Reusable Rockets to Orbit

louis · 2012-01-07 21:11:09

With non-reusability, and a figure of $5000 per kg to LEO I think we are "good to go" for Mars. So I am not too hung up on it, although obviously it's desirable if the economics work out. I've never liked sea recovery. Seems so intrinsically costly and problematic. But if we are being creative...maybe a giant inflatable out at sea would help...so the returning rocket stage lands on the huge inflatable. The inflatable could be towed into position by fast motor boats launched from an ocean going ship...

JoshNH4H · 2012-01-08 01:00:16

Well I for one don't want to lose Hop's input. I've been flamed by him - and lived to tell the tale. [ And I'm still awaiting his answer on why with more meteorites we have higher meteorite prices on Earth...] Best to ignore the flaming or respond in kind - banning would diminish the site.

Just a quick reply, because I'm about to go to bed: It is incredibly difficult to get yourself banned from newmars. I can only think of one non-spammer user who has ever received a ban, and that was Zydar, after violating the "No Intelligent Alien Life Conspiracies" rule something like 10 times after having been repeatedly informed that he was not allowed to do so, including several post deletions. Even then his ban was not permanent. Had mod powers been meted out, I would either go in to the offending post and delete the offending content, or report it to a moderator and ask them to deal with it.

Nevertheless, flames are not an acceptable way to interact with other members of the forum. They degrade the level of discussion and turn threads into contests of wit instead of collaborative efforts. It is no great leap of imagination to figure which is best for the forum, most productive, and most enjoyable to read and participate in.

GW Johnson · 2012-01-08 13:10:42

Now, now, guys! Civility!

I just posted on another thread some of the cost figures per pound of payload to LEO from Spacex's page a few months ago. We can do a lot better than $5000/kg I saw posted here a conversation or two back. That corresponds roughly to $2400-2500/pound. That's Atlas-5 at max capability 20-25 metric tons to LEO, and the same for Falcon-9 at 10 metric tons to LEO.

Spacex's projections just a couple of months ago for Falcon-Heavy are 53 metric tons to LEO, which works out to $800-1000/pound (roughly $1600-2000/kg). I doubt very seriously the new government design could ever even possibly approach that cost figure, in spite of being 100+ tons, and there is a scale effect. It's shuttle derived hardware and ways of operating, which derive from shuttle at $1.5 billion per launch of 25 tons max. Not carrying an orbiter will help, but not all that much. You work it out. Ridiculously expensive.

I'm for using Falcon-Heavy, not screwing around with some ridiculously expensive government design.

GW

JoshNH4H · 2012-01-08 16:57:03

Firstly, on the cost of SpaceX rockets: Just check out their website. For all the technical information on stage mass etc. that SpaceX doesn't give, they are very clear about how much one of their rockets costs. As of right now, you can buy a Falcon 1e (1,010 kg to LEO) for $10.9 million, a Falcon 9 (10,450 kg to LEO) for $54 million to $59.5 million, and a Falcon Heavy (53,000 kg to LEO) for $80 million-$125 million. That's a price to LEO of $10,792/kg, $5,167-$5,694/kg, and $1,509-$2358/kg, respectively.

My goal here is to talk about designs that can enable costs to LEO below $100/kg. Now, that's still not cheap; It would still cost multiple tens of thousands of dollars for a ticket to orbit. However, given the prices as they currently exist it is very cheap by comparison.

I don't know where government designed rockets are coming into this, but I have no doubt that the currently planned heavy lift rockets will fail to lower costs significantly in much the same way as past rockets have.

Terraformer · 2012-01-09 09:59:30

Had mod powers been meted out, I would either go in to the offending post and delete the offending content, or report it to a moderator and ask them to deal with it.

Ahem. The latter course of action is the best way, because you're not impartial...

Hop does have a point - you can't ignore the trajectory when designing a rocket, since if you're using one set of rocket engines to provide both horizontal and vertical thrust, you're going to have to vector them, which means your rocket is going to be at an angle, which determines the mass distribution you can have...

GW Johnson · 2012-01-09 12:37:29

Try graphing Spacex's costs as $/kg payload to LEO vs kg payload to LEO. You get a decreasing curve of costs, not linear. Then spot Atlas-5 20-25 tons at about same per-launch cost as Falcon-9. Bigger tends to be lower cost but Atlas-5 falls well above the Spacex curve. This in spite of Atlas flying in one form or another since 1956. That will illustrate best what I have been saying about the importance of small logistical tail and a not-gigantic (bloated) company. If Spacex were to close up shop today, Atlas's prices would at least double tomorrow.

