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  1. #11
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    Quote Originally Posted by DaveW View Post
    I think you're being a little optimistic there. I don't think developing a SSTO capable of transporting a couple hundred pounds to LEO could be designed, built and launched for less than $10M. The SpaceShipOne that won the X Prize cost $25M to develop, only had to reach 100km, with a payload of a single person.

    I definitely think the SSTO concept is interesting and appealing, but anyone wanting to pursue it had better have deep pockets.
    In a follow up post I'll show why I think it could be developed for just a few million dollars. This is just for an expendable SSTO. But its payload will be high enough for reentry/landing systems to be added and still have a positive payload. Developing the reusable version would cost more.
    Doing this low cost expendable test vehicle is important just to make it apparent doing an SSTO really isn't that hard, and also to show that their payload can be large enough to have reentry/landing systems and still carry significant payload.


    Bob Clark

  2. #12
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    Quote Originally Posted by DaveW View Post
    The SpaceShipOne that won the X Prize cost $25M to develop, only had to reach 100km, with a payload of a single person.
    And the altitude is the least of it. SS1 made no attempt to get anything resembling an orbit. Throwing a ball up in the air and catching it again is quite a bit less challenging than throwing it around the world.

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  3. #13
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    The point of the matter is that the many small spacecraft and suborbital craft of lightweight composite design become high Mach suborbital, a la the X-33, when switched to using high efficiency engines. And moreover if they are scaled up by a factor of 2, then these larger versions become now fully orbital vehicles.

    I discuss this in regard to the Air Forces's X-37B and Sierra Nevada's Dream Chaser here:

    Newsgroups: sci.space.policy, sci.astro, sci.physics, sci.space.history
    From: Robert Clark <rgregorycl...@yahoo.com>
    Date: Fri, 22 Jul 2011 15:09:14 -0700 (PDT)
    Subject: Re: A kerosene-fueled X-33 as a single stage to orbit vehicle.
    A kerosene-fueled X-33 as a single stage to orbit vehicle. - sci.space.policy | Google Groups

    This is also true of the X-34 and SpaceShipOne: they become high Mach suborbital, as a single stage, when switched to high efficiency engines. And when scaled up twice as large with the high efficiency engines, they become now fully orbital single stage vehicles.
    The case of SpaceShipOne is especially interesting because the twice scaled up vehicle is already built in SpaceShipTwo. Then swapping out the hybrid engines of SpaceShipTwo for high efficiency liquid fueled engines produces a SSTO.


    Bob Clark

  4. #14
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    Quote Originally Posted by RGClark View Post
    The point of the matter is that the many small spacecraft and suborbital craft of lightweight composite design become high Mach suborbital, a la the X-33, when switched to using high efficiency engines. And moreover if they are scaled up by a factor of 2, then these larger versions become now fully orbital vehicles.

    I discuss this in regard to the Air Forces's X-37B and Sierra Nevada's Dream Chaser here:

    Newsgroups: sci.space.policy, sci.astro, sci.physics, sci.space.history
    From: Robert Clark <rgregorycl...@yahoo.com>
    Date: Fri, 22 Jul 2011 15:09:14 -0700 (PDT)
    Subject: Re: A kerosene-fueled X-33 as a single stage to orbit vehicle.
    A kerosene-fueled X-33 as a single stage to orbit vehicle. - sci.space.policy | Google Groups

    This is also true of the X-34 and SpaceShipOne: they become high Mach suborbital, as a single stage, when switched to high efficiency engines. And when scaled up twice as large with the high efficiency engines, they become now fully orbital single stage vehicles.
    The SpaceShipOne isn't SSTO even if it reaches orbit, it's an air launched plane. Just because the first stage is a reuseable plane instead of a rocket stage doesn't mean it doesn't count.

    Quote Originally Posted by RGClark View Post
    The case of SpaceShipOne is especially interesting because the twice scaled up vehicle is already built in SpaceShipTwo. Then swapping out the hybrid engines of SpaceShipTwo for high efficiency liquid fueled engines produces a SSTO.


    Bob Clark
    I don't think this is really something you can claim. For one, SpaceShipTwo is designed to be air launched. I think it's reasonable to assume structural changes would be needed to deal with the higher max q that is involved with a ground launch. It also performs a feathered reentry, which wouldn't be feasible for orbital reentry. It would need to be redesigned to handle the higher heat loads reentering from orbital speeds. Just swapping out the engines and looking at thrust to weight doesn't mean it will work as an SSTO.

