SUPERSONIC AIRCRAFT WITH AERODYNAMIC TAIL STRUCTURE

BACKGROUND OF THE INVENTION The global economy makes long range business travel more essential than ever. However, other than Concorde, with presence declining as transatlantic flights have discontinued, the pace of business travel remains at 1960's-era speeds. Technology advances have produced longer range, safer, and more comfortable aircraft-but not faster flights.

Supersonic overland capability and range are drivers of market potential for aircraft in the commercial and business sector. Buyers of supersonic commercial aircraft are expected to be from entities such as corporations, governments and government agencies, and high net-worth individuals. Most operators are expected to be large organizations, for example corporations and governments, with sophisticated flight departments that can manage multiple aircraft types.

Flights are expected to depart and arrive in a wide range of environments, from large international and national airports to small local airfields or suburban airports, with or without substantial service capabilities.

Although a supersonic aircraft for usage in commercial and business environments is to have many characteristics of a high-performance military aircraft, flight characteristics, operations, maintenance, and cost should be compatible to a business or commercial realm. The aircraft should be compatible with the infrastructure, servicing and operations experience base, and air traffic control system of the extant civil business jet.

The user community expects the aircraft to be usable not only in large, urban international hubs but also in suburban airports so that compatibility with shorter runway lengths, narrower taxiways, and lower maximum gross weight surfaces is desirable. Servicing and maintenance compatibility with personnel, equipment, and capabilities found at well-equipped fixed based operators (FBOs) and maintenance facilities is highly useful.

Many of the desirable features of supersonic civilian aircraft, particularly low-boom performance and long range, are very difficult to attain. Bill Sweetman in"Flights of fancy take shape-from Jane's (www janes. com) ", 21 July 2000, discusses the United States Defense

Advanced Research Projects Agency (DARPA) Quiet Supersonic Platform (QSP) program that is aimed at developing an efficient supersonic-cruise aircraft that does not produce a sonic boom.

The difficulty of such a result is indicated by the agency's admission that only a revolutionary design will meet the goal, and that incremental application of new technologies, or integration of existing technologies, is expected to be insufficient to attain the reduced boom goal.

Extension of aircraft range involves balancing of fuel capacity, payload volume, fuel consumption at desired speeds, aerodynamic, and other factors. Reduction of aerodynamic drag can assist in extending range, reducing sonic boom, and improving aircraft performance.

SUMMARY OF THE INVENTION In accordance with an embodiment of the illustrative system, a supersonic aircraft comprises a wing and at least two engine nacelles coupled to the lower surface of the wing on the trailing edge. The supersonic aircraft further comprises an inverted V-tail abutting to the upper side of the wing comprising a central vertical stabilizer, at least two inverted stabilizers coupled to sides of the central vertical stabilizer and coupled to the wing and supporting at least two engine nacelles, and at least two ruddervators respectively pivotally coupled to at least two inverted stabilizers. The supersonic aircraft also comprises a controller coupled to at least two ruddervators and capable of adjusting the aircraft longitudinal lift distribution throughout a flight envelope to maintain a reduced sonic boom and reduced drag trim condition.

BRIEF DESCRIPTION OF THE DRAWINGS Embodiments of the invention relating to both structure and method of operation, may best be understood by referring to the following description and accompanying drawings: FIGUREs 1A, 1B, and 1C are schematic pictorial diagrams respectively showing side, front, and top views of an embodiment of a supersonic aircraft capable of reducing drag through channel relief; FIGUREs 2A, 2B, and 2C are schematic pictorial diagrams respectively showing side, front, and top views of another embodiment of a supersonic aircraft that has a T-tail geometry; FIGURES 3A and 3B are schematic pictorial diagrams illustrating side and bottom perspective views of another embodiment of a supersonic aircraft in a configuration that enables channel relief; FIGURE 4 is a schematic pictorial diagram illustrating an embodiment of a channel control system that can be used in a supersonic aircraft;

FIGUREs 5A and 5B are two schematic pictorial diagrams illustrating the channel region within the wing and empennage in more detail; FIGUREs 6A and 6B are two pictorial diagrams depicting different views of a trailing edge flap that can be used in an embodiment of an aircraft capable of channel relief control; FIGUREs 7A, 7B, 7C, and 7D are four schematic perspective pictorial views showing detailed diagrams of portions of the tail structure; FIGURE 8 is a graph illustrating a drag rise plot indicative of pressure drag in a supersonic/transonic aircraft; FIGURE 9 is a schematic block diagram showing an example a flight control actuation architecture embodiment that can be used as the controller; FIGURE 10 is a schematic block diagram that depicts an embodiment of a suitable hydraulic power and distribution system architecture for supplying actuating power to the control effectors and systems; FIGURE 11 is a schematic pictorial structural diagram illustrating an example of a supersonic aircraft with an inverted V-tail structure and relatively large rudder in proportion to the tail; FIGUREs 12A and 12B are schematic pictorial views that illustrate an embodiment of a tail/nacelle integration; FIGUREs 13A, 13B, 13C, and 13D, are diagrams showing front, bottom, and side pictorial structural views of an example of a nacelle, wing, and tail configuration; FIGURE 14 is a pictorial diagram shows a frontal view of a wing and nacelle geometry in an illustrative low sonic boom aircraft; FIGUREs 15A, 15B, and 15C depict multiple schematic pictorial diagrams illustrating an embodiment of a supersonic aircraft; FIGURE 16 is a schematic pictorial diagram that illustrates an embodiment of a structural support member that attaches to a wing and to an inverted stabilizer; FIGUREs 17A and 17B are a pictorial diagram showing a side view, and a plurality of cross-sectional views, of an embodiment of a structural support member; and FIGUREs 18A and 18B are pictorial diagrams illustrating frontal and side views, respectively, of a structural support member.

DETAILED DESCRIPTION OF THE EMBODIMENTS For most supersonic and transonic aircraft, wave drag is a major component of total drag so that wave drag reduction is an objective of many design methodologies. Some aircraft have empennages that encompass control surfaces that are sufficiently close to form a channel, forming complex shock patterns which often manifest as extra drag and tend to appear in operation beyond a particular Mach number. The shock patterns can choke the channel at transonic conditions. In an aerodynamic analysis, a drag rise curve becomes more peaked, forming a transonic thrust pinch point, in the presence of channel choking. In an aerodynamic analysis, the presence of channel choking may cause the transonic drag rise curve to become more peaked, thus forming a transonic pinch point as depicted in FIGURE 8.

What is desired is a relief mechanism that can clear the transonic pinch point and reduce or eliminate channel choking.

In accordance with some embodiments, a supersonic aircraft comprises a fuselage extending forward and aft along a longitudinal axis, the fuselage having a lower surface and an upper surface, a highly swept low aspect ratio wing coupled to the fuselage and having a forward leading edge and an aft trailing edge, an effector flap coupled to the wing trailing edge, and a tail empennage. The tail empennage is coupled to the fuselage aft of the wing on the fuselage upper surface in a position high relative to the wing. The tail empennage forms a channel region subject to complex shock patterns for transonic flight conditions. The aircraft further comprises an effector coupled to the tail empennage and a controller coupled to the effector flaps and the effectors. The controller further comprises a control process that reduces drag through channel relief by deflecting both the effector flap down and the effector up. In a particular embodiment, the tail can be an inverted V-tail, and the controller deflects a trailing edge flap on the wing down while deflecting a ruddervator on an inverted stabilizer upward.

In accordance with other embodiments, a supersonic aircraft comprises an aircraft body extending forward and aft, a highly swept low aspect ratio wing coupled to the body and having a forward leading edge and an aft trailing edge, an effector flap coupled to the trailing edge of the wing, and an inverted V-tail. The inverted V-tail is coupled at the aft portion of the aircraft body and coupled to the wing in a braced wing configuration. The inverted V-tail forms a channel region that can generate complex shock patterns. The aircraft further comprises ruddervator control surfaces coupled to the inverted V-tail and a controller. The controller is coupled to the effector flap and the ruddervator control surfaces. The controller comprises a control process that reduces drag through channel relief by deflecting both the effector flap down and the ruddervator control surfaces up.

In further embodiments, a channel control system is used in a supersonic aircraft including a fuselage, wings, a tail empennage, and a plurality of control effectors coupled to the wings and the tail empennage. The empennage and wings form a channel region that can form complex shock patterns at transonic speeds. The channel control system comprises a plurality of actuators coupled to the control effectors. The effectors include a flap coupled to the wing and an effector coupled to the tail empennage. The channel control system further comprises at least one vehicle management computer coupled to the plurality of actuators. The vehicle management computers further comprise a process for managing the control effectors in a drag reduction mode through channel relief by deflecting both the flap downward and the tail empennage effector upward.

Referring to FIGUREs 1A, 1B, and 1C, schematic pictorial diagrams respectively showing side, front, and top views of an embodiment of a supersonic aircraft 100. The supersonic aircraft 100 comprises a fuselage 101 extending forward and aft along a longitudinal axis, a highly swept low aspect ratio wing 104 coupled to the fuselage 101 with an effector flap 106 coupled to the wing trailing edge, and a tail empennage 108. The tail empennage 108 is coupled to the fuselage 101 aft of the wing 104 on the fuselage upper surface in a high position relative to the wing 104. The tail empennage 108 and wing 104 form a channel region 102 that is subject to complex shock patterns. The aircraft further comprises an effector 110 coupled to the tail empennage 108 and a controller 111 communicatively coupled to the effector flap 106 and the effector 110. The controller 111 implements a control process that reduces drag through channel relief by deflecting both the effector flap 106 down and the effector 110 up.

