Streamline airframe with boundary ingestion fluidic propulsive elements

Abstract

To provide an ejector system that provides compressed fluid selectively taken in from thrusters according to positions of the respective thrusters in an aircraft equipped with a plurality of thrusters on wings thereof.SOLUTION: A vehicle includes a main body, and at least one wing coupled to the main body. A source of compressed fluid is coupled to the main body. The vehicle further includes first and second thrusters, each of said first and second thrusters having an intake structure and each of said first and second thrusters being in fluid communication with a source. The first thruster is coupled to the main body and the second thruster is coupled to the at least one wing. The first and second thrusters are positioned, when in a first configuration, such that at least a portion of a boundary layer produced due to motion of the vehicle is ingested by intake structures of the first and second thrusters. The vehicle further includes a system for selectively providing the compressed fluid to the first and second thrusters.SELECTED DRAWING: Figure 1

Description

[Technical field]

[0001] Copyright Notice

[0001] This disclosure is protected by U.S. and / or international copyright laws. (Copyright) 2019 Jetoptera, Inc. All Rights Reserved. A portion of the disclosure of this patent document contains material that is subject to copyright protection. The copyright owner has no objection to the facsimile reproduction by any person of the patent document or the patent disclosure, as it appears in the Patent and Trademark Office patent file or records, but otherwise reserves all copyright rights.

[0002] Priority claim

[0002] This application claims priority to U.S. Provisional Patent Application No. 62 / 677,419, entitled "Streamline Airframe with Boundary Ingestion Fluidic Propulsive Elements," filed May 29, 2018, the contents of which are incorporated by reference in their entirety into this specification. [Background technology]

[0003]

[0003] Among the many objectives in aircraft design are the minimization or elimination of major rotating parts, reducing the overall weight of the aircraft, and lowering the overall drag profile of the aircraft. [Brief description of the drawings]

[0004] [Figure 1] 1 is a cross-sectional view of one embodiment of the present invention depicting the top half of an ejector and the velocity and temperature profiles within the internal flow. [Diagram 2] 2 illustrates surface features of the ejector of FIG. 1 according to one embodiment. [Diagram 3] 1 illustrates a partial perspective view of an air intake structure according to one or more embodiments. [Figure 4] 1 illustrates a partial perspective view of an air intake structure according to one or more embodiments. [Diagram 5]1 illustrates cross-sectional variations in ejector internal geometry according to one embodiment. [Figure 6] 1 illustrates a top view of an aircraft according to one embodiment. [Figure 7] 1 illustrates a thruster in a stowed position within a wing or fuselage, according to one embodiment. [Figure 8] 1 illustrates a thruster in a stowed position within a wing or fuselage, according to one embodiment. [Figure 9] 1 illustrates a thruster in a deployed position according to one embodiment. [Figure 10] 1 illustrates a thruster in a deployed position according to one embodiment. DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS

[0005]

[0011] This patent application is intended to describe one or more embodiments of the present invention. The use of absolute terms such as "shall" and "will" as well as specific quantities should be interpreted as applicable to one or more of such embodiments, but not necessarily to all such embodiments. Thus, embodiments of the present invention may omit or include modifications of one or more features or functions described in the context of such absolute terms.

[0006]

[0012] FIG. 1 illustrates a cross section of the upper half of a thruster such as an ejector 200, which may be mounted on a vehicle (not shown), such as a UAV or manned aircraft, such as an airplane, as a non-limiting example. The plenum 211 is supplied with air (i.e., a pressurized motif gas flow) that is hotter than the surroundings, for example, from a combustion engine that may be used by the vehicle. This pressurized motif gas flow, indicated by arrow 600, is introduced into the interior of the ejector 200 through at least one conduit, such as a primary nozzle 203. More specifically, the primary nozzle 203 is configured to accelerate the motif fluid flow 600 to a variable predetermined desired velocity directly on the convex Coanda surface 204 as a wall jet. Additionally, the primary nozzle 203 provides an adjustable amount of the fluid flow 600. This wall jet then serves to entrain through the intake structure 206 a secondary fluid, such as ambient air, indicated by arrow 1, which may be stationary or approach the ejector 200 at a non-zero velocity from the direction indicated by arrow 1. In various embodiments, the nozzles 203 may be arranged in an array, in a curved orientation, a spiral orientation, and / or a zigzag orientation.

