A kerosene-fueled X-33 as a single stage to orbit vehicle.
in [Physics]
|
Prev: JSH:Twin primes probability correlation
Next: SpaceX says Falcon 9 rocket test fire is a success
From: Pat Flannery on 2 May 2010 06:02 On 5/1/2010 7:47 AM, Androcles wrote: > > I'm looking for a numerical integration naval chart program if anyone > has access to one. > These are the conditions under which I want to estimate the required > delta-X to Antarctica: > > 1.)use a light propellant such as coal/air heated steam; steamship > propellers are > known to reduce wind losses. > > 2.)use a moderate to low buoyancy ratio, say, 1.0 and > below; low buoyancy ratios also reduce wind losses. > > 3.)launch near Plymouth to get the ca. 460 mile anthracite boost. > > 4.)only get to Australia, the place considered the southern continent, > to just launch rowing boats or make ship-to-ship transfers, not for long > term voyages. > > If anyone needs the right information to supply such a program please see > Robert Clark. Extremely funny and 100% on-target. Now, if we took all the food and water off of the Santa Maria and replaced it with a equal weight of gunpowder and a huge aft-facing cannon, we could fire the cannon several thousand times and move the ship from Spain to the new world in a matter of only a few days via direct Newtonian action-reaction propulsion via the recoil on the cannon. In fact, if we were to install hydrofoils on the ship to reduce its hydrodynamic drag, it would probably be able to hit around 80 knots on the way, especially if we were to cut hull weight by replacing the wood with carbon-graphite epoxy, and the gunpowder with nitroglycerin. For that matter, a aft-facing turbofan engine would weigh far less than the propulsive cannon and also give more steady thrust. To fuel it, I suggest methane gas extracted from the rotting corpses of fish caught on the way in a towed net that would serve to generate drag to keep its course true to westwards... I'm still working on the details of this, but give me a week or so and I'll have that thing traveling through the water so fast that by the time they realize they've reached their destination they will already be near Sumatra with dead Aztecs wrapped around the hydrofoils. Pat
From: Robert Clark on 3 May 2010 04:54 On May 1, 11:47 am, "Androcles" <Headmas...(a)Hogwarts.physics_z> wrote: > "Robert Clark" <rgregorycl...(a)yahoo.com> wrote in message >... > > > I'm looking for a numerical trajectory integration program if anyone > > has access to one. > > These are the conditions under which I want to estimate the required > > delta-V to orbit: > > > 1.)use a dense propellant such as kerosene/LOX; dense propellants are > > known to reduce gravity losses. > > > 2.)use a moderate to high liftoff thrust/weight ratio, say, 1.4 and > > above; high liftoff T/W also reduces gravity losses. > > > 3.)launch near equator to get the ca. 460 m/s tangential boost. > > > 4.)only get to 100 km, the altitude considered space, to just launch > > satellites or make orbital transfers, not for long term orbits. > > > Bob Clark > > I'm looking for a numerical integration naval chart program if anyone > has access to one. > These are the conditions under which I want to estimate the required > delta-X to Antarctica: > > 1.)use a light propellant such as coal/air heated steam; steamship > propellers are > known to reduce wind losses. > > 2.)use a moderate to low buoyancy ratio, say, 1.0 and > below; low buoyancy ratios also reduce wind losses. > > 3.)launch near Plymouth to get the ca. 460 mile anthracite boost. > > 4.)only get to Australia, the place considered the southern continent, > to just launch rowing boats or make ship-to-ship transfers, not for long > term voyages. > > If anyone needs the right information to supply such a program please see > Robert Clark. I agree this is counter intuitive. And the reason behind it is even more counter intuitive: it's because the gravity loss is reduced by using a lower Isp propellant. Since dense fuels have a lower Isp than hydrogen, their gravity loss will be reduced. Here's one explanation of this effect: Hydrogen delta-V (Henry Spencer; Mitchell Burnside Clapp). http://yarchive.net/space/rocket/fuels/hydrogen_deltav.html Here's a simplified case that makes it believable. The gravity loss occurs when the vehicle is traveling vertically to achieve the necessary altitude for orbit. This is done with orbital rockets by getting to some particular vertical velocity component so that the rocket's momentum will allow it to coast in the vertical direction to reach that altitude. The rest of the trip after the vertical portion is used to make the horizontal thrust needed to get to the necessary tangential velocity for orbit. So in this simplified case I'll just look at the case where your rocket only needs to get to the needed vertical speed for the altitude for orbit, say 1,400 m/s, not to achieve orbital velocity. Suppose now you have two vehicles one hydrogen-fueled the other kerosene or other dense propellant fueled. Let them both have the same liftoff thrust/ weight ratio, say, 1.4. Note that because the dense propellant has a lower Isp its mass ratio will be higher than the hydrogen to get to