Control system for aircraft

Abstract

An aircraft includes a first leading edge defining a forward edge of a left aircraft wing, a second leading edge defining a forward edge of a right aircraft wing, a plurality of plasma actuators disposed along the first and second leading edges, a control processing unit communicatively coupled to each plasma actuator, and at least one flight stability sensor communicatively coupled to the control processing unit. The control processing unit commands the at least one plasma actuator to generate the plasma in response to a signal from the flight stability sensor.

Description

[0001] This application is a divisional application of the patent application with application number 201910657338.7 and invention title "Control System for an Aircraft", filed on July 19, 2019.

[0002] Priority Claim

[0003] This application claims the priority of U.S. Application No. 62 / 700,462, filed on July 19, 2018. The disclosure of U.S. Application No. 62 / 700,462 is hereby incorporated by reference. Technical Field

[0004] The subject matter disclosed herein relates to aircraft and methods of controlling an aircraft. Background Art

[0005] Supersonic and hypersonic aircraft typically use control surfaces as a means of control. Actuators and other mechanisms for positioning the control surfaces are typically used to control the control surfaces.

[0006] Flight stability and control of aircraft at supersonic and hypersonic speeds are multi-faceted areas that include a balance of several factors, largely due to the speed at which the aircraft is flying. At supersonic and hypersonic speeds, the aircraft is subject to high-frequency disturbances and may require a faster response rate than can be achieved with traditional control surfaces (such as ailerons, elevators, and rudders). Additionally, even at lower speeds, aircraft control surfaces (such as a movable supersonic engine exhaust nozzle) are very heavy, which reduces aircraft efficiency. Summary of the Invention

[0007] Aspects of the present embodiments are outlined below. These embodiments are not intended to limit the scope of the claimed embodiments; rather, these embodiments are only intended to provide a brief overview of possible forms of the embodiments. Additionally, the embodiments may include various forms that may be similar or different from the embodiments set forth below, corresponding to the scope of the claims.

[0008] In one embodiment, an aircraft includes a first leading edge defining a forward edge of a left aircraft wing, a second leading edge defining a forward edge of a right aircraft wing, a plurality of plasma actuators disposed along the first and second leading edges, a control processing unit communicatively coupled to each plasma actuator, and at least one flight stability sensor communicatively coupled to the control processing unit. The control processing unit commands at least one plasma actuator to generate plasma in response to a signal from the flight stability sensor.

[0009] In another embodiment, an aircraft control system includes a control processing unit, at least one sensor communicatively coupled to the control processing unit, and at least one plasma actuator disposed near a leading edge of an aircraft wing, the plasma actuator communicatively coupled to the control processing unit. The control processing unit commands the plasma actuator to generate plasma in response to at least one signal from the at least one sensor. BRIEF DESCRIPTION OF THE DRAWINGS

[0010] These and other features, aspects, and advantages of the present disclosure will be better understood when the following detailed description is read with reference to the accompanying drawings, in which like characters represent like parts throughout the drawings, wherein:

[0011] Figure 1 is a side schematic view of a plasma ignition combustion system;

[0012] Figure 2 is a side schematic view of a plasma ignition combustion system having a schematic view of a control system;

[0013] Figure 3 is a side schematic view of a plasma ignition combustion system mounted on a wing;

[0014] Figure 4 is a side schematic view of a plasma ignition combustion system mounted on an engine;

[0015] Figure 5 is a rear-to-front view of an engine exhaust ring including a plasma ignition combustion system;

[0016] Figure 6 is a top view of a subsonic aircraft including a plasma ignition combustion system;

[0017] Figure 7 is a top view of a supersonic aircraft including a plasma ignition combustion system;

[0018] Figure 8 is a top view of a hypersonic aircraft including a plasma ignition combustion system;

[0019] Figure 9 is a front view of a hypersonic aircraft including a plasma ignition combustion system;

[0020] Figure 10 is a side view of a hypersonic aircraft including a plasma ignition combustion system;

[0021] Figure 11 is a front view of a hypersonic aircraft including a plasma-assisted control system;

[0022] Figure 12Is a perspective view of a hypersonic vehicle including a plasma-assisted control system;

[0023] Figure 13 Is a side view of a hypersonic vehicle including a plasma-assisted control system;

[0024] Figure 14 Is a schematic diagram of a control system for a plasma ignition combustion system; and

[0025] Figure 15 Is a schematic diagram of a control system for a plasma-assisted control system.

[0026] Unless otherwise indicated, the drawings provided herein are intended to illustrate the features of embodiments of the present disclosure. It is believed that these features are applicable to various systems including one or more embodiments of the present disclosure. Accordingly, the drawings are not meant to include all conventional features known to those of ordinary skill in the art for practicing the embodiments disclosed herein. Detailed Description

[0027] In the following specification and claims, many terms will be referenced, which shall be construed to have the following meanings.

[0028] Unless the context clearly dictates otherwise, the singular forms "a", "an", and "the" include plural referents.

[0029] "Optional" or "optionally" means that the subsequent described event or circumstance may or may not occur, and the description includes instances where the event occurs and instances where the event does not occur.

[0030] Approximating language, as used throughout the specification and claims, may be used to modify any quantitative representation that can permissibly vary without resulting in a change in the basic function to which it is related. Accordingly, a value modified by one or more terms, such as "about" and "substantially", is not limited to the precise value specified. In at least some instances, the approximating language may correspond to the precision of the instrument used for measuring the value. Here and throughout the specification and claims, range limitations may be combined and / or interchanged, such ranges are identified and include all the sub-ranges subsumed therein unless the context or language indicates otherwise.

[0031] As used herein, the term "axial" refers to a direction aligned with the center axis or shaft of a gas turbine engine, or a direction aligned with the center axis of a propulsion engine and / or an internal combustion engine. The axially forward end of a gas turbine engine is the end closest to the fan and / or compressor inlet, where air enters the gas turbine engine at the compressor inlet. The axially rearward end of a gas turbine engine is the end of the gas turbine closest to the engine exhaust, where low pressure combustion gases leave the engine via the low pressure (LP) turbine at the engine exhaust. In a non-turbine engine, axially rearward is toward the exhaust and axially forward is toward the inlet.

[0032] As used herein, the term "circumferential" refers to one or more directions (and tangent thereto) around the circumference of a ring of a burner or a circle defined, for example, by the swept area of a turbine blade. As used herein, the terms "circumferential" and "tangent" are synonymous.

[0033] As used herein, the term "radial" refers to a direction moving outward away from the center axis of a gas turbine or the center axis of a propulsion engine. The "radially inward" direction is aligned toward the center axis moving toward a decreasing radius. The "radially outward" direction is aligned away from the center axis moving toward an increasing radius.

[0034] As used herein, the term "plasma" refers to a gas that has been made conductive by heating or subjecting it to an electromagnetic field, where long-range electromagnetic fields dominate the behavior of the matter.

[0035] As used herein, the term "cold plasma" refers to a plasma in which the characteristic temperature of electrons is much higher than the characteristic temperature of "heavy" particles (i.e., neutral and ionized molecules and atoms), as opposed to being in thermal equilibrium (i.e., a "hot" plasma).