In some configurations, the controller 111 may also deflect other effectors, such as ailerons, while deflecting the effector flap 106 and the effector 110 to counter any resulting pitching moment. In other embodiments, examples of effectors 110 may include leading edge or trailing edge surfaces of the wing 104. Similarly, examples of effector flaps 106 can include a leading edge device for the horizontal tail or the entire slat of the tail.

In the illustrative aircraft 100 the tail empennage 108 has an inverted V-tail geometry coupled to the wing 104 in a braced wing configuration. The V-tail 108 further connects to a vertical stabilizer 112, lateral inverted stabilizers 114, and inverted V-tail control surface ruddervators 116 capable of adjusting the aircraft longitudinal lift distribution throughout the flight envelope to maintain a low-boom, low-drag trim condition.

In some embodiments, the controller 111 implements a control process that increases drag by choking the channel, deflecting both the effector flap 106 up and the effector 110 down. In the deployment for increased drag, the control effectors can operate as a speedbrake, for example for usage in emergency descents.

The aircraft embodiment 100 further comprises engines 118 coupled to the aft portion of the wing lower surface with the inverted V-tail geometry empennage 108 coupled to the wing in the braced wing configuration carrying lift at the aft portion of the aircraft 100 on a high mounted tail 108. The length of the aircraft is effectively extended for shock waves below the aircraft 100, thereby further reducing sonic boom. The inverted V-tail 108 carries tail lift high to maintain a continuous lift distribution and structurally bracing the wing 104 and engines 118.

Referring to FIGURE 1C, control effectors are shown for the supersonic aircraft 100.

Two sets of surfaces are available for pitch control including canards 120 and ruddervators 116.

Roll control uses ailerons 128 and high speed spoilers 130. Yaw control is supplied by a rudder 140, ruddervators 116, and differential canard 120.

In combination with the canards 120, the supersonic aircraft 100 has multiple stability and control effectors. The canard 120 and symmetric defections of the ruddervators 116 control pitch power. A vertical rudder 140 controls yaw. Inboard, midboard and outboard ailerons 128, and the high speed roll spoilers 130 control roll. The segmented ailerons 128 provide both roll control power and automatic wing camber control to optimize lift and drag throughout the flight envelope. The roll spoilers 130 are configured to control roll at supersonic Mach numbers. High- speed spoilers 130 supplement aileron roll power at transonic and supersonic speeds where Mach number and aeroelastic effects reduce aileron effectiveness.

Some embodiments of an aircraft may include, in combination with the canards 120, a leading edge device to be used in conjunction with an outboard trailing edge flap deflection, which can be a more effective roll device from both drag and aeroelastic perspectives.

In an illustrative embodiment, trailing edge (TE) flaps 132 are effectors 106 than can be deployed 30° down to generate additional lift during landing. TE flap deployment reduces angle- of-attack specifications by approximately 2° during landing. During second-segment climb, the TE flaps 132 are extended 10° to improve the lift-to-drag ratio for better climb performance.

Leading edge (LE) Krueger flaps 134 are extended 130° for low speed operations including takeoff, approach and landing. The LE Krueger flaps 134 improve lift-to-drag ratio by 1.5, resulting in better climb performance that facilitates second-segment climb in case of engine malfunction.

In some embodiments, the aircraft 100 can be configured with a high lift system that includes simple inboard trailing edge flaps 132 and full-span leading edge Krueger flaps 134.

The multiple control surfaces of the supersonic aircraft 100, for example the ruddervators 116 inboard and outboard design, enable continued operation and landing following single actuator failure or a single control surface jamming. Differential canard deflection can generate a

yawing moment to counter a jammed rudder. Ailerons 128 and ruddervators 116 include multiple surfaces, increasing fault tolerant capability and supplying redundant elements for improved reliability.

Referring again to FIGURES 1A, 1B, and 1C, in the illustrative aircraft 100, shaping of the wing 104, body 101, empennage 108, and the integration of the propulsion system 116 are configured to produce a shaped sonic signature and control supersonic cruise drag. An inverted V-tail geometry 108 facilitates the overall low-boom design and supports nacelles 122 in an appropriate position relative to the wing 104, as well as enabling for trim to attain a low sonic- boom lift distribution. Inverted V-tail control surfaces, called ruddervators 116, adjust the aircraft longitudinal lift distribution throughout the flight envelope to maintain a low-boom, low-drag trim condition. The canard 120 supplies additional trim control and augments longitudinal control power.

In various embodiments, the illustrative aircraft 100 may include one or more of several advancements including addition of an all-flying canard 120, an optimized wing 104, incorporation of leading edge flaps 134 and spoilers 130, and a reconfigured body or fuselage 101. The canard 120 improves takeoff rotation and high-speed control. Wing planform and airfoil shapes are configured to assist high-speed performance, low-speed performance, low sonic boom, stability and control, and structural mass fraction characteristics. Sizes of the inverted V-tail 108 and fins can be configured to improve both structural and aerodynamic integration, benefiting both weight and drag characteristics. Flaps 134 improve takeoff performance. Spoilers 130 assist high-speed roll control.

The illustrative aircraft 100 has a twin-engine, slender-body configuration with a highly swept low aspect ratio wing 104, a configuration highly appropriate for low-boom performance based on volume-based sonic boom ruling. Equivalent area minimization can be used to reduce sonic boom signature. When equivalent area due to geometric area and lift sum to the minimized distribution, a minimized ground sonic boom occurs. The aft engine location beneath the wing 104, in combination with a highly integrated wing/inlet geometry, produces both low-boom compatibility and low inlet/nacelle installation drag. The inverted V-tail geometry 108 supplies both a low sonic-boom performance while generating longitudinal trim in cruise, and structural support for the engine/nacelle installation.

Some embodiments of the aircraft 100 implement one or more of several features including a multi-spar wing 104, a fuselage structure 101 with stringer-stiffened skins supported by frames, canards 120 that are integrated with the pressurized fuselage cabin structure, and aft- located engines 116 supported by a torque-box structure that extends aft of the wing 104 and is attached to the inverted V-tails 108.

Referring to FIGUREs 2A, 2B, and 2C, schematic pictorial diagrams respectively showing side, front, and top views of another embodiment of a supersonic aircraft 200 with a T- tail geometry 208. The aircraft 200 tail empennage has a supersonic T-tail geometry 208 that includes a vertical stabilizer 212, a lateral horizontal stabilizer 214, and a control surface elevator 216 capable of adjusting the aircraft longitudinal lift distribution throughout the flight envelope to maintain a low-boom and/or low-drag trim condition. The horizontal stabilizer 214 is the entire horizontal tail, an all-moving tail.

Referring to FIGURES 3A and 3B, schematic pictorial diagrams show side and bottom perspective views of an embodiment of a supersonic aircraft 300. The supersonic aircraft 300 comprises an aircraft body 302 extending forward and aft, a highly swept low aspect ratio wing 304 coupled to the body 302 and having a forward leading edge and an aft trailing edge. An effector flap 306 is coupled to the trailing edge of the wing 304. The aircraft has an inverted V- tail 308. The inverted V-tail 308 is coupled at the aft portion of the aircraft body 302 and coupled to the wing 304 in a braced wing configuration. The inverted V-tail 308 forms a channel region 310 that can generate complex shock patterns. The aircraft 300 further comprises a ruddervator control surfaces 312 coupled to the inverted V-tail 308 and a controller 314. The controller 314 is coupled to the effector flap 306 and the ruddervator control surfaces 312.

The controller 314 implements a control process that reduces drag through channel relief by deflecting both the effector flap 306 down and the ruddervator control surfaces up 312. In some embodiments, the controller 314 can also implement a control process that increases drag by choking the channel 310, deflecting both the effector flap 306 up and the effector 312 down. The control process that increases drag can be used, for example, for operation as a speedbrake.

The controller 314 implements a control process that adjusts aircraft longitudinal lift distribution throughout a flight envelope to maintain a low sonic boom, low drag-trim condition.

The illustrative aircraft 300 has two wing-mounted engines 316 positioned beneath the wing 304 at an aft location. The braced wing V-tail 308 supports the engines 316 and enables trim for a low sonic boom lift distribution. In some embodiments, the engines 316 have a highly integrated wing/inlet geometry that enables low-boom compatibility and low inlet/nacelle installation drag. In some embodiments, the body 302 has a slender body configuration that equalizes area forward and aft and facilitates sonic boom reduction. A canard 318 coupled to the body 302 supplies additional trim and augments longitudinal control power.

Other embodiments may utilize a different tail configuration, for example a T-tail or other forms. The illustrative inverted V-tail 308 has a central vertical stabilizer 320, inverted stabilizers 322 coupled to sides of the central vertical stabilizer 320 and also coupled to the fuselage 302.

The inverted stabilizers 322 and assist the fuselage 302 in supporting engine nacelles 324. The

inverted V-tail 308 also includes ruddervators 312 that are pivotally coupled to the inverted stabilizers 322 and can have operations managed by the controller 314. Generally, the controller 314 controls the ruddervators 312 to move up and down together for longitudinal control.