    This isn't to say that the SSTO isn't a good concept. I'm kind of partial to the X-34 after spending a couple weeks at Dryden and talking with some of the people who worked on the project, but that doesn't mean it was a success. It didn't reach orbit, and the project cost was closing in on $200M. It'd love to see a SSTO craft developed, but I think expecting it to be done even in the tens of millions of dollars range would be a stretch.
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  5. #15
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    Quote Originally Posted by RGClark View Post
    The point of the matter is that the many small spacecraft and suborbital craft of lightweight composite design become high Mach suborbital, a la the X-33, when switched to using high efficiency engines. And moreover if they are scaled up by a factor of 2, then these larger versions become now fully orbital vehicles.
    ...
    This is also true of the X-34 and SpaceShipOne: they become high Mach suborbital, as a single stage, when switched to high efficiency engines. And when scaled up twice as large with the high efficiency engines, they become now fully orbital single stage vehicles.
    The case of SpaceShipOne is especially interesting because the twice scaled up vehicle is already built in SpaceShipTwo. Then swapping out the hybrid engines of SpaceShipTwo for high efficiency liquid fueled engines produces a SSTO.
    SpaceShipeOne is given a dry mass of 1,200 kg:

    SpaceShipOne - Wikipedia, the free encyclopedia

    We'll fill the entire fuselage aft of the pilot's cabin up until the nozzle with kerosene/LOX propellant. I'll estimate dimensions from the image attached below. The cylindrical portion of the fuselage is about 10 feet long. After this there is a tapered portion of the fuselage that extends up to the nozzle, about 7.5 feet long. The cylindrical portion is about 5 feet wide. The narrow end of the tapered portion is about 1.5 feet wide.
    The tapered portion is in the shape of a frustum:

    Volume of a Frustum of a Cone

    By the volume formula on that page, it's volume will be (1/3)*Pi*(7.5)(2.5^2 + 2.5*.75 + .75^2) = 68.2 cu. ft.
    The volume of the cylindrical portion of the fuselage will be Pi*10*(2.5)^2 = 196.34 cu. ft., for a total of 264.57 cu. ft., or 7.5 cubic meters. The overall density of kerolox is about 1,000 kg/m^3. So this will have about 7,500 kg of propellant.
    We need to replace the hybrid engine and tanks with kerolox engines and tanks. Astronautix gives the hybrid engine of SpaceShipOne a mass of 300 kg:

    SpaceDev Hybrid

    Removing this gives the engine-less SpaceShipOne a mass of 900 kg. For a replacement kerolox engine we'll use the RD-0242-HC at 120 kg:

    RD-0242-HC

    As I mentioned before the high chamber pressure suggests this is a high performance engine. With altitude compensation it should get a vacuum Isp in the range of 360 s. As a point of comparison the rather low efficiency Merlin 1C just by using a longer, vacuum optimized nozzle increases its vacuum Isp from 305 s to 342 s. The high efficiency Russian engines also can get a sea level Isp in the range of 331 s. So we'll take this as the sea level Isp using altitude compensation. Using the estimate of Ed Kyle of the trajectory averaged Isp being 2/3rds of the way from the sea level value to the vacuum value, we'll take the average Isp as 350 s.
    We also have to add the mass of the kerolox tanks. Their mass will be about 1/100th that of the mass of propellant so at 75 kg. The total dry mass will now be 1,095 kg.
    To this we add thermal protection. The advanced ceramics used on the Air Force's X-37B mass about 12 kg/m^2. The cross-sectional area to be covered on the bottom of the vehicle from the tip of the nose cone to the end of the tapered section of the fuselage will be about 8.5 square meters. This gives a thermal protection mass of 102 kg for the fuselage. For the wings, the wing area is 15 m^2 resulting in a thermal protection mass of 180 kg for the wings. So the total mass is now 1,377 kg, call it 1,380 kg.
    Then the delta-V will be 350*9.8ln(1 + 7500/1380) = 6,385 m/s. Even adding the total mass of two pilots at 200 kg, the delta-V would still be 5,998 m/s.
    So it will be a high Mach suborbital craft. As I'll show in a following post the twice scaled up SpaceShipTwo will be a fully orbital craft when switched out to use high efficiency liquid fueled engines.
    At this high a delta-V though this reconfigured SpaceShipOne could also be used for the Air Force's Reusable Booster System program. This is intended to cut launch costs by using a reusable booster and an expendable upper stage. This will give a small low cost proof of principle version of the system.