[0036] As used herein, the term "plasma actuator" refers to a plasma generating device that produces a plasma acting on a control surface of an aircraft, related to having fuel (plasma assisted combustion) or no fuel (plasma assisted control). Plasma actuators can contribute to stabilizing and / or enhancing combustion, and can also produce a plasma acting on one or more control surfaces of an aircraft, and interact with the aerodynamic conditions of the aircraft via flight. For example, by using swirl vanes or bluff bodies in an air flow that create a recirculation zone that stabilizes the flame position, the combustion flame can be spatially stabilized. By modulating or adjusting the fuel flow, an unstable (time-varying) flame can be temporally stabilized. Plasma can locally enhance combustion, stabilize the flame at a given location, and / or can be adjusted to manage unstable (time-varying) flame characteristics. Plasma can also be used, for example, during supersonic flight, to modify how shock waves act on the surface of an aircraft.

[0037] As used herein, the term "ramjet" refers to an air-breathing jet engine that uses the forward motion of the engine to compress the incoming air without an axial compressor or a centrifugal compressor.

[0038] As used herein, the term "scramjet" refers to a variant of a ramjet engine in which combustion occurs in a supersonic airflow therein.

[0039] As used herein, the term "subsonic" refers to a speed less than the speed of sound, less than about Mach 1. As used herein, the term "transonic" refers to a speed of from about 0.8 Mach to about 1.2 Mach. As used herein, the term "supersonic" refers to a speed greater than the speed of sound, more specifically, a speed from about Mach 1 to about Mach 5. As used herein, the term "hypersonic" refers to a speed above about Mach 5.

[0040] Embodiments of the present disclosure may relate to subsonic, supersonic, and hypersonic aircraft employing a plasma ignition combustion system in cooperation with aircraft control surfaces. The embodiments disclosed herein illustrate enhanced and simplified control of an aircraft using control surfaces.

[0041] Figure 1 A plasma ignition combustion system 10 using a control surface 12 of the present embodiment is shown. The control surface 12 may be a substrate provided with components of the claimed embodiment. Additionally, fluid may flow over the substrate or the outer surface of the control surface 12. Fuel is injected into the airflow at A through the control surface 12, and the gas flows in direction B. The combustion process is initiated by at least one plasma actuator 14 located downstream of the injection location 16. Combustion occurs in a combustion zone C and generates a force on the control surface 12, which can be used for the stability and control of the aircraft.

[0042] Several plasma actuator arrangements are possible. By injecting microwave power into a gas (such as air or a fuel-air mixture), "microwave plasma" can be generated, where the microwave power is preferentially coupled to a region of the gas that is already ionized and conducting, such as in front of a flame, thereby adding energy to in front of the flame and increasing the local heat release rate.

[0043] Microwave plasma can also be generated upstream of the flame zone, in air or an air-fuel mixture, where it can act as a plasma source for generating reactive radicals that flow in and enhance the combustion process, without having to deposit energy into normal gas heating. The generated plasma can be cold or hot. Gas can be introduced into the combustion zone (such as from the sidewalls of a combustion chamber) through the plasma, and this device is sometimes referred to as a "plasma tube". The microwave frequency can be in the range of from about 0.3 GHz to about 300 GHz.

[0044] The plasma tube plasma actuator can also be powered by other means such as radio frequency induction (in the range of about 3 kHz to about 0.3 GHz), or by electrodes driven by direct current or alternating current. Thermal jets occur in the combustion chamber to stabilize and control the flame. Radio frequency or microwave energy can be generated by power electronics or magnetrons and transmitted to the desired area in the engine through a transmission line such as a coaxial cable, or through other appropriately shaped structures such as waveguides or "applicators".

[0045] Spark plasmas can be generated to stabilize the flame in a manner similar to a diffusion ignition flame in a burner, where the total fuel-air ratio is lean (i.e., oxygen remains after complete fuel combustion). In such an arrangement, the plasma acts as a local heat source. Such plasmas can be generated by intermittent "spark" plasmas (e.g., spark plug igniters), or by a continuous "arc" plasma maintained between two electrodes by controlling the current flowing through the circuit. Spark plasmas can also be achieved via intermittent laser spark plasmas (or continuous laser arc plasmas) generated by focusing laser power into a gas volume.

[0046] Cold plasmas can be maintained in a gas by controlling power deposition such that energy does not transfer from the electrons to the heavy particles because of low pressure, low power density, or the energy is applied for a short time (pulsed). The generated plasma produces reactive radicals that flow into and enhance the combustion process without depositing energy into ordinary gas heating. Nanosecond plasmas can also be configured with an air flow as a plasma tube.

[0047] Figure 2 A plasma ignition combustion system 10 using a control surface 12 of this embodiment is shown. Fuel is injected through at least one fuel injector 18 at an injection location 16. One or more fuel injectors 18 are in fluid communication with a fuel control valve 20 that controls the amount of fuel flowing to the fuel injector 18. A fuel supply 22 is in fluid communication with the fuel control valve 20 and upstream of the fuel control valve 20. In some embodiments, a fuel injector biter 24 can be mechanically coupled to the fuel injector 18 to adjust, as needed, the angle at which fuel is dispersed from the fuel injector 18 based on the ambient air flow and operating conditions of the aircraft. The fuel flows downstream through the control surface 12 to a plasma location 26 that is adjacent to one or more plasma actuators 14. The plasma location 26 is downstream of the injection location 16 with respect to the air flow direction B. The fuel is ignited and a combustion zone C is formed adjacent to the downstream portion 12' of the control surface.

[0048] Still referring to Figure 2, the fuel injector 18 and the plasma actuator 14 are disposed substantially flush with the control surface 12, limiting potential detrimental effects, such as an increase in drag, when they are not in operation. The plasma ignition combustion system 10 may include a flow surface 28 near the plasma location 26. The flow surface 28 may be used to enhance the surface on which the forces from the combustion zone C act. For example, in Figure 2 an embodiment, the flow surface 28 may be a thin half nozzle or crescent shaped. In other embodiments, the flow surface 28 may be semi-cylindrical (i.e., "half tube") shaped, hemispherical, or semi-elliptical, conical, semi-conical, frustoconical, sinusoidal, and other contoured shapes. In other embodiments, the flow surface 28 may be planar and may be inclined or angled relative to the control surface 12. In other embodiments, the flow surface 28 may be segmented planar, including multiple planar surfaces assembled from individual planar segments and arranged at various angles. In other embodiments, multiple flow surfaces 28 may be used. In other embodiments, a separate flow surface 28 may not be required. In other embodiments, the profile or shape of the control surface 12 will be designed to obviate the need for a separate flow surface 28. The flow surface 28 may generally be open at one end and shaped at the end proximate the control surface 12. The flow surface 28 may enhance the transfer of forces generated by plasma ignition combustion to the control surface 12.

[0049] Still referring to Figure 2 , the plasma ignition combustion system 10 may include a power source 30 electrically coupled to the plasma actuator 14 for generating plasma. The plasma ignition combustion system 10 may include a control processing unit 34. The control unit or control processing unit 34 may be communicatively coupled to each of the fuel supplier 22, the fuel control valve 20, the fuel injector latch 24, the plasma actuator 14, and the power source 30. The control unit 34 may also be communicatively coupled to the aircraft controller 32 and the local airspeed indicator 36 and / or the aircraft airspeed indicator 38. In Figure 2 an embodiment, components that may be communicatively coupled to each other are connected by a dashed line. However, other communication couplings among the components are possible.

