Abstract:
A compound exhaust system including two or more stages, preferably three, uses the force of exhaust gases traveling from a thrust generating source through the exhaust system to maximize thrust and minimize wasted energy consumption. In particular, the compound exhaust system for a thrust generating source comprises at least a first stage exhaust and a smaller, second stage exhaust. Various reflections of high speed traveling gases are provided and the resultant pressures induced during this travel supplement the thrust provided by the thrust generating source, which can be a jet engine. The invention also relates to a novel airframe that uses multiple engines having the inventive compound exhaust system.

Description:
BACKGROUND OF THE INVENTION 
     1. Field of Invention 
     The invention relates to an improved compound exhaust system for engines in flying craft capable of increased energy efficiency. 
     2. Description of Related Art 
     Numerous aircraft and spacecraft utilize solid, gas or liquid fossil fuels in jet or rocket engines to provide thrust necessary for flight of the craft. While numerous improvements in engine efficiency have been achieved over the years, the main focus in further efficiency has been in the engine design itself and much energy is still wasted or needlessly expelled out the exhaust of such conventional jet or rocket engine exhaust systems. 
     There is a need for a more energy conserving exhaust system that can minimize wasted fuel by effectively obtaining more thrust for a given input supply of fuel. 
     SUMMARY OF THE INVENTION 
     Applicant has overcome the above long felt needs and desires by inventing a novel compound exhaust system that replicates additional thrust by reusing the exhaust several times over before it is spent out the final stage of the exhaust system. 
     The invention relates to a compound exhaust system including two or more stages, preferably three, that use the force of the exhaust gases traveling through the exhaust system to maximize thrust and minimize wasted energy consumption. In particular, the compound exhaust system for a thrust generating source comprises at least a first stage exhaust and a smaller, second stage exhaust, the first stage exhaust including 
     an inlet control that receives exhaust gases travel at a high speed from a thrust generating source, the inlet control having a predetermined inlet diameter at a centerline of the exhaust; 
     a diverging conical wall extending from the inlet control to a lower support of the first stage, 
     an inverted cone impact area in-line with the centerline, the inverted cone impact area having a diameter substantially the same as the diameter of the inlet control and being located between the lower support and the inlet control, 
     a canalled solid impact area surface extending from the inverted cone impact area to the lower support, the canalled solid impact area and the diverging conical wall defining a first gas expansion area therebetween, the canalled solid impact area surface having a series of deep annular canals at an angle of about 45° to the centerline and a series of high speed jet nozzles extending between the canals, 
     a first upper cone-shaped reaction area surface defined on a bottom surface of the canalled solid impact area surface and defining an upper boundary of a first upper cone-shaped gas reaction area, the series of high speed jet nozzles running parallel to the canals and extending through the canalled solid impact area to define a flow path between the first gas expansion area and the first upper cone-shaped gas reaction area, 
     a lower high impact reaction area surface having a lower cone-shaped gas reaction area surface located along the centerline and defining a lower boundary of the upper cone-shaped gas reaction area, the lower high impact reaction area surface also having a diverging conical element extending from the lower cone-shaped gas reaction area surface to the lower support, and 
     a first low pressure gas overflow channel in fluid communication with the first upper cone-shaped reaction area, the first low pressure gas overflow channel being defined along a periphery of the lower support of the first stage and channeling gases from the first upper cone-shaped reaction area to the second stage of the exhaust system, and 
     the second stage exhaust including 
     a converging conical wall extending from the lower support of the first stage to a lower support of the second stage, 
     a second inverted cone impact area in-line with the centerline, the second inverted cone impact area having a diameter smaller than the diameter of the inlet control and being located between the lower support of the first stage and the lower support of the second stage, 
     a second canalled solid impact area surface extending from the second inverted cone impact area to the lower support of the second stage, the canalled solid impact area and the converging conical wall defining a second gas expansion area therebetween, the second canalled solid impact area surface having a second series of deep annular canals at an angle of about 45° to the centerline and a second series of high speed jet nozzles extending between the second canals and also angled at about 45°, 
     a second upper cone-shaped reaction area surface defined on a bottom surface of the second canalled solid impact area surface and defining an upper boundary of a second upper cone-shaped gas reaction area, the second series of high speed jet nozzles running parallel to the second canals and extending through the second canalled solid impact area surface to define a flow path between the second gas expansion area and the second upper cone-shaped gas reaction area, 
     a second lower high impact reaction area surface having a second lower cone-shaped gas reaction area surface located along the centerline defining a lower boundary of the second upper cone-shaped gas reaction area, the lower high impact reaction area surface also having a diverging conical element extending from the lower cone-shaped gas reaction area surface to the lower support, and 
     a second low pressure gas overflow channel in fluid communication with the second upper cone-shaped reaction area, the second low pressure gas overflow channel being defined along a periphery of the lower support of the second stage and channeling gases from the second upper cone-shaped reaction area toward a thrust vectoring nozzle located downstream from the second stage of the compound exhaust system, which forms an exit from the compound exhaust system. 
     The invention also relates to a novel airframe that uses multiple engines having the inventive compound exhaust system. 
    
