A kerosene-fueled X-33 as a single stage to orbit vehicle.
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From: Robert Clark on 14 Mar 2010 21:24 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 a kerosene-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 Russian kerosene engines 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.htm 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 the kerosene-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 an airplane and doesn't fall in the ocean after one flight?" "The answer to this question is: All of these vehicles were fantasy projects. They violated basic laws of physics and engineering. They were impossible with current technology, or any technology we can afford to develop on the timescale and budgets available to NASA. They were doomed attempts to avoid the Cold Equations of Spaceflight." http://www.spacedaily.com/news/oped-05zy.html 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. Bob Clark
From: Me on 15 Mar 2010 10:02 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. This is a sign of a mental problem.
From: Robert Clark on 16 Mar 2010 12:37 On Mar 14, 9:24 pm, Robert Clark <rgregorycl...(a)yahoo.com> wrote: > ... > 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 >... This is another key advantage of dense propellant vehicles for the SSTO application, that the delta-V to orbit would be about 300 m/s less than for a hydrogen-fueled SSTO vehicle. The main idea behind this is that dense propellant vehicles burn mass so much more quickly that they achieve the speed needed to attain the right altitude for orbit more quickly. Since the gravity loss is dependent on the time spent on this vertical portion of the trip, dense propellant vehicles experience less gravity loss. Still, the explanation is probably not easy to grasp unless you do the actual numerical calculations over the trajectory of the flight. However, I can show an approximate calculation that makes the idea more understandable below. This Wikipedia article also mentions the fact that dense propellant vehicles require 300 m/s less delta-V to orbit than hydrogen vehicles: Single-stage-to-orbit. # 4 Dense versus hydrogen fuels. "The end result is the thrust/weight ratio of hydrogen-fueled engines is 3050% lower than comparable engines using denser fuels." "This inefficiency indirectly affects gravity losses as well; the vehicle has to hold itself up on rocket power until it reaches orbit. The lower excess thrust of the hydrogen engines due to the lower thrust/weight ratio means that the vehicle must ascend more steeply, and so less thrust acts horizontally. Less horizontal thrust results in taking longer to reach orbit, and gravity losses are increased by at least 300 meters per second. While not appearing large, the mass ratio to delta-v curve is very steep to reach orbit in a single stage, and this makes a 10% difference to the mass ratio on top of the tankage and pump savings." http://en.wikipedia.org/wiki/Single-stage-to-orbit#Dense_versus_hydrogen_fuels However, the explanation given here is not quite correct. This rather implies it is a function of greater thrust/weight ratio only. But in actual fact the lowered delta-V required for dense fuels applies *even when the hydrogen and the dense fuel vehicles have the same thrust/weight ratio*. For the calculation of the delta-V savings for dense fuels, suppose both the dense-fueled and hydrogen-fueled vehicles have a initial T/W of, say, 1.3. Let Mi be the initial gross mass of the vehicle, r the constant propellant flow rate, Ve the exhaust velocity, a(t) the acceleration, changing with time, of the vehicle due to the thrust, and g the acceleration due to gravity 9.8 m/s^2. Then the mass of the vehicle at time t is Mi-r*t, and the thrust force is (Mi-r*t)a(t). We'll use the fact that the thrust of a rocket is (propellant flow rate)x(exhaust velocity) to get the equation (Mi-r*t)a(t) = r*Ve. We can solve this for the acceleration to get a(t) = r*Ve/(Mi-r*t). Now because we set the initial thrust/weight ratio as 1.3 we know that thrust = r*Ve = 1.3(g*Mi), so Mi = r*Ve/(1.3g). Then plug this into the equation for acceleration to get: a(t) = r*Ve/(r*Ve/(1.3g) - r*t) = Ve/(Ve/(1.3.g) - t). Quite notable here is that the propellant flow rate cancels out and the acceleration due to the thrust depends only on the exhaust velocity Ve, or equivalently, only on the Isp. Then for the vertical portion of the trip where gravity drag takes place, the rocket's acceleration will be Ve/(Ve/(1.3g)-t) - 9.8. Now it may not be apparent at first glance but this formula says the acceleration is greater for a smaller exhaust velocity Ve, so for a smaller Isp. To make it clearer multiply top and bottom of the expression for a(t) by 1.3g to bring it to 1.3g*Ve/(Ve-1.3g*t). Then if you do the division this becomes 1.3g + t*(1.3g)^2/(Ve-1.3g*t). Now you see because the Ve is in the denominator the expression is larger when Ve is *smaller*. So a dense propellant with a lower Isp will accelerate faster during this vertical portion of the trip meaning it spends less time when gravity drag is operating so that gravity drag is reduced. You couldn't make the Isp be arbitrarily small though because that would result in huge fuel loads and tanks, and, most importantly, engines to get the vehicle off the ground. Bob Clark
