SUPERSONIC AIRCRAFT

BACKGROUND OF THE INVENTION The world of aviation has held out the promise of widespread supersonic commercial aviation for decades only to be denied by technological, economic, and political roadblocks.

Although beginning operation over a quarter century ago, the Concorde remains the only commercial aircraft that travels at supersonic speeds while fighting technological obsolescence.

Fuel consumption and maintenance requirements of the Concorde make it commercially difficult to operate in today's competitive environment. Possibly overshadowing all other technological and economic shortcomings is the Concorde's thunderous sonic boom that is capable of shattering windows in buildings under the flight path, a fault that has restricted the Concorde to routes over oceans.

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. ianes. com)", 21 July 2000, discusses the United States Defense Advanced Research Projects Agency (DARPA) Quiet Supersonic Platform (QSP) program that is intended to develop 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.

The sonic boom creates a major practical risk of commercial supersonic aviation. So long as commercial supersonic aircraft are prohibited from flying over populated land masses, commercialization of such aircraft is financially unsound. 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 rises sharply, gradually declines, then rapidly returns to ambient atmospheric pressure. A wall of compressed air that moves at the airplane speed spreads from the wave and, in passing over the 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.

SUMMARY OF THE INVENTION In an illustrative embodiment, an aircraft designed for transonic speed comprises an airfoil, a nacelle, an engine, and an inverted V-tail. The nacelle is mounted on the lower aerodynamic airfoil surface and behind the trailing edge of the airfoil. The engine is enclosed within and structurally supported by the nacelle. The inverted-V tail is coupled to the airfoil at a position on the upper aerodynamic surface directly across the airfoil surface so that the inverted-V tail structurally supports the nacelle and engine in combination with support from the airfoil.

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 that respectively illustrate side, front, and top views of an embodiment of a supersonic aircraft; FIGUREs 2A and 2B are respective perspective top and bottom views of the supersonic aircraft; FIGUREs 3A, 3B, 3C, and 3D are a group of schematic pictorial diagrams showing multiple aspects of an automated fuel transfer system; FIGURE 4 is a schematic block diagram that illustrates an embodiment of an aircraft lift device including control actuators that can be operated in combination with center of gravity management to control the aircraft; FIGUREs 5A and 5B are schematic pictorial diagrams showing an embodiment of a Krueger flap that can be controlled in combination with center of gravity management to control the aircraft; FIGUREs 6A, 6B, 6C, 6D, and 6E are multiple schematic pictorial diagrams illustrating various control effectors that can be controlled in combination with center of gravity management;

FIGUREs 7A and 7B are two schematic pictorial diagrams depicting views of canard control effectors that can operate in conjunction with center of gravity management; FIGURE 8 is a schematic block diagram showing an example a flight control actuation architecture embodiment that can be used as a controller to manage aerodynamic effectors and center of gravity management; FIGURE 9 is a schematic block diagram showing an embodiment of a suitable hydraulic power and distribution system architecture for supplying actuating power to effectors and the center of gravity management system; FIGURE 10 is a graph that shows a center-of-gravity envelope for an aircraft embodiment; FIGURE 11 is a graph showing an example of a control configuration that can be controlled by the Vehicle Management Computers in an embodiment of the supersonic aircraft to attain longitudinal stability and control during takeoff and landing; FIGURE 12 is a graph illustrating an example of a control configuration that can be controlled by the Vehicle Management Computers in an embodiment of the supersonic aircraft to attain longitudinal stability and control during supersonic cruise; FIGURE 13 is a schematic block diagram illustrating an embodiment of a fly-by-wire (FBW) flight control system that attains aircraft controllability via management of control effectors and center of gravity; FIGURE 14 is a schematic pictorial diagram illustrating an embodiment of an inlet for an aircraft engine including an axisymmetric translating cowl and a fixed geometry axisymmetric inlet spike; FIGURE 15 is a schematic pictorial and airflow diagram that illustrates airflow in an embodiment of an inlet; FIGUREs 16A and 16B are schematic pictorial diagrams that respectively show the engine inlet with the translating cowl retracted and extended;

FIGURE 17 is a schematic pictorial diagram that depicts aerodynamic characteristics of the external compression, fixed compression geometry inlet; FIGURE 18 is a schematic pictorial diagram showing a side view of a propulsion system installed under an aircraft wing; FIGURE 19 is a schematic pictorial diagram that shows an example of a propulsion system integrated into an aircraft; FIGUREs 20A and 20B are pictorial diagrams respectively showing frontal and side views of a wing/nacelle/inlet geometry; FIGUREs 21A, and 21B show alternative views of fixed compression geometry, translating-cowl, isentropic compression engine inlets; FIGURE 22 is a perspective pictorial diagram showing an example of an inlet mounted beneath a wing segment ; FIGURE 23A is a chart representing a typical Mach 1.8 mission output with an engine with fixed compression geometry external compression inlet; FIGURE 23B is a graph showing Mach number flight profiles for subsonic and supersonic performance of a selected engine and inlet configuration; FIGUREs 24A and 24B are respectively a pictorial frontal view and a three-dimensional view from the rear of the aircraft showing structural integration of a propulsion system into an aircraft; FIGURE 25 is a schematic pictorial structural diagram illustrating an example of a supersonic aircraft including the engine structural integration system; FIGURE 26 is a schematic pictorial diagram that depicts the integration of a left nacelle, left wing, and left inverted V-tail stabilizer; FIGUREs 27A, 27B, and 27C are front, bottom, and side pictorial structural views showing an example of a nacelle, wing, and tail configuration;

FIGUREs 28A and 28B are side and front schematic views respectively showing an embodiment of an engine attachment/nacelle concept; FIGUREs 29A, 29B, 29C, and 29D are multiple views showing nacelle integration into an aircraft; FIGUREs 30A and 30B are computational fluid dynamic (CFD) images respectively showing aerodynamic characteristics of two nacelle configurations; FIGURE 31 is a schematic block diagram showing an embodiment of an automatic thrust management system that can be used in an aircraft; FIGURE 32 is a three-dimensional graph showing a takeoff-climb profile definition used for acoustic noise determination; FIGURE 33 is a graph that illustrates a notational time history of engine thrust for multiple scenarios that may occur during the take-off climb profile; FIGURE 34 is a schematic block diagram illustrating an"aircraft-centric"automatic takeoff thrust management system that implements a programmed lapse rate; FIGURE 35 is a schematic block diagram illustrating an"engine-centric"automatic takeoff thrust management system that implements a programmed lapse rate; and FIGURE 36 is a schematic block diagram showing an example of a thrust command logic structure.

