US 2015.0007549A1 (19) United States (12) Patent Application Publication (10) Pub. No.: US 2015/0007549 A1 BOSSard (43) Pub. Date: Jan. 8, 2015

(54) ROTARY TURBO (52) U.S. Cl. CPC. F02K9/00 (2013.01); F02K3/00 (2013.01) (71) Applicant: John Bossard, Huntsville, AL (US) USPC ...... 60/204; 60/246 (72) Inventor: John Bossard, Huntsville, AL (US) (57) ABSTRACT (21) Appl. No.: 14/191,926 (22) Filed: Feb. 27, 2014 A is combined with a co-axially integrated rotary 9 rocket to form a propulsion system called a Rotary Turbo Related U.S. Application Data Rocket that can function as a turbojet, as an afterburning turbojet, as an Air Turbo Rocket, or as a . The (60) Provisional application No. 61/772,821, filed on Mar. Rotary Turbo Rocket can operate in any of these propulsion 5, 2013. modes singularly, or in any combination of these propulsion Publication Classification modes, and can transition continuously or abruptly between operating modes. The Rotary Turbo Rocket can span the Zero (51) Int. Cl. to orbital flight Velocity speed range and/or operate continu FO2K9/00 (2006.01) ously as it transitions from atmospheric to space flight by FO2K 3/00 (2006.01) transitioning between operating modes.

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ROTARY TURBO ROCKET both air-breathing and rocket propulsion modes, no advan tage is taken of combining the required turbomachinery of the BACKGROUND OF THE INVENTION jet and rocket engines, and thus the concept requires separate 0001 1. Field of the Invention turbomachinery assemblies for the jet and rocket engines, 0002 The current invention relates to a class of propulsive which greatly increases complexity and mass of the overall engines known as turbine-based-combined-cycle (TBCC) engine. engines and more particularly to a turbojet combined with a 0008 Rotary rocket engines use the rotation of a combus rotary rocket. tion chamber, powered by the release of exhaust gases from 0003 2. Description of the Prior Art the combustion chamber, to pump and mix propellants. The 0004. The need for endo-atomospheric, trans-atmo rotational force comes from a tangential component of rocket spheric, and exo-atomospheric Vehicles has created interestin thrust, which is created by canting one or more rocket nozzles propulsion systems that can operate in, through, and out of the in the rotary rocket assembly. U.S. Pat. No. 2,479,829 and atmosphere. A number of propulsion systems designed to U.S. Pat. No. 2,395,114 disclose rotating combustion cham operate in atmosphere have been proposed. Turbine-based bers in which rotation is produced by the passage of exhaust combined-cycle (TBCC) engines, for example, are propul gases through curved nozzles and against Vanes. U.S. Pat. No. sive engines that have elements of both turbojet engines, 6.212.876 discloses a rocket motor with an ultracentrifugal which use atmospheric air as the main constituent of their liquid pump driven by a tangential component of the primary reaction mass, and rocket engines, which carry all of their thrust from multiple combustion chambers by means of tilted reaction mass. nozzles or vanes. Rotation of combustion chambers and 0005. An Air Turbo Rocket (ATR) is a type of TBCC pump enhances both propellant mixing and combustion engine which combines a conventional Solid or liquid propel chamber cooling by the Coriolis-effect and centripetal accel lant rocket-type gas generator with a conventional turbojet in eration. Rotary provide the advantage of eliminating Such a way that heated, fuel rich gas from the rocket is used to the weight and complexity of separate, non-integrated pro power the mechanical compression of air from the atmo pellant pumps compared to pump-fed combustion chambers, sphere. This fuel rich gas and the compressed air are then and the weight of high-pressure propellant tanks when com mixed and burned in a turbojet-type afterburner. The air pared to pressure-fed combustion chambers. Because the breathing portion of the ATR uses atmospheric oxygen to rocket chambers are integral with the pump elements, rotary burn the product of fuel rich combustion by the rocket in a rockets also eliminate the need for high-pressure, high-speed combustion chamber that is analogous to an afterburner in a rotating seals, which is a very challenging and expensive turbojet engine. An advantage of an ATR is relatively high issue for conventional -fed rocket engines. One thrust generation at high speeds, high thrust per unit frontal disadvantage of a rotary rocket is that there is no rotation at area, and a compact and lightweight engine configuration. A start-up So Some means other than rotation must be provided primary disadvantage of an ATR is its high propellant con for pumping propellants at engine start. Another disadvantage sumption rate, relative to a turbojet. U.S. Pat. No. 4,096,803 of rotary rockets is that the rocket chambers are heavier than discloses a solid propellant air turbo rocket that can operate in conventional rocket chambers because they must be designed a mode by jettisoning a solid propellant rocket-driven to withstand the centrifugal forces generated during the rota turbo-compressor. tion of the rocket chambers. 