Re: A kerosene-fueled X-33 as a single stage to orbit vehicle.
Robert Clark
here it is again:
The Coming SSTO'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 [i]currently existing[/i] engines and
stages the size of the smallest of the very light, or personal, jets
[1], 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.
We'll base it on the SpaceX Falcon 1 first stage. According to the
Falcon 1 Users Guide on p.8 [2], 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, [3].
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, [4]. 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 [5], [6].
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 [7] 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 [8]. I've also been informed by email that engine performance
programs such as Propep [9] 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.
For the sea level Isp of the RD-0242-HC, again the version of the
high performance, high chamber pressure, RD-0124 with a shortened
nozzle optimized for sea level operation gets a 331 s Isp. So I'll
take the sea level Isp as this value using altitude compensation that
allows optimized performance at all altitudes.
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 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
[10]. 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. Then with a 360 s vacuum Isp we get a
delta-V of 360*9.8ln(1 + 21,540/950) = 11,160 m/s. So we can add on
payload mass: 360*9.8ln(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. Dr. Bruce Dunn wrote a report showing
the payload that could be delivered using high energy density
hydrocarbon fuels other than kerosene [11]. For methylacetylene he
gives an ideal vacuum Isp of 391.1 s. High performance engines can get
get ca. 97% and above of the ideal Isp so I'll take the vacuum Isp
value as 384 s. 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. Then the payload will be 1,120 kg:
384*9.8ln(1 + 21,540/(950 + 1,120)) = 9,160 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 1,400 kg: 384*9.8ln(1 + 21,540/(670 +
1,400)) = 9,160 m/s.
Note this 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 [i]dry mass[/i]. 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 [12].
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 [13]. 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 [14]. 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 [15 ]. With modern materials
this probably can be reduced to half that.
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 [16], and the NK-43 at a 400,000 lbs
vacuum thrust costs about $4 million [17]. 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
need two of them for $600,000.
So we can estimate the cost of the reusable version as significantly
less than $10,600,000 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.
REFERENCES.
1.)List of very light jets.
http://en.wikipedia.org/wiki/List_of_very_light_jets
2.)Falcon 1 Users Guide.
http://www.spacex.com/Falcon1UsersGuide.pdf
3.)Merlin (rocket engine)
4 Merlin 1C Engine specifications
http://en.wikipedia.org/wiki/Merlin_(rocket_engine)#Merlin_1C_Engine_specifications
4.)RD-0242-HC.
http://www.astronautix.com/engines/rd0242hc.htm
5.)LR-87.
http://en.wikipedia.org/wiki/LR-87
6.)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
7.)RD-0124.
http://www.astronautix.com/engines/rd0124.htm
8.)Merlin (rocket engine).
2.5 Merlin Vacuum
http://en.wikipedia.org/wiki/Merlin_(rocket_engine)#Merlin_Vacuum
9.)Propep
http://www.spl.ch/software/index.html
10.)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/945baea33c95a22?hl=en
11.)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
12.)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
13.)Reusable Launch System.
http://en.wikipedia.org/wiki/Reusable_launch_system#Horizontal_landing
14.)Landing gear weight (Gary Hudson; George Herbert; Henry Spencer).
http://yarchive.net/space/launchers/landing_gear_weight.html
15.)Mercury Capsule.
http://www.astronautix.com/craft/merpsule.htm
16.)Wired 9.12: From Russia, With 1 Million Pounds of Thrust.
http://www.wired.com/wired/archive/9.12/rd-180.html
17.)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
Robert Clark
doesn't carry much payload. It wouldn't have to be anything extensive
like perhaps what Boeing is planning with their X-37B derived SSTO.
A small one could be demonstrated by amateur science or technical
organizations, for instance by the British Interplanetary Society, or
the Planetary Society.
The Planetary Society is spending about $5.8 million total on their
two attempts at solar sail demonstators:
Cosmos 1.
http://en.wikipedia.org/wiki/Cosmos_1
LightSail-1.
http://en.wikipedia.org/wiki/LightSail-1#Creation
A small SSTO demonstrator that could carry a few hundred pound
payload could be developed for less than this amount and would be far
more important for it would show that low cost SSTO's are possible.
In fact the organization developing it could even make money on it
because they could use it to launch small scientific payloads.
For the purpose of just making the demonstration it might work to
make the vehicle half the size of the one I described in the post
below.
So it would use one RD-0242 engine, have a propellant load about
10,000 kg, and, perhaps, have a dry weight of 475 kg. However, vehicle
dry weights don't scale linearly. Scaling a vehicle up actually
improves your mass ratio. So by making the vehicle half-scale we
probably would not get as good a mass ratio, i.e., the dry mass would
likely be more than just half that of the full sized vehicle.
In addition to the amateur science organization funded test SSTO's,
it might be funded as an X-prize competition. This might have the same
effect as the Ansari X-Prize had in spurring commercial suborbital
ventures. It would spur manned commercial orbital ventures.
However, these would need high performance turbopump fed engines.
This is an entire level of difficulty above that of the suborbital
rockets which just use pressure-fed engines. In fact the complexity of
turbopump fed engines have led rocket engineers to opine "orbital
launchers are turbopump developments with rockets attached".
I recommend teams attempting the venture engage in partnerships with
Aerojet or Pratt & Whitney who have experience with high chamber
pressure, turbopump-fed engines, especially of the Russian type. They
both also have experience in converting an engine from one fuel to
another, Aerojet with the conversion of the Titan II engines from
kerosene to hypergolics, and Pratt & Whitney more recently with the
conversion of the Apollo lunar lander engines from hypergolics to
methane.
Their costs would be partially defrayed by the amount of the X-prize.
This prize amount should at least be that of the $30 million total
prize money offered for the Google Lunar X-Prize competition, since
its importance greatly exceeds it. Note too for such prize
competitions the amount spent by the teams often exceeds that offered
by the prize. They could also be offered a portion of the profits that
would come from development of the vehicles as small payload orbital
launchers.
For this prototype test vehicle you probably would not need to use
the SpaceX weight optimized Falcon 1 first stage since you just want
to get positive payload to orbit. Interestingly I found that Armadillo
Aerospace has successfully used common bulkhead design which saves
significantly on tank weight for their suborbital test rockets. They
would be a good choice for a low cost stage.
However, Armadillo has not been successful in their last two
suborbital test flights, apparently due to failures in guidance and
control. Though Armadillo apparently has solved this for hovering
vehicles, it is a significantly more difficult problem for a vehicle
traveling at high speed. I recommend a partnership with the MIT Draper
labs. They did the G & C for the Apollo missions. More recently they
are engaged in partnerships to win the Google Lunar X-Prize.
Bob Clark
> 2.)Falcon 1 Users Guide.http://www.spacex.com/Falcon1UsersGuide.pdf
>
> 3.)Merlin (rocket engine)
> 4 Merlin 1C Engine specificationshttp://en.wikipedia.org/wiki/Merlin_(rocket_engine)#Merlin_1C_Engine_...
>
> 4.)RD-0242-HC.http://www.astronautix.com/engines/rd0242hc.htm
>
> 5.)LR-87.http://en.wikipedia.org/wiki/LR-87
>
> 6.)Pratt and Whitney Rocketdyne's RS-18 Engine Tested With Liquid
> Methane.
> by Staff Writers
> Canoga Park CA (SPX) Sep 03, 2008http://www.space-travel.com/reports/Pratt_and_Whitney_Rocketdyne_RS_1...
>
> 7.)RD-0124.http://www.astronautix.com/engines/rd0124.htm
>
> 8.)Merlin (rocket engine).
> 2.5 Merlin Vacuumhttp://en.wikipedia.org/wiki/Merlin_(rocket_engine)#Merlin_Vacuum
>
> 9.)Propephttp://www.spl.ch/software/index.html
>
> 10.)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/3d9...
>
> 11.)Alternate Propellants for SSTO Launchers
> Dr. Bruce Dunn
> Adapted from a Presentation at:
> Space Access 96
> Phoenix Arizona
> April 25 - 27, 1996http://www.dunnspace.com/alternate_ssto_propellants.htm
>
> 12.)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-r...
>
> 13.)Reusable Launch System.http://en.wikipedia.org/wiki/Reusable_launch_system#Horizontal_landing
>
> 14.)Landing gear weight (Gary Hudson; George Herbert; Henry Spencer).http://yarchive.net/space/launchers/landing_gear_weight.html
>
> 15.)Mercury Capsule.http://www.astronautix.com/craft/merpsule.htm
>
> 16.)Wired 9.12: From Russia, With 1 Million Pounds of Thrust.http://www.wired.com/wired/archive/9.12/rd-180.html
Robert Clark
difficult than suborbital flight, as measured for example by the
energy requirements. On that basis a prize for a commercial manned
flight to orbit might be 10 times of the suborbital X-Prize, so to
$100 million. This is actually probably doable considering that the
required engines and stages already exist to do it as an SSTO.
Another source for funding might be Bigelow Aerospace. Bigelow had
offered a prize in 2004 of $50 million for a commercial reusable
manned launcher to orbit. The prize though expired in January 2010
with no takers:
America's Space Prize.
http://en.wikipedia.org/wiki/America's_Space_Prize
However, the original Orteig Prize for a non-stop cross Atlantic
flight also expired with no takers. It was the second offer of the
prize for an additional 5 year period which was won by Lindbergh.
Then Bigelow could offer the manned space flight prize for an
additional 5 year period. But the original conditions for the prize
were probably too ambitious. Bigelow appeared to want manned transport
craft to his Bigelow space hotels to be fully developed from the
winner of the prize in his requiring a 5 man vehicle. However,
following the example of the suborbital X-Prize, just accomplishing a
small 1 man test flight would be sufficient to serve as an impetus for
commercial ventures to invest in developing such launchers aside from
the prize.
Then I suggest Bigelow lower the requirement to only needing a single
crew member. This would allow multiple test flights before a manned
flight is attempted.
Bob Clark
On Aug 24, 3:17 pm, Robert Clark <rgregorycl...@yahoo.com> wrote:
> It would be a truly watershed moment just creating a SSTO even if it
> doesn't carry much payload. It wouldn't have to be anything extensive
> like perhaps what Boeing is planning with their X-37B derived SSTO.
> A small one could be demonstrated by amateur science or technical
> organizations, for instance by the British Interplanetary Society, or
> the Planetary Society.
> The Planetary Society is spending about $5.8 million total on their
> two attempts at solar sail demonstators:
>
> Cosmos 1.http://en.wikipedia.org/wiki/Cosmos_1
>
> LightSail-1.http://en.wikipedia.org/wiki/LightSail-1#Creation
Keith Henson
> A common estimate is that orbital flight is an order of magnitude more
> difficult than suborbital flight,
snip
I appreciate what you are proposing here. The problem is that low
exhaust velocity (compared to mission velocity) leads to high mass
ratios and very little margin for structure and engines, not to
mention payload.
It's physically possible to get exhaust velocity comparable to LEO (9
km/s) using hydrogen heated by beamed energy to 2700 deg K or higher.
Unfortunately, it takes very large laser or microwave generators to
get ten ton class payload to orbit and the only way they make economic
sense is for the beamed energy sources to be operating most of the
time. Which is just fine if you need 500,000 tons per year to GEO.
I have given some thought to the problems here:
http://htyp.org/Beamed_energy_propulsion#Numerical_investigation_of_a_Gausian_power_beam_heating_a_surface_cooled_with_hydrogen_and_re-radiating
Keith Henson
Sylvia Else
> It would be a truly watershed moment just creating a SSTO even if it
> doesn't carry much payload.
I assume you're talking about a reusable SSTO. It would be a watershed
moment if there were a credible path to creating something useful by
incremental steps. There would then be reason to attempt the required
improvements.
But a vehicle powered by an ordinary rocket has significant theoretical
limitations, and improvements would have to take the form of reducing
the dry weight, or improving the ISP of the engine, and there is already
plenty of motive for doing both even without an SSTO.
So on the whole, I don't think much would be achieved by constructing a
minimal payload SSTO using technology that's already been demonstrated.
Sylvia.
Keith Henson
Good analysis.
Reusable SSTO with the current state of the art in materials seems to
take higher ISP than chemical rockets can give. :-(
Skylon (by Reaction Engines) gets a really remarkable 1100 second or
so for the part of the flight in the atmosphere. This raises the
average exhaust velocity to around 6000 m/s so they get 49 tons of
vehicle and 15 tons of payload to LEO (or so the analysis indicates).
Beamed energy with 9000 m/s exhaust velocity makes reusable SSTO easy,
but it doesn't come in a small size.
Keith
Pat Flannery
> Skylon (by Reaction Engines) gets a really remarkable 1100 second or
> so for the part of the flight in the atmosphere.
I'll believe their isp figures when they build a Skylon, or even an
engine for a Skylon, and test it in flight.
Pat
Marvin the Martian
This whole engineering thing -- where you design to do the most with the
least, it just doesn't get to you guys does it? It's all "kewlness
factor".
Sylvia Else
It's not so clear how one goes about transferring the energy from the
beam to the reaction mass. In a conventional rocket the energy is being
produced throughout the mass as it burns.
If the reaction mass is transparent to the beam, then the energy will
have to be deposited on a surface, and then heat the mass by conduction.
This implies a surface capable of operating at the exhaust gas
temperature, which even conventional LOX/LH2 rockets don't have.
If the reaction mass is opaque to the beam, then only the surface will
be heated.
Either way, getting the energy distributed effectively through the
reaction mass seems like a challenge.
