CONTROLLING THE FLOW RATIO IN A MULTIFLOW JET ENGINE

Abstract: The system constitutes an additional aircraft propulsion system comprising an outer duct designed to wholly enclose each separate engine powering the aircraft whether it be either, a pure, bypass, or fanjet type of jet engine. This ducting has included within it an arrangement of variable annulli located between it and the engine it encloses. These annulli are caused to move in conjuction with the axial movement of either the inlet to, and or, the exit from the outer ducting. This system arrangement is caused to funtion as an "Air Velocity Gearbox", whereby, the air mass flow velocities both entering and leaving the jet engine are caused to combine with the additional air mass flow passing through the ducted system. The purpose is to cause the resultant air velocity leaving the combined system to be not less than "V" (the aircraft velocity) and not greater than "Vj" (the engine propelling nozzle velocity. The optimum desired resultant velocity will be dependent upon the interactive gearing (mixing) respective air velocity ratios determined as most suitable to the attainment of the optimum propulsive efficiency at any specific condition during the aircraft's flight mission.

Description: Controlling the flow ratio in a multiflow jet engine This invention relates to an additional propulsion system for jet powered aircraft and to an aircraft provided with such a system and is called an Air Velocity Gearbox. A jet powered aircraft is primarily a Motion Vehicle which is designed and constructed to carry a payload that is caused to fly tlirough the atmosphere. Such an aircraft represent the best known application of the 3rd Law of Motion which states, "To every action there is an equal and opposite reaction". It is conventionally accepted tlat in accordance with this Law that the Jet Engine provides the Action part of the application, the thrust so provided being harnessed to propel the aircraft and in so doing, comply with the Reaction part of the Law thereby completing this particular application of the 3rd law of motion. This invention is in no way concerned with improving the efficiency of any existing system comprising a conventional Jet Powered Aircraft. A conventional Jet Powered Aircraft in the sense used here is that which is propelled by either a pure, bypass or fan type of Jet Engine. This invention is solely concerned with esablishing an improvement in the Propulsive Effectiveness of a conventionally jxwered aircraft in flight in receipt of its jet engine thrust. A jet powered aircraft in flight, although forming one single physical Motion Vehicle, is in reality caused to use two totally different interacting Air Motion Systems at the same time in doing so. Essentially, a Jet Engine or that which we technically regard as a Jet Engine in the free standing state, in terms of Motion Physics, is no longer a Jet Engine in the conventional sense when propelling an Aircraft in Flight, it is AN AIR MOTION SYSTEM. IT MOVES AIR. Furthermore, a Jet Powered Aircraft in Flight also in Motion Physics is no longer a Jet Powered Aircraft, it is AN AIR MOTION VEHICLE/SYSTEM, IT MOVES AIR and, as a consequence, is caused to fly. Hereafter referred to as Motion Systen No.2. The Case. This invention concerns itself with only one aspect of a Jet Powered Aircraft in flight, namely, its Propulsive Effectiveness as a Motion Vehicle that is caused to fly. Aerodynamics, thermodynamics, turbine, compressor, combustion, mectianical efficiencies and the like are not the issue here. We are in no way concerned with the technology or efficiencies of any of the existing physical system comprising a modern jet powered aircraft. The one and only issue we are concerned wit is the Motion Effectiveness in terms of the Velocity attainable by the Aircraft upon receipt of its engine thrust, therefore, the effectiveness of the work done upon the aircraft as a Motion Vehicle that is caused to fly utilising Newton's 3rd Law of Motion in this particular application. We are concerned only with those factors which govern the Motion Efficiency of the Jet Powered Aircraft in flight, ie. , the Velocity attainable in response to the receipt of a specific thrust Force. Note: It is assumed that both the design and construction of any existing or future aircraft considered is technologically the best possible. The task is how to improve the Propulsive Effectivensss of even that regarded as the best attainable in terms of Conventional thinking. The only laws we are concerned with are Newton's three Laws of Motion. The only science that we need concern ourselves with is that of fundamental Dynamics, i.e that branch of science dealing with Force as producing or effecting Motion. Fundamentals A .let Powered Aircraft in flight is no longer primarily a Jet Powered Aircraft in terms of Motion Physics, it is a Motion Vehicle that is caused to fly thereby requiring it to be designed and constructed in this way in order to utilise Newton's 3rd Law of Motion in this particular manner. A Jet Powered Aircraft in Flight, in terms of Motion Law is not simply a Motion Vehicle that it caused to fly, it is primarily an AIR MOTION SYSTEM. IT MOVES AIR. IT ACELLERATES THE AIR THROUGH WHICH IT IS PASSING thereby, making sustained Jet Powered Flight as it is understood, possible. Hereafter referred to as Motion System No.2. A Jet Engine when used to power an aircraft in flight, IN MOTION PHYSICS is no longer a Jet Engine in the conventional sense it is regarded but, is primarily, an AIR MOTION SYSTEM; IT ALSO MOVES AIR. Hereafter referred to as Motion System No.l. IT IS THE INTERACTING REACTION OF THE AIR VEIOCITIES ACTING BETSJEEN THESE TOO TOTALLY DIFFERENT AIR MOTION SYSTEMS THAT IS THE PRINCIPLE FACTOR GOVERNING THE PROPULSIVE EFFECTIVENESS OF THE ENGINE THRUST UPON THE AIRCRAFT THAT IS CAUSED TO FLY BEIWEEN THEM, AND THEREFORE, ITS RESULTANT VELOCITY. Because the medium through which the Aircraft Motion takes place is Air, the Propulsive Effectiveness of it as a Motion Vehicle that is caused to Fly, will have two components: firstly, that of its aerodynamics which is determined by the physical characteristics of Air In motion around physical structures secondly, the differences between the interacting AIR Velocities acting between both Air Motion Systems an the Aircraft at the same time. The first of these we can do nothing about and is the price we have to pay for the privilege of flying through the air. Not so the second, it is here, the Propulsive Effectiveness of the Jet Powered Aircraft as a Motion Vehicle that is caused to fly is determined. In Motion Physics, the Jet Engine when forming an integral part of an aircraft in flight does not propel the aircraft in the direct manner of say, the car engine wMch transmits a direct physical torque tlirust through its transmission system to the Earth Mass. Such a tlirust comprises essentially of a single point direct tlirust. Not so a jet engine. The tlirust of a jet engine is the product of a continously