GW

JoshNH4H · 2012-01-09 16:38:38

Terraformer wrote:

Had mod powers been meted out, I would either go in to the offending post and delete the offending content, or report it to a moderator and ask them to deal with it.
Ahem. The latter course of action is the best way, because you're not impartial...
Hop does have a point - you can't ignore the trajectory when designing a rocket, since if you're using one set of rocket engines to provide both horizontal and vertical thrust, you're going to have to vector them, which means your rocket is going to be at an angle, which determines the mass distribution you can have...

I was on the fence. Seeing that nobody has mod powers anyway at this point with the exception of joshcryer and jburk, it's not as if I have any choice. Nevertheless, I am capable of identifying a flame, regardless of whom it is against.

I was never claiming that it has no effect. Rather, my claim was that the variance in trajectory is not that significant if you design your rocket to have sufficient thrust and therefore its effect on the basic design choices is not significant.

I don't quite understand what you're talking about in terms of mass distribution, though... Is it possible that you have fallen prey to the Pendulum Rocket Fallacy?

louis · 2012-01-09 18:03:22

JoshNH4H wrote:

Firstly, on the cost of SpaceX rockets: Just check out their website. For all the technical information on stage mass etc. that SpaceX doesn't give, they are very clear about how much one of their rockets costs. As of right now, you can buy a Falcon 1e (1,010 kg to LEO) for $10.9 million, a Falcon 9 (10,450 kg to LEO) for $54 million to $59.5 million, and a Falcon Heavy (53,000 kg to LEO) for $80 million-$125 million. That's a price to LEO of $10,792/kg, $5,167-$5,694/kg, and $1,509-$2358/kg, respectively.
My goal here is to talk about designs that can enable costs to LEO below $100/kg. Now, that's still not cheap; It would still cost multiple tens of thousands of dollars for a ticket to orbit. However, given the prices as they currently exist it is very cheap by comparison.
I don't know where government designed rockets are coming into this, but I have no doubt that the currently planned heavy lift rockets will fail to lower costs significantly in much the same way as past rockets have.

I do think that $100 per kg is actually very cheap, unbelievably cheap, given what you are doing...you wouldn't expect to get to the top of Mount Everest for $10 would you?

My own view is that if you get down to $1000-2000 per kg with reliable rockets, a whole new world opens up - we just don't realise it yet because investors are conservative when it comes to billion dollar investments! But eventually someone is going to get the formula right.

GW Johnson · 2012-01-09 21:09:45

I did exactly what I suggested in my previous posting, and created the launch costs plot. I did it for the 3 Spacex Falcon birds, 3 of the Atlas-5 family, and for Delta-4 heavy. I plotted the data in metric units as $/kg vs metric tons delivered to LEO from Canaveral. I also replotted in US customary, for folks who know those units better: $/lb vs US tons delivered to LEO from Canaveral. Those data are public view over at http://exrocketman.blogspot.com

GW

Hop · 2012-01-10 07:55:18

JoshNH4H wrote:

It is my understanding that several off-the-shelf rocket stages have the mass ratio to function as SSTO launch vehicles. At the very least I can point to the Centaur upper stage as having this theoretical capability, seeing as I've actually done the calculations for it, though this is saying nothing of thrusting capabilities).

Doing my own calculations...

Centaur V1
Gross mass: 22825 kg
Empty mass: 2026 kg
ISP: 451 seconds

Given that ISP and mass ratio, you have a delta V of 11.15 km/s. More than enough to achieve LEO, right?

But let's look at thrust to weight ratio.
Thrust: 99190 newtons.
The initial thrust to mass ratio is 99190 newtons/22825 kg. Or 4.35 newtons per kilogram. Gravity exerts a force of 9.8 newtons per kilogram.

T/W < 1.

This stage would not even get off the ground.

JoshNH4H · 2012-01-10 08:52:36

My response to this is quite simple: I was talking about Mass Ratio and nothing else. If you take a look at my post (the exact same text which you quoted, no less!) I specifically stated that whether the rockets actually had the thrust to make it to orbit was a different topic entirely. To quote the relevant text (it's the last third or so of the last sentence in the passage which you quoted) "though this is saying nothing of thrusting capabilities."

