阅读说明：本技术 用于飞行器的控制系统 (Control system for an aircraft ) 是由 尼古拉斯·威廉·雷泰 纳伦德拉·迪甘伯·乔希 于 2019-07-19 设计创作，主要内容包括：一种飞行器,包括限定左飞行器机翼的前向缘的第一前缘,限定右飞行器机翼的前向缘的第二前缘,沿第一和第二前缘设置的多个等离子体致动器,通信地联接到每个等离子体致动器的控制处理单元,和通信地联接到控制处理单元的至少一个飞行稳定性传感器。控制处理单元响应于来自飞行稳定性传感器的信号,命令至少一个等离子体致动器产生等离子体。(an aircraft includes a 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 and second leading edges, a control processing unit communicatively coupled to each plasma actuator, and at least flight stability sensors communicatively coupled to the control processing unit.)

An aircraft of the type , comprising:

an leading edge, the 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 th leading edge and the second leading edge;

a control processing unit communicatively coupled to each plasma actuator of the plurality of plasma actuators; and

at least flight stability sensors, the at least flight stability sensors communicatively coupled to the control processing unit,

wherein the control processing unit commands at least plasma actuators of the plurality of plasma actuators to generate plasma in response to signals from the at least flight stability sensors.

2. The aircraft of claim 1 wherein the at least flight stability sensors detect at least aerodynamic properties during at least of supersonic flight and hypersonic flight.

3. The aircraft of claim 2 wherein the at least aerodynamic properties include at least of shock wave size and shock wave frequency.

4. The aircraft of claim 1 wherein step comprises at least aircraft orientation sensors, the at least aircraft orientation sensors comprising at least of a GPS sensor, a lidar, and a gyroscope.

5. The aircraft of claim 1, wherein the at least flight stability sensors include at least of turbulence sensors, strain gauges, and microphones.

6. The aircraft of claim 1, wherein the at least flight stability sensors include at least of a vibration sensor, a static pressure sensor, and a differential pressure sensor.

7. The aircraft of claim 1, wherein the at least flight stability sensors are disposed on an underside of the aircraft.

8. The aircraft of claim 7 wherein step comprises at least two flight stability sensors, wherein at least flight stability sensors are disposed on the underside of the aircraft, near an aircraft nose.

9. The aircraft of claim 3, wherein the control processing unit adjusts at least of plasma frequency and plasma magnitude based on at least of the shock wave magnitude and the shock wave frequency.

10. The aircraft of claim 4 wherein the at least aircraft orientation sensors are disposed on at least of the aircraft left wing and the aircraft right wing.

Technical Field

The subject matter disclosed herein relates to an aircraft and a method of controlling an aircraft.

Background

Supersonic and hypersonic aircraft typically use control surfaces as the control means actuators and other mechanisms for positioning the control surfaces are typically used to control the control surfaces.

Aircraft flight stability and control at supersonic and hypersonic speeds is a multifaceted area that involves a balance of several factors, largely due to the speed at which the aircraft is flying. At supersonic and hypersonic speeds, aircraft are subject to high frequency disturbances and may require faster response rates than can be achieved by conventional control surfaces (e.g., ailerons, elevators, and rudders). Additionally, even at lower speeds, aircraft control surfaces (e.g., movable supersonic engine exhaust nozzles) are very heavy, which reduces aircraft efficiency.

Disclosure of Invention

Aspects of the present embodiment are summarized below. These embodiments are not intended to limit the scope of the claimed embodiments, but rather, they are intended to provide only a brief summary of possible forms of embodiments. Furthermore, embodiments may include various forms that may be similar to or different from the embodiments set forth below, corresponding to the scope of the claims.

In embodiments, the aircraft includes a th 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 th and second leading edges, a control processing unit communicatively coupled to each of the plasma actuators, and at least flight stability sensors communicatively coupled to the control processing unit.

In another embodiments, the aircraft control system includes a control processing unit, at least sensors communicatively coupled to the control processing unit, and at least plasma actuators disposed proximate to a leading edge of the aircraft wing, the plasma actuators communicatively coupled to the control processing unit.

Drawings

These and other features, aspects, and advantages of the present disclosure will become 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:

FIG. 1 is a side schematic view of a plasma-ignited combustion system;

FIG. 2 is a side schematic view of a plasma-ignited combustion system with a schematic view of a control system;

FIG. 3 is a side schematic view of a wing mounted plasma ignition combustion system;

FIG. 4 is a side schematic view of a plasma-ignited combustion system mounted on an engine;

FIG. 5 is a rear-to-front view of an engine exhaust ring including a plasma-ignited combustion system;

FIG. 6 is a top view of a subsonic vehicle including a plasma ignition combustion system;

FIG. 7 is a top view of a supersonic aircraft including a plasma-ignited combustion system;

FIG. 8 is a top view of a hypersonic aerial vehicle including a plasma ignition combustion system;

FIG. 9 is a front view of a hypersonic aerial vehicle including a plasma ignition combustion system;

FIG. 10 is a side view of a hypersonic aircraft including a plasma ignition combustion system;

FIG. 11 is a front view of a hypersonic aerial vehicle including a plasma-assisted control system;

FIG. 12 is a perspective view of a hypersonic aerial vehicle including a plasma-assisted control system;

FIG. 13 is a side view of a hypersonic aerial vehicle including a plasma-assisted control system;

FIG. 14 is a schematic diagram of a control system of the plasma-ignited combustion system; and

FIG. 15 is a schematic diagram of a control system for a plasma-assisted control system.