accelerating air mass passing through it, but, it is no longer a jet engine in the conventional sense, it is an integral part of the aircraft structure. Accelerate a car engine and the increase tlurust force is transmitted as a direct physical ACCELERATION TO THE EARTH MASS through its direct physical shaft transmission system and, the CAR MASS ITSELF, DIRECTLY ACCELERATES. Accelerate a jet engine and its AIR MASS FLOW ACCELERATES RELATIVE TO THE STILL AIR MASS OF THE ATMOSPHERE AND IN SO DOING CAUSES THE AIRCRAFT TO ACCELERATE: THIS IS A VITALLY IMPORTANT DIFFERENCE IN HOW THRUST FORCE IS UTILISED TO CAUSE VEHICLE MOTION OR ACCELERATION IN THESE TWO TOTALLY DIFFERENT APPLICATIONS OF THE 3RD LAW OF MOTION. In Motion Physics, it is the accelerating air mass flow passing tlrough the engine of a jet powered aircraft in flight that is the both the source and cause of the tlirust force that propels the aircraft. The fact that it is the jet engine that causes this to take place, in Dynamics, is an irrelevance. The Engine is simply the means of accomplishing the momentum change in the air mass flow passing tlirough it as an integral part of the aircraft structure. In the steady flight state the Aircraft is already in receipt of a forward Tlirust Force resulting from the rearwardly accelerating Air Mass passing tlirough its engines which, are caused to form an integral part of its structure. A Simple Analogy of The Principle of this Invention. Consider; in the case of the motor car it was discovered that by introducing a gearbox into its transmission system between the engine and the transmission points (its drive wheels) of contact with the EARTH MASS we could more efficiently utilise the power delivered to achieve a particular Velocity state relative to the EARTH MASS and, thereby, improve the Propulsive Effectiveness of the power required to accomplish a particular motion state/velocity. This is only possible because the connection between the source of power and the physical Earth Mass is tlirough a physically direct shaft/wheel transmission system. The point is, the gearbox, to be effective, has to be logically interposed between both the source of power, the engine, and THE EARTH MASS direct. Consider, a motor car is a Motion Vehicle in the same way as a Jet Powered Aircraft is, both represent utilisations of the same Motion law, only the application of the overall principle is different. Therefore, if the principle of a gearbox and its Propuslive Effectiveness benefits holds good for the motor car it must also hold good, in principle, for any motion system e.g., a Jet Powered Aircraft in Flight. Consider, a modern Fan Jet powered Aircraft in terms of Propulsive Efficiency i.e., Motion Efficiency terms, can be likened to a modern motor car with only two gears. If my logic holds good, we should somehow be able to apply the gearbox principle to jet powered flight. The question of course is, How? The answer is given above. It simply requires recognising that the source of tlirust in a Jet Powered Aircraft is not the Jet Engine itself, this is simply the means of producing the tlirust, it is the actual air mass flow itself. This then takes the place of the source engine power in the car, in the car gearbox logic above. And, in place of the physical Earth Mass we have the Air Mass surrounding the Earth, The Atmosphere, which, in physics, is simply a direct extention of the Earth Mass. Consider, in the car gearbox case, the gearbox is logically interposed directly between the two physical systems involved, namely, the engine itself and the Earth Mass. And so, any Propulsive Efficiency system for a Jet Powered Aircraft must logically be interposed between the equivalent two systems which are, in fact, the engine air mass flow and the surrounding extension of the Earth Mass i.e., The Atmosphere itself. This is what my invention is all about. I have simply taken the principle of the propulsion efficiency system of the gearbox of the motor car and applied it to a Jet Powered Aircraft; both being different applications of exactly the same Law. If you accept the Propulsive Effectiveness of the gearbox in the motor car which Is a Reality then, it must hold good in principle for all similar applications in principle of Newton's 3rd Law, and it does. An aircraft in flight, is a Motion Vehicle in the same way that a car is, only in this instance it is the form of an aeroplane that is caused to fly. In level flight it is already in receipt of the Tlirust Force required to maintain the steady state. It is the effectiveness of this Tlirust Force upon the aircraft in terms of its Resulting Velocity of the Aircraft that we are solely concerned with. In terms of Dynamics it is not simply a Motion Vehicle; it is the reaction to the acceleration of the air around the aircraft caused by its passage through the Air that makes sustained flight possible. It is no longer simply a Motion Vehicle that happens to be constructed in the form of a Jet Powered Aircraft it is primarily an AIR MOTION SYSTEM. IT MOVES AIR. IT ACCELERATES THE AIR through which it is passing, (Motion System No.2). The Aerodynamics part of the Propulsive Losses is caused by the RE-ACTION part of the application of the 3rd Motion Law as it applies to the Aircraft as a whole. The differences between the interacting AIR VEIJOCITIES acting between the engine as an integral part of the aircraft structure and the aircraft as a whole, also results from the RE-ACTION part of the utilisation of the Law and, is directly reflected in the AIRCRAFT'S RESULTANT VELOCITY, RELATIVE TO THE RATE OF CHANGE IN AIR MASS FLOW ACCELERATION (VELOCITY INCREASE) TAKING PLACE THROUGHOUT THE ENGINE CYCLE AS A PHYSICALLY INTEGRAL, PART OF THE AIRCRAFT STRUCTURE AND, TO THE RELATIVE STILL AIR CONDITION OF THE ATMOSPHERE THROUGH WHICH IT IS PASSING, AT THE SAME INSTANT. To maintain the steady flight state in compliance with Newton's First Law of Motion, the Force necessary is that required to overcome, what is conventionally regarded as simply Aircraft Drag, i.e the Forces acting against the maintainance of this state which lias already been paid for in specific fuel consumption.. The actual causes of these RE-ACTION Thrust Losses are totally different. There are three principle comirønents to the drag force ACTING BETWEEN THESE WO DIFFERENT AIR VELOCITY/MOTION SYSTEMS and the aircraft flying between them these are:(l),aerodynamnic drag, Da (profile and form) and (2), Cx,pr the engine drag coefficient and (3), Apr, qo engine intalce momentum drag. In conventional practice these losses are regarded as simply drag losses requiring a S{>ecific tlirust force to overcome them or in the steady flight case to balance them. In performance calculation there is no distinction made between these tliree totally different forms of drag losses that are caused by tliree quite different factors. The first, by the aerodynamics of the aircraft and its RESULTANT VELOCITY; the second principally caused by the high compressor compression ratios, third by the differences