Hop · 2012-01-10 10:41:42

JoshNH4H wrote:

I was talking about Mass Ratio and nothing else.

You were talking about mass ratio to function as a SSTO vehicle.

If your T/W is low, you will incur excessive gravity loss and your needed delta V will be higher than 9.4 km/s.

So no, the Centaur IV does not have the mass ratio to function as a SSTO vehicle.

GW Johnson · 2012-01-10 11:03:41

The mass ratio and the thrust/weight has to be there for surface launch. Adding engines usually drives the mass ratio down, unless you scale up the tankage a bit. That's why most first stages are so large.

First stage engines are usually of different design than upper stage engines: shorter bells. It really needs to operate perfectly expanded at launch level (usually sea level). She'll be underexpanded as you climb, which costs performance, but then so does over-expansion, and it's far worse. Usually first stages leave the sensible atmosphere, so it's in vacuum by burnout. Upper stage engines can be designed to be optimal in vacuum-only, with very long bells.

So, it's not exactly the same engines. Even for the same propellants, first stage Isp is a lot lower than "typical" vacuum Isp designs. Be careful trying to scale from one application to another.

GW

Last edited by GW Johnson (2012-01-10 11:04:56)

RGClark · 2012-01-10 15:27:10

GW Johnson wrote:

To answer a question Josh asked earlier, I spent some time at what was then LTV Aerospace in Dallas working on the old "Scout" launcher. I used a combination of jiggered rocket equation stuff and motor manufacturer catalogue data to set up the real trajectory code stuff. The "gold standard" was (of course) the trajectory code. My job was to determine feasible advanced configurations for "Scout", and feasibility of some really unusual missions for it to do. "Scout" was a 4 stage solid propellant vehicle. They lost 1 of 4 in flight test, then never another one in 30-some years.
For Bob Clark: airbreather thrust, particularly ramjet, is very strongly (dominantly) dependent upon flight speed and altitude air density. The nozzle thrust is calculated same way as a rocket (chamber total pressure, gas properties, pressure ratio across the nozzle, and nozzle geometry), the pressure is just lower and the expansion ratio a lot less. You do need to worry about the difference between static and total chamber pressure, unlike most rockets.
The ram drag is the drag of decelerating the ingested stream of air into the vehicle. Its massflow multiplied by its freestream velocity (in appropriate units of measure) is the way that is done. But, nozzle force minus ram drag is only "net jet" thrust. There are several more propulsion-related drag items to account.
There is spillage drag for subcritical inlet operation (which also means reduced inlet massflow!), additive or pre-entry drag for ingested stream tubes in contact with the vehicle forebody, and the drag of boundary layer diverters or bleed slots, quite common with supersonic inlets. None of those are simple to calculate "from scratch" (we use wind tunnel test data to correlate empirically a coefficient for each as a function of Mach and vehicle attitude angles), and taken together they are often quite a significant force.
If you subtract that sum of drags from net jet thrust, you have the "local" or "installed" thrust, corresponding with just plain airframe drag. Most airframers work in that definition. If you don't, then you have to add that sum of propulsive drags to the airframe drag to get the corresponding proper drag for "net jet" thrust-drag accounting (not very popular outside the propulsion community).

Thanks for the detailed response. That's actually a little too much detail for what I need. I read your post on ramjet boosters:

Sunday, August 22, 2010
Two Ramjet Aircraft Booster Studies
http://exrocketman.blogspot.com/2010/08 … e-boe.html

I noted that you were able to get better payload with more shallow launch angle but it created a problem for retrieving the first stage booster, since it went so far downrange. If I'm reading it correctly you were able to double the payload mass with the shallow angle, presumably using aerodynamic lift.
What I'm trying to determine if I can increase my payload just going to the range turbojets can get to, ca. Mach 3+. I intend to use the jets to get to medium altitude for a turbojet, ca. 15,000 m. But I need to get to a good angle as well as reaching its max speed. Another problem is that I don't know if it can get to max speed while climbing.
I looked at the case of the SR-71 and the XB-70 Valkyrie. These had more thrust than I wanted but that added weight because jet engines are so heavy. In any case I noted the climbing rate. From that it seemed doable, considering the high effective Isp, that you could reduce propellant mass that way. The problem is this is for a SSTO application and I can't afford the weight. What I wanted was the engines to put out in the range of 1/7th the vehicle weight to reduce the jet engine mass. What I don't know is how will that effect the climb rate, and will it even be able to reach supersonic now.
Note that an advantage of the SSTO is that you can get the better payload by flying a shallow angle and not have to worry about recovering the booster stage.

