VentureStar is it possible now

RobertDyck · 2016-04-20 06:01:33

Tom Kalbfus wrote:

What about the Orion? Is that reusable? For the amount of money NASA is spending on it, it should be!

No. Not reusable.

Tom Kalbfus · 2016-04-20 06:07:53

It could be replaced with something that is. Problem with Orion is that it is designed to reenter the Earth's atmosphere at Earth's escape velocity+, we don't have permanent heat shields that can do that over and over again. If we want a reusable interplanetary vehicle, we need something that slows to low Earth orbit first and then reenter's the atmosphere. I don't know about grazing the Earth's atmosphere to enter low Earth orbit. We haven't returned many things to Earth from interplanetary space.

RobertDyck · 2016-04-20 06:16:44

kbd512 wrote:

Not quite. The materials that the prototype was made of, when scaled out, would not have provided the lift capability the LM was contractually obligated to provide. As a result, someone at LM decided to try using composite tanks. The sub-scale test articles to prove that the composite tanks could withstand the structural loads applied to them and adequately contend with the environment created by the LH2 worked, but LM had a hell of a time making that work. When the tech was scaled up to the sizes required of the X-33 demonstrator, IIRC they had three successive pressure test failures. In other words, while it may have been possible to fabricate really small LH2 tanks, the technology wasn't there to fabricate larger LH2 tanks.

The original proposal included solid wall composite tanks. So composite tanks were included from the beginning. DC-XA used solid wall composite LH2 tank, and lithium-aluminum for LOX. So working size solid wall composite LH2 tank was already demonstrated.

Lockheed-Martin tried to change the solid wall design to a hollow wall. That is a double wall tank, each wall as thin as a piece of paper, with honeycomb structure between the walls. The honeycomb structure made of more composite material, also as thin as a piece of paper, but oriented on edge between the walls. Theoretically the hollow all design is lighter for the same strength, but when they tested it, the paper thin walls developed hairline cracks due to thermal stress. They tested it with liquid nitrogen, because 78% of Earth's air is nitrogen, and it isn't flammable, so it's safe. Liquid nitrogen seeped through the cracks, filling the cells in the hollow wall. When they drained the tank, liquid nitrogen remained within cells of the hollow wall. When it warmed, thermal expansion sealed the cracks. As it continued to warm, liquid nitrogen boiled. Phase change from liquid to gas caused expansion, but since cracks were sealed, that built pressure. As the tank warmed, the pressure burst the hollow wall, the tank disintegrated.

This demonstrates a hollow wall composite tank is not compatible with cryogenic propellant. That's why they tested it in a hanger. But DC-XA already proved a solid wall composite tank does work with LH2. So they should have stuck with that.

They had 2 failures that I recall. The first was they built a tank, but left it outside the hanger to cure. It emitted toxic and flammable vapours as it cured, so didn't want to contain those vapours within a sealed hanger. Unfortunately the janitorial crew thought everything left outside was garbage, so hauled it away. The second was the disintegration failure described above. Was there a third?

Disintegration had another problem. Analysis showed a student who helped assemble the hollow wall tank had used a piece of tape to hold down a part while epoxy cured. The tape was never removed, it became integrated within the wall. When the tank disintegrated, it failed at that tape. The piece of tape was a major flaw; however, even if the tape was not there, the tank would have failed anyway.

Last edited by RobertDyck (2016-04-20 06:22:14)

SpaceNut · 2016-04-20 17:10:27

Lots of drift since my last post but we are back....

Orion was originally targetted for 10 re-uses but due to rising mass and cost that was cut out of the design.

I am thinking that the newer stir weld tanks and 3 d printing using lasers might be the way to make an intergrated tank and airframe for multiple use if we get away from LH2 as a fuel that makes the tanks larger and harder to control at temperatures when trying to reduce mass....

Antius · 2016-04-21 05:06:55

SpaceNut wrote:

Lots of drift since my last post but we are back....
Orion was originally targetted for 10 re-uses but due to rising mass and cost that was cut out of the design.
I am thinking that the newer stir weld tanks and 3 d printing using lasers might be the way to make an intergrated tank and airframe for multiple use if we get away from LH2 as a fuel that makes the tanks larger and harder to control at temperatures when trying to reduce mass....

