BR112018004262B1 - VEHICLES WITH GAS GENERATOR, FRONT EJECTORS AND AT LEAST ONE FLUIDLY COUPLED REAR EJECTOR - Google Patents

Abstract

VEHICLES WITH GAS GENERATOR, FRONT EJECTORS AND AT LEAST ONE FLUIDLY COUPLED REAR EJECTOR. Vehicles that include a main body and a gas generator that produces a flow of gas. At least one rear conductor and one front conductor are fluid coupled to the generator. First and second front ejectors are fluidly coupled to the at least one front driver. At least one rear ejector is fluidly coupled to the at least one rear conductor. The front ejectors respectively include an outlet structure through which gas from the at least one front conductor flows. The at least one front ejector includes an outlet structure through which gas from the at least one rear conduit flows. The elements of the first and second primary airfoils have guiding edges respectively located directly downstream of the first and second ejectors. At least one member of the secondary airfoil has a guiding edge located directly downstream of the exit structure of the at least one rear ejector.

Description

COPYRIGHT NOTICE

[0001] This disclosure is protected by United States and international copyright laws. © 2016 Jetoptera. All rights reserved. Part of the disclosure of this patent document contains material subject to copyright protection. The copyright owner has no objection to facsimile reproduction by any person of the patent document or patent disclosure as it appears in the patent file or register of the Patent and Trademark Office, but otherwise reserves all rights copyright.

PRIORITY CLAIM

[0002] This patent application claims priority to United States Provisional Patent Application No. 62/213,465, filed September 2, 2015, the entire disclosure of which is incorporated by reference as if it were fully included in the present patent application .

BACKGROUND OF THE INVENTION

[0003] Aircraft that can hover, take off and land vertically are commonly called vertical takeoff and landing (VTOL) aircraft. This classification includes fixed-wing aircraft as well as helicopters and aircraft with tiltable motorized rotors. Some VTOL aircraft can also operate in other modes, such as short runway takeoff and landing (STOL). VTOL is a subset of V/STOL (vertical and/or short runway takeoff and landing).

[0004] For illustrative purposes, an example of a current aircraft that has VTOL capability is the F-35 Lightning. Conventional methods of vectoring the vertical flow of lift air include the use of nozzles that can be rotated in a selected direction, along with the use of two sets of flat flap blades arranged at 90 degrees to each other and located on the outer nozzle. The F-35 Lightning's propulsion system similarly provides vertical lift force through the use of a combination of turbine engine vector thrust and a vertically oriented lift propeller. The support propeller is located behind the cockpit in a compartment with upper and lower clamshell doors. The engine emits exhaust gases through a three-bearing rotating nozzle that can divert thrust from horizontal to just ahead of vertical. Roll control ducts extend outward on each wing and receive their thrust from air from the engine propeller. Pitch control is affected by the division of thrust between the support/engine propeller. Yaw control occurs through the yaw movement of the engine's rotating nozzle. Rolling control is provided by differentially opening and closing the nozzles at the ends of the two rolling control ducts. The support propeller has a "D" shaped telescoping nozzle to provide thrust deflection in both the forward and aft directions. The D-nozzle has fixed blades at the exit opening.

[0005] The design of an aircraft or drone generally consists of its propellant elements and the structure into which these elements are integrated. Conventionally, the propulsive device in aircraft may be a turbojet, turbofan, turboprop or turboshaft, piston engine or electric motor equipped with a propeller. The propulsion system (propeller) in small unmanned aerial vehicles (UAVs) is conventionally a piston engine or an electric motor that supplies power through a shaft to one or more propellers. The propellant for a larger aircraft, manned or unmanned, is traditionally a jet engine or a turboprop. The propellant is generally attached to the fuselage or the body or wings of the aircraft through aerodynamic supports or supports capable of transmitting force to the aircraft and sustaining the loads. The emerging mixed jet (jet stream) of air and gases is what propels the aircraft in the opposite direction to the jet's effluent flow.

[0006] Conventionally, the effluent airflow from a large propeller is not used for lift purposes in level flight and therefore a significant amount of kinetic energy is not utilized to the benefit of the aircraft unless it is rotated as in some of the applications that exist today (namely the Osprey Bell Boeing V-22). Instead, lift in most existing aircraft is created by the wings and tail. Furthermore, even in these specific VTOL uses (e.g., takeoff to transition to level flight) found on the Osprey, the lift produced by the propeller itself is minimal during level flight, and most of the lift force is not nevertheless, provided by the wings.

[0007] The current state of the art for creating lift in aircraft is to generate a high-speed airflow over the wings and wing elements, which are generally airfoils. Aerodynamic profiles are characterized by chord lines extending mainly in the axial direction, from the leading edge to the trailing edge of the aerodynamic profile. Based on the angle of attack formed between the incident airflow and the chord line, and in accordance with the principles of aerodynamic profile lift generation, the low pressure air flows over the suction (upper) side and vice versa, by Bernoulli's law, moving at higher speeds than the lower side (pressure side). The lower the speed of the plane, the lower the lift force and the greater the wing surface area, or higher angles of incidence, will be required, including for takeoff.

[0008] Large unmanned aerial vehicles (UAV) are no exception to this rule. Lift is generated by designing the aerodynamic profile with the appropriate angle of attack, chords, wingspan and camber. Flapes, grooves and many other devices are other conventional tools used to maximize lift by increasing the lift coefficient and wing surface area, but will generate lift corresponding to the airspeed relative to the aircraft. (Increasing the area (S) and lift coefficient (CL) allows a similar amount of lift to be generated at a lower airspeed (V0) according to the formula, but at the cost of greater drag and weight.) These techniques current ones also perform poorly with a significant drop in efficiency in strong crosswind conditions.

[0009] While smaller UAVs use thrust generated by propellers to lift the vehicle, current technology relies solely on controlling electric motor speeds, and the smaller UAV may or may not have the ability to pivot the motors to generate thrust and lift , or transition to level flight by tilting the propellers. Additionally, smaller UAVs using these propulsion elements suffer from inefficiencies related to batteries, power density, and large propellers, which may be efficient when hovering, but are inefficient in level flight and create difficulties and danger when operating due to the rapid tilting of the blades. Most current quadcopters and other electrically powered aerial vehicles are only capable of very short flight periods and cannot lift or carry large payloads, as the weight of the electrical system and battery can already exceed 70% of the vehicle's weight in every moment of the flight. A similar vehicle using jet fuel or any other hydrocarbon fuel typically used in transportation will carry more usable fuel by at least an order of magnitude. This can be explained by the much higher energy density of hydrocarbon fuel compared to battery systems (by at least an order of magnitude), as well as the lower weight to total vehicle weight ratio of a hydrocarbon fuel-based system. .

[0010] Consequently, there is a need for greater efficiency, improved capabilities and other technological advances in aircraft, particularly for UAVs and certain manned aerial vehicles.

BRIEF DESCRIPTION OF THE DRAWINGS

[0011] FIGS. 1A - 1C illustrate some of the differences in structure, forces and controls between a conventional electric quadcopter and an embodiment of the present invention.

[0012] FIG. 2A is a top view of a conventional wing and airplane structure; FIG. 2B is a front view of a conventional wing and airplane structure;

[0013] FIG. 3 is a cross section of an embodiment of the present invention depicting the upper half of an ejector and velocity and temperature profiles within the internal flow;

[0014] FIG. 4 is an embodiment of the present invention that represents a propeller/ejector placed in front of an airfoil;

[0015] FIG. 5 is another embodiment of the present invention in which the thruster/ejector is placed in front of a control surface as part of the airfoil of another wing.

[0016] FIGS. 6A-6C illustrate the present invention shown in FIG. 5 from different points of view.

[0017] FIG. 7A is another embodiment of the present invention that uses a jet effluent and airfoil towards it to push the aircraft forward and generate lift, which replaces the engine in the wing.

[0018] FIG. 7B is the front view of the present invention shown in FIG. 7A.

[0019] FIG. 7C is another embodiment of the present invention that features tandem wings;

[0020] FIG. 8A is a side view of another embodiment of the present invention, with the tandem thrust/lift generation system in which the front thrust increasing ejectors produce thrust with a wing in canard configuration and the rear thrust increasing ejectors produce thrust and support in the rear region.

[0021] FIG. 8B is the perspective view of the present invention shown in FIG. 8A.

[0022] FIG. 9 is a perspective view of the present invention shown in FIGS. 8A and 8B and presents the aircraft tail arrangements and gas generator assembly.

[0023] FIGS. 10A - 10E show variations in the lift coefficient at a constant speed of an airfoil, as a function of the angle of incidence and shows the stall angle of attack.

[0024] FIGS. 11A - 11B show the improvement in stall margin with different positions of the present invention.

[0025] FIGS. 12A - 12C are yet another embodiment of the present invention that presents the propellant ejector component in a position relative to the wing.

[0026] FIGS. 13A - 13C illustrate how the present invention can control the pitch, roll and yaw of the aircraft using the ejectors, which increase thrust, in conjunction with the thin airfoils placed in the wake of the ejectors.

[0027] FIG. 14 is an embodiment of the present invention with flap-shaped elements on the walls of the diffuser of a Coanda ejector that is divided into two parts.

[0028] FIGS. 15A - 15C illustrate, from different points of view, the 3D characteristics of an embodiment of the invention.

[0029] FIG. 16A shows another embodiment of the present invention for improving performance and stall margin.

[0030] FIGS. 16B - 16D illustrate, from different points of view, the present invention shown in FIG. 16A.

[0031] FIGS. 17A - 17C illustrate yet another embodiment of the present invention.

[0032] FIGS. 18A - 18C show typical conventional arrangements for Coanda-type ejectors.

[0033] FIG. 18D is an embodiment of the present invention depicting the circular Coanda ejector with simple primary nozzle elements.

[0034] FIGS. 19A - 19D represent different embodiments of the present invention with better performing primary nozzles.

[0035] FIG. 19E illustrates flow over the delta wing obstruction placed in the center of the interior of the primary nozzle.

[0036] FIG. 20 explains the thermodynamics of an embodiment of the present invention.

[0037] FIG. 21 is yet another embodiment of the present invention with features that improve flow separation delay.

[0038] FIGS. 22A to 22F illustrate different 3D features and embodiments of the present invention.

[0039] FIG. 23 illustrates some features according to an embodiment of the present invention.

[0040] FIG. 24 demonstrates a Coanda-type ejector applied to a VTOL-only aircraft.

[0041] FIG. 25 shows an alternative ejector arrangement as another embodiment of the present invention.

[0042] FIG. 26A shows a high contour turbofan.

[0043] FIG. 26B shows a modified turbofan to serve as gas generator as one embodiment of the present invention.

[0044] FIG. 27A is an embodiment of the present invention that shows the bleed network and ducts.

[0045] FIG. 27B is another embodiment of the bleed network and ducts.

[0046] FIG. 27C is another embodiment of a bleed network and ducts that shows the controller and sensors.

[0047] FIG. 27D is yet another embodiment of a bleed and duct network that shows the identified controller and sensors.

[0048] FIGS. 28A-28E are possible forms of propellants of the present invention.

[0049] FIG. 29 is a possible arrangement of the propulsion system in takeoff or floating in an embodiment of the present invention.

[0050] FIGS. 30A - 30B illustrate the thermodynamic cycles of a jet engine.

[0051] FIG. 31 is an embodiment of the present invention.

DETAILED DESCRIPTION

[0052] This patent application seeks to describe one or more embodiments of the present invention. It should be understood that the use of absolute terms such as "shall", "may" and the like, as well as specific quantities, should be interpreted as applying to one or more of these embodiments, but not necessarily to all of these embodiments. As such, embodiments of the invention may omit or include modification of one or more features or functionalities described in the context of such absolute terms. Furthermore, the directions in this patent application are for reference purposes only and shall in no way affect the meaning or interpretation of the present invention.

[0053] The present inventions disclosed in this patent application, whether working independently or together, allow a UAV to perform the maneuvers of an electric UAV without the use of large propellers or fans, while maximizing autonomy, range and vehicle payload to total weight ratio. Electric UAVs such as quadcopter can hover, take off vertically and land and as such perform loops, etc., by simply controlling the rotational speed of the propellers attached to it. The present invention eliminates the need for large propellers or fans and replaces the logic of controlling the rotational speed of the propellers with mainly fluidic control of the rotating thrust-increasing ejectors fed with motive fluid from a gas generator on board the vehicle. Non-electric UAVs that employ jet engines typically do not operate at low speeds or efficiently and are limited in their maneuverability when compared to electric UAVs. FIGS. 1A to 1C illustrate some of the differences in structure, forces and rotational speeds between a conventional electric quadcopter and a fluidic quadcopter, one of the embodiments of the present invention.

[0054] The present invention presents several elements that significantly increase the maneuverability of a non-electric UAV. For example, one embodiment of the present invention discloses a new propulsion device (thrust) that can be deployed on an aircraft. Another embodiment describes new 3D elements implemented in ejectors as part of the thruster. Yet another embodiment discloses a tandem system that combines a thrust generator (propeller) and a thin airfoil wing (lift element) that can be deployed on an aircraft. Yet another embodiment discloses a specific tandem system consisting of ejector nozzles and thin airfoil placed in the wake of the nozzle and uses jet effluent from the nozzle to generate thrust and lift. Another embodiment reveals a new placement of an ejector on a wing to allow a high angle in incidence flight. Another embodiment discloses the application of a thermodynamic cycle of the propulsion system with optionally advantageous characteristics that increase efficiency and reduce the total weight of the propulsion system. Finally, another embodiment describes a thrust generation system that combines VTOL capabilities with turbo engines and pitch, roll and yaw control of an aerial vehicle. Each of the aforementioned embodiments and many other embodiments of the present inventions disclosed in this patent application will be explained in the following sections.

