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New smallsat launcher start-up.
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Robert Clark
Sep 18, 2014, 9:12:13 AM9/18/14
to
SpaceX Alum Goes After Falcon 1 Market With Firefly.
Aerospike revival, advances in composite structures are shaping design of
low-cost smallsat launch vehicle.
Aug 25, 2014 Frank Morring, Jr. | Aviation Week & Space Technology
http://aviationweek.com/space/spacex-alum-goes-after-falcon-1-market-firefly
Excellent news! As described in the article it will use both lightweight
composite tanks and altitude compensating aerospike nozzles.
This small launcher is to be two stage. But scaling it up using both of
these high efficiency methods at the same time could result in a SSTO.
Bob Clark
------------------------------------------------------------------
Single-stage-to-orbit was already shown possible 50 years ago
with the Titan II first stage.
In fact, contrary to popular belief SSTO's are actually easy.
Just use the most efficient engines and stages at the same time,
and the result will automatically be SSTO.
Blog: Http://Exoscientist.blogspot.com
------------------------------------------------------------------
Aerospike revival, advances in composite structures are shaping design of
low-cost smallsat launch vehicle.
Aug 25, 2014 Frank Morring, Jr. | Aviation Week & Space Technology
http://aviationweek.com/space/spacex-alum-goes-after-falcon-1-market-firefly
Excellent news! As described in the article it will use both lightweight
composite tanks and altitude compensating aerospike nozzles.
This small launcher is to be two stage. But scaling it up using both of
these high efficiency methods at the same time could result in a SSTO.
Bob Clark
------------------------------------------------------------------
Single-stage-to-orbit was already shown possible 50 years ago
with the Titan II first stage.
In fact, contrary to popular belief SSTO's are actually easy.
Just use the most efficient engines and stages at the same time,
and the result will automatically be SSTO.
Blog: Http://Exoscientist.blogspot.com
------------------------------------------------------------------
J. Clarke
Sep 18, 2014, 9:30:55 AM9/18/14
to
In article <lvelni$r3v$1...@dont-email.me>,
rgrego...@gmSPAMBLOCKail.com says...
"showed possible" 50 years ago? I have seen no analysis that suggests
that it comes even close to having enough delta-v to achieve orbit.
rgrego...@gmSPAMBLOCKail.com says...
> Single-stage-to-orbit was already shown possible 50 years ago
> with the Titan II first stage.
> In fact, contrary to popular belief SSTO's are actually easy.
> Just use the most efficient engines and stages at the same time,
> and the result will automatically be SSTO.
Do you have a source for the contention that the Titan II first stage
> with the Titan II first stage.
> In fact, contrary to popular belief SSTO's are actually easy.
> Just use the most efficient engines and stages at the same time,
> and the result will automatically be SSTO.
"showed possible" 50 years ago? I have seen no analysis that suggests
that it comes even close to having enough delta-v to achieve orbit.
Jeff Findley
Sep 18, 2014, 10:03:16 AM9/18/14
to
In article <MPG.2e84a8669...@news.newsguy.com>,
jclark...@cox.net says...
http://en.wikipedia.org/wiki/Single-stage-to-orbit
From above:
While single-stage rockets were once thought to be beyond reach,
advances in materials technology and construction techniques have
shown them to be possible. For example, calculations show that
the Titan II first stage, launched on its own, would have a
25-to-1 ratio of fuel to vehicle hardware.[10] It has a
sufficiently efficient engine to achieve orbit, but without
carrying much payload.[11]
11. Mitchell Burnside-Clapp (February 1997). "A LO2/Kerosene
SSTO Rocket Design". Retrieved 2009-09-14.
This analysis was discussed many times in the "old days" in the
sci.space newsgroups.
Jeff
--
"the perennial claim that hypersonic airbreathing propulsion would
magically make space launch cheaper is nonsense -- LOX is much cheaper
than advanced airbreathing engines, and so are the tanks to put it in
and the extra thrust to carry it." - Henry Spencer
jclark...@cox.net says...
>
> Do you have a source for the contention that the Titan II first stage
> "showed possible" 50 years ago? I have seen no analysis that suggests
> that it comes even close to having enough delta-v to achieve orbit.
This is *very* easy to find.
> Do you have a source for the contention that the Titan II first stage
> "showed possible" 50 years ago? I have seen no analysis that suggests
> that it comes even close to having enough delta-v to achieve orbit.
http://en.wikipedia.org/wiki/Single-stage-to-orbit
From above:
While single-stage rockets were once thought to be beyond reach,
advances in materials technology and construction techniques have
shown them to be possible. For example, calculations show that
the Titan II first stage, launched on its own, would have a
25-to-1 ratio of fuel to vehicle hardware.[10] It has a
sufficiently efficient engine to achieve orbit, but without
carrying much payload.[11]
11. Mitchell Burnside-Clapp (February 1997). "A LO2/Kerosene
SSTO Rocket Design". Retrieved 2009-09-14.
This analysis was discussed many times in the "old days" in the
sci.space newsgroups.
Jeff
--
"the perennial claim that hypersonic airbreathing propulsion would
magically make space launch cheaper is nonsense -- LOX is much cheaper
than advanced airbreathing engines, and so are the tanks to put it in
and the extra thrust to carry it." - Henry Spencer
J. Clarke
Sep 18, 2014, 2:06:12 PM9/18/14
to
In article <MPG.2e84b0325...@reader80.eternal-september.org>,
jeff.f...@siemens.nospam.com says...
wikepedia does not mean that it is true, or that it is any kind of
consensus, or that it is even supported by the references given.
Reference 10 has a good deal of information about the Titan II but
nothing about its first stage being able to achieve orbit. Reference 11
has only one sentence concerning the Titan II, "The historical examples
of the extraordinary mass fractions of the Titan II first stage,
the Atlas, and the Saturn first stage are all persuasive."
I'm sorry but I am not seeing anything persuasive there.
Then it should be easy to find a discussion with some kind of
calculation or test results to support the contention.
According to someone who claims to have done flight mechanics on the
program at Lockheed and who thus may be taken to be authoritative about
the real-world performance of the Titan II, the actual payload that a
Titan II first stage could deliver to a 90 mile orbit is _negative_ 1400
pounds.
<https://groups.google.com/d/msg/sci.space.policy/F3G37kuJdjA/pjFf2fh_H9
gJ>
In other words it ain't gonna come close.
jeff.f...@siemens.nospam.com says...
>
> In article <MPG.2e84a8669...@news.newsguy.com>,
> jclark...@cox.net says...
> >
> > Do you have a source for the contention that the Titan II first stage
> > "showed possible" 50 years ago? I have seen no analysis that suggests
> > that it comes even close to having enough delta-v to achieve orbit.
>
> This is *very* easy to find.
