Towards a combined-cycle SSTO.

[Robert Clark]

Looking at the numbers I'm now convinced you can make a single stage
to orbit vehicle with a combined ramjet/rocket engine, and without
having to use scramjets.

The idea is to combine the turbo-ramjet/rocket into a single engine.
This is what Skylon wants to do with their Sabre engine. But the Sabre
will use hypersonic airbreathing propulsion up to Mach 5.5 before the
rockets take over. This will require complicated air-cooling methods
using heat exchangers with flowing liquid hydrogen for the Skylon.

However, just being able to get to say the Mach 3.2 reached by the
SR-71 would take a significant amount off the delta-V required for
orbit. Of course if the ramjet could get to Mach 5 that would be even
better but key this would be doable with the existing engines of the
SR-71. Note too the engines of the XB-70 Valkyrie bomber could operate
at Mach 3 and as far as I know they didn't have ramjet operation mode.
So it might not even be necessary for the engines to have a ramjet
mode, turbojet might be sufficient.

The problem with using jets for the early part of the flight of an
SSTO has been they are so heavy for the thrust they produce, generally
in the T/W range of around 5 to 10. While rocket engines might have a
T/W ratio in the range of 50 to 100. But a key point is the jet engine
will be operating during the aerodynamic lift portion of the flight
where the L/D ratio of perhaps 7. The XB-70 for instance had a L/D of
about 7 during cruise at Mach 3. So if we take the T/W of the jet
engine to be say 7 and the L/D to be 7, then the thrust to lift-off
weight ratio might be about 50 to 1 comparable to that of rockets.

BTW, it is surprising there has been so little research on this type
of combination with the jet and rocket combined into one. You hear
alot about turbine-based-combined-cycle (TBCC) where it combines
turbo- and scram-jets and rocket-based-combined-cycle (RBCC) , where
the exhaust from a rocket is used to provide the compression for a
ramjet. But not this type of combined turbojet/rocket engine. It
doesn't seem to have an accepted name for example. It would not seem
to be too complicated. You just use the same combustion chamber for
rocket as for the jet. Probably also you would want to close off the
inlets when you switch to rocket mode.

For the calculation the delta-V and propellant load would be feasible,
note that for a dense propellant SSTO might require as much as 300 m/s
lower delta-V than a hydrogen fueled SSTO, in the range of about 8,900
m/s, so I'll use kerosene as the fuel. Hydrogen might have an
advantage though in being light-weight if what you wanted was
horizontal launch. Say you were able to get to Mach 3+ with the jets,
1,000 m/s. The delta-V to supplied by the rocket-mode is then 7,900 m/
s. But note also you can get to high altitude say to 25,000 m. This
might subtract another 300 m/s from the required rocket-mode delta-V,
so now to 7,600 m/s.

A bigger advantage than this of the altitude is the fact that you get
the full vacuum Isp during rocket-mode, call it an exhaust velocity of
3,600 m/s for kerosene rockets. Note this results in a mass-ratio for
the rocket mode portion of e^(7,600/3,600) = 8.3, less than half that
usually cited for a kerosene-fueled all rocket SSTO. Note the fuel
required for the jet-powered portion would only be a fraction of the
dry mass rather than multiples of it based on the fact the 1,000 m/s
jet-powered speed is only a fraction of the 10,000 m/s or so effective
exhaust speed of jet engines.

Note this brings the kerosene fuel load to be about that of hydrogen
fueled SSTO's, except you still have the high density of kerosene.
With modern lightweight materials this should be well doable.

Bob Clark

c.f.,

Newsgroups: sci.space.policy, sci.astro, sci.physics,
sci.space.history, rec.arts.sf.science
From: Keith Henson <hkeithhen...@gmail.com>
Date: Thu, 17 Nov 2011 16:26:16 -0800 (PST)
Subject: Re: A kerosene-fueled X-33 as a single stage to orbit
vehicle.
http://groups.google.com/group/rec.arts.sf.science/msg/f7d229b2f96f222c?hl=en&

[Jeff Findley]

In article <3b2f7eed-d8a7-44be-ae29-
6d1e68...@s18g2000vby.googlegroups.com>, rgrego...@yahoo.com
says...

>
> Looking at the numbers I'm now convinced you can make a single stage
> to orbit vehicle with a combined ramjet/rocket engine, and without
> having to use scramjets.

While this seems like there might be advantages, it would be very
telling to compare this design to an all liquid fueled rocket engine
powered vehicle. While comparing to another SSTO, I would compare your
vehicle with two different engine cycles to a TSTO with optimized liquid
fueled rocket engines on each stage.

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

[Sylvia Else]

On 27/01/2012 12:18 AM, Robert Clark wrote:

> For the calculation the delta-V and propellant load would be feasible,
> note that for a dense propellant SSTO might require as much as 300 m/s
> lower delta-V than a hydrogen fueled SSTO, in the range of about 8,900
> m/s, so I'll use kerosene as the fuel. Hydrogen might have an
> advantage though in being light-weight if what you wanted was
> horizontal launch. Say you were able to get to Mach 3+ with the jets,
> 1,000 m/s. The delta-V to supplied by the rocket-mode is then 7,900 m/
> s. But note also you can get to high altitude say to 25,000 m. This
> might subtract another 300 m/s from the required rocket-mode delta-V,
> so now to 7,600 m/s.

Given that SSTO seems to be a decidedly marginal proposition, words like
"might" really have no place in the discussion. To judge a proposal one
needs a detailed numerical analysis, not just handwaving.

Sylvia.

[Brad Guth]

But your Skylon hand waving was always sooooo impressive.

http://translate.google.com/#
Brad Guth, Brad_Guth, Brad.Guth, BradGuth, BG / “Guth Usenet”

[Jack Tingle]

On 1/26/2012 7:09 PM, Sylvia Else wrote:

> Given that SSTO seems to be a decidedly marginal proposition, words like
> "might" really have no place in the discussion. To judge a proposal one
> needs a detailed numerical analysis, not just handwaving.

Nope, I think at this point in the history of aerospace we can probably
write it off without further discussion. It's pretty much been beaten to
death.

Skeptically,
Jack Tingle

[Jonathan]

"Sylvia Else" <syl...@not.here.invalid> wrote in message
news:9oe89g...@mid.individual.net...

Presidend Obama already solved NASA's cost to orbit problem.
The solution is called...unmanned.



s



>
> Sylvia.

[trigonometry1972@gmail.com |]

On Jan 28, 1:41 pm, Fred J. McCall <fjmcc...@gmail.com> wrote:

> "Jonathan" <Callinst...@gmail.com> wrote:
>
> >Presidend Obama already solved NASA's cost to orbit problem.
> >The solution is called...unmanned.
>

> Uh, how does that reduce launch costs, again?
>
> --
> "Some people get lost in thought because it's such unfamiliar
>  territory."
>                                       --G. Behn

Less dead weight, less spam in the can plus more room
for those leam mean machines/instruments. Of course,
it doesn't improve the transport tech.

I'd assume a SSTO for Mars would be more doable, if humans
could get it there.

Mars the chance to live in a cave and to operate local remotes for
exploration and
extraction...................................................Trig

[alie...@gmail.com]

On Jan 26, 5:18 am, Robert Clark <rgregorycl...@yahoo.com> wrote:
> Looking at the numbers I'm now convinced you can make a single stage
> to orbit vehicle with a combined ramjet/rocket engine, and without
> having to use scramjets.

AFAIK the only advantage to a jet mode is not having to carry
oxidizer for that part of the ascent within atmosphere dense enough
*at a specific velocity* to obtain sufficient oxygen to take the place
of the oxidizer. "Sufficient" tails off with height of course, so you
have to tailor velocity profile to density (modulo oxygen
concentration per altitude).

Rather than favoring a particular design a priori, it would seem
that the first step is determining the maximum altitude at which any
kind of air-breather will work *better than a rocket* at that
altitude. That's your potential final engine configuration before
going to pure rocket.

Anybody done that?

Then, determine the altitude and velocity domains in which a
specific type of jet is most efficient. Then, determine the
feasibility of combining the winners so they can transition from one
mode to the next *without awkward loss of thrust* during transition.

IOW, work backwards from a high-altitude efficiency benchmark to
determine what sort of engine you use to launch with. I'll mention
recalling reading that at sea level, piston engines are more efficient
than any kind of jet. I'd love to see a SSTO with props...

Heard of the Pulse Detonation Engine?

http://en.wikipedia.org/wiki/Pulse_detonation_engine

Still under research but very promising.

> The idea is to combine the turbo-ramjet/rocket into a single engine.
> This is what Skylon wants to do with their Sabre engine. But the Sabre
> will use hypersonic airbreathing propulsion up to Mach 5.5 before the
> rockets take over. This will require complicated air-cooling methods
> using heat exchangers with flowing liquid hydrogen for the Skylon.

ISTM that as long as the air-cooling tech is passive and doesn't
introduce enough drag to offset the no-oxidizer-aboard advantage, and
the hydrogen cooling tech isn't as heavy as the oxidizer would have
been, then fine. Otherwise, no point.

But the above suggested analytical path may indicate it isn't
worthwhile to include either turbojet or ramjet mode.


Mark L. Fergerson

[Jens Egon Nyborg]

Den 29-01-2012 01:07, Fred J. McCall skrev:

> "trigonom...@gmail.com |"<trigonom...@gmail.com> wrote:
>
>> On Jan 28, 1:41 pm, Fred J. McCall<fjmcc...@gmail.com> wrote:
>>> "Jonathan"<Callinst...@gmail.com> wrote:
>>>
>>>> Presidend Obama already solved NASA's cost to orbit problem.
>>>> The solution is called...unmanned.
>>>
>>> Uh, how does that reduce launch costs, again?
>>>
>>

>> Less dead weight, less spam in the can plus more room
>> for those leam mean machines/instruments. Of course,
>> it doesn't improve the transport tech.
>>
>

> It also doesn't reduce launch costs. Launch costs are measured in
> Dollars Per POUND. It reduces PAYLOAD, but so would any other
> reduction in mission capability.
>
> Not sending anything up reduces launch costs to zero?
>
> Oh, and I wasn't aware you were Jonathan, although you're both nutty
> as fruitcakes...
>

Some people measure launch cost in bang/bucks. Depends on your purpose I
suppose.

Anyone know the relevant SI units?

[Sylvia Else]

On 29/01/2012 5:14 PM, nu...@bid.nes wrote:
> On Jan 26, 5:18 am, Robert Clark<rgregorycl...@yahoo.com> wrote:
>> Looking at the numbers I'm now convinced you can make a single stage
>> to orbit vehicle with a combined ramjet/rocket engine, and without
>> having to use scramjets.
>
> AFAIK the only advantage to a jet mode is not having to carry
> oxidizer for that part of the ascent within atmosphere dense enough
> *at a specific velocity* to obtain sufficient oxygen to take the place
> of the oxidizer. "Sufficient" tails off with height of course, so you
> have to tailor velocity profile to density (modulo oxygen
> concentration per altitude).
>
> Rather than favoring a particular design a priori, it would seem
> that the first step is determining the maximum altitude at which any
> kind of air-breather will work *better than a rocket* at that
> altitude. That's your potential final engine configuration before
> going to pure rocket.

