Tuesday, July 10, 2018

DARPA's Spaceplane: an X-33 version, Page 3.

Copyright 2018 Robert Clark

 In the previous post "DARPA's Spaceplane: an X-33 version, Page 2"I discussed some recent high strength metal alloys that might give the X-33 and VentureStar even lighter weight propellant tanks than originally envisioned. Still, because of the conformal, noncylindrical shape of the tanks, they would still not have the weight efficiency of a cylindrically shaped rocket. Then I'll discuss some methods that will come close to having the weight efficiency of cylindrical tanks.

Multi-tubed Propellant Tanks.
 One possibility would be to achieve the same lightweight tanks as cylindrical ones by using multiple, small diameter, cylindrical tubes. 

 A similar idea is described here:

Assemblies of Conformal Tanks
Space is utilized efficiently and sloshing is reduced.
This Prototype Assembly of Conformal Tanks was built to demonstrate the feasibility of building such an assembly to fit an approximately toroidal available volume.

 You could get the same volume by using varying lengths and diameters of the multiple cylinders to fill up the volume taken up by the tanks. The cylinders would not have to be especially small. In fact they could be at centimeter to millimeter diameters, so would be of commonly used sizes for tubes and pipes.

 The weight of the tanks could be brought down to the usual 35 to 1 ratio for aluminum-lithium hydrogen/oxygen propellant to tank mass. Then the mass of the tanks on the X-33 would be 210,000 lbs/35 = 6,000 lbs, saving 9,200 lbs off the vehicle dry weight. This would allow the hydrogen-fueled X-33 to achieve its originally desired Mach 15 maximum velocity.

 But note that since now the tanks are composed of cylindrically-shaped tubes, we no longer have the problem of the conformally-shaped carbon composite tanks failing. Then we could get in the range of 50 to 1 propellant to tank mass ratio by using carbon composites for the cylindrically-shaped tubes, and we could reduce the dry mass due to the tanks by an additional 2,000 lbs.

 The same multi-tube approach applied to the full-scale VentureStar would allow it to significantly increase its payload carrying capacity. At a 35 to 1 aluminum-lithium ratio of propellant mass to tank mass for cylindrical tanks, the 1,929,000 lbs propellant mass would now require a mass of only 1,929,000/35 = 55,000 lbs for the tanks, a saving of 83,000 lbs off the original tank mass. This could go to extra payload, so from 45,000 lbs max payload to 128,000 lbs max payload.

 But again we could now use carbon-composites for the cylindrical tubes. This would shave an additional 16,000 lbs off the weight of the tanks, and increase the payload now to 144,000 lbs.

 An analogous possibility might be to use a honeycombed structure for the entire internal makeup of the tank. The X-33's carbon composite tank was to have a honeycombed structure for the skin alone. Using a honeycomb structure throughout the interior might result in a lighter tank in the same way as does multiple cylinders throughout the interior.

 For the multi-tube approach and the honeycombed variant there would be a significant problem for maintenance however. As this is intended to be reusable, we would have difficulty examining all the interior parts of the tanks for cracks or punctures. 

 If you removed the many layers of the multi-tube method layer by layer for examination that would involve significant time and expense. Even more importantly by so many times having to physically move one of the layers of tubes you run an increased risk of damaging one of the tubes. 

 One possibility is that since there would be a gap space between one tube and a tube positioned diagonally to it, we could use this space to insert high resolution imaging equipment. It might also work to insert x-ray imaging devices.

Partitioned Propellant Tanks to Save Weight.
  A different approach to getting near cylindrical-tube weight efficiency, might be to model the tanks, viewing them vertically, as conical but with a flat front and back, and rounded sides. Then the problem with the front and back naturally trying to balloon out to a circular cross section might be solved by having supporting flat panels at regular intervals within the interior. 

 The X-33 composite tanks did have support arches to help prevent the tanks from ballooning but these only went partially the way through into the interior. You might get stronger a result by having these panels go all the way through to the other side.
These would partition the tanks into portions. This could still work if you had separate fuel lines, pressurizing gas lines, etc. for each of these partitions and each got used in turn sequentially. A preliminary calculation based on the deflection of flat plates under pressure shows with the tank made of standard aluminum alloy and allowing deflection of the flat front and back to be only of millimeters that the support panels might add only 10% to 20% to the weight of the tanks, while getting similar propellant mass to tank mass ratio as cylindrical tank. 

 Note you might not need to have a partitioned tank, with separate fuel lines, etc., if the panels had openings to allow the fuel to pass through. These would look analogous to the wing ribs in aircraft wings that allow fuel to pass through. You might have the panels be in a honeycomb form for high strength at lightweight that still allowed the fuel to flow through the tank. Or you might have separate beams with a spaces between them instead of solid panels that allowed the fuel to pass through between the beams

 We'll view the X-33 hydrogen tanks standing vertically as conical with flattened front and back and rounded sides. This report on page 19 by the PDF file page numbering gives the dimensions of the X-33 hydrogen tanks as 28.5 feet long, 20 feet wide and 14 feet high:

Final Report of the X-33 Liquid Hydrogen Tank Test Investigation Team.

 Call it 9 meters long, 6 meters wide, and 4.3 meters deep for this calculation. I'll simplify the calculation by approximating the shape as rectangular, i.e., uniformly 6 meters wide. Note that the rounded portions of the sides, top, and bottom will be considered separately. I'll call the vertical length of each section x, and the bulkhead thickness h. Since the length of the tank is 9m, the number of sections is 9/x.

 Typically propellant tanks are pressurized in the 20-40 psi range. I'll take it as 30 psi; call it 2 bar, 2x10^5 Pa. Referring to the drawing of the tank, each bulkhead takes part in supporting the internal pressure of the two sections on either side of it. This means for each section the internal pressure is supported by one-half of each bulkhead on either side of it, which is equivalent to saying each bulkhead supports the internal pressure of one section.

