Vapor as rocket propellant and coolant

The cheapest rocket launcher that I can imagine has a
pressure-fed, liquid fuel engine and self pressurized
propellant tanks. The self pressurized tanks are
pressurized with the vapor pressure of the propellants
instead of the helium gas.

Conventional pump-fed engines are cooled with liquid
propellants. Cooling the self pressurized, pressure-fed
engine with the liquid propellants may be troublesome
because the liquid may boil in the cooling pipes.
Raising the pressure and temperature above the critical
point is not practicable because the tanks would be too
heavy and the propellant density would be too low.

An interesting and probably unexplored option is
vaporizing all the propellants in the tanks and using
the vapors as propellants and coolants. If the
pressure-fed engine is an integral part of the tank,
the direct contact between the engine and the tank
will provide additional cooling. Actually, the amount
of heat needed to boil all the propellant is so large
that the direct contact cooling may suffice to cool
the engine.

How easy would it be to mix gaseous oxygen and gaseous
methane in the combustion chamber? It seems that the
gaseous propellants swirling in the chamber would be
easy to mix because they would move faster than the
liquid propellants. If this is true, it means that the
gaseous propellants do not need very narrow injection
nozzles, which are difficult to fabricate. Furthermore,
the pressure drop in a wide gaseous injection nozzle
is smaller than the pressure drop in the narrow, liquid
injection nozzle.

It seems that the pressure-fed, gaseous propellant
engine is the winner -- no other engine is cheaper or
more reusable. The tanks would be as heavy as the engine
and strong enough to survive the reentry. An all aluminum
rocket would have the mass ratio of about 6. A rocket
made of titanium tank and aluminum engine would be more
difficult to fabricate, it but would have the mass ratio
of about 8.

The heat transfer coefficient of cooling gas is
enhanced by roughness of the surface. For helium gas
flowing along a smooth surface the coefficient
is only 5,000 W/(m^2K), but rough surface raises
the coefficient to 50,000 W/(m^2K). If the temperature
difference is 400 K, the latter coefficient
corresponds with heat flux of 2 MW/m^2. (Liquid water
can transfer heat flux of up to 20 MW/m^2.) Source:
http://www.ubka.uni-karlsruhe.de/cgi...page=41#page41

Thermal flux of gaseous oxygen along a very rough
surface is about 0.33 MW/m^2 at 400 K temperature
difference. Thermal conductivity of gaseous methane
is about 0.4 MW/m^2 at 400 K temperature difference.
These values are one order of magnitude too low!
Transpiration cooling is too far fetched, so the
engine must be cooled with liquefied rather than
gaseous propellants.

The critical point of oxygen is at 5.08 MPa and 154.7 K.
Its density is 1.14 g/ml at the boiling point and
0.427 g/ml at the critical point.

The critical point of methane is at 4.60 MPa and 190.6 K.
Its density is 0.423 at the boiling point and
0.162 g/ml at the critical point.

If we cool the engine only with liquids, their
temperatures must not exceed the critical temperatures.
Low tank pressure is desirable because it reduces
the mass of vapor left in the tank when it dries up.
this means that low tank pressure improves mass
ratio.

Suppose that tank pressure is 1 MPa and its final
temperature is almost critical. How much propellant
mass will be left in the tanks when they dry up?
I do not have gas tables so I do not know. My wild
guess is seven percent of the original mass. The
residual vapor may be used for cooling the rocket
during its reentry.

Andrew Nowicki wrote in message An interesting and probably unexplored option is...

How easy would it be to mix gaseous oxygen and gaseous
methane in the combustion chamber? It seems that the
gaseous propellants swirling in the chamber would be
easy to mix because they would move faster than the
liquid propellants. If this is true, it means that the
gaseous propellants do not need very narrow injection
nozzles, which are difficult to fabricate. Furthermore,
the pressure drop in a wide gaseous injection nozzle
is smaller than the pressure drop in the narrow, liquid
injection nozzle.

It seems that the pressure-fed, gaseous propellant
engine is the winner -- no other engine is cheaper or
more reusable. rocket would have the mass ratio of about 6. A rocket
made of titanium tank and aluminum engine would be more
difficult to fabricate, it but would have the mass ratio
of about 8.

