Advanced Automated Fabrication System And Methods For Thermal And Mechanical Components Utilizing Quadratic Or Squared Hybrid Direct Laser Sintering, Direct Metal Laser Sintering, CNC, Thermal Spraying, Direct Metal Deposition And Frictional Stir Welding. Cross-reference To Related Applications

Abstract

ADVANCED AUTOMATED FABRICATION SYSTEM AND METHODS FOR THERMAL AND MECHANICAL COMPONENTS UTILIZING QUADRATIC OR SQUARED HYBRID DIRECT LASER SINTERING, DIRECT METAL LASER SINTERING, CNC, THERMAL SPRAYING, DIRECT METAL DEPOSITION AND FRICTIONAL STIR WELDING. CROSS-REFERENCE TO RELATED APPLICATIONS

Claims

1. An automated fabrication system with methods for producing thermal and mechanical fabrications, the system and methods comprising:

2. An enclosed automated apparatus for producing a component from a powder, comprises at least one of a: a) Means for consecutively dispensing a plurality of layers of powder within a boundary to a target surface; and b) An energy source; and c) Means for beam management utilizing mirrors on X axis, Y axis and focusing lens system for Z axis beam diameter control; and d) Means of a transparent thermal barrier between beams and build area e) A computer control system with artificial intelligence using machine learning for monitoring, analysis of 3d object and 2d sliced layers to include controlling the system; and f) Scanning system consisting of at least one method of 3D object electromagnetic radiation scanning used with 3d object data and 2d sliced layer analysis with fabricated layer; and g) A method for directing the energy source at locations of each dispensed layer of powder at the target surface corresponding to cross-sections of the component to be produced therein and fusing the powder thereof; and h) Means for a counter rotating barrel dispensing powder near said target surface comprises at least one of a: i. A thermally controlled counter rotating barrel; ii. Means for moving said counter rotating barrel across said target surface in contact with said powder; and iii. Means for rotating said counter rotating barrel to a direction of said movement of said counter rotating barrel across said target surface; iv. wherein said movement and said counter-rotation of said barrel distribute a layer of powder over said target surface. i) Thermal control means via gas exchange for moderating the temperature difference between unfused powder in a top layer of powder at the target surface and the material holding sump(s) and laser fused monolithic component in the one of the plurality of layers of powder immediately beneath the topmost layer j) Means for cartridge based build area and transfer method thereof and comprises at least one of a: i. Build platform assembly; ii. An actuator; said actuator comprised by a lift mechanism; iii. An enclosure; said enclose comprising metal supports, metal casing, metal sheets; iv. Thermal communication channeling medium; v. Carriage for transfer means; k) Means for sealing and pressurizing fabrication system;

3. An enclosed automated apparatus for producing a component from a powder and wire, comprises at least one of a: a) Means for consecutively dispensing a plurality of layers of powder and wire within a boundary to a target surface; and b) An energy source; and c) Means for beam management utilizing mirrors on X axis, Y axis and focusing lens system for Z axis beam diameter control; and d) Means for transparent thermal barrier between beams and build area e) A computer control system with artificial intelligence using machine learning for monitoring, analysis and controlling the system; and f) Scanning system consisting of at least one method of 3D object electromagnetic radiation scanning used with 3d object data and 2d sliced layer analysis with fabricated layer; and scanning system consisting of at least one type of 3D object scanner, thermal or optical or light based sensor, x-ray, sonic scanning; and g) A method for directing the energy source at locations of each dispensed layer of powder at the target surface corresponding to cross-sections of the part to be produced therein and fusing the powder thereof; and h) A method for directing the energy source and wire at locations of each targeted layer at the target surface corresponding to cross-sections of the part to be produced therein and fusing the powder at the target location; and l) Means for a counter rotating barrel dispensing powder near said target surface comprises at least one of a: i. A thermally controlled counter rotating barrel; ii. Means for moving said counter rotating barrel across said target surface in contact with said powder; and iii. Means for rotating said counter rotating barrel to a direction of said movement of said counter rotating barrel across said target surface; iv. wherein said movement and said counter-rotation of said barrel distribute a layer of powder over said target surface. i) A method for directing removal of powder material at locations of each targeted layer at the target surface corresponding to cross-sections of the part to be produced therein; and j) Thermal control means via gas exchange for moderating the temperature difference between unfused powder in a top layer of powder at the target surface and the material holding sump(s) and laser fused monolithic component in the one of the plurality of layers of powder immediately beneath the topmost layer; and m) Means for portable cartridge based build area with transfer method thereof and comprises at least one of a: i. Build platform assembly ii. An actuator; said actuator comprised by a lift mechanism; iii. An enclosure; said enclose comprising metal supports, metal casing, metal sheets; iv. Thermal communication channeling medium v. Carriage for transfer means n) Means for sealing and pressurizing fabrication system

4. An enclosed automated apparatus for producing a component from a powder, comprises at least one of a: o) Means for consecutively dispensing a plurality of layers of powder within a boundary to a target surface; and p) An energy source; and q) Means for beam management utilizing mirrors on X axis, Y axis and focusing lens system for Z axis beam diameter control; and r) Means of a transparent thermal barrier between beams and build area s) A computer control system with artificial intelligence using machine learning for monitoring, analysis and controlling the system; and t) Scanning system consisting of at least one method of 3D object electromagnetic radiation scanning used with 3d object data and 2d sliced layer analysis with fabricated layer; and u) A method for directing the energy source at locations of each dispensed layer of powder at the target surface corresponding to cross-sections of the component to be produced therein and fusing the powder thereof; and v) Means for a scraper to dispense powder near said target surface comprises at least one of a: i. A scraper; ii. Means for moving said scraper across said target surface in contact with said powder; and wherein said movement of said scraper distribute a layer of powder over said target surface. w) Thermal control means via gas exchange for moderating the temperature difference between unfused powder in a top layer of powder at the target surface and the material holding sump(s) and laser fused monolithic component in the one of the plurality of layers of powder immediately beneath the topmost layer x) Means for cartridge based build area and transfer method thereof and comprises at least one of a: i. Build platform assembly; ii. An actuator; said actuator comprised by a lift mechanism; iii. An enclosure; said enclose comprising metal supports, metal casing, metal sheets; iv. Thermal communication channeling medium; v. Carriage for transfer means; y) Means for sealing and pressurizing fabrication system;

5. The apparatus of claim 2, wherein said thermal control means further comprises: heater, cooling, heat exchanger to transfer thermal energy for thermal control of a gas; and means for directing the thermal controlled gas at the target surface and exhaust means for exhausting directed thermally controlled gas from the vicinity of the target surface.

6. The apparatus of claim 2, wherein said energy source comprises a quad laser array; and wherein said controller comprises: a computer; and lens and mirrors controlled by said computer to direct the width of the beams and aim and focus of the beams from the quad array of lasers.

7. The apparatus of claim 6, wherein said controller further comprises: interface hardware, coupled to said computer, to enable and disable the quad laser array as its targeted energy beam is moved across the targeted surface.

8. The apparatus of claim 7, wherein the computer is programmed with the defined boundaries of each cross-section of the part.

9. The apparatus of claim 7, wherein the computer comprises means for determining the defined boundaries of each layer of the part from the overall dimensions of the part.

10. The apparatus of claim 6, wherein said controller further comprises: interface hardware, coupled to said computer, to enable and disable the direct material depositing as its target is moved across the targeted surface.

11. The apparatus of claim 10, wherein the computer is programmed with the defined boundaries of each cross-section of the part whereas computer comprises means for determining the defined boundaries of each layer of the part from the overall dimensions of the part.

12. The apparatus according to claim 2, wherein the automated Computer Numerical Control (CNC) is the automation of machine tools by means of computers executing pre-programmed sequences of machine control commands whereas performing CNC finalization process comprises means for cutting, smoothing, polishing, spraying, coating or joining components; An automated CNC machine tool control system for a CNC machine tool of the type comprising a controllable, movable tool for processing a fabrication component, means for receiving control instructions describing processing functions to be performed on the fabrication component, a processing unit and memory means, comprises at least one of a: a) means for receiving and storing in the memory means fabrication component shaping instructions from 3 dimensional computer aided design data; b) means for transmitting command signals to a movable tool to thereby cause the movable tool to move; and c) means for generating control signals, said generating means including an object oriented software program comprising a plurality of objects, each said object including a plurality of instructions and associated data, said generating means including message means for transmitting information between said objects, at least one of said objects including a model of the processes to be performed on a fabrication component by the movable tool, said generating means coupled to said message means, said generating means generating control signals responsive to messages from said processing objects, said generating means communicating said control signals to said transmitting means.

13. Fabrication means utilizing the apparatus according to claim 11, wherein layers are fused to form a monolithic heat exchanger comprised by: a) Fusing layers of at least one type of materials consisting of powder or wire; and b) A manifold extending between axially opposed ends and having first inlet means and first outlet means for respectively permitting the ingress and egress of a first heat exchange fluid; and c) A pair of end members fused to the axially-opposed ends of the manifold to define an internal chamber therein having an intermediate region disposed between two opposite non intermediate end regions of the chamber, the end members having second inlet means and second outlet means for respectively permitting the ingress and egress of a second heat exchange fluid; and d) A plurality of uniformed rounded zig-zag channels extending from an end member to an adjacent end member for the first heat exchange fluid; and e) A plurality of uniformed rounded zig-zag channels extending from an end member to an adjacent end member for the second heat exchange fluid

14. The apparatus of claim 11, wherein means for fabrication of a supercritical, transcritical and subcritical carbon dioxide turbine system, wherein said supercritical, transcritical and subcritical carbon dioxide turbine system comprises a plurality of turbines, compressors, evaporators, absorbers, heat exchangers and condensers.

15. The apparatus of claim 11, wherein provides means for fabrication of a monolithic axial turbine rotor with internal cooling channels, wherein said axial flow turbine comprising: a. rotor blades fabricated via powder bed with internal cooling channels further comprised by smoothing and polishing fabrication comprised by method of claim 12; and b. rotor hub fabricated via at least one method: i. powder bed ii. cnc machined iii. casting

16. The apparatus of claim 11, wherein provides means for fabrication of a monolithic radial flow turbine impeller with internal cooling channels, wherein finalization of said radial flow turbine impeller fabrication is comprised by smoothing and polishing fabrication comprised by method of claim 12.

17. Fabrication means utilizing the apparatus according to claim 14, wherein layers are fused to form monolithic component builds of a supercritical, transcritical, subcritical turbine system comprises at least one of a: a. Carbon dioxide storage, pump and valve: and b. A high temperature recuperator; and c. A medium temperature recuperator; and d. A low temperature recuperator; and e. A Heat Exchanger; and f. A precooler; and g. A condenser; and h. An evaporator; and i. An impeller and/or propeller with internal cooling channels; and j . A modular sealing and bearing cartridge; and k. A compressor 1. A turbine

18. The turbine system of claim 17, wherein the supercritical, transcritical, subcritical carbon dioxide turbine operates at a temperature of at least approximately 250 degrees Fahrenheit.

