The 100/10/1 Rule.

In article ,
Alan Jones wrote:
I think the Russians designed or developed a tri-propellant engine
that would initially burn LOX and a dense hydrocarbon fuel at
relatively high thrust, and then switch to LH fuel at a lower thrust
but higher ISP...

"Designed" is the right word -- they may have done a bit of early testing
on the basic concept, but the RD-701 was nowhere near fully developed.

This was supposed to be good for nearly SSTO
vehicles. This did not pan out, and I have not read anything about
this concept in a long time. Can you provide some history on this?
In view of kT's fixation on SSTO, is there any merit to revisiting
that concept?

In the early 90s, if I'm remembering correctly, there was a burst of
enthusiasm for tripropellant systems. However, not everybody agreed that
they were a good idea -- a fair number of analyses optimized with the
cutover point so early or late that one of the two fuels essentially
disappeared. (Which one depended on the assumptions and the optimization
criteria.) Most of the more optimistic analyses eventually ended up
going the same way, when some bugs were gotten out and the underlying
assumptions were carefully scrutinized. In the end, pretty much everyone
agreed that the extra complexity didn't seem worth it.
--
spsystems.net is temporarily off the air; | Henry Spencer
mail to henry at zoo.utoronto.ca instead. |

Pat Flannery wrote:

There were two NASA studies on it, but neither of them are on the web:
http://tinyurl.com/2243lo
http://tinyurl.com/263dkc
Now that's odd... those were supposed to be these:

Title: Application of dual-fuel propulsion to a single stage AMLS
vehicle
Author(s): Lepsch, Roger A., Jr.; Stanley, Douglas O.; Unal, Resit
Abstract: As part of NASA's Advanced Manned Launch System (AMLS)
study to determine a follow-on, or complement, to the Space Shuttle, a
reusable single-stage-to-orbit concept utilizing dual-fuel rocket
propulsion has been examined. Several dual-fuel propulsion concepts were
investigated. These include a separate engine concept combining Russian
RD-170 kerosene-fueled engines with SSME-derivative engines the kerosene
and hydrogen-fueled Russian RD-701 engine concept and a dual-fuel,
dual-expander engine concept. Analysis to determine vehicle weight and
size characteristics was performed using conceptual level design
techniques. A response surface methodology for multidisciplinary design
was utilized to optimize the dual-fuel vehicle concepts with respect to
several important propulsion system and vehicle design parameters in
order to achieve minimum empty weight. Comparisons were then made with a
hydrogen-fueled reference, single-stage vehicle. The tools and methods
employed in the analysis process are also summarized.
NASA Center: Langley Research Center
Publication Date: Jun 1, 1993
Document Source: Other Sources
No Digital Version Available: Go to Tips On Ordering
Document ID: 19930066069
Accession ID: 93A50066
Report Number: AIAA PAPER 93-2275
Meeting Information: 29th; AIAA, SAE, ASME, and ASEE, Joint
Propulsion Conference and Exhibit; June 28-30, 1993; Monterey, CA;
United States
Keywords: HYDROGEN KEROSENE PROPULSION SYSTEM CONFIGURATIONS
REUSABLE ROCKET ENGINES SINGLE STAGE ROCKET VEHICLES SINGLE STAGE TO
ORBIT VEHICLES OPTIMIZATION PERFORMANCE PREDICTION
Notes: AIAA, SAE, ASME, and ASEE, Joint Propulsion Conference and
Exhibit, 29th, Monterey, CA, June 28-30, 1993
Accessibility: Unclassified; Copyright; Unlimited; Publicly available;
Updated/Added to NTRS: 2004-11-03


