A kerosene-fueled X-33 as a single stage to orbit vehicle.
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From: Robert Clark on 21 Mar 2010 10:17 On Mar 15, 10:02 am, Me <charliexmur...(a)yahoo.com> wrote: > On Mar 14, 9:24 pm, Robert Clark <rgregorycl...(a)yahoo.com> wrote: > > > Then it is important that such a SSTO vehicle be produced even if > > first expendable to remove the psychological barrier that it can not > > be done. Once it is seen that it can be done, and in fact how easily > > and cheaply it can be done, then there it will be seen that in fact > > the production of SSTO vehicles are really no more difficult than > > those of multistage vehicles. > > Then will be opened the floodgates to reusable SSTO vehicles, and low > > cost passenger space access as commonplace as trans-oceanic air > > travel. > > More clueless BS. Clark thinks he is smarter than everyone else. [re-posted to correct typos.] No. I'm reporting what some experts in the field have said, that it is easier to produce a SSTO vehicle with dense fuels rather than with hydrogen. Some examples: Single Stage To Orbit Mass Budgets Derived From Propellant Density and Specific Impulse. John C. Whitehead, Lawrence Livermore National Laboratory. 32nd AIAA/ASME/SAE/ASEE Joint Propulsion Conference. Lake Buena Vista, FL July 1-3, 1996 Abstract "The trade between specific impulse and density is examined in view of SSTO requirements. Mass allocations for vehicle hardware are derived from these two properties, far several propellant combinations and a dual-fuel case. This comparative analysis, based on flight-proven hardware, indicates that the higher density of several alternative propellants compensates for reduced Isp, when compared with cryogenic oxygen and hydrogen. Approximately half the orbiting mass of a rocket-propelled SSTO vehicle must be allocated to propulsion hardware and residuals. Using hydrogen as the only fuel requires a slightly greater fraction of orbiting mass for propulsion, because hydrogen engines and tanks are heavier than those for denser fuels. The advantage of burning both a dense fuel and hydrogen in succession depends strongly on tripropellant engine weight. The implications of the calculations for SSTO vehicle design are discussed, especially with regard to the necessity to minimize non-tankage structure." http://www.osti.gov/bridge/servlets/purl/379977-2LwFyZ/webviewable/379977.pdf A Single Stage to Orbit Rocket with Non-Cryogenic Propellants. Clapp, Mitchell B.; Hunter, Maxwell W. AIAA, SAE, ASME, and ASEE, Joint Propulsion Conference and Exhibit, 29th, Monterey, CA, June 28-30, 1993. Abstract "Different propellant combinations for single-stage-to-orbit-rocket applications were compared to oxygen/hydrogen, including nitrogen tetroxide/hydrazine, oxygen/methane, oxygen/propane, oxygen/RP-1, solid core nuclear/hydrogen, and hydrogen peroxide/JP-5. Results show that hydrogen peroxide and JP-5, which have a specific impulse of 328 s in vacuum and a density of 1,330 kg/cu m. This high-density jet fuel offers 1.79 times the payload specific energy of oxygen and hydrogen. By catalytically decomposing the hydrogen peroxide to steam and oxygen before injection into the thrust chamber, the JP-5 can be injected as a liquid into a high-temperature gas flow. This would yield superior combustion stability and permit easy throttling of the engine by adjusting the amount of JP-5 in the mixture. It is concluded that development of modern hydrogen peroxide/JP-5 engines, combined with modern structural technology, could lead to a simple, robust, and versatile single-stage-to-orbit capability." http://www.erps.org/docs/SSTORwNCP.pdf Alternate Propellants for SSTO Launchers. Dr. Bruce Dunn Adapted from a Presentation at: Space Access 96 Phoenix, Arizona April 25 =96 27, 1996 Introduction "The most commonly proposed propellant combination for an SSTO launcher is liquid oxygen and liquid hydrogen, at a mixture ratio of approximately 6.0. There have been a number of studies of alternate fuels for SSTO launchers, but they have been limited. To date, most studies have concentrated on methane, propane and RP-1 burned with liquid oxygen to the exclusion of other oxidizers and other fuels. These studies have often, but not always shown lower vehicle dry masses for hydrocarbon propellants (for the same payload size). The lowest dry masses of all are found in dual-fuel vehicles, using dense hydrocarbons early in the flight and hydrogen late in the ascent. These vehicles however suffer from mechanical and structural complexity over their single-fuel cousins, and are unlikely to represent the least expensive way to get a defined payload to orbit." http://www.dunnspace.com/alternate_ssto_propellants.htm This is certainly a minority opinion that dense fuels are better for a SSTO than hydrogen, but it has occurred numerous times in science that the minority opinion turns out to be the correct one. The argument for why dense propellants are better for a SSTO is quite simple and can be understood by anyone familiar with the "rocket equation" that describes the relationship between the exhaust velocity and the mass of propellant for a rocket. Indeed the argument is as about as close to a mathematical proof as you can get in engineering. First two key facts have to be kept in mind: 1.) the tank mass scales by volume, *NOT* by the mass of the fluid contained. This means that the same size and *same mass* tanks can hold about 3 times as much kero/LOX as