between the interacting air velocities acting across the entire aircraft inclusive of its engines i.e., (Vj -V), V and, the relative still air condition of the atmosphere tlirough which it is passing at any specific instant "Vo". The first is a physical friction drag loss, the second, an air mass flow velocity loss, and third, a motion/velocity loss ex p erienced by the Aircraft in its passage tlirough the air, caused by the loss of work done upon the Aircraft due to the uncontrolled interaction of the air velocities acting between the engine and aircraft AIR MOTION SYSTEM'S and the atmosphere through which the Aircraft as a Motion VEHICLE is passing. Propulsion System Drag is traditionally determined by the formula: Dpr - Cx,pr Apr, qo. where; Cx,pr is the engine drag coefficient, Apr, the greatest cross sectional area of the engine with cowling and qo, the dynamic pressure of flight speed. But, Cx,pr is the com{x>nent largely due to the compression pressure ratios of the engine which is a design factor, whereas, Apr, qo are the component parts caused by the Air Motion Velocity loss suffered by the aircraft as a Motion Vehicle and is caused as a direct result of the bastard Motion Systems marriage between its jet engines and the aircraft they i∞wer. The above equation in this form clearly implies tliat the Apr,qo are drag comjκπients of the engine and that by implication are caused by the engine. That the loss manifests itself at the engine is correct but, that the engine is the cause, is not correct. The Range Factor of the engine: " The following factor Kr, represents the connection between fuel mass, engine mass and net thrust and can also be expressed in terms of the specific values". In the simplest case, in which the flight mission is to be accomplished at constant speed and altitude, the engine data for these conditions can be taken from the performance specification. The following Figure shows the range factor Kr, where:- mpr/Fpr: is the unit mass of the propulsion system, as a primarily technological criterion. fr Fpr: is the specific fuel consumption, a primarily thermo-dynamic criterion. Dpr/Fpr: is an expression containing tlirust per frontal area, therefore a primarily aerodynamic criterion. mpr . mfr Kr = mpr + mfr = Fpr Fpr wo Fpr - Dpr 1 - Dpr Fpr Where: mpr Mass of Propulsion System Fpr Tlirust of Propulsion System. Dpr Drag of Propulsion System. Apr Cross section of Propulsion System. Cx,pr Drag Coefficient of Propulsion System. qo Dynamic Pressure. r Range. wo Flight Speed. mfr Fuel Mass for flight Mission Range. Cx,pr Drag Coefficient of Propulsion The Range Factor Kr combines with the specific values of the engine in a meaningful manner and also defines the mission. The lower the value of Kr the better the engine is likely to be. This invention enables us to virtually eliminate the Apr, qo components from the Dpr equation and as a bonus also makes possible the possibility of practially reducing Cx,pr at any equivalent tlirust condition thereby, appreciably reducing the Kr factor. The utilisation of Newton's 3rd law in THE STEADY FLIGHT STATE for example, is in the form of a continuous cycle of interaction between two major Air Motion Systems that are caused to take place one after the other as part of a continuous process. These are respectively, "THE ACTION & RE-ACTION of the Engine as an AIR MOTION SYSTEM (Motion System No.l.) in creating a Tlirust Force resulting in the ACTION that is The Aircraft's Resultant Velocity. The RE-ACTION to which, is to cause the Acceleration of the surrounding Atmosphere (Motion System No.2) thereby, forming, the two major Air Motion Systems involving this particular application of The I_aw. It is the Air Velocity Interaction between these two AIR MOTION SYSTEMS that makes the flight of a Jet Powered Aircraft between them jκjssible in this particular application of The Law. In Dynamics, it is a physical impossibility to recover a loss of MOTION VELOCITY of any lcind incurred in the RE-ACTION half of the cycle in the ACTION part of the same MOTION SYSTEM CYCLE, especially, as in this specific instance, where the ACTION part and the RE-ACTION parts res p ectively, belong to two totally different AIR MOTION SYSTEMS that are physically forced to ACT as IF THEY ARE ONE PHYSICAL MOTION SYSTEM. In the level flight steady state condition the resultant AIRCRAFT VELOCITY attainable in receipt of a specific engine tlirust represents, the test Velocity resultant attainable from all our technology. And, in terms of systems efficiencies all oxir ingenuity in the design and construction of the aircraft and its engines in this particular application of Newton's 3rd Law of Motion. It is the ultimate measure of the velocity efficiency of this particular application of the 3rd Law of Motion. How good it is, is directly reflected in how fast it can fly, ie, its Velocity, in receipt of the optimum Tlirust that can le economically justified for a specific mission. Virtually all of our technology is invested in the design and construction of the aircraft which regrettably, lias principally concerned itself with meeting the scientific requirements that make possible the utilisation of the ACTION part of the 3rd Law in this particular instance. Comparatively little thought has leen devoted to understanding the RE¬ ACTION part of the Law as if, having achieved the ulitimate teclinologically in aircraft and engine design, the result is regarded as the lest that can be achieved in terms of the physical utilsation of the 3rd Law. This is not so. The Propulsive Efficiency of the Aircraft as a Motion Vehicle is determined in the RE-ACTION part of the utilisation of the Law which, in accord with the Law, is governed by the DIFFERENCES BETWEEN THE RESULTANT INTERACTING AIR VELOCITIES acting across the entire aircraft inclusive of its engine/engines tlat is caused to react as one physically combined MOTION SYSTEM IN THE FORM OF ONE COMBINED PHYSICAL MOTION VHHCLE/AIRCRAFT, i.e, the difference between Engine (Vj - V) and the resultant aircraft veloctiy V where Vj is the final jet VELOCITY and V the resultant AIRCRAFT VEIJOCITY and the relative still Air Velocity of the Atmosphere Vo. In a Jet Powered Aircraft in terms of Motion Physics, whilst the Jet Engine is the means whereby the Aircraft tlirust is produced, it is not in fact, in Motion Phusics j the engine that propels the aircraft but, the reaction to its ACCEIJERATING AIR MASS FLOW while forming an integral physical part of the Aircraft Structure. By definition, the Unit Dimension of a physical quantity of Acceleration is LT" 2 , i.e. Length x Time Squared, the Units of which are, ms "" * , metres per second, squared. Consequently, no Acceleration can take place between the Engines Mounting Bolts and the Aircraft it is bolted to because there is no Length to the Dimension between them. By definition, the physical quantity of Force, Thrust Force (mass x acceleration), lias the Dimension of MLT" , where M is Mass, L is Length and T is Time squared per second. Here again, there can