Propulsion Device and Thrust System.

[0055] FIGS. 2A and 2B describe a conventional airplane with wing-mounted engines that produce thrust, which generates acceleration and speed of the aircraft, resulting in the generation of lift on the wings; the function of the engine is to create thrust and the effluent from the engine jet is not used for further generation of lift, but is lost to ambient temperature. The jet effluent has a greater speed than that of the aircraft and as such the lift generated by the wing is a function of the aircraft speed and not the local speed of the engine jet effluent, which is the subject of the patent application. current.

[0056] One embodiment of the present invention includes a thruster that uses fluidics to entrain and accelerate ambient air and provides a high-velocity jet effluent mixture of high-pressure gas (fed to the thruster from a gas generator) and the ambient air drawn in in an engineered manner directly towards the airfoil placed exactly behind the propellant, in the wake of the propellant jet, and in a symmetrical or non-symmetrical manner.

[0057] FIG. 3 illustrates a cross-section of just the upper half of the ejector 200. The chamber 211 is supplied with air hotter than the environment. The pressurized motive gas flow 600, which may be produced by, as a non-limiting example, a fossil fuel engine or an electrically driven compressor, communicates with the internal side of the ejector through ducts with primary nozzles 203. The primary nozzles 203, due to their design, accelerate the motive fluid 600 to the speed required by the ejector performance. The primary (motive) fluid 600 emerges at high speed over the surface of Coanda 204 as a wall jet, entraining ambient air 1, which may be at rest or approaching the ejector at a non-zero velocity from the left of the figure. The mixture of flow 600 and ambient air 1 moves purely axially in the throat section 225 of the ejector. By diffusion in the diffuser 210, the mixing and smoothing process continues, so that the temperature profiles (750) and the velocity in the axial direction (700) no longer present high and low values, as occurs in the throat section 225, but become more uniform at the ejector exit. As the mixture of 1 and 600 approaches the exit plane, the temperature and velocity profiles are nearly uniform; in particular, the temperature of the mixture is low enough to be directed toward an airfoil such as a wing or control surface.

[0058] In FIG. 4, another embodiment of the present invention is illustrated, with the propeller/ejector 200 placed in front of the airfoil 100 and generating a lifting force 400. The local flow over the airfoil 100 is at a speed greater than the speed of the aircraft, due to the greater speed of 300 of the exit jet effluent of the propellant 200 compared to the air speed of the aircraft 500. The propellant vigorously mixes the hotter engine stream provided by the gas generator with the inlet stream of cold ambient air at a high rate of drag. The mixture is homogeneous enough to reduce the hot motive stream 600 from the ejector temperature to the mixing temperature profile 750 which will not impact the airfoils 100 or 150 mechanically or structurally. The velocity profile of the effluent jet exiting the propellant 200 is such that it will allow more lift 400 to be generated by the airfoil 100 due to the higher local speeds. Additional control surfaces may be implemented on the airfoil 100, such as the stabilizer surface 150 shown in the figure. By changing the angle of such surfaces 150, the attitude of the aircraft can be quickly changed with little effort given the greater local velocity of the effluent jet 300.

[0059] FIG. 5 illustrates that the thruster/ejector 200 may also be placed in front of a control surface 152 as part of another wing airfoil 101. The thruster may be non-axis-symmetric in shape, and the control surface may be placed exactly on the wake. of said thruster 200. The thruster vigorously mixes the hotter driving stream provided by the gas generator, with the incoming cold ambient air stream having a high drag rate. Likewise, the mixture is homogeneous enough to reduce the hot driving current 600 from the ejector temperature to a mixture temperature profile that will not mechanically or structurally impact the control surface. In this embodiment, yaw can be controlled by changing the orientation of the control surface 152. The main function of the thruster 200 is to generate thrust, but also to control lift or attitude. In this embodiment, the yaw control is in orientation 151 creating a rotation about axis 10 of the aircraft.

[0060] FIGS. 6A to 6C show the illustration in FIG. 5 from different points of view.

[0061] For example, the emerging jet with a rectangular pattern due to the propeller's rectangular exhaust plane can also be oriented much more easily and in more directions than by a propeller and electric motor. In another example, an emerging jet with a rectangular pattern due to the propeller's rectangular exhaust plane is directed to the leading edge of a short wing placed some distance behind the propellant to maximize the lift benefit. As described in the present invention, the propellant can therefore generate the thrust necessary for the aircraft to move forward, mainly in the direction opposite to the direction of the jet effluent. Furthermore, jet effluent moving at a speed higher than the speed of the aircraft and resulting from said propellant or ejector will be used to increase the lift force resulting from its flow over the airfoil placed behind said propellant or ejector. . The jet speed must always exceed the aircraft speed and the difference between the two speeds will need to be minimal to maximize propulsive efficiency. It follows that the greater the flow mass providing thrust at lower speeds, even greater than the speed of the aircraft, the greater the propulsive efficiency. For example, using a propellant efficiency equation known to those familiar with the art: PE = 270/(7 + 70) where V is the speed of the propellant exit jet and V0 is the speed of the aircraft, if the speed of the jet propulsion is 150% of the aircraft speed, the airplane speed will be 50% of the propellant's emergent speed, and the propulsion efficiency will be 80%. After leaving the exhaust section of an aircraft's propellant, the exhaust stream from most conventional jet aircraft is lost to the environment and no benefit is derived from the residual jet, although the jet from e.g. still have energy on the treadmill. The exhaust flow is typically a round jet at higher speeds (and therefore energy), mixing with a parallel flow at lower speed and ultimately mixing with the aircraft's pair of trailing vortices. Once it leaves the aircraft engine as exhaust, the jet stream no longer benefits the aircraft and the higher the speed of the exhaust jet, the lower the propulsive efficiency and waste of energy to the environment.

[0062] An embodiment of the present invention utilizes the mixed current emerging from the thruster of the present invention, which would otherwise be lost to the environment in conventional aircraft, to generate lift or create direction change capability by directing it directly at a thin airfoil or other surface placed directly behind this propeller, to generate lift or changes in the aircraft's attitude. Since the pressurized gas supply can also be modulated or used in a segmented manner through a network contained in the aircraft's fuselage and wings, the drag and speed of the effluent jet can be adjusted through primary or secondary methods. The main method refers to the modulation of pressure, flow, temperature and/or segmentation (multiple feeds to multiple propellants distributed throughout the aircraft). The concept of segmentation involves the use of multiple propellant elements conveniently placed throughout the aircraft, that is, segmenting the function of a single, large propellant into multiple smaller ones that are fed with pressurized gas through a network of ducts. The secondary method may involve changing the geometry or position of the thruster relative to the neutral position of that thruster. For example, in horizontal flight, providing adequate gas pressure and flow to the propellant can result in jet effluent at 125% of the aircraft speed. In the case of a jet effluent axial velocity 125% greater than the aircraft speed, the propulsive efficiency becomes 88%. If the emergent velocity becomes 110% at higher speeds with the same level of thrust generated through ambient air drag, the propulsive efficiency improves to 95%.

Thrust and Lift Generator

[0063] Another embodiment of the present invention generally relates to a combination of thrust and lift obtained through a tandem system composed of a thrust generating element that directs the majority of a non-circular high-velocity jet effluent in the axial direction to a thin airfoil located downstream of the jet effluent. The local high axial velocity of this jet effluent generates lift at considerably higher levels than regular wing lift at aircraft speed because ~(Jet flow velocity)2. The jet effluent is a mixture of hot, high-energy gases supplied to the thrust generator through high-pressure gas generator outlet ducts and entrained ambient air. The entrained air is brought into a high energy flow of kinetic energy through a momentum transfer from the high pressure gases supplied to the thrust generator within the thrust generating element. The resulting mixture of air and gas emerges from the thrust generator and can be directed by pointing primarily in the axial direction, downward, toward the thin leading edge of the airfoil and/or the pressure side of the airfoil.

[0064] In most conventional aircraft, it is not currently possible to direct jet effluent to an airfoil or wing in order to utilize its lost energy. In the case of turbojets, the high temperature of the jet effluent actually precludes its use for generating lift through an aerodynamic profile. Typical jet exhaust temperatures are 1000 degrees Celsius and sometimes higher when afterburner is used to increase thrust, as in most military aircraft. When turbofans are used, despite the use of high bypass in modern aircraft, there is still a significant residual element of non-axial direction, due to fan rotation, despite the vanes that direct exhaust fluids from the fan and core, in generally, axially. The presence of the hot core gases at very high temperatures and the residual rotational motion of the emerging mixture, in addition to the cylindrical nature of the jets in the downdraft, make the use of airfoils directly placed behind the turbofan engine impractical. Furthermore, the mixing length of hot and cold streams from jet engines, such as turbofans, is in miles, not inches. On the other hand, the current use of larger turboprops generates large cylindrical downdraft flows, the size of the propeller diameter, with greater speed of the rotational component behind the propeller and moving large amounts of air at lower speeds. The rotational component makes it difficult to utilize the downstream kinetic energy for purposes other than propulsion, and therefore some of the kinetic energy is lost and not used efficiently. Some of the air moved by the large propellers is also directed to the engine core. In summary, the jet stream of current propulsion systems has residual energy and currently untapped potential.

[0065] In this embodiment of the present invention, the current can be used as a lift-generating flow if directed directly to a thin planar profile for lift generation. For example, when the axial speed of the jet effluent is 125% greater than the speed of the airplane, the portion of the wing that receives the jet effluent flow can generate lift greater than 50% for the same wingspan compared to the case in which the air, at the speed of the aircraft, sweeps across the entire length of the wings. Using this example, if the jet's effluent speed is increased to 150%, lift increases more than 45% than with the original wing at aircraft speed, including a density drop effect if the pressurized exhaust gas from a turbine were, for example, used.

[0066] Alternatively, a wing, such as a thin lightweight wing, could be deployed directly behind the propellant ejector exit plane immediately after the vehicle has completed takeoff maneuvers and is transitioning to level flight. , helping to generate more lift with less engine energy.

[0067] Alternatively, using this embodiment of the present invention, the wing does not need to have as long a wingspan, and for the same chord, the wingspan can be reduced by more than 40% to generate the same lift. In this lift equation known to those familiar with the technique: where S is the wing surface, p is the density, V is the speed of the aircraft (wing) and CL is the lift coefficient. A UAV with a wingspan of, for example, 3 m can reduce the wingspan to just 1.8 m as long as the jet is oriented directly towards the wing at all times during level flight, with a thin, tethered wing, bulging and CL similar to the original wing. The detrimental impact of temperature on density is much smaller if the mixing ratio (or drag index) is large and therefore the jet is only slightly hotter.

[0068] FIG. 7A describes an alternative approach to having the jet engine placed on the wing and independently producing thrust. In FIG. 7A, the jet engine no longer produces a jet effluent that pushes the aircraft forward, but is instead used as a gas generator and produces a motive air stream that powers a series of ejectors built into the wing to forward propulsion. In this embodiment, the gas generator (not shown) is fitted to the fuselage of the aircraft, and the green portion represents the inlet, the gas generator and the ducts leading to the red ejectors, which are flat, similar in shape to flaps or ailerons. , and can be activated to control the aircraft's attitude in addition to providing the necessary thrust. FIG. 7A shows yet another wing (secondary) that is placed in tandem with the first (main) wing, which contains the thrust increase ejectors, and just behind said ejectors. The secondary wing therefore receives a speed much greater than the speed of the aircraft, and as such creates a high lift force, since this force is proportional to the airspeed. In this embodiment of the present invention, the secondary wing will see a moderately higher temperature due to the mixing of the motive fluid produced by the gas generator (also known as the primary fluid) and the secondary fluid, which is ambient air, entrained by the motive fluid at a rate between 5-25 parts of secondary fluid for each part of primary fluid. As such, the temperature that the secondary wing sees is slightly higher than ambient temperature, but significantly lower than the motive fluid, allowing the secondary wing materials to support and support lift loads, according to the formula : on what is the final temperature of the fluid mixture of the jet effluent leaving the ejector, ER is the entrainment rate of the parts of the ambient air entrained by the driving air, is the hottest temperature of the driving or primary fluid, and is the approaching ambient air temperature.

[0068] FIG. 7B represents the front view of the aircraft shown in FIG. 7A with arrows illustrating the additional lift force generated by the shorter, tandem wings and the lack of wing engines.

[0069] FIG. 7C represents another embodiment of the present invention with tandem wings. In this embodiment, the thrust increasing ejectors 701 that are part of the thruster system are placed on the main wings (front wings) 703 connected through ducts and receive the motive fluid from a gas generator placed inside the fuselage. The ejector generates thrust and transmits force mechanically to the aircraft. The jet effluent generates a constant high-velocity flow that is used by the secondary wing (gray wing) 702 to produce additional lift. The combination of the two shorter wings produces more lift than that of a wing of much greater wingspan, without the ejector thrust amplifiers that rely on a jet engine attached to said larger wing to produce thrust.