The ruddervators 312 can be configured with sufficient torsional stiffness to reduce or minimize flutter resulting from ruddervator rotation coupling with V-tail bending and torsion.

Ruddervators 312 have appropriate actuator stiffness and ruddervator torsional stiffness, along with a V-tail mass distribution controlled using ballast weight to manage ruddervator rotation coupling with V-tail bending and torsion. The ruddervators 312 can be symmetrically deflected in combination with the canards to supply pitch control power. The vertical rudder 326 supplies yaw control with roll control supplied by inboard, outboard, and midboard ailerons, and high speed roll spoilers.

The controller 314 also manages other control effectors in combination with the canards 318 and the ruddervators 312, including leading edge Krueger flaps 332, trailing edge flaps 306, ailerons 328, and spoilers 330.

Referring to FIGURE 4 in combination with FIGUREs 3A and 3B, a schematic pictorial diagram illustrates an embodiment of a channel control system 400 is used in the supersonic aircraft 300 that includes a fuselage 302, wings 304, a tail empennage 308, and a plurality of control effectors coupled to the wings 304 and the tail empennage 308. The empennage 308 and wings 304 form a channel region 310 that can form complex shock patterns at transonic speeds.

The channel control system 400 comprises a plurality of actuators 402 coupled to the control effectors. The effectors include a flap 404 coupled to the wing 304 and an effector 406 coupled to the tail empennage 308. The channel control system 400 further comprises at least one vehicle management computer 408 coupled to the plurality of actuators. The vehicle management computers implement a process for managing the control effectors in a drag reduction mode through channel relief by deflecting both the flap 404 downward and the tail empennage effector 406 upward. In some embodiments, the vehicle management computers also manage the control effectors in a speedbrake mode that increases drag by choking the channel and deflecting the flap 404 upward and the tail empennage effector 406 downward.

In the illustrative embodiment, the wing 304 is a highly swept low aspect ratio wing with the effector flap 404 mounted on the trailing edge of the wing. The tail empennage 308 is in a configuration of an inverted V-tail coupled at the aft portion of the aircraft body 302 and coupled to the wing 304 in a braced wing configuration. The tail empennage effector 406 is a ruddervator that is coupled to an inverted stabilizer in the inverted V-tail 308.

The illustrative aircraft 300 further includes a differential canard 318 that is mounted on the body 302. The canard 318 has an actuator 302 communicatively coupled to a vehicle

management computer that can deflect the canard 318 in combination with the control effectors to facilitate drag reduction and drag augmentation for usage as a speed brake.

Referring to FIGUREs 5A and 5B, two schematic pictorial diagrams illustrate the channel region 310 within the wing 304 and empennage 308 in more detail. The empennage 308 includes a tail structure section 502, a vertical stabilizer to inverted stabilizer joint section 504, and an inverted stabilizer to nacelle joint section 506. The tail structure section 502 includes the vertical stabilizer 320, and a pair of inverted stabilizers 322. Control effectors include the rudder 326 pivotally connected to the trailing edge of the vertical stabilizer 320 and ruddervators 312 pivotally connected to the trailing edge of the inverted stabilizers 322.

The vertical stabilizer 320 is attached to the top of the aircraft center body and aft section 508. The top of the vertical stabilizer 320 is attached to the tops of the left and right inverted stabilizers 322.

FIGURE 5B shows a view of a right nacelle structure 510 including the right engine nacelle 324, a right structural support member 512, such as a torque box or torsion box, and wing spars 514 within the right wing 516. The right nacelle structure 418 is attached to the right wing section 516 and the lower right inverted stabilizer 410.

The wing 516 includes multiple support spars or ribs 514 within an airfoil that supports the airfoil structural support members 512 on the right and left sides of the aircraft. The airfoil structural support members 512 have a configuration that reduces body freedom flutter by increasing chordwise wing bending by engine rib enhancement.

Referring to FIGUREs 6A and 6B, two pictorial diagrams illustrate different views of a trailing edge flap 600 that can be used in an embodiment of an aircraft capable of channel relief control. The trailing edge flap 600 is located between the engine nacelle 602 and the fuselage 604. The trailing edge flap surface can rotate in a downward direction in combination with an upward deflection of the ruddervator in a controlled angle to reduce drag. In an illustrative embodiment, the flap surface can rotate down to a maximum predetermined angle, for example 20°. The engine nacelle 602 has sufficient clearance for the flap 600 to deflect to a maximum deflection angle.

An actuator 606 drives deflection of the trailing edge flap 600. In an illustrative embodiment, the actuator 606 is an electro-mechanical rotary actuator with integral motors, brakes, and a speed sensor.

FIGURE 6B shows the trailing edge flap 600 attached at the inboard wing rear spar 608.

The trailing edge flap 600 is pivotally connected to the wing rear spar 608 via a flap hinge 610 and an actuator hinge 612.

Referring to FIGURES 7A, 7B, 7C, and 7D, four schematic perspective pictorial views show detailed diagrams of portions of the tail structure 700. FIGURE 7A depicts a view of the interface between a vertical stabilizer 702 and inverted stabilizers 704L, R. The top of the right 704R and left 704L inverted stabilizers are attached to the top of the vertical stabilizer 702. A rudder 706 is attached to the aft end 708 of the vertical stabilizer 702. The illustrative inverted stabilizers 704L, R couple to the vertical stabilizer 702 using left 710L and right 710R stabilizer upper lugs. Also shown in a left ruddervator surface 712L pivotally coupled to the left inverted stabilizer 704L.

The right inverted stabilizer 704R attaches to the wing adjacent to the right nacelle 714R.

FIGURE 7B shows a ruddervator section 716 including the left inverted stabilizer 704L coupling between the vertical stabilizer 702 and the left wing adjacent to the left nacelle 714L.

The illustrative configuration includes two ruddervators on each side, each of which is coupled to the inverted stabilizer. In the depicted view, a left outboard ruddervator 718LO and a left inboard ruddervator 718LI are shown coupled to the left inverted stabilizer 704L using ruddervator hinges 720 and actuator hinges 722 that control movement of the ruddervators.

FIGURE 7C illustrates baseline actuators 724 for the ruddervators 718. In the illustrative embodiment, the actuators 724 are electro-mechanical rotary-type actuators. When integrated into the inverted V-tail, the rotary actuators 724 do not protrude into the airstream and thereby avoid increases in aerodynamic drag. The illustrative baseline actuators 724 have integral motor drivers, brakes, and a speed sensor. Also in an illustrative embodiment, the ruddervator surface rotates approximately i35° about the hinge line. An aircraft can include multiple ruddervators 718 on each side, a redundancy that is useful to maintain aircraft control in a jammed-surface condition. The ruddervators 718 in FIGURE 7B include redundant inboard and outboard sections.

FIGURE 7D illustrates the rudder section 725 including the rudder 706 that is pivotally attached to the trailing edge 726 of the vertical stabilizer 702 by rudder hinges 728. The view also shows the right wing 730R and right nacelle 714R. In the illustrative embodiment, the aircraft fuselage 732 extends aft with the aft portion of the fuselage forming a fuselage tail cone 734. The rudder 706 is merged with the fuselage tail cone 734 so that the rudder and tail cone rotate pivotally with respect to the central vertical stabilizer 702 and the forward portion of the fuselage 732.

Referring to FIGURE 8, a graph illustrates a drag rise plot 800 indicative of pressure drag in a supersonic/transonic aircraft. The plot 800 shows the relationship of pressure drag to Mach number as the aircraft flies at different speeds. A baseline 802 shows drag rise performance for a supersonic aircraft embodiment without effector flap and effector deflection. Below

supersonic speeds, pressure drag gradually rises with Mach number. At transonic Mach levels, the baseline drag rise plot 802 peaks, rapidly rising then falling at supersonic speeds. At high Mach numbers, the baseline pressure drag levels. With channel relief 804, an effector flap is deflected down while the effector is deflected up, relieving the channel and reducing the pressure drag at the transonic levels. During speed brake application 806, the effector flap and effector are deflected in opposite directions in comparison to channel relief deployment, choking the channel and creating a larger pressure drag peak in the transonic region. In the subsonic and supersonic regions, the baseline 802, the channel relief line 804, and the speed brake line 806 are approximately the same.

Referring to FIGURE 9, a schematic block diagram shows an example a flight control actuation architecture embodiment 900 that can be used to control effectors in the empennage to operate in a drag reduction mode and, conversely, to increase drag in operation as a speedbrake..

In the illustrative example, primary flight control actuation uses"Fly-by-Wire"dual tandem linear hydraulics with triple electronic redundancy. Dual tandem actuation 902 is powered by two independent hydraulic systems 904 and 906 and sized for full rated performance based on a single system operation. The flight control system is closed-loop and commanded by the Vehicle Management Computers 908. The flight control system 900 performs control law implementations to produce aircraft handling qualities throughout flight. The system 900 can implement outer loop control modes such as Autopilot, Autolanding, and Auto collision avoidance control. The flight control actuation system 900 can also execute system integrity and health management functions. Various types of actuators can be implemented including, for example, Dual Tandem hydraulic actuators, Simplex hydraulic actuators, Rotary vane hydraulic actuators, multiple cylinders hydraulic actuators, integrated rotary electromechanical actuators (IREMA), and the like.