There doesn't appear to be a theoretical limit, so at any altitude,
there is a velocity at which an airbreathing engine will out perform a
rocket.

In practice, that velocity becomes unmanageably (and indeed unreachably)
high as the atmosphere thins into the interplanetary medium, but its
existence means that your approach isn't going to work - there is no
altitude at which no kind of air breather willl work better than a rocket.

It's an engineering problem, not a theoretical one.

Sylvia.

[Thomas Womack]

In article <o2nai7tneuj8s0bja...@4ax.com>,
Fred J. McCall <fmc...@gmail.com> wrote:

>As I said, if you reduce the mission capability by reducing the
>payload, that is NOT 'reduced launch cost'. It's meaningless.

If, instead of launching a twelve-ton satellite, you launch a
hundred-ton orbiter with a twelve-ton satellite in its belly, release
the satellite, and land the orbiter, you're launching eight times as
much as you needed to, and spending an awful lot more on launch costs.

There /may/ be single satellites expensive enough that it's worth
launching a hundred-ton orbiter with a crew of satellite repairmen,
fixing the satellite, releasing it and landing the orbiter - but that
was pretty marginal for Hubble and there aren't many satellites more
expensive than Hubble.

Tom

[Jens Egon Nyborg]

Den 29-01-2012 15:51, Fred J. McCall skrev:

> Jens Egon Nyborg<je...@invalid.invalid> wrote:

>>
>> Anyone know the relevant SI units?
>>
>

> Just change dollars per pound to dollars per kilogram. That wasn't
> too hard....
>

Serves me right for being facetious, I suppose. (SI dollars? really?)

Looking beyond that, you seem to be arguing, that's we shouldn't look at
other ways of viewing than from a weight/expence angle.

We are some that lok at it from an value/expence angle, however.

This approach is much more complicated, of course, but, in my opinion,
not more error prone.

[Howard Brazee]

On Sun, 29 Jan 2012 10:12:06 +0100, Jens Egon Nyborg
<je...@invalid.invalid> wrote:

>Some people measure launch cost in bang/bucks. Depends on your purpose I
>suppose.

For general purposes, that is a fine measurement. There are specific
purposes where better measurements are available, but none have been
mentioned in this thread.

--
"In no part of the constitution is more wisdom to be found,
than in the clause which confides the question of war or peace
to the legislature, and not to the executive department."

- James Madison

[me]

On Sun, 29 Jan 2012 23:27:46 +1100, Sylvia Else
<syl...@not.here.invalid> wrote:


>There doesn't appear to be a theoretical limit, so at any altitude,
>there is a velocity at which an airbreathing engine will out perform a
>rocket.

Actually, there is. Consider the simple ratio of the energy available
from the combustion process to the energy of the captured airstream.
On a per mass basis this is the heat of combustion divided by the
total enthalpy of captured airstream.

http://preview.tinyurl.com/87gz29g

See 3.1 Airframe Integrated Scramjet Design Challenges 1st paragraph.

[Greg Goss]

Fred J. McCall <fjmc...@gmail.com> wrote:

>Jens Egon Nyborg <je...@invalid.invalid> wrote:

>>Serves me right for being facetious, I suppose. (SI dollars? really?)
>

>Yes, SI dollars. Really. Other than SDRs on the IMF, what other
>'International Currency' is there? The Euro? Can't use it for lots
>and lots of things that you can use dollars for. And dollars are
>already 'metric'; the decidollar is called a dime, the centidollar is
>called a penny....

And my house taxes are defined by a "mill" rate. But a hundred
dollars is a C-note and a thousand are a G-note or a grand. So it's
not really SI.
--
"Recessions catch what the auditors miss." (Galbraith)

[Sylvia Else]

All that means, to my mind, is that the performance of an ideal air
breathing engine aproaches that of a rocket engine asymptotically from
above as velocity increases.

Sylvia.

[Jeff Findley]

In article <NKadnd6O2cVkn7nS...@giganews.com>,
Calli...@gmail.com says...

> Presidend Obama already solved NASA's cost to orbit problem.
> The solution is called...unmanned.

This statement is so devoid of content that it's worthless. Just what
are you trying to say? Please be specific.

[Brian Thorn]

On Sat, 28 Jan 2012 09:15:12 -0500, "Jonathan" <Calli...@gmail.com>
wrote:

>> Given that SSTO seems to be a decidedly marginal proposition, words like
>> "might" really have no place in the discussion. To judge a proposal one
>> needs a detailed numerical analysis, not just handwaving.
>
>
>Presidend Obama already solved NASA's cost to orbit problem.
>The solution is called...unmanned.

We've had unmanned spaceflight since 1957, but we still have the cost
to orbit problem.

Brian

[Robert Clark]

On Jan 26, 9:25 am, Jeff Findley <jeff.find...@nospam.ugs.com> wrote:
> In article <3b2f7eed-d8a7-44be-ae29-
> 6d1e68509...@s18g2000vby.googlegroups.com>, rgregorycl...@yahoo.com

> says...
>
>
>
> > Looking at the numbers I'm now convinced you can make a single stage
> > to orbit vehicle with a combined ramjet/rocket engine, and without
> > having to use scramjets.
>
> While this seems like there might be advantages, it would be very
> telling to compare this design to an all liquid fueled rocket engine
> powered vehicle.  While comparing to another SSTO, I would compare your
> vehicle with two different engine cycles to a TSTO with optimized liquid
> fueled rocket engines on each stage.
>
> Jeff
> --

The point in its favor is that you can save on the fuel required for
the rocket propulsion mode. There are a few disadvantages though.
First because you are launching horizontally, both the wings and
landing gear have to be sized to support the full gross weight of the
vehicle. If you are launching vertically as with all rocket power then
even with horizontally landing, these have to be sized just for the
*dry mass* of the vehicle.
Also, if you did use separate jet engines for the airbreathing
portion then those heavy jet engines are sitting there doing nothing
for the longest portion of the flight when you are under rocket mode.
If you do combine the jet engines and rocket engines into one as I was
suggesting, then that is also a research project to get that to work.

The biggest advantage for the partial airbreathing method for a SSTO
might be though that rocket engineers have the idea that all rocket
propulsion for a SSTO won't work (despite the fact the rocket equation
says that it can.) So if you present them a method that can get a very
Isp for some portion of the trip they may be more inclined to believe
it can work(eventhough it might turn out the extra weight required
means an all rocket method can still get more payload to orbit.)



Bob Clark

[Jeff Findley]

In article <e1afdd9e-bd60-45bd-863f-
5dc8fc...@eb6g2000vbb.googlegroups.com>, rgrego...@yahoo.com
says...

>
> On Jan 26, 9:25ï¿œam, Jeff Findley <jeff.find...@nospam.ugs.com> wrote:
> > In article <3b2f7eed-d8a7-44be-ae29-
> > 6d1e68509...@s18g2000vby.googlegroups.com>, rgregorycl...@yahoo.com
> > says...
> >
> >
> >
> > > Looking at the numbers I'm now convinced you can make a single stage
> > > to orbit vehicle with a combined ramjet/rocket engine, and without
> > > having to use scramjets.
> >
> > While this seems like there might be advantages, it would be very
> > telling to compare this design to an all liquid fueled rocket engine

> > powered vehicle. ï¿œWhile comparing to another SSTO, I would compare your

> > vehicle with two different engine cycles to a TSTO with optimized liquid
> > fueled rocket engines on each stage.
>

> The point in its favor is that you can save on the fuel required for
> the rocket propulsion mode.

Fuel is an *extremely* small portion of today's launch costs.
Specifically what is saved is oxidizer, not what is traditionally
thought of as fuel (kerosene, LH2, and etc). I'd also like to note that
LOX is literally pennies per pound in industrial quantities. If saving
a tiny amount of LOX is the only advantage, why bother at all?

> There are a few disadvantages though.
> First because you are launching horizontally, both the wings and
> landing gear have to be sized to support the full gross weight of the
> vehicle. If you are launching vertically as with all rocket power then
> even with horizontally landing, these have to be sized just for the
> *dry mass* of the vehicle.
> Also, if you did use separate jet engines for the airbreathing
> portion then those heavy jet engines are sitting there doing nothing
> for the longest portion of the flight when you are under rocket mode.
> If you do combine the jet engines and rocket engines into one as I was
> suggesting, then that is also a research project to get that to work.
>
> The biggest advantage for the partial airbreathing method for a SSTO
> might be though that rocket engineers have the idea that all rocket
> propulsion for a SSTO won't work (despite the fact the rocket equation
> says that it can.) So if you present them a method that can get a very
> Isp for some portion of the trip they may be more inclined to believe
> it can work(eventhough it might turn out the extra weight required
> means an all rocket method can still get more payload to orbit.)

Companies like SpaceX are proving every day that it's possible to launch
payloads into orbit at far less cost than what NASA and the traditional
aerospace companies would have us believe. The way to prove an SSTO
would work is to build and fly an SSTO.

Another approach is to pursue reusability using a TSTO approach. The
Falcon launch vehicles get to orbit just fine using two stages and
they're using relatively inefficient (as measured by ISP) LOX/kerosene
propellants.

[Robert Clark]

On Jan 30, 5:14 pm, Jeff Findley <jeff.find...@nospam.ugs.com> wrote:
> ...

>
> Fuel is an *extremely* small portion of today's launch costs.
> Specifically what is saved is oxidizer, not what is traditionally
> thought of as fuel (kerosene, LH2, and etc).  I'd also like to note that
> LOX is literally pennies per pound in industrial quantities.  If saving
> a tiny amount of LOX is the only advantage, why bother at all?
>

I was referring to the fact that the problem with getting a SSTO is
the large amount of propellant that is needed for the structural mass
of the vehicle to hold all that mass, not the cost of the propellant.
It was thought such a large amount of propellant couldn't be carried
for the low amount of dry mass required. If the amount of propellant
could be reduced by using a partial airbreathing trajectory, then that
would bring the ratio of the propellant load to the dry mass within a
more reasonable range.


Bob Clark

[Jeff Findley]

In article <93f75b9f-7c1c-4652-aba4-c606b8c945b9
@l16g2000vbl.googlegroups.com>, rgrego...@yahoo.com says...

All the more reason to pursue a reusable TSTO before attempting a
reusable SSTO. Staging is a key way to drop dry mass when it's not
needed anymore. Yes it adds some complexity, but it's the sort of
complexity that's been around since the first successful launch to LEO.

Despite the hype, air breathing engines which transition to rocket mode
are a research project, not a proven technology. One could easily pour
billions of dollars into such a research topic without producing any
useful results. I was around when NASP was being pushed as the next
generation in reusable launch vehicles *and* hypersonic transports. I'm
still waiting...

[Greg Goss]

Jeff Findley <jeff.f...@nospam.ugs.com> wrote:
>5dc8fc...@eb6g2000vbb.googlegroups.com>, rgrego...@yahoo.com
>says...