DETAILED DESCRIPTION OF THE EMBODIMENTS In some embodiments, a supersonic aircraft comprises a wing, a fuselage, a plurality of fuel tanks contained within the wing and/or fuselage, and a fuel transfer system communicatively coupled to the plurality of fuel tanks and capable of transferring fuel among the plurality of fuel tanks. The aircraft further comprises at least one sensor capable of indicating a flight parameter and a controller. The controller is coupled to the one or more sensors and to the fuel transfer system. The controller can transfer fuel among the plurality of fuel tanks and adjust the aircraft center of gravity to reduce trim drag and increase aircraft range.

In other embodiments, an automated fuel transfer system can be used in a supersonic aircraft. The aircraft includes a fuselage and wing. The automated fuel transfer system comprises a plurality of fuel tanks distributed within the wing and/or the fuselage, a plurality of pumps coupled to the plurality of fuel tanks and capable of transferring fuel among the plurality of fuel tanks, at least one sensor capable of indicating a flight parameter, and a controller. The controller is coupled to the sensors and the plurality of pumps, and transfers fuel among the plurality of fuel tanks to modify the aircraft lift distribution, thereby reducing or minimizing aircraft sonic boom.

According to further embodiments, an aircraft control system may be used in a supersonic aircraft. The aircraft includes a fuselage and wing. The control system comprises a plurality of control effectors coupled to the wing, a plurality of fuel tanks distributed within the wings and/or the fuselage, a plurality of pumps coupled to the plurality of fuel tanks and capable of transferring fuel among the plurality of fuel tanks, and a plurality of actuators coupled to the control effectors.

The control system further comprises at least one sensor capable of indicating a flight parameter, and at least one vehicle management computer coupled to the at least one sensor, the plurality of pumps, and the plurality of actuators. The vehicle management computer can manage the control effectors and transfer fuel among the plurality of fuel tanks to adjust aircraft trim and center of gravity position to operate the aircraft in at least two flight modes. The variable flight modes have different trim drag and sonic boom performance.

An aircraft may use a set of sensors to indicate current and projected flight parameters of the aircraft during flight. The parameters are supplied to a computer or controller that executes an optimization algorithm to obtain an appropriate aircraft center of gravity to attain one or more objectives including: (1) reducing trim drag and thereby increase aircraft range, (2) changing the lift distribution to attenuate, reduce, minimize, or otherwise optimize aircraft sonic boom, (3) reducing the trim requirement to increase aircraft controllability, and (4) assisting in maintaining stability during flight. The optimal center of gravity for desired performance is attained by an automated system that actively transfers fuel among multiple aircraft internal fuel tanks.

Active center of gravity management is desirable, particularly in a supersonic aircraft, since the aircraft's aerodynamic center moves significantly from subsonic to supersonic flight regimes. Automated control of center of gravity management increases aircraft performance, stability, controllability, and reduces or minimizes sonic boom signature, while reducing workload and attention demands on the flight engineer or pilot. An automated system reduces demands on the pilot that may be overwhelming in some conditions if the system were manual.

The automated center of gravity management system improves flight performance and flight safety while reducing aircraft operating costs.

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 an aerodynamic body 102, a plurality of fuel tanks 106 contained within the aerodynamic body 102, and a fuel transfer system 108 communicatively coupled to the plurality of fuel tanks and capable of transferring fuel among the plurality of fuel tanks 106. In an illustrative embodiment, the aerodynamic body 102 includes a fuselage 101 and a wing 104, and the plurality of fuel tanks 106 can be distributed in suitable areas of the fuselage 101 and wing 104. The aircraft further comprises at least one sensor 110 that is capable of indicating a flight parameter, and a controller 112. The controller 112 is coupled to the sensors 110 and to the fuel transfer system 108. The controller 112 can transfer fuel among the plurality of fuel tanks 106 and adjust the aircraft center of gravity to reduce trim drag and increase aircraft range.

The various sensors 110 indicate various flight parameters, current and/or projected. The sensors measure and supply one or more items of information selected from control effector positions and settings, aircraft angle of attack (a), weight, dynamic pressure, and the like. Side stick and rudder pedal controls can be connected to sensors to indicate pilot and copilot control of flight control surfaces from the flight deck. Sensors can also indicate positioning of dual rudder pedals with brake control for control of the rudder. In an illustrative embodiment, the rudder pedals can be adjusted individually and, for example, can have separate sensors to indicate settings. One or more sensors can be connected to a pitch trim control switch located on the stick controller to supply information indicative of lateral roll and directional yaw trim.

Sensors can be implemented within a flap control system that is indicative of flap position. An asymmetric brake can be installed on the ends of the flap drive system with a sensing device used to sense a predetermine condition and, in response, actuate the brakes in the event of a torque shaft failure.

The aircraft 100 can also include a aircraft status sensor or set of sensors capable of detecting establishment of takeoff climb conditions, and engine failure detectors respectively coupled to the at least two engines and capable of detecting engine failure. For example, sensors 110 can be included for detecting various control parameters such as engine speed or Mach number, engine inlet temperature, engine revolutions per minute, engine inlet pressure, weight on wheels, and others.

Sensors 110 indicate status of the aircraft, for example whether the aircraft has established takeoff climb conditions. Status sensors 110 are used in the illustrative embodiment to automatically control takeoff thrust. The sensors 110 can also be used to detect status for

purposes of stability augmentation for pitch handling during flight, and for proper elevator adjustment.