19. The turbine system of claim 17, wherein the supercritical, transcritical, subcritical carbon dioxide turbine comprises a supercritical carbon dioxide Brayton power conversion cycle utilizing heat exchangers.

20. A modular sealing and bearing turbine cartridge comprises: a) At least one Primary Shaft Sleeve b) At least one Intermediate Sleeve c) At least one Inner Sleeve d) At least one Adjustable Threaded Collar e) At least one Upper Lock Collar f) At least one Upper Lock Ring g) At least one Lower Lock Collar h) At least one Lower Lock Ring i) At least one Outer Labyrinth j) At least one Optional Inner Labyrinth(s) k) At least one Intermediate Labyrinth l) At least one Outer Leveling Pad m) At least one Inner Leveling Pad n) At least one Outer Stationary Seal Bearing o) At least one Inner Thrust Bearing p) At least one Thrust Ring q) At least one Outer Thrust Bearing r) At least one Stationary Seal s) At least one Tilting Journal Pad t) At least one Spring u) At least one Inner Stationary Seal Bearing v) At least one Inner Journal Bracket w) At least one Outer Journal Bracket x) At least one monolithic channeled housing

21. The process of claim 14 wherein said supercritical, transcritical and subcritical carbon dioxide turbine system further comprising: a. External thermal input connects to primary heat exchanger HX1 that converts and transfers external generated thermal energy input to inject thermal energy into the carbon dioxide Brayton top cycle b. Ducting from HX1 connects to the primary turbine T1 connected to generator/alternator 1 connected to main compressor MC and ducting to provide input to secondary turbine T2 and generator/alternator 2 connected to recompressor RC c. Gas film compressor BC provides pressure boost to gas ducted to gas supported bearings (turbine bearings) connected to at least one: motor, engine, turbine d. Ducting from turbine T1 and turbine T2 connects thermal input to high temperature recuperator/heat exchanger HX2 then ducted to low temperature recuperator/heat exchanger HX3 e. Ducting from HX3 connects thermal input to gas pre-cooler/heat exchanger HX4 then ducted connects thermal input to condenser, f. Ducting from HX3 then connects thermal input to transcritical turbine 3 connected to generator/alternator 3 g. Pump P1 is connected to at least one: transcritical turbine 3, at least one individual standalone motor, engine, turbine h. Secondary compressor SC is connected to at least one: transcritical turbine 3, at least one individual standalone motor, engine, turbine i. Duct to connect between transcritical turbine 3 and heat exchanger HX5 connected to heat exchanger HX6 that is connected to Heat exchanger HX7 connected to an expansion valve and then connected to an evaporator. j. Pump P2 is connected to cot expansion tank and accepts input from CO2 Storage k. CO2 expansion tank refills the CO2 cycles via ducting to upper Brayton and lower Brayton cycles.

22. A method of generating electricity, heating and cooling with a supercritical, transcritical, subcritical carbon dioxide turbine, the method comprising: Transfer of thermal energy heating a heat transfer fluid to a temperature of at least about 250 degrees Fahrenheit from the thermal energy source; transporting energy from the heat transfer fluid to heat a Brayton cycle working fluid of the supercritical, transcritical, subcritical carbon dioxide turbine system; passing the heated Brayton cycle working fluid through the supercritical Brayton cycle; and thermal energy communication of the Brayton cycle working fluid from the supercritical carbon dioxide turbine with a high temperature recuperator to the transcritical Brayton cycle: and thermal energy communication of the Brayton cycle working fluid from the medium temperature recuperator to the subcritical Brayton cycle;

23. The method of claim 21, wherein carbon dioxide thermal transfer fluid transports the thermal energy for the Brayton cycle working fluid of the supercritical, transcritical, subcritical carbon dioxide turbine system comprises using a heat exchanger.

24. The method of claim 21, wherein the supercritical carbon dioxide turbine system comprises a supercritical, transcritical, subcritical carbon dioxide Brayton power conversion cycle. a) A Brayton cycle working fluid for providing energy to the supercritical, transcritical and subcritical carbon dioxide turbines; and b) a high temperature recuperator that receives the Brayton working fluid from the supercritical carbon dioxide turbine and thermal energy communicates it; and c) A medium temperature recuperator that receives the Brayton working fluid from the high temperature recuperator and thermal energy communicates it; and d) a low temperature recuperator that receives the Brayton working fluid from the high temperature recuperator and thermal energy communicates it; e) A precooler;

25. Fabrication means utilizing the apparatus according to claim 11, wherein layers are fused in a single build of monolithic components to form high temperature fuel cell system comprised by: A cooling channel, an anode channel, an anode inlet and an anode outlet, a first anode channel portion proximal to the anode inlet, a second anode channel portion proximal to the anode outlet, and a gas separation means operable to enrich a hydrogen gas component of an anode exhaust gas exiting the anode outlet to produce a first product gas enriched in the said hydrogen gas component such that at least a portion of the first product gas enriched in the hydrogen gas component can be provided as a portion of a fuel mixture supplied to the anode inlet

26. The high temperature fuel cell system according to claim 25 wherein the high temperature fuel cell comprises HDLS fused monolithic plates and monolithic ends to form components of a solid oxide fuel cell.

27. The high temperature fuel cell system according to claim 25, wherein the anode and cathode channels are arranged such that allows uniform placement cooling channels to moderate excessive heat and reduction of thermal hot spots within the fuel cell.

28. The high temperature fuel cell system according to claim 27, wherein the anode and cathode channels are arranged such that the fuel gas mixture in the anode channel is capable of flowing in a direction countercurrent to a flow of the oxygen-enriched gas in the cathode channel.

29. The high temperature fuel cell system according to claim 25, wherein the first anode channel portion comprises an anode material mixture thereof, and the second anode channel portion comprises a selected anode material.

30. The high temperature fuel system according to claim 25, wherein the high temperature fuel cell engages an internal thermal management system to moderate thermal energy from within the fuel cell assembly

31. A thermal energy management system for solid oxide fuel cells comprising: a monolithic heat exchanger comprising a coolant inlet port, a coolant outlet port, and a plurality of cell channels for passing a flow of coolant there through; said monolithic heat exchanger being connected to an SOFC stack; and a seal material disposed between said SOFC stack and said heat exchanger to control thermal connection and coolant between said SOFC stack and said heat exchanger; wherein in operation, a flow of inlet coolant having a selected temperature is passed through said heat exchanger cell channels and thermal energy flowing into and out of said SOFC stack is managed primarily by a thermal transfer fluid connection between said SOFC stack and said heat exchanger.

32. The thermal energy management system of claim 31, wherein said heat exchanger preheats or cools one or a combination of input fuel stream and oxidizing gas stream feeding said SOFC stack.

33. The thermal energy management system of claim 31, further comprising: communication provided between said SOFC stack and said heat exchanger and configured to control thermal coupling between said SOFC stack and said heat exchanger.

34. The thermal energy management system of claim 31, further comprising: a sealing material between said SOFC stack and said heat exchanger to control thermal connection between said SOFC stack and said heat exchanger; and connection between said SOFC stack and said heat exchanger and configured to control thermal connection between said SOFC stack and said heat exchanger.

35. A method for managing the thermal energy flowing into and out of an SOFC system comprising: connecting a heat exchanger to an SOFC stack, said monolithic heat exchanger comprising a coolant inlet side for introducing a flow of coolant, a plurality of cells for passing a flow of coolant there through, and a coolant outlet side for discharging said flow of coolant; seal material between said SOFC stack and said heat exchanger and configuring said seal material to control thermal connection between said SOFC stack and said heat exchanger; and transfer of said coolant having a selected temperature through said heat exchanger cell channels so as to manage thermal energy flowing into and out of said SOFC stack primarily by coolant connection between said SOFC stack and said heat exchanger.

36. The method of claim 35, further comprising: preheating or cooling one or a combination of input fuel stream and oxidizing gas stream feeding said SOFC stack with said heat exchanger.

37. A thermal energy management system for solid oxide fuel cells comprising: an HDLS fused monolithic heat exchanger comprising a coolant inlet side, a coolant outlet side, and a plurality of cells for passing a flow of coolant there through; said heat exchanger being coupled to an SOFC stack; and a material disposed between said SOFC stack and said monolithic heat exchanger to control thermal coupling between said SOFC stack and said heat exchanger; wherein in operation, a flow of inlet air having a selected temperature is passed through said heat exchanger cells and thermal energy flowing into and out of said SOFC stack is managed primarily by radiation coupling between said SOFC stack and said heat exchanger.

38. The thermal energy management system of claim 37, further comprising: an air gap disposed between said SOFC stack and said monolithic heat exchanger and configured to control thermal coupling between said SOFC stack and said monolithic heat exchanger.

39. The thermal energy management system of claim 37, wherein said material is selected from the group consisting of a high emissivity material, a metal wall, metal media, or particles or a combination thereof.

40. The thermal energy management system of claim 37, wherein said monolithic heat exchanger is a HDLS fused monolithic heat exchanger.

41. A method for managing the thermal energy flowing into and out of an SOFC system comprising: A. Connection of a HDLS fused monolithic heat exchanger to an SOFC stack, said heat exchanger comprising a coolant inlet side for introducing a flow of coolant, a plurality of cells for passing a flow of coolant there through, and a coolant outlet side for discharging said flow of coolant; B. disposing a material between said SOFC stack and said heat exchanger and configuring said material to control thermal coupling between said SOFC stack and said heat exchanger; and C. passing said coolant having a selected temperature through said heat exchanger cell channels so as to manage thermal energy flowing into and out of said SOFC stack primarily by coolant connection between said SOFC stack and said heat exchanger.

42. Fabrication means utilizing the apparatus according to claim 11, wherein layers are fused to form a monolithic advanced gas cooled fast nuclear reactor comprised by: a. A monolithic pressure reactor vessel adapted to contain nuclear fuel therein, said monolithic vessel being adapted for operation with said advanced gas cooled fast nuclear reactor whereby it will become radioactively contaminated in the course of its operative life; and b. A shield structure including: 1. a hdls fused monolithic reactor chamber for housing the reactor vessel during its operative life; and 2. a hdls fused monolithic extraction chamber above the reactor chamber in communicating relationship with the reactor chamber and capable of receiving the reactor control rods during transfer, maintenance and at the expiration of its operative life for at least a time sufficient to permit the thermal generation to decay to acceptable levels; and 3. hdls fused monolithic pressure vessel with a plurality of isolated heat exchanger cores c. A platform supporting the hdls fused monolithic reactor vessel within the monolithic reactor chamber, said platform being capable of permitting upward movement of said reactor control rods into the monolithic extraction chamber; and d. means for supporting said platform; e. means for engaging support from said platform; and f. means for engaging the reactor core rods from the hdls monolithic reactor chamber to the monolithic extraction chamber at the expiration of said operative life.

43. An advanced gas cooled fast nuclear reactor according to claim 42 wherein: a. said platform supporting the monolithic pressure reactor vessel within the monolithic reactor chamber is adapted for upward movement of the control rods into the extraction chamber at the expiration of said operative life, maintenance or shipping; and b. said means for engaging the monolithic reactor core rods from the monolithic reactor chamber to the extraction chamber includes a mechanical system operatively connected to said platform whereby at the expiration of the operative life of the reactor vessel, maintenance or shipping the mechanical system may be activated to cause upward movement of said platform and said exhausted monolithic reactor control rods into the extraction chamber.