>> The point in its favor is that you can save on the fuel required for
>> the rocket propulsion mode.
>
>Fuel is an *extremely* small portion of today's launch costs.
>Specifically what is saved is oxidizer, not what is traditionally
>thought of as fuel (kerosene, LH2, and etc). I'd also like to note that
>LOX is literally pennies per pound in industrial quantities. If saving
>a tiny amount of LOX is the only advantage, why bother at all?

Oxygen is heavy.

[Jeff Findley]

In article <9oqaep...@mid.individual.net>, go...@gossg.org says...

So are air breathing engines and wings.

Look, we're trying to minimize the cost per lb to LEO, not the wet mass
or the dry mass of the launch vehicle. Focusing on optimizing the wrong
parameter results in a vehicle which is not suited for the desired
purpose.

Optimizing for minimum wet mass (or dry mass) is fine, for expendable
missiles where the carrier (submarine, B-52, missile silo, etc.) can't
grow to carry a bigger (in size or mass) missile. Henry Spencer called
this mindset the "performance uber alles" mindset, which the US
aerospace industry seems to have obtained from its captured German
"rocket scientists". It's a perfectly valid metric for designing
missiles, but not so valid for designing reusable launch vehicles where
low cost is the goal.

Again, note the low cost that SpaceX is able to achieve even though they
are using ("heavy") LOX and relatively low ISP kerosene propellants in
their two stage launch vehicles.

[me]

Not sure by what rules your ideal system plays, but there is nothing
which says an air breather need make positive thrust.

[Sylvia Else]

In an ideal system there are no entropy gains. The captured airstream
undergoes adiabatic compression with an associated temperature increase.
This process is completely reversible, so that the the air can be
expanded with a lowering of temperature and return to its original
state. There would be no net force on the engine.

If heat is added between the compression and expansion, then the
expansion starts from a higher pressure, which can clearly be used to
generate thrust.

Achieving, or even approaching, an ideal system is not easy, of course,
but that is, as I said, and engineering problem.

Sylvia.

[alie...@gmail.com]

On Jan 31, 7:27 am, Fred J. McCall <fjmcc...@gmail.com> wrote:
> Greg Goss <go...@gossg.org> wrote:
> >Jeff Findley <jeff.find...@nospam.ugs.com> wrote:
> >>5dc8fc57e...@eb6g2000vbb.googlegroups.com>, rgregorycl...@yahoo.com

> >>says...
> >>>  The point in its favor is that you can save on the fuel required for
> >>> the rocket propulsion mode.
>
> >>Fuel is an *extremely* small portion of today's launch costs.
> >>Specifically what is saved is oxidizer, not what is traditionally
> >>thought of as fuel (kerosene, LH2, and etc).  I'd also like to note that
> >>LOX is literally pennies per pound in industrial quantities.  If saving
> >>a tiny amount of LOX is the only advantage, why bother at all?
>
> >Oxygen is heavy.
>

> And if you have to accelerate air (including all the parts you don't
> need) to something close to the velocity of your vehicle so that you
> can burn it in an engine, all you're saving is a little bit of tankage
> while accepting the weight of inlets.

Makes scramjets look better and better.


Mark L. Fergerson

[Greg Goss]

Fred J. McCall <fjmc...@gmail.com> wrote:

>Greg Goss <go...@gossg.org> wrote:
>

>And if you have to accelerate air (including all the parts you don't
>need) to something close to the velocity of your vehicle so that you
>can burn it in an engine, all you're saving is a little bit of tankage
>while accepting the weight of inlets.

I'm not claiming that air-breathing is the solution. I was just
disputing the assumption that the simpler solution was the best.

[Jeff Findley]

In article <9orqe5...@mid.individual.net>, go...@gossg.org says...

For low to moderate flight rates, the simpler solution generally costs
less to develop, is less error prone, and costs less to fly. That's the
world we live in today. This is NOT the time to propose uber complex
solutions that would require an extremely high flight rate in order to
justify their extremely high cost.

[Derek Lyons]

Robert Clark <rgrego...@yahoo.com> wrote:

>On Jan 30, 5:14 pm, Jeff Findley <jeff.find...@nospam.ugs.com> wrote:
>> ...
>>
>> Fuel is an *extremely* small portion of today's launch costs.
>> Specifically what is saved is oxidizer, not what is traditionally
>> thought of as fuel (kerosene, LH2, and etc).  I'd also like to note that
>> LOX is literally pennies per pound in industrial quantities.  If saving
>> a tiny amount of LOX is the only advantage, why bother at all?
>>
>
> I was referring to the fact that the problem with getting a SSTO is
>the large amount of propellant that is needed for the structural mass
>of the vehicle to hold all that mass, not the cost of the propellant.

But the problem with an air breathing SSTO is that you don't 'save'
structural mass - you move it from the tankage to the fuselage, wings,
landing gear/equipment, and engines. And worse yet, now you've got
haul all that mass all the way to orbit and back to the ground.

>It was thought such a large amount of propellant couldn't be carried
>for the low amount of dry mass required. If the amount of propellant
>could be reduced by using a partial airbreathing trajectory, then that
>would bring the ratio of the propellant load to the dry mass within a
>more reasonable range.

There's a problem here too - it's Very Bad Engineering to oprimize for
a single factor without adressing the costs of that optimization.
Especially when you're replacing something fairly cheap and simple (a
structural tube) with multiple somethings that are each, in and of
themselves, more expensive and not simple.

D.
--
Touch-twice life. Eat. Drink. Laugh.

http://derekl1963.livejournal.com/

-Resolved: To be more temperate in my postings.
Oct 5th, 2004 JDL

[Sylvia Else]

On 8/02/2012 6:37 AM, Derek Lyons wrote:
> Robert Clark<rgrego...@yahoo.com> wrote:
>
>> On Jan 30, 5:14 pm, Jeff Findley<jeff.find...@nospam.ugs.com> wrote:
>>> ...
>>>
>>> Fuel is an *extremely* small portion of today's launch costs.
>>> Specifically what is saved is oxidizer, not what is traditionally
>>> thought of as fuel (kerosene, LH2, and etc). I'd also like to note that
>>> LOX is literally pennies per pound in industrial quantities. If saving
>>> a tiny amount of LOX is the only advantage, why bother at all?
>>>
>>
>> I was referring to the fact that the problem with getting a SSTO is
>> the large amount of propellant that is needed for the structural mass
>> of the vehicle to hold all that mass, not the cost of the propellant.
>
> But the problem with an air breathing SSTO is that you don't 'save'
> structural mass - you move it from the tankage to the fuselage, wings,
> landing gear/equipment, and engines. And worse yet, now you've got
> haul all that mass all the way to orbit and back to the ground.

On the plus side, you get the mass back, still in its original
structural form, ready to be used again.

>
>> It was thought such a large amount of propellant couldn't be carried
>> for the low amount of dry mass required. If the amount of propellant
>> could be reduced by using a partial airbreathing trajectory, then that
>> would bring the ratio of the propellant load to the dry mass within a
>> more reasonable range.
>
> There's a problem here too - it's Very Bad Engineering to oprimize for
> a single factor without adressing the costs of that optimization.
> Especially when you're replacing something fairly cheap and simple (a
> structural tube) with multiple somethings that are each, in and of
> themselves, more expensive and not simple.

That's not bad engineering. It's an application of economics. The primay
goal is to reduce the cost per unit payload put into orbit. The engineer
is tasked with achieving that, not with complying with some arbitrary
notion about what constitutes good engineering.

Sylvia.

[Greg Goss]

fair...@gmail.com (Derek Lyons) wrote:

>Robert Clark <rgrego...@yahoo.com> wrote:
>
>>On Jan 30, 5:14 pm, Jeff Findley <jeff.find...@nospam.ugs.com> wrote:
>>> ...
>>>
>>> Fuel is an *extremely* small portion of today's launch costs.
>>> Specifically what is saved is oxidizer, not what is traditionally
>>> thought of as fuel (kerosene, LH2, and etc).  I'd also like to note that
>>> LOX is literally pennies per pound in industrial quantities.  If saving
>>> a tiny amount of LOX is the only advantage, why bother at all?
>>>
>>
>> I was referring to the fact that the problem with getting a SSTO is
>>the large amount of propellant that is needed for the structural mass
>>of the vehicle to hold all that mass, not the cost of the propellant.
>
>But the problem with an air breathing SSTO is that you don't 'save'
>structural mass - you move it from the tankage to the fuselage, wings,
>landing gear/equipment, and engines. And worse yet, now you've got
>haul all that mass all the way to orbit and back to the ground.

Is there any real reason to avoid 2STO with both stages re-usable?
Put the air breathing engines in the air, and the rockets up where it
gets thin.

[Sylvia Else]

Making air breathing engines that can operate at high enough speeds to
be useful is already problematic.

Additional complexity is involved in separation; not just a separation
mechanism is required, but either the 2nd stage motors have to start
before separation, or ullage motors, or some other system, will be
required to ensure that the second stage propellants are properly
located in their tanks. Failure to separate, or failure in the ullage
system, will likely cause loss of vehicle. Starting the 2nd stage motors
before separation pretty much implies a piggy-back style configuration,
which creates other problems.

Recovering the first stage to the launch point will be difficult because
of the speed and down range distance at separation. The first stage may
have to land somewhere else and then be returned to the launch location
as a separation operation. (Could have multiple launch locations, spread
round the world, with the first stage being reused where it lands - but
the market would have to be large to justify multiple launch sites).

It may be more trouble than it's worth, economically, if a larger SSTO,
albeit with a lower payload fraction, can be built to put the same
payload into orbit.

Sylvia.

[Greg (Strider) Moore]

"Sylvia Else" wrote in message news:9pebh1...@mid.individual.net...

>
>On 8/02/2012 6:37 AM, Derek Lyons wrote:
.
>>
>> But the problem with an air breathing SSTO is that you don't 'save'
>> structural mass - you move it from the tankage to the fuselage, wings,
>> landing gear/equipment, and engines. And worse yet, now you've got
>> haul all that mass all the way to orbit and back to the ground.
>
>On the plus side, you get the mass back, still in its original structural
>form, ready to be used again.
>

You have the same plus with a non-airbreathing SSTO. So you really haven't
gained anything here.

>
>> There's a problem here too - it's Very Bad Engineering to oprimize for
>> a single factor without adressing the costs of that optimization.
>> Especially when you're replacing something fairly cheap and simple (a
>> structural tube) with multiple somethings that are each, in and of
>> themselves, more expensive and not simple.
>
>That's not bad engineering. It's an application of economics. The primay
>goal is to reduce the cost per unit payload put into orbit. The engineer is
>tasked with achieving that, not with complying with some arbitrary notion
>about what constitutes good engineering.

Right, so how does replacing a large O2 fuel tank, one of the cheapest
things there is with a complex engine reduce costs again?