[0070] FIGS. 8A and 8B illustrate yet another embodiment of the present invention. As shown in FIGS. 8A and 8B, the tandem thrust/lift generation system is connected to an aerial vehicle 804, where the front thrust augmentation ejectors 801, which include leading edges and inlet portions for upstream air intake, produce thrust just behind a wing in canard configuration, with each of these ejectors positioned on the starboard and port side of the vehicle. The canard wing is oriented at a high angle of incidence and close to stall when in level flight, wherein the presence of the thrust-increasing ejector extends the stall margin of said canard wing 803. The thrust-increasing ejector 801 mechanically transmits the thrust force to structure 804 and produces a downstream jet effluent consisting of well-mixed primary and secondary air streams, which in turn are used to generate significantly greater lift on wing 802. The system is also replicated in the tail of the aircraft in a similar way. The thrust boost ejectors 801 receive a compressor bleed stream from the gas generator 800, while the tail thrust boost ejectors receive the pressurized hot gases exiting the gas turbine from the gas generator 800. The combination The use of compressor bleed air for the 801 ejectors and the use of hot exhaust gases for the rear ejectors as primary fluids, respectively, result in (1) increased thrust in level flight due to the entrainment of ambient air by the ejectors and (2) additional lift generated on surfaces placed behind said ejectors, such as wing 802 with leading edges. These elements placed behind the ejector are generally thin structures and can be constructed from composite materials, including but not limited to ceramic matrix composites (CMCs). This arrangement provides greater flexibility to switch during the transition from takeoff to hover and level flight.

[0071] FIG. 9 provides more details of the back (or hot) section of the illustration in FIGS. 8A and 8B. The slender structures 904 having leading edges are placed following a set of hot thrust increasing ejectors 901 having leading edges and receiving the primary (motive) fluid as hot exhaust gas from the gas generator 800 located near the cockpit 805 and which entrains air into the inlet 902 of the ejector 901. The duct connecting the gas generator exhaust 800 to the element 901 is incorporated into the structure of the vertical vane 950. The ejectors 901 entrain the incoming ambient air into the inlet areas 902 and expel an entrained mixture of air and motive gas at high speed through the outlet 903 and mainly to the thin tail structure 904 which, in turn, generates additional lift. Both elements 801 in FIGS. 8A and 8B and 901 in FIG. 9 can rotate around its main axis for VTOL and hover control. Furthermore, each ejector of the ejector assembly 901 may rotate about the same axis with and/or independently of the other ejector.

[0072] FIGS. 10A to 10E show the various flows as well as the lift in relation to the angle of incidence with the point corresponding to the angle of incidence highlighted in each case. As the angle of incidence of an airfoil is increased, lift increases until boundary layer separation in the airfoil determines stall, just past the point of maximum lift (see Figure 10D).

[0073] FIG. 10A demonstrates the relationship between lift and angle of incidence of the canard wing structure 803 illustrated in FIGS. 8A and 8B at an angle of incidence of zero degrees (0°), where the point represents the lift force and the streamlines represent the flows around the canard airfoil. FIGS. 10B to 10D show the result of increasing the lift force of structure 803 as the angle of incidence or angle of attack increases to the stall point, fully represented in FIG. 10D. In addition to the position of the airfoil (relative to the angle of incidence) as shown in FIG. 10D, e.g., in the position depicted in FIG. 10E, lift decreases rapidly as the flow becomes turbulent, may separate, and the aerodynamic lines are no longer smooth. Lift increases almost linearly as the angle of incidence increases, but at the angle of incidence shown in FIG. 10D reaches the maximum value, beyond which the flow separates on the upper side of the airfoil. In FIG. 10E, there is recirculation and increased drag, loss of lift generated by opposing flows and separation of boundary layers. This causes the lift force to drop significantly and results in a stall.

[0074] FIGS. 11A and 11B show the characteristic lift curve of FIGs. 10A to 10E, with a second curve showing an extension of the stall margin, which demonstrates the improvement in lift relative to the angle of incidence beyond the stall point if the ejector is placed relative to the wing so that it delay separation and facilitate boundary layer ingestion at high angles of attack. In FIG. 11B, the lift continues to increase without stopping with the angle of incidence, due to the presence of the ejector. Placing the ejector beyond the apex of the airfoil allows reattachment or avoidance of flow separation from the upper boundary layer that would otherwise separate in the absence of the ejector ingesting said boundary layer, due to the high angle of incidence of said boundary layer. airfoil. The ejector is introducing a local area of low pressure at its inlet, forcing ingestion of the boundary layer developed over the upper side of the airfoil. The stall margin is made much greater by placing the thrust-increasing ejector beyond the apex of the canard wing or airfoil 803 of FIGs. 7C, 8A and 8B. These results indicate that the presence of the ejector extends the stall margin and allows the generation of greater lift forces by increasing the angle of attack beyond the value of the stall angle of attack of said airfoil without the presence of the ejector. Furthermore, FIGS. 11A and 11B illustrate possible placements of the ejector relative to the airfoil chord to re-optimize flow around the airfoil.

[0075] FIGS. 12A to 12C represent yet another embodiment of the present invention. The main wing and thrust augmentation ejector system produces forward thrust and high-speed jet effluent conditioning that can be used to generate additional lift when coupled with a secondary airfoil (not shown, but can be placed in the wake or effluent downstream of the ejectors). As illustrated in FIG. 12A, the ejectors are formed by two (2) air knife halves that together generate drag, transfer momentum and acceleration from the ambient air through the use of the primary or motive fluid and the ejection of the final fluid mixture primary and secondary at high speeds. The two halves 1201 and 1202 can rotate and translate independently to position themselves relative to the wing in such a way that they are optimizing boost at any given time, based on the aircraft's attitude and mission (or point in the mission), condition of the primary fluid (flow rate, pressure and temperature). This allows the throat formed by the two halves to have, in one case, a certain value, and in another case, a greater or lesser value. For example, on takeoff, the two halves may both point downward to allow the aircraft to take off vertically. The two halves can move independently and position themselves relative to each other to maximize thrust with a maximum primary fluid flow rate and a maximum drag rate, generating a certain inlet-to-throat area ratio that is conducive to maximizing thrust. impulse. However, when in level flight, the two halves of the ejector can be horizontal and streamlined with the wing, with a smaller throat area for lower pressures, temperature and primary fluid flow rate, again maximizing the increase in thrust. The throat area, exit area, inlet area and their proportions can also be adjusted according to a thrust-maximizing algorithm. Both 1201 and 1202 contain a chamber 1211 and 1212, respectively, connected to a duct and receiving said primary fluid from, for example, a compressor bleed port of a gas generator. The two halves together form a variable inlet area 1201a and a variable outlet area 1201b and a diffusion form formed by walls 1213 and 1214, respectively, to optimally diffuse the flow and maximize said thrust. The primary flow is introduced from chambers 1211 and 1212, respectively, into the throat area through a number of specially designed nozzles 1203 and 1204, respectively, in a continuous or pulsed manner.

[0076] FIG. 12C further describes the arrangement of this ejector for level flight of an aircraft. FIG. 12C shows that the flat ejector can be inserted within the thickness of the airfoil when all elements described in this disclosure are used for greater efficiency. FIG. 12C shows the outline of said inner and outer surfaces of the ejector and FIG. 12B shows the 3D model of the lower and upper halves 1201 and 1202 of the revealed flat Coanda ejector, integrated with the wing. The two halves, which can be driven independently, together form an input 1201a and an output 1201b; they allow the high-velocity introduction of a primary fluid through the primary nozzles 1203 onto the Coanda surfaces 1204.

[0077] FIGS. 13A to 13C describe how the present invention can control the pitch, roll and yaw of the aircraft using the thrust increasing ejectors in conjunction with the thin airfoils placed in the wake of the ejectors. As far as pitching is concerned, the cold and hot ejectors can be rotated independently around their main axis to cause the aircraft to pitch forward or backward. Pitch control is affected by the front/rear split of ejector thrust and/or by modulating the flow of motive fluid supplied to the ejectors. Regarding rolling, the ejectors can be rotated independently so that the aircraft rolls. With regard to yaw, a combination of additional rotation about a perpendicular axis with the positioning of the thin airfoils in the wake of the jet effluent can be used to cause a change in the aircraft's attitude. This embodiment of the present invention makes these maneuvers possible with the use of special joints that can rotate, transmit loads and allow the passage of the primary fluid to said ejectors.

Coanda Device

[0078] In yet another embodiment of the present invention, the thruster and/or thrust generator of the tandem system has the ability to entrain large quantities of air and accelerate it to the jet's effluent speed. This is achieved by employing a Coanda device. These flow enhancing devices have been generally described by different publications which will be discussed in greater detail below. For example, in his article "Theoretical Remarks on Thrust Augmentation", (Reissner Anniversary Volume, Contributions to Applied Mechanics, 1949, pp. 461-468), von Karman describes in detail why a Coanda device results in significantly greater elevation of thrust through multiple jets. Similarly, US Patent No. 3,795,367 (Mocarski) describes a device for entraining air with high boost ratios greater than 1.8, while US Patent No. 4,448,354 (Reznick) applies a linear Coanda device to the VTOL capability of a jet engine. In these aforementioned publications and other references not mentioned in this document, the application of Coanda devices was limited and described only for VTOL and not for level flight. One of the main lessons was that scalability and application for horizontal flight was not practical, particularly for axis-symmetric devices of the Coanda type, where their size would induce increased drag for larger aircraft. An application to a small UAV, however, may be more suitable with a greater degree of integration. Embodiments of the present invention are capable of integrating the ejectors with the fuselage and the engine or propulsion system because the vehicle does not need to take into account the need for a large number of seats. Integration as disclosed in these embodiments is not currently practical or commercially reasonable on large commercial flights.

[0079] This embodiment of the present invention improves the Coanda device and applies it using new techniques for better drag and delay or avoidance of separation in its aggressive turns within the device. Although the compactness of these devices is critical for their deployment in aviation and other fields, the inlet portion needs to be large to improve air drag. Reznick argues that a circular element is more efficient than a linear one. Mocarski shows that drag is critical to increasing thrust. The diffuser portion needs to be long enough to ensure that no boundary layer separation occurs within the device and mixing is complete at the device exit. Conventionally, these diffusers have been long with a very gentle slope to minimize the risks of boundary separation.

[0080] The present invention shows improved drag in devices through new elements that depend on 3D geometry and fluid flow effects and the use of separation avoidance techniques in the Coanda device. The preferred embodiment of the present invention has a drag ratio between 3-15, preferably higher. In another embodiment of the present invention, the device will receive the motive gas from a pressurized source, such as a gas generator, a piston engine (for pulsed operations) or a compressor or turbocharger. Another feature of the present invention is the ability to change the shape of the diffuser walls of the flat ejector used for propulsion by retracting and extending the surfaces to change the geometry so that maximum performance is obtained at all points in the aircraft's mission. Additionally, the need to rotate the entire ejector 90 degrees for VTOL and hover is no longer necessary when deployed diffuser walls are used to direct the jet flow downward.

[0081] Another embodiment of the present invention presents flap-shaped elements for the diffuser walls of a Coanda ejector that is divided into two parts, illustrated in FIG. 14, such as the upper 1401 and lower 201 ejector half that are similar to an air knife. Elements 115 and 215 are actuators or rods that allow movement of said surfaces to the desired diffuser position at both 110a and 210a, respectively.

[0082] Yet another embodiment of the present invention discloses how the positioning of 3D inlet geometries and/or primary fluid gap features, independently or together, significantly improve thruster performance, along with the introduction of separation avoidance patterns. of flow in the propellant. For example, as illustrated in FIGS. 15A to 15C, the 2D input is replaced by a 3D input. FIGS. 15A to 15C further illustrate the revealed multiple 3D ejector elements that improve its performance over baseline and 2D ejectors with inlet, throat and diffuser in the same planes, respectively.

[0083] The inlet may further match the profile shape of the boundary layer formed behind the apex of an aircraft's main airfoil (as illustrated in Figure 16A), thus helping to ingest the boundary layer and delay the overall stall (which improves in all margins) further illustrated in FIG. 25 for the position relative to the airfoil. FIGS. 11A and 11B illustrate the benefits of placing it this way in relation to the airfoil and its boundary layer profile.

[0084] FIG. 16A shows an embodiment of the flat ejector for a wing structure to improve performance at its high angle of incidence and stall margin. The ejector is fed with a primary fluid, for example from a gas generator, and is positioned so as to streamline the flow over said airfoil to delay stalling.

[0085] FIGS. 16B to 16D show different angles of the illustration shown in FIG. 16A, with details of the positioning of the ejector on the wing, the chambers that supply the primary fluid to the ejector and their relative position to each other and to the airfoil.

[0086] The ejector described in FIG. 14 has planar geometry and contains an upper and a lower portion, both introducing the motive fluid as wall jets into a plurality of slits and generally perpendicular to the direction of the jet flow effluent or flow lines, of elements 1401 and 201 that can rotate independently around axes 102 and 202. The curved walls called Coanda walls 104 and 204 allow the primary jets to follow the curvature and drag in the process, and with a ratio greater than 3:1, secondary air, generally arriving from the flow above an airfoil, as the boundary layer of the upper surface of the wing. The primary nozzles 103 and 203 have various shapes with various 3D effects to maximize the drag ratio such as mini delta wings 212 in Fig. 22B, or they may be fluidic oscillators powered by said chambers supplied with the motive fluid to generate a pulsed injection operation. of motive fluid on the walls of Coanda. The fluid mixture reaches the throat area (minimum ejector area) in a pure axial direction. In addition to this point, the current invention introduces a segmented and movable diffuser section, such as a single flap, which plays an important role in the performance of said ejector through vectorization and/or maximization of its performance.

