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Tuesday, May 29, 2012

Re: SSTO's would have made possible Arthur C. Clarke's vision of 2001.

Copyright 2012 Robert Clark

 Very interesting article on The Space Review discussing a recently
discovered copy of a 1963 TV interview with Arthur C. Clarke:

The perils of spaceflight prediction.
by Jeff Foust
Monday, December 5, 2011

http://thespacereview.com/article/1981/1

 In the interview Clarke gives some predictions of the future of space
exploration. From the standpoint of the beginnings of human
spaceflight, he suggests a manned Mars mission within 25 years, which
would have been by 1988, and Moon bases by the end of the 20th
century.
 This turned out to be too optimistic. But as I argued in this blog, this
could indeed have been technically and even financially feasible: if
it had been recognized that reusable SSTO's are possible and in fact
aren't even really hard, we would have had routine, private
spaceflight by the 1970s.
 Such wide spread, frequent launches using reusable spacecraft would
have cut the costs to space by two orders of magnitude, at least. This
would then have made the costs of lunar bases and manned Mars missions
well within the affordability range.
 The important point is that the required high efficiency engines and
lightweight stages for SSTO's already exist and have for decades. All
that is required is to marry the two together. An expendable test SSTO
could be produced, like, tomorrow. Just this one simple, cheap test
would finally make clear the fact that routine spaceflight is already
doable.


Bob Clark

SSTO's would have made possible Arthur C. Clarke's vision of 2001.

                                                      Copyright 2012 Robert Clark




Space Travel: The Path to Human Immortality?
Space exploration might just be the key to human beings surviving mass
genocide, ecocide or omnicide.
July 24, 2009
On December 31st, 1999, National Public Radio interviewed the futurist and science fiction 
genius Arthur C. Clarke. Since the author had forecast so many of the 20th Century's most 
fundamental developments, the NPR correspondent asked Clarke if anything had happened 
in the preceding 100 years that he never could have anticipated. 'Yes, absolutely,' Clarke 
replied, without a moment's hesitation. 'The one thing I never would have expected is that, 
after centuries of wonder and imagination and aspiration, we would have gone to the moon ... 
and then stopped.'
http://www.alternet.org/news/141518/space_travel:_the_path_to_human_immortality/
  I remember thinking when I first saw 2001 as a teenager and could 
appreciate it more, I thought it was way too optimistic. We could
never have huge rotating space stations and passenger flights to orbit
and Moon bases and nuclear-powered interplanetary ships by then.
 That's what I thought and probably most people familiar with the space
program thought that. And I think I recall Clarke saying once that the
year 2001 was selected as more a rhetorical, artistic flourish rather
than being a prediction, 2001 being the year of the turn of the
millennium (no, it was NOT in the year 2000.)
 However, I've now come to the conclusion those could indeed have been 
possible by 2001. I don't mean the alien monolith or the intelligent computer, 
but the spaceflights shown in the film.
 It all comes down to SSTO's. As I argued previously [1] these could
have led and WILL lead to the price to orbit coming down to the $100
per kilo range. The required lightweight stages existed since the 60's
and 70's for kerosene with the Atlas and Delta stages, and for
hydrogen with the Saturn V upper stages. And the high efficiency
engines from sea level to vacuum have existed since the 70's with the
NK-33 for kerosene, and with the SSME for hydrogen.
 The kerosene SSTO's could be smaller and cheaper and would make
possible small orbital craft in the price range of business jets, at a
few tens of millions of dollars. These would be able to carry a few
number of passengers/crew, say of the size of the Dragon capsule. But
in analogy with history of aircraft these would soon be followed by
large passenger craft.
 However, the NK-33 was of Russian design, while the required
lightweight stages were of American design. But the 70's was the time
of detente, with the Apollo-Soyuz mission. With both sides realizing
that collaboration would lead to routine passenger spaceflight, it is
conceivable that they could have come together to make possible
commercial spaceflight.
 There is also the fact that for the hydrogen fueled SSTO's, the
Americans had both the required lightweight stages and high efficiency
engines, though these SSTO's would have been larger and more
expensive. So it would have been advantageous for the Russians to
share their engine if the American's shared their lightweight stages.
 For the space station, many have soured on the idea because of the ISS 
with the huge cost overruns. But Bigelow is planning on "space hotels"
derived from NASA's Transhab[2] concept. These provide large living
space at lightweight. At $100 per kilo launch costs we could form
large space stations from the Transhabs linked together in modular
fashion, financed purely from the tourism interests. Remember the low
price to orbit allows many average citizens to pay for the cost to LEO.
 The Transhab was developed in the late 90's so it might be
questionable that the space station could be built from them by 2001.
But remember in the film the space station was in the process of being
built. Also, with large numbers of passengers traveling to space it
seems likely that inflatable modules would have been thought of
earlier to house the large number of tourists who might want a longer
stay.
 For the extensive Moon base, judging from the Apollo missions it might
be thought any flight to the Moon would be hugely expensive. However,
Robert Heinlein once said: once you get to LEO you're half way to
anywhere in the Solar System. This is due to the delta-V requirements
for getting out of the Earth's gravitational well compared to reaching
escape velocity.
 It is important to note then SSTO's have the capability once refueled
in orbit to travel to the Moon, land, and return to Earth on that one
fuel load. Because of this there would be a large market for passenger
service to the Moon as well. So there would be a commercial
justification for Bigelow's Transhab motels to also be transported to
the Moon [3].
 Initially the propellant for the fuel depots would have to be lofted
from Earth. But we recently found there was water in the permanently
shadowed craters on the Moon [4]. Use of this for propellant would
reduce the cost to make the flights from LEO to the Moon since the
delta-V needed to bring the propellant to LEO from the lunar surface
is so much less than that needed to bring it from the Earth's surface
to LEO.
 This lunar derived propellant could also be placed in depots in lunar
orbit and at the Lagrange points. This would make easier flights to
the asteroids and the planets. The flights to the asteroids would be
especially important for commercial purposes because it is estimated
even a small sized asteroid could have trillions of dollars worth of
valuable minerals [5]. The availability of such resources would make
it financially profitable to develop large bases on the Moon for the
sake of the propellant.
 Another possible resource was recently discovered on the Moon: uranium
[6]. Though further analysis showed the surface abundance to be much
less than in Earth mines, it may be that there are localized
concentrations just as there are on Earth. Indeed this appears to be
the case with some heavy metals such as silver and possibly gold that
appear to be concentrated in some polar craters on the Moon [7].
 So even if the uranium is not as abundant as in Earth mines, it may be
sufficient to be used for nuclear-powered spacecraft. Then we wouldn't
have the problem of large amounts of nuclear material being lofted on
rockets on Earth. The physics and engineering of  nuclear powered
rockets have been understood since the 60's [8]. The main impediment
has been the opposition to launching large amounts of radioactive
material from Earth into orbit above Earth. Then we very well could
have had nuclear-powered spacecraft launching from the Moon for
interplanetary missions, especially when you consider the financial
incentive provided by minerals in the asteroids of the asteroid belt.


