Hybrid drive engine

• US 9,709,069 B2
• Filed: 10/22/2013
• Issued: 07/18/2017
• Est. Priority Date: 10/22/2013
• Status: Active Grant

• Current Assignee: Dayspring Church Of God Apostolic
• Original Assignee: Dayspring Church Of God Apostolic
• Inventors: Bosley, David C.
• Primary Examiner(s): Edgar, Richard
• Application Number:
  US14/059,765
• Publication Number:
  US 20150110614A1
• Time in Patent Office: 1,365 Days
• Field of Search: 415 90
• CPC Class Codes: F01D 1/36 using fluid friction
• CPC Class Codes: F04D 17/161 Shear force pumps
• CPC Class Codes: F04D 29/083 especially adapted for elastic fluid pumps
• CPC Class Codes: F04D 29/288 Part of the wheel having an ejecting effect, e.g. being bladeless diffuser
• CPC Class Codes: Y02T 50/60 Efficient propulsion technologies, e.g. for aircraft

Abstract

A hybrid drive engine uses air foil shaped disks of a first configuration for a compressor portion thereof and air foil shaped disks of a second configuration for a turbine portion thereof, whereby the disks exhibit aerodynamic effects of lift. Particularly, the compressor disks are configured to cause aerodynamic lift off of a periphery of the disks, while the turbine disks are configured to cause aerodynamic lift off of an inner hole of the disks. The aerodynamic nature of the disks cause each disk thereof to form two opposing airfoil shapes either head to head or trailing edge to trailing edge across the through hole.

Claims

1. A disk stack for an engine comprising:

a plurality of disks;

each disk having a body defining a closed-figure airfoil comprising;

an upper surface and a lower surface;

a central opening;

an inner perimeter defined by the central opening; and

an outer perimeter radially outward of the inner perimeter;

the closed-figure airfoil having a cross-section comprising;

a line defining the lower surface; and

a convex line defining the upper surface and reaching a zenith that is a highest point on the airfoil section of the disk;

a spacer situated within the central opening of each disk;

an axis of revolution substantially normal to first and second planes defined by the inner perimeter and the outer perimeter of the disk; and

a separator lip on the upper surface of the disk and located proximate the outer perimeter of the disk, the separator lip extending along a narrow ridge that is higher than the immediately adjacent portion of the upper surface.

2. The disk stack of claim 1, wherein the inner perimeter is higher than the outer perimeter.

3. The disk stack of claim 1, wherein the airfoil section has a downwardly depending flap adjacent to the outer perimeter.

4. The disk stack of claim 1, wherein a line tangent to the outer surface of the separator lip is within plus or minus 45 degrees of being parallel to the axis of revolution.

5. The disk stack of claim 1, wherein the inner and outer perimeters are circles described about the axis of revolution.

6. The disk stack of claim 1, wherein the disk has an angular moment capable of assisting aerodynamic lift.

7. The disk stack of claim 1, wherein the upper and lower surfaces are textured to improve aerodynamic performance and boundary layer entrainment.

8. The disk stack of claim 1, wherein the convex line defining the upper surface reaches the zenith at a location that is substantially one third of the distance from the inner perimeter to the outer perimeter.

9. The disk stack of claim 1, wherein the separator lip is lower than the zenith of the upper surface.

10. The disk stack of claim 9, further comprising a downwardly extending flap adjacent the outer perimeter of the lower surface.

11. The disk stack of claim 10, wherein the outer surface of the outer rim is substantially conical such that the diameter of the top edge of the rim is less than the diameter of the bottom edge of the rim.

12. A disk stack for an engine comprising:

a plurality of disks;

each disk having an annular body defining an airfoil with;

an upper surface and a lower surface;

an axis of revolution;

a projected reference plane that is normal to the axis of revolution;

an inner and outer perimeter; and

a reference chord line passing through the inner and outer perimeters;

each airfoil configured with a negative airfoil angle such that revolution of the reference chord defines an angled surface of a frustum of a cone, whereby during revolution a forward portion of the airfoil is at a lower angle of incidence to incoming air than a remainder of the airfoil thereby compensating for air downwash effects from the forward portion of the airfoil and balancing the aerodynamic lift fore and aft on the disk, the airfoil having a cross section presenting a low aerodynamic drag to a flow of air generally parallel to the projected reference plane.