44. at least one movable support column positioned within the monolithic extraction chamber for supporting the platform during the operative life, maintenance and shipping of the reactor vessel in a position defining the top of reactor chamber;

45. at least one spring positioned such that upon activation thereby allowing the reactor control rods to elevate into the monolithic extraction chamber.

46. An advanced gas cooled fast nuclear reactor according to claim 42 wherein said monolithic extraction chamber includes:

47. An advanced gas cooled fast nuclear reactor according to claim 46 wherein said mechanical system comprises:

48. An monolithic advanced gas cooled fast nuclear reactor according to claim 47 including: a. movable support means positioned within the monolithic extraction chamber for supporting the platform during the operative life, maintenance and shipping of the reactor vessel in a position defining the top of the reactor chamber; and b. an access way leading into the upper portion of the extraction chamber to permit access into the extraction chamber for the purpose of removing said support means in preparation for replacement of said exhausted reactor vessel core material.

49. A monolithic advanced gas cooled fast nuclear reactor according to claim 48 including: a. an access way leading into the upper portion of the monolithic extraction chamber through which locking mechanism may be engaged to allow removal and replacement of the monolithic advanced gas cooled fast nuclear reactor.

50. An advanced gas cooled fast nuclear reactor according to claim 49 wherein a. said monolithic reactor chamber is located above ground level, and B. said monolithic extraction chamber is located above the reactor chamber level.

51. Fabrication utilizing the apparatus according to claim 11, wherein layers are fused to form a monolithic build liquid rocket engine components consisting of a thrust chamber, throat, exhaust and pintle injector comprised by: For a space vehicle a single build monolithic constructed rocket engine body with no welds or bolted connections for providing propulsion force, said rocket engine having an pintle injector for feeding oxygen and hydrogen into a thrust producer means consisting of a single thrust chamber, a turbopump supplied source of liquid methane connected via coolant channels within the rocket body and a turbopump source of liquid oxygen connected to said injector and being located relative to said thrust chamber so that the center of the rocket engine body forms a mounting and sealing system for said pintle injector.

52. A rocket engine as in claim 51 wherein said connecting means includes pintle injector providing flow and pressure control and shutoff of fuel and oxidizer for said rocket engine.

53. A rocket engine as in claim 52 wherein a turbopump means includes a turbine driven by methane and oxygen exhaust after being in indirect heat exchange relationship with a prebumer, a first pump for oxygen and a second pump for hydrogen, and said turbopump impellers powering said first pump and said second pump.

54. A liquid fuel rocket engine having a turbopump for boosting the pressure of fuel component and for boosting the pressure of oxidizer component, two pressure driving means for pressurizing said fuel and said oxidizer, a combustor wherein said pressurized fuel and oxidizer are fed through a pintle injector into the combustion chamber to produce a mixed fuel and oxidizer combustion gas to be discharged outwardly, a combustor chamber cooling jacket mounted operatively around the circumference of said combustion chamber means, a throat area connected to a high expansion nozzle extending from said combustor, and an expansion nozzle cooling jacket disposed operatively around the circumference of said high expansion nozzle means, respectively;

55. The liquid fuel rocket engine of claim 51, wherein said engine is further characterized in that a direct fuel based cooling channel is disposed between said turbopump means and said rocket engine body.

56. The liquid fuel rocket engine of claim 51, wherein said engine is further characterized in that a direct oxidizer cooling channel is disposed between said turbopump means and said pintle injector.

57. Fabrication utilizing the apparatus according to claim 11, wherein layers are fused to form an Axisymmetric Rocket-Based Air-augmented Combined Cycle propulsion system rocket engine comprised by: A monolithic rocket engine with scram engine thrust producing engine that has either rocket engine operation, air breathing operation with assistance of the rocket engine or continuous air breathing comprising of: an outer frame to connect the following components, symmetrical annular air intake compression ramps attached to the outer edges of the aerospike ramps center, an axial flow air diffuser area, flame area and compressor area, annular aerospike thrust cells connected to the annular thrust wall which provides the exhaust expansion ramp for the engine, air breathing combustors located at the beginning of the air compression ramp, a liquid fuel turbopump, liquid oxygen turbopump turbine, several linear actuators to change the air compression ramp geometry for thrust vectoring, and a control system to control basic engine functions, such as throttles for both air and fuel, air intake ramp shape, output ramp shape, fuel and oxidizer supply valves and an ignition system.

58. A turbopump turbine described in claim 57, which either drives liquid fuel pump or a liquid oxygen pump for the rocket that is controlled by a fuel and oxidizer valves, but does not drive both pump and air compressor simultaneously.

59. An air supply from claim 57 consisting of several movable annular cone mechanically connected in which compress the incoming air by a ram effect and can be moved by the attached linear actuators to change the air compression ramp geometry.

60. An annular arrangement of thrust cell which form an annular rocket thrust in claim 57 that serves as an exhaust expansion ramp for both said air breathing scram jets and said liquid rocket thrust cells while having the ability to thrust vector the rocket thrust without changing the air breathing exhaust ramp geometry.

Description

BRIEF DESCRIPTION OF THE DRAWINGS

[0287] The present disclosure is best understood from the following detailed description when read with the accompanying Figures. It is emphasized that, in accordance with the standard practice in the industry, various features are not drawn to scale. In fact, the dimensions of the various features may be arbitrarily increased or reduced for clarity of discussion. In order that the invention may be readily understood, one embodiment of the invention is illustrated by way of example in the accompanying drawings. Further details of the invention and its advantages will be apparent from the detailed description included below.

[0288] FIG. 1 is an internal view of the present inventions quadratic or squared HDLS with 3D object scanners shown

[0289] FIG. 2 is an internal view of the present inventions quadratic or squared HDLS with 3D object scanners, thermal vents and transparent thermal barrier shown

[0290] FIG. 3 is an internal view of the present inventions quadratic or squared HDLS with 3D object scanners, thermal vents, transparent thermal barrier and cartridge carriage/rails/track shown

[0291] FIG. 4 is an external view of the present inventions quadratic or squared HDLS with support services moved external to the main machine

[0292] FIG. 5 is an external view of the present inventions quadratic or squared HDLS with example method of a potential door and seal

[0293] FIG. 6 is an external view of the present inventions quadratic or squared HDLS with open CNC Processing System

[0294] FIG. 7 is an external view of the present inventions quadratic or squared HDLS with enclosed. CNC Processing System

[0295] FIG. 8 is an external view of the present inventions quadratic or squared HDLS with open CNC Processing System and Manual/Robotic Transport.

[0296] FIG. 9 is an external view of the present inventions quadratic or squared HDLS with open CNC Processing System with Tank or Vessel Construction System and Manual/Robotic Transport.

[0297] FIG. 10 is an external top view of the present inventions quadratic or squared HDLS with enclosed CNC Processing Stations and. Manual/Robotic Transport, view of the present inventions quadratic or squared HDLS and enclosed work area capable of single and multiple arm with CNC system and tool system to contain and exhaust gases and fumes additionally contain materials within the CNC work area.

[0298] FIG. 11 is an internal view of the present inventions quadratic or squared HDLS cartridge

[0299] FIG. 12 is an internal view of the present inventions quadratic or squared HDLS cartridge with thermal transfer with heated bed and completing the HDLS 3D thermal management system

[0300] FIG. 13 is an internal view of the present inventions quadratic or squared HDLS cartridge with thermal venting and material capture

[0301] FIG. 14 is an exposed view of the present inventions quadratic or squared HDLS cartridge with thermal management ducting and material capture shown

[0302] FIG. 15 is a view of a HDLS Quad Laser Array, HDLS Quad Laser Array with w/DMD Insert and Repair System, HDLS Quad Laser Array with multiple DMD Insert and Repair System

[0303] FIG. 16 HDLS Quad Laser Array with thermal gas vents layout and individual laser setup with focus lens system and targeting X axis and Y axis galvo with thermal barrier shown

[0304] FIG. 17 is an external view of the present inventions quadratic or squared HDLS with internal and CNC mounted DMD injection and repair attachment

[0305] FIG. 18 is a view of the CNC Processing System friction stir welding attachment

[0306] FIG. 19 is a view of examples of honeycomb design techniques

[0307] FIG. 20 is a view of example of friction stir welding techniques

[0308] FIG. 21 is a schematic of a Supercritical, Transcritical and Subcritical Energy Conversion System according to one or more aspects of the present disclosure.

[0309] FIG. 22 is a schematic of a Supercritical, Transcritical and Subcritical Energy Conversion System according to one or more aspects of the present disclosure.

[0310] FIG. 23 is a view of the example of Dual CO2 Cooling as optional to Supercritical, Transcritical and Subcritical Energy Conversion System FIG. 24 is a view of the configuration of a LABSLinear Advanced Bearing and Seal System: 1. Primary Shaft Sleeve 2. Intermediate Sleeve 3. Inner Sleeve 4. Adjustable Threaded Collar 5. Upper Lock Collar 6. Upper Lock Ring 7. Lower Lock Collar 8. Lower Lock Ring 9. Outer Labyrinth 10. Optional Inner Labyrinths 11. Intermediate Labyrinth 12. Inner Labyrinth 13. Outer Level Pad 14. Inner Level Pad 15. Outer Stationary Seal Bearing 16. Inner Thrust Bearing 17. Thrust Ring 18. Outer Thrust Bearing 19. Stationary Seal 20. Tilting Journal Pad 21. Spring 22. Inner Stationary Seal Bearing 23. Inner Journal Bracket 24. Outer Journal Bracket 25 Secondary Labyrinth

[0311] FIG.25 is a view of the configuration of a LABSLinear Advanced Bearing and Seal System

[0312] FIG. 26 is a view of the configuration of a LABSLinear Advanced Bearing and Seal System

[0313] FIG. 27 is a view of journal pads, journal pad surface and pivot placements, thrust bearing example and thrust bearing design example

[0314] FIG. 28 is a view of the examples of Multistage CO2/Steam TurbineAlternate Cooling/Lubrication System Designs with internal cooling channels

[0315] FIG. 29 is a view of Multistage CO2/Steam Turbine and example of component placements

[0316] FIG. 30 is an example of heat exchanger flow patterns 1600

[0317] FIG. 31 is an example of a printed circuit heat exchanger, component build and channels 1602

[0318] FIG.32 is a view of heat exchanger comparison, General Shell, Tube and Plate Heat Exchangers 1604 with Typical 80-90% Effectiveness, Printed Circuit Heat Exchangers (PCHE) 1606 Typical 90-97% Effectiveness and DMLS fabricated state of the art and highly optimized Printed Design Heat Exchanger (PDHE) 1608 Optimized for 99% Effectiveness with Highest Efficiency, Highest Surface Area, Least Use of Materials

[0319] FIG. 33 is a view of an HDLS fabricated Printed Design Heat Exchanger (PDHE) with example of multiple input 16010a or output capability 1610b, 1610c, 1610d, 1610e