>
>Sylvia.
>
>

--
Greg D. Moore http://greenmountainsoftware.wordpress.com/
CEO QuiCR: Quick, Crowdsourced Responses. http://www.quicr.net

[Sylvia Else]

On 8/02/2012 10:41 PM, Greg (Strider) Moore wrote:
> "Sylvia Else" wrote in message news:9pebh1...@mid.individual.net...
>>
>> On 8/02/2012 6:37 AM, Derek Lyons wrote:
> .
>>>
>>> But the problem with an air breathing SSTO is that you don't 'save'
>>> structural mass - you move it from the tankage to the fuselage, wings,
>>> landing gear/equipment, and engines. And worse yet, now you've got
>>> haul all that mass all the way to orbit and back to the ground.
>>
>> On the plus side, you get the mass back, still in its original
>> structural form, ready to be used again.
>>
>
> You have the same plus with a non-airbreathing SSTO. So you really
> haven't gained anything here.

No credible design for a non-airbreathing reusable SSTO exists.

>
>>
>>> There's a problem here too - it's Very Bad Engineering to oprimize for
>>> a single factor without adressing the costs of that optimization.
>>> Especially when you're replacing something fairly cheap and simple (a
>>> structural tube) with multiple somethings that are each, in and of
>>> themselves, more expensive and not simple.
>>
>> That's not bad engineering. It's an application of economics. The
>> primay goal is to reduce the cost per unit payload put into orbit. The
>> engineer is tasked with achieving that, not with complying with some
>> arbitrary notion about what constitutes good engineering.
>
> Right, so how does replacing a large O2 fuel tank, one of the cheapest
> things there is with a complex engine reduce costs again?

It's not just that the oxygen tank is reduced in the air-breather, the
oxygen itself used during the air-breathing phase doesn't have to be
carried. This feeds through into the vehicle size. How big is the
pure-rocket reusable vehicle? Can you build it at all?

You can't even make the cost comparison unless you have a credible
design for a reusable pure-rocket SSTO, and, as I said above, there is
no credible design in existence.

Sylvia.

[Derek Lyons]

Then why optimize for things that are *INCREASING* the cost per unit
payload?

>The engineer is tasked with achieving that, not with complying with some
>arbitrary notion about what constitutes good engineering.

Had I insisted he complied with some arbitrary notion, you'd have a
point. But I did no such thing. In fact, if you actually read my
message you'd note the emphasis on cost.

[Jeff Findley]

In article <9pebh1...@mid.individual.net>, syl...@not.here.invalid
says...

>
> On 8/02/2012 6:37 AM, Derek Lyons wrote:
> > Robert Clark<rgrego...@yahoo.com> wrote:
> >
> >> On Jan 30, 5:14 pm, Jeff Findley<jeff.find...@nospam.ugs.com> wrote:
> >>> ...
> >>>
> >>> Fuel is an *extremely* small portion of today's launch costs.
> >>> Specifically what is saved is oxidizer, not what is traditionally
> >>> thought of as fuel (kerosene, LH2, and etc). I'd also like to note that
> >>> LOX is literally pennies per pound in industrial quantities. If saving
> >>> a tiny amount of LOX is the only advantage, why bother at all?
> >>>
> >>
> >> I was referring to the fact that the problem with getting a SSTO is
> >> the large amount of propellant that is needed for the structural mass
> >> of the vehicle to hold all that mass, not the cost of the propellant.
> >
> > But the problem with an air breathing SSTO is that you don't 'save'
> > structural mass - you move it from the tankage to the fuselage, wings,
> > landing gear/equipment, and engines. And worse yet, now you've got
> > haul all that mass all the way to orbit and back to the ground.
>
> On the plus side, you get the mass back, still in its original
> structural form, ready to be used again.

But the logical failure you are making is that reusing this hardware is
possible with an air breathing SSTO but not possible with something more
conventional such as a TSTO with liquid fueled rocket engines for each
stage. There is nothing magical about an air breathing SSTO in this
regard.

When it comes to cost, simpler is cheaper. Cost scales much more
closely to complexity than it does to dry mass.

> >> It was thought such a large amount of propellant couldn't be carried
> >> for the low amount of dry mass required. If the amount of propellant
> >> could be reduced by using a partial airbreathing trajectory, then that
> >> would bring the ratio of the propellant load to the dry mass within a
> >> more reasonable range.
> >
> > There's a problem here too - it's Very Bad Engineering to oprimize for
> > a single factor without adressing the costs of that optimization.
> > Especially when you're replacing something fairly cheap and simple (a
> > structural tube) with multiple somethings that are each, in and of
> > themselves, more expensive and not simple.
>
> That's not bad engineering. It's an application of economics. The primay
> goal is to reduce the cost per unit payload put into orbit. The engineer
> is tasked with achieving that, not with complying with some arbitrary
> notion about what constitutes good engineering.

Given today's launch markets, the economics simply don't work for an air
breathing SSTO. Those fancy engines are an R&D project which will cost
billions to develop. On top of that, they require a *much* more complex
vehicle to be designed around them. Requiring wings and aerodynamics
optimized for maximum intake of air into the engines is a great way to
blow billions upon billions on aerodynamics optimization, which leads to
huge headaches for the structural engineers due to the resulting complex
geometries.

Complex geometries typically mean more structural mass, so the
structural engineers typically resort to exotic structural materials in
an attempt to keep mass down. That means bleeding edge composites
instead of more conventional metallic alloys. Those bleeding edge
composites in conjunction with complex geometries really drive up costs.

Off the shelf liquid fueled rocket engines coupled with (simple and
cheap reusable LOX tanks) makes far more economic sense. The vehicle
ends up being much more simple in terms of aerodynamics which results in
a far more simple vehicle structurally. Simple structures are easier to
design, build, and fly.

If using liquid fueled rocket engines means that you can't do SSTO, and
need a TSTO instead, then so be it. Multi-stage launch vehicles have
been the norm since Sputnik started the beginning of the space age.

[Jeff Findley]

In article <9pf876...@mid.individual.net>, syl...@not.here.invalid
says...

>
> On 8/02/2012 10:41 PM, Greg (Strider) Moore wrote:
> > "Sylvia Else" wrote in message news:9pebh1...@mid.individual.net...
> >>
> >> On 8/02/2012 6:37 AM, Derek Lyons wrote:
> > .
> >>>
> >>> But the problem with an air breathing SSTO is that you don't 'save'
> >>> structural mass - you move it from the tankage to the fuselage, wings,
> >>> landing gear/equipment, and engines. And worse yet, now you've got
> >>> haul all that mass all the way to orbit and back to the ground.
> >>
> >> On the plus side, you get the mass back, still in its original
> >> structural form, ready to be used again.
> >>
> >
> > You have the same plus with a non-airbreathing SSTO. So you really
> > haven't gained anything here.
>
> No credible design for a non-airbreathing reusable SSTO exists.

Who cares about SSTO? What about a reusable TSTO? Are you saying a
reusable TSTO with liquid fueled rocket engines isn't credible?

TSTO means your first stage engines can be optimized for sea level
operation and a high thrust to weight ratio while your upper stage
engines can be optimized for high ISP and vacuum operation. Engines
optimized for each role can be bought off the shelf since every single
launch vehicle currently in operation uses such engines.

There is nothing inherent in a liquid fueled rocket engine that makes it
disposable. Countless test stand runs and a few uses in test vehicles
(e.g. DC-X) and operational launch vehicles (space shuttle) have shown
this to be true.

> >>> There's a problem here too - it's Very Bad Engineering to oprimize for
> >>> a single factor without adressing the costs of that optimization.
> >>> Especially when you're replacing something fairly cheap and simple (a
> >>> structural tube) with multiple somethings that are each, in and of
> >>> themselves, more expensive and not simple.
> >>
> >> That's not bad engineering. It's an application of economics. The
> >> primay goal is to reduce the cost per unit payload put into orbit. The
> >> engineer is tasked with achieving that, not with complying with some
> >> arbitrary notion about what constitutes good engineering.
> >
> > Right, so how does replacing a large O2 fuel tank, one of the cheapest
> > things there is with a complex engine reduce costs again?
>
> It's not just that the oxygen tank is reduced in the air-breather, the
> oxygen itself used during the air-breathing phase doesn't have to be
> carried. This feeds through into the vehicle size. How big is the
> pure-rocket reusable vehicle? Can you build it at all?

Cost scales much more closely with complexity than it does with vehicle
size. Again, the way to reduce the size of a reusable launch vehicle is
to drop unnecessary dry mass along the way to orbit. Even though I
don't think that the metric makes sense, a reusable TSTO is easier to
design when it comes to minimizing vehicle size and payload mass
fraction.

What really matters is cost per pound to orbit. Who really cares if the
vehicle is big, as long as it's cheap to operate?

> You can't even make the cost comparison unless you have a credible
> design for a reusable pure-rocket SSTO, and, as I said above, there is
> no credible design in existence.

Again, why the focus on SSTO? I don't even want to see a fully reusable
TSTO as a "next step". What I'd like to see is a TSTO with a fully
reusable first stage and an expendable upper stage. Why? Because
*everything* is easier on a reusable first stage when compared to a
reusable SSTO.

Also note that on a Falcon 9, the first stage has *nine* engines while
the upper stage only has *one*. A reusable Falcon 9 first stage would,
at a minimum, save the costs of throwing away nine times the engines
that are thrown away by the upper stage. If such a reusable first stage
is necessarily bigger and needs more tankage, engines, and etc. so be
it. At least they're all being reused instead of thrown away.

[Jeff Findley]

In article <9pehil...@mid.individual.net>, go...@gossg.org says...

If you do that, you still have the problem that the upper stage is a
near SSTO, unless your first stage flies very high and very fast. Air
breathing vehicles which fly very high and very fast have very complex
air breathing engines to the point that they are a research project
(i.e. hypersonic air breathing engines). Air breathing engines which
switch to a pure rocket mode are still more complex and more expensive.

I don't see how you can reduce costs by increasing complexity. Cost
scales with complexity far more closely than it scales with dry mass of
the vehicle.

[Jeff Findley]

In article <9penvm...@mid.individual.net>, syl...@not.here.invalid
says...

>
> On 8/02/2012 5:03 PM, Greg Goss wrote:
> > fair...@gmail.com (Derek Lyons) wrote:
> >
> >> Robert Clark<rgrego...@yahoo.com> wrote:
> >>
> >>> On Jan 30, 5:14 pm, Jeff Findley<jeff.find...@nospam.ugs.com> wrote:
> >>>> ...
> >>>>
> >>>> Fuel is an *extremely* small portion of today's launch costs.
> >>>> Specifically what is saved is oxidizer, not what is traditionally
> >>>> thought of as fuel (kerosene, LH2, and etc). I'd also like to note that
> >>>> LOX is literally pennies per pound in industrial quantities. If saving
> >>>> a tiny amount of LOX is the only advantage, why bother at all?
> >>>>
> >>>
> >>> I was referring to the fact that the problem with getting a SSTO is
> >>> the large amount of propellant that is needed for the structural mass
> >>> of the vehicle to hold all that mass, not the cost of the propellant.
> >>
> >> But the problem with an air breathing SSTO is that you don't 'save'
> >> structural mass - you move it from the tankage to the fuselage, wings,
> >> landing gear/equipment, and engines. And worse yet, now you've got
> >> haul all that mass all the way to orbit and back to the ground.
> >
> > Is there any real reason to avoid 2STO with both stages re-usable?
> > Put the air breathing engines in the air, and the rockets up where it
> > gets thin.
> >
>
> Making air breathing engines that can operate at high enough speeds to
> be useful is already problematic.