Sylvia.
Keith Henson
I thought so too. Was considering several kinds of heaters where, for
example, the heat was transferred through a window to an absorbent
like fine soot in the hydrogen stream.
But it turned out (if I didn't make an error in the model) that tubes
were good enough.
http://htyp.org/Beamed_energy_propulsion
Keith
Keith Henson
Have you studied the design? Do you have any specific reason the
Skylon design will not work?
Being a skeptic is just fine, but you should have reason(s) for it.
ESA did an extensive review of Skylon and could not find any problems
with it.
Keith
Quadibloc
> On Aug 26, 2:18 pm, Pat Flannery <flan...@daktel.com> wrote:
> > On 8/26/2011 9:09 AM, Keith Henson wrote:
>
> > > Skylon (by Reaction Engines) gets a really remarkable 1100 second or
> > > so for the part of the flight in the atmosphere.
>
> > I'll believe their isp figures when they build a Skylon, or even an
> > engine for a Skylon, and test it in flight.
>
> Have you studied the design? Do you have any specific reason the
> Skylon design will not work?
>
> Being a skeptic is just fine, but you should have reason(s) for it.
The fact that their specific impulse figure is "really remarkable" -
so far beyond the current state of the art - is, I would think,
sufficient reason for hesitancy. It isn't unfair to the people behind
Skylon to decline to base one's plans on something that ambitious
being successful until the success is in fact proven.
John Savard
Wayne Throop
::: engine for a Skylon, and test it in flight.
:: Do you have any specific reason the Skylon design will not work?
: Quadibloc <jsa...@ecn.ab.ca>
: The fact that their specific impulse figure is "really remarkable" -
: so far beyond the current state of the art - is, I would think,
: sufficient reason for hesitancy. It isn't unfair to the people behind
: Skylon to decline to base one's plans on something that ambitious
: being successful until the success is in fact proven.
I was assuming their number was so remarkable because they only need
to count the mass of the fuel in their calculation of isp, and not the
mass of the oxygen, as long as they are breathing air. I'd almost think
the numbers are too *low*, except that they are suffering drag to get
that oxygen.
But I dunno how they actually cacklated it.
Robert Clark
We'll know for sure in the (very) near term:
Boeing proposes SSTO system for AF RBS program.
The new issue of Aviation Week has a brief blurb about a Boeing
proposal for the Air Force's Reusable Booster System (RBS) program:
Boeing Offers AFRL Reusable Booster Proposal - AvWeek - June.13.11
(subscription required).
"Darryl Davis, who leads Boeing's Phantom Works, tells AvWeek that
they are proposing a 3-4 year technology readiness assessment that
would lead up to a demonstration of a X-37B type of system
but would be smaller. Wind tunnel tests have been completed. Davis
says the system would be a single stage capable of reaching low Earth
orbit and, with a booster, higher orbits. The system would return to
Earth as a glider.
Davis says "that advances in lightweight composites warrant another
look" at single-stage-to-orbit launchers."
http://www.hobbyspace.com/nucleus/index.php?itemid=30110
Bob Clark
Robert Clark
> ::: I'll believe their isp figures when they build a Skylon, or even an
> ::: engine for a Skylon, and test it in flight.
>
> :: Do you have any specific reason the Skylon design will not work?
>
> : The fact that their specific impulse figure is "really remarkable" -
> : so far beyond the current state of the art - is, I would think,
> : sufficient reason for hesitancy. It isn't unfair to the people behind
> : Skylon to decline to base one's plans on something that ambitious
> : being successful until the success is in fact proven.
>
> I was assuming their number was so remarkable because they only need
> to count the mass of the fuel in their calculation of isp, and not the
> mass of the oxygen, as long as they are breathing air. I'd almost think
> the numbers are too *low*, except that they are suffering drag to get
> that oxygen.
>
> But I dunno how they actually cacklated it.
Yes. That is how they calculate it. The Isp numbers are well in the
range for airbreathing engines such as jets. And for the rocket only
portion their Isp numbers are also well within the range of hydrogen
fueled rocket engines.
Bob Clark
Sylvia Else
Maybe I'm missing something, but if your heating is by transfer of heat
from the tubes to the propellant, then the tubes must be hotter than the
reaction mass at the point where it leaves the heater. This limits the
temperature achieved in the reaction mass, and thus the pressure and
exhaust speed.
However, since your reaction mass is not a combustion product, you have
much greater choice. It appears to me that the higher the molecular
weight, the better, as long as you avoid dissociation. An obvious
candidate is N2, or perhaps Argon, whose atomic weight (being as it's
inert) is higher then the molecular weight of N2).
Sylvia.
Greg Goss
>> But it turned out (if I didn't make an error in the model) that tubes
>> were good enough.
>Maybe I'm missing something, but if your heating is by transfer of heat
>from the tubes to the propellant, then the tubes must be hotter than the
>reaction mass at the point where it leaves the heater. This limits the
>temperature achieved in the reaction mass, and thus the pressure and
>exhaust speed.
>
>However, since your reaction mass is not a combustion product, you have
>much greater choice. It appears to me that the higher the molecular
>weight, the better, as long as you avoid dissociation. An obvious
>candidate is N2, or perhaps Argon, whose atomic weight (being as it's
>inert) is higher then the molecular weight of N2).
Does the starting point need to be a gas? (I seem to recall that
Heinlein's Rocket Ship Gallileo used zinc or something for reaction
mass.)
--
"If the Gods Had Meant Us to Vote They Would Have Given Us Candidates" (Jim Hightower)
Sylvia Else
If it's not a gass, then some of the energy is used to meet its latent
heats of fusion and vaporisation. Whether this matters will be an
engineering consideration.
Still, perhaps you're thinking of having the solid metal directly
vaporised by the beam. I suppose that could work in principle.
Sylvia.
Keith Henson
Carbon is the material of choice for the tubes. Same material that
was used in the NERVA engines.
> However, since your reaction mass is not a combustion product, you have
> much greater choice. It appears to me that the higher the molecular
> weight, the better, as long as you avoid dissociation. An obvious
> candidate is N2, or perhaps Argon, whose atomic weight (being as it's
> inert) is higher then the molecular weight of N2).
It turns out that hydrogen by far the best reaction material. Being
light it moves really fast.
Keith
> Sylvia.
Pat Flannery
> Have you studied the design? Do you have any specific reason the
> Skylon design will not work?
Yes, I've studied the design, which has a souped-up version of the
engine HOTOL was supposed to use and which the British government
canceled when they found that it wouldn't work.
But the real clue though is that little book SKYLON's designers wrote
about the destruction of Sodom and Gomorrah by the asteroid that skipped
off of the Alps:
http://www.skeptic.com/eskeptic/09-02-04/
If that doesn't set your bullshit sensors off, nothing will.
These guys are either ineffectual dreamers, or more likely, con artists
- trying to get people to invest in the next Moller flying car or Rotan
rocket.
Their engine will need just "a little more work" for the next decade or
two to be the greatest thing ever, and you can get in on the ground
floor of the operation and reap great monetary rewards when it takes off.
> Being a skeptic is just fine, but you should have reason(s) for it.
>
> ESA did an extensive review of Skylon and could not find any problems
> with it.
Show me a working SABRE engine that generates that kind of isp, then
we'll talk SKYLON.
Pat
Pat Flannery
>> However, since your reaction mass is not a combustion product, you have
>> much greater choice. It appears to me that the higher the molecular
>> weight, the better, as long as you avoid dissociation. An obvious
>> candidate is N2, or perhaps Argon, whose atomic weight (being as it's
>> inert) is higher then the molecular weight of N2).
>
> It turns out that hydrogen by far the best reaction material. Being
> light it moves really fast.
You can shoot the laser up the tailpipe at liftoff, but how are you
going to do that when it's arcing over on a ballistic trajectory to
enter orbit, and moving at a nearly 90 degrees to the beam emitter?
Also, what about clouds between the emitter and SSTO vehicle?
Pat
Keith Henson
"Really remarkable" isp in the context of rockets translates to
"expected" to "low" in powered flight.
If you back calculated from the altitude and velocity reached in air
breathing mode Skylon has an equivalent exhaust velocity of 10.5 km/s,
an ISP of ~1100.
Good for a rocket, poor for a plane.
Keith
Brad Guth
Actually radon (Rn222) would offer a way better value as an ion cannon
thruster. It's a use it or lose it kind of element.
http://translate.google.com/#
Brad Guth, Brad_Guth, Brad.Guth, BradGuth, BG / “Guth Usenet”
Greg Goss
>> However, since your reaction mass is not a combustion product, you have
>> much greater choice. It appears to me that the higher the molecular
>> weight, the better, as long as you avoid dissociation. An obvious
>> candidate is N2, or perhaps Argon, whose atomic weight (being as it's
>> inert) is higher then the molecular weight of N2).
>
>It turns out that hydrogen by far the best reaction material. Being
>light it moves really fast.
But why is "fast" an advantage? You're trying to transfer momentum
(proportional to V) using energy (proportional to V^2). Wouldn't
heavier be better, in which case the slower speed is made up for by
more momentum transfer per mile per hour?
Brad Guth
Good point, as long as there's not too much gravity holding you back.
It still requires energy to alter the velocity, and the greater the
mass the greater the energy requirement.
Once past the point of no return is where everything should really
speed up, such as once past a quarter of the distance to Sirius.
Wayne Throop
Specifically, for a thermal exhaust at a given temperature
(and hence for a given total energy), it has the largest velocity.
Of course, the photon is even better.
: Greg Goss <go...@gossg.org>
: But why is "fast" an advantage? You're trying to transfer momentum
: (proportional to V) using energy (proportional to V^2). Wouldn't
: heavier be better, in which case the slower speed is made up for by
: more momentum transfer per mile per hour?
You must keep distinct the total mass of propellant/exhaust, and the mass
of the individual particles in it. And since you get more velocity for
a fiven temperature from less massive particles, it's more efficient.
All you get from more massive molecles is increased density. Which
doesn't help as much as you might wish. Sometimes it can be nice,
but not as often as you might wish.
Mind you, that's only for thermal rockets.
Milage for other schemes may vary.
The cases where you may want a more massive exhaust particle,
even in a thermal rocket, involve tradeoffs of thrust vs total delta-v.
That's why photons are not considered much, despite being "the best"; you
get lots and lots of total delta-v (if you have total conversion, or
equivalent), but it's not all that easy to get high thrust. Plus the
kzinti lesson implies you may not want to use them to launmch from
the surface, even if you get the thrust.
And of course, if you have total conversion, you have to count the energy
you are throwing away if your exhaust has mass. You're better to convert
it to photons. But then, that's only if you have total conversion.
Jeff Findley
fb7877...@e34g2000prn.googlegroups.com>, hkeith...@gmail.com
says...
>
> Reusable SSTO with the current state of the art in materials seems to
> take higher ISP than chemical rockets can give. :-(
This is an opinion, not a fact. One of the reasons we have not really
tried to create a fully reusable SSTO is the cost of development. This
would surely be in the billions of dollars range, and if NASA does it,
surely in the tens of billions of dollars range.
For that sort of investment to pay off, you've got to have a very high
demand for launches. In other words, you need a market for frequent
launches. That market currently does not exist.
NASA half heartedly tried to go down this path with X-33 with Lockheed,
but neither party was truly committed to the project. NASA saw X-33 as
a technological sandbox to play in and therefore picked the most
technically challenging design of the three proposed. Either of the
other two designs had a much higher chance of successfully flying, but
with the shuttle program still flying, X-33 didn't "need" to succeed.
For Lockheed, they won a fairly lucrative R&D contract. Even though
they were putting up some of "their own money" for X-33, it was still a
win for them, keeping many people employed for several years. Best of
all, the competition didn't win the contract, so any progress which was
made would be kept from their competitors.
> Skylon (by Reaction Engines) gets a really remarkable 1100 second or
> so for the part of the flight in the atmosphere. This raises the
> average exhaust velocity to around 6000 m/s so they get 49 tons of
> vehicle and 15 tons of payload to LEO (or so the analysis indicates).
Skylon gets no such thing. It's as much of a paper vehicle as all of
the SSTO's proposed through the years (and there have been many).
> Beamed energy with 9000 m/s exhaust velocity makes reusable SSTO easy,
> but it doesn't come in a small size.
In terms of the ground infrastructure required, no it certainly doesn't.
When you compute the (beamed) power requirements for even a modest sized
SSTO, you end up needing a huge ground infrastructure.
Jeff
--
" Ares 1 is a prime example of the fact that NASA just can't get it
up anymore... and when they can, it doesn't stay up long. ;) "
- tinker
Jeff Findley
<7rSdnfKl4oj0fsrT...@posted.northdakotatelephone>,
fla...@daktel.com says...
Paper engines are *always* better than existing engines. If they
weren't, why fund a very costly research and development program. ;-)
Seriously though, even if this engine is successful, I doubt it could
compete with a similar vehicle powered by modern liquid fueled rocket
engines. I'd like to see Reaction Engines do a fair comparison between
Skylon and an all rocket powered vehicle with the same payload.
Pat Flannery
> Seriously though, even if this engine is successful, I doubt it could
> compete with a similar vehicle powered by modern liquid fueled rocket
> engines. I'd like to see Reaction Engines do a fair comparison between
> Skylon and an all rocket powered vehicle with the same payload.
Assuming they actually got around to building a SABRE engine, it would
be fun to hear how they intended to flight test it, as that would
require an aircraft something like a X-15 to carry it, which Britain
doesn't have.