be no Tlirust, between the engines mounting bolts and the aircraft because, there is no Mass movement between them, nor is there any LENGTH, distance. By definition, Force is defined as that which would produce an Acceleration if acting alone, but, while bolted to the Aircraft the engine Tlirust Force is no longer acting alone because, of the Re-Action Forces acting not, upon the Jet Engine itself but, upon the Aircraft as a whole. The Dimension of a quantity of Force in Dynamics expressed in terms of Mass, Length and Time is given as MLT "2 , where T is per second, per second. The Unit of wliich is the Newton [l Newton = Ikgms ]. When bolted to an aircraft there is no tlirust force between the engine mounting bolts and the aircraft as such in the conventional sense but, letween the engine as an integral part of the aircraft structure and the air mass flow passing tlirough it which, in Motion Physics is a different matter. The Mass Flow is caused to take place between the relatively still Air of the surrounding Atmosphere and the Aircraft itself of which the engine simply happens to le an integral structural part:- the function of wliich happens to be to cause the air mass flow PASSING THROUGH THE AIRCRAFT TO ACCELERATE. A Jet Powered Aircraft in the steady Flight state is a physical mass that is caused to fly, between a horizontally accelerating column of air and the relatively still air of the surrounding atmosphere. Consequently, the Propulsive Effectivenss of the work done upon the intervening AIRCRAFT MASS in its passage tlirough the air will le governed firstly, by the the quality and efficiency of its aerodynamic design and secondly, but, more importantly, by the difference letween the interacting air velocities acting between the two Air Motion Systems and the aircraft at all mission conditions. The first we can do nothing about technology having virtually- reached the limit of its potential, not so, the second. By introducing the means whereby we can control and smoothe the interacting VELOCITY difference between the air mass flow passing tlirough the aircraft, and the surrounding atmosphere tlirough which it is passing, we can dramatically improve its Propulsive Efficiency as a Motion Vehicle tliat is caused to fly. This is what my invention does. My invention comprises an infinitely variable AIR VELOCITY GEARBOX interposed between the AIRCRAFT'S AIR MASS FLOW and the surrounding atmosphere thereby enabling us to dramatically reduce the Engine Range Factor Kr. The question of course is How? VELOCITY by definition is the rate of change of displacement, expressed as either M.P.H, ft. per second or metres, per second, etc. The Dimension of a quantity of VELOCITY is given as LT , i.e., Length x Time, per second. Logically therefore, if we are to control the motion interaction letween the Air Velocities of these two Air Mass Systems, i.e., the continuous air mass flow passing tlirough the aircraft relative to the Velocity state of the surrounding atmosphere, we have only LT to play with. The difficulty is, if we wish to AVOID interfering with the power invested in the system at any specfic condition in a mission, and we do, then, all we have to play with is the dimensions of Velocity itself namely, displacement i.e., Length, or Time itself. Again logically having gone to so much trouble to accomplish the desired Motion State of Velocity, we cannot interfere with the aircraft's displacement "L" unless of course, if by so doing, we can increase it. The question is still, HOW? The answer is simple, the only remaining dimension we can play with is TIME itself. Fortunately, in Dynamics, the dimensions of Quantities in terms of I._ength, Mass and Time and their inter relationship to one another applies to each single MOTIOM FORCE system. In this specific instance we have not one, but two interacting AIR MOTION SYSTEMS to play with. The trouble has teen caused because in terms of the Dimension of TIME, both AIR MOTION SYSTEMS are forced to ACT in Time as one but to RE-ACT physically as two, because their respective physical systems are bolted together. Therefore, we can cheat the limitation of the dimension of TIME which applies to each separate Motion System according to Dynamics by, causing the Re-Action of one Air Motion System Cycle to take place wholly within the other. This is what I have done. T s is wliat I have invented. This will le in effect An Infinitely Variable, Fully Automatic, Air Velocities Gear Box making possible a dramatic improvement in the operational economics of many existing and all future jet powered aircraft so equipped 1. The overall response Velocity of the outside system can be used to govern the overall Velocity of the air mass flow passing tlirough the engine (Motion System No.l). 2. By enclosing the whole of the working cycle of the engine, Motion System Nol, within that of Motion System No2, the Air Velocity Gearbox", we have in effect, enclosed, one within the other, in terms of Motion Time. 3. Such an arrangement makes possible the establisliment of an infinitely variable control over the gearing of the interacting Air Velocities of both Motion Systems relative to the resultant aircraft velocity at all mission conditions. Panel B It is no longer necessary to continue to accept the existing interactive Propulsive Velocity penalty incurred by the uncontrolled interaction of the two Air Motion Systems acting upon a Jet Powered Aircraft as a Motion Vehicle in flight. By causing the controlled integration of both Air Motion Systems No.s 1 and 2 by means of the Air Velocity Gearbox System in the manner described here under described, the existing Aircraft Velocity loss incurred because of the differences letween the Aircraft's Resultant Velocity in flight, the Engine Air Mass Flow Velocities, and those experienced between Air Motion Systems 1 & 2 at the same instant, can be appreciably reduced thereby markedly improving the overall Propulsive Effectiveness of the aircraft as a Motion Vehicle. According to present invention there is provided an additional Aircraft Propulsion System comprising a conventional Jet Engine located within a Ducted System surrounding the engine. The duct leing connected to the Aircraft or its Engine/Engines and supported in such a way as to, permit its axial movement relative to the Engine/Aircraft and so arranged to receive it air mass flow. The duct to le constructed in such a manner as to permit control over the inlet and outlet Air Velocities entering and leaving the Engine relative to that of the Aircraft itself at all forward Flight conditions. The Duct System may also include its own combustion system which may le selectively controlled to work independently of that of the Jet Engine or inconjunction in the manner descriled. In flight and in terms of Motion Law, the Jet Engine is required to te considered as an integral part of the Aircraft structure. Any moverment caused to take place letween the outer Duct Annulus and the inlet to and exit from the .let Engine itself, its mounting pylon or the