Title: Thermo-Kinetics Characterization of Kerosene RP-1
Combustion for Tri-Propellant Engine Design Calculations
Author(s): Wang, Ten-See
Abstract: A one-formula surrogate fuel C12H24 and its quasi-global
kinetics have been developed to support the conceptual design of the
injectors and thrust chambers of the tri-propellant engines. This
surrogate fuel represents a fuel blend that properly depicts the
physical and chemical properties of kerosene RP-l. The accompanying
thermodynamics for the substitute fuel is generated and is anchored with
the heat of formation of kerosene RP-1. The surrogate fuel model and its
thermodynamics are verified by comparing a series of one dimensional
rocket thrust chamber shifting-equilibrium calculations under a
kerosene-fueled Russian Engine (RD-170) operating conditions. The models
are then tested in conjunction with a computational fluid dynamics
formulation to predicting the thermo-flowfield of two single-element
shear tri-axial injectors a gaseous oxygen kerosene Hydrogen injector
and a liquid oxygen liquid kerosene gaseous hydrogen injector. The
mixture ratio of those injectors is that of a proposed tri-propellant
engine RD-704. Reasonable flame structure is predicted for those injectors.
NASA Center: NASA (non Center Specific)
Publication Date: Oct. 1995
Document Source: CPIA, 10630 Little Patuxent Pkwy., Suite 202,
Columbia, MD 21044-3264
No Digital Version Available: Go to Tips On Ordering
Document ID: 19960026197
Accession ID: 96N27925
Keywords: ENGINE DESIGN THRUST CHAMBERS ROCKET THRUST COMPUTATIONAL
FLUID DYNAMICS THERMODYNAMICS PREDICTIONS LIQUID OXYGEN LIQUID HYDROGEN
KINETICS KEROSENE HEAT OF FORMATION CHEMICAL PROPERTIES
Accessibility: Unclassified; No Copyright; Unlimited; Publicly
available;
Updated/Added to NTRS: 2005-10-10

Pat

"Paul F. Dietz" wrote in
:

Herb Schaltegger wrote:
On Sat, 10 Mar 2007 00:54:39 -0600, Henry Spencer wrote
(in article ):

Also note that the mass-ratio disparity is not as large as you'd think,
because a dense-propellant SSTO needs less delta-V to reach orbit.

Um, what?

He explained that. What part of the explanation didn't you understand?

He kinda sorta explained that. The actual delta-V required to reach orbit
is purely a function of orbital mechanics and is independent of propellant
density. But the rocket equation does not account for things like gravity
and drag losses, and those things do depend on propellant density. Rather
than expand the rocket equation to explicitly include those terms, the
convention is to apply a fudge factor to the delta-V term. Dense-propellant
SSTOs have lower gravity losses so they need a smaller fudge factor. (They
also typically have lower drag losses since the dense propellants allow
smaller tanks, but that's not as significant as the lower gravity losses.)

So a dense-propellant SSTO doesn't really need less delta-V to reach orbit,
but you use a smaller delta-V term when modelling one using the rocket
equation.

--
JRF

Reply-to address spam-proofed - to reply by E-mail,
check "Organization" (I am not assimilated) and
think one step ahead of IBM.

richard schumacher wrote:

He will start by dropping the pointless qualifier "unmanned".


I had a thought here... if NASA built something along these lines it
would be spread all over the US:
Engines built here, tankage built over there, electronics somewhere
else, then assembled at yet another place, and taken from there to a
launch site.
But if done commercially...since it doesn't stage, you could build it at
one place, stick the payload on it, and roll it a mile or so away and
launch it. Without staging, one of the main arguments for a seacoast
launch site vanishes, and you now have all sorts of options open to you
as to where you want your launchpad at.
For GEO you still want it as far south as possible, but for polar orbits
you can put it just about wherever you want, so why not right next to
the rocket factory?
You could make it a manned launcher, but NASA would try to get their
hands on it, and it would end up being dragged down to KSC.
If you do it for commercial and military launches, then it might have a
chance of sneaking under the NASA radar.

Pat

In article ,
Pat Flannery wrote:
...keeping the ET's changing center of mass within the
limited gimbal range of the SSMEs absolutely dictated putting the LOX tank
as far away from the SSMEs as humanly possible.

I thought the gimbaling would probably be the main problem... the thing
is already fairly unstable without getting it completely unstable by
sticking the LH2 on top and having it constantly trying to go out of
control.

As I understand it, aerodynamic stability did not end up being an
overwhelming concern, because the SRBs are such a large part of the
stack's mass for most of the atmospheric flight. (Their nozzle gimbals
also handle most of the control then.) By the time they depart, the air
is thinning out fast and the importance of the aerodynamics is dwindling.

The main issue was reducing movement of the center of mass and keeping it
within SSME gimbaling range, so that the off-center SSMEs can always have
their thrust vector pointed toward it. The SRBs and the ET LOX tank
totally dominate the stack's center of mass for most of the flight, so
their centers *had* to be in roughly the same direction, as seen from the
SSMEs.

Was there any particular reason the LOX ended up on top in the Atlas?
Gimbaling limits again?

Gimbal range wasn't an issue with the engines directly under the tanks.
Reducing aerodynamic instability, to go easier on the control system,
might have been. I don't think I've ever seen a discussion of exactly why
Atlas has the LOX on top.

I note that both Jupiter and Thor had it on the bottom.