>
> http://en.wikipedia.org/wiki/Single-stage-to-orbit
>
> From above:
>
> While single-stage rockets were once thought to be beyond reach,
> advances in materials technology and construction techniques have
> shown them to be possible. For example, calculations show that
> the Titan II first stage, launched on its own, would have a
> 25-to-1 ratio of fuel to vehicle hardware.[10] It has a
> sufficiently efficient engine to achieve orbit, but without
> carrying much payload.[11]
>
> 11. Mitchell Burnside-Clapp (February 1997). "A LO2/Kerosene
> SSTO Rocket Design". Retrieved 2009-09-14.
>
I am aware of the wikipedia statement. That a statement is made in
> In article <MPG.2e84a8669...@news.newsguy.com>,
> jclark...@cox.net says...
> >
> > Do you have a source for the contention that the Titan II first stage
> > "showed possible" 50 years ago? I have seen no analysis that suggests
> > that it comes even close to having enough delta-v to achieve orbit.
>
> This is *very* easy to find.
>
> http://en.wikipedia.org/wiki/Single-stage-to-orbit
>
> From above:
>
> While single-stage rockets were once thought to be beyond reach,
> advances in materials technology and construction techniques have
> shown them to be possible. For example, calculations show that
> the Titan II first stage, launched on its own, would have a
> 25-to-1 ratio of fuel to vehicle hardware.[10] It has a
> sufficiently efficient engine to achieve orbit, but without
> carrying much payload.[11]
>
> 11. Mitchell Burnside-Clapp (February 1997). "A LO2/Kerosene
> SSTO Rocket Design". Retrieved 2009-09-14.
>
wikepedia does not mean that it is true, or that it is any kind of
consensus, or that it is even supported by the references given.
Reference 10 has a good deal of information about the Titan II but
nothing about its first stage being able to achieve orbit. Reference 11
has only one sentence concerning the Titan II, "The historical examples
of the extraordinary mass fractions of the Titan II first stage,
the Atlas, and the Saturn first stage are all persuasive."
I'm sorry but I am not seeing anything persuasive there.
Then it should be easy to find a discussion with some kind of
calculation or test results to support the contention.
According to someone who claims to have done flight mechanics on the
program at Lockheed and who thus may be taken to be authoritative about
the real-world performance of the Titan II, the actual payload that a
Titan II first stage could deliver to a 90 mile orbit is _negative_ 1400
pounds.
<https://groups.google.com/d/msg/sci.space.policy/F3G37kuJdjA/pjFf2fh_H9
gJ>
In other words it ain't gonna come close.
Greg Goss
Sep 18, 2014, 2:45:55 PM9/18/14
to
"J. Clarke" <jclark...@cox.net> wrote:
>jeff.f...@siemens.nospam.com says...
In addition to dropping the weight of the first stage, multi-stage
rockets allow the engine bell to be designed for the atmospheric
pressure it is working in. A first-level analysis, perhaps what was
done for the Wikipedia article might just take the sea-level thrust
and calculate that to orbit, while a more professional calculation
(like your Lockheed source) might calculate in efficiency losses as
the engine becomes mismatched to the low air pressure higher up.
--
We are geeks. Resistance is voltage over current.
>jeff.f...@siemens.nospam.com says...
rockets allow the engine bell to be designed for the atmospheric
pressure it is working in. A first-level analysis, perhaps what was
done for the Wikipedia article might just take the sea-level thrust
and calculate that to orbit, while a more professional calculation
(like your Lockheed source) might calculate in efficiency losses as
the engine becomes mismatched to the low air pressure higher up.
--
We are geeks. Resistance is voltage over current.
Jeff Findley
Sep 18, 2014, 3:22:08 PM9/18/14
to
In article <MPG.2e84e8dc6...@news.newsguy.com>,
jclark...@cox.net says...
of Mitchell Burnside-Clapp. Besides, it's easy enough to run the
numbers and see that the fuel to dry mass ratio is exactly what is
claimed. The difficult part is running a simulation to prove that you
can actually get it into orbit with the existing engines.
> Reference 10 has a good deal of information about the Titan II but
> nothing about its first stage being able to achieve orbit. Reference 11
> has only one sentence concerning the Titan II, "The historical examples
> of the extraordinary mass fractions of the Titan II first stage,
> the Atlas, and the Saturn first stage are all persuasive."
>
> I'm sorry but I am not seeing anything persuasive there.
Sorry you missed this debate on sci.space, but it was had maybe 25 years
ago. That an expendable SSTO is possible is not the question, because
it is absolutely physically possible. History has examples of engines
and stages which would have made it possible, in theory.
Such a vehicle was never actually flown because what would be the point?
The engineers, at the time, came mostly from the ICBM derived vehicle
era and believed in "performance uber alles" (as Henry Spencer liked to
say). So, naturally they'd add at least one extra stage to the vehicle
to boost the payload into orbit.
> > This analysis was discussed many times in the "old days" in the
> > sci.space newsgroups.
>
> Then it should be easy to find a discussion with some kind of
> calculation or test results to support the contention.
Knock yourself out. I'm not going to dig for 25 year old Usenet News
postings because I've already read them. I've been here (sci.space
groups) since I had a purdue.edu account starting in 1988.
Here are Clapp's numbers for "ideal delta-V" for several stages. At
least on paper, several stages have SSTO class performance.
> According to someone who claims to have done flight mechanics on the
> program at Lockheed and who thus may be taken to be authoritative about
> the real-world performance of the Titan II, the actual payload that a
> Titan II first stage could deliver to a 90 mile orbit is _negative_ 1400
From dim memory, the issue wasn't to do with the mass fraction of the
stage, the ISP of the engines, and etc. (see the link above) but the
fact that the Titan II first stage engines could not throttle, at least
"off the shelf".
jclark...@cox.net says...
>
> In article <MPG.2e84b0325...@reader80.eternal-september.org>,
> jeff.f...@siemens.nospam.com says...
> >
> > In article <MPG.2e84a8669...@news.newsguy.com>,
> > jclark...@cox.net says...
> > >
> > > Do you have a source for the contention that the Titan II first stage
> > > "showed possible" 50 years ago? I have seen no analysis that suggests
> > > that it comes even close to having enough delta-v to achieve orbit.
> >
> > This is *very* easy to find.
> >
> > http://en.wikipedia.org/wiki/Single-stage-to-orbit
> >
> > From above:
> >
> > While single-stage rockets were once thought to be beyond reach,
> > advances in materials technology and construction techniques have
> > shown them to be possible. For example, calculations show that
> > the Titan II first stage, launched on its own, would have a
> > 25-to-1 ratio of fuel to vehicle hardware.[10] It has a
> > sufficiently efficient engine to achieve orbit, but without
> > carrying much payload.[11]
> >
> > 11. Mitchell Burnside-Clapp (February 1997). "A LO2/Kerosene
> > SSTO Rocket Design". Retrieved 2009-09-14.
> >
>
> I am aware of the wikipedia statement. That a statement is made in
> wikepedia does not mean that it is true, or that it is any kind of
> consensus, or that it is even supported by the references given.
It is when the author is well respected in the field, which is the case
> In article <MPG.2e84b0325...@reader80.eternal-september.org>,
> jeff.f...@siemens.nospam.com says...