    Bob Clark

REFERENCES

1.)The Coming SSTO's. http://exoscientist.blogspot.com/2012/05/coming-sstos.html 2.)TransHab. http://en.wikipedia.org/wiki/TransHab 3.)Private Moon Bases a Hot Idea for Space Pioneer. by Leonard David, SPACE.com's Space Insider Columnist Date: 14 April 2010 Time: 02:23 PM ET http://www.space.com/8217-private-moon-bases-hot-idea-space-pioneer.html 4.)Mining the Moon's Water: Q & A with Shackleton Energy's Bill Stone. by Mike Wall, SPACE.com Senior WriterDate: 13 January 2011 Time: 03:57 PM ET http://www.space.com/10619-mining-moon-water-bill-stone-110114.html 5.)Riches in the Sky: The Promise of Asteroid Mining. Mark Whittington, Nov 15, 2005 http://www.associatedcontent.com/article/11560/riches_in_the_sky_the_promise_of_asteroid_pg2.html?cat=58 6.)Uranium could be mined on the Moon. Uranium could one day be mined on the Moon after a Japanese spacecraft discovered the element on its surface. By Julian Ryall in Tokyo 4:58PM BST 01 Jul 2009 http://www.telegraph.co.uk/science/space/5711129/Uranium-could-be-mined-on-the-Moon.html 7.)Silver, Gold, Mercury and Water Found in Moon Crater Soil by LCROSS Project. Catherine Dagger, Oct 22, 2010 http://www.associatedcontent.com/article/5922906/silver_gold_mercury_and_water_found_pg2.html?cat=15 8.)NERVA. http://en.wikipedia.org/wiki/NERVA

Saturday, May 26, 2012

The Coming SSTO's.


Copyright  2012  Robert Clark

Credit: Gary Hudson

 

Preface

Arthur C. Clarke said:
Every revolutionary idea seems to evoke three stages of reaction. They may be summed up by the phrases:

1- It's completely impossible.
2- It's possible, but it's not worth doing.
3- I said it was a good idea all along.
 SSTO's are at level 2 in Clarke's quote. Soon they will be at level 3.

Highlights

  • Required high efficiency engines and lightweight stages for SSTO’s have existed since the 1970’s.
  • The key figure of merit for a launch vehicle should not be payload to gross mass ratio, but payload to dry mass ratio.
  • SSTO’s can have this payload to dry mass ratio greater than 1, better than any previous launcher.
  • Reusable SSTO’s can be made the size of the very light (personal) jets, thus making routine manned spaceflight possible. 

Abstract

___________________________________________________________________

 Contrary to popular belief, SSTO's are actually easy - if done in the right way. The right way
can be summarized in one sentence:
If you use the most weight efficient stages and the most fuel efficient engines at the same time,
then the result will be SSTO capable whether you intend to or not.
___________________________________________________________________

Introduction 

We all know that to get a good payload to space you want a high
efficiency engine. And we all know we want to use lightweight
structures so the weight savings can go to increased payload. So you
would think it would be obvious to use both these ideas to maximize
the payload to orbit, right?

 And indeed both have been used together - for upper stages. Yet this
fundamentally obvious concept still has not been used for first
stages
. It is my thesis that if you do this, then what you wind up
with will automatically be SSTO capable. This is true for either
kerosene fueled or hydrogen fueled stages.

 Part of the misinformation that has been promulgated is that the mass
ratio for SSTO's is some impossible number. This is false. We've had
rocket stages with the required mass ratio's since the 60's, nearly 50
years, both for kerosene and hydrogen fueled. Another part of the
misinformation is that it would require some unknown high energy fuel
and engine to accomplish. This is false. The required engines have
existed since the 70's, nearly 40 years, both for kerosene and
hydrogen fueled.

 What has NOT been done is to marry the two concepts together for
first stages. All you need to do is swap out the low efficiency
engines that have been used for the high mass ratio stages and replace
them with the high efficiency engines. It really is that simple.
This makes possible small, low cost orbital vehicles that could
transport the same number of passengers as the space shuttle, about 7,
but would have a comparable cost to a mid-sized business jet, a few
tens of millions of dollars.

 Then once you have the SSTO's they make your staged vehicles even
better because you can carry greater payload when they are used for
the individual stages of the multi-staged vehicle.

 In disseminating the false dogma that SSTO's are not possible it is
sometimes said instead that they are not practical because the payload
fraction is so small. Even this is false. And indeed this is just as
damaging as making the false statement they are not possible because
the statements are often conflated into meaning the same thing. So
when those in the industry make the statement they are not
"practical", meaning actually they are doable but not economical, this
becomes interpreted among many space enthusiasts and even many in
the industry as meaning it would require some revolutionary advance to
make them possible.

 The fact that you can carry significant payload to orbit using SSTO's
can be easily confirmed by anyone familiar with the rocket equation.
To get a SSTO with significant payload using efficient kerosene
engines you need a mass ratio of about 20 to 1. And to get a SSTO with
significant payload using efficient hydrogen engines you need a mass
ratio of about 10 to 1. Both of these the high mass ratio stages and
the high efficiency engines for both kerosene and hydrogen have
existed for decades now.

 See this list of rocket stages:

Stages - Alphabetical Index.
http://www.friends-partners.org/partners/mwade/stages/staindex.htm

 Among the kerosene-fueled stages you see that several among the Atlas
and Delta family have the required mass ratio. However, for the early
Atlas stages you have to be aware of the type of staging system they
used. They had drop-off booster engines and a main central engine,
called the sustainer that continued all the way to orbit. But even
when you take this into account you see these highly weight optimized
stages had surprisingly high mass ratios.