This is also a problem with an air breathing SSTO. In order to "pay"
for the extra mass of the air breathing engines, they need to operate as
high and fast as possible, otherwise you're just hauling all that extra
engine mass all the way to LEO for no real benefit.

> Additional complexity is involved in separation; not just a separation
> mechanism is required, but either the 2nd stage motors have to start
> before separation, or ullage motors, or some other system, will be
> required to ensure that the second stage propellants are properly
> located in their tanks. Failure to separate, or failure in the ullage
> system, will likely cause loss of vehicle. Starting the 2nd stage motors
> before separation pretty much implies a piggy-back style configuration,
> which creates other problems.

This complexity has been in every single successful launch to LEO, so
it's a sort of complexity that is handled with time proven technologies.
To me, this is better than depending on an engine design which has never
flown.

> Recovering the first stage to the launch point will be difficult because
> of the speed and down range distance at separation. The first stage may
> have to land somewhere else and then be returned to the launch location
> as a separation operation. (Could have multiple launch locations, spread
> round the world, with the first stage being reused where it lands - but
> the market would have to be large to justify multiple launch sites).

That all depends on the launch trajectory and when the first stage
separates. To minimize this problem, the first stage could launch on a
trajectory which is mostly vertical in order to get the second stage out
of the atmosphere (where its vacuum optimized upper stage engines can
operate). The resulting horizontal velocity on reentry of the first
stage would be close to zero.

The other way to handle this is to do a propulsive burn post-separation
which would put the nearly empty first stage on a return trajectory.
The fuel and oxidizer to do this would be far less than that burned
prior to second stage separation because the nearly empty first stage
mass is *far* less than that of a full first stage with a full second
stage on top.

Even if the first stage reenters at some "high" velocity, the reentry
conditions will still be far less demanding than those experienced by an
reusable SSTO (or reusable upper stage) which is reentering at nearly
orbital velocity.

> It may be more trouble than it's worth, economically, if a larger SSTO,
> albeit with a lower payload fraction, can be built to put the same
> payload into orbit.

But the problem we're trying to solve is lowering launch costs when
compared to today's expendables. A TSTO with a reusable first stage and
an expendable second stage would be a step in the right direction.
Neither a fully reusable SSTO nor a fully reusable TSTO is necessary to
solve the problem at hand.

[Greg Goss]

Sylvia Else <syl...@not.here.invalid> wrote:

>On 8/02/2012 5:03 PM, Greg Goss wrote:
>> Is there any real reason to avoid 2STO with both stages re-usable?
>> Put the air breathing engines in the air, and the rockets up where it
>> gets thin.
>>
>
>Making air breathing engines that can operate at high enough speeds to
>be useful is already problematic.

This is probably the killer. Getting altitude above much of the
atmosphere is "nice", and allows a better optimization of the rocket
bell of the second stage, but you're limited to how much speed can be
provided. Blackbird, which I think runs just under Mach 3 has much
cleaner lines than some kind of siamese twin structure.

>Additional complexity is involved in separation; not just a separation
>mechanism is required, but either the 2nd stage motors have to start
>before separation, or ullage motors, or some other system, will be
>required to ensure that the second stage propellants are properly
>located in their tanks.

I'd like to see further discussion of this point. Can you recommend a
suitable web page?

I'm not a "rocket scientist", so all I know of "ullage" is a quickie
discussion on wikipedia. You're still in gravity at separation, so
any empty space in the tank is at a known position. With full tanks,
even a rear central draw point (halfway up the tank when horizontal,
"bottom" of the tank under thrust) would be fully immersed. It's not
like restarting liquid fuelled engines from orbit.

I'm visualizing it as a piggyback, which you don't like, but I don't
really know the plusses or minuses of doing some kind of
"disintegrating roman candle" in this context.

>Failure to separate, or failure in the ullage
>system, will likely cause loss of vehicle. Starting the 2nd stage motors
>before separation pretty much implies a piggy-back style configuration,
>which creates other problems.
>
>Recovering the first stage to the launch point will be difficult because
>of the speed and down range distance at separation. The first stage may
>have to land somewhere else and then be returned to the launch location

>as a separat[e] operation.

The first stage presumably acts like an airplane. Return to base
should be a lot simpler than the Space Shuttle's. Wouldn't you just
refuel it and drive back?

>It may be more trouble than it's worth, economically, if a larger SSTO,
>albeit with a lower payload fraction, can be built to put the same
>payload into orbit.

[Derek Lyons]

Jeff Findley <jeff.f...@nospam.ugs.com> wrote:

>I don't see how you can reduce costs by increasing complexity.

Yet, increased complexity is part of the very definition of a
re-useable.

[Sylvia Else]

The complexity has already cause the loss of missions, even though the
total number of missions flown is really not that large.

When you enter the realms of reusable designs, it's not just the failure
rate that has to be looked at, but the consequences of failures. Failure
of of the separation mechanism will probably result in loss of vehicle.
Failure, other than catastrophic failure, of an engine in an
appropriately designed reusable system should only result in a mission
abort, with both the vehicle and payload surviving for another attempt
after the fault is repaired.

That is to say, a higher failure rate is acceptable if the consequences
are not so serious.

>
>> Recovering the first stage to the launch point will be difficult because
>> of the speed and down range distance at separation. The first stage may
>> have to land somewhere else and then be returned to the launch location
>> as a separation operation. (Could have multiple launch locations, spread
>> round the world, with the first stage being reused where it lands - but
>> the market would have to be large to justify multiple launch sites).
>
> That all depends on the launch trajectory and when the first stage
> separates. To minimize this problem, the first stage could launch on a
> trajectory which is mostly vertical in order to get the second stage out
> of the atmosphere (where its vacuum optimized upper stage engines can
> operate). The resulting horizontal velocity on reentry of the first
> stage would be close to zero.

It would increase the gravity losses. I'm doubtful of whether getting
the second stage out of the atmosphere that way is worthwhile.

>
> The other way to handle this is to do a propulsive burn post-separation
> which would put the nearly empty first stage on a return trajectory.
> The fuel and oxidizer to do this would be far less than that burned
> prior to second stage separation because the nearly empty first stage
> mass is *far* less than that of a full first stage with a full second
> stage on top.

So now the first stage also needs ullage motors, and restartable engines.

>
> Even if the first stage reenters at some "high" velocity, the reentry
> conditions will still be far less demanding than those experienced by an
> reusable SSTO (or reusable upper stage) which is reentering at nearly
> orbital velocity.
>
>> It may be more trouble than it's worth, economically, if a larger SSTO,
>> albeit with a lower payload fraction, can be built to put the same
>> payload into orbit.
>
> But the problem we're trying to solve is lowering launch costs when
> compared to today's expendables. A TSTO with a reusable first stage and
> an expendable second stage would be a step in the right direction.
> Neither a fully reusable SSTO nor a fully reusable TSTO is necessary to
> solve the problem at hand.
>
> Jeff

Sylvia.

[Sylvia Else]

But what are you comparing the cost with? Replacing a cheap tank by an
expensive engine may indeed be economic if the cheap tank gets thrown
away each time, but the expensive engine isn't.

If you're not throwing away the cheap tank, then you need a credible way
of recovering it and reusing it at a cost that doesn't negate its
cheapness. The obvious answer is a reusable pure rocket SSTO, but none
exists, even on paper.

Sylvia.

[Sylvia Else]

On 9/02/2012 3:37 AM, Greg Goss wrote:
> Sylvia Else<syl...@not.here.invalid> wrote:
>
>> On 8/02/2012 5:03 PM, Greg Goss wrote:
>>> Is there any real reason to avoid 2STO with both stages re-usable?
>>> Put the air breathing engines in the air, and the rockets up where it
>>> gets thin.
>>>
>>
>> Making air breathing engines that can operate at high enough speeds to
>> be useful is already problematic.
>
> This is probably the killer. Getting altitude above much of the
> atmosphere is "nice", and allows a better optimization of the rocket
> bell of the second stage, but you're limited to how much speed can be
> provided. Blackbird, which I think runs just under Mach 3 has much
> cleaner lines than some kind of siamese twin structure.
>
>> Additional complexity is involved in separation; not just a separation
>> mechanism is required, but either the 2nd stage motors have to start
>> before separation, or ullage motors, or some other system, will be
>> required to ensure that the second stage propellants are properly
>> located in their tanks.
>
> I'd like to see further discussion of this point. Can you recommend a
> suitable web page?

Sorry, I have no references.

>
> I'm not a "rocket scientist", so all I know of "ullage" is a quickie
> discussion on wikipedia. You're still in gravity at separation,

The gravitational field remains allmost unchanged even when you're in
LEO. It's not what counts. At the point of separation, and with the
engines shut down, the vehicle is pretty much in free-fall, with only
the residual atmospheric drag acting on the vehicle. In that situation,
the fuel will not have any preffered position in the tanks (except for a
slight tendency to move to the front because of the residual drag).

so
> any empty space in the tank is at a known position. With full tanks,
> even a rear central draw point (halfway up the tank when horizontal,
> "bottom" of the tank under thrust) would be fully immersed. It's not
> like restarting liquid fuelled engines from orbit.
>
> I'm visualizing it as a piggyback, which you don't like, but I don't
> really know the plusses or minuses of doing some kind of
> "disintegrating roman candle" in this context.

>
>> Failure to separate, or failure in the ullage
>> system, will likely cause loss of vehicle. Starting the 2nd stage motors
>> before separation pretty much implies a piggy-back style configuration,
>> which creates other problems.
>>
>> Recovering the first stage to the launch point will be difficult because
>> of the speed and down range distance at separation. The first stage may
>> have to land somewhere else and then be returned to the launch location
>> as a separat[e] operation.
>
> The first stage presumably acts like an airplane. Return to base
> should be a lot simpler than the Space Shuttle's. Wouldn't you just
> refuel it and drive back?

Perhaps, but then you need a place for it to land and a way of
refuelling it. If it runs on LH2, you'll need to install the
infrastructure, and that doesn't come cheap.

Indeed, you'll probably have to build the landing strip, because using
an existing strip will be very disruptive to the normal use. That's
because the vehicle will have no, or very limited, ability, to fit into
a normal traffic pattern.

These issues are by no means insurmountable, but they all add to the cost.

>
>> It may be more trouble than it's worth, economically, if a larger SSTO,
>> albeit with a lower payload fraction, can be built to put the same
>> payload into orbit.
>
>

Sylvia.

[Sylvia Else]

On 9/02/2012 1:27 AM, Jeff Findley wrote:
> Requiring wings and aerodynamics
> optimized for maximum intake of air into the engines is a great way to
> blow billions upon billions on aerodynamics optimization, which leads to
> huge headaches for the structural engineers due to the resulting complex
> geometries.

