Re: Hypersonic scramjet transports imminent.
Robert Clark
> ...
> A Mach 6.5+ transport could make transatlantic and trans continental
> U.S. flights in less than an hour, compared to 6 hour flights now. The
> financial incentive will make it likely that a commercial transport
> would be produced within a few years of the first manned prototype
> vehicle.
>
The reluctance to use airbreathing engines for part of the time to
reach orbit is due in large part to the fact that jet engines are
heavy compared to the thrust they can produce. See the list of thrust-
to-weight ratios for some engines here:
Thrust-to-weight ratio.
http://en.wikipedia.org/wiki/Thrust-to-weight_ratio#Engines
The thrust-to-weight ratio for turbojets might be only 5 to 6, where
as for rocket engines such as the space shuttle main engines might be
73 or above. A big part of this poor thrust-to-weight ratio for jets
is the complexity and weight of the compressors and turbines jet
engines have to carry:
Jet engine.
http://en.wikipedia.org/wiki/Jet_engine
However, the thrust-to-weight ratio for ramjets because of their
simplicity can be quite high:
Ramjet Performance Primer.
"There are no physical limits to the minimum weight of a ramjet other
than design and materials. The 1950's Marquardt RJ43-MA-7 had a thrust/
weight (T/W) ratio of about 40. With today's engineering and materials
that could probably be brought up to 150-200 without too much effort.
Such T/W ratios would make ramjet powered vehicles excellent
accelerators."
http://www.alt-accel.com/ramjet2.htm
Airbreathing engines need compression of the air to create high
thrust. Turbojets use compressors. Ramjets are able to get high
compression from the high velocity of the incoming air alone,
dispensing with the compressors and accompanying turbines. Then the
suggestion is to replace the compressors/turbines in turbojets with
other means to achieve this high compression. One method that has been
tested is the ejector ramjet or rocket-based combined cycle engine,
where rocket exhaust is used to accelerate air into the intake of a
ramjet thus allowing the ramjet to operate even at zero speed.
This method still needs to use onboard oxidizer, for the rocket, to
operate. An ideal method would only use the burnable fuel to operate,
as do ramjets and turbojets. For a turbojet/ramjet intended as the
first phase of a SSTO vehicle that uses rockets at the end stage, what
might work is to use the very high pressure turbopumps that high
performance rocket engines such as the space shuttle main engines use.
Since these high pressure turbopumps are needed to be carried along
for the rocket phase anyway perhaps they can be used as well during
the airbreathing portion of the trip.
There are several ways this might be accomplished. The shuttle liquid
hydrogen turbopumps can produce 500 bars of pressure of the liquid
hydrogen with a through put of 73 kg/sec each. I'm imagining this high
pressure liquid hydrogen be directed into the intake of the jet
engine. You want to do this in a way to compress the air. One way
might be that the turbopump outlet into the jet engine be in the form
of an annular (ring) opening all around the inlet, some distance into
the inlet. This would tend to compress the air together as the liquid
comes out directed inward to the center. You also want the air to be
forced back to the rear of the engine so the liquid hydrogen would
need to be angled somewhat also backwards towards the rear.
The liquid would tend to spread out however, and for a large intake
say a meter across or more for the large supersonic turbojet inlets,
it's not certain how far the liquid would go to penetrate into the
middle portion of the air to achieve the high compression needed here
as well, not just the outer air. You want most of the air to be
compressed at least to the 20 bar range commonly seen with turbojets
in order to achieve the high thrust achieved by the means of
compressors.
Then another possibility would be to use an analogue of the ejector
ramjet compression method. This works by using supersonic exhaust from
a rocket to force the air into the intake, thus being compressed as is
the case with ramjets flying at supersonic speeds. Then what we could
do with the turbopump's output, is to use the Bernoulli principle to
convert the very high pressure into a supersonic velocity:
Bernoulli's principle.
Incompressible flow equation.
http://en.wikipedia.org/wiki/Bernoulli%27s_principle#Incompressible_flow_equation
For a streamline at constant height, (1/2)(velocity)^2 + pressure/
density = constant. With pipelines leading out of the liquid hydrogen
turbopumps of about 30 cm wide, a density of liquid hydrogen of 72 kg/
m^3, and mass flow rate of 73 kg/sec, I calculate the flow speed as 33
m/s. Then if we want to convert the pressure of 500 bar = 50,000,000
pascals to high velocity we would get a speed of 1180 m/s, about Mach
3. Then this supersonic flow could be directed into the intakes to
accelerate and thereby compress the air as is down with ejector
ramjets.