Bob Clark

Last edited by RGClark (2012-01-11 00:57:59)

RGClark · 2012-01-11 01:12:39

Hop wrote:

JoshNH4H wrote:
It is my understanding that several off-the-shelf rocket stages have the mass ratio to function as SSTO launch vehicles. At the very least I can point to the Centaur upper stage as having this theoretical capability, seeing as I've actually done the calculations for it, though this is saying nothing of thrusting capabilities).
Doing my own calculations...
Centaur V1
Gross mass: 22825 kg
Empty mass: 2026 kg
ISP: 451 seconds
Given that ISP and mass ratio, you have a delta V of 11.15 km/s. More than enough to achieve LEO, right?
But let's look at thrust to weight ratio.
Thrust: 99190 newtons.
The initial thrust to mass ratio is 99190 newtons/22825 kg. Or 4.35 newtons per kilogram. Gravity exerts a force of 9.8 newtons per kilogram.
T/W < 1.
This stage would not even get off the ground.

What would be the mass ratio and delta-V if you added engines to achieve lift off?
BTW, Dr. John Schilling has an online payload estimator that allows you to estimate the
the payload you can lift to orbit with a launch vehicle. You enter in the propellant and dry
mass and the vacuum thrust and vacuum Isp. The program does give a warning if your
thrust is less than your launch weight.

Launch Vehicle Performance Calculator.
http://www.silverbirdastronautics.com/LVperform.html

Bob Clark

GW Johnson · 2012-01-11 15:07:30

Bob Clark:

What I had posted for the two-stage airplane was not the final form, just what I had done at that time, and where it pointed. I tried to climb really high before pulling over and accelerating the ramjet first stage airplane, and the air was too thin for thrust-minus-drag to accelerate the mass at a practical rate. That's why the first stage went too far downrange to fly back.

Numbers down closer to 60,000 feet for the staging altitude look better. I need to re-run that same study with 60,000 feet staging, and see if the booster flyback becomes practical. I simply haven't done it yet. But I'm pretty sure it would work.

From what I'm told, there are 3 variables of importance to selecting the staging of a HTOL 2-stage vehicle. They are, in order of importance, speed, path angle, and altitude. The most important is speed. That's why I picked ramjet: with external or mixed-compression inlets, it's capable of useful thrusts to M5.5 to 6, and can take over as low as M1.6-ish. I went with separate rocket and ramjet engines that can be burned in parallel, that's how I achieve about 45-degree path angle at staging in a sudden pull-up transient without deceleration (combined cycle probably won't be able to do that).

I just couldn't make it work right at 100,000 feet. Frontal thrust densities are an order of magnitude higher at 60,000 feet. Altitude is the least important of the three variables, so I feel pretty good about the HTOL 2-stage approach with a rocket+ramjet airplane 1st stage. The 2nd stage can be a rocket ballistic pod, or a rocket airplane, whatever the mission needs.

Configuration design for drag reduction and for impinging-shock avoidance are the truly critical issues. Shock impingement heating can cut through structures in a second or two at M6+. We saw that on the X-15 flight that carried the scramjet test article.

