Microturbomachinery

• US 5,932,940 A
• Filed: 11/15/1996
• Issued: 08/03/1999
• Est. Priority Date: 07/16/1996
• Status: Expired due to Term

• Current Assignee: Massachusetts Institute of Technology
• Original Assignee: Massachusetts Institute of Technology
• Inventors: Jacobson, Stuart A., Lang, Jeffrey H., Schmidt, Martin A., Nagle, Steven F., Senturia, Stephen D., Epstein, Alan H., Spearing, Mark S., Ananthasuresh, G. K., Ehrich, Fredric F., Waitz, Ian A., Breuer, Kenneth S.
• Primary Examiner(s): Jackson, Jr., Jerome
• Application Number:
  US08/749,454
• Time in Patent Office: 991 Days
• Field of Search: 257/414, 257/415, 60/39.34, 60/39.35, 60/39.36, 60/39.37, 60/330, 60/364, 290/52, 310/40 MM
• US Class Current: 310/40.0MM
• CPC Class Codes: F01D 15/06 Adaptations for driving, or combinations with, hand-held tools or the like control thereof
• CPC Class Codes: F01D 25/20 using lubrication pumps
• CPC Class Codes: F01D 5/28 Selecting particular materials; Particular measures relating thereto; Measures against erosion or corrosion
• CPC Class Codes: F01K 13/00 General layout or general methods of operation of complete plants
• CPC Class Codes: F02C 3/05 the compressor and the turbine being of the radial flow type
• CPC Class Codes: F02K 9/48 driven by a gas turbine fed by propellant combustion gases or fed by vaporized propellants or other gases
• CPC Class Codes: F05D 2220/50 for auxiliary power units (APU's)
• CPC Class Codes: F05D 2230/14 Micromachining
• CPC Class Codes: F05D 2250/82 Micromachines
• CPC Class Codes: F05D 2250/84 Nanomachines
• CPC Class Codes: F05D 2300/2261 of silicon
• CPC Class Codes: F05D 2300/2283 of silicon
• CPC Class Codes: F05D 2300/607 Monocrystallinity
• CPC Class Codes: F23C 2900/03001 Miniaturized combustion devices using fluid fuels
• CPC Class Codes: F23R 3/42 characterised by the arrangement or form of the flame tubes or combustion chambers

Abstract

The invention provides a micro-gas turbine engine and associated microcomponentry. The engine components, including, e.g., a compressor, a diffuser having diffuser vanes, a combustion chamber, turbine guide vanes, and a turbine are each manufactured by, e.g., microfabrication techniques, of a structural material common to all of the elements, e.g., a microelectronic material such as silicon or silicon carbide. Vapor deposition techniques, as well as bulk wafer etching techniques, can be employed to produce the engine. The engine includes a rotor having a shaft with a substantially untapered compressor disk on a first end, defining a centrifugal compressor, and a substantially untapered turbine disk on the opposite end, defining a radial inflow turbine. The rotor is preferably formed of a material characterized by a strength-to-density ratio that enables a rotor speed of at least about 500,000 rotations per minute. An annular, axial-flow combustion chamber is provided that is located axially between the compressor and turbine disks and that has a ratio of annular height to axial length of at least about 0.5. The micro-gas turbine engine can be configured with an integral microgenerator as a source of electrical power, and can be employed for a wide range of power, propulsion, and thermodynamic cycle applications.

Claims

1. A gas turbine engine comprising:

a compressor having an engine inlet port and a rotatable compressor disk for compressing inlet air;

a diffuser located to receive compressed air from the compressor and produce a corresponding rise in static air pressure;

a combustor comprising a combustion chamber located to receive the compressed air from the diffuser through a combustor inlet to combust the compressed air with a fuel;

a plurality of turbine nozzle guide vanes located to receive combusted air exhausted from the combustor;

a turbine having a rotatable turbine disk located to receive combusted air from the turbine nozzle guide vanes, and having an engine exhaust port; and

bearing surfaces of the rotatable elements;

wherein the compressor, diffuser, combustion chamber, turbine nozzle guide vanes, and turbine elements and the bearing surfaces each substantially comprise a structural material common to all of the elements.