0006 EP 0403372 B 1 discloses a combined turbo-rocket 0009. Despite the development of the aforementioned pro ramjet engine capable of operating in a combined turbo pulsion technologies, high speed endo-atomospheric, trans rocket mode or a ramjet mode. The engine uses to atmospheric and exo-atomospheric Vehicles powered by pump fuel and oxidizer to a gas generator that drives a com TBCC engines have not yet come into widespread use. TBCC pressor for turbo-rocket mode operation or to radial injectors engines which are only air-breathing are limited in their maxi for ramjet mode operation. The engine operates in turbo mum flight speeds and altitudes, and cannot fly exo-atmo rocket mode at low Mach speeds and Switches to ramjet mode spherically. Alternatively, rocket engines have too high of a at high Mach speeds. Although the engine uses stored oxygen propellant-usage rate for long-range endo or trans-atmo as an oxidizer, this engine uses atmospheric oxygen in all spheric flights. Thus, there is a continuing need for a new operational modes and is therefore not capable of operation TBCC engine capable of continuously providing propulsion from take-off to landing for endo, trans, and exo-atomo outside the atmosphere. spheric flight operations. 0007 U.S. Pat. No. 5,159,809 discloses a propulsion engine designed to reversibly change from air-breathing, non-air-breathing, and combined modes. The engine com SUMMARY OF THE INVENTION prises a having an annular combustion chamber 0010. Accordingly, embodiments of the present invention disposed inside the central body and downstream from the preferably seek to mitigate, alleviate or eliminate one or more air-breathing combustion chamber of a . The rocket deficiencies, disadvantages or issues in the art, Such as those engine has a streamlined central body portion extending the identified above, singly or in any combination by providing a central body of the air-breathing combustion chamber to form Rotary Turbo Rocket (RTR) comprising a turbojet co-axially a spike that penetrates into the throat of the external nozzle. integrated with a rotary rocket. The turbojet may operate in The combustion chamber of the rocket engine and the asso any mode available to turbojet engines, including SubSonic, ciated streamlined central body are integrated in the main transonic, and SuperSonic flight speeds, by adapting the air nozzle and penetrate into the main nozzle at the throat while inlet and turbomachinery operation according to known being shaped to ensure aerodynamic continuity of the air methods. The rotary rocket may be supplied with fuel and breathing stream from the air-breathing combustion chamber oxidizer at any mixture ratio for use at any time whether in being ejected through the main noZZle. Although this engine atmosphere or not. When operated with a fuel-rich mixture concept combines a jet engine with a rocket engine, allowing ratio, the rotary rocket functions as a fuel-injector for an US 2015/0007549 A1 Jan. 8, 2015 afterburning turbojet mode. Furthermore, when using a fuel nying drawings is not intended to be limiting of the invention. rich mixture ratio, the rotary rocket can produce shaft power In the drawings, like numbers refer to like elements. to drive the compressor, burning the fuel-rich mixture in the 0020 FIG. 1 illustrates the basic configuration of one afterburner, and thus allowing the RTR to operate in an Air embodiment of a Rotary Turbo Rocket (RTR) comprising a Turbo Rocket mode. Accordingly, the RTR is able to operate turbojet and a rotary rocket 7 contained in an outer casing 71. as a turbojet, an afterburning turbojet, an Air Turbo Rocket, a The turbojet comprises a compressor 11, a combustion cham rotary rocket, together or separately, in any combination. ber 10, turbine 6, and bearings 5. The rotary rocket 7 is Furthermore, the RTR is able to transition between any of positioned aft of (behind) the turbojet. The compressor 11, these modes abruptly or in a variable or continuous or gradual turbine 6, bearings 5, and rotary rocket 7 are mounted on a a. main shaft 4 through which oxidizer12 and fuel 13 are deliv 0011. According to one aspect of the invention, a rotary ered through oxidizer and fuel feed lines 112,113 to the rotary turbo rocket is provided, in which a turbojet engine is coaxi rocket 7. A seal housing 2 contains the front end of the main ally integrated with a rotary rocket engine. According to shaft 4 where inlets into oxidizer and fuel feed lines 112,113 another aspect of the invention, a method is provided for are located. Bearings 5 allow the main shaft 4 to rotate with operating a RTR as a turbojet, an afterburning turbojet, an air respect to the combustion chamber 10 of the turbojet. An turbo rocket, and/or as a rotary turbo rocket. According to optional afterburner 9 may be positioned aft of the rotary another aspect of the invention, a method is provided for rocket 7. An exhaust nozzle 8 is positioned at the aftend of the operating these modes together or separately, and for transi turbojet and rotary rocket 7. A Turbojet fuel line 114 provides tioning continuously or abruptly between these modes. One fuel to the turbojet 10. or more operational parameters of the RTR may be monitored 0021. An air inlet 1 positioned at the front end of the RTR and/or controlled by a computer or processor connected to is configured to allow air to enter and provide oxygen to be one or more components of the RTR. Methods for operating used by the combustor 10 and the afterburner 9 of the turbojet. the RTR may involve method steps and/or processes that are During spaceflight or SuperSonic flight in rotary rocket mode, monitored, controlled, and/or performed by a computer or the air inlet may be partially or completely closed. For processor. embodiments in which the turbojet may operate at SubSonic or SuperSonic speeds, the air inlet 1 may have a variable BRIEF DESCRIPTION OF THE DRAWINGS geometry that reduces the speed of superSonic incoming air to SubSonic speeds. This can be accomplished using known 0012. These and other aspects, features and advantages of geometries such as conic spikes or wedges that produce shock which embodiments of the invention are capable of will be waves which form on the inlet and slow incoming air to apparent and elucidated from the following description of SubSonic speeds. The shock waves may be oblique or normal embodiments of the present