> >
> > In article <MPG.2e84a8669...@news.newsguy.com>,
> > jclark...@cox.net says...
> > >
> > > Do you have a source for the contention that the Titan II first stage
> > > "showed possible" 50 years ago? I have seen no analysis that suggests
> > > that it comes even close to having enough delta-v to achieve orbit.
> >
> > This is *very* easy to find.
> >
> > http://en.wikipedia.org/wiki/Single-stage-to-orbit
> >
> > From above:
> >
> > While single-stage rockets were once thought to be beyond reach,
> > advances in materials technology and construction techniques have
> > shown them to be possible. For example, calculations show that
> > the Titan II first stage, launched on its own, would have a
> > 25-to-1 ratio of fuel to vehicle hardware.[10] It has a
> > sufficiently efficient engine to achieve orbit, but without
> > carrying much payload.[11]
> >
> > 11. Mitchell Burnside-Clapp (February 1997). "A LO2/Kerosene
> > SSTO Rocket Design". Retrieved 2009-09-14.
> >
>
> I am aware of the wikipedia statement. That a statement is made in
> wikepedia does not mean that it is true, or that it is any kind of
> consensus, or that it is even supported by the references given.
of Mitchell Burnside-Clapp. Besides, it's easy enough to run the
numbers and see that the fuel to dry mass ratio is exactly what is
claimed. The difficult part is running a simulation to prove that you
can actually get it into orbit with the existing engines.
> Reference 10 has a good deal of information about the Titan II but
> nothing about its first stage being able to achieve orbit. Reference 11
> has only one sentence concerning the Titan II, "The historical examples
> of the extraordinary mass fractions of the Titan II first stage,
> the Atlas, and the Saturn first stage are all persuasive."
>
> I'm sorry but I am not seeing anything persuasive there.
ago. That an expendable SSTO is possible is not the question, because
it is absolutely physically possible. History has examples of engines
and stages which would have made it possible, in theory.
Such a vehicle was never actually flown because what would be the point?
The engineers, at the time, came mostly from the ICBM derived vehicle
era and believed in "performance uber alles" (as Henry Spencer liked to
say). So, naturally they'd add at least one extra stage to the vehicle
to boost the payload into orbit.
> > This analysis was discussed many times in the "old days" in the
> > sci.space newsgroups.
>
> Then it should be easy to find a discussion with some kind of
> calculation or test results to support the contention.
postings because I've already read them. I've been here (sci.space
groups) since I had a purdue.edu account starting in 1988.
Here are Clapp's numbers for "ideal delta-V" for several stages. At
least on paper, several stages have SSTO class performance.
> According to someone who claims to have done flight mechanics on the
> program at Lockheed and who thus may be taken to be authoritative about
> the real-world performance of the Titan II, the actual payload that a
> Titan II first stage could deliver to a 90 mile orbit is _negative_ 1400
stage, the ISP of the engines, and etc. (see the link above) but the
fact that the Titan II first stage engines could not throttle, at least
"off the shelf".
J. Clarke
Sep 18, 2014, 5:51:49 PM9/18/14
to
In article <MPG.2e84fada4...@reader80.eternal-september.org>,
Please present your evidence that the wikipedia article was written by
Mitchell Burnside-Clapp.
> > Reference 10 has a good deal of information about the Titan II but
> > nothing about its first stage being able to achieve orbit. Reference 11
> > has only one sentence concerning the Titan II, "The historical examples
> > of the extraordinary mass fractions of the Titan II first stage,
> > the Atlas, and the Saturn first stage are all persuasive."
> >
> > I'm sorry but I am not seeing anything persuasive there.
>
> Sorry you missed this debate on sci.space, but it was had maybe 25 years
> ago.
There is something called "google groups" . . .
> That an expendable SSTO is possible is not the question, because
> it is absolutely physically possible.
Then why are you even mentioning this?
> History has examples of engines
> and stages which would have made it possible, in theory.
However you have not shown this possibility with anything but a
wikipedia entry that references two articles neither of which contain
anything that supports the assertion.
> Such a vehicle was never actually flown because what would be the
point?
> The engineers, at the time, came mostly from the ICBM derived vehicle
> era and believed in "performance uber alles" (as Henry Spencer liked to
> say). So, naturally they'd add at least one extra stage to the vehicle
> to boost the payload into orbit.
Sez you.
> > > This analysis was discussed many times in the "old days" in the
> > > sci.space newsgroups.
> >
> > Then it should be easy to find a discussion with some kind of
> > calculation or test results to support the contention.
>
> Knock yourself out. I'm not going to dig for 25 year old Usenet News
> postings because I've already read them. I've been here (sci.space
> groups) since I had a purdue.edu account starting in 1988.
In other words you can't be assed to support your assertion.
> Here are Clapp's numbers for "ideal delta-V" for several stages. At
> least on paper, several stages have SSTO class performance.
Ideal delta-v doesn't make a rat's ass. We're talking about real-world
hardware here.
> > According to someone who claims to have done flight mechanics on the
> > program at Lockheed and who thus may be taken to be authoritative about
> > the real-world performance of the Titan II, the actual payload that a
> > Titan II first stage could deliver to a 90 mile orbit is _negative_ 1400
>
> From dim memory, the issue wasn't to do with the mass fraction of the
> stage, the ISP of the engines, and etc. (see the link above) but the
> fact that the Titan II first stage engines could not throttle, at least
> "off the shelf".
I'm aware of the issue. Until the issue is actually addressed in the
real world you do not know how the fix will affect the performance of
the engine. You might end up making things worse instead of better.
I still don't see any evidence that the Titan II first stage is capable
of orbit.
Mitchell Burnside-Clapp.
> > Reference 10 has a good deal of information about the Titan II but
> > nothing about its first stage being able to achieve orbit. Reference 11
> > has only one sentence concerning the Titan II, "The historical examples
> > of the extraordinary mass fractions of the Titan II first stage,
> > the Atlas, and the Saturn first stage are all persuasive."
> >
> > I'm sorry but I am not seeing anything persuasive there.
>
> Sorry you missed this debate on sci.space, but it was had maybe 25 years
> ago.
> That an expendable SSTO is possible is not the question, because
> it is absolutely physically possible.
> History has examples of engines
> and stages which would have made it possible, in theory.
wikipedia entry that references two articles neither of which contain
anything that supports the assertion.
> Such a vehicle was never actually flown because what would be the
point?
> The engineers, at the time, came mostly from the ICBM derived vehicle
> era and believed in "performance uber alles" (as Henry Spencer liked to
> say). So, naturally they'd add at least one extra stage to the vehicle
> to boost the payload into orbit.
> > > This analysis was discussed many times in the "old days" in the
> > > sci.space newsgroups.
> >
> > Then it should be easy to find a discussion with some kind of
> > calculation or test results to support the contention.
>
> Knock yourself out. I'm not going to dig for 25 year old Usenet News
> postings because I've already read them. I've been here (sci.space
> groups) since I had a purdue.edu account starting in 1988.
> Here are Clapp's numbers for "ideal delta-V" for several stages. At
> least on paper, several stages have SSTO class performance.
hardware here.