the same velocity, so it will have a higher propellant load. Then near the end of the trip when most of the propellant is burned off, note that the dense propellant vehicle will have a higher thrust/ weight ratio than the hydrogen one because its thrust needed to be several times higher compared to its dry mass because of the higher initial propellant load. This turns out to be the case through out the trip after the initial liftoff: the acceleration will be greater for the dense propellant vehicle. Therefore its gravity loss will be reduced because the time of the vertical trip is reduced. Here's another way of seeing the needed burn time will be smaller for the dense propellant case. Recall that the thrust of a rocket is (thrust) = (propellant flow rate)x(exhaust velocity). We know there will be a greater amount of propellant for the dense case compared to the hydrogen case. So the thrust will have to be proportionally larger as well, and so the propellant flow rate will also need to be greater. Now since the flow rate for the dense propellant will also be higher does that mean they will both burn up all their fuel in the same length of time? The key point is they won't. The dense propellant vehicle will burn up all its fuel in a shorter period of time (we're still looking at the simplified case of only looking at a vertical trip to 1,400 m/s.) The reason for this is because of that exhaust velocity term in the equation for thrust. If the propellant load is some multiple times that of the hydrogen case the thrust will have to be similarly multiply times higher. But the increased propellant flow rate that gives that multiple times greater thrust will have to be higher than the multiple of the amount of greater propellant because that exhaust velocity term is smaller. For instance if the dense fuel thrust has to be, say, twice that of the hydrogen case, then the propellant flow rate will have to be more than twice as great to make up for the lower exhaust velocity term. This means the length of time for the dense case to burn up its fuel will be shorter, resulting in a reduced gravity loss. The actual scenario for an orbital flight where there is also a horizontal thrust portion is only slightly more complicated but it also leads to the conclusion the gravity drag is reduced for the dense propellant case. If you want to see an equation that expresses the idea the acceleration is greater for the dense propellant case, even with the same initial T/W ratio, see equation 4 on page 14 here: A flexible reusable space transportation system. by Dr. Steven Pietrobon Journal of the British Interplanetary Society. vol. 53, pp. 276-288, May/June 2000 http://www.sworld.com.au/steven/pub/nsto.pdf Bob Clark
From: Androcles on 3 May 2010 06:31 "Robert Clark" <rgregoryclark(a)yahoo.com> wrote in message news:c4afdc78-b39e-46b4-bffd-0433eaaa296e(a)b18g2000yqb.googlegroups.com... On May 1, 11:47 am, "Androcles" <Headmas...(a)Hogwarts.physics_z> wrote: > "Robert Clark" <rgregorycl...(a)yahoo.com> wrote in message >... > > > I'm looking for a numerical trajectory integration program if anyone > > has access to one. > > These are the conditions under which I want to estimate the required > > delta-V to orbit: > > > 1.)use a dense propellant such as kerosene/LOX; dense propellants are > > known to reduce gravity losses. > > > 2.)use a moderate to high liftoff thrust/weight ratio, say, 1.4 and > > above; high liftoff T/W also reduces gravity losses. > > > 3.)launch near equator to get the ca. 460 m/s tangential boost. > > > 4.)only get to 100 km, the altitude considered space, to just launch > > satellites or make orbital transfers, not for long term orbits. > > > Bob Clark > > I'm looking for a numerical integration naval chart program if anyone > has access to one. > These are the conditions under which I want to estimate the required > delta-X to Antarctica: > > 1.)use a light propellant such as coal/air heated steam; steamship > propellers are > known to reduce wind losses. > > 2.)use a moderate to low buoyancy ratio, say, 1.0 and > below; low buoyancy ratios also reduce wind losses. > > 3.)launch near Plymouth to get the ca. 460 mile anthracite boost. > > 4.)only get to Australia, the place considered the southern continent, > to just launch rowing boats or make ship-to-ship transfers, not for long > term voyages. > > If anyone needs the right information to supply such a program please see > Robert Clark. I agree this is counter intuitive. And the reason behind it is even more counter intuitive: it's because the gravity loss is reduced by using a lower Isp propellant. Since dense fuels have a lower Isp than hydrogen, their gravity loss will be reduced. Here's one explanation of this effect: Hydrogen delta-V (Henry Spencer; Mitchell Burnside Clapp). http://yarchive.net/space/rocket/fuels/hydrogen_deltav.html Here's a simplified case that makes it believable. The gravity loss occurs when the vehicle is traveling vertically to achieve the necessary altitude for orbit. This is done with orbital rockets by getting to some particular vertical velocity component so that the rocket's momentum will allow it to coast in the vertical direction to reach that altitude. The rest