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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
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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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