Two rockets and a ramjet to orbit

Here is my latest proposal on how to get to space cheaply. It consists
of three stages, the first of which is a "platform" powered by a CH4-
LOX rocket; the second of which has a CH4-LOX rocket and a simple
ramjet (no scramjets needed); and the third stage of which is an LH2-
LOX rocket.

I assumed CH4-LOX rockets have an exhaust velocity of 3000 m/s and LH2-
LOX rockets have an exhaust velocity of 4400 m/s, operating at high
altitude which is nearly a vacuum.

Stage 1 has a delta-V of 2000 m/s, 1000 m/s of which is used to
accelerate and 1000 m/s of which is used to decelerate. It is a flying
platform, that rises to perhaps 100 meters and then accelerates
forward with 2.5 g's net acceleration imparted; it actually
accelerates at 2 to 2.29 g's forward and 1.0 g's vertically. This
platform is powered by many CH4-LOX rockets. It only reaches 858 m/s
horizontal speed, then releases its payload and decelerates.

The platform was originally going to be a rocket-sled on rails, but
the long track (nearly 100 km) would be more expensive than increasing
the mass ratio so it can fly on its own power.

So stage 1 is 50% propellant, 25% structure, and 25% payload. Its
payload is Stage-2, which is 1,000,000 kg. Thus stage 1 is 4,000,000
kg and this includes 2,000,000 kg of CH4-LOX.

Once it reaches Mach 2.5 (at only 0.1 km altitude), which is 858 m/s,
Stage-2 separates and activates its RAMJET. The Ramjet breathes air
and is powered by hydrogen. It flies from 0.1 km and 858 m/s up to 40
km altitude (131,000 ft) where it reaches Mach 5 (1715 m/s). All this
is done with only 250,000 kg ramjet propellant (25% the mass of stage
2 is ramjet propellant).

After burning ramjet propellant, Stage 2 now has a mass of 750,000 kg.
280,000 kg is CH4-LOX which accelerates stage-2 by 1375 m/s. Stage-2
is 33% structural mass, 14% payload, 28% CH4-LOX, and 25% LH2 (ramjet
propellant).

Stage-3 then separates. This stage is 53% LH2-LOX propellant, 17%
structure, and 30% payload.

So how did we do? On the "launchpad" the whole vehicle has a mass of
4,000,000 kg. But this includes the flying platform, which doesn't get
very far; it moves perhaps 100 km away horizontally but never rises
more than 100 m vertically. The total payload is 42,000 kg. This is
only 1.05% the total vehicle mass, which is somewhat low, but the
system is highly reusable.

How would it really fly? Imagine a platform in Mojave that takes off
and in 60 seconds reaches Mach 2.5. Then Stage-2 (1,000,000 kg) takes
off from the platform, powered by Ramjets. The platform then
decelerates and lands about 100 km away horizontally. Then Stage-2
blasts off under Ramjet power until it reaches Mach 5, then it ignites
its CH4-LOX rockets, reaching the equivalent (at sea level) of Mach 9.
Then Stage-3 separates off (at high altitude) and fires its LH2-LOX
rocket. Now Stage 2 glides to a landing in the Pacific Ocean where
boats retrieve it (it now has a mass of 330,000 kg). Stage 3's LH2-LOX
rockets fire it to a 400 km circular orbit!

The idea is the WHOLE SYSTEM is supposed to be reusable. I realize
1.05% is a low net payload fraction, but I think that being reusable
would lower the cost.

I want it to cost $10 million per flight, plus $1.5 billion start-up
costs (starting with nothing at all). Flying 100 flights a year and
carrying 20 tourists at $950,000 per ticket, one reaches financial
break-even in 3 years. Is $10 million per flight a bit optimistic? Is
100 flights a year per ship optimistic too?

I hope you all enjoyed this post, it has my usual optimism. I assumed
the Ramjet would operate at 50% efficiency on average to come up with
these figures.

On Sun, 10 May 2009 08:46:20 -0700 (PDT), Willow
wrote:

After burning ramjet propellant, Stage 2 now has a mass of 750,000 kg.
280,000 kg is CH4-LOX which accelerates stage-2 by 1375 m/s. Stage-2
is 33% structural mass, 14% payload, 28% CH4-LOX, and 25% LH2 (ramjet
propellant).

I'm sure that others here will contest your figures... but I'm
wondering why you decided to use a hydrogen-fuelled ramjet in the same
stage as a methane-fuelled rocket? Seems to be an unnecessary
complexity, when you can just increase the size of your CH4 tank and
fuel the rocket and the ramjet from the same source...