[0320] FIG. 34 is a view of an HDLS fabricated Printed Design Heat Exchanger (PDHE) with example usage for molten salt 1612 and CO2 channels 1613

[0321] FIG. 35 is a view of an HDLS fabricated Printed Design Heat Exchanger (PDHE) with example usage for CO2 input 1614aand CO2 output channels 1614b

[0322] FIG. 36 is the view of typical zigzag channels 1616 showing the sharp connections that create turbulence and pressure drops and HDLS fabricated rounded zigzag channels 1618 promoting smooth flow while demonstrating high surface area

[0323] FIG. 37 is a view of examples of HDLS fabricated monolithic PDHE heat exchanger with dual filter ports 1620 and comparison to the many parts and components required for the labor intensive and multi-process construction and fabrication of a PCHE heat exchanger 1622

[0324] FIG. 38 is a view of examples of HDLS fabricated monolithic PDHE heat exchanger with single filter port

[0325] FIG. 39 is a view of the Modular Solid Oxide Fuel Cell

[0326] FIG. 40 is a view of the Modular Solid Oxide Fuel Cell with Flow Supply Duct to Cell from Supply Channel Tunneled Inside the Plate under the Seal for Optimized Channel Sealing Between Plates with male and female connections, compression connections and fuel/oxygen sub-channel examples

[0327] FIG. 41 is a view of the Modular Solid Oxide Fuel Cell with channel layout, seals, connections, channels and cell area

[0328] FIG. 42 is a view of the Modular Solid. Oxide Fuel Cell with seals, connections, channels and cell area

[0329] FIG. 43 is a view of the HDLS fused monolithic component structure dual isolated cooling heat exchanger Advanced Gas Modular Fast Reactor-Generation VSingle Stage utilizing internal and external heat exchanger sleeves from HDLS fused monolithic component builds

[0330] FIG. 44 is a view of the HDLS fused monolithic component structure dual isolated cooling heat exchanger Advanced Gas Modular Fast Reactor-Generation VDual Stage utilizing internal and external heat exchanger sleeves from HDLS fused monolithic component builds

[0331] FIG. 45 is a view of the Toroidal ARBACC

[0332] FIG. 46 is a view of the Toroidal ARBACC w/Turbopump and output manifold

[0333] FIG. 47 is a view of the HDLS fabricated monolithic Aerospike Engine with attached HDLS fabricated. Thrust Cells, HDLS fabricated monolithic Thrust cell with pintle injector

[0334] FIG. 48 is a view of the ARBACC Rocket Engine with Gas Generator Cycle

[0335] FIG. 49 is a view of the HDLS fabricated monolithic Rocket Engine with Gas Generator Cycle single and multi-engine configuration examples

[0336] FIG. 50 is a view of the space vehicle with Toroidal ARBACC mounted

[0337] FIG. 51 is a view of the space vehicle with Toroidal ARBACC mounted

[0338] FIG. 52 is a view of the space vehicle with Toroidal ARBACC mounted

[0339] FIG. 53 is a view of the space vehicle with Toroidal ARBACC mounted

[0340] FIG. 54 is a view of the space vehicle with Toroidal ARBACC mounted

[0341] FIG. 55 is a view of the Linear ARBACC configuration

[0342] FIG. 56 is a view of the Linear ARBACC w/Turbopump and manifold configuration

[0343] FIG. 57 is a view of the space vehicle with Linear ARBACC configuration

[0344] FIG. 58 is a view of the space vehicle with Linear and Toroidal ARBACC configuration

[0345] FIG. 59 is a view of the space vehicle with Linear ARBACC configuration

[0346] FIG. 60 is a view of the space vehicle with Linear ARBACC configuration

[0347] It is to be understood that the following disclosure describes several exemplary embodiments for implementing different features, structures, or functions of the invention. Exemplary embodiments of components, arrangements, and configurations are described. below to simplify the present disclosure, however, these exemplary embodiments are provided merely as examples and are not intended to limit the scope of the invention. Additionally, the present disclosure may repeat reference numerals and/or letters in the various exemplary embodiments and across the Figures provided herein. This repetition is for the purpose of simplicity and clarity and does not in itself dictate a relationship between the various exemplary embodiments and/or configurations discussed in the various Figures.

[0348] Moreover, the formation of a first feature over or on a second feature in the description that follows may include embodiments in which the first and second features are formed in direct contact, and may also include embodiments in which additional features may be formed interposing the first and second features, such that the first and second features may not be in direct contact. Finally, the exemplary embodiments presented below may be combined in any combination of ways, i.e., any element from one exemplary embodiment may be used in any other exemplary embodiment, without departing from the scope of the disclosure.

[0349] Additionally, certain terms are used throughout the following description and claims to refer to particular components. As one skilled in the art will appreciate, various entities may refer to the same component by different names, and as such, the naming convention for the elements described herein is not intended to limit the scope of the invention, unless otherwise specifically defined herein. Further, the naming convention used herein is not intended to distinguish between components that differ in name but not function. Further, in the following discussion and in the claims, the terms including and comprising are used in an open-ended fashion, and thus should be interpreted to mean including, but not limited to. All numerical values in this disclosure may be exact or approximate values unless otherwise specifically stated. Accordingly, various embodiments of the disclosure may deviate from the numbers, values, and ranges disclosed herein without departing from the intended scope.

[0350] Additionally, the system includes methods, processes and applications for fabrication of a supercritical, transcritical and subcritical carbon dioxide energy system including its turbine, compressors, heat exchangers, thermal components and pumping systems with methods of fabrication and manufacturing. Various embodiments of the present invention may include carbon dioxide handling equipment, that may include, for example, a carbon dioxide source or carbon dioxide generator, a pressurizing apparatus or compressor, one or more pressure vessels, various interconnecting piping, valves, one or more vent pipes, or some combination of these items. Various embodiments of the present invention also include enclosures or enclosing walls or structure. Another embodiment will allow power, heating and cooling generation.

[0351] An embodiment presented in FIG. 21 shows a schematic and paths of direction and connection of the present inventions methods and applications of an energy conversion system, which preferably will include renewable energy generated from wind and solar but alternatively can use fossil fuels and nuclear fission or fusion thermal generation for input to the supercritical, transcritical and subcritical carbon dioxide energy conversion system. Thermal storage integration may be used to provide thermal energy to a supercritical, transcritical and subcritical carbon dioxide energy conversion system up to 24 hours a day 7 days a week with energy storage reserves used for peaker power generation.

[0352] Integrating renewable energy systems for input in conjunction with a combined cycle supercritical, transcritical and subcritical carbon dioxide energy conversion system allows for efficient use of supercritical, transcritical and subcritical carbon dioxide energy conversion system and increases the electric conversion efficiency of a combined cycle supercritical, transcritical and subcritical carbon dioxide energy conversion system to approximately 63-69% and above 80% when using recycled thermal waste heat for general heating and cooling applications. A supercritical, transcritical and subcritical carbon dioxide energy conversion system can provide power generation, heating and cooling from a single system utilizing complete energy cycles of available energy. This greatly increases the overall efficiency of energy system, thereby reducing plant capital costs, lowers recurring maintenance costs and total costs of electricity production.

[0353] A supercritical, transcritical and subcritical carbon dioxide energy conversion system generally includes carbon dioxide storage and pump P2 and motor/engine/turbine powered to introduce carbon dioxide into the system at high pressure to establish and maintain adequate carbon dioxide charge and by replacement of carbon dioxide lost to leakage. High pressure piping via Ducts D1-D32, valves and other type of connectors connect the system components to circulate the gas and liquids amongst the various components and loops of the system. The turbine, generator/alternator and compressor is shown inside the dashed area can be interchanged with the various configurations shown for scaling the system up or down.

[0354] A primary heat exchanger HX1 is used for transfer of external generated thermal energy input to inject thermal energy into the carbon dioxide top cycle for input to the primary turbine T1 and generator/alternator 1 and main compressor MC, secondary turbine T2 and generator/alternator 2 and recompressor RC, gas film compressor BC (turbine bearings) and motor/engine/turbine, high temperature recuperator/heat exchanger HX2, low temperature recuperator/heat exchanger HX3, gas precooler/heat exchanger HX4, condenser, transcritical turbine 3 and generator/alternator 3, pump P1 and motor/engine/turbine, secondary compressor SC and motor/engine/turbine, heat exchanger HX5, heat exchanger HX6. Heat exchanger HX7, expansion valve and evaporator.

[0355] An embodiment presented in FIG. 2 shows a schematic of energy conversion system that with conversion of the secondary system using turbine T3 shaft as the input source for the transcritical and subcritical compressor SC and pump P1. This allows the middle and bottom cooling cycle to greatly improve system efficiency by optimizing utilization from a greater percentage of the thermal energy input.

[0356] An embodiment presented in FIG. 3 shows a diagram of a multistage bearing and seal array but is only a reference as other bearing, lubrication and sealing methods that may be used.

[0357] An embodiment presented in FIG. 4 shows a diagram of a multistage bearing and seal array but is only a reference as other bearing, lubrication and sealing methods that may be used.

[0358] An embodiment presented in FIG. 5 shows a diagram of a multistage bearing and seal array but is only a reference as other bearing, lubrication and sealing methods that may be used.

[0359] An embodiment presented in FIG. 6 shows a diagram of a multistage bearing and seal array but is only a reference as other bearing, lubrication and sealing methods that may be used.

[0360] An embodiment presented in FIG. 7 shows a diagram of a gas film titling pad bearing and a gas film thrust bearing but is only a reference of a method that may be used.

[0361] An embodiment presented in FIG. 8 shows a diagram of a single or multistage cooling channel and duct system and components as a reference to the communication of cooling gas between components and the rotor and blade or impeller cooling methods that may be used.

[0362] An embodiment presented in FIG. 9 shows a diagram example for a single or multistage turbine and compressor with bearing and seal array arrangements but is only a reference as other rotor and blade or impeller, bearing, lubrication and sealing methods that may be used.

[0363] An embodiment presented in FIG. 10 shows a diagram of a prior art and the present inventions improvements with massive surface area ratio enhancement.

[0364] An embodiment presented in FIG. 11 shows a diagram of a heat exchanger using a singular input and multiple thermal transfer loops utilizing an optimized surface area ratio.

[0365] An embodiment presented in FIG. 12 shows a diagram of a heat exchanger using a molten salt input design and a steam or CO2 transfer loop utilizing an optimized surface area ratio.

[0366] An embodiment presented in FIG. 13 shows a diagram of a heat exchanger using a steam or CO2 input design and a steam or CO2 transfer loop utilizing an optimized surface area ratio.

[0367] An embodiment presented in FIG. 14 shows a diagram of an electronic arc furnace for melting materials to allow normal or special alloy matrixes.

[0368] An embodiment presented in FIG. 15 shows a diagram of a water based atomization process to process materials melted in the electronic arc furnace for atomized materials for use in the quadratic or squared high density laser sintering (HDLS) process as input materials to allow normal or special alloy matrixes usage.

