GB2610014A - Turbofan aerospike engine - Google Patents

Turbofan aerospike engine Download PDF

Info

Publication number
GB2610014A
GB2610014A GB2205620.4A GB202205620A GB2610014A GB 2610014 A GB2610014 A GB 2610014A GB 202205620 A GB202205620 A GB 202205620A GB 2610014 A GB2610014 A GB 2610014A
Authority
GB
United Kingdom
Prior art keywords
aerospike
air
turbofan
cooling
engine
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Pending
Application number
GB2205620.4A
Other versions
GB202205620D0 (en
Inventor
Lynley Ashley Adrian
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
Individual
Original Assignee
Individual
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by Individual filed Critical Individual
Priority to GB2205620.4A priority Critical patent/GB2610014A/en
Publication of GB202205620D0 publication Critical patent/GB202205620D0/en
Publication of GB2610014A publication Critical patent/GB2610014A/en
Pending legal-status Critical Current

Links

Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
    • F23R3/02—Continuous combustion chambers using liquid or gaseous fuel characterised by the air-flow or gas-flow configuration
    • F23R3/16—Continuous combustion chambers using liquid or gaseous fuel characterised by the air-flow or gas-flow configuration with devices inside the flame tube or the combustion chamber to influence the air or gas flow
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02K—JET-PROPULSION PLANTS
    • F02K3/00—Plants including a gas turbine driving a compressor or a ducted fan
    • F02K3/02—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber
    • F02K3/04—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low pressure outputs, for augmenting the jet thrust, e.g. of double-flow type
    • F02K3/06—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low pressure outputs, for augmenting the jet thrust, e.g. of double-flow type with front fan
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02K—JET-PROPULSION PLANTS
    • F02K9/00—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof
    • F02K9/97—Rocket nozzles
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23D—BURNERS
    • F23D14/00—Burners for combustion of a gas, e.g. of a gas stored under pressure as a liquid
    • F23D14/20—Non-premix gas burners, i.e. in which gaseous fuel is mixed with combustion air on arrival at the combustion zone
    • F23D14/22—Non-premix gas burners, i.e. in which gaseous fuel is mixed with combustion air on arrival at the combustion zone with separate air and gas feed ducts, e.g. with ducts running parallel or crossing each other
    • F23D14/24—Non-premix gas burners, i.e. in which gaseous fuel is mixed with combustion air on arrival at the combustion zone with separate air and gas feed ducts, e.g. with ducts running parallel or crossing each other at least one of the fluids being submitted to a swirling motion
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23M—CASINGS, LININGS, WALLS OR DOORS SPECIALLY ADAPTED FOR COMBUSTION CHAMBERS, e.g. FIREBRIDGES; DEVICES FOR DEFLECTING AIR, FLAMES OR COMBUSTION PRODUCTS IN COMBUSTION CHAMBERS; SAFETY ARRANGEMENTS SPECIALLY ADAPTED FOR COMBUSTION APPARATUS; DETAILS OF COMBUSTION CHAMBERS, NOT OTHERWISE PROVIDED FOR
    • F23M5/00—Casings; Linings; Walls
    • F23M5/08—Cooling thereof; Tube walls
    • F23M5/085—Cooling thereof; Tube walls using air or other gas as the cooling medium
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
    • F23R3/02—Continuous combustion chambers using liquid or gaseous fuel characterised by the air-flow or gas-flow configuration
    • F23R3/04—Air inlet arrangements
    • F23R3/10—Air inlet arrangements for primary air
    • F23R3/12—Air inlet arrangements for primary air inducing a vortex
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
    • F23R3/02—Continuous combustion chambers using liquid or gaseous fuel characterised by the air-flow or gas-flow configuration
    • F23R3/26—Controlling the air flow
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
    • F23R3/42—Continuous combustion chambers using liquid or gaseous fuel characterised by the arrangement or form of the flame tubes or combustion chambers
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
    • F23R3/42—Continuous combustion chambers using liquid or gaseous fuel characterised by the arrangement or form of the flame tubes or combustion chambers
    • F23R3/58—Cyclone or vortex type combustion chambers
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R2900/00—Special features of, or arrangements for continuous combustion chambers; Combustion processes therefor
    • F23R2900/03042—Film cooled combustion chamber walls or domes

Landscapes

  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)

Abstract

A turbofan engine has an internal aerospike combustor 13 located in the combustion chamber, wherein compressor air supplied into the combustor air tube 14 circulates clockwise or anticlockwise within the aerospike combustion chamber, compressor air supplied into the cooling air tube 15 exits through the cooling apertures 16 in the spike and accelerates over the aperture curved surface 17 towards the latter end, vortex generators 18 enable a boundary layer between the hot exhaust gas from the aerospike and the exiting gas from the said cooling apertures. The engine may use a zero carbon (carbon free) fuel, e.g. liquid hydrogen fuel.

