Ceramic radial turbine

Abstract

A small gas turbine engine with a ceramic turbine to allow for higher turbine inlet temperatures, where a metallic compressor is secured to a ceramic shaft extending from a ceramic turbine to form a single piece ceramic shaft and turbine, where a threaded nut secures a split ring retainer on the compressor end of the ceramic shaft. A hollow thrust runner is compressed between the compressor disk and the turbine disk by the threaded nut to secure rotor together. A centering spring forms a tight fit between the metallic thrust runner and the ceramic shaft on the turbine side.

Claims

1. A rotor for a small gas turbine engine comprising: a metallic compressor; a ceramic turbine with a ceramic shaft forming a single piece ceramic shaft and ceramic turbine; a metallic thrust runner secured between the metallic compressor and the ceramic turbine with the ceramic shaft extending within the metallic thrust runner and forming a cooling air passage; a split ring retainer and a threaded nut secured over the ceramic shaft to secure the metallic compressor to the ceramic shaft; and a centering spring secured between an inner surface of the metallic thrust runner and an outer surface of the ceramic shaft near to the ceramic turbine, the centering spring includes a plurality of radial outward projections and a plurality of radial inward projections offset to produce a spring effect between the metallic thrust runner and the ceramic shaft.

2. The rotor for the small gas turbine engine of claim 1, wherein the ceramic shaft on the compressor end has a recess on an outer surface; and the split ring retainer has an inner projecting piece that fits within the recess of the ceramic shaft.

3. The rotor for the small gas turbine engine of claim 1, wherein the split ring retainer includes threads on an outer surface; and the retainer nut includes threads on an inner surface to engage the threads on the split ring retainer.

4. The rotor for the small gas turbine engine of claim 1, wherein the metallic thrust runner includes an annular thrust bearing disk extending outward.

5. The rotor for the small gas turbine engine of claim 1, wherein the threaded nut retains the split ring retainer in place on the ceramic shaft and applies axial force to the compressor to compress the metallic thrust runner between the compressor and the turbine.

6. A rotor for a small gas turbine engine comprising: a metallic compressor; a ceramic turbine with a ceramic shaft forming a single piece ceramic shaft and ceramic turbine; a metallic thrust runner secured between the metallic compressor and the ceramic turbine with the ceramic shaft extending within the metallic thrust runner and forming a cooling air passage; a split ring retainer and a threaded nut secured over the ceramic shaft to secure the metallic compressor to the ceramic shaft; and a plurality of torque keys between an inner surface of a compressor disk of the metallic compressor and an outer surface of the ceramic shaft provide a torque transfer from the ceramic shaft to the compressor disk.

Description

BRIEF DESCRIPTION OF THE SEVERAL VIEWS OF THE DRAWINGS

(1) FIG. 1 shows an isometric view of a rotor with a compressor and a turbine of the present invention.

(2) FIG. 2 shows a side view of the rotor of FIG. 1.

(3) FIG. 3 shows a cross section angled view of the rotor of FIG. 1.

(4) FIG. 4 shows a cross section side view of the rotor of FIG. 1.

(5) FIG. 5 shows an isometric view of a nut used to secure the compressor disk to the shaft of the present invention.

(6) FIG. 6 shows an unassembled view of a split ring collet used to secure the compressor disk to the shaft of the present invention.

(7) 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

(8) 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.

(9) 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.

(10) The radial flow gas turbine 12 and ceramic shaft 13 are both formed as a single piece and from Si.sub.3N.sub.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).

(11) 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.

(12) 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.

(13) 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.