    Bob Clark


  6. #16
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    Not to continue to be a party pooper, but don't you think that simply scaling dimensions and then looking at fuel volume and thrust/weight ratio and then declaring that such a ship could function as a SSTO when it was designed as a low Mach tourist zero-g ride is a bit of a gross simplification of the Engineering process?

    I got a chance to tour SpaceX a couple years ago, and the amount of work they do to produce a human rated capsule is incredible. I think the idea that you can take something like the SpaceShipTwo and just double the size and switch engines to produce something that can take a human into orbit and back is kind of ridiculous. In my field (Electrical Engineering) it would be like saying we're going to take an existing 1000W switching power supply, increase the dimensions by 2 and replace the components with higher rated ones to produce a 10,000W power supply. You could try, but you'd either run out of money or burn down your lab before you succeed.
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  7. #17
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    Yeah, there really are pretty severe scaling issues that are likely to crop up - not least would be that I think I remember decades ago working through a problem which pretty much showed that you could continue to scale up the size of a rocket/booster to the point that the thing would barely move.

    What is proposed is not precisely scaling up the rocket propulsion system itself, but you're still going to have problems by trying to scale everything up by a change of componentry.

    I don't see a trip to Mars by humans as being anything other than excessively costly and dangerous. Maybe once I'm long dead and gone. . . The Moon might make more sense for long-term occupancy but even then I see it as a stretch.

    Oh, and I see little excuse for the ISS to remain. Seems to me that it has been more of a political exercise than a scientific endeavor. Too costly for the return - so kill ISS and plow that money into the Hubble replacement.

    But I've never been an engineer. . .

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  8. #18
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    In regards to getting the most economical delivery of payload to orbit. Quite key here is that if you use the principle of using both the most lightweight stages and the most efficient engines at the same time then you can loft even more payload to orbit with your mult-stage launchers. Plus, the individual stages can now be used as SSTO's to loft smaller payloads at a lower cost than using the full multi-stage launchers.
    I mentioned before that SpaceX is using weight optimized design for their Falcon 9 launcher. They are getting a 20 to 1 mass ratio for the Falcon 9 first stage. And they expect to achieve a 30 to 1 mass ratio for the side boosters on their Falcon Heavy. If they had used high efficiency engines such as the NK-33 or the RD-180 instead of the Merlins on their Falcons they could loft even more payload to orbit as well as using the first stages or boosters alone as SSTO's to launch smaller payloads.
    It is notable that Elon Musk this week announced that SpaceX will be working on a "super efficient" engine which he says will allow reusable launchers that can bring the price to orbit down to $50 to $100 per pound, in the range of what I was saying. The key point is this is doable now with the high efficiency engines already existing and the lightweight stages already existing.

    August 03, 2011
    Looking at Spacex plans for Making Falcon Rockets Reusable to get to $50 per pound launch costs

    August 02, 2011
    Elon Musk of Spacex talks about a Reusable Falcon Heavy to get to $50 a pound to space.
    Two technology areas Musk didn’t like were lifting bodies/wings and nuclear rockets.
    On the former, he said he was a “vertical takeoff, vertical landing” type guy and eschewed wings since they had to be tailored for each planet’s atmosphere and were useless on airless bodies such as the Moon.
    Drawbacks to nuclear power included the need for shielding (heavy), water (heavy), and public objections against launching nuclear fuel on a rocket. “It’s a tricky thing getting a reactor up there with a ton of uranium,” Musk said and went on to say while nuclear power would be useful for Mars or lunar operations, he implied that some assembly (i.e., mining and processing fuel off planet) would be required.
    http://nextbigfuture.com/2011/08/elo...lks-about.html