    
     BRIEF DESCRIPTION OF THE DRAWINGS 
     The invention will be described with reference to the following drawings wherein: 
     FIG. 1 is a cross-sectional view of a compound exhaust system according to a first embodiment of the invention; 
     FIG. 2 is another cross-sectional view of the compound exhaust system showing exhaust flow through the system; 
     FIG. 3 is a cross-sectional view of a first stage of the exhaust system according to the invention; 
     FIGS. 4A-D show various sections of an exemplary stage according to the invention; 
     FIG. 5 is a top internal view of the exhaust system shown in FIGS. 1-2; 
     FIG. 6 is a cross-sectional view of a first stage of the exhaust system showing dimensioning according to the invention; 
     FIG. 7 is a perspective view of an exemplary fuselage that can achieve VTOL flight using several thrust sources with the inventive compound exhaust system; 
     FIG. 8 is another perspective view of another exemplary fuselage that can achieve VTOL flight using several thrust sources with the inventive compound exhaust system; and 
     FIG. 9 is a cross-sectional view of an alternative embodiment of the invention using a non-air breathing thrust generation source. 
    
    
     DETAILED DESCRIPTION OF PREFERRED EMBODIMENTS 
     With reference to FIG. 1, input to the inventive compound exhaust system  200  is a supply of high speed gases G that release energy and generate thrust to propel a vehicle, such as a flying craft, spacecraft or other vehicle. The high speed gases are created by a thrust source  100  that can include most any known conventional jet or rocket engine, such as a turbo-fan jet engine, or subsequently developed technology that achieves thrust and generates similar exhaust gases. This is not limited to liquid, gas or solid fossil fuel engines, but can include alternative thrust sources such as the water and air propulsion system disclosed in my U.S. patent application Ser. No. 09/200,703, now U.S. Pat. No. 6,290,184, the subject matter of which is incorporated herein in its entirety. However, the actual method or apparatus for initial thrust generation is not the primary subject of this invention. Rather, this invention primarily focuses on exhausting of these gases G to increase thrust and efficiency. 
     Thrust source  100  generates exhaust gases G that travel towards compound exhaust system  200 , which as shown is a three-stage exhaust system including first stage  200 A, second stage  200 B and third stage  200 C. Each stage includes a rim  236 A,  236 B, and  236 C, respectively, that seals and spaces the stage from adjacent stages. The first stage  200 A has a relatively large gas expansion area (GEA)  204 A formed in the shape of an inverted cone. The second stage  200 B has a smaller GEA  204 B. The third stage  200 C has an even smaller GEA  200 C, which makes the accelerated gases travel faster due to the high pressure after the expanding gases flow from the second stage  200 B in greater volume into the third stage  200 C. More detailed views of the exhaust system are illustrated in FIG.  2 . While three stages are preferred, the invention can be practiced with as few as two stages or could include four or more stages depending on the desired application. 
     As shown in FIGS. 1-2, air duct control  202  at the inlet of exhaust system  200  has a diameter that is equal to the diameter of an inverted cone impact area (ICIA)  212 A of the first stage  200 A of the exhaust system. Highly accelerated gases rushing down the ICIA  212 A from air duct control  202  have a compound tubular (solid) shape and upon impact with the ICIA  212 A and directional pin  206  rise upward within gas expansion area (GEA)  204 A toward the height of the air duct control  202  where pressure is greater due to a tapered conical top wall  216  of the exhaust system  200 . This augments the original thrust of the system by pushing the new supply of gases G upward, as well as providing lift forces to wall  216 . 
