Some proposals for low cost heavy lift launchers.

[RGClark]

Here are some possibilities for lower cost super heavy lift launchers, in the 100,000+ kg payload range. As described in this article the proposals for the heavy lift launchers using kerosene-fueled lower stages are focusing on using diameters for the tanks of that of the large size Delta IV, at 5.1 meters wide or the even larger shuttle ET, at 8.4 meters wide:

All-Liquid: A Super Heavy Lift Alternative?
by Ed Kyle, Updated 11/29/2009
http://www.spacelaunchreport.com/liquidhllv.html

The reason for this is that it is cheaper to create new tanks of the same diameter as already produced ones by using the same tooling as those previous ones. This is true even if switching from hydrogen to kerosene in the new tanks.
However, I will argue that you can get super heavy lift launchers without using the expensive upper stages of the other proposals by using the very high mass ratios proven possible by SpaceX with the Falcon 9 lower stage, at above 20 to 1:

SPACEX ACHIEVES ORBITAL BULLSEYE WITH INAUGURAL FLIGHT OF FALCON 9 ROCKET.
Cape Canaveral, Florida – June 7, 2010
"The Merlin engine is one of only two orbit class rocket engines developed in the United States in the last decade (SpaceX’s Kestrel is the other), and is the highest efficiency American hydrocarbon engine ever built.
"The Falcon 9 first stage, with a fully fueled to dry weight ratio of over 20, has the world's best structural efficiency, despite being designed to higher human rated factors of safety."
http://www.spacex.com/press.php?page=20100607

We will use tanks of the same size as these other proposals but will use parallel, "bimese", staging with cross-feed fueling. This method uses two copies of lower stages mated together in parallel with the fueling for all the engines coming sequentially from only a single stage, and with that stage being jettisoned when it's expended its fuel. See the attached images below for how parallel staging with cross-feed fueling works.
Do the calculation first for the large 8.4 meter wide tank version. At the bottom of Kyle's "All-Liquid: A Super Heavy Lift Alternative?" article is given the estimated mass values for the gross mass and propellant mass of the 8.4 meter wide core first stage. The gross mass of this single stage is given as 1,423 metric tons and the propellant mass as 1,323 metric tons, so the empty mass of the stage would be approx. 100 metric tons (a proportionally small amount is also taken up by the residual propellant at the end of the flight.) Then the mass ratio is 14 to 1. However, the much smaller Falcon 9 first stage has already demonstrated a mass ratio of over 20 to 1.
A key fact about scaling is that you can increase your payload to orbit more than the proportional amount indicated by scaling the rocket up. Said another way, by scaling your rocket larger your mass ratio in fact gets better. The reason is the volume and mass of your propellant increases by cube of the increase and key weight components such as the engines and tanks do also, but some components such as fairings, avionics, wiring, etc. increase at a much smaller rate. That savings in dry weight translates to a better mass ratio, and so a payload even better than the proportional increase in mass.
This is the reason for example that proponents of the "big dumb booster" concept say you reduce your costs to orbit just by making very large rockets. It's also the reason that for all three of the reusable launch vehicle (RLV's) proposals that had been made to NASA in the 90's, for each them their half-scale demonstrators could only be suborbital.
Then we would get an even better mass ratio for this "super Evolved Atlas" core than the 20 to 1 of the Falcon 9 first stage, if we used the weight saving methods of the Falcon 9 first stage, which used aluminum-lithium tanks with common bulkhead design. It would also work to get a comparable high mass ratio if instead the balloon tanks of the earlier Atlas versions prior to the Atlas V were used.
So I'll use the mass ratio 20 to 1 to get a dry mass of 71.15 mT, call it 70,000 kg, though we should be able to do better than this. We'll calculate the case where we use the standard performance parameters of the RD-180 first, i.e., without altitude compensation methods. I'll use the average Isp of 329 s given in the Kyle article for the first leg of the trip, and 338 s for the standard vacuum Isp of the RD-180. For the required delta-V I'll use the 8,900 m/s often given for kerosene fueled vehicles when you take into account the reduction of the gravity drag using dense propellants. Estimate the payload as 115 mT. Then the delta-V for the first leg is 329*9.8ln(1 + 1,323/(2*70 + 1*1,323 + 115)) = 1,960 m/s. For the second leg the delta-V is 338*9.8ln(1 + 1,323/(70 + 115)) = 6,950 m/s. So the total delta-V is 8,910 m/s, sufficient for LEO with the 115 mT payload, by the 8,900 m/s value I'm taking here as required for a dense propellant vehicle.
Now let's estimate it assuming we can use altitude compensation methods. We'll use performance numbers given in this report:

Alternate Propellants for SSTO Launchers.
Dr. Bruce Dunn
Adapted from a Presentation at:
Space Access 96
Phoenix Arizona
April 25 - 27, 1996
http://www.dunnspace.com/alternate_ssto_propellants.htm

In table 2 is given the estimated average Isp for a high performance kerolox engine with altitude compensation as 338.3 s. We'll take the vacuum Isp as that reached by high performance vacuum optimized kerolox engines as 360 s. Estimate payload as 145,000 kg. For the first leg, the delta-V is 338.3*9.8ln(1 + 1,323/(2*70 + 1*1,323 + 145)) = 1,990 m/s. For the second leg the delta-V is 360*9.8ln(1 + 1,323/(70 + 145)) = 6,940 m/s, for a total delta-V of 8,930 m/s, sufficient for orbit with the 145,000 kg payload.
Now we'll estimate the payload using the higher energy fuel methylacetylene. The average Isp is given as 352 s in Dunn's report. The theoretical vacuum Isp is given as 391 s. High performance engines can get quite close to the theoretical value, at 97% and above. So I'll take the vacuum Isp as 380 s. Estimate the payload as 175,000 kg. Then the delta-V over the first leg is 352*9.8ln(1 + 1,323/(2*70 + 1*1,323 + 175)) = 2,040 s. For the second leg the delta-V will be 380*9.8ln(1 + 1,323/(70 + 175)) = 6,910 s, for a total delta-V of 8,950 m/s, sufficient for orbit with the 175,000 kg payload.


Bob Clark

 

[Urwumpe]

Correction - the practical limit for high-performance vacuum optimized kerolox rocket engines is fixed at 350 s (~3400 m/s), about 100 m/s lower. The problem is not, that you can drive the chamber pressure higher with the energy available in the fuel and still gain a higher specific impulse. By only looking at what happens in the power head and by assuming simple gas dynamics relations between chamber pressure and specific impulse, you are right, 3500 m/s is really the logical result by applying the theoretical maximum chamber pressure that can be reached without higher tank head (pressure at the outlet of the propellant tanks) on known engine data.

The problem is the needed expansion ration for getting the optimum, and the resulting nozzle mass, which degrades rocket stage performance. Unless you have a very heavy upper stage, that has a very long burn time, the larger nozzle will eat the performance. Many such engines will increase the negative effect.

Just on the paper: 100 m/s more specific impulse mean about twice the chamber pressure, so the combustion chamber and thrust can be only half as large for the same thrust, but the nozzle has to be a bit more than 2 times longer than the original and still reach almost the same exit diameter (so you get the same thrust). Including all weight reductions by the smaller power head, your engine will weight over 5 times more than a less powerful variant, only for getting 100 m/s more specific impulse.

You can of course also reduce the nozzle size and let it be underexpanded, but then, instead of reaching the theoretical gain of 100 m/s, maybe (nozzle aerodynamics are more complex than fighter jet aerodynamics) 20 m/s will eventually leave the engine.

That combined with an explosion in production costs for an expendable engines.

Things will be even worse for an SSTO, because you enter hells kitchen for the many trade-offs and optimizations and experimental measurements for getting essentially an RD-170 (which is about the maximum you can sequence out of kerosene and LOX, there is not much left to be gained at the throat of the engines) to be at optimal expansion for a wide range of pressures. For lift-off it is great to have a high chamber pressure, because you gain by it in any case. In the vacuum, high chamber pressures are suddenly annoying: As tempting and good as they are, in practical engineering, there can be a too good rocket engine.

If you think this is now nitpicking...well, I was just counter-nitpicking at you. You pick a few numbers of a friendly favorable calculation (Dunn is a BDB apostle, his numbers have always been too good to be true.), so your big dumb booster becomes good - as long as nobody looks to closely at the numbers. Is a bit like the past in Europe when you wanted to buy an airline ticket. You can get from Berlin to London for 45 €. Plus fees, taxes, fuel costs and the price of an attractive competent stewardess. Don't just look too closely at the price.

Also... putting a high performance engine on something that was planned to become an BDB is some sort of ironic, since this cost factor should have been killed by BDB... now it is embraced as necessary evil for reuse. When the launch ground infrastructure comes into place, the calculations will get along "Well, maybe we can do things a bit smaller now, and save a lot of money here" - the evolution of rocket feasibility studies from 1950 to today, repeated on the paper.

