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.

Description:
FIELD OF THE INVENTION 
     The present disclosure relates to devices for increasing the pressure of fluids such as air, or changing a fluid&#39;s thermodynamic energy into mechanical energy for propulsion and, more particularly, to a propulsion system that utilizes aerodynamically configured spinning disks to yield a dynamic response from the resulting slipstream. 
     BACKGROUND 
     Propulsion systems that utilize compressors and turbines are well known. Previously, compressors or fluid compression systems more or less impart sudden changes, vibrations and shocks to redirect and or compress the fluid. One object of the present invention is to move and provide a motive force to a fluid without imparting cavitation similar shocks. While turbines or turbine systems have been designed to take advantage of the flow characteristics of hot fluids by utilizing a series of aerodynamic vanes that extend radially outward from and are evenly spaced on a rotatable axis such as a shaft. Hot fluid flow in the turbine exchanges its thermodynamic energy by impinging upon the vanes so as to rotate the shaft. However, because of excessive heat of the fluid, some turbine versions suffer problems such as stretching, melting and otherwise stressing of the vanes. These can be especially dangerous circumstances that not only could damage the turbine, but could cause components to break or free themselves from the turbine, especially when loaded with centrifugal energy. 
     Further, the methodology to attach the aerodynamic vanes to the shaft is very time consuming, complex, cumbersome and thus expensive. Even so, compressors and turbines with aerodynamic vanes are routinely used and are designed in an attempt to overcome their centripetal or thermal deficiencies and avoid the aforementioned complications. As such, some compressors employ pistons, paddles, vanes or blades to impart energy to the fluid flow. Additionally, turbine vanes are designed with holes, bores, and/or other configurations that allow cooling air or other fluid(s) to flow through them to help maintain their resistance to metal fatigue. This however, adds to the cost of production, maintenance, and complicates the arrangement and layout of the compressor/turbine system. 
     Vane-less or variant compressor/turbine systems have been developed that do not suffer from the complications of compressor/turbine systems with vanes. One such vane-less design utilizes Prandtl Layer compressor/turbine systems of smooth disks to utilize adhesion and viscosity of fluids to exchange the fluid movement, to increase the pressure of the final outlet fluid or in the turbine utilize thermodynamic energy from smooth disks into rotation of a central shaft. This arrangement, however, fails to afford a high efficiency method of energy exchange. 
     Another vane-less arrangement by Nikola Tesla employs a plurality of substantially planar parallel disks between which fluid is directed. However, this arrangement fails to efficiently contend with turbulent losses that exist at both the input and output of the compressor/turbine systems due to the use of input nozzles and spider-mounted shaft disks. This arrangement also suffers losses resulting from “scrubbing” of the peripheral turbine casing. These, along with end-wall flow coupling illustrate just a few of the known problems of such compressor/turbine systems. 
     Overall, these various arrangements and the prior art as a whole, fail to implement aerodynamically efficient compressor/turbine systems designs. At best, it can be said that some compressor/turbine systems provide at least neutral aerodynamic features. Consequently, these designs suffer in overall efficiency from this shortcoming alone. Often to successfully overcome these shortcomings requires technological mechanism(s) that significantly over complicate compressor/turbine systems arrangement, layout, and capital outlay required for implementation. 
     In view of the above, it can be appreciated that a mechanism which imparts simplicity of design and implementation along with a methodology to significantly increase the efficiency of the compressor/turbine system is highly desirable as well as attractive financially. Moreover, a mechanism that increases the mass of air or other fluid within a compression system within a given volume or as well in converting heat energy to mechanical energy while eliminating the typical turbine system is also highly desirable as well as attractive financially. 
     It would therefore be desirable to have a vane-less compressor/turbine system that overcomes the above cited deficiencies. 
     It would therefore also be desirable to have a vane-less compressor/turbine system that provides a more aerodynamic energy conversion mechanism. 
     The present invention sufficiently accomplishes these means. 
     SUMMARY OF THE INVENTION 
     A hybrid drive engine is provided having a compressor section, a turbine section and a central shaft that mechanically links both the compressor and turbine sections, the compressor and turbine sections each having a plurality of symmetrical annular disks (disk stack) that are each aerosculpted (i.e. aerodynamically configured) for aerodynamic effects of lift. The compressor disks of the disk stack are each configured to cause aerodynamic lift off of a periphery of the disks/disk stack, while the turbine disks of the disk stack are configured to cause aerodynamic lift off of an inner (through) hole of the disks/disk stack. The aerodynamic nature of the disks/disk stack cause each disk thereof to form two opposing airfoil shapes either head to head or trailing edge to trailing edge across the through hole. The present hybrid drive engine increases efficiency and advantage of the shear disk design that capitalizes on boundary layer air flow and adhesion with the Bernoulli advantage of lift without over complicating the design of the disks/disk stack. 
     In one form, the compressor section has first and second compressor portions each having a compressor disk stack of a plurality of aerosculpted compressor disks, while the turbine section has first and second turbine portions each having a turbine disk stack of a plurality of aerosculpted turbine disks. 
     In a particular form, an aerosculpted disk has an annular body defining an inner perimeter, an outer perimeter, and an airfoil having a cross-section comprising a line defining a lower surface, a convex line defining an upper surface and reaching a zenith that is the highest point on the airfoil section of the aerosculpted disk, and a separator lip on the upper surface and located on or proximate the outer perimeter, the separator lip extending to a narrow peak that is higher than an immediately adjacent portion of the upper surface. The separator lip may be lower than the zenith of the convex surface. 