 The force on each section is the cross-sectional area times the internal pressure, so 6m*x*(2*10^5 Pa), with x as in the diagram the vertical length of each section. The bulkhead cross-sectional area is 6m*h, with h the thickness of the bulkheads. Then the pressure the bulkheads have to withstand is 6m*x*(2*10^5 Pa)/6m*h = (2*10^5 Pa)*x/h.

 The volume of each bulkhead is 6m*h*4.3m. The density of aluminum-lithium alloy is somewhat less than aluminum, call it 2,600 kg/m^3. So the mass of each bulkhead is (2,600 kg/m^3)*6m*h*4.3m = 67,080*h. Then the total mass of all the 9/x bulkheads is (9/x)*67080*h = 603,720*(h/x).

 Note that additionally to the horizontal bulkheads shown there will be vertical bulkheads on the sides. These will have less than 1/10 the mass of the horizontal bulkheads because the length of each section x will be small compared to the width of 6m, and will have likewise small contribution to the support of the internal pressure.

 The tensile strength of some high strength aluminum-lithium alloys can reach 700 MPa, 7*10^8 Pa. Then the pressure the bulkheads are subjected to has to be less than or equal to this: (2*10^5 Pa)*x/h <= 7*10^8 Pa, so x/h <= 3,500, and h/x => 1/3,500. Therefore the total mass of the bulkheads = 603,720*(h/x) => 172.5 kg. Note we have not said yet how thick the bulkheads have to be only that their total mass is at or above 172.5 kg, for one of the twin rear tanks. It's twin would also require 172.5 kg in bulkhead mass. The third, forward, tank had about 2/3rds the volume of these twin rear tanks so I'll estimate the bulkhead mass it will require as 2/3rds of 172.5 kg, 115 kg. Then the total bulkhead mass would be 460 kg, about 15% of the 3,070 kg tank mass I calculated for the reconfigured X-33.

 This is for the bulkheads resisting the outwards pressure of the sections. Notice I did not calculate the pressure inside the tank on the bulkheads from the propellant on either side. This is because the pressure will be equalized on either side of the bulkheads. However, we will have to be concerned about the pressure on the rounded right and left sides of the tank, and the rounded top and bottom of the tank, where the pressure is not equalized on the outside of the tank.

 Before we get to that, remember the purpose of partitioning the tank was to minimize the bowing out of the front and back sides from the internal pressure. Consider this page then that calculates the deflection of a flat plat under a uniform load:

eFunda: Plate Calculator -- Clamped rectangular plate with uniformly distributed loading.
This calculator computes the maximum displacement and stress of a clamped (fixed) rectangular plate under a uniformly distributed load.

 In the data input boxes, we'll put 200 kPa for the uniform load, 6 meters for the horizontal distance, .3 m, say, for the vertical distance, and 6 mm for the thickness of the plate. For the vertical distance x I'm taking a value proportionally small compared to the tank width, but which won't result in an inordinate number of partitioned sections of the tank. For the thickness I'm taking a value at 1/1000th the width of the tank, which is common for cylindrical tanks. For the material specifications for aluminum-lithium we can take the Young's modulus as 90 GPa. Then the calculator gives the deflection as only 2.35mm, probably adequate.

 However, we still have to consider what happens to the rounded sides and the bottom and top. Look at the last figure on this page:

Thin-Walled Pressure Vessels.

 It shows the calculation for the hoop stress of a cylindrical pressure vessel. The calculation given is 2*s*t*dx = p*2*r*dx, using s for the hoop stress. This implies, s = p*r/t, or equivalently t = p*r/s. So for a given material strength s, the thickness will depend only on the radius and internal pressure.

 However, what's key here is the same argument will apply in the figure if one of the sides shown is flat, instead of curved. Therefore in our scenario, the rounded sides, top and bottom, which we regard as half-cylinders, will only need the thickness corresponding to a cylinder of their same diameter, i.e., one of a diameter of 4.3m. 

 So the rounded portions actually require a smaller thickness than what would be needed for a cylinder of diameter of the full 6m width of the tank.

 This means the partitioned tank requires material of somewhat less mass than a cylindrical tank of dimension the full width of the tank plus about 15% of that mass as bulkheads.

 The new high strength metal alloys might also save further on this weight. However, we now have to consider the Young's modulus of the alloys, because of the deflection of the plates calculation, and not just the tensile strength.

A Key Advantage of Partitioned Tanks.
 There is an another advantage of using partitioned tanks in addition to the weight savings. A problem with weight growth of the X-33/VentureStar arose in regards to the size of the wings. For stability reasons, you would want the center of gravity (CG) to remain ahead of the center of pressure (CP) during the entire flight. But as the propellant is burned off, the propellant mass near the front will be decreased and this will increase the effect of the heavy engines at the rear on the moving the CG rearward. To deal with this problem during the X-33 development, the wings kept getting larger and larger. But this cancels out the advantage the X-33 had in its dry mass in its original design in not needing heavy wings.

 The original X-33/VentureStar was supposed to look like this:

 But in the later incarnations, it looked like this:

 The added wing size was to move the CP rearward to keep the CG ahead of it. 

 Partitioned propellant tanks are nothing new actually in aerospace. They are quite commonly used on jet airliners to deal with the problem of CG shift as fuel is burned off:

Balancing by Fuel-Pumping.
The Concorde Tank-Schematic:

"1 + 2 + 3 + 4 are the Collector-Tanks, feeding the engines directly. Usually they feed there counterpart engines – but they can be cross-switched to feed more and/or other engines at the same time.
5 + 7 and 8 + 6 are the Main-Transfer Tanks, feeding the 4 Collector-Tanks. Initially 5 + 7 are active. If those are empty 6 + 8 take over (or must be activated from the Engineering Panel!).
5a + 7a are Auxiliary-Tanks (to 5 and 7).
9 + 10 are the Trim-Tanks for balancing forward
11 is the Trim-Tank for balancing afterward"

 Then the partitioned tanks could solve two problems of the dry mass of the X-33/VentureStar: weight growth in the tanks and in the wings.