Willow wrote:
Here is my latest proposal on how to get to space cheaply. It consists
of three stages, the first of which is a "platform" powered by a CH4-
LOX rocket; the second of which has a CH4-LOX rocket and a simple
ramjet (no scramjets needed); and the third stage of which is an LH2-
LOX rocket.


Faster you light the ramjet(s) after liftoff, the better - as it's going
to be a lot more effective at the higher atmospheric pressures at
low/medium altitudes than higher in the atmosphere in regards to saving
oxidizer weight; see "Gnom":
http://www.astronautix.com/lvs/gnom.htm

Pat

On May 10, 2:40 pm, Pat Flannery wrote:
[snip]
Faster you light the ramjet(s) after liftoff, the better - as it's going
to be a lot more effective at the higher atmospheric pressures at
low/medium altitudes than higher in the atmosphere in regards to saving
oxidizer weight; see "Gnom":http://www.astronautix.com/lvs/gnom.htm

Pat

Hey, thanks for the link! After a little searching I found out someone
else proposed a similar idea he
http://pdf.aiaa.org/preview/1995/PV1995_6103.pdf

At least we know now that the idea has merit! In response to an
earlier post, I've revised the second stage to be all hydrogen-fueled,
like the second stage. The first stage is still CH4-LOX however
because it is denser and delivers greater thrust per unit power.

The first stage is now the same as before but I made changes so it now
reads like this:

Stage-2: Ramjet's have an effective exhaust velocity of 7800 m/s
(burning LH2). We fly to 40 km altitude, Mach 5, from Mach 2.5 at 100
meters altitude!
Delta-V is 3300 m/s including drag (real change in velocity realized
is only 858 m/s). 3300=7800 ln R = 35% LH2 needed. This leaves us
with 65% stage-2 mass.

Then we fire the LH2-LOX rocket on Stage-2, with a delta-V of 1375 m/
s. I assume a specific impulse of 4250 m/s for LH2-LOX. So 1375=4250
ln R = 28% LH2-LOX. But 28% of 65% is 18.2%.

We now have this for Stage-2:
) 35% LH2 (ramjet propellant)
) 18.2% LH2-LOX (rocket propellant)
) 32.8% structure
) 14% payload

Stage-3 now has a delta-V of 3330 m/s to reach orbit. 3330=4250 ln R
= 55% propellant. Stage-3:
) 55% LH2-LOX (rocket propellant)
) 15% structure
) 30% payload

Stage-3 now uses an external tank which is separated before orbital
insertion, to contain the 55% LH2-LOX propellant. Thus the third stage
is reusable, as it is 1/3 structure and 2/3 payload. The structural
mass now includes wings for reentry and glide back to the runway.

Willow

Willow wrote:

Hey, thanks for the link! After a little searching I found out someone
else proposed a similar idea he
http://pdf.aiaa.org/preview/1995/PV1995_6103.pdf

Oh, wait till you see _this_ thing:
http://www.astronautix.com/graphics/n/n1mok2.jpg
http://www.astronautix.com/stages/n1mok.htm
And the Soviets weren't the only ones to come up with a giant
ramjet-assisted booster to get things into space; we made that N-1 look
minor with one of the "Nova" studies for a follow-on to the Saturn V;
check out the Martin-Marietta R10R2 design from September of 1963:
http://www.astronautix.com/graphics/n/novammad.gif
From he http://www.astronautix.com/lvs/nova.htm:

"Expendable version of most exotic Martin Nova variant; single stage to
orbit, 30 cd module air augmented engines in annular shroud. Operational
date would have been October 1980.
Manufacturer: Martin. LEO Payload: 596,000 kg (1,313,000 lb). to: 185 km
Orbit. Liftoff Thrust: 140,540.000 kN (31,594,640 lbf). Total Mass:
9,189,200 kg (20,258,700 lb). Core Diameter: 21.30 m (69.80 ft). Total
Length: 59.00 m (193.00 ft).

* Stage1: 1 x Nova MM R10E-2. Gross Mass: 8,474,000 kg (18,681,000 lb).
Empty Mass: 1,188,000 kg (2,619,000 lb). Motor: 40 x CD Module. Thrust
(vac): 140,539.000 kN (31,594,424 lbf). Isp: 620 sec. Burn time: 310
sec. Length: 49.10 m (161.00 ft). Diameter: 21.30 m (69.80 ft).
Propellants: Air/Lox/LH2.