of the trip after the vertical portion is used to make the horizontal thrust needed to get to the necessary tangential velocity for orbit. So in this simplified case I'll just look at the case where your rocket only needs to get to the needed vertical speed for the altitude for orbit, say 1,400 m/s, not to achieve orbital velocity. Suppose now you have two vehicles one hydrogen-fueled the other kerosene or other dense propellant fueled. Let them both have the same liftoff thrust/ weight ratio, say, 1.4. Note that because the dense propellant has a lower Isp its mass ratio will be higher than the hydrogen to get to the same velocity, so it will have a higher propellant load. Then near the end of the trip when most of the propellant is burned off, note that the dense propellant vehicle will have a higher thrust/ weight ratio than the hydrogen one because its thrust needed to be several times higher compared to its dry mass because of the higher initial propellant load. This turns out to be the case through out the trip after the initial liftoff: the acceleration will be greater for the dense propellant vehicle. Therefore its gravity loss will be reduced because the time of the vertical trip is reduced. Here's another way of seeing the needed burn time will be smaller for the dense propellant case. Recall that the thrust of a rocket is (thrust) = (propellant flow rate)x(exhaust velocity). We know there will be a greater amount of propellant for the dense case compared to the hydrogen case. So the thrust will have to be proportionally larger as well, and so the propellant flow rate will also need to be greater. Now since the flow rate for the dense propellant will also be higher does that mean they will both burn up all their fuel in the same length of time? The key point is they won't. The dense propellant vehicle will burn up all its fuel in a shorter period of time (we're still looking at the simplified case of only looking at a vertical trip to 1,400 m/s.) The reason for this is because of that exhaust velocity term in the equation for thrust. If the propellant load is some multiple times that of the hydrogen case the thrust will have to be similarly multiply times higher. But the increased propellant flow rate that gives that multiple times greater thrust will have to be higher than the multiple of the amount of greater propellant because that exhaust velocity term is smaller. For instance if the dense fuel thrust has to be, say, twice that of the hydrogen case, then the propellant flow rate will have to be more than twice as great to make up for the lower exhaust velocity term. This means the length of time for the dense case to burn up its fuel will be shorter, resulting in a reduced gravity loss. The actual scenario for an orbital flight where there is also a horizontal thrust portion is only slightly more complicated but it also leads to the conclusion the gravity drag is reduced for the dense propellant case. If you want to see an equation that expresses the idea the acceleration is greater for the dense propellant case, even with the same initial T/W ratio, see equation 4 on page 14 here: A flexible reusable space transportation system. by Dr. Steven Pietrobon Journal of the British Interplanetary Society. vol. 53, pp. 276-288, May/June 2000 http://www.sworld.com.au/steven/pub/nsto.pdf Bob Clark =========================================== No matter which way you slice it, lifting a mass against gravity will give it potential energy to fall again. Putting it in orbit means giving it kinetic energy as well. During re-entry both the KE and PE are lost as heat, which is why all vehicles designed for re-entry have heat shields. So at the end of the excursion, with energy being a conserved quantity, all of the energy in the fuel has been converted to heat which is then radiated into space. Your problem isn't to find a trajectory, it is to find how to put mass into orbit at minimum cost. That is, find joules per dollar for the type of fuels and oxidants and the technology available, whether it is kerosene-oxygen, hydrogen-oxygen, old rubber tyres, nitromethane for "Top Fuel" drag racing or nitro-glycerine! Airliners burn two tons of fuel on take-off and throttle back to cruise. They have the advantage of using atmospheric oxygen up to 50,000 feet (10 miles) and normally cruise economically at 30,000 feet where the air is thin enough to reduce drag and thick enough to breathe (they are pressurised to 8000 feet, compressors force air into the cabin). If you've ever flown you'll know your ears pop. But to reach the ISS an altitude of 200 miles is needed, so an oxidant needs to be lifted along with the fuel for the other 190 miles. http://en.wikipedia.org/wiki/Space_Shuttle_external_tank "A Space Shuttle External Tank (ET) is the component of the Space Shuttle launch vehicle that contains the liquid hydrogen fuel and liquid oxygen oxidizer." Nowhere in your analysis above have you considered that. The laws of physics and chemistry cannot be defeated, it is only economics that you can meddle with, joules per dollar. I have to tell you honestly that you'll never convince an aeronautical engineer of your quick fix, you've left out far too much reality.