I have done limited work calculating mass ratios for different vehicle and fuel combinations. I think it will be very difficult to use anything other than LOX/H2 for a propellant in an Earth-launch SSTO. The problem is that dV to reach orbit is ~9500m/s when gravity losses and air resistance are taken into account. The exhaust velocity of LOX/CH4 (the next best propellant combination) is ~3800m/s. So structural and payload mass fraction at take-off must be <8% take-off mass. That's a tall order indeed, especially if you are talking about lifting bodies and non-spherical tank configurations.

One of the reasons I favour TSTO is that it allows the use of LOX-Kerosine for lower stage and LOX-LPG for upper stage, both of which are cheap, dense and easy to handle (O2 is a soft cryogen), and still allows decent structural mass fractions for both stages. That means you can use low-cost pressure fed engine and propellant feed systems, high-strength steels for tank and structure and about 80% of your vehicle can be recovered from the sea using simple drag-chutes to reduce impact velocity. That was the whole idea behind Sea Dragon, which remains the best reusable space vehicle design produced to date.

Tom Kalbfus · 2016-04-21 06:07:58

Antius wrote:

SpaceNut wrote:
Lots of drift since my last post but we are back....
Orion was originally targetted for 10 re-uses but due to rising mass and cost that was cut out of the design.
I am thinking that the newer stir weld tanks and 3 d printing using lasers might be the way to make an intergrated tank and airframe for multiple use if we get away from LH2 as a fuel that makes the tanks larger and harder to control at temperatures when trying to reduce mass....
I have done limited work calculating mass ratios for different vehicle and fuel combinations. I think it will be very difficult to use anything other than LOX/H2 for a propellant in an Earth-launch SSTO. The problem is that dV to reach orbit is ~9500m/s when gravity losses and air resistance are taken into account. The exhaust velocity of LOX/CH4 (the next best propellant combination) is ~3800m/s. So structural and payload mass fraction at take-off must be <8% take-off mass. That's a tall order indeed, especially if you are talking about lifting bodies and non-spherical tank configurations.
One of the reasons I favour TSTO is that it allows the use of LOX-Kerosine for lower stage and LOX-LPG for upper stage, both of which are cheap, dense and easy to handle (O2 is a soft cryogen), and still allows decent structural mass fractions for both stages. That means you can use low-cost pressure fed engine and propellant feed systems, high-strength steels for tank and structure and about 80% of your vehicle can be recovered from the sea using simple drag-chutes to reduce impact velocity. That was the whole idea behind Sea Dragon, which remains the best reusable space vehicle design produced to date.

Here is the Sea Dragon
https://en.wikipedia.org/wiki/Sea_Drago … SVSD-4.png
This article states that it was designed to haul 550 tons of cargo into Earth orbit. As you can see by this diagram, the Sea Dragon as designed would have been a huge rocket, it would have had 5 times the lift capacity as a Saturn V rocket, no doubt it would have been expensive to built, but if we could reuse the bottom stage, we might just be able to justify its cost, and 550 tons to low Earth orbit, might be enough for a manned mission to Mars.

Last edited by Tom Kalbfus (2016-04-21 06:09:50)

Antius · 2016-04-21 15:07:55

Might not be so expensive as it sounds. A517 pressure vessel steels have yield strength 900MPa and cost $10/kg. The materials cost for the whole stage is therefore $20million. Nitrous oxide propane propellant can safely be stored in a single vessel, is self-pressurising and has vacuum isp of 312s. A pressure-fed engine can have fewer than 100 parts. Bottom line is that a technologically simple lower stage could be very cheap.

SpaceNut · 2016-04-21 18:42:15

Here is a search for all topics that are comments with Sea Dragon within them use one of those or create a new one....as we are very off topic....for Venture star

Antius I do not know beans about burn rates or delta V produced but the reaso for any multi stage to work better is that you are dropping the dead mass off from the nexts stage burning....

To do something simular would be to stuff an srb into a chamber and once its spent allow it to dead fall out of the remaining rocket once the next set of engines begin to fire...

kbd512 · 2016-04-21 20:08:02

Antius wrote:

If you air-to-air refuel at 40,000 feet, you shave about 0.5km/s off of your delta-V. That doesn't sound like much. On the plus side, if that 0.5km/s had to be provided by rocket thrust from ground level without refuel, both engine efficiency and propulsive efficiency would be low, due to the need to operate engines at 1 bar and the simple fact that the rocket is accelerating from zero. So inflight refuelling might make the difference between an SSTO that is workable and one that is not. Even with that advantage, it would be difficult to sacrifice the need for LOX/H2 fuel if you wanted to achieve a vehicle with a half decent payload ratio. A compromise might be to top up with LOX only, which is much easier to transfer than LH2.