The drawings provided herein are intended to illustrate features of embodiments of the disclosure, unless otherwise indicated, these features are believed to be applicable to various systems including or more embodiments of the disclosure.

Detailed Description

In the following specification and claims, reference will be made to a number of terms, which shall be defined to have the following meanings.

The singular forms "", "" and "the" include plural referents unless the context clearly dictates otherwise.

"optional" or "optionally" means that the subsequently described event or circumstance may or may not occur, and that the description includes instances where the event occurs and instances where it does not.

Thus, a value modified by or more terms (e.g., "about" and "substantially") is not to be limited to the precise value specified.

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

As used herein, the term "circumferential" refers to directions or directions around (and tangential to) the circumference of a ring of a combustor or, for example, a circle defined by the swept area of a turbine blade.

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

As used herein, the term "plasma" refers to a gas that has been electrically conductive by heating or subjecting it to an electromagnetic field, where the long range electromagnetic field dominates the behavior of the species.

As used herein, the term "cold plasma" refers to a plasma that: where the characteristic temperature of the electrons is much higher than that of the "heavy" particles (i.e., neutral and ionized molecules and atoms), rather than being in thermal equilibrium (i.e., "hot" plasma).

As used herein, the term "plasma actuator" refers to a plasma generating device that generates plasma that acts on the control surfaces of an aircraft, in connection with either fueled (plasma-assisted combustion) or unfueled (plasma-assisted control). the plasma actuator may help stabilize and/or enhance combustion, and may also generate plasma that acts on or more control surfaces of the aircraft, as well as interact with the aerodynamic conditions of the aircraft via flight.

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

As used herein, the term "scramjet engine" refers to a variation of ramjet engines in which combustion occurs in a supersonic gas stream therein.

As used herein, the term "subsonic" refers to a velocity less than the speed of sound, less than about mach 1. As used herein, the term "transonic speed" refers to a speed of about mach 0.8 to about mach 1.2. As used herein, the term "supersonic" refers to velocities greater than sonic velocity, and more specifically, velocities from about Mach 1 to about Mach 5. As used herein, the term "hypersonic" refers to velocities above about mach 5.

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

FIG. 1 illustrates a plasma-ignited combustion system 10 of the present embodiment that utilizes a control surface 12, the control surface 12 may be a substrate provided with components of the claimed embodiment, additionally, fluid may flow over the outer surface of the substrate or control surface 12.

Several plasma actuator arrangements are possible. A "microwave plasma" may be generated by injecting microwave power into a gas (e.g., air or a fuel-air mixture), wherein the microwave power is preferentially coupled to a region of the gas that has been ionized and conducted, e.g., in front of a flame, thereby adding energy to the front of the flame and increasing the local heat release rate.

Microwave plasma can also be generated in air or air-fuel mixtures upstream of the flame zone, where it can act as a plasma source generating reactive radicals that flow into and enhance the combustion process without having to deposit energy into the normal gas heating. The plasma generated may be cold or hot. Gas may be introduced into the combustion zone by the plasma (e.g. from the side wall of the combustion chamber), this arrangement sometimes being referred to as a "plasma tube". The microwave frequency may be in the range of about 0.3GHz to about 300 GHz.

The plasma tube plasma actuator may also be powered by other means such as radio frequency induction (in the range of about 3kHz to about 0.3 Ghz), or by electrodes driven by direct current or alternating current. Thermal spraying occurs in the combustion chamber to stabilize and control the flame. Radio frequency or microwave energy may be generated by power electronics or magnetrons and delivered to a desired area in the engine via a transmission line, such as a coaxial cable, or other suitably shaped structure such as a waveguide or "applicator".

A spark plasma may be generated to stabilize the flame in a manner similar to a diffusion pilot flame in a burner, where the overall fuel-air ratio is lean (i.e., where oxygen remains after the fuel is fully combusted). In this arrangement, the plasma acts as a local heat source. Such plasmas may be generated by an intermittent "spark" plasma (e.g., a spark plug igniter), or by a continuous "arc" plasma maintained between two electrodes by controlling the current through the circuit. Spark plasma can also be achieved via intermittent laser spark plasma or (continuous laser arc plasma) generated by focusing laser power into the gas volume.

The cold plasma can be kept in the gas by controlling the power deposition so that energy is not transferred from the electrons to the heavy particles because of low pressure, low power density, or energy is applied for a short time (pulsed). The resulting plasma creates reactive radicals that flow into and enhance the combustion process without having to deposit energy into the normal gas heating. The nanosecond plasma may also be configured with a gas flow as a plasma tube.

FIG. 2 illustrates the plasma-ignited combustion system 10 of the present embodiment using a control surface 12, fuel is injected at an injection location 16 through at least fuel injectors 18. one or more fuel injectors 18 are in fluid communication with a fuel control valve 20, the fuel control valve 20 controlling the amount of fuel flowing to the fuel injectors 18. A fuel supply 22 is in fluid communication with the fuel control valve 20 and upstream of the fuel control valve 20. in 83 embodiments, a fuel injector clutch 24 may be mechanically coupled to the fuel injectors 18 to adjust the angle at which fuel is dispersed from the fuel injectors 18 as needed depending on the ambient airflow and operating conditions of the aircraft.