2. A gas turbine engine comprising:

a compressor having an engine inlet port and a rotatable compressor disk for compressing inlet air;

a diffuser located to receive compressed air from the compressor and produce a corresponding rise in static air pressure;

a combustor comprising a combustion chamber located to receive the compressed air from the diffuser through a combustor inlet to combust the compressed air with a fuel;

a plurality of turbine nozzle guide vanes located to receive combusted air exhausted from the combustor; and

a turbine having a rotatable turbine disk located to receive combusted air from the turbine nozzle guide vanes, and having an engine exhaust port;

wherein the compressor diffuser, combustion chamber, turbine nozzle guide vanes, and turbine elements each substantially comprise silicon carbide.

3. The gas turbine engine of claim 2 wherein the silicon carbide comprises vapor deposited silicon carbide.

4. A gas turbine engine comprising:

a compressor having an engine inlet port and a rotatable compressor disk for compressing inlet air;

a diffuser located to receive compressed air from the compressor and produce a corresponding rise in static air pressure;

a combustor comprising a combustion chamber located to receive the compressed air from the diffuser through a combustor inlet to combust the compressed air with a fuel;

a plurality of turbine nozzle guide vanes located to receive combusted air exhausted from the combustor; and

a turbine having a rotatable turbine disk located to receive combusted air from the turbine nozzle guide vanes, and having an engine exhaust port;

wherein the compressor, diffuser, combustion chamber, turbine nozzle guide vanes, and turbine elements each substantially comprise silicon.

69. The gas turbine engine of claim 4 wherein the silicon elements are coated with silicon carbide.

70. The gas turbine engine of claim 4 wherein the silicon elements are coated with silicon nitride.

5. A gas turbine engine comprising:

a compressor having an engine inlet port and a rotatable compressor disk for compressing inlet air;

a diffuser located to receive compressed air from the compressor and produce a corresponding rise in static air pressure;

a combustor comprising a combustion chamber located to receive the compressed air from the diffuser through a combustor inlet to combust the compressed air with a fuel;

a plurality of turbine nozzle guide vanes located to receive combusted air exhausted from the combustor; and

a turbine having a rotatable turbine disk located to receive combusted air from the turbine nozzle guide vanes and having an engine exhaust port;

wherein the compressor, diffuser, combustion chamber, turbine nozzle guide vanes, and turbine elements each substantially comprise silicon nitride.

6. A gas turbine engine comprising:

a compressor having an engine inlet port and a rotatable compressor disk for compressing inlet air;

a diffuser located to receive compressed air from the compressor and produce a corresponding rise in static air pressure;

a combustor comprising a combustion chamber located to receive the compressed air from the diffuser through a combustor inlet to combust the compressed air with a fuel;

a plurality of turbine nozzle guide vanes located to receive combusted air exhausted from the combustor; and

a turbine having a rotatable turbine disk located to receive combusted air from the turbine nozzle guide vanes, and having an engine exhaust port;

wherein the compressor, diffuser, combustion chamber, turbine nozzle guide vanes, and turbine elements each substantially comprise a semiconductor material.

7. A gas turbine engine comprising:

a compressor having an engine inlet port and a rotatable compressor disk for compressing inlet air;

a diffuser located to receive compressed air from the compressor and produce a corresponding static rise in air pressure;

a combustor comprising a combustion chamber located to receive the compressed air from the diffuser through a combustor inlet to combust the compressed air with a fuel;

a plurality of turbine nozzle guide vanes located to receive combusted air exhausted from the combustor; and

a turbine having a rotatable disk located to receive combusted air from the turbine nozzle guide vanes, and having an engine exhaust port;

wherein the compressor, diffuser, combustion chamber, turbine nozzle guide vanes, and turbine elements each comprise a microfabricated element, and a ratio of engine radial height to engine axial length is at least about 1.

8. The gas turbine engine of claim 7 wherein the wherein the compressor, diffuser, combustion chamber, turbine nozzle guide vanes, and turbine elements each substantially comprise a structural material common to all of the elements.

9. The gas turbine engine of claim 7 wherein the compressor and turbine each comprise single crystal silicon disks.

10. The gas turbine engine of claim 9 wherein the compressor and turbine each comprise microfabricated bulk-etched silicon wafers.

11. The gas turbine engine of claim 7 wherein the combustion chamber comprises microfabricated vapor-deposited silicon carbide.