GW

RGClark · 2012-01-12 14:45:32

Elon Musk has said he wants to cut the costs to space to the $100 to $200 per kg range by reusability. This is about a two order of magnitude reduction in cost. To put this in perspective, this is like a trans-atlantic flight that costs $1,000 suddenly being cut to cost $10 to $20.
Musk has said this transformation of the Falcon 9 to full reusability will be very hard. I don't believe it will be. But first, keep in mind how important that reduction in cost will be if it succeeds. If it succeeds then SpaceX will monopolize the launch business if the other launch companies do not field their own reusable vehicles. So there is a tremendous financial incentive for SpaceX to invest in reusability. Now, most in the industry believe reusability is very difficult for orbital vehicles and not even worth the expense. So if Musk reinforces that idea then he has a better chance at being able to field one without the other launch providers having one. And since they will not have even started to develop one, it will take them some time to catch up. The effect is that Musk will have a monopoly on all launches for at least a few years.
I don't know if that is Musk's intent in saying reusability is very hard. Actually I'm inclined to believe he is just saying what most in the industry believe including his own engineers. But a key reason why reusability is not very hard is because the cost in mass in reentry and landing systems is surprisingly low. In regards to the technical difficulty, there is none. We know how to do it as the shuttle orbiter and the X-37B and Dragon spacecraft has shown. I include the Dragon in the list of reusables because its heat shield showed minimal degradation on return. Musk has said the same heat shield could make hundreds of flights, at least to LEO.
I made an estimate before of about 28% of the landed mass has to go to reentry/landing systems. This was based on estimates of 15% for thermal protection, 10% for wings or for propellant for vertical landing, and 3% for landing gear. However, I said likely with modern materials this could be cut to half that. In fact, it might even be lower than 10%.

1.)Weight of thermal protection.

Robert Zubrin has given an estimate of 15% of the landed weight for the weight of thermal protection systems(TPS):

Reusable launch system.
http://en.wikipedia.org/wiki/Reusable_l … at_shields

However, I gather this was in relation to the older capsules, Mercury, Gemini, Apollo, etc. Indeed the weight of the ablative heat shield on the Apollo capsule was about 15%:

Apollo Command/Service Module.
2.7 Specifications
http://en.wikipedia.org/wiki/Apollo_Com … ifications

However, the space shuttle with its mostly silica tiles was able to reduce the TPS weight to about 8% of the maximum landing weight of 104,000 kg:

Space Shuttle thermal protection system.
3.3 Weight considerations.
http://en.wikipedia.org/wiki/Space_shut … iderations

Also, for the X-37B the TUFROC leading edge material instead of the shuttles RCC and the TUFI AETB material instead of the shuttles silica tiles are either of equal or lower weight than the shuttles TPS materials while being tougher and requiring less maintenance:

X-37B Orbital Test Vehicle.
http://www.boeing.com/defense-space/ic/ … b_otv.html

For ablative TPS, the PICA-X material used on the Dragon capsule weights about half the weight of the AVCOAT material used on the Apollo heat shield:

Re: Dragon v/s Orion.
http://forum.nasaspaceflight.com/index. … #msg754168

while being able to still survive lunar and even Martian reentry speeds.

SpaceX has found that at least for LEO reentry speed judging from the minimal degradation on the Falcon 9/Dragon test flight, the PICA-X heat shield could be reused hundreds of times.

Also, for vertical powered landings a la the DC-X, you might not even need an extra heat shield for base first landings. One proposal for a VTVL SSTO uses low thrust during the descent as well as a high temperature-resistant aerospike nozzle to serve as the reentry thermal protection. You would need to retain more mass in propellant or some inert gas for this purpose though.
Another idea for a vertical landing vehicle would be to reenter head first. This was the preferred method of the Air Force since it provided increased cross-range. In that case you would have the blunt heat shield at the top of each stage. I thought this method would be unstable with the heavy engines now at the top during reentry, but since this was considered for the orbital version of the DC-X presumably this was solved.

2.)Weight of the wings and the landing gear.

For horizontal landing, a common estimate is that the weight of wings is 10% of the landed weight. This comes from aircraft examples though where the wings have to carry the weight of the fuel which can be as much as the dry weight of the aircraft itself or more.
An example where the propellant will not be carried in the wings and lightweight composites will be used is the Skylon. According to their released specifications the wing weight will be less than 2% of the take-off weight, which is the appropriate weight to compare to for a horizontal take-off vehicle:

The SKYLON Spaceplane.
by Richard Varvill and Alan Bond
Journal of the British Interplanetary Society, Vol. 57. pp. 22–32, 2004
p. 32.
http://www.reactionengines.co.uk/downlo … _22-32.pdf

On that same page the landing gear weight is the only 1.5% of the take-off weight.
Then for a vertical take-off vehicle these low weight proportions should apply to the dry, landing weight.

Bob Clark

JoshNH4H · 2012-01-14 21:16:28

Hop wrote:

JoshNH4H wrote:
I was talking about Mass Ratio and nothing else.
You were talking about mass ratio to function as a SSTO vehicle.
If your T/W is low, you will incur excessive gravity loss and your needed delta V will be higher than 9.4 km/s.
So no, the Centaur IV does not have the mass ratio to function as a SSTO vehicle.