Still referring to fig. 2, the fuel injector 18 and the plasma actuator 14 are disposed nearly flush with the control surface 12, limiting potential detrimental effects such as increases in drag when they are not functioning the plasma-ignited 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 forces from the combustion zone C act the flow surface 28 may be, for example, a thin half nozzle or crescent in the embodiment of fig. 2 the flow surface 28 may be semi-cylindrical (i.e., "half pipe") shaped, hemispherical, or semi-elliptical, conical, semi-conical, frusto-conical, sinusoidal, and other contour 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 piecewise planar including a plurality of planar surfaces assembled from respective planar segments and arranged at various angles in other embodiments the plurality of flow surfaces 28 may be used in other embodiments the separate flow surface 28 may not be needed in other embodiments the control surface 12 or the control surface profile may be designed to avoid the need for the flow surface 28 to be shaped near the end of the plasma-ignited combustion zone C.

Still referring to FIG. 2, the plasma-ignited combustion system 10 may include a power supply 30 that is electrically coupled to a plasma actuator 14 for generating a plasma, the plasma-ignited combustion system 10 may include a control processing unit 34, the control unit or control processing unit 34 may be communicatively coupled to the fuel supply 22, the fuel control valve 20, the fuel injector bite 24, each of the plasma actuator 14 and the power supply 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 the embodiment of FIG. 2, the components that may be communicatively coupled to each other are connected via dashed lines.

For example, in embodiments, the plasma will be generated such that it exists near the plasma location 26 just prior to the arrival of fuel from the injection location 16. accordingly, it may be desirable to time the injection of fuel by the fuel injector 18 with the plasma generation by the plasma actuator 14 to minimize energy losses (via fuel losses and unused plasma). The local airspeed indicator 36 may be used to roughly estimate the time of flight of fuel from the injection location 16 to the plasma location 26, since the th distance 40 between the injection location 16 and the plasma location 26 may be fixed and therefore a known amount, since the local indicator 36 is disposed at the control surface 12 downstream of the injection location 16 and upstream of the plasma location 26, since boundary layer and other fluid effects may exist near the control surface 12, and since these effects may vary with operating and environmental conditions, the local airspeed indicator 36 may be able to accurately determine how quickly the fuel will travel between the 362 nd distance 3640 and the plasma location th distance 3640, since these effects may vary with operating and environmental conditions.

The local airspeed indicator 36 shown in FIG. 2 may be an ultrasonic sensor, or a calibrated static pressure type sensor for rough estimation of airflow. Local airspeed indicator 36 may also be other types of sensors, including a flat probe sensor, a pitot tube sensor, a differential pressure sensor, and/or any other sensor that may be used to measure flow across a surface. Ultrasonic sensors are capable of distinguishing between the fuel velocity and the air velocity flowing through the sensor under conditions where there is a velocity difference between the two fluids. Other sensors that do not distinguish between fuel velocity and air velocity flowing therethrough may still accurately predict the time of flight of fuel flowing from the injection site 16 upstream of the plasma site 26 by correlating fuel velocity with air velocity. The plasma-ignited combustion system 10 may also include airflow indications 38 from different locations and/or from the aircraft controller 32. As described above, the local airspeed indicator 36 may have the benefit of accounting for boundary layer conditions. However, in embodiments where the airspeed of the aircraft is highly correlated with the time of flight of the fuel flowing upstream from the injection location 16 to the plasma location 26, an airflow indication 38 from the aircraft controller 32 may be sufficient. Under flight conditions when the airflow direction is misaligned from the line connecting the injection location 16 to the 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 the fuel injectors 18 may be adjusted by the fuel injector occluder 24 to ensure that fuel dispersed by the fuel injectors 18 reaches the plasma location 26. Additionally, guides, tubes, vanes, and/or other devices (not shown) may be employed to direct the fuel dispersed by the fuel injectors 18 to the plasma location 26.

Fig. 3 shows an embodiment of a plasma ignition combustion system 10 on an airfoil control surface 12 the airfoil control surface 12 shown in fig. 3 may be a wing of an aircraft, other airfoil structures on an aircraft, an airfoil aircraft, and other surfaces serving as control surfaces 12 the embodiment of fig. 3 includes fuel injected at an injection location 16 upstream of a plasma location 26 via at least fuel injectors 18, wherein plasma is generated via at least plasma actuators 14 at least plasma actuators 14 ignite the fuel, creating a combustion zone c at a control surface downstream end 12 'air flows over the control surface 12 in direction B the embodiment of fig. 3 may also include several other system components of fig. 2 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 occluder 24 in other embodiments, components of the plasma ignition combustion system 10 will be disposed on the underside 12 "of the control surface, but not on the control surface 12" on the topside of the control surface 12 ", or on the topside control surface 12"', components of the plasma ignition control surface 12, or on the topside control surface 12.