Atlas rocket derived SSTO

 See for instance the Atlas Agena SLV-3:

SLV-3 Atlas / Agena B.
Family: Atlas. Country: USA. Status: Hardware. Department of
Defence Designation: SLV-3.
Standardized Atlas booster with Agena B upper stage.
Specifications
Payload: 600 kg. to a: 19,500 x 103,000 km orbit at 77.5 deg
inclination trajectory.
Stage Number: 0. 1 x Atlas MA-3 Gross Mass: 3,174 kg. Empty Mass:
3,174 kg. Thrust (vac): 167,740 kgf. Isp: 290 sec. Burn time: 120 sec.
Isp(sl): 256 sec. Diameter: 4.9 m. Span: 4.9 m. Length: 0.0 m.
Propellants: Lox/Kerosene No Engines: 2. LR-89-5
Stage Number: 1. 1 x Atlas Agena SLV-3 Gross Mass: 117,026 kg.
Empty Mass: 2,326 kg.
Thrust (vac): 39,400 kgf. Isp: 316 sec. Burn
time: 265 sec. Isp(sl): 220 sec. Diameter: 3.1 m. Span: 4.9 m. Length:
20.7 m. Propellants: Lox/Kerosene No Engines: 1. LR-105-5
Stage Number: 2. 1 x Agena B Gross Mass: 7,167 kg. Empty Mass: 867
kg. Thrust (vac): 7,257 kgf. Isp: 285 sec. Burn time: 240 sec. Isp(sl): 0
sec. Diameter: 1.5 m. Span: 1.5 m. Length: 7.1 m. Propellants: Nitric
acid/UDMH No Engines: 1. Bell 8081

http://www.friends-partners.org/partners/mwade/lvs/slvgenab.htm

 Looking at only the gross/empty mass of stage 1, you would think this
stage had a mass ratio close to 50 to 1. But that is only including the
sustainer engine. The more relevant ratio would be when you add in the
mass of the jettisonable booster engines to the dry mass since they are
required to lift the vehicle off the pad. These are contained within the
stage 0 mass at 3,174 kg. This makes the loaded mass now 117,026 +
3,174 = 120,200 and the dry mass 2,326 + 3,174 = 5,500 kg, for a mass
ratio of 21.85.

 But this was using the low efficiency engines available in the early
60's. Let's swap these out for the high efficiency NK-33 [1]. The
sustainer engine used was the LR-105-5 [2] at 460 kg. At 1,220 kg the
NK-33 weighs 760 kg more. So removing both the sustainer and
booster engines to be replaced by the NK-33 our loaded mass becomes
117,786 kg and the dry mass 3,086 kg, and the mass ratio 38.2 (!)

How to calculate the delta-v the rocket can achieve

 A problem with doing these payload to orbit estimates is the lack of a
simple method for getting the average Isp over the flight for an engine,
which inhibits people from doing the calculations to realize SSTO is
possible and really isn't that hard. To calculate the delta-V achievable
I'll follow the suggestion of Mitchell Burnside Clapp who spent many
years designing and working on SSTO projects including stints with the
DC-X and X-33 programs. He argues that instead of using the average
Isp you should use the vacuum Isp and just use 30,000 feet per second,
about 9,150 m/s, as the required delta-V to orbit for dense propellants.
The reason for this is that you can just regard the reduction in Isp at sea
level and low altitude as a loss and add onto the required delta-V for
orbit this particular loss just like you add on the loss for air drag and
gravity loss [3].

 Using the vacuum Isp of 331 s and a 9,150 m/s delta-V
for a flight to orbit, we can lift 4,200 kg to orbit:

(1.) 331*9.81ln((117,786+4,200)/(3,086+4,200)) = 9,150 m/s.
This is a payload fraction of 3.4%, comparable to that of many multistage
rockets.

 Note in fact that this has a very good value for a ratio that I
believe should be regarded as a better measure, i.e., figure of merit,
for the efficiency of a orbital vehicle. This is the ratio of the
payload to the total dry mass of the vehicle. The reason why this is a
good measure is because actually the cost of the propellant is a minor
component for the cost of an orbital rocket. The cost is more
accurately tracked by the dry mass and the vehicle complexity. Note
that SSTO's in not having the complexity of staging are also good on
the complexity scale.

 For the ratio of the payload to dry mass you see this is greater than
1 for this SSTO. This is important because for every orbital vehicle I
looked at, and possibly for every one that has existed, this ratio is
going in the other direction: the vehicle dry mass is greater than the
payload carried. Often it is much greater. For instance, for the space
shuttle system, the vehicle dry mass is more than 12 times that of the
payload.

 This good payload fraction and even better payload to dry mass ratio
was just by using the engine in its standard configuration, no
altitude compensation. However, for a SSTO you definitely would want
to use altitude compensation. Dr. Bruce Dunn in his report "Alternate
Propellants for SSTO Launchers" [4] estimates an ideal vacuum Isp of
375.9 s for high performance kerosene engines. Using altitude
compensation we may suppose our engine can achieve this while still
getting good performance at sea level. Modern engines can reach
efficiencies of 97% and above of their ideal Isp. Then I’ll take the Isp
as 365 s. Then we could lift 6,500 kg to orbit:

(2.) 365*9.81ln((117,786+6,500)/(3,086+6,500)) = 9,175 m/s.

Higher payload possible with more energetic hydrocarbon fuels

 But kerosene is not the most energetic hydrocarbon fuel you could
use. Dunn in his report estimates an ideal vacuum Isp of 391.1 s for
methylacetyene. Dunn notes that Methyacetylene/LOX when densified
by subcooling gets a density slightly above that of kerolox, so I'll keep
the same propellant mass. Using altitude compensation and 97%
efficiency, I’ll take the vacuum Isp as 380 s. This would allow a
payload of 7,600 kg :

(3.) 380*9.81ln((117,786+7,600)/(3,086+7,600)) = 9,180 m/s.

 Quite key for why reusable SSTO's will make manned space travel
routine is the small size and low cost they can be produced. A manned
SSTO can be produced using currently existing engines and
stages the size of the smallest of the very light, or personal, jets
[5], except it would use rocket engines instead of jet engines, and
the entire volume aft of the cockpit would be filled with propellant,
i.e., no passenger cabin. So it would have the appearance of a fighter jet.