[0050] Since generating plasma consumes energy, it is desirable to generate plasma only when needed. For example, in one embodiment, plasma is generated such that it is present near plasma location 26 just before fuel from injection location 16 arrives. Thus, it may be desirable to time the fuel injection through fuel injector 18 together with the plasma generation through plasma actuator 14 to minimize energy losses (via fuel losses and unused plasma). Local airspeed indicator 36 can be used to roughly estimate the flight time of fuel from injection location 16 to plasma location 26, since the first distance 40 between injection location 16 and plasma location 26 may be fixed and thus a known quantity. Since local airspeed indicator 36 is disposed at control surface 12 downstream of injection location 16 and upstream of plasma location 26, and since boundary layers and other fluid effects may be present in the vicinity of control surface 12, and since these effects may vary with operating and environmental conditions, local airspeed indicator 36 may be able to accurately determine how fast the fuel will travel the first distance 40, which is between injection location 16 and plasma location 26.

[0051] Figure 2 The local airspeed indicator 36 shown in FIG. may be an ultrasonic sensor, or a calibrated static pressure type sensor for roughly estimating air flow. Local airspeed indicator 36 may also be other types of sensors, including flat probe sensors, pitot sensors, differential pressure sensors, and / or any other sensors that can be used to measure flow across a surface. Ultrasonic sensors are capable of distinguishing fuel velocity and air velocity flowing past under conditions where there is a velocity difference between two fluids. Other sensors that do not distinguish fuel velocity and air velocity flowing past can still accurately predict the flight time of fuel flowing from injection location 16 to upstream of plasma location 26 by correlating fuel velocity with air velocity. Plasma ignition combustion system 10 may also include air flow indication 38 from different locations and / or from aircraft controller 32. As described above, local airspeed indicator 36 may have the benefit of taking boundary layer conditions into account. However, in embodiments where aircraft airspeed and the flight time of fuel flowing from injection location 16 to upstream of plasma location 26 are highly correlated, air flow indication 38 from aircraft controller 32 may be sufficient. Under flight conditions where the air flow direction is not aligned with the line connecting injection location 16 to plasma location 26 (e.g., due to the formation of a lateral boundary layer and / or other aerodynamic effects or aircraft maneuvers), the orientation of fuel injector 18 can be adjusted by fuel injector snapper 24 to ensure that fuel dispersed by fuel injector 18 reaches plasma location 26. Additionally, guide devices, tubes, vanes, and / or other devices (not shown) may be employed to direct the fuel dispersed by fuel injector 18 to plasma location 26.

[0052] Figure 3 An embodiment of a plasma ignition combustion system 10 on an airfoil control surface 12 is shown. Figure 3 The airfoil control surface 12 shown in may be a wing of an aircraft, other airfoil structures on the aircraft, an airfoil aircraft, and other surfaces used as the control surface 12. Figure 3 The embodiment includes fuel injected at an injection location 16 upstream of the plasma location 26 via at least one fuel injector 18, wherein plasma is generated via at least one plasma actuator 14. At least one plasma actuator 14 ignites the fuel, creating a combustion zone C at the downstream end 12' of the control surface. Air flows over the control surface 12 in direction B. Figure 3 The embodiment may also include Figure 2 a number of other system components, including but not limited to a power supply 30, a local airspeed sensor 36, a flow surface 28, a fuel supply 22, a fuel control valve 20, a control processing unit 34, an aircraft controller 32, an aircraft airspeed indicator 38, and a fuel injector latch 24. In other embodiments, the components of the plasma ignition combustion system 10 will be disposed on the underside 12” of the control surface rather than on the top side of the control surface 12, or in addition to being disposed on the top side of the control surface 12, the components of the plasma ignition combustion system 10 will also be disposed on the underside 12” of the control surface. In other embodiments, the components of the plasma ignition combustion system 10 will be disposed near the upstream end 12”' of the control surface rather than on the top surface 12 of the control surface and / or on the underside 12” of the control surface, or in addition to being disposed on the top surface 12 of the control surface and / or on the underside 12” of the control surface, the components of the plasma ignition combustion system 10 will also be disposed near the upstream end 12”' of the control surface.

[0053] Figure 4 An embodiment of the plasma ignition combustion system 10 in a scramjet engine 41 application is shown. Figure 4 The scramjet engine 41 shown may include an air inlet 42 that provides a main burner section 42 upstream of a flared exhaust section 48 and upstream of a divergent section 46. The scramjet engine 41 may generally be axisymmetric about an engine centerline CL. The flared exhaust section 48 may include one or more control surfaces 12 that form an annular exhaust port and diverge radially outward from the engine centerline CL as they extend rearward in direction B. Figure 4 The embodiment includes fuel injected at an injection location 16 upstream of the plasma location 26 via at least one fuel injector 18, wherein plasma is generated via at least one plasma actuator 14. At least one plasma actuator 14 ignites the fuel, creating a combustion zone C at the control surface 12. Figure 4Embodiments may also include Figure 2 several other system components, including but not limited to power supply 30, flow surface 28, local airspeed sensor 36, fuel supply 22, fuel control valve 20, control unit 34, aircraft controller 32, aircraft airspeed indicator 38, and fuel injector latch 24. In other embodiments, the components of the plasma ignition combustion system 10 will be arranged in various orientations around the annular exhaust port to allow force vectors to be applied to the control surface 12 at different angles required to control the aircraft. Figure 4 Embodiments can reduce system complexity because the fuel delivery and handling system may already be in place due to fuel combustion at the main burner section 42. Additionally, by using the surface in the engine exhaust system or exhaust nozzle as a control surface in conjunction with plasma ignition combustion, thrust can be extracted from the plasma ignition combustion, thereby increasing the thrust from the supersonic combustion engine 41 and / or reducing the fuel flow required at the main burner section 42. This embodiment is similar to Figure 4 arrangements that are also possible in subsonic combustion and / or conventional gas turbine aircraft engine configurations.

[0054] Figure 5 shows a rear-to-front view embodiment of the plasma ignition combustion system 10 in a supersonic combustion engine 41 application similar to Figure 4 . In other embodiments, the plasma ignition combustion system 10 can be in a gas turbine engine or other subsonic engine. Figure 5 The embodiment is viewed from the rear end of the supersonic combustion engine 41 through the flared exhaust section 48. A plurality of plasma ignition combustion systems 10 are circumferentially spaced around the exhaust ring of the supersonic combustion engine 41. In Figure 5 embodiments, each of the plurality of plasma ignition combustion systems 10 includes a plasma actuator 14, a flow surface 28, and Figure 2 the other system components shown in Figure 5Embodiments include eight plasma ignition combustion systems 10 that are spaced approximately evenly around a loop of a supersonic combustion engine 41 at intervals of about 45 degrees. In other embodiments, other numbers of plasma ignition combustion systems 10 and other spacing arrangements may be used. Additionally, the flow surfaces 28 may not be needed due to the curvature of the engine loop and / or a different number of flow surfaces 28 than the number of plasma actuators 14 may be used. By operating the plasma ignition combustion systems 10 asymmetrically, a net force in any desired direction is possible. This force can act on the control surfaces 12 within the engine. In other embodiments, the force can act on surfaces within the engine, which in turn can act on the control surfaces 12 of the aircraft.