     In a particular form, an aerosculpted disk has an annular airfoil defined by an upper surface and a lower surface, an axis of revolution, a projected reference plane that is normal to the axis of revolution, and an inner perimeter and outer perimeter, the annular airfoil configured such that at least a portion of the annular airfoil has a negative airfoil angle relative to the projected reference plane such that at, least in the portion, the outer perimeter is lower than the inner perimeter when the projected reference plane is horizontal and the disk is oriented with the upper surface uppermost, thereby compensating for air downwash and balancing the lift fore and aft in the disk. 
     The present hybrid drive engine operates in the following manner. Air (fluid) enters an inlet pulled by a fan. This air is then drawn in by the compressor disks which compress the air. The compressed air is then ported through a one way flow diffuser before being channeled into combustors and then being ported into tangential inlets to the turbine section. Upstream of the tangential inlets to the turbine are inlet ports for a combustor, igniter, and flame. The net result of the combustor flame, the high density compressor air and the adhesion-viscosity is that the now expanding tangential exhaust fluid onto the turbine stack turns the shaft. The shaft is physically linked to the compressor and fan. When the shaft turns it provides power to sustain operation of the hybrid drive engine. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       The above mentioned and other features of this invention, and the manner of attaining them, will become apparent and the invention itself will be better understood by reference to the following description of embodiments of the invention taken in conjunction with the accompanying drawings, wherein: 
         FIG. 1  is a side plan view of a hybrid drive engine fashioned in accordance with the principles of the present invention; 
         FIG. 2  is a front plan view of the hybrid drive engine of  FIG. 1  taken along line  2 - 2  thereof; 
         FIG. 3  is an enlarged sectional view of the hybrid drive engine of  FIG. 1 ; 
         FIG. 4  is a top plan view of an end disk for a compressor disk stack of the present invention; 
         FIG. 5  is a top plan view of an end disk for a turbine disk stack of the present invention; 
         FIG. 6  is an exploded view of a compressor or turbine disk stack of the present invention; 
         FIG. 7  is a top plan view of a compressor disk stack situated within a scroll duct of the present invention; 
         FIG. 8  is a side perspective view of the compressor disk stack situated within the scroll duct as shown in  FIG. 7 ; 
         FIG. 9A  is a top plan view of an aerodynamically shaped compressor disk of the compressor disk stack; 
         FIG. 9B  is a side sectional view of the aerodynamically shaped compressor disk of the compressor disk stack as shown in  FIG. 9A  taken along line  9 B- 9 B thereof; 
         FIG. 10  is a top plan view of a turbine disk stack situated within a scroll duct of the present invention; 
         FIG. 11  is a side perspective and exploded view of the turbine disk stack and scroll duct as shown in  FIG. 10 ; 
         FIG. 12A  is a top plan view of an aerodynamically shaped turbine disk of the turbine disk stack; 
         FIG. 12B  is a side sectional view of the aerodynamically shaped turbine disk of the compressor disk stack as shown in  FIG. 12A  taken along line  12 B- 12 B thereof; 
         FIG. 13  is an enlarged side view of labyrinth seal disks of the present invention that prevent compressed air or exhaust gas from escaping the present hybrid drive engine; 
         FIG. 14  is an exemplary embodiment of a torrid section or combustor for the present hybrid drive engine; 
         FIG. 15A  is a side sectional view of an embodiment of an aerodynamically shaped turbine disk in accordance with the present principles; 
         FIG. 15B  is a side sectional view of an embodiment of an aerodynamically shaped turbine disk in accordance with the present principles; 
         FIG. 15C  is a side sectional view of an embodiment of an aerodynamically shaped turbine disk in accordance with the present principles; 
         FIG. 15D  is a side sectional view of an embodiment of an aerodynamically shaped turbine disk in accordance with the present principles; 
         FIG. 16  is an enlarged partial, sectional view of an embodiment of an aerodynamically shaped turbine disk in accordance with the present principles; 
         FIG. 17A  is a side sectional view of an embodiment of an aerodynamically shaped compressor disk in accordance with the present principles; 
         FIG. 17B  is a side sectional view of an embodiment of an aerodynamically shaped compressor disk in accordance with the present principles; 
         FIG. 17C  is a side sectional view of an embodiment of an aerodynamically shaped compressor disk in accordance with the present principles; 
         FIG. 17D  is a side sectional view of an embodiment of an aerodynamically shaped compressor disk in accordance with the present principles; 
         FIG. 18A  is a plan view of a labyrinth end disk in accordance with the present principles; 
         FIG. 18B  is a sectional view of the labyrinth end disk of  FIG. 18A  taken along line  18 B- 18 B thereof; 
         FIG. 18C  is an enlarged view of a portion of the sectional view of the labyrinth end disk of  FIG. 18B  taken around circle  18 C- 18 C thereof; 
         FIG. 19A  is a top plan view of a labyrinth wall plate in accordance with the present principles; 
         FIG. 19B  is a sectional view of the labyrinth wall plate of  FIG. 19A  taken along line  19 B- 19 B thereof; 
         FIG. 20A  is a top plan view of a shaft wall plate/seal in accordance with the present principles; and 
         FIG. 20B  is a sectional view of the shaft wall plate/seal of  FIG. 20A  taken along line  20 A- 20 A thereof. 