 Bob Clark

Sunday, June 10, 2018

DARPA's Spaceplane: an X-33 version, Page 2.

Copyright 2018 Robert Clark

 The OldSpace companies had always discounted the viability of reusable launchers on the grounds that the launch market was not enough to pay for it. However, a new market will soon be opening up for hundreds to thousands of launches required for the impending satellite megaconstellations. Now even the OldSpace company ArianeSpace is speaking of transitioning to reusability.

 So with reusability soon to become prevalent we have now further justification for resurrecting the X-33. Boeing supported by a DARPA grant is developing a reusable, spaceplane first stage, the XS-1, then Lockheed with the X-33 would have a competing reusable launcher.

 In the blog post DARPA's Spaceplane:an X-33 version, I discussed that the X-33 used as a reusable first stage has importance beyond that of just a test stage of an operational SSTO, the VentureStar. For the X-33 could be its own operational vehicle, cutting costs in its own right as a reusable first stage.  But intriguingly the problems that originally doomed the X-33 and its SSTO follow-on the VentureStar may also be solvable.

 As discussed in that earlier post, it was the failure of the composite tanks that caused the X-33 program to be cancelled. But some new high strength aluminum alloys may have the comparable lightweight characteristics as carbon composite tanks.

 Carbon composite propellant tanks are a pretty well developed technology, as long as they are cylindrically shaped. But the unusual conformal shape of the composite tanks on the X-33 caused them to fail.

 Carbon composite saves about half-off the weight of standard aluminum tanks. But interestingly some new aluminum alloys have comparable high strength at lightweight as carbon composite and therefore could be used to give the lightweight tanks needed. 

 See for example the graphic here:

  The 7075 T6 alloy has nearly twice the strength per weight as the standard 6061 T6 alloy, and the 7068 T6 was nearly 2.5 times better. 

 A consideration as described on that page is that 7075 is 2 to 3 times more expensive than the standard 6061 and the 7068 is 3 to 4 times more expensive. But considering that because of their higher strength, smaller amounts of the material by a factor of 2 to 2.5 would be needed the price difference in practice would not be as great.

 Note also since it was the inability to produce the composite tanks in the X-33 at the needed lightweight that caused the program to be cancelled, existence of the high strength aluminum alloys make the SSTO VentureStar once again viable.

 Development Cost.

 The cost of carbon fiber is about twice that of standard aluminum, so the cost of the tanks with high strength aluminum would not be much more than the cost of the carbon fiber X-33. Since Lockheed would be paying this itself, it might want first to do a smaller version of the X-33.

 In the earlier "DARPA's SpacePlane" post, I suggested a smaller version half-size in linear dimensions of the X-33 might cost ca. $45 million to build. This would test the technology and moreover using it as an upper stage of the X-33 would give a fully reusable system.

   Bob Clark

UPDATE 7/4/2018: 

 I've been informed of other other high strength, lightweight metal alloys that could also allow VentureStar to achieve its goal of a being a reusable SSTO, and allow the X-33 to be able to serve as a low cost reusable first stage.

 The alloys have various strengths and weaknesses. For example some are are just now being experimented with but their measured strength-to-weight ratio is more than 3 times better than standard aluminum. Some are steel alloys which have better weldability than the aluminum alloys, etc.

 For instance in the graphic above, the titanium 6Al-4V alloy is a little better than the 7075 and is already used in rockets for example for solid motor casings.

 There is also a high strength steel alloy, the 17-7 PH stainless steel CH 900:

Re: SpaceX second stage secret sauce?

 It has comparable strength-to-weight as the 7068, i.e., nearly 2.5 times better than standard aluminum. It also has better weldability than the aluminum alloys.

 A recent report shows some high strength aluminum alloys such as the 7075 can be 3D-printed:

Engineers Have Found a Way to 3D Print Super Strong Aluminum.
B. Ferguson/HRL Laboratories
by Dom Galeon September 22, 2017 Hard Science

 This is useful since the high strength aluminum alloys such as the 7075 have poor weldability. But the conformal shapes of the X-33/VentureStar tanks would be difficult to make without welding.

 Ti 5553 alloy is another ultra strong titanium alloy, even better than the Ti 6Al-4V. It has a max tensile strength in the range of 1,400 MPa. At a density of 4.64 gm/cc, this puts it in strength-to-weight ratio at even better than the 7068 alloy, and nearly 3 times better than standard aluminum:

Processing of a metastable titanium alloy (Ti-5553) by selective laser melting.
November 2016Ain Shams Engineering Journal 8(3)

Finally, a titanium alloy known as the Ti185 was long known but it was difficult to produce it so it had uniform strength throughout. A new method of producing it using titanium hydride powder can produce it so it is uniformly strong:

Low-cost and lightweight: Strongest titanium alloy aims at improving vehicle fuel economy and reducing CO2 emissions
April 1, 2016, Pacific Northwest National Laboratory

 Approaching 1,700 MPa in tensile strength, it would be 3.5 times better on strength-to-weight than standard aluminum. Because it is made of titanium hydride powder, it may also be possible to make it by 3D-printing, which would solve the problem of producing a conformal shape for the tanks of the X-33/VentureStar.

Thursday, June 7, 2018

Half-size Ariane core stage for a reusable launcher.

Copyright 2018 Robert Clark

 Long-time space advocates will recall back in the late 90's there was a push for large numbers of communication satellites for the purposes of cell-phone communication. This led to the creation of several private launch companies then to serve what was expected to be hundreds to thousands of required launches.

 However, it turned out the great majority of cell-phone communications could be served by terrestrial cell towers. The large satellite constellation plans were then abandoned, and those private launch companies then collapsed.

 But now once again there are renewed plans for satellite megaconstellations containing hundreds to thousands of satellites, such as OneWeb or SpaceX's StarLink. This time it is primarily for high speed internet service. This time there is billion dollar backing for the projects and there have been preliminary launches to test the idea.

