Abstract:
An orbiting engine includes a plurality of combustion chambers ( 52 ) rigidly connected to and disposed equidistantly from a central shaft ( 54 ). The combustion chambers ( 52 ) are oriented so that exhaust gases emerge therefrom tangentially to a circle concentric with the shaft ( 54 ), thereby causing the shaft to rotate. In preferred embodiments, a centrifugal compressor ( 62 ), radially inward from the combustion chambers ( 52 ) and driven by the shaft ( 54 ), compresses air that serves to oxidize fuel injected into the combustion chambers ( 52 ).

Description:
CROSS REFERENCE TO RELATED APPLICATIONS 
     This application is a 371 of PCT/IB97/01462 filed on Sep. 29, 1997 
    
    
     FIELD AND BACKGROUND 
     The present invention relates to rotary engines, and more particularly, to an engine whose power is provided by rocket motors orbiting a central power shaft. 
     Through the Twentieth Century, the efficiencies and power to weight ratios of engines have improved steadily. Modern powerful aerospace gas turbine engines attain thermal efficiencies of 35% and a power to weight ratio of 3 to 4 kW per kg at a cost of $250 per kg. This high cost excludes these engines from the low power (under 500 kW) aircraft vehicular and industrial applications. For example, small airplanes and helicopters are powered only by prison engines to this day. The main factor contributing to this high cost is the complexity of a rotating turbine operating at high speed and high temperature. The main factor contributing to the low efficiency of low power turbine engines is the small dimension of the rotating compressors and turbines. 
     A major limitation of a rotating turbine is the creep characteristics of the turbine blade material. At present, the temperature limit at the turbine inlet of a gas turbine engine is 1500° C. because of creep, although the combustion temperature is much higher (around 2500° C.). This prevents turbine engines, especially small turbine engines, from approaching their theoretical thermodynamic efficiencies. 
     Thus, there is a widely recognized need for, and it would be highly advantageous to have, a low power rotary engine that would overcome these disadvantages of present known turbine engines. 
     SUMMARY OF THE INVENTION 
     According to the present invention there is provided an engine for driving a load shaft, comprising: (a) a power shaft; and (b) a plurality of combustion chambers, each of the chambers being substantially equidistant from the power shaft and rigidly connected to the power shaft, each of the chambers being provided with a nozzle oriented so that exhaust gases exit the chamber therethrough in a direction substantially tangential to a circle concentric with the power shaft. 
     The earliest known reaction motor is the device shown in FIG. 1, invented more than 2000 years ago by Hero of Alexandria for sprinkling water of purification on worshippers in a pagan temple. A spherical boiler  10  is provided with two tangential nozzles  12  at opposite sides of boiler  10  and pointing in opposite directions. Boiler  10  is mounted on a stand  16  so that boiler  10  is free to rotate about an axis perpendicular to the plane defined by nozzles  12 . Boiler  10  is partially filled with water and fire  18  is placed under boiler  10  to boil the water. Water and steam  14  emerging from nozzles  12  causes boiler  10  to rotate. Hero&#39;s device was intended to provide a spray of water, not to provide power. An ordinary garden water sprinkler functions similarly, also to provide a spray of water rather than providing power. The device of the present invention uses reaction motors, disposed circumferentially like nozzles  12  of Hero&#39;s device, to provide power to a central shaft. 
     About fifty year ago, a system similar to Hero&#39;s was tried for helicopter propulsion. In this system, hot gases were forced radially through the blades of the helicopter and exhausted circumferentially from the blade tips. The idea was to rotate the blades using the reaction force of the circumferentially exhausted gas. This system was unable to give sufficient propulsive power at a reasonable efficiency, because the blades could not withstand the high temperature of the hot gases needed for the desired performance. That system differs from the present invention in that that system uses hot gases generated elsewhere to drive the rotation of the blades, whereas the present invention drives a central shaft using reaction motors deployed around a circle concentric with the central shaft. 
     FIG. 2 shows the conceptually simplest, albeit least preferred, embodiment of the present invention. Two reaction motors, shown in FIG. 2 as solid fuel rocket motors  10 , are connected rigidly by struts  22  to a power shaft  24 . Rocket motors  20  are on opposite sides of shaft  24 , and point in opposite directions. As is well known in the art, the bodies of rocket motor  20  serve as combustion chambers  26  for the solid fuel (actually a fuel-oxidizer mixture) contained therein. The exhaust gases produced by burning the solid fuel exit combustion chambers  26  via nozzles  28 . All four struts  22  are of equal length, so that the exhaust gases emerge from nozzles  28  in opposite directions substantially tangent to a circle  30  concentric with shaft  24 . Shaft  24  is supported by bearings (not shown) that allow shaft  24  to rotate about the longitudinal axis thereof, but allow no other motion, so that the reactive forces associated with the emergence of exhaust gases from nozzles  28  apply a torque to shaft  24 , causing shaft  24  to rotate about the longitudinal axis thereof. 
     The embodiment of FIG. 2 illustrates the principle of the present invention, but suffers from several practical deficiencies. The embodiment of FIG. 2 operates only as long as the solid fuel in rocket motors  20  lasts. Furthermore, it is difficult if not impossible to regulate the power output from this embodiment. Therefore, preferred embodiments of the device of the present invention provide continuous, easily regulated flows of fluid fuel and oxidizer to circumferentially mounted combustion chambers. The preferred oxidizer is air, compressed by one or more centrifugal compressors coaxial with shaft  24  and driven by shaft  24 . 
    