[0087] For example, at takeoff, the inlet of said ejector is fixed and above the airfoil 1700 in fig. 17A pointing forward. FIG. 17A describes the deployment of such an ejector formed by an upper semi-ejector (1401) and a lower ejector (201) and in conjunction with the main wing of an aircraft or a drone. The two semi-ejectors can rotate around axes 102 and 202, respectively, and can also translate according to mission requirements. FIGS. 17B and 17C show the case where only the upper semi-ejector 201 is actively used with a primary fluid, while 1401 is replaced by a simple flap. As before, 201 can rotate around axis 202 and translate relative to the axial position. Semi-ejectors receive the primary fluid under pressure from, for example, a gas generator such as a gas turbine, and allow it to pass through primary nozzles, which may use fluidic oscillators (i.e., pulse at certain frequencies, such as up to 2000 Hz inclusive, to generate pulsating drag of the secondary flow).

[0088] In FIG. 14, the diffuser 210 of the upper half of the ejectors is extended to form a curved surface 210a and guide the mixed primary and secondary flows downward. At the same time, the lower diffuser 110 is also extended to 110a, maintaining the appropriate area growth ratio and mixing characteristics to obtain the maximum thrust required by the aircraft. Some portions of 110a and 210a may not be deployed and 110 and 210 may be controlled independently according to an appropriate schedule. Furthermore, the upper element 201 may or may not be moved axially to meet mission needs. In one embodiment, different amounts of primary fluid and/or delivered under different conditions may be supplied to the upper elements 201 or lower elements 1401 in a continuous or pulsed manner. Diffusing surfaces 110a and 210a may contain cavities and other elements that delay or prevent boundary layer separation. Additional secondary nozzles may also be opened if the fully extended 110a is utilized, at specific and potentially staggered locations, and may be pulsed in accordance with the operating modes of the fluidic oscillators to provide a pulsed mode of operation to the ejector.

[0089] When fluid is received from a compressor bleed, the driving air has a lower temperature. The exhaust gas from the hot end of the gas generator (turbine exhaust), for example, for motive gas temperatures of 816 °C at a pressure of 207 kP the air compressor discharges and the drag ratio of 5:1 and temperature ambient temperature of 138 °C, the temperature of the mixture becomes 168 °C, for which the air density is 0.82 kg/m3, a ~30% drop from ambient. In this way, the overall wingspan can be reduced by ~10%, even taking into account the effects of density reduction, when an airfoil is deployed behind the main booster. For drag ratios of 10:1 (better than the design 5:1), for similar conditions and an emerging jet at 125% aircraft speed, the lift benefit is greater as the mixture density is now greater, at ~90 °C mixture temperature and the lift generated over the wingspan swept by the jet is ~16%. In this example, the wing length can be reduced accordingly.

Thrust Generator

[0090] Another embodiment of the present invention generally refers to a new 3D thrust generator that is capable of receiving pressurized gases from a chamber, entraining still or moving ambient air (including, among others, conditions greater than 0.05 Mach), accelerate the air through momentum and energy transfer with the high-pressure gases, and direct the well-mixed fluids to a high-velocity, non-circular jet effluent with primarily steering velocity component axial. The jet effluent may be a mixture of hot, high-energy gases supplied to the thrust generator through ducting from a high-pressure gas generator outlet and ambient air entrainment. The entrained air can be brought to a high kinetic energy flow level by transferring momentum with the high pressure gases supplied to a propulsive device within the thrust generator. The resulting mixture of air and gas emerges from the thrust generator and points mainly in the axial, downstream direction, opposite to the direction of the vehicle's trajectory. The well-mixed flow provides a mainly unidirectional flow of cooler gas at high speed, which can be used for propulsion, buoyancy, lift generation and attitude control via airfoils placed in the wake of the cooler jet. This is not seen in any vehicle powered by a conventional jet engine. This thrust generator can be free-standing outside the fuselage, built into the fuselage at the front or rear of the vehicle, and/or built into the wings to improve stall margin.

[0091] Reznick invented a circular device with the primary nozzles detached from the Coanda surface and therefore not generating wall jets. Although Reznick teaches that additional secondary fluid is being admitted due to displacement to the Coanda surface, his application, however, is strictly circular in form and therefore cannot be scaled up into a more practical application for larger flow aircraft and , for example, must still be integrated into a wing, as the drag becomes increasingly greater. Furthermore, the slits also appear to be of simple geometry and do not feature any specific 3D features for blending enhancements. The present invention introduces an aerodynamic propellant that generates a rectangular-shaped effluent in the exit plane, to use the energy to generate additional lift on the thin airfoil, an improvement and departure from the circular Reznick application, which cannot be effectively used throughout of an airfoil longer than its own diameter for generating lift in level flight and cannot be deployed over a wing to ingest the boundary layer of a wing, as one of the embodiments of the present invention.

Primary Nozzle Geometry.

[0092] Note that, in all of the patents described, the inventors are not employing any features that increase the area of the primary jet for the secondary flow and, therefore, there are limitations in the described inventions. Furthermore, there is no staggering of primary nozzles in the Coanda device, with the exception of the presence in the Throndson of central primary nozzles, which are not placed in the Coanda device, but rather in the center of the inlet perimeter of the Coanda primary nozzles. The primary nozzles are therefore generally placed in the same axial plane and not staggered, nor are they of different size from adjacent ones, but of the same size and shape. If for a circular Coanda device this is optionally advantageous, for a non-circular device that has a constant spacing between opposing sides of the Coanda primary nozzles along the longest dimension of its inlet plane, the resulting thrust shape would be equally distributed. In an ideal situation, but during level flight, if this device is used for thrust generation, the secondary incoming air will be unevenly admitted into the device and therefore thrust generation will pose challenges to the wing structure and its design. . This is mainly because, in the aforementioned prior art, it was envisaged that these devices would be used in the initial and final stages of an aircraft's flight and not as a single thrust-generating propellant for the entire mission, from takeoff to landing and including floating flight and level flight. In fact, Throndson's invention is only applicable for vertical takeoff, landing and floating, with the power plant assuming the function of providing, in level flight, thrust through a turbojet or turbofan. Thus, in his invention, the devices, including the Coanda ejector, are turned off and form the profile of the wing airfoil in level flight, that is, they become non-operational or inactive during level flight, after the takeoff transition. On the other hand, Reznick teaches a circular device with primary thrust-increasing nozzles, but without incorporating them into a wing for level flight and exploring the input and output of the device beyond thrust generation, which is the present invention. .

[0093] FIGS. 18A to 18D show conventional arrangements for Coanda-type ejectors. FIG. 18A represents a traditional circular-shaped Coanda ejector of the prior art. FIG. 18B shows a flat Coanda-type ejector embedded in a prior art wing. The primary fluid source is a gas turbine and the ejector is optionally advantageously intended to be used for vertical takeoff and turned off in level flight. FIG. 18B covers the elements revealed as Throndson variables, including diameters, angles, and lengths.

[0094] FIG. 18C is Reznick figure 3 and shows another circular embodiment where hypermixing nozzles are used and the primary fluid nozzles are away from the ejector walls. Therefore, the primary jets are no longer wall jets. Reznick only covers the circular geometries of the ejector, clearly intended to be used for takeoff assistance due to scalability limitations.

[0095] FIG. 18D represents an embodiment of the present invention with circular Coanda nozzle elements, with a chamber 211 supplied with primary fluid, which is accelerated through the primary nozzles 203 and injected as wall jets onto the surface 204.

[0096] Throndson uses the non-circular shape of ejector and also rectangular grooves. A rectangular slot is useful in such an application, but it produces a limited surface area for the shear jet drag of oncoming secondary air. In fact, a rectangular slot described by the inventors above produces a jet drag characteristic for a rectangular slot perimeter of the given dimensions, 2L + 2h = 2(L + h) where L is the length and h is the height of each slot. A much greater amount of secondary flow is entrained if a larger primary nozzle perimeter is used, which includes the impact of 3D characteristics. Axially scaling the vertices of a zigzag or wavy (sinusoidal) walls of a primary nozzle as shown in FIGs. 18A to 18D greatly increases secondary air drag as taught in the present disclosure. Pulsed operation by incorporating fluidic oscillators with the primary nozzles further increases the efficiency and drag characteristic of the ejector.

[0097] FIGS. 19A through 19D describe some of the proposed changes to the primary nozzles for improved performance. FIG. 19A shows a zigzag configuration of the primary nozzles along the circumference of the ejector inlet area, while the perimeter of the primary jet exposed to the secondary flow is doubled compared to a single slit perimeter, thus increasing drag through layers of Turbulent shear developed between said zigzag walls of the primary nozzles. FIG. 19B shows a rectangular groove with an increased perimeter and wrinkled to generate additional turbulence and therefore increase drag 1.5 to 4 times compared to the original smooth walls of a rectangular groove. FIG. 19B schematically depicts the increase in surface area from the primary jet to the secondary or entrained air, with the 3D structure of the tips shown in the axial direction. Although typically in a rectangular slot arrangement, the secondary air is entrained mainly between two adjacent slots and on the outer radius side of the slot, the drag is largely enhanced now by surface and 3D effects. FIG. 19C explains the primary jets and secondary jets, respectively, with the turbulence generated by the 3D features, greatly improving the mixing and momentum sharing of the primary flow to the secondary flow over a shorter distance. FIG. 19C shows the interaction and resulting flows of said adjacent wrinkle-wall cracks, with the red arrows showing the primary fluid and the blue arrows the entrained secondary fluid. Shear layers are formed along the walls and the increased perimeter results in significantly greater drag of secondary flows for the same primary inlet flow conditions. Pulsed operation of the primary nozzles further increases the drag ratio.

[0098] Another feature of an embodiment of the present invention is the introduction of advantageous features within the primary nozzle (see FIG. 19E). It is well known that a flow over a delta wing produces opposing vortices towards the center of the delta wing. A miniature feature is placed in some or all of the primary nozzles to generate such vortices that emerge from the primary nozzle. In this case, the vortices advantageously drag significantly larger quantities of secondary air into the ejector, increasing its mixture and giving it momentum carried by the primary fluid that leaves the primary nozzles in a continuous or pulsed manner.

[0099] FIG. 19E shows flow over the delta wing obstruction placed within the primary nozzles at their center, which alters the flow pattern so that it significantly increases the drag ratio of a normal primary nozzle slot without requiring changes in flow rates. , pressure and temperatures. In particular, the primary vortex cores are opposite in direction to the center of said groove or delta wing, entraining significant secondary fluid from the area between the grooves.

[00100] FIG. 20 displays the thermodynamic cycle of the current development with the evolution of working and entrained fluids to obtain high thermodynamic efficiency. FIG. 20 demonstrates that air drag will determine the movement of point D, which represents the condition of the mixture of primary gas and secondary air, in a diagram of the thermodynamic propulsion cycle, to the left and at lower values of temperature and entropy. This is optionally advantageous for a high propulsive efficiency device, in which massive quantities of air are entrained and accelerated at relatively low exit jet velocities, maintaining a high level of thrust due to higher mass flow rates, a key ingredient for achieving high propulsive efficiency. In FIG. 19A, an increase in perimeter by a factor of 2 is shown by replacing each length of a normal rectangular groove perimeter with an equilateral triangle of 2x longer perimeter in the same plane. The perimeter can be increased further by staggering all the vertices of the wall grooves in several planes (see figure 19B). The result of this primary nozzle is to increase the amount of fluid entrained as secondary fluid by at least 15-50% through deep mixing within the formed shear layers. If the initial condition of the secondary air is low velocity, then the performance of rectangular and non-rectangular perimeter shapes may not be very different, however, when the ejector advances and the velocity of the approaching secondary air is significantly greater, such as between Mach 0.0 and Mach 0.25, then the pointed profile shape of the primary nozzle can also be significantly improved by placing the innermost and outermost peaks of the primary nozzle in front of and behind the axial plane of said rectangular groove. In other words, each primary nozzle now becomes a 3D structure that will efficiently delay or advance secondary air drag to improve the overall drag rate. In a Coanda ejector, it is optionally advantageous that the secondary air entrainment and mixing with the primary air for momentum transfer happens quickly and over a short distance. Adding this and other 3D elements to the primary nozzle helps improve the performance of this ejector.

[00101] Another feature related to primary nozzles, as used in this embodiment, is the introduction of fluidic oscillators within the flow channel of the primary nozzles. These fluidic oscillators provide, for example, switching of up to 2000 Hz between two adjacent primary nozzles to alternate wall jet flows and improve drag rates through pulsed operation of the motive fluid.