[0055] Figure 6 A top view of an exemplary subsonic aircraft 51 is shown. The plasma ignition combustion system 10 (not shown) of the present embodiment can be used in subsonic aircraft 51 applications. For example, the plasma ignition combustion system 10 can be disposed on the surface of the subsonic aircraft 51, including but not limited to the right wing 50, left wing 52, right engine nacelle 54, left engine nacelle 56, right horizontal stabilizer 58, left horizontal stabilizer 60, aircraft fuselage 62, vertical stabilizer 64 (left side and / or right side), right winglet 66, and / or left winglet 68. Additionally, the plasma ignition combustion system 10 can be disposed on the surfaces (and other surfaces) corresponding to those described above on the underside of the subsonic aircraft 51.

[0056] Figure 7 A top view of an exemplary supersonic aircraft 61 is shown. The plasma ignition combustion system 10 (not shown) of the present embodiment can be used in supersonic aircraft 61 applications. For example, the plasma ignition combustion system 10 can be disposed on the surface of the supersonic aircraft 61, including but not limited to the left control surface 70, right control surface 72, left wing 74, right wing 76, left engine 78, right engine 80, central aircraft body portion 82, and tail 84. Additionally, the plasma ignition combustion system 10 can be disposed on the surfaces (and other surfaces) corresponding to those described above on the underside of the supersonic aircraft 61.

[0057] Figure 8A top view of an exemplary hypersonic vehicle 71 is shown. The plasma ignition combustion system 10 (not shown) of the present embodiment can be used in hypersonic vehicle 71 applications. For example, the plasma ignition combustion system 10 can be disposed on the surface of the first hypersonic vehicle 71, including but not limited to the right horizontal surface 86, the left horizontal surface 88, the right vertical surface 90 (right outer side and / or left inner side), the left vertical surface 91 (outer left side and / or right inner side), the rear part 92 of the vehicle body, the middle part 94 of the vehicle body, and the front part 96 of the vehicle body. Additionally, the plasma ignition combustion system 10 can be disposed on the corresponding surfaces (and other surfaces) on the lower side of the first hypersonic vehicle 71.

[0058] Figure 9 A front view of the hypersonic vehicle 71 is shown, including an air inlet 98 disposed on the lower side of the first hypersonic vehicle 71. The plasma ignition combustion system 10 (not shown) of the present embodiment can be used in the first hypersonic vehicle 71 applications. For example, the plasma ignition combustion system 10 can be disposed on the surface of the first hypersonic vehicle 71, including but not limited to the right horizontal surface 86, the left horizontal surface 88, the right vertical surface 90 (either side and / or both sides), and the left vertical surface 91 (either side and / or both sides).

[0059] Figure 10 A side view of the first hypersonic vehicle 71 is shown, including an air inlet 98 disposed on the lower side of the first hypersonic vehicle 71. The plasma ignition combustion system 10 (not shown) of the present embodiment can be used in the first hypersonic vehicle 71 applications. For example, the plasma ignition combustion system 10 can be disposed on the surface of the first hypersonic vehicle 71, including but not limited to the left horizontal surface 88, the left vertical surface 91 (either side and / or both sides), the upstream part 100 of the lower side, and the downstream part 102 of the lower side. The air inlet 98 is disposed between the upstream part 100 of the lower side and the downstream part 102 of the lower side.

[0060] Figure 11Shows a front view of a second hypersonic vehicle 300 having a different configuration from the first hypersonic vehicle 71. The second hypersonic vehicle 300 includes a left leading edge 304 that defines the leading edge of the left vehicle wing 312. The left leading edge 304 may extend forward to the vehicle nose 311, where the left leading edge 304 may converge with the right leading edge 306, which defines the leading edge of the vehicle right wing 310. The second hypersonic vehicle 300 may include an inlet 302 disposed on the lower side 308 of the vehicle. A plurality of plasma actuators 14 may be disposed along each of the left leading edge 304 and the right leading edge 306. The plurality of plasma actuators 14 may be flush with the leading edges 304, 306 such that they do not extend or protrude from the vehicle into the oncoming airflow. Additionally, the plurality of plasma actuators 14 may be disposed at or near the leading edges 304, 306 such that they are positioned to generate plasma at the leading edges 304, 306 where shockwaves are most likely to exist. In other words, the plurality of plasma actuators 14 do not need to be precisely disposed at the leading edges 304, 306 as long as they are close enough to generate plasma at the leading edges 304, 306. For example, in one embodiment, the plurality of plasma actuators 14 may be located within about 5% of the vehicle length of at least one of the leading edges 304, 306, where the vehicle length is defined by the length of the vehicle body vertex line 316.

[0061] In operation, when the second hypersonic vehicle 300 reaches supersonic and / or hypersonic speeds, shockwaves may propagate along the lower side 308 of the vehicle, along the tops and bottoms of the left wing 310 and the left wing 312, and along other surfaces of the second hypersonic vehicle 300. The shockwaves (not shown) may provide lift and / or actuation forces on various surfaces of the second hypersonic vehicle 300 where recovery or cancellation of control forces is desired in order to stably control the second hypersonic vehicle 300. Thus, the plasma actuators 14 may be used to generate plasma along each of the left leading edge 304 and the right leading edge 306 between the vehicle and the shockwaves. This may result in an accidental effective shockwave propagation angle. Additionally, this may also change the propagation region to redirect downstream towards the rear end (not shown) of the vehicle. Similarly, using the plasma actuators 14 to generate plasma between the vehicle and the shockwaves may buffer the vehicle from the shockwaves, change the shockwave angle, and / or change the forces acting on the control surfaces 12 of the vehicle.

[0062] Still referring to Figure 11, the second hypersonic vehicle 300 may include one or more flight stability sensors 301 disposed on the lower side 308 of the vehicle. The one or more flight stability sensors 301 may be used to sense at least one aerodynamic characteristic of the second hypersonic vehicle 300 under given operating conditions. For example, the one or more flight stability sensors 301 may consist of an airspeed indicator that indicates when there is a condition in supersonic flight and thus the presence of a shock wave is evident. In another embodiment, the one or more flight stability sensors 301 may include a static pressure sensor that indicates the presence and / or magnitude of the shock wave, as well as the frequency at which the shock wave propagates along the lower side 308 and / or the left leading edge 304 and right leading edge 306 of the vehicle. The one or more flight stability sensors 301 may also be disposed along the left leading edge 304 and right leading edge 306, where the shock wave is most likely to form and / or act under the action of force.

