GB2383823A

GB2383823A - Jet propulsion engine

Info

Publication number: GB2383823A
Application number: GB0223247A
Authority: GB
Inventors: Harold Birkett
Original assignee: Individual
Current assignee: Individual
Priority date: 2001-10-09
Filing date: 2002-10-08
Publication date: 2003-07-09
Anticipated expiration: 2022-10-08
Also published as: GB0223247D0; GB2383823B

Abstract

A thermo-dynamic air induction, compression, combustion and ejection air breathing and rocket type propulsion engine comprised of modules each of which in air breathing mode uses the high velocity low pressure exhaust gasses ejected from a combustion chamber outlet for the purpose of ambient air induction which operates on a pulse combustion cycle during the start up period then on a resonance cycle and finally on a continuous combustion cycle and is also able to function in rocket engine mode whilst the arrangement is modular construction with each relative to each other to suit the use and structure into which they are installed and not restricted to any particular profile. In figure 1, the stream of expanded gases from the forward combustion chamber assembly 3 passes through injector nozzles 4 and over profiled surface 5 thus entraining air from intake 1. The resulting mixture passes into middle chamber 6 from which one stream passes via rear combustion chamber 7 to outlet 8 while the remainder returns via passage 9 and flap valves (33, fig.2) to combustion chamber 3. The engine operates on a pulse cycle during start-up, then on a resonance cycle and finally on a continuous combustion cycle.

Description

IMPROVED PROPULSION ENGINE by Harold Birkett DESCRIPTION.

Technical Field.

This invention relates to a thermo-dynamic jet propulsion engine which is able to function as an air breathing or rocket type engine or a combination of both and is more versatile in its use due to modular construction Background Art.

Pulsejet, gas turbine, ramjet, rocket engine and gas ejector.

Disclosure Of Invention.

A thermo-dynamic air induction compression combustion and ejection air breathing/rocket type propulsion engine derived from proven technologies of modular construction whose arrangement are integrated relative to each other but not restricted to any particular profile or shape and can be arranged to suit the aerodynamic or other profile into which they are fitted comprising of at least one combustion chamber which uses the ejected high velocity low pressure exhaust gasses from each combustion chamber outlet for the purpose of air induction with the resulting air/exhaust gas mixture being divided into at least two streams with one stream recirculated back via mechanical (or aerodynamic) non-return valve means as required into combustion chamber which operates on a pulse combustion cycle during start up then on a resonance cycle and finally on a continuous combustion cycle as the fuel supply velocity pressure and volume of air/exhaust gasses passing through the engine increases so that the non-return means remain open with a steady flow passing through whilst another stream is ejected rearward via a further combustion chamber if used where further fuel is added ignited and expelled whilst a third stream if required passes through a pressure relief or other control means into a by-pass and is also expelled. The engine has a multi type fuel/gas and injection facility enabling the use of liquefied gasses such as oxygen, hydrogen, air, powdered fuel used in combination with each other or with the petrocarbon fuel during the air breathing mode which may be required during start up in order to achieve the required static pressure for continuous combustion or used to supplement the reduced air supply to engine in rarefied atmosphere during the transition period from air breathing to rocket engine mode required for propulsion in space.

Fig 1. Example Of An Engine With Modules Located Equally Around Axis.

Shows a sixteen module assembly comprising of :-Outer enclosure 1, annular air inlet suction chamber 2, forward combustion chamber assembly 3, ejector nozzle 4, aerodynamic profile surface 5, middle chamber 6, rear combustion chamber assembly 7, multiple fuel injectors 13, ignitions 14, outlet nozzle 8, passageway 9, pressure relief or controlled type gas/air inlet 11, into by-pass 12, (if required), with all located around central vertebrae tube assembly 10.

Fuel (and oxidant or other gas if required) is injected into each of the forward combustion chamber assemblies 3, and each ignited in an empirically determined sequence when combustion takes place with each of the resulting stream of expanded gasses passing through ejector nozzles 4, and over suitably profiled surface 5, resulting in the high velocity low pressure streams of gasses inducing ambient air through air intake 2, which is entrained into and around the ejected gasses with the resulting air/gas mixture passing into the middle chamber 6, with decreasing velocity and increasing pressure then out again, with one stream passing into the rear combustion chamber 7, (where when required further fuel is added via

13, ignited 14, ) and ejected rearwards through the outlet nozzle 8, whilst the other stream passes forward through passageway 9, and into the forward combustion chamber 3, via non-return means together with air sucked back through the ejector nozzle 4, due to the negative pressure created inside the combustion chamber at the end of each combustion cycle when the gasses are purged, recharged, compressed and ignited.