The control effector configuration, controlled by the Vehicle Management Computers 908, uses redundant control surfaces, enabling continued safe flight and landing in event of a single actuator failure or mechanically-jammed control surface. Redundancy is extended to the ailerons and ruddervators, which are also designed into multiple surfaces for increased fault tolerance and improved overall safety.

The Vehicle Management Computers 908 implement processes for controlling the effectors, including the canards to distribute lift to reduce or minimize sonic signature and to drive the aircraft to relaxed stability. In an illustrative embodiment, two electronic flight control systems are used to give superior handling qualities and optimal performance throughout the flight envelope. The first system is a full-authority Fly-By-Wire system designed for stability and handling qualities and determining the basic dynamic response of the aircraft.

The second flight control system is an active center-of-gravity (CG) management system.

As fuel is burned throughout the mission, the CG management system redistributes the remaining fuel to maximize range and reduce or minimize sonic boom signature. The CG management system also enables the canard, wing and inverted V-tail to interact in harmony to lift the vehicle efficiently for maximum range while producing a low sonic boom signature.

Referring to FIGURE 10, a schematic block diagram shows an embodiment of a suitable hydraulic power and distribution system architecture 1000 for supplying actuating power to the effectors and systems. For high reliability, the system 1000 is highly redundant with a hydraulic system supplying three independent sources 1002,1004, 1006 of hydraulic power to operate primary flight controls, landing gear 1014, nose wheel steering 1016, wheel brakes 1018, and thrust reversers 1020. The three independent systems 1002,1004, and 1006 give triple redundancy for continued safe flight and landing.

Hydraulic power for the systems is supplied by two engine driven pumps 1022 and an AC motor pump 1024 on system 1 1002 and system 2 1004. The engine driven pumps 1022 can operate continuously while the AC motor pumps 1024 operate on demand basis. Additionally, the AC motor pumps 1024 are an extra source of hydraulic power that gives redundancy within each system. The AC motor pumps 1024 can be operated on the ground for system checkout without running the engines or using a hydraulic ground carts.

System 3 1006 has two air driven pumps 1026 and an AC motor pump 1024. One air driven pump 1026 operates continuously while the other air driven pump 1026 and the AC motor pump 1024 operate on a demand basis. The AC motor pump 1024 in system 3 1006 can also be operated on the ground for system checkout without running the engines or using a hydraulic ground cart. System 3 1006 also includes a ram air turbine 1028 for emergency hydraulic and electrical power in the event of dual engine flameout. The ram air turbine 1028 is sized to supply hydraulic and electrical power to essential equipment from the certified altitude to safe landing for level 3 handling quality.

Supersonic flight over the United States and other countries is a challenging environmental issue for a viable supersonic commercial aircraft. Current FAA regulations prohibit civil flights at Mach numbers greater than one without case-by-case exceptions approved by the Administrator. Many other countries have similar restrictions.

Previous research has shown that the highly impulsive nature of the"N-wave"sonic- boom signatures of all existing supersonic aircraft is the primary cause of negative response and regulatory limitations on supersonic travel. Conclusions of NASA research further indicate the exceptional difficultly of designing an aircraft with an"N-wave"signature of sufficiently low amplitude for general public acceptance. However, the research also found that a"shaped"

signature was less objectionable and that a reasonably achievable amplitude wave could meet Committee on Hearing and Bioacoustics of the National Research Council (CHABA) guideline for acceptable noise impact to the general public, depending on frequency of exposure.

A sonic boom occurs due to pressure waves that occur when an aircraft moves at supersonic speeds. During subsonic flight, air displaced by a passing plane flows around the plane in the manner water flows around an object in a stream. However, for a plane flying at supersonic speeds, the air cannot easily flow around the plane and is instead compressed, generating a pressure pulse through the atmosphere. The pressure pulse intensity decreases as a consequence of movement from the airplane, and changes shape into an N-shaped wave within which pressure raises sharply, gradually declines, then rapidly returns to ambient atmospheric pressure. A wall of compressed air that moves at airplane speed spreads from the wave and, in passing over ground, is heard and felt as a sonic boom. The rapid changes in pressure at the beginning and end of the N-wave produce the signature double bang of the sonic boom.

Research has recently shown that boom intensity can be reduced by altering aircraft shape, size, and weight. For example, small airplanes create a smaller amplitude boom due to a lower amount of air displacement. Similarly, a lighter aircraft produces a smaller boom since an airplane rests on a column of compressed air and a lighter plane generates a lower pressure column. An aircraft that is long in proportion to weight spreads the N-wave across a greater distance, resulting in a lower peak pressure. Furthermore, wings that are spread along the body and not concentrated in the center as in a conventional aircraft produces a pressure pulse that is similarly spread, resulting in a smaller sonic boom.

Shaping of a sonic boom refers to a technique of altering source pressure disturbance such that a non-N-wave shape is imposed on the ground. Shaping sonic boom can reduce loudness by 15-20 dB or higher with no added energy beyond that to sustain flight. Shaping to minimize loudness is based on insight regarding changes in aircraft pressure disturbances during propagation to the ground.

Shaped sonic booms are only achieved deliberately. No existing aircraft creates a shaped sonic boom that persists for more than a fraction of the distance to the ground while flying at an efficient cruise altitude since non-shaped pressure distributions quickly coalesce into the fundamental N-wave shape. The N-wave form generates the largest possible shock magnitude from a particular disturbance. The N-wave shape results because the front of a supersonic aircraft generates an increase in ambient pressure while the rear generates a decrease in pressure.

Variation in propagation speed stretches the disturbance during propagation to the ground.

Shaped boom techniques typically attempt to prevent coalescing of the pressure disturbance by adding a large compression at the aircraft nose and an expansion at the tail with pressure in

between constrained between the compression and expansion. The shaped boom stretches the ends of the signature faster than the in-between pressures, creating a non-N-wave sonic boom at the ground.

Boom reduction makes a supersonic aircraft less objectionable by minimizing the loudness of a sonic boom. Audible frequencies in a sonic boom occur in the rapid pressure changes, or shocks, at the beginning and end of the typical N-waveform. More quiet shocks have decreased pressure amplitudes and increased pressure change time durations.

Although sonic boom reduction is an important design criterion for a supersonic aircraft, other considerations always impact design decisions. For example, a useful aircraft will have an appropriate capacity for holding passengers and/or cargo and be a suitable configuration for safe operation. Some design aspects include integration of landing gear and airframe.

What is desired is a supersonic aircraft with tail and control structures that effectively control the aircraft in subsonic, transonic, and supersonic flight, and enable sonic boom reduction or minimization.

In accordance with some embodiments of the disclosed aeronautical system, a supersonic aircraft comprises a wing having upper and lower surfaces and extending from a leading edge to a trailing edge and at least two engine nacelles coupled to the lower surface of the wing on the trailing edge. The supersonic aircraft further comprises an inverted V-tail coupled to the wing comprising a central vertical stabilizer, at least two inverted stabilizers coupled to sides of the central vertical stabilizer and coupled to the wing and supporting at least two engine nacelles, and at least two ruddervators respectively pivotally coupled to at least two inverted stabilizers. The supersonic aircraft also comprises a controller coupled to at least two ruddervators and capable of adjusting the aircraft longitudinal lift distribution throughout a flight envelope to maintain a reduced sonic boom and reduced drag trim condition.

According to other embodiments, a supersonic aircraft comprises a wing having upper and lower surfaces and extending forward from a leading edge aft to a trailing edge, and an inverted V-tail coupled to the wing comprising a central vertical stabilizer with leading and trailing edges, and at least two inverted stabilizers coupled to sides of the central vertical stabilizer and coupled to the wing. The aircraft further comprises a rudder pivotally mounted on the trailing edge of the central vertical stabilizer. The rudder has a sufficient area and rudder control sizing to enable adequate yaw acceleration to achieve at least 8 degrees of yaw angle change within four seconds for decrab and a rudder actuator rate less than 60 degrees/second.

In accordance with other embodiments, a supersonic aircraft comprises a fuselage extending forward and aft about a longitudinal axis. The fuselage has upper and lower surfaces.

The lower surface has a general axial curvature about the longitudinal axis and a local aft

flattening. The aft flattening of the fuselage adds lateral stiffening to the aircraft structure The aircraft further comprises a wing coupled inboard to the fuselage and extending outboard, and having a forward leading edge to an aft trailing edge. The aircraft also comprises an inverted V- tail coupled to the wing and fuselage comprising a central vertical stabilizer, at least two inverted stabilizers coupled to sides of the central vertical stabilizer and to the wing outboard of the fuselage. Furthermore, the aircraft comprises a strake coupled to and extending from the central vertical stabilizer through the fuselage interior and coupling to the lower fuselage surface at the position of local aft flattening.

According to further additional embodiments, a supersonic aircraft comprises a wing having upper and lower surfaces and extending from a leading edge to a trailing edge, at least two engine nacelles coupled to the lower surface of the wing on the trailing edge, and an inverted V- tail coupled to the wing. The inverted V-tail comprises a central vertical stabilizer, at least two inverted stabilizers coupled to sides of the central vertical stabilizer and coupled to the wing and supporting at least two engine nacelles. The aircraft further comprises at least two wing structural support members coupled to the upper surface of the wing generally overlying at least two engine nacelles. The wing structural support members couple between the inverted stabilizers and the wing and extend from the wing trailing edge forward. The structural support members add support to assist carrying engine nacelles weight.