Antius,

I'm not worried so much about shaving dV off the requirement as creating an affordable and practical SSTO by using a reasonable takeoff weight. This vehicle is highly dependent upon its inert mass and designing a vehicle that has to takeoff or land anywhere near its fully fueled mass is going to make the vehicle so heavy as to be impractical. GW has a pretty good article on why VTHL vehicles are so impractical using chemical propellants.

Let's use VentureStar (VTHL) and Star Raker (HTHL) as the examples of impractical SSTO designs:

VentureStar had to have an empty weight of just 62,700 lb for a vehicle with a span of 68 ft and length of 127 ft. I can't think of any other aerospace vehicles nearly so large with empty weights that low. The Isp requirement was so high and the empty mass requirement so low because it had to carry every drop of fuel required to take off and ascend to orbit with its 20t payload. That's 1000t GLOW to put a 20t payload into orbit. For comparison, a Falcon 9R (v1.1?) has a 505t GLOW to orbit a 13t payload using engines with substantially lower Isp and a nearly identical inert mass.

Star Raker was designed to use ten 140K lb thrust turbojets, which didn't exist and still don't exist, and three RS-25 thrust class engines. The vehicle was larger and heavier than the AN-225. It had to use jettisonable landing gear. It had lower thermal protection requirements due to lower wing loading, but that was its only design advantage. Every other aspect of its operation was considerably more complicated than the Space Shuttle or VentureStar. Star Raker also took off with every drop of fuel required to attain orbit, although in-flight refueling was proposed. As proposed, it would have had a 2270t GLOW and 100t payload. That's a better payload mass fraction than VentureStar or STS and a far more useful heavy lift vehicle, but still at staggering size and weight.

Is there any way to use production F135 engines to climb to altitude, load the majority of the fuel after the vehicle is already between 40k and 60k, perhaps using StratoLaunch to pull double duty as a tanker, and then make the ascent using rocket engines? We obviously need full-flow staged combustion LOX / RP1 engines with aerospike nozzles.

Is this a workable solution?

kbd512 · 2016-04-21 20:30:11

RobertDyck wrote:

The original proposal included solid wall composite tanks. So composite tanks were included from the beginning. DC-XA used solid wall composite LH2 tank, and lithium-aluminum for LOX. So working size solid wall composite LH2 tank was already demonstrated.

My recollection of the composite tank issues with this program is not the best. I thought they had three successive pressure test failures, but I could be mistaken.

Regarding the use of aluminum-lithium alloys, I believe my recollection is correct.

From NASA's website: X-33/VentureStar

Such was the scale of the initial protest, go-ahead was given to build the LOX (liquid oxygen) tank out of the same aluminium-lithium alloy that is currently used on the external tanks for the Space Shuttle, a small but important victory for the protesting engineers at the time. The LOX tank passed testing and was installed with plumbing and electronics around the front third of the vehicle’s structure.
...
Such was the consensus that a failure was the only outcome for the composite tank, MAF (Michoud Assembly Facility) engineers had already started the process on having their own Al-Li LH2 tank ready for fabrication, pre-empting the call for change of plan.
Faced with a project failure, Lockheed Martin and X-33 NASA managers gave the green light to proceed with the fabrication of the new tank. Ironically this new tank weighed in less than the composite tank – disproving one of the reasons for going with a composite tank in the first place.
...
While the aluminium LH2 tank was much heavier than the composite tank in the skins, the joints were much lighter, which was where all the weight in the composite tank was, due to the multi-lobed shape of the tank requiring a large amount of surrounding structure, such as the joints. Ironically, the original design of the X-33 on the drawing board had the tanks made out of aluminium for this reason – but the cost played a factor for the potential customer base.

RobertDyck wrote:

Lockheed-Martin tried to change the solid wall design to a hollow wall. That is a double wall tank, each wall as thin as a piece of paper, with honeycomb structure between the walls. The honeycomb structure made of more composite material, also as thin as a piece of paper, but oriented on edge between the walls. Theoretically the hollow all design is lighter for the same strength, but when they tested it, the paper thin walls developed hairline cracks due to thermal stress. They tested it with liquid nitrogen, because 78% of Earth's air is nitrogen, and it isn't flammable, so it's safe. Liquid nitrogen seeped through the cracks, filling the cells in the hollow wall. When they drained the tank, liquid nitrogen remained within cells of the hollow wall. When it warmed, thermal expansion sealed the cracks. As it continued to warm, liquid nitrogen boiled. Phase change from liquid to gas caused expansion, but since cracks were sealed, that built pressure. As the tank warmed, the pressure burst the hollow wall, the tank disintegrated.