In various embodiments, different sensors 110 and sensor combinations can be used. A highly useful status sensor is a weight-on-wheels detector. A typical weight-on-wheels detector uses one or more strain gages mounted on landing gear posts to determine whether a force is above or below a predetermined threshold force indicative of aircraft contact with the ground.

Other typical weight-on-wheels detectors use position sensors or switches to detect compression of the landing gear as it supports the aircraft weight.

Other embodiments may supplement or replace the weight-on-wheels detector with other sensors 110. A landing gear truck attitude sensor can detect truck angle relative to landing gear posts, monitoring for an upward swinging tilt of the gear indicative of takeoff or a downward shift of the gear truck at touchdown.

Some embodiments may use other or additional sensors for redundancy and self- checking. Multiple or dual sensors may improve reliability by avoiding false positive and false negative indications.

The set of sensors 110 indicate current and/or projected flight parameters of the aircraft 100 during operation. The controller 112 uses the parameters in a control process to obtain an appropriate center of gravity location for the aircraft to: (1) reduce trim drag and thereby increase aircraft range, (2) modify lift distribution to minimize or reduce sonic boom, (3) reduce the trim criteria to increase aircraft controllability, and (4) assist in maintaining stability during flight. The controller 112, upon determining the appropriate center of gravity location, actively transfers fuel among the aircraft's multiple fuel tanks 106 to automatically change the aircraft center of gravity to the determine location.

For example, the controller 112 can transfer fuel among the plurality of fuel tanks 106 to adjust the aircraft center of gravity to modify the aircraft lift distribution to reduce or minimize the aircraft sonic boom by pumping fuel to a maximum aft position. Moving the center of gravity aft trims the aircraft for a low sonic boom condition without significant trim drag penalty.

Accordingly, the controller 112 is capable of transferring fuel among the plurality of fuel tanks 106 to adjust the aircraft center of gravity and thus the aircraft longitudinal lift distribution throughout the flight envelope to maintain a low-boom, low-drag trim condition. Similarly, the controller 112 can transfer fuel among the plurality of fuel tanks 106 to adjust the aircraft center of gravity to a more aft location to reduce trim criteria to increase aircraft controllability.

In another example, the controller 112 can transfer fuel among the plurality of fuel tanks 106 to adjust the aircraft center of gravity laterally in the event of an engine failure that tends to pull the aircraft 100 to one side to help maintain aircraft stability and controllability during flight.

The controller 112 transfers fuel among the plurality of fuel tanks 106 to adjust the aircraft center of gravity in compliance with control laws to stabilize the aircraft 100 and provide satisfactory handling qualities to a pilot by evaluating closed-loop aircraft responses to atmospheric disturbance. The illustrative aircraft has an empennage 114 in the form of an inverted V-tail that includes a vertical stabilizer 120 and inverted stabilizers 121.

Referring to FIGURE 1C, control effectors that are controlled in combination with center-of-gravity control are shown for the supersonic aircraft 100. Two sets of surfaces are available for pitch control including the canards 118 and ruddervators 124. Roll control uses ailerons 128 and high speed spoilers 130. Yaw control is supplied by a rudder 140, ruddervators 124, and differential canard 118.

In combination with the canards 118, the supersonic aircraft 100 has multiple stability and control effectors. The canard 118 and symmetric defections of the ruddervators 124 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.

In an illustrative embodiment, trailing edge (TE) flaps 132 are deployed to generate additional lift during landing. TE flap deployment reduces angle-of-attack specifications during landing. During second-segment climb, the TE flaps 132 are extended to improve the lift-to-drag ratio for better climb performance. Leading edge (LE) Krueger flaps 134 are extended for low speed operations including takeoff, approach and landing. The LE Krueger flaps 134 improve lift-to-drag ratio, 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 a full-span leading edge Krueger flaps 134. Aircraft center-of-gravity can be controlled concurrently with the TE flaps 132 and Krueger flaps 134 to move the center-of-gravity more aft during supersonic operation and more forward during takeoff, approach, and landing.

The multiple control surfaces of the supersonic aircraft 100, for example the ruddervators 124 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 124 include multiple surfaces, increasing fault tolerant capability and supplying redundant elements for improved reliability.

Referring again to FIGURES lA, 1B, and 1C, in the illustrative aircraft 100, shaping of the wing 104, body 101, empennage 114, 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 114 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 124, adjust the aircraft longitudinal lift distribution throughout the flight envelope to maintain a low-boom, low-drag trim condition. The canard 118 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 118, an optimized wing 104, incorporation of leading edge flaps 134 and spoilers 130, and a reconfigured body or fuselage 101. The canard 118 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 114 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.

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 114 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 118 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 114.

In some embodiments, an aircraft designed for transonic speed comprises an airfoil, a nacelle, an engine, and an inverted V-tail. The airfoil has a leading edge, a trailing edge, upper and lower aerodynamic surfaces, and inboard and outboard ends. The nacelle is mounted on the lower aerodynamic airfoil surface and behind the trailing edge of the airfoil. The engine is enclosed within and structurally supported by the nacelle. The inverted-V tail is coupled to the airfoil at a position on the upper aerodynamic surface directly across the airfoil surface so that the inverted-V tail structurally supports the nacelle and engine in combination with support from the airfoil.

According to another embodiment, an aircraft comprises at least one engine, at least one nacelle encasing the engine, a fuselage, an inverted V-tail coupled to the fuselage, and wings.

The wings are bilaterally coupled to the fuselage and coupled to the nacelle. The wings having a leading edge, a trailing edge, upper and lower aerodynamic surfaces, inboard ends attached to the fuselage, and outboard ends. The wings have a dihedral gull incorporated inboard of the nacelles, the dihedral being most pronounced at the wing trailing edge.

In additional embodiments, an aircraft comprises a fuselage, a wing coupled to the fuselage, a nacelle capable of containing and supporting an engine, and a diverter. The diverter is coupled between the nacelle and the wing. The wing has a portion inboard toward the fuselage that integrates with the nacelle and the diverter in a configuration that follows the fuselage contour with a substantially normal intersection so that interference drag is reduced.