12. The gas turbine engine of claim 7 wherein compressor and turbine each comprise silicon carbide disks.

13. The gas turbine engine of claim 12 wherein the compressor and turbine each comprise microfabricated bulk-etched silicon carbide wafers.

14. A gas turbine engine comprising:

a rotor comprising a shaft having a substantially untapered compressor disk on a first end of the shaft with compressor blades defining a centrifugal compressor and a substantially untapered turbine disk on an opposite end of the shaft with turbine blades defining a radial inflow turbine, the rotor further comprising a material characterized by a strength-to-density ratio enabling a rotor speed of at least about 500,000 rotations per minute;

an annular, radial out-flow diffuser located radially outward of the compressor disk for receiving compressed air from the compressor;

an annular combustor located radially outward of the rotor shaft, having a combustor inlet radially outward of the diffuser, and having an exhaust outlet radially outward of the turbine disk; and

a plurality of radial inflow turbine nozzle guide vanes located radially outward of the turbine disk and having an inlet for accepting combustion exhaust and directing the exhaust to the turbine.

15. The gas turbine engine of claim 14 wherein the compressor and turbine disks each comprise a disk diameter of between about 1 mm and 20 mm and wherein the combustor comprises a combustion chamber having an annular height of between about 2 mm-10 mm and an axial length of between about 0.5 mm-12 mm.

16. The gas turbine engine of claim 14 wherein the combustor comprises a combustion chamber having dimensions characterized by a ratio of chamber annular height to chamber axial length that is at least about 1.

17. The gas turbine engine of claim 14 wherein the rotor shaft is journalled for rotation by a radial gas journal bearing having a clearance of between about 2 μ

m-25 μ

m between the rotor shaft and an engine housing wall.

18. The gas turbine engine of claim 17 wherein the radial gas journal bearing has a tangential, circumferentially asymmetric pressurization seal.

19. The gas turbine engine of claim 14 wherein the compressor disk and the turbine disk are each journalled for rotation by a radial gas journal bearing having a clearance between a radially peripheral edge of the disk and a corresponding engine housing wall.

20. The gas turbine engine of claim 14 wherein the compressor disk and the turbine disk are located back-to-back on the rotor.

21. The gas turbine engine of claim 14 wherein the compressor blades comprise a plurality of radial-flow blades together of a common axial height and each of a constant axial height along blade radial extent, and wherein the turbine blades comprise a plurality of radial-flow blades together of a common axial height and each of a constant axial height along blade radial extent.

22. The gas turbine engine of claim 21 wherein the diffuser comprises a plurality of radial-flow diffuser vanes together of a common axial height with the plurality of compressor blades and each of a constant axial height along vane radial extent, and wherein the turbine guide vanes comprise a plurality of radial-flow vanes together of a common axial height with the plurality of turbine blades and each of a constant axial height along vane radial extent.

23. The gas turbine engine of claim 14 further comprising a motor having a motor rotor defined by the compressor and having stator electrodes defined by conducting regions located on an engine endwall opposite the compressor disk and interconnected to define a stator excitation configuration of at least two excitation phases.

24. The gas turbine engine of claim 23 wherein the motor rotor is defined by the compressor blades.

25. The gas turbine engine of claim 23 wherein the compressor blades comprise a plurality of radial-flow blades together of a common axial height and each of a constant axial height along blade radial extent, the compressor further comprising a blade shroud connected on the axial height of the blades and having an upper conducting layer defining an electric induction motor rotor.

26. The gas turbine engine of claim 14 further comprising a generator having a generator rotor defined by the compressor and having stator electrodes defined by conducting regions located on an engine endwall opposite the compressor disk and interconnected to define a stator excitation configuration of at least one excitation phase.

27. The gas turbine engine of claim 14 further comprising a generator having a generator rotor defined by the turbine and having stator electrodes defined by conducting regions located on an engine endwall opposite the turbine disk and interconnected to define a stator excitation configuration of at least one excitation phase.

28. The gas turbine engine of claim 14 further comprising a motor having a motor rotor defined by the turbine and having stator electrodes defined by conducting regions located on an engine endwall opposite the turbine disk and interconnected to define a stator excitation configuration of at least two excitation phases.