 It's very likely now that the projects will take place. For space advocates, an important result of the large numbers of launches required is that it provides a clear advantage for low cost reusable launchers.

 SpaceX always believed reusable launchers could be financially feasible. But other space launch providers were skeptical. They didn't think the number of launches under the current market would pay for reusability.

 But now with the advent of the new megaconstellation plans even previously skeptical Arianespace plans to transition to reusability. See for example the articles here:

 One project Arianspace is planning is called Callisto. It is to be a small sized hydrolox test vehicle to test reusable, vertical landing boosters. It is to be analogous to the SpaceX Grasshopper tests.

 However, unlike the SpaceX Grasshopper that used the original Merlin engines and the same F9 propellant tanks, though perhaps only partially filled, the Callisto plan is to use an entirely newly designed and built stage.

 I see a problem with this. For the money spent on Callisto, it will not be an actual operational vehicle. This mirrors a problem with the X-33 test vehicle that was supposed to test the technologies for an operational SSTO vehicle. But for all the money spent on the X-33, it itself would not have been an operational vehicle.

 I believe this was a mistake. It would have been better if the X-33 itself was to be used as an operational vehicle. It could have been used as a reusable first stage booster to cut costs for a two-stage to orbit system, a la the SpaceX plan:

DARPA's Spaceplane: an X-33 version.

 Then my recommendation is not to repeat the mistake of the X-33 program by instead actually using operational stages to test reusability and vertical landing.

 This could be done with two existing Arianespace stages. The Ariane 5 core stage and the Ariane H10-3 cryogenic upper stage. In both cases you would use partially filled tanks, approx. half-filled so that the stage could lift-off on their single engines.

 For the operational versions, you would make the tanks themselves half-size, instead of half-filling a full-sized tank, to save dry mass, at least for the Ariane 5 core. For the Ariane H10-3 for the upper stage use it might be able to carry its full propellant load dependent on the propellant load on the Ariane 5 core to be able to lift off on its single Vulcain engine.

 Another advantage of this approach is that it would finally provide Europe with an independent manned spaceflight capability.

 There is a key problem that would need to be solved. Discussion on on a space forum was that the Vulcain II is not throttleable. The HM7-B engine used on the Ariane H10-3 upper stage is also not throttleable. Then both engines would need to be upgraded to be throttleable. As support for the idea this should be feasible, it should be noted the original versions of the SpaceX Merlin engines prior to the Merlin 1D were not throttleable. SpaceX has also shown with its "hoverslam" approach to vertical landing, it would not have to have a high degree of throttleability. Probably the degree of throttleability common to liquid fuel engines in the range of 60% would be sufficient.
See discussion here:

A half-size Ariane for manned spaceflight.

 Bob Clark

UPDATE: 6/10/2018

 In the discussion above I forgot a key point. The most important factor in
regards to cost is not the development cost.
The key cost factor is what they would charge per flight for a reusable
launcher. Robert Zubrin made this point insightfully in one of his books. He
recounts that he made the argument for reusable launchers in his former job
with one of the big launch companies.

 He argued that they could cut the cost of launch by an order of magnitude.
The company execs responded: why would we do that? Their view was their
revenue would then be slashed by a factor of ten. They were assuming the
market would still be the same but they would be getting one-tenth the

 So the OldSpace companies were acting quite rationally in a business sense
in discounting reusability. They were saying the market was not enough to
make it advantageous to them.

 But if there were a large market then they would make more money making more launchers at the lower price. That is, the price would be reduced by a factor of ten but the number of launches would be increased by more than a factor of ten.

 Also, the importance of the large market and lowered prices for satellite
launches extends beyond that of just the satellite market. By making
launches at such reduced prices, that increases the possible market for
passenger flights to space. So the impending megaconstellation launches may
also bring to fruition the long desired routine passenger flights to space.

Sunday, February 18, 2018

Multi-Vulcain Ariane 6.

Copyright 2018 Robert Clark

 The worldwide space community was amazed by the success at which SpaceX was able to launch the Falcon Heavy and land the two side boosters back at the launch site: 

SpaceX Falcon Heavy Test Flight - Launch and Landing.

 This led the current head of ESA Jan Woerner to suggest that ESA should promote reusability for future launchers. But, Woerner argues, the currently planned version of the Ariane 6 using two solid side boosters and with no reusability is so far along it should probably be completed:

Posted on 11/02/2018 by Jan Woerner

Posted on 15/02/2018 by Jan Woerner

 However SpaceX is progressing so rapidly with reusability, and reduced costs, that by the time the Ariane 6 is expected to reach full operation in 2023 it's high cost may make it obsolete.

 But in fact a faster, and much cheaper alternative is possible for the Ariane 6 that would have also have the advantage it could be made reusable: and that is to give the Ariane 5 core an additional Vulcain 2 engine so it can lift off without needing the solid side boosters. 

 According to studies by the French space agency CNES this could be done for only a $200 million development cost:

On the lasting importance of the SpaceX accomplishment, Page 3: towards European human spaceflight.

 This is compared to the multibillion dollar development cost for the Ariane 6 version with side boosters. 

 I had previously written about the possibility of giving the Ariane 5 a second Vulcain engine here:

The Coming SSTO's: multi-Vulcain Ariane.

 However, that was in the context of producing a SSTO. But SSTO's are still controversial. And in any case it would have a lower payload capacity than the solid booster version of the Ariane 6, and couldn't, in itself, fulfill the important role of launching GEO satellites.

 So I'll discuss the case of giving the Ariane 5 core a second Vulcain engine while keeping an upper stage. I was surprised that my calculation showed this version without side boosters was able to get close to the 6.5 ton to GTO requirement ESA set for the Ariane 6 just with the two liquid stages, no solid side boosters.