Referring again to FIGUREs 1A, 1B, and 1C, schematic pictorial diagrams respectively showing side, front, and top views of an embodiment of a supersonic aircraft 100 with an inverted V-tail configuration 108. The aircraft 100 comprises a airfoil 103 formed of a wing 104 and fuselage 101 and having upper and lower surfaces and extending from a leading edge 105 to a trailing edge 107. The aircraft 100 further comprises at least two engine nacelles 122 coupled to the lower surface of the airfoil 103 on the trailing edge 107, and an inverted V-tail 108 coupled to the airfoil 103. The inverted V-tail 108 comprises a central vertical stabilizer 112, inverted stabilizers 114 coupled to sides of the central vertical stabilizer 112 and coupled to the airfoil 103 and supporting at least two engine nacelles 122. The aircraft 100 further comprises at least two wing structural support members 115 coupled to the upper surface of the airfoil 103 generally overlying at least two engine nacelles 122. The wing structural support members 113 couple between the inverted stabilizers and the airfoil 103 and extend from the wing trailing edge 107 forward. The structural support members 113 add support to assist carrying weight of the engine nacelles 122.

The wing structural support member 115 can be configured as a spine or support attached to the top of the wing as a protrusion forward of the inverted-V tail 108. The wing structural support member 115 generally extends a portion of the distance to the leading edge. The wing

structural support member 115 may also be termed a"chunnel."The chunnel can be configured to reduce or minimize aerodynamic wave drag.

In the illustrative embodiment, the inverted V-tail 108 is integrated into the wing trailing edge 107. The wing 104 has a gull or dihedral 109 inboard of the engine nacelles 122. The dihedral 109 is configured in a manner sufficient to increase take-off roll at the fuselage tip 117 and to extend lifting length and reduce sonic boom effects.

The illustrative aircraft 100 has two main landing gear 119 coupled to a lower surface of the wing 104 respectively inboard of the engine nacelles 122. The main landing gear 119 retract into the wing 104 and fuselage 101. The wing inboard portion is configured to integrate with the nacelle 122 and forms the dihedral gull 109 that enhances low-sonic-boom signature by vertically staggering wing longitudinal lift distribution. The dihedral gull 109 is formed by twisting and cambering the wing 104 for low sonic boom and low induced drag while preserving a tailored local wing contour at a location of main landing gear retraction.

In some examples, the illustrative aircraft arrangement 100 has twin non-afterbuming turbofan engines 118 set below and behind the wing 104. The non-afterburning turbofan engines 118 operate behind simple fixed-geometry axisymmetric external compression inlets 121.

Considerations of community noise and takeoff, transonic, and cruise thrust specifications determine engine cycle selection and engine sizing. Nacelles 122 enclose the engines 118 and are coupled to lower surface the wings 104 at the wing trailing edge 107.

The shaping of the supersonic aircraft 100 including aspects of the wing 104, the tail assembly or empennage 123, and the integration of wing, nacelle, and landing gear are adapted according to sonic boom signature and supersonic cruise drag considerations. The empennage or tail system 123 includes stabilizers, elevators, and rudders in the inverted V-tail geometry 108.

The inverted V-tail geometry 108 supports nacelles 122 in highly suitable positions relative to the wing 104 to suppress boom, and trims the supersonic aircraft 100 in cruise to attain an improved low-boom lift distribution. Panels of the inverted V-tail 108 support the nacelles 122 and non- afterburning turbofan engines 118 in combination with support of the wing 104 to handle flutter.

Inverted V-tail control surfaces, termed ruddervators 116, adjust aircraft longitudinal lift distribution throughout the flight envelope to maintain a low boom, low drag trim condition.

In the illustrative embodiment, the aircraft wings 104, empennage, and propulsion system integration can be configured for reduced sonic boom signature and supersonic cruise drag. The aircraft 100 further includes an inverted V-tail geometry that reduces boom amplitude, supports engine nacelles 122 in appropriate positions relative to the wings 104, and facilitates aircraft trimming in cruise to attain an optimum low-boom lift distribution. Usage of the V-tail geometry to supplement the wings'support of the engine nacelles improves flutter performance.

Inverted V-tail control surfaces 116, termed"ruddervators,"adjust aircraft longitudinal lift distribution throughout the flight envelope to maintain a low boom, low drag trim condition.

The wings 104 have a substantial dihedral, or"gulling", incorporated into the wing inboard of the engines 118, a geometry that is most pronounced at the wing trailing edge. The gull 109 is produced by twisting and cambering the wing to produce low-boom and low induced drag while preserving a tailored local wing contour at the location of main landing gear retraction.

In some embodiments, the wing leading edge 105 has a substantially straight geometry to accommodate a simple hinge line 124 for a Krueger flap 134 that extends for the full length of the wings 104. The inboard wing integrates with the nacelle and diverter geometry, and follows the fuselage contour with a substantially normal intersection to reduce interference drag. An inboard wing flap hinge line is fully contained within the wing contour with upper and lower wing surfaces configured as planar as possible to facilitate low speed aerodynamic performance.

The wing gull 109 raises the engines 118 to increase available tip-back angle and reduce thrust-induced pitching moments. The wing gull 109 lowers the aircraft body to reduce the cabin door height above the ground and reduce entry stair length. The low fuselage 101 sets a low aircraft center of gravity, reducing tip-over angle and promoting ground stability. The gull 109 tends to"wrap"the wing around the nacelle 122, enhancing favorable interference between the inlets 121 and the wings 104, so that the resulting wing/body/nacelle geometry facilitates successful ditching and gear-up landings. In addition, the wing gull 109 enhances the aircraft low-boom signature by vertically staggering the longitudinal lift distribution of the wings 104.

Favorable interference may also be achieved by wave cancellation or induced drag reduction due to nacelle lift.

In some embodiments, the supersonic aircraft 100 can include a canard 120 that operates primarily as a longitudinal power control device, particularly effectively during takeoff and in high-speed flight. The canard 120 also functions to fine tune the aircraft longitudinal trim condition. The canard 120 augments rudder operation by supplying yaw control power when left and right canard surfaces are deflected differentially.

The supersonic aircraft 100 includes segmented ailerons 128 that supply roll control power and automatic wing camber control to improve lift and drag conditions through the flight envelope. High-speed spoilers 130 supplement aileron roll power at transonic and supersonic speeds where Mach and aeroelastic effects reduce aileron effectiveness. The supersonic aircraft 100 has a high lift system including an inboard trailing edge flap 132 and a full-wingspan leading edge Krueger flap 134.

The supersonic aircraft 100 has multiple stability and control effectors. The canard 120 and symmetric defections of the ruddervators 116 control pitch power. A vertical rudder 140

controls yaw. Inboard, midboard and outboard ailerons 128, and the high speed roll spoilers 130 control roll. The roll spoilers 130 are configured to control roll at supersonic Mach numbers. In an illustrative embodiment, trailing edge (TE) flaps 136 are deployed 30° down to generate additional lift during landing. TE flap deployment reduces angle-of-attack specifications by approximately 2° during landing. During second-segment climb, the TE flaps 136 are extended 10° to improve the lift-to-drag ratio for better climb performance. In addition, trailing edge flaps 132 can be used in conjunction with ailerons 128 for drag reduction at transonic conditions.

Leading edge (LE) Krueger flaps 134 are extended 130° for low speed operations including takeoff, approach and landing. The LE Krueger flaps 134 improve lift-to-drag ratio by 1.5, resulting in better climb performance that facilitates second-segment climb in case of engine malfunction.

The supersonic aircraft 100 includes multiple control surfaces, for example the ruddervators 116 inboard and outboard design, to enable continued operation and landing following single actuator failure or a single control surface jamming. Differential canard deflection can generate a yawing moment to counter a jammed rudder. Ailerons 128 and ruddervators 116 include multiple surfaces, increasing fault tolerant capability and supplying redundant elements for improved reliability.

The supersonic aircraft 100 has a fuselage 101 with a geometry configured to address multiple different objectives. The basic fuselage camber line and volume distribution address suitable crew vision concerns. The fuselage 101 is shaped to enable a level cabin floor with near- constant cabin height and a close-to-the-ground cabin door 125 with a relatively short entry stairway. The fuselage 101 has an internal volume suitable to contain multiple subsystems and a suitable amount of fuel in the body to attain an extended range. The fuselage shape integrates well with the wing 104 and empennage 123, with the entire aircraft configuration being conducive to attaining a low-boom signature and supersonic cruise drag levels. The supersonic aircraft 100 has a relatively short nose landing gear 126 and a main landing gear 119 that stow in a compact stowage volume. The wing 104 and/or fuselage 101 form a wing having aerodynamic characteristics appropriate for low-boom supersonic and transonic flight.

In some embodiments, the aircraft 100 can have a blunted nose 127 with or without a conical tip 129 and an inverted V-tail surface 108 that overlaps the wing 104, features that facilitate low-sonic-boom aircraft performance. The configuration suppresses features of a sonic boom pressure waveform that otherwise would make the boom sound level unacceptable. The supersonic aircraft 100 creates an N-shaped pressure wave caused by overpressure at the nose 127 and underpressure at the tail 108. Pressure rises rapidly at the nose 127, declines to an underpressure condition at the tail 108, and then returns to ambient pressure. Rapid pressure rises

at the front and rear of the pressure wave producing the characteristic double explosion of the sonic boom.