[0369] The present invention allows single component fabrication of impellers and rotors with blades without using joints, welds and other types of connections. This will allow high pressure gases and/or supercritical fluids and fluids to be used as lubrication within the component build while reducing the number of seals or the size of the seals while still allowing seals to substantially limit leaks. The present invention provides for channels both conformed and nonconformed for the purpose of mixing which may include mixing systems such as vortex generators to create turbulence within the cooling channels, this may be established within the fabrication of the blades, vanes and injectors of both radial and axial turbine components to assist in cooling effectiveness not achievable from prior art. The present invention provides the ability for complex geometries within channels and ducts of the component build that along with scalable methods for fabrication enable previous inaccessible and unavailable complex scalable designs to create optimal design specifications monetizing previous prior art advances into a single fabricated component.

[0370] The present invention with its ability to scale the component builds provides for cooling and lubrication channels and systems design build within a single component build that has no joints, welds or connections thereby offering the highest performance and efficient possible. The present invention will provide for higher turbine temperatures allow for higher turbine efficiencies. The present invention will provide for higher thermal cooling efficiency while reducing thermal stresses to the components to a minimum. The present invention provides turbine manufacturing with greater efficiency and greater power potential through scaling.

[0371] The present invention provides for heat exchangers to fabricated to higher levels of pressure capability while greatly surpassing prior art. For example, prior art typical heat exchanger design provided for 80-90% effectiveness, the newest technique referred to as Printed Circuit Heat Exchangers (PCHE) typically provide 90-96% effectiveness whereas the present invention is capable of fabricating a Printed Design Heat Exchanger (PDHE) with effectiveness as high as 99% without prior art deficiencies that had joints, welds, fusions, material usage limitations due to manufacturing issues and sizing constraints.

[0372] The present invention provides for optimizing the component design for optimal contact surface area for maximizing thermal energy transfer, reduction of parasitic losses and reducing material requirements thereby also reducing the weight and space requirements while maintaining the optimal material characteristics from the chosen material used for the fabricated component.

[0373] The present invention provides for the targeted component to be easily designed and then fabricated with a number of materials for a customized solution for gases and/or supercritical fluids or liquids, clean or fouling and even corrosive on a beneficial and cost effective basis advantage of prior art.

[0374] The present invention provides for fabrication with capabilities of the highest thermal effectiveness, highest temperatures, highest pressures, lowest pressure drops, highest compactness, highest erosion resistance, highest corrosion resistance and longest life advantages over any prior art and is only limited by the material characteristics chosen for fabrication.

[0375] The present invention provides for predetermined estimates for component replacement and repair by selective material choice and material thickness prior to fabrication. The present invention requires no special orders of materials as such provides for lower material costs, short lead times greatly reducing downtime and project delays hence greater cost reductions and lower cost of energy and cost of ownership.

[0376] The present invention provides for inclusion of an electric arc furnace integration into the system processes to provide the ability to create special alloy metallurgy that matches exactly the specific design needs for strength, corrosion resistance, psi tensile strength and temperature requirements within the upper safe limits of special alloy materials for the purpose usage in component builds.

[0377] The present invention provides for a novel advantage especially concerning aerospace, heavy equipment and mining equipment and other mobile based industries with weight to energy issues, weight is an ability the present invention largest advantage in the mobile sector that the prior art isn't and/or can't change in designs for weight saving concerns using honeycomb and other supported void types of volume yet light weight designs that that the present invention can incorporate and fabricate. This allows the present invention to use fabrications with the least amount of weight like a honeycomb void design would allow while retaining very high tensile strength.

[0378] Referring now to the drawings in detail, wherein like numbers are used to indicate like elements throughout, there is illustrated in FIG. 1 is a supercritical, transcritical, subcritical CO2 combined cycle energy system and with supporting hardware. This arrangement allows use of the topping cycle for high temperature thermal energy use for power generation, reuse of the reduced temperature thermal energy for a middle cycle for power generation and a bottom cycle using low temperature thermal energy for cooling generation and processing the recycled thermal energy for water heating thereby utilizing energy to upwards of 90%+ efficiency.

[0379] The Linear Advanced Bearing and Seal (LABS) array system forms a replaceable cartridge for easy maintenance in the field. After removal the cartridge with its tight tolerances can be sent in for repair for inspection to determine failure and provide data for product enhancements. This process will also allow proper servicing in a sealed environment to prevent contamination and further damage.

[0380] The Linear Advanced Bearing and Seal (LABS) array assembly provides a cartridge based system as illustrated in FIG. 24-FIG. 26 that quantifies the quintessential nature of the bearing and seal assembly. This assembly also includes the integration of the cooling gas input channel within the assembly to further reduce sealing requirements simply due to its integration within LABS assembly.

[0381] The Linear Advanced Bearing and Seal (LABS) array assembly of the present invention as illustrated in FIG. 26 demonstrates the assembly can be contained in a clam shell enclosure for added sealing capabilities and higher pressures with additional protection provided for the assembly. As illustrated in FIG. 24-FIG. 26 bearing and seal assembly according to one of more aspects of the present invention through this disclosure is only an example and different configuration and placement of the bearings and seals is possible.

[0382] As illustrated in FIG. 24-FIG. 26 bearing and seal assemblies according to one or more aspects of the present disclosure may be used in conjunction with a turbomachinery with the assembly enclosed and/or connected to a casing and having a balance pressure side, low-pressure gas side a high-pressure gas side. In an exemplary embodiment, the turbomachinery may consist of a high-pressure turbo-compressor. The turbomachinery may also include a rotor shaft configured to extend through the turbomachinery and exit one or both sides of the casing into a housing that may consist of a single or multiple rotor and blades and/or impeller(s). The rotor shaft may use a be journal bearing at each end by employing suitable bearings. In alternative embodiments, the casing and the housing may include the same overall structure, or otherwise the casing and housing may each be enclosed by a separate overall casing structure.

[0383] As illustrated in FIG. 24-FIG. 26, one embodiment a LABS assembly as discussed herein may be utilized effectively on a single sided turbo machine (e.g., machines of the overhang type).

[0384] Relative to the housing, the rotor shaft may be supported via bearings to provide the shaft free rotation and sealed via a series of seals to reduce process gas leakage from the inner area of the turbo machine. In particular, the turbomachinery requires a LABS assembly configured to supported the rotor shaft via bearings to provide the shaft free rotation while reducing unwanted movement and to prevent process gas or liquids from escaping from the turbomachinery inner or outer casing and system housing, thereby entering the atmosphere.

[0385] In an exemplary embodiment, the LABS assembly on the gas exit side may include a high-pressure seal, a high-pressure labyrinth seal, a single seal, a labyrinth seal, a tandem seal including an intermediate labyrinth seal, and a separation labyrinth seal. Each bearing and seal may extend circumferentially around the rotating shaft and be sequentially mounted longitudinally outward relative to the housing. The bearing and seal assembly may be similar to the bearing and seal assembly on the opposite side of the turbomachinery.

[0386] Referring to FIG. 24-FIG. 26, illustrated is an exemplary embodiments of the bearing and sealing assembly. As illustrated, the high-pressure seal may be situated on the high-pressure process gas or liquids of the turbo machine, and radially coupled to an outer edge of the interior of the turbo machine casing. The high-pressure seal typically is utilized reduce the pressure of any process gas or liquids escaping the turbomachinery casing to a lower inner-stage pressure. This may be done to create a delta pressure that serves to balance axial theist forces generated inside the turbo machine. In one embodiment, a portion of this reduced-pressure process gas may be collected via a duct and re-injected at process gas or liquid side to be re-pressurized by the turbomachinery or external input. The high-pressure labyrinth seal, located coaxially adjacent to the high-pressure seal, may be configured to separate any escaping process gas from the high-pressure seal.

[0387] Traditionally, a labyrinth-type seal has been employed coaxially adjacent the high-pressure labyrinth seal and the potential for a secondary labyrinth seal array can be configured to further reduce pressure of any process gas escaping the high-pressure labyrinth seal to a level that a tandem seal can physically accept. However, in high-pressure, low-flow applications, using the traditional labyrinth-type blow-down seal may cause up to 10-15% efficiency losses in power and total process flow of the turbomachinery. According to the present disclosure, to decrease these efficiency losses, the pressure reduction process may instead be handled by a single pressure reduction seal. It has been shown that using a single pressure reduction seal may reduce total efficiency loss from 10-15% to about 2-5%, and even less than about 1% in some applications.

[0388] Therefore, an exemplary embodiment of the present disclosure may include the combination of a single pressure reduction seal and a tandem seal; thus taking advantage of the current tandem experience while benefiting from the bearing pressurizations yet still efficiently reducing pressure efficiency losses. This combination is not necessarily configured but can be seen as a triple or quadruple seal system.

[0389] During typical operation of a dry gas face seal, a portion of the high-pressure process gas is cleaned and introduced to the gas seal to help maintain a high-pressure sealing effect, and also to prevent potential contamination of the seals. Prior to cleaning, this process gas may contain foreign matter such as dirt, iron filings, and other solid particles which can contaminate the seals. Therefore, cleaned seal gas, including filtered process gas or an inert gas from an external source, may be injected at each gas seal at a predetermined pressure higher than the pressures in the preceding inner-areas of the housing in order to block process gas leakage. In operation, the cleaned gas may be pressurized by a small reciprocating compressor, or may utilize pressurized gas from an alternative turbo machine application.

[0390] Likewise, externally pressured gas may be injected at the tandem in a similar fashion. In particular, cleaned seal gas may be injected via a duct between the labyrinth seal and the primary gas seal at a pressure in excess of the pressure incident in reduced pressure at the primary vent duct. In an exemplary embodiment, the majority of the seal gas injected via a duct may flow across the labyrinth seal and into a seal duct via reduced pressure at the primary vent duct. However, a small portion of the seal gas may flow across the primary seal as leakage loss, which may either be collected or discharged to flare via the tandem primary vent duct.

[0391] The foregoing has outlined features of several embodiments so that those skilled in the art may better understand the detailed description that follows. Those skilled in the art should appreciate that they may readily use the present disclosure as a basis for designing or modifying other processes and structures for carrying out the same purposes and/or achieving the same advantages of the embodiments introduced herein. Those skilled in the art should also realize that such equivalent constructions do not depart from the spirit and scope of the present disclosure, and that they may make various changes, substitutions and alterations herein without departing from the spirit and scope of the present disclosure.

[0392] The Linear Advanced Bearing and Seal (LABS) array assembly of the present invention through utilization of the cooling gas channel input. FIG. 29 illustrates an axial half cross-section example through the relevant components of prior art conventional turbomachinery. A separate flow metering aperture with an accurately sized opening can be used to control the flow and pressure that are delivered to the lubricating gas and cooling gas plenum, as desired given the necessary design parameters.

[0393] The inner chambers of the bearing and seal gallery is supplied with lubricating gas and/or fluids via supply ducts and gas and/or liquids are removed via a drain scavenge duct. An outer most chamber of the gallery is ventilated with low pressure compressed cooling gas and sealed with seals. Compressed cooling gas is delivered to the outer chamber of the bearing and seal gallery is provided through a low pressure gas supply duct (not shown) communicating between the low pressure stage compressor and/or regulator (not shown) and the bearing and seal gallery chamber.

