conical nose cone fairing suggests the interstage > > should weigh more. So I'll keep the dry mass as 1,751 kg. > > Now considering that we only need 8,500 m/s delta-V we can add 636 kg > > of payload. But this is even higher than the payload capacity of the > > two stage Falcon 1! > > We saw that the thrust value of the RD-0124 is not much smaller than > > the gross weight of the Falcon 1 first stage. So we can get a vehicle > > capable of being lifted by a single RD-0124 by reducing the propellant > > somewhat, say by 25%. This reduces the dry weight now since one > > RD-0124 weighs less than a Merlin 1C and the tank mass would also be > > reduced 25%. Using an analogous calculation as before, the payload > > capacity of this SSTO would be in the range of 500 kg. > > We can perform a similar analysis on the Falcon 1e first stage that > > uses the upgraded Merlin 1C+ engine. Assuming the T/W ratio of the > > Merlin 1C+ is the same as that of the Merlin 1C, the mass of the two > > of the RD-124's would now be only 100 kg more than the Merlin 1C+. > > The dry mass and total mass numbers on the SpaceLaunchReport page for > > the Falcon 1e are estimated. But accepting these values we would be > > able to get a payload in the range of 1,800 kg. This is again higher > > than the payload capacity of the original two stage Falcon 1e. In fact > > it could place into orbit the 1-man Mercury capsule. > > The launch cost of the Falcon 1, Falcon 1e is only about $8 million - > > $9 million. So we could have the first stage for that amount or > > perhaps less since we don't need the engines which make up the bulk of > > the cost. How much could we buy the Russian engines for? This article > > says the much higher thrust RD-180 cost $10 million: > > > From Russia, With 1 Million Pounds of Thrust. > > Why the workhorse RD-180 may be the future of US rocketry. > > Issue 9.12 | Dec 2001 > > "This engine cost $10 million and produces almost 1 million pounds of > > thrust. You can't do that with an American-made engine."http://www.wired.com/wired/archive/9.12/rd-180.html > > > This report gives the price of the also much higher thrust AJ26-60, > > derived from the Russian NK-43, as $4 milliion: > > > A Study of Air Launch Methods for RLVs. > > Marti Sarigul-Klijn, Ph.D. and Nesrin Sarigul-Klijn, Ph.D. > > AIAA 2001-4619 > > "The main engine is currently proposed as the 3,260 > > lb. RP-LOX Aerojet AJ26-60, which is the former > > Russian NK-43 engine. Thrust to weight of 122 to > > 1 compares to the Space Shuttle Main Engines > > (SSME) 67 to 1 and specific impulse (Isp = 348.3 > > seconds vacuum) is 50 to 60 seconds better than > > the Atlas II, Delta II, or Delta III RP-LOX engines. > > A total of 831 engines have been tested for > > 194,000 seconds. These engines are available for > > $4 million each, which is about 10% the cost of a > > SSME."http://mae.ucdavis.edu/faculty/sarigul/aiaa2001-4619.pdf > > > Then the much lower thrust RD-0124 could quite likely be purchased > > for less than $4 million. So the single RD-0124 powered SSTO could be > > purchased for less than $12 million. > > > Even though the mathematics says it should be possible, and has been > > for decades, it is still commonly believed that SSTO performance with > > chemical propulsion is not possible even among experts in the space > > industry: > > > Space Tourism is a Hoax > > By Fredrick Engstrom and Heinz Pfeffer > > 11/16/09 09:02 AM ET > > "In 1903, the Russian scientist Konstantin Tsiolkovsky established the > > so-called rocket equation, which calculates the initial mass of a > > rocket needed to put a certain payload into orbit, given that the > > orbital speed is fixed at 28,000 kilometers per hour, and that the > > maximum speed of the gas exhausted from the rocket that propels it > > forward is also fixed. > > "You quickly find that the structure and the tanks needed to contain > > the fuel are so heavy that you will never be able to orbit a > > significant payload with a single-stage rocket. Thus, it is necessary > > to use several rocket stages that are dumped on the way up to get any > > net mass, i.e. payload, into orbit. > > "Let us look at the most successful rocket on the market the > > European Ariane 5. Its start weight is 750 tons, of which 650 tons are > > fuel, 80 tons are structure and around 20 tons are left for low Earth > > orbit payload. > > "You can have a different number of stages, and you can look for minor > > improvements, but you can never get around the fact that you need big > > machines that are staged to reach orbital speed. Not much has happened > > in propulsion in a fundamental sense since Wernher von Brauns Saturn > > rocket. And there is nothing on the horizon, if you discount > > controlling gravity or some exotic technology like that. In any case, > > it is not for tomorrow."http://www.spacenews.com/commentaries/091116-space-tourism-hoax.html > > > The Cold Equations Of Spaceflight. > > by Jeffrey F. Bell > > Honolulu HI (SPX) Sep 09, 2005 > > "Why isn't Mike Griffin pulling out the blueprints for X-30/NASP, DC-X/ > > Delta Clipper, or X-33/VentureStar? Billions of dollars were spent on > > these programs before they were cancelled. Why aren't we using all > > that research to design a cheap, reusable, Single-Stage-To-Orbit > > vehicle that operates just like > > ... > > read more »
