Patent Publication Number: US-11028698-B1

Title: Ceramic radial turbine

Description:
CROSS-REFERENCE TO RELATED APPLICATIONS 
     This Application claims the benefit to U.S. Provisional Application 62/688,819 filed on Jun. 22, 2018 and entitled CERAMIC RADIAL TURBINE. 
    
    
     GOVERNMENT LICENSE RIGHTS 
     None. 
     BACKGROUND OF THE INVENTION 
     Field of the Invention 
     The present invention relates generally to a gas turbine engine, and more specifically to a small gas turbine engine for a Unmanned Aerial Vehicle (UAV) with a high turbine inlet temperature. 
     Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98 
     In a gas turbine engine, a gas turbine drives a compressor to supply compressed air to a combustor where a fuel is burned to produce a hot gas flow that is passed through the gas turbine to drive the compressor and a fan to propel the vehicle. The efficiency of the engine can be increased by passing a higher temperature gas flow into the turbine. However, the turbine inlet temperature is limited to what the turbine materials can withstand. Nickel super alloys are typically used as a material for the gas turbine. Turbine airfoil cooling is performed to allow for even higher turbine inlet temperatures. However, for a small gas turbine engine of the type used to power a UAV, the airfoils of the gas turbine are too small for cooling passages. 
     BRIEF SUMMARY OF THE INVENTION 
     A small gas turbine engine for a UAV, where the engine includes a radial flow gas turbine and rotor shaft both made as a single piece and from a ceramic material so that an increased firing temperature can be used that will allow for a power to weight ratio of the engine to be more than double that from an all-metal gas turbine engine. 
     A metallic shaft thrust runner forms an annular cooling passage with the ceramic shaft to pass cooling air. A compliant spacer star centering ring is located adjacent to the radial flow gas turbine between the ceramic shaft and the metallic shaft thrust runner. 
    
    
     
       BRIEF DESCRIPTION OF THE SEVERAL VIEWS OF THE DRAWINGS 
         FIG. 1  shows an isometric view of a rotor with a compressor and a turbine of the present invention. 
         FIG. 2  shows a side view of the rotor of  FIG. 1 . 
         FIG. 3  shows a cross section angled view of the rotor of  FIG. 1 . 
         FIG. 4  shows a cross section side view of the rotor of  FIG. 1 . 
         FIG. 5  shows an isometric view of a nut used to secure the compressor disk to the shaft of the present invention. 
         FIG. 6  shows an unassembled view of a split ring collet used to secure the compressor disk to the shaft of the present invention. 
         FIG. 7  shows an isometric view of a centering springs used to connect a metal thrust runner to a ceramic shaft of the present invention. 
     
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
     The present invention is a small gas turbine engine used to power an unmanned aero vehicle (UAV) in which the gas turbine is a radial flow gas turbine made of a ceramic material along with a ceramic shaft connected to a metal compressor, where the ceramic radial flow gas turbine is without cooling and the ceramic shaft includes an metal outer sleeve that forms a cooling passage for the turbine shaft. The ceramic radial flow turbine and the ceramic shaft are formed as a single piece. The ceramic radial turbine and ceramic shaft of the present invention will allow for a combustor firing temperature (T4) of around 2,400 degrees F. which will more than double the power to weight capability of the engine over a prior art all metal gas turbine engine. 
     The small gas turbine engine includes a radial flow compressor  11  and a radial flow gas turbine  12  both supported by air foil bearings. A reverse flow combustor is integrated within the structure of a high effectiveness recuperator. The engine powers a high speed electric generator that is also supported on air foil bearings. The electric generator can be directly driven by the shaft of the engine, or can be driven through an oil-less gearbox for shaft drive applications. Use of the integrated recuperator with this engine will allow for a compressor pressure ratio of 5 to 6 which will avoid the historic issues of environmental effects causing ceramic surface degradation seen in APU (Auxiliary Power Unit) applications and stationary industrial gas turbines. 
     The radial flow gas turbine  12  and ceramic shaft  13  are both formed as a single piece and from Si 3 N 4  monolithic ceramic material. With this monolithic ceramic material, it is feasible to increase the relative rotor inlet temperature to 2,250 degrees F. equivalent to around 2,400 degrees F. firing temperature (T4). 
       FIG. 1  shows an assembled rotor of the present invention with a metal compressor  11  and a ceramic turbine  12 , a threaded nut  17  that secures the front side of the compressor  11  to the shaft  13 , and a thrust runner  14  with a thrust bearing disk  18  for air foil thrust bearings to make contact with.  FIG. 2  shows a side view of the assembled rotor with the compressor  11  and the turbine  13  with the thrust runner  14  in-between the two disks  11  and  13 .  FIG. 3  shows a cross section view of the rotor with the ceramic shaft  13  extending from the ceramic turbine  13  through the thrust runner  14  to the metallic compressor  11 .  FIG. 4  shows a cross section side view of the rotor from  FIG. 3 . The threaded retention nut  17  is secured over the split ring retainer  21  ( FIG. 6 ) to secure the compressor  11  to the shaft  13 . Torque keys  19  in  FIG. 4  are used to provide torque between the shaft  13  and the compressor disk  11 . 
     The radial flow compressor  11  made from a non-ceramic material is secured to the ceramic shaft  13  using the threaded split ring retainer  21  held in place by the single piece threaded retention nut  17 . At the compressor end, the ceramic shaft  13  is ground with a double conical recess where the threaded split ring retainer  21  is inserted and compressed by a retention nut  17 . Small flats are ground on the ceramic shaft  13  that interface with corresponding flats on the interior of a split retention ring  21 .  FIG. 5  shows the retention nut  17 .  FIG. 6  shows the two-piece threaded split ring retainer  21  with an inner annular protecting part  22  that fits within double conical recess formed on the outer surface of the shaft  13  to retain the compressor  11  to the ceramic shaft  13 . The inner surface of the retention nut  17  and the outer surface of the split ring retainer  21  have threads that engage to secure the retention nut over the split ring retainer  23 . With this arrangement, the high temperature turbine rotor is thermally isolated from the bearing shaft runner with the interrupted conduction path of the compliant spacer star of the turbine side of the ceramic shaft  13 . The temperature drop in the ceramic shaft  13  is allowed to utilize the entire shaft length to the compressor  11  end and will minimize the shaft thermal stresses, and reduce the heat load to the bearing runner. 
     On the ceramic turbine shaft  13 , a metallic shaft runner  14  is positioned with an interference fit compliant spacer star centering ring  23  situated between the shaft runner  14  and the ceramic turbine shaft  13 . The centering ring  23  provides for a tight fit between the metal thrust runner  14  and the ceramic shaft  13  so that a tight fit is formed even when the metallic thrust runner  14  expands with respect to the ceramic shaft  13  under high temperatures. An annular cooling flow passage  15  is formed between the ceramic shaft  13  and the metallic thrust runner  14  in which cooling air is passed through the annular passage  15  and through the compliant spacer star centering ring  23 .  FIG. 7  shows an isometric view of the compliant spacer star centering spring  23  with radial outward projections  24  abutting an inner surface of the metallic shaft runner  14  and radial inward projections  25  abutting an outer surface of the ceramic shaft  13 . The radial outward projections  24  are offset from the radial inward projections  25  to produce a spring effect between the metallic thrust runner  14  and the ceramic shaft  13 .