For a planned "microlauncher" (just starting www.microlaunchers.com ),
I am considering NH3 vapor as regenerative coolant for a low pressure
3rd stage engine. NH3 gas decomposes endothermically and should be a
good coolant for a low pressure (Pc about 1-2 atm.) engine made of
electroformed nickel.

An all electroformed stage can have a very high mass ratio--over 20 if
the pressures are low enough.

Charles Pooley

I think some existing engines have two-phase flow for the
coolant.

Hydrogen-burning pump engines do not -- the pump raises
the hydrogen above its critical point (-400.3F, 187.5 psi)
and it stays there until it burns. Note that the fluid
in these systems essentially behaves as a gas, as in your
lower pressure idea.

I like the idea of a methane-oxygen vapor-phase engine
embedded in the fuel tank. I like it because the flame
inside the engine looks more like a bunsen burner and less
like a traditional rocket engine.

There are some engineering challenges though. Here I've
assumed tank pressure of 1.5 MPa (225 psi) and bulk launch
temperature of about 85 K.

Fuel:

The following discussion uses methane. I looked at
propane, but hot propane cannot vaporize all the oxygen
necessary to burn it without autoigniting first.

Propane mix: C3H8 + 7O2 is 44g to 224g (1-to-5.1)
Propane drops 637 K to boil oxygen
Methane mix: CH4 + 2O2 is 16g to 64g (1-to-4)
Methane drops 390 K to boil oxygen
Propane:
0.075 kJ/mol-K = 1.7045 kJ/kg-K
Methane:
0.035 kJ/mol-K = 2.1875 kJ/kg-K
vaporization: 510 kJ/kg
Oxygen:
0.029 kJ/mol-K = 0.9062 kJ/kg-K
vaporization: 213 kJ/kg

Nozzle / combustion chamber cooling:

I've assumed here that just the methane is used as a
coolant, and LOX is injected into the hot methane before
ignition. This avoids dealing with hot oxygen and two
coolant streams. Some sort of tank-pressure loop will
need to supply heat to the oxygen tank to make it self-
pressurizing.

Tank pressurization:

Self-pressurizing with 32g and 16g molecular weight gases
is heavy. People don't use helium for no reason.
Here I've assuming CF tanks with a working strain of
200MPa and a density of 1.57 g/cc.

Spherical carbon fiber tanks:11.8e-6 kg/pressure-volume
Spherical aluminum tanks: 46.1e-6 kg/pressure-volume
Spherical titanium tanks: 14.0e-6 kg/pressure-volume
Methane at 500K: 3.9e-6 kg/pressure-volume
Oxygen at 500K: 7.8e-6 kg/pressure-volume
Helium at 500K: 1.0e-6 kg/pressure-volume

1.5 MPa CF tanks with self-pressurized oxygen and methane
will have a burnout mass of (just the tanks and
pressurant, nevermind the rest of the rocket) of about
28.2 kg/m^3, versus a propellant load of about 1000
kg/m^3. The mass ratio of just the tank then is 35.
This will get much worse once the engine is inserted in
the tank, and will get somewhat worse for cylindrical
tanks.

Pressurization control:

You have only crude control over the heat input to the
tank, so you'll have only crude control over the boiling
rate. To regulate the tank pressure, you'll have a
pressure regulator bleed pressurization gas off the top
directly into the rocket injector. Most of the
propellant will boil inside a low pressure shroud inside
the tank around the nozzle and combustion chamber. Some
propellant will boil on the back side of this shroud,
and this boiling will form the pressurization for the
tank. I would retard the mixing of this in-tank boiling
propellant with the rest of the fluid with yet another
shroud, this one all the way up the interior of the tank
to the gas surface. This way the bulk of the propellant
can be subcooled for improved density and engine cooling.

If bleeding into the injector makes the thrust too
variable, you can desensitize the motor (and throw away
some Isp) by bleeding pressurant gas into the nozzle.

Pressurization just after launch is going to be
inconsistant since the gas phase portion is expanding a
lot and the heat path from engine to boiling propellant
hasn't reached equilibrium yet.

On the ground, I'd pressurize the tanks with helium and
chill them with an LN2 loop unil just before launch.