In a further embodiment, an aircraft comprises a fuselage, a wing coupled to the fuselage, a nacelle, and an inverted V-tail. The wing has a leading edge, a trailing edge, upper and lower aerodynamic surfaces, and inboard and outboard ends. The nacelle is mounted to the lower wing surface behind the trailing edge. The inverted-V tail is coupled to the fuselage and forms a structural coupling between the fuselage that is capable of supporting the nacelle in a position that reduces flutter.

According to another embodiment, an aircraft designed for transonic speed comprises an airfoil and a nacelle. The airfoil has a leading edge, a trailing edge, upper and lower aerodynamic surfaces, and inboard and outboard ends. The nacelle is mounted on the lower aerodynamic airfoil surface and has an upper surface with a dip that generates a reduction in drag.

In still another embodiment, a supersonic aircraft comprises an airfoil, an inverted V-tail coupled to the airfoil capable of generating trim in cruise with reduced sonic boom impact, an engine including an inlet, and a nacelle. The nacelle encases the engine and is coupled to the

airfoil lower surface. The nacelle and inlet have an integrated geometry that improves low-sonic- boom compatibility and reduced inlet/nacelle drag.

Various embodiments can be implemented to reduce sonic boom, increase aircraft performance, and reduce systems packaging complexity. Aerodynamic effects associated with deflecting wing, V-tail control surfaces and ruddervators, and canards enable aerodynamic performance suitable for supersonic and transonic flight with a reduced sonic boom.

In some embodiments, the aircraft includes a wing with a relatively large surface area sized to attain a selected cruise performance based on a cruise drag reduction level relative to status drag, propulsion system lapse rate characteristics, and aircraft design cruise Mach number.

The aircraft wing surface area can be selected based on low-speed considerations including takeoff and landing performance; and wing/engine ratio is selected based on engine scale factor, design cruise Mach number, and a goal level of drag reduction.

Some of the disclosed embodiments include wings that have a substantial dihedral, or "gulling"incorporated into the wing inboard of the engines. The gull or dihedral results from twisting and cambering the wing for low-boom and low induced drag while preserving a tailored local wing contour in the position of main landing gear retraction. In some embodiments, the dihedral geometry is most pronounced at the wing trailing edge. In some embodiments, the wing includes Krueger flaps and the leading edge of the wing extends in an essentially straight line to facilitate using a simple hinge line that accommodates the Krueger flaps. Some embodiments have a wing with reduced leading and trailing edge sweeps.

Several of the disclosed embodiments have an inverted V-tail that structurally couples the wing and fuselage for low-sonic-boom performance. The inverted V-tail functions in combination with the wings to support nacelles in a highly appropriate position relative to the wing. The added nacelle support of the inverted V-tail reduces or eliminates difficulties with flutter, facilitates vehicle trimming, and enables optimum low-sonic-boom lift distribution. The inverted-V tail can include control surfaces, for example ruddervators, that adjust aircraft longitudinal lift distribution throughout the flight envelope to maintain a low sonic boom and a low drag trim condition.

Another distinctive characteristic of some disclosed embodiments is the integration of wing, nacelle, and diverter to follow the contour of a low-sonic-boom fuselage with as close a normal intersection as possible to attain a low interference drag. Some embodiments have an inboard flap hinge line that is fully contained within the wing contour with the wing upper and lower surfaces held as planar as is possible to facilitate seal design.

In some embodiments, the wing has a dihedral or gulling that enhances low-sonic-boom signature by vertically staggering the wing longitudinal lift distribution. In some examples the wing gull wraps around the nacelle and enhances favorable interference between the engine inlets and the wing, resulting in a wing/body/nacelle geometry that is conducive to successful ditching and gear-up landings. In some embodiments, the gulling lowers the aircraft body or fuselage to reduce the height of the aircraft cabin above the ground and reduce entry stair length.

Referring to FIGUREs 2A and 2B, perspective top and bottom views, respectively, of the supersonic aircraft 100, showing an airframe 200 including a body structure 201, a wing structure 202, a tail structure 203, a nacelle structure 204, and control surfaces 205.

FIGUREs 2A and 2B in combination with FIGUREs lA, 1B, and 1C show various structural aspects of the engine structural integration system 103. The wings 104 have a substantial dihedral, or"gulling"208 incorporated into the wings 104 inboard of the engines 116.

The dihedral geometry is most pronounced at the wing trailing edge 210. The gull or dihedral results from twisting and cambering the wing 104 for low-boom and low induced drag while preserving a tailored local wing contour in the position of main landing gear retraction.

In some embodiments, the inboard portion 214 of the wing 104 is configured to integrate with the nacelle 122 and a diverter 216 formed between the nacelle 122 and the wing 104 to follow the contour of a low-sonic-boom fuselage 142 with as close a normal intersection as possible to attain low interference drag. In some embodiments, an inboard flap hinge line 218 is fully contained within the wing contour with the wing upper 220 and lower 222 surfaces held as planar as possible to facilitate seal design.

With the resulting wing configuration, the wing gull 208 raises the engines 116 to increase available tipback angle and reduce thrust-induced pitching moments. The gull 208 enhances low-boom signature by vertically staggering the wing longitudinal lift distribution and lowers the aircraft body or fuselage 142 to reduce the height of the cabin door 138 above the ground, thereby reducing entry stair length. The low fuselage 142 assists in maintaining a low aircraft center of gravity, reducing tipover angle and promoting ground stability. The wing gull 208 forms a wrapping of the wing 104 around the nacelle 122 that enhances favorable interference between the inlets 119 and the wing 104, resulting in a wing/body/nacelle geometry conducive to successful ditching and gear-up landings.