 This article discussed the various options that were being considered for the Ariane 6 including a version that added a second Vulcain engine to the Ariane 5 core:

CNES, ASI Favor Solid-Rocket Design For Ariane 6.
By Amy Svitak
Source: Aviation Week & Space Technology
October 15, 2010

You see it only lists 2,200 kg to GTO with the two stage H2C all-liquid configuration of the Ariane 6.

 And even more surprisingly, my calculation shows my two-stage version could get the high GTO payload without using the expensive new upper stage being planned for the Ariane 6 with the new Vinci engine. It could do it with an already existing cryogenic Ariane H10 upper stage.

 This indicates why the development cost for this version would be so low: no new solids need be developed, already existing Vulcain 2 engines are used, no new expensive upper stage and Vinci upper stage engine need be developed, and already existing Ariane upper stages are used.

 I'll use Dr. John Schilling's launch performance calculator to estimate the payload possible:

Launch Vehicle Performance Calculator.

 Use publicly available sources for the specifications of the Ariane 5G core, Vulcain 2 engine, and Ariane H10 upper stage to enter in data into the Schilling calculator. The calculator takes the vacuum values for the Isp and thrust even for first stage engines, since it already takes into account the diminution of these values at sea level. Then with two Vulcain engines the input page looks like this:

 Some quirks of the Schilling calculator you need to be aware of though if you use it. First, always select "No" for the "Restartable Upper Stage" option, otherwise the payload will be reduced. Second, always set the "Inclination" to match the launch site latitude otherwise the payload will be reduced.

 Note that while the thrust indicated for the first stage at 2,700 kilonewtons (vacuum) is well above the rocket gross mass, the sea level thrust for the two Vulcains is about 190,000 kilogram-force. With payload, this would not be enough to lift off. A couple of solutions. One would be to change the mixture ratio at lift off to increase the oxidizer to fuel ratio. This decreases Isp, but increases thrust. Another possibility is that liquid fuel engines can be run at some percentage above their rated values. For instance the space shuttle main engines could be run at 9% above their rated thrust values and the latest version of the SpaceX Merlin engines run at about 15% above the Merlins rated thrust value.   Another possibility would be to reduce the propellant load. I'll look at this last solution in a moment.

 In any case with these specifications the calculator gives this payload to LEO:

 Now to get the GTO payload, change the "Apogee" entry from 185 km to 35,000 km. Then the calculator gives the result:

Altitude Compensation case. 
 We can get even better performance by using altitude compensation on the Vulcains. As discussed on this blog various methods exist to give standard bell nozzle engines, altitude compensating nozzles. These new additions would not be heavy or expensive.

 As shown by the RL10-B2 engine, with a nozzle extension a hydrogen-fueled engine can get a vacuum Isp in the range of 465 s. So change the first and second stage Isp's to 465 s and increase the vacuum thrust proportionally. Then the payload for the LEO case would be:

  And the payload to GTO would be:

 This indicates the importance of altitude compensation even for multi-stage vehicles. The payload can be increased in the range of 25%. However, it could be argued the Schilling calculator was not formulated to deal with the altitude compensating case, so these estimates are less reliable. A legitimate concern. Accurate trajectory simulations need to be done to determine if altitude compensation can really increase payload to this extent even with multi-stage rockets.

Reduced Propellant Case.
 As mentioned above another method to deal with the lift off thrust not being able to loft the rocket with payload, is to reduce the propellant load. In fact the currently adopted version of the Ariane 6 is planned to reduce the propellant on the core down to 140,000 kg.

 If we do this to our all liquid version, this brings the LEO payload to 10,430 kg  and the GTO payload to 4,890 kg.
 And with the altitude compensation, the LEO payload comes to 12,800 kg and alt. comp. GTO payload to 6,370 kg.

And Yet Even More.
 Beyond the importance of giving Europe a reusable, low cost launcher, there is also the fact that being all liquid-fueled, no solid motor boosters, Europe now would have an orbital launcher that could be used for manned spaceflight. Europe would finally have its own independent manned spaceflight capability.

   Bob Clark

Thursday, December 28, 2017

Altitude compensation attachments for standard rocket engines, and applications, Page 6: space shuttle tiles and other ceramics for nozzles. UPDATED: 3/6/2018

Copyright 2017 Robert Clark

Some possibilities for altitude compensating nozzles include actual adjustable-sized nozzles but using non-flexible materials such as ceramics. In such cases the high expansion ratios needed for optimized expansion at high altitude would give nozzles of impractical size, for instance they wouldn't fit within the width of the rocket stage.

A couple of possibilities for dealing with this eventuality:


Vacuum optimized nozzles, whether altitude compensating or not can take a great deal of mass of an engine or stage. See for example the specifications of the Star 48B here:

You see the weight of the nozzle assembly is nearly that of the case assembly. This would become even more of a weight problem with extreme expansion ratios of hundreds to one. 

 The lightweight space shuttle underside tiles may provide a solution. Their volume density is only 0.144 gm/cm^3. And according to this report their areal density is 1.19 gm/cm^2:

TPS Materials and Costs for Future Reusable Launch Vehicles.

 The AETB-8/TUFI listed is a toughened tile material that has higher impact resistance while maintaining the same temperature resistance.

 Judging from the size of the size of the Star 48B nozzle, the nozzle weight might be reduced from 90 lbs to ca. 20 lbs using the AETB-8/TUFI tiles.

 A possibly even more lightweight material was developed by aerospace engineer/mathematian GW Johnson. He described it in this video presentation at the 16th Mars Society conference in 2013:

Reusable Ceramic Heat Shields - GW Johnson - 16th Mars Society Convention.

 Johnson estimated the volume density as only 0.03 gm/cm^3, a third that of the shuttle tiles. Also interesting is that Johnson originally developed the material to act as insulation for ramjet combustion chambers. 

 Note that even the insulation in the Star 48B solid motor is a sizable weight at 60 lbs, compared to the 129 lbs casing weight. Then Johnson's ceramic might also be able to be used as a lightweight replacement for solid motor insulation.