As current example here, about how different real rockets are to blue print or Colliers Magazine rockets:

Just look at Sea Dragon - great idea, the ocean costs nothing and it needs no VAB, since all assembly and a lot of welding will happen on the water in a small protected lagoon. And then you read how many buoys, barges, ships and tugs it needs for each step of the operations. Ooops.

Not to mention that the idea of larger combustion chambers for simplicity had been found to be opposing to real conditions in the 1960s, when the F1 engine was designed - larger chambers mean much more trouble for getting stable thrust, downsizing of thrust chambers by higher chamber pressure was also a relief of this problem, since you had been able to get the same thrust ranges with chamber and injector geometries that had already been investigated.

SpaceX is also a poor example of simplicity. Their rocket is maybe marketed as simpler as other designs and in terms of engine technology, this is true, but practically, it is not different to LMM or Boeing products. The rocket needed less testing by having less powerful engines, but they needed more painful testing and paid a lot of teaching money for also reducing the amount of engineers on the ground for launching. The Falcon 9 is only as much cheaper as other designs, as it is also less powerful - and it is doubtful at the current development speed, that the Falcon 9 will get the needed number of flights before an successor is already thrown into the field... It is a first one, but not really a good one (The Falcon 9 is actually in one league as the good old Soyuz rocket - and guess which one has currently the better market position).

 

[T.Neo]

Just look at Sea Dragon - great idea, the ocean costs nothing and it needs no VAB, since all assembly and a lot of welding will happen on the water in a small protected lagoon. And then you read how many buoys, barges, ships and tugs it needs for each step of the operations. Ooops.

Indeed. I looked into Sea Dragon (and sea launch in general) a while back, and you really struggle with things that you don't get with a land launch, that outweigh the gains of sea launch/assembly.

Putting together such a large vehicle would be very difficult regardless, even moderately sized launchers are no easy feat to assemble.

 

[RGClark]

Indeed. I looked into Sea Dragon (and sea launch in general) a while back, and you really struggle with things that you don't get with a land launch, that outweigh the gains of sea launch/assembly.
Putting together such a large vehicle would be very difficult regardless, even moderately sized launchers are no easy feat to assemble.

You can get really large payloads with the 8.4 meter wide super "Evolved Atlas" stage by using parallel, "trimese", staging with cross-feed fueling. This would use now three copies of the lower stages mated together in parallel with the fueling for all the engines coming sequentially from only a single stage, and with that stage being jettisoned when its fuel is expended.
Again we'll calculate first the case where we use the standard performance parameters of the RD-180, i.e., without altitude compensation methods. I'll use the average Isp of 329 s given in the Kyle article for the first leg of the trip, and for the required delta-V, again the 8,900 m/s often given for kerosene fueled vehicles when you take into account the reduction of the gravity drag using dense propellants. Estimate the payload as 200 mT. Then the delta-V for the first leg with all three super Evolved Atlas's attached will be 329*9.8ln(1+1,323/(3*70 + 2*1,323 + 200)) = 1,160 m/s. For the second leg we'll use the vacuum Isp of 338 s, then the delta-V will be 338*9.8ln(1 + 1,323/(2*70 + 1*1,323 + 200)) = 1,940 m/s. And for the final leg 338*9.8ln(1 + 1,323/(70 +200)) = 5,880 m/s. So the total delta-V is 8,980 m/s, sufficient for orbit with the 200,000 kg payload.
Now let's estimate it assuming we can use altitude compensation methods. We'll use performance numbers given in this report:

Alternate Propellants for SSTO Launchers.
Dr. Bruce Dunn
Adapted from a Presentation at:
Space Access 96
Phoenix Arizona
April 25 - 27, 1996
http://www.dunnspace.com/alternate_ssto_propellants.htm

In table 2 is given the estimated average Isp for a high performance kerolox engine with altitude compensation as 338.3 s. We'll take the vacuum Isp as that reached by high performance vacuum optimized kerolox engines as 360 s. Estimate the payload now as 250 metric tons. Then the delta-V during the first leg will be 338.3*9.8ln(1+1,323/(3*70 + 2*1,323 + 250)) = 1,180 m/s. For the second leg the delta-V will be 360*9.8ln(1 + 1,323/(2*70 + 1*1,323 + 250)) = 2,020 m/s. For the third leg the delta-V will 360*9.8ln(1 + 1,323/(70 + 250)) = 5,770 m/s. So the total will be 8,970 m/s, sufficient for orbit with the 250,000 kg payload.
Now we'll estimate the payload using the higher energy methylacetylene. The average Isp is given as 352 s in Dunn's report. The theoretical vacuum Isp is given as 391 s. High performance engines can get quite close to the theoretical value, at 97% and above. So we'll take the vacuum Isp as 380 s. Estimate the payload now as 300 mT. The first leg delta-V will now be 352*9.8ln(1 + 1,323/(3*70 + 2*1,323 +300)) =1,210 m/s. For the second leg 380*9.8ln(1 + 1,323/(2*70 + 1*1,323 + 300)) = 2,080 m/s. For the third leg 380*9.8ln(1 + 1,323/(70 + 300)) = 5,660 m/s. So the total is 8,950 m/s, sufficient for orbit with the 300,000 kg payload.