> > According to someone who claims to have done flight mechanics on the
> > program at Lockheed and who thus may be taken to be authoritative about
> > the real-world performance of the Titan II, the actual payload that a
> > Titan II first stage could deliver to a 90 mile orbit is _negative_ 1400
>
> From dim memory, the issue wasn't to do with the mass fraction of the
> stage, the ISP of the engines, and etc. (see the link above) but the
> fact that the Titan II first stage engines could not throttle, at least
> "off the shelf".
real world you do not know how the fix will affect the performance of
the engine. You might end up making things worse instead of better.
I still don't see any evidence that the Titan II first stage is capable
of orbit.
Jeff Findley
Sep 19, 2014, 6:59:43 AM9/19/14
to
> Mitchell Burnside-Clapp
>
Somehow the link showing "ideal delta-V" for various stages didn't get
pasted in my prior post, so here it is:
https://groups.google.com/forum/#!topic/sci.space.policy/PZgWB9WWhNw
>
Somehow the link showing "ideal delta-V" for various stages didn't get
pasted in my prior post, so here it is:
https://groups.google.com/forum/#!topic/sci.space.policy/PZgWB9WWhNw
Jeff Findley
Sep 19, 2014, 7:15:14 AM9/19/14
to
In article <MPG.2e851dd11...@news.newsguy.com>,
jclark...@cox.net says...
that the only thing which would satisfy you is actually flying an
expendable SSTO to LEO. No point in debating this. Been there, done
that a quarter of a century ago. We heard the same back then about
rapid turn-around of rocket engines isn't possible along with vertical
landing using rockets isn't possible, and etc. until DC-X proved that it
could be done. Yet the nay-sayers kept moving the goal posts with each
"first" that was proven.
The bottom of the line is that SSTO hasn't been definitively proven by
an actual vehicle flying to LEO because no one has tried. Quite
frankly, no one is going to try anytime soon due to NASA's X-33 debacle.
NASA's official line about that massive failure is that the technology
wasn't ready, which is utter b.s. considering they deliberately chose
the most technically challenging of the three proposals, setting
themselves up for failure.
Today, there are no less than three companies/government agencies
working on reusable rocket stages which would be part of a partially
reusable TSTO. Re-flight of a reusable, liquid rocket powered, first
stage will be the next big "first" in aerospace, IMHO. This is yet
another area where I hear the "old guard" saying it can't be done, even
in the face of mounting empirical evidence that it will be done, and
quite possibly will be done within the next two years, which is
relatively "soon" in terms of history.
And once a reusable first stage is proven, the nay-sayers will no doubt
point out the obvious that reusing an upper stage is much harder and
can't be done for this reason or that.
Fine. I've waited more than 25 years for a sanely designed reusable
launch vehicle. I can wait a bit more for history to catch up with the
vision of a world where we have cheap access to space despite it being
"hard".
jclark...@cox.net says...
> I still don't see any evidence that the Titan II first stage is capable
> of orbit.
OMFG, it's the late '80s, early '90s, calling... I can see in your mind
> of orbit.
that the only thing which would satisfy you is actually flying an
expendable SSTO to LEO. No point in debating this. Been there, done
that a quarter of a century ago. We heard the same back then about
rapid turn-around of rocket engines isn't possible along with vertical
landing using rockets isn't possible, and etc. until DC-X proved that it
could be done. Yet the nay-sayers kept moving the goal posts with each
"first" that was proven.
The bottom of the line is that SSTO hasn't been definitively proven by
an actual vehicle flying to LEO because no one has tried. Quite
frankly, no one is going to try anytime soon due to NASA's X-33 debacle.
NASA's official line about that massive failure is that the technology
wasn't ready, which is utter b.s. considering they deliberately chose
the most technically challenging of the three proposals, setting
themselves up for failure.
Today, there are no less than three companies/government agencies
working on reusable rocket stages which would be part of a partially
reusable TSTO. Re-flight of a reusable, liquid rocket powered, first
stage will be the next big "first" in aerospace, IMHO. This is yet
another area where I hear the "old guard" saying it can't be done, even
in the face of mounting empirical evidence that it will be done, and
quite possibly will be done within the next two years, which is
relatively "soon" in terms of history.
And once a reusable first stage is proven, the nay-sayers will no doubt
point out the obvious that reusing an upper stage is much harder and
can't be done for this reason or that.
Fine. I've waited more than 25 years for a sanely designed reusable
launch vehicle. I can wait a bit more for history to catch up with the
vision of a world where we have cheap access to space despite it being
"hard".
J. Clarke
Sep 19, 2014, 7:23:10 AM9/19/14
to
In article <MPG.2e850ffb9...@reader80.eternal-september.org>,
jeff.f...@siemens.nospam.com says...
itself deliver 1400 pounds to low earth orbit as it sits" when the
person who did the flight mechanics analysis for Lockheed says that "as
it sits" it can deliver ****>>>NEGATIVE<<<**** 1400 pounds to LEO.
Seems that your esteemed Mr. Clapp is merely parroting something he read
on the internet and getting it _wrong_.
jeff.f...@siemens.nospam.com says...
>
> In article <MPG.2e84fada4...@reader80.eternal-september.org>,
> jeff.f...@siemens.nospam.com says...
> > Mitchell Burnside-Clapp
> >
>
> Somehow the link showing "ideal delta-V" for various stages didn't get
> pasted in my prior post, so here it is:
>
> https://groups.google.com/forum/#!topic/sci.space.policy/PZgWB9WWhNw
I find it interesting that he states "the Titan II first stage can
> In article <MPG.2e84fada4...@reader80.eternal-september.org>,
> jeff.f...@siemens.nospam.com says...
> > Mitchell Burnside-Clapp
> >
>
> Somehow the link showing "ideal delta-V" for various stages didn't get
> pasted in my prior post, so here it is:
>
> https://groups.google.com/forum/#!topic/sci.space.policy/PZgWB9WWhNw
itself deliver 1400 pounds to low earth orbit as it sits" when the
person who did the flight mechanics analysis for Lockheed says that "as
it sits" it can deliver ****>>>NEGATIVE<<<**** 1400 pounds to LEO.
Seems that your esteemed Mr. Clapp is merely parroting something he read
on the internet and getting it _wrong_.
J. Clarke
Sep 19, 2014, 7:57:17 AM9/19/14
to
In article <MPG.2e85da341...@reader80.eternal-september.org>,
jeff.f...@siemens.nospam.com says...
And you have still presented no evidence that the Titan II first stage
is capable of achieving orbit. I don't _care_ if Delta Clipper or DC-X
or Orion or the Starship Enterprise is capable of achieving orbit. My
question is about the Titan II first stage. If you cannot present
evidence in the form of either an actual test or a full 3DOF trajectory
model taking into consideration the actual properties of the actual
vehicle then you are not providing anything other than arm-waving.
jeff.f...@siemens.nospam.com says...
is capable of achieving orbit. I don't _care_ if Delta Clipper or DC-X
or Orion or the Starship Enterprise is capable of achieving orbit. My
question is about the Titan II first stage. If you cannot present
evidence in the form of either an actual test or a full 3DOF trajectory
model taking into consideration the actual properties of the actual
vehicle then you are not providing anything other than arm-waving.