  Bob Clark 

UPDATE, 3/6/2018

 A problem though is the shuttle tiles and also the Johnson tiles have low tensile strength, only about 1 bar. This probably can be improved by adding strengthening bars to the ceramic as steel rebar is used to improve the strength of concrete.

 Additionally, as the exhaust travels down a nozzle towards the exit the pressure also decreases. So perhaps it could be used as the nozzle extension for the nozzle below where the exhaust gas falls below 1 bar pressure. This could result in a significant reduction in weight. See for instance the comparison size of the RL10 nozzle used on the DC-X optimized for low altitude flight compared to one for high altitude, near vacuum flight:

 This might be hard to test though for amateurs. Amateurs don't have access to the expensive test equipment used to test engine performance under high altitude, near vacuum conditions.

 To do the test at sea level we would need to make the attachment so that the length of the nozzle would make the exit pressure to be quite low. This then entails dangerous loads on the nozzle due to the overexpansion at sea level ambient pressure. This would be made even worse due to the brittle nature of the ceramic.

 The most effective test would be on the rockets by university and amateur teams whose rockets have reached 100K feet. The pressure there is only about 1% that of sea level.

 The problem with getting a material to work is that it has to be both high temperature and lightweight. But what if it didn't have to be high temperature?

 I was puzzled by these images showing the nozzle extensions on rocket engines glowing red hot while firing:

RL-10B2 engine with nozzle extension.

Merlin Vacuum engine with nozzle extension.

  But when the exhaust extends down a nozzle it both expands and cools. So why were these nozzle extensions glowing red hot?

 Perhaps they get red hot because they are heated by conduction by the upper parts of the nozzle where the exhaust gas is still hot.

 If so, then we could use the ceramic tiles as insulation between the upper part of the nozzle, and the attached nozzle extension. In that case the nozzle extension would need to be as heat resistant and there would be many more materials that would have the lightweight characteristic required.

Tuesday, October 24, 2017

A Small Raptor Spaceship.

Copyright 2017 Robert Clark

 In the blog post "SpaceX BFR tanker as an SSTO", I suggested the BFR tanker could as an expendable SSTO get a comparable payload to orbit as the Falcon Heavy at ca. 50 metric tons.  SpaceX however wants to move to reusables. But Elon in his presentation on the BFR suggested the BFR tanker as a reusable SSTO might get less than 15 metric tons to orbit.

 This large loss in payload when switching to reusable is due to the large amount of propellant that must be kept on reserve for the vertical, propulsive landing. I argued then for using winged, horizontal landing to retain most of the payload of the expendable case, perhaps only a 10% drop from ca. 50 tons to ca. 45 tons.

 This would be different from the SpaceX preferred vertical landing method. But the production of a routine manned orbital flight capability is so important it should be implemented even if it would require completely different spaceships for the orbital flight and interplanetary flight uses.

 There is also the fact that if it does prove to be the case that switching to a winged, horizontal landing allows most of the expendable SSTO payload to be retained, then this may also be the case for the full two-stage BFR. Then instead of losing 100 tons off the expendable 250 ton payload to only 150 tons as reusable, a 40% loss, perhaps only 10% would be lost, so a 225 ton reusable payload. 

 This would be important for maintaining high payload with reusability both for the SSTO and full two-stage cases. But both of these are high payload launchers at ca. 45 and 225 tons. But SpaceX has spoken of moving all their launchers to the Raptor engines. In that case SpaceX needs a small launcher.

 We could make it half-size to the BFR upper stage. However, to give it flexibility to also be used as a upper stage we'll make it one-quarter size, at a ca. 275 ton propellant load. A first level estimate would put the dry mass at 1/4th of the BFR tanker dry mass so at 50/4 = 12.5 metric tons. But as mass ratio improves as you scale up a rocket, so also does it reduce as you scale a rocket down. Then as a second level estimate we'll take the scaling relationship as Elon did in his presentation on the BFR.

 Elon cited the dry mass of the BFR spaceship as 85 tons, a factor of 85/75 = 1.13 times more than just by proportional scaling. However, this small size stage is 1/4th size, not just half-size. So we'll apply this scaling factor twice, i.e., by the factor 1.13^2 = 1.28. Then we'll take the dry mass as 1.28*12.5 = 16 tons. 

 This can be lofted by two of the Raptor engines with a total sea level thrust of 2*1,700,000 N = 3,400,000 N = 347,000 kilogram-force. Now use the Schilling launch vehicle performance calculator to estimate the payload. The estimator takes the vacuum Isp and thrust values so 375 s Isp and 2*1,900,000 N = 3,800,000 N total thrust.

 Then the input screen appears as:

 And the results appear as:

A payload of ca. 8,100 kg as an expendable SSTO. Using the 10% payload loss estimate for a winged reusable, this would be ca. 7,300 kg as a reusable SSTO. 

 This might be enough to carry the manned Dragon 2 to orbit as a reusable SSTO. Based on its size being 1/8th that of the ITS upper stage, which has a $130 million production cost, we can estimate the production cost of this SRS as $16 million, and keep in mind it is intended to have hundreds of flights. We can imagine then private individuals purchasing their own small, reusable launchers to orbit.

 Moreover, Elon has spoken of doing short hops of the BFR upper stage to test the point-to-point transport capability. A quarter-scale version would allow this to be tested with less financial risk. Such a smaller test vehicle would also allow you to test more cheaply alternative return methods such as winged, horizontal landing. This test vehicle if successful could then go right to production for use by SpaceX for launches they sell, or for selling to individuals for conducting their own launches.

 To increase the payload, we may want to add another engine. After giving the dry mass an additional 1,000 kg, and raising the total thrust to 5,700,000 N because of the third engine, the input screen appears as:

 And the results are:

 A 9,700 kg expendable SSTO payload. After a 10% reusable payload loss, the reusable SSTO payload would be 8,700 kg. This higher payload may be necessary for the Dragon 2 to also carry the launch abort system.