I believe my meaning was excessively clear and not arguable, given that you accept the single premise on which it is based.

RGClark and GW Johnson-

It looks like we're approaching some kind of consensus in that, for current state-of-the-art capsules (and the space shuttle), 8% of the mass being re-entered is heat shielding. Could we get that down to 5%, with more advanced technology like inflatables and new materials? Especially given the lower mass per area involved in sending down an empty tank, and therefore heating as well as structural stresses, it seems possible.

Wrt wings, I'm not sure how much these wings would really be doing in terms of lifting vs. acting as a triangular heat shield. If your wings aren't a lifting body, do you lose the ability to navigate to a specific location? In any case, 2% of the mass to be landed isn't too bad, though I am thinking about a vertical landing if possible in order to minimize stresses as well as landing gear mass.

I don't know if I ever actually said it, but I think that it's logical for this kind of rocket to have an aerospike engine instead of the traditional bell nozzle. It just seems to makes more sense given that the engine will have to operate all the way down from ground level up to space. I suppose the other option is to continue what seems to have been a love affair with inflatables for this reusable rocket and posit an inflatable extension to the nozzle; it could be insulated by running temporarily heavy on Methane to get it covered in Carbon soot. Thoughts?

GW Johnson · 2012-01-15 11:21:54

I may be the victim of past thinking, but I think the main decision to make is whether you want lifting flight or ballistic flight post re-entry. Putting the weight off-center a bit in a traditional capsule gets you side forces of some significant fraction of the drag force during re-entry hypersonics. This allows you to shift the location of the landing ellipse by a couple of ellipse dimensions L-R, or downrange-uprange. But it's worthless as a directional control once the hypersonics are over. From that point, you're ballistic. Control there requires real wings.

Wings are pretty useless during re-entry because they're fragile and vulnerable to overheat, as in the shuttle design, the X-20 we abandoned, and the X-15 which was considered for orbital flight using a Titan-3 booster, but which plan was abandoned because it could not survive re-entry. In the X-15's case, this was because of wings burning away really fast at nose-first attitudes, but inadequate structural strength to re-enter more nose-high ("semi-broadside") like the shuttle. The X-20 was to have addressed that weakness with experimental ablatives, and/or transpiration cooling, and/or heat-sinking. Shuttle did it with slow ablatives (carbon-carbon LE's)and refractories (low density ceramics).

There are a couple of things we haven't tried yet: some sort of ablative-protected inflatable structures, and protection by an engine plume between vehicle and hypersonic shock layer. There is great potential in both ideas, but I see nothing at all going on around me. Not a space agency anywhere is exploring this, and private concerns “typically” do not invest their own money in exploring radically new technologies (that’s why there have to be government agencies, no one else will do that job).

The aerospike nozzle is indeed what you might want for engines of any type that must function across varying backpressure going up. Coming down, I fear the sharp aerospike might be rather vulnerable, if you used the engine plume as your heat shield. But I don't know. No one does, not yet. Aerospike would have a lot of application to first stages going up, except our current methods for building them end up being heavier and inconveniently-packaged for a cylindrical shape, and the nozzle efficiencies typically run 2 or more % lower than a traditional bell. It only looks better at off-design backpressure. It's a non-trivial trade-off, and unfamiliar territory, so no one has really done it. X-33 was supposed to do it, but never flew for a whole plethora of reasons, some technical, some not.

Recovering a upper rocket stage in its entirety would probably be best done with a blunt ablative on the front end, and some ablative-protected inflatables around the engine(s) to double as floats, and as water-impact shields for the engine bells. Take the main sea impact on the heat shield, we already know that works fine. You just need to add a series of drogues and chutes that deploy out the rear, around the engine(s). Biggest risk I see is shroud line fouling on the engine(s). But it's doable.

That's a lot of extra gear and equipment. Plus, the tankage has to take the "whack" when it hits the sea at dozens to a hundred gees. You're just not going to build a thing like that at 5-8% inert fractions. Steel, aluminum, and titanium are "only so strong". I really don't see organic composites as a tankage option for something intended to survive Mach 25-class re-entry, either. It's just too thermally harsh. Something that could be re-flown a lot of times is just going to be heavier. 10+% inert, probably 15+%. Just guessing.