Airframe direct is merely a question of application. The requirement is met if such movement be caused relative to the Engine as an integral part of the aircraft structure. The System may further comprise an arrangement of inner annulli located between the outer annulus and the inlet to and exit from the Jet Engine which it enlcoses in the manner described. The purpose and function of the inner annulli is to provide the means of smoothing the air flow acting across both Motion Systems and thereby ensure the optimum Velocity matching possible of the respective airstreams at the inlet to and exit from the Secondary Motion System. Some embodiements of the invention will now le descibed, by way of example, with reference to the accompanying drawings, in wliich:- Figure Nol. is a diagramatic longtitudinal section tlirough the first embodiment of a system according to present invention. Figure No2. is a diagrammatic longitudinal section tlirough the second embodiment; Figure No3. is a diagrammitic longtidunal section tlirough a third embodiment; and Figure No . is a diagramatic longtitudinal section througth a fourth embodiment. In each of the embodiments descriled similar parts are indentified by tte same reference numerals in each of the Figures. The propulsion system shown in Figure 1 comprises a Jet Engine 10 which may le a gas turbine jet propulsion engine with or without a by-pass system. The engine 10 is supported from the Aircraft structure by a pylon 11. The engine is located within a Duct 12 which surrounds the Engine and which is mounted on a mounting rail 13 to allow axial movement relative to the engine 10. The duct 12 is moved by an actuating ram or rams 14. Provided with the duct 12 aliead of the engine 10 is an annulus 15 which can le moved axially within the duct 12. An instrument 16 is provided for sensing the total air pressure at the inlet to the engine 10. A second instrument 17 is located at the inlet to the part of the duct 12 surrounding the engine to measure total air pressure of the air mass flow entering this part of the duct 12. A third instrument 18 is located in the duct 12 in close proximity to the final propelling nozzle of the engine 10 to measure air velocity and a fourth instrument 19 is located downstream of the propelling nozzle of the engine 10 to measure jet stream velocity. The annulus 15 direct airflow to the engine 10 and lias an inner diameter/area approximately 80% of that at the inlet to the engine 10. At high forward aircraft velocities the air mass induced into the secondary motion system formed by the duct 12 does not cause any starvation of the air mass required by the engine 10. The annulus 15 can be moved relative to the duct 12 by a ram or rams 20. The engine 10 is provided with a shield 21 to insulate the engine 10 from the heat generated in the duct 12 and to smoothe the air flow around the outside of the engine 10. During take off of the aircraft, the duct 12 will be moved forward relative to the engine 10 as the air mass flow increases and the air flowing tlirough the duct 12 is expanded to do work on the aircraft via the duct 12. Thus duct 12 constitutes a secondary propulsion system seperate from the engine 10. As shown in Figure 2, the duct 12 las a variable inner annulus 22 located between the inlet to the engine 10 and the inlet to the secondary system formed by the duct 12. A variable inner annulus 23 is located letween the exit from the engine 10 and the outer main annulus formed by duct 12. Tlis inner annulus 23 may be mounted on the engine 10 or on the duct 12. These variable annulli 22 and 23 are used to smoothe the air flow when entering and leaving the secondary system. Annulus 23 is also used to provide means of mixing and balancing the velocities of the exit airstreams. In the embodiment of Figure 3, a combustion system 2k is provided in duct 12. The combustion system 2k is operated in parallel wit that of engine 10. The embodiment of Figure k is similar to that of Figure 3 but in this emobodiment the inlet area of the outer annulus formed by the duct 12 is greater than that of the engine inlet. It is envisaged that the ratio of the secondary system inlet area to that of the let engine 10 may be anything from 1:1 upwards. In the event that the secondary system's inlet ratio to that of the inlet to the jet engine 10 is greater than 1:1, then this annulus must have an inner annulus, with an inlet diameter at least equal to the inlet area of the jet engine 10 itself and, must also be free to move axially relative to the outer annulus of duct 12. It is envisaged that these area to area relationships will le set higher for supersonic and hypersonic velocities, for high cruise and long range performance a ratio as high as the existing fan jet engine by-pass ratios would probably prove to te desirable. Here again, the choice is very much that of the designers, depending upon what kind of purposethe aircraft is intended for. An inner annulus is located at the inlet to the secondary system having an inner diameter equal to that of the jet engine/core engine inlet. In each application it is envisaged that the propelling nozzle of -the outer secondary system will be approximately one third of the overall length of the jet engine 10 leyond the exit from the final propelling nozzle of the jet engine 10 itself. This relative distance is very much a matter for the engine/aircraft designers to determine. It is envisaged that this distance must le sufficient to provide time to optimise the velocity mixing between the respective airstreams of each motion system before leaving the final propelling nozzle, of the combined fully intergrated motion systems. The total pressure sensing instruments 16, and 17, located at the inlets to the jet engine 10 and the secondary system of this invention, are required to have a reset pressure bias in favour of the inlet to the jet engine 10. It is assumed that the designers will determine a positive total pressure balance in favour of the jet engine 10 for all aircraft velocities in order to avoid any starvation condition being created at its inlet. However, as the objective is to eliminate the velocity/motion/propulsive force penalty loss, it is envisaged that this bias will be small, but this is very much a matter for the repsective designers and users to decide. Similarily, at the exit from the respective systems, the velocity instruments 18, and 19 will be required to be set at some predetermined air velocity ratios between the secondary airstream exit velocity and that of the jet engine 10. It is conceivable that tlis ratio may be required to le variable at different conditions, it is for this reason that the annulus 23 should be required to be variable. Additionally, the means of variability is required to be such that its downstream to unpstream angle relative to the secondary airstream flow can be altered. This is necessary in order to le able to control the degree of volume mixing to take place letween the respective airstreams in order to establish an optimum exit velocity