invention, reference being made shock waves that form at an oblique angle to or are perpen to the accompanying drawings, in which dicular to the flow of air. AS flight speeds change, it may be 0013 FIG. 1 is a basic schematic illustrating an embodi necessary to change the geometry of the inlet. This may be ment of a Rotary Turbo Rocket comprising a rotary rocket accomplished, for example, by translating an intake spike assembly co-axially integrated into a turbojet; forward or backward or by changing the wedge angle of an 0014 FIG. 2 illustrates one embodiment of a set of axi intake wedge in a manner analogous to known SuperSonic symmetric components which fit together and are connected inlets. Likewise, the exit nozzle 8 may also have a variable to an aft end of a main shaft to form one embodiment of a geometry to accommodate and optimize engine performance rotary rocket assembly; over the operating range of the engine. One of the advantages 0015 FIG. 3 is a side and end view illustrating a variable of the RTR over known is the ability to use the rotary cant angle C. of an embodiment of a rotary rocket combustion rocket engine to restart the turbojet in the event that the chamber(s) and rocket nozzles; turbojet combustor 10 extinguishes, or flames out. 0016 FIGS. 4A and B are schematics illustrating two con 0022. A seal housing 2 of the RTR in FIG. 1 is a system figurations for propellant flow passages within a main shaft which allows two or more propellants, typically a fuel and an assembly of different embodiments of a Rotary Turbo oxidizer, to pass into the rotating components of the RTR Rocket; without mixing and without leaking The seal housing 2 com 0017 FIG. 5 is a side view illustration of one embodiment prises seals 52, 53 that allow the flow of propellants from the of a seal housing assembly; and fuel and oxidizer Sources 12, 13 to channels or passages in the 0018 FIG. 6 is a schematic illustrating an embodiment of main shaft 4 that deliver propellants to the rotary rocket 7. a Rotary Turbo Rocket in which the turbojet and the rotary Any suitable seal technologies may be used including, for rocket can be mechanically coupled or de-coupled. example, metallic labyrinth seals and/or Teflon face seals. Metallic labyrinth seals may provide optimal sealing and DETAILED DESCRIPTION OF THE INVENTION minimal drag, but are expensive and possibly less tolerant of contamination or shaft dynamic upsets. Teflon face seals may 0019 Specific embodiments of the invention now will be be used and are inexpensive and robust, but provide more drag described with reference to the accompanying drawings. This upon the rotating shafts than metallic labyrinth seals. FIG. 5 invention may, however, be embodied in many different shows a more detailed view of one embodiment of a seal forms and should not be construed as limited to the embodi housing. A purged space 56 between the fuel inlet 58 and ments set forth herein; rather, these embodiments are pro oxidizer inlet 54 may additionally be provided in the seal vided so that this disclosure will be thorough and complete, housing to purge propellants that may leak past the fuel seal and will fully convey the scope of the invention to those 52 and oxidizer seal 53, and thus avoid and/or preclude inad skilled in the art. The terminology used in the detailed Vertent mixing. A purge gas can enter the purged space 56 via description of the embodiments illustrated in the accompa the purge gas inlet 55. Seal spacers 57 position the fuel seals US 2015/0007549 A1 Jan. 8, 2015

52 and oxidizer seal 53 at their correct axial positions on the duced vibrations produced by the rotating elements. Axial main shaft 4, FIG. 1, and relative to the fuel, oxidizer, and loads come from the propulsive loads produced by the RTR purge inlets, 58,54, and 55. The seal housing is structurally engine, and from air drag loads which are generated during Supported and positioned at its correct axial position relative flight through the atmosphere. The number and locations of to the outer casing 71 using support brackets 59. These brack the main Support bearings are determined by the types and ets must attach to the seal housing Such that they maintain magnitude of loads that they support, and also by the vibra clearance for the fuel, oxidizer, and purge inlets. The seals 52. tional dynamics of the rotating assembly. 53 and the seal spacers 57 are secured within the seal housing 0026 Bearings may be positioned between the compres using a seal housing cover 60. This cover also provides any sor 11 and turbine 6, as shown in FIG. 1. Bearings may pre-load which may be required for proper seal operation. additionally be positioned within the seal housing 2, as well The seal housing 2, FIG. 1, can be placed in any of a number as aft of the rotary rocket 7, as necessary to resolve vibrational of different positions relative to the rotary rocket and/or the and load issues which may emerge from specific RTR con turbojet. The configuration shown in FIG. 1 is for one exem figurations. There must be sufficient space to accommodate plary embodiment of the RTR in which propellants enter the the type of bearing required as well as means for lubricating main shaft of the RTR at the front, just in front of the com and cooling the bearings. The bearing position may addition pressor. The seals need not be positioned as shown but may ally be determined by the need to support or eliminate rota also be positioned in other locations such as between the tion-induced vibrational modes from the rotating elements of compressor 11 and turbine 6, between the turbine 6 and the the RTR. Typical bearing types suitable for use in the RTR rotary rocket 7, or aft of the rotary rocket 7. The numbers and include ball and roller element bearings and foil-type bear locations