Two-stage case.
 We'll calculate now the payload using the BFR tanker as the booster and the SRS as the upper stage.
The input page now looks like:

And the results page looks like:

Methods of increasing take-off thrust.
 One issue with this two-stage rocket though is the low liftoff thrust/weight ratio when carrying the upper stage might make the actual payload less than that indicated by the calculator. So we'll explore some possibilities of increasing the take-off thrust. One possibility is a thrust scale up. For instance the Merlin Full Thrust is at about a 15% increase above the rated thrust value of the Merlin 1D. And the SSME had a maximum thrust value 9% above its rated value.

 However, another possibility is a recent research advance in engines known as "thrust augmentation nozzle", TAN. It's sort of like an afterburner for rocket engines. What it does is inject propellant into the nozzle and ignites it to generate additional thrust. See discussion here:

Thrust Augmented Nozzles
Posted on November 12, 2007 by Jonathan Goff

 In experiments the researchers were able to increase thrust by up to 70%.

 It is notable that TAN also serves as a method of altitude compensation for it allows larger, vacuum optimized nozzles to also be used at sea level by preventing separation by conducting combustion also in the nozzle. Some method of altitude compensation should be used to optimize performance both at sea level and vacuum rather than making trades of which combination of sea level, mid level, and vacuum engines to use. Some possibilities to do the altitude compensation, though not the liftoff  thrust increase, are discussed at, [1], [2], [3], [4], [5].

 A variation on TAN may allow also to increase the effective Isp at sea level. The idea behind the variation is to use atmosphere air as the oxidizer for the augmented thrust combustion.

We will need to bring the atmospheric air into the nozzle to burn with the fuel. One possibility is indicated here:

Rocket motor thrust nozzle with means to direct atmospheric air into the interior of the nozzle.
US 3469787 A.

 We would open up vents on the nozzle to allow air to flow in, then burn it with the fuel. We would have to insure the vents we opened were further down on the nozzle so that the reduced pressure of the exhaust flow further down would allow the atmospheric air to enter in. We also don't want after we ignite the fuel with the air for this exhaust to exit back out the vents, further reason for making the opened vents to be further down the sides of the nozzle.

  Another implementation of this idea would use the aerospike nozzle.

 The fuel would be emitted from the sides of the aerospike lower down on the spike where the exhaust pressure is lower and the ambient air pressure would constrain the combustion.

 A problem with both the vented nozzle and aerospike implementations though is the pressure of combustion would be at most one bar. This would limit the thrust produced. Still, it may be the mass of nitrogen heated along with the oxygen might permit sufficient thrust production.

 Another possibility is to use a vapor-air detonation as the combustion method. This will permit high exhaust speed for the combustion:

Methane-Air Detonation Experiments at NIOSH Lake Lynn Laboratory.

  Bob Clark

1.)Altitude compensation attachments for standard rocket engines, and applications.

2.)Altitude compensation attachments for standard rocket engines, and applications, Page 2: impulse pressurization methods.

3.)Altitude compensation attachments for standard rocket engines, and applications, Page 3: stretchable metal nozzles.

4.)Altitude compensation attachments for standard rocket engines, and applications, Page 4: the double aerospike.

5.)Altitude compensation attachments for standard rocket engines, and applications, Page 5: metal foil expandable nozzles.

Sunday, October 22, 2017

SpaceX BFR tanker as an SSTO.

Copyright 2017 Robert Clark

 Elon Musk has suggested the development of orbital point-to-point manned transport may pay for the development of his Mars colonization plans, the idea being there would be a great market for such manned flights. Peter Diamandis has made this point as well, that there would be a great market for such flights to orbit:

Peter Diamandis: Taking the next giant leap in space.

 Elon suggests such orbital transports would be best implemented as two-stage vehicles. However, this simulation shows the upper stage of the SpaceX Interplanetary Transport System (ITS) introduced by Elon Musk in 2016 could get a total 190 metric tons to orbit as an expendable SSTO, including the stage and the payload:

 Since the stage was estimated to weigh 90 tons, this would mean 100 metric tons payload as an expendable SSTO. Then the question is how much mass would be taken up for propellant for return using vertical landing.

 However, in the latest incarnation introduced in 2017 the upper stage of the now BFR is about half as large in propellant load and with 6 engines instead of 9. Here are screen grabs from the video on the latest version:

  So we may estimate this half-size version to have half the dry mass at ca. 45 metric tons and could get approx. 50 metric tons to LEO as an expendable SSTO.

 As justification for this dry mass estimate, here's the description of the original ITS upper stage, both spaceship and tanker versions:

 And here's the description of the BFR spaceship, half size to the ITS version:

 You see the BFR spaceship is at about half the listed dry mass value as the ITS spaceship.  Actually during the video Musk says the design mass was just at half at 75 tons, but the 85 tons mass was allowing for weight growth. So it is plausible the BFR tanker is ca. half the mass of the ITS version or a little more, ca. 45+ tons.

 We'll give it 9 engines instead of 6 so it'll have enough thrust to lift off with the heavy payload, and give it a dry mass of 50 metric tons with weight growth. Now do a payload estimate using Dr. John Schilling's launch performance estimator. The high expected thrust/weight ratio for the Raptors means they'll weigh perhaps 1,000 kg each. So the addition of 3 will add perhaps 3 tons to the dry mass. Since the payload is so high this will be a relatively small payload loss.

 The Schilling estimator takes the vacuum values for the Isp and thrust, so enter 375 s as the Isp and 9*1,900 kN= 17,100 kN for the total thrust.

 Insure that the "Restartable upper stage" option is set to "No" otherwise the payload will be reduced. And set the launch inclination to match the launch site, so at 28.5 degrees for Cape Canaveral:

 Then the result is:

 Confirming the ca. 50 metric ton payload as an expendable SSTO.

