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Lens Turbojet for Acceleration



 
 
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  #1  
Old July 10th 03, 11:53 AM
johnhare
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Posts: n/a
Default Lens Turbojet for Acceleration

Turbojets historically have been constantly improved in the direction
of improved specific fuel consumption (SFC vs Isp for rockets). Thrust
to weight has been third of more down the list after SFC, reliability,
maintainability, etc. For aircraft, this is proper.

For spacecraft, or their carriers, the trades are different in many cases.
T/W and wide range of mach is far more criticle than several percentage
points of SFC. Spacecraft accelerator engines can trade away certain
capabilities to achieve the required results. Subsonic flight at high
altitudes
requires compressor blading that does not get excessively low Renolds
numbers, which means they can't be small chord in the low pressure
compressor end. By controling the flight profile, the accelerator can
control the Renolds numbers on the compressor blading. Micro chord
blades can be used if they happen to optimize that way.

Also, accelerator engines for our purposes don't absolutely have to
have tens of thousands of hours between overhauls, nice but not
required. Low hundreds would work for the near future for a sufficiently
usefull engine.

In looking over schematics and hardware on existing (50s-60s) jet
engines, I am struck by the large numer and mass of parts other than
the actual blades and burners. My guess is that less than 10% of the mass
of the engines I have looked at is blading, with the rest being axels, hubs,
casing, seals, stator controls, etc. This, and the limits imposed by the
turbine inlet temperatures led me to eventually conclude that the basic
design concepts are wrong for our purposes.

In looking at a large number of airbreathing accelerator proposals,
it appears to me that the turbojet cycle is still the best if the mass
and turbine temp problems can be resolved. Most of the alternatives
seem to give away so much Isp or SFC that they become more trouble
than they are worth compared to a normal rocket engine.

So I worked out a new layout for the turbojet that addresses some
of the issues. This first description is an early simplified version for
clarity. My purpose in posting is to start getting it into the public
domain to prevent being halted by someone elses' patent again. The
Carmack open source concept has some hard business sense behind
it.

Picture a hollow globe with the latitude and longitude markings on it.
Flatten the globe along the axis until the polar diameter is a third or
so of the equatorial diameter. Starting at the poles, the hub extends from
90 degrees to 70 degrees latitude north and south. Compressor blades
pull air into the lens both north and south from 70 degrees to 10 degrees
latitude north and south. Turbine blades extend from a sealing ring 10
degrees north and south of the equator. The turbine is blocked off from
90 degrees west longitude for 270 degrees back to 0 longitude. The
exhaust is through that 90 degrees of longitude by 20 degrees of latitude.

The burner inside the lens is fed with air from all sides through the
stator,
which eliminates most cooling concerns for the main burner. The turbine
is operating on a 25 % duty cycle with some cooling during the off cycle
by air leaking through the seals. controling the leakage/cooling rate allows
the burner to operate at an equivelence ratio of 1, or stochiometric.
This eliminates the afterburner, and allows a very low compression ratio
compressor to function.

The above would not have enough compression ratio to be effective at
low speeds, although it would be very effective supersonically. By
introducing a second, contrarotating lens inside the first one, a
compression
ratio of a bit over 2 can be had with modest velocities of the rotating
lens turbocompressors. This should allow a reasonable thrust after the
turbine pressure drop. Static fuel Isp should be in the 800 to 1200 range
depending on assumptions. Supersonic fuel Isp should exceed 2,000
without much problem. Maximum mach is limited by turbine temperatures
which is limited by the compression heating of the air which cools
the turbine, the materials of the turbine blades, and the benifits of the
25% duty cycle. Mach 3 is a guess.

Structure of the blading is a continous band along longitude lines with
each strip being compressor blade-turbine blade-compressor blade.
The blade material is in a tension arch with stiffiners at 10 degrees
north and south latitude. The axel is stationary and experiences
no torque. The main load to the bearings in this configuration is
the off center loading of the turbine. Future configurations address this.
I believe composites can be used for this type structure.
Structural mass appears to be minimal, with some of my estimaes
making it comperable in T/W to a rocket engine.

More later.