for both airstreams before leaving the final propelling nozzle of the combined system. It is also envisaged that in the cold running condition of duct 12 tliat the exit airstream energy from the jet engine 10 will be utilised by the judicious mixing of the airstreams from duct 12 and engine 10 by using the variable inner annulus 23 located between the exists from these respective systems thereby ensuring tliat the secondary exit air flow velocity is always greater than both its inlet velocity and tliat of the aircraft at any instant in flight. In the hot running condition it is envisaged tliat both the velocity and pressure of the respective airstreams can le balanced lefore leaving the final propelling nozzle of the combined motion system. It is envisaged tliat it will le desirable in the hot running condition that the outer airstream should be kep)t at a higher relative pressure tlian the inner air stream from engine 10. In tlis way the Higher velocity energy of the inner airstream can be used to accelerate the ourer airstream thereby clianging its prøsitive p>ressure bais to matching velocity of the inner airstream. Here again it is very much a question of designers choice, the op)tion are many and varied. The designers liaving determined the desired relative ratios of velocity and pressure at the inlet to and exit from the respective systems it is envisaged that a simple integrated analogue to digital compκiter control device could be readily designed to optimise the integration letween the respective velocities letween the combined motion systems and thereby the overall p)rop>ulsive efficiency of the aircraft in flight. The designers liaving determined the control criteria, operation is extremely simple. In the cold running condition the aircraft after talce-off will reach a state of choke at the inlet to the jet engine, the event, as soon as the total air pressure begins to rise, this will le sensed by the air pressure measuring devices located at the respective motion systems inlets. As soon as the total piressure begins to exceed the desired pressure ratio for tliat condition of flight, an electrical signal is passed to an electronic control system, which actuates ram control system to move the outer annulus 12 forward relative to the engine 10 thereby allowing the main inlet airstream to spill into the secondary system causing the reversal of the tendancy for the total inlet pressure to rise. The opening letween the inlet to the secondary system and the inlet to the jet negine 10 will te maintained so that the pre programmed total air pressure acceptable at tliat condition is maintained. In p>arallel with this operation, the inner annilli 22 and 23 respectively, will be caused to smoothe the flow of air tlirough the secondary system such that the downstream exit velocity is at all time greater than the inlet velocity to the duct system and higher tlian the relative aircraft velocity at any instant thereby ensuring the elimination of the prevoius interactive ram drag propulsive force loss/aircraft velocity loss. In this instance the higher downstream velocity of the hot airstream leaving the jet engine 10 will be used to induce the required increase in the exit velocity of the secondary system prior to being required to be mixed with the engine jet stream thereby maintaining the required pκ>sitive velocity balance across the secondary system. The inner annulus 23 is located so tliat the exit from the secondary system will also be caused to create the degree of airstream mixing necessary to induce the required velocity balance across the secondary system. This velocity balance across the secondary system will at all times be greater than the aircraft velocity at any instant thereby ensuring the aintaince of the aircraft's momentum state at an op>timum state and overall propϊulsive efficiency. The system will le particularly advantageous in the decent phase because we can trade and focus the aircraft's energy height and total momentum to act upon the air mass flow entering tlie inlet to the Air Velocity Gearbox System. Tle Jet Engine Tlirust required in decent and indeed at all flight conditions will le appreciably less tlian is required to maintain an equivalent velocity/motion state by any exisitng form of jet powered aircraft in p>owered by the same engines. The required pxisitive velocity balance across the secondary duct prropmlsion system will be maintained regardless of whether the aircraft is accelerating or decelerating. Tle only condition necessary to activate the secondary propulsion duct system Is the total inlet pressure ratio limit set at the inlet to the jet engine 10 for all inlet air velocity conditions as determined by the repective engine and aircraft designers of any particular aircraft. Tlie operation for the hot condition (Figures 3 & k) is as for the cold condition excep>t tliat the introduction of heat into the secondary system will involve a rep>rograming of the parameters of velocity and pressure acting across the secondary system. Essentially, this will involve no more tlian a redefining of the relative angle changes to the inner annulus 23 located at the exit to the secondary system to optimise the velocity balance between both airstreams before reaching the final variable propelling nozzle. Tle introduction of heat energy into the secondary system can usefully be brought about as soon as the control choke state is exceeded at the engine inlet. It is envisaged tliat the inital acceleration towards take off velocity will be initiated in the normal way using the pxiwer of the jet engine 10 but the secondary combustion system 24 can be usefully activated at any conventient time or condition, here again, it is very much a matter for the designers of the pjartlcular application. Tlie secondary combustion system 24 will be particularily useful at supersonic and hypersonic speeds. To compensate for the loss of power during take-off at high ISA conditions and or at altitudes above l,000ft, the bias at the inlets to the respective systems should require the inlet gap) letween them to be relatively increased in order to optimise the conversion of the higher airstream temperature to velocity energy. The secondary system is not mechanical and therefore, is less effected by outside ISA temperatures and in fact does not incur as great a propxilsive thrust penalty because of it. Consequently, the piayload penalty incurred because of high ISA temperature conditions prevailing at the peint of take off can le reduced. All volume/density/mass flow loss/penalties, incurred by a jet engine at high ISA conditions will be more tlian fully compensated for by tle compensating velocity increase in tlie air mass flow passing tlirough the secondary motion system the "Air Velocity Gearbox". At high aircraft speeds, above supersonic speed, it would be more benificial to switch the primary roles of both systems, the secondary system being used to produce the primary pxiwer and the engine talcing on the role of the secondary back up) or, compilementary