FIG. 4 illustrates an embodiment of a plasma-ignited combustion system 10 in an application of supersonic combustion engine 41 the supersonic combustion engine 41 illustrated in FIG. 4 may include an air tube inlet 42 that provides a main burner section 42 upstream of a flared exhaust section 48, upstream of a diverging section 46. supersonic combustion engine 41 may be generally axisymmetric about an engine centerline CL. the flared exhaust section 48 may include or more control surfaces 12 that form an annular exhaust port and diverge radially outward from the engine centerline CL as they extend aft in direction B. the embodiment of FIG. 4 includes fuel injected via at least fuel injectors 18 at an injection location 16 upstream of a plasma location 26, where plasma is generated via at least plasma actuators 14. at least plasma actuators 14 ignite the fuel, generate a combustion zone C at the control surfaces 12. the embodiment of FIG. 4 may also include several other system components of FIG. 2, including but not limited to power supply 30, flow surface 28, local airspeed sensor 36, fuel supply valve 22, fuel control valve 20, combustion control unit 20, combustion nozzle arrangement to reduce the thrust vector of the combustion engine, and/combustion engine arrangement of the aircraft may be configured to reduce the combustion engine thrust forces required in order to reduce combustion engine efficiency, and/combustion engine efficiency, and/combustion engine efficiency, and to reduce the combustion engine efficiency of the aircraft.

Fig. 5 shows a backward-forward looking embodiment of the plasma-ignited combustion system 10 in a supersonic combustion engine 41 application similar to that of fig. 4. In other embodiments, the plasma-ignited combustion system 10 may be in a gas turbine engine or other subsonic engine. The embodiment of fig. 5 is viewed from the aft end of the supersonic combustion engine 41 through a flared exhaust portion 48. A plurality of plasma-ignited combustion systems 10 are circumferentially spaced around the exhaust annulus of the supersonic combustion engine 41. In the embodiment of fig. 5, each of the plurality of plasma-ignited combustion systems 10 includes a plasma actuator 14, a flow surface 28, and other system components shown in fig. 2, including, but not limited to, a power supply 30, a local airspeed sensor 36, 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 clutch 24. The embodiment of fig. 5 includes 8 plasma-ignited combustion systems 10 that are substantially evenly spaced around the ring of supersonic combustion engines 41 at about 45 degree intervals. In other embodiments, other numbers of plasma-ignited 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 ring, and/or a different number of flow surfaces 28 than the number of plasma actuators 14 may be used. By asymmetrically operating the plasma-ignited combustion system 10, a net force in any desired direction is possible. This force may act on the control surface 12 within the engine. In other embodiments, the force may act on surfaces within the engine, which in turn may act on the control surface 12 of the aircraft.

FIG. 6 illustrates a top view of an exemplary subsonic vehicle 51. The plasma ignition combustion system 10 (not shown) of the present embodiment may be used in subsonic aircraft 51 applications. For example, the plasma-ignited combustion system 10 may be disposed on a surface of a subsonic vehicle 51, including, but not limited to, a right wing 50, a left wing 52, a right nacelle 54, a left nacelle 56, a right horizontal stabilizer 58, a left horizontal stabilizer 60, a vehicle fuselage 62, a vertical stabilizer 64 (left and/or right), a right winglet 66, and/or a left winglet 68. In addition, the plasma ignition combustion system 10 may be provided on surfaces (and other surfaces) on the underside of the subsonic vehicle 51 corresponding to those described above.

Fig. 7 shows a top view of an exemplary supersonic aircraft 61. The plasma-ignition combustion system 10 (not shown) of the present embodiment may be used in supersonic aircraft 61 applications. For example, plasma-ignited combustion system 10 may be disposed on a surface of supersonic aircraft 61, including, but not limited to, 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 portion 84. In addition, the plasma ignition combustion system 10 may be provided on surfaces (and other surfaces) on the underside of the ultrasonic vehicle 61 corresponding to those described above.

FIG. 8 illustrates a top view of an exemplary hypersonic aircraft 71. the plasma-ignited combustion system 10 of the present embodiment (not shown) may be used in hypersonic aircraft 71 applications, for example, the plasma-ignited combustion system 10 may be disposed on th hypersonic aircraft 71 surface, including but not limited to right horizontal surface 86, left horizontal surface 88, right vertical surface 90 (right outboard and/or left inboard), left vertical surface 91 (outer left and/or right inboard), aircraft body aft portion 92, aircraft body mid portion 94, and aircraft body forward portion 96. additionally, the plasma-ignited combustion system 10 may be disposed on corresponding surfaces (among others) on the underside of th hypersonic aircraft 71 corresponding to those described above.

FIG. 9 illustrates a front view of a hypersonic aircraft 71 including an air inlet 98 disposed on the underside of the hypersonic aircraft 71 the plasma ignition combustion system 10 (not shown) of the present embodiment may be used in hypersonic aircraft 71 applications, for example, the plasma ignition combustion system 10 may be disposed on the surface of the hypersonic aircraft 71 including, but not limited to, the right horizontal surface 86, the left horizontal surface 88, the right vertical surface 90 (on and/or to either side ) and the left vertical surface 91 (on and/or to either side ).

FIG. 10 illustrates a side view of the hypersonic aerial vehicle 71 including an air inlet 98 disposed on the underside of the hypersonic aerial vehicle 71. the plasma ignition combustion system 10 (not shown) of the present embodiment may be used in the hypersonic aerial vehicle 71 application, for example, the plasma ignition combustion system 10 may be disposed on the surface of the hypersonic aerial vehicle 71 including, but not limited to, the left horizontal surface 88, the left vertical surface 91 (on either or both sides ), the lower upstream portion 100 and the lower downstream portion 102. the air inlet 98 is disposed between the lower upstream portion 100 and the lower downstream portion 102.