     Upon reaching air duct control  202 , the rising gases are pushed down by the greater forces of the new gases coming from the thrust source  100  and then forced left, partially due to the directional pin  206 , to spin downward towards canalled solid impact area (CSIA)  214 A, which includes deep canals  215  angled 45° towards the center, causing the spinning gases to continue spinning through all the CSIA  214 A. This spinning creates a tornado effect as the gases G rise once more within the circumferential area of the tapered conical top part  216  of the exhaust system  200 , which creates additional upward thrust while the bulk of the spinning accelerated gases are forced downward by the rush of new gases from air duct control  202  toward angled high speed jet nozzles (HSJN)  218 A provided in CSIA  214 A at a 45° angle. A central lower side of the CSIA  214 A is dome-shaped. 
     After the gases travel through the high speed jet nozzles  218 , the gases exit and converge toward a center line of a first lower cone-shaped gas reaction area surface (LCSGRAS)  233 A, which includes a centrally located raised dome portion and a conical section extending from the dome portion to a lower support  222 A of the first stage  200 A. High speed gases coming from nozzles  218 A nearest the ring  236 A (i.e., outer peripheral nozzles) will ride on top of highly accelerated gases coming from high speed nozzles  218 A located on top of an upper cone-shaped gas reaction area (first UCSGRA)  224 A, which is defined between a first upper cone-shaped gas reaction area surface (first UCSGRAS)  225 A formed on a lower side of the canalled solid impact area  214  and the first LCSGRAS  233 A. 
     These gases travel toward the center of the first UCSGRA  224 A which, by the time these gases reach a spot where high speed gases from the first UCSGRA  224 A meet, form a solid wall of high speed gases. Thus, all accelerated gases coming from all sectors of the first stage exhaust hit a lower cone-shaped gas reaction area (first LCSGRA  234 A) and curve upwards in a spiral continuously moving towards the first UCSGRA  224 A. Upon impact with the much stronger expanded gases coming out from the top portion of the first UCSGRA  224 A, an upward pressure ensues at the center of the first UCSGRA  224 A that produces a large upward push while the spiraling widens its curvature and rushes towards and out an angled gas overflow chamber (GOC)  226 A located around the periphery of the bottom of the first stage  200 A. Thus, the gas flow has an involute action, spiraling continuously around a point in constantly increasing curvature until it reaches the GOC  226 A. 
     GOC  226 A is preferably louvered at a 45° angle towards a center of second stage  200 B of the compound exhaust system. GOC  226 A is located a little lower than a top portion of the lower section  233 A. GOC  226 A can be louvered by carving louvers from a solid wall or by boring oversized gas nozzles angled 45° towards the center of the lower stage. The louvers and/or gas nozzles are formed of a suitable size relative to the other dimensions of the exhaust system. When a large volume of accelerated gases pass through the louvered low pressure gas overflow channel  226 , the pressure at the first UCSRA  224 A becomes greater producing additional thrust. Thus, the angled high speed jet nozzles  218  and angled gas overflow channel  226 A contribute to thrust. 
     Gases exiting the first stage  200 A through gas overflow channel  226 A descend downward into gas expansion area (GEA)  204 B toward a second ICIA  212 B of the second stage  200 B, where the gases replicate the action and reaction forces created in the first stage by way of involute action until they exit from gas overflow chamber (GOC)  226 B of the second stage  200 B into the third stage  200 C, where the gases again encounter the same forces. 
     That is, gases from GEA  204 B are funneled toward ICIA  212 B and upon impact rise upward toward the lower side of LCSGRA  234 A where they are pushed down by new gases flowing from the GOC  226 A and forced to spin downward toward CSIA  214 B, which causes the gases to rise once more in a tornado effect. These gases then spin toward and through nozzles  218 B, where upon exit the gases converge toward LCSGRAS  233 B. From here, the gases travel toward the center of the second UCSGRA  224 B. The gases then curve upward in a spiral where they meet new gases and are forced to spiral in a widening curvature until they rush out gas overflow chamber (GOC)  226 B into the third stage  200 C. Gas flow through the third stage is the same as through the second stage. Upon exit of the gases through gas overflow chamber  226 C of the third stage, the gases travel into an exhaust chamber (EC)  210 , which is located between the third stage  200 C and thrust vectoring nozzle (TVN)  34 . Gases entering EC  210  travel through TVN  34  at very high speeds. The TVN  34  has a diameter equal to the gas entry point  202 . 