Lower stages are far easier. Hypersonic heating under Mach 10 or 12 is a whole lot less intense. Organic composites in some areas become feasible. You could even put flyback wings on it, perhaps. But whatever you do, it won't be 5-8% inert. The way around higher inerts at fixed rocket performance is more stages. That, too, is well-known and well-proven. More stages does not necessarily mean more cost. But you do have to keep it simple so the logistical support tail can stay small. The rocket / system designs need to look more like battlefield weapons than our traditional launch rockets. Trying the first steps into that model is the real secret of Spacex's lower costs. They've re-used nothing yet. And at 5% inerts, they won't. You can bet any flyback first stage with landing legs or whatever, will be a lot heavier than the stage they're flying now.

As for the suborbital rocket planes, those typically leave the air and return at Mach 3 or so. That’s actually very little transient heating, and it is a short transient. Not at all the same problem as sustained flight at Mach 3. Even with organic composites as exposed structure, it is possible to heat-sink your way through that short re-entry. There is no radical new heat protection or structural technology there with Spaceship Two or Lynx. The “radical” item in Spaceship Two is the folding tail. Lynx’s “radical” technology is the long life / low maintenance engine.

GW

Hop · 2012-01-15 19:49:50

JoshNH4H wrote:

Hop wrote:
JoshNH4H wrote:
I was talking about Mass Ratio and nothing else.
You were talking about mass ratio to function as a SSTO vehicle.
If your T/W is low, you will incur excessive gravity loss and your needed delta V will be higher than 9.4 km/s.
So no, the Centaur IV does not have the mass ratio to function as a SSTO vehicle.
I believe my meaning was excessively clear and not arguable,

Mass ratio and T/W are a trade space.

For a given payload you can have better T/W by having more rocket engines, but this harms your mass ratio.

You can also improve T/W by using lower ISP but higher thrust engines. However lower ISP will also hurt your mass ratio.

Another trade space is comfort and spaciousness vs good mileage for motor vehicles. You can have the good mileage of a moped or the comfort of a limo but not both.

Saying a Centaur has the mass ratio to be a SSTO is like like saying a limo get 90 miles to a gallon. A limo doesn't get a moped's mileage. And with a very poor thrust to weight ratio, a Centaur's delta V budget for reaching LEO will not be 9.4 km/s.

The notion that a Centaur has the mass ratio to be an SSTO is not only arguable, it's flat out wrong.

Hop · 2012-01-15 20:51:15

RGClark wrote:

What would be the mass ratio and delta-V if you added engines to achieve lift off?

As I mentioned to Josh, there's a trade. You can improve T/W and thus reduce gravity loss by adding more engines. But this increases your dry mass and harms mass ratio.

RGClark wrote:

BTW, Dr. John Schilling has an online payload estimator that allows you to estimate the
the payload you can lift to orbit with a launch vehicle. You enter in the propellant and dry
mass and the vacuum thrust and vacuum Isp. The program does give a warning if your
thrust is less than your launch weight.
Launch Vehicle Performance Calculator.
http://www.silverbirdastronautics.com/LVperform.html
Bob Clark

Yes, a good resource. I've linked to Schilling's calculator several times in this thread. As you mentioned it warns you if your T/W is less than one.

Schilling also posted some explanation of his calculator: Launch Methodology

He talks about the benefit of higher thrust during early ascent. He writes, in part "If dissimilar staging is allowed, the optimal design is generally one which uses a high-thrust first stage (often augmented by strap-on boosters) to rapidly escape the early high-loss flight regime, and a relatively low thrust but high Isp upper stage to most efficiently add delta-V once most loss terms have become irrelevant."

That is why I like two or even three stages to orbit. The lower stage can be high thrust kerosene to get out of the early high loss regime faster. Then the upper stages that don't need high T/W ratios can use better ISP stages. Besides having better ISP, the upper stage can have fewer rocket engines, thus helping with the mass ratio.

The major obstacle for reusing the upper stages is getting past the 8 km/s re-entry. As I mentioned earlier, propellant in LEO could provide a means for upper stages to shed re-entry velocity and thus avoid the temperature and stress of 8 km/s re-entry.

Another possibility is momentum exchange tethers.

The tether foot in the above illustration moves 6.9 km/s at 300 km altitude. Shaving 1 km/s not only helps with the mass fraction but lessens the re-entry abuse.