From: Robert Clark on 13 Jun 2010 11:36 It is known that you can reduce the fuel requirements to orbit by using a lifting trajectory. For instance it was explored for use on a reusable orbital craft in the 80's: Fire in the sky: the Air Launched Sortie Vehicle of the early 1980s (part 1) by Dwayne Day Monday, February 22, 2010 http://www.thespacereview.com/article/1569/1 In the first post in this thread I suggested this fact be used for a lifting-body SSTO. Here is the argument I made: ------------------------------------------------------------------ "Then the gravity losses could be further reduced by flying a lifting trajectory, which would also increase the payload capability by a small percentage. "The trajectory I'll use to illustrate this will first be straight- line at an angle up to some high altitude that still allows aerodynamic lift to operate. At the end of this portion the vehicle will have some horizontal and vertical component to its velocity. "We'll have the vertical component be sufficient to allow the vehicle to reach 100 km, altitude. The usual way to estimate this vertical velocity is by using the relation between kinetic energy and potential energy. It gives the speed of v = sqrt(2gh) to reach an altitude of h meters. At 100,000 m, v is 1,400 m/s. "Now to have orbital velocity you need 7,800 m/s tangential, i.e., horizontal velocity. If you were able to fly at a straight-line at a constant angle to reach 7,800 m/s horizontal velocity and 1,400 m/s vertical velocity and such that the air drag was kept at the usual low 100 to 150 m/s then you would only need sqrt(7800^2 + 1400^2) = 7,925 m/s additional delta-v to reach orbit. Then the total delta-v to orbit might only be in the 8,100 m/s range. Note this is significantly less than the 9,200 m/s delta-v typically needed for orbit, including gravity and air drag. "The problem is with usual rocket propulsion to orbit not using lift the thrust vector has to be more or less along the center-line of the rocket otherwise the rocket would tumble. You can gimbal the engines only for a short time to change the rocket's attitude but the engines have to be then re-directed along the center line. However, the center line has to be more or less pointing into the airstream, i.e., pointing in the same direction as the velocity vector, to reduce aerodynamic stress and drag on the vehicle. But the rocket thrust having to counter act gravity means a large portion of the thrust has to be in the vertical component which means the thrust vector has to be nearly vertical at least for the early part of the trip when the gross mass is high. Then the thrust vector couldn't be along the center line of a nearly horizontally traveling rocket at least during the early part of the trip." ------------------------------------------------------------------ The key fact I want to focus on is that sqrt(7800^2 + 1400^2) = 7,925 m/s number. Actually if you add on the ca. 460 m/s velocity you get for free from the Earth's rotation you might be able to reduce this to sqrt(7400^2 + 1400^2) = 7,531 m/s. So I what I want to investigate is if it is possible for a rocket that does not have wings or lifting surfaces to travel at such a straight-line trajectory at an angle from lift-off so that the achieved velocity will be in this range (but keeping in mind this might not be the same as the equivalent "delta-V" that actually has to be put out by the engines.) But I have question: if you angle the rocket launch from the start with the thrust vector along the center line with the trajectory angle such that the vertical component of the thrust equals the rocket weight could you have the rocket travel at a straight-line all the way to orbit? I'm inclined to say no because the gravity is operating at the center of gravity of the rocket not at the tail where the thrust is operating. This would certainly work if you had a point particle, but I'm not sure if it would work when your body has some linear extent. This method for traveling at a straight-line at an angle for some or all of the trip would make my calculation easier. However, I'll show in a following post there is another way to do it even if this first method doesn't work. The second method though would require some modification to the usual design of rockets and is more computationally complicated. Bob Clark