 But this puts it as an expendable SSTO in the payload range of the expendable Falcon Heavy while being also in the same size range of the Falcon Heavy. So this SSTO would get the same payload fraction as a 2 and 1/2 stage vehicle. Moreover, judging from the fact the ITS tanker upper stage was to cost $130 million production cost, the half size BFR tanker might only be $65 million, so it would be half the cost of the Falcon Heavy. But the Falcon Heavy as an expendable launcher already would be a significant cut in the cost to orbit. So the BFR tanker as an expendable SSTO could be a great reduction in the cost to space, compared to current values.

 I had earlier done a calculation that showed the Falcon Heavy as an expendable with 53 metric ton payload capacity and $125 million launch cost could be financially feasible as a tourism vehicle to orbital space or transport to orbital space hotels:

Falcon Heavy for Orbital Space Tourism.
 So this BFR tanker could likewise be feasible financially as an expendable SSTO, as the price should be well less than $125 million as an expendable. But of course SpaceX wants to make it reusable. The reusability should cut the launch cost multiple times. Then the question is how much will reusability cut into the payload mass?

Reusable SSTO case.
  In the presentations on the ITS and BFR both the spaceship and tanker versions of the upper stage were always presented as reusable. So it is likely the heat shield mass is already included in the cited vehicle dry mass values. I'm estimating though surprisingly high values for the thermal protection of the BFR upper stages, either spaceship or tanker versions. I'm using the fact as indicated in the wiki page on the BFR that it will use the PICA-X thermal protection material. Several references give the PICA-X density as about 0.25 gm/cc = 250 kg/m3, and the thickness as on the Dragon 2 as 7.5 cm, 0.075 m, about 3 inches.

 The BFR upper stage has a length of 48 meters and a width of 9 meters. The top part of the stage is conical. Visually, this top portion is about 1/3rd the vehicle length, so about 16 meters long. So I'll approximate the bottom area to be covered by thermal protection covered as 32*9 + (1/2)*16*9= 360 m2.

 Then the volume of the thermal protection material is 360 * 0.075 = 27 m3. At a density of 250 kg/m3, that amounts to a mass for the thermal protection of 27 * 250 = 6,750 kg, which is a surprisingly high included mass in the dry mass of 85 tons for the spaceship upper stage or the 50 tons for the tanker upper stage.

 One possibility, is the thickness of the PICA-X for the Dragon 2 is coming from the fact it is doing a ballistic reentry, thus generating high heat. However, the BFR upper stage will be doing a more gentle gliding reentry. So perhaps the thermal protection will only need to be half as thick, so weigh half as much. For instance, the heat shield tiles on the underside on the space shuttle only weighed half the PICA-X tiles.

 And new versions of these space shuttle tiles used on the X-37B are more durable while remaining lightweight:

The X-37B stands in front of part of the fairing that protects it during launch, showing off the silica tiles on its underside. Those TUFI (toughened uni-piece fibrous insulation) tiles are said to be more durable than their counterparts on the space shuttle. On the leading edge of the wings, meanwhile, are TUFROC (toughened uni-piece fibrous refractory oxidation-resistant ceramic ) tiles, which NASA named the government winner of its 2011 Invention of the Year Award.

 Elon has implied the reusable version of the BFR upper stage would only get perhaps in the range of 10 to 15 metric tons payload (by saying it's an order of magnitude less than the full BFR 150 ton reusable payload.) That loss in payload is not coming from the heat shield mass since that's already included in the vehicle dry mass. The loss in payload is high though, 40 tons, nearly the size of the entire vehicle dry mass, presumably because of the amount of the propellant that needs to be kept on reserve for landing on return.

 I'd like to see a trade study of the payload of instead going with wings for a horizontal landing. See for example the discussion here:

 Wings typically take up only 10% of an aircraft dry mass. Then with carbon composites, that would be cut to less than 5% of the landed (dry) mass. Keep in mind the loss in payload with vertical, propulsive landing is nearly 100% of the vehicle dry mass. Also, going with short, stubby wings as with the X-37B, you can make the wing weight even less:

 The areal size of the wings in that case would also be less than that of bottom area of the BFR tanker, perhaps only 1/4th to 1/3rd the areal size. So the increase in heat shield mass would only be at most 1/3 that of the approx. 6,750 kg mass of the current heat shield, so perhaps an extra 2,250 kg. But actually the addition of wings gives a gentler glide slope so probably the heat shield thickness could be reduced. The result might even be the total heat shield mass would be reduced by adding wings.

 Elon has spoken of preferring vertical landing because it could be used generally on both worlds with and without atmospheres. However, to achieve his desired goal of making mankind multiplanetary, making human orbital spaceflight commonplace is an important part of that goal. A lower development and production cost BFR upper stage acting as a reusable SSTO would go a long way towards that goal. So even if the optimized orbital or point-to-point transport looks completely different than the interplanetary lander, such as requiring wings for example, it would still be important to develop it.

 Advantage of altitude-compensation.
 In the discussion after the introduction of the BFR, Elon Musk and other commenters on various space forums, engaged in alot of speculation on the optimal combination of sea level engines, vacuum engines, and some medium, intermediate area ratio engines. This is necessitated because the highest Isp engines, optimized for vacuum, can not operate reliably, and safely at sea level. But then using sea level engines or even intermediate level engines would subtract from the Isp possible.

 This illuminates once again the importance of implementing altitude compensation engines in space flight, at least for Earth launch. This would permit the high thrust needed at launch at sea level, as well as the high Isp needed at near vacuum.

 There are many ways to implement the altitude compensation and none are particularly difficult to do. Some methods are discussed here [1], [2], [3], [4], [5].


1.)Altitude compensation attachments for standard rocket engines, and applications.

2.)Altitude compensation attachments for standard rocket engines, and applications, Page 2: impulse pressurization methods.

3.)Altitude compensation attachments for standard rocket engines, and applications, Page 3: stretchable metal nozzles.

4.)Altitude compensation attachments for standard rocket engines, and applications, Page 4: the double aerospike.

5.)Altitude compensation attachments for standard rocket engines, and applications, Page 5: metal foil expandable nozzles.