Still another possibility to get the air to flow at high speed to
induce similar compression as with a ramjet might be to ionize the air
and accelerate it by electromagnetic fields. The turbopumps use a
turbine which is a key means by which electric power is generated. The
SSME turbopumps operate at 70,000 horsepower while weighing only about
700 pounds. There is pretty high efficiency conversion of turbine
mechanical power to electrical power. However, we would need a
lightweight means of ionizing and electromagnetically accelerating the
air. A couple of possibilities for the ionization might be by using a
microwave generator or electrically charged wires running throughout
the inner volume of the intakes.
In any case some of the exhaust from the jet would have to be bled off
to run the turbopump. This might seem to reduce the performance of the
jet engine but actually quite a large proportion of the power
generated in usual jet engines is used just to run the compressors:
What is a Gas Turbine Engine and How Does it Work?
"The cycle that governs the operation of a gas turbine engine is
referred to as the Brayton constant pressure cycle. The engine
compressor typically requires about 2/3 (!) of the usable heat energy
produced in the burner to turn at maximum speed; the remaining energy
can then be used to produce thrust or mechanical power, or a
combination of the two."
http://www.turbokart.com/gasturbine.htm
To get an idea of the power we need, we'll use as a model the J58
engine which powered the SR-71 to Mach 3+. I haven't seen any numbers
on the horsepower generated by the J58 but I'll estimated it from the
1 horsepower per 2.5 pounds thrust common for turbojets:
Turbojet.
Thrust to power ratio.
http://en.wikipedia.org/wiki/Turbojet#Thrust_to_power_ratio
The J58 generated about 25,000 lbs thrust in usual turbojet mode, so a
horsepower of 10,000 hp. Note though that fuel needed to run the J58
is much less than the 73 kg/sec liquid hydrogen put out by the shuttle
turbopump. This page gives its fuel use in the usual turbojet mode as
0.9 lb/(lbf-h), i.e., .9 lbs/hour for each pound of thrust:
Pratt & Whitney J58.
Specification of J58-P4.
http://en.wikipedia.org/wiki/J58#Specification_of_J58-P4
This is 22,500 lbs/hr of fuel at 25,000 lbs thrust, or 6.25 lbs/sec,
2.8 kg/sec. This is in jet fuel. Hydrogen would give higher thrust and
indeed will use about half the fuel for the same thrust as shown in
the attached diagram of turbojet/ramjet/scramjet Isp's. So this would
be 1.4 kg/sec of hydrogen. This is 1/52nd the usual mass flow rate of
the SSME turbopump of 73 kg/sec. The power used by a turbopump is
proportional to the mass flow rate, so the power needed would be
70,000 hp/52 = 1,346 hp
This about 1/7th the power output of the J58 engine. A problem though
is whether this would supply sufficient compression for the high air
inflow of the jet. We might need to flow more fuel through the
turbopumps than is burned by the engines. But this would mean we are
running the jet engine fuel rich. However, the Isp for jet engines is
so high we could afford to run fuel rich and still have a
significantly better Isp than rockets.
Bob Clark
Robert Clark
>
> The reluctance to use airbreathing engines for part of the time to
> reach orbit is due in large part to the fact that jet engines are
> heavy compared to the thrust they can produce. See the list of thrust-
> to-weight ratios for some engines here:
>
> Thrust-to-weight ratio.http://en.wikipedia.org/wiki/Thrust-to-weight_ratio#Engines
>
> The thrust-to-weight ratio for turbojets might be only 5 to 6, where
> as for rocket engines such as the space shuttle main engines might be
> 73 or above. A big part of this poor thrust-to-weight ratio for jets
> is the complexity and weight of the compressors and turbines jet
> engines have to carry:
>
> Jet engine.http://en.wikipedia.org/wiki/Jet_engine
>
> However, the thrust-to-weight ratio for ramjets because of their
> simplicity can be quite high:
>
> Ramjet Performance Primer.
> "There are no physical limits to the minimum weight of a ramjet other
> than design and materials. The 1950's Marquardt RJ43-MA-7 had a thrust/
> weight (T/W) ratio of about 40. With today's engineering and materials
> that could probably be brought up to 150-200 without too much effort.
> Such T/W ratios would make ramjet powered vehicles excellent
> accelerators."http://www.alt-accel.com/ramjet2.htm
>
> Airbreathing engines need compression of the air to create high
> thrust. Turbojets use compressors. Ramjets are able to get high
> compression from the high velocity of the incoming air alone,
> dispensing with the compressors and accompanying turbines. Then the
> suggestion is to replace the compressors/turbines in turbojets with
> other means to achieve this high compression. One method that has been
> tested is the ejector ramjet or rocket-based combined cycle engine,
> where rocket exhaust is used to accelerate air into the intake of a
> ramjet thus allowing the ramjet to operate even at zero speed.