FIG. 11 illustrates a front view of a second hypersonic aerial vehicle 300 having a different configuration than the hypersonic aerial vehicle 71. the second hypersonic aerial vehicle 300 includes a left leading edge 304 defining a forward edge of a left aerial vehicle wing 312. the left leading edge 304 may extend forward to an aircraft nose 311, where the left leading edge 304 may converge with a right leading edge 306, and the right leading edge 306 defines a forward edge of an aircraft right wing 310. the second hypersonic aerial vehicle 300 may include an inlet 302 disposed on an aircraft underside 308. a plurality of plasma actuators 14 may be disposed along each of the left leading edge 304 and the right leading edge 306. a plurality of plasma actuators 6314 may be flush with the leading edges 304,306 such that they do not extend or protrude from the aerial vehicle into the oncoming airstream. additionally, a plurality of plasma actuators 14 may be disposed at the leading edges 304,306 or near the leading edges 304,306 such that they are positioned to generate plasma at the leading edges 304,306, where shock waves are most likely to be present at the leading edges 304, 306. in other words, a plurality of plasma actuators 14 need not be disposed precisely at the length of the leading edges 304,306, as long as the length of the aircraft nose 304, 316 is defined by at least about five percent of the length of the aircraft nose 304, 316.

In operation, shock waves may propagate along the aircraft underside 308, along the top and bottom of the left wing 310 and the left wing 312, and along other surfaces of the second hypersonic aircraft 300 as the second hypersonic aircraft 300 reaches supersonic and/or hypersonic speeds.

Still referring to FIG. 11, the second hypersonic aerial vehicle 300 may include or more flight stability sensors 301 disposed on the vehicle underside 308. or more flight stability sensors 301 may be used to sense at least aerodynamic characteristics of the second hypersonic aerial vehicle 300 under given operating conditions. for example, or more flight stability sensors 301 may be comprised of an airspeed indicator that indicates when a supersonic flight condition exists and, therefore, the presence of a shockwave is apparent. in another embodiments, or more flight stability sensors 301 may include static pressure sensors that indicate the presence and/or magnitude of a shockwave and the frequency at which the shockwave propagates along the vehicle underside 308 and/or the left and right leading edges 304 and 306. or more flight stability sensors 301 may also be disposed along the left and right leading edges 304 and 306 where the shockwave is most likely to form and/or act upon a force.

Still referring to FIG. 11, the aircraft control system may use or more flight stability sensors 301 to control the frequency and/or magnitude of plasma generated by plasma actuators 14. for example, where the magnitude of the shock wave is proportional to the aircraft airspeed, or more flight stability sensors 301 may be used as static pressure sensors to measure the magnitude of the shock wave, resulting in an approximation of the airspeed.

FIG. 12 illustrates a perspective view of a second hypersonic aircraft 300 including a left wing 312, a right wing 310, a left leading edge 304, a right leading edge 306, an aircraft 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 aircraft 300 may also include an aircraft body apex line 316 extending the length of the aircraft. the aircraft body apex line 316 may define an intersection between the right wing 310 and the left wing 312. the aircraft body apex line 316 may be defined by a single line, or alternatively, may be a curved and/or slightly smooth or rounded portion of the top of the second hypersonic aircraft 300 where the left wing 312 and the right wing 310 intersect or intersect. the second hypersonic aircraft 300 also includes a trailing end 318 defined by a left trailing edge 322 and a right trailing edge 324. the left trailing edge 322 and a right trailing edge 324 also define the trailing edges of the left and right wings 310, 312. an exhaust port 320 may also be disposed in the trailing end 318. the aircraft 326 defines the left wing 312, the intersection between the right wing 310 and the right trailing edges 306 and the intersection 306, and the right leading edge 306 may form an acute angle of about 35 in an embodiment of about 35, degrees between the left leading edge 306 and right leading edge 306. the left leading edge 306. the angle may form an acute angle of about an angle of about angle between about an acute angle of about in an embodiment of about degrees in the right angle between the left leading edge 306 and a right angle of the right leading edge 306. the right angle of about 3630 leading edge 306. the leading edge 306 and a right angle of about a leading edge 304.

Still referring to fig. 12, the second hypersonic aircraft 300 may include a th sensor 328 disposed at or near the aircraft nose 311, a second sensor 330 disposed at or near the aircraft apex 326 (i.e., located on the top surface of the aircraft near the center of the rear end of the aircraft), 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 aircraft 300 may also include other sensors 336 at other locations including corresponding locations on the bottom surface of the aircraft. sensors 328,330,332,334,336 may be used to establish various orientations and frames of reference of the aircraft during flight. for example, sensors 328,330,332,334,336 may be used to establish aircraft angle of attack 116, aircraft yaw 126, aircraft angular acceleration 130, aircraft vertical acceleration 132, aircraft vibration, aircraft attitude 120, aircraft altitude 122, and other parameters.each of sensors may be a gyroscope, a GPS sensor, an accelerometer, a laser radar, a proximity sensor, a communication device for establishing a location relative to the frame of other than satellites, a barometer, a navigation light, a gyro light guide height 122, and other parameters.3. each of the sensors may be a gyro, a gyro sensor, a GPS, a gyro sensor, a gyro.