[0394] Referring to FIG. 29, a plurality of turbine blades is mounted to a rotor or an impeller in single stage high work high pressure turbine of the turbomachinery. The gas or liquid has a pressure (typically seen during operation of the engine) downstream of the turbine which is lower in conventional turbomachinery because of the effect of the high work high pressure turbine on the gas or liquid there through the turbomachinery.

[0395] Referring to FIG. 29, a plurality of turbine blades is mounted to a multiple of rotors or impellers in multistage stage high work high pressure turbine of the turbomachinery. The gas or liquid has a pressure (typically seen during operation of the engine) downstream of the turbine which is lower in conventional turbomachinery because of the effect of the high work high pressure turbine on the gas or liquid there through the turbomachinery.

[0396] Referring to FIG. 29, alternatively in a turbomachinery compressor a plurality of turbine blades is mounted to a rotor or an impeller in single stage high work high pressure compressor of the turbomachinery. The gas or liquid has a pressure (typically seen during operation of the engine) downstream of the compressor which is higher in conventional turbomachinery because of the effect of the high work high pressure turbine on the gas or liquid there through the turbomachinery.

[0397] Referring to FIG. 29, alternatively in a turbomachinery compressor a plurality of turbine blades is mounted to a multiple of rotors or impellers in a multiple single stage high work high pressure compressor of the turbomachinery. The gas or liquid has a pressure (typically seen during operation of the turbomachinery) downstream of the compressor which is higher in conventional turbomachinery because of the effect of the high work high pressure turbine on the gas or liquid there through the turbomachinery.

[0398] The turbine has a rotation about the turbomachinery axis is shown in FIG. 28 in a circular motion about the periphery of a blade. Each component including the rotor of impeller has its blades quadratic or squared HDLS printed with the chosen material as a single component free of welds, fusions or mechanical connections. The blade or air foil whether axial turbine with a rotor and blades or radial with an impeller communicates cooling gas flow through its ducts that each blade or airfoil include a cooling duct communicating at one end with the cooling duct transfer channels cooling plenum within connecting components that also connect to the shaft for rotation and at another end with the blade interior, as described below. Relatively cool gas is provided via the compressor and supplied to cooling gas plenum through conventional duct means. Compressed cooling gas may also be delivered to the gas cooling chambers of the bearing and seal gallery through a pressure gas supply duct from the LABS bearing and gear assembly which also provides lubrication to the bearings and cooling to the gallery, as described in more detail below.

[0399] The rotor or impeller that includes blade air foils that extend radially from the root which connects to the shaft and include cooling gas channels communicating between the cooling gas inlet duct and the cooling gas channel of the turbomachinery engine as shown in FIG. 8 and exiting through intermediate (lower radius) ducts of the components along the shaft. Each cooling gas duct through the rotor and impeller extending into the airfoil and includes connection to the gas cooling ducts between components having cooling gas communicated from the gas cooling duct input.

[0400] A high work single or multiple stage turbomachinery experiences a relatively large pressure drop across the turbine because of the amount of work extracted from the flow. The resulting pressure of the gas path downstream of the turbine is therefore markedly reduced compared to the turbomachinery input pressure. Due to the implied high pressure characteristic of a high work single or multiple stage high pressure turbomachinery creates a pressure drop on the output of duct of the turbomachinery. This allows the cooling gas plenum, channels and ducts to provide sufficient intake flow of cooling gas flow to cool the turbine blades or airfoils and vanes or nozzles without requiring the mechanical complexity of the prior art.

[0401] As shown in FIG. 26, cooling gas enters through a duct in the LABS bearing and seals gallery is then exhausted into the cooling gas plenum, channels and ducts through components in FIG. 28. A separate flow metering aperture and/or temperature sensor with an accurately sized opening can be used to control the flow, pressure and temperature that are delivered to and through the cooling gas plenum, as desired given the necessary design parameters. Optimization of the ducting shape, aperture size, vortex generators and orientation depends on the turbine radius, speed of rotation and the parameters of the cooling gas plenum, as well will be apparent by one skilled in the art in regards of disclosure for the present invention.

[0402] With regards to past prior art, many attempts have been employed to enhance rocket engines, mostly this was through the delivery of energy and choice of energy for rocket engines. The energy potential of the propellant and oxidizers have been gradually increased to reduce their weight and the operating pressures have been increased to enhance total thrust.

[0403] A liquid fuel rocket engine construction includes the usual main combustion chamber having a nozzle discharge. One or more gas generator is used to generate shaft energy input to the turbine section of the turbopump and to discharge exhausted combustion gases and/or supercritical fluids external to primary turbopump exhaust. The primary turbopump typically drives at least one separate fuel component pumps and one separate oxidizer component pumps.

[0404] A liquid fuel rocket engine construction includes a typical main combustion chamber having a nozzle discharge.

[0405] The preferred method of the present invention provides for a fabrication and construction method and application of rocket engines and, in particular, to new and useful liquid fuel rocket engines having at least one first stage and provides the ability for additional stages of a space launch system.

[0406] The preferred method of the present invention provides for injection of the regenerative cooling agent consisting of fuel or oxidizer according to the invention which provides for the highest ratio of surface area to flow volume of the injection cooling agent in addition to the highest ratio of contact area for optimization of regenerative cooling efficiency. Due to the uniform distribution of the regenerative cooling agent, a rapid and thus advantageous cooling and combustion thereby thrust is achieved. Thus, it is then possible to obtain either a reduction of the length or weight or the potential of both for the combustion chamber weight and cooling and better combustion with extended burn time capabilities and engine reusability.

[0407] The preferred method of the present invention using quadratic or squared fabrication system provides for optimized regenerative cooling within in the context of rocket engine design. The preferred method provides for a method and process of fabrication to allow maximum performance, maximum chamber pressure and optimal cooling.

[0408] The preferred method of the present invention provides for optimization through which the regenerative cooling agent is communicated through channels which are formed via the inner and outer jacket or skins of the engine.

[0409] The preferred method of the present invention utilizes the quadratic or squared fabrication system through with the inner jacket having optimal thickness for thermal energy transfer via conduction and the thickness for sides of each duct optimally built for the required strength and tensile strength and an outer jacket all optimized with a safety margin for minimal pressure drop, maximized pressure and temperature and optimized weight reduction. An example of the duct and channels within a wall section is shown.

[0410] The preferred method of the present invention fabrication of the complete rocket engine and pintle injector component along with the combustion chamber, exhaust nozzle bell and regenerative cooling channels within a single component build with no seams, joints or connections while able to scale to hundreds of thousands foot pounds of thrust compared to the limited build, non-optimal chamber pressure, thrust limitation due to cooling constraints imposed by prior art fabrication methods.

[0411] The preferred method of the present invention greatly exceeds the ability of prior art methods and beyond the capabilities and potential of prior art fabrication thereof by enabling the present invention novel fabrication methods for novel applications with higher pressure, temperature, cooling capacity and scaling than was previously available from prior art.

[0412] The preferred method of the present invention provides for some or all the fuel and/or oxidizer is communicated through ducts, channels, or in a jacket around the combustion chamber and/or exhaust nozzle to cool the engine. This is effective because the fuel and/or oxidizer are effective coolants. The heated fuel and/or oxidizer is then communicated into a special gas generator to power the turbopump or directly injected into the main combustion chamber

[0413] In accordance with the method of the invention, a liquid fuel rocket engine is operated with at least one pre-combustion chamber for burning fuel components arranged to discharge the combustion gases and/or supercritical fluids through an auxiliary turbine. The turbine is connected to drive the fuel component pumps, and it is arranged to discharge the gases and/or supercritical fluids directly into the main combustion chamber. The fuel components are directed into the pre-combustion chamber and the pre-combustion chamber is operated with an excess of either oxidizer or fuel component so that there will be a completed burning of the fuel combustion products after they are discharged through the turbine external from the assembly through an exhaust duct.

[0414] A further object of the invention is to provide a liquid-fuel rocket engine which is simple in design, rugged in construction and economical to manufacture and reusable.

[0415] This invention relates in general to the construction of combustion chambers and in particular, to a new and useful method and construction of a combustion chamber and exhaust nozzle particularly of a rocket engine which includes a collecting or distributing channel for interconnecting a plurality of channels or ducts such as for cooling with additional fuel or oxidizer distribution purposes.

[0416] Combustion chambers and exhaust thrust nozzles for rocket engines which are propelled by liquid propellants are typically subject to extremely high thermal stresses in addition to very high compressive stresses. In order to control the great amounts of thermal energy which is generated by the combustion chambers and the thrust nozzles are frequently made of a special alloy material since such materials have thermal conductivity yet are capable of handling high thermal stresses which is facilitated by removal of the thermal energy by the regenerative cooling system provided.

[0417] A further object of the invention is to provide a combustion chamber construction with an annular collecting or feeding duct which is simple in design, rugged in construction, and economical to manufacture.

[0418] A premise on which the present invention includes a rocket engine method called Air-augmented aerospike or ducted aerospike, which utilizes additional mass air flow via an inlet that collects external mass flow and passes the flow through ducts, whereby the use of atmospheric air reduces oxidizer requirements which then combines with the propellant gases and/or supercritical fluids to increase the specific impulse of the propellant. While ducted rockets have been investigated they previously posed difficulties and complex to design efficiently. The present invention, an Air-augmented aerospike rocket engine amalgamated with a scramjet engine, improves the delivered energy density of rocket engines, with less complexity of prior art.

[0419] A standard scramjet (supersonic combusting ramjet) is a valiant of a ramjet air breathing jet engine that requires high vehicle speed in which injection and then combustion takes place in supersonic airflow. Typically, when operational the airflow in a scramjet is supersonic throughout the entire engine. This allows the scramjet to operate efficiently at extremely high speeds which can go up to Mach 25.

[0420] Scramjet engines are a unique type of jet engine, and rely on the combustion of fuel and atmospheric air as an oxidizer to produce thrust. Similar to conventional jet engines, scramjet-powered aircraft carry the fuel on board, and obtain the oxidizer by the induction of atmospheric oxygen (as compared to conventional rockets, which carry both fuel and an oxidizing agent). This requirement limits scramjets to suborbital atmospheric propulsion, where the oxygen content of the air is sufficient to maintain combustion.

[0421] The scramjet is composed of four basic components: a converging inlet, where incoming air is compressed; an injector, a combustor, where gaseous or atomized liquid fuel is burned with atmospheric oxygen to produce heat; flame-holder, and a diverging nozzle, where the heated air is accelerated to produce thrust. An injector with preference for a pintle injector is designed for use with a scramjet engine and provides high combustion efficiency and pressure recovery for length-to-diameter (L/D) ratios tunable over a wide range of operating conditions.

[0422] The present invention method comprises use of a flame holder that typically involves an injector which will provide excellent performance over a wide range of conditions of L/D ratios. A flame holder is a component of a jet engine designed to help maintain continual combustion.