Tech just hadn't caught up to where it need to be to make this feasible.

RobertDyck wrote:

This demonstrates a hollow wall composite tank is not compatible with cryogenic propellant. That's why they tested it in a hanger. But DC-XA already proved a solid wall composite tank does work with LH2. So they should have stuck with that.

At that time, that was true.

RobertDyck wrote:

They had 2 failures that I recall. The first was they built a tank, but left it outside the hanger to cure. It emitted toxic and flammable vapours as it cured, so didn't want to contain those vapours within a sealed hanger. Unfortunately the janitorial crew thought everything left outside was garbage, so hauled it away. The second was the disintegration failure described above. Was there a third?

I thought there was, but I could be incorrect.

RobertDyck wrote:

Disintegration had another problem. Analysis showed a student who helped assemble the hollow wall tank had used a piece of tape to hold down a part while epoxy cured. The tape was never removed, it became integrated within the wall. When the tank disintegrated, it failed at that tape. The piece of tape was a major flaw; however, even if the tape was not there, the tank would have failed anyway.

All in all, the tech just wasn't ready at that time. Today, it might be a completely different story. I really don't know.

SpaceNut · 2016-04-22 16:45:26

http://www.3ders.org/articles/20140430- … llite.html
Lockheed Martin & RedEye 3D print giant fuel tank models for satellite

http://www.designnews.com/author.asp?doc_id=272890
Composite Rocket Fuel Tank Will Explore Deep Space

http://www.jeccomposites.com/news/compo … honeycombs
3D-printing of carbon-fiber epoxy honeycombs
Harvard engineers use carbon fiber-epoxy resin inks and 3D printing to construct lightweight cellular composites that mimic the material performance of balsa wood

GW Johnson · 2016-04-22 17:34:50

Composites are inherently porous. You must have an impermeable liner of some sort, even with LOX, if you choose to use them. With LH2, the diffusivity problem right through solid materials is far worse. I myself have seen a steel welding gas bottle of H2 gas leak down by 50% in just a few weeks. Among other issues, that is why most LH2 tanks down here on the ground are made of stainless steel. And they still leak.

GW

GW Johnson · 2016-04-23 09:15:49

VentureStar was a concept aimed at lowering cost of space access by means of some degree of reusability. Same as was shuttle before it. Expecting the first examples to fully succeed when you do something new is not reasonable, as we learned with shuttle. People learn by doing, again and again and again....

To my way of thinking, the space access problem divides quite naturally into systems that get you to LEO, and systems that go anywhere else from LEO. The requirements for those two flight regimes are so drastically different that it seems nonsensical to try to design one item to do both, at least at this time in history.

We should have learned by now not to spend money on major projects that aim to build one-size-fits-all hardware for incompatible design requirements: example 1 TFX (later F-111) circa 1970, example 2 F-35 today. That same lesson applies in all areas of human endeavor, including spaceflight.

There is one exception to what I just said: the smaller one-way payload to other bodies in the solar system. That is where direct launch to transfer trajectory makes sense, as long as you have a rocket big enough to send the payload on its way. Which in turn is why such payloads have such severe weight and volume restrictions on their designs. Otherwise, you just dock together tinkertoys in LEO to send whatever you want wherever you want. Doing that is cheaper than gigantic rocket development.

Now to reusability: we have learned that we can build airframes capable of flying into LEO and back that have significant reusability. There's more than one way to do this, but the shuttle heat protection devices were the first to fly. It's stuff like that which can protect leeward surfaces, and windward surfaces that are not stagnation zones or shock-impingement zones. We still require ablatives for stagnation zones, and you must avoid shock-impingement zones at all costs. That last prevents parallel placement of nacelles for entry.

The picture is not so clear with respect to rocket engines, nearly all of which are turbopumped. While thrust chambers, injector plates, and exit bells can all be made to fire lots of times, the turbopumps cannot. At this time in history we cannot build long life turbopump assemblies for rocket engines.

Much was made of tile replacement troubles on shuttle; not much was ever said about replacing turbopump assemblies on the engines, but they did. And guess what is still the most expensive part of the rocket engine!