29. The gas turbine engine of claim 27 wherein the generator rotor is defined by the turbine blades.

30. The gas turbine engine of claim 27 wherein the turbine blades comprise a plurality of radial-flow blades together of a common axial height and each of a constant axial height along blade radial extent, the turbine further comprising a blade shroud connected on the axial height of the blades and having an upper conducting layer defining an electric induction generator rotor.

31. The gas turbine engine of claim 14 further comprising:

an annular fuel plenum located radially outward of the compressor disk; and

an array of fuel injectors connected between the fuel plenum and a location on an engine endwall along internal engine air flow upstream of the combustor inlet.

32. The gas turbine engine of claim 31 further comprising a fuel tank, a motor connected to the rotor, an auxiliary power source for starting the motor, and control electronics, and wherein the gas turbine engine is housed in an engine package having a air inlet port connected to the compressor, an air exhaust port connected to the turbine, a fuel port connected to the fuel tank, a fuel line connecting the fuel tank to the fuel plenum, an electrical port connected to the control electronics and auxiliary power source, and an isolation wall isolating the fuel tank from other engine components.

33. The gas turbine engine of claim 32 wherein the engine package comprises a volume less than about 1 liter.

34. The gas turbine engine of claim 32 wherein the fuel tank comprises a volume sufficient for storing between about 5 ml-5000 ml of liquid engine fuel.

35. The gas turbine engine of claim 32 wherein the engine package comprises a plastic package.

36. The gas turbine engine of claim 14 further comprising an annular recuperator connected to receive compressed air from the diffuser and to receive air exhausted from the turbine to heat the compressed air and discharge the heated compressed air to the combuster.

37. The gas turbine engine of claim 36 wherein the recuperator comprises planar hot air channels connected to receive air exhausted from the turbine and located in alternating fashion with cold air channels connected to receive compressed air from the diffuser.

38. The gas turbine engine of claim 14 further comprising:

a bypass fan located axially forward of the compressor disk and having bypass fan blades; and

an axial bypass channel located radially outward of the combustor and providing an exhaust port axially aft of the turbine.

39. The gas turbine engine of claim 38 wherein the compressor blades comprise a plurality of radial-flow blades together of a common axial height and each of a constant axial height along blade radial extent, the compressor further comprising a blade shroud connected on the axial height of the blades and supporting the bypass fan blades, together of a common axial height and each of a constant axial height along blade radial extent.

40. The gas turbine engine of claim 14 wherein the rotor, diffuser, combustor, and turbine nozzle guide vanes elements each substantially comprise a structural material common to all of the elements.

41. The gas turbine engine of claim 14 wherein the rotor, diffuser, combustor, and turbine nozzle guide vane elements each comprise a microfabricated element, and a ratio of engine radial height to engine axial length is at least about 1.

65. The gas turbine engine of claim 21 wherein the diffuser comprises a plurality of radial-flow diffuser vanes together of a common axial height that is distinct from axial height of the compressor blades, and wherein the turbine nozzle guide vanes comprise a plurality of radial-flow vanes together of a common axial height that is distinct from axial height of the turbine blades.

42. A gas turbine engine comprising:

a rotor comprising a shaft having a compressor disk on a first end of the shaft with compressor blades defining a centrifugal compressor and a turbine disk on an opposite end of the shaft with turbine blades defining a radial inflow turbine;

an annular, radial out-flow diffuser located radially outward of the compressor disk for receiving compressed air from the compressor;

an annular, linerless combustor having a combustion chamber located radially outward of the rotor shaft, having a combustor inlet radially outward of the diffuser, and having an exhaust outlet radially outward of the turbine disk, the combustion chamber having dimensions characterized by a ratio of chamber annular height to chamber axial length that is at least about 1; and

a plurality of radial in-flow turbine nozzle guide vanes located radially outward of the turbine disk and having an inlet for accepting combustion exhaust from the combustion chamber and directing the exhaust to the turbine.

43. The gas turbine engine of claim 42 wherein the compressor and turbine disks each comprise a disk diameter of between about 1 mm and 20 mm and wherein the combustor comprises a combustion chamber having an annular height of between about 2 mm-10 mm and an axial length of between about 0.5 mm-12 mm.

44. The gas turbine engine of claim 42 wherein the compressor disk and the turbine disk are located back-to-back on the rotor.