     
    
    
     Although the drawings represent embodiments of various features and components according to the present invention, the drawings are not necessarily to scale and certain features may be enhanced in order to better illustrate and explain the present invention. The exemplifications set out herein thus illustrate embodiments of the invention, and such exemplifications are not to be construed as limiting the scope of the invention in any manner. 
     DETAILED DESCRIPTION 
     Those of skill in the art will understand that various details of the invention may be changed without departing from the spirit and scope of the invention. Furthermore, the foregoing description is for illustration only, and not for the purpose of limitation, the invention being defined by the claims. 
       FIGS. 1 and 2  depict two views of an exemplary configuration or embodiment of a hybrid drive (“hydrive”) engine generally designated  10 , fashioned in accordance with the present principles. The hybrid drive engine  10  is a “bladeless” or “vane-less” propulsion unit that utilizes aerodynamically configured (“aerosculpted”), annular spinning disks to yield a dynamic response from the resulting slipstream as described herein. 
     The hybrid drive engine  10  includes a housing  12  fashioned from a suitable material that can withstand the various pressures and other parameters of an engine. Without being exhaustive, suitable materials include aluminum, plastic, steel, titanium, other metal, metal alloy, or ceramic. Other non-listed materials may be used and are contemplated. The hybrid drive engine  10  includes an air inlet or fan intake  13  that is shaped to receive and funnel air into a fan  14  that directs the incoming air into a first stage compressor  16 . The first stage compressor  16  compresses air via a first compressor disk stack  29  of aerodynamically configured (“aerosculpted”) disks  30  (see e.g.  FIG. 3 ). The first compressor disk stack  29  receives an airflow normal to the plane of the disk stack  29  and provides a tangential flow of compressed air. As shown in  FIG. 2 , the hybrid drive engine  10  includes a starter motor  15  that is coupled to a central core shaft  24  on which the fan  14  is situated for initially rotating the core shaft  24  and the fan  14  to provide initial air flow into the core. The starter motor  15  is preferably, but not necessarily, an electric motor that is connected to a battery or other electricity source (not shown). It should be appreciated that the fan starter can be gear driven, belt driven, link driven, a combination thereof or driven by other means. The fan  14  could be sized small for low airflow into the core, sized medium for intermediate airflow into the core, or sized large for high airflow into the core and bypass air. Alternatively, the present hybrid drive engine  10  may be designed without a fan for self-entrained airflow into the core. 
     The hybrid drive engine  10  further includes a second stage compressor  18  that receives air compressed by and from the first stage compressor  16  via a duct  11 . The duct  11  is situated to receive the tangential flow of fluid (e.g. airflow) from the first stage compressor  16  and provide that airflow normal to the second stage compressor  18 . The second stage compressor  18  compresses the air previously compressed by the first stage compressor via a second compressor disk stack  38  of the aerodynamically configured (“aerosculpted”) compressor disks  30  (again, see e.g.  FIG. 3 ). The second stage compressor  18  is in communication with a scroll duct  17  that receives two tangential flows of further compressed air from the second stage compressor  18 . 
     The scroll duct  17  directs a first further compressed fluid (e.g. airflow) from the second stage compressor  18  to a first combustor or torrid section  20   a , while as second further compressed fluid (e.g. airflow) from the second stage compressor  18  to a second combustor or torrid section  20   b . The first combustor  20   a  includes a first valve  19   a  that valves fuel from a fuel source (not shown) into the combustor  20   a  in order to affect combustion therein and heat the portion of fluid tangentially flowing from the second compressor  18  into the first combustor  20   a . The second combustor  20   b  includes a second valve  19   b  that valves fuel from a fuel source (not shown) into the combustor  20   b  in order to affect combustion therein and heat the fluid tangentially flowing from the second compressor  18  into the second combustor  20   b.    
     An exemplary combustor or igniter generally designated  20  representing first and second combustors/igniters  20   a ,  20   b  is shown in  FIG. 14 , the combustor/igniter  20  provides or is a part of a high energy ignition system to ignite the mixed air/fuel mixture. The combustor  20  has a body  110  that is mounted to a frame that preferably may, but may not, be connected to the housing or casing  12  of the hybrid drive engine  10 . Both fuel and air enter through duct  17  and enter the igniter at  111 . Inside the body  110  the air is entrained and mixed with the fuel. The heated airflow (hot air/gas) exits via the outlet duct  25  and is further mixed with the air. The heater airflow is then directed into the ducting leading into the first or primary turbine section  44 . It should be appreciated that the combustor  20  may take different forms and configurations depending on the type of fuel used for combustion (i.e. heating the airflow). 
     Referring back to  FIGS. 1 and 2 , the first combustor  20   a  is in communication with a first outlet duct  25   a  which is in communication with a first stage turbine  26 , while the second combustor  20   b  is in communication with a second outlet duct  25   b  which is likewise in communication with the first stage turbine  26 . The first stage turbine  26  expands air via a first turbine disk stack  40  of aerodynamically configured (“aerosculpted”) turbine disks  41  (see  FIG. 3 ). The heated and further compressed air from the first combustor  20   a  is provided tangentially to the first stage turbine  26 , while the heated and further compressed air from the second combustor  20   b  is provided tangentially to the first stage turbine  26  preferably, but not necessarily, opposite to the tangential airflow from the first outlet duct  25   a . Air ducting  27 , in communication with the first stage turbine  26 , receives two tangential flows of expanded air from the first stage turbine  26  and provides that airflow tangentially to a second stage turbine  28 . The second stage turbine  28  further expands the an via a second turbine disk stack  44  of aerodynamically configured (“aerosculpted”) turbine disks  41  (again, see e.g.  FIG. 3 ). The air ducting  27  provides the two flows of expanded air from the first stage turbine  26  tangentially into the second stage turbine  28 . Output from the second stage turbine  28  is provided normal to the second turbine disk stack  44  via outlet  21 . 