The conical tip 129 of the nose 127 can be configured to create a pressure spike ahead of the aircraft forward shock, raising local temperature and sound velocity, thereby extending the forward shock and slowing the pressure rise. The supersonic aircraft 100 has a sharply swept arrow wing configuration 104 that reduces peak overpressure in the wave by spreading wing lift along the aircraft length. The wing configuration 104 has reduced wing leading and trailing edge sweeps. The inverted V-tail 108 can generate additional lift near the tail to improve aerodynamics and reduce boom.

Referring again to FIGUREs 3A and 3B, perspective top and bottom pictorial views, respectively, of an embodiment of a supersonic aircraft 300 capable of adjusting the aircraft lift distribution to maintain reduced drag trim and reduced or minimized sonic boom. The supersonic aircraft 300 comprises a wing 304 that extends from a leading edge 301 to a trailing edge 302.

The illustrative aircraft 300 has two engine nacelles 324 attached to the lower surface 2332 of the wing 304 on the trailing edge 302. The aircraft also includes an inverted V-tail 308 attached to the wing 304. The inverted V-tail 308 has a central vertical stabilizer 320, inverted stabilizers 322 coupled to sides of the central vertical stabilizer 320 and also coupled to the wing 304. The inverted stabilizers 322 and assist the wing 304 in supporting the engine nacelles 324. The inverted V-tail 308 also includes ruddervators 312 that are pivotally coupled to the inverted stabilizers 322.

The aircraft 300 further comprises a controller 314 that is communicatively coupled to the ruddervators 312 and can adjust the aircraft longitudinal lift distribution throughout a flight envelope to maintain a reduced sonic boom and reduced drag trim condition. Generally, the controller 314 controls the ruddervators 312 to move up and down together for longitudinal control. The controller 314 can also control asymmetric deflection of the ruddervators for roll control in synchronization with the rudder 326 for directional control.

The ruddervators 312 can be configured with sufficient torsional stiffness to reduce or minimize flutter resulting from ruddervator rotation coupling with V-tail bending and torsion.

The inverted V-tail geometry is useful for overall low-sonic boom performance. The ruddervators 312, inverted V-tail control surfaces, adjust aircraft longitudinal lift distribution throughout the flight envelope to maintain a low boom, low drag trim condition.

Ruddervators 312 have appropriate actuator stiffness and ruddervator torsional stiffness, along with a V-tail mass distribution controlled using ballast weight to manage ruddervator rotation coupling with V-tail bending and torsion.

The ruddervators 312 can be symmetrically deflected in combination with the canards to supply pitch control power. The vertical rudder 326 supplies yaw control with roll control supplied by inboard, outboard, and midboard ailerons, and high speed roll spoilers.

In an illustrative embodiment, the ruddervators 312 effectively control pitch using maximum deflections of 30 degrees for low speeds, and il0 degrees for high speeds. With 10 degrees deflection, the ruddervator effectiveness reduces as Mach number is increased beyond 0.9. The ruddervator may be less effective in the higher Mach numbers. Loss of effectiveness as Mach number approaches supersonic speed is common for trailing edge control devices. Flexible effects due to structural bending also contributed to the loss of ruddervator control effectiveness at high dynamic pressure conditions. The ruddervator 312 is an effective pitch control device at the subsonic speeds, providing approximately the same pitch control capability as the canard.

Optimal trimming surfaces may be a combination of both the canard and the ruddervator based on the least impact to trim drag increment.

Referring to FIGURE 11, a schematic pictorial structural diagram illustrates an example of a supersonic aircraft 1100 with an inverted V-tail structure 1102 and relatively large rudder 1104 in proportion to the tail 1102. The aircraft 1100 comprises a wing 1106 having upper 1108 and lower 1110 surfaces and extending forward from a leading edge 1112 aft to a trailing edge 1114. The aircraft 1100 further comprises the inverted V-tail 1104 coupled to the wing 1106 that has a central vertical stabilizer 1116 with leading 1118 and trailing 1120 edges, and inverted stabilizers 1122L, R coupled to sides of the central vertical stabilizer 1116 and coupled to the wing 1106. The rudder 1104 is pivotally mounted on the trailing edge 1120 of the central vertical stabilizer 1116. The rudder 1104 has a relatively large size in proportion to the central vertical stabilizer 1116. Specifically, the rudder 1104 has a sufficient area and rudder control sizing to enable adequate yaw acceleration to achieve at least 8 degrees of yaw angle change within four seconds for decrab and a rudder actuator rate less than 60 degrees/second.

In some embodiments, the rudder 1104 has an area that is greater than 60% of the area of the central vertical stabilizer 1116, an appropriate rudder area and rudder control sizing to counteract asymmetric engine thrust in the event of a single engine failure.

In some embodiments, the rudder 1104 has a sufficient area and rudder control sizing to attain a minimum control speed in air (Vmca) of approximately 165 knots. Vmca is defined as the speed at which the rudder is adequate to counteract asymmetric engine thrust with a bank angle less than or equal to 5 degrees.

In some embodiments, the rudder 1104 is sufficiently large and the inverted V-tail 1102 is configured at a position sufficiently aft with respect to the aircraft 1100 to attain rudder yaw control for single engine failures.

The supersonic aircraft 1100 includes a left 1124L and right 1124R wing sections respectively attached to the left 1126L and right 1126R sides of a center body/inboard wing section 1128 and a lower part 1130 of the left 1122L and right 1122R inverted stabilizers. Left 1132L and right 1122R leading edge flaps and left 1134L and right 1134R ailerons are attached to the left 1124L and right 1124R wing section forward spars 1136. Wing skins 1138 have integral stiffeners 1140 machined in a panel 1142 that runs between the wing spars 1136L, R.

The inverted V-tail structure 1102 includes three sections, a tail structure section 1144, a vertical stabilizer to inverted stabilizer joint section 1146, and inverted stabilizer to nacelle joint section 1148. The vertical stabilizer 1116 is attached to the top of the center body and aft body section 1150. The top of the vertical stabilizer 1116 is attached to the top of left 1122L and right 1122R inverted stabilizers. The lower end of left inverted stabilizer 1122L is attached to the surface of a left wing or wing structural support member 1152L, which may otherwise be termed a torque box, torsion box, or similar terminology. Left 1154L and right 1154R ruddervators are respectively attached to the aft of the left 1122L and right 1122R inverted stabilizers. The rudder 1104 is pivotally attached to the end of the vertical stabilizer 1116.

The illustrative embodiment of the aircraft 1100 further comprises engine nacelles 1156 coupled to the lower surface of the wing 1106 on the wing trailing edge 1114. In some embodiments, the aircraft further includes a controller 1158 coupled to the ruddervators 1154L, R. The controller 1158 can control the ruddervators 1154L, R to adjust yaw axis using sideslip command control law, and roll axis using sideslip command control law. In the yaw axis, pilot pedal input is interpreted as sideslip angle demand and pilot roll stick input is interpreted as roll rate demand.

The aircraft 1100 also comprises a fuselage 1160 merged with the wing 1106 and extending forward and aft along a longitudinal axis 1162. The aft portion of the fuselage 1160 forms a fuselage tail cone 1164. In the illustrative embodiment, the rudder 1104 is merged with the fuselage tail cone 1164 so that the rudder and tail cone rotate pivotally with respect to the central vertical stabilizer 1116 and the fuselage 1160.

Single engine failure minimum control speeds are used for rudder control sizing. Rudder control sizing is designed based on single engine failure analysis so that yaw control is adequate to maintain control of directional flight path angle. Control sizing is sufficient to yaw the aircraft into the direction of the operative engine, and the direction of the inoperative engine. For example in a 25 knots crosswind landing, the rudder can be designed to supply adequate yaw acceleration to achieve 8.5 degrees of yaw angle change within 3 seconds for decrab.

In another example, the rudder can be sized based on analysis of minimum control speed on the ground (Vmcg) with one engine failure during takeoff. The rudder size and rudder control

can be sized to supply adequate control margin for single engine failure during takeoff on the ground according to Vmcg, minimum controlled ground speed. At Vmcg, lateral deviation from runway centering is constrained to less than 30 feet. Aerodynamic moments balance engine thrust with one engine out and creating windmilling drag, and the other engine at max thrust plus a thrust bump for a"hot"engine. Moment balance can be done about the aircraft center of gravity considering main gear reactions caused by rudder sideforce.

In other embodiments, moment is balanced about the main gear center, which lies in line with the gear and halfway equidistant between the gear. Engine thrust imbalance is controllable with full rudder deflection. Vmcg, primarily a balance of engine thrust imbalance with the rudder, is relatively independent of flap setting or aircraft weight.

The vertical rudder is sized to counteract the asymmetric engine thrust in the event of single-engine failure.. The resultant side force generated by deflecting the rudder causes the aircraft to deviate directionally from the intended course. Rudder side force is countered by the gravity force generated by banking the vehicle. A maximum bank angle limit of 5 degrees is imposed by FAA for certification demonstration. In the condition where 5 degrees bank is inadequate to counteract the aerosurface side force, steady sideslip can be used. The direction of steady sideslip, however, creates additional yawing moment in the direction of the operative engine that increases sizing requirements of the rudder. When 5 degrees bank angle generates more than enough counteractive side force, steady sideslip generates yawing moment in the direction that reduces the appropriate rudder deflection.