Description

Turbofan Aerospike Engine
FIELD OF THE INVENTION
This invention relates to a zero carbon turbofan aerospace engine for removing fossil fuels from present jet engine.
BACK GROUND OF THE INVENTION
Present aerospace engine use carbon based fuels that produce green house gasses which increase the carbon in the atmosphere leading to global warning. Greenhouse gasses trap the heat of the sun in the atmosphere and subsequently raise the global temperature adversely effecting the climate globally.
The history of the aerospike originates in the 1950's by Rocketdyne the exhaust is more efficient then the traditional bell designed for a particular atmospheric pressure and the hence the altitude. At low attitude there is high atmospheric pressure the bell shape rocket exhaust is small. At high altitude there is low atmospheric pressure approaching vacuum the bell shape rocket exhaust is large. The Aerospike exhaust gas is shaped by the enteral air pressure. The Aerospike has many advantages however the Aerospike have problems of excessive temperatures on the spike where the exhaust makes contact with the spike.
STATMENT OF INVENTION
To overcome air pollution and enable a zero carbon flight from the present jet engines the present invention proposes a turbofan aerospike engine comprising; a plurality of radially mounted variably hinged inlet flaps 1, plurality of radially adjustable curved surface 2 wherein the curvature is of said adjustable curve surface is adjusted in addition the aperture of the said inlet flap 1 by computational judgement 4 means using data from at least one airflow speed sensor 3 this cooperative arrangement will regulate and increase the air mass flow rate in the bypass duct of the turbofan aerospike engine, bleed air from the low pressure compressor 5 is guided by ducting 6 means is used in cooling the outer combustion cans of the aerospike combustor 13, bleed air is guided by ducting 7 to an air gas accumulator 8, at least one air cooling 9 means is used to reduce the volume of the said air gas accumulator 8, at least one air flow regulator 10, at least one air supply ducting 11 in orders to supply pressurised air to the at least one external aerospike 24 external to the turbofan preferably wing mounted, internal said aerospike combustor 13 is located in the combustion camber of a turbofan engine such that the internal stagnation pressure is sufficiently high such that the exhaust gas from the aerospike is maintained within and the said aerospike combustor 13 casing without touching the container, high stagnation pressure compressor air supplied into the combustor air tube 14 and circulates clockwise or anticlockwise within the aerospike combustion chamber, high pressure compressor air supplied into the cooling air tube 15 and exits through the cooling aperture's 16 in the spike and accelerated over the aperture curved surface 17, towards the latter end, vortex generators 18 enable a boundary layer between the hot exhaust gas from the aerospike and the exiting gas from the said cooling aperture's 16 which is supplied by the said cooling air tube 15, in addition guide fins 19 guide the air flow and enable conduction / convention cooling, the generator 12 is powered by the rotating shaft of the turbofan aerospike engine supplies electrical power to the brushless electric motor 20 turns the fuel turbo pump 21which pumps the carbon free fuel to the aerospike combustion chamber 18 and circulates clockwise or anticlockwise within the aerospike combustion chamber, the carbon free fuel supplied by the turbo pump 21 and compressed air is supplied by the high pressure compressor combustor air tube 14 and mixed in the circulation and ignited by the combustion chamber igniter 22, at sufficient high altitude and speed the propulsion system maybe changes to a ram jet 23 which compressed air guided by said ducting 7 to said air gas accumulator 8, at least one said air cooling 9 means is used to reduce the volume of the air accumulator, at least one said air flow regulator 10, at least one said air supply ducting 11 in order to supply pressurised air to the at least one said external aerospike 24 external to the turbofan preferably wing mounted,
BRIEF DESCRIPTION OF THE DRAWINGS
For a better understanding of the invention, an embodiment of the invention will now be described by way of a non-limiting example with reference to the accompanying drawing which: Fig 1. Shows a view of the turbofan aerospike engine DETAILED DESCRIPTION OF THE INVENTION The following description is a preferred embodiment of the turbofan aerospike engine The description does not limit the application exclusively to the aviation industry or powers stations.
A turbofan aerospike engine comprising a plurality of radially mounted variably hinged inlet flaps 1, plurality of radially adjustable curved surface 2 wherein the curvature is of said adjustable curve surface is adjusted in addition the aperture of the said inlet flap 1 by computational judgement 4 means using data from at least one airflow speed sensor 3 this cooperative arrangement will regulate and increase the air mass flow rate in the bypass duct of the turbofan aerospike engine, bleed air from the low pressure compressor 5 is guided by ducting 6 means is used in cooling the outer combustion cans of the aerospike combustor 13, bleed air is guided by ducting 7 to an air gas accumulator 8, at least one air cooling 9 means is used to reduce the volume of the said air gas accumulator 8, at least one air flow regulator 10, at least one air supply ducting 11 in orders to supply pressurised air to the at least one external aerospike 24 external to the turbofan preferably wing mounted, internal said aerospike combustor 13 is located in the combustion camber of a turbofan engine such that the internal stagnation pressure is sufficiently high such that the exhaust gas from the aerospike is maintained within and the said aerospike combustor 13 casing without touching the container, high stagnation pressure compressor air supplied into the combustor air tube 14 and circulates clockwise or anticlockwise within the aerospike combustion chamber, high pressure compressor air supplied into the cooling air tube 15 and exits through the cooling aperture's 16 in the spike and accelerated over the aperture curved surface 17, towards the latter end, vortex generators 18 enable a boundary layer between the hot exhaust gas from the aerospike and the exiting gas from the said cooling aperture's 16 which is supplied by the said cooling air tube 15, in addition guide fins 19 guide the air flow and enable conduction / convention cooling, the generator 12 is powered by the rotating shaft of the turbofan aerospike engine supplies electrical power to the brushless electric motor 20 turns the fuel turbo pump 21which pumps the carbon free fuel to the aerospike combustion chamber 18 and circulates clockwise or anticlockwise within the aerospike combustion chamber, the carbon free fuel supplied by the turbo pump 21 and compressed air is supplied by the high pressure compressor combustor air tube 14 and mixed in the circulation and ignited by the combustion chamber igniter 22, at sufficient high altitude and speed the propulsion system maybe changes to a ram jet 23 which compressed air guided by said ducting 7 to said air gas accumulator 8, at least one said air cooling 9 means is used to reduce the volume of the air accumulator, at least one said air flow regulator 10, at least one said air supply ducting 11 in order to supply pressurised air to the at least one said external aerospike 24 external to the turbofan preferably wing mounted.