propiulsion system. In this way, the limitations iπ osed upen the thermal efficency of the aircraft as a Motion Vehicle by the pihysical limitations of the systems comprisng it's jet engine/engines, in the primary mode, is appreciably reduced. At speeds above Mach 2 the acceleration respκπise capability of the aircraft using the secondary system as the primary source of p>ower would make px>ssible a tremendous improvement in its acceleration capability. Because the drag coefficient of the Air Velocity Gearbox, as Air Motion System, is considerably less than that of the Jet Engine it serves, its through velocities are more easily achieved, consequently, in combination with the jet engine the altitude performance of the aircraft could le dramatically improved on any mission compared with the aircraft {lowered by conventional engines delivering the same tlirust. The altitude penalty incurred, in terms of motion law, is primarily caused by the reduction in air density and consequently of air mass, this factor plus the fact tliat becuase the motion of the aircraft is primarily mechanically induced the engine tlirough velocities remain relatively unclianged and may well become a limiting factor at speeds above Mach 4. Tlie onset of the decrease in overall tlirust with altitude can te appreciably delayed because it can te compensated for by increasing the air velocity tlirough the Air Velocity Gearbox System. The altitude performance of an aircraft powered by this combined/integrated propiulsion motion system over the equivalent conventional jet engine producing the same power would le, quite dramatic. Tle hot running condition of the secondary system can usefully le initiated at any time, here again it is very much a designers opjtion deciding when and under what conditions tliis should te done. Tlie volicity control in the hot condition is exercised in exactly the same way as for the cold running condition excep)t that the final exit temperature of the airstream leaving the secondary system should not le allowed to exceed tliat of the jet engine when it is in use as the primary source of tlirust because this would cause serious air flow disturbance and therefore thrust/velocity loss to tle combined system and therefore to tle aircraft. Here again, a pre programmed exit temperature relationship) letween both exit airstreams can be chosen by the designers to cover the entire flight envelope of any particular aircraft application. Tle flight envelope of any existing jet pκwered aircraft can te dramatically extended for no increase in tle mechanical efficiences of the systems comprising its jet engine/engines. It is envisaged that the final propelling nozzle of the combined system would also be variable, maling pessible a final influence upen the relative airstream velocities before leaving tlie aircraft's propulsive system. It is further envisaged tliat the whole operation whether in cold running condition, with or without the combustion op)tion, can readily te controlled using a simpile analogue/digital coπputerised control device. Although the secondary combustion system at velocities/speeds of less than Mach 2 will be relativley low pressure nontheless, efficient combustion can still le obtained because of the longer time cycle/spece in which the conversion of the fuel to heat can take p)lace. The operation and control in this mode is exactly the same as for the cold and hot conditions Figures 1,inclusiveof Figures 2 and 3 respectively. The only difference being the additional control requirement to move duct 15 relative to the outer annulus. Duct 15 as explained, stiall te required to function as the inlet to the outer annulus did in the application where the inlet areas to both systems are the same, as in Figure No.l. In this instance some form of ram arrangment, located within the Nacelle forming the inlet to the outer annulus/Nacelle connected to duct 15 the inner annulus located at the enterance to the jet engine, in such a way, that it is caused to move axially relatively to both the outer annulus/Nacelle itself and the inlet to the jet engine. From an economic commercial operation peint of view, it would le desirable to keep) the overall velocitiy differences letween the combined system relative to the Resultant Aircraft Velocity as low as possible. This condition would in turn make possible the attainment of very high aircraft propxilsion/motion systems efficiences. This condition would also require tliat the air mass flow velocity p)assing tlrough the secondary system le as close as p>ossible to the resultant aircraft velocity. In such an application it is envisaged tliat a by p)ass ratio as high as that used in existing fan jet engines could be accomodated to advantage. For fighter or any military applications where a high acceleration with high through air velocity performance is critical, then the by p)ass ratio could be somewhat less, probably somewhere between 2 and 4 to 1, the lower figure for the fighter application and the later for the bomber or long range surveillance aircraft. Panel C Tlie aircraft propulsive system so designed as to fully enclose an installed p)ure jet, bypass, or fan jet type of engine. Ηie system comprises an outer annulus, the inlet diameter to which may le either equal to or greater tlian the inlet diameter of the jet engine tliat the annulus encloses. Tle exit diameter of the annulus may not be less than the exit diameter of the jet engine's final propelling nozzle. Tlie distance between the final propelling nozzle of the jet engine and the system is to be sufficient to ensrue adequate mixing of the respective airstreams in order to optimise the airstream velocities tefore the combined airstreams leave the final propelling nozzle of the combined systems. Tlie outer annulus is required to te connected to the aircraft structure in a manner sufficient to ensure that is can te caused to move axially relative to the jet engine tliat it encloses and the aircraft structure to which the jet engine is secured. Movement of the outer annulus is by means of a suitable mounting and ram arrangement that may either le located on the jet engine or to some suitable point on the airframe depending ipon the particular form of jet engine installation. Tlie outer Annulus/Nacelle includes an arrangement of tlree inner annulli. Tlie first inner annulus, which may le of different inlet diameter/area relative to the inlet to the jet engine it serves, is located in tle inlet to the outer annulus and in front of the inlet to the jet engine. This annulus may le either fixed or free to move axially relative to the outer annulus. Should movement of the inner annulus relative to the outer annulus le required, provision is made for a suitable mounting and ram arrangement located within the front end of the outer annulus. Ηie second inner annulus, which is optional, is to te located letween the inlet to the jet enfine and the inlet to the duct system. This annulus which is to te connected to tlie outer annulus, is so arranged that both its inner area