     In summary, the gas flow through the system enters the exhaust through gas entry point  202  into the first stage  200 A in a solid tubular shape, rushing towards ICSIA  212 A and upon hitting the same will rise upwards, but since the new incoming gases have greater speed and pressure, the gases are pushed back downwards in a spinning condition towards the canalled area, which creates a tornado effect inside the gas expansion area continuously pushing the device attached to the exhaust upward. Other gases within the gas expansion area begin to exit the nozzles  218  towards a lower section of the first section. As the nozzles  218  are angled at 45 degrees towards the lower cone-shaped gas reaction area will develop an involute action around the hollow disc-shaped lower section of the first stage. See FIG. 3, which better illustrates the gas flow through the first stage where the curvature of the accelerating gases in the involute state increase in width and hit the edges of the gas overflow channel, which is angled 45 degrees towards the gas reaction area of the second stage. With this construction, accelerated gases will flow out of the gas overflow channel in greater volume, thus producing an increased upward pressure against the upper cone-shaped gas reaction area pushing the craft upward. All highly accelerated gases emerging from the gas overflow channel mass together toward the center of the second stage, where the process is repeated until it exits out TVN  34 . 
     FIGS. 4A-D show various sections of the first stage. FIG. 4A shows an upper section having the ICIA  212 A, canals  215  and first upper cone-shaped gas reaction area  224 A. The upper section is solid with high speed jet nozzles  218  except for the upper cone-shaped gas reaction area  224 A. FIG. 4B shows a lower section, which is hollow and includes the lower cone-shaped gas reaction area  233 A. FIG. 4C shows a gas overflow channel  226 A having louvers angled at 45 degrees. Alternatively, as shown in FIG. 4D, the gas overflow channel  226 A can have oversize nozzles to handle gas flow. 
     Thrust can be supplemented by chilled air entry valves (CAEV)  228  (see FIG. 1) which are high pressure valves that spray a fine pressurized chilled air from a source and deliver the chilled air within the second gas expansion area  204 B. The chilled air rapidly expands when heated, causing additional pressure within the second gas expansion area  204 B and additional thrust. One or more electronic sparkplugs  242  may also be provided in the second stage when the power source is an air-breathing engine or rocket to provide continuous or intermittent sparks around the gas expansion area of the second stage to further burn all fine combustibles coming from the gas overflow channel of the first stage to further assist in trust generation. 
     The inventive compound exhaust system can be used in conjunction with various power sources, such as air-breathing engines, rockets, or other combustion-related engines. Various dimensions of the elements will vary depending on the particular application, such as engine capacity, cargo capacity, whether the exhaust will be used in normal atmosphere or in space, and other considerations. 
     In a first embodiment, the upper section of each stage  200 A,  200 B and  200 C is divided into four sections, labeled A, B, C and D. Each section is subdivided into two hemispheres to provide eight regions A 1 , A 2 , B 11 , B 2 , C 1 , C 2 , D 1  and D 2  as shown in FIGS. 5 and 6. The four sections A, B, C and D are separated by canalled impact areas (CIA)  214 . Each section is made up of a solid metal material having a plurality of high speed nozzles extending therethrough at a 45 degree angle. The nozzle diameter preferably varies from one stage to the next. For example, in a first embodiment, the regions within the first through third stages may have the following sizes: 
     