Falcon 1 first stage based SSTO

 We'll base it on the SpaceX Falcon 1 first stage. According to the
Falcon 1 Users Guide on p.8 [6], the first stage has a dry mass of
3,000 lbs, 1,360 kg, and a usable propellant mass of 47,380 lbs,
21,540 kg. We need to swap out the low efficiency Merlin engine for a
high efficiency engine. However, SpaceX has not released the mass for
the Merlin engine. We'll estimate it from the information here, [7].
From the given T/W ratio and thrust, I'll take the mass as 650 kg.
We'll replace it with the RD-0242-HC, [8]. This is a proposed
modification to kerosene fuel of an existing hypergolic engine. This
type of modification where an engine has been modified to run on a
different fuel has been done before so it should be doable [9], [10].
The engine mass is listed as 120 kg. We'll need two of them to loft
the vehicle. So the engine mass is reduced from that of the Merlin
engine mass by 410 kg, and the dry mass of the stage is reduced down
to 950 kg. Note that the mass ratio now becomes 23.7 to 1.

 We need to get the Isp for this case. For a SSTO you want to use
altitude compensation. The vacuum Isp of the RD-0242-HC is listed as
312 s. However, this is for first stage use so it's not optimized for
vacuum use. Since the RD-0242-HC is a high performance, i.e., high
chamber pressure, engine with altitude compensation it should get
similar vacuum Isp as other high performance Russian engines such as
the RD-0124 [11] in the range of 360 s. As a point of comparison the
Merlin Vacuum is a version of the Merlin 1C optimized for vacuum use
with a longer nozzle. This increases its vacuum Isp from 304 s to 342 s
[12]. I've also been informed by email that engine performance
programs such as Propep [13] give the RD-0242-HC an ideal vacuum
Isp of 370 s. So a practical vacuum Isp of 360 s should be reachable
using altitude compensation.

 Then with a 360 s vacuum Isp we get a delta-V of:

(4.) 360*9.81ln(1 + 21,540/950) = 11,160 m/s.

 So we can add on payload mass:

(5.) 360*9.81ln(1+21,540/(950 + 790)) = 9,150 m/s,

allowing a payload of 790 kg.

 To increase the payload we can use different propellant combinations
and use lightweight composites. For methylacetylene again, I’ll take the
vacuum Isp value as 380 s. Then the payload will be 1,070 kg:

(6.) 380*9.81ln(1 + 21,540/(950 + 1,070)) = 9,157 m/s.

 We can get better payload by reducing the stage weight by using
lightweight composites. The stage weight aside from the engines is 710
kg. Using composites can reduce the weight of a stage by about 40%.
Then adding back on the engine mass this brings the dry mass to 670
kg. So our payload can be 1,350 kg:

(7.) 380*9.81ln(1 + 21,540/(670 + 1,350)) = 9,157 m/s.

Why payload to dry mass ratio is a better launcher figure of merit

 Note this again has a very high value for what is now regarded as a
key figure of merit for the efficiency of a launch vehicle: the ratio of
the payload to the dry mass. The ratio of the payload to the gross mass
is now recognized as not being a good figure of merit for launch
vehicles. The reason is that payload mass is being compared then to
mostly what makes up only a minor proportion of the cost of a
launch vehicle, the cost of propellant. By comparing instead to the
dry mass you are comparing to the expensive components of the
vehicle, the parts that have to be constructed and tested [14].

 This vehicle in fact has the payload to dry mass ratio over 2. Every
other launch vehicle I looked at, and possibly every other one that
has ever existed, has the ratio going in the other direction, i.e.,
the dry mass is greater than the payload mass. Often it is much
greater. For example for the space shuttle system the dry mass is over
12 times that of the payload mass, undoubtedly contributing to the
high cost for the payload delivered.

 Because of this high value for this key figure of merit, this
vehicle would be useful even as a expendable launcher. However, a
SSTO is most useful as a reusable vehicle. This will be envisioned as a
vertical take-off vehicle. However, it could use either a winged
horizontal landing or a powered vertical landing. This page gives the
mass either for wings or propellant for landing as about 10% of the
dry, landed mass [15]. It also gives the reentry thermal protection
mass as 15% of the landed mass. The landing gear mass is given as 3%
of the landed mass here [16]. This gives a total of 28% of the landed
mass for reentry/landing systems. With lightweight modern materials
quite likely this could be reduced to half that.

 If you use the vehicle just for a cargo launcher with cargo left in
orbit, then the reentry/landing system mass only has to cover the dry
vehicle mass so with lightweight materials perhaps less than 100 kg
out of the payload mass has to be taken up by the reentry/landing
systems. For a manned launcher with the crew cabin being returned, the
reentry/landing systems might amount to 300 kg, leaving 1,100 kg for
crew cabin and crew. As a mass estimate for the crew cabin, the single
man Mercury capsule only weighed 1,100 kg [17]. With modern
materials this probably can be reduced to half that.

Cost estimates comparable to a mid size business jet

 For the cost, the full two stage Falcon 1 launcher is about $10
million. The engines make up the lion share of the cost for launchers.
So probably much less than $5 million just for the 1st stage sans
engine. Composites will make this more expensive but probably not
much more than twice as expensive. For the engine cost, Russian
engines are less expensive than American ones. The RD-180 at
1,000,000 lbs vacuum thrust costs about $10 million [18], and the NK-
43 at a 400,000 lbs vacuum thrust costs about $4 million [19]. This is in
the range of $10 per pound of vacuum thrust. On that basis we might
estimate the cost of the RD-0242-HC of about 30,000 lbs vacuum thrust
as $300,000. We would need two of them for $600,000.

 I'm informed though this was based on ca. year 2000 prices and the
prices have approximately doubled since then, [20]. Even so the price
for two of these engines is likely to be less than $2 million.

 So we can estimate the cost of the reusable version as significantly
less than $12,000,000 without the reentry/landing system costs. These
systems added on for reusability at a fraction of the dry mass of the
vehicle will likely also add on a fraction on to this cost. Keep in
mind also that the majority of the development cost for the two stage
Falcon 1 went to development of the engines so in actuality the cost
of just the first stage without the engine will be significantly less
than half the full $10 million cost of the Falcon 1 launcher. The cost
of a single man crew cabin is harder to estimate. It is possible it
could cost more than the entire launcher. But it's likely to be less
than a few 10's of millions of dollars.


    Bob Clark

UPDATE, Sept. 26, 2013:

 See more accurate calculations using Dr. John Schillings Launch Performance Calculator here:

The Coming SSTO's: Page 2.
http://exoscientist.blogspot.com/2013/08/the-coming-sstos-page-2.html

REFERENCES.