Referring to FIGUREs 3A, 3B, 3C, and 3D in combination with FIGUREs lA, 1B, and 1C, a group of schematic pictorial diagrams illustrate multiple aspects of an automated fuel transfer system 300 that can be used in a supersonic aircraft 100. The aircraft 100 includes a

fuselage 101 and wing 104. The automated fuel transfer system 300 comprises a plurality of fuel tanks 302 distributed within the wing 104 and/or the fuselage 101, a plurality of pumps 304 coupled to the plurality of fuel tanks 302 and capable of transferring fuel among the plurality of fuel tanks 302, at least one sensor 110 capable of indicating a flight parameter, and a controller 112. The controller 112 is coupled to the sensors 110 and the plurality of pumps 304, and controls fuel transfer among the plurality of fuel tanks 302 to modify the aircraft lift distribution, thereby reducing or minimizing aircraft sonic boom.

FIGURE 3A is a perspective view of a fuel tank arrangement in a system embodiment, showing an example of positioning and capabilities of the tanks. The fuel tanks 302 include wing tanks 310, a forward fuselage tank 312, an aft fuselage tank 314, and a vertical tail tank 316. The fuel tanks 302 are divided into left and right tanks by a center keel 318 inside the forward 312 and aft 314 fuselage tanks. The forward 312 and aft 314 fuselage tanks are separated by the landing gear bay 320. The illustrative system 300 includes two feed tanks 322, one vent tank 324, and one dry bay (not shown) for the left and right fuel tanks respectively. In the example, vent tanks 324 are located on the outer edge of each wing 104. In other embodiments, for example for configurations in which fuel is carried in the vertical tail 120, the vent tanks 324 may be in another position.

In an illustrative embodiment, the fuel system 300 has three component subsystems including a feed system, a refuel/defuel system and a vent/pressure system. The fuel system supplies and delivers an appropriate fuel amount to aircraft main engines and an Auxiliary Power Unit (APU) for a flight mission. In the illustrative embodiment, the fuel system 300 includes four major fuel tanks, pumps, valves and tubing connections between the fuel tanks, engines and the APU. The forward body fuel tank 312 is located between the nose landing gear wheel well and the main landing gear wheel well. The aft body fuel tank 314 is located between the main landing gear wheel well and the APU. The left and right wing fuel tanks 310 are located from the wing root out to the wing tip of each wing 104. The vertical tail fuel tank 316 can operate as a fuel tank and can be plumbed into the aft body fuel tank 314.

The fuel feed system 300 can supply a substantial fuel flow to each engine 116 for takeoff using two pumps per side at sea level and standard conditions. The fuel transfer system 300 can jettison a substantially larger amount of fuel, if desired. The fuel transfer system 300 enables center of gravity control for selected conditions including emergency flight scenarios. The re- fueling system assists rapid refueling of an aircraft, even if the tanks are fully empty, within a short time using multiple fuel trucks. The fuel feed system 300 incorporates feed tanks 302 that can supply a small amount of fuel to the engine 116 from the forward or aft feeder tank 322

without increasing pump power. The fuel vent system pressurizes the fuel tanks with nitrogen enriched air and maintains constant fuel tank pressure while the outside air pressure is changing due to changes in altitude.

The total fuel volume is the sum of the volumes in the forward body 312, aft body 314, left wing 310, right wing 310, and vertical tail 316 fuel tanks. Fuel tank partitions separate fuel tanks in the wing, the body and the vertical tail. A vertical keel partition separates the left side and the right side of both the forward body 312 and aft body 314 fuel tanks.

FIGURE 3B is a schematic block diagram of the fuel feed system 300 showing a pair of fuel feed pumps 304 in a forward feed tank 326 and a pair 304 in the aft feed tank 328 to supply fuel to each engines 116 and to the APU 330. Ejector pumps and fuel feed lines are integrated to supply fuel to the engines 116 and the APU 330. The fuel system 300 enables fuel from each tank to be burned in sequence to control aircraft center of gravity. Most fuel system components are positioned inside the tanks 302. Boost pumps 304 are located outside the tanks 302 to facilitate accessibility and maintenance without defueling the aircraft 100. Dual boost pumps 304 in the forward 326 and aft 328 fuselage feed tanks generate fuel flow. Fuel from the forward fuselage tank 312 is supplied to each engine 116 first to begin shifting the aircraft center of gravity aft in preparation for supersonic flight. Once fuel in the forward fuselage tank 312 is consumed to a predetermined level, the aft fuselage dual boost pumps 304 continue supplying fuel to respective engines 116. A fuel scavenge system removes the remaining fuel in the forward tanks 312. A cross feed valve 332 connects the left and right fuel feed manifold 318 in the event of total fuel failure on either side. An intertank shut off valve 334 between the forward 312 and Aft 314 fuselage tanks transfers fuel from one side to the other during the flight in response to fuel imbalance.

The fuel feed system 300 has ejector pumps 304 that assist fuel transfer from forward section to the middle section of the wing fuel tank 310, from the middle section to the aft section, from the aft section of the wing tank 310 to the aft feed tank 314, from the aft body front tank to the aft feed tank and from the forward body front tank 312 to the forward feed tank. One-way flapper valves in the fuel tank partitions enable fuel to flow in the same direction as the ejector pumps transfer fuel. One-way flapper valves are also located in the forward and aft body partitions, enabling fuel to flow from the rear part to the front part of both body tanks 312 and 314. Both fixed and movable capacitance probes are located in the wing fuel tanks 310 and indicate fuel level in each fuel compartment. Float switches at the forward body tank deactivate the forward feed pump and activate the aft feed pump when fuel levels are low in the forward body tanks.

The fuel feed system 300 has independent left and right side systems to feed fuel through fuel lines from the forward 312 and aft 314 fuel tanks to the engines 116. The fuel feed for the Auxiliary Power Unit (APU) 330 is from the left side feed system through a feed line. Two fuel- to-hydraulic heat exchangers are located in each of the aft body fuel tanks 314. A left to right cross feed shut off valve 334 is located between the right aft feed pumps and left aft feed pumps.

When the cross feed valve 334 is in the open position, cross feed of fuel from either the left or right feed pumps can occur. One-way check valves in the outlet of each feed pump 304 to prevent back flow of fuel from other feed pumps.

FIGURE 2C shows a schematic block diagram of a refuel/defuel system 336 including re-fueling and de-fueling receptacles 338 located in the right outboard wing lower leading edge.