[0007]

[0013] The mixture of flow 600 and air 1 may move purely axially at the throat section 225 of the ejector 200. As the mixing and smoothing process continues due to diffusion in a diffusing structure such as the diffuser 210, the axial temperature (800) and velocity (700) profile of the ejector 200 no longer has the high and low values ​​present at the throat section 225, but becomes more uniform at the end 100 of the diffuser 210. As the mixture of flow 600 and air 1 approaches the exit face of the end 100, the temperature and velocity profile becomes nearly uniform. In particular, the temperature of the mixture is low enough to be directed towards an airfoil such as a wing or control surface.

[0008]

[0014] 2, the V-shaped vortex generating secondary nozzles 205 are staggered when compared to the regular rectangular primary nozzles 203 to inject at least 25% of the total fluid flow 600 before the remainder of the mass flow rate of the fluid flow is injected by the nozzles 203 at a later time. This injection by the nozzles 205 prior to the injection of the nozzles 203 results in a higher entrainment rate sufficient to significantly improve the performance of the ejector 200. The secondary nozzles 205 introduce a more favorable entrainment of the secondary flow through the shear layer and are staggered both axially and circumferentially with respect to the primary nozzles 203.

[0009]

[0015] The primary nozzle 203 may include a delta wing structure 226 with support legs connected at their innermost to a midpoint of the primary nozzle 203 structure, with the apex of the delta wing structure pointing into the flow of the fluid stream 600. This then generates two vortices with opposite directions, strongly mixing the already mixed mixture of primary and secondary fluid streams coming from the nozzle 205 from either side of the primary nozzle 203.

[0010]

[0016] Additionally, one embodiment enhances the surface for flow separation delay via elements such as dimples 221 disposed on the Coanda surface 204. The dimples 221 prevent flow separation and significantly improve the performance of the ejector 200. Additionally, the surface of the diffuser 210 (see FIG. 1 ) may also include dimples 222 and / or other elements that delay or prevent boundary layer separation.

[0011]

[0017] In one embodiment, the intake structure 206 may be circular in configuration. However, in various embodiments, and as best shown in Figures 3-4, the intake structure 206 may be non-circular, and may in fact be asymmetric (i.e., not identical on either side of at least one, or any given plane that bisects the intake structure). For example, as shown in Figure 3, the intake structure 206 may include first and second opposing edges 301, 302, where the second opposing edge includes a curved portion that protrudes toward the first opposing edge. As shown in Figure 4, the intake structure 206 may include first and second lateral opposing edges 401, 402, where the first lateral opposing edge has a larger radius of curvature than the second lateral opposing edge.

[0012]

[0018] 5, an embodiment may include at least one internal actuation element (e.g., actuator and / or linkage) 601, 602 disposed between the exterior surface 603, 604 and the interior surface 605, 606 of the ejector 200. In the illustrated embodiment, the actuator 601 is configured to move the first surface 605 relative to the second surface 606 (e.g., toward and away from the central axis of the ejector 200) when the second surface is not moving. Similarly, the second actuator 602 is configured to move the second surface 606 relative to the first surface 605 when the first surface is not moving. This ability to change the internal geometry of the ejector 200 to multiple configurations allows the ejector to optimally operate in multiple flight conditions (e.g., liftoff, takeoff, cruise flight, etc.).

[0013]

[0019] One or more embodiments primarily use fluid propulsion ejector / thruster system (FPS) propulsion. Exemplary FPS systems are described, for example, in U.S. Patent Application Nos. 15 / 456,450, 15 / 221,389, and 15 / 256,178, which are incorporated by reference herein as if fully set forth herein. One or more embodiments combine one or more fluid propulsion ejector / thruster systems (FPS) and distributed propulsion to eliminate "bulks" protruding from the vehicle, fully streamline the airframe, and actively capture / control the drag-generating boundary layer, thereby saving fuel and increasing range.

[0014]

[0020] Referring to FIG. 6, one embodiment uses one or more gas generators 610a-c to feed Coanda effect based ejectors 200 mounted on major bodies such as a fuselage 612 and wings 614 via a series of conduits 616, thus providing both suction jets and wall jets to reduce drag, delay flow separation (which causes drag and premature stall), and include distributed propulsion throughout the fuselage and wings.