    Bob Clark

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  10. #20
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    Quite key for why reusable SSTO's will make manned space travel routine is the small size and low cost they can be produced. A manned SSTO can be produced using currently existing engines and stages the size of the smallest of the very light, or personal, jets [1], except it would use rocket engines instead of jets, and the entire volume aft of the cockpit would be filled with propellant, i.e., no passenger cabin. So it would have the appearance of a fighter jet.
    We'll base it on the SpaceX Falcon 1 first stage. According to the Falcon 1 Users Guide on p.8 [2], the first stage has a dry mass of 3,000 lbs, 1,360 kg, and a usable propellant mass of 47,380 lbs, 21,540 kg. We need to swap out the low efficiency Merlin engine for a high efficiency engine. However, SpaceX has not released the mass for the Merlin engine. We'll estimate it from the information here, [3]. From the given T/W ratio and thrust, I'll take the mass as 650 kg.
    We'll replace it with the RD-0242-HC, [4]. This is a proposed modification to kerosene fuel of an existing hypergolic engine. This type of modification where an engine has been modified to run on a different fuel has been done before so it should be doable [5], [6]. The engine mass is listed as 120 kg. We'll need two of them to loft the vehicle. So the engine mass is reduced from that of the Merlin engine mass by 410 kg, and the dry mass of the stage is reduced down to 950 kg. Note that the mass ratio now becomes 23.7 to 1.
    We need to get the Isp for this case. For a SSTO you want to use altitude compensation. The vacuum Isp of the RD-0242-HC is listed as 312 s. However, this is for first stage use so it's not optimized for vacuum use. Since the RD-0242-HC is a high performance, i.e., high chamber pressure engine, with altitude compensation it should get similar vacuum Isp as other high performance Russian engines such as the RD-0124 [7] in the range of 360 s. As a point of comparison the Merlin Vacuum is a version of the Merlin 1C optimized for vacuum use with a longer nozzle. This increases its vacuum Isp from 304 s to 342 s [8]. I've also been informed by email that engine performance programs such as Propep [9] give the RD-0242-HC an ideal vacuum Isp of 370 s. So a practical vacuum Isp of 360 s should be reachable using altitude compensation.
    For the sea level Isp of the RD-0242-HC, again the version of the high performance, high chamber pressure, RD-0124 with a shortened nozzle optimized for sea level operation gets a 331 s Isp. So I'll take the sea level Isp as this value using altitude compensation that allows optimized performance at all altitudes.
    To calculate the delta-V achievable I'll follow the suggestion of Mitchell Burnside Clapp who spent many years designing and working on SSTO projects including stints with the DC-X and X-33 programs. He argues that you
    should use the vacuum Isp and just use 30,000 feet per second, about 9,150 m/s, as the required delta-V to orbit for dense propellants [10]. The reason for this is that you can just regard the reduction in Isp at sea level and low altitude as a loss and add onto the required delta-V for orbit this particular loss just like you add on the loss for air drag and gravity loss. Then with a 360 s vacuum Isp we get a delta-V of 360*9.8ln(1 + 21,540/950) = 11,160 m/s. So we can add on payload mass: 360*9.8ln(1+21,540/(950 + 790)) = 9,150 m/s, allowing a payload of 790 kg.
    To increase the payload we can use different propellant combinations and use lightweight composites. Dr. Bruce Dunn wrote a report showing the payload that could be delivered using high energy density hydrocarbon fuels other than kerosene [11]. For methylacetylene he gives an ideal vacuum Isp of 391.1 s. High performance engines can get get ca. 97% and above of the ideal Isp so I'll take the vacuum Isp value as 384 s. Dunn notes that Methyacetylene/LOX when densified by subcooling gets a density slightly above that of kerolox, so I'll keep the same propellant mass. Then the payload will be 1,120 kg: 384*9.8ln(1 + 21,540/(950 + 1,120)) = 9,160 m/s.
    We can get better payload by reducing the stage weight by using lightweight composites. The stage weight aside from the engines is 710 kg. Using composites can reduce the weight of a stage by about 40%. Then adding back on the engine mass this brings the dry mass to 670 kg. So our payload can be 1,400 kg: 384*9.8ln(1 + 21,540/(670 + 1,400)) = 9,160 m/s.
    Note this has a very high value for what is now regarded as a key figure of merit for the efficiency of a launch vehicle: the ratio of the payload to the dry mass. The ratio of the payload to the gross mass is now recognized as not being a good figure of merit for launch vehicles. The reason is that payload mass is being compared then to mostly what makes up only a minor proportion of the cost of a launch vehicle, the cost of propellant. By comparing instead to the dry mass you are comparing to the expensive components of the vehicle, the parts that have to be constructed and tested [12].
    This vehicle in fact has the payload to dry mass ratio over 2. Every other launch vehicle I looked at, and possibly every other one that has ever existed, has the ratio going in the other direction, i.e., the dry mass is greater than the payload mass. Often it is much greater. For example for the space shuttle system the dry mass is over 12 times that of the payload mass, undoubtedly contributing to the high cost for the payload delivered.
    Because of this high value for this key figure of merit, this vehicle would be useful even as a expendable launcher. However, a SSTO is most useful as a reusable vehicle. This will be envisioned as a vertical take-off vehicle. However, it could use either a winged horizontal landing or a powered vertical landing. This page gives the mass either for wings or propellant for landing as about 10% of the dry, landed mass [13]. It also gives the reentry thermal protection mass as 15% of the landed mass. The landing gear mass is given as 3% of the landed mass here [14]. This gives a total of 28% of the landed mass for reentry/landing systems. With lightweight modern materials quite likely this could be reduced to half that.
    If you use the vehicle just for a cargo launcher with cargo left in orbit, then the reentry/landing system mass only has to cover the dry vehicle mass so with lightweight materials perhaps less than 100 kg out of the payload mass has to be taken up by the reentry/landing systems. For a manned launcher with the crew cabin being returned, the reentry/landing systems might amount to 300 kg, leaving 1,100 kg for crew cabin and crew. As a mass estimate for the crew cabin, the single man Mercury capsule only weighed 1,100 kg [15 ]. With modern materials this probably can be reduced to half that.
    For the cost, the full two stage Falcon 1 launcher is about $10 million. The engines make up the lion share of the cost for launchers. So probably much less than $5 million just for the 1st stage sans engine. Composites will make this more expensive but probably not much more than twice as expensive. For the engine cost, Russian engines are less expensive than American ones. The RD-180 at 1,000,000 lbs vacuum thrust costs about $10 million [16], and the NK-43 at a 400,000 lbs vacuum thrust costs about $4 million [17]. This is in the range of $10 per pound of vacuum thrust. On that basis we might estimate the cost of the RD-0242-HC of about 30,000 lbs vacuum thrust as $300,000. We need two of them for $600,000.
    So we can estimate the cost of the reusable version as significantly less than $10,600,000 without the reentry/landing system costs. These systems added on for reusability at a fraction of the dry mass of the vehicle will likely also add on a fraction on to this cost. Keep in mind also that the majority of the development cost for the two stage Falcon 1 went to development of the engines so in actuality the cost of just the first stage without the engine will be significantly less than half the full $10 million cost of the Falcon 1 launcher. The cost of a single man crew cabin is harder to estimate. It is possible it could cost more than the entire launcher. But it's likely to be less than a few 10's of millions of dollars.