John Hare



  #2  
Old July 12th 03, 10:08 AM
James Wentworth
external usenet poster
 
Posts: n/a
Default Lens Turbojet for Acceleration

I believe Pabst von Ohain built a conceptually-similar "pancake" turbojet in
the late 1930s. If I recall correctly, it was a centrifugal compressor with
a "mirror image" radial inflow turbine whose blades were on the back of the
compressor disc. Due to poor fuel-air mixing in the combustor, it wouldn't
quite run on its own (combustion continued outside the nozzle), but as long
as the electric starter motor was kept on the "pancake" turbojet did run and
produced significant thrust. -- Jason

johnhare wrote in message
om...
Turbojets historically have been constantly improved in the direction
of improved specific fuel consumption (SFC vs Isp for rockets). Thrust
to weight has been third of more down the list after SFC, reliability,
maintainability, etc. For aircraft, this is proper.

For spacecraft, or their carriers, the trades are different in many cases.
T/W and wide range of mach is far more criticle than several percentage
points of SFC. Spacecraft accelerator engines can trade away certain
capabilities to achieve the required results. Subsonic flight at high
altitudes
requires compressor blading that does not get excessively low Renolds
numbers, which means they can't be small chord in the low pressure
compressor end. By controling the flight profile, the accelerator can
control the Renolds numbers on the compressor blading. Micro chord
blades can be used if they happen to optimize that way.

Also, accelerator engines for our purposes don't absolutely have to
have tens of thousands of hours between overhauls, nice but not
required. Low hundreds would work for the near future for a sufficiently
usefull engine.

In looking over schematics and hardware on existing (50s-60s) jet
engines, I am struck by the large numer and mass of parts other than
the actual blades and burners. My guess is that less than 10% of the mass
of the engines I have looked at is blading, with the rest being axels,

hubs,
casing, seals, stator controls, etc. This, and the limits imposed by the
turbine inlet temperatures led me to eventually conclude that the basic
design concepts are wrong for our purposes.

In looking at a large number of airbreathing accelerator proposals,
it appears to me that the turbojet cycle is still the best if the mass
and turbine temp problems can be resolved. Most of the alternatives
seem to give away so much Isp or SFC that they become more trouble
than they are worth compared to a normal rocket engine.

So I worked out a new layout for the turbojet that addresses some
of the issues. This first description is an early simplified version for
clarity. My purpose in posting is to start getting it into the public
domain to prevent being halted by someone elses' patent again. The
Carmack open source concept has some hard business sense behind
it.

Picture a hollow globe with the latitude and longitude markings on it.
Flatten the globe along the axis until the polar diameter is a third or
so of the equatorial diameter. Starting at the poles, the hub extends from
90 degrees to 70 degrees latitude north and south. Compressor blades
pull air into the lens both north and south from 70 degrees to 10 degrees
latitude north and south. Turbine blades extend from a sealing ring 10
degrees north and south of the equator. The turbine is blocked off from
90 degrees west longitude for 270 degrees back to 0 longitude. The
exhaust is through that 90 degrees of longitude by 20 degrees of latitude.

The burner inside the lens is fed with air from all sides through the
stator,
which eliminates most cooling concerns for the main burner. The turbine
is operating on a 25 % duty cycle with some cooling during the off cycle
by air leaking through the seals. controling the leakage/cooling rate

allows
the burner to operate at an equivelence ratio of 1, or stochiometric.
This eliminates the afterburner, and allows a very low compression ratio
compressor to function.

The above would not have enough compression ratio to be effective at
low speeds, although it would be very effective supersonically. By
introducing a second, contrarotating lens inside the first one, a
compression
ratio of a bit over 2 can be had with modest velocities of the rotating
lens turbocompressors. This should allow a reasonable thrust after the
turbine pressure drop. Static fuel Isp should be in the 800 to 1200 range
depending on assumptions. Supersonic fuel Isp should exceed 2,000
without much problem. Maximum mach is limited by turbine temperatures
which is limited by the compression heating of the air which cools
the turbine, the materials of the turbine blades, and the benifits of the
25% duty cycle. Mach 3 is a guess.

Structure of the blading is a continous band along longitude lines with
each strip being compressor blade-turbine blade-compressor blade.
The blade material is in a tension arch with stiffiners at 10 degrees
north and south latitude. The axel is stationary and experiences
no torque. The main load to the bearings in this configuration is
the off center loading of the turbine. Future configurations address this.
I believe composites can be used for this type structure.
Structural mass appears to be minimal, with some of my estimaes
making it comperable in T/W to a rocket engine.

More later.

John Hare





 




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