Note that Skylon avoids that problem by having intakes that are not part
of the wing, and are clear of all vehicle generated shock waves except
the nose, and they're completely inside the shock cone for that.

Sylvia.

[Jeff Findley]

In article <4f35de80...@news.supernews.com>, fair...@gmail.com
says...

>
> Jeff Findley <jeff.f...@nospam.ugs.com> wrote:
>
> >I don't see how you can reduce costs by increasing complexity.
>
> Yet, increased complexity is part of the very definition of a
> re-useable.

True, but unnecessary complexity should be avoided.

Suitable liquid fueled rocket engines for use in a TSTO already exist.
The sort of engine like SABRE which has to operate as an air breather
from subsonic speeds to hypersonic speeds and then operate as a liquid
fueled rocket engine from there to orbit do not exist. I'd trade in a
SABRE based reusable SSTO design for a liquid fueled rocket engine based
reusable TSTO any day.

The sort of "complexity" that is involved in stage separation is the
sort of "complexity" which has existed in every single launch vehicle
that's successfully placed a payload in LEO since Sputnik.

I'd rather a design rely on proven complexity than on the SABRE sort of
complexity which makes aerospace propulsion and aerodynamics researchers
giddy with joy while simultaneously giving the structural engineers
headaches while trying to design a complex structure and simultaneously
meet aggressive mass fraction goals.

[Jeff Findley]

In article <9pgoc2...@mid.individual.net>, syl...@not.here.invalid
says...

True, which is why SSTO supporters don't like TSTOs dropping stages.

But, I'm trying to find a middle ground between current expendable TSTOs
and a fully reusable SSTO. The gap between the two is so great that it
might span tens of billions of dollars in R&D and perhaps decades of
time. I'd really like to find something which costs far less and takes
less time to develop. If it's better than today's expendable TSTOs,
that's all that seems to matter.

Replacing the first stage of an expendable TSTO with a reusable first
stage seems a logical next step down the path of reusability. In terms
of failure modes, it shouldn't be any worse than the expendable first
stage, because all of the same failure modes caused by events such as
stage separation are exactly the same.

In terms of reliability, the reusable first stage ought to win simply
because it's reusable. A reusable first stage only makes its first
flight once, preferably on a test flight which isn't carrying payload
for a paying customer. If it survives its first test flight, any bugs
found are fixed and it can be test flown again to prove itself at least
somewhat reliable. An expendable first stage is *always* making its
first flight. Any bugs found on that first flight potentially impact
the payload being carried.

Cost isn't the only reason reusable hardware is preferable to
expendable. Customers really don't want to see their payloads crashing
into the ocean. They'll often pay a bit of a premium in cost to gain a
better chance at mission success.

> >> Recovering the first stage to the launch point will be difficult
because
> >> of the speed and down range distance at separation. The first stage may
> >> have to land somewhere else and then be returned to the launch location
> >> as a separation operation. (Could have multiple launch locations, spread
> >> round the world, with the first stage being reused where it lands - but
> >> the market would have to be large to justify multiple launch sites).
> >
> > That all depends on the launch trajectory and when the first stage
> > separates. To minimize this problem, the first stage could launch on a
> > trajectory which is mostly vertical in order to get the second stage out
> > of the atmosphere (where its vacuum optimized upper stage engines can
> > operate). The resulting horizontal velocity on reentry of the first
> > stage would be close to zero.
>
> It would increase the gravity losses. I'm doubtful of whether getting
> the second stage out of the atmosphere that way is worthwhile.

For one thing, it's worthwhile because it makes the design of the second
stage easier. The second stage can have engines optimized for vacuum
for optimal ISP. Designing an engine which operates from sea level
pressure to vacuum introduces compromises which impact performance,
cost, and reliability. The SSME is one fine example of such an engine,
and look at how many redesigns it took to make it as reliable as it is
today. This sort of engine requirement introduces complexities into the
engine itself.

Secondly, and more importantly, such a design would allow you to start
off with a reusable first stage and an expendable second stage. Again,
I see this as a middle ground between today's expendable TSTOs,
tomorrow's fully reusable TSTOs, and finally (hopefully) a fully
reusable SSTO.

> > The other way to handle this is to do a propulsive burn post-separation
> > which would put the nearly empty first stage on a return trajectory.
> > The fuel and oxidizer to do this would be far less than that burned
> > prior to second stage separation because the nearly empty first stage
> > mass is *far* less than that of a full first stage with a full second
> > stage on top.
>
> So now the first stage also needs ullage motors, and restartable engines.

A reusable anything is going to either need restartable engines or
additional engines. Otherwise, how do you deorbit? How would a
reusable vehicle maintain its orientation in orbit after main engine
shutdown? This criticism would apply to any reusable vehicle and is
therefore an unfair attack on a reusable first stage.

In a Falcon 9 like design, you would not need all nine engines firing to
land. At this point the vehicle is very light having burned nearly all
its propellant and having released its payload (the second stage). If
an engine fails to restart, try another.

Besides, from the payload's point of view a complete failure to restart
the first stage engines for recovery of the stage is a non-event. The
paying customer still gets its payload into orbit. From this point of
view, the failure is no worse than a successful, expendable, first
stage, so such a failure in no way jeopardizes the mission.


In conclusion, a reusable first stage only needs to be better than the
expendable stage its replacing to be a success. That's a far easier
problem to solve than the problems facing an SSTO.

[Rick Jones]

In sci.space.history Jeff Findley <jeff.f...@nospam.ugs.com> wrote:
> Replacing the first stage of an expendable TSTO with a reusable
> first stage seems a logical next step down the path of reusability.
> In terms of failure modes, it shouldn't be any worse than the
> expendable first stage, because all of the same failure modes caused
> by events such as stage separation are exactly the same.

But the first stage of a TSTO has "new and improved" failure modes of
its own right? I mean it has more "functionality" than the previous,
expendable first stage and so more things to potentially go wrong. Is
it really going to be the case that none of those are worse? Can you
really assume the "new" failures in the reusable first stage are all
less severe than before?

rick jones
--
a wide gulf separates "what if" from "if only"
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...

[Sylvia Else]

On 10/02/2012 12:47 AM, Jeff Findley wrote:
> In article<4f35de80...@news.supernews.com>, fair...@gmail.com
> says...
>>
>> Jeff Findley<jeff.f...@nospam.ugs.com> wrote:
>>
>>> I don't see how you can reduce costs by increasing complexity.
>>
>> Yet, increased complexity is part of the very definition of a
>> re-useable.
>
> True, but unnecessary complexity should be avoided.
>
> Suitable liquid fueled rocket engines for use in a TSTO already exist.
> The sort of engine like SABRE which has to operate as an air breather
> from subsonic speeds to hypersonic speeds and then operate as a liquid
> fueled rocket engine from there to orbit do not exist. I'd trade in a
> SABRE based reusable SSTO design for a liquid fueled rocket engine based
> reusable TSTO any day.
>
> The sort of "complexity" that is involved in stage separation is the
> sort of "complexity" which has existed in every single launch vehicle
> that's successfully placed a payload in LEO since Sputnik.

None of which have been reusable. That complexity has also been part of
every failed launch, and in some cases has been the cause of the failure.

Sylvia.

[Sylvia Else]

On 10/02/2012 1:21 AM, Jeff Findley wrote:

> But, I'm trying to find a middle ground between current expendable TSTOs
> and a fully reusable SSTO. The gap between the two is so great that it
> might span tens of billions of dollars in R&D and perhaps decades of
> time. I'd really like to find something which costs far less and takes
> less time to develop. If it's better than today's expendable TSTOs,
> that's all that seems to matter.
>
> Replacing the first stage of an expendable TSTO with a reusable first
> stage seems a logical next step down the path of reusability. In terms
> of failure modes, it shouldn't be any worse than the expendable first
> stage, because all of the same failure modes caused by events such as
> stage separation are exactly the same.

They're exactly the same if the separation mechanism is the same, but
that implies that the first stage sits below the second stage as in
current launcher designs, which in turn implies a vertical launch, and
associated difficulties with managing failures so that they do not cause
loss of vehicle and payload.

>
> In terms of reliability, the reusable first stage ought to win simply
> because it's reusable. A reusable first stage only makes its first
> flight once, preferably on a test flight which isn't carrying payload
> for a paying customer. If it survives its first test flight, any bugs
> found are fixed and it can be test flown again to prove itself at least
> somewhat reliable. An expendable first stage is *always* making its
> first flight. Any bugs found on that first flight potentially impact
> the payload being carried.

I've wondered from time to time to what extent launcher failures are
caused by an inability to construct the launcher properly given a
working design, and to what extent they're do to an apparent inability
to launch the same design twice. It seems there's always some tinkering
of the design going on, so that every launch is a design test flight.

Deorbit is achieved using a separate orbital maneovering system (OMS),
rather than the main engines. It doesn't take much delta-v to leave LEO,
since all you need is to change your orbit so that it enters the
atmosphere. Consequently, the main engines of an SSTO do not need to be
restartable, nor do they need to be started in free-fall.

The space shuttle didn't use its main engines for deorbit. I can't
imagine that any resuable system would.

I should observe that Skylon's OMS is intended itself to be LH2,LOX
based, to avoid the need to manage highly toxic fuels on the ground. I
don't know how they intend to handle the free-fall problem, but probably
not by way of ullage motors. The system is much smaller than the tanks
for the main engines.

> How would a
> reusable vehicle maintain its orientation in orbit after main engine
> shutdown? This criticism would apply to any reusable vehicle and is
> therefore an unfair attack on a reusable first stage.

Using the OMS.

> In a Falcon 9 like design, you would not need all nine engines firing to
> land. At this point the vehicle is very light having burned nearly all
> its propellant and having released its payload (the second stage). If
> an engine fails to restart, try another.
>
> Besides, from the payload's point of view a complete failure to restart
> the first stage engines for recovery of the stage is a non-event. The
> paying customer still gets its payload into orbit. From this point of
> view, the failure is no worse than a successful, expendable, first
> stage, so such a failure in no way jeopardizes the mission.
>
>
> In conclusion, a reusable first stage only needs to be better than the
> expendable stage its replacing to be a success. That's a far easier
> problem to solve than the problems facing an SSTO.

If a TSTO only has a reusable first stage, then any failure in the
second stage results in loss of the payload. That risk has to be
included in the launch costs (typically born by the owner of the payload
by way of insurance premiums).

Sylvia.

[Greg Goss]

Sylvia Else <syl...@not.here.invalid> wrote:

>> A reusable anything is going to either need restartable engines or
>> additional engines. Otherwise, how do you deorbit?
>
>Deorbit is achieved using a separate orbital maneovering system (OMS),
>rather than the main engines. It doesn't take much delta-v to leave LEO,
>since all you need is to change your orbit so that it enters the
>atmosphere. Consequently, the main engines of an SSTO do not need to be
>restartable, nor do they need to be started in free-fall.
>
>The space shuttle didn't use its main engines for deorbit. I can't
>imagine that any resuable system would.