This trimese version of the vehicle would be huge however. For instance it would weigh more than the Saturn V. One of the big cost factors for the development of some of the super heavy lift launchers is that they are so heavy they would require the construction of new and expensive launch platforms. Undoubtedly, the bimese version would be the one to be built first if this launch system is selected.



Bob Clark

---------- Post added at 06:48 AM ---------- Previous post was at 06:35 AM ----------

Correction - the practical limit for high-performance vacuum optimized kerolox rocket engines is fixed at 350 s (~3400 m/s), about 100 m/s lower. The problem is not, that you can drive the chamber pressure higher with the energy available in the fuel and still gain a higher specific impulse. By only looking at what happens in the power head and by assuming simple gas dynamics relations between chamber pressure and specific impulse, you are right, 3500 m/s is really the logical result by applying the theoretical maximum chamber pressure that can be reached without higher tank head (pressure at the outlet of the propellant tanks) on known engine data.
The problem is the needed expansion ration for getting the optimum, and the resulting nozzle mass, which degrades rocket stage performance. Unless you have a very heavy upper stage, that has a very long burn time, the larger nozzle will eat the performance. Many such engines will increase the negative effect.
...

You may be right in regards to using a very high expansion ratio in order to get those high vacuum Isp's while using a standard nozzle. However, an aerospike nozzle may be able to work.
Also, note that the first calculation did not require altitude compensation or the very high vacuum Isp's. It just used the standard performance parameters of the RD-180 to get a 115 metric ton payload.


Bob Clark

 

[Urwumpe]

You may be right in regards to using a very high expansion ratio in order to get those high vacuum Isp's while using a standard nozzle. However, an aerospike nozzle may be able to work.

No change then - the mass is then inside the (truncated) spike.

 

[RGClark]

The previous posts in this thread were about using liquid-fueled only stages. However, the shuttle derived heavy lift vehicles proposed that use SRB's can also reduce costs by being made fully reusable.

Some have soured on the idea of reusability because of the case of the space shuttle. But the problem with the shuttle was that that spacecraft that had to be carried to orbit was so heavy. It was nearly four times the weight of the payload you could carry. I quoted before a statement by Robert Zubrin in one of his books that emphasized this point, which he argued contributed to the shuttle being a fiscal disaster.

But actually for most launch vehicles the upper stage that actually makes it to orbit along with the payload is usually comparatively small, in fact, frequently smaller than the payload. And this gets even better the larger the launcher gets. Note then for a large launcher since the dry mass of the upper stage will be a fraction of the payload mass and the reentry/landing systems will be a fraction of the upper stage dry mass, the extra weight to make the upper stage reusable is actually a small fraction of the payload mass, so will only subtract a small amount from the payload.

This point is well illustrated by the DIRECT Jupiter-246 heavy lift launcher. See the specifications in the diagram attached below.

Note that the dry mass of the upper stage is less than 12,000 kg. But the payload mass is in the range of 105,000 kg to 117,000 kg. However, the extra mass for reentry/landing systems for an orbital stage is commonly estimated to total about 28% of the dry mass of the stage (see below), which is about 3,400 kg for this upper stage. This would subtract off a comparatively small amount from the payload mass.

However, the biggest cost saving would not be in making the upper stage reusable but in the reusability of the expensive core stage with its 4 SSME engines and shuttle ET-derived propellant tank. This has a dry mass of 66,895 kg. So if we used the 28% estimate for reentry/landing systems this would be an extra 18,730 kg added to this stage weight. So it would conceivable subtract this amount from the payload weight.

But there are two key reasons why it will likely not have to be this high an amount that has to be subtracted off from the payload weight. First, another point Zubrin makes in that passage I quoted from his book Entering Space is that for a first stage every extra kilo added to the first stage weight generally will only subtract about .1 of a kilo from the payload weight. However, this core stage is not quite a first stage; it's closer to being a second stage. The amount of payload that has to be subtracted off will be somewhat more than 1/10th though not the full amount of this extra weight, depending on how much delta-V this stage makes up.