of seals and purged spaces shown in the figures are ings, as well as other types of bearings. Bearings within the exemplary and may be varied. RTR must be of radial and/or axial load bearing capability, 0023. An advantage of locating the seals and seal housing depending on bearing location and function. in front of the compressor is that it positions these compo 0027. The turbojet part of the RTR comprises a compres nents away from the hot areas of the engine. Furthermore, Sor 11, which may be axial, radial, or mixed-flow in configu heat transfer to the propellant can be managed effectively ration, and may be comprised of one or more compressor because there is more room for insulation, as needed, and stages and their associated Stators and diffusers. The turbojet because surface areas where heat transfer can occur can be also comprises a combustor 10, which burns fuel 114 injected reduced to that of a cylindrical shape. The seals and seal into the combustor 10 with air from the compressor 11, and a housing may also form an important component pertaining to turbine 6. The turbine 6 may be axial, radial, or mixed-flow in the rotational dynamics of the RTR's rotating elements. Their configuration and may be comprised of one or more turbine placement on one end of the rotating elements provides a stages, and turbine nozzle guide vanes. means of controlling and mitigating rotation-induced vibra 0028. The RTR may possess an afterburner 9, in which tions in the system. fuel and air are burned to produce thrust. Air to the afterburner 0024. As shown in FIG. 1, the main shaft 4 supports and comes from the compressor 11, and may comprise excess air aligns the main rotating components of the turbojet and the from the turbojet's combustor 10, or unreacted air from the rotary rocket. These components include the compressor 11, compressor during operating modes in which the turbojet's turbine 6, bearings 5, rotary rocket 7, rocket combustion combustor is not in operation. Fuel to the afterburner may chamber(s) 73, rocket nozzles 72, pump 75, forward closure come from the RTR's rotary rocket 7, in which the rotation of 75 (FIG. 4), and fuel and oxidizer feed lines 112, 113. In the rotary rocket pumps the fuel into the afterburner. This fuel addition, the main shaft accommodates the flow of fuel and may be a fuel-rich effluent from a fuel-rich combustion pro the oxidizer to the rotary rocket 7. This may be accomplished cess within the rotary rocket, or unreacted fuel during oper through any of a number of possible configurations. For ating modes in which shaft power to the rotary rocket is being example, in FIG. 4, a set of coaxial tubes 41, 42 is used to supplied from the turbojet's turbine. Additionally or alterna provide flow passages for the fuel and oxidizer, with the tively, fuel may be injected directly into the afterburner with oxidizerpassing down a central passage in the main shaft, and out having passed through the rotary rocket through an after fuel passing down an annular passage formed from the out burner fuel line 115. side wall of an inner tube 112, and the inside wall of the main (0029. As shown in FIG.2, the rotary rocket part of the RTR shaft 4. Alternatively, discrete axial channels or passages can may comprise a collection of rocket chambers 73, throats, and be formed in the main shaft, alternating between fuel and nozzles 72, typically positioned at a common radius from the oxidizer. These channels can then be accessed with radial main shaft 4. The rocket chambers 73 can be individual cham holes around the perimeter of the shaft, with the hole sets 43, bers, connected chambers, or one continuous annular cham 44 positioned at different axial positions along the shaft, as ber. The rocket throats and nozzles 72 can also be individual shown in FIG. 4. The main shaft 4 also provides the contact elements, or continuous annular elements. In either configu and alignment surfaces for the inner races of the RTR's bear ration, however, the rocket elements must provide a compo ings. nent of thrust in the azimuthal direction, i.e. at a tangent to the 0025. The rotating assembly of the RTR comprises the rotary rockets axis of rotation. This thrust component gen main components of the turbojet (compressor 11, turbine 6, erates shaft power in the rotary rocket, and also to any other bearings 5), and of the rotary rocket 7, and feedlines 112,113, operating modes of the RTR which might require shaft power. and the main shaft 4 as shown in FIG.1. The bearings 5 of the In some operational modes, the main shaft 4 may be driven by RTR allow the rotating elements of the RTR to turn freely, the turbojet to pump propellants 12, 13 into the rotary rocket while simultaneously supporting radial and axial loads from combustion chamber(s) 73 for starting the rotary rocket. One these elements. Radial loads come from the masses of the or more segments 22 of the main shaft 4 may have an rotating elements, from flight-induced loads when the RTR is expanded diameter that forms one or more stops which may producing thrust for a flight vehicle, and from rotation-in be used to prevent one or more bearings 5 from sliding along US 2015/0007549 A1 Jan. 8, 2015