Vaughn
Sep 19, 2014, 9:05:00 AM9/19/14
to
On 9/18/2014 9:12 AM, Robert Clark wrote:
> SpaceX Alum Goes After Falcon 1 Market With Firefly.
> Aerospike revival, advances in composite structures are shaping design
> of low-cost smallsat launch vehicle.
> Aug 25, 2014 Frank Morring, Jr. | Aviation Week & Space Technology
> http://aviationweek.com/space/spacex-alum-goes-after-falcon-1-market-firefly
>
>
> Excellent news! As described in the article it will use both lightweight
> composite tanks and altitude compensating aerospike nozzles.
> This small launcher is to be two stage.
Interesting! But why does the world need another Falcon 1 class vehicle?
> SpaceX Alum Goes After Falcon 1 Market With Firefly.
> Aerospike revival, advances in composite structures are shaping design
> of low-cost smallsat launch vehicle.
> Aug 25, 2014 Frank Morring, Jr. | Aviation Week & Space Technology
> http://aviationweek.com/space/spacex-alum-goes-after-falcon-1-market-firefly
>
>
> Excellent news! As described in the article it will use both lightweight
> composite tanks and altitude compensating aerospike nozzles.
> This small launcher is to be two stage.
Is my memory correct that SpaceX gave up on the Falcon 1 because of a
lack of launch customers for that size payload? The current trend is to
piggyback smaller payloads on larger rockets to utilize their excess
capacity which would otherwise go to waste.
ra...@darkstar.arc.nasa.gov
Sep 19, 2014, 4:52:30 PM9/19/14
to
In article <MPG.2e85da341...@reader80.eternal-september.org>,
up were the two booster engines.
http://en.wikipedia.org/wiki/SCORE_(satellite)
(Yes, it's Wikipedia. But the caption under the picture is still correct.)
--
Kathy Rages
Jeff Findley <jeff.f...@siemens.nospam.com> wrote:
>In article <MPG.2e851dd11...@news.newsguy.com>,
>jclark...@cox.net says...
>> I still don't see any evidence that the Titan II first stage is capable
>> of orbit.
>
>In article <MPG.2e851dd11...@news.newsguy.com>,
>jclark...@cox.net says...
>> I still don't see any evidence that the Titan II first stage is capable
>> of orbit.
>
>The bottom of the line is that SSTO hasn't been definitively proven by
>an actual vehicle flying to LEO because no one has tried.
Project Score came close. The only thing the Atlas dropped on the way
>an actual vehicle flying to LEO because no one has tried.
up were the two booster engines.
http://en.wikipedia.org/wiki/SCORE_(satellite)
(Yes, it's Wikipedia. But the caption under the picture is still correct.)
--
Kathy Rages
J. Clarke
Sep 19, 2014, 5:49:23 PM9/19/14
to
In article <E9ydnXeBYpcTCoHJ...@posted.internetamerica>,
ra...@darkstar.arc.nasa.gov says...
There's no question that Atlas can achieve orbit, but "all it dropped on
the way up were the two booster engines" ignores the fact that those two
engines are nearly half the dry mass of the entire vehicle.
ra...@darkstar.arc.nasa.gov says...
the way up were the two booster engines" ignores the fact that those two
engines are nearly half the dry mass of the entire vehicle.
Spac...@hotmail.com
Sep 19, 2014, 10:01:37 PM9/19/14
to
best apaceship earth theory is by a Purdue alum
(he was the top of his ME class, and
worked for an oilco. for instance,
there is a machine tool that makes a good analogy
for plate tectonics (as opposed to "currents
in the mantle, although I have since come
to believe in that, for a particular aspect
of geology, that is not in the geophysics theory mainstream
(he was the top of his ME class, and
worked for an oilco. for instance,
there is a machine tool that makes a good analogy
for plate tectonics (as opposed to "currents
in the mantle, although I have since come
to believe in that, for a particular aspect
of geology, that is not in the geophysics theory mainstream
Robert Clark
Sep 20, 2014, 1:52:04 AM9/20/14
to
Perhaps the discrepancy is coming from the different versions of the Titan
II. This page gives its first stage a quite high mass ratio of 29 to 1 :
Titan.
http://www.braeunig.us/space/specs/titan.htm
But the version described on this Astronautix page only has a mass ratio of
17.5 to 1:
Titan 2.
http://www.astronautix.com/lvs/titan2.htm
Using the 29 to 1 mass ratio on the www.braeunig.us page and a 296 s
vacuum Isp for the LR-87 first stage engine we can calculate a stage delta-v
of 296*9.81ln(29) = 9,780 m/s, which is sufficient for orbit.
You can also try Dr. John Schilling's launch performance calculator:
http://www.silverbirdastronautics.com/LVperform.html
Select 1 for the number of stages in the options for the calculator, and
4,220 kg and 118,300 kg for the dry mass and propellant mass respectively.
Enter the vacuum Isp of the engine, 296 s. On the www.braeunig.us page, it
only gives the sea level thrust of 1,913 kN, but the calculator normally
works on the vacuum thrust. So enter in an estimate of 2,100 kN for the
vacuum thrust.
Select Yes for "Default Propellant Residuals?" and No for "Restartable Upper
Stage?" options. Choosing Yes for the "Restartable Upper Stage?" options
would wind up reducing your payload in the calculator.
Choose Cape Canaveral as the launch site, with a launch inclination of 28.5
degrees to match the launch site latitude. Then the calculator gives a
result of:
Mission Performance:
Launch Vehicle: User-Defined Launch Vehicle
Launch Site: Cape Canaveral / KSC
Destination Orbit: 185 x 185 km, 28 deg
Estimated Payload: 1374 kg
95% Confidence Interval: 286 - 2711 kg
I had previously found that the "estimated payload" by the calculator was in
the range of 10% accurate for known rockets I had tried it on.
Bob Clark
------------------------------------------------------------------
------------------------------------------------------------------
==============================================================
"J. Clarke" wrote in message
news:MPG.2e84e8dc6...@news.newsguy.com...
==============================================================
II. This page gives its first stage a quite high mass ratio of 29 to 1 :
Titan.
http://www.braeunig.us/space/specs/titan.htm
But the version described on this Astronautix page only has a mass ratio of
17.5 to 1:
Titan 2.
http://www.astronautix.com/lvs/titan2.htm
Using the 29 to 1 mass ratio on the www.braeunig.us page and a 296 s
vacuum Isp for the LR-87 first stage engine we can calculate a stage delta-v
of 296*9.81ln(29) = 9,780 m/s, which is sufficient for orbit.
You can also try Dr. John Schilling's launch performance calculator:
http://www.silverbirdastronautics.com/LVperform.html
Select 1 for the number of stages in the options for the calculator, and
4,220 kg and 118,300 kg for the dry mass and propellant mass respectively.