1.)NK-33.
http://www.friends-partners.org/partners/mwade/engines/nk33.htm

2.)LR-105-5.
http://www.friends-partners.org/partners/mwade/engines/lr1055.htm

3.) Newsgroups: sci.space.policy
From: Mitchell Burnside Clapp <cla...@plk.af.mil>
Date: 1995/07/19
Subject: Propellant desity, scale, and lightweight structure.
http://groups.google.com/group/sci.space.policy/browse_frm/thread/3d981607d59684dc/

4.)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

5.)List of very light jets.
http://en.wikipedia.org/wiki/List_of_very_light_jets

6.)Falcon 1 Users Guide.
http://www.spacex.com/Falcon1UsersGuide.pdf

7.)Merlin (rocket engine)
4. Merlin 1C Engine specifications
http://en.wikipedia.org/wiki/Merlin_%28rocket_engine%29#Merlin_1C_Engine_specifications

8.)RD-0242-HC.
http://www.astronautix.com/engines/rd0242hc.htm

9.)LR-87.
http://en.wikipedia.org/wiki/LR-87

10.)Pratt and Whitney Rocketdyne's RS-18 Engine Tested With Liquid Methane.
by Staff Writers
Canoga Park CA (SPX) Sep 03, 2008
http://www.space-travel.com/reports/Pratt_and_Whitney_Rocketdyne_RS_18_Engine_Tested_With_Liquid_Methane_999.html

11.)RD-0124.
http://www.astronautix.com/engines/rd0124.htm

12.)Merlin (rocket engine).
2.5 Merlin Vacuum
http://en.wikipedia.org/wiki/Merlin_(rocket_engine)#Merlin_Vacuum

13.)Propep
http://www.spl.ch/software/index.html

14.)A Comparative Analysis of Single-Stage-To-Orbit Rocket and Air-Breathing Vehicles.
p. 5, 52, and 67.
http://govwin.com/knowledge/comparative-analysis-singlestagetoorbit-rocket-and/15354

15.)Reusable Launch System.
http://en.wikipedia.org/wiki/Reusable_launch_system#Horizontal_landing

16.)Landing gear weight (Gary Hudson; George Herbert; Henry Spencer).
http://yarchive.net/space/launchers/landing_gear_weight.html

17.)Mercury Capsule.
http://www.astronautix.com/craft/merpsule.htm

18.)Wired 9.12: From Russia, With 1 Million Pounds of Thrust.
http://www.wired.com/wired/archive/9.12/rd-180.html

19.)A Study of Air Launch Methods for RLVs.
Marti Sarigul-Klijn, Ph.D. and Nesrin Sarigul-Klijn, Ph.D.
AIAA 2001-4619 , p.13
http://mae.ucdavis.edu/faculty/sarigul/aiaa2001-4619.pdf

20.)Personal communication, Gary Hudson.

Thursday, May 10, 2012

On Commercial Flights to the ISS and "space tugs".

Copyright 2012 Robert Clark 


The SpaceX flight to the International Space Station has been delayed multiple times:

1st Private Rocket Launch to Space Station Delayed. 
by Denise Chow, SPACE.com Staff Writer 
Date: 16 January 2012 Time: 01:51 PM ET
http://www.space.com/14251-launch-delay-spacex-dragon-spaceflight.html

SpaceX launch to station faces delay.
By WILLIAM HARWOOD 
CBS News 
"KENNEDY SPACE CENTER, FL--The long-awaited launch of a commercial 
cargo ship bound for the International Space Station almost certainly 
will be delayed from May 7 to at least May 10 and possibly longer, 
sources said late Tuesday, to give company engineers additional time 
to complete pre-flight tests and checkout..." 
05/01/2012 11:18 PM Filed in: Space News | Commercial Space | International Space Station 
http://www.cbsnews.com/network/news/space/home/spacenews/files/3f8a43db532f88e3f0ed6b8fe9876eeb-403.html

 SpaceX has said they need to do additional testing of the software controlling the link-up, understandable when a $100 billion asset in the ISS could be at risk.
 It would seem prudent for SpaceX to do several test runs rendezvousing with orbiting satellites before attempting the link up with the ISS. The current plan is not for Dragon to perform the final link-up with ISS under its own power and navigation capabilities but just to get close enough for the robot arm to grapple it and pull it to the station to dock with it. So to test this, all the Dragon has to do is demonstrate the ability to get close enough to some orbiting satellite without colliding with it, to within a similar distance as it would be to the ISS. It might be able to do this several times with different satellites to further demonstrate this capability.  
 With several commercial companies proposing private spacecraft to dock with the ISS this would appear to be a common problem. Robert Zubrin proposed an interesting solution in his book Entering Space: Creating a Spacefaring Civilization. He suggests having a "space tug" attached to the the station specifically to ferry in approaching cargo and crew spacecraft:

 "NASA managers are understandably jittery about the idea of just anybody flying up and attempting to dock their Radio Shack-wired cheapo spacecraft with one of these assets.
 "One solution that might get past this problem would be the development, by NASA, of a proximity operations vehicle (a "proxops" stage) that would hang around the station and grab resupply payloads delivered to the general vicinity by commercial suppliers."
Entering Space, p. 68.

 This would provide another use of the "space tugs" that have been proposed for orbital satellite servicing:

Satellite Refueling in Orbit, Coming Soon? 
By Steve Rousseau 
October 17, 2011 5:00 PM 
http://www.popularmechanics.com/technology/military/satellites/satellite-refueling-in-orbit-coming-soon

Space Infrastructure Servicing. 
http://en.wikipedia.org/wiki/Space_Infrastructure_Servicing 