Pat
Keith Henson
> In article <e249bbd7-b640-492d-9f70-
> fb78775b3...@e34g2000prn.googlegroups.com>, hkeithhen...@gmail.com
> says...
>
> > Reusable SSTO with the current state of the art in materials seems to
> > take higher ISP than chemical rockets can give. :-(
>
> This is an opinion, not a fact.
It's more of a reflection on the present state of the art. Eric
Drexler has described nanotechnology based SSTO made of grown diamond
that should reach orbit on (as I recall) under 100 kg of dry mass.
> One of the reasons we have not really
> tried to create a fully reusable SSTO is the cost of development. This
> would surely be in the billions of dollars range, and if NASA does it,
> surely in the tens of billions of dollars range.
>
> For that sort of investment to pay off, you've got to have a very high
> demand for launches. In other words, you need a market for frequent
> launches. That market currently does not exist.
I agree.
> NASA half heartedly tried to go down this path with X-33 with Lockheed,
> but neither party was truly committed to the project. NASA saw X-33 as
> a technological sandbox to play in and therefore picked the most
> technically challenging design of the three proposed. Either of the
> other two designs had a much higher chance of successfully flying, but
> with the shuttle program still flying, X-33 didn't "need" to succeed.
>
> For Lockheed, they won a fairly lucrative R&D contract. Even though
> they were putting up some of "their own money" for X-33, it was still a
> win for them, keeping many people employed for several years. Best of
> all, the competition didn't win the contract, so any progress which was
> made would be kept from their competitors.
>
> > Skylon (by Reaction Engines) gets a really remarkable 1100 second or
> > so for the part of the flight in the atmosphere. This raises the
> > average exhaust velocity to around 6000 m/s so they get 49 tons of
> > vehicle and 15 tons of payload to LEO (or so the analysis indicates).
>
> Skylon gets no such thing. It's as much of a paper vehicle as all of
> the SSTO's proposed through the years (and there have been many).
I qualified it "or so the analysis indicates" but I have no reason to
doubt their figures. The design they use for air breathing recovers
more energy from the hydrogen than just burning it. Hydrogen has an
energy content of around 50 kWh/kg. It also takes 20 kWh to liquefy
the stuff. The SABRE engine gets some of the 20 kWh/kg back by using
the temperature difference between ram air and the liquid hydrogen to
compress (cooled) air to rocket chamber pressure.
The hardest part of the design is the 2 GW heat exchanger on each
engine. Miles and miles of fine tubing. Recognizing this as the
hardest problem, they have built substantial sections of the heat
exchanger and tested them. To my amazement, they work just fine.
> > Beamed energy with 9000 m/s exhaust velocity makes reusable SSTO easy,
> > but it doesn't come in a small size.
>
> In terms of the ground infrastructure required, no it certainly doesn't.
> When you compute the (beamed) power requirements for even a modest sized
> SSTO, you end up needing a huge ground infrastructure.
It depends partly on how you partition the velocity gain. If you use
the Skylon in maximum sub orbital mode, then 500 MW of laser with
bounce mirrors in GEO are enough for a 500,000 ton per year parts
stream for power satellites. http://www.theoildrum.com/node/7898
It's large, but in the context of building $160 B worth of power
satellites a year, $4 to $10 B is small enough.
Keith
Rick Jones
> NASA saw X-33 as a technological sandbox to play in and therefore
> picked the most technically challenging design of the three
> proposed.
Isn't that what many (most?) pundits are saying should be NASA's focus
- the most technically challenging and so greater chance of failure
things, rather than the mundane but most likely to be a success in
other than engineering lessons learned terms?
rick jones
--
The computing industry isn't as much a game of "Follow The Leader" as
it is one of "Ring Around the Rosy" or perhaps "Duck Duck Goose."
- Rick Jones
these opinions are mine, all mine; HP might not want them anyway... :)
feel free to post, OR email to rick.jones2 in hp.com but NOT BOTH...
Jeff Findley
says...
>
> In sci.space.history Jeff Findley <jeff.f...@ugs.nojunk.com> wrote:
> > NASA saw X-33 as a technological sandbox to play in and therefore
> > picked the most technically challenging design of the three
> > proposed.
>
> Isn't that what many (most?) pundits are saying should be NASA's focus
> - the most technically challenging and so greater chance of failure
> things, rather than the mundane but most likely to be a success in
> other than engineering lessons learned terms?
Yes, but a sane way to approach technology development is to pick one
technology to test on an X-vehicle. X-33 tested quite a few more than
that:
1. Lifting body aerodynamics suitable for an SSTO
2. Linear aerospike liquid fueled rocket engine
3. Conformal (to the aerodynamic shape), composite liquid hydrogen tank
There are at least three new technologies there, the failure of any one
would doom the program to failure.
The lifting body shape kept changing during the program. It changed so
much (for the follow-on orbital design) that what started out as small
winglets grew to such a large size that they couldn't reasonably be
called anything but wings.
The failure of the liquid hydrogen tank during testing was a very
obvious indication that the technology was having at least some serious
teething problems.
The linear aerospike seemed to be doing well, but from what I remember
reading, it was experiencing some problems with respect to weight
related to the exhaust ramps. Still, the engine development seemed to
be doing better than the rest of the vehicle.
Jeff Findley
<Uoqdna6U6dliMcbT...@posted.northdakotatelephone>,
fla...@daktel.com says...
Lack of an obvious way to perform incremental testing on a SABRE engine
is a problem. Launching it on top of expendable missiles or expendable
launch vehicles becomes very expensive very quickly.
Sylvia Else
> In article
> <Uoqdna6U6dliMcbT...@posted.northdakotatelephone>,
> fla...@daktel.com says...
>>
>> On 8/29/2011 5:07 AM, Jeff Findley wrote:
>>
>>
>>> Seriously though, even if this engine is successful, I doubt it could
>>> compete with a similar vehicle powered by modern liquid fueled rocket
>>> engines. I'd like to see Reaction Engines do a fair comparison between
>>> Skylon and an all rocket powered vehicle with the same payload.
>>
>> Assuming they actually got around to building a SABRE engine, it would
>> be fun to hear how they intended to flight test it, as that would
>> require an aircraft something like a X-15 to carry it, which Britain
>> doesn't have.
>
> Lack of an obvious way to perform incremental testing on a SABRE engine
> is a problem. Launching it on top of expendable missiles or expendable
> launch vehicles becomes very expensive very quickly.
>
> Jeff
It's worth noting that the core of the Sabre engine operates entirely in
subsonic airflow. RE intend to construct testbeds that produce the
entire range of core inlet conditions on the ground.
The nacelle cannot be tested that way, and there is a proposal to build
a subscale "Nacelle Test Vehicle". It's not clear whether this will be
reusable, but they won't be putting the very expensive core Sabre
hardware on it, only a scaled nacelle. Part of the testing will
presumably involve verifying that the core inlet conditions correspond
to those used in the testing of the core.
Sylvia.
Pat Flannery
>
> It's more of a reflection on the present state of the art. Eric
> Drexler has described nanotechnology based SSTO made of grown diamond
> that should reach orbit on (as I recall) under 100 kg of dry mass.
Describing it and building it (or it even being possible to build) are
two different things.
As long as Reaction Engines wants to build a super engine, how about a
giant diameter flat jet engine?:
http://civilianmilitaryintelligencegroup.com/6476/sketched-designs-for-avro-silverbug-vstol-craft
Just because that couldn't be made to work in the 1950's, think how much
more advanced the state of the art is now? :-D
Pat
Pat Flannery
> Yes, but a sane way to approach technology development is to pick one
> technology to test on an X-vehicle. X-33 tested quite a few more than
> that:
>
> 1. Lifting body aerodynamics suitable for an SSTO
> 2. Linear aerospike liquid fueled rocket engine
> 3. Conformal (to the aerodynamic shape), composite liquid hydrogen tank
>
> There are at least three new technologies there, the failure of any one
> would doom the program to failure.
When Kelly Johnson was running the Skunk works, they had a rule about
not more than one major new technology per aircraft; they found all
about that the hard way with the A-12 with its titanium structure,
turboramjets, and stealth design.
> The lifting body shape kept changing during the program. It changed so
> much (for the follow-on orbital design) that what started out as small
> winglets grew to such a large size that they couldn't reasonably be
> called anything but wings.
>
> The failure of the liquid hydrogen tank during testing was a very
> obvious indication that the technology was having at least some serious
> teething problems.
>
> The linear aerospike seemed to be doing well, but from what I remember
> reading, it was experiencing some problems with respect to weight
> related to the exhaust ramps. Still, the engine development seemed to
> be doing better than the rest of the vehicle.
IIRC, the aft ramps on the engine weren't cooling
enough and getting burned during firing also.
Pat
Jeff Findley
cHTnZ2dnU...@posted.northdakotatelephone>, fla...@daktel.com
says...
>
> On 8/29/2011 12:56 PM, Jeff Findley wrote:
>
> > Yes, but a sane way to approach technology development is to pick one
> > technology to test on an X-vehicle. X-33 tested quite a few more than
> > that:
> >
> > 1. Lifting body aerodynamics suitable for an SSTO
> > 2. Linear aerospike liquid fueled rocket engine
> > 3. Conformal (to the aerodynamic shape), composite liquid hydrogen tank
> >
> > There are at least three new technologies there, the failure of any one
> > would doom the program to failure.
>
> When Kelly Johnson was running the Skunk works, they had a rule about
> not more than one major new technology per aircraft; they found all
> about that the hard way with the A-12 with its titanium structure,
> turboramjets, and stealth design.
And yet the same organization deliberately made that very same
"mistake" on X-33. The sad thing is that NASA actually wanted them to
make that mistake. NASA picked the winning X-33 design that
incorporated the highest number of new technologies. Lockheed didn't
care if X-33 succeeded nearly as much as it cared about winning the
contract.
The incentives of "business as usual" for NASA contractors are all
screwed up.
> > The lifting body shape kept changing during the program. It changed so
> > much (for the follow-on orbital design) that what started out as small
> > winglets grew to such a large size that they couldn't reasonably be
> > called anything but wings.
> >
> > The failure of the liquid hydrogen tank during testing was a very
> > obvious indication that the technology was having at least some serious
> > teething problems.
> >
> > The linear aerospike seemed to be doing well, but from what I remember
> > reading, it was experiencing some problems with respect to weight
> > related to the exhaust ramps. Still, the engine development seemed to
> > be doing better than the rest of the vehicle.
>
>
> IIRC, the aft ramps on the engine weren't cooling
> enough and getting burned during firing also.
It would have been nice to keep that engine program going to the point
it actually produced an operational engine suitable for use on an actual
launch vehicle.
I'd like to see SLS killed and individual technologies developed by
different X vehicles. This what NACA did for aircraft. NACA didn't
build operational vehicles and operate them like NASA has been doing
with the space shuttle and wants to do with SLS.
Harsh? Maybe, but shuttle workers are already getting pink slips. Now
would be the opportunity to transform NASA back into an organization
that actually develops new technologies instead of developing yet
another launch vehicle based primarily on technologies which were new
back in the 1960's (J-2) and 1970's (SSME and shuttle SRB).
Thomas Womack
Jeff Findley <jeff.f...@ugs.nojunk.com> wrote:
>And yet the same organization deliberately made that very same
>"mistake" on X-33. The sad thing is that NASA actually wanted them to
>make that mistake. NASA picked the winning X-33 design that
>incorporated the highest number of new technologies. Lockheed didn't
>care if X-33 succeeded nearly as much as it cared about winning the
>contract.
>
>The incentives of "business as usual" for NASA contractors are all
>screwed up.
How far should NASA take things in-house? I can see an argument for
going the whole way, for having it be NASA employees in NASA-owned
buildings who bend the metal and lay out the carbon fibre and grind
the silicon carbide and run the wafers through the lithography
machines to designs made by NASA-employed engineers - but I'm not sure
you can make that argument to the senators for Boeing and
Lockheed-Martin, even if you argue that it provides the commercial
aerospace industry in the US with a ready stream of well-trained new
recruits to whom they merely have to supply more interesting jobs and
higher pay than NASA.
Tom
Pat Flannery
>>
>> IIRC, the aft ramps on the engine weren't cooling
>> enough and getting burned during firing also.
I was surprised to find out we once had a working aerospike engine that
even used liquid fluorine and LH-2 for propellants, the Rocketdyne AMPS-1
There's some info on it here:
http://picturetrail.com/sfx/album/view/23272850
http://picturetrail.com/sfx/album/main/23272850/389399611
http://picturetrail.com/sfx/album/main/23272850/389399612
http://picturetrail.com/sfx/album/main/23272850/389400205
Pat
Keith Henson
> Keith Henson <hkeithhen...@gmail.com> wrote:
>
> >It turns out that hydrogen by far the best reaction material. Being
> >light it moves really fast.
>
> But why is "fast" an advantage? You're trying to transfer momentum
> (proportional to V) using energy (proportional to V^2). Wouldn't
> heavier be better, in which case the slower speed is made up for by
> more momentum transfer per mile per hour?
What you are normally trying to do with a rocket is to make some delta
V, like surface to LEO and still have payload.
Round numbers, the SSME exhaust velocity is around 4.5 km/s, LEO takes
about 9 km/s.
From the rocket equation, going twice the exhaust velocity takes a
mass ratio of 7.4, leaving 13.5% of lift off mass for structure,
engines and payload. That's not easy for expendable rockets and given
the current state of the art, probably impossible for a reusable
vehicle.