13. The disk stack of claim 12, wherein at least a portion of each disk has a textured surface with features having both circumferential and radial discontinuities.

14. The disk stack of claim 12, wherein each disk is made from a high temperature material.

15. The disk stack of claim 14, wherein the high temperature material comprises a rigid metallic material.

16. The disk stack of claim 14, wherein the high temperature material comprises a high impact thermoplastic.

17. The disk stack of claim 12, wherein the negative airfoil angle is computed by α

_p=α

S_t/S_p, where α

_p=degrees of airfoil angle in those portions of the negative airfoil that are angled, S₁=total airfoil area, S_p=area of the angled portions of the airfoil, α

=(K·

W/D²) where K=45±

15, W=weight of the disk in ounces, and D=mean diameter of the annulus in inches.

18. The disk stack of claim 12, wherein the negative airfoil angle is computed by α

=K W/V² D², where α

=airfoil angle in degrees, K=45±

15, W=the weight of the compressor or turbine disk in ounces, and D=mean diameter of the annulus in inches.

19. The disk stack of claim 12, wherein the negative airfoil angle is computed by α

=K W/D², where α

=airfoil angle in degrees, W=the weight of the compressor or turbine disk in ounces, V=intended flight (rotation) velocity in feet per second, and D=mean diameter of the annulus in inches.

20. The disk stack of claim 12, wherein the negative airfoil angle is computed by α

_p=α

S_t/S_p, where α

_p=degrees of airfoil angle in those portions of the airfoil that are angled, S_t=total airfoil area, S_p=area of the angled portions of the airfoil, α

=(K·

W/D²) where K=45±

15, W=the weight of the compressor or turbine disk in ounces, and D=mean diameter of the annulus in inches.

21. An engine comprising:

a housing;

a compressor section disposed in the housing and comprising a compressor disk stack of a plurality of annular compressor disks having a compressor disk body defining a closed-figure airfoil;

a turbine section disposed in the housing and in communication with the compressor section, the turbine section comprising a turbine disk stack of a plurality of annular turbine disks having a turbine disk body defining a closed-figure airfoil; and

a central shaft mechanically linking the turbine disk stack with the compressor disk stack

wherein the closed-figure airfoil defining the body of each disk of the compressor and turbine disk stacks comprises;

an upper surface and a lower surface;

a central opening;

an inner perimeter defined by the central opening;

an outer perimeter radially outward of the inner perimeter;

a cross-section comprising;

a line defining the lower surface; and

a convex line defining the upper surface and reaching a zenith that is a highest point on the airfoil of the disk; and

a separator lip on the upper surface of the disk and located proximate the outer perimeter of the disk, the separator lip extending along a narrow ridge that is higher than the immediately adjacent portion of the upper surface.

22. The engine of claim 21, wherein the inner perimeter of each compressor and turbine disk is higher than the outer perimeter thereof.

23. The engine claim 21, wherein the airfoil section has a downwardly depending flap adjacent to the outer perimeter.

24. The engine of claim 21, wherein a line tangent to the outer surface of the separator lip is within plus or minus 45 degrees of being parallel to the axis of revolution.

25. The engine of claim 21, wherein the inner and outer perimeters of each compressor and turbine disks are circles described about the axis of revolution.

26. The engine of claim 21, wherein each compressor and turbine disk has an angular moment capable of assisting aerodynamic lift.

27. The engine of claim 21, wherein the upper and lower surfaces of each compressor and turbine disk are textured to improve aerodynamic performance and boundary layer entrainment.

28. The engine of claim 21, wherein the convex line defining the upper surface reaches the zenith at a location that is substantially one third of the distance from the inner perimeter to the outer perimeter.

29. The engine of claim 21, wherein the separator lip is lower than the zenith of the upper surface.

30. The engine of claim 29, further comprising a downwardly extending flap adjacent the outer perimeter of the lower surface.

31. The engine of claim 21, further comprising:

a second compressor section disposed in the housing and comprising a second compressor disk stack of a plurality of annular compressor disks having a compressor disk body defining a closed-figure airfoil; and

a second turbine section disposed in the housing and comprising a second turbine disk stack of a plurality of annular turbine disks having a turbine disk body defining a closed-figure airfoil;

an outer shaft mechanically linking the second compressor disk stack to the second turbine disk stack.