thus providing another income stream to the company(s) that produce the tugs.
  As for the satellite servicing business plan, according to the "Satellite Refueling in Orbit, Coming Soon?" article a satellite may last 10 to 15 years. And according to the "Space Infrastructure Servicing" wikipedia page, 200 kg of fuel may provide an additional 2 to 4 years of life. So it might take 100 kg per year for fuel, and over 10 years would require 1,000 kg.
 The cost to get anything to GEO, including this fuel, is in the range of $20,000 to $25,000 per kg. So for 1,000 kg of fuel to get to GEO for satellite refueling it would cost perhaps $25 million. But this would double the life of the satellite since it would again have a full fuel load for 10 years. So for $25 million you saved the satellite companies from paying, say, $300 - $500 million, to purchase and launch a new satellite.
 So even if you charged 4 times the usual price to get to GEO for this fuel, the satellite companies could still consider this a bargain.
 You would need a small reusable servicing spacecraft to be launch from the payload bay of the launch vehicle to transport the fuel to GEO. If you use LH2/LOX propellant for this spacecraft like the Centaur upper stages, then it takes about the same amount of propellant to get from LEO to GEO, as the mass of the (spacecraft + payload), the payload being the refueling fuel in this case. The dry mass of the spacecraft is only a small proportion of the propellant as indicated by the Centaur upper stage, about 1/10th.
 So a ca. 20,000 kg cargo capacity of the current largest launchers could be made up of half LH2/LOX propellant for the space tug and half the refueling fuel for the satellites. That's 10,000 kg of refueling fuel. If you do charge the satellites companies 4 times the usual rate to GEO to $100,000 per kg for this fuel, then that's potentially $1 billion revenue from that one launcher flight.
 The estimate of a charge of $100,000 per kg of the refueling fuel to be delivered to satellites in GEO may seem high but it's actually less than a price that has been quoted by a company planning on doing such refueling missions. On that "Space Infrastructure Servicing" wikipedia
page is given this ref. to an article on satellite servicing:

de Selding, Peter B. (2011-03-14). "Intelsat Signs Up for Satellite Refueling Service".
Space News. Retrieved 2011-03-15. "If the MDA 
spacecraft performs as planned, Intelsat will be paying a total of
some $200 million to MDA. This assumes that four or five satellites
are given around 200 kilograms each of fuel. ... The maiden flight of
the vehicle would be on an International Launch Services Proton
rocket, industry officials said. One official said the MDA spacecraft,
including its 2,000 kilograms of refueling propellant, is likely to
weigh around 6,000 kilograms at launch."

http://www.spacenews.com/satellite_telecom/intelsat-signs-for-satellite-refueling-service.html

 So for this company they are actually charging $200,000 per kg of the refueling fuel.




      Bob Clark 

Monday, May 7, 2012

Low Cost HLV, page 2: Comparison to the S-IC Stage.

Copyright 2012 Robert Clark 

The dry mass of the first stage of the vehicle described in the Low Cost HLV post was 1/20th the gross mass of the stage at 110,000 kg. Interestingly the propellant mass of the S-IC first stage of the Saturn V was about the same as in this HLV proposal, about 2,100,000 kg. Then it will be interesting to make a comparison to the dry mass of the S-IC stage. This page gives it as 130,000 kg:

Ground Ignition Weights.
http://history.nasa.gov/SP-4029/Apollo_18-19_Ground_Ignition_Weights.htm

 However, the Saturn V had a quite heavy first stage thrust structure:

SP-4206 Stages to Saturn.
7. The Lower Stages: S-IC and S-II.
Rosen apparently took the lead in pressing for the fifth engine, consistent with his obstinate push for a "big rocket." The MSFC contingent during the meetings included William Mrazek, Hans Maus, and James Bramlet. Rosen argued long and hard with Mrazek, until Mrazek bought the idea, carried the argument to his colleagues, and together they ultimately swayed von Braun. Adding the extra power plant really did not call for extensive design changes; this was Rosen's most convincing argument. Marshall engineers had drawn up the first stage to mount the original four engines at the ends of two heavy crossbeams at the base of the rocket. The innate conservatism of the von Braun design team was fortunate here, because the crossbeams were much heavier than required. Their inherent strength meant no real problems in mounting the fifth powerplant at the junction of the crossbeams, and the Saturn thus gained the added thrust to handle the increasingly heavy payloads of the later Apollo missions. "Conservative design," Rosen declared, "saved Apollo."2
http://history.nasa.gov/SP-4206/ch7.htm

Indeed it was heaviest single component of the S-IC stage, and so of the Saturn V:

S-IC.
2. Components.
http://en.wikipedia.org/wiki/S-IC#Components

 This online lecture of Dr. David Akin of the University of Maryland gives mass estimating relationships for various rocket components, taken from the reports NASA uses in designing rockets:


Mass Estimating Relations.
• Review of iterative design approach
• Mass Estimating Relations (MERs)
• Sample vehicle design analysis
http://spacecraft.ssl.umd.edu/academics/483F09/483F09L13.mass_est/483F09L13.MER.pdf

 On page 17 is given a relation between the thrust of the stage in Newtons and the mass of the thrust structure in kilograms:

MThrust structure(kg) = 2.55×10−4T(N)

 For 4 RD-171's at 7,900 kN each, this would be 8,000 kg. So we can subtract off 13,000 kg from that S-IC dry mass to get 117,000 kg. We're also using one less engine, so subtract off 8,350 kg for the one less engine. However, the RD-171 weighs about 1,000 kg more than the F-1, so add on 4,000 kg to get about 113,000 kg for the stage, quite close to the 20 to 1 mass ratio estimate. Note this is even without the weight saving alloys and composites now used.




  Bob Clark


Wednesday, May 2, 2012

SpaceX Dragon spacecraft for low cost trips to the Moon.

Copyright 2012 Robert Clark
OTV - 1986 
Space Tug depicted in NASA's 'Pioneering Space Frontier'. A space tug from the 1986 Pioneering the Space Frontier policy report. 
Credit: NASA via Marcus Lindroos


 SpaceX has said two Falcon Heavy launches would be required to carry a manned Dragon to a lunar landing. However, the 53 metric ton payload capacity of a single Falcon Heavy would be sufficient to carry the 40 mT (Earth departure stage + lunar lander) system described below. This would require 30 mT and 10 mT gross mass Centaur-style upper stages. This page gives the cost of a ca. 20 mT Centaur upper stage as $30 million:

Centaur IIA.

 A 30 mT Centaur would be somewhat more than this and a 10 mT Centaur-style stage would be somewhat less, so the total for both about $60 million. Note it could also be done with two 20 mT Centaurs at slightly reduced performance.

 The 53 mT to LEO capacity of the Falcon Heavy would also allow large lunar cargo transport using two of the 20 mT gross mass Centaurs that already exist either using the Dragon to carry the cargo or by carrying somewhat more cargo just within a lightweight container.

 An important cargo delivery to the Moon would be in-situ resource utilization (ISRU) equipment, specifically for producing propellant from the water discovered to lie within the shadowed craters near the lunar poles. Elon Musk has said a key goal of his is to mount a manned Mars mission within 1 to 2 decades. Such a mission could be mounted more cheaply if the large amount of propellant required did not have to be lofted from the Earth's deep gravity well but could be taken from the Moon.