The second key reason is that this core stage will not have to reach all the way to orbit so its reentry regime will not be as severe as for an orbital stage, so the reentry systems not as heavy. To see why, notice that unlike the shuttle ET, this ET-syle propellant tank will be carrying the 200,000 kg gross weight of the upper stage plus the ca. 100,000 kg payload, much more mass to loft before staging than for the shuttle. So it will reach significantly lower velocity.

For the 28% of the landing mass for reentry/landing systems, first Robert Zubrin gives an estimate of about 15% for reentry thermal protection.
Secondly, in a discussion between Henry Spencer and Mitchell Burnside Clapp on the relative benefits of horizontal vs. vertical landing, the extra mass for winged landing or a powered descent is about 10%.
Finally, in a discussion on yarchive.net/space, the landing gear weight is given as about 3%.

However, note with modern materials quite likely this 28% estimate for the reentry/landing systems can be cut in half.

Bob Clark

 

[RGClark]

Recently there has been research on reducing the structural weight of rockets, which can go to lofting additional payload mass. Tests and theoretical work suggests the dry mass of rockets can be reduced by 10% to 20%:

NASA engineers say 'shell-buckling' test supports thinner rocket skins.
Published: Wednesday, March 23, 2011, 5:48 PM
By Lee Roop, The Huntsville Times
http://blog.al.com/space-news/2011/03/nasa_engineers_say_shell-buckl.html

I wrote NASA engineers about how this research can improve the
payload of HLV's in an email copied below:

=================================================
Subject : RS-84 powered SpaceX "Falcons" for heavy lift vehicles.
Date : Sat, Feb 12, 2011 09:00 AM EST
From : "Robert Clark" <****@****>
To : ****@nasa.gov, ****@nasa.gov, ****@nasa.gov

Hello. I was very interested to read about your research on minimizing structural mass on launch vehicles. I thought you might be interested in this post to a space bbs copied below. It is based on a simple and indeed obvious idea: to maximize payload use *both* a high efficiency engine *and* a low
inert mass fraction vehicle. The high efficiency engines give you a high Isp and the low inert mass fraction gives you a high mass ratio, the two components that go into the rocket equation to calculate the delta-V.
Optimizing both of these will maximize your delta-V and payload.
This has been done for upper stages. For instance if you read the history of the development of the Apollo vehicles a great deal of the development time went to making the S-IVB upper stage use a common bulkhead design to minimize the dry weight as well as getting the highest performance possible from the
J-2 engines. Also, the Centaur upper stages use a "balloon tank" design to minimize tank weight, and high Isp vacuum optimized engines.
What I have been discussing on various forums and in correspondence with individuals in the industry is to also apply this really fundamentally clear idea to first stages. But it still is not being done even among the "NewSpace" companies that you would expect to be innovative in their design. For
instance, SpaceX deserves great credit in being able to get their Falcon 9 first stage to have a better than 20 to 1 mass ratio, but then the engine they use on it is the Merlin which is no more efficient than the engines used on the original Atlas rocket launched in 1960. And Orbital Sciences will be using the NK-33 derived AJ-26 engines from Aerojet that have the highest average Isp for kerosene engines, but then they use heavy Russian-design for their structures that result in a poor mass ratio.
These companies have their own contracts and heavy financial investments in their own structures and engines. What is needed is someone outside the box, and someone who can think outside the box, such as NASA to make the request to combine the best features of both and leave out the bad features.
The post below discusses using RS-84 engines for the SpaceX heavy lift proposal rather than the still low efficiency Merlin 2 and improving the mass ratio in the heavy lift vehicle to be at or above the 20 to 1 mass ratio of their Falcon 9 first stage, which should be possible from principles of scaling. This would have the additional advantage of the engine being reusable which gives you also an additional option of reducing costs even further by making some or all of the stages be reusable.
Note as well that if as your research shows the structural mass can be reduced by as much as 20% on launch vehicles then that would raise the SpaceX HLV proposal's mass ratio from the 16.7 to range to close to the 20 to 1 range.


