New Shuttle-derived booster squabbles

On 5/13/2011 2:02 PM, Rick Jones wrote:
In sci.space.history Pat wrote:
What's interesting about SpaceX is that everyone thought that a
reduction in launch costs would come from some new technology, and
Musk returned to the most basic of rocket technologies, like
clustered mass produced Lox/kerosene engines and no cryogenics in
the upper stage to get Falcon 9's costs down

I guess because it is the oxidizxer rathe than the fuel LOX doesn't
count as a cryogenic?

In common usage, it means when you start using LH2 as fuel.


What is the freezing point of RP-1? http://en.wikipedia.org/wiki/RP-1
doesn't metion it, but http://en.wikipedia.org/wiki/Jet_fuel talks
about temps below -40. I suppose for short durations that is a don't
care in an upper stage?

The N-1 ran into a strange problem regarding that. Korolev thought he
was being clever by designing it so that the bigger spherical Lox tank
was on the bottom of the smaller diameter kerosene tank, leading to a
conical shape for the whole rocket.
Bad move.
As soon as the propellants began to flow, the kerosene would have to
pass through a pipe that had been chilled for many minutes by immersion
in the interior of the Lox tank as the rocket was fueled. Despite
insulating it, such long immersion would lower its temperture so far
that the kerosene would freeze as it went through the pipe; so instead
it had to pass through multiple pipes on the exterior of the big Lox
tank, greatly adding to both the weight and complexity of the propellant
feed system. Flip it around with the Lox at the top like in the Saturn V
first stage, and all you have to do is insulate the Lox feed pipe
through the kerosene tank that won't have Lox in it till things get
moving at launch. The upper stages of the Saturn V used the this
approach of the colder LH2 propellant at the top for this reason, and
the Shuttle ET uses the Lox on top fed to the engine via a external pipe
in the same way, to move the stack's CG forward for stability during ascent.

Pat

In article h4mdndW_
hdakotatelephone,
says...

On 5/13/2011 11:32 AM, David Spain wrote:
To be totally fair to the Direct approach, their approach and cost
figures derive from the goal being the original VSE (read Lunar
missions) in mind not LEO to ISS. Which is all SpaceX is on the hook to
deliver (for now).

Still SpaceX is setting the bar very low... :-)

What's interesting about SpaceX is that everyone thought that a
reduction in launch costs would come from some new technology, and Musk
returned to the most basic of rocket technologies, like clustered mass
produced Lox/kerosene engines and no cryogenics in the upper stage to
get Falcon 9's costs down

Actually, Henry Spencer long argued that low launch costs hadn't yet
been achieved not because of lack of technology, but because no one
(with enough money) tried seriously to make it work.

This was the argument for "new space" all along, that we'd never get low
cost launches from the existing launch providers, because there simply
was no economic incentive for them to pursue such a thing. You'd be
asking them to shrink their revenue stream and profits. Why in the
world would a large corporation do this to themselves when they have a
large cash cow of a business in doing things "the old way"?

That's why we needed a start-up to achieve low cost. They have the
motivation to achieve low cost, not maximum revenue and profit margins.

Jeff
--
" Solids are a branch of fireworks, not rocketry. :-) :-) ", Henry
Spencer 1/28/2011

In article ,
says...

In sci.space.history Pat Flannery wrote:
What's interesting about SpaceX is that everyone thought that a
reduction in launch costs would come from some new technology, and
Musk returned to the most basic of rocket technologies, like
clustered mass produced Lox/kerosene engines and no cryogenics in
the upper stage to get Falcon 9's costs down

I guess because it is the oxidizxer rathe than the fuel LOX doesn't
count as a cryogenic?

Look at the temperatures. By comparison, LOX is "mildly" cryogenic,
while LH2 is "deeply" cryogenic. Plus, LH2 and H2 loves to leak out of
the tiniest cracks imaginable. Remember the not too infrequent aft
compartment H2 leaks in the shuttle? They were definitely a pain to
deal with.