    
     BRIEF DESCRIPTION OF THE DRAWINGS 
     The invention is herein described, by way of example only, with reference to the accompanying drawings, wherein: 
     FIG. 1 (prior art) is a schematic perspective view of the reaction motor of Hero of Alexandria; 
     FIG. 2 is a schematic perspective view of the simplest embodiment of the present invention. 
     FIG. 3 is a schematic axial cross-sectional view of a preferred embodiment of the present invention; 
     FIG. 4 is a close-up schematic view of cut A—A of FIG. 3; 
     FIG. 5 is a schematic view of the axis of the device of FIG. 3, showing the differential planetary transmission; 
     FIG. 6 is a diagram of the thermodynamic cycle of the present invention. 
    
    
     DESCRIPTION OF THE PREFERRED EMBODIMENTS 
     The present invention is of an engine in which power is supplied to a central shaft by circumferentially mounted reaction motors. 
     The principles and operation of an orbiting engine according to the present invention may be better understood with reference to the drawings and the accompanying description. 
     Referring now to the drawings, FIG. 3 is a schematic axial cross-section of a preferred embodiment of the present invention. A circular disk  50  is mounted on a hollow power shaft  54 . Around the circumference of disk  50  are mounted combustion chambers  52 . Through hollow shaft  54  runs a compressor shaft  68 . Compressor shaft  68  terminates in several compressor blades  61 . Compressor shaft  68  is supported within hollow shaft  54  and within a compressor housing  60  by several bearings  70 . Bearings  70  allow shaft  68  to turn freely about the longitudinal axis thereof within hollow shaft  54 . Compressor blades  61  together with compressor housing  60  constitute a centrifugal compressor rotor  62 . A radial extension  63  of compressor housing  60  defines a diffuser  64  between radial extension  63  and disk  50 . Diffuser  64  is terminated by a circumferential array of vanes  66  around the periphery thereof. 
     FIG. 4 is a partial schematic cross-sectional view of the preferred embodiment of FIG. 3 along the section A—A. In FIG. 4 are shown two combustion chambers  52 . Combustion chambers  52  are provided with fuel injectors  56 , with longitudinal slits  58  and with nozzles  53 . Nozzles  53  point substantially tangentially to the circle defined by the outer rim of disk  5 . Together, combustion chambers  52  and nozzles  53  constitute reaction motors  51 . 
     In operation, ambient air is sucked in and pressurized by compressor rotor  62  and further pressurized by diffuser  64 . Compressor rotor  62  is designed for an optimal pressure ratio, typically between about 10:1 and about 15:1. The flow of air out of compressor rotor  62  is matched to the power requirements of the engine: roughly 1 kg/sec for 250 sec kW output. Exiting diffuser  64 , the air expands past vanes  66  and enters combustion chambers  52  via slits  58 . Fuel is impelled by centrifugal force via conduits (not shown) in disk  50  to fuel injectors  56 . The compressed air and the fuel are mixed and ignited in combustion chambers  52 , forming a hot exhaust gas that expands through nozzles  53 . The reaction force thus generated turns disk  50  and hollow shaft  54  about the longitudinal axis of hollow shaft  54 . Thus, reaction motors  51  travel in a circular orbit about the common longitudinal axis of shafts  54  and  68 ; hence the name “orbiting engine” for the present invention. The flow of fuel into combustion chambers  52  via fuel injectors  56  is controlled by conventional means to be proportional to the amount of compressed air entering combustion chambers  52  via slits  58 , thereby regulating the combustion temperature. Nozzles  53  are of the convergent-divergent type operating at a relatively high pressure ratio to cause the exit velocity to be supersonic, typically about 1250 meters per second. The number of reaction motors  51  is chosen according to the airflow and power requirements. A typical circumferential speed of reaction motors  51  is about 700 meters per second. 
     Enclosing the circular orbit along which reaction motors  51  travel is a stationary, substantially toroidal housing  80 , sealed from disk  50  by labyrinths  82 . Toroidal housing  80  has two functions: to serve as a heat exchanger to preheat the compressed air entering combustion chamber  52  using the heat of the exhaust gases exiting nozzles  53 , and to insulate the noise generated in combustion chambers  52 . The exhaust gases exit toroidal housing  80  via an exhaust port  84 . 
     FIG. 5 is a schematic diagram of the extension of shafts  54  and  68  past the bottom of FIG. 3, showing how power shaft  54  and compressor shaft  68  are coupled to each other and to a load shaft  90  by a differential planetary transmission  100 , so that the torque generated by the exhaust gases expanding through nozzles  53  is used to drive both compressor rotor  62  and the load placed on load shaft  90 . Power shaft  54  terminates in planetary carrier  102 , to which are attached planetary gears  104 . Compressor shaft  68  terminates in sun gear  106 . Load shaft  90  terminates in ring gear  108 . Together, sun gear  106 , planetary gears  104  and ring gear  108  constitute differential planetary transmission  100 . Power shaft  54  drives compressor rotor  62  through planetary gears  104 , sun gear  106  and compressor shaft  68 . Power shaft  54  drives the load placed on the invention through planetary gears  104 , ring gear  108  and load shaft  90 . This transmission is illustrative: power shaft  54  may be coupled to compressor shaft  68  and load shaft  90  using one of many other transmissions well-known in the art, although differential planetary transmission  100  is the most compact transmission. 
     FIG. 6 is a diagram of the thermodynamic cycle of the preferred embodiment of FIG.  3 . The abscissa of FIG. 6 is entropy S, and the ordinate is temperature T. The dashed lines are isobars. From point  1  to point  2 , air at ambient pressure and temperature is compressed and heated by compressor rotor  62  and diffuser  64 , and enters combustion chambers  52  via vanes  66  and slits  58 . From point  2  to point  3 , air and fuel are burned in combustion chambers  52 , creating a hot exhaust gas. From point  3  to point  4 , the exhaust gas expands through nozzles  53  to near ambient pressure. From point  4  to point  1 , the exhaust gas cools down to ambient temperature. 
     Assuming matched nozzle expansion conditions, the gross power output from the engine, per unit mass flow (energy per unit mass per unit time) is: 
     