Propellant mixtu

You need some way to regulate the mixture going into the
combustion chamber. Maybe you can do this with an orifice
in the cold LOX line directly after the pickup point,
where the transition into two-phase will have progressed
the least.

Mixing:

If you can inject LOX and warm (800K) gaseous methane,
the propellant velocities will be very different and give
a lot of mixing. Basically, each drop of LOX is a little
snowball in hell, with the methanee blasting the surface
liquid off. The methane will lose 390 K evaporating the
LOX, and more if the LOX is subcooled.

I suspect that if you delay burning until after you've
mixed, the burn will be more stable.

There is about 330 degrees K of margin between the
temperature at which methane will autoignite in oxygen
(868 K in air) and the temperature at which the methane
will liquify before vaporizing all the LOX (540 K).
Running a rich mixture will give more margin (and better
Isp too).

Engine size:

The gas flow is physically large. A 200 kN methane-oxygen
engine would need to burn 60 kg of mix per second. The
12 kg of methane per second, at 1.5 MPa and 800 K will
be 3.3 m^3/sec. Across a 30 cm orifice that's 46 m/s.
At this velocity there will be large drag pressure losses.

Startup problems:

A lightweight engine designed for gas-phase combustion
will explode if you fill it with LOX and liquid methane
and ignite it. So at startup you'd have to release small
amounts of gaseous oxygen and methane into the combustion
chamber to get it going. If you want to avoid propellant
valves that can throttle the engine (which sound hard to
get right), you'll want to just bang open the propellant
valves.

Before the propellant flow starts, how do you keep the
engine from overheating? When the propellant flow starts,
how do you ensure that the propellant vaporizes on the
way to the engine?

I think the answer is at the tank pick-up points, which
are at the bottom of the nozzle. [I think of the nozzle
rim as being slightly pointy, having two points which
extend lower than the rest.] At engine startup,
propellants are supplied from the ground in cold gas
form. The gas pressure is ramped up until it's just a
little lower than the tank pressure, at which point the
onboard valves are banged open, the external supply
is closed, and the holddown bolts blow to allow takeoff.
Restarts don't happen.

Thanks -- this was a fun little exercise.

Andrew Nowicki wrote in message ...
The cheapest rocket launcher that I can imagine has a
pressure-fed, liquid fuel engine and self pressurized
propellant tanks. The self pressurized tanks are
pressurized with the vapor pressure of the propellants
instead of the helium gas.

Are you sure the cost of large, high-strength fuel tanks
and the associated performance penalties are cheaper than
a low-cost turbopump? Fabrication of large, lightweight
pressure vessels can be a headache, and pressure-fed
rockets tend to call for strong alloys that are even
harder to weld.

Conventional pump-fed engines are cooled with liquid
propellants. Cooling the self pressurized, pressure-fed
engine with the liquid propellants may be troublesome
because the liquid may boil in the cooling pipes.

I got the impression that the coolant boiled quite
frequently - do SSMEs keep their hydrogen coolant liquid
throughout the engine?

It seems that the pressure-fed, gaseous propellant
engine is the winner -- no other engine is cheaper or
more reusable.

That depends on a lot of factors outside of just how
easy it is to manufacture a single engine. Mass-production
of a turbopump-fed engine may result in an engine less
costly per unit than a pressure-fed engine that's built
once per year.

The tanks would be as heavy as the engine
and strong enough to survive the reentry.

Strong enough, or heat resistant enough?

Mike Miller, Materials Engineer

Charles Pooley wrote:

For a planned "microlauncher" (just starting www.microlaunchers.com ),
I am considering NH3 vapor as regenerative coolant for a low pressure
3rd stage engine. NH3 gas decomposes endothermically...

Above 700 degrees Celsius.

...and should be a good coolant for a low pressure
(Pc about 1-2 atm.) engine made of electroformed nickel.

Is nickel strong enough?
Does it react with ammonia above 700 degrees Celsius?

Iain McClatchie wrote:

There are some engineering challenges though. Here I've
assumed tank pressure of 1.5 MPa (225 psi) and bulk launch
temperature of about 85 K.

Makes sense.

Methane:
0.035 kJ/mol-K = 2.1875 kJ/kg-K
vaporization: 510 kJ/kg
Oxygen:
0.029 kJ/mol-K = 0.9062 kJ/kg-K
vaporization: 213 kJ/kg

Latent heat of vaporization depends on the pressure.