     A sectional view of the present hybrid drive engine  10  without the combustors and associated ducting or the starter motor is depicted in  FIG. 3 . As seen, a central or core shaft  24  extends the length of the housing  12  from the fan  14  through the second stage turbine  28 , while an outer shaft  23  is provided between the compressor section  18  and the turbine section  28 . Without being exhaustive, the core shaft  24  and/or shaft  23  may be made from aluminum, steel, titanium, other metal, and/or metal alloy. Other non-listed materials may be used and are contemplated. 
     As indicated above, the present hybrid drive engine  10  has a first or primary compressor section  16 , a second or secondary compressor section  18 , a first or primary turbine section  26 , and a second or secondary turbine section  28  each of which defines a respective internal chamber that holds a compressor disk stack or a turbine disk stack. It is important for the chambers of each section to be sealed with respect to each other and to the ambient. In furtherance of this, each chamber includes a chamber seal structure  55  and a shaft seal structure  106  each of which has labyrinth structures. These seal structures prevent the air or hot gases from exiting the casing/housing  12  in an improper manner (e.g. following the paths of least resistance to escape the casing of the respective section). 
     With additional reference to  FIGS. 13, 18A -C, and  19 A-B, the seal structure  55  and its associated components are shown. The seal structure  55  includes a wall plate  102  that is formed, made or provided in the chamber casing (wall) and an end disk  100  that is a first disk of the first and second compressor disk stacks  29 ,  38  and of the first and second turbine disk stacks  40 ,  44 . Labyrinth features of the seal structure  55  (i.e. the end disk  100  and the wall plate  102 ) helps to prevent build-up of air frictional forces that reduce the efficiency of the hybrid drive engine  10 . 
     As particularly seen in  FIGS. 19A and 19B , the wall plate  102  includes a central opening  120  for allowing the central shaft  24  to pass through the wall plate  102 . A sealing structure  105  comprising a plurality of perpendicular fingers  121  extend radially outwardly from and annularly about the opening  120 . A first annular projection  122  is radially outwardly adjacent the fingers  121  with a first annular trough  123  radially outwardly adjacent the first annular projection. A second annular projection  124  is radially outwardly adjacent the first annular trough  123 , while a second annular trough  125  is radially outwardly adjacent the second annular projection  124 . A third annular projection  126  is radially outwardly adjacent the second annular trough  125 , while a third annular trough  127  is radially outwardly adjacent the third annular projection  126 . Lastly, a fourth annular projection  128  is radially outwardly adjacent the third annular trough  127 . The fourth annular projection  128  provides a peripheral rim to the wall plate  102 . 
     As particularly seen in  FIGS. 18A, 18B and 18C , the end disk  100  of the seal structure  55  has a generally annular shape having a central opening  133  with a spoke structure  130  extending from the sides of the opening  133  into the center. A configured bore  131  is provided in the center or the spoke structure  130  for receiving the central shaft  24 . The end disk  100  also has labyrinth features that cooperate with the labyrinth features of the wall plate  102  or shaft seal  106  as appropriate. Adjacent the opening  120  is a sealing structure  104  consisting of a plurality of annual fingers  136  that project normal to the plane of the disk  100 . While six (6) fingers  136  are shown, there may be more or less fingers as desired and/or necessary to provide sealing. The fingers  136  and the fingers  121  of the wall plate  102  are spaced to mesh with one another (see  FIG. 13 ). The end disk  100  further has a planar portion  132  radial to the sealing structure  104 . The planar portion  132  extends from the radially outmost finger  136  to the outer periphery of the end disk  100 . 
     In  FIG. 13 , an enlarged portion of the wall plate  102  and the end plate  100  of the sealing structure  55  is shown. This enlargement shows how the fingers of the two sealing structures  104 ,  105  mesh as well as the interaction of the planar portion  132  of the end disk  100  and the configured troughs and projections of the wall plate. Without being exhaustive, the two sealing structures themselves can be made of metal or ceramic or some such abraidable material, or other metal, or metal alloy. Other non-listed materials may be used and are contemplated. The arrows  101  and  103  indicate entrained air that is prevented from escaping the engine section due to the labyrinth seals/plates  100 ,  102 . 
     The shaft seal or disk  106  is particularly shown in  FIGS. 20A and 20B . The disk  106  has a central bore  151  through which the shaft  24  extends. The disk  106  has a series  150  of annular projections  152  and annular notches  153  that extend radially outwardly from the central bore  151 . An outermost annular projection  158  is provided at the periphery of the disk  106 . The projection  158  is larger in width than the other projections  152 . 
     The top casing for the compressor sections and the end casing for the turbine sections have the matching labyrinth seals to prevent the compressed air or hot gasses from exiting the casing. Each labyrinth end disk  104  is thicker than the rest of the disks. The exhaust nozzle  21  section has a sealing structure  55  as seen in  FIG. 3 . However, the top or start of the turbine disk stack which is forward most on the stack, has only a thick wall shaft plate  106  made into the casing and a regular end disk without labyrinth features. 