LH2/LOX. This is extremely important because the propellant tanks make up the single biggest component of the dry weight of a rocket, typically 30% to 40%, even more than that of the engines. And 2.) dense propellant engines such as kerosene ones typically have thrust/weight ratios twice as good as hydrogen ones. This is key because switching to kerosene means your fuel load and therefore gross mass will be greater. But because of the kerosene engines better T/W ratio, the increase in engine weight will be relatively small. Many people get the second of these points. It's the reason why first stages generally use kerosene or other dense propellant for example. However, the first point most people are not as familiar with. But it's the more important of the two because the increase in propellant being carried far exceeds the increase needed to overcome the lowered Isp of the dense propellants. To see why tank mass scales with volume, take a look at the equations for tank mass here: Pressure vessel. http://en.wikipedia.org/wiki/Pressure_vessel#Scaling Note it depends only on tank dimensions, internal pressure, and strength and density of the tank material. Then because the internal pressure of the tanks will be about the same for the hydrogen case as for the kerosene case, for proper operation of the turbopumps, the kerosene filled tanks will hold about 3 times more propellant at the same size and weight of the tanks. Now for the calculation that switching to kerosene can result in multiple times greater payload. The vacuum Isp for good hydrogen engines is about 450 s, and for good kerosene ones about 350 s. This means the mass ratio for a hydrogen SSTO is about 10 and for a kerosene one it's about 20. These values are higher than what you would expect based just on the vacuum Isp alone because you also have to consider gravity and air drag, and the fact that the Isp is decreased at sea level and low altitude. Now suppose we switch our hydrogen-fueled SSTO for a kerosene-one using the same sized tanks. The volume stays the same so the mass of the tanks stays the same. But the amount of propellant is now about 3 times larger. For the engines, since propellant mass makes up almost all the vehicle gross weight, the gross weight will be about 3 times larger too. So the engines will need about 3 times the thrust. For the original hydrogen-engines the thrust/weight ratio was about 50 to 1. And since the gross mass was about 10 times the dry mass for the hydrogen vehicle, this means the engine mass was about 1/5, or 20%, of the dry weight. Now switching to kerosene makes the gross weight about 3 times larger. If the kerosene engines had only a 50 to 1 T/W ratio then you would need 3 times heavier engines so they would be at 3/5 of the dry weight. But since the thrust/weight of the kerosene engines is twice that of the hydrogen ones, the engine weight is 1.5/5, 30%, of the dry weight so the vehicle dry weight is increased only by 10%, due to the heavier engines. Now since the mass ratio is 10 for the hydrogen case but 20 for the kerosene, you normally need about twice the kerosene propellant for the same sized vehicle+payload total to reach orbit. But what we actually have is about 3 times more propellant in our kerosene vehicle, 1.5 times more than is necessary to get the same vehicle size and payload to orbit. The vehicle does weigh about 10% more in dry weight, so then the total vehicle+payload weight that can now be lifted to orbit will be 1.5/1.1 = 1.364 times higher than for the hydrogen case. Now for the hydrogen powered SSTO vehicles that have been proposed the payload is a fraction of the vehicle dry weight. The 100,000 kg dry weight of the VentureStar compared to the 20,000 kg payload capacity is typical. Then the kerosene version of such a vehicle could loft (1.364)*(120,000 kg) = 164,000 kg to orbit. Or considering that our vehicle is at a dry weight of 110,000 kg with the kerosene-engine change, the payload would be 54,000 kg, 2.7 times the payload weight of the hydrogen case. As I said this is an easy calculation to do. But many people simply won't do it. They have been so conditioned to think that Isp is the most important thing that the assumption is hydrogen must be used for an SSTO. It probably doesn't help matters the fact that the gross mass becomes about 3 times as great with the dense propellants. Gross mass has been frequently used as the measure of the cost of a launch vehicle, which I like to call "the hegemony of the GLOW weight". But this is actually a very poor measure to use. The reason is propellant cost is a trivial component of the launch cost to orbit. More important is the dry mass and complexity of the launch vehicle for the payload that can be orbited. Then what's important is switching to a dense propellant allows multiple times greater payload at the same sized and similarly dry-massed vehicle. Bob Clark
From: hallerb on 21 Mar 2010 10:42 Why take along EVERYTHING for a SSTO when the vehicle could use a airplane to get the in orbit portion to at least 50,000 feet above most of the atmosphere, not tied to a single launch location, fly the airplane to a convenient launch location, fuel to get to 50,000 feet can be from tanker refueling along the way.......... granted for a really large payload a BIG HUGGER AIRLINER might need to be a custom build, but the upsides are huge. no risky loaded bomb launch being the first. SSTO is just a distraction from the more important...... LOW COST TO ORBIT!!