The above work cycle is repeated with increasing pressure, velocity, volume of air/gasses, optimum fuel input, and number of cycles per second, starting with pulse, with the valves opening and closing then resonance with the valves opening and partially closing and finally continuous combustion and flow of air/gas when the required velocity and pressure differential to the flow of air/gas into the forward combustion chambers is achieved and the non-return valves remain in a near static open position. During the start up period the required velocity at some position between the ejector nozzle outlet and combustion chamber inlet has not necessarily to be that which will generate a pressure wave in order to achieve continuous combustion but is desirable in order to avoid combustion perturbation.

Although Sect C-C shows the orientation of the ejector nozzles 4, as circumferential segments around the axis of the engine assembly these will be of a shape and orientation which will maximize air induction and gas/air mixing required to support combustion.

Noise emission.

Suppressed radially by reflection and insulation of the outer enclosure 1.

Reduced from the front air inlet 2, with the forward sound emitted from the rearward facing ejector nozzles 4, traveling upstream against the flow of induced air.

Reduced from rear outlet nozzle 8, due to the sound waves being reflected from the curved rear surface of middle chamber 6, whilst the remainder is refracted from and dispersed around central tube 10, before being emitted through outlet 8, with no suppression of the noise emitted by the rear combustion chamber 7, when in use.

The total sound emitted from forward combustion chamber 3, during pulse combustion will be from each combustion chamber at different stages of the combustion cycle due to each being fired separately in a pre-defined sequence controlled by an electronic closed loop ignition system which will result in a wave canceling or muffling effect upon each other when they converge and mix inside middle chamber 6.

Transition time from pulse to continuous combustion.

Is empirical dependent upon the size of the combustion chambers but would quickly achieved due to the low moment of inertia of the thin metal non-return valves and gasses, and rapid increase in pressure, velocity and flow of gasses. This time may be may be interpolated from the German VI-launch speed 150 m. p. h, 45 pulses per second at 360 m. p. h. compressed air 50 litres/mile inside a I I ft long x 1.9 ft dia formed tube with metal leaf type non-return valves at the front end. Also from drone and model aircraft (restricted in use due to noise).

The central vertebrae tube can be used as required to house the fuel, gas and electrical supply services and controls for the engine or as an engine by-pass or for atmos. air supply to rear combustion chamber or as a ram jet with tube suitably profiled or an ion or other space propulsion assembly.

Fig 2. Example Of A Typical Combustion Chamber Assembly 3, Outer tubular case assembly 31, non return air/gas inlet holes 32, shown covered by thin spring steel strips attached between holes to the inside of 31, which form non-return flaps 33, which due to being sprung from flat to an arc result in the flaps exerting a non-return sealing pressure around the holes, fuel and or gas injector assemblies 34, together with ignition spark/glow plug 35, which may be axially adjustable, inner flame tube 36, exhaust

gas ejector nozzle outlet 4. Required connections for parts comprising this assembly will be to suit.

Operation.

Start up fuel (and gas if required) is injected through multi injectors 34, into the combustion chamber 31, ignited by spark ignition means 35, (adjustable if required) when combustion takes place resulting in the expanding gasses traveling under pressure along the chamber and out through ejector nozzle 4, which results in the chamber being purged. Fuel is injected into the air/gas entering into the combustion chamber through the non return inlets 32, and back again through the ejector nozzle outlet 4, due to the negative pressure at the end of each exhaust cycle of combustion, when the air/gas/fuel mixture is compressed due to the dynamic pressure caused by the wave action of the flow of gasses into the chamber from each end of the chamber then ignited ideally at the point of maximum compression when the above cycle is repeated. but with an increase in pressure, velocity, quantity of air/gas flow and cycles/sec dependent upon the quantity of fuel injected until continuous combustion is achieved as previously described.

The dimensions, profiles and material of construction are empirical dependent upon requirements.

aa : Fig 2AFurtherexampleOfA Typical Combustion Chamber Assembly3, Showing circumferential non-return flaps punched or cut through a rectangular flat thin spring metal sheet 33, which upon assembly would be flexed into a spiral and inserted inside combustion chamber 31, then sprung outwards and attached around the inside eqi-distance between alternate line of holes 32, thus forming spring pre-loaded non-return flaps over each hole. End air/gas inlet 37, with axially sliding non-return valve 38, (may require to be spring loaded). Orientations'a'and'b'each showing empirical examples of combustion chamber arrangements be dependent upon the most efficient air induction, gas/air mixing and compression capabilities.

Fig 3. Example Of A Twin Engine Aerospace Craft.

Showing each engine assembly with modules around an open ended vertebrae tube which can act as a by pass or secondary air supply to rear combustion chambers) at subsonic speed or as a ram jet profiled combustion chamber with fuel injection means at supersonic speed and/or as an enclosure for services or controls or other assembly which could be installed if used for propulsion in space.

Fig 4. Example Of A Single Engine Aerospace Craft.

Showing engine module assembly at rear of craft with fuselage profiled in such a manner so as to act as a ram jet pressure head to the engine air inlet which will maximize air intake area and minimize surface drag area.