Still referring to FIG. 12, each of the sensors 328,330,332,334,336 may be used alone or in conjunction with one another to establish at least aspects of the aircraft orientation.Sensors 328,330,332,334,336 may be tuned such that they operate over a frequency range of 1kHz to 5MHz, producing 1000 to millions of orientation signals per second.the orientation signals from sensors 328,330,332,334,336 may be used by an aircraft control system to adjust the orientation of the aircraft via a plurality of plasma actuators 14. by asymmetrically activating plasma actuators 14, the aircraft control system may cause a net force to act on the aircraft resulting in a 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 may generate a net force on the aircraft resulting in a change or adjustment of the aircraft yaw 126 (not shown), or a roll 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 aft end 318, the control system may generate a net force on the aircraft resulting in a change or adjustment of the aircraft angle of attack 116 (not shown).

Fig. 13 illustrates a side view of a second hypersonic aircraft 300 including 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.

FIG. 14 shows a control system 200 that may be used to control plasma-ignited combustion system 10. the control system includes a control unit 34, the control unit 34 receiving at least airspeed indications 106, which airspeed indications 106 may be from ultrasonic sensors 104, aircraft airspeed indicators 38, and/or local airspeed indicators 36. the control unit 34 also receives input from at least flight commands 108, which flight commands 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 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, vibration sensors 114, angle of attack indications 116, flight segment indications 118, aircraft attitude 120, aircraft altitude 122, gyroscopes 123, turbulence sensors 124, aircraft indications 126, aircraft control modes 128, aircraft angle 130, and aircraft acceleration vertical acceleration 132. the control unit 34 may use the plurality of aircraft sensors and parameters 110 to determine which actions to perform and devices for performing, for example, if excessive vibration or activation of a device for performing, may reduce the effects on the plasma ignition system 10, or on the plurality of aircraft acceleration control elements 6335, which may be dependent on the turbulence control system or on other factors 3612.

Still referring to fig. 14, the control unit may determine a plurality of control target values including, but not limited to, a target injection angle 134 (i.e., angle of injection), a target fuel mass flow rate 135, a target fuel pulse rate 138, a target duration 140 (i.e., duration over which one or more plasma-ignited combustion systems 10 may 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 time of flight (or estimated time of flight) of fuel flowing from injection location 16 to plasma location 26), and a target plasma size 146. these target values may be communicated to fuel injector bite 24, fuel injector 18, and/or plasma actuator 14, as shown in fig. 14. after a lapse of periods of time (T ═ D1, where D1 may equal the , second delay, etc. determined by control processing unit 34 as target delay 144), the control unit may evaluate the control surface orientation at 150, which determination may depend on input control surface orientation signals from or a plurality of input control surface sensors 148, may in turn evaluate the yaw control surface orientation of aircraft 116, and/or may be sent back to control unit 126, for example, to determine whether yaw control system input control parameters at , and/or to be sent back to the yaw control unit 116.

Control system 200 may also include other components not shown in FIG. 14, such as fuel control valve 20 and power supply 30. furthermore, control system 200 may include communication connections not shown in FIG. 14. the components of control system 200 operate within a frequency range of about 1Hz to about 1000 Hz. for example, both plasma actuator 14 and fuel injector 18 may operate within a frequency range of about 1Hz to about 200Hz, or about 10Hz to 150Hz, or about 25Hz to 100Hz, or about 50Hz to 75 Hz. other sensors of control system 200 (e.g., plurality of aircraft sensors and parameters 110) as well as airspeed indicator 38 and/or ultrasonic sensor 104 may operate within a range of about 50Hz to about 1000 Hz. control system 100 operates within a frequency range equal to or higher than system components, such as within a range of about 200Hz to about 1000 Hz. in some embodiments of , control system 100 operates within a frequency range of greater than 1000 Hz.

In operation, the plasma ignition combustion system and control system 200 of the present embodiment is used to balance thrust, horizontal acceleration, vertical acceleration, and angular acceleration by providing a restoring force on the control surface 12 of the aircraft and its structure. As shown in fig. 2-10 of the present embodiment, the plasma-ignited combustion system may be used on various surfaces of aircraft of different architectures and configurations, including, but not limited to, subsonic, supersonic, and hypersonic, including airfoils, engines, exhaust nozzles for supersonic engines, and the like, as well as on their structures.

FIG. 15 shows a control system 400, which control system 400 may be used to control a hypersonic aerial vehicle, such as the second hypersonic aerial vehicle 300 in FIGS. 11-13, and other supersonic and hypersonic aerial vehicles, such as those in FIGS. 7-10. 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), the control processing unit 34 receiving at least airspeed indicators 106, which may be from ultrasonic sensors 104 (not shown). the control unit 34 also receives inputs from at least flight commands 108, which flight commands 108 may include commands such as various aircraft maneuvers or commands to stabilize flight due to turbulence or changing environmental and/or operating conditions.

Still referring to FIG. 15, 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 control processing unit 34, including, but not limited to, turbulence sensor 124, vibration sensor 114, static pressure sensor 103, differential pressure sensor and/or indicator 105, strain gauge 101 and microphone 107, and other sensors and parameters.

Still referring to FIG. 15, control system 400 may also include a plurality of aircraft control parameters 420, including but not limited to: ambient humidity 112, flight segment indicator 118, ambient temperature 119 (and/or free air temperature), aircraft altitude 122, and aircraft control mode 128, among other control parameters. Each of 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 other purposes in addition to those shown in FIG. 15. For example, 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 height 122 may also be used to correct or adjust other parameters as needed.