     Referring now to  FIG. 4 , there is depicted an end disk for the compressor disk stacks  29 ,  38 , generally designated  31 . The end disk  31  is defined by a generally planar disc of a particular thickness having a central opening  33  for the central shaft  24  and three (3) configured openings  32   a ,  32   b ,  32   c  that are radially outwardly positioned relative to and from the central opening  33 . The openings  32   a ,  32   b ,  32   c  allow incoming air to flow through the disc. If desired, the openings  32   a - c  may be shaped differently if desired and/or there may be more or less openings. Additionally, the end disk  31  has three (3) radially inward threaded bores  34   a ,  34   b ,  34   c  and three (3) radially outward threaded bores  36   a ,  36   b ,  36   c  that are all configured to receive threaded fasteners for attaching the end disk  31  to the compressor disk stack. 
       FIG. 5  depicts an end disk for the turbine disk stacks  40 ,  44  generally designated  42 . The end disk  42  is defined by a generally planar disc of a particular thickness having a central opening  43  for the central shaft  24  and three (3) configured openings  32   a ,  32   b ,  32   c  that are radially outwardly positioned relative to and from the central opening  43 . The openings  32   a ,  32   b ,  32   c  allow incoming air to flow through the disc. If desired, the openings  32   a - c  may be shaped differently if desired and/or there may be more or less openings. Additionally, the end disk  31  has three (3) radially inward threaded bores  34   a ,  34   b ,  34   c  and three (3) radially outward threaded bores  36   a ,  36   b ,  36   c  that are all configured to receive threaded fasteners for attaching the end disk  42  to the turbine disk stack. 
       FIG. 6  shows an exploded view of the compressor disk stacks  29 ,  38  and the turbine disk stacks  40 ,  44 . Each disk stack includes a respective end disk  31 ,  42  on the top and bottom of the disk stack. Situated between the end disks  31 ,  42  is a plurality of aerosculpted compressor disks  30  (for the compressor disk stack) or aerosculpted turbine disks  41  (for the turbine disk stack) details of which are described below. In general, each aerosculpted disk  30 ,  41  includes three (3) radially inward threaded bores  48   a ,  48   b ,  48   c  that correspond in placement to the three (3) radially inward threaded bores  34   a ,  34   b ,  34   c  of the end disks  31 ,  42 , as well as three (3) radially outward threaded bores  46   a ,  46   b ,  46   c  that correspond in placement to the three (3) radially outward threaded bores  36   a ,  36   b ,  36   c  of the end disks  31 ,  42 . Spacers  60  are provided over each radially inward and outward threaded bores  48   a - c ,  46   a - c  and thus between each disk. Threaded fasteners (not shown) are used to attach the disks of the disk stack together. Other fastening means may be used. It should also be appreciated that while nine (9) aerosculpted compressor/turbine disks  30 ,  41  are shown, more or less aerosculpted disks may be used. 
       FIGS. 7 and 8  depict two views of either one of the compressor disk stacks  29 ,  38  situated within the engine casing and particularly with respect to either one of the respective scroll ducts  16 ,  17 . The compressor disk stack  29 ,  38  receives a normal airflow through the three (3) configured openings  32   a ,  32   b ,  32   c  thereof. The aerosculpted disk stack  29 ,  38  then compressed the entrained air and exhausts it through duct portion  16   a ,  17   a  which is then, in the case of scroll duct  6 , outlet into air duct  11  and ported to the inlet of the compressor disk stack  38 . A compressed airflow from the compressor stack  38  is then tangentially discharged to and through duct portions  17   a ,  17   b . This compressed airflow may be provided to the combustor/igniter  20  or directly to the first turbine stage  26  (in bypass air flow operation). 
       FIGS. 9A and 9B  depict two vie of an aerodynamically shaped (“aerosculpted”) compressor disk  30  of either one of the compressor disk stacks  29 ,  38 . The aerosculpted compressor disk  30  is defined by an annular disc having a central opening  45 . An annular planar portion  49  is provided radially adjacent the central opening  45 . Three (3) radially inward threaded bores  48   a ,  48   b ,  48   c  are positioned about the planar portion  49  for receiving threaded fasteners not shown). The compressor disk  30  further includes a slanted or angled peripheral portion  47 . Three (3) radially outward threaded bores  46   a ,  46   b ,  46   c  are positioned about the angled peripheral portion  47  for receiving threaded fasteners (not shown). Further features and characteristics of the aerosculpted compressor disk are described below with reference to  FIGS. 17A-D . The arrows shown in  FIG. 9B  represent air pressure exerted upon the present aerosculpted compressor disk  30 . 
       FIGS. 10 and 11  depict two views of either one of the turbine disk stacks  40 ,  44  situated within the engine casing and particularly with respect to the tangential ducting  25  thereof. The turbine disk stack  40  receives a tangential airflow through duct portion  25   a ,  25   b  which is then expanded by the aerosculpted disks of the turbine disk stack  40 . An expanded airflow from the turbine stack  40  is then normally discharged to and through duct portion  27 . This expanded airflow is then provided tangentially to the second turbine stage  44 . 
       FIGS. 12A and 12B  depict two views of an aerodynamically shaped (“aerosculpted”) turbine disk  41  of either one of the turbine disk stacks  40 ,  44 . The aerosculpted turbine disk  41  is defined by an annular disc having a central opening  45   a . An annular planar portion  52  is provided radially adjacent the central opening  45   a . Three (3) radially inward threaded bores  48   a ,  48   b ,  48   c  are positioned about the planar portion  52  for receiving threaded fasteners (not shown). The turbine disk  41  further includes a slanted or angled peripheral portion  51 . Three (3) radially outward threaded bores  46   a ,  46   b ,  46   c  are positioned about the angled peripheral portion  51  for receiving threaded fasteners (not shown). Further features and characteristics of the aerosculpted turbine disk are described below with reference to  FIGS. 15A-D . The arrows shown in  FIG. 12B  represent air pressure exerted upon the present aerosculpted turbine disk  41 . 