Fig 5. Example Of A Multi-Engine Aerospace Craft.

Showing engine module assemblies at rear of a rectangular wedge (or other profile to suit) shaped craft together with innovations as required consisting of :Outer fuselage 1, profiled to provide lift/or ram jet pressure head to upper and lower air inlets 3, into propulsion engines 4, vertical members profiled to suit projecting above and below which act as air retainers, provide structural stability, house services and controls as required 2, wing assemblies 5, jet powered elevator/ailerons 6, and rudders 7, upper and lower thrust divertor/air brakes 8, at least one pair of lift inparting aerofoils/ailerons or slats 9, powered vertical lift means 10, pressurized interior fuselage 11, thermal/acoustic/fire/impact protection provided in between structures by high strength

balloon like pockets (integrated fire retardant powder), of different sizes inflated with fire suppressant gas such as carbon dioxide 12, retractable wings 13.

Fig 6. Example Of A Single Annular Combustion Chamber Engine Similar in principle as previous but with one annular combustion chamber with a continuous ejector nozzle (with spacers inside outlets if required). The combustion zone could be further back towards the air/exhaust gas inlet however.

Enclosed Fig. sizes, profiles, proportionality, material, construction etc. are empirical and diagrammatic only.

It may be that after continuous combustion and/or a pressure wave has been established by the method described then a static heat source such as nuclear or other radiant could be used located between the ejector outlets and combustion chamber inlets and used in a closed loop gas turbine assembly for instance.

Although ram and pulse jets have proven unsatisfactory when used to rotate helicopter blades due to sound and cavitation it may be the propulsion modules described could function satisfactorily when integrated into or attached to helicopter blades Dependent upon results it may be that non-return means could be aerodynamic in some instances.

Claims

IMPROVED JET PROPULSION ENGINE.

1. A thermo-dynamic air induction compression combustion and ejection air breathing propulsion engine derived from proven technologies of modular construction whose arrangement are integrated relative to each other but not restricted to any particular profile or shape and can be arranged to suit the aerodynamic or other profile into which they are fitted each of which uses the high velocity low pressure exhaust gasses ejected from each combustion chamber outlet for the purpose of air induction with the resulting air/exhaust gas mixture being divided into at least two streams with one stream recirculated back via mechanical (or aerodynamic) non-return means as required into the combustion chamber which operates on a pulse combustion cycle during start up then on a resonance cycle and finally on a continuous combustion cycle as the fuel supply velocity pressure volume of air/exhaust gasses passing through the engine increases so that a pressure wave may or may not be generated in order that the non-return means remain open with a steady flow passing through whilst another stream is ejected rearward via a further combustion chamber if used where further fuel or gas is added ignited and expelled whilst a third stream if used passes through a by-pass and is expelled from the rear of the engine which has a multi type fuel injection facilities enabling the use of liquefied air or reactive gasses such as oxygen and hydrogen, or powdered fuel or in conjunction with each other or with the petrocarbon fuel both in air breathing mode which may be required during start up in order to achieve the required static pressure for continuous combustion and to supplement the air supply in rarefied atmosphere during the transition period from air breathing to rocket engine mode for propulsion in space.

2. A thermo-dynamic air breathing propulsion engine as claimed in claim 1, where the combustion chambers are arranged around an open ended central vertebrae tube assembly enclosing electrical, mechanical and control equipment with a conical pressure head covering the front end of the tube.

3. A thermo-dynamic air breathing propulsion engine as claimed in claim 1 and 2 but with an open ended central vertebrae tube assembly through which air can pass from the front to the rear of the engine assembly and act as an engine by pass.

4. A thermo dynamic air breathing propulsion engine as claimed in claim 1 and 3, where air can pass from the front directly into the rear combustion chamber.

5. A thermo-dynamic air breathing propulsion engine as claimed in claim 1, and 2, where the open ended central tube is suitably profiled and functions as a ram jet engine when the required forward air inlet velocity is achieved..

6. A thermo-dynamic air breathing propulsion engine as claimed in claim 1 and 2, with modules located at the rear end of an aircraft whose profile is such that the aircraft fuselage acts as a ram jet pressure head into the air intake of the engine resulting in maximum air intake area with minimum surface area drag.

7. A thermo-dynamic air breathing propulsion engine as previously claimed which is able to function using or in conjunction with different types of fuel including powders, liquefied environmentally friendly reactive gasses and oxidants.

8. A thermo-dynamic air breathing propulsion engine as previously claimed which is able to function using a static radiant heat source in middle chamber after required gas flow through engine has been established.

9. A thermo-dynamic air breathing propulsion engine as claimed in claim 1, and 2, where an ion or other space propulsion system can be installed inside the central open ended tube.

10. A thermo-dynamic air breathing propulsion engine substantially as described herein with reference to Figures 1-4 of the accompanying drawings.