From: hallerb on 13 Jun 2010 18:33 On Jun 13, 11:36�am, Robert Clark <rgregorycl...(a)yahoo.com> wrote: > �It is known that you can reduce the fuel requirements to orbit by > using a lifting trajectory. For instance it was explored for use on a > reusable orbital craft in the 80's: > > Fire in the sky: the Air Launched Sortie Vehicle of the early 1980s > (part 1) > by Dwayne Day > Monday, February 22, 2010http://www.thespacereview.com/article/1569/1 > > In the first post in this thread I suggested this fact be used for a > lifting-body SSTO. Here is the argument I made: > > ------------------------------------------------------------------ > "Then the gravity losses could be further reduced by flying a lifting > trajectory, which would also increase the payload capability by a > small percentage. > �"The trajectory I'll use to illustrate this will first be straight- > line at an angle up to some high altitude that still allows > aerodynamic lift to operate. At the end of this portion the vehicle > will have some horizontal and vertical component to its velocity. > �"We'll have the vertical component be sufficient to allow the vehicle > to reach 100 km, altitude. The usual way to estimate this vertical > velocity is by using the relation between kinetic energy and potential > energy. It gives the speed of v = sqrt(2gh) to reach an altitude of h > meters. At 100,000 m, v is 1,400 m/s. > �"Now to have orbital velocity you need 7,800 m/s tangential, i.e.., > horizontal velocity. If you were able to fly at a straight-line at a > constant angle to reach 7,800 m/s horizontal velocity and 1,400 m/s > vertical velocity and such that the air drag was kept at the usual low > 100 to 150 m/s then you would only need sqrt(7800^2 + 1400^2) = 7,925 > m/s additional delta-v to reach orbit. Then the total delta-v to orbit > might only be in the 8,100 m/s range. Note this is significantly less > than the 9,200 m/s delta-v typically needed for orbit, including > gravity and air drag. > �"The problem is with usual rocket propulsion to orbit not using lift > the thrust vector has to be more or less along the center-line of the > rocket otherwise the rocket would tumble. You can gimbal the engines > only for a short time to change the rocket's attitude but the engines > have to be then re-directed along the center line. However, the center > line has to be more or less pointing into the airstream, i.e., > pointing in the same direction as the velocity vector, to reduce > aerodynamic stress and drag on the vehicle. But the rocket thrust > having to counter act gravity means a large portion of the thrust has > to be in the vertical component which means the thrust vector has to > be nearly vertical at least for the early part of the trip when the > gross mass is high. Then the thrust vector couldn't be along the > center line of a nearly horizontally traveling rocket at least during > the early part of the trip." > ------------------------------------------------------------------ > > �The key fact I want to focus on is that sqrt(7800^2 + 1400^2) = 7,925 > m/s number. Actually if you add on the ca. 460 m/s velocity you get > for free from the Earth's rotation you might be able to reduce this to > sqrt(7400^2 + 1400^2) = 7,531 m/s. So I what I want to investigate is > if it is possible for a rocket that does not have wings or lifting > surfaces to travel at such a straight-line trajectory at an angle from > lift-off so that the achieved velocity will be in this range (but > keeping in mind this might not be the same as the equivalent "delta-V" > that actually has to be put out by the engines.) > But I have question: if you angle the rocket launch from the start > with the thrust vector along the center line with the trajectory angle > such that the vertical component of the thrust equals the rocket > weight could you have the rocket travel at a straight-line all the way > to orbit? I'm inclined to say no because the gravity is operating at > the center of gravity of the rocket not at the tail where the thrust > is operating. This would certainly work if you had a point particle, > but I'm not sure if it would work when your body has some linear > extent. > This method for traveling at a straight-line at an angle for some or > all of the trip would make my calculation easier. However, I'll show > in a following post there is another way to do it even if this first > method doesn't work. > The second method though would require some modification to the usual > design of rockets and is more computationally complicated. > > �Bob Clark the math is over my head, but if the vehicle can refuel along the way, with air to air refueling done in the military all the time the payload can be a lot more
From: Androcles on 13 Jun 2010 19:03