Minimum control speed in the air (Vmca) is determined with the aircraft at maximum sea level takeoff thrust, maximum takeoff gross weight, takeoff flap configuration, gears up, and with one engine failed. Vmca is the minimum airspeed at which the rudder is adequate to counteract the asymmetric engine thrust with bank angle less than or equals to 5 degrees.

Minimum control speed during approach and landing (Vmcl) is the minimum calibrated airspeed at which the vehicle is controllable with one engine failed and the operative engine is set at the go-around power. Vmcl must be less than or equal to the approach speed (Vapp) and is determined with the aircraft in the landing configuration. The aircraft is designed to have adequate rudder and aileron control power to handle single engine failure for go-around with the vehicle in the landing configuration.

Furthermore, the rudder is sized to maintain directional and lateral control of the aircraft following a single engine failure for the entire flight envelope. Adequate yaw and roll control power is made available throughout the operational flight envelope to maintain a given course, with the remaining engine at maximum continuous power and bank angle less than 5 deg.

The illustrative aircraft is designed with the empennage positioned sufficiently aft to increase the moment arm and the rudder area is configured to a relatively large size. In addition, some embodiments utilize a high speed roll spoiler to increase the roll controllability of the vehicle for supersonic conditions. The spoiler is scheduled with the ailerons on the opposite wing for best performance.

A leading edge flap, either a simple flap for Krueger flap, on the outboard wing may be used for structural torsion alleviation and thus maintain reasonable wing thicknesses and wave drag levels as a consequence.

Rudder and aileron control power and actuator rate are sized adequately to handle gust upset for a side discrete gust. For example, in an aircraft at maximum takeoff weight with takeoff speed of 200 knots, the aircraft can be designed to be neutrally stable in the yaw axis. The rudder control power is sized to provide augmentation for stabilization. Additionally, aileron deflections are sized to maintain wing level at a selected gust frequency, for example based on gust length of 700 feet and a severe gust intensity of 30 knots (50 fps) corresponding to 7 degrees of sideslip upset.

In a particular embodiment, rudder size and rudder yaw control effectiveness are determined by measurements of yawing moment coefficient ACn for various angles-of-attack and Mach numbers, including flexible effects due to structural bending. Maximum rudder deflections of 30 degrees are used for low speeds, and 10 degrees are used for high speeds. Rudder yaw effectiveness generally decreases at increasing Mach number due to the flexible effects as dynamic pressure increases at higher Mach numbers. A relatively large rudder area and relatively aftward positioning of the vertical tail improve rudder yaw control effectiveness.

Referring again to FIGURES 5A and 5B in combination with FIGURE 3, two schematic pictorial diagrams show an example of an embodiment of a tail structure 308 for usage with the described supersonic aircraft. The tail structure 308 includes a tail structure section 502, a vertical stabilizer to inverted stabilizer joint section 504, and an inverted stabilizer to nacelle joint section 506. The tail structure 308 includes a vertical stabilizer 320, and a pair of inverted stabilizers 322. Control structures include a rudder 326 pivotally connected to the trailing edge of the vertical stabilizer 320 and ruddervators 312 pivotally connected to the trailing edge of the inverted stabilizers 322.

In FIGURE 5B, the wing 516 includes multiple support spars or ribs 514 within a wing that support the wing structural support members 512 on the right and left sides of the aircraft.

The wing structural support members 512 have a configuration that reduces body freedom flutter by increasing chordwise wing bending by engine rib enhancement. The wing ribs 514 are capable of supporting the wing structural support members 512 and reducing and/or eliminating nacelle

structural torsion. The wing structural support members add volume that integrates with a lowest far-field wave drag penalty and blends, as a fillet, with the inverted V-tail 308.

FIGUREs 12A and 12B are schematic pictorial views that illustrate an embodiment of a tail/nacelle integration. Referring to FIGURE 12A, a schematic pictorial diagram depicts the integration of a left nacelle 1200, left wing 1202, and left inverted V-tail stabilizer 1204. The top of the right and left 1204 inverted stabilizers are attached to the top of the vertical stabilizer.

FIGURE 12B is a view aft at the nacelle integration to the wing 1202. The top of torque box is removed for clarity. An accessory access panel is shown on the bottom of nacelle 1206 and the nacelle skin is removed for clarity. The torque box 1208 includes a left inboard torque box channel 1210, left torque box support 1212, and left outboard torque box channel 1214. The torque box 1208 also includes ducts for carrying fluids for the aircraft environmental control system. A diverter 1216 is positioned between the torque box 1208 and the left outboard wing 1202.

Referring to FIGUREs 13A, 13B, 13C, and 13D, front, bottom, and side pictorial structural views show an example of a nacelle, wing, and tail configuration. A nacelle structure 1300 includes a right nacelle structure 1302 and left nacelle structure 1304. The right nacelle structure 1302 is attached to the right wing section 1306 and the lower right inverted stabilizer 1308. The left nacelle structure 1304 is attached to the left wing section 1310 and the lower left inverted stabilizer 1312. A left structural support member or torque box 1316 is attached to the top of the left wing surface 1310 and engine inboard 1318 and outboard 1320 diverters. The left engine outboard diverter 1320 attached to the lower surface of the left wing 1310 and the top of the engine nacelle 1304. The left engine outboard diverter 1320 attaches the frames of the left engine nacelle 1304.

FIGURE 13D shows the left nacelle, wing, and tail configuration for an embodiment with the fuselage 1322 merged with the wing 1304 and extends and aft to a fuselage tail cone 1324. The tail includes a vertical stabilizer 1326 and a rudder 1328 pivotally attached to the aft edge of the vertical stabilizer 1326. The rudder 1328 is merged with the fuselage tail cone 1324 so that the rudder and tail cone rotate pivotally with respect to the central vertical stabilizer 1326 and the fuselage 1322 Referring to FIGURE 14, a pictorial diagram shows a frontal view of a wing and nacelle geometry 1400 in an illustrative low sonic boom aircraft. The wing 1402 has a gull or dihedral 1404 inboard of the engine nacelles 1406. The wing inboard dihedral 1404 integrates with the nacelles 1406 and enhances low-sonic-boom signature by vertically staggering wing longitudinal lift distribution. The dihedral gull 1404 is formed by twisting and cambering the wing 1402 for low sonic boom and low induced drag.

In some embodiments the wing trailing edge can be integrated to optimally relieve the diverter channel 1408 so that the wing 1402 wraps around the nacelle 1406 so that the trailing edge is not constrained to be linear in the vertical direction. The trailing edge can be shaped to relieve the area that diverges due to the dihedral 1404.

Referring to FIGUREs 15A, 15B, and 15C, multiple schematic pictorial diagrams depict an embodiment of a supersonic aircraft 1500 comprising a fuselage 1502 extending forward and aft about a longitudinal axis 1504. The fuselage 1502 has upper surface 1506 and lower surface 1508. The lower surface 1508 has a general axial curvature about the longitudinal axis and a local aft flattening 1510. The flattened fuselage adds lateral stiffening to the aircraft structure. The aircraft 1500 further comprises a wing 1512 coupled inboard to the fuselage 1502 and extending outboard. The wing 1512 also extends from a forward leading edge 1514 to an aft trailing edge 1516. The aircraft 1500 has an inverted V-tail 1518 coupled to the wing 1512 and fuselage 1502 comprising a central vertical stabilizer 1520, at least two inverted stabilizers 1522 coupled to sides of the central vertical stabilizer 1520 and to the wing 1512 outboard of the fuselage 1502.

The aircraft 1500 also has a strake 1524 coupled to and extending from the central vertical stabilizer 1520 through the fuselage interior and coupling to the lower fuselage surface, for example a keel running through the center of the fuselage 1502, at the position of local aft flattening 1510.

The fuselage/fin keel structure that includes the fin strake 1524 improves aircraft fuselage stiffness. A reduced volume in the fuselage 1502 facilitates sonic boom reduction and control.

The added strength from the dorsal or strake 1524 enables a desired reduction in fuselage volume and compensates for any reduction in stiffness that results from the flattened fuselage 1510.

The inverted V-tail 1518 is integrated into the wing trailing edge 1516. The wing 1512 has a gull or dihedral inboard of the couplings of the inverted stabilizers 1522 to the wing 1512.

The dihedral increases take-off roll at the fuselage tip, extends lifting length, and reduce sonic boom effect.

The supersonic aircraft 1500 has engine nacelles 1526 coupled beneath the wing 1512 at the wing trailing edge 1516, and two main landing gear 1528 coupled to the wing lower surface, inboard of the engine nacelles 1526. The main landing gear 1528 retract into the wing 1512 and fuselage 1502. The wing 1512 inboard portion integrates with the nacelle 1526 and forms the dihedral gull that enhances low-sonic-boom signature by vertically staggering longitudinal lift distribution. The dihedral gull is formed by twisting and cambering the wing 1512 for low sonic boom and low induced drag while preserving a tailored local wing contour at a location of main landing gear retraction.

In some embodiments, the aircraft can have Krueger flaps mounted on the leading edge 1514 of the wing 1512. The wing leading edge 1514 is configured sufficiently straight to accommodate a simple hinge line for the Krueger flap. The inboard wing 1512 integrates with the engine nacelles 1526 and follows the low sonic boom fuselage 1502 contour with a sufficiently normal configuration to attain low interference drag. The wing 1512 has an inboard flap hinge line fully contained within the wing contour with the wing 1512 upper and lower surfaces having an essentially planar form.