Claims (5)

  1. Claims 1. A turbofan aerospike engine comprising a plurality of radially mounted variably hinged inlet flaps, plurality of radially adjustable curved surface wherein the curvature is of said adjustable curve surface is adjusted in addition the aperture of the said inlet flap by computational judgement means using data from at least one airflow speed sensor this cooperative arrangement will regulate and increase the air mass flow rate in the bypass duct of the turbofan aerospike engine, bleed air from the low pressure compressor is guided by ducting means is used in cooling the outer combustion cans of the aerospike combustor, bleed air is guided by ducting to an air gas accumulator, at least one air cooling means is used to reduce the volume of the said air gas accumulator, at least one air flow regulator, at least one air supply ducting in orders to supply pressurised air to the at least one external aerospike external to the turbofan preferably wing mounted, internal said aerospike combustor is located in the combustion camber of a turbofan engine such that the internal stagnation pressure is sufficiently high such that the exhaust gas from the aerospike is maintained within and the said aerospike combustor casing without touching the container, high stagnation pressure compressor air supplied into the combustor air tube and circulates clockwise or anticlockwise within the aerospike combustion chamber, high pressure compressor air supplied into the cooling air tube and exits through the cooling aperture's in the spike and accelerated over the aperture curved surface, towards the latter end, vortex generators enable a boundary layer between the hot exhaust gas from the aerospike and the exiting gas from the said cooling aperture's which is supplied by the said cooling air tube, in addition guide fins guide the air flow and enable conduction / convention cooling, the generator is powered by the rotating shaft of the turbofan aerospike engine supplies electrical power to the brushless electric motor turns the fuel turbo pump which pumps the carbon free fuel to the aerospike combustion chamber and circulates clockwise or anticlockwise within the aerospike combustion chamber, the carbon free fuel supplied by the turbo pump and compressed air is supplied by the high pressure compressor combustor air tube and mixed in the circulation and ignited by the combustion chamber igniter, at sufficient high altitude and speed the propulsion system maybe changes to a ram jet which compressed air guided by said ducting to said air gas accumulator, at least one said air cooling means is used to reduce the volume of the air accumulator, at least one said air flow regulator, at least one said air supply ducting in order to supply pressurised air to the at least one said external aerospike external to the turbofan preferably wing mounted.
  2. 2. A turbofan aerospike engine according to claim 1, in which the carbon free fuel means maybe liquid hydrogen or other environmentally safe fuel.
  3. 3. A turbofan aerospike engine according to claim 1, in which the high altitude propulsion means is enables by means If said a ram jet and said at least one said external aerospike external to the turbofan preferably wing mounted.
  4. 4. A turbofan aerospike engine according to claim 1, in which the said generator and said brushless electric motor is preferably using superconducting material.
  5. 5. A turbofan aerospike engine according to claim 1, in which the said computational judgement means is provided to control the aerospike engine combustion.
GB2205620.4A 2022-04-14 2022-04-14 Turbofan aerospike engine Pending GB2610014A (en)