the main shaft and/or hold one or more bearings 5 at a fixed canted rockets equals to total torque multiplied by the angular position along the main shaft. The one or more segments 22 speed of the RTR's rotating components. This torque power is may have various lengths in the axial direction depending on the output power of the rotary rocket, and provides the the needed bearing spacing, and may be disk-shaped, coni required power to pump the propellants, as well as overcome cally-shaped, or cylindrical as shown in FIG. 2. any bearing and seal losses. This torque power can also be 0030 High rocket chamber pressures are achieved in the used to provide power to the compressor 11 when operating in rotary rocket because propellants 12, 13 are pumped to high Air Turbo Rocket mode, or to supplement power produced by pressures as they flow radially outward from the main shaft 4 the turbine 6 when operating in turbojet mode. During opera to the rocket chambers 73. The pump component 75 of the tion of the rotary rocket, the rotating components spin up and rotary rocket 7 is located between the main shaft 4 and the reach an rpm, or rotary speed, in which the torque power rocket chambers 73, and contains passages 75a allowing fuel output equals the sum of the pump and bearing/seal losses, and oxidizer to flow radially outwards. The passages can be, which is a power input. This power balance is also achieved for example, alternating passages located in the same plane, when the RTR is operating in other propulsion modes, such as or they can be located on different sides of the propellant the Air Turbo Rocket, and turbojet modes. pump 75, such as its forward and aft faces as shown in FIG. 2. 0033. Variations in rotary rocket torque power output can The shape of each passage, both its cross-sectional area and beachieved by varying the propellant flow rate and/or chang its path to the rocket chambers, can be configured to provide ing the cantangle C. of the rocket nozzles 72, as shown in FIG. optimal pumping efficiency, which may depend on the rotat 3. Variations in the cant angle can be achieved by using a ing speed and radius of the rotary rocket, and also on the types mechanical, electromechanical, magnetic, electromagnetic, of propellants being pumped. The propellant pump 75 is hydraulic, or other linkage to translate or gimbal the rocket capable of pumping any fuel or oxidizer found in regular use nozzle or nozzles, through a rotating connection. in rocket engines, including cryogenic and hypergolic pro pellants. The rotary rocket may also be configured for use 0034. When the RTR operates in the turbojet mode, its with liquid monopropellants and a catalyst that causes rapid steady state operation is essentially the same as that of a decomposition of the monopropellant and the catalyst can be turbojet without an integrated rotary rocket. Fuel 114 is sup plied to the combustor 10 of the turbojet where it mixes and located in any of a number of locations including the fuel or burns with air from the compressor 11. The combustion prod oxidizer feed lines 112, 113, the propellant pump 75, or the ucts flow through the turbojet's turbine 6, producing shaft combustion chamber(s) 73. power to drive the compressor. When these combustion prod 0031 FIG. 3 shows the combustion chamber configura ucts are exhausted through the RTR exit nozzle 8, jet thrust is tion of one embodiment on the invention in detail. Propellants produced. The transient operation of the turbojet differs from exit from the propellant pump 75 at a high velocity and that of the RTR. The increased mass of the RTR means that its pressure and enter into the combustion chamber or chambers rotational inertia is larger than that of the turbojet, which 73 of the rotary rocket 7. There, the propellants vaporize, mix, increases the time needed to bring the RTR up to, or down to, and combust to form a hot, high pressure gas which exits a given rpm relative to a conventional turbojet engine. The through the nozzle or nozzles 72 of the rotary rocket, and increased angular inertia of the RTR is actually a benefit to thereby produces thrust. Because the propellants enter the overall turbojet operation because it dampens thermody combustion chamber at high velocity, the walls of the cham namic and aerodynamic perturbations during turbojet and air ber 73 are preferably curved, allowing the flow direction of the propellants to be directed into each other to promote turbo rocket operation. atomization and mixing, and also preventing or eliminating 0035. For the RTR to operate in afterburning turbojet chamber erosion caused by the high-velocity impact of pro mode, fuel is supplied to the rotary rocket 7, but with little or pellants into the chamber walls. The combustion chamber 73 no oxidizer. The fuel 13 is pumped into the afterburner region of the rotary rocket 7 may comprise individual chambers, 9 of the RTR through the rotary rocket 7, with pump power interconnected chambers, or a single continuous annular being Supplied predominately or completely from the turbo chamber. By using a single annular chamber, propellant flow jet's turbine 6. Fuel pumped into the afterburner 9 is then rate variations among the propellant pump passages will not ignited and burned as per normal afterburner operation. The result in thrust variations among the rocket chambers 73. The turbojet operates as per normal, with any unreacted air from throats, or Smallest cross-sectional flow areas, of the rotary the turbojet mixing and burning with the fuel emerging from rocket typically lie at a constant radial distance from the axis the rotary rocket. Alternatively or additionally, fuel can be of rotation. They are sized based on the total propellant flow provided to the afterburner 9 using direct injection of fuel rate entering the combustion chamber(s) 73, and the number through one or more afterburner fuel lines 115. of rocket nozzles 72 used in the rotary rocket design. The 0036. For the RTR to operate in Air Turbo Rocket mode, rocket nozzles 72 extend from the throats to the aft face of the fuel 13 and some oxidizer12 are supplied to the rotary rocket rotary rocket. Each of the nozzles is of standard rocket nozzle 7, where the fuel-rich mixture is burned. As this mixture design and can, for example be in the form of a regular cone exhausts from the rotary rocket 7, it supplies torque power to with a 15° half-angle, or a