Didn't the Russian shuttle have a third set of engines, a pair of
flip-out air breathers that would allow a go-around on landing?

[Jeff Findley]

In article <jh1cs9$c15$1...@usenet01.boi.hp.com>, rick....@hp.com
says...

>
> In sci.space.history Jeff Findley <jeff.f...@nospam.ugs.com> wrote:
> > Replacing the first stage of an expendable TSTO with a reusable
> > first stage seems a logical next step down the path of reusability.
> > In terms of failure modes, it shouldn't be any worse than the
> > expendable first stage, because all of the same failure modes caused
> > by events such as stage separation are exactly the same.
>
> But the first stage of a TSTO has "new and improved" failure modes of
> its own right? I mean it has more "functionality" than the previous,
> expendable first stage and so more things to potentially go wrong. Is
> it really going to be the case that none of those are worse? Can you
> really assume the "new" failures in the reusable first stage are all
> less severe than before?

The devil is in the details, but to a first order approximation,
everything which needs to function *prior to first stage separation* on
a reusable (VTVL) first stage is the same as on an expendable first
stage. One reason why I prefer VTVL over HTHL is the similarity in
complexity with existing expendables. A HTHL first stage (e.g. powered
by combined cycle engines like SABRE) would introduce failure modes
related to its air breathing engines, wings, ailerons, rudder, landing
gear, and other hardware needed for takeoffs and landings.

A a reusable VTVL first stage will have all the same failure modes of a
conventional, expendable, first stage. This includes hardware like
liquid fueled rocket engines, hydraulic or electric actuators, valves,
pressurization systems, stage separation hardware, gyros, radios, flight
control computers, and etc.

The added complexity, that is the hardware which makes it reusable, is
only used *after* first stage separation. Therefore, failures in the
additional hardware which makes it reusable would not put the payload in
jeopardy.

[Jeff Findley]

In article <9pj5a7...@mid.individual.net>, syl...@not.here.invalid
says...

>
> On 10/02/2012 1:21 AM, Jeff Findley wrote:
>
> > But, I'm trying to find a middle ground between current expendable TSTOs
> > and a fully reusable SSTO. The gap between the two is so great that it
> > might span tens of billions of dollars in R&D and perhaps decades of
> > time. I'd really like to find something which costs far less and takes
> > less time to develop. If it's better than today's expendable TSTOs,
> > that's all that seems to matter.
> >
> > Replacing the first stage of an expendable TSTO with a reusable first
> > stage seems a logical next step down the path of reusability. In terms
> > of failure modes, it shouldn't be any worse than the expendable first
> > stage, because all of the same failure modes caused by events such as
> > stage separation are exactly the same.
>
> They're exactly the same if the separation mechanism is the same, but
> that implies that the first stage sits below the second stage as in
> current launcher designs, which in turn implies a vertical launch, and
> associated difficulties with managing failures so that they do not cause
> loss of vehicle and payload.

VTVL would be preferred because it is most similar to the existing first
stage. Furthermore, VTVL does not generally introduce additional
failure modes during launch, as HTHL does. For a HTHL vehicle, there
are a myriad of systems introduced related to the landing gear, wings,
and aerodynamic control surfaces. For control, a rocket powered VTVL
vehicle need only rely on engine gimbaling to control its orientation
and its flight path, which is exactly the same as today's expendable
launch vehicles.

> > In terms of reliability, the reusable first stage ought to win
simply
> > because it's reusable. A reusable first stage only makes its first
> > flight once, preferably on a test flight which isn't carrying payload
> > for a paying customer. If it survives its first test flight, any bugs
> > found are fixed and it can be test flown again to prove itself at least
> > somewhat reliable. An expendable first stage is *always* making its
> > first flight. Any bugs found on that first flight potentially impact
> > the payload being carried.
>
> I've wondered from time to time to what extent launcher failures are
> caused by an inability to construct the launcher properly given a
> working design, and to what extent they're do to an apparent inability
> to launch the same design twice. It seems there's always some tinkering
> of the design going on, so that every launch is a design test flight.

The recent slew of Russian launch failures seem to be exactly what you
describe. Their launch vehicle designs have remained largely the same
for decades. What remains to cause failures are related to either
changes in the design or to manufacturing defects. This is exactly the
"infant mortality" problem which is solved by a reusable launch vehicle
proving itself through test flights *before* flying a payload for a
paying customer.

As far as big engines go, the J-2 was restartable, and it was the main
engine for the Saturn IB upper stage, the Saturn V second stage, and the
Saturn V upper stage. On today's upper stages, I believe that the RL-10
engines are restartable. There is *nothing* fundamentally difficult
about restarting a liquid fueled rocket engine, even what I would call
large liquid fueled rocket engines (the J-2 was no small engine). It's
certainly a much easier problem to solve than creating hypersonic air
breathing engines, which simply do not yet exist.

> I should observe that Skylon's OMS is intended itself to be LH2,LOX
> based, to avoid the need to manage highly toxic fuels on the ground. I
> don't know how they intend to handle the free-fall problem, but probably
> not by way of ullage motors. The system is much smaller than the tanks
> for the main engines.
>
> > How would a
> > reusable vehicle maintain its orientation in orbit after main engine
> > shutdown? This criticism would apply to any reusable vehicle and is
> > therefore an unfair attack on a reusable first stage.
>
> Using the OMS.

In other words, a resatartable liquid fueled rocket engine.

> > In a Falcon 9 like design, you would not need all nine engines firing to
> > land. At this point the vehicle is very light having burned nearly all
> > its propellant and having released its payload (the second stage). If
> > an engine fails to restart, try another.
> >
> > Besides, from the payload's point of view a complete failure to restart
> > the first stage engines for recovery of the stage is a non-event. The
> > paying customer still gets its payload into orbit. From this point of
> > view, the failure is no worse than a successful, expendable, first
> > stage, so such a failure in no way jeopardizes the mission.
> >
> >
> > In conclusion, a reusable first stage only needs to be better than the
> > expendable stage its replacing to be a success. That's a far easier
> > problem to solve than the problems facing an SSTO.
>
> If a TSTO only has a reusable first stage, then any failure in the
> second stage results in loss of the payload. That risk has to be
> included in the launch costs (typically born by the owner of the payload
> by way of insurance premiums).

True, but that's also true for today's TSTO expendables. Since I'm
proposing replacing the first stage of an existing TSTO expendable, the
risk is *exactly the same*.

Again, I'm looking for a way to improve the current situation which does
*not* require tens of billions of dollars of R&D and decades before
hardware flies on today's launch vehicles. A reuable first stage topped
by an existing second stage certainly seems like a logical next step.

[Sylvia Else]

Again, we have to make a clear distinction between failures that result
in loss of vehicle (and payload), and failures that merely result in a
mission abort.

Landing gear and aerodynamic control systems are mature technologies
used every day in the carrying of people in airliners. While failures do
occur, they are managed, and very rarely cause loss of vehicle.

It's not that restartable engines are difficult to build. In a sense,
all liquid fuel engines are restartable, since I believe they're ground
fired before being attached to a launch vehcile. The issue relates to
the additional equipment required to restart them in a zero-gee environment.

Sylvia.

[Sylvia Else]

On 11/02/2012 12:29 AM, Jeff Findley wrote:

> The added complexity, that is the hardware which makes it reusable, is
> only used *after* first stage separation. Therefore, failures in the
> additional hardware which makes it reusable would not put the payload in
> jeopardy.

That's not entirely true. Things have been known to trigger when they're
not meant to, sometimes with disasterous consequences.

Sylvia.

[Snidely]

Sylvia Else <syl...@not.here.invalid> scribbled something like ...

> It's not that restartable engines are difficult to build. In a sense,
> all liquid fuel engines are restartable, since I believe they're
> ground fired before being attached to a launch vehcile. The issue
> relates to the additional equipment required to restart them in a
> zero-gee environment.
>

Which had been worked out by the time Gemini was docking with Agena stages.
Some of those engines are fired after years in space (viz Cassini).

/dps

[Jeff Findley]

In article <9pm0v6...@mid.individual.net>, syl...@not.here.invalid
says...

>
> On 11/02/2012 12:57 AM, Jeff Findley wrote:
> >
> > VTVL would be preferred because it is most similar to the existing first
> > stage. Furthermore, VTVL does not generally introduce additional
> > failure modes during launch, as HTHL does. For a HTHL vehicle, there
> > are a myriad of systems introduced related to the landing gear, wings,
> > and aerodynamic control surfaces. For control, a rocket powered VTVL
> > vehicle need only rely on engine gimbaling to control its orientation
> > and its flight path, which is exactly the same as today's expendable
> > launch vehicles.
>
> Again, we have to make a clear distinction between failures that result
> in loss of vehicle (and payload), and failures that merely result in a
> mission abort.
>
> Landing gear and aerodynamic control systems are mature technologies
> used every day in the carrying of people in airliners. While failures do
> occur, they are managed, and very rarely cause loss of vehicle.

While true, the supersonic and hypersonic speeds coupled with the high
temperatures of reentry experienced by a reusable winged launch vehicle
means that the (mostly subsonic) experience you're talking about does
not necessarily directly transfer. The X-15, SR-71, and etc. have a
more similar flight regime, so the data pool there is actually much more
shallow than that which would contain all aircraft.

As an example, despite the successes of prior Pegasus launch vehicles,
the first Pegasus XL was lost due to errors in predictions of vehicle
response to various aerodynamic forces:

PEGASUS XL ANOMALY REVIEW COMPLETED
http://sunland.gsfc.nasa.gov/smex/fast/mission/pegfail.html

In other words, the sorts of things you're asserting are mature
technologies are the very things that did in the first Pegasus XL
launch.

> > As far as big engines go, the J-2 was restartable, and it was the
main
> > engine for the Saturn IB upper stage, the Saturn V second stage, and the
> > Saturn V upper stage. On today's upper stages, I believe that the RL-10
> > engines are restartable. There is *nothing* fundamentally difficult
> > about restarting a liquid fueled rocket engine, even what I would call
> > large liquid fueled rocket engines (the J-2 was no small engine). It's
> > certainly a much easier problem to solve than creating hypersonic air
> > breathing engines, which simply do not yet exist.
>
> It's not that restartable engines are difficult to build. In a sense,
> all liquid fuel engines are restartable, since I believe they're ground
> fired before being attached to a launch vehcile. The issue relates to
> the additional equipment required to restart them in a zero-gee environment.

True, which is why I gave the J-2 and RL-10 as examples of high
performance restartable liquid fueled rocket engines. This is 1960's
tech we're talking about.

There is a requirement to settle the propellants in the tanks prior to
ignition, but this does not necessarily result in additional hardware.
I would give that task to the vehicle's RCS system, which is always
going to be there due to the requirement to provide proper orientation
in zero gravity during periods when the main engines are not firing. If
you don't have an RCS, you can't control the orientation of your vehicle
prior to reentry.

[Jeff Findley]

In article <9pm14n...@mid.individual.net>, syl...@not.here.invalid
says...