[00102] Yet another feature implemented in this invention is the staggering of the nozzles with their characteristics, by placing them in various locations along the Coanda surface and, therefore, by introducing the primary flow in multiple axial locations, adjacent to the wall in the form of a wall jet and in a pattern that increases drag and mixing of the secondary fluid. For example, FIG. 21 shows such an embodiment in which a V-shaped vortex generating feature is scaled compared to a normal rectangular slot and injects at least 25% of the total primary fluid before the remainder of the primary fluid mass flow is injected moments later. This injection prior to the rectangular slots results in a higher drag rate sufficient to significantly increase ejector performance. Furthermore, in FIG. 21, the nozzles 205 inject the primary fluid before the primary nozzle 203. The nozzle 205 has a feature that introduces more favorable drag of the secondary flow through shear layers and these nozzles 205 are staggered both axially and circumferentially when compared to the nozzles 205. primary nozzles 203. The primary nozzles 203 have a delta wing feature which is provided with a support leg connected to the midpoint of the primary groove structure on the innermost side thereof and with the delta wing structure pointing against the flow of primary fluid to generate two vortices opposite in direction and strong drag on both sides of the primary nozzle 203, the already entrained mixture of primary and secondary fluid flows resulting from the nozzles 205. The vortices and V-structure of the primary nozzles result in improved drag from 10 to 100% compared to rectangular, non-staggered slots and an overall improvement of momentum transfer from primary to secondary flow.

[00103] Furthermore, FIG. 21 depicts a simpler construction using delta wings placed within smooth wall grooves to form specific delta wing flows and shear layers that advantageously increase the drag ratio by more than 2 times compared to the rectangular primary grooves of smooth wall. All of these elements can be combined for the best drag ratio. The present invention improves the surface for delaying flow separation by means of elements 221. By placing cavities in the Coanda surface 204, said Coanda surface 204 has a relatively aggressive curvature for the shortest distance change of direction of the primary flow, originating radially from the primary nozzles 203, to the axial direction opposite the thrust direction, towards the throat 225. The cavities prevent flow separation and significantly increase the ejector performance, in conjunction with the delta turbulators shown in FIGS. 19D and 19E.

[00104] FIG. 23 illustrates some features according to an embodiment of the present invention. In particular, FIG. 23 compares the similar primary groove height used by Throndson with the proportions used by Throndson to demonstrate the improvement of this embodiment of the present invention. The radius of curvature for the slot height of this embodiment is less than 5:1 with improved separation delay cavities placed on the Coanda surface. In FIG. 23, the radius R' is about 2-3 times smaller than the radius R of the Throndson patent, for similar crack heights. It follows that a ratio of less than 5:1 is possible, due to the use of a logarithmic profile in conjunction with the employment of cavities in the curved Coanda surface, to make the flow from radially pure at the exit of the primary groove to axially more aggressive. pure in the throat. As a result, much faster curvature follows without flow separation, so the device throat can be larger than specified by Throndson by at least 25-100%. The half angle of the diffuser can also be made significantly more aggressive than in the prior art, allowing a much shorter diffuser to be implemented and faster momentum transfer between the primary and secondary flows. As such, FIG. 23 highlights the differences between the present invention and the prior art, especially when more aggressive elements, such as turbulators, primary nozzles, cavities and movable walls, improve the prior art.

Coanda Ejector

[00105] In general, the design of a Coanda ejector applied to an aircraft has been described by many publications. For example, U.S. Patent No. 3,664,611 (Harris) teaches a Coanda-type ejector built into the wing for vertical takeoff and landing purposes. Device is not active during cruise, see FIG. 24. Harris is silent on the use of effluent to generate more lift in a tandem-type arrangement. Furthermore, Harris does not apply the device for use in level flight conditions. Instead, in accordance with conventional practices, the device collapses into an aerodynamic profile wing during level flight conditions.

[00106] Mocarski, on the other hand, teaches that in a Coanda ejector, the high-energy primary fluid, also called motive fluid, is injected as a wall jet and the principle of such a device is to determine a low pressure zone where ambient air is drawn in, followed by a mixing and convergence zone toward a throat, followed by a diffuser to expand the mixture back to ambient pressure at high speed. US Patent No. 3,819,134 (Throndson) modifies and improves the concept described in Mocarski.

[00107] Throndson describes an enhancement to the technology, adding a primary flow in the center of the Coanda-type ejectors to entrain the secondary fluid and improve nozzle performance, with the primary center nozzle using 30-70% of the total primary fluid and the remainder being used in Coanda type parametric nozzles. Throndson states that the increase in thrust is greatly enhanced by this combination, and does not comment on the geometry of the primary fluid nozzles, which appear as simple grooves or holes. Furthermore, the grooves appear to be continuous or discontinuous and without special features. Throndson is silent on the use of jet effluent for downstream lift generation and, in fact, he only employs the device in takeoff, transition and landing and not for cruising conditions, like Harris.

[00108] The present invention further improves the Coanda ejector, generating thrust in all flight conditions by rotating the Coanda device and placing a thin airfoil in the jet effluent, thus generating more lift. Such turbofans generally employ at least two advantages over Harris and Throndson:

[00109] Firstly, the use of the ejector downstream of a wing so as to ingest its damaging boundary layer under cruising conditions, improving the aerodynamic performance of the wing and allowing high angles of incidence, thereby increasing overall performance. The present invention also allows operation of the ejector in all phases of flight, from takeoff to float, transition, cruise and landing. One embodiment also allows the use of a semi-ejector (1/2 of a flat ejector as described by Throndson or Harris) in conjunction with a wing flap to form a non-symmetrical Coanda ejector with boundary layer drag only at the outer edge of the wing. said boundary layer and forming a diffuser with said wing flap, including vectoring the thruster by moving the flap and the Coanda ejector-type air knife in coordination for takeoff and landing.

[00110] Secondly, by using a thin airfoil downstream of (in the ejector sweep) at least in level flight, but also for other flight conditions to generate additional lift at a higher speed flow, allows that the tandem airfoil and propeller are more compact and efficient while generating considerable lift, compared to those described in the aforementioned patents. In this embodiment of the present invention, the shape and profile of the propellant jet effluent are critical to achieving its new efficiency and functionality. Said thin airfoil is placed at a convenient distance from the exit plane of said ejector/propeller to maximize lift, but also before the energy of the jet effluent is dissipated into the environment. This is convenient and practical, as the energy of any jet propulsion device generally dissipates only over very long distances behind the aircraft.

[00111] It is also important to understand that both elements of said tandem need to work together efficiently and optimally, including movement at certain angles and rates favorable to the concept. The thruster/propeller mechanically transmits the thrust component to the aircraft fuselage or its main wing, while the slender airfoil downstream of the thruster is in mechanical contact with the fuselage and not the thruster, still receiving the jet effluent in such a manner that it maximizes the aircraft's lift and allows maneuvering through movements of certain surfaces on the aforementioned thin airfoil.

[00112] Another feature of the present invention provides the ability to use the same nozzles for sustaining the aircraft, floating and landing, as well as for cruising purposes. The lift system described in U.S. Patent No. 8,910,464 (Ambrose) represents the common backbone of VTOL jet fighters. It has limitations due to the extra weight carried in cruise mode, namely the support propeller and its auxiliaries. Under current VTOL technology, the cold nozzles (front nozzles) and the lift propeller are turned off during level flight, which leaves the main exhaust nozzle to provide the reaction force and propel the aircraft forward in forward conditions such as cruise flight. An embodiment of the present invention combines thrust-generating elements of the Coanda nozzle with the aircraft's propulsion system, allowing the ejector to be used in all stages of the flight with minimization of weight and elimination of moving parts. Additionally, it allows the use of this ejector to uniquely minimize drag and maximize lift during level flight.

[00113] Mocarski presents the same technology for the Coanda device with continuous or discontinuous primary fluid grooves, mainly circular or linear. In all of these patents, the Coanda surface is a smooth circular or 2D profile to determine a simple boundary layer attachment without particular elements that could increase drag, increase the aggressive curvature of the Coanda surface, or delay its separation. In Coanda-type ejectors, it is critical that the surface curvature that allows the boundary layer of the wall jet grows and mixes with the secondary air and does not separate. Once the primary jet flow from the primary nozzles separates, the Coanda ejector will not function efficiently or at all. Therefore, it is essential that the curvature of the surface is such that it allows maximum growth of the boundary layer and the entrainment of the secondary fluid and mixing with it, without separation at the wall.

[00114] On the other hand, if the curvature is too great, the device becomes impractically long and large in diameter, which also restricts the amount of secondary fluid drag and mixing and inducing a very long diffusing portion of the device. The groove to radius ratio of the Coanda bend is described by Throndson as being between 1:5-1:15, but a ratio less than 1:5 may be ideal for rapid bending. The curvature of the Coanda curve is clearly stated by Throndson to ideally be between 30-110 degrees compared to the axis of the device. If the diffusion section is becoming too large, this is a major limitation in deploying the technology to an aircraft in level flight, as the length of the diffuser imposes significant additional drag and weight on the aircraft. If the curve becomes >110 degrees, the diffuser can become shorter and improve mixing over a much shorter distance, ensuring thorough mixing and the transfer of energy and momentum to the secondary flow before the mixture exits the device. Note that the diffuser walls are also flat and without 3D elements to improve the mixing process. An embodiment of the invention features movable walls after the throat section, especially in the ejector diffusion area, in a manner favorable to the vertical take-off and landing of an aircraft, without the need to move the entire ejector around its horizontal axis, but rather by extending the segmented diffuser surfaces in the manner described below.

Coanda surface.

[00115] The Coanda surface, as taught by Reznick, Mocarski, and Throndson, should be a rounded curvature, with Throndson providing even finer detail than the range of groove-to-radius height ratios of 1:5 to 1:15 . A logarithmic profile is preferred by those skilled in the art as it provides the fastest growth of the boundary layer without separation from the wall jet. However, one embodiment of the present invention achieves a much more aggressive curvature by introducing cavities in the Coanda surface to significantly improve the curvature of the surface to keep the flow turned on when mixing and moving the mixture to the throat and diffuser. An aggressive curvature is preferred because it allows the ability to quickly mix and rotate the flow in the axial direction, through the throat and into the diffuser section. The turnover of fast-moving fluids, in fact, can keep the boundary layer aggregated, while the boundary layer grows and mixes with the core flow.

[00116] The cavities in the present invention can be of different sizes, can be staggered or aligned, can be located in areas where the curvature is more aggressive and not in areas where the curvature of the fluid is less aggressive. The cavities can also be used in a more aggressive diffuser, in which the half angle of the diffuser is not constant, but variable, grows and then reduces to 0 as represented by element 105 in FIG. 14.

[00117] FIG. 14 represents one of the improvements achieved by the present invention, especially when compared to Throndson. FIG. 14 compares the height of the similar primary groove used by Throndson with the proportions given in Throndson to demonstrate the improvement of the present invention. In particular, FIG. 14 shows the upper and lower halves of semi-ejectors that together form a better, more flexible, and higher-performing ejector that can be suitable for both vertical takeoff and level flight in cruising conditions. The lower wall (1401) of the semi-ejector more aggressively utilizes the nozzle wall's primary jets to more aggressively rotate the flow about axis 102 and over surface 103, a Coanda inlet surface. The point of maximum height of this curve is positioned axially at a point about the distance "G" from the lowest similar position of the curved wall (closes to the axis shown in blue) of element 201. Thus, the two semi-ejectors (or walls of air knife ejectors) 1401 and 201 are staggered, that is, their inlets are not axially positioned in the same location. Similarly, the minimum axial distance of 1401 and 201 is scaled by the distance 'G'. Its diffusers 110 and 210 can change shape through actuators 115 and 215, respectively, in which the flat segmented surfaces forming 110 and 210 become the curved cross-section 110a and 210a, respectively, which direct the flow within said ejector downward or in various directions as dictated by the mission. FIG. 14 also shows the changes in proportions compared to the prior art.

[00118] Furthermore, the ratio of the radius of curvature to the height of the groove of the present invention is less than 5:1 with improved separation delay cavities placed on the Coanda surface. As a result, much faster curvature follows without flow separation, so the device throat can be larger than specified by Throndson by at least 25-100%. Furthermore, by applying a constant variation of the half angle of the diffuser part (i.e. the non-linear growth of the wall away from the center line) and employing the surface with cavities in said diffuser, its dimension can grow without separation from the flow of more aggressively, resulting in a shortening of the total length of the device.

[00119] Furthermore, if the upper and lower halves of the ejector act separately with respect to fluid delivery and functionality, but are capable of working together for the entrainment, mixing and diffusion of the mixture to the exit chamber, then the performance It is greatly improved by the additional movable diffuser walls on the top and bottom surfaces of the flat diffuser. This, in turn, also allows the more compact device to be implemented in conjunction with a wing structure for propulsion purposes in level flight or vertical takeoff, hovering and landing without the need to rotate the entire structure.

[00120] Furthermore, the use of cavities allows the wall to vary from the starting point to the exit around the perimeter of said ejector, which thus allows good integration with the wing structure. A different structure is used on the rounded ends of the ejector, where cavities or special features on the primary nozzles may not be necessary for the ejector to perform well. FIG. 22A shows an ejector having significant 3D features as described in this invention. Furthermore, in the most aggressive areas of the curvatures on the lower wall of the diffuser, the use of cavities (elements 221 in figure 21) will also allow greater rotation of up to 90 degrees downwards or even greater. This is more aggressive than the prior art (e.g., Throndson; see Fernholz, HH "Z. Flugwiss. 15, 1967, Heft 4, pp136-142). FIG. 14 shows a cross section of an ejector having features significant 3D images as described in this disclosure. FIG. 14 also shows mostly segmented walls in the diffuser, capable of redirecting ejector jet effluent and maximizing its performance from changes in the diffuser areas and mixing zones. The ejector inlet plane is not accommodated in a plane (it is not planar) and therefore it is possible to place the ejector above a wing structure so that the boundary layer ingestion improves the performance of the wing airfoil, as may be seen in FIG. 14 with dimension G (separation) between the two inlets of the upper and lower ejector halves. The Coanda surface (103 in FIG. 14; 204 in FIG. 22A) therefore does not admit the primary fluid wall jets at the same axial position, but initially close to the airfoil surface and later away from the wing airfoil surface, in the direction of the axial time history.