[0063] Still referring to Figure 11 , the vehicle control system may use the one or more flight stability sensors 301 to control the frequency and / or magnitude of the plasma generated by the plasma actuator 14. For example, in the case where the magnitude of the shock wave is proportional to the airspeed of the vehicle, the one or more flight stability sensors 301 may be used as static pressure sensors to measure the magnitude of the shock wave, thereby obtaining an approximation of the airspeed. Similarly, the one or more flight stability sensors 301 may be used to sense the shock wave frequency, and the vehicle control system may use this shock wave frequency to control the cancellation and / or stable activation of at least one plasma actuator 14. By measuring the frequency of the pulses caused by the pressure wave, a single flight stability sensor 301 may be used to determine the shock wave frequency, which is sensed by the single flight stability sensor 301. In other embodiments, multiple flight stability sensors 301 located at multiple positions on the vehicle may be used to determine the shock wave frequency, and the multiple flight stability sensors 301 sense the flight time it takes for a single shock wave to propagate from a first flight stability sensor 301 to a second flight stability sensor 301.

[0064] Figure 12A perspective view of a second hypersonic vehicle 300 is shown, which includes a left wing 312, a right wing 310, a left leading edge 304, a right leading edge 306, a vehicle nose 314, and a plurality of plasma actuators 14 disposed along the left leading edge 304 and the right leading edge 306. The second hypersonic vehicle 300 may further include a vehicle body apex line 316 extending the length of the vehicle. The vehicle body apex line 316 may define an intersection between the right wing 310 and the left wing 312. The vehicle body apex line 316 may be defined by a single line or, alternatively, may be a curved and / or slightly smoothed or rounded portion of the top of the second hypersonic vehicle 300 where the left wing 312 and the right wing 310 intersect or cross. The second hypersonic vehicle 300 further includes a rear end 318 defined by a left trailing edge 322 and a right trailing edge 324, which also define the trailing edges of the left and right wings 310, 312. A vehicle exhaust port 320 may also be disposed in the rear end 318. A vehicle apex 326 defines an intersection of the left wing 312, the right wing 310, and the rear end 318. The left leading edge 304 and the right leading edge 306 may be acute angles at an intersection that may be located on or in front of the vehicle. For example, in one embodiment, the left leading edge 304 and the right leading edge 306 form an angle less than about 60 degrees. In another embodiment, the left leading edge 304 and the right leading edge 306 form an angle between about 5 degrees and about 45 degrees. In another embodiment, the left leading edge 304 and the right leading edge 306 form an angle between about 9 degrees and about 35 degrees. In another embodiment, the left leading edge 304 and the right leading edge 306 form an angle between about 15 degrees and about 25 degrees. In another embodiment, the left leading edge 304 and the right leading edge 306 form an angle between about 17 degrees and about 23 degrees.

[0065] Still referring to Figure 12, the second hypersonic vehicle 300 may include a first sensor 328 disposed at or near the nose 311 of the vehicle, a second sensor 330 disposed at or near the apex 326 of the vehicle (i.e., located at the center of the top surface of the vehicle near the rear end), a third sensor 332 disposed on the right wing 310 near the rear end 318, and a fourth sensor 334 disposed on the left wing 312 near the rear end 318. The second hypersonic vehicle 300 may also include other sensors 336 at other locations, including corresponding locations on the bottom surface of the vehicle. The sensors 328, 330, 332, 334, 336 may be used to establish various orientations and reference frames of the vehicle during flight. For example, the sensors 328, 330, 332, 334, 336 may be used to establish the angle of attack 116 of the vehicle, the yaw 126 of the vehicle, the angular acceleration 130 of the vehicle, the vertical acceleration 132 of the vehicle, the vibration of the vehicle, the attitude 120 of the vehicle, the altitude 122 of the vehicle, and other parameters. Each of the sensors 328, 330, 332, 334, 336 may be a gyroscope, a GPS sensor, an accelerometer, a lidar, a proximity sensor, a communication device for establishing a position relative to a reference frame other than a satellite, a barometer, a navigation compass, a quantum gyroscope, a MEMS gyroscope, an optical fiber gyroscope, a gyrocompass, a heading indicator, a gyro, a Foucault pendulum, a hemispherical resonator gyroscope, a vibrating structure gyroscope, a dynamically tuned gyroscope (DTG), a ring laser gyroscope, a London moment gyroscope, an optical accelerometer, and other types of sensors. In one embodiment, the first sensor 328 will be located within approximately 10% of the length of the vehicle at the nose 314 of the vehicle, where the length of the vehicle is defined by the length of the vehicle body apex line 316. In another embodiment, the first sensor 328 will be located within approximately 5% of the length of the vehicle at the nose 311 of the vehicle, where the length of the vehicle is defined by the length of the vehicle body apex line 316. In another embodiment, the second sensor 330 will be located within approximately 10% of the length of the vehicle at the rear end 318 of the vehicle, where the length of the vehicle is defined by the length of the vehicle body apex line 316. In another embodiment, the second sensor 330 will be located within approximately 5% of the length of the vehicle at the rear end 318 of the vehicle, where the length of the vehicle is defined by the length of the vehicle body apex line 316.

[0066] Still referring to Figure 12, each of the sensors 328, 330, 332, 334, 336 can be used individually or in cooperation with each other to establish at least one aspect of the aircraft orientation. The sensors 328, 330, 332, 334, 336 can be tuned such that they operate in a frequency range of 1 kHz to 5 MHz and generate 1000 to millions of orientation signals per second. The orientation signals from the sensors 328, 330, 332, 334, 336 can be used by the aircraft control system to adjust the orientation of the aircraft via the plurality of plasma actuators 14. By asymmetrically activating the plasma actuators 14, the aircraft control system can cause a net force to act on the aircraft, resulting in the desired target orientation of the aircraft. For example, by activating more plasma actuators 14 along the left leading edge 304 than the right leading edge 306, the control system can generate a net force on the aircraft that causes a change or adjustment in the yaw 126 (not shown) of the aircraft, or a rolling force on the aircraft 300. Similarly, by activating more plasma actuators 14 at or near the aircraft nose 311 than at or near the aircraft rear end 318, the control system can generate a net force on the aircraft that causes a change or adjustment in the angle of attack 116 (not shown) of the aircraft.

[0067] Figure 13 A side view of a second hypersonic aircraft 300 is shown, which includes a left wing 312, a left leading edge 304, an aircraft nose 311, an aircraft body apex line 316, a left trailing edge 322, an aircraft apex 326, a plurality of plasma actuators 14, one or more flight stability sensors 301, and a plurality of aircraft orientation sensors 328, 330, 334.

[0068] Figure 14Shows a control system 200 that can be used to control a plasma ignition combustion system 10. The control system includes a control unit 34 that receives at least one airspeed indication 106, which can be from an ultrasonic sensor 104, an aircraft airspeed indicator 38, and / or a local airspeed indicator 36. The control unit 34 also receives an input from at least one flight command 108, which may include commands such as various aircraft maneuvers or commands to stabilize flight due to turbulence or changing environmental and / or operating conditions. The control processing unit 34 may also receive input signals from a plurality of aircraft sensors and parameters 110, which include but are not limited to ambient humidity 112, a vibration sensor 114, an angle of attack indication 116, a flight segment indication 118, an aircraft attitude 120, an aircraft altitude 122, a gyroscope 123, a turbulence sensor 124, an aircraft yaw indication 126, an aircraft control mode 128, an aircraft angular acceleration 130, and an aircraft vertical acceleration 132. The control unit 34 may use the plurality of aircraft sensors and parameters 110 to determine what actions to perform and the means for performing them. For example, if excessive vibration or turbulence is sensed, the control unit may activate one or more plasma ignition combustion systems 10 to act on one or more control surfaces 12 to relieve forces, and the execution of this operation may depend on the altitude 122, the angle of attack 116, the vertical acceleration 132, and / or other factors.