You have 3 choices to solve the poorly-reusable turbopump problem: (1) pressure fed with concomittantly-heavy tankage, (2) use solids with low impulse and heavy cases, or (3) find some other way to pump your propellants. XCOR Aerospace has pioneered a solution to that third choice: their piston-pumped engines.

They have finally demonstrated heat-engine driven piston pumping using a third fluid as the engine coolant/source of heat for their pumping rig. But, they have also demonstrated that this technology solves more problems than it creates only in smaller sizes of engines.

That being the case, I can forsee small spaceplanes being practical with fully-reusable engines, but not large ones like VentureStar. I also forsee definite limitations to the cost reductions Spacex might achieve by recovering first stage boosters. The engines will need new turbopump assemblies every flight, or at best every few flights.

What I forsee is a future with partly-reusable launch rockets not too different from what we are beginning to see today, augmented with small spaceplanes that ride the same rockets as boosters. Big items get launched on the rockets, anything big coming home does so in a capsule. Very small items and a small crews can ride the spaceplane, or the returnable capsules.

The vehicles that go elsewhere from LEO will look very unlike anything flying today, and resemble more some of the old science fiction, in that habitats will resemble the old Skylab, except likely made of inflatables. These modules plus propellant and supply modules, plus some engines, are your means to send crews beyond the moon. These will get built to fly many missions, not just one. Cost will drive that.

GW

Last edited by GW Johnson (2016-04-23 09:24:52)

SpaceNut · 2016-04-23 19:54:23

Its the new players in the game that are re-writing the rules for the future but only so long as there is continued lowering of cost, increasing of reliability and a gradual but steadiliy increasing capability of large mass to orbits... but there is more to this as the old guards are showing signs that they can still adapt and as they go through the growing pains of learning the new rules...as cost plus will be no more....

kbd512 · 2016-04-28 21:43:46

I think robots are the answer to the issues with the engines. A robot needs to fabricate the turbo pumps in a time and materials efficient manner. Forget about reusability of all engine components. Can we design turbo pumps that are easy to manufacture, simple to assemble, and fast to replace? Maybe we need other robots to assemble the engine components.

If I have a robot that builds composite tanks, a robot that fabricates engine components, a robot that inspects the components, a robot that can assemble the fabricated components, then I should be able to get away with the inefficiency of trashing some of the components after each flight.

We really need an all-in-one assembly plant for fabrication of commodity rocket components and engine components for reusable spacecraft. The plant must be run more like an automotive assembly line rather than a clean room.

GW Johnson · 2016-04-29 13:05:49

For the smaller-thrust applications, I think XCOR's piston-pumped technology is the way to go. Those guys advertise TBO's that look like state-of-the-art piston and turbine engines in the civil aircraft industry. There is no way a turbo-pumped rocket will ever approach TBO's of a few thousand hours/few thousand ignitions like that.

For the larger thrust sizes, I think kbd5212's notion of easily-replaced mass-produced turbo-pump assemblies, assisted by robotics, is likely the right answer. I don't know much about robotics, but I recognize and understand designs that easily worked-upon, versus those that are nightmares.

I do tend to cuss the engineers who design servicing nightmares. It is quite clear they never got their fingernails dirty actually physically building or repairing anything. That is a critical and inexcusable lack in an engineer's education, says this engineer, who has practiced for over 4 decades, in several fields ranging from aerospace defense weapons to farm implements.

GW

Last edited by GW Johnson (2016-04-29 13:09:51)

Antius · 2016-04-30 09:32:42

Maybe the trick is to produce a simplified turbopump, with fewer parts. Tank weight can be reduced by a factor of three or more compared to pressure fed concepts, so there is plenty of mass margin to play with. A self-contained electrically powered pump would be excellent, as there would be no need for complicated gas generators and turbines, just lithium batteries. That way you can flange the pumps into the propellant lines and replace them rapidly after each flight. For upper stages, where combustion chamber pressures can be lower, a jet pump might suffice.

Antius · 2016-04-30 11:59:34

From memory, a liquid fuelled engine injector needs to inject propellant at a pressure 5bar greater than chamber pressure in order to avoid burnout from acceleration induced combustion instability. If the tanks are pressurised to 10bar and lower stage chamber pressure is 20bar, then the pump must provide a pressure head of 15 bar. This is just about achievable using a single-stage centrifugal pump. If fuel is kerosene or propane, then the pump can be mineral oil lubricated, alloy steel and essentially bought off the shelf. The same is true for the oxidiser pump if nitrous oxide is used, as it is stable in the presence of hydrocarbons until temperatures are high enough to begin dissociation.