What is the freezing point of RP-1? http://en.wikipedia.org/wiki/RP-1
doesn't metion it, but http://en.wikipedia.org/wiki/Jet_fuel talks
about temps below -40. I suppose for short durations that is a don't
care in an upper stage?

Short durations shouldn't be a problem because of the large thermal mass
of the tank of kerosene. Besides, if it does become somewhat of a
problem, just insulate the tank a bit and add heaters.

Jeff
--
" Solids are a branch of fireworks, not rocketry. :-) :-) ", Henry
Spencer 1/28/2011

The savings in structural mass and cost using LOX/RP1 isn't sufficient
to pay for the increase in size when compared to LOX/LH2 launching the
same payload. I gave a detailed analysis supporting that elsewhere.

Given that we're moving toward a hydrogen economy costs will be even
more favorable going forward.

On May 14, 8:00*pm, William Mook wrote:
The savings in structural mass and cost using LOX/RP1 isn't sufficient
to pay for the increase in size when compared to LOX/LH2 launching the
same payload. *I gave a detailed analysis supporting that elsewhere.

Given that we're moving toward a hydrogen economy costs will be even
more favorable going forward.

Do you have a link to that calculation? Kerosene engines are
typically simpler and easier to produce than hydrogen ones. It's no
coincidence that SpaceX chose to go with kerosene to produce its low
cost launchers.
Note also that for the heavy lift launchers being considered, the
lower stages are either hydrogen fueled plus solid rocket boosters or
all kerosene fueled. None of them use all hydrogen for the lower
stages. Hydrogen is most useful for upper stages because of the
lighter weight of the propellant load that has to be lifted by the
lower stages.
And several authors have noted that to produce a SSTO it's actually
easier to do using dense propellants rather than hydrogen.

Bob Clark

On May 14, 8:52*pm, Robert Clark wrote:
On May 14, 8:00*pm, William Mook wrote:

The savings in structural mass and cost using LOX/RP1 isn't sufficient
to pay for the increase in size when compared to LOX/LH2 launching the
same payload. *I gave a detailed analysis supporting that elsewhere.

Given that we're moving toward a hydrogen economy costs will be even
more favorable going forward.

*Do you have a link to that calculation? Kerosene engines are
typically simpler and easier to produce than hydrogen ones. It's no
coincidence that SpaceX chose to go with kerosene to produce its low
cost launchers.
*Note also that for the heavy lift launchers being considered, the
lower stages are either hydrogen fueled plus solid rocket boosters or
all kerosene fueled. None of them use all hydrogen for the lower
stages. *Hydrogen is most useful for upper stages because of the
lighter weight of the propellant load that has to be lifted by the
lower stages.
*And several authors have noted that to produce a SSTO it's actually
easier to do using dense propellants rather than hydrogen.

* * Bob Clark


I'm not quibbling with SpaceX here. There are many dimensions. One
is are you willing to bet your company on the result? That's
something SpaceX had to do. NASA did not.

* * * *

LOX is a cryogen, but not one that's as difficult as cryogenic
hydrogen. Hydrogen isn't that difficult today either. Handling
hydrogen is the basis of its use as a terrestrial fuel. So, it need
not be expensive. In fact, the attraction I have toward liquid
hydrogen is that its also the basis of a hydrogen economy, so its a
technology driver in the energy field assuming low cost hydrogen
production with solar or nuclear sources.

Still lets compare LOX/H2 with LOX/RP1

Hydrogen/Oxygen combinations have 4.5 km/sec exhaust speeds with a 6:1
oxygen fuel ratio by weight. So,

GAS DENSITY MASS VOLUME
Lox 1.14 kg/liter 6 kg 5.26 liters
H2 0.07 kg/liter 1 kg 14.29 liters

0.36 kg/liter 7 kg 19.55 liters

The S-IVB attains an 11.6% structure fraction in 1960s. This was an
upper stage and needn't be built as heavily. So, today we might
attain for a lower stage, something in this range. Let's say 12%
structure in the first stage and a 10% structure in the upper stage.
Or 15% structure for a reusable stage.