Nozzle / combustion chamber cooling:

I've assumed here that just the methane is used as a
coolant, and LOX is injected into the hot methane before
ignition. This avoids dealing with hot oxygen and two
coolant streams.

OK

Some sort of tank-pressure loop will need to supply
heat to the oxygen tank to make it self- pressurizing.

I like the idea of physical contact between the oxygen
tank and the methane tank. Fans inside both tanks churn
the propellants.

Tank pressurization:

Self-pressurizing with 32g and 16g molecular weight gases
is heavy. People don't use helium for no reason.
Here I've assuming CF tanks with a working strain of
200MPa and a density of 1.57 g/cc.

Aluminum-lithium 1460 is weldable. Its yield tensile
strength is 470 MPa at 293 K and 560 MPa at 20 K.
http://www.matweb.com/search/Specifi...bassnum=MA146B

1.5 MPa CF tanks with self-pressurized oxygen and methane
will have a burnout mass of (just the tanks and
pressurant, nevermind the rest of the rocket) of about
28.2 kg/m^3, versus a propellant load of about 1000
kg/m^3.

The overall density of 3.5 kg of liquid oxygen and 1 kg
of liquid methane is 828 kg/m^3 at 90K, and 890 kg/m^3
at 64K.
http://www.dunnspace.com/alternate_ssto_propellants.htm

How did you get the 28.2 kg/m^3? Do you have the gas
tables for oxygen and methane?

The mass ratio of just the tank then is 35.

Almost too good to be true!

Pressurization control:

You have only crude control over the heat input to the
tank...

Fans inside both tanks churn the propellants and thus
increase the heat flux between the engine and the
propellants.

On the ground, I'd pressurize the tanks with helium and
chill them with an LN2 loop unil just before launch.

Good idea!

Propellant mixtu

You need some way to regulate the mixture going into the
combustion chamber. Maybe you can do this with an orifice
in the cold LOX line directly after the pickup point,
where the transition into two-phase will have progressed
the least.

Do you think that cryogenic valves would be too expensive?

Mixing:

If you can inject LOX and warm (800K) gaseous methane,
the propellant velocities will be very different and give
a lot of mixing. Basically, each drop of LOX is a little
snowball in hell, with the methanee blasting the surface
liquid off. The methane will lose 390 K evaporating the
LOX, and more if the LOX is subcooled.

OK

Engine size:

The gas flow is physically large. A 200 kN methane-oxygen
engine would need to burn 60 kg of mix per second. The
12 kg of methane per second, at 1.5 MPa and 800 K will
be 3.3 m^3/sec. Across a 30 cm orifice that's 46 m/s.
At this velocity there will be large drag pressure losses.

You exaggerate. 46 m/s is close to average car speed on
a highway. 1.5 MPa equals 153 meters of water head.

A lightweight engine designed for gas-phase combustion
will explode if you fill it with LOX and liquid methane
and ignite it. So at startup you'd have to release small
amounts of gaseous oxygen and methane into the combustion
chamber to get it going. If you want to avoid propellant
valves that can throttle the engine (which sound hard to
get right), you'll want to just bang open the propellant
valves.

Very good point!
The engine and tank must be warmed up to get
them going. Maybe they can get the heat from
direct contact with a lower rocket stage?
Or maybe additional liquid propellant pipes
and additional liquid injection nozzles would
be best solution of this problem?

It would also be good idea to have a snorkel-like
valve to prevent flow of liquid propellant into
the engine.

Andrew Nowicki wrote:

AN It seems that the pressure-fed, gaseous propellant
AN engine is the winner -- no other engine is cheaper or
AN more reusable.

Mike Miller wrote:

MM That depends on a lot of factors outside of just how
MM easy it is to manufacture a single engine. Mass-production
MM of a turbopump-fed engine may result in an engine less
MM costly per unit than a pressure-fed engine that's built
MM once per year.

AN The tanks would be as heavy as the engine
AN and strong enough to survive the reentry.

MM Strong enough, or heat resistant enough?

I believe that rocket launcher must be reusable to
be economical. (The Shuttle is merely salvageable.)
Suppose that the launcher has three stages.