       
           E   gross   =W   e   V    (1)  
       
     
     where W e  is the nozzle exit velocity and V is the peripheral engine speed. W e  is given by:                W   e     =       2                   η   n          C   p            T   c          (     1   -     1     δ   r       γ   -   1     γ           )                   (   2   )                                
     where η n  is the nozzle adiabatic efficiency, T c  is the combustion exit stagnation temperature (the temperature at point  3  of FIG.  6 ), δ r  is the combustor pressure ratio, C p  is the specific heat at constant pressure of the exhaust gas, C v  is the specific heat at constant volume of the exhaust gas and γ is C p /C v . The energy E c  required to compress the air is taken from the total output of the engine E gross .                E   c     =         C   p12                       (       T   2     -     T   1       )       η   m         =         C   p12         η   m          η   c              [       δ       γ   -   1     γ       -   1     ]                 (   3   )                                
     where δ is the total pressure ratio of compressor rotor  62 , δ m  is the compressor mechanical efficiency, η c  is the compressor adiabatic efficiency, and C p12  is the specific heat of air at constant pressure going from point  1  to point  2  on FIG.  6 . Strictly speaking C p12  is a function of temperature; but for the purpose of estimating the efficiency of the embodiment of FIG. 3, C p12  can be approximated as a constant. Thus, the net output power per unit mass is:                E   net     =         W   e        V     -       C   p12                       (       T   2     -     T   1       )       η   m                   (   4   )                                
     and the thermal efficiency of the embodiment of FIG. 3 is:                η   th     =         E     net                 _                 output         E   input       =           W   e        V     -         C   p12          (       T   2     -     T   1       )       /     η   m               C   p23          (       T   c     -     T     2      R         )       /     η   com                   (   5   )                                
     where T 2R  is the total rotor exit temperature (relative to the rotor), C p23  is the specific heat of combustion air at constant pressure going from point  2  to point  3  on FIG. 6, and η com  is the combustor efficiency. 
     When recuperation of the exhaust gas heat is applied using a heat exchanger, then the temperature immediately prior to combustion increases from T 2  to T 2A , and the exhaust temperature increases to T 4A . Consequently, the new exit velocity is reduced to W eA  and the new thermal efficiency is:                η   th     =           W   eA        V     -         C   p12          (       T   2     -     T   1       )       /     η   m               C   p2A3          (       T   c     -     T     2      R         )       /     η   com                 (   6   )                                
     Assuming the following parameter values: 
     T 1 =300° K. 
     compressor pressure ratio δ r =15:1 
     compressor adiabatic efficiency η c =0.88 
     compressor mechanical efficiency η m =0.98 
     combustor efficiency=0.96 
     nozzle efficiency η n =0.95 
     T 3 =1773° K. 
     heat exchange efficiency=0.5 
     orbiting speed V=700 m/sec 
     the following estimates were obtained: 
     Net Output: 
     For a non-recuperated cycle: 226 kW/kg/sec 
     For a recuperated cycle: 221 kW/kg/sec 
     Thermal Efficiency: 
     For a non-recuperate cycle: η th =26.0% 
     For a recuperated cycle: η th =35.0% 
     A competitive gas turbine engine working with the same cycle pressure ratio and the same T 3  delivers only 163 kW/kg/sec with a thermal efficiency of only 23.2%. 
     While the invention has been described with respect to a limited number of embodiments, it will be appreciated that many variations, modifications and other applications of the invention may be made.