Bob Clark
===================================================


The space bbs post I was referring to in the email is here:

Newsgroups: sci.space.policy, sci.astro, sci.physics,
sci.space.history
From: Robert Clark <rgregorycl...@yahoo.com>
Date: Mon, 24 Jan 2011 12:16:47 -0800 (PST)
Subject: Re: Some proposals for low cost heavy lift launchers.
http://groups.google.com/group/sci.space.policy/msg/23586fe490c09d2d?hl=en


Video of an experimental test on a life-size structure (shuttle ET size tank) of the new mass reduction methods has been posted to the net:

World's largest can crusher. [highlights]
http://www.ustream.tv/recorded/13514449/highlight/159241

World's Largest Can Crusher @ 9:30AM CST - 1.
March 23, 2011 | Length: 56:47
http://www.ustream.tv/recorded/13512828

World's Largest Can Crusher @ 9:30AM CST - 2.
March 23, 2011 | Length: 62:26
http://www.ustream.tv/recorded/13513586

World's Largest Can Crusher @ 9:30AM CST - 3.
March 23, 2011 | Length: 52:12
http://www.ustream.tv/recorded/13514449

World's Largest Can Crusher @ 9:30AM CST - 4.
March 23, 2011 | Length: 64:52
http://www.ustream.tv/recorded/13515162

World's Largest Can Crusher @ 9:30AM CST - Post Test Briefing.
March 23, 2011 | Length: 10:33
http://www.ustream.tv/recorded/13516463

---------- Post added 04-29-11 at 04:49 AM ---------- Previous post was 04-28-11 at 07:17 PM ----------

Below is a follow up email I sent the NASA engineers after the SpaceX
plans on the Falcon Heavy were announced. I mention in the letter
SpaceX expects to achieve a 30 to 1(!) mass ratio with their side
boosters. See here for that:

FALCON HEAVY OVERVIEW.
http://www.spacex.com/falcon_heavy.php


Bob Clark


==========================================
Subject : About the Falcon Heavy.
Date : Wed, Apr 06, 2011 11:46 PM EDT
From : "Robert Clark" <****@****>
To : ****@nasa.gov, ****@nasa.gov, ****@@nasa.gov

Hello. There is alot of buzz about the Falcon Heavy and its lift capabilities
using cross-feed fueling, approx. 50% higher than that of the Falcon 9 Heavy without it.
I've discussed this before but I think it should be emphasized again: using
BOTH the lightweight strucural designs of SpaceX AND the highest efficiency kerosene engines available now such as the NK-33 or RD-180, you can lift markedly higher payload at relatively low development costs.
In the news about the Falcon Heavy the mass ratio of the side boosters has been given as better than 30 to 1(!) As you know the first stage of the Falcon 9 is already a quite good 20 to 1. I am curious how they are able to improve the mass ratio of the boosters to this degree. Perhaps in the Falcon 9 first stage they are including the mass of the interstage in the dry mass number and the boosters don't have this so so have a smaller dry mass number? That would be surprising if that is the explanation since you wouldn't expect the interstage to make up such a large proportion of the dry mass of a first stage. Another possibility is that the side boosters don't have to be carried as far as the main core stage so are not subjected to the higher acceleration loads and the structural elements such as the propellant tanks can be made lighter and/or thinner.
SpaceX does deserve kudos for developing such high mass ratios. However, they are still using the Merlin engines of no better efficiency than those of the original Atlas rocket engines developed 50 years ago. It is understandable that SpaceX would want to stick with their engines since they spent alot of development time and money on them. They understandably want to get a return on that investment. But you would think someone at NASA would have the creativity to do the calculation to show how much better a launcher you can get if you combine the SpaceX structural designs with the much better engines we have available now.
Perhaps this is a calculation you could do using the 30 to 1 mass ratio of
the side boosters and the 20 to 1 mass ratio of the core first stage and using NK-33's or the RD-180 for the engines.
It might interesting as well to do the calculation for how these side
boosters with better engines and the 30 to 1 mass ratio can be SSTO's with
significant payload capability. Even if, for example, they have thinner tank
walls you might still be able to get it to work with 3 NK-33's by shutting
down 1 or 2 engines later in the flight to reduce the acceleration loads.