Still referring to FIG. 15, the control processing unit 34 may use an airspeed indicator 106 (which may include an indicated airspeed and/or a corrected or true airspeed) as an indication of the presence of a shockwave. For example, when airspeed indicator 106 signals that the aircraft is traveling at supersonic speeds, it may be possible to infer that a shockwave is present even without direct shockwave measurements or indications from, for example, a plurality of flight stability sensors and parameters 430. The control processing unit 34 may or may not have input from flight commands 108. For example, where the desired heading and/or control mode includes maintaining a current heading, there may be no input from flight command 108, but control processing unit 34 will continue to actively control the aircraft, e.g., maintain flight stability and aircraft orientation.

Still referring to fig. 15, control processing unit 34 determines a plasma actuator target for each plasma positions of the th plasma position 15A, the second plasma position 15B, the third plasma position 15C, and any other plasma positions on the aircraft based on the several inputs of fig. 15 and possibly other inputs, for each plasma position of the plurality of plasma positions 15A-15C, control processing unit 34 determines a target duration 140, a target plasma frequency, pulse rate, and/or waveform 142, a target plasma delay 144 (and/or sequence timing, e.g., when a pattern or sequence is desired to activate the plurality of plasma actuators 14), and a target plasma size 146 for each plasma position 15A determined plasma target value of th plasma position 15A is then communicated to th plasma actuator 14A, th plasma actuator 14A in turn performs the desired target plasma actuation and/or routine, in fig. 15, a target plasma stability value is shown only for the plasma position 15A, however the third plasma position 15B, a plasma orientation parameter may be communicated to a flight control system that evaluates the target plasma stability of the aircraft surface stability based on at least the third plasma position 3512, the flight parameter of the third plasma position 15B, the aircraft, the flight parameter of the aircraft, and/or the flight parameter of the aircraft, the flight parameter of the aircraft, the flight parameter of the aircraft.

In operation, plasma-assisted control system 400 of FIG. 15 may operate at a frequency of about 500Hz to about 50kHz based on input from sensors that may operate at frequencies of tens of Hz to tens of megahertz, for example, control system 400 may operate at about 5kHz to about 15kHz, based on input from sensors having varying operating frequencies, the entire control scheme or portions thereof and/or the module may be executed about 5,000 times to about 15,000 times per second, in other embodiments, control system 400 may operate at about 500Hz to about 50kHz, sensors may have time lags due to thermal lags associated with the time taken for the temperature sensor to heat up or cool down, for example, other sensors, such as electronic GPS or laser radar sensors, etc., may transmit and receive millions of signals per second, portions or modules of control system 400 may operate at different frequencies from other portions, for example, due to the continuously varying aerodynamic perturbations experienced under supersonic and high supersonic and hypersonic conditions, multiple plasma actuators 14 may operate at higher frequencies to maintain a constant operating frequency with a stable plasma processing signal, such as a stable plasma processing signal, and/or a stable flight control system may be adapted to at least provide a stable plasma processing response to a stable plasma processing characteristic, such as a stable plasma at least a stable shock signal, such as a stable flight response, such as a stable plasma processing system 14, a stable flight control system may be suitable for example, and/or stable flight control system, such as a stable flight control system may be adapted to provide a stable high frequency, such as a stable flight control signal suitable for a stable flight control system at least suitable for a stable flight response to a stable flight control system at least stable flight response to a stable shock signal suitable for a stable flight parameter of a stable shock signal of a stable flight parameter of a stable flight.

The control systems of fig. 14 and 15 may be used on subsonic, transonic, supersonic and hypersonic aircraft, such as those shown in fig. 6-13. Additionally, the plasma ignition combustion system 10 and the plasma-assisted control system 400 may be combined into a single system. For example, under supersonic flight conditions, the plasma may be activated alone without fuel injection when flight stability regulation and/or high frequency aircraft control regulation is desired or required. In other embodiments, the plasma may be used to ignite the fuel when a higher magnitude of control adjustment is required and/or when various aircraft maneuvers are requested from flight commands 108. Activating the plasma actuator 14 alone without fuel injection may be performed at a higher frequency than plasma-ignited combustion. The fuel delivery system that delivers fuel to, for example, an aircraft engine or an aircraft that is a coolant for the control system, may be combined as much as possible with the systems and components of the plasma ignition combustion system 10 (fuel supply 22, fuel control valve 20, fuel injectors 18, etc.).

Conventional aircraft may have movable surfaces for thrust vectors in the exhaust nozzle, and/or as control surfaces. However, these mechanical systems are heavy and relatively slow to respond (about 25Hz for conventional hydraulic actuators). In contrast, the plasma-ignited combustion system and the plasma-assisted control system of the present embodiment may alternatively be used on the exterior surface of an aircraft, for example on the wing and tail, to provide control forces without the need for movable surfaces and associated systems. The plasma-ignition combustion system and the plasma-assisted control system of the present embodiment can also be operated at higher frequencies in the range of about 500Hz to 15kHz, thereby achieving stable hypersonic flight.

The advantages of the present embodiments are that they enable aircraft control at higher speeds (100s Hz, rather than-10 Hz), which may be necessary at hypersonic conditions, and the present embodiments may be lighter in weight than conventional control surfaces, which will improve aircraft efficiency the fuel injectors 18 and plasma actuators 14 are synchronized such that each pulse and/or dispersion of fuel from the fuel injector 18 travels downstream to the plasma location 26 as plasma is formed, igniting the fuel.