contoured, bell-type nozzle. pump the propellants, and also provides power to drive the 0032 To produce the torque required to operate in various compressor 11. The fuel-rich rotary rocket exhaust mixes and modes of the RTR, some or all of the rocket nozzles 72 may burns with air from the compressor in the afterburner 9, thus be canted, or set at an angle C. relative to the axis of rotation producing thrust. Fuel through fuel lines 114 to the turbojet (FIG. 3). The magnitude of the total torque produced by the combustor 10 can be turned off in this mode, since all requi rotary rocket equals the magnitude of the thrust component in site power to drive the compressor can come from the rotary the tangential direction multiplied by the radius to the center rocket 7. There may also be operating modes unique to the line of the throat, and multiplied by the number of rocket RTR in which fuel and/or oxidizer may be supplied simulta noZZles possessing a cant angle. The power produced by the neously to some combination to the turbojet combustor 10, to US 2015/0007549 A1 Jan. 8, 2015

the afterburner 9, and to the rotary rocket 7, thus allowing purposes of operational efficiency and optimization. For engine performance and operational modes which have not example, during rotary rocket operation, the rotating ele heretofore been possible. ments of the turbojet are not required, and it may be desirable 0037. When the RTR is in rocket mode, fuel 13 and oxi to mechanically decouple the rotary rocket 7 from the turbojet dizer12 enter the rocket chambers 73 at a near-stoichiometric so that the turbojet's rotating components are not rotating. A mixture ratio, where they mix and burn. Combustion products similar condition may exist when the turbojet is operating and exit through the rocket nozzles 72 producing rocket thrust. the rotary rocket is not. However, there may be operating Pump power is supplied completely from the power produced modes in which it is desirable or required to have the rotary from the torque of the rotary rockets canted rocket nozzles. rocket 7 and turbojet mechanically coupled. Such as during No fuel is supplied to the turbojet combustor 10 or afterburner afterburner operation, or when operating in Air Turbo Rocket 9, and the turbojet's compressor 11 and turbine 6 may or may mode, or in other possible operating modes of the RTR. In not be mechanically coupled to the main shaft 4. order to be able to mechanically couple or de-couple the 0038 Because the shaft power produced by the rotary rotating components of the rotary rocket from those of the rocket may depend only on the propellant flow rate and rocket turbojet, the rotary rocket must be able to rotate indepen noZZle cant angle, the RTR can transition continuously or dently of the turbojet. One possible configuration is shown in abruptly between operating modes. This also allows multiple FIG. 6, wherein the main shaft 4 of the rotary rocket 7 pos modes to operate together or individually. Generally, the RTR sesses its own first set of bearings 5a, and the turbojet assem is operated in an air breathing mode for as long as possible to bly possesses its own second set of bearings 5b. These two take advantage of atmospheric air as a significant portion of assemblies can be mechanical coupled or de-coupled using a the reaction mass as well as a source of oxidizer. This mini coupling device 67. This configuration allows the rotary mizes propellant usage, resulting in longer flight ranges and/ rocket and the turbojet assembly to rotate separately or or, for suborbital or orbital flight, a higher maximum velocity. together, depending upon the commanded State of the cou A transition from SubSonic to SuperSonic flight may be pling device 67. The coupling device may comprise, for achieved in any operating mode. AS flight Velocity increases example, a clutch-type device another mechanism configured at SuperSonic speeds, air becomes shock-heated and more to alternately engage and disengage rotatable elements of the difficult to use. For this reason, the RTR can transition to pure turbojet with and from the rotatable elements of the rotary rocket mode at a flight speed that optimizes propellant con rocket. Sumption. The rotary rocket, however, may be engaged at any 0041 While several embodiments of the present invention time and at any flight speed, independently of the air-breath have been described and illustrated herein, those of ordinary ing mode in operation. For instance, engaging the rotary skill in the art will readily envision a variety of other means rocket can provide significant thrust augmentation in air and/or structures for performing the functions and/or obtain breathing modes at SubSonic or SuperSonic speeds. ing the results and/or one or more of the advantages described 0039. One important feature of RTR operation is the herein. Each of Such variations and/or modifications is power balance during the various operating modes. The deemed to be within the scope of the present invention. More power balance refers to the fact that the power input needed to generally, those skilled in the art will readily appreciate that operate Some of the components of the engine must come all parameters, dimensions, materials, and configurations from the power output generated by other components. To described herein are meant to be exemplary and that the actual reach a steady state of dynamic equilibrium, the total power parameters, dimensions, materials, and/or configurations will input required must be equal to the power output. Compo depend upon the specific application or applications for nents which require power input include the turbojet com which the teachings of the present invention is/are used. pressor 11, the rotary rocket pump 75, and losses from bear Those skilled in the art