Bob Clark
==========================================

 

[RGClark]

NASA appears to be leaning to a 70 mt payload shuttle-derived
launcher as an interim solution to developing a heavy lift vehicle.
This would use two 4-segment SRM's as does the shuttle and an ET. But
it would not have a shuttle orbiter, nor would this Phase I vehicle
have an upper stage:

SLS planning focuses on dual phase approach opening with SD HLV.
April 25th, 2011 by Chris Bergin
http://www.nasaspaceflight.com/2011/04/sls-planning-dual-phase-approach-opening-sd-hlv/

However, built into this plan is that at most 4 flights of this
vehicle will be made before it is discontinued in favor of a more
expensive, 130 mt payload upgrade. These 4 flights are to regarded as
"test flights" according to the Bergin article. They will use 3 SSME's
at a time and only 12 of those will be available including those taken
from the retired space shuttles, thus allowing only 4 flights.
Presumably after that either the production of new SSME's will be
started or their expendable versions will be, or NASA will choose
instead to use kerosene fueled engines for the core stage.
However, a better plan in my view would be to explore methods in
which this Phase I vehicle could be reusable. Then this low cost HLV
could have many more missions as well as cutting costs in being
reusable. This would give you more options as to when and if the more
expensive vehicle needed to be developed.
Many at NASA are not favorably inclined towards reusable systems
because of the experience of the shuttle. However, as I mentioned
before in the post #6 above, a key reason for why the shuttle was not
economical would not hold in this case: it would not have to carry the
80 mt orbiter that took out most of the vehicles payload capacity.
Another reason why a reusable vehicle could be done better now is
because of the research that has already been done to address the
failings of the shuttle system. For instance, for the X-33/VentureStar
program the metallic shingles to be used for thermal protection have
confirmed in testing they would require less maintenance than the
ceramic tiles of the shuttle.
The advanced ceramics used on the Air Force's X-37B were also
expected to cut maintenance on thermal protection. It would be useful
to find out if they have been successful in that regard.
The X-37B may also serve as a good model to use for the reentry
system for the ET tank to be used on the Phase I vehicle. Note that
the X-37B's short stubby wings are much smaller in proportion to the
size of the vehicle than those of the space shuttle. That and the
composite materials used for the wings would result in much reduced
mass used for the wings for the ET tank.
Other lightweight reentry systems would be the German IXV program
which does not use wings:

IXV Program Aims to Put ESA at Cutting Edge of Re-entry Technology
Posted by Doug Messier on September 18, 2010, at 4:08 am in ESA.
http://www.parabolicarc.com/2010/09/18/ixv-program-aims-put-esa-cutting-edge-reentry-technology/

and the inflatable one NASA is investigating:

NASA Successfully Tests Vacuum-Packed Inflatable Heat Shield.
A vacuum-packed inflatable shroud could enable future spacecraft
reentry on both Earth and Mars.
By Jeremy HsuPosted 08.17.2009 at 3:00 pm
http://www.popsci.com/military-avia.../nasa-puts-inflatable-heat-shield-flight-test


Bob Clark

---------- Post added at 11:40 PM ---------- Previous post was at 01:56 PM ----------

The above discusses that maintenance costs for thermal protection
should be significantly less than for the space shuttle. But another
significant recurring cost for the shuttle program was for maintenance
on the engines. Now, the SSME's have to be overhauled after every
flight, costing ten's of millions of dollars. However, Henry Spencer a
highly regarded expert on the history of space flight has said
Rocketdyne studies show that with a lot of work to upgrade it,
maintenance could be reduced to $750K per flight per engine:

Engine reusability (Henry Spencer)
http://yarchive.net/space/rocket/engine_reusability.html

Spencer here said this would not be satisfactory for really large
reductions in space costs. But this would be a reduction in SSME
maintenance costs by 1 to 2 orders of magnitude, a major reduction in
the costs for using the engine. A key question though is how much
would be the cost to make the necessary upgrades to the engine.
I also did not estimate the extra mass of the reentry/landing
systems. Here's a diagram showing the specifications for the DIRECT
team's version of this Phase I ca. 70 mt launcher:

DIRECTv3 Jupiter-130 - LEO Cargo Launch Vehicle Configuration.
http://www.launchcomplexmodels.com/....4000.10051_CaLV_30x130nmi_29.0deg_090606.jpg

The dry mass of the core stage is given as 63.7 mt. This mass can be
reduced by going to a common bulkhead design for the propellant tanks
that for instance SpaceX was able to use to reduce dry mass for its
Falcon vehicles. As it is now, the intertank on the shuttle ET actually weighs
more than the oxygen tank:

External Tank.
http://science.ksc.nasa.gov/shuttle/technology/sts-newsref/et.html

Going to a common bulkhead design would eliminate this mass, reducing
the dry mass by about 5 mt. Also recent research has shown that dry
mass of rocket vehicles in general can be reduced by 10% to 20%. This
would take off about another 5 mt to 10 mt.
I gave an estimate before in this thread of about 28% of the dry mass
for reentry/landing systems. However, as I said probably with modern
materials we can cut this in half. Then with all these reductions
together the extra mass for reentry/landing systems might only be in
the range of 7,000 kg. So we would still maintain 90% of the payload
mass while gaining reusability and a longer useful life for this low
cost heavy lift launcher.