Four fuel transfer pumps 340 are located at the rear body tank 314. Each fuel transfer pump 340 can transfer a substantial flow rate from the aft tank 314 to either of wing tanks 310, the forward body tank 312, or jettison the fuel overboard. In an illustrative embodiment, the re-fuel and de- fuel system 336 is integrated into the aircraft with appropriate sizing for rapidly refueling the aircraft with reasonable turnaround time for re-fuel operations using a commercially available fuel source and fuel handling support equipment. The refuel/defuel system includes a jettison capability.

In the illustrative aircraft 100, the refuel/defuel station 338 has two refueling and one defueling receptacle. The system 336 has one fuel valve 342 for the forward 312, aft 314, and wing 310 tanks. The aircraft refueling system 336 refuels the fuel tanks sequentially. Refueling begins with the forward fuselage tank 312. Once the forward fuselage tank 312 is full, fuel begins spilling to the forward wing tanks 310 through standpipes. The aft fuselage tank 314 beings filling next and, when full, fuel spills to the wing tanks 310. Wing tanks 310 are the last tanks to refuel. A fueling float switch in the wing tank 310 automatically closes the fueling valve when the fuel quantity reaches capacity. Receptacles are spaced such that two fuel trucks can supply fuel to the aircraft simultaneously.

The defuel system enables removal of fuel from each tank and enables transfer of fuel between tanks on the ground. The fuel boost pumps 304 are used to get fuel out of the tanks 302 and into the feed manifold. With the defuel valve open, fuel transfers to the defuel station.

The transfer system transfers fuel from the aft fuselage tanks 314 to the forward fuselage tanks 316. The transfer system is configured and sized to transfer fuel forward during supersonic to subsonic transition and dual engine flame out where the center of gravity is modified to control the aircraft 100.

FIGURE 2D is a side pictorial view of the aircraft 100 showing the aircraft fuel envelope profile 250. The fuel system 300 has multiple fuel tank spillover lines including a forward body tank spill-over line 252, aft body tank spill-over line 260, and vertical tail vertical tail tank spill- over 262 which allow fuel from one tank with a supply spout to flow to the adjacent fuel tank.

Fuel from the vertical tail fuel compartments 316 flow to the aft body rear fuel compartment 314 by gravity drains, the vertical tail to body tank feed lines 258 and 264. Two fuel transfer pumps per side, for example the illustrative left aft body transfer pumps 274, are located in the aft body front fuel compartment. Each transfer pump can transfer fuel at a substantial rate to the forward body front fuel compartment 312, the aft wing fuel compartment 310, or jettison fuel out aft of the wing rear spare.

Fuel can be transferred from the forward 312 and aft 314 body fuel tanks to other fuel tanks with a spout, specifically the forward body tank transfer line and spout 266, the aft body tank spout 276, and the vertical tail tank transfer line and spout 280. A shut off valve in each spout, in particular the forward body tank transfer shut-off valve 254, the forward inter-tank shut- off valve 268, the aft inter-tank shut-off valve 272, and the vertical tail tank transfer shut-off valve 278, enables fuel to be transferred to the tank with the shut off valve in the open position. The fuel transfer line 270 carries fuel throughout the aircraft 100. FIGURE 2D also shows the refuel and defuel panel 256.

The controller 112 transfers fuel among the plurality of fuel tanks 302 to adjust the aircraft center of gravity so that fuel in the forwardmost tanks is consumed first, configuring the aircraft trim on attaining cruise condition at a maximum aft center-of gravity for a reduced sonic boom condition.

Referring again to FIGUREs 1A, 1B, and 1C, and FIGURES 2A, 2B, 2C, and 2D, the schematic pictorial and block diagrams illustrate a control system that can be used in a supersonic aircraft 100. The aircraft 100 includes a fuselage 101 and wing 104. The control system comprises a plurality of control effectors coupled to the wing 104, a plurality of fuel tanks 106 distributed within the wing 104 and/or the fuselage 101, and a plurality of pumps 304 coupled to the plurality of fuel tanks 302 and capable of transferring fuel among the plurality of fuel tanks 302. In the illustrative embodiment, the control effectors include canards 118, ruddervators 124, ailerons 128, high speed spoilers 130, and rudder 140. The system further comprises a plurality of actuators coupled to the control effectors, at least one sensor 110 capable of indicating a flight parameter, at least one vehicle management computer 112 coupled to the sensors 110, the plurality of pumps 304, and the plurality of actuators. The vehicle management computer 112 can manage the control effectors and transfer fuel among the plurality of fuel tanks 302 to adjust

aircraft trim and center of gravity position to operate the aircraft 100 in at least two flight modes.

The variable flight modes have different trim drag and sonic boom performance.

The vehicle management computer or computers 112 can operate the aircraft 100 in a maximum range, maximum Mach over water mode with control effectors deployed for relatively reduced trim drag and center of gravity positioned relatively forward. In another mode, the vehicle management computers 112 can operate the aircraft 100 in a slightly reduced range, relatively lower Mach over land mode with control effectors deployed for a slight increase in trim drag and center of gravity positioned relatively aft to attain suitable aerodynamics at a reduced sonic boom level.

In some embodiments, the vehicle management computer or computers 112 can control the fuel tanks 302 to burn in sequence for aircraft center of gravity so that fuel in the forwardmost tanks is consumed first, and configuring the aircraft trim on attaining cruise condition at a maximum aft center-of-gravity for appropriate aerodynamic flight at an attenuated sonic boom condition.

In some embodiments, the aircraft 100 has multiple fuel boost pumps 304 positioned outside of the fuel tanks 302 for the ease of accessibility and maintenance without defueling the aircraft. The fuel boost pumps include dual boost pumps in forward and aft fuselage feed tanks 302. Fuel from the forward fuselage tank is supplied to engines 116 first to begin shifting the aircraft center of gravity aft in preparation for supersonic flight. When fuel in the forward fuselage tank is consumed to a predetermined level, aft fuselage dual boost pumps continuing supplying fuel to the engines 116.