[00121] FIGS. 22A to 22F illustrate different embodiments of the present invention, where 3D features are used in the input. FIG. 22F has an upper semi-ejector. FIG. 22B shows element 212 as a delta turbulator placed within a primary nozzle, which considerably improves the drag of secondary flows and cavities such as 222 that improve retention and prevent separation, even in the most aggressive curvatures of the curved wall of Coanda 803 of FIGS. 22A to 22F, and FIG. 14.

[00122] In FIGS. 22A to 22F, element 212 is also introduced into the primary nozzles to improve drag and they may or may not be employed on the opposite side of the ejector depending on conditions and to improve performance. Cavities 221 are placed in contour 204 and in the diffuser to ensure good momentum transfer in the shortest length and generate the exit velocity and temperature profile of the primary and secondary fluid mixture as uniform as possible and prevent flow separation. These primary fluids may be pressurized air from a compressor bleed or pressurized exhaust gas from a gas turbine or a mixture of both and may be fed separately to the upper and lower halves 1401 and 201 of the ejector, which adds another degree of freedom to maximize thrust generation efficiency.

[00123] In FIG. 22A, 3D features increase the perimeter exposed to the flow and allow for a higher drag ratio. In FIGS. 22B and 22C, specially designed turbulators, such as a delta wing placed in the center of the primary grooves, cause flow from the primary fluid chamber, constantly supplied by, for example, a gas generator, to be accelerated in one pass and forced to flow over said delta turbulator 212. Element 212 forces flow in patterns that greatly improve secondary flow drag through a series of mechanisms that include shear layers, rotating and counter-rotating turbulent flows, and increasing the wetted perimeter of said primary nozzles 203 The incorporation of fluidic oscillators within the primary nozzles also provides additional drag capacity through pulsed operation at adjacent primary nozzles.

[00124] FIG. 25 shows an ejector arrangement with the advantage of having a 3D inlet with the lower lip (22) of the ejector being closer to the side 20 of the wall of the upper air profile and beyond the apex of said aerodynamic profile, axially stepped and positioned at the front of the upper lip (23) of said ejector and further away from the surface of the airfoil 20. The position of the lips is modeled to match the most likely boundary layer velocity profile (21) resulting from the airflow near the airfoil. By anticipating flow drag closer to the airfoil wall with lip 22, compared to boundary layer lip 23 drag, better inlet distribution and ejector performance is achieved. In one embodiment of the present invention, the ejector can be moved up and down (in a direction vertical to the upper airfoil wall) to optimize performance. This will allow for better airfoil performance at higher angles of attack and better performance of the ejector itself, due to a better drag process. The 3D elements that differentiate this description from the prior art include the position of the inlet lips 22 and 23, the relative position of the curved walls 204, the positioning of the throat area 24 and the diffuser walls 25. In one embodiment, the two halves of the ejector can move independently in relation to each other and the airfoil, resulting in a constantly optimized position in relation to the aircraft's performance.

[00125] FIGS. 17A to 17C show a flat ejector placed on the airfoil that forms a wing and with two halves that can move independently (see figure 17A) and also an embodiment in which only the upper half of the ejector is used similar to an air knife, but forming a throat and diffuser with a wing flap on the airfoil, which matches the required performance (see figure 17B). In these embodiments, the flap may or may not contain primary nozzles, and the flap moves independently of the upper ejector half, also described as an air knife. The advantage of such a system is that it is simpler; it still allows for high angles of incidence on the wing, as explained in this disclosure, by preventing wing stall from boundary layer ingestion, and the potential to rotate independently of the flap and air knife for optimized performance and maneuverability.

[00126] In particular, FIG. 17A shows embodiments of a flat ejector as described above with elements such as cavities on the wing behind the apex of the wing and such that it mainly ingests the boundary layer above the upper surface of the wing. FIGS. 17A to 17C, on the other hand, show the use of a type of air knife ejector that forces ingested air (as secondary fluid) from above the wing to accelerate and propel the aircraft forward. In all of these embodiments, the ejectors can rotate. Furthermore, in another embodiment, the ejector inlets may also rotate to a limited extent, but their diffuser walls may extend as further explained in FIG. 14, changing angles, escape areas, etc., as dictated by flight conditions. 1401 and 201 can rotate independently about axes 102 and 202, respectively.

Fluidic Propulsion Cycle and System.

[00127] Yet another embodiment of the present invention relates, in general, to a cycle and propulsive system that provides thrust through the transfer of fluidic momentum. The propulsion system consists of a gas generator 1) that provides various sources of high pressure air or gas currents to 2) a network of ducts that direct said compressed fluids to 3) increase the thrust generation of the elements installed in various aircraft stations. Thrust generation enhancing elements direct a high-velocity jet effluent, whose largest velocity component is primarily axial, in the desired direction, thereby generating an opposing thrust force. Jet effluent is a mixture of hot, high-energy gases supplied to the thrust generating element through ducts from high-pressure gas generator locations such as compressor bleeds, combustion bleeds, turbine bleeds and/or or exhaust nozzle, and is designed to drag surrounding air at very high drag speeds. The entrained air is transformed into a flow of high kinetic energy level by transferring momentum with the high pressure gases supplied to said thrust generating element, within the thrust generating element; the resulting mixture of air and gas emerges from the thrust generating element and points mainly in the axial direction, towards said leading edge of the airfoil and mainly on the pressure side of the airfoil preferably in the direction that maximizes the lift of said airfoil downstream.

[00128] FIG. PA2-7 illustrates an embodiment of the present invention, showing the propulsion device in VTOL configuration. Ejectors 201 and 301 are oriented downward and thrust is moving the air vehicle upward. The ejectors are rotating in synchrony and the primary fluid flow is modulated to match the thrust requirement of the front and rear ejectors by compressor bleeds to ejectors 201 and exhaust gases to ejectors 301.

[00129] The most efficient conventional propulsion system for medium and long-haul aircraft engines is the high-bypass turbofan. Conventional turbofans employ at least two shafts, one common to the fan and low-pressure turbine, and one common to the core, which may consist of an intensifier, a high-pressure compressor, and a high-pressure turbine. The high efficiency of a turbofan is brought about by the high bypass ratio, low fan pressure ratios to determine the high propulsion efficiency; and by high global pressure indexes, for high thermal efficiency. The specific fuel consumption of the aircraft is inversely proportional to the product of the thermal and propulsive efficiencies. Thermal losses in a turbofan are mainly due to combustion and thermodynamic losses in components such as the compressor, turbine and mechanical efficiencies that are less than 100%. The irreversibility of the combustion process is generally the main component that leads to lower thermal efficiency and typical high pressure power generation facilities are only 40% thermal efficient. Practical and other limitations of aircraft (weight, drag, etc.) prevent the implementation of methods known in the art for improving thermal efficiency, such as interleaved radiator, heat recovery and others.

[00130] Propulsion efficiency is, on the other hand, maximized when the propellant accelerates a greater amount of air mass flow at small axial speed, immediately above and as close as possible to the aircraft speed. This results in the need to have very large fan diameters and high fan speeds, which increases the drag and weight of the aircraft. Currently, the largest and most efficient turbofan is very large, as the fan diameter exceeds 3 m in size. While increasing fan diameter improves propulsive efficiency, drag increases due to hood size and a compromise is often made to obtain the ideal system. Current thruster efficiency levels exceed 85% and efforts are dedicated to distributing thrusters across the wings to maximize it. A popular idea in the art is the concept of distributed driving elements. The propellants can be distributed across the wings and fuselage of the aircraft. They are mainly electrically or mechanically driven fans on the wings and receiving mechanical work or electrical power from a central unit. Such concepts are difficult to implement due to the complexity of the network involved, the weight of the electric motors and their operability at high altitudes, and in the case of mechanical transmission networks, efficiency, complexity and weight. The dominant design remains that of two engines.

[00131] A disadvantage of the current dominant design is that the turbofan is heavy and complex. More than 30% of its total weight is the fan system alone, which includes the fan accessories and the low-pressure turbine that drives it. Large rotating parts mean that there are additional design limitations, including limitations on tip speed, restrictions on the weight and dimensions of the low pressure turbine, as well as the inlet temperatures to the low pressure turbine. The fan blade needs to qualify and be certified in dedicated Fan Blade Ingestion and Fan Blade Expulsion tests. Additionally, the fan casing must be capable of containing the ejection of such fan blades and protecting the integrity of the aircraft. With smaller systems, the challenge of downsizing a complex turbofan system is significant if efficiency is to be maintained. Particularly for UAVs and small aircraft, Derivation Ratio (BPR) levels are much lower due to material limitations. As their diameters are reduced, fans need to spin faster to maintain their efficiency, and tip losses occur at higher speeds, which generates lower efficiency. For small turbofans, the challenge is that reducing the fan (and compressor) means that the rotational speed must increase dramatically. Those familiar with the art understand that the diameter of a fan scales directly proportional to the square root of the fluid mass flow, while the speed of the fan blade tip is directly proportional to the product of the diameter and the rotational speed (e.g. Pi * Diam * RPM). Therefore, if the fan diameter is significantly reduced, then, conversely, the rotational speed needs to increase to preserve the same tip speed (for mechanical and compressibility reasons), otherwise the losses in performance increase significantly. For example, if a 50-inch diameter fan spins at 2000 RPM, for the same tip speed, a 20-inch fan needs to spin at 5000 RPM, and a 10-inch fan would spin at 10,000 RPM, and so on. This also implies that the Fan Pressure Ratio (FPR) would increase accordingly, leading to lower efficiency of the smaller diameter fan. Furthermore, containment of such a highly stressed fan component would be difficult to achieve and would require a thicker fan casing, increasing weight and causing significant complications regarding the dynamics of the system rotor and its bearing subsystem. This is why large fans are much more efficient than smaller fans. The status quo of small turbofans is significantly lower performance than larger systems, at least 3-4 times lower than the BPRs of large fans and with higher FPRs, which generates low efficiencies (high fuel burn), high rotational speeds (high stress and maintenance) and challenging operability and thermal management. turboprops face the same challenge, although for really small systems they have the best propulsion efficiency. Its main disadvantage is the large size of the propellers required to move large amounts of air and makes its implementation in systems with VTOL capabilities difficult. The modern turboprop uses a low pressure turbine to drive the propeller and employs additional auxiliary systems, such as gears and bearings and their subsystems, pitch control and others.

[00132] Another element of modern aircraft propulsion jets, such as turbofans and turboprops, is that a certain amount of compressor bleed air is required for cabin pressurization, cooling, and turbine discharge to the environment for operability. of the engine itself. The compressor bleed air of a typical modern jet engine is up to 20% of the total compressor discharge air. Compressor bleeds intended for cabin pressurization are not required if the aircraft flies at low altitudes or is unmanned, and this portion constitutes at least 10% of the total bleed. If the turbine is not cooled, then about 10% more compressor air can be extracted before reaching combustion, at the expense of lower firing temperatures and therefore cycle efficiency. However, with the advancement of new non-metallic materials and their high temperature and stress capabilities, the turbine and indeed the majority of the hot section can be manufactured from ceramic matrix composites, not only eliminating the need for cooling air , but also allowing for higher firing temperatures. For example, while limiting the turbine inlet temperature with current, uncooled metallic components is known to those familiar with the topic to be about 950 °C, current CMC materials can withstand the firing temperature uncooled turbine temperature of 1090 °C or more. This results in much greater cycle efficiency and in most cases reduced engine weight, with overall benefits for the aircraft. If the 950°C firing temperature cycle, uncooled all-metal component engine with 20% compressor air bleed is replaced with a 50% air bleed compressor firing at 1090°C with ceramic components , the cycle efficiency can be comparable, despite the 50% of compressed air being made available for other purposes, at the compressor unloading station.

[00133] TABLE 1 shows this comparison for various air bleeds in two uncooled engines with the same unit flow (i.e. 1kg/s) and various bleed percentages and the same turbine output power providing the power. input required for the compressor. The first line shows the cycle pressure ratio, the second line shows the compressor bleed, the thermal efficiency calculation of the metallic motor is shown in line 3 and the thermal efficiency of the CMC versions with similar bleeds and maximum bleed with the same efficiency compared to the metallic version is shown in the last two lines. The general assumption is that the turbine is not cooled in all cases, but air is bled from the compressor for other purposes. The table shows that if the bleed percentage is maintained between the metallic and CMC versions, then with a cycle pressure ratio beyond 8, the CMC engine becomes more efficient. On the other hand, if engine efficiency is kept similar to metallic, the percentage of bleed air can be increased dramatically. This can also be explained by maintaining the same fuel flow for combustion but reducing air flow until the firing temperature becomes 2000F (uncooled CMC technology) from 1090°C (maximum for uncooled metallic technology ). By producing the same turbine power output to balance the compressor power input, more compressor air can be made available at the high firing temperatures. Based on this, the inventor devised a cycle that allows a large amount of compressor bleed to be routed through a flow network and to provide an array of distributed thrust-enhancing devices placed in locations that increase and improve the propulsive efficiency of the compressor. current technique, at a thermal and overall efficiency similar to or better than the aircraft. TABLE 1

[00134] Consequently, conventional thrusters cannot be reduced without significantly compromising their efficiencies. An embodiment of the present invention overcomes the current deficiency by utilizing an improved cycle that eliminates the ventilating subsystem in conjunction with the low pressure turbine. Therefore, this embodiment of the present invention is a system particularly suitable for smaller aircraft systems and UAVs, particularly those that need to be capable of VTOL and STOL operation, because of its efficient, compact and highly integrated power generators.