Bob Clark

 

[RGClark]

NASA appears to be leaning to a 70 mt payload shuttle-derived
launcher as an interim solution to developing a heavy lift vehicle.
This would use two 4-segment SRM's as does the shuttle and an ET. But
it would not have a shuttle orbiter, nor would this Phase I vehicle
have an upper stage:

SLS planning focuses on dual phase approach opening with SD HLV.
April 25th, 2011 by Chris Bergin
http://www.nasaspaceflight.com/2011/04/sls-planning-dual-phase-approach-opening-sd-hlv/

However, built into this plan is that at most 4 flights of this
vehicle will be made before it is discontinued in favor of a more
expensive, 130 mt payload upgrade. These 4 flights are to regarded as
"test flights" according to the Bergin article. They will use 3 SSME's
at a time and only 12 of those will be available including those taken
from the retired space shuttles, thus allowing only 4 flights.
Presumably after that either the production of new SSME's will be
started or their expendable versions will be, or NASA will choose
instead to use kerosene fueled engines for the core stage.
However, a better plan in my view would be to explore methods in
which this Phase I vehicle could be reusable. Then this low cost HLV
could have many more missions as well as cutting costs in being
reusable. This would give you more options as to when and if the more
expensive vehicle needed to be developed...

This is for the interim, Phase I, 70 mt launcher. This is to use two
SRB's and an external tank as with the shuttle system, but no orbiter
and no upper stage. However, the possibilities become especially
interesting when we look at the case of making the Phase II, 100+ mt
payload launcher reusable. This vehicle will have an additional upper
stage. This has a significant advantage for the lightness of the
reentry/landing systems in that only the upper stage at a small dry
mass would have to have the full reentry systems of an orbiting
vehicle. The upper stage of the DIRECT teams's Jupiter-246 for
instance weighs less than 12,000 kg. Also, for this case the ET would
reach a much reduced velocity and would not reach orbit so its reentry
systems would be much simpler and lighter.[1]
To get the full benefits of reusability we'll switch out the RL-10's
or J-2X engines used on the upper stage for SSME(s). This does have a
problem though in that the SSME would have to be made air startable.
The benefits for reusability are so significant that costs estimates
for this upgrade should be made.
However, a different potential solution would also reap additional
benefits. If instead of placing this stage atop the ET tank, we put it
in parallel with it, then the stage could also be started on the
ground.
This has a benefit because now we could use cross-feed fueling
between the ET and upper stage tanks. Cross-feed fueling with parallel
staging is known to be able to increase your payload. For instance by
using it for their Falcon Heavy vehicle SpaceX was able to increase
its payload by 50%.
Note also that we wouldn't have the development cost for a new 4 SSME
engine core stage, as is currently planned for the Phase II vehicle.
We would use the same 3-engine core stage as used for the Phase I
vehicle. The extra thrust for the Phase II vehicle would come from the
upper stage now firing in parallel from the start.
However, another potentially game changing effect of doing this is
that if you look at the mass ratio of this reconfigured upper stage
with SSME(s) you see it has SSTO capability. This is because it has
the weight optimization of an upper stage and now using an engine
optimized to be most efficient during the entire flight to orbit it
can reach orbit in a single stage with significant payload.[2]
In fact not just the Jupiter-246 upper stage would have this
capability, but in fact the Ariane 5's upper stage, the Apollo's S-II
and S-IVB, and the planned Ares I upper stage would as well if
switched out to use SSME(s).
This is important because we will have already existing stages as
well as the engines to make at least an initial version of this upper
stage. This means we could have a significant cost reduction on an
initial version of the upper stage.
Another very key fact is because this upper stage can be used as a
separate launcher and even manned launcher, thus with its own market,
we could initiate it's development and production in parallel to the
low cost Phase I vehicle. So we would get in fact not only a 70+ mt
vehicle, but we would get the 100+ mt launcher and a manned launcher
in just the same short time frame of the Phase I launcher and at a
smaller cost than now planned for the Phase II, 100+ mt launcher.
Indeed because there would be such a significant market for this
manned SSTO vehicle, NASA might not have to pay for its development at
all.


Bob Clark


1.)http://www.orbiter-forum.com/showthread.php?p=249968&postcount=6

2.)http://www.orbiter-forum.com/showthread.php?p=243924&postcount=34