Enter the vacuum Isp of the engine, 296 s. On the www.braeunig.us page, it
only gives the sea level thrust of 1,913 kN, but the calculator normally
works on the vacuum thrust. So enter in an estimate of 2,100 kN for the
vacuum thrust.
Select Yes for "Default Propellant Residuals?" and No for "Restartable Upper
Stage?" options. Choosing Yes for the "Restartable Upper Stage?" options
would wind up reducing your payload in the calculator.
Choose Cape Canaveral as the launch site, with a launch inclination of 28.5
degrees to match the launch site latitude. Then the calculator gives a
result of:
Mission Performance:
Launch Vehicle: User-Defined Launch Vehicle
Launch Site: Cape Canaveral / KSC
Destination Orbit: 185 x 185 km, 28 deg
Estimated Payload: 1374 kg
95% Confidence Interval: 286 - 2711 kg
I had previously found that the "estimated payload" by the calculator was in
the range of 10% accurate for known rockets I had tried it on.
Bob Clark
------------------------------------------------------------------
Single-stage-to-orbit was already shown possible 50 years ago
with the Titan II first stage.
In fact, contrary to popular belief SSTO's are actually easy.
Just use the most efficient engines and stages at the same time,
and the result will automatically be SSTO.
Blog: Http://Exoscientist.blogspot.com
with the Titan II first stage.
In fact, contrary to popular belief SSTO's are actually easy.
Just use the most efficient engines and stages at the same time,
and the result will automatically be SSTO.
------------------------------------------------------------------
==============================================================
"J. Clarke" wrote in message
news:MPG.2e84e8dc6...@news.newsguy.com...
Robert Clark
Sep 21, 2014, 8:04:39 AM9/21/14
to
Richard Branson has claimed to be a proponent of governmental initiatives
aimed at reducing CO2 emissions. If so he should try to convert his entire
Virgin Atlantic fleet to electric or hydrogen powered.
How efficient could such jet aircraft be compared to kerosene fueled jets?
Bob Clark
------------------------------------------------------------------
Single-stage-to-orbit was already shown possible 50 years ago
with the Titan II first stage.
In fact, contrary to popular belief SSTO's are actually easy.
Just use the most efficient engines and stages at the same time,
and the result will automatically be SSTO.
Blog: Http://Exoscientist.blogspot.com
------------------------------------------------------------------
"William Mook" wrote in message
news:b8887a67-de0f-4977...@googlegroups.com...
https://www.youtube.com/watch?v=XzeCQblYHic
https://www.youtube.com/watch?v=P8Pb_psj1A8
https://www.youtube.com/watch?v=WGPBsLLAHU8
https://www.youtube.com/watch?v=oD8TfjDMYT0
The Sikorsky Firefly is interesting. At 1100 lbs (500 kg) the Li-ion
batteries contain 64 kWh of electricity. At 140 kW this is enough for 24.7
minutes of flight. Since its impossible to discharge batteries 100% - this
accounts for the 16 minutes of flight.
A hydrogen fuel cell that broke down water into hydrogen and oxygen, and
stored it on board the aircraft, to make water (also stored) again to
produce power - and allocating 250 kg for the propellant storage and 250 kg
for the associated hardware, we have 17.1x the energy stored on board as the
electric battery system. That is 422.4 minutes. (about 7 hours) At 159
km/hr cruise speed this is a range of 1,119.4 km between charges!
140 kW Proton Exchange Membrane fuel cell, at $63 per kW costs $8820. This
is 1/3 the cost of a Lycoming 360 engine it replaces! A 140 KW brushless DC
motor combined with a solid state controller costs $2200 - so overall, the
system is half the cost of the Lycoming.
The system requires 235 kW of laser energy to charge at the same rate it
discharges. Quadrupling the size of the Proton Exchange Membrane system,
allows it to charge in one quarter the time it flies. 500 litres of water
are converted to a full charge of hydrogen and oxygen in 1 hour and 45
minutes of beam time.
The ability to receive power in flight permits reception of power during the
two hours surrounding either sunrise or sunset. The aircraft is then
capable of flying 7 hours after spending 1.75 hours charging.
Automated electric helicopters that are charged with a power beaming set up
provides flight on demand.
https://www.youtube.com/watch?v=KzWwGvAalRk
https://www.youtube.com/watch?v=undX_rxY-dQ
https://www.youtube.com/watch?v=NevgqMqWf5Y
A total of 111 of the C300 helicopters are supported per satellite, which
provide 14 hours of flight service out of every 24 hours. At $0.11 per kWh
the power costs are $215 per day per vehicle. That's $15.35 per hour - or
9.65 cents per km.
At $250,000 and an 8% discount rate, with a 20 year life span, and 4%
maintenance cost, we have $35,463 per year. Dividing by 5113.5 hours per
year obtains $6.94 per hour. That's another 4.35 cents per km. A total of
$0.14 per km. Dividing by three people, that's less than $0.05 per
passenger km.
The MD-500 helicopter has a 207 kW engine, a 48% increase in power, and
lifts 100% increase in weight! 76 of these ships can be supported per
satellite. Greater lift capacity combined with higher engine cost, means
this five passenger system, is charged in two hours (even in flight) and
operates 10 hours nonstop. This permits 24 hours of flight every 24 hours.
So, this system is always on the go!
Crew: 1-2
Capacity: 5 total
Length: 30 ft 10 in (9.4 m)
Rotor diameter: 26 ft 4 in (8.03 m)
Height: 8 ft 2 in (2.48 m)
Empty weight: 1,088 lb (493 kg)
Max. takeoff weight: 2,250 lb (1,157 kg)
Powerplant: 1 � Allison 250-C20 Turboshaft, 278 hp (207 kW)
Performance
Maximum speed: 152 knots (175 mph, 282 km/h)
Cruise speed: 125 kn (144 mph, 232 km/h)
Range: 375 mi (605 km)
Service ceiling: 16,000 ft (4,875 m)
Rate of climb: 1,700 ft/min (8.6 m/s)
A speed of 282 km/h x 24 hours = 6768 km/day. That's 2,472,012 km per
year - and with five passengers - that's 12,360,060 passenger-km per year.
With a 20% service cycle time - this translates to 9,888,048 passenger-km
per vehicle per year - with 95 ships per satellite.
The cost of power at $0.11 per kWh is $22.77 per hour. This is $4.56 per
passenger. At 282 km/hr this is 1.62 cents per passenger-km!!
The MD-520N costs $1.3 million used. The MD-6M Little Bird is $3.6 million.
The Allison 250 C engine is $200,000+. That's enough to buy 3 MW of Proton
Exchange Membranes - and at 207 kW use rate, you can have as much as 6:1
advantage during high speed charge. With a $3.6 million price tag and 80%
utilization rate, and a 4% maintenance cost we have $144,000 per year in
maintenance, spread over 7,020.8 hours. That's $20.51 per hour. Financing
$3.6 million over 20 years at 8% costs $366,667 per hour. Dividing across
7,020.8 hours that's $52.23 per hour. A total cost of $95.51 per hour.