Using Sea Dragon as an example: The first stage would have burned about 10,000 tonnes of propellant in about 90 seconds. Total pump head would need to be something like 17bar when friction in pipes is accounted for. Thats a head of 170m and a total work requirement of 1.7KJ/Kg propellant, so 17GJ of pump work overall. If pump efficiency is taken to be 80%, then we need about 20GJ of input electrical energy. Pump input power would be 220MW. To power the pump using lithium ion batteries with a discharge rate of 340W/kg, would require about 650 tonnes of batteries. That's about 5% of Sea Dragon lower stage weight, so it would appear to be doable.

A diesel generator or open cycle gas turbine would offer a much better power to weight ratio and waste heat could be used to maintain tank pressurisation. Not sure how they would work out cost wise in the power range needed, but they would be reusable and should be commercially available equipment.

Last edited by Antius (2016-04-30 12:55:39)

SpaceNut · 2016-05-01 08:33:48

Antius IF a simplified turbopump, with fewer parts reduces mass and costs I am all for it. Also if we are using a liquid fuel such as kerosene or RP-1 the electrical pump can be put into the tank just like all cars are as it cools the pump while in operation.
I see venturestar as a simplified shuttle launch system that narrows up the scope of operation and use to just a single thing, transport of crews and no cargo just life support that is short term....

GW Johnson · 2016-05-02 13:28:41

VentureStar was to be a reusable spaceplane. They hoped they had answers to the between-flight refurbishment costs to help lower its per-launch cost. It was a very large vehicle.

The problem with spaceplane versus expendable launcher is the deliverable payload as a fraction fraction of total launch weight (including any and all boosters). It is inherently a far smaller fraction with the spaceplane, because you must deliver payload+return vehicle to orbit. While, with the expendable, you essentially just deliver payload to orbit. So, the unit cost of the delivered payload will always be much higher with the spaceplane.

Because of that, I don't really believe there is a real niche for giant spaceplanes delivering mass cargo. They will always be way too expensive to use in that way. At least as long as we are restricted to chemical propulsion.

But, for small items, or small crews, unit cost doesn't drive you, some other necessity does, so it is just the gross launch price you worry about. Looked at that way, smaller is inherently cheaper.

If the craft can really be turned around fast between launches, then a small spaceplane is likely the most desirable answer for smallish items. That's why I think the Sierra Nevada Dreamchaser is a far more practical design than the Lockheed VentureStar ever could have been. Somewhere between X-20 DynaSoar and Dreamchaser is just about the right size range for spaceplanes, in my opinion.

GW

RobertDyck · 2016-05-02 14:03:23

One Russian company proposed an interesting alternative. I've mentioned it before. When Russian politicians wanted a copy of Shuttle, the company Molniya proposed an intermediate size shuttle called "Multipurpose Aerospace System (MAKS)". Launched from the back of the largest cargo plane ever flown: Antonov AN-225. That meant it could launch from any airport that could service that aircraft. A cockpit for 2 crew, plus cargo hold that could carry 7 metric tonnes to ISS. And they designed a passenger module that would fill the cargo hold, with 4 more crew. Engines would use RP1/LOX with a tiny bit of LH2 during initial launch, then the same engines would shift to LH2/LOX for the final push to orbit. External tank was expended.

Antius · 2016-05-02 17:14:17

An expendable upper stage could be made very cheaply. Alloy steel tanks, compressed gas reaction manoeuvring, low chamber pressure ablatively lined pressure fed engines. It is also the lightest stage and therefore least valuable.

Trying to recover a stage at orbital velocity makes as much sense as combing Afghanistan for spent bullets. Like the bullet, the upper stage is impracticable to recover, so we make it as cheap and technologically simple as possible. No one tries to build techology into a bullet. We bring down costs through economy of scale and we don't bother recovering them for the same reason. That doesn't mean we shouldn't bother trying to reuse an upper stage at all. But why not take advantage of the huge energy investment placed in the stage and recycle it in orbit?