[00135] The propellant consists of a "pruned" fan and a high pressure compressor placed on the same shaft with a high pressure turbine to form a gas generator, a network of ducts connected to ejectors and the thrust increasing ejectors. The cycle uses a compression system consisting of a trimmed fan (pre-compression only in the core) and a high-pressure, single-stage or multi-stage compressor, preferably a centrifugal compressor with several bleed holes. Compressor bleed holes can bleed up to 50% of the total airflow in the system, with the remainder directed to a combustor system. Combustion adds heat in the form of fuel at constant pressure or volume, and generates a hot flow of gas that is directed to the turbine. The high-pressure turbine expands the hot fluid at a pressure and temperature lower than the turbine inlet pressure and temperature in a conventional expansion process. Preferably, the turbine and combustion materials are high temperature materials that need little or no cooling flow, like modern CMCs. The turbine, which can be centripetal or axial in at least one stage, provides the work necessary to drive the compression system. The exhaust gas leaving the turbine is at lower pressures and temperatures than turbine inlet conditions, but at least twice the ambient air pressure and at temperatures typical of current low-pressure turbine turbofan levels, or that is, 815 °C to 980 °C. Thus, the high-pressure turbine expansion process still results in a high-energy, high-temperature, high-pressure flow that, instead of being directed to a low-pressure turbine, is directed through ducts to various locations on the aircraft at a thrust-generating fluidic propellant.

[00136] Ducts can also be insulated and use high temperature materials such as CMCs. The propulsive elements that receive pressurized air or hot gases use fluidics to drag and accelerate the ambient air in a first section; and after mixing the motive fluid and ambient air and completing the integral transfer of momentum from high energy to low energy (ambient air), accelerating the mixed flow in a second diffuser section, and thereby providing a high-velocity jet effluent as a mixture of high pressure gas (supplied to the propellant from a gas generator) and ambient air entrained at high speeds, preferably and mainly in the axial direction, with a certain axial velocity profile known in the art. The drag rates of said driving elements are between 3-15 and up to 25 for each part of high pressure fluid delivered. Due to the high drag and turbulent and complete mixing of the flows, the jet effluent therefore has a much lower temperature. Following the laws of physics regarding mixing and momentum transfer, the speed of the jet effluent from the propellant element is close to but exceeds the speed of the aircraft. Jet effluent is also non-circular in nature, with few or no rotating components (as opposed to the large propellers of a turboprop or even a turbofan) and can be directed to airfoils to further recover its energy after producing thrust, e.g. example, directed towards the leading edge of a short wing placed some distance behind the booster to generate additional lift. In all embodiments, the gas generator is a modified turbofan in which the fan has been cut to provide core flow only.

[00137] FIG. 26A shows a traditional bypass turbofan that bypasses most of the flow and mixes the core and bypass flows at the turbofan exit. FIG. 26B shows the turbofan with a trimmed fan that takes into account only the core and the bleed flows from the compressor to produce the thrust necessary for propulsion. The compressor bleeds are conveniently located throughout the cold section 2601 of the gas generator 800, as opposed to the hot section 2602 of the generator, so as to allow for maximum efficiency at any time during operation. For example, on takeoff, more compressor bleeding may be required and a higher rotor speed may be optionally advantageous. In this part of the mission, the bleed holes are open so that compressor operation is away from the surge line in a more advantageous condition than if there was no bleed. A current turbofan, like the one shown in FIG. 26A can only allow a maximum of 15% bleed over the course of the mission, but by altering the pruned fan design favorably, more core flow and more compressor bleed can be induced in the present invention, up to and including 50% bleed of the total air passing through the engine. Multiple bleeds may also be involved to increase the efficiency of the system, which maximizes bleed at the lower stages and minimizes bleed at higher pressures, as those skilled in the art would recognize. However, the amount of flow purposefully involved in the bleeding cycle is particularly applicable in this invention. Bleed air leaving the compressor bleed holes is directed through ducts to thrust-increasing thrusters placed in conjunction with the upper surface of an airfoil or immediately behind an airfoil with a high angle of incidence.

[00138] FIG. 27A illustrates an example of a bleed network and ducts embodied in the present invention. The network contains a gas generator 800 that feeds multiple cold thrust boost ejectors 801 and hot thrust boost ejectors 901 through compressor bleed ports 251 and 351, respectively. The pressure and temperature sensors can take signal flow measurements from elements 1702 (cold) and 1707 (hot) that are fed to the microcontroller 900 (not shown). Flow from gas generator 800 to thrust booster ejectors 801 through compressor bleed ducts 251 is dictated by control valves 1703, controlled by microcontroller 900. The same controller determines the actuation of rotary joints 1701 (for element 801) and 1705 (for element 901.) FIG. 27A further shows a series of four (4) cold thrust booster ejectors 801 which are fed from the same compressor orifices 251 of said gas generator 800 and controlled by the microcontroller 900.

[00139] FIG. 27B represents a network similar to that illustrated in FIG. 27A, but with only two cold thrust booster ejectors that are fed from the bleed holes outside the compressor of the gas generator 800. The swivel joints 1701 allow rotation of the elements 801 in multiple directions and can also allow the passage of fluid to said ejector 801. The position of said ejector relative to the aircraft is controlled by the use of electrical or pneumatic or mechanical means from a microcontroller 900. Sensors 1702 that measure the flow rate, pressures and temperatures in the duct to downstream of the bleed hole 251 are used to feed information to the microcontroller. In turn, the microcontroller controls the rotation of the elements 801 around the swivel joints 1701 while also commanding the flow adjustment through the control valves 1703. Likewise, the magnitude and orientation of the thrust of the elements 901 can be adjusted through adjustments to flow rates through control valves 1707 and orientation adjustments about swivel joints 1705 until the position of the aircraft is acceptable. The controller is thus being fed with information from the propulsion system's duct network, in addition to the operational parameters of the gas generator and the orientation and magnitude of the thrust in each of the thrust-increasing ejectors.

[00140] FIG. 27C shows the microcontroller 900 and its network, representing at least twelve (12) inputs and at least four (4) outputs. The outputs primarily control flow rate and thrust (ejector) orientation to control the aircraft's attitude at any time in its mission.

[00141] FIG. 27D offers more network details. Flows from compressor bleed ports 251 and exhaust 351 are fed to thrust boost ejectors 801 and 901. Inputs to microcontroller 900 include gas generator parameters (rpm, compressor bleed air temperature and pressure). , exhaust pressure and temperature, etc.) through input 11. Inputs 26 include power supply to the accelerometers included in the system. Input 30 is the gyroscope. Input 40 is an ultrasonic or barometric altitude sensor input that signals the aircraft's altitude. Input 50 is the GPS input. Inputs 70 are the Bluetooth inputs. Inputs 80 are the R/C receiver.

[00142] Furthermore, FIG. 27D illustrates the feedback from sensors in the ducts, shown as power information 2702 and 2706, to the controller for flow adjustments through actuation of control valves 1703 and 1707. The control valves are connected to the controller via cables 2703 and 2707, respectively, and they adjust based on input received from the controller. The flow is adjusted and measured accordingly by sensors 1702 and 1706, which provide feedback information to the controller to signal that the adjustment has been made accordingly. Likewise, the other sensors 11, 26, 30, 40, 50, 60, 70 and 80, individually or together, are processed by the controller and the adjustments of the ejectors 801 and 901 are transmitted by cables 2701 and 2705.

[00143] In another embodiment of the present invention, the thruster can be rotated downward and direct the thrust to change the attitude of the aircraft in vertical takeoff or short takeoff. FIGS. 28A - 28E show possible forms of a propellant in the present invention. FIG. 28A illustrates a relatively simple quad-ejector system with the ejectors powered by a gas generator in the center. Two of the (cold) ejectors are fed from the compressor bleed ports with compressed air as motive (or primary) air, while the hot ejectors receive exhaust gases from the gas generator exhaust port. All four ejectors point downwards, in hover mode, but aircraft attitude adjustments are possible through changes in the parameters mentioned in FIGs. 27A - 27D. FIG. 28B shows an embodiment of the present invention in which the device(s) can be incorporated into the aircraft. In FIG. 28B, only the two cold ejectors are shown (on all fours). The ejector is flat and placed behind the apex of the main wing, working to expand the main wing's stall margin and generate a high-velocity effluent for use in downstream lift-generating structures. The flat ejector can also rotate along the main axis of the wing for hovering (pointing downward toward the ground, mostly) or to adjust altitude in flight. FIG. 28C illustrates a more complex canard-type system composed of four ejectors (two hot located in the tail and two cold placed behind the canard airfoils) and the ejector embedded inside the system housing next to the (cold) ejector. The ejectors are flat and produce a rectangular-shaped jet effluent in their exit plane, mainly in the axial direction of flight. Alternatively, an ejector with 3D elements is shown in FIG. 28D, in which the intake and throat and diffuser are 3D in nature (rather than 2D), which improves drag and overall performance. Only the upper half of a flat ejector is shown above the wing in FIG. 28E, pairing with the flap of said wing to form a complete structure with the primary fluid introduced only into the upper half of the ejector (the ejector half, which meets the wing flap).

[00144] FIG. 29 shows a possible arrangement of the propulsion system in takeoff or floating in an embodiment of the present invention. The ejectors point downward to suspend the structure and keep it in a floating position.

[00145] FIG. 13 shows the maneuverability of a UAV equipped with a propulsion system. The positioning of the ejectors for pitch, roll and yaw is shown, for both rear (hot) and cold (canard) ejectors.

[00146] In one embodiment of the present invention, the thrusters that receive the compressor discharge air (at at least twice the ambient air pressure) and the hot gas effluent from the high pressure turbine (at at least twice the ambient air pressure) environment) are, in this embodiment, pointing downwards on takeoff, therefore generating opposing thrust that exceeds the weight of the UAV for takeoff. U.S. Patent No. 8,087,618 (Shmilovich et al.) describes the use of such a device embedded in the wing system and utilizing the turbojet exhaust to direct exhaust air or compressed air only at takeoff and small portions of the compressor bleed air (less than 15% is mentioned) for additional flow control over the wing. In particular, they are not increasing thrust, but simply redirecting the exhaust flow by controlling it with compressed air during takeoff. An embodiment of the present invention utilizes a specially designed power plant that specifically extracts bleed air, in excess of 20%, from the compressor and directs it to said propellant throughout the flight, from takeoff to landing. The particular way this is achieved is by designing a compressor with more open first stages that can accommodate more flow, followed by bleeding a portion of the flow in large quantities, such as up to 50% of the total airflow, and directing that portion at all times to the cold gas thrusters, and utilizing the entire portion of the flow remaining for the thermodynamic cycle with the residual energy of the high-pressure post-turbine flow directed to a hot gas thruster. Compressor bleed flows can also be modulated by employing flow controls such as control valves or fluidic valves that modulate flow delivery to the thrusters. Both cold and hot thruster types can be rotated at least from 90 degrees to 120 degrees, pointing downwards and upwards compared to the direct direction of flight and independently. The cold gas thruster can be incorporated or hidden in the wings or, preferably, in the wake of a first wing with a very high angle of incidence (canard wing) and significantly increasing its stall margin by placing the thruster inlet in close proximity. of the airfoil canard and preferably in the last third of its chord and closer to the trailing edge. The high incidence angle can cause separation and stall, but the addition of the thruster at that location will extend its operability well beyond the stall point.

[00147] In another embodiment, the injection of fluids such as water or liquid nitrogen to cool the hot gases sent to the hot propellant can increase the takeoff thrust generated by said propellant by increasing the mass flow rate of the motive air. If the propulsion system is built into a UAV, the amount of water on board the aircraft may be such that after takeoff and the end of the mission, when the fuel on board has almost been consumed, the landing will not require the additional thrust in at least 25% and up to 50%.

[00148] In yet another embodiment, the use of high pressure turbine exhaust gases as the primary/motive fluid for the hot ejector can be increased with additional bleeding from the cold air compressor, particularly to maintain a cooler temperature in the mixture feed. to the hot main propellant primary nozzles, particularly during level flight. In this way, with constant pressure mixing and a reduction in the temperature of the gas mixture, longer-lasting and/or cheaper materials can be used in the pipelines. Modulation of the air bleed from the cold compressor can be carried out through a valve that changes the supply flow from the cold propellant to the duct that feeds the hot propellant, or through a secondary inlet to the hot propellant chamber to extend its life useful. In this case, the cold thrusters are aligned with the main wing system or can be retracted into the fuselage and therefore do not participate in thrust generation.