       
         
               
               
               
               
             
           
               
                   
                 TABLE 1 
               
               
                   
                   
               
               
                   
                 STAGE 
                 REGION 
                 SIZE 
               
               
                   
                   
               
             
             
               
                   
                 FIRST STAGE 
                 A1, A2 
                 ⅛″ 
               
               
                   
                   
                 B1, B2 
                 {fraction (1/16)}″ 
               
               
                   
                   
                 C1, C2 
                 {fraction (1/16)}″ 
               
               
                   
                   
                 D1, D2 
                 ¼″ 
               
               
                   
                 SECOND STAGE 
                 A1, A2 
                 ¼″ 
               
               
                   
                   
                 B1, B2 
                 ⅛″ 
               
               
                   
                   
                 C1, C2 
                 ⅛″ 
               
               
                   
                   
                 D1, D2 
                 ½″ 
               
               
                   
                 THIRD STAGE 
                 A1, A2 
                 ½″ 
               
               
                   
                   
                 B1, B2 
                 ¼″ 
               
               
                   
                   
                 C1, C2 
                 ¼″ 
               
               
                   
                   
                 D1, D2 
                 1″ 
               
               
                   
                   
               
             
          
         
       
     
     In an alternative embodiment, the nozzle sizes are uniform in all of the first, second and third stages. The size is proportional to the size and type of fuselage used. 
     The three stages  200 A,  200 B and  200 C can be affixed to each other in several ways. In the embodiment shown, a circular rim  210  encircles the adjoining edges of adjacent stages and couples them together. As shown in FIG. 7, the inventive compound exhaust system can be used with a fuselage  300  to provide a vertical take-off and landing (VTOL) craft. While shown in a saucer-shaped configuration, the exhaust system is adapted for use with any type of vehicle fuselage. 
     In this configuration, the fuselage  300  includes five fixed turbo-fanjet engines  100 A, two forward turbo-fan jet engines  100 B, two forward reversible turbo-fan jet engines  100 C, and two backward reversible turbo-fan engines  100 D, along with drift control nozzles  400  located on each of the five fixed jet engines  100 A as well as on upper, lower and side portions of fuselage  300  for a total of 24 drift control nozzles. Drift control nozzles  400  can be operated in pairs to correct for drift of the fuselage in flight. 
     During take-off, each of the five fixed jet engines  100 A together with the jets  100 B and  100 C are throttled to provide thrust. Engines  100 B and  100 C are positioned vertically during take-off to assist in vertical thrust. Once a proper altitude is reached, the two forward reversible engines  100 B and the two backward reversible engines  100 C can be shut down and returned to a horizontal orientation. At this time, any of engines  100 B,  100 C or  100 D can be throttled up to propel the craft forwards or backwards. Similarly, rotation or drift correction can be achieved by the drift correction nozzles  400 . During forward flight, power to the fixed jets can be decreased. 
     Turning is achieved by relative control of the various jets. For example, steering or banking left can be achieved by making the fixed jets  100 A on the left side decrease in power output (throttle down) while power to the fixed jets  100 A on the right side is increased (throttle up). This results in the craft lowering its left side while the right side is raised. Downward turning can be achieved by reducing the power to the jet  100 A at the front of the craft and/or increasing power to the jet  100 A at the rear side of the craft. 
     Each of the engines  100 A-D include the inventive compound exhaust system for improved efficiency. Each of the engines  100 A-D may be 18,000 pound thrust jet engines. 
     FIG. 8 shows a VTOL airframe  300  according to another embodiment of the invention. Like numerals refer to like elements. In this embodiment, a cockpit  500  is provided on top of the saucer-shaped fuselage. Below cockpit  500  is air inlet  600 . A series of doors  700  and windows  800  can be provided around the craft. The jet engines preferably are 18,000 pound thrust engines except the central engine  100 A, which can have 36,000 pound thrust. 
     The inventive compound exhaust system has been shown in the previous examples to be useable with various power sources, such as air-breathing engines, rockets, or other combustion-related engines. However, the invention is also applicable to non-air breathing thrust generating power sources. An exemplary embodiment of such is illustrated in FIG.  9 . The thrust generating source  100  may use, for example, water (H 2 O) and/or liquid hydrogen (H 2 ) as fuel. The compound exhaust system in this embodiment is provided with an automatic retractable air tight locking device  310  at the periphery of exhaust chamber  210  that will trap expanding gases. Multiple similar vacuum-locking doors  320  may be provided above locking device  310  around the exhaust chamber  210  that are in communication with multiple efficient compressors  330  that are capable of sucking out the trapped gases and pump them into multiple condensers  340  that convert the gases back to a liquid state. Thereafter, the converted liquid is again pumped back to a fuel cell  350  to be recycled and later fed to the thrust source  100 . The vacuum-doors  320  are preferably located about one foot above the air tight locking device  310 . 
     In operation, gases pass from the thrust source  100  through the three (or more) stages of the compound exhaust system  200  up to the exhaust chamber  210  where the gases are contained by the locking device  310 . Thereafter, the multiple vacuum-doors  320  may be opened to allow sucking of the gases by the multiple compressors  330  to the multiple condensers  340  where the gases are converted back to their liquid state and thereafter pumped back to a fuel cell  350 . 
     The locking device  310  at the periphery of exhaust chamber  210  may be openable to allow maintenance. The locking device  310  may also be openable to allow the superheated exiting gases to be released to the atmosphere. This may form a thick cloud of vapors, which may be a harmless byproduct or a desirable stealth mechanism that can at least partially conceal the craft. 
     In another embodiment, the flying craft can be used for interplanetary flight. The fuselage in such an embodiment would be designed to withstand the pressures and temperatures encountered when exiting or entering earth&#39;s atmosphere. Additionally, the fuselage would be pressurized. Optionally, when traveling through space, the exhaust gases can be vented back into the craft by a vacuum system powered by compressors so that the exhaust gases can be recycled. 
     To handle the forces generated by the novel propulsion system, the exhaust should be made from a suitable high strength, high heat metal. Numerous suitable metals or composite materials would be known to one of ordinary skill in the art. One such suitable material is KM-1557. The remainder of the spacecraft can be made from suitable materials based on desired requirements for each component, based on its size, strength, weight, and intended use of the flying craft. 
     While specific aspects of the invention have been described with respect to preferred embodiments of the invention, these are not intended to be limiting. Various modifications can be made without departing from the scope of the appended claims.