and axial angle to the system's airstream can itself te automatically altered and controlled. Ηie tliird inner annulus, if required, is to te located letween the exit from the final propelling nozzle of the jet engine and the exit from the duct system lefore its airstream mets the jet engine airstream. This annulus may also le fixed or variable in diameter and angular movement relative ot the airstream. This annulus is required to le connected to the outer annulus. Ηie system also includes an inner annulus sliield which surrounds the jet engine from its inlet to its pxrpelling nozzle. Tliis shield smoothes the air flow both around and across the length of the engine and seperates the system air flow from the jet engine and may also le required to serve as a heat sliield between the engine and the systems air flow. This system further comprises an arrangement of air pressure, velocity and temperature instrumentation devices sufficient to establish the state of the system air flow in these terms throughout the systems operation. Tle system further comprises a final propelling nozzle wliich may te either fixed or variable depending upon the particular application. Tle entire system as described in Panels A to B inclusive of C is required to be controlled be a suitable integrated analogue to digital computerised control system designed to control the movement of the outer annulus relative to the aircraft in flight and regulate the air flow tlrough the system utilising the variable inner annulli to control the velocity of the airstream passing tlrough the system and to ensure tliat the exit velocity of the airstream from the combined systems is at all times greater than the inlet air velocity to the Air Velocity Gearbox system and as close to tliat of the resultant aircraft velocity as is necessary to ensure optimum Aircraft Propulsive Efficiency at tliat condition. Panel E Drawings Figures Key & Drawings Figure Nol. 10. Jet Engine 11. Pylon - Connection to aircraft. 12. Duct - Enlosing Jet Engine. 13. Mounting Rail. Ik . Actuating Rams. 15. Annulus Directing Airflow into 10. 16. Instrument for sensing total air pressure at the inlet ro 10. 17. Instrument for sensing total air pressure in Duct 12. 18. Instrument for measuring air velocity in Duct 12. 19. Instrument downstream of 10 to measure engine jet velocity. Figure No2. As Figure Nol Plus. 20. Actuating Rams. 21. Heat Shield. 22. Variable Inner Annulus located letween 10 & 12. 23. Variable Inner Annulus located letween the exit from 10 & propelling nozzle of 12. Figure No.3. As Figures 1 & 2 Plus. 24. Combustion System. Inclusive of an analogue digital control computer. Figure No.4. As Figures 1, 2 & 3 with the inlet area to 12 greater than the inlet area to 10.

Claims: Panel C Claims 1. An additional aircraft propulsion system described as an "Air Velocity Gearbox", suitable for use by many existing and all future jet pewered aircraft that are, px>wered by either, pπαre, bypass, or fan jet type of engines and wliich is required to operate at either subsonic, supersonic, or hypersonic speeds, wherein the engine is enclosed within an adjustible system of ducting that includes an arrangement of variable annulli controlled by compxiter, the whole arrangement of wliich is designed to permit the establishment of control over both the engine inlet velocities and, propelling nozzle air velocities leaving the combined system, relative to, the aircraft velocity at any peint in a flight mission and at all flight conditions 2. An additional aircraft propnilsion as claimed in 1 inclusive of a combustion system tliat is additional to that of the jet engines propelling the aircraft and which may le operated in p rallel with or independanlty from such engines depending up)on whatever the aircraft performance demand requirements may dictate. 3. An additional aircraft propxilsion system such as claimed in claims 1 & 2, wherein, the inlet & exit ducting can te caused to move forwards or backwards relative to the jet engine it surrounds thereby providing the principle means of influencing and hence governing the air pressure and or the air velocity interchange of the air mass flow passing letween the inlet to the jet engine and, the exit from the final propelling nozzle of the combined "Air Velocity Gearbox & Jet Engine System." k . An additional aircraft prop ilsion system as claimed in claims 1, 2 & 3 wherin provision is made for a movable inlet annulus located at the inlet to the "Air Velocity Gearbox" system its axial pesition realative to the engine inlet is serves leing acconplished by means of an arrangement of actuating rams wliich are controlled by compxiter in response an air pressure instrument signal located at the engine inlet, thereby, permitting control to be exercised over the total air pressure entering the engine at all flight conditions. 5. An additional aircraft propxilsion system as claimed in claims 1,2,3 and 4 wherein an arrangement of two variable annulli items No.s 22 & 23 are respectively located in close proximity to the inlet to and exit from the engine and within the "Air Velocity Gearbox" system. Ηie throat areas of these annulli are controlled by compxiter in respxπise to signals received from air pressure instruments (17) and air velocity instruments (18). The angular change of the annulli relative to the ducted airstream causes their respective throat areas to change thereby causing a desired cliange in the airmass flow velocity passing tlirough the combined systems. 6. An additional aircraft propxilsion system as claimed in claims 1,2,3,4 & 5 wherein a heat shield (21) is located around the jet engine forming both a heat sliield in the event of the secondary combustion system leing brought into use, it also provides the means of smootliing the airflow around the jet engine. 7. An additional aircraft propxilsion system as claimed in claims 1,2,3,4,5,6 & 7, where the jet engine is p)ylon mounted to the aircraft and the inlet and exit ducting is caused to move relative to the engine. 8. An additional aircraft propxilsion system as claimed in claims 1,2,3,4,5, & 6 where the jet engine is p)ylon mounted and the inlet and exit ducting is caused to move relative to the aircraft. 9 An additional aircraft propxilsion system as claimed in claims 1,2,3,4,5 & 6 where the combined system inclusive of the "Air Velocity Gearbox System inclusive of the jet engine it serves is located either wholly or pertly within the aircraft fuselage, wings or tail structure. 10. An additional aircraft propxilsion system such as claimed in claim 9 where the inlet and exit ducting of the "Air Velocity Gearbox System",is caused to move relative to the aircraf . 11. An additional aircraft propxilsion system such as claimed in claim 9 where the inlet and exit ducting of the "Air Velocity Grearbox System", is caused to move relative to the engine.

Inventors: CRAIG ALFRED COWIE, CRAIG, ALFRED, COWIE
Applicants: CRAIG ALFRED COWIE, CRAIG, ALFRED, COWIE
Application year: 1990
Main classification: F02K7
IPC codes: F02K3, F02K7