For 9 km/s exhaust velocity, which they came close to with NERVA, the
mass ratio is only 3. So 33% of the liftoff mass can be structure and
payload. For a 300 ton takeoff mass, a 50 ton vehicle could put 50
tons of payload in LEO.
For a given temperature, the exhaust velocity is proportional to one
over the square root of the molecular mass. So if hydrogen gave 9 km/
s at some temperature, nitrogen, with 14 times the mass would reduce
the exhaust velocity by a factor of 3.75 or 2400 m/s which is worse
than a SSME.
Isn't math fun. :-)
Keith
Derek Lyons
That's an engineering detail Pat - and Keith isn't interested much in
the details. So long as his equations produce sufficiently huge
numbers, he's happy.
D.
--
Touch-twice life. Eat. Drink. Laugh.
http://derekl1963.livejournal.com/
-Resolved: To be more temperate in my postings.
Oct 5th, 2004 JDL
Jeff Findley
two...@chiark.greenend.org.uk says...
You snipped the rest of my reply, but I'd like to see NASA only doing
R&D for new technologies and let them buy launches on commercial
vehicles (i.e. Atlas, Delta, Falcon, Taurus, etc.). Let the commercial
launch providers pick and choose the new technologies (developed by
NASA) to incorporate into their next generation of launch vehicles. Let
competition in the open market decide what technologies "win".
In this sort of a world, NASA would only be allowed to build and fly X-
vehicles, not an operational system like the space shuttle or the
proposed SLS. I think it's just fine for a pure R&D organization like
this to do everything that's new in house. For everything that's "off
the shelf", they ought to buy components from existing suppliers.
For example, on a vehicle intended to flight test a linear aerospike
engine, NASA could potentially build and test the entire engine
themselves. But, when it came to building the rest of the vehicle
(using conventional materials and methods), that could be subcontracted
out to a winning bidder. The vehicle would be flown enough times to
prove the engine worked and to improve the engine design. Such an X-
vehicle would not be allowed to fly "operational" missions which would
compete with the commercial sector.
I'm not sure how X-vehicles, like I describe above, would take jobs away
from aerospace contractors. Any existing services NASA could buy from
them, like cargo or crew deliveries to ISS, NASA would not be allowed to
do with their X-vehicles.
Derek Lyons
>I'd like to see SLS killed and individual technologies developed by
>different X vehicles. This what NACA did for aircraft. NACA didn't
>build operational vehicles and operate them like NASA has been doing
>with the space shuttle and wants to do with SLS.
That's because the NACA wasn't charged with any operational function.
But what folks don't realize is that NASA isn't just the NACA on
steroids, and never has been, which means means that any comparision
between the two fails right out of the gate.
NASA is basically flawed by design because it's expected to both
support operations *and* to develop nifty new technologies - without a
firewall between the two sides.
>Harsh? Maybe, but shuttle workers are already getting pink slips. Now
>would be the opportunity to transform NASA back into an organization
>that actually develops new technologies instead of developing yet
>another launch vehicle based primarily on technologies which were new
>back in the 1960's (J-2) and 1970's (SSME and shuttle SRB).
One could argue that the reason we're trapped in a backwater is
because we keep insisting on chasing shiny new technologies rather
than iterating on what we have. It's the latter that has historically
lead to reliable and cheap technology. It was also a major, and
largely invisible, function of the NACA.
Jeff Findley
<StednTdO390Bb8HT...@posted.northdakotatelephone>,
fla...@daktel.com says...
Sadly, NASA has become an organization obsessed with missions rather
than with technology development like NACA did for aircraft. Such an
organization tends to be risk averse since any new technology could fail
and jeopardize "the mission". Is it any wonder that SLS would use 60's
and 70's technologies which were matured by the shuttle program?
Derek Lyons
>Sadly, NASA has become an organization obsessed with missions rather
>than with technology development like NACA did for aircraft.
Because missions are what the Administration supports and Congress
funds. It's been pretty much like that way since Day One. NASA
technology development, especially the big ticket items, has largely
always been either in support (at least vaguely) of a Mission or a
politically attractive Cause (or Congressional district).
>Such an organization tends to be risk averse since any new technology
>could fail and jeopardize "the mission".
NASA has *always* been risk averse - that's one of the reasons they
originally had a whole raftload of (NACA style) test flights planned
before any lunar orbital attempt, let alone any landing attempt. That
scheme got shelved only when it appeared the Russians would beat us
politically and money and time began running out.
Marvin the Martian
> On Aug 28, 7:51 pm, Greg Goss <go...@gossg.org> wrote:
>> Keith Henson <hkeithhen...@gmail.com> wrote:
>> >> However, since your reaction mass is not a combustion product, you
>> >> have much greater choice. It appears to me that the higher the
>> >> molecular weight, the better, as long as you avoid dissociation. An
>> >> obvious candidate is N2, or perhaps Argon, whose atomic weight
>> >> (being as it's inert) is higher then the molecular weight of N2).
>>
>> >It turns out that hydrogen by far the best reaction material. Being
>> >light it moves really fast.
>>
>> But why is "fast" an advantage? You're trying to transfer momentum
>> (proportional to V) using energy (proportional to V^2). Wouldn't
>> heavier be better, in which case the slower speed is made up for by
>> more momentum transfer per mile per hour? --
>> "If the Gods Had Meant Us to Vote They Would Have Given Us Candidates"
>> (Jim Hightower)
>
> Good point,
No it isn't. You're trying to maximize acceleration, not force.
> as long as there's not too much gravity holding you back.
Which has nothing at all to do with it.
Jeff Findley
says...
Sometimes this is true. It was true in spades when NASA picked the
winning X-33 design.
But sometimes I feel that the pendulum swings too far the other
direction. SLS has been mandated to use as much shuttle derived
technology as possible. The thought of incorporating SRB's derived from
the shuttle SRB's into SLS gives me the creeps. I do not want NASA
flying astronauts on large segmented solids. They're fine for missiles,
not so much for manned vehicles or even for launch vehicles carrying
satellites.
Jeff Findley
says...
>
> Jeff Findley <jeff.f...@ugs.nojunk.com> wrote:
>
> >Sadly, NASA has become an organization obsessed with missions rather
> >than with technology development like NACA did for aircraft.
>
> Because missions are what the Administration supports and Congress
> funds. It's been pretty much like that way since Day One. NASA
> technology development, especially the big ticket items, has largely
> always been either in support (at least vaguely) of a Mission or a
> politically attractive Cause (or Congressional district).
>
> >Such an organization tends to be risk averse since any new technology
> >could fail and jeopardize "the mission".
>
> NASA has *always* been risk averse - that's one of the reasons they
> originally had a whole raftload of (NACA style) test flights planned
> before any lunar orbital attempt, let alone any landing attempt. That
> scheme got shelved only when it appeared the Russians would beat us
> politically and money and time began running out.
They've arguably become more risk averse. Mercury and Gemini flew on
"man rated" missiles (Redstone, Atlas, Titan II). Back then, Titan II
was a new missile with a very short track record (including new engines
using new, storable, toxic, hypergolic propellants).
Fast forward to ESAS when NASA argued that EELV's weren't safe enough
for Orion, that only Ares I would be safe enough. Ares I being a design
using 60's and 70's era engines.
Derek Lyons
>In article <4e5d2d02....@news.supernews.com>, fair...@gmail.com
>says...
>>
>> Jeff Findley <jeff.f...@ugs.nojunk.com> wrote:
>>
>> >Such an organization tends to be risk averse since any new technology
>> >could fail and jeopardize "the mission".
>>
>> NASA has *always* been risk averse - that's one of the reasons they
>> originally had a whole raftload of (NACA style) test flights planned
>> before any lunar orbital attempt, let alone any landing attempt. That
>> scheme got shelved only when it appeared the Russians would beat us
>> politically and money and time began running out.
>
>They've arguably become more risk averse. Mercury and Gemini flew on
>"man rated" missiles (Redstone, Atlas, Titan II). Back then, Titan II
>was a new missile with a very short track record (including new engines
>using new, storable, toxic, hypergolic propellants).
I'm not debating that they've become more risk averse as it's the
stone cold truth, just noting that the difference over time is one of
degree rather than one of kind. Public acceptance of the risk, which
also plays into the institutional attitude, has (on the other hand)
undergone a change of both degree and kind.
>Fast forward to ESAS when NASA argued that EELV's weren't safe enough
>for Orion, that only Ares I would be safe enough.
That's as much politics as risk aversion.
>Ares I being a design using 60's and 70's era engines.
I don't view that as bad thing necessarily. One doesn't develop
reliable tech by chasing the will 'o the wisp of the shiniest and
newest. One developes reliable tech by iteration and through
operations.
(quoting from another message)
Jeff Findley <jeff.f...@ugs.nojunk.com> wrote:
>In article <4e601bbc....@news.supernews.com>, fair...@gmail.com
>says...
>
>> One could argue that the reason we're trapped in a backwater is
>> because we keep insisting on chasing shiny new technologies rather
>> than iterating on what we have. It's the latter that has historically
>> lead to reliable and cheap technology. It was also a major, and
>> largely invisible, function of the NACA.
>
>Sometimes this is true. It was true in spades when NASA picked the
>winning X-33 design.
>
>But sometimes I feel that the pendulum swings too far the other
>direction. SLS has been mandated to use as much shuttle derived
>technology as possible. The thought of incorporating SRB's derived from
>the shuttle SRB's into SLS gives me the creeps. I do not want NASA
>flying astronauts on large segmented solids. They're fine for missiles,
>not so much for manned vehicles or even for launch vehicles carrying
>satellites.
Why? From an engineering point of view, they are no more or less
reliable than any other propulsion system available. (That is to say,
by normal standards they are abysmally unreliable.)
Jeff Findley
Because large segmented solids have failure modes which are best
described as "ugly". Case rupture is one of the worst of these modes.
Case rupture can happen with very little warning and can be quite
violent.
Also, since shuttle heritage SRB's have a thrust termination system
which unzips the SRB along its length, its use would result in quite a
bit of debris. I doubt if this system would be changed for SLS.
Either of the above makes a launch escape system for a manned capsule
quite challenging to engineer and quite heavy.
Pat Flannery
>> Why? From an engineering point of view, they are no more or less
>> reliable than any other propulsion system available. (That is to say,
>> by normal standards they are abysmally unreliable.)
>
> Because large segmented solids have failure modes which are best
> described as "ugly". Case rupture is one of the worst of these modes.
> Case rupture can happen with very little warning and can be quite
> violent.
>
> Also, since shuttle heritage SRB's have a thrust termination system
> which unzips the SRB along its length, its use would result in quite a
> bit of debris. I doubt if this system would be changed for SLS.
They stayed pretty intact until ocean impact when the destruct system
ripped them open on the Challenger flight.
They were originally going have a nose blow-off panel type thrust
termination venting system, like was invented for the Titan IIIC when it
was going to carry MOL and Dyna-Soar, but studies showed the ET was too
fragile to take the blast effects when it was activated.
During the development of the Titan III SRBs, it was found that
activating the venting system caused internal casing pressure to drop so
fast that combustion of the solid fuel stopped.
If they had gone ahead with Ares-1, the fact that it had just the SRB as
the first stage and no tankage next to it would have allowed them to
install a venting system if desired.
Pat
Marvin the Martian
> On 8/27/2011 7:04 AM, Keith Henson wrote:
>> Have you studied the design? Do you have any specific reason the
>> Skylon design will not work?
>
> Yes, I've studied the design, which has a souped-up version of the
> engine HOTOL was supposed to use and which the British government
> canceled when they found that it wouldn't work. But the real clue though
> is that little book SKYLON's designers wrote about the destruction of
> Sodom and Gomorrah by the asteroid that skipped off of the Alps:
> http://www.skeptic.com/eskeptic/09-02-04/ If that doesn't set your
> bullshit sensors off, nothing will. These guys are either ineffectual
> dreamers, or more likely, con artists - trying to get people to invest
> in the next Moller flying car or Rotan rocket.
> Their engine will need just "a little more work" for the next decade or
> two to be the greatest thing ever, and you can get in on the ground
> floor of the operation and reap great monetary rewards when it takes
> off.
>
>
>> Being a skeptic is just fine, but you should have reason(s) for it.
>>
>> ESA did an extensive review of Skylon and could not find any problems
>> with it.
>
> Show me a working SABRE engine that generates that kind of isp, then
> we'll talk SKYLON.
>
> Pat
Skylon reaches 1.7 km/s as an air breather[1, 29] , but to reach LEO you
need 9.3 to 10 km/s.[2] Your savings as a air breather is only
(1.7/9.3)^2 = 3%.
It is a flying single stage rocket.
Sorry to burst people's bubbles, but for added wing weight and the
complexity of the engine, you'll lose that 3% advantage.
We can build single stage to orbit rockets now, it's just not worth it.
[1] http://en.wikipedia.org/wiki/Skylon_%28spacecraft%29#SABRE_Engines
[2] http://en.wikipedia.org/wiki/Delta_v#Delta-vs_around_the_Solar_System
Sylvia Else
Another person who thinks they can rule Skylon out by means of simple
arithmetic. Have you looked at the actual numbers for Skylon? Can you
point to a specific 'error' made by Reaction Engines.
Did you allow for the effect that a wing has on gravity losses?
On its being "worth it" did you look at the economics of having a fully
reusable craft doesn't have to be reassembled before each launch, nor
lifted into a vertical position on a launch pad?
Sylvia.