It is easy to make a reusable first stage. A pump-fed
rocket with a flimsy tank and wings can land like
an airplane. This is how Baikal is going to work:
http://www.spacedaily.com/news/rocketscience-03j.html
Baikal has lots of expensive hardwa foldable wings,
jet engines, landing gear... A pressure-fed first stage
without wings would be cheaper and strong enough to
survive splashdown.

If the second stage is pressure-fed, it will certainly
be strong enough to survive the reentry. High temperature
is a more difficult problem. I can imagine two solutions:
- The so called engine cluster is made of a multitude
of decimeter-size engines carved in a monolithic
slab of aluminum alloy, which is also the bottom
wall of the liquid propellant tank. The slab is
used as a heat sink during reentry. It is cooled
by vapor left in the tank.
http://www.islandone.org/LEOBiblio/SPBI1018.JPG
- The second stage of a rocket launcher spins about
its axis of symmetry, which is perpendicular to
its trajectory. The spinning rocket is relatively
cool because it radiates heat in all directions.
http://www.islandone.org/LEOBiblio/SPBI1010A.GIF

A pump-fed, second stage of the rocket launcher cannot
survive reentry because its tank is too weak.

I do not know how to make a reusable third stage.

MM Are you sure the cost of large, high-strength fuel tanks
MM and the associated performance penalties are cheaper than
MM a low-cost turbopump? Fabrication of large, lightweight
MM pressure vessels can be a headache, and pressure-fed
MM rockets tend to call for strong alloys that are even
MM harder to weld.

I do not believe that a flimsy tank can be reusable.
Welding is not rocket science. If you do not have
much confidence in the strength of the weld, weld
extra sheet metal to the weak weld to make it
stronger; then run hydrostatic test.

High performance, pump-fed rocket is good choice for
the expendable third stage.

In article ,
wrote:
...NH3 gas decomposes endothermically and should be a
good coolant for a low pressure (Pc about 1-2 atm.) engine made of
electroformed nickel.

A small caution: NH3 decomposition involves transient formation of atomic
hydrogen, which makes decomposing ammonia quite corrosive to some
materials. (The early-60s nuclear-rocket guys discovered that running
ammonia through a graphite-core reactor was going to require really good
protective coatings for the graphite.) I would guess that nickel would
be okay, but it's an issue to watch out for.
--
"Think outside the box -- the box isn't our friend." | Henry Spencer
-- George Herbert |

Andrew Nowicki wrote in message ...

I do not believe that a flimsy tank can be reusable.

You're probably using an overly strict definition of "flimsy." Fuel
tanks for pump-fed rockets do survive re-entry without a heat shield
(usually pancaking into some frightened farmer's field).

If you're worried about the fragility of the tankage, put some extra
framework around it. With a pump-fed engine a lightweight tank made of
a highly weldable alloy, you'll save more than enough weight to
justify the changes.

Welding is not rocket science.

Well, if you're using high-strength alloys with little tolerance
toward welding, it can be as complicated as rocket science.

Alloy segregation, grain size, heat affected zones, fusion, lack of
fusion, slag inclusion, the effects of multiple weld passes, the
problems with work piece orientation...welding is COMPLICATED,
especially when dealing with high strength alloys fit for large
pressure-fed rockets. It's so complicated that it's not yet possible
to automate the welding of large submarine hulls, which use more
forgiving alloys than the alloys considered for pressure-fed big dumb
boosters.

If you do not have much confidence in the strength
of the weld, weld extra sheet metal to the
weak weld to make it stronger; then run hydrostatic
test.

No, it's not that easy, especially not for pressure vessels that will
be loaded in tension (i.e., internally pressurized tanks). If you
leave a crack or flaw in the weld, you've basically guaranteed that
crack will grow explosively.

A sheet of extra metal just means that much heat affected zone around
the bad weld, plus the original crack is still waiting to "unzip" the
rest of the weld. Hydrostatic testing is just a way to make that crack
fail on the ground rather than in the air.

By the time you're done with this "simple, robust" pressure fed
booster, you're likely to have gone through an unnecessarily
complicated and expensive manufacturing process that might eat up the
savings you're hoping for.

Mike Miller, Materials Engineer