From: Robert Clark on 3 May 2010 11:17 According to SpaceX they have a vacuum version of the Merlin engine with a vacuum Isp of 342 s: New Merlin Vacuum engine demonstrates highest efficiency for an American hydrocarbon rocket engine. "McGregor, TX. March 10, 2009 Space Exploration Technologies (SpaceX) successfully conducted a full mission duration firing of its new Merlin Vacuum engine on March 7, at SpaceX's Test Facility in McGregor, Texas. The engine fired for a full six minutes, consuming 100,000 pounds of liquid oxygen and rocket grade kerosene propellant. "The new engine, which powers the upper stage of SpaceX's Falcon 9 launch vehicle, demonstrated a vacuum specific impulse of 342 seconds the highest efficiency ever for an American hydrocarbon rocket engine. Thrust was measured at approximately 92,500 lb of force in vacuum conditions and the engine remained thermally stable over the entire run." http://www.spacex.com/press.php?page=20090310 To have this efficiency while being optimized for vacuum operation a rocket engine has to have a longer and heavier nozzle. This greatly reduces its efficiency at sea level. However, a version of this vacuum Merlin that uses instead a aerospike nozzle might be able to maintain its high efficiency at sea level as well. This might also make the engine lighter but I've seen conflicting reports on whether using an aerospike nozzle really makes an engine lighter. Though not yet used on actual orbital launch vehicles, the aerospike has been flight tested: Aerospike. http://en.wikipedia.org/wiki/Aerospike_engine#Performance And according to these refs the aerospike gets high efficiency from sea level to high altitude conditions: Aerospikes (Henry Spencer; Ben Muniz; Jeff Greason) http://yarchive.net/space/rocket/aerospike.html Nozzle Design. by R.A. O'Leary and J. E. Beck Threshold Magazine, Spring 1992 http://www.pwrengineering.com/articles/nozzledesign.htm They do mention there is some diminution in efficiency in the velocity range of Mach 1 to Mach 3, but overall its performance is well above that of the usual bell nozzle. I had suggested that a Falcon 1 first stage that used a high performance Russian engine in stead of the Merlin 1C could make a SSTO. But of course SpaceX itself would be disinclined to use a different engine from the one they have spent millions of dollars developing. But if an aerospike Merlin could indeed maintain correspondingly high performance at sea level, then SpaceX could have an SSTO from an aerospike Merlin Falcon 1 first stage as well. Bob Clark
From: Androcles on 3 May 2010 11:21
"Robert Clark" <rgregoryclark(a)yahoo.com> wrote in message news:0d4a435d-a1c0-4452-a138-93c432165188(a)k29g2000yqh.googlegroups.com... According to SpaceX they have a vacuum version of the Merlin engine with a vacuum Isp of 342 s: New Merlin Vacuum engine demonstrates highest efficiency for an American hydrocarbon rocket engine. "McGregor, TX. � March 10, 2009 � Space Exploration Technologies (SpaceX) successfully conducted a full mission duration firing of its new Merlin Vacuum engine on March 7, at SpaceX's Test Facility in McGregor, Texas. The engine fired for a full six minutes, consuming 100,000 pounds of liquid oxygen and rocket grade kerosene propellant. "The new engine, which powers the upper stage of SpaceX's Falcon 9 launch vehicle, demonstrated a vacuum specific impulse of 342 seconds � the highest efficiency ever for an American hydrocarbon rocket engine. Thrust was measured at approximately 92,500 lb of force in vacuum conditions and the engine remained thermally stable over the entire run." http://www.spacex.com/press.php?page=20090310 To have this efficiency while being optimized for vacuum operation a rocket engine has to have a longer and heavier nozzle. This greatly reduces its efficiency at sea level. However, a version of this vacuum Merlin that uses instead a aerospike nozzle might be able to maintain its high efficiency at sea level as well. This might also make the engine lighter but I've seen conflicting reports on whether using an aerospike nozzle really makes an engine lighter. Though not yet used on actual orbital launch vehicles, the aerospike has been flight tested: Aerospike. http://en.wikipedia.org/wiki/Aerospike_engine#Performance And according to these refs the aerospike gets high efficiency from sea level to high altitude conditions: Aerospikes (Henry Spencer; Ben Muniz; Jeff Greason) http://yarchive.net/space/rocket/aerospike.html Nozzle Design. by R.A. O'Leary and J. E. Beck Threshold Magazine, Spring 1992 http://www.pwrengineering.com/articles/nozzledesign.htm They do mention there is some diminution in efficiency in the velocity range of Mach 1 to Mach 3, but overall its performance is well above that of the usual bell nozzle. I had suggested that a Falcon 1 first stage that used a high performance Russian engine in stead of the Merlin 1C could make a SSTO. But of course SpaceX itself would be disinclined to use a different engine from the one they have spent millions of dollars developing. But if an aerospike Merlin could indeed maintain correspondingly high performance at sea level, then SpaceX could have an SSTO from an aerospike Merlin Falcon 1 first stage as well. Bob Clark ============================================= Dodging the issue and spouting jargon doesn't make you smart. |