Derek Lyons
>Skylon reaches 1.7 km/s as an air breather
Skylon doesn't reach anything - as it doesn't exist except as a stack
of papers and some magnetic domains on a hard drive.
Derek Lyons
>In article <4e5e5cea....@news.supernews.com>, fair...@gmail.com
>says...
>> Jeff Findley <jeff.f...@ugs.nojunk.com> wrote:
>> >But sometimes I feel that the pendulum swings too far the other
>> >direction. SLS has been mandated to use as much shuttle derived
>> >technology as possible. The thought of incorporating SRB's derived from
>> >the shuttle SRB's into SLS gives me the creeps. I do not want NASA
>> >flying astronauts on large segmented solids. They're fine for missiles,
>> >not so much for manned vehicles or even for launch vehicles carrying
>> >satellites.
>>
>> Why? From an engineering point of view, they are no more or less
>> reliable than any other propulsion system available. (That is to say,
>> by normal standards they are abysmally unreliable.)
>
>Because large segmented solids have failure modes which are best
>described as "ugly". Case rupture is one of the worst of these modes.
>Case rupture can happen with very little warning and can be quite
>violent.
Which of course happens very rarely.
>Also, since shuttle heritage SRB's have a thrust termination system
>which unzips the SRB along its length, its use would result in quite a
>bit of debris. I doubt if this system would be changed for SLS.
Better take that card back out of your palm - because the subject of
the discussion isn't "shuttle heritage" or "SLS".
>Either of the above makes a launch escape system for a manned capsule
>quite challenging to engineer and quite heavy.
So then we grow up and accept that failures are occasionally going to
happen and people are going to die as a result of them.
Keith Henson
> Marvin the Martian <mar...@ontomars.org> wrote:
>
> >Skylon reaches 1.7 km/s as an air breather
>
> Skylon doesn't reach anything - as it doesn't exist except as a stack
> of papers and some magnetic domains on a hard drive.
That's true. So was the Saturn V and the Space Shuttle at one time.
^^^^^^^^^^^^^^^^
ESA: British Skylon spaceplane seems perfectly possible
Wizzo robot runway rocketplane cleared to proceed
By Lewis Page
Posted in Space, 24th May 2011 12:50 GMT
The clever part of SABRE is its ability to use oxygen from the air and
burn it in a normal rocket back end. This is achieved by taking air at
the front and chilling it down incredibly fast using very, very
powerful refrigeration gear running on a closed loop of liquid helium,
which dumps the resulting heat into the cryogenic liquid-hydrogen
fuel.
The trouble with this is that air contains water vapour, and in the
normal course of events chilling it down like this would soon block up
the SABRE with ice. Preventing frost buildup is one of Reaction
Engines' main special sauces, and they have demonstrated that they can
do it in the lab to the ESA engineers' satisfaction:
As part of the ESA technical evaluation of the SABRE engine, the
design and operating principles of the frost control mechanism were
explained to ESA. In addition a number of tests were performed at
laboratory scale on request of ESA to demonstrate the repeatability of
the frost control. ESA can confirm that the frost control mechanism of
the SABRE engine, (at laboratory scale), works and is repeatable. In
addition ESA expects these positive results to be repeated on the
planned tests of the heat exchanger when it is tested on a VIPER jet
engine.
These larger-scale ground tests are planned for this summer, as the
Reg previously reported.
Assuming that a SABRE nacelle can be successfully built and flight
tested aboard the proposed Nacelle Test Vehicle aeroplane - the ESA
endorses this plan - Skylon isn't out of the woods yet. It will still
be necessary to build the huge main fuselage and wings, which need to
be light, strong, able to resist massive heating, and able to hold
hundreds of tonnes of explosive cryogenic-liquid fuel.
The ESA structures team think that Reaction Engines have a decent shot
at doing this, however:
Structural design work undertaken by REL does not demonstrate any
areas of implausibility, given the relatively benign environment of
the flight trajectory.
According to the ESA, the cigar-shaped main fuselage of Skylon is
"more akin to that of an Airship than a conventional launcher or
aircraft". This makes sense as its designers are facing similar
problems to those that the long-ago engineers who built the great
rigid airships of the 1930s had to tackle.
Like their illustrious predecessors, the Reaction Engines team need to
enclose as much volume as possible with as little weight and internal
structure as possible. The old-timers were even - mostly - trying to
enclose the same stuff, hydrogen, though in their case in gaseous form
rather than liquid.
Overall the ESA can't see right off any reason that a Skylon-style
aeroshell, wings etc can't be built.
^^^^^^^^^^^^^^^
Read the rest here: http://www.theregister.co.uk/2011/05/24/skylon_esa_report/
And there is a lot more if you Google for Skylon ESA.
Jeff Findley
says...
>
> Another person who thinks they can rule Skylon out by means of simple
> arithmetic. Have you looked at the actual numbers for Skylon? Can you
> point to a specific 'error' made by Reaction Engines.
>
> Did you allow for the effect that a wing has on gravity losses?
>
> On its being "worth it" did you look at the economics of having a fully
> reusable craft doesn't have to be reassembled before each launch, nor
> lifted into a vertical position on a launch pad?
You remind me of annoying people at work who use the very same argument
to justify the complete lack of a sane design for their code. When
accused of writing spaghetti code, their defense is, "Show me one line
of code I don't need".
Well that's not the point, now is it? The point is why go down this
insane path in the first place? Whether it's spaghetti code or air
breathing engines on a launch vehicle doesn't matter. The problem is
the initial assumptions are crap.
The development costs for Skylon/SABRE are far too high for far too
little payoff. Since SABRE has to work as a rocket engine for a good
portion of Skylon's launch trajectory, just ditch the air breathing
engine and use a rocket engine the entire way. This results in a
simpler design, which ultimately means lower costs. The added LOX
needed costs pennies per pound, but the big win is losing the complex
engines, complex aerodynamics (wings and engine intakes), and etc.
If you're concerned that a conventional rocket engine won't have the
performance needed, then work on finishing development of aerospike
engines. That technology is *far* more mature than SABRE and would cost
*far* less to develop.
Stop promoting using a shiny new hammer for pounding screws into wood.
It's simply not the right tool for the job.
Jeff Findley
<1b6dnTwUJsAIc8PT...@posted.northdakotatelephone>,
fla...@daktel.com says...
I know the history. The point is that the current destruct system is
violent and results in a lot of debris. Changing to a nose blow off
system would help, but that's an expensive design change which partially
negates using "shuttle derived" hardware.
Better to kill SLS entirely and buy commercial launches. For the tens
of billions which would be spent on SLS, you could buy *a lot* of EELV
launches.
Jeff Findley
Failures are inevitable, but we should be smart enough to pick the
mature technologies which have more benign failure modes. Large
segmented solids have drawbacks which make them undesirable for use on
launch vehicles. They were picked for shuttle because they were cheaper
to develop. Continuing to go down this failed path is dumb.
Robert Clark
> On Aug 24, 12:30 pm, Robert Clark <rgregorycl...@yahoo.com> wrote:
>
> > A common estimate is that orbital flight is an order of magnitude more
> > difficult than suborbital flight,
>
> snip
>
> I appreciate what you are proposing here. The problem is that low
> exhaust velocity (compared to mission velocity) leads to high mass
> ratios and very little margin for structure and engines, not to
> mention payload.
>
> It's physically possible to get exhaust velocity comparable to LEO (9
> km/s) using hydrogen heated by beamed energy to 2700 deg K or higher.
>
> Unfortunately, it takes very large laser or microwave generators to
> get ten ton class payload to orbit and the only way they make economic
> sense is for the beamed energy sources to be operating most of the
> time. Which is just fine if you need 500,000 tons per year to GEO.
>
> I have given some thought to the problems here:http://htyp.org/Beamed_energy_propulsion#Numerical_investigation_of_a...
>
> Keith Henson
I think these two ideas for low cost space access can benefit each
other.
As Robert Heinlein observed, once you get to low Earth orbit, you are
half way to anywhere in the Solar System. This is because of the
relatively small amount of extra velocity you need to reach escape
velocity compared to that needed to get out of Earth's gravity well.
Because of this, it has been noted that an SSTO if refueled in orbit,
could travel to the Moon, land, takeoff, and return to Earth on that
one single refueling.
This would clearly provide an even greater impetus for such SSTO's to
be produced, IF we had that fuel in orbit. However, the cost of
getting that fuel in orbit would be comparatively high even if you
brought the costs to orbit down to the $100/kilo range. This is
because the amount of fuel needed for your SSTO might be 10, 20 or
more times that of the mass of the vehicle.
This vastly reverses the cost structure for launch vehicles to LEO
where the costs for the propellant are literally just at a few parts
per thousand. Then an advantage of beamed propulsion is that it can
bring the cost to orbit down close to that of just the energy cost.
This page estimates that energy cost to be about $1 kilo:
The Cost of Launch to Orbit.
http://home.earthlink.net/~kstengel226/astro/cost2orbit.html
You would have to add on the power cost of accelerating the
propellant. But at the high Isp values you say you can get from the
laser-heated hydrogen, which thus requires a low propellant load, this
might only be an additional $1 to $2 per kilo.
The problem with the beamed propulsion is the large initial outlay
for the high power lasers to launch payloads at thousands of kilo
sizes. However, this becomes much more economically feasible if all
you're launching is fuel continually at only say 1 to 10 kilo size.
So production of such small scale beam propulsion would be beneficial
to insure the development of the SSTO's. And once it is seen how
beneficial beam propulsion can be at the small scale being able to
reduce the cost to orbit to the $2 to $3 per kilo range, this would
provide great impetus to invest the large initial outlay needed to
produce the beam propulsion systems for large cargo and manned vehicle
flights.
So what do you estimate would be the required laser power to launch 1
to 10 kilo payloads? How much do you estimate this would cost?
Bob Clark
Robert Clark
> ...
>
> The hardest part of the design is the 2 GW heat exchanger on each
> engine. Miles and miles of fine tubing. Recognizing this as the
> hardest problem, they have built substantial sections of the heat
> exchanger and tested them. To my amazement, they work just fine.
> ...
>
> Keith
>
This might be another example of SSTO's and beamed propulsion helping
each other.
Bob Clark
Robert Clark
> In article
> <Uoqdna6U6dliMcbTnZ2dnUVZ_rydn...@posted.northdakotatelephone>,
> flan...@daktel.com says...
>
> > On 8/29/2011 5:07 AM, Jeff Findley wrote:
>
> > > Seriously though, even if this engine is successful, I doubt it could
> > > compete with a similar vehicle powered by modern liquid fueled rocket
> > > engines. I'd like to see Reaction Engines do a fair comparison between
> > > Skylon and an all rocket powered vehicle with the same payload.
>
> > Assuming they actually got around to building a SABRE engine, it would
> > be fun to hear how they intended to flight test it, as that would
> > require an aircraft something like a X-15 to carry it, which Britain
> > doesn't have.
>
> Lack of an obvious way to perform incremental testing on a SABRE engine
> is a problem. Launching it on top of expendable missiles or expendable
> launch vehicles becomes very expensive very quickly.
>
> Jeff
Remember the airbreather only has to work around Mach 5 to 6.
Suborbital sounding rockets reach these speeds routinely. The V2
reached Mach 5+ in the 40's.
Likely the commercial reusable suborbital vehicles could also test it.
Bob Clark
Jeff Findley
37776a...@f41g2000yqh.googlegroups.com>, rgrego...@yahoo.com
says...
If it's only going to work up to Mach 5 or Mach 6, what's the point?
Orbital velocity is far higher than that!
SABRE and Skylon are cool research topics, but it's very unlikely to
produce hardware which measures up to the hype.
Pat Flannery
>> If they had gone ahead with Ares-1, the fact that it had just the SRB as
>> the first stage and no tankage next to it would have allowed them to
>> install a venting system if desired.
>
> I know the history. The point is that the current destruct system is
> violent and results in a lot of debris. Changing to a nose blow off
> system would help, but that's an expensive design change which partially
> negates using "shuttle derived" hardware.
>
> Better to kill SLS entirely and buy commercial launches. For the tens
> of billions which would be spent on SLS, you could buy *a lot* of EELV
> launches.
I agree completely; the whole Shuttle-derived concept sounded good when
proposed, but next thing you knew, nothing was going to be stock,
particularly the number of segments on the SRB for Ares.
The only thing that it accomplishes is keeping the way-too-expensive
Shuttle infrastructure and jobs around, which is not the right way to go
when new boosters are now available that are far less complex to support.
Pat
Derek Lyons
>On Aug 31, 10:54 pm, fairwa...@gmail.com (Derek Lyons) wrote:
>> Marvin the Martian <mar...@ontomars.org> wrote:
>>
>> >Skylon reaches 1.7 km/s as an air breather
>>
>> Skylon doesn't reach anything - as it doesn't exist except as a stack
>> of papers and some magnetic domains on a hard drive.
>
>That's true. So was the Saturn V and the Space Shuttle at one time.
The difference is I can tell the difference and act as if it exists.
You cannont and act as if it does not.
>ESA: British Skylon spaceplane seems perfectly possible
<snippage breatheless and wordy press release/marketing droid speech.>
>Read the rest here: http://www.theregister.co.uk/2011/05/24/skylon_esa_report/
Why would I want to? I've been reading bullcrap like that for decades
now. More of the same is still just more of the same.
>And there is a lot more if you Google for Skylon ESA.
Again, why would I want to?
Sylvia Else
> In article<9c8c76...@mid.individual.net>, syl...@not.here.invalid
> says...