Version:

Nova MM R10R-2. Status: Study 1963. Other Designations: Renova.

Reusable version of most exotic Martin Nova variant; single stage to
orbit, 30 cd module air augmented engines in annular shroud. Operational
date would have been October 1980.

Martin Marietta’s "Renova" design featured rocket engines enclosed in an
air duct equipped with adjustable inlets. The conical payload fairing
would serve as an inlet spike during the ascent through the atmosphere.
The air (which is heated by the rocket exhaust) would contrubute
additional thrust as it expands past the plug-shaped afterbody. The
inlets would be closed after leaving the atmosphere.

Manufacturer: Martin. LEO Payload: 423,000 kg (932,000 lb). to: 185 km
Orbit. Liftoff Thrust: 140,540.000 kN (31,594,640 lbf). Total Mass:
9,154,600 kg (20,182,400 lb). Core Diameter: 21.30 m (69.80 ft). Total
Length: 59.00 m (193.00 ft).

* Stage1: 1 x Nova MM R10R-2. Gross Mass: 8,647,000 kg (19,063,000 lb).
Empty Mass: 1,361,000 kg (3,000,000 lb). Motor: 40 x CD Module. Thrust
(vac): 140,539.000 kN (31,594,424 lbf). Isp: 620 sec. Burn time: 310
sec. Length: 49.30 m (161.70 ft). Diameter: 21.30 m (69.80 ft).
Propellants: Air/Lox/LH2.

At least we know now that the idea has merit! In response to an
earlier post, I've revised the second stage to be all hydrogen-fueled,
like the second stage. The first stage is still CH4-LOX however
because it is denser and delivers greater thrust per unit power.


That actually might work; you'd have to weigh off the total tankage
weight of doing it with methane rather than kerosene versus its specific
impulse and total fuel cost. Liquid methane is going to cause problems
in keeping it cold before launch or making pressure-resistant fuel
tankage if you wanted to do it at normal surface temperatures - which
kerosene doesn't suffer from.
Me, I'm old-fashioned and would get the ramjet stage up to ignition
speed via a giant honking solid propellant stage. :-)
Minimal ramjet ignition speed is around 300-400 mph, and without even
going into the aerodynamic design of a scramjet, you should be able to
get the ramjet stage up to around Mach 4-5 using 1950's technology and
design - which used subsonic combustion due to dual-shock intake cones
and good internal design to reduce the intake air's velocity to subsonic
speeds once it hit the flame holder ignition point deep inside the engine.

Pat

Willow wrote:

Here is my latest proposal on how to get to space cheaply. It consists
of three stages, the first of which is a "platform" powered by a CH4-
LOX rocket; the second of which has a CH4-LOX rocket and a simple
ramjet (no scramjets needed); and the third stage of which is an LH2-
LOX rocket.

Making things more complicated rarely makes them cheaper.

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

http://derekl1963.livejournal.com/

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

Derek Lyons wrote:

Making things more complicated rarely makes them cheaper.




Assuming that stage one only needs to get stage two up to ramjet
ignition speed, stage one should be pretty easy to make recoverable.
Assuming that stage two shuts down at around 100,000 feet and Mach 5,
when its ramjets stop generating thrust, then it can descend back into
the atmosphere, and should be able to fly back to the liftoff point via
using its ramjets with only a little extra fuel - as then it would then
be in a power-assisted glide.
At speeds of over Mach 3 it doesn't even need wings; "body lift" alone
will keep it generating enough lift to keep it airborne and return it to
its launch site, particularly given its low size-to-weight ratio from
the expendature of its fuel.
Empty, it might well be able to glide-land like a lifting body at a
fairly low speed.

Pat

Pat Flannery wrote:

Derek Lyons wrote:

Making things more complicated rarely makes them cheaper.

Assuming that stage one only needs to get stage two up to ramjet
ignition speed, stage one should be pretty easy to make recoverable.
Assuming that stage two shuts down at around 100,000 feet and Mach 5,
when its ramjets stop generating thrust, then it can descend back into
the atmosphere, and should be able to fly back to the liftoff point via
using its ramjets with only a little extra fuel - as then it would then
be in a power-assisted glide.
At speeds of over Mach 3 it doesn't even need wings; "body lift" alone
will keep it generating enough lift to keep it airborne and return it to
its launch site, particularly given its low size-to-weight ratio from
the expendature of its fuel.
Empty, it might well be able to glide-land like a lifting body at a
fairly low speed.