 Another important cargo delivery would be to carry a rover that could do a sample return mission from the near polar locations. Lunar orbiter observations suggest there may be valuable minerals concentrated in such locations:

SCIENCE -- October 21, 2010 at 2:05 PM EDT
Moon Blast Reveals Lunar Surface Rich With Compounds.
BY: JENNY MARDER
"There is water on the moon ... along with a long list of other compounds,
including, mercury, gold and silver. That's according to a more detailed
analysis of the chilled lunar soil near the moon's South Pole, released as six papers by a large team of scientists in the journal, Science Thursday."

 If these tentative detections could be confirmed then that could possibly form a commercial market for flights to the Moon.

 In this vein note there is even stronger evidence for large amounts of valuable minerals on asteroids. Observations suggest that even a small size asteroid could contain trillions of dollars (that's trillions with a 't') worth of valuable minerals:

Riches in the Sky: The Promise of Asteroid Mining.
Mark Whittington, Nov 15, 2005

 It is quite important to note then that since the delta-V requirements to some near Earth asteroids is less than that to the Moon, that the sample return version of the lunar lander could also be used to return samples from the near Earth asteroids. If these asteroidal detections could be definitively confirmed by a sample return mission then that would provide further justification for private investment in lunar propellant production installations.

 SpaceX expects to launch the first Falcon Heavy in 2013. Because the required Centaur stages already exist it is possible that a lunar lander could be formed from such mated together stages within this time frame at least for a unmanned cargo version.

 It is important though that such a lander be privately financed. Because the required stages already exist I estimate a lander could be formed from them for less than a $100 million development cost. This is based on the fact that SpaceX was able to develop the Falcon 9 launcher for about $300 million development cost. And this required development of both the engines and the stages for a 300 mT gross mass and 30 mT dry mass launcher. But for this lunar lander, the engines and stages already exist for a total 40 mT gross mass and 4 mT dry mass system.

 If the system were to be government financed then based on the fact that SpaceX was able to develop the Falcon 9 for 1/10th the development cost of usual NASA financed systems, the cost of the lander would suddenly balloon to a billion dollar development.

 Note that while the evidence for valuable minerals in the lunar shadowed craters is not yet particularly strong, the evidence for such minerals in the asteroids is. So there is a strong financial incentive for forming such a lunar lander as it could also be used for the asteroidal lander.

 But asteroidal mineral retrieval flights could be launched much more cheaply if the propellant could be obtained from the Moon. Then there is a strong financial incentive to produce ISRU installations on the Moon which would require lunar return missions from the shadowed crater regions to assess the best means of harvesting this lunar water for propellant. If such return missions also confirm the presence of valuable minerals in the shadowed craters then that would be like icing on the cake for justification of private investment in such missions.



    Bob Clark


=====================================================================
 The Orion spacecraft and Altair lunar lander intended for a manned Moon mission are large craft that would require a heavy lift launcher for the trip. However the Dragon capsule is a smaller capsule that would allow lunar missions with currently existing launchers.
The idea for this use would be for it to act as a reusable shuttle only between LEO and the lunar surface. This page gives the dry mass of the Dragon capsule of 3,180 kg:

SpaceX reveals first Dragon engineering unit.
DATE:16/03/07
By Rob Coppinger

The wet mass with propellant would be higher than this but for use only as a shuttle between LEO and the Moon, the engines and propellant would be taken up by the attached propulsion system. With crew and supplies call the capsule mass 4,000 kg.
On this listing of space vehicles you can find that the later versions of the Centaur upper stage have a mass ratio of about 10 to 1:


The Isp's given for the RL-10A engines used on these stages are around 450 s, but an updated version with a longer, extensible nozzle has an Isp of 465.5 s:

RL10B-2.

This page gives the delta-V's needed for trips within the Earth-Moon system:

Delta-V budget.
Earth–Moon space.


The architecture will be to use a larger Centaur upper stage to serve as the propulsion system to take the vehicle from LEO to low lunar orbit. This larger stage will not descend to the surface, but will remain in orbit. A smaller Centaur stage will serve as the descent stage and will also serve as the liftoff stage that will take the spacecraft not just back to lunar orbit, but all the way to back to LEO. The larger Centaur stage will return to LEO under its own propulsion, to make the system fully reusable. Both stages will use aerobraking to reduce the delta-V required to return to LEO.
For the larger Centaur, take the gross mass of the stage alone as 30,000 kg, and its dry mass as 1/10th of that at 3,000 kg. For the smaller Centaur stage take the gross mass as 10,000 kg and the dry mass as 1,000 kg. The "Delta-V budget" page gives the delta-V from LEO to low lunar orbit as 4,040 m/s. In calculating the delta-V provided by the larger Centaur stage we'll retain 1,000 kg propellant at the end of the burn for the return trip of this stage to LEO: 465.5*9.8ln((30,000 + 10,000 + 4,000)/(3,000 +10,000 + 4,000 + 1,000)) = 4,077 m/s, sufficient to reach low lunar orbit. For this stage alone to return to LEO, 1,310 m/s delta-V is required. The 1,000 kg retained propellant provides 465.5*9.8ln((3,000 + 1,000)/3,000) = 1,312 m/s, sufficient for the return.
The delta-V to go from low lunar orbit to the Moon's surface is 1,870 m/s. And to go from the Moon's surface back to LEO is 2,740 m/s, for a total of 4,610 m/s. The delta-V provided by this smaller Centaur stage is 465.5*9.8ln((10,000 + 4,000)/(1,000 + 4,000)) = 4,697 m/s, sufficient for lunar landing and the return to LEO.
The RL-10 engine was proven to be reusable for multiple uses with quick turnaround time on the DC-X. The total propellant load of 40,000 kg could be lofted by two 20,000+ kg payload capacity launchers, such as the Atlas V, Delta IV Heavy, Ariane 5, and Proton.
The price for these launchers is in the range of $100-140 million according to the specifications on this page:

Expendable Launch Vehicles.

So two would be in the range of $200-$280 million. The Dragon spacecraft and Centaur stages being reusable for 10+ uses would mean their cost per flight should be significantly less than this. This would bring the cost into the range affordable to be purchased by most national governments.
Still, it would be nice to reduce that $200 million cost just to bring the propellant to orbit. One possibility might be the heavy lift launchers being planned by NASA. One of the main problems in deciding on a design for the launchers is that there would be so few launches the per launch cost would be too high. However, launching of the propellant to orbit for lunar missions would provide a market that could allow multiple launches per year thus reducing the per launch cost of the heavy lift launchers. For instance, the Direct HLV team claims their launcher would cost $240 million per launch if they could make 12 launches per year:

JULY 23, 2009
Interview with Ross Tierney of Direct Launch by Sander Olson.