>>
>> Another person who thinks they can rule Skylon out by means of simple
>> arithmetic. Have you looked at the actual numbers for Skylon? Can you
>> point to a specific 'error' made by Reaction Engines.
>>
>> Did you allow for the effect that a wing has on gravity losses?
>>
>> On its being "worth it" did you look at the economics of having a fully
>> reusable craft doesn't have to be reassembled before each launch, nor
>> lifted into a vertical position on a launch pad?
>
> You remind me of annoying people at work who use the very same argument
> to justify the complete lack of a sane design for their code. When
> accused of writing spaghetti code, their defense is, "Show me one line
> of code I don't need".
>
> Well that's not the point, now is it?
The point is that Marvin the Martian though that some entirely spurious
argument based on kinetic energy could demonstrate that the benefits of
using the Sabre engine will amount to very little.
> The point is why go down this
> insane path in the first place? Whether it's spaghetti code or air
> breathing engines on a launch vehicle doesn't matter. The problem is
> the initial assumptions are crap.
What are the initial assumptions that you claim are crap? If you're
going to characterise them that way, you might at least have the
intellectual honesty to state what exactly it is that you're criticising.
>
> The development costs for Skylon/SABRE are far too high for far too
> little payoff. Since SABRE has to work as a rocket engine for a good
> portion of Skylon's launch trajectory, just ditch the air breathing
> engine and use a rocket engine the entire way. This results in a
> simpler design, which ultimately means lower costs. The added LOX
> needed costs pennies per pound, but the big win is losing the complex
> engines, complex aerodynamics (wings and engine intakes), and etc.
But the added LOX increases the required tankage, which increases the
dry mass directly. It also increases the mass of the rest of the
structure, which has to carry the LOX and tankage. The extra LOX, by
virtue of being very cold, and liquid, contains less energy than oxygen
removed from the air.
> If you're concerned that a conventional rocket engine won't have the
> performance needed, then work on finishing development of aerospike
> engines. That technology is *far* more mature than SABRE and would cost
> *far* less to develop.
Those who believe that an aerospike engine based pure rocket SSTO can
work are of course free to try to develop such a thing.
Sylvia.
Sylvia Else
The point is that it gives a reasonable payload fraction.
Sylvia.
Marvin the Martian
Yes, I do think it is ruled out by simple math. Math is good. People who
deny math don't make sense to me.
> Have you looked at the actual numbers for Skylon? Can you
> point to a specific 'error' made by Reaction Engines.
It doesn't matter what the engines do up until they reach their maximum
air breathing speed, once that "phase" is over, you have the basic rocket
problem; a flying rocket, but the same problem - except you're starting
at 1.7 km/sec and you have a great deal of extra mass which makes for a
bad mass ratio.
> Did you allow for the effect that a wing has on gravity losses?
Irrelevant.
> On its being "worth it" did you look at the economics of having a fully
> reusable craft doesn't have to be reassembled before each launch, nor
> lifted into a vertical position on a launch pad?
Economics is what drives the use of multistage rockets. Chemical rockets
are restricted to an I_sp < 350 1/sec.
Now, if you were talking a nuclear rocket, then it can be done.
> Sylvia.
Marvin the Martian
> On 2/09/2011 12:22 AM, Jeff Findley wrote:
>> In article<9c8c76...@mid.individual.net>, syl...@not.here.invalid
>> says...
>>>
>>> Another person who thinks they can rule Skylon out by means of simple
>>> arithmetic. Have you looked at the actual numbers for Skylon? Can you
>>> point to a specific 'error' made by Reaction Engines.
>>>
>>> Did you allow for the effect that a wing has on gravity losses?
>>>
>>> On its being "worth it" did you look at the economics of having a
>>> fully reusable craft doesn't have to be reassembled before each
>>> launch, nor lifted into a vertical position on a launch pad?
>>
>> You remind me of annoying people at work who use the very same argument
>> to justify the complete lack of a sane design for their code. When
>> accused of writing spaghetti code, their defense is, "Show me one line
>> of code I don't need".
>>
>> Well that's not the point, now is it?
>
> The point is that Marvin the Martian though that some entirely spurious
> argument based on kinetic energy could demonstrate that the benefits of
> using the Sabre engine will amount to very little.
It wasn't "spurious", it has a fundamental flaw in the concept. There is
a whole lot of effort to gain the first 1.7 km/s with some economy, and
then drag all that air breathing structure to LEO.
Once it is out of the air breathing phase, it is just the rocket equation
with a starting velocity of 1.7 km/s and a very bad thrust ratio.
>> The point is why go down this
>> insane path in the first place? Whether it's spaghetti code or air
>> breathing engines on a launch vehicle doesn't matter. The problem is
>> the initial assumptions are crap.
>
> What are the initial assumptions that you claim are crap? If you're
> going to characterise them that way, you might at least have the
> intellectual honesty to state what exactly it is that you're
> criticising.
>
>
>> The development costs for Skylon/SABRE are far too high for far too
>> little payoff. Since SABRE has to work as a rocket engine for a good
>> portion of Skylon's launch trajectory,
3% is not a "good portion". Energy is everything. Distance means nothing.
Marvin the Martian
> Marvin the Martian <mar...@ontomars.org> wrote:
>
>>Skylon reaches 1.7 km/s as an air breather
>
> Skylon doesn't reach anything - as it doesn't exist except as a stack of
> papers and some magnetic domains on a hard drive.
>
> D.
The Skylon proposal is to reach 1.7 km/s as an air breather.
Sylvia Else
> On Fri, 02 Sep 2011 11:41:57 +1000, Sylvia Else wrote:
>
>> On 2/09/2011 12:22 AM, Jeff Findley wrote:
>>> In article<9c8c76...@mid.individual.net>, syl...@not.here.invalid
>>> says...
>>>>
>>>> Another person who thinks they can rule Skylon out by means of simple
>>>> arithmetic. Have you looked at the actual numbers for Skylon? Can you
>>>> point to a specific 'error' made by Reaction Engines.
>>>>
>>>> Did you allow for the effect that a wing has on gravity losses?
>>>>
>>>> On its being "worth it" did you look at the economics of having a
>>>> fully reusable craft doesn't have to be reassembled before each
>>>> launch, nor lifted into a vertical position on a launch pad?
>>>
>>> You remind me of annoying people at work who use the very same argument
>>> to justify the complete lack of a sane design for their code. When
>>> accused of writing spaghetti code, their defense is, "Show me one line
>>> of code I don't need".
>>>
>>> Well that's not the point, now is it?
>>
>> The point is that Marvin the Martian though that some entirely spurious
>> argument based on kinetic energy could demonstrate that the benefits of
>> using the Sabre engine will amount to very little.
>
> It wasn't "spurious", it has a fundamental flaw in the concept. There is
> a whole lot of effort to gain the first 1.7 km/s with some economy, and
> then drag all that air breathing structure to LEO.
Your argument was based on kinetic energy per unit mass. Then you
expressed the savings as a ration of the kinetic energy per unit mass at
1.7km/s and the kinetic energy per unit mass at orbital speed.
The result is meaningless. The vehicle's total mass at 1.7 km/s is quite
different from its total mass at orbital speed. You might as well be
trying to put apples and oranges into numerical order.
>
> Once it is out of the air breathing phase, it is just the rocket equation
> with a starting velocity of 1.7 km/s and a very bad thrust ratio.
So now it's just back to the usual handwaving assertions. The thrust
ratio is lower than it would have been had the craft been magically put
at that point in its trajectory, but without the Sabre overhead. But
whether the overhead is worth accepting can only be determined by
examining the numbers in detail.
>
>
>>> The point is why go down this
>>> insane path in the first place? Whether it's spaghetti code or air
>>> breathing engines on a launch vehicle doesn't matter. The problem is
>>> the initial assumptions are crap.
>>
>> What are the initial assumptions that you claim are crap? If you're
>> going to characterise them that way, you might at least have the
>> intellectual honesty to state what exactly it is that you're
>> criticising.
>>
>>
>>> The development costs for Skylon/SABRE are far too high for far too
>>> little payoff. Since SABRE has to work as a rocket engine for a good
>>> portion of Skylon's launch trajectory,
>
> 3% is not a "good portion". Energy is everything. Distance means nothing.
Your 3% figure is meaningless, as I indicated above.
In addition, if you want to do energy calculations, you have to look at
all of the energy transfers, not just the energy in the vehicle. A lot
of the energy of originally in the fuel ends up in the atmosphere.
If your approach is valid, it should give an observer independent
measure of the relative merits of a launch system. Try running it from
the point of view of a person already in the destination orbit.
Sylvia.
Sylvia Else
I'm not denying the math. I saying you've got it wrong.
Sylvia.
Derek Lyons
>In addition, if you want to do energy calculations, you have to look at
>all of the energy transfers, not just the energy in the vehicle. A lot
>of the energy of originally in the fuel ends up in the atmosphere.
True of the SABRE as well - and in spades considering how much time it
spends in the atmosphere. But the amount of energy that ends up in
the atmosphere is utterly irrelevant.
Derek Lyons
>On 2/09/2011 12:22 AM, Jeff Findley wrote:
>
>> The development costs for Skylon/SABRE are far too high for far too
>> little payoff. Since SABRE has to work as a rocket engine for a good
>> portion of Skylon's launch trajectory, just ditch the air breathing
>> engine and use a rocket engine the entire way. This results in a
>> simpler design, which ultimately means lower costs. The added LOX
>> needed costs pennies per pound, but the big win is losing the complex
>> engines, complex aerodynamics (wings and engine intakes), and etc.
>
>But the added LOX increases the required tankage, which increases the
>dry mass directly.
So what? Tankage is *cheap*. Complex engines are *expensive*.
>It also increases the mass of the rest of the structure, which has to
>carry the LOX and tankage.
So what?
>The extra LOX, by virtue of being very cold, and liquid, contains less
>energy than oxygen removed from the air.
So what? The amount of energy needed to bring it to same temperature
as that captured by the SABRE engine is but a trivial fraction of the
total involved.
Sylvia Else
> Sylvia Else<syl...@not.here.invalid> wrote:
>
>> In addition, if you want to do energy calculations, you have to look at
>> all of the energy transfers, not just the energy in the vehicle. A lot
>> of the energy of originally in the fuel ends up in the atmosphere.
>
> True of the SABRE as well - and in spades considering how much time it
> spends in the atmosphere. But the amount of energy that ends up in
> the atmosphere is utterly irrelevant.
>
The point was that if one wants to draw inferences from an analysis of
energy transfers, one has to identify them. It's not sufficient to
identify only some of them.
The starting point for a launch vehicle is the energy in its fuel just
before it starts its takeoff/liftoff. The end point is the kinetic and
potential energy in the vehicle in orbit. A lot of the energy initially
in the fuel doesn't end up as kinetic energy and potential energy in the
vehicle, but as heat energy in the atmosphere (eventually - some takes
other forms beforehand). The initial energy required in the fuel is the
sum of the kinetic and potential energies in the vehicle in orbit and
the heat energy left in the atmosphere. The latter is therefore not in
the least bit irrelevant.
Sylvia.
Sylvia Else
> Sylvia Else<syl...@not.here.invalid> wrote:
>
>> On 2/09/2011 12:22 AM, Jeff Findley wrote:
>>
>>> The development costs for Skylon/SABRE are far too high for far too
>>> little payoff. Since SABRE has to work as a rocket engine for a good
>>> portion of Skylon's launch trajectory, just ditch the air breathing
>>> engine and use a rocket engine the entire way. This results in a
>>> simpler design, which ultimately means lower costs. The added LOX
>>> needed costs pennies per pound, but the big win is losing the complex
>>> engines, complex aerodynamics (wings and engine intakes), and etc.
>>
>> But the added LOX increases the required tankage, which increases the
>> dry mass directly.
>
> So what? Tankage is *cheap*. Complex engines are *expensive*.
The cost is not the issue. The mass is.
>
>> It also increases the mass of the rest of the structure, which has to
>> carry the LOX and tankage.
>
> So what?
It cuts into the payload fraction.
>
>> The extra LOX, by virtue of being very cold, and liquid, contains less
>> energy than oxygen removed from the air.
>
> So what? The amount of energy needed to bring it to same temperature
> as that captured by the SABRE engine is but a trivial fraction of the
> total involved.
It's not a large amount, but it's certainly enough to be significant.
At the start of the takeoff run, say 20C, it's about 12.6kJ/mol. Towards
the end of the air-breathing ascent, the recovered intake temperature
reaches about 1100C. Heating the LOX to that temperature takes about
43kJ/mol.
The heat of formation of water in the gaseous state is about 242kJ/mol.
A mol of oxygen is used to form two mols of water, so one has to compare
the 43kJ/mol with 2*242kJ/mol.
So towards the end of the air-breathing ascent, capturing the oxygen
from the air gives about a 9% energy gain over carrying LOX.
The greatest benefit, though, comes from using the air as reaction mass.
Since the exhaust is slower, it carries away less of the available energy.
Sylvia.
Derek Lyons
>On 2/09/2011 5:11 PM, Derek Lyons wrote:
>> Sylvia Else<syl...@not.here.invalid> wrote:
>>
>>> In addition, if you want to do energy calculations, you have to look at
>>> all of the energy transfers, not just the energy in the vehicle. A lot
>>> of the energy of originally in the fuel ends up in the atmosphere.
>>
>> True of the SABRE as well - and in spades considering how much time it
>> spends in the atmosphere. But the amount of energy that ends up in
>> the atmosphere is utterly irrelevant.