[0015]

[0021] One or more embodiments of the ejector 200 may be configured in shapes other than circular, and may use a primary fluid to entrain (suck or take in) a large volume of air and accelerate this air to higher velocities by mixing the primary and secondary (entrained) fluids (e.g., gases from the turbine and ambient air). One embodiment includes a gas turbine supply (via a pneumatic network, rather than a mechanical network as in most distributed propulsion systems) to a number of these ejectors 200 mounted approximately "flush" with the airplane fuselage 612 and / or wings 614, and may receive compressed hot fluid from inside the fuselage or other onboard locations and use it to entrain more air from the boundary layer formed over the fuselage in front of the ejector, mix with the hot gases in the ejector, and discharge it downstream parallel to or tangent to the fuselage in a wall jet manner.

[0016]

[0022] The distribution of the hot fluid to the ejectors 200 can be done in a controllable way using control valves. These ejectors 200, which can be numerous and made of thin metal ceramic composites, are arranged on the fuselage 612 and wings 614 like bird wings, complementing each other for suction / entrainment inlet and for exhaust, in wall jets that are intelligently distributed (i.e. staggered) to cover a large area of ​​the aircraft. The ejectors 200 according to various embodiments can be given any shape (rectangular, circular, crescent, curved, etc.) and therefore can cover any part of the aircraft (as opposed to placing the suction side at the end of the fuselage of a large plane, which would limit its effect to local conditions and would be ineffective, and would further introduce large rotating parts (rotors / fans) that would induce RPM limitations for noise and efficiency). Furthermore, at high Reynolds number and high temperature conditions of the gas supplied to the ejectors 200 by the distribution network, losses are minimal and the FPS system actually eliminates major rotating parts while reducing weight. The network of thermal conduits 616 can be insulated with ultra-lightweight materials and valves can be actuated to allow or deny flow to the ejector 200 as and when required.

[0017]

[0023] As shown in Figures 6-10, an embodiment includes one or more gas generators 610a-c in a first configuration, best shown in Figures 7-8, inactive and fluidly connected with a plurality of ejectors 200 located within the interior of the fuselage 612 or wing 614. As best shown in Figures 9-10, the ejectors 200 can be triggered by suitable actuation means emerging from the interior of the fuselage 612 or wing 614 as needed, causing the gas flow generated by the gas generators 610a-c to entrain a large amount of air and generate an area of ​​suction over a large part of the aircraft's fuselage. A particular advantage of this configuration is that the boundary layer is "re-energized" and turns into a jet capable of generating thrust. Additionally, the thrust is now in fact "distributed" over the entire aircraft, ensuring a very large propulsive efficiency. This, combined with the high thermal efficiency from the gas generators, which have a pressure ratio of at least 20:1, makes the aircraft very efficient, including also the lower drag guaranteed by the distribution of the ejectors 200.

[0018]

[0024] Although preferred embodiments of the present disclosure have been illustrated and described, as above, many changes can be made without departing from the spirit and scope of the present disclosure. Thus, the scope of the described systems and techniques is not limited by the disclosure of the preferred embodiments. Instead, the described systems and techniques should be determined entirely by reference to the following claims.

[0019]

[0025] The embodiments of the disclosure in which an exclusive property or privilege is claimed are defined as follows:

Claims

1. A flying device, comprising: The main body, at least one wing coupled to the main body; a source of compressed fluid coupled to the main body; first and second thrusters, each of the first and second thrusters having an intake structure, each of the first and second thrusters in fluid communication with the source, the first thruster coupled to the main body and the second thruster coupled to the at least one wing, the first and second thrusters positioned such that when the first and second thrusters are in a first configuration, at least a portion of a boundary layer generated due to motion of the flight device is captured by the intake structure of the first and second thrusters, the first and second thrusters controllably configured to emerge from and be contained within at least one of the main body and the at least one wing; A flying device comprising:

2. 2. The flight device of claim 1, wherein in a second configuration, at least one of the first and second thrusters is completely contained within the main body and at least one of the at least one wing.

3. The first and second thrusters each include: A convex surface; a diffusing structure coupled to the convex surface; at least one conduit coupled to the convex surface and configured to introduce the compressed fluid to the convex surface.

4. 4. The flight device of claim 3, wherein for each of the first and second thrusters, the intake structure is coupled to the convex surface and the diverging structure comprises a termination configured to provide an outlet for the compressed fluid and a boundary layer.

5. The flight device of claim 3 , wherein the convex surface of at least one of the first and second thrusters includes a plurality of concave portions.

6. The flight device of claim 1 , wherein the intake structure of at least one of the first and second thrusters is asymmetric.

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