The lower stages are a different story. They are both easier to recover and larger and therefore more valuable. It makes more sense to recover them and reuse them in the same way. And the easiest way to do it is with drag skirt and ocean recovery. There are other drivers of economy. Economy of scale applies to the size of the rocket as engineering and launch costs are spread across larger payloads. If we can use electrical feedpumps in low chamber pressure lower stages, then tankage can be made from standard range alloy steels rather than aluminium or maraging alloys. That brings down cost massively, as even very large marine steel structures can be cheap (just look at oil tankers). In addition, steel has a much more forgiving stress cycle than aluminium.

At what point does reusability no longer make sense? The point at which you have to add large amounts of weight and/or technological sophistication and cost to achieve it. For an intermediate stage, this is probably the point at which simple drag skirts cease to be sufficient to allow recovering and we need more complex re-entry shields. This suggests to me that optimum number of stages to reach orbit is probably more than two. If the booster stage has dV of 3km/s then it doesn't make sense trying to make a cheap expendable upper stage with dV of 6km/s if we can build a reusable intermediate stage. The question then is, at what sort of velocity does a Sea Dragon style drag skirt cease to be a workable solution? For a three stage vehicle, this will determine the dV that the 3rd stage must be designed to provide, since the 2nd stage should be designed for recovery and the 3rd should not.

Playing with the rocket equation and working out mass ratios for pressure fed stages, leads me to believe that dense low vapour pressure propellants like propane and nitrous oxide are most compatible with alloy steel stages and the dV of each stage should be about the same as its average exhaust velocity. Hence, using sea storable and self-pressurising propellants like propane-n2o, we would want three stages each providing about 3km/s of dV. The lower two should be reusable, the upper stage initially expandable, though ultimately it could be recycled in orbit. The upper stage could be further simplified by keeping payload faring on the 2nd stage, as stage 3 is deployed in vacuum. Lower stage chamber pressure should be about 20bar, intermediate stage the same or less, upper stage should be lower, probably no more than 5-7bar. Lower stage engine chamber and throat should be cooled with ablative liquids. Upper stage should solid ablative and refractory linings. The lower stages should be built in shipyard type facilities. The upper stage, with its simple mechanical components is compatible with 3d printing and car factory style manufacturing.

Last edited by Antius (2016-05-02 18:08:32)

SpaceNut · 2016-05-02 18:05:21

There is no reason to have a hug monster venturestar when we can follow the sppace x rocket modeling for how to judge the mass restriction to its shape....design for seperate cargo and crew versions to make it a better buy for how its to be used.

I think the Sierra Nevada Dreamchaser mounted on a reuseable Falcon 9 is the best way to go forward for a near complete resuseable here and now as dreamchaser would need a smal boosting stage to finish getting it to orbit........

RobertDyck · 2016-05-02 20:32:17

SpaceNut wrote:

I think the Sierra Nevada Dreamchaser mounted on a reuseable Falcon 9 is the best way to go forward for a near complete resuseable here and now as dreamchaser would need a smal boosting stage to finish getting it to orbit........

DreamChaser is too big for Falcon 9, and too small for Falcon Heavy. But current plans are to launch unmanned with a non-reusable cargo pod on Atlas V. That's why I suggested DreamChaser with a larger cargo pod on Falcon Heavy.

::Edit:: Actually, Falcon 9 Full Thrust might be able to do it. Accurate dry mass figures for Dream Chaser are hard to get. If they transport crew and no cargo, remove the cargo pod, construct the lifting body with integrated wings (no hinge), and launch without fairing. That relies on keeping launch mass of Dream Chaser down, and lift mass of Falcon 9 up. I couldn't land at the launch site, it would require the barge. Getting aerodynamic data for launch without a fairing is key.

Model preparation with multiple pressure transducers: Transonic Dynamics Tunnel at NASA Langley.

Last edited by RobertDyck (2016-05-02 21:50:35)

Tom Kalbfus · 2016-05-03 10:00:46

Presumably it would be used to exchanges crews for the space station. I wonder how many of the orbiters they plan to make? The Dreamchaser appears small, it may be cheaper to build than the shuttle. I wonder how many uses we would expect to get out of each one. I think the production line should be kept open, instead of limiting it to a fleet of four. Since it is small in comparison to the launcher, as opposed to the shuttle, there might not be that much cost savings in using it over and over and over again ad infinitum, in some cases it might prove cheaper just to build another copy than to reuse it too many times, until perhaps it fails and gets someone killed! for example. Perhaps the newer ones would be used to transport astronauts and the older ones for cargo that is easily replaced.

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