In some embodiments, the aircraft 100 can comprise a fuel scavenge system that removes remaining fuel in fuel tanks using a cross feed valve connecting left and right fuel feed manifold in the event of total fuel failure on either side. An intertank shut off valve between forward and aft fuselage tanks for transferring fuel from one side to the other during the flight due in case of fuel imbalance.

In a particular embodiment, the vehicle management computer or computers 112 can perform multiple fuel management operations, many of which affect or control center of gravity.

In one operation, the computers 112 transfer fuel among the plurality of fuel tanks to adjust the aircraft center of gravity to adjust the aircraft center of gravity and reduce trim drag, increasing aircraft range. Management of center of gravity can be used to adjust the aircraft center of gravity to reduce trim criteria to increase aircraft controllability. Adjusting center of gravity is performed to maintain aircraft stability during flight. In some examples, adjusting aircraft center of gravity

can be used to adjust the aircraft longitudinal lift distribution throughout the flight envelope to maintain a low-boom, low-drag trim condition.

Referring to FIGURE 4, a schematic block diagram illustrates an embodiment of an aircraft lift device 400 including control actuators that can be operated in combination with center of gravity management to control the aircraft. The aircraft lift device 400 comprises an aircraft wing 104 that can mount onto an aircraft fuselage. The aircraft lift device 400 has a leading edge 402 extending along the wing inboard 404 to outboard 406, and a strake 401 that can couple to the aircraft fuselage and extend from the fuselage to the leading edge 402 of the wing 104. In the illustrative embodiment, the leading edge 402 is formed as a Krueger flap 408 that is outboard of the strake 401 and inboard of a simple flap 410. The Krueger flap 408 and the simple flap 410 generally have different leading edge structures. In other embodiments, the entire leading edge may be a single structure or may have multiple leading edge segments. For example, in some embodiments, the Krueger flap 408 can extend from the strake 401 to the wing tip. The aircraft lift device 400 further comprises a Krueger flap 408 coupled to the leading edge 402 at a relatively inboard portion of the wing adjacent the strake 401, and a simple leading edge flap 410 coupled to the leading edge 402 of the wing 104 and extending from a junction 412 at the Krueger flap 408 to an outboard portion 406 of the wing 104.

The aircraft uses the active center-of-gravity (CG) management system in combination with the control effectors. As fuel is burned throughout the mission, the CG management system redistributes the remaining fuel to maximize range and minimize sonic boom signature. The CG management system enables the canard, wing and inverted V-tail to work in harmony to lift the vehicle efficiently for maximum range while producing a low sonic boom signature. In operation, the leading edge flaps, including the Krueger flaps 408 and the leading edge flaps 410, are extended for low speed operations during takeoff, approach, and landing while the CG management system positions the center-of-gravity in a relatively forward position in comparison to the aft position during supersonic flight that reduces sonic boom level.

Referring to FIGUREs 5A and 5B, two schematic pictorial diagrams show an embodiment of a Krueger flap 500 that can be controlled in combination with center of gravity management to control the aircraft. Krueger flaps 500 are aerodynamically-effective movable components on the leading edge of the airfoil, high-lift devices that supply additional lift in certain configurations and under certain flight attitudes. Krueger flaps 500 are connected to the leading edge 502 of the wing 508 and extend from the wing lower surface 504 to increase lift capability during low-speed operation. High-lift devices, such as Krueger flaps 500, facilitate

lift-off and landing at low speeds, and maintain undisturbed wing root airflow over the wing upper surface 506 without separation at the transition from fuselage to wing 508.

From the stowed position, the rotary actuators 510 can rotate the Kruger flap 500 downward and forward from the lower surface 504 of the wing 508. The illustrative Krueger flap 500 shows one example of a suitable rotary actuator 510 that is suitable for usage in a wing 508 or other airfoil. In general, any Krueger flap with appropriate configuration, aerodynamic configuration, and actuating mechanism can be used. Generally, a suitable Krueger flap has an actuating mechanism capable of forming the wing leading edge configuration into a rigid airfoil structure at multiple different operating positions maintaining short and efficient load paths.

Furthermore, a suitable Krueger flap has a control linkage mechanism that is stable at the different operating positions and deflects downward when actuated through a range of selected rotational angles while maintaining a substantially smooth wing surface with an aerodynamic, relatively constant radius of curvature. The actuating linkage operates to controllably stow and deploy the flap 500 during takeoff and landing, and for usage as a speed brake, if desired, during either high or low-speed in-flight operating conditions.

FIGURE 5B shows a close-up view of a portion of the Krueger flap 500 in greater detail.

Details shown include a left wing front spar 512, left Krueger flap hinge point 514, a flight spoiler hinge beam 516, left leading edge rib 518, and left outboard flight spoiler 520.

Referring to FIGURE 5A and 5B in combination with FIGURE 4, the rounded form 404 of the inboard portion of the leading edge flap 410 smoothly transitions to the form of the Kruger flap at the Krueger flap junction 406 to reduce or minimize any gap in the wing leading edge.

Referring again to FIGURE 4, the leading edge 402 of the wing 104 is configured so that the shape of the leading edge flap 410 merges into the form of the Krueger flap 408. In particular, the structure and configuration of the leading edge flap 410 and the Krueger flap 408 are arranged so that when the Krueger flap 408 is deployed, air flow separation over the wing 104 is reduced or minimized. The cross-sectional morphology of the leading edge flap 410 is matched to the Krueger flap 408 to avoid structural discontinuities, protrusions, or gaps that can create a vortex at a position along the leading edge 402, such as at the junction between the Krueger flap 408 and the leading edge flap 410. A vortex formed at the top of the wing 104 corrupts the flow field.

The leading edge flap 410 avoids flow field corruption via usage of rounded edges and structures in the Krueger flap 408 and the leading edge flap 410, particularly in the vicinity of the junction.