     Referring to  FIGS. 15A-D  various embodiments of the aerodynamically configured turbine disk  41  are shown in cross section to illustrate an area extending from at outer perimeter  61  to an inner perimeter  64 . Each figure illustrates an upper convex surface  63  originating at the outer perimeter  61  that joins to a leading edge  59  of aerodynamically protruding annular fin  62 . The leading edge  59  is shown as having a juncture with a descending edge  58  which, turn, has a juncture with a straight upper  65 . The straight upper surface  65  terminates at the inner perimeter  64 . An outer skin  66  is also illustrated in these figures. 
     The aerodynamically configured turbine disk  41   a  as shown in  FIG. 15A  particularly illustrates the underside of the annular, aerodynamically configured turbine disk having a lower convex surface  67 , a straight vertical undercut  68 , and a lower cutaway surface  69 , wherein the lower cutaway surface  69  is parallel to the horizontal plane. A central convex surface  71  terminates at the inner perimeter  64 . 
     The aerodynamically configured turbine disk  41   b  as shown in  FIG. 15B  particularly illustrates the underside of the annular, aerodynamically configured turbine disk having a lower convex surface  67 , a straight vertical undercut  68 , and a lower cutaway surface  69 , wherein the lower cutaway surface  69  is at an angle to the horizontal plane. A central convex surface  71  terminates at the inner perimeter  64 . 
     The aerodynamically configured turbine disk  41   c  as shown in  FIG. 15C  particularly illustrates the underside of the annular, aerodynamically configured turbine disk having a lower convex surface  67 , a curved vertical undercut  68 , and a lower cutaway surface  69 , wherein the lower cutaway surface  69  is at an angle to the horizontal plane. A central convex surface  71  terminates at the inner perimeter  64 . 
     The aerodynamically configured turbine disk  41   d  as shown in  FIG. 15D  particularly illustrates the underside of the annular, aerodynamically configured turbine disk having a lower convex surface  67 , an angled vertical undercut  68 , and a lower cutaway surface  69 , wherein the lower cutaway surface  69  is parallel to the horizontal plane. A central convex surface  71  terminates at the inner perimeter  64 . 
     It should be appreciated that other embodiments of an annular, aerodynamically configured turbine disk may be fashioned in accordance with the present principles. For instance, and without being exhaustive, an annular, aerodynamically configured turbine disk may include an underside having a lower convex surface, a curved vertical undercut, and a lower cutaway surface, where the lower cutaway surface is parallel to the horizontal plane. Another annular, aerodynamically configured turbine disk may include an underside having a lower convex surface, an angled vertical undercut, and a lower cutaway surface, where the lower cutaway surface is parallel to the horizontal plane. A further annular, aerodynamically configured turbine disk may include an underside having a lower convex surface, an angled vertical undercut, and a lower cutaway surface, where the lower cutaway surface is at an angle to the horizontal plane. A yet further annular, aerodynamically configured turbine disk may include an underside having a straight lower surface, an angled vertical undercut, and a lower cutaway surface, where the lower cutaway surface is at an angle to the horizontal plane. A still further annular, aerodynamically configured turbine disk may include an underside having a straight lower surface, a curved vertical undercut, and a lower cutaway surface, where the lower cutaway surface is parallel to the horizontal plane. An even further annular, aerodynamically configured turbine disk may include an underside having a straight lower surface, a curved vertical undercut, and a lower cutaway surface, where the lower cutaway surface is at an angle to the horizontal plane. A yet further annular, aerodynamically configured turbine disk may include an underside having a straight lower surface, a straight vertical undercut, and a lower cutaway surface where the lower cutaway surface is parallel to the horizontal plane. A still further annular, aerodynamically configured turbine disk may include an underside having a straight lower surface, a straight vertical undercut, and a lower cutaway surface, where the lower cutaway surface is at an angle to the horizontal plane. 
     Referring to  FIG. 16 , a section of the turbine disk  41  as illustrated in  FIG. 12A  is shown in cross-section, taken along line  16 - 16  thereof. The section shows the outer perimeter  61 , an inner perimeter  64 , and an outer skin  66 . The upper surface of the turbine disk  41  has a convex upper surface  63  that extends from the outer perimeter  61  to a juncture with the leading edge  59  of aerodynamic protruding annular fin  62 . Leading edge  59  extends to a juncture with descending edge  58  of the aerodynamic protruding annular fin  62 . The descending edge  58  terminates at a juncture with a straight upper surface  65  which terminates at an inner perimeter  64 . The underside of the turbine disk  41  has a lower convex surface  67  that originates at an outer perimeter  61  and terminates at a juncture with a straight vertical undercut  68 . The straight vertical undercut  68  connects the lower convex surface  67  with a lower cutaway surface  69 . The lower cutaway surface  69  originates at the upper extremity of the straight vertical undercut  68  and terminates at a juncture with central convex surface  71 . The central convex surface  71  terminates at the inner perimeter  64 . 