Designers with different priorities, probably. The Atlas's balloon tanks
minimize the structural-mass penalty of putting the heavy LOX high up, so
the more conventional structures might be expected to encourage putting it
on the bottom. Max Hunter, who was chief engineer for Thor, in later
years was big on saving structural mass by putting the LOX right above the
engines, so that may have been the issue for Thor.
--
spsystems.net is temporarily off the air; | Henry Spencer
mail to henry at zoo.utoronto.ca instead. |

On Sat, 10 Mar 2007 14:52:12 -0600, in a place far, far away, Pat
Flannery made the phosphor on my monitor glow in
such a way as to indicate that:



richard schumacher wrote:

He will start by dropping the pointless qualifier "unmanned".


I had a thought here... if NASA built something along these lines it
would be spread all over the US:

Before or after the launch?

On Sat, 10 Mar 2007 14:47:25 -0600, in a place far, far away, "Jorge
R. Frank" made the phosphor on my monitor glow
in such a way as to indicate that:

"Paul F. Dietz" wrote in
:

Herb Schaltegger wrote:
On Sat, 10 Mar 2007 00:54:39 -0600, Henry Spencer wrote
(in article ):

Also note that the mass-ratio disparity is not as large as you'd think,
because a dense-propellant SSTO needs less delta-V to reach orbit.

Um, what?

He explained that. What part of the explanation didn't you understand?

He kinda sorta explained that. The actual delta-V required to reach orbit
is purely a function of orbital mechanics and is independent of propellant
density. But the rocket equation does not account for things like gravity
and drag losses, and those things do depend on propellant density. Rather
than expand the rocket equation to explicitly include those terms, the
convention is to apply a fudge factor to the delta-V term. Dense-propellant
SSTOs have lower gravity losses so they need a smaller fudge factor. (They
also typically have lower drag losses since the dense propellants allow
smaller tanks, but that's not as significant as the lower gravity losses.)

So a dense-propellant SSTO doesn't really need less delta-V to reach orbit,
but you use a smaller delta-V term when modelling one using the rocket
equation.

It depends on how you use the terms. *Ideal* delta V is the same for
all vehicles from a given location, but "actual" is not.

On Sat, 10 Mar 2007 14:47:25 -0600, Jorge R. Frank wrote
(in article ):

"Paul F. Dietz" wrote in
:

Herb Schaltegger wrote:
On Sat, 10 Mar 2007 00:54:39 -0600, Henry Spencer wrote
(in article ):

Also note that the mass-ratio disparity is not as large as you'd think,
because a dense-propellant SSTO needs less delta-V to reach orbit.

Um, what?

He explained that. What part of the explanation didn't you understand?

He didn't really. See Jorge's explanation below.

He kinda sorta explained that. The actual delta-V required to reach orbit
is purely a function of orbital mechanics and is independent of propellant
density.

Yep. Hence my comment/question.

But the rocket equation does not account for things like gravity
and drag losses, and those things do depend on propellant density. Rather
than expand the rocket equation to explicitly include those terms, the
convention is to apply a fudge factor to the delta-V term. Dense-propellant
SSTOs have lower gravity losses so they need a smaller fudge factor. (They
also typically have lower drag losses since the dense propellants allow
smaller tanks, but that's not as significant as the lower gravity losses.)

So a dense-propellant SSTO doesn't really need less delta-V to reach orbit,
but you use a smaller delta-V term when modelling one using the rocket
equation.

Thank you, Jorge.

--
You can run on for a long time,
Sooner or later, God'll cut you down.
~Johnny Cash

Pat Flannery wrote:

:Without staging, one of the main arguments for a seacoast
:launch site vanishes, and you now have all sorts of options open to you
:as to where you want your launchpad at.

Actually, no. You still have all the usual downrange range safety
issues. This is why launches over water are preferred. If something
goes wrong, you're less likely to hit something you don't own.

The real answer to this is to get reliability provably up to the point
where people think no more of the 'range safety' issues for a launch
than they do for those connected with an airplane taking off.

--
"Rule Number One for Slayers - Don't die."
-- Buffy, the Vampire Slayer

Rand Simberg wrote:
Before or after the launch?


You probably want to launch it in some fairly desolate area to be on the safe side.
But you do get the advantage of getting all your engines firing on the pad prior to liftoff, so you can be sure they are operating correctly.
Of course as soon as an engineer sees that they are going to suggest the Atlas engine drop, then suggest dropping the associated tankage as well as the engines, then since that will come to Earth somewhere, it should be on a seacoast... :-D

Pat