Dividing by five passengers that's $19.10 per passenger hour. Dividing by
282 km/hr that's 6.77 per passenger km.
Charging $1.50 for the first 5 km and $0.15 per km thereafter, (with
distance calculated by straight line from point of pick up to point of
departure, with total paid upon entry, so no extra charges are incurred for
pick up and drop off of other passengers) a system of electric helicopter
drones would already be competitive with any other system of transport.
With five hours at either end - to get to and from airports - combined with
cancellations and other factors - we can beat any travel of any sort in
terms of price, quality and speed of service with any distance of less than
1,410 km for $211.50. This is the distance from Rome to Paris. It costs
315.92 euro by car ($405.36) in fuel and takes 14 hours. A point to point
electric drone heli at these prices, makes a lot of sense!
http://www.flightstats.com/go/Media/stats.do
Paris to New York would take 20.7 hours point to point - and cost $875 each
way. Not bad.
Boeing has a Quad Tilt Rotor
http://i799.photobucket.com/albums/yy279/The_terran_empire/hvtol.jpg
Which ups the ante - by converting to an airplane type configuration -
giving a much higher speed and better efficiency than a pure helicopter.
Quad Rotors have a long history
http://illumin.usc.edu/assets/media/152/4589757501_315a2bc282_o.jpg
And a brilliant future
http://realitypod.com//HLIC/8e62034e75dabb79bdb0a8bd278eb44d.jpg
We should be able to cut the costs to about 1/10th the costs calculated
here, for the hardware, and cut the costs of the energy by about half in the
short term, while increasing speeds to 1,000 kph - typical of airliners
today. This radically reduces cost per passenger km (or tonne km for
cargo).
aimed at reducing CO2 emissions. If so he should try to convert his entire
Virgin Atlantic fleet to electric or hydrogen powered.
How efficient could such jet aircraft be compared to kerosene fueled jets?
Bob Clark
------------------------------------------------------------------
Single-stage-to-orbit was already shown possible 50 years ago
with the Titan II first stage.
In fact, contrary to popular belief SSTO's are actually easy.
Just use the most efficient engines and stages at the same time,
and the result will automatically be SSTO.
Blog: Http://Exoscientist.blogspot.com
------------------------------------------------------------------
news:b8887a67-de0f-4977...@googlegroups.com...
https://www.youtube.com/watch?v=XzeCQblYHic
https://www.youtube.com/watch?v=P8Pb_psj1A8
https://www.youtube.com/watch?v=WGPBsLLAHU8
https://www.youtube.com/watch?v=oD8TfjDMYT0
The Sikorsky Firefly is interesting. At 1100 lbs (500 kg) the Li-ion
batteries contain 64 kWh of electricity. At 140 kW this is enough for 24.7
minutes of flight. Since its impossible to discharge batteries 100% - this
accounts for the 16 minutes of flight.
A hydrogen fuel cell that broke down water into hydrogen and oxygen, and
stored it on board the aircraft, to make water (also stored) again to
produce power - and allocating 250 kg for the propellant storage and 250 kg
for the associated hardware, we have 17.1x the energy stored on board as the
electric battery system. That is 422.4 minutes. (about 7 hours) At 159
km/hr cruise speed this is a range of 1,119.4 km between charges!
140 kW Proton Exchange Membrane fuel cell, at $63 per kW costs $8820. This
is 1/3 the cost of a Lycoming 360 engine it replaces! A 140 KW brushless DC
motor combined with a solid state controller costs $2200 - so overall, the
system is half the cost of the Lycoming.
The system requires 235 kW of laser energy to charge at the same rate it
discharges. Quadrupling the size of the Proton Exchange Membrane system,
allows it to charge in one quarter the time it flies. 500 litres of water
are converted to a full charge of hydrogen and oxygen in 1 hour and 45
minutes of beam time.
The ability to receive power in flight permits reception of power during the
two hours surrounding either sunrise or sunset. The aircraft is then
capable of flying 7 hours after spending 1.75 hours charging.
Automated electric helicopters that are charged with a power beaming set up
provides flight on demand.
https://www.youtube.com/watch?v=KzWwGvAalRk
https://www.youtube.com/watch?v=undX_rxY-dQ
https://www.youtube.com/watch?v=NevgqMqWf5Y
A total of 111 of the C300 helicopters are supported per satellite, which
provide 14 hours of flight service out of every 24 hours. At $0.11 per kWh
the power costs are $215 per day per vehicle. That's $15.35 per hour - or
9.65 cents per km.
At $250,000 and an 8% discount rate, with a 20 year life span, and 4%
maintenance cost, we have $35,463 per year. Dividing by 5113.5 hours per
year obtains $6.94 per hour. That's another 4.35 cents per km. A total of
$0.14 per km. Dividing by three people, that's less than $0.05 per
passenger km.
The MD-500 helicopter has a 207 kW engine, a 48% increase in power, and
lifts 100% increase in weight! 76 of these ships can be supported per
satellite. Greater lift capacity combined with higher engine cost, means
this five passenger system, is charged in two hours (even in flight) and
operates 10 hours nonstop. This permits 24 hours of flight every 24 hours.
So, this system is always on the go!
Crew: 1-2
Capacity: 5 total
Length: 30 ft 10 in (9.4 m)
Rotor diameter: 26 ft 4 in (8.03 m)
Height: 8 ft 2 in (2.48 m)
Empty weight: 1,088 lb (493 kg)
Max. takeoff weight: 2,250 lb (1,157 kg)
Powerplant: 1 � Allison 250-C20 Turboshaft, 278 hp (207 kW)
Performance
Maximum speed: 152 knots (175 mph, 282 km/h)
Cruise speed: 125 kn (144 mph, 232 km/h)
Range: 375 mi (605 km)
Service ceiling: 16,000 ft (4,875 m)
Rate of climb: 1,700 ft/min (8.6 m/s)
A speed of 282 km/h x 24 hours = 6768 km/day. That's 2,472,012 km per
year - and with five passengers - that's 12,360,060 passenger-km per year.
With a 20% service cycle time - this translates to 9,888,048 passenger-km
per vehicle per year - with 95 ships per satellite.
The cost of power at $0.11 per kWh is $22.77 per hour. This is $4.56 per
passenger. At 282 km/hr this is 1.62 cents per passenger-km!!
The MD-520N costs $1.3 million used. The MD-6M Little Bird is $3.6 million.
The Allison 250 C engine is $200,000+. That's enough to buy 3 MW of Proton
Exchange Membranes - and at 207 kW use rate, you can have as much as 6:1
advantage during high speed charge. With a $3.6 million price tag and 80%
utilization rate, and a 4% maintenance cost we have $144,000 per year in
maintenance, spread over 7,020.8 hours. That's $20.51 per hour. Financing
$3.6 million over 20 years at 8% costs $366,667 per hour. Dividing across
7,020.8 hours that's $52.23 per hour. A total cost of $95.51 per hour.