Mostly true (there's some quibbling on details), but having exactly
zip point nada to do with my statement. Making things more
complicated rarely makes thing cheaper.

Reusability isn't a magic wand that makes things cheaper
automagically.

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

http://derekl1963.livejournal.com/

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

On May 11, 6:44 am, (Derek Lyons) wrote:
Pat Flannery wrote:

Derek Lyons wrote:

Making things more complicated rarely makes them cheaper.

[snip]
Mostly true (there's some quibbling on details), but having exactly
zip point nada to do with my statement. Making things more
complicated rarely makes thing cheaper.

---
Here is the latest version, which gives a little more payload to
orbit:
Stage 1: 4,000,000 kg.
Vex=2600 m/s. 50% propellant, 30% payload, 20% structure.
Releases a 1,200,000 kg second stage at Mach 2.5.
Powered by kerosene and oxygen.
Stage 2:
Ramjet: Vex=8000 m/s. Flies from mach 2.5 to 5.5 at 40 km.
33% is LH2 (ramjet propellant), 3000 m/s is imparted delta-V.
Realized delta-V is 1915 m/s.
Rocket: 27.5% LH2-LOX (rocket propellant), Vex=4250 m/s.
delta-V is 2245.666 m/s.
20% structure. 19.5% payload.
Stage 3:
Rocket: 60.5756% propellant, 19.4244% payload. 20% structure.
Net payload: 45,453.17 kg to LEO. This is 1.1363% Stage-1 mass.
20% structure.
---

I switched the first stage to kerosene from methane and revised the
estimates. There is now 45,453.17 kg payload (we'll say 45,000 kg)
from a 4,000,000 kg rocket. The first stage doesn't get very far
however and is easily recovered.

I think this is a simple design. In relation to what do you say it's
complicated? Also it is supposed to be 100% reusable, which means the
spacecraft need not be paid for every flight. The ship is supposed to
cost $400 million (estimated by taking the weight of 4,000,000 kg and
multiplying by $100/kg). If it flew just once, then we would get
45,000 kg for $400 million or $8889/kg. On the other hand if a mission
costs $10 million, $2.5 million of which goes towards paying for the
$400 million ship (which only has to be paid for once every life
cycle), then after 160 flights we have the ship paid off and we've
spent $1.6 billion lifting 7,200,000 kg.

This means that if the ship is good for 160 flights, then we get $222/
kg.

Each mission needs:
) 2,000,000 kg kerosene
) 396,000 kg LH2 (ramjet fuel)
) 330,000 kg LH2-LOX (second-stage rocket)
) 141,747 kg LH2-LOX (third-stage rocket)

Estimating $1.00/kg for all propellants, we need $2,867,747 per
flight. We'll leave $5.0 million for the actual cost of propellant for
a flight. We set aside $2.5 million for paying for the spacecraft.
This leaves $2.5 million per flight for operations.

[snip]
Reusability isn't a magic wand that makes things cheaper
automagically.

The above numbers are simplified, but not bogus. I really think re-
usability is the trick that can get costs down from about $9,000 per
kg to about $250 per kg. I also think a low-flying rocket followed by
a ramjet/rocket stage and finally an all-rocket stage is the way to go.

Willow wrote:

On May 11, 6:44 am, (Derek Lyons) wrote:
Pat Flannery wrote:

Derek Lyons wrote:

Making things more complicated rarely makes them cheaper.

[snip]
Mostly true (there's some quibbling on details), but having exactly
zip point nada to do with my statement. Making things more
complicated rarely makes thing cheaper.

---
Here is the latest version, which gives a little more payload to
orbit:

(sigh) Do you think repeating yourself, especially when it fails to
adress the point, actually accomplishes anything?

I think this is a simple design. In relation to what do you say it's
complicated?

Try comparing it to virtually any other rocket ever flown. Then
consider why NASA abandoned the fully reusable (and quite complex)
schemes as the Shuttle evolved towards it's final configuration.

Reusability isn't a magic wand that makes things cheaper
automagically.

The above numbers are simplified, but not bogus.

What then are they based on? You've already had to be corrected on
operational issues (methane fuel) which tells me you aren't really
paying attention to to the whole package, which makes the balance
suspect.

I really think re-usability is the trick that can get costs down from
about $9,000 per kg to about $250 per kg.

If current launches were expensive mostly because we don't reuse the
rockets, you'd be onto something.

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

http://derekl1963.livejournal.com/

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