I thought I prefaced what I said with "to a first order approximation".
Obviously there are additional failure modes introduced.

Note that wings, control surfaces, and landing gears also introduce
additional failure modes of their own. If your landing gear deploys
during launch, that would be a bad thing too.

Some of those sorts of failures have been really weird and took a lot of
time and effort to solve. Here is a crazy example of what has gone
wrong with flight control surfaces in the past:

http://en.wikipedia.org/wiki/United_Airlines_Flight_585

[Sylvia Else]

On 14/02/2012 6:47 AM, Jeff Findley wrote:
> In article<9pm14n...@mid.individual.net>, syl...@not.here.invalid
> says...
>>
>> On 11/02/2012 12:29 AM, Jeff Findley wrote:
>>
>>> The added complexity, that is the hardware which makes it reusable, is
>>> only used *after* first stage separation. Therefore, failures in the
>>> additional hardware which makes it reusable would not put the payload in
>>> jeopardy.
>>
>> That's not entirely true. Things have been known to trigger when they're
>> not meant to, sometimes with disasterous consequences.
>
> I thought I prefaced what I said with "to a first order approximation".
> Obviously there are additional failure modes introduced.
>
> Note that wings, control surfaces, and landing gears also introduce
> additional failure modes of their own. If your landing gear deploys
> during launch, that would be a bad thing too.
>
> Some of those sorts of failures have been really weird and took a lot of
> time and effort to solve. Here is a crazy example of what has gone
> wrong with flight control surfaces in the past:
>
> http://en.wikipedia.org/wiki/United_Airlines_Flight_585

But it should be noted that despite the design flaw, occurences were
still rare when expressed as a probability per flight.

The same is true of the design flaw that caused the crash landing of BA
Flight 38

http://en.wikipedia.org/wiki/British_Airways_Flight_38

The type had been in service for twelve years before that event.

Sylvia.

[Jeff Findley]

In article <9ptpek...@mid.individual.net>, syl...@not.here.invalid
says...

True, but there are design flaws related to how parts wear time. I
can't remember if age and/or wear was a factor in this failure. But
it's easy to come up with examples where fatigue, age, or wear caused a
failure.

I'm just pointing out that there are many complexities in air breathing
vehicles. Since subsonic jet aircraft have been around for decades, we
do have a lot of experience with designing, building, and flying them,
so reliability tends to remain high despite the complexity. However,
this complexity still drives up costs (manufacturing and maintenance).

It's still my belief that any HTHL air breathing SSTO is going to be
more complex than a (sane) VTVL TSTO. The engines alone are far more
complex than liquid fueled rocket engines. The aerodynamics and
airframe are absolutely more complex.

Complexity drives up cost and tends to drive down reliability. In order
to improve reliability of a complex piece of hardware, you have to do
rigorous (expensive, time consuming) testing. And as BA Flight 38
showed, sometimes lab testing doesn't duplicate the conditions in flight
and you still run into failures.

[Derek Lyons]

If you think Skylon doesn't have expensive and headache inducing
intakes, I'll have some of what you're smoking. (Hint: The expensive
and headache inducing parts come of having intakes in the first place,
and are largely irrelevant to their location.)

[Derek Lyons]

Jeff Findley <jeff.f...@nospam.ugs.com> wrote:

>In article <4f35de80...@news.supernews.com>, fair...@gmail.com
>says...
>>
>> Jeff Findley <jeff.f...@nospam.ugs.com> wrote:
>>
>> >I don't see how you can reduce costs by increasing complexity.
>>
>> Yet, increased complexity is part of the very definition of a
>> re-useable.
>

>Suitable liquid fueled rocket engines for use in a TSTO already exist.

For a suitable level of handwaving, yeah. But, absent handwaving,
you're left with just the SSME. Engine, not engines. The DC-X
engines weren't orbit capable. And extrapolating from the performance
of disposable engines that are swaddled in cotton from cradle to
watery grave to the much more rigorous environment of a reuseable is
questionable at best.

>The sort of "complexity" that is involved in stage separation is the
>sort of "complexity" which has existed in every single launch vehicle
>that's successfully placed a payload in LEO since Sputnik.

Your failure and unwillingness to discuss the remaining complexities,
of great interest to the operators and all occuring after seperation
is noted.

>I'd rather a design rely on proven complexity than on the SABRE sort of
>complexity which makes aerospace propulsion and aerodynamics researchers
>giddy with joy while simultaneously giving the structural engineers
>headaches while trying to design a complex structure and simultaneously
>meet aggressive mass fraction goals.

Yeah, adding on the weight and complexity if re-use to a "traditional"
rocket stage won't cause anyone any headaches, or require changes to
structure, or change the mass fraction, etc...

In other words, after your extensive handwaving - we're right back
where we started.

[Jeff Findley]

In article <4f425dc3....@news.supernews.com>, fair...@gmail.com

says...
>
> Jeff Findley <jeff.f...@nospam.ugs.com> wrote:
>
> >In article <4f35de80...@news.supernews.com>, fair...@gmail.com
> >says...
> >>
> >> Jeff Findley <jeff.f...@nospam.ugs.com> wrote:
> >>
> >> >I don't see how you can reduce costs by increasing complexity.
> >>
> >> Yet, increased complexity is part of the very definition of a
> >> re-useable.
> >
> >Suitable liquid fueled rocket engines for use in a TSTO already exist.
>
> For a suitable level of handwaving, yeah. But, absent handwaving,
> you're left with just the SSME. Engine, not engines. The DC-X
> engines weren't orbit capable. And extrapolating from the performance
> of disposable engines that are swaddled in cotton from cradle to
> watery grave to the much more rigorous environment of a reuseable is
> questionable at best.

Actually, the SSME isn't restartable, so it's right out for a VTVL TSTO
stage, unless you're willing to pay the price for separate landing
engines. Remember that the SSME was originally spec'ed to power their
HLV's upper stage, but wsa swapped for the J-2X. Once fully developed,
the J-2X would be a suitable upper stage engine.

SpaceX has engines which are much more suitable for a reusable TSTO.
Falcon 9's upper stage Merlin engine is (in flight) restartable. I
doubt it would take much to make the first stage engines (in flight)
restartable, if they already aren't.

For a TSTO, I don't see why existing RL-10's (currently used in the
upper stages of the EELV's) would not be suitable for powering a
reusable upper stage.

> >The sort of "complexity" that is involved in stage separation is the
> >sort of "complexity" which has existed in every single launch vehicle
> >that's successfully placed a payload in LEO since Sputnik.
>
> Your failure and unwillingness to discuss the remaining complexities,
> of great interest to the operators and all occuring after seperation
> is noted.

I don't see how this is evasive when we've been doing stage separations
on every single successful launch to LEO or beyond. That's decades of
experience compared to zip for SABRE.

> >I'd rather a design rely on proven complexity than on the SABRE sort of
> >complexity which makes aerospace propulsion and aerodynamics researchers
> >giddy with joy while simultaneously giving the structural engineers
> >headaches while trying to design a complex structure and simultaneously
> >meet aggressive mass fraction goals.
>
> Yeah, adding on the weight and complexity if re-use to a "traditional"
> rocket stage won't cause anyone any headaches, or require changes to
> structure, or change the mass fraction, etc...
>
> In other words, after your extensive handwaving - we're right back
> where we started.

The size of the stages of a resuable TSTO will be bigger and heavier
than an expendable TSTO with the same payload. But since you're not
throwing the hardware away after every single flight, and because fuel
and oxidizer are such a small percentage of total launch costs, you'll
still come out ahead.

Also, as Henry Spencer repeatedly said in these groups, cost scales
closely with complexity but rather loosely with size. Anyone who has
looked at a Big Dumb Booster type of design in detail knows this.

Unfortunately the typical aerospace engineer gets caught in what Henry
Spencer called the "performance uber alles" mindset inherited from the
German "rocket scientists". This is an excellent mindset for designing
missiles like the V-2, but not so much for a resuable launch vehicle.

[Sylvia Else]

On 17/02/2012 6:38 AM, Derek Lyons wrote:
> Sylvia Else<syl...@not.here.invalid> wrote:
>
>> On 9/02/2012 1:27 AM, Jeff Findley wrote:
>>> Requiring wings and aerodynamics optimized for maximum intake of
>>> air into the engines is a great way to blow billions upon billions
>>> on aerodynamics optimization, which leads to huge headaches for
>>> the structural engineers due to the resulting complex
>>> geometries.
>>
>> Note that Skylon avoids that problem by having intakes that are not part
>> of the wing, and are clear of all vehicle generated shock waves except
>> the nose, and they're completely inside the shock cone for that.
>
>
> If you think Skylon doesn't have expensive and headache inducing
> intakes, I'll have some of what you're smoking. (Hint: The expensive
> and headache inducing parts come of having intakes in the first place,
> and are largely irrelevant to their location.)
>
> D.

Of course there are still intakes, but positioning the intakes so that
the wing has no effect on them makes life a lot easier. In particular,
there are no "wings and aerodynamics optimized for maximum intake of
air into the engines" leading to the blowing of billions apon billions.

Sylvia.

[Jeff Findley]

In article <9q5s0a...@mid.individual.net>, syl...@not.here.invalid
says...

>
> On 17/02/2012 6:38 AM, Derek Lyons wrote:
> > Sylvia Else<syl...@not.here.invalid> wrote:
> >
> >> On 9/02/2012 1:27 AM, Jeff Findley wrote:
> >>> Requiring wings and aerodynamics optimized for maximum intake of
> >>> air into the engines is a great way to blow billions upon billions
> >>> on aerodynamics optimization, which leads to huge headaches for
> >>> the structural engineers due to the resulting complex
> >>> geometries.
> >>
> >> Note that Skylon avoids that problem by having intakes that are not part
> >> of the wing, and are clear of all vehicle generated shock waves except
> >> the nose, and they're completely inside the shock cone for that.
> >
> > If you think Skylon doesn't have expensive and headache inducing
> > intakes, I'll have some of what you're smoking. (Hint: The expensive
> > and headache inducing parts come of having intakes in the first place,
> > and are largely irrelevant to their location.)
>

> Of course there are still intakes, but positioning the intakes so that
> the wing has no effect on them makes life a lot easier. In particular,
> there are no "wings and aerodynamics optimized for maximum intake of
> air into the engines" leading to the blowing of billions apon billions.

The intakes they show on the notional Skylon design look awfully small
to the point of being laughable. I really don't see how they're going
to be effective in the thin upper atmosphere.

NASP and other wave rider type designs maximize intake efficiency by
essentially turning the entire underside of the vehicle (before the
engines) into an intake. They also tend to turn the underside of the
aft part of the vehicle into part of the "engine bell" in order to
maximize the amount of thrust extracted from the exiting exhaust gas.

[Sylvia Else]

Effective in what sense? Having enough cross-sectional area to capture
the required amount of air? I doubt that RE would have made such a basic
mistake.

The engine sizes are somewhat deceptive, given that they're longer than
the payload bay which is itself 13 metres long. The intakes are about 5
metres across.

Sylvia.