From: Robert Clark on 18 Mar 2010 15:02 In the first post of this thread I calculated that switching to kerosene would allow the hydrogen-fueled suborbital X-33 to now become an orbital craft. However, I thought it would be able to carry minimal payload if any. However, I realize I used too low a value for the density of chilled LOX at 1,160 kg/m^3. It should be actually about 10% higher than the usual 1,142 kg/m^3. This is described here: Alternate Propellants for SSTO Launchers. Dr. Bruce Dunn Adapted from a Presentation at: Space Access 96 Phoenix Arizona April 25 - 27, 1996 http://www.dunnspace.com/alternate_ssto_propellants.htm In table 2 it gives the densities of some chilled fuels including kerosene, i.e., RP-1, and of LOX. The density given for the chilled kerosene is 867 kg/m^3, and for LOX 1,262 kg/m^3. So for the 296 m^3 volume I was taking for the X-33 propellant tanks and a 2.7 mixture ratio for the NK-33 engine, this gives a kero/LOX propellant mass of 332,600 kg. Now taking the average Isp of the NK-33 as 315 s, this gives a delta-V for the 21,700 kg dry mass, reconfigured X-33 of 8,797 m/s. But when you take into account you get a 462 m/s velocity boost for free from launching at the equator, you only need about 8,500 m/s delta-V to be provided by the rocket to reach orbit. This allows us to add payload. Adding 2,300 kg payload, the delta-V becomes 8,500 m/s, sufficient for orbit. We can actually get higher payload than this by using more energetic hydrocarbons than kerosene. For instance in table 2 of Dunn's paper on alternate SSTO propellants, he gives the payload for chilled methylacetylene/LOX as 24% higher than for chilled kero/LOX. This would be a payload of 2,850 kg. These payload amounts would also allow the X-33 to carry a 2 man crew in its 5 by 10 foot payload bay in a tandem arrangement a la the F-14 seating arrangement. So you could get a fully reusable, SSTO vehicle at much reduced price than the full-sized VentureStar. This article gives the price to build a new X-33 as $360 million in 1998 dollars: Adventure star. http://www.flightglobal.com/pdfarchive/view/1998/1998%20-%203141.html Even taking into account inflation, the cost to build the kerosene- fueled version should be comparable or perhaps even less because of the drop in prices for carbon composites and because kerosene engines are generally cheaper than hydrogen ones. The launch preparation costs should also be low since the X-33 was expected to be operated by only a 50 man ground crew compared to the 18,000 required for the shuttle system: Lockheed Secret Projects: Inside the Skunk Works. By Dennis R. Jenkins http://books.google.com/books?id=DUkl5bH6k6EC&lpg=PA95&dq=x-33%20venturestar&lr=&pg=PA106#v=onepage&q=&f=true Say the builder expected 25% profit over costs of the vehicle over 100 flights. That would be a charge of $4.5 million per flight. At a 2,850 kg payload capacity that would be $1,580 per kilo, or $720 per pound, to orbit. Not as good as the full-sized VentureStar but still significantly better than current launch prices. Note that the other half-scale suborbital demonstrators for the NASA RLV program by Rockwell and McDonnell-Douglas (see images linked below) could be built for comparable prices and would likewise become full orbital craft by switching to kerosene or other dense propellant. Then we could have 3 separate designs for fully reusable SSTO vehicles at costs that could allow fully private financing that would significantly reduce launch costs and would allow manned flights. Successful operation of these X-33-sized orbital vehicles at a profit would encourage private financing to build the full-scale VentureStar- sized RLV's that could bring launch costs down to the $100 to $200 per kilo range. Bob Clark http://www.astronautix.com/nails/x/x33rock.jpg http://www.astronautix.com/graphics/x/x33p4.jpg
From: Robert Clark on 21 Mar 2010 04:06
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. 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. 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 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. Its 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 its 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%, from 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 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 with the kerosene-engine change, the payload would be 54,000, 2.7 times the payload weight of the hydrogen case. As I said this is an easy calculation to do. But many people simply wont 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 doesnt 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". 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 whats important is switching to a dense propellant allows multiple times greater payload at the same sized and similarly dry-massed vehicle. Bob Clark |