Embodiments herein may improve combustion stability and enable plasma stabilized combustion systems to be used to control aircraft with little or no moving parts (see U.S. application 15/979,217, assigned to general electric company, starkecadi, n.y.). embodiments herein may also be used on the leading edges, trailing edges, and/or other surfaces of at least vanes of supersonic and/or hypersonic projectiles.

Exemplary embodiments of plasma-ignited combustion systems, plasma-assisted control systems, and related components are described above in detail. The system is not limited to the specific embodiments described herein, but rather, components of systems and/or steps of the methods may be utilized independently and separately from other components and/or steps described herein. For example, the configurations of components described herein may also be used in combination with other processes and are not limited to practice with the systems and related methods as described herein. Rather, the exemplary embodiment can be implemented and utilized in connection with many applications where supersonic combustion and/or supersonic aircraft control is desired.

Although specific features of various embodiments of the disclosure may be shown in drawings and not in others, this is for convenience only as any feature of any drawing may be referenced and/or claimed in combination with any feature of any other drawing, in accordance with the principles of the disclosure.

This written description uses examples to disclose embodiments of the disclosure, including the best mode, and also to enable any person skilled in the art to practice the disclosure, including making and using any devices or systems and performing any incorporated methods. 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. Such other examples are intended to be within the scope of the claims if they have structural elements that do not differ from the literal language of the claims, or if they include equivalent structural elements with insubstantial differences from the literal languages of the claims.

A further aspect of the invention is provided by the subject matter of the following clauses:

the aircraft includes a 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 leading edge and the second leading edge, a control processing unit communicatively coupled to each of the plurality of plasma actuators, and at least flight stability sensors communicatively coupled to the control processing unit, wherein the control processing unit commands at least of the plurality of plasma actuators to generate plasma in response to signals from the at least flight stability sensors.

2. The aircraft according to any preceding item, wherein the at least flight stability sensors detect at least aerodynamic properties during at least of supersonic flight and hypersonic flight.

3. The aircraft according to any preceding item, wherein the at least aerodynamic properties include at least of shock wave size and shock wave frequency.

4. The aircraft of any preceding item, further steps include at least aircraft orientation sensors, the at least aircraft orientation sensors including at least of GPS sensors, lidar and gyroscopes.

5. The aircraft of any preceding item, wherein the at least flight stability sensors include at least of turbulence sensors, strain gauges, and microphones.

6. The aircraft of any preceding item, wherein the at least flight stability sensors include at least of a vibration sensor, a static pressure sensor, and a differential pressure sensor.

7. The aircraft of any preceding claim, wherein the at least flight stability sensors are disposed on an underside of the aircraft.

8. The aircraft of any preceding claim, further comprising at least two flight stability sensors, wherein at least flight stability sensors are disposed on the underside of the aircraft, near an aircraft nose.

9. The aircraft of any preceding item, wherein the control processing unit adjusts at least of plasma frequency and plasma size based on at least of the shock wave magnitude and the shock wave frequency.

10. The aircraft of any preceding claim, wherein the at least aircraft orientation sensors are disposed on at least of the aircraft left wing and the aircraft right wing.

11. The aircraft of any preceding claim, wherein the at least aircraft orientation sensors are disposed near at least of an aircraft nose and an aircraft aft end.

12. The aircraft of any preceding item, further comprising at least aircraft orientation sensors, the at least aircraft orientation sensors comprising at least of a GPS sensor, a lidar and a gyroscope, wherein the at least flight stability sensors comprise at least of a turbulence sensor, a strain gauge and a microphone, wherein at least flight stability sensors are disposed on the underside of the aircraft, near an aircraft nose, and wherein the aircraft is capable of supersonic or hypersonic flight.

An aircraft control system comprising a control processing unit, at least sensors communicatively coupled to the control processing unit, and at least plasma actuators, the at least plasma actuators disposed proximate to a leading edge of an aircraft wing, the at least plasma actuators communicatively coupled to the control processing unit, wherein the control processing unit commands the at least plasma actuators to generate plasma in response to at least signals from the at least sensors.

14. The control system of any preceding item, wherein the at least signals represent at least aerodynamic properties during supersonic or hypersonic flight.

15. The control system of any preceding item, wherein the at least sensors include at least of a turbulence sensor, a strain gauge, a microphone, a vibration sensor, a static pressure sensor, and a differential pressure sensor.

16. The control system of any of the preceding items, further steps including at least aircraft orientation sensors, wherein the at least orientation sensors include at least of GPS, lidar and gyroscopes.

17. The control system of any of the preceding items, further comprising an airspeed indicator and at least of a temperature sensor, a humidity sensor, and an altimeter.

18. The control system of any preceding item, wherein the at least signals are representative of at least of shock wave magnitude and shock wave frequency.

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

20. The control system of any preceding item, further comprising an airspeed indicator and at least aircraft orientation sensors, wherein the at least orientation sensors comprise at least of GPS, lidar and gyroscopes, wherein the control system is operable in a range of about 500Hz to about 50kHz, wherein the at least sensors comprise at least of strain gauges, microphones and vibration sensors, and wherein the at least signals represent shock wave magnitude and frequency.