[0423] Generally, all commercial continuous-combustion jet engines require a flame holder. A flame holder creates a low-speed eddy in the engine to prevent a flameout scenario. The design of the flame holder is an issue of balance between a stable eddy and drag, this becomes more critical as flow speeds increase.

[0424] Typical effective designs are the H and V flame holders. One method is the H-gutter flame holder, which is shaped like a letter H with a curve facing and opposing the flow of air. The most effective and most widely used method however, is the V-gutter flame-holder, which is shaped like a V with the point in the direction facing the flow of air. Many reviews have shown that adding a small amount of base bleed from the V in the V shaped-gutter helps reduce drag without reducing effectiveness.

[0425] An injector is provided and flame-holders are provided to enhance and maintain combustion efficiency. Different type of flame-holders may be selected to achieve particular targeted results. For example, toroidal ring flame-holders provide a nearly symmetrical spreading of the fuel-air mixture, toroidal ring of injector ports and flame-holders will increase secondary flows in regions of the combustor dome between ports, and perpendicular toroidal ring flame-holders will increase secondary flows outboard of the ports as well as in the center of the engine.

[0426] Additional features may include the use of electronic igniters and/or pilot shrouds for lowering both the lean limits, fuel detonations and pressure oscillations and the use of flow dividers for raising the rich operating limits. It is therefore a general object of the present invention to provide an injector dump combustor which will have good performance over a wide range of conditions of L/D ratios.

[0427] The method of the present invention comprises an air-augmented aerospike rocket engine that is mounted in the center of a long duct. As the rocket engine vehicle moves through the atmosphere the air enters the inlet of the ducts, where it is compressed via a ram effect. It transverses down the throat of the duct it is further compressed, it is at this stage, usage of the scramjet injectors enables scramjet functionality with its higher Specific Impulse (Isp), this supersonic flow can then mixed with the fuel-rich exhaust from the aerospike rocket engine. These advantages provide supersonic flow for the establishment of high speed thermal expanded flow to encourage of pressure fields to originate containment in creating enhanced virtual bell nozzle effect to effectuate altitude-compensating for a net thrust benefit.

[0428] Another advantage is diverting part of the mass flow through the aerospike plug base to enable a pressure bleed which will reduce drag on the engine and its attached vehicle. In this fashion a smaller rocket engine in conjunction with the scramjet component to accelerate a much larger working mass flow leading to significantly higher thrust within the atmosphere. When leaving the atmosphere, primary propulsion would be transferred and maintained solely from the rocket engine thrust.

[0429] The method of the present invention comprises an air-augmented aerospike rocket engine incorporating scramjet technology providing use of the similar high-energy propellant and cooling schemes and techniques to maintain sustained operation and uses. The inclusion of the scramjet engine will allow a vehicle voyage through the atmosphere to benefit from the atmospheric air flow to reach Mach 25 while conserving propellant, meanwhile transferring primary propulsion to the rocket engine for accelerating to Mach 25 and above with the air-augmented aerospike rocket engine with enough speed necessary for leaving earth orbit.

[0430] The method of the present invention may be comprised with the inclusion of a ballute. This device is an amalgamation of a balloon and parachute and its function is a parachute-like braking device optimized for use at high altitudes and supersonic velocities. A ballute typically is an inflated structure intended to ensure flow separation which stabilizes the intended target as it decelerates through different flow regimes slowing from supersonic to subsonic speeds.

[0431] Space transportation architecture covers a wide range of launch concepts all proposed. as options for a space launch vehicle. The commercial market demands that any option and meet an additional set of challenging requirements be met to fulfill current and future commercial space lift needs.

[0432] One of the most demanding of the system design requirements for a launch system that is capable of performing space missions is to limit the development time and cost for the complete system. Typically, there is a long duration of stages between initial investment, design, development, prototyping and finally the operational system with associated revenues for a return on investment.

[0433] In addition to targeting low development costs, any system that is to be developed must also be extremely reliable and enable safety of manned human flight missions. Development would include an extremely reliable design that precludes, within practicality, catastrophic system failures. Examples of design features that help to prevent such catastrophic failures include: full engine shutdown from liftoff, full vehicle abort capability throughout the entire mission, robust design and standard operating margins, and integrated vehicle health monitor and control system. The integrated vehicle health monitor and control system within an ideal concept should also include a constant evolving control system that is relatively tolerant of many critical failures or malfunctions of key flight systems.

[0434] The invention will be better understood, and further objects, novel features, and advantages thereof will become more apparent from the following description of the preferred embodiments, taken in conjunction with the accompanying drawings.

[0435] An object of the invention is to provide a space launch vehicle maneuvering thrusters would be placed on each side of the vehicle and would be used by the reaction control system having an efficient fuel usage for payload to orbit.

[0436] Another object of the invention is to provide a space launch vehicle having an air augmented aerospike rocket engine coupled to a space craft for efficient delivery of a payload into orbit.

[0437] Yet another object of the invention is to provide a space launch vehicle having an external tank coupled to a space craft for efficient delivery of a payload into orbit.

[0438] Still another object of the invention is to provide a launch vehicle having an external tank and air augmented aerospike rocket engines coupled to a space craft for efficient delivery of a payload into orbit and to provide flight reentry of an orbiter.

[0439] A further object of the invention is to provide a launch vehicle having an external tank and multiple air augmented aerospike rocket engines coupled to a space craft for efficient delivery of a payload into orbit.

[0440] The present invention has three primary classes of space launch vehicles characterized by one or more rocket stages having air augmented aerospike rocket engines for propulsion, a space craft with flight control and/or stabilization surfaces, and characterized by an attached propellant booster stage for the efficient delivery of a payload into space. The rocket stage(s) can be in the preferred forms an orbiter, standard rocket, a booster, multiple boosters or any combinations thereof. The orbiter and/or standard rocket that preferably includes a payload bay and/or payload module. The propellant feeding stage in the preferred forms can be an external tank (ET) or a core stage the latter of which preferably includes air augmented aerospike rocket engines and a payload bay. The use of these components provides a variety of launch systems having a wide variety of capabilities.

[0441] These and other advantages will become more apparent from the following detailed description of the preferred embodiment.

[0442] In accordance with the invention, an altitude-compensating, Rocket-Based Air-augmented Combined Cycle propulsion system rocket engine assembly is provided for horizontal and vertically launched vehicles which offers substantial advantages over prior art engine assemblies such as standard aerospike and bell nozzles. Space craft performance is improved 12-20% over prior art engines using conventional engine and nozzle arrangement results in a light weight, high performance space launch system.

[0443] In accordance with a first aspect of the invention, there is provided a rocket engine housing duct including at an inlet, injector, combustion area, flame-holder, outlet, an injector, least two combustion chambers each including an outlet end defining a throat exhaust area; means for supplying a propellant to said at least two combustion chambers including throttling injector means, associated with each of said at least two combustion chambers and located upstream of said throat area, for receiving said propellant and for injecting said propellant into the associated combustion chamber; and control means for selectively controlling the throttling injector means for each of said at least two combustion chambers so that said at least two chambers enabling thrust vectoring capable propulsion.

[0444] Preferably, the rocket engine assembly further comprises expansion means located downstream of said throat exhaust area for providing expansion of combustion gases and/or supercritical fluids produced by said at least two combustion chambers so as to increase the net propulsion. In one preferred embodiment, the expansion means comprises an expansion nozzle. In an alternative preferred embodiment, the expansion means comprises an aerospike body. In one preferred implementation, the expansion means comprises a fixed position exhaust nozzle but, as described below, a movable nozzle can also be employed.

[0445] In one preferred embodiment, the at least two chambers are disposed in side-by-side relation. In an advantageous implementation, multiple combustion chambers arranged in a cluster in side-by-side relation.

[0446] The injector means preferably comprises a coaxial pintle injector disposed coaxial with the associated combustion chamber. Advantageously, the injector means comprises at least one movable element for providing flow regulation of the propellant.

[0447] According to a second aspect of the invention, there is provided a rocket engine assembly for a space launched vehicle with maneuvering thrusters would be placed on each side of the vehicle and would be used by the reaction control system, comprising a rocket engine housing duct with an inlet, injector, combustion area, flame-holder, outlet, including at least two combustion chamber disposed in side-by-side relation and each including an outlet; means defining a throat exhaust area at the outlet of each the at least two combustion chambers; propellant supply means for separately supplying an oxidizer and fuel to said combustion chambers; throttling injector means, associated with each of said combustion chambers located downstream of said throat exhaust area, for receiving said oxidizer and fuel and for injecting said oxidizer and fuel into the associated combustion chamber; and control means for selectively controlling said throttling injector means of each of said combustion chambers to enabling thrust vectoring capable propulsion.

[0448] According to a third aspect of the invention, there is provided a rocket engine assembly for a space launched vehicle with a maneuvering thrusters would be placed on each side of the vehicle and would be used by the reaction control system, comprising a rocket engine housing duct with an inlet, injector, combustion area, flame-holder, outlet, including at least two combustion chamber disposed in side-by-side relation and each including an outlet; means defining a throat exhaust area at the outlet of each the at least two combustion chambers; propellant supply means for separately supplying an oxidizer and fuel to said combustion chambers; throttling injector means, associated with each of said combustion chambers located downstream of said throat exhaust area, for receiving said oxidizer and fuel and for injecting said oxidizer and fuel into the associated combustion chamber; and control means for selectively controlling said throttling injector means of each of said combustion chambers to enabling thrust vectoring capable propulsion.

[0449] As indicated above, the assembly preferably comprises expansion means located downstream of said throat exhaust area for providing expansion of combustion gases and/or supercritical fluids produced by said at least two combustion chambers. As also was described previously, expansion means comprises an expansion nozzle or an aerospike body, and can comprise a fixed position exhaust nozzle.

[0450] In accordance with yet another aspect of the invention, there is provided a rocket engine assembly for a space launched rocket vehicle, comprising a rocket engine housing duct with an inlet, injector, combustion area, flame-holder, outlet, including at least two combustion chambers each including an outlet end defining a throat exhaust area; propellant supply means for supplying a oxidizer and fuel to said at least two combustion chambers, said propellant supply means including injector means, associated with each of said at least two combustion chambers and located upstream of said throat exhaust area, for receiving said propellant and for injecting said propellant into the associated combustion chamber; regulator means for regulating the flow rate of said oxidizer and fuel to each of said at least two combustion chambers; control means for selectively controlling said regulator means for each of said at least two combustion chambers so that said at least two chambers enabling thrust vectoring capable propulsion; and expansion means, such as an expansion body or an aerospike body, located downstream of said throat exhaust area for providing expansion of combustion gases and/or supercritical fluids produced by said at least two combustion chambers so as to increase said net propulsion.

[0451] In one preferred implementation, regulator means comprises a control valve located in a propellant supply pipe upstream of said injector means. In another preferred implementation, regulator means comprises a control valve located in a oxidizer supply pipe upstream of said injector means. In another preferred implementation, the regulator means comprises a movable element of said injector means which is controlled by said control means.

[0452] Further features and advantages of the present invention will be set forth in, or apparent from, the detailed description of preferred embodiments thereof which follows.