     In operation, air (or other fluid) is directed into the turbine disks/disk stack where the convex upper surface  63  diverts airflow in an upward direction, thereby increasing the speed at which the air is traveling. This results in a decrease in air pressure above the annular, aerodynamically configured turbine disk  41 . When this airflow strikes aerodynamic protruding annular fin  62 , it is now more deflected upward, but more sharply than the first deflection. This diversion increases air speed and reduces air pressure once more. At the same time, air passing on the lower side of the annular, aerodynamically configured turbine disk  41 , which includes the lower convex surface  67 , the straight vertical undercut.  68 , and the lower cutaway surface  69 , is captured beneath the unit, thereby reducing speed and increasing upward air pressure. 
     The turbine disk  41  may be made from various materials. Without being exhaustive, these include aluminum, plastic, steel, titanium, other metals, metal alloys, ceramic, glass, and/or a combination of these. The turbine disk  41  may be manufactured by a HIP (Hot Iostatic Press) method. 
     Referring to  FIGS. 17A-D , various embodiments of the aerodynamically configured annular compressor disk  30  are shown in cross section to illustrate the airfoil effects thereof in accordance with the present principles. The annular, aerodynamically configured compressor disk  30   a  of  FIG. 17A  illustrates a preferred embodiment. The cross-section embodies a line  76  defining a lower surface and a convex line  75  defining an upper surface. In accordance with the present principles, a separator lip  77  is provided on the outer perimeter of the upper surface. The lip  77  extends upward to a narrow peak  78  that is higher than the immediately adjacent portion of the upper surface  75  of the airfoil (disk). The lip  77  provides stability during rotation over a wide range of air velocities. 
     Moreover, the lip  77  is termed a separator lip in that it is believed that the lip causes the airflow to separate from the leading edge of the forward portion of the airfoil. It is further believed that the separator lip  77  reduces the lift slope of the forward portion of the airfoil so that it becomes balanced with the lift slope of the aft portion or the disk. The lift slope is the rate of change of lift versus angle of incidence or dL/dA (ΔL/ΔA) where L=lift and A=angle of incidence. It is further believed that the lift slopes of the forward aft sections of the aerodynamically configured annular disk have become matched (due to the action of the separator lip) because the aerodynamically configured annular disk is stable over a wide range of airflow velocities and angle of incidence. 
     It has been also been discovered by the inventor that an important parameter of the separator lip  77  is that it must have a narrow peak  78  in order to produce stable rotation as described above. A preferred width of the peak is less than one millimeter (1 mm). However, other widths may be used. A preferred embodiment has the peak  78  is substantially defined by the joining together of the surfaces  79  and  80  immediately adjacent to the peak  78 . For stable spinning (or flight), the angle  81  between the adjacent surfaces  79 ,  80  should be less than 60 degrees (60°). 
     It has also been discovered by the inventor that an important parameter of the separator lip  77  is the angle  82  formed between a line tangent to an outer surface  79  of the lip  77  and the axis of revolution of the disk. If this angle is too great, stable spinning will not be maintained over a wide range of velocities. As the angle  82  is increased, there is a reduction of its stability. For example, a disk with an angle of 45 degrees (45°) was found to have less stability than other disks with smaller angles. In a preferred embodiment, this angle is approximately 30 degrees (30°). 
     Other angles  82  are illustrated in  FIGS. 17B and 17C .  FIG. 17B  shows a compressor disk  30   b  with an angle  82  of zero degrees (0°) while  FIG. 17C  shows a compressor disk  30   c  with an angle  82  of minus 30 degrees (−30°). These disks ( 30   b ,  30   c ) are stable but have a less stable rotation than the preferred embodiment disk  30   a  of  FIG. 17A . While the sectional views illustrate a straight line defining the outer edge of the lip  77 , which creates a conical surface, it is believed that stable spinning can also be achieved if this line was curved—provided that the peak of the lip  77  is narrow. 
     Another important parameter of the compressor disks is the line defining the upper surface  75  of the airfoil section is convex in order to develop adequate lift combined with stability and low drag. In a preferred embodiment, the zenith of the convex upper surface  75  is the highest point on the airfoil section. It was determined that best results were achieved when this zenith is closer to the inner perimeter than to the outer perimeter. The preferred location for this zenith was discovered to be about one-third (⅓) of the distance from the inner perimeter to the outer perimeter. 
     As shown in  FIG. 17A , the airfoil section of the compressor disk  30   a  has a substantially straight line  76  defining a substantially flat lower surface except for a downwardly depending flap  83  in the region of the outer perimeter of the lower surface. It has been determined that this flap aids in achieving balanced spinning. The flap  83  is also illustrated in the alternative sections of the compressor disks  30   b  and  30   c  shown in  FIGS. 17B and 17C  respectively. 
       FIG. 17D  illustrates a compressor disk  30   d  having an alternative to the flap  83 . Particularly, the compressor disk  30   d  defines an angled airfoil in which the inner perimeter is higher than the outer perimeter. It has been discovered that either this higher inner perimeter, or the flap  83 , or a combination of these features is needed to achieve stable spin. 
     It is alternatively correct in describing the separator lip  77  and the flap  83  of the compressor disk  30  to indicate that the compressor disk  30  includes an outer rim  84  adjacent to its outer perimeter. This rim  84  is comprised of an outer rim surface  79  extending from a bottom edge  83  below the lower airfoil surface  76  to a top edge  78  above the outer portion of the upper airfoil section  75 , an upper-inner rim surface  80  extending downward from the top edge  78  to the outer portion of the upper airfoil surface  75 , and a lower rim surface  85  extending upward from the bottom edge  83  to the lower airfoil surface  76 . 