    REFERENCES.
    1.)List of very light jets.
    List of very light jets - Wikipedia, the free encyclopedia

    2.)Falcon 1 Users Guide.
    http://www.spacex.com/Falcon1UsersGuide.pdf

    3.)Merlin (rocket engine)
    4 Merlin 1C Engine specifications
    Merlin (rocket engine) - Wikipedia, the free encyclopedia

    4.)RD-0242-HC.
    RD-0242-HC

    5.)LR-87.
    LR-87 - Wikipedia, the free encyclopedia

    6.)Pratt and Whitney Rocketdyne's RS-18 Engine Tested With Liquid Methane.
    by Staff Writers
    Canoga Park CA (SPX) Sep 03, 2008
    Pratt and Whitney Rocketdyne's RS-18 Engine Tested With Liquid Methane

    7.)RD-0124.
    RD-0124

    8.)Merlin (rocket engine).
    2.5 Merlin Vacuum
    Merlin (rocket engine) - Wikipedia, the free encyclopedia

    9.)Propep
    Software

    10.)Newsgroups: sci.space.policy
    From: Mitchell Burnside Clapp <cla...@plk.af.mil>
    Date: 1995/07/19
    Subject: Propellant desity, scale, and lightweight structure.
    sci.space.policy | Google Groups

    11.)Alternate Propellants for SSTO Launchers
    Dr. Bruce Dunn
    Adapted from a Presentation at:
    Space Access 96
    Phoenix Arizona
    April 25 - 27, 1996
    Alternate SSTO Propellants

    12.)A Comparative Analysis of Single-Stage-To-Orbit Rocket and Air-Breathing Vehicles.
    p. 5, 52, and 67.
    A Comparative Analysis of Single-Stage-To-Orbit - Learn More at GovWin

    13.)Reusable Launch System.
    Reusable launch system - Wikipedia, the free encyclopedia

    14.)Landing gear weight (Gary Hudson; George Herbert; Henry Spencer).
    Landing gear weight (Gary Hudson; George Herbert; Henry Spencer)

    15.)Mercury Capsule.
    Mercury Capsule

    16.)Wired 9.12: From Russia, With 1 Million Pounds of Thrust.
    Wired 9.12: From Russia, With 1 Million Pounds of Thrust

    17.)A Study of Air Launch Methods for RLVs.
    Marti Sarigul-Klijn, Ph.D. and Nesrin Sarigul-Klijn, Ph.D.
    AIAA 2001-4619
    p.13
    http://mae.ucdavis.edu/faculty/sarig...a2001-4619.pdf

 

 
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