45. The gas turbine engine of claim 42 wherein the compressor blades comprise a plurality of radial-flow blades together of a common axial height and each of a constant axial height along blade radial extent, and wherein the turbine blades comprise a plurality of radial-flow blades together of a common axial height and each of a constant axial height along blade radial extent.

46. The gas turbine engine of claim 42 wherein the diffuser comprises a plurality of radial-flow diffuser vanes together of a common axial height with the plurality of compressor blades and each of a constant axial height along vane radial extent, and wherein the turbine nozzle guide vanes comprise a plurality of radial-flow vanes together of a common axial height with the plurality of turbine blades and each of a constant axial height along vane radial extent.

47. The gas turbine engine of claim 42 further comprising:

an annular fuel plenum located radially outward of the compressor disk; and

an array of fuel injectors connected between the fuel plenum and a location on an engine endwall along internal engine air flow upstream of the combustor inlet.

48. The gas turbine engine of claim 47 further comprising a fuel tank, a motor connected to the rotor, an auxiliary power source for starting the motor, and control electronics, and wherein the gas turbine engine is housed in an engine package having a air inlet port connected to the compressor, an air exhaust port connected to the turbine, a fuel port connected to the fuel tank, a fuel line connecting the fuel tank to the fuel plenum, an electrical port connected to the control electronics and auxiliary power source, and an isolation wall isolating the fuel tank from other engine components.

49. The gas turbine engine of claim 42 wherein the rotor, diffuser, combustion chamber, and turbine nozzle guide vane elements each substantially comprise a structural material common to all of the elements.

50. The gas turbine engine of claim 42 wherein the rotor, diffuser, combustion chamber, and turbine nozzle guide vane elements each comprise a microfabricated element, and a ratio of engine radial height to engine axial length is at least about 1.

66. The gas turbine engine of claim 42 wherein the diffuser comprises a plurality of radial-flow diffuser vanes together of a common axial height that is distinct from axial height of the compressor blades, and wherein the turbine nozzle guide vanes comprise a plurality of radial-flow vanes together of a common axial height that is distinct from axial height of the turbine blades.

51. A gas turbine engine comprising:

a compressor disk having radial-flow compressor blades defining a centrifugal compressor;

a turbine disk located in a plane coincident with the compressor disk and having radial-flow turbine blades defining a radial inflow turbine;

an annular, radial out-flow diffuser located radially outward of the compressor disk for receiving compressed air from the compressor;

a combustor having a combustion chamber with a combustor inlet in communication with the diffuser for receiving compressed air from the diffuser and an exhaust vent in the plane of the turbine disk; and

a plurality of radial inflow turbine nozzle guide vanes located radially outward of the turbine disk and having an inlet for accepting combustion exhaust and directing the exhaust to the turbine disk.

52. The gas turbine engine of claim 51 wherein the gas turbine engine has a thickness of less than about 1 centimeter.

53. The gas turbine engine of claim 51 further comprising a fuel manifold located adjacent to the diffuser and connected to the diffuser by a row of spaced apart fuel injectors.

54. The gas turbine engine of claim 51 further comprising forward and aft cooling channels each located adjacent to a corresponding forward and aft combustion chamber wall and connected between the diffuser exhaust vent and the combustor inlet.

55. The gas turbine engine of claim 51 wherein the diffuser comprises a plurality of diffuser vanes, and wherein the compressor blades, the turbine blades, the diffuser vanes, and the turbine nozzle guide vanes comprise radial-flow blades and vanes together of a common axial height and each of a constant axial height along corresponding blade and vane radial extent.

56. The gas turbine engine of claim 51 further comprising a motor having a motor rotor defined by the compressor disk and stator electrodes defined by conducting regions located on an engine endwall opposite the compressor disk and interconnected to define a stator excitation configuration of at least two excitation phases.

57. The gas turbine engine of claim 51 further comprising a generator having a generator rotor defined by the turbine disk and having stator electrodes defined by conducting regions located on an engine endwall opposite the turbine disk and interconnected to define a stator excitation configuration of at least one excitation phase.