Dividing by five passengers that's $19.10 per passenger hour. Dividing by
282 km/hr that's 6.77 per passenger km.
Charging $1.50 for the first 5 km and $0.15 per km thereafter, (with
distance calculated by straight line from point of pick up to point of
departure, with total paid upon entry, so no extra charges are incurred for
pick up and drop off of other passengers) a system of electric helicopter
drones would already be competitive with any other system of transport.
With five hours at either end - to get to and from airports - combined with
cancellations and other factors - we can beat any travel of any sort in
terms of price, quality and speed of service with any distance of less than
1,410 km for $211.50. This is the distance from Rome to Paris. It costs
315.92 euro by car ($405.36) in fuel and takes 14 hours. A point to point
electric drone heli at these prices, makes a lot of sense!
http://www.flightstats.com/go/Media/stats.do
Paris to New York would take 20.7 hours point to point - and cost $875 each
way. Not bad.
Boeing has a Quad Tilt Rotor
http://i799.photobucket.com/albums/yy279/The_terran_empire/hvtol.jpg
Which ups the ante - by converting to an airplane type configuration -
giving a much higher speed and better efficiency than a pure helicopter.
Quad Rotors have a long history
http://illumin.usc.edu/assets/media/152/4589757501_315a2bc282_o.jpg
And a brilliant future
http://realitypod.com//HLIC/8e62034e75dabb79bdb0a8bd278eb44d.jpg
We should be able to cut the costs to about 1/10th the costs calculated
here, for the hardware, and cut the costs of the energy by about half in the
short term, while increasing speeds to 1,000 kph - typical of airliners
today. This radically reduces cost per passenger km (or tonne km for
cargo).
Orval Fairbairn
Sep 21, 2014, 2:37:40 PM9/21/14
to
In article <lvmesu$s4j$1...@dont-email.me>,
1. Electric jets are a non-starter, since electricity storage defeats it
because of weight.
2. Even liquid hydrogen has such poor energy density that storage tanks
would occupy the entire passenger cabin -- and then some.
Sorry, but hydrocarbons present, by far, the best bets for fuel.
"Robert Clark" <rgrego...@gmSPAMBLOCKail.com> wrote:
> Richard Branson has claimed to be a proponent of governmental initiatives
> aimed at reducing CO2 emissions. If so he should try to convert his entire
> Virgin Atlantic fleet to electric or hydrogen powered.
>
> How efficient could such jet aircraft be compared to kerosene fueled jets?
>
> Bob Clark
Not very!
> Richard Branson has claimed to be a proponent of governmental initiatives
> aimed at reducing CO2 emissions. If so he should try to convert his entire
> Virgin Atlantic fleet to electric or hydrogen powered.
>
> How efficient could such jet aircraft be compared to kerosene fueled jets?
>
> Bob Clark
1. Electric jets are a non-starter, since electricity storage defeats it
because of weight.
2. Even liquid hydrogen has such poor energy density that storage tanks
would occupy the entire passenger cabin -- and then some.
Sorry, but hydrocarbons present, by far, the best bets for fuel.
Robert Clark
Sep 27, 2014, 11:22:40 AM9/27/14
to
=================================================================
In article <d405c11e-51ef-4df6...@googlegroups.com>,
William Mook <mokme...@gmail.com> wrote:
>
> The weight of the tank is such that a tank that carries 11.9 kg masses 6.1
> kg. So, to carry 171 kg of hydrogen requires 2,442.9 litres of storage
> volume in a take that weighs 87.7 kg!
>
> http://www.the-linde-group.com/internet.global.thelindegroup.global/en/images/
> HydrogenBrochure_EN14_10196.pdf
>
> Now, on the other hand, a Lycoming 0-360 masses 117 kg and has another 86
> kg
> of other machinery attached to it. A total of 203 kg. A 270 kW DC
> electric
> motor masses only 65 kg, has none of the machinery attached to it, and
> thus,
> is 138 kg lighter! More than making up the 87.7 kg weight of the fuel
> tanks.
> ...
===================================================================
===================================================================
>
>
> Overlooked is the size of the tankage required to pack the LH2. As I
> posted earlier,
Nonsense. The tankage was included in the calculation.
> just the tanks would occupy the useful part of the
> aircraft.
Real engineers that have designed built and tested real hydrogen fuelled
aircraft and hydrogen fuelled vehicles, like the BMW Hydrogen 7 and Boeing's
Phantom Eye and Boeing's Hydrogen Fuel Cell aircraft, find the overall mass
is less since electric motors and fuel cells mass vastly less than thermal
engines when combined with the mass of their air handling and exhaust
systems.
===============================================================================
That ratio of the weight of the fuel carried compared to the weight of the
tank for hydrogen at 2 to 1 is quite poor. For kerosene it's in the range of
100 to 1.
The Lycoming O-360 piston-engine is 1950's technology, not very efficient
in regards to the weight of the engine compared to the horsepower. Modern
jet engines are an order of magnitude better in regards to the power they
put out compared to their weight. It's the weight of the systems for a
hydrogen-fueled aircraft that is the problem.
In article <d405c11e-51ef-4df6...@googlegroups.com>,
William Mook <mokme...@gmail.com> wrote:
>
> The weight of the tank is such that a tank that carries 11.9 kg masses 6.1
> kg. So, to carry 171 kg of hydrogen requires 2,442.9 litres of storage
> volume in a take that weighs 87.7 kg!
>
> http://www.the-linde-group.com/internet.global.thelindegroup.global/en/images/
> HydrogenBrochure_EN14_10196.pdf
>
> Now, on the other hand, a Lycoming 0-360 masses 117 kg and has another 86
> kg
> of other machinery attached to it. A total of 203 kg. A 270 kW DC
> electric
> motor masses only 65 kg, has none of the machinery attached to it, and
> thus,
> is 138 kg lighter! More than making up the 87.7 kg weight of the fuel
> tanks.
> ...
===================================================================
===================================================================
"William Mook" wrote in message
news:48287366-d47c-464d...@googlegroups.com...
>
>
> Overlooked is the size of the tankage required to pack the LH2. As I
> posted earlier,
Nonsense. The tankage was included in the calculation.
> just the tanks would occupy the useful part of the
> aircraft.
Real engineers that have designed built and tested real hydrogen fuelled
aircraft and hydrogen fuelled vehicles, like the BMW Hydrogen 7 and Boeing's
Phantom Eye and Boeing's Hydrogen Fuel Cell aircraft, find the overall mass
is less since electric motors and fuel cells mass vastly less than thermal
engines when combined with the mass of their air handling and exhaust
systems.
===============================================================================
That ratio of the weight of the fuel carried compared to the weight of the
tank for hydrogen at 2 to 1 is quite poor. For kerosene it's in the range of
100 to 1.
The Lycoming O-360 piston-engine is 1950's technology, not very efficient
in regards to the weight of the engine compared to the horsepower. Modern
jet engines are an order of magnitude better in regards to the power they
put out compared to their weight. It's the weight of the systems for a
hydrogen-fueled aircraft that is the problem.
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