<hallerb(a)aol.com> wrote in message news:e183261a-8f6a-4b59-b411-f4416c22b558(a)20g2000vbi.googlegroups.com... On Jun 13, 11:36�am, Robert Clark <rgregorycl...(a)yahoo.com> wrote: > �It is known that you can reduce the fuel requirements to orbit by > using a lifting trajectory. For instance it was explored for use on a > reusable orbital craft in the 80's: > > Fire in the sky: the Air Launched Sortie Vehicle of the early 1980s > (part 1) > by Dwayne Day > Monday, February 22, 2010http://www.thespacereview.com/article/1569/1 > > In the first post in this thread I suggested this fact be used for a > lifting-body SSTO. Here is the argument I made: > > ------------------------------------------------------------------ > "Then the gravity losses could be further reduced by flying a lifting > trajectory, which would also increase the payload capability by a > small percentage. > �"The trajectory I'll use to illustrate this will first be straight- > line at an angle up to some high altitude that still allows > aerodynamic lift to operate. At the end of this portion the vehicle > will have some horizontal and vertical component to its velocity. > �"We'll have the vertical component be sufficient to allow the vehicle > to reach 100 km, altitude. The usual way to estimate this vertical > velocity is by using the relation between kinetic energy and potential > energy. It gives the speed of v = sqrt(2gh) to reach an altitude of h > meters. At 100,000 m, v is 1,400 m/s. > �"Now to have orbital velocity you need 7,800 m/s tangential, i.e., > horizontal velocity. If you were able to fly at a straight-line at a > constant angle to reach 7,800 m/s horizontal velocity and 1,400 m/s > vertical velocity and such that the air drag was kept at the usual low > 100 to 150 m/s then you would only need sqrt(7800^2 + 1400^2) = 7,925 > m/s additional delta-v to reach orbit. Then the total delta-v to orbit > might only be in the 8,100 m/s range. Note this is significantly less > than the 9,200 m/s delta-v typically needed for orbit, including > gravity and air drag. > �"The problem is with usual rocket propulsion to orbit not using lift > the thrust vector has to be more or less along the center-line of the > rocket otherwise the rocket would tumble. You can gimbal the engines > only for a short time to change the rocket's attitude but the engines > have to be then re-directed along the center line. However, the center > line has to be more or less pointing into the airstream, i.e., > pointing in the same direction as the velocity vector, to reduce > aerodynamic stress and drag on the vehicle. But the rocket thrust > having to counter act gravity means a large portion of the thrust has > to be in the vertical component which means the thrust vector has to > be nearly vertical at least for the early part of the trip when the > gross mass is high. Then the thrust vector couldn't be along the > center line of a nearly horizontally traveling rocket at least during > the early part of the trip." > ------------------------------------------------------------------ > > �The key fact I want to focus on is that sqrt(7800^2 + 1400^2) = 7,925 > m/s number. Actually if you add on the ca. 460 m/s velocity you get > for free from the Earth's rotation you might be able to reduce this to > sqrt(7400^2 + 1400^2) = 7,531 m/s. So I what I want to investigate is > if it is possible for a rocket that does not have wings or lifting > surfaces to travel at such a straight-line trajectory at an angle from > lift-off so that the achieved velocity will be in this range (but > keeping in mind this might not be the same as the equivalent "delta-V" > that actually has to be put out by the engines.) > But I have question: if you angle the rocket launch from the start > with the thrust vector along the center line with the trajectory angle > such that the vertical component of the thrust equals the rocket > weight could you have the rocket travel at a straight-line all the way > to orbit? I'm inclined to say no because the gravity is operating at > the center of gravity of the rocket not at the tail where the thrust > is operating. This would certainly work if you had a point particle, > but I'm not sure if it would work when your body has some linear > extent. > This method for traveling at a straight-line at an angle for some or > all of the trip would make my calculation easier. However, I'll show > in a following post there is another way to do it even if this first > method doesn't work. > The second method though would require some modification to the usual > design of rockets and is more computationally complicated. > > �Bob Clark the math is over my head, but if the vehicle can refuel along the way, with air to air refueling done in the military all the time the payload can be a lot more ==================================================== Altitude of International Space Station = 200 miles. Altitude of Mount Everest = 6 miles. Speed of International Space Station in orbit = 17,000 mph Speed of Military Tanker with fuel = 500 mph Let's refuel at 30,000 feet, 500 mph. Only 194 miles and 16,500 mph to go. What is your idea of a "lot" more? Are we there yet? Are we there yet? Are we there yet? |