[0069] Still referring to Figure 14 , the control unit can determine a plurality of control target values, including but not limited to a target injection angle 134 (i.e., the angle of injection of the injection), a target fuel mass flow rate 135, a target fuel pulse rate 138, a target duration 140 (i.e., the duration for which one or more plasma ignition combustion systems 10 can be activated), a target plasma pulse rate and / or plasma waveform 142, a target delay 144 (i.e., the time difference from when fuel is injected to when plasma is generated based on the flight time (or estimated flight time) of the fuel flowing from the injection position 16 to the plasma position 26), and a target plasma size 146. These target values can be transmitted to the fuel injector actuator 24, the fuel injector 18, and / or the plasma actuator 14, as Figure 14As shown. After a period of time (T = D1, where D1 can be equal to a first delay, second delay, etc. determined by the control processing unit 34 as the target delay 144), the control unit evaluates the control surface orientation at 150, the determination of which can depend on inputs from one or more control surface gauges 148, which in turn can receive inputs from a plurality of aircraft sensors and parameters 110, such as angle of attack 116 and / or aircraft yaw 126. After the control system evaluates the control surface orientation at 150, a signal can be sent back to the control processing unit 34 to determine if further action is required.

[0070] The control system 200 may also include Figure 14 other components not shown, such as fuel control valve 20 and power supply 30. Additionally, the control system 200 may include Figure 14 communication connections not shown. The components of the control system 200 operate in a frequency range of about 1 Hz to about 1000 Hz. For example, both the plasma actuator 14 and the fuel injector 18 can operate in a frequency range of about 1 Hz to about 200 Hz, or about 10 Hz to 150 Hz, or about 25 Hz to 100 Hz, or about 50 Hz to 75 Hz. Other sensors of the control system 200 (such as the plurality of aircraft sensors and parameters 110) as well as the airspeed indicator 38 and / or the ultrasonic sensor 104 can operate in a range of about 50 Hz to about 1000 Hz. The control system 100 operates in a frequency range equal to or higher than that of the system components, such as in a range of about 200 Hz to about 1000 Hz. In some embodiments, the control system 100 operates in a frequency range greater than 1000 Hz.

[0071] In operation, the plasma ignition combustion system and the control system 200 of this embodiment are used to balance thrust, horizontal acceleration, vertical acceleration, and angular acceleration by providing restoring forces on the control surfaces 12 of the aircraft and its structure. As shown in this embodiment Figure 2-10 the plasma ignition combustion system can be used on various surfaces of and on the structure of aircraft of different architectures and configurations, different architectures and configurations including but not limited to subsonic, supersonic, and hypersonic, and its structure includes wings, engines, exhaust nozzles of supersonic engines, and so on.

[0072] Figure 15 A control system 400 is shown, which can be used to control hypersonic aircraft such as Figure 11-13 the second hypersonic aircraft 300 in, and such as Figure 7-10Among those other supersonic and hypersonic aircraft. The control system 400 includes a control processing unit 34, an aircraft airspeed indicator 38 (not shown) and / or a local airspeed indicator 36 (not shown), and the control processing unit 34 receives at least one airspeed indication 106 that may be from an ultrasonic sensor 104 (not shown). The control unit 34 also receives an input from at least one flight command 108, which flight command 108 may include commands such as various aircraft maneuvers or commands to stabilize flight due to turbulence or changing environmental and / or operating conditions. The control unit 34 may also receive input signals from a plurality of aircraft orientation sensors and parameters 410, the plurality of aircraft orientation sensors and parameters 410 including but not limited to: angle of attack indication 116, aircraft attitude 120, gyroscope 123, aircraft yaw indication 126, aircraft angular acceleration 130, aircraft pitch angle indication 109, aircraft roll indication 111, lidar sensor 113, GPS sensor 115, navigation compass 117, and aircraft vertical acceleration 132. The control unit 34 may use the plurality of aircraft orientation sensors and parameters 410 to determine what actions to perform and the means for performing them. For example, if the aircraft deviates from a target control setting or orientation, or if a new heading is desired, the plurality of aircraft orientation sensors and parameters 410 may be used to determine, establish, and / or re-establish a new and / or desired heading.

[0073] Still referring to Figure 15 , the control system 400 may also include a plurality of flight stability sensors and parameters 430. The plurality of flight stability sensors and parameters 430 may transmit signals to the control processing unit 34, including but not limited to: turbulence sensor 124, vibration sensor 114, static pressure sensor 103, differential pressure sensor and / or indication 105, strain gauge 101, and microphone 107, as well as other sensors and parameters. The plurality of flight stability sensors and parameters 430 may be used to characterize various aerodynamic and acoustic aspects of flight, especially during supersonic flight. For example, the turbulence indicator 124 may indicate the presence of unstable conditions, crosswinds, and / or environmental disturbances; the static pressure sensor 103 and the microphone 107 may be used to characterize the magnitude and frequency of shock waves, as well as other characteristics such as the incident shock wave angle and shock wave geometry; the differential pressure indication 105 may be used to evaluate different shock wave characteristics at different locations on the aircraft; the strain gauge 101 may be disposed on or within various control surfaces 12 of the aircraft in order to evaluate the magnitude, frequency, and propagation mode of shock waves acting on the various control surfaces 12 of the aircraft, thereby deflecting and / or deforming them; and the vibration sensor 114 may be used to sense vibrations in the aircraft and its surfaces and components in order to evaluate at least one flight characteristic such as shock wave frequency and / or shock wave magnitude.

[0074] Still referring to Figure 15, the control system 400 may also include a plurality of aircraft control parameters 420, including but not limited to: ambient humidity 112, flight segment indication 118, ambient temperature 119 (and / or free air temperature), aircraft altitude 122, and aircraft control mode 128, as well as other control parameters. Each parameter and / or sensor among the plurality of aircraft orientation sensors and parameters 410, the plurality of aircraft control parameters 420, and the plurality of flight stability sensors and parameters 430 may also be used in association with other control modules and / or for purposes other than Figure 15 those shown in. For example, the aircraft altitude 122 may also be used to determine and / or establish flight stability and / or aircraft orientation. Additionally, and by way of non-limiting example, the aircraft altitude 122 may also be used to correct or adjust other parameters as needed.

[0075] Still referring to Figure 15 , the control processing unit 34 may use the airspeed indication 106 (which may include indicated airspeed and / or corrected or true airspeed) as an indication of the presence of a shock wave. For example, when the airspeed indication 106 signals that the aircraft is traveling at supersonic speed, a shock wave may be presumed to exist even without a direct shock wave measurement or indication from, for example, the plurality of flight stability sensors and parameters 430. The control processing unit 34 may or may not have an input from the flight command 108. For example, in cases where the desired heading and / or control mode includes maintaining the current heading, there may be no input from the flight command 108, but the control processing unit 34 will continue to actively control the aircraft, such as maintaining flight stability and aircraft orientation.