FIGURE 15A is a perspective side view of the landing gear system 1530 with the nose landing gear 1532 and the dual main landing gear 1528 extended and fuselage landing gear doors 1534 closed. The landing gears 1532 and 1528 are retracted during cruise flight. The landing gear 1532 and 1528 are extended an in a locked position. The landing gear 1532 and 1528 supply sufficient clearance between the fuselage tail 1536 and the lower aft part of the engine nacelle 1526 during rotate takeoff and high angle flare landing operations.

The main landing gear 1528 is located forward and inboard of the engine nacelles 1526.

The described main landing gear 1528 and integration of the landing gear configuration within the aircraft structure facilitate ground stability. In the illustrative embodiment, the aircraft 1500 has a structure that increases stability by increasing tail scrape angle and widening the wheel base.

The tail scrape angle is increased by lowering the fuselage 1502 and raising the engines and nacelles 1526, for example by virtue of a gulling of the wing 1512 that relatively raises the engine.

The retracted main landing gear 1528 fits into the wing 1512 at an angle that matches the wing dihedral. Wing dihedral increases the aerodynamic stability of the aircraft 1500 and benefits engine/wing iteration to reduce drag. In some embodiments, the wing 1512 includes Krueger flaps and the leading edge 1514 of the wing 1512 extends in an essentially straight line to facilitate using a simple hinge line that accommodates the Krueger flaps. Some embodiments have a wing 1512 with reduced leading and trailing edge sweeps.

Wheels of a four-wheel truck 1538 are aligned fore and aft in the fuselage 1502 to reduce or minimize cross-sectional volume and compactly stored in a landing gear compartment or wheel well 1540. The main landing gear geometry integrates into the inboard wing dihedral of the aircraft 1500 and has a size that fits into a compact fuselage volume by virtue of the forward and inboard retraction. The main landing gear geometry also conforms to available load paths of the aircraft structure to react to landing gear loads.

The main landing gear 1528 have sufficient length to supply ground clearance between the engine nacelles 1526, aft-body of the fuselage 1502, wingtips, and engine nacelles with

respect to the runway or static ground line. To reduce or minimize the risk that the landing gear will fail to extend, the tall main landing gear 1528 omits shortening mechanisms including folding and hinge structures. The aircraft accommodates the tall main landing gear 1528 by retracting the gear forward and inboard into the main landing gear compartments 1540 that similarly angle in a forward and inboard direction.

In some embodiments, the inboard portion of the wing 1512 is configured to integrate with the nacelle and a diverter formed between the nacelle and the wing 1512 to follow the contour of a low-sonic-boom fuselage 1502 with as close to normal intersection as possible to attain low interference drag. In some embodiments, an inboard flap hinge line is fully contained within the wing contour with the wing 1512 upper and lower surfaces held as planar as possible to facilitate seal design.

With the resulting wing configuration, the wing gull raises the engines and nacelles 1526 to increase available tip back angle and reduce thrust-induced pitching moments. The gull enhances low-boom signature by vertically staggering the wing longitudinal lift distribution and lowers the aircraft body or fuselage 1502 to reduce the height of the cabin door above the ground, thereby reducing entry stair length. The low fuselage 1502 assists in maintaining a low aircraft center of gravity, reducing tip over angle and promoting ground stability. The wing gull forms a wrapping of the wing 1512 around the nacelle that enhances favorable interference between the inlets and the wing 1512, resulting in a wing/body/nacelle geometry conducive to successful ditching and gear-up landings.

FIGURE 15B shows a bottom view of the left main landing gear 1528 during retraction.

In the illustrative conditions, the left landing gear strut 1542 is partially retracted at about a 60° angle from the fully extended position. A high axial load is imposed on the main gear trunnion link 1544 due to non-normal angle between the main strut 1542 and the trunnion axis which is reacted into the wing structure 1512. The landing gear 1528 retracts to an angle that follows wing contours of an inboard dihedral. The landing gear 1528 fits within local wing 1512 contours and is stored in the landing gear compartment 1540 within the wings 1512 and fuselage 1502.

FIGURE 15C is a cross-sectional view of the fuselage 1502 at the position of the landing gear compartment 1528. The strake 1524 extends from the dorsal fin vertical stabilizer 1520 entirely through the fuselage 1502 from the upper surface 1506 to the lower surface 1508. The strake 1524 extends through the main landing gear compartment 1528 to the flattened portion 1510 of the fuselage 1502. The strake 1524 carries the inverted V-tail 1518 bending loads through the aft fuselage 1502, reducing the effects of body flutter and facilitating improvement of sonic boom reduction performance by enabling a smaller cross-section in the fuselage 1502.

Referring to FIGURE 16, a schematic pictorial diagram illustrates an embodiment of a structural support member 1600 that attaches to a wing 1602 and to an inverted stabilizer 1604.

The structural support member 1600 firmly attaches to structural ribs 1606 that supply structural support to the wing 1602 and connect the diverter to the wing 1602. Typically, the support ribs in a wing are distributed at regular intervals along the wing. In the illustrative embodiment, the outboard rib 1606 is moved inboard so that the structural support member 1600, in combination with the inverted stabilizer 1604 and the wing 1602, supply increased support to the engine, nacelle, and diverter. Accordingly two diverter ribs 1606 are closely spaced with a narrow structural support member 1600 extending and affixed to the ribs 1606. The structural support member 1600 adds volume to increase support to the main spar or rib 1606 in the wing 1602. The structural support member 1600 also wraps around the intersection of the inverted stabilizer 1604 and the wing 1602 to form a fairing that improves aerodynamics.

Referring to FIGUREs 17A and 17B, a pictorial diagram shows a side view, and a plurality of cross-sectional views, of an embodiment of a structural support member 1700. The views show an aircraft tail 1708 with the structural support member 1700 connected to a wing 1702 and inverted stabilizer 1704. The structural support member 1700 has an aerodynamic structure, shown in the forward 1712, medial 1714, and aft 1716 cross-sectional views. The aerodynamic form enables the structural support member 1700 to have additional volume that increases strength while reducing drag.

Referring to FIGUREs 18A and 18B, pictorial diagrams illustrate frontal and side views, respectively, of a structural support member 1800. The structural support member 1800 attaches to a wing 1802 and inverted stabilizer 1804 and couples to ribs or spars 1806 that support the wing 1802 and a diverter 1806 beneath the wing 1802. The diverter 1808 is attached between the wing 1802 and a nacelle 1810 that encases an engine 1802. The diverter 1808 improves aerodynamics of the connection between the nacelle 1810 and the wing 1802.

The illustrative structural support member 1800 adds volume to the connection of the nacelle 1810, the diverter 1808, and the wing 1802, supported by the inverted stabilizer 1804 while maintaining an aerodynamic form. The structural support member 1800 forms an aerodynamic fairing that wraps around the junction of the wing 1802 and the inverted stabilizer 1804. The structural support member 1800 is connected to the diverter ribs 1806 and, in the illustrative embodiment, the diverter 1808 has a swept leading edge that improves aerodynamics of the connection between the nacelle 1810 and the wing 1802. The diverter 1808 also has a relatively large depth, for example in the range of six to twelve inches from the wing 1802 to the nacelle 1810 to add stiffness to the structure. In one embodiment, the diverter 1808 has a depth of about eight inches. Support supplied by the structural support member 1800 enables the diverter

1808 to have an increased depth and to be moved forward relative to the wing 1802 and nacelle 1810 to improve the strength and aerodynamics of the wing-nacelle-inverted stabilizer structure.

In an illustrative embodiment, the diverter 1808 couples the engine nacelle 1810 to the wing 1802 with a pair of ribs 1806 extending through the wing 1802 and diverter 1808. The ribs 1806 support the engine nacelle 1810 and are closely spaced with a first rib approximately aligned with the center of the nacelle 1810 and a second rib inboard of the first rib approximately to the inboard edge of the engine nacelle 1810. The structural support member 1800 extends essentially between the first and second ribs 1806 and extends vertically upward to the inverted stabilizer 1804 to add volume for strength while wrapping about the connection of the wing 1802 and inverted stabilizer 1804 as an aerodynamic fairing.

While the present disclosure describes various embodiments, these embodiments are to be understood as illustrative and do not limit the claim scope. Many variations, modifications, additions and improvements of the described embodiments are possible. For example, those having ordinary skill in the art will readily implement the steps necessary to provide the structures and methods disclosed herein, and will understand that the process parameters, materials, and dimensions are given by way of example only. The parameters, materials, and dimensions can be varied to achieve the desired structure as well as modifications, which are within the scope of the claims. Variations and modifications of the embodiments disclosed herein may also be made while remaining within the scope of the following claims. For example, although a particular aircraft geometry and configuration is described, the channel configuration and techniques for controlling the channel can be utilized in aircraft with different geometries. In particular, although the described aircraft has an inverted V-tail configuration, other tail configurations such as T-tail configurations and others may be used. The described propulsion configuration includes two engines mounted at aft positions in a highly swept wing. Other suitable embodiments may have different engine configurations with fewer or more engines, with engines mounted on the fuselage or tail rather than on the wing, or mounted above rather than beneath the wing.