[00149] The thermodynamic cycle of a typical jet engine is shown in FIG. 30A. The evolution of the working fluid is described from the end of the inlet (point 2) to 3 in a process of compression, fuel addition and combustion at constant pressure through an isobaric process of 3-4, expansion over a turbine of 4- 5. The latter provides the work required by the compressor and additional energy available to drive a fan (through a turbine connected to the fan) of a turbofan or direct expansion through a nozzle to the atmosphere in the case of a turbojet. This embodiment of the present invention eliminates the free turbine required to drive the fan, connects the trimmed fan to the main shaft of the engine, and utilizes the energy of the combustion gases at point 45 to entrain and increase thrust by means of a specially designed ejector built into the aircraft, at all points during the aircraft's mission (takeoff, transition, level flight, bank and landing). The thermodynamic evolution of the invented cycle takes gases 45 through a nearly isentropic expansion at a lower pressure with high efficiency, much greater than expansion through a turbine. Process 45-A' describes the evolution and can be recognized as nearly isentropic, as nozzle expansions of this type are known to have very high efficiency. The evolution of the working fluid 45-A' occurs through a multiplicity of primary grooves located within the ejectors described above. Expansion continues by constant pressure or constant area mixing with approaching ambient air under p2static conditions. The evolution follows A'-D' for the working fluid, while ambient air is brought at constant pressure from the inlet condition point C to D'. In this constant pressure mixing process, the final temperature of the mixture depends on the ambient air drag ratio in the ejector. As described below, a specially designed ejector is used in this invention that maximizes the drag ratio through various elements of said primary nozzles and the ejector mixing section to values greater than 5:1 (five parts of drag air per each part of the primary working fluid). A pumping effect follows, raising the temperature and pressure of the mixture of hotter primary fluid and ambient air drawn into the air. It is , respectively, higher than pamb. This is point D on the diagram in FIG. 30B. A nearly isentropic diffusion and ejection of the mixture is the DE evolution, with the final temperature and pressure of Texit and Pexit, respectively, where Pexit is equal to the ambient pressure at the aircraft speed. The location of point D is between point 2 and point 5, closer to point 2 for higher drag ratios. The advantage of such systems is then obvious, noting that a large quantity of air can be entrained and energized to produce thrust at a lower temperature and mixing speed. This, in turn, allows the exhaust from such a thermodynamic cycle to be utilized not only for thrust generation, but also to direct it advantageously over various airfoils for additional lift generation or to vector for VTOL and STOL capabilities of the aircraft, in cycle exit. Furthermore, the placement of the drag inlet of said ejectors in some embodiments may be such that ingestion of the resulting boundary layer from an airfoil such as a wing results in additional benefits to the stall margin of the wing's assisting airfoil. In one embodiment, a first set of wing airfoils is positioned such that it operates at a very high angle of incidence, with a very low stall margin. By placing said ejector just behind said apex point of the airfoil, in the area prone to developing boundary layer separation, the suction side (i.e., the drag side or the inlet side of the ejector) determines that the margin of stall is significantly improved, allowing very high lift generation on said airfoil/wing to be obtained without stall.

[00150] Additionally, in some cases where a short takeoff distance is desirable, exhaust gases from a turbofan can be directed to the suction side of an aerodynamic profile (such as a flap). Although several concepts have used this technique, there have been limited results. In this embodiment of the present invention, there is higher lift proportional to the square of the highest local velocity, at least for the portion of the wing exposed to the jets emerging from the thrust element, because it utilizes the benefits of directing the higher kinetic energy fluid ( exhaust gas mixture air) directly to the pressure side of the wing or flap, rather than the suction side, or directly to the leading edge much in the manner of an airfoil of turbocharged machines (such as a turbine).

[00151] Furthermore, the ejector exhaust being at a significantly lower temperature and yet an average exit velocity greater than the air speed, can be directed to a secondary thin airfoil, downstream. The ejectors guide their jet effluents towards airfoils that are slender and can be manufactured from reinforced composite materials. The higher speed of the jet effluent determines high lift force on said airfoil compared to the lift generated by said airfoil that would only receive the aircraft's speed flow. On the other hand, the size and shape of airfoils can be significantly reduced to produce lift similar to that of a very large wing. Returning now to the inlet of the ejectors and observing their placement just above the airfoil 803, the ingestion of the boundary layer has developed after the apex of said airfoil 803 in the ejector and this boundary layer being sucked by said ejector determines a better stall margin of the airfoil 803 and allows it to operate efficiently at a greater angle of incidence.

[00152] In another embodiment illustrated in FIG. 31, said cold ejectors 801 are placed behind the airfoil 803 and in front of the airfoils 802, which impacts both, increasing the stall margin on 803 due to the suction of the boundary layer of said airfoil 803 and the lift force due to the greater velocity of the jet effluent at exit 801 directed efficiently towards airfoil 802. This allows for more aggressive level flight position of the two airfoils, 803 and 802, shorter airfoils for the same lift and ejector rotation for vertical takeoff, hover and maneuvering of the aircraft. It also allows for more beneficial use of jet effluent from the thermodynamic device that can be used for lift generation and not wasted into the environment, as is the case with current jet engines.

[00153] Although this is not an exhaustive list, different embodiments of the present invention are designed to provide some or all of the following improvements and advantages:

[00154] Improve the ability to maximize thrust increase and jet effluent vectorization of a Coanda-type flat ejector in all flight conditions;

[00155] Improve efficiency and shorten the device for better integration with the aircraft wing or fuselage by introducing specific 3D features into the primary nozzle and Coanda surface;

[00156] Incorporating such a device into a wing to exploit specific wing geometries in order to increase the efficiency of the aircraft;

[00157] Improve the efficiency of the primary nozzle to draw and mix the secondary fluid in the shortest period and shortest length of the device through additional features;

[00158] Increase the general geometry in a non-circular manner to allow efficient operation in level flight of the aircraft, in addition to taking off, hovering and landing, while improving the propulsive efficiency of the aircraft and eliminating the presence of nacelles and propulsion engines main ones on the wings and fuselage of the aircraft;

[00159] Generate additional thrust and lift due to the greater local speed of the jet over the wing using the residual kinetic energy of the jet effluent that generally only generates thrust through mechanical connection;

[00160] Shorten the wings while preserving the same lift by extending the walls of the propellant diffuser as a lift and propulsion generation device;

[00161] Improve the ejector to work better in conditions less than the ideal conditions of a fixed geometry ejector (for example, optimizing operations and thermodynamic propulsion cycles by using 2 halves of an ejector, which can be moved relative to each other and add a flap-like feature capable of completely expanding and collapsing the ejector diffuser walls);

[00162] Increase lift as a proportion of wingspan due to the relatively low temperature of the jet effluent mixtures emerging from the propellant and the axial velocity component with values higher than the aircraft speed;

[00163] Include composites as a type of material used in thin airfoils due to their ability to withstand higher temperatures of the emerging jet effluent mixture;

[00164] Reduce the dimensions and total weight of the aircraft because the airfoil can be thinner in width and shorter in wingspan with high mechanical resistance to stress;

[00165] Significantly improve the maneuverability and versatility of an aircraft, including allowing V/STOL and buoyancy, through rotation and modulation of the flow of both the propellant and the airfoil; and/or

[00166] Improve aircraft attitude control, buoyancy and VTOL capabilities by enabling a compact system with small oscillations and a distributed propulsion system, particularly in UAVs, UASs and drones.

[00167] Furthermore, in addition to the many features mentioned above, different embodiments of the present invention may also include some or all of the following improvements and advantages:

[00168] The thermodynamic cycle will be simpler, with an ejector/eductor type element that replaces the entire functionality of the low pressure turbine and fan subsystem, thus reducing the weight of the system by at least 30%. This is particularly advantageous for smaller UAV-type systems where turbofans are not efficient due to the reasons explained above;

[00169] The potential to rotate or vector the eductor-type thrusters independently and allow vertical takeoff and landing without moving large rotating parts;

[00170] The potential to modulate the flow to these thrusters during takeoff and level flight, as well as upon landing and in emergencies by applying different levels of thrust at various locations on the aircraft, and completely isolating any number of such thrusters;

[00171] The potential to eliminate large components of rotating parts with non-moving parts of the same functionality, i.e., the fan replaced by fluidic thruster/edductor; a direct improvement of component life is expected from non-moving parts versus rotating parts, especially for small UAVs and aircraft where fan dimensions require very high speeds;

[00172] The potential to use lightweight, high-temperature materials such as composite materials, carbon fiber-based materials, and CMCs for ducts and thrusters;

[00173] The potential to modulate the bleeds so that, in level flight, only the hot thrusters receive hot gas or a mixture of hot exhaust gases and cooler air from the gas generator compressor bleed;

[00174] The benefit that the gas generator is optionally operated at the same rotational speed, without large excursions in RPMs between takeoff and cruise, away from the surge or stall line;

[00175] The benefit of giving any shape to the propellant and being able to integrate it considerably with the fuselage and wings of the aircraft;

[00176] The benefit of having large drag and turbulent mixing within said propellants, so that the jet effluent from their exhaust is sufficiently low in temperature to allow an aircraft lift or attitude control airfoil to survive and function properly including generating more lift using the jet's higher speed; and/or

[00177] The benefit of incorporating the propellant into the wing behind the apex of the wing bulge where the boundary layer would separate at high angles of incidence, and ingesting said boundary layer to delay its separation and increase the stall margin of said wing in the level flight.

[00178] It should be noted that any of the ejectors 701, 801, 901 can be configured using any ejector geometry described in the present patent application.

[00179] Although the preceding text sets out a detailed description of numerous different embodiments, it should be understood that the scope of protection is defined by the words of the following claims. The detailed description should be interpreted as exemplary only and does not describe all possible embodiments as describing all possible embodiments would be impractical, if not impossible. Numerous alternative embodiments could be implemented, using current technology or technology developed after the filing date of this patent, which would still fall within the scope of the claims.

[00180] Thus, many modifications and variations can be made to the techniques and structures described and illustrated here without departing from the spirit and scope of the present claims. Accordingly, it should be understood that the methods and apparatus described herein are illustrative only and do not limit the scope of the claims.

Claims (10)

1. Vehicle (804), comprising: a main body having a front portion, a rear portion, a starboard side and a port side; a gas generator (800) coupled to the main body and producing a gas flow; at least one front conductor fluidly coupled to the generator (800); at least one rear conductor fluidly coupled to the generator (800); first and second front ejectors (801) fluidly coupled to the at least one front conductor, coupled to the front portion and respectively coupled to the starboard and port sides, the front ejectors (801) respectively comprising an outlet structure through which gas from the at least a front driver flows at a predetermined adjustable speed; at least one rear ejector (901) fluidly coupled to the at least one rear conductor and coupled to the rear portion and respectively coupled to the starboard and port sides, comprising an outlet structure through which gas from the at least one rear conductor flows at a speed adjustable preset; elements of the first and second primary airfoils (802), both with guiding edges, and respectively coupled to the starboard and port sides, and further with their guiding edges respectively located directly downstream of the first and second ejectors (801), such that gas from the front ejectors (801) flows over the guiding edges of the primary airfoil elements (802); at least one secondary airfoil member (904) having a guiding edge and coupled to the main body, with the guiding edge located directly downstream of the exit structure of the at least one rear ejector (901), such that the flow of gas from the rear ejector (901) flows over the leading edge of the at least one secondary airfoil (904); and first and second canard configuration wings (803) coupled to the forward portion and, respectively, coupled to the starboard and port sides, the canard configuration wings (803) being specifically to develop boundary layers of ambient air that circulate over them when the vehicle is in motion and respectively located directly upstream of the first and second front ejectors (801), so that they are fluidly coupled to the boundary layers, characterized in that the first and second front ejectors (801) respectively comprise first and second inlet portions, and the first and second front ejectors (801) are positioned such that the boundary layers are ingested by the inlet portions. 2. Vehicle (804) according to claim 1, characterized in that: the gas flow produced by the generator (800) is the only means of propulsion of the vehicle (804). 3. Vehicle according to claim 1, characterized in that: the generator (800) comprises a first region in which the gas flow is at a low temperature and a second region in which the gas flow is at an elevated temperature; the at least one rear conductor supplying gas from the first region to the first and second front ejectors (801); and the at least one rear conductor supplies gas from the second region to the at least one rear ejector (901). 4. Vehicle (804) according to claim 1, characterized in that: the gas generator (800) is located in the main body. 5. Vehicle (804) according to claim 1, characterized in that: the first and second front ejectors (801) each have a guiding edge and the entirety of each of the first and second front ejectors (801) is rotatable in around an axis oriented parallel to the guide edge. 6. Vehicle (804) according to claim 1, characterized in that: the first and second front ejectors (801) each have a guiding edge and the entirety of each of the first and second front ejectors (801) is rotatable in around an axis oriented perpendicular to the guiding edge. 7. Vehicle (804) according to claim 1, characterized in that: the at least one rear ejector (901) has a guiding edge and is rotatable about an axis oriented parallel to the guiding edge. 8. Vehicle (804) according to claim 1, characterized in that: the at least one rear ejector (901) has a guiding edge and is rotatable about an axis oriented perpendicular to the guiding edge. 9. Vehicle (804) according to claim 1, characterized in that: at least one of the exit structures is non-circular. 10. Vehicle (804) according to claim 1, characterized by: additionally comprising a cockpit portion (805) configured to allow manned operation of the vehicle (804).