     The compressor disk  30  may be made from various materials. Without being exhaustive, these include aluminum, plastic, steel, titanium, other metals, metal alloys, ceramic, glass, and/or a combination of these. The compressor disk  30  may be manufactured by the HIP method. 
     A disk stack in accordance with the present principles may have an inner perimeter that is higher than the outer perimeter. The airfoil section of an aerosculpted compressor or turbine disk may have a downwardly depending flap adjacent the outer perimeter. A line tangent to the outer surface of the separator hp is within plus or minus 45 degrees (+/−45°) of parallelism to the axis of revolution of the disk. The inner and outer perimeters of the aerosculpted disk may be circles described about the axis of revolution. Moreover, an aerosculpted disk has an angular moment capable of assisting its aerodynamic lift. Furthermore, the upper and/or lower surfaces of an aerosculpted turbine and/or compressor disk may be textured in order to improve aerodynamic performance and boundary layer entrainment. Still further, a convex line of an aerosculpted compressor and/or turbine disk, defining the upper surface thereof, reaches a zenith at a location that is substantially one third (⅓) of the distance from the inner perimeter to the outer perimeter. 
     With respect to the airfoil of the present aerosculpted compressor and turbine disks, in one form the annular airfoil angle is computed from the following formula: 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/D 2 ), where K=45±15, W=weight of the disk in ounces, and D=mean diameter of the annulus in inches. 
     In one form, the annular airfoil aerosculpted disk) is configured with a negative airfoil angle such that the revolution of a chord length of a chord line passing through an inner and outer perimeter of the disk defines the angled surface of a frustum of a cone. In this manner, in rotation, the forward portion of the annular airfoil is at a lower angle of incidence to the airflow path than the remainder of the annular airfoil, thereby compensating for air downwash effects from the forward portion and balancing the aerodynamic lift fore and aft in the compressor/turbine disk or disk stack. The aerosculpted disks typically, but not necessarily, have a weight of less than 2.0 ounces per square inch of projected area, thereby permitting a substantially level spin at speeds below 100 feet per second. 
     In another form, the airfoil angle is determined by the following formula: α=K W/D 2 , 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. For example, a compressor or turbine disk may have the following dimensions: weight=2 to 4 ounces, mean diameter=8 to 12 inches, chord length=1 to 3 inches, thickness of 0.05 to 0.20 inches, and airfoil angle=1 to 2 degrees. 
     In another form, the airfoil angle is determined by the following formula: α=K W/V 2  D2, where α=airfoil angle in degrees, W=the weight of the compressor or turbine disk in ounces, V=intended rotation velocity in feet per second, and D=mean diameter of the annulus in inches. 
     In another form, the airfoil angle is determined by the following formula: 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/D 2 ) where K=45±15, W=the weight of the compressor or turbine disk in ounces, and D=mean diameter of the annulus in inches. 
     While not shown in the figures, and without being exhaustive, the present hybrid drive engine  10  could be alternately designed with low bypass air into the combustor for cooling, with intermediate bypass air into the combustor for cooling and thrust augmentation, and/or with high bypass air into the combustor for cooling, additional compressed air requirements and thrust augmentation. Additionally, the present hybrid drive engine  10  may be designed with reverse flow combustors, forward flow combustors, two (2) combustors, multiple combustors, can-ular combustors, and/or can-annular combustors. Moreover, the present hybrid drive engine  10  may be designed with a single compressor diverter to the combustor, with multiple compressor diverters to the combustor, or with two (2) compressor diverters to the combustor. Furthermore, the present hybrid drive engine  10  may be designed with a single stage compressor, a dual stage compressor, a multiple stage compressor, a single stage turbine, a dual stage turbine, or a multiple stage turbine. Still further, the present hybrid drive engine  10  may be designed with memory metals in each stage of the turbine, only aerosculpted memory metals in each stage of the turbine, with only aerosculpted disks in each stage of the turbine, with some aerosculpted disks in each stage of the turbine, and/or with convex disks in each stage of the turbine. Even further, the present hybrid drive engine  10  may be designed with only aerosculpted disks in each stage of the compressor; with only some aerosculpted disks in each stage of the compressor, with aerosculpted memory metals in each stage of the compressor, and/or with convex disks in each stage of the compressor. 
     It should be appreciated by those skilled in the art that the present hybrid drive engine  10  has potentially different uses, materials, sizes, methods of operation, and forms/embodiments than those explicitly shown and/or described herein. Additionally, the present hybrid drive engine may be used with gases other than air, fluids and/or a combination thereof. The present hybrid drive engine may be designed for a large or small thrust output, a large or small torque output, a large or small specific impulse output, a mega or micro traction output, or a large or small Primary Take-off Shaft output. Uses of the present hybrid drive engine include, without being exhaustive, an aircraft jet engine, an automobile engine, a watercraft engine, a locomotive engine, a space craft engine, an underwater craft engine, a power generation engine, a construction implement engine, an agricultural implement engine, a snow or ice implement engine, a land-craft engine, a medical implement engine, a robotic engine and power source, a hydraulic engine and power source, a prosthetic engine and power source, a well engine and power source, and a transportation engine and power source. 
     While the invention has been illustrated and described in detail in the foregoing drawings and description, the same is to be considered as illustrative and not restrictive in character, it being understood that only illustrative embodiments thereof have been show and described and that all changes and modifications that are within the scope of the following claims are desired to be protected. 
     All references cited in this specification are incorporated herein by reference to the extent that they supplement, explain, provide a background for or teach methodology or techniques employed herein.