From: J. Clarke on 21 Mar 2010 11:49 On 3/21/2010 10:42 AM, hallerb(a)aol.com wrote: > Why take along EVERYTHING for a SSTO when the vehicle could use a > airplane to get the in orbit portion to at least 50,000 feet above > most of the atmosphere, not tied to a single launch location, fly the > airplane to a convenient launch location, fuel to get to 50,000 feet > can be from tanker refueling along the way.......... > > granted for a really large payload a BIG HUGGER AIRLINER might need to > be a custom build, but the upsides are huge. > > no risky loaded bomb launch being the first. > > SSTO is just a distraction from the more important...... > > LOW COST TO ORBIT!! Why do people think that launching from 50,000 feet will help somehow? Going into orbit is not a matter of going high, it's a matter of going _fast_. Launching from 50,000 feet or from sea level you still need to impart 18,000 miles an hour of delta-v. That's the hard part.
From: hallerb on 21 Mar 2010 14:39 On Mar 21, 11:49�am, "J. Clarke" <jclarke.use...(a)cox.net> wrote: > On 3/21/2010 10:42 AM, hall...(a)aol.com wrote: > > > Why take along EVERYTHING for a SSTO when the vehicle could use a > > airplane to get the in orbit portion to at least 50,000 feet above > > most of the atmosphere, not tied to a single launch location, fly the > > airplane to a convenient launch location, fuel to get to 50,000 feet > > can be from tanker refueling along the way.......... > > > granted for a really large payload a BIG HUGGER AIRLINER might need to > > be a custom build, but the upsides are huge. > > > no risky loaded bomb launch being the first. > > > SSTO is just a distraction from the more important...... > > > LOW COST TO ORBIT!! > > Why do people think that launching from 50,000 feet will help somehow? > Going into orbit is not a matter of going high, it's a matter of going > _fast_. �Launching from 50,000 feet or from sea level you still need to > impart 18,000 miles an hour of delta-v. �That's the hard part. the hardest part is having enough fuel onboard to get you thru the dense lower atmosphere.those large tanks weigh more. with a aircraft first stage that part is taken care of by a mature well understood technology, and since in air refueling to release altitude would be used lots of unnecessary mass wouldnt need lifited off the pad.. plus the aircraft with space plane could be released at the equator gaing some margins too. and bad weather would be much less of a issue. no more storm clouds nearing pad troubles. just fly a few hundred miles away to a nice clear area. its far easier to accelerate a lower mass object, the
From: Greg D. Moore (Strider) on 21 Mar 2010 17:40
J. Clarke wrote: > On 3/21/2010 10:42 AM, hallerb(a)aol.com wrote: >> Why take along EVERYTHING for a SSTO when the vehicle could use a >> airplane to get the in orbit portion to at least 50,000 feet above >> most of the atmosphere, not tied to a single launch location, fly the >> airplane to a convenient launch location, fuel to get to 50,000 feet >> can be from tanker refueling along the way.......... >> >> granted for a really large payload a BIG HUGGER AIRLINER might need >> to be a custom build, but the upsides are huge. >> >> no risky loaded bomb launch being the first. >> >> SSTO is just a distraction from the more important...... >> >> LOW COST TO ORBIT!! > > Why do people think that launching from 50,000 feet will help somehow? > Going into orbit is not a matter of going high, it's a matter of going > _fast_. Launching from 50,000 feet or from sea level you still need > to impart 18,000 miles an hour of delta-v. That's the hard part. Because 50,000 feet gets you above the bulk of the atmosphere which provides a decent bonus. -- Greg Moore Ask me about lily, an RPI based CMC. |