RP-1/Oxygen combinations have 3.4 km/sec exhast speeds with a 2.56:1
oxygen fuel ratio by weight. So,

PROP DENSITY MASS VOLUME

Lox 1.14 kg/liter 2.56 kg 2.25 liters
RP1 0.81 kg/liter 1.00 kg 1.23 liters

1.02 kg/liter 3.56 kg 3.48 liters

The S1-C has a structure fraction of 12.8% - but it carried the entire
Saturn V through Max Q. The Falcon Heavy attains a 9% structure
fraction today - so we'll use that for the first stage 7% structure in
the upper stage, and 11% structure for a reusable system.

Let's look at putting up 10 tons with a two stage to orbit rocket with
various combinations. Our goal is to impart an idealized 9.2 km/sec
to the payload to attain a 330 km orbit from Cape Canaveral with two
stages imparting 4.6 km/sec each.

Propellant Fraction
Type H2 64.02% RP1 74.15%

Stage Structure Fraction & Payload Fraction
Low 10% 25.98% 7% 18.85%
High 12% 23.98% 9% 16.85%
Reusable 15% 20.98% 11% 14.85%

Orbiter Stage Structure and Total Weight
Low 3.85 38.49 3.71 53.06
High 5.00 41.70 5.34 59.35
Reusable 7.15 47.67 7.41 67.35

Booster Stage Structure and Total Weight
Low 10.97 148.16 15.99 281.50
High 15.86 173.91 26.36 352.30
Reusable 26.93 227.20 42.49 453.60

So, we can see that an a two-stage system to place 10 tonnes into LEO
from Cape Canaveral will have the following characteristics;

LOX/LH2 - Low - 148.16 TOW - 14.82 structure
LOX/LH2 - High- 173.91 TOW - 20.86 structure
LOX/LH2 - Reus-227.20 TOW - 24.12 structure

LOX/RP1 - Low - 281.50 TOW - 19.70 structure
LOX/RP1 - High-352.30 TOW - 31.70 structure
LOX/RP1 - Reus-453.60 TOW - 49.90 structure

The cost of LOX/LH2 is slightly higher in structure than LOX/RP1, is
the same in avionics and control, and higher in propulsion, and higher
in infrastructure. The cost of launch infrastructure scales with Take
off weight (TOW). On average we have the following ratio, looking at
the cost history of various systems like the Saturn V and Atlas;

COMPARISON OF COSTS

Vehicle
LOX/LH2 vs LOX/RP1: 1.0x avionics x mass - 15% total
LOX/LH2 vs LOX/RP1: 1.1x structure x mass - 35% total
LOX/LH2 vs LOX/RP1: 1.8x propulsion x TOW - 50% total

Launch Center
LOX/LH2 vs LOX/RP1: 1.5x infrastructure take off weight

RESULTS

Vehicle

LOX/LH2 Light 0.94
LOX/RP1 Light 1.00
LOX/LH2 Heavy 1.11
LOX/RP1 Heavy 1.27
LOX/LH2 Reuse 1.48 (0.296 x 5)
LOX/RP1 Reuse 1.67 (0.334 x 5)

Infrastructure (100 launches over life)

LOX/LH2 Light 0.79 (0.008 x 100)
LOX/LH2 Heavy 0.93 (0.009 x 100)
LOX/RP1 Light 1.00 (0.010 x 100)
LOX/LH2 Reuse 1.21 (0.012 x 100)
LOX/RP1 Heavy 1.25 (0.013 x 100)
LOX/RP1 Reuse 1.61 (0.016 x 10))

A reusable LOX/LH2 system that is used at least 5x is superior to any
expendable system, and higher specific impulse beats out lower cost
lower specific impulse every time due to the larger size of a low
specific impulse system for a given payload to orbit.

As the 1960s studies of Sea Dragon shows, very large systems have the
capacity to reduce costs by operation of economies of scale to change
these factors. Larger systems are preferred for any real industrial
use of outer space to resolve environmental and resource problems for
Earth's population today in any case.