>>
>
>The point was that if one wants to draw inferences from an analysis of
>energy transfers, one has to identify them. It's not sufficient to
>identify only some of them.
In other words "having been caught with my hand in the cookie jar, now
I'm going to claim I was never in the kitchen in the first place".
>The starting point for a launch vehicle is the energy in its fuel just
>before it starts its takeoff/liftoff. The end point is the kinetic and
>potential energy in the vehicle in orbit. A lot of the energy initially
>in the fuel doesn't end up as kinetic energy and potential energy in the
>vehicle, but as heat energy in the atmosphere (eventually - some takes
>other forms beforehand). The initial energy required in the fuel is the
>sum of the kinetic and potential energies in the vehicle in orbit and
>the heat energy left in the atmosphere. The latter is therefore not in
>the least bit irrelevant.
Since the amount of energy radiating from the nozzle is the same (I.E.
within the atmosphere and without), where it's radiated to is indeed
irrelevant. That you ignore the 'without the atmosphere' part, which
makes up most of powerd flight time, is telling.
Derek Lyons
>On 2/09/2011 5:12 PM, Derek Lyons wrote:
>> Sylvia Else<syl...@not.here.invalid> wrote:
>>
>>> On 2/09/2011 12:22 AM, Jeff Findley wrote:
>>>
>>>> The development costs for Skylon/SABRE are far too high for far too
>>>> little payoff. Since SABRE has to work as a rocket engine for a good
>>>> portion of Skylon's launch trajectory, just ditch the air breathing
>>>> engine and use a rocket engine the entire way. This results in a
>>>> simpler design, which ultimately means lower costs. The added LOX
>>>> needed costs pennies per pound, but the big win is losing the complex
>>>> engines, complex aerodynamics (wings and engine intakes), and etc.
>>>
>>> But the added LOX increases the required tankage, which increases the
>>> dry mass directly.
>>
>> So what? Tankage is *cheap*. Complex engines are *expensive*.
>
>The cost is not the issue. The mass is.
Fit the booster with sufficient engines and the mass is no longer a
problem. (That's SpaceX's big breakthrough - new engines.)
>>
>>> It also increases the mass of the rest of the structure, which has to
>>> carry the LOX and tankage.
>>
>> So what?
>
>It cuts into the payload fraction.
Of course - the Holy Number must be preserved at all costs. No matter
how ludicrous the dollars and man hours spent to maintain it, the Holy
Tradition says it Must Be So.
Seriously, maintaining mass faction to the near exclusion of all else
makes boosters far more expensive than they would be if the designers
just let the vehicle grow some.
>>> The extra LOX, by virtue of being very cold, and liquid, contains less
>>> energy than oxygen removed from the air.
>>
>> So what? The amount of energy needed to bring it to same temperature
>> as that captured by the SABRE engine is but a trivial fraction of the
>> total involved.
>
>It's not a large amount, but it's certainly enough to be significant.
>
>At the start of the takeoff run, say 20C, it's about 12.6kJ/mol. Towards
>the end of the air-breathing ascent, the recovered intake temperature
>reaches about 1100C. Heating the LOX to that temperature takes about
>43kJ/mol.
>
>The heat of formation of water in the gaseous state is about 242kJ/mol.
>A mol of oxygen is used to form two mols of water, so one has to compare
>the 43kJ/mol with 2*242kJ/mol.
>
>So towards the end of the air-breathing ascent, capturing the oxygen
>from the air gives about a 9% energy gain over carrying LOX.
Of course, you ignore the immense amount of time before the end of the
atmopsheric run when the advantage is much less and the immense amount
of time after the end of the atmospheric run when the advantage is
non-existent. These periods are inconvient to your arguement and you
thus dismiss them.
>The greatest benefit, though, comes from using the air as reaction mass.
>Since the exhaust is slower, it carries away less of the available energy.
Or course, much of the energy retained is either lost to drag or
otherwise contributes nothing to the actual goal of reaching orbit.
Keith Henson
> On 2/09/2011 4:17 AM, Jeff Findley wrote:
>
>
>
> > In article<78a8b18c-c520-433d-bfbf-
It's still not a lot, 15 tons out of 300 is 5%.
But it will put 30 tons of second stage in a high suborbital. (See
page ten? of the Skylon User's Manual) 500 MW of lasers heating
hydrogen in that stage would put 20 of the 30 tons in GEO. See The
Oil Drum article I wrote a few months ago. Or ask me for the spread
sheet.
Keith Henson
PS payload is *very* sensitive to exhaust velocity.
> Sylvia.
Marvin the Martian
Then show me the math.
Marvin the Martian
That's not correct. I compared the K.E. of the orbiting payload to the
K.E. of the payload at 1.7 km/s.
>> Once it is out of the air breathing phase, it is just the rocket
>> equation with a starting velocity of 1.7 km/s and a very bad thrust
>> ratio.
>
> So now it's just back to the usual handwaving assertions.
You don't like or understand math, apparently. IT appears that you're
mindlessly dismissive of any argument against your religious beliefs.
Ergo, YOU don't matter. I cannot reason with the unreasonable. Sorry - I
was clear enough for a rational person, but not a goat or a pig, and I
don't worry about the irrational opinions of farm animals.
Chris
> But the added LOX increases the required tankage, which increases the
> dry mass directly.
Dry mass is generally seen as bad because it is related to the factor
we actually care about: the number of hours it takes to design the
damn thing, which is a significant cost driver. Spending a few
gigabucks (last I saw Skylon's plan involved spending 10 billion USD)
to get lower dry mass seems counterproductive unless you are planning
on flying damn near daily for years and years.
If we're living in a dream world (and any plan where Step 2 is
'Someone gives us 10 billion dollars' is indistinguishable from a
dream world) I'd rather spend 10 billion on financing 100 different
programs and ideas with 100 million dollars, rather than spend
everything on getting one engine that might or might not work and
might or might not be in the right direction. 95 of those 100 ideas
will probably fail. 3 of those ideas will be okay. One ought to be
pretty good. And one will actually be brilliant. The problem is, it is
difficult to know ahead of time which is which, hence trying a lot.
> It also increases the mass of the rest of the
> structure, which has to carry the LOX and tankage.
But does it increase the mass by more than the airbreathing engine
does is the really important question. Yes, Skylon intends to use the
same engine, but it will still weigh more and be less efficient than a
pure-rocket engine, for any given technology level, simply by virtue
of the basic design requirements.
> The extra LOX, by
> virtue of being very cold, and liquid, contains less energy than oxygen
> removed from the air.
But at a much higher density.
Chris Manteuffel
Patok
> On Fri, 02 Sep 2011 13:39:54 +1000, Sylvia Else wrote:
>
>>
>> The result is meaningless. The vehicle's total mass at 1.7 km/s is quite
>> different from its total mass at orbital speed. You might as well be
>> trying to put apples and oranges into numerical order.
>
> That's not correct. I compared the K.E. of the orbiting payload to the
> K.E. of the payload at 1.7 km/s.
Indeed you did. But why? The kinetic energy in orbit is not what we're
interested in, it is the energy spent getting there. I hope you understand that
they are not the same.
--
You'd be crazy to e-mail me with the crazy. But leave the div alone.
*
Whoever bans a book, shall be banished. Whoever burns a book, shall burn.
Marvin the Martian
> Marvin the Martian wrote:
>> On Fri, 02 Sep 2011 13:39:54 +1000, Sylvia Else wrote:
>>
>>
>>> The result is meaningless. The vehicle's total mass at 1.7 km/s is
>>> quite different from its total mass at orbital speed. You might as
>>> well be trying to put apples and oranges into numerical order.
>>
>> That's not correct. I compared the K.E. of the orbiting payload to the
>> K.E. of the payload at 1.7 km/s.
>
> Indeed you did. But why? The kinetic energy in orbit is not what
> we're
> interested in, it is the energy spent getting there. I hope you
> understand that they are not the same.
All I needed to do is show that the advantage of being an air breather to
1.7 km/s was a trivial advantage.
Even that much math was dismissed with no rational reason. The rebuttal
seems to be I should have spent much more time making my argument better.
Of course, it would still have been dismissed for no rational reason if
I had spent more time on it.
Sylvia Else
> Since the amount of energy radiating from the nozzle is the same (I.E.
> within the atmosphere and without), where it's radiated to is indeed
> irrelevant. That you ignore the 'without the atmosphere' part, which
> makes up most of powerd flight time, is telling.
Radiated, for crying out loud?
Look at how the kinetic energy in the exhaust varies for a given
momentum change, depending on the mass of the exhaust.
Sylvia.
Sylvia Else
Which would be great if the payload were self launching. But there's
this huge vehicle around it, and right up until the moment when orbit is
reached, varying amounts of propellant.
So now all you need to do is show how comparing the kinetic energy of
the payload alone at different points in the ascent has any bearing on
the matter.
>>> Once it is out of the air breathing phase, it is just the rocket
>>> equation with a starting velocity of 1.7 km/s and a very bad thrust
>>> ratio.
>>
>> So now it's just back to the usual handwaving assertions.
>
> You don't like or understand math, apparently. IT appears that you're
> mindlessly dismissive of any argument against your religious beliefs.
I haven't seen you offer any mathematics that's applied properly.
>
> Ergo, YOU don't matter. I cannot reason with the unreasonable. Sorry - I
> was clear enough for a rational person, but not a goat or a pig, and I
> don't worry about the irrational opinions of farm animals.
So you're reduced to insults. That didn't take long.
Sylvia.
Sylvia Else
> On Sep 1, 9:41 pm, Sylvia Else<syl...@not.here.invalid> wrote:
>
>> But the added LOX increases the required tankage, which increases the
>> dry mass directly.
>
> Dry mass is generally seen as bad because it is related to the factor
> we actually care about: the number of hours it takes to design the
> damn thing, which is a significant cost driver. Spending a few
> gigabucks (last I saw Skylon's plan involved spending 10 billion USD)
> to get lower dry mass seems counterproductive unless you are planning
> on flying damn near daily for years and years.
>
> If we're living in a dream world (and any plan where Step 2 is
> 'Someone gives us 10 billion dollars' is indistinguishable from a
> dream world)
I don't know which step 2 you're referring to. Certainly RE's step two
is not to be given 10 billion dollars.
Step 2 is to get a $350 million in investor funding, based on the
success of a test of a core part of the engine.
http://www.space.com/11414-skylon-space-plane-british-engine-test.html
>
> But does it increase the mass by more than the airbreathing engine
> does is the really important question. Yes, Skylon intends to use the
> same engine, but it will still weigh more and be less efficient than a
> pure-rocket engine, for any given technology level, simply by virtue
> of the basic design requirements.
The really important question is whether the vehicle reaches orbit for a
given payload with a lower dry mass than it would have had to have if it
had been purely rocket driven. RE's calculations say that it will.
Skylon's ascent data is given in
http://www.reactionengines.co.uk/downloads/C1_trajectory_output.xls
Those claiming it won't work should point to where they believe that
data cannot be realised.
>
>> The extra LOX, by
>> virtue of being very cold, and liquid, contains less energy than oxygen
>> removed from the air.
>
> But at a much higher density.
The density isn't paticularly interesting. The pressure is more
important, since the pressure has to be increased to get the propellant
into the cumbustion chamber. Increasing the pressure of a liquid
requires little energy, whereas increasing the pressure of a gas
requires quite a lot. However, the essence of the Sabre cycle is to take
the incoming air at its recovered temperature, extract the heat energy
from it, thus cooling it down, and then use that energy to compress it,
thus increasing is pressure.
Sylvia.
Sylvia Else
I wasn't the one purporting to prove that the benefit was only 3%. Since
what you offered was inapplicable, I can't tell what it was you were
trying to offer.
Sylvia.
Sylvia Else
The main issue with payload fraction is that the smaller it gets the
more sensitive the viability of the system gets to dry weight. This
doesn't matter so much once you have a launcher that is known to deliver
some payload fraction, even if it's small. But during development, a low
payload fraction gives you a huge risk that mass overruns on the vehicle
will wipe out a large part, or even all, of the payload.
Even new airliner designs are routinely overweight when they enter
service, with weight reduction programs being required to get the dry
weight down to the levels promised on subsequent deliveries.
Lower than expected engine performance can do the same of course, but at
least you can go a good way towards checking that without having to
build the complete vehicle.
Sylvia.
Derek Lyons
>The really important question is whether the vehicle reaches orbit for a
>given payload with a lower dry mass than it would have had to have if it
>had been purely rocket driven.
That's only an important question when comparing the dry mass of an
airbreathing system and a rocket propelled system. It ignore a whole
host of other comparisons, It's also roughly as useful a measurement
as comparing the weight of rice the designer of each system consumed
in the past annum.
Derek Lyons
>500 MW of lasers heating hydrogen in that stage would put 20 of
>the 30 tons in GEO. See The Oil Drum article I wrote a few
>months ago. Or ask me for the spread sheet.
If I want to read fantasy, I'll re-read Lord of The Rings.
Keith Henson
Exactly. The dry vehicle is 49 tons. It would not take a lot of
overruns to eat the 15 ton payload.
But ESA seems to think Reaction Engines knows what they are doing and
have enough mass margin to hit their target performance.
> Even new airliner designs are routinely overweight when they enter
> service, with weight reduction programs being required to get the dry
> weight down to the levels promised on subsequent deliveries.
>
> Lower than expected engine performance can do the same of course, but at
> least you can go a good way towards checking that without having to
> build the complete vehicle.
Well stated.
Keith
> Sylvia.