will recognize, or be able to ascertain ings 5 and seals 52,53. Components which produce a power using no more than routine experimentation, many equiva output include the turbojet turbine 6, and the torque power lents to the specific embodiments of the invention described produced by the rotary rocket 7. The power balance varies herein. The present invention is directed to each individual depending on the operating mode of the RTR. In turbojet and feature, system, article, material, and/or method described turbojet/afterburner modes, the power balance is achieved herein. In addition, any combination of two or more Such when the power output of the turbine 6 equals the sum of the features, systems, articles, materials, and/or methods, if Such power input to the compressor 11 plus bearing and seal losses. features, systems, articles, materials, and/or methods are not In Air Turbo rocket mode, power balance is achieved when mutually inconsistent, is included within the scope of the the power output from the rotary rocket 7 equals the sum of present invention. the power input to the compressor 11 plus the pump power to 0042 Unless otherwise defined, all terms (including tech operate the rotary rocket 7, plus bearing and seal losses. In nical and Scientific terms) used herein have the same meaning rocket mode, power balance is achieved when the power as commonly understood by one of ordinary skill in the art to output from the rotary rocket 7 equals the sum of the pump which this invention belongs. It will be further understood power required by the rotary rockets pump 75, plus bearing that terms, such as those defined in commonly used dictio and seal losses. And there may be unique RTR operating naries, should be interpreted as having a meaning that is modes in which the power balance is achieved through some consistent with their meaning in the context of the relevant art combination of power output from the turbine 6 and the rotary and will not be interpreted in an idealized or overly formal rocket 7 which then equals the power input to the compressor sense unless expressly so defined herein. 11, the pump power required by the rotary rocket pump 75, 0043. As used herein, the singular forms “a”, “an and and the bearings and seal losses. “the are intended to include the plural forms as well, unless 0040. Because of the co-axial configuration of the RTR, it expressly stated otherwise. It will be further understood that is possible to mechanically couple and uncouple the rotating the terms “includes.” “comprises.” “including and/or “com elements of the turbojet with those of the rotary rocket for prising, when used in this specification, specify the presence US 2015/0007549 A1 Jan. 8, 2015

of stated features, integers, steps, operations, elements, and/ 2. The rotary turbo rocket of claim 1, wherein said main or components, but do not preclude the presence or addition shaft is reversibly rotationally coupled to one or both of the of one or more other features, integers, steps, operations, turbojet engine and rotary rocket engine. elements, components, and/or groups thereof. It will be 3. The rotary turbo rocket of claim 1, and further compris understood that when an element is referred to as being “con ing a mechanism for coupling or de-coupling a rotation of the nected' or “coupled to another element, it can be directly rotary rocket engine to or from a rotation of the turbojet connected or coupled to the other element or intervening compressor. elements may be present. Furthermore, “connected” or 4. The rotary turbo rocket of claim 1, and further compris “coupled as used herein may include wirelessly connected ing a mechanism for coupling or de-coupling a rotation of the or coupled. As used herein, the term “and/or” includes any rotary rocket engine to or from a rotation of the turbojet and all combinations of one or more of the associated listed turbine. items. 5. The rotary turbo rocket of claim 1, wherein the rotary 0044) Multiple elements listed with “and/or should be rocket, main shaft and compressor are configured such that construed in the same fashion, i.e., "one or more' of the the compressor may be driven by power delivered from the elements so conjoined. Other elements may optionally be rotary rocket through the main shaft to the compressor. present other than the elements specifically identified by the 6. The rotary turbo rocket of claim 5, and further compris “and/or clause, whether related or unrelated to those ele ing a mechanism for mechanically coupling or de-coupling ments specifically identified. Thus, as a non-limiting the delivery of power from the rotary rocket to the compres example, a reference to “A and/or B, when used in conjunc tion with open-ended language Such as “comprising can SO. refer, in one embodiment, to A only (optionally including 7. The rotary turbo rocket of claim 1, wherein the rotary elements other than B); in another embodiment, to B only rocket, main shaft and compressor are configured such that a (optionally including elements other than A); in yet another rotation of the rotary rocket may be driven by power delivered embodiment, to both A and B (optionally including other to the rotary rocket through the main shaft from one or both of elements); etc. As used herein in the specification and in the the turbine and the compressor. claims, 'or' should be understood to have the same meaning 8. The rotary turbo rocket of claim 1, and further compris as “and/or as defined above. For example, when separating ing an afterburner positioned between said rotary rocket items in a list, 'or' or “and/or shall be interpreted as being engine and said rear facing nozzle outlet, and wherein said inclusive, i.e., the inclusion of at least one, but also including afterburner is configured to combust fuel in