On May 14, 10:46*pm, William Mook wrote:
...
The S1-C has a structure fraction of 12.8% - but it carried the entire
Saturn V through Max Q. *The Falcon Heavy attains a 9% structure
fraction today - so we'll use that for the first stage 7% structure in
the upper stage, and 11% structure for a reusable system.
...

That's the main problem there. SpaceX deserves major kudos for showing
you can get a mass ratio above 20 to 1 with a kerosene-fueled first
stage in the Falcon 9:

SpaceX Achieves Orbital Bullseye With Inaugural Flight of Falcon 9
Rocket
Source: SpaceX Posted Monday, June 7, 2010
http://www.spaceref.com/news/viewpr.html?pid=30992

This corresponds to a structure fraction of 5%. And SpaceX reports
that the kerosene side boosters of its Falcon Heavy will have a mass
ratio of 30 to 1:

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

This corresponds to a structure fraction of only 3%.


Bob Clark

On Sat, 14 May 2011 02:27:39 -0800, Pat Flannery
wrote:

I guess because it is the oxidizxer rathe than the fuel LOX doesn't
count as a cryogenic?

In common usage, it means when you start using LH2 as fuel.

No, it doesn't. LOX is always referred to as cryogenic as well.

Brian

On Sat, 14 May 2011 01:56:37 -0800, Pat Flannery
wrote:


It's somewhat like the way the British airline industry died. Rather
than a company like Boeing or Douglas going to the airlines and asking
them what they wanted, the British government would order to be built
what they thought the airlines needed, and then try to force it down
their throats.

That's part of it, but pre-Airbus, Britain just couldn't compete with
American industry, and Boeing and Douglas simply outbid them on most
airline RFPs. They held out a while longer, but they were essentially
a British version of Lockheed, and eventually left the market.
Ironically, Lockheed's departure from the market was due in large part
to the collapse of British owned Rolls Royce, which doomed their
L-1011 (an otherwise much better plane than Douglas's rival DC-10).

Brian

On May 15, 1:33*am, Robert Clark wrote:
On May 14, 10:46*pm, William Mook wrote:

...
The S1-C has a structure fraction of 12.8% - but it carried the entire
Saturn V through Max Q. *The Falcon Heavy attains a 9% structure
fraction today - so we'll use that for the first stage 7% structure in
the upper stage, and 11% structure for a reusable system.
...

That's the main problem there. SpaceX deserves major kudos for showing
you can get a mass ratio above 20 to 1 with a kerosene-fueled first
stage in the Falcon 9:

SpaceX Achieves Orbital Bullseye With Inaugural Flight of Falcon 9
Rocket
Source: SpaceX Posted Monday, June 7, 2010http://www.spaceref.com/news/viewpr.html?pid=30992

*This corresponds to a structure fraction of 5%. And SpaceX reports
that the kerosene side boosters of its Falcon Heavy will have a mass
ratio of 30 to 1:

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

*This corresponds to a structure fraction of only 3%.

* Bob Clark

I don't know if its a problem. I agree that 3% and 5% are quite low
and easily achieved. I've been attacked for saying that in the past.
You can put these lower figures in the calculation and see what effect
it has. In all fairness, 5% and 7% are achievable by deep cryogenics
applying the same improvements. These structure fractions were
achieved by the Atlas system as well as the Saturn IV-B in their
tankage subsystems.

When you start hanging things like engines, fins, sensors, avionics,
controls on them you start having fraction creep! lol. Modern MEMS
technology and IC technology allows huge improvements in this regard.
There are also some very nice structured coatings now that allow very
light weight solutions to aerodynamic heating problems. Aerogel
filled structures with chemically milled skins in tension make very
lightweight wings and fins possible. Layered aerogel metal structures
provide very efficient lightweight insulation. So, all these issues
are addressed taking a new look at things with modern tech - and
SpaceX is doing all the right things in that regard.