Last edited by Hop (2012-01-15 20:53:40)

JoshNH4H · 2012-01-18 21:03:33

Hop-

You knew what I was talking about. I know you knew what I was talking about because you hyperfocused on the deficiency which I excepted using the word "theoretically" to say that the rocket could not in reality make it to orbit. If it will end this nonsense, I will also offer up the Saturn V second stage, which has enough T/W to get off the ground as well as the delta-V to make it to orbit, if you assume a trajectory averaged Isp for the J-2 engine of at least 380 s, which seems reasonable seeing as you get out of the lower atmosphere soon enough.

Hop · 2012-01-19 16:04:12

JoshNH4H wrote:

I will also offer up the Saturn V second stage, which has enough T/W to get off the ground as well as the delta-V to make it to orbit,

I'm at a disadvantage here. I will offer links and math to back up my assertions as well as drawings to convey an idea. You as well as others just write from your gut. You proclaim the math exercises not worth your time. So you can pop out posts in 3 minutes when my posts can consume 15 minutes or 20 minutes. Time is a precious commodity for me and this lopsided exchange is annoying.

Anyway...

Saturn II C-5A

Gross mass: 384,057 kg
Dry mass: 31,740 kg
Thrust: 4446.65 kN
Specific Impulse sea level 200 s.

You can try inputting these quantities to Schilling's launch vehicle calculator

And you continue to ignore that an SSTO RLV would also have to endure re-entry. Adding TPS and other accommodations for re-entry will also harm mass fraction.

Last edited by Hop (2012-01-19 16:05:07)

JoshNH4H · 2012-01-19 20:54:36

Hop, actually I did do the math. I did the math, in fact, on several different stages, because obviously not all are capable of making orbit, in theory or otherwise. A short post does not necessarily indicate a small amount of work having gone into it. The astronautix article on the Saturn V listed information for the Saturn V second stage that is different from that in the article to which you linked. I suspect that my link is correct because of the following text in yours:

Final common second stage design for Saturn C-3, C-4 and C-5 (November 1961). Developed into Saturn V second stage.

Also, it is silly to use the sea level Isp to approximate the Isp for the entire trajectory. As I said, it could in theory function as an SSTO vehicle so long as the trajectory averaged Isp were at least 380 s. This is pretty reasonable, IMO.

By the way, I'm not "continually ignoring" anything. I made a correct statement. You went off on a tangent for no reason. I offered something that would allow us to get back on track. You want to continue this pointless tangent. Why?

edit: Found an incomplete sentence

Last edited by JoshNH4H (2012-01-20 13:54:17)

Rune · 2012-01-20 07:29:40

GW Johnson wrote:

I may be the victim of past thinking, but I think the main decision to make is whether you want lifting flight or ballistic flight post re-entry. Putting the weight off-center a bit in a traditional capsule gets you side forces of some significant fraction of the drag force during re-entry hypersonics. This allows you to shift the location of the landing ellipse by a couple of ellipse dimensions L-R, or downrange-uprange. But it's worthless as a directional control once the hypersonics are over. From that point, you're ballistic. Control there requires real wings.

Or a reaction control system. I think you might be a little too much focused on either runway horizontal landing or sea landing. What's wrong with vertical powered landing? Slowing and landing a subsonic, mostly empty, stage propulsively is a piece of cake, both thrust- and fuel-wise. And, you don't have to add either wings and undercarriage (10% empty weight? Plus heavier TPS on the sharp leading edges), or floats and the structure and corrosion treatment to survive the ocean (bad company to reusability, ocean immersion). You just need lots of engines or deep throttling, and bigger tanks. Yet I don't recall to read you ever suggesting it, except for Mars landers.

Rune. We pretty much agree on everything else.

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Re: Reusable Rockets to Orbit

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Re: Reusable Rockets to Orbit

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Re: Reusable Rockets to Orbit

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Re: Reusable Rockets to Orbit

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Re: Reusable Rockets to Orbit

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Re: Reusable Rockets to Orbit

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Re: Reusable Rockets to Orbit

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Re: Reusable Rockets to Orbit

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Re: Reusable Rockets to Orbit

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Re: Reusable Rockets to Orbit

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Re: Reusable Rockets to Orbit

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Re: Reusable Rockets to Orbit

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Re: Reusable Rockets to Orbit

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Re: Reusable Rockets to Orbit

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Re: Reusable Rockets to Orbit

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