[0453] An embodiment of the invention is described with reference to the figures using reference designations as shown in the figures. Referring to FIG. 1, a first preferred embodiment a single rocket engine without a launch vehicle. The preferred embodiment comprises a rocket engine housing duct with an inlet, injector, combustion area, flame-holder, outlet, including at least two combustion chambers each including an outlet end defining a throat exhaust area; propellant supply means for supplying a oxidizer and fuel to said at least two combustion chambers, said propellant supply means including injector means, associated with each of said at least two combustion chambers and located upstream of said throat exhaust area, for receiving said propellant and for injecting said propellant into the associated combustion chamber; regulator means for regulating the flow rate of said oxidizer and fuel to each of said at least two combustion chambers; control means for selectively controlling said regulator means for each of said at least two combustion chambers so that said at least two chambers enabling thrust vectoring capable propulsion; and expansion means, such as an expansion body or an aerospike body, located downstream of said throat exhaust area for providing expansion of combustion gases and/or supercritical fluids produced by said at least two combustion chambers so as to increase said net propulsion. For each application, it is anticipated that minor modifications will be required to tailor the rocket engine to specific applications.

[0454] An embodiment of the invention is described with reference to the figures using reference designations as shown in the figures. Referring to FIG. 2, a first preferred embodiment also known as Prometheus Mark 1 consists of a launch vehicle, includes a two engine orbiter having Axisymmetric Rocket-Based Air-augmented Combined Cycle propulsion system rocket engines (ARBACC). The orbiter has internal stored propellant. The twin engine orbiter includes a canard, left orbiter flight control surface and a right orbiter flight control surface, an orbiter shell, orbiter body flaps including orbiter left orbiter body flaps and right orbiter body flaps. For each application, it is anticipated that minor modifications will be required to tailor the launch vehicle to specific applications.

[0455] An embodiment of the invention is described with reference to the figures using reference designations as shown in the figures. Referring to FIG. 3, a first preferred embodiment also known as Cronus Mark 1 consists of a launch vehicle, includes a center body with an Axisymmetric Rocket-Based Air-augmented Combined Cycle propulsion system rocket engine (ARBACC). The orbiter has internal propellant tanks. The liquid rocket boosters have internal propellant tanks.

[0456] The twin engine orbiter includes a left orbiter flight control surface and a right orbiter flight control surface, an orbiter shell, orbiter body flaps including orbiter left orbiter body flaps a and right orbiter body flaps. For each application, it is anticipated that minor modifications will be required to tailor the launch vehicle to specific applications.

[0457] An embodiment of the invention is described with reference to the figures using reference designations as shown in the figures. Referring to FIG. 4, a first preferred embodiment also known as Proteus Mark 1 consists of a Fly-back Air-augmented Liquid Booster (FALB), includes an Axisymmetric Rocket-Based Air-augmented Combined Cycle propulsion system rocket engine (ARBACC). The FALB has internal stored propellant. Maneuvering thrusters would be placed on each side of the vehicle and would be used by the reaction control system. The FLAB is an Unmanned Aerial Vehicle (UAV), which is a craft with no pilot on hoard. An FLAB can be remote controlled aircraft by a pilot at a ground control station or can fly autonomously based on pre-programmed flight plans with a dynamic automation system. The FALB includes a canard, left flight control surface and a right flight control surface, an orbiter shell, body flaps including left body flaps and right body flaps. For each application, it is anticipated that minor modifications will be required to tailor the launch vehicle to specific applications.

[0458] An embodiment of the invention is described with reference to the figures using reference designations as shown in the figures. Referring to FIG. 5, a first preferred embodiment also known as Helios Mark 1 consists of a shuttle like launch system which includes an Axisymmetric Rocket-Based Air-augmented Combined Cycle propulsion system rocket engine (ARBACC). This embodiment has internal and external stored propellant. Maneuvering thrusters would be placed on each side of the vehicle and would be used by the reaction control system. This embodiment would include the use of FLAB and/or Solid Rocket Boosters (SBR) and/or Liquid Rocket Boosters (LBR) and an External Tank (ET). This embodiment consists of a left flight control surface and a right flight control surface, an orbiter shell, body flaps including left body flaps and right body flaps. For each application, it is anticipated that minor modifications will be required to tailor the launch vehicle to specific applications.

[0459] An embodiment of the invention is described with reference to the figures using reference designations as shown in the figures. Referring to FIG. 6, a first preferred embodiment also known as Aurora Mark 1 consists of a launch vehicle, includes a quad engine orbiter having Axisymmetric Rocket-Based Air-augmented Combined Cycle propulsion system rocket engines (ARBACC). The orbiter has internal stored propellant. The quad engine orbiter includes a primary wing and lift body, left orbiter flight control surface and a right orbiter flight control surface, an orbiter shell, orbiter body flaps including orbiter left orbiter body flaps and right orbiter body flaps. For each application, it is anticipated that minor modifications will be required to tailor the launch vehicle to specific applications.

[0460] An embodiment of the invention is described with reference to the figures using reference designations as shown in the figures. Referring to FIG. 7, a first preferred embodiment a single rocket engine without a launch vehicle. The preferred embodiment comprises a rocket engine housing duct with an inlet, injector, combustion area, flame-holder, outlet, including at least two combustion chambers each including an outlet end defining a throat exhaust area; propellant supply means for supplying a oxidizer and fuel to said at least two combustion chambers, said propellant supply means including injector means, associated with each of said at least two combustion chambers and located upstream of said throat exhaust area, for receiving said propellant and for injecting said propellant into the associated combustion chamber; regulator means for regulating the flow rate of said oxidizer and fuel to each of said at least two combustion chambers; control means for selectively controlling said regulator means for each of said at least two combustion chambers so that said at least two chambers enabling thrust vectoring capable propulsion; and expansion means, such as an expansion body or an aerospike body, located downstream of said throat exhaust area for providing expansion of combustion gases and/or supercritical fluids produced by said at least two combustion chambers so as to increase said net propulsion. For each application, it is anticipated that minor modifications will be required to tailor the rocket engine to specific applications.

[0461] An embodiment of the invention is described with reference to the figures using reference designations as shown in the figures. Referring to FIG. 8, a first preferred embodiment also known as Perseus Mark 1 consists of a launch vehicle, includes a two engine orbiter having Axisymmetric Rocket-Based Air-augmented Combined Cycle propulsion system rocket engines (ARBACC). The orbiter has internal stored propellant. The twin engine orbiter includes a canard, left orbiter flight control surface and a right orbiter flight control surface, an orbiter shell, orbiter body flaps including orbiter left orbiter body flaps and right orbiter body flaps. For each application, it is anticipated that minor modifications will be required to tailor the launch vehicle to specific applications.

[0462] An embodiment of the invention is described with reference to the figures using reference designations as shown in the figures. Referring to FIG. 9, a first preferred embodiment also known as Perseus Mark 2 consists of a launch vehicle, includes a two engine orbiter having Axisymmetric Rocket-Based Air-augmented Combined Cycle propulsion system rocket engines (ARBACC). The orbiter has internal stored propellant. The twin engine orbiter includes a canard, left orbiter flight control surface and a right orbiter flight control surface, an orbiter shell, orbiter body flaps including orbiter left orbiter body flaps and right orbiter body flaps. For each application, it is anticipated that minor modifications will be required to tailor the launch vehicle to specific applications.

[0463] An embodiment of the invention is described with reference to the figures using reference designations as shown in the figures. Referring to FIG. 10 and FIG. 11, a first preferred embodiment also known as Hermes Mark 1 also known as the StarCruiser concept consists of a two-stage, vertical or horizontal takeoff, horizontal landing configuration with a large unmanned booster and a manned stage designed for up to 150 passengers and 5 crew members. The fully reusable system is accelerated by a multiple engine orbiter having Axisymmetric Rocket-Based Air-augmented Combined. Cycle propulsion system rocket engines (ARBACC). The orbiter has internal stored propellant. The ARBACC engine orbiter includes a canard, left orbiter flight control surface and a right orbiter flight control surface, an orbiter shell, orbiter body flaps including orbiter left orbiter body flaps and right orbiter body flaps.

[0464] After primary rocket engine cut-off the passenger stage will enter a high-speed powered/gliding flight phase and shall be capable of traveling long intercontinental distances within an extremely short time. Altitudes of approximately 200 kilometers and Mach numbers beyond 24 are projected, depending on the mission and the associated flight path flown. For each application, it is anticipated that minor modifications will be required to tailor the launch vehicle to specific applications.

[0465] Although specific shapes and geometries have been illustrated in the drawings, it is also to be understood that the throat exhaust cross sections can be of various different shapes and sizes, and can be arranged in various different geometric locations with respect to each other. However, in each case, the throat exhaust section should be positioned so as to communicate combustion chamber gases and/or supercritical fluids to a single downstream expansion nozzle or body so as to create hypersonic expansion and increased thrust as described above.

[0466] Although the invention has been described above in connection with preferred embodiments thereof, it will be understood by those skilled in the art that variations and modifications can be effected in these preferred embodiments without departing from the scope and spirit of the invention.

[0467] The various features of novelty which characterize the invention are pointed out with particularity in the claims annexed to and forming a part of this specification. For a better understanding of the invention, its operating advantages and specific objects attained by its use, reference should be had to the accompanying drawings and descriptive matter in which there is illustrated and described a preferred embodiment of the invention.

[0468] Although the above description relates to a specific preferred embodiment as presently contemplated by the inventor, it will be understood that the invention in its broad aspect includes mechanical and functional equivalents of the elements described herein.

[0469] Although various representative embodiments of this invention have been described above with a certain degree of particularity, those skilled in the art could make numerous alterations to the disclosed embodiments without departing from the spirit or scope of the inventive subject matter set forth in the specification and claims. Joinder references (e.g. attached, adhered, joined) are to be construed broadly and may include intermediate members between a connection of elements and relative movement between elements. As such, joinder references do not necessarily infer that two elements are directly connected and in fixed relation to each other. Moreover, network connection references are to be construed broadly and may include intermediate members or devices between network connections of elements. As such, network connection references do not necessarily infer that two elements are in direct communication with each other. In some instances, in methodologies directly or indirectly set forth herein, various steps and operations are described in one possible order of operation, but those skilled in the art will recognize that steps and operations may be rearranged, replaced or eliminated without necessarily departing from the spirit and scope of the present invention. It is intended that all matter contained in the above description or shown in the accompanying drawings shall be interpreted as illustrative only and not limiting. Changes in detail or structure may be made without departing from the spirit of the invention as defined in the appended claims.

[0470] Although the present invention has been described with reference to the embodiments outlined above, various alternatives, modifications, variations, improvements and/or substantial equivalents, whether known or that are or may be presently foreseen, may become apparent to those having at least ordinary skill in the art. Listing the steps of a method in a certain order does not constitute any limitation on the order of the steps of the method. Accordingly, the embodiments of the invention set forth above are intended to be illustrative, not limiting. Persons skilled in the art will recognize that changes may be made in form and detail without departing from the spirit and scope of the invention. Therefore, the invention is intended to embrace all known or earlier developed alternatives, modifications, variations, improvements and/or substantial equivalents.