Keith Henson
> On Fri, 02 Sep 2011 16:41:17 -0400, Patok wrote:
> > Marvin the Martian wrote:
> >> On Fri, 02 Sep 2011 13:39:54 +1000, Sylvia Else wrote:
>
> >>> The result is meaningless. The vehicle's total mass at 1.7 km/s is
> >>> quite different from its total mass at orbital speed. You might as
> >>> well be trying to put apples and oranges into numerical order.
>
> >> That's not correct. I compared the K.E. of the orbiting payload to the
> >> K.E. of the payload at 1.7 km/s.
>
> > Indeed you did. But why? The kinetic energy in orbit is not what
> > we're
> > interested in, it is the energy spent getting there. I hope you
> > understand that they are not the same.
>
> All I needed to do is show that the advantage of being an air breather to
> 1.7 km/s was a trivial advantage.
It's not. It makes the difference between significant payload for a
reusable SSTO and none (or less than none). From 26 km up it's around
8 km/s to orbit if you start with no velocity (ignoring gravity and
drag loss).
8/4.5 is a delta V of 1.77 x Ve, giving a mass ratio of 5.9, 17%
payload plus dry mass.
6.3/4.5 is 1.4, a mass ratio of 4.1 or 24% payload plus dry mass.
Subtract out dry mass of 15% (very optimistic for reusable) and the
payload goes from 2% to 9%. An increase of payload better than 4 x is
not a trivial advantage.
Keith
PS, launch to the east from the equator and Uabs is 2 km/s when going
to rocket mode. The wings go a long way to counter gravity loss as
well.
Keith Henson
> On Sep 1, 9:41 pm, Sylvia Else <syl...@not.here.invalid> wrote:
>
> > But the added LOX increases the required tankage, which increases the
> > dry mass directly.
>
> Dry mass is generally seen as bad because it is related to the factor
> we actually care about: the number of hours it takes to design the
> damn thing, which is a significant cost driver. Spending a few
> gigabucks (last I saw Skylon's plan involved spending 10 billion USD)
> to get lower dry mass seems counterproductive unless you are planning
> on flying damn near daily for years and years.
If you use them to build power satellites, the *minimum* rate is 3
flights per hour.
And that requires several billion in lasers to get the second stage to
GEO.
Keith Henson
Thomas Womack
Keith Henson <hkeith...@gmail.com> wrote:
>On Sep 2, 11:21=A0am, Chris <cmant...@gmail.com> wrote:
>> On Sep 1, 9:41=A0pm, Sylvia Else <syl...@not.here.invalid> wrote:
>>
>> > But the added LOX increases the required tankage, which increases the
>> > dry mass directly.
>>
>> Dry mass is generally seen as bad because it is related to the factor
>> we actually care about: the number of hours it takes to design the
>> damn thing, which is a significant cost driver. Spending a few
>> gigabucks (last I saw Skylon's plan involved spending 10 billion USD)
>> to get lower dry mass seems counterproductive unless you are planning
>> on flying damn near daily for years and years.
>
>If you use them to build power satellites, the *minimum* rate is 3
>flights per hour.
>
>And that requires several billion in lasers to get the second stage to
>GEO.
Is there anything else you can do with a laser field? Jordin Kare has
a very enthusiastic presentation in which he reckons the adaptive
optics can be made cheap enough to use fields of one-metre mirrors
with adaptive secondaries; this has the advantage that the laser +
one-metre mirror + mount fits in a shipping container and is
portable rather than being a huge inconvenient installation in the
wrong place.
Tom
Robert Clark
> On Sep 1, 6:43 pm, Sylvia Else <syl...@not.here.invalid> wrote:
>
> > The point is that it gives a reasonable payload fraction.
>
> It's still not a lot, 15 tons out of 300 is 5%.
>
> But it will put 30 tons of second stage in a high suborbital. (See
> page ten? of the Skylon User's Manual) 500 MW of lasers heating
> hydrogen in that stage would put 20 of the 30 tons in GEO. See The
> Oil Drum article I wrote a few months ago. Or ask me for the spread
> sheet.
>
> Keith Henson
>
I gather you are saying that 25,000 watts of laser heating can send 1
kg from high suborbital to GEO. That amount of wattage is not
particularly high. It's the amount of transmission power of a medium
sized radio station. You would lose power though in conversion to
laser transmission.
How much would be the laser power for 1 kg of (vehicle + payload) to
launch from the ground but just to LEO?
This can be profitable for launching propellant currently, aside for
any usage of refueling SSTO's. The cost of getting a satellite to GEO
is about twice that of getting to LEO. If you could cut the cost of
the LEO to GEO portion of the trip by 1 to 2 orders of magnitude, then
that would result in a significant savings to the satellite companies.
In this scenario you would keep a reusable space transfer tug in
orbit to carry the satellite from LEO to GEO. The laser launch system
would refuel the space tug with propellant. The Centaur V1 upper stage
could serve as the hydrogen fueled space tug. It uses a version of the
RL10 engine. This engine in a sea level version was shown to be highly
reusable when used on the DC-X test rocket:
Engine reusability.
http://yarchive.net/space/rocket/engine_reusability.html
This method would also allow satellite servicing and refueling in
GEO, which is now being considered:
Wed, 24 February, 2010
DLR Takes Step Toward In-orbit Servicing Demonstration.
By Peter B. de Selding
http://www.spacenews.com/civil/022410-dlr-takes-step-toward-in-orbit-servicing-demo.html
Fri, 2 April, 2010
NASA Plans To Refuel Mock Satellite at the Space Station.
By Debra Werner
http://www.spacenews.com/civil/100402-nasa-plans-refuel-mock-satellite.html
Bob Clark
bob haller
>
> This method would also allow satellite servicing and refueling in
> GEO, which is now being considered:
>
> Wed, 24 February, 2010
> DLR Takes Step Toward In-orbit Servicing Demonstration.
>
> Fri, 2 April, 2010
> NASA Plans To Refuel Mock Satellite at the Space Station.
>
> Bob Clark
as long as its cheaper to launch a brand new satellite than repair one
already in orbit the service business is a loser.
besides new satellites that last at least 10 years usualy come with
state of the art upgrades, like electronics etc.
most in orbit birds fail from running out of station keeping fuel or
fail to make the proper orbit.
the company that wants to dock with a existing satellite and use the
servicing vehicle for station keeping and perhaps power has a
excellent idea.
all new birds would be built with a universal docking ability and
perhaps a plug in power port.
service vehicle docks provides all new fuel supply to station keep and
perhaps move bird in orbit, could also plug in and provide power.
the cost of having a station worker service a satellite would truly be
astronomical at its best.
has nasa ever released a cost per hour for working hours of a
astronaut in ISS?
Robert Clark
> On Sep 4, 2:19 am, Robert Clark <rgregorycl...@gmail.com> wrote:
As indicated in that "NASA Plans To Refuel Mock Satellite at the
Space Station" article I cited, the plan is for it to be robotic
refueling and serving. DEXTRE is mentioned in that article. Robonaut
could also be used.
DARPA has also already successfully tested remote controlled
refueling of satellites:
In-space satellite servicing tests come to an end
BY STEPHEN CLARK
SPACEFLIGHT NOW
Posted: July 4, 2007
http://spaceflightnow.com/news/n0707/04orbitalexpress/
You are correct though that most of the satellite makers are cool to
the idea, though Intellsat has purchased such a refueling operation
for some of its satellites:
Wed, 16 March, 2011
Satellite Builders Not Enthusiastic About In-orbit Servicing Project.
By Peter B. de Selding
"WASHINGTON — The world’s principal commercial satellite manufacturers
on March 15 unanimously dismissed a proposed satellite in-orbit
service project backed by Intelsat and Canada’s MDA, saying the idea,
while intriguing, is unlikely to make business sense for years.
"Addressing the Satellite 2011 conference here, officials from these
companies said that while there appears to be no technical obstacle to
robotically refueling satellites or performing minor repairs, closing
the business case appears next to impossible without heavy government
backing.
"Many of these officials had not yet been fully briefed on plans by
Luxembourg- and Washington-based Intelsat — the biggest commercial
customer for these manufacturers — to begin use of MDA’s Space
Infrastructure Services (SIS) robotic servicing unit as soon as 2015.
"Intelsat has contracted with Richmond, British Columbia-based MDA to
purchase 1,000 kilograms of fuel to extend the lives of several
Intelsat satellites by between three and five years each. Intelsat has
agreed to pay $280 million for the service assuming that all the
ordered fuel is successfully transferred to aging Intelsat
satellites."
http://www.spacenews.com/satellite_telecom/110316-sat-builders-refueling-plans.html
Undoubtedly the cost would be reduced significantly if the cost of
getting the fuel to orbit was reduced by 3 orders of magnitude and the
refueling vehicle was reusable.
Bob Clark
Keith Henson
wrote:
> In article <9e1b5991-46a4-42a1-949a-99465748c...@x21g2000prd.googlegroups.com>,
> Keith Henson <hkeithhen...@gmail.com> wrote:
>
>
>
> >On Sep 2, 11:21=A0am, Chris <cmant...@gmail.com> wrote:
> >> On Sep 1, 9:41=A0pm, Sylvia Else <syl...@not.here.invalid> wrote:
>
> >> > But the added LOX increases the required tankage, which increases the
> >> > dry mass directly.
>
> >> Dry mass is generally seen as bad because it is related to the factor
> >> we actually care about: the number of hours it takes to design the
> >> damn thing, which is a significant cost driver. Spending a few
> >> gigabucks (last I saw Skylon's plan involved spending 10 billion USD)
> >> to get lower dry mass seems counterproductive unless you are planning
> >> on flying damn near daily for years and years.
>
> >If you use them to build power satellites, the *minimum* rate is 3
> >flights per hour.
>
> >And that requires several billion in lasers to get the second stage to
> >GEO.
>
> Is there anything else you can do with a laser field?
Especially in combination with some redirection mirrors in orbit (GEO
is best in my opinion) you can make very short work of cleaning up the
space junk. Short pulse ablation is better, but even CW head on to
the direction of travel at apogee will drop space junk into the
atmosphere or vaporize it.
> Jordin Kare has
> a very enthusiastic presentation in which he reckons the adaptive
> optics can be made cheap enough to use fields of one-metre mirrors
> with adaptive secondaries; this has the advantage that the laser +
> one-metre mirror + mount fits in a shipping container and is
> portable rather than being a huge inconvenient installation in the
> wrong place.
Jordin's main reason for putting the lasers in shipping containers is
so they can be built in a factory. Much less expensive than field
installations.
Keith
> Tom
Keith Henson
> On Sep 2, 11:35 am, Keith Henson <hkeithhen...@gmail.com> wrote:
>
> > On Sep 1, 6:43 pm, Sylvia Else <syl...@not.here.invalid> wrote:
> >...
>
> > > The point is that it gives a reasonable payload fraction.
>
> > It's still not a lot, 15 tons out of 300 is 5%.
>
> > But it will put 30 tons of second stage in a high suborbital. (See
> > page ten? of the Skylon User's Manual) 500 MW of lasers heating
> > hydrogen in that stage would put 20 of the 30 tons in GEO. See The
> > Oil Drum article I wrote a few months ago. Or ask me for the spread
> > sheet.
>
> > Keith Henson
>
> I gather you are saying that 25,000 watts of laser heating can send 1
> kg from high suborbital to GEO.
Right. The you can use a relatively low acceleration for a long time.
> That amount of wattage is not
> particularly high. It's the amount of transmission power of a medium
> sized radio station. You would lose power though in conversion to
> laser transmission.
> How much would be the laser power for 1 kg of (vehicle + payload) to
> launch from the ground but just to LEO?
The rule of thumb is a MW per kg, GW per ton. It's the limited amount
of time you have with the vehicle in view of the beamed energy
source. That dictates high acceleration and needs lots of power. I
have been working on various concepts to extent the time and reduce
the power needed.
> This can be profitable for launching propellant currently, aside for
> any usage of refueling SSTO's. The cost of getting a satellite to GEO
> is about twice that of getting to LEO. If you could cut the cost of
> the LEO to GEO portion of the trip by 1 to 2 orders of magnitude, then
> that would result in a significant savings to the satellite companies.
At 60 tons per hour, the power cost is around 17 kWh/kg. The capital
cost is not very high, but only if it is operated close to full time.
http://lists.extropy.org/pipermail/extropy-chat/2011-April/065859.html
> In this scenario you would keep a reusable space transfer tug in
> orbit to carry the satellite from LEO to GEO.
I don't think there is a business case for a tug, unfortunately. In a
power satellite production project, the second stages go one way and
are used up as parts for making power satellites.
Keith Henson
> The laser launch system
> would refuel the space tug with propellant. The Centaur V1 upper stage
> could serve as the hydrogen fueled space tug. It uses a version of the
> RL10 engine. This engine in a sea level version was shown to be highly
> reusable when used on the DC-X test rocket:
>
> Engine reusability.http://yarchive.net/space/rocket/engine_reusability.html
>
> This method would also allow satellite servicing and refueling in
> GEO, which is now being considered:
>
> Wed, 24 February, 2010
> DLR Takes Step Toward In-orbit Servicing Demonstration.
> By Peter B. de Seldinghttp://www.spacenews.com/civil/022410-dlr-takes-step-toward-in-orbit-...
>
> Fri, 2 April, 2010
> NASA Plans To Refuel Mock Satellite at the Space Station.
> By Debra Wernerhttp://www.spacenews.com/civil/100402-nasa-plans-refuel-mock-satellit...
>
> Bob Clark