air delivered from more than one, of a number or list of elements, and, option the compressor, and optionally, fuel from the turbojet com ally, additional unlisted items. Only terms clearly indicated to bustor, or from fuel injected into the afterburner, or from the contrary, such as “only one of or “exactly one of” or, excess fuel emerging from the rotary rocket. when used in the claims, “consisting of will refer to the 9. The rotary turbo rocket of claim 1, further comprising a inclusion of exactly one element of a number or list of ele seal assembly and wherein: ments. In general, the term 'or' as used herein shall only be said main shaft comprises one or more propellant feedlines interpreted as indicating exclusive alternatives (i.e. “one or which provide a flow path for propellant to one or more the other but not both') when preceded by terms of exclusiv rotary rocket combustion chambers and ity, such as “either,” “one of “only one of or “exactly one said seal assembly is configured to provide a flow path of “Consisting essentially of when used in the claims, shall between one or more propellant storage tanks and said have its ordinary meaning as used in the field of patent law. one or more propellant feed lines. 0045. As used herein in the specification and in the claims, 10. The rotary turbo rocket of claim 9, wherein said one or the phrase “at least one.” in reference to a list of one or more more propellant feed lines comprises a central fluid passage elements, should be understood to mean at least one element and an annular fluid passage. selected from any one or more of the elements in the list of elements, but not necessarily including at least one of each 11. The rotary turbo rocket of claim 9, wherein: and every element specifically listed within the list of ele said one or more propellant feed lines comprises one or ments and not excluding any combinations of elements in the more axial channels in fluid communication with radial list of elements. holes around a perimeter of the shaft and said seal assembly is configured to provide a flow path 1. A rotary turbo rocket comprising: between one or more propellant storage tanks and said a forward facing air inlet; one or more axial channels through said radial holes. a rear facing nozzle outlet; 12. The rotary turbo rocket of claim 9, wherein said seal assembly comprises a fuel inlet, a fuel seal, an oxidizer inlet, a turbojet engine having alongitudinal axis and comprising an oxidizer seal a purged space, and a purge gas inlet posi a compressor, a combustion chamber, and a turbine; tioned such that fuel and/or oxidizer that leak past the fuel seal a rotary rocket engine having a longitudinal axis and com and/or the oxidizer seal may be purged from said purged prising a combustion chamber and one or more rocket space by a purge gas delivered through said purge gas inlet to outlets; and prevent inadvertent mixing a fuel and an oxidizer. a main shaft Supporting rotating components of the turbojet 13. The rotary turbo rocket of claim 1, wherein said rocket and rotary rocket engine outlets comprise rocket nozzles positioned at a common wherein the turbojet and rotary rocket engines are arranged radius from the main shaft and wherein said nozzles are coaxially with the turbojet engine positioned behind the configured to provide a component of thrust at a tangent to the air inlet and in front of the rotary rocket and the nozzle longitudinal axis of rotation of said rotary rocket such that outlet positioned behind the rotary rocket engine. thrust produced by the rotary rocket rotates said main shaft. US 2015/0007549 A1 Jan. 8, 2015

14. The rotary turbo rocket of claim 13, further comprising rotary turbo rocket is transferred to the rotary rocket, and a linkage between the main shaft and the nozzles of the rotary thereby pumps fuel and oxidizer into the afterburner of the rocket and wherein said linkage is configured to change a cant rotary turbo rocket. angle C. of the rocket nozzles. 19. A method for operating a rotary turbo rocket, said 15. The rotary turbo rocket of claim 1, wherein said rotary method comprising: rocket engine comprises a housing containing said rotary varying a fuel and/or an oxidizer flow to a turbojet engine rocket combustion chamber and said housing comprises one and a rotary rocket engine to change a mode of said or more passages configured to deliver one or more propel rotary turbo rocket turbojet between a rotary rocket lants from the main shaft radially to said combustion chamber mode, an air turbo rocket mode, and/or a turbojet mode and wherein rotation of the rotary rocket engine provides a wherein the turbojet engine and rotary rocket engine are pumping force Sufficient to pump said propellant into said arranged coaxially with the turbojet engine positioned in combustion chamber. front of the rotary rocket engine and wherein the rotary 16. The rotary turbo rocket of claim 15, wherein said hous rocket engine and the turbojet engine are rotationally ing comprises separate passages configured to deliver fuel coupled through a main shaft positioned alone a center and oxidizer radially outwards from fuel and oxidizer pas line of a longitudinal axis of the turbojet engine and a sages in the main shaft to one or more combustion chambers center line of a longitudinal axis of the rotary rocket contained in the housing. engine. 17. The rotary turbo rocket of claim 1, further comprising a mechanism for changing a cant angle C. of the rotary rocket 20. The method of claim 19, and further comprising chang noZZle. ing a power output from the rotary rocket by changing a cant 18. The rotary turbo rocket of claim 1, wherein shaft power angle C. of one or more rocket nozzles on said rotary rocket. produced by the turbine from the turbojet component of the k k k k k