58. A combustion power generator comprising:

a first pump disk having pump blades defining a rotary fuel pump;

a second pump disk located in a plane coincident with the first pump disk and having pump blades defining a rotary oxidizer pump;

a turbine disk located in a plane coincident with the first and second pump disks and having turbine blades defining a radial inflow turbine;

a first annular, radial out-flow diffuser located radially outward of the first pump disk for receiving fuel from the fuel pump;

a second annular, radial out-flow diffuser located radially outward of the second pump disk for receiving oxidizer from the oxidizer pump;

a combustor comprising a combustion chamber with a combustor inlet in communication with the first and second diffusers for receiving fuel and oxidizer, and an exhaust vent in the plane of the turbine disk;

a plurality of radial inflow turbine nozzle guide vanes located radially outward of the turbine disk and having an inlet for accepting combustion exhaust and directing the exhaust to the turbine disk; and

a generator having a generator rotor defined by the turbine and having stator electrodes defined by conducting regions located on a generator endwall opposite the turbine disk and interconnected to define a stator excitation configuration of at least one excitation phase.

59. A combustion power generator comprising:

a pump disk having pump blades defining a rotary fuel pump;

a turbine disk located in a plane coincident with the pump disk and having turbine blades defining a radial inflow turbine;

an annular, radial out-flow diffuser located radially outward of the pump disk for receiving fuel from the fuel pump;

a combustor comprising a combustion chamber with a combustor inlet in communication with the diffuser for receiving fuel, and an exhaust vent in the plane of the turbine disk;

a plurality of radial inflow turbine nozzle guide vanes located radially outward of the turbine disk and having an inlet for accepting combustion exhaust and directing the exhaust to the turbine disk; and

a generator having a generator rotor defined by the turbine and having stator electrodes defined by conducting regions located on a generator endwall opposite the turbine disk and interconnected to define a stator excitation configuration of at least one excitation phase.

60. The combustion power generator of claim 59 wherein the power generator has a thickness of less than about 1 centimeter.

61. The combustion power generator of claim 59 wherein the diffuser comprises a plurality of diffuser vanes, and wherein the fuel pump blades, the turbine blades, the diffuser vanes, and the turbine nozzle guide vanes comprise radial-flow blades and vanes together of a common axial height and each of a constant axial height along corresponding blade and vane radial extent.

62. The combustion power generator of claim 59 further comprising:

a fuel pump motor having a motor rotor defined by the fuel pump and stator electrodes defined by conducting regions located on an endwall opposite the fuel pump disk.

63. The combustion power generator of claim 59 wherein the fuel pump, diffuser, combustion chamber, turbine nozzle guide vanes, and turbine elements each substantially comprise a structural material common to all of the elements.

64. The gas turbine engine of claim 59 wherein the fuel pump, diffuser, combustion chamber, turbine guide vanes, and turbine elements each comprise a microfabricated element, and a ratio of engine radial height to engine axial length is at least about 1.

67. A gas turbine engine comprising:

a silicon compressor having an engine inlet port and a rotatable compressor disk for compressing inlet air;

a silicon diffuser located to receive compressed air from the compressor and produce a corresponding rise in static air pressure;

a silicon carbide combustor comprising a combustion chamber located to receive the compressed air from the diffuser through a combustor inlet to combust the compressed air with a fuel;

a plurality of silicon carbide turbine nozzle guide vanes located to receive combusted air exhausted from the combustor; and

a silicon carbide turbine having a rotatable turbine disk located to receive combusted air from the turbine nozzle guide vanes, and having an engine exhaust port.

68. A gas turbine engine comprising:

a rotor comprising a shaft having a compressor disk, of between about 1 mm and about 10 mm in diameter, on a first end of the shaft with compressor blades defining a centrifugal compressor, and a turbine disk, of between about 1 mm and about 10 mm in diameter, on an opposite end of the shaft with turbine blades defining a radial inflow turbine;

an annular, radial out-flow diffuser located radially outward of the compressor disk for receiving compressed air from the compressor;

an annular combustor located radially outward of the rotor shaft, having a combustor inlet radially outward of the diffuser, having an exhaust outlet radially outward of the turbine disk and having a combustion chamber with an annular height of between about 2 mm and about 20 mm and an axial length of between about 0.5 mm and about 12 mm; and

a plurality of radial inflow turbine nozzle guide vanes located radially outward of the turbine disk and having an inlet for accepting combustion exhaust and directing the exhaust to the turbine.