> This method still needs to use onboard oxidizer, for the rocket, to
> operate. An ideal method would only use the burnable fuel to operate,
> as do ramjets and turbojets.
> ...
Since some thrust of the jet engine would be bled off to
run the turbopump, what might work instead would be to have
some proportion of the exhaust from the jet engine be sent
directly through piping back toward the entrance of the
intakes. This would create a supersonic flow into the
intakes that would draw air in as is down with ramjets
moving at high speed and with the rocket ejector ramjet that
can work at zero speed. Note that since with usual turbojet
engines as much of 2/3 of the power produced just goes to
run the compressor, we could bleed off quite a bit of the
exhaust to get good compression of the large amounts of
incoming air and still achieve both high thrust and high
Isp.
For either method of using the turbopump fuel output or
the engine exhaust directed into the intakes, rather than
having it being directed in a parallel direction into the
intakes, we could also have this supersonic fuel or exhaust
flow be directed in a cross-wise direction across the
intakes. This would also draw in the air by the Venturi
effect, as was being discussed here.
We can calculate how fast the air would be brought in
using this Venturi method. Since the cross-flow creates a
near vacuum, internal pressure would be initially zero, and
the force to move the air inward will be p*A, with p the
ambient external pressure and A the cross-sectional area of
the
intake. On the other hand the force is also (mass flow
rate)*v, with v
the velocity. The mass flow rate will be r*A*v, with r the
ambient air
density. So the force is r*A*v^2. Setting the two equations
for force
equal we get: p = r*v^2, and v = sqrt(p/r).
At sea level the pressure is 100,000 pascals = 100,000
N/m^2, and the
density is about 1 kg/m^2. As an approximation I'll take
the pressure
and density decreasing to the same extent at altitude, so v
will be v
= sqrt(p/r) = sqrt(100,000/1) = 316 m/s, which is Mach 1.
This would then act similarly to a ramjet flying at Mach
1. Ramjets can be optimized to get quite good fuel
efficiency and thrust at Mach 1.
Bob Clark
Robert Clark
>...
> Since some thrust of the jet engine would be bled off to
> run the turbopump, what might work instead would be to have
> some proportion of the exhaust from the jet engine be sent
> directly through piping back toward the entrance of the
> intakes. This would create a supersonic flow into the
> intakes that would draw air in as is down with ramjets
> moving at high speed and with the rocket ejector ramjet that
> can work at zero speed. Note that since with usual turbojet
> engines as much of 2/3 of the power produced just goes to
> run the compressor, we could bleed off quite a bit of the
> exhaust to get good compression of the large amounts of
> incoming air and still achieve both high thrust and high
> Isp.
Found this after a web search:
Gas heat engine.
Patent number: 7111449
Filing date: Nov 14, 2002
Issue date: Sep 26, 2006
Inventor: David W. Stebbings
http://www.google.com/patents?id=29B6AAAAEBAJ
It describes the idea of using some of the exhaust from the nozzle to
be routed back to the inlet to induce the compression of the incoming
air, thus dispensing with the compressor/turbine of a usual jet
engine.
In the citations of the prior art, it also mentions a patent going
back to 1950 also based on this idea:
ASPIRATOR COMPRESSOR TYPE JET.
Patent number: 2502332
Filing date: Apr 12, 1945
Issue date: Mar 1950
Inventor: McCollum
http://www.google.com/patents?id=DnVWAAAAEBAJ
The "Gas heat engine" patent argues this engine is more energy
efficient than a usual turbojet as measured by the power output by the
engine compared to the power available in the fuel burned. It is also
more light weight as not requiring the compressors and turbines.
Then the question is why is it not being used for jet aircraft since
the idea goes back more than fifty years?
Perhaps the answer lies in this fact mentioned in the "Gas heat
engine" patent:
"The engine uses more fuel per unit thrust than the gas turbine
because the output power varies as Ve^2 while the thrust is
proportional to Ve."
p. 46.
Calculating from the data in the table accompanying this passage, the
Isp for the new engine is perhaps 15% worse than a comparable turbojet
engine. For jet airline industries where conserving fuel is a major
concern, this would make it less appealing to that industry. Still its
simplicity of operation and lowering of maintenance costs and its
lightweight could conceivably make up for this loss.
And as a first phase in a combined-cycle engine for a SSTO vehicle it
would still be quite useful. As shown in this image of the Isp's for
different types of engines:
http://www.lorrey.biz/images/pde-performance.gif
a hydrogen powered turbojet could have an Isp in the range of 5,000 s
to 7,000 s. Even an Isp 85% of this value would have a tremendous
advantage over a rocket engine. Then the fact that this new type of
jet dispenses with the heavy compressors/turbines of usual jet engines
makes it potentially very useful for a SSTO vehicle.
Bob Clark