Priority Applications (1)

Application Number Priority Date Filing Date Title
GB2205620.4A GB2610014A (en) 2022-04-14 2022-04-14 Turbofan aerospike engine

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
GB2205620.4A GB2610014A (en) 2022-04-14 2022-04-14 Turbofan aerospike engine

Publications (2)

Publication Number Publication Date
GB202205620D0 GB202205620D0 (en) 2022-06-01
GB2610014A true GB2610014A (en) 2023-02-22

Family

ID=81753194

Family Applications (1)

Application Number Title Priority Date Filing Date
GB2205620.4A Pending GB2610014A (en) 2022-04-14 2022-04-14 Turbofan aerospike engine

Country Status (1)

Country Link
GB (1) GB2610014A (en)

Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20180313542A1 (en) * 2015-10-29 2018-11-01 Safran Aircraft Engines Aerodynamic injection system for aircraft turbine engine, having improved air/fuel mixing
US20200049103A1 (en) * 2018-06-29 2020-02-13 Christopher Craddock Aerospike Rocket Engine
US20200173662A1 (en) * 2018-11-29 2020-06-04 General Electric Company Premixed Fuel Nozzle
CN113864819A (en) * 2021-09-14 2021-12-31 南京航空航天大学 Afterburner with air cooling structure

Patent Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20180313542A1 (en) * 2015-10-29 2018-11-01 Safran Aircraft Engines Aerodynamic injection system for aircraft turbine engine, having improved air/fuel mixing
US20200049103A1 (en) * 2018-06-29 2020-02-13 Christopher Craddock Aerospike Rocket Engine
US20200173662A1 (en) * 2018-11-29 2020-06-04 General Electric Company Premixed Fuel Nozzle
CN113864819A (en) * 2021-09-14 2021-12-31 南京航空航天大学 Afterburner with air cooling structure

Also Published As

Publication number Publication date
GB202205620D0 (en) 2022-06-01

Similar Documents

Publication Publication Date Title
US10654579B2 (en) Cooled cooling air system for a gas turbine
CN108137164B (en) Auxiliary power unit with combined generator cooling
US9429072B2 (en) Return fluid air cooler system for turbine cooling with optional power extraction
US7690188B2 (en) Combination engines for aircraft
US11674443B2 (en) Hydrogen fuel system
US7485981B2 (en) Aircraft combination engines complemental connection and operation
US7690189B2 (en) Aircraft combination engines inlet airflow control system
US2514513A (en) Jet power plant with boundary layer control for aircraft
US20080314573A1 (en) Aircraft combination engines thermal management system
BR102017004993A2 (en) ENGINE BLEEDING CONTROL SYSTEM, AND METHOD FOR CONTROLLING A MOTOR BLEEDING SYSTEM
RU2013126294A (en) COMBINED TURBO-RINGER REACTIVE ENGINE
EP1992788B1 (en) Aircraft combination engines plural airflow conveyances system
CN106762221B (en) Turbo ramjet engine ram-air turbine heat to electricity conversion and forecooling method
US6499286B1 (en) Method for the operation of a supersonic jet-engine combined power unit
Kang et al. Multi-stage turbocharger system analysis method for high altitude UAV engine
EP3995679A1 (en) Hydrogen fuel system
RU2561757C1 (en) Three-component air-jet engine
Chen et al. The installation performance control of three ducts separate exhaust variable cycle engine
CN108843460A (en) Heat to electricity conversion and pushing method is pre-chilled in turbo ramjet engine
US7762056B2 (en) Aircraft combination engines exhaust thrust recovery
US20220065163A1 (en) Ground operations of a hybrid electric propulsion system
GB2610014A (en) Turbofan aerospike engine
RU135000U1 (en) HYDROCARBON RECTANGULAR ENGINE
US11827370B1 (en) Aircraft propulsion system with intermittent combustion engine and electric transmission system and method for operating the same
CN110023593A (en) Turbogenerator and its cooling means