In various embodiments, the junction between the leading edge flap 410 and the Krueger flap 408 can have some structural discontinuity. For example, the junction can include a step

variation, although a gap in flap continuity between the Krueger flap 408 and leading edge flap 410 segments can impact aerodynamic characteristics. In some embodiments, a structural element that smoothes the transition between segments can be used to improve aerodynamic performance. In some embodiments, the structural material can be a flexible material such as rubber, plastic, a synthetic, and the like.

The particular structure of the Krueger flap 408 and the leading edge flap 410 can vary depending on the wing configuration. For example, whether the leading edge 402 is a true supersonic leading edge. In particular, whether the leading edge is contained within the Mach cone of the aircraft. If the leading edge 402 is inside the Mach cone, structural discontinuities, protrusions, and gaps are to be avoided. For a leading edge 402 that is outside the Mach cone, the leading edge flap 410 can include more irregular structures such as a sharp edge transitioning to a Krueger flap structure.

Any suitable element or structure can be used to mate the Krueger flap 408 and the leading edge flap 410 when either stowed or deployed. Generally, the portions of the Krueger flap 408 and the leading edge flap 410 at the junction can be formed so that the edges of each have similar shape, thereby reducing or eliminating structural discontinuity at the junction.

Referring to FIGUREs 6A, 6B, 6C, 6D, and 6E, multiple schematic pictorial diagrams illustrate various control effectors that can be controlled in combination with center of gravity management. FIGURE 6A depicts control effectors of the wing 104 and empennage 114 in more detail. The empennage 114 includes a tail structure section 602, a vertical stabilizer to inverted stabilizer joint section 604, and an inverted stabilizer to nacelle joint section 606. The tail structure section 602 includes the vertical stabilizer 120, and a pair of inverted stabilizers 121.

Control effectors include the rudder 140 pivotally connected to the trailing edge of the vertical stabilizer 120 and ruddervators 124 pivotally connected to the trailing edge of the inverted stabilizers 121. The vertical stabilizer 120 is attached to the top of the aircraft center body and aft section 608. The top of the vertical stabilizer 120 is attached to the tops of the left and right inverted stabilizers 121.

Referring to FIGUREs 6B and 6C, two pictorial diagrams illustrate different views of a trailing edge flap 132 that can be used in an embodiment of an aircraft capable of channel relief control. The trailing edge flap 132 is located between the engine nacelle 122 and the fuselage 101. 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. The engine nacelle 122 has sufficient clearance for the flap 132 to deflect to a maximum deflection angle. An actuator 610 drives deflection of the trailing edge flap 132.

FIGURE 6C shows the trailing edge flap 132 attached at the inboard wing rear spar 612.

The trailing edge flap 132 is pivotally connected to the wing rear spar 612 via a flap hinge 614 and an actuator hinge 616.

Referring to FIGUREs 6D and 6E, two schematic perspective pictorial views show detailed diagrams of portions of the tail structure 602. FIGURE 6D shows a ruddervator section 618 including the left inverted stabilizer 121L coupling between the vertical stabilizer 120 and the left wing adjacent to the left nacelle 122L. 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 124LO and a left inboard ruddervator 124LI are shown coupled to the left inverted stabilizer 121L using ruddervator hinges 620 and actuator hinges 622 that control movement of the ruddervators. FIGURE 6E illustrates baseline actuators 624 for the ruddervators 124. In the illustrative embodiment, the actuators 624 are electro-mechanical rotary- type actuators that, when integrated into the inverted V-tail, do not protrude into the airstream and thereby avoid increases in aerodynamic drag. The illustrative baseline actuators 624 have integral motor drivers, brakes, and a speed sensor. An aircraft can include multiple ruddervators 124 on each side, a redundancy that is useful to maintain aircraft control in a jammed-surface condition.

Referring to FIGUREs 7A and 7B, two schematic pictorial diagrams depict views of additional control effectors on the aircraft 100, canards 118 coupled to the fuselage 101 forward of the wing 104. FIGURE 7A shows a top, cut-away view of the aircraft 100 embodiment in the vicinity of the canard 118. The canard 118 can be particularly effective during takeoff and in high-speed flight. The canard 118 augments the rudder 140 by supplying substantial yaw control power when the left and right canard surfaces are deflected differentially. The diagram shows left and right canard control surfaces 702L and 702R, canard leading edges 704L and 704R, and canard rotation joints 706L and 706R. Also shown is the body or fuselage 101 enclosing a flight crew compartment 708 and a passenger compartment 710. The left and right canard control surfaces 702L and 702R can pivot about the rotation joints 706L and 706R.

Referring to FIGURE 7B, a schematic pictorial diagram shows a top, cut-away view of a left canard 118. The canards 118 are each driven by a linear electromechanical actuator (EMA) 712. The canard 118 is used to control pitch and can also be dithered for yaw. In alternative embodiments, a hydraulic actuator can be used to drive motion of the canard. The actuators 712 to multiple canards 118 enable differential control of the canards 118 to induce lift on the fuselage 101 and the wing 104 on opposing sides of the body 101 to cause canard lift and body lift to blend into lift produced by the wing 104.

Referring to FIGURE 8, a schematic block diagram shows an example a flight control actuation architecture embodiment 800 that can be used as the controller 112 for concurrently managing control effectors and fuel system pumps to manipulate aircraft center-of-gravity. In the illustrative example, primary flight control actuation uses"Fly-by-Wire"dual tandem linear hydraulics with triple electronic redundancy. Dual tandem actuation 802 is powered by two independent hydraulic systems 804 and 806 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 808. The flight control system 800 performs control law implementations to produce aircraft handling qualities throughout flight. The system 800 can implement outer loop control modes such as Autopilot, Autolanding, and Auto collision avoidance control. The flight control actuation system 800 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 flight management computers 808 can implement a process that differentially controls the canards 118 to induce lift on the fuselage 101 and the wing 104 on respective opposing sides of the fuselage body 101 to cause lift from the canard and body lift to blend into lift produced by the wing. The computers 808 further control the canards 118 to stretch the aircraft lifting length and tailor the effective area distribution to produce a shaped sonic boom signature. Differential control of the canards 118 can be used to offset effects of the canard dihedral. The flight management computers 808 control the canards