This launcher would have a 70,000 kg payload capacity. However, if you removed the payload fairing and interstage and just kept the propellant to be launched to orbit in the ET itself and considering the fact that the shuttle system was able to launch 100,000+ kg to orbit with the shuttle and payload, it's possible the propellant that could be launched to orbit could be in the range of 100,000 kg. Then the cost per kg to orbit would be $2,400 per kg, or about a $100 million cost for the propellant to orbit.
Reduction of the per launch cost for the heavy lift launchers would then allow affordable launches of the larger spacecraft and landers for lunar missions.


Bob Clark

=====================================================================

Tuesday, May 1, 2012

Low Cost HLV.

Copyright 2012 Robert Clark
Credit: Modified from:
NASA's Space Launch System - Winners and Losers
by Ed Kyle, 06/17/2011

http://www.spacelaunchreport.com/sls4.html

 The announcement by two separate teams backed by highly regarded scientists and entrepreneurs for asteroidal or lunar mining means that quite likely there will be a significant market for super heavy lift. Note too that there were separate shuttle privatization plans with business models that involved privately investing perhaps $2 billion to produce a "shuttle 2.0". Quite key here though is a vehicle this size could serve as a super HLV without the ca. 80 mT shuttle orbiter. 

  I think at this point it is abundantly clear NASA can not be expected to make a cost effective launcher. An internal NASA estimate put the total development and launch cost of just four of the interim 70 mT SLS vehicles as $41 billion, which amounts to over $10 billion per launch.

 SpaceX has shown by using good cost-saving business practices to be able to produce a launcher at greatly reduced costs. They estimate their upcoming Falcon Heavy will break the $1,000 per pound barrier, or $2,000 per kg. Keep in mind then that increasing the size of your launcher is supposed to reduce your per kg costs. So likewise using good business practices, a super heavy lift launcher privately developed should be able to at least match this or exceed it. This would be in the range of $200 million per launch for a ca. 100 mT launcher, a radically reduced cost over that of the SLS. 

  I think consideration should also be given to an all liquid system. You would use the DIRECT team's Jupiter HLV hydrogen-fueled upper stage but instead of using the shuttle ET for the first stage to hold hydrolox, use the same sized tank for kerolox. This would give a super heavy lift vehicle without the SRB's.

 The DIRECT team wanted to use the same size tank to save on costs since you can use the same existing tooling in this case. However, a key fact is this will still be the case even if you switch to kerolox propellant. You would have to change the location of the divider between the fuel and oxidizer of course, but this is comparatively low cost compared to producing whole new tooling for a different size tank.

Now kerolox is a denser propellant so you are going to get a higher propellant load in this case. The density is about 3 times that of hydrolox, so lets say the propellant load of the first stage is now 2,100 mT. What about the dry mass? 

 At this point I think we should take note of the lightweight characteristics of the Falcon 9 that SpaceX was able to achieve. SpaceX has said the first stage of the Falcon 9 has a mass ratio better than 20 to 1. SpaceX did this by using well known techniques such as a common bulkhead design for the tanks. So we could follow this also to minimize first stage dry mass.

 Also, note that by scaling our propellant tanks up, we actually improve our mass ratio. So likely we can get an even better mass ratio than this for our large first stage. But using the 20 to 1 figure, or 19 to 1 for propellant to dry mass ratio, we get a dry mass of 110 mT.

 We need heavy thrust kerosene engines. I'll use the RD-171, with a sea level thrust of about 1,700,000 lbs, and vacuum Isp of 337 s. This will require 4 of the engines. This could be replaced later with the F-1 but using the RD-171 would allow you to start now on the vehicle development with better performance.

 For the specifications of the upper stage, the DIRECT team went through several versions of their Jupiter super heavy lift vehicle. I'll use the one they referred to as Jupiter-246 Heavy, LV 41.5004.08001. For whatever reason, the DIRECT team no longer has the specifications for their vehicles up on their web site. This version's specifications are within this post to the SpaceFellowship.com forum:


Re: An SSTO as "God and Robert Heinlein intended".
Posted on: Sat Mar 12, 2011 9:49 pm
http://spacefellowship.com/Forum/viewtopic.php?p=44979&sid=3d11bfaff22840cdcd239a4c452c48d6#p44979

though you'll have to register on that forum to view it.

 This version used a propellant mass of 190,849 kg, a dry mass of 11,825 kg and 6 of the RL-10B2 engines, with an Isp of 459 s. Other versions have used the new J-2X engine, but just 6 RL-10's are likely to be cheaper. 

 This version's interstage and payload fairing were at about 4,000 kg each. We'll round off the upper stage propellant mass to 190 mT and dry mass to 11 mT. Then using a 9,150 m/s delta-V to orbit we can estimate the payload to orbit as 145 mT:

337*9.81ln(1+2100/(110+201+4+4+145)) + 459*9.81ln(1+190/(11+4+145))=9,176 m/s

 Admittedly, this payload estimate seems high so I plugged some numbers into John Schilling's "Launch Vehicle Performance Calculator" and got:

Mission Performance:Launch Vehicle: User-Defined Launch Vehicle
Launch Site: Cape Canaveral / KSC
Destination Orbit: 185 x 185 km, 28 deg
Estimated Payload: 102573 kg
95% Confidence Interval: 86317 - 121993 kg

 So it's still, likely, ca. 100 mT.

 There are several variations on this theme. For example to save on development costs we could use the Ariane 5 core stage as the upper stage. Since the ESA was amenable to using it for an upper stage for a re-booted Ares I, i.e., ATK's "Liberty" rocket, they would likely be amenable to this as well. You could also make this be parallel staging with cross feed fueling to improve performance.

 Another possibility would be to make the upper stage also be kerosene-fueled. Say you used the same light-weight tooling and tank diameter for the hydrogen fueled upper stage but using now kerolox propellant. Again, you could improve performance by making it parallel staged with cross-feed fueling. But this has an additional advantage in that you could take one of the engines off the first stage to use it for the upper stage. This would result in a lower dry weight for the first stage. The upper stage though would then be somewhat overpowered using a RD-171, so it may suffice instead to use a RD-180 just for the upper stage.

 



   Bob Clark