[0076] Still referring to Figure 15 , the control processing unit 34 determines plasma actuator targets for each of the first plasma position 15A, the second plasma position 15B, the third plasma position 15C, and any other plasma positions on the aircraft based on Figure 15 a number of inputs and possibly other inputs. For each of the plurality of plasma positions 15A - 15C, the control processing unit 34 determines a target duration 140, a target plasma frequency, a pulse rate and / or waveform 142, a target plasma delay 144 (and / or sequence timing, e.g., when a pattern or sequence for activating multiple plasma actuators 14 is desired), and a target plasma size 146. Each determined plasma target value for the first plasma position 15A is then transmitted to the first plasma actuator 14A, which in turn performs the desired target plasma actuation and / or routine. In Figure 15In [the figure], the target plasma value is shown only for the first plasma position 15A. However, the second plasma position 15B, the third plasma position 15C, and the fourth through Nth plasma positions will also have target plasma values, which are similarly transmitted to the corresponding plasma actuators 14B, 14C, etc. After a duration equal to the first duration (T = D1), the control system 400 evaluates the orientation of at least one control surface 12 and at least one parameter representative of flight stability at 440. The evaluation of flight stability can be at least partially based on a plurality of flight stability sensors and parameters 430, while the evaluation of the control surface 12 and / or the aircraft orientation can be at least partially based on a plurality of aircraft orientation sensors and parameters 410.

[0077] In operation, Figure 15The plasma-assisted control system 400 can operate at a frequency of about 500 Hz to about 50 kHz based on inputs from sensors that can operate at frequencies from dozens of Hz to dozens of MHz. For example, the control system 400 can operate at about 5 kHz to about 15 kHz, executing the entire control scheme or portions and / or modules thereof about 5,000 to about 15,000 times per second based on inputs from sensors having varying operating frequencies. In other embodiments, the control system 400 can operate at about 500 Hz to about 50 kHz. Some sensors may have a time delay, for example, due to a thermal lag associated with the time it takes for a temperature sensor to heat or cool. Other sensors, such as electronic GPS or lidar sensors, etc., can transmit and receive millions of signals per second. Some parts or modules of the control system 400 can operate at a different frequency than other parts. For example, due to continuously varying aerodynamic perturbations experienced under supersonic and hypersonic flight conditions, multiple plasma actuators 14 can operate at a higher frequency to accommodate the high frequencies associated with continuously maintaining stable flight. The multiple plasma actuators 14 can be actuated to generate plasma based on an electrical input signal that can be modulated very quickly. In other words, the control system 400 must operate at a high enough frequency to allow the system to respond appropriately and quickly to maintain vehicle stability. In one embodiment, the plasma-assisted control system 400 can receive at least one signal (the at least one signal indicating at least one flight characteristic, e.g., shock wave frequency and / or shock wave magnitude) from multiple flight stability sensors and parameters 430 at a control processing unit 34, and, based on the at least one flight characteristic, command at least one plasma actuator to generate plasma in response to the signal and be customized to provide stable flight. For example, the control processing unit 34 can command at least one of the multiple plasma actuators 14 to be actuated with a reaction force and / or stabilizing force having a magnitude and frequency commensurate with the corresponding shock wave magnitude and frequency sensed by the multiple flight stability sensors and parameters 430.

[0078] Figure 14 and 15 The control system can be used on subsonic, transonic, supersonic, and hypersonic vehicles, such as Figure 6-13Those shown in. Additionally, the plasma ignition combustion system 10 and the plasma assist control system 400 can be combined into a single system. For example, under supersonic flight conditions, when flight stability regulation and / or high-frequency aircraft control regulation is desired or required, the plasma can be actuated alone without fuel injection. In other embodiments, when a higher magnitude of control regulation is needed and / or when various aircraft maneuvers are requested from the flight command 108, the plasma can be used to ignite fuel. The activation of the plasma actuator 14 alone without fuel injection can be performed at a higher frequency than plasma ignition combustion. The fuel delivery system that delivers fuel to, for example, an aircraft engine or as a coolant for the control system on the aircraft can be combined with the systems and components (fuel supply 22, fuel control valve 20, fuel injector 18, etc.) of the plasma ignition combustion system 10 as much as possible.

[0079] Conventional aircraft may have movable surfaces for thrust vectoring in the exhaust nozzle and / or for use as control surfaces. However, these mechanical systems are heavy and relatively slow to respond (about 25 Hz for conventional hydraulic actuators). In contrast, the plasma ignition combustion system and the plasma assist control system of the present embodiment can alternatively be used on the outer surface of an aircraft, such as on the wings and tail, to provide control forces without the need for movable surfaces and associated systems. The plasma ignition combustion system and the plasma assist control system of the present embodiment can also operate at higher frequencies in the range of about 500 Hz to 15 kHz, thereby enabling stable hypersonic flight.

[0080] The advantages of the present embodiments are that they are able to achieve aircraft control at higher speeds (100s of Hz instead of ~10 Hz), which may be necessary in hypersonic situations. Also, the present embodiments may be lighter in weight than traditional control surfaces, which will improve aircraft efficiency. The fuel injector 18 and the plasma actuator 14 are synchronized such that each pulse and / or dispersion of fuel from the fuel injector 18 travels downstream to the plasma location 26 just as the plasma is formed, thus igniting the fuel. The synchronized activation of the fuel injector 18 and the plasma actuator 14 can occur dozens, hundreds, thousands, or even more times per second. For example, in some embodiments, the synchronized activation of the fuel injector 18 and the plasma actuator 14 can occur at an operating frequency of 10 kHz. In other embodiments, the synchronized activation of the fuel injector 18 and the plasma actuator 14 occurs between approximately 5 kHz and 15 kHz. In other embodiments, the synchronized activation of the fuel injector 18 and the plasma actuator 14 occurs between approximately 1 kHz and 5 kHz. In other embodiments, the synchronized activation of the fuel injector 18 and the plasma actuator 14 occurs between approximately 100 Hz and 1 kHz. By using multiple pairs of fuel injectors 18 and plasma actuators 14 arranged at different positions and orientations on one or more control surfaces 12, and by activating different pairs at different times, the aircraft can be controlled to address overcompensation of one pair by having the second pair provide a restoring force.

[0081] Embodiments herein can improve combustion stability and enable a plasma-stabilized combustion system to be used to control an aircraft with few or no moving parts (see U.S. Application 15 / 979,217 assigned to General Electric Company of Schenectady, New York). Embodiments herein can also be used on the leading edge, trailing edge, and / or other surfaces of at least one fin of a supersonic and / or hypersonic projectile. For example, plasma actuators and systems similar to those in the foregoing figures can be disposed along one or more leading edges of the fins of a hypersonic missile to control and / or stabilize its flight.

[0082] Exemplary embodiments of a plasma ignition combustion system, a plasma-assisted control system, and related components have been described in detail above. The system is not limited to the specific embodiments described herein, but rather the components of the system and / or the steps of the method can be used independently of and separately from other components and / or steps described herein. For example, the construction of the components described herein can also be used in combination with other processes and is not limited to practicing with the systems and related methods as described herein. Instead, the exemplary embodiments can be implemented and utilized in conjunction with many applications that desire supersonic combustion and / or supersonic aircraft control.

[0083] Although specific features of various embodiments of the present disclosure may be shown in some figures and not in others, this is merely for convenience. In accordance with the principles of the present disclosure, any feature of any figure may be referenced and / or claimed in combination with any feature of any other figure.

[0084] This written description uses examples to disclose embodiments of the present disclosure, including the best mode, and also enables those skilled in the art to practice the present disclosure, including making and using any device or system and performing any combined method. The patentable scope of the embodiments described herein is defined by the claims and may include other examples that occur to those skilled in the art. If these other examples have structural elements that are not different from the literal language of the claims, or if they include equivalent structural elements that are not materially different from the literal language of the claims, then these other examples are intended to fall within the scope of the claims.

[0085] Further aspects of the invention are provided by the subject matter of the following clauses:

[0086] 1. An aircraft, comprising: a first leading edge that defines a forward edge of a left aircraft wing; a second leading edge that defines a forward edge of a right aircraft wing; a plurality of plasma actuators disposed along the first leading edge and the second leading edge; a control processing unit communicatively coupled to each of the plurality of plasma actuators; and at least one flight stability sensor communicatively coupled to the control processing unit, wherein the control processing unit commands at least one of the plurality of plasma actuators to generate plasma in response to a signal from the at least one flight stability sensor.

[0087] 2. The aircraft according to any preceding clause, wherein the at least one flight stability sensor detects at least one aerodynamic characteristic during at least one of supersonic flight and hypersonic flight.

[0088] 3. The aircraft according to any preceding clause, wherein the at least one aerodynamic characteristic includes at least one of shock wave magnitude and shock wave frequency.

[0089] 4. The aircraft according to any preceding clause, further comprising at least one aircraft orientation sensor including at least one of a GPS sensor, lidar, and a gyroscope.

[0090] 5. The aircraft according to any preceding clause, wherein the at least one flight stability sensor includes at least one of a turbulence sensor, a strain gauge, and a microphone.

[0091] 6. An aircraft according to any preceding item, wherein the at least one flight stability sensor includes at least one of a vibration sensor, a static pressure sensor, and a differential pressure sensor.

[0092] 7. An aircraft according to any preceding item, wherein the at least one flight stability sensor is disposed on the underside of the aircraft.

[0093] 8. An aircraft according to any preceding item, further comprising at least two flight stability sensors, wherein at least one flight stability sensor is disposed on the underside of the aircraft, near the nose of the aircraft.

[0094] 9. An aircraft according to any preceding item, wherein the control processing unit adjusts at least one of a plasma frequency and a plasma size based on at least one of the shock wave magnitude and the shock wave frequency.

[0095] 10. An aircraft according to any preceding item, wherein the at least one aircraft orientation sensor is disposed on at least one of a left wing of the aircraft and a right wing of the aircraft.

[0096] 11. An aircraft according to any preceding item, wherein the at least one aircraft orientation sensor is disposed near at least one of a nose of the aircraft and a rear end of the aircraft.

[0097] 12. An aircraft according to any preceding item, further comprising at least one aircraft orientation sensor, the at least one aircraft orientation sensor including at least one of a GPS sensor, a lidar, and a gyroscope, wherein the at least one flight stability sensor includes at least one of a turbulence sensor, a strain gauge, and a microphone, wherein at least one flight stability sensor is disposed on the underside of the aircraft, near the nose of the aircraft, and wherein the aircraft is capable of supersonic flight or hypersonic flight.

[0098] 13. An aircraft control system, comprising: a control processing unit; at least one sensor communicatively coupled to the control processing unit; and at least one plasma actuator disposed near a leading edge of an aircraft wing, the at least one plasma actuator communicatively coupled to the control processing unit, wherein the control processing unit commands the at least one plasma actuator to generate plasma in response to at least one signal from the at least one sensor.

[0099] 14. A control system according to any preceding item, wherein the at least one signal represents at least one aerodynamic characteristic during supersonic flight or hypersonic flight.

[0100] 15. According to the control system of any preceding item, wherein the at least one sensor includes at least one of a turbulence sensor, a strain gauge, a microphone, a vibration sensor, a static pressure sensor, and a differential pressure sensor.

[0101] 16. According to the control system of any preceding item, further comprising at least one aircraft orientation sensor, wherein the at least one orientation sensor includes at least one of a GPS, a lidar, and a gyroscope.

[0102] 17. According to the control system of any preceding item, further comprising: an airspeed indicator; and at least one of a temperature sensor, a humidity sensor, and an altimeter.

[0103] 18. According to the control system of any preceding item, wherein the at least one signal represents at least one of a shock wave magnitude and a shock wave frequency.

[0104] 19. According to the control system of any preceding item, wherein the control system is capable of operating in a range of about 500 Hz to about 50 kHz.

[0105] 20. According to the control system of any preceding item, further comprising: an airspeed indicator; and at least one aircraft orientation sensor, wherein the at least one orientation sensor includes at least one of a GPS, a lidar, and a gyroscope, wherein the control system is capable of operating in a range of about 500 Hz to about 50 kHz, wherein the at least one sensor includes at least one of a strain gauge, a microphone, and a vibration sensor, and wherein the at least one signal represents a shock wave magnitude and frequency.

Claims

1. An aircraft, characterized in that, Comprising: A combustion engine, the combustion engine including a fuel injector; A plurality of plasma actuators, the plurality of plasma actuators being disposed in the combustion engine, downstream of the fuel injector; A control processing unit, the control processing unit being communicatively coupled to each of the plurality of plasma actuators; And At least one sensor, the at least one sensor being communicatively coupled to the control processing unit, Wherein, in response to a signal from the at least one sensor, the control processing unit commands the fuel injector and at least one of the plurality of plasma actuators to generate plasma.

2. The aircraft according to claim 1, wherein, Wherein the at least one sensor detects at least one aerodynamic characteristic during at least one of supersonic flight and hypersonic flight.

3. The aircraft according to claim 2, characterized in that, Wherein the at least one aerodynamic characteristic includes at least one of shock wave magnitude and shock wave frequency.

4. The aircraft according to claim 1, characterized in that, Wherein the control processing unit commands the fuel injector to adjust at least one of the injection angle, the fuel mass flow rate, or the fuel pulse rate.

5. The aircraft according to claim 1, characterized in that, Wherein the at least one sensor includes at least one of a turbulence sensor, a humidity sensor, or a gyroscope.

6. The aircraft according to claim 1, characterized in that Wherein the at least one sensor includes at least one of a vibration sensor, a static pressure sensor, and a differential pressure sensor.

7. The aircraft according to claim 1, characterized in that, Wherein the at least one sensor is disposed on the underside of the aircraft.

8. The aircraft according to claim 7, characterized in that, Further comprising at least two sensors, wherein at least one of the at least two sensors is disposed on the underside of the aircraft, near the aircraft nose.

9. The aircraft according to claim 1, characterized in that, Wherein the control processing unit is configured to adjust at least one of the plasma delay, the plasma pulse rate, or the plasma size.

10. The aircraft according to claim 1, characterized in that, Wherein the control processing unit is configured to determine a plurality of control target values, the plurality of control target values including at least one of a target injection angle, a target fuel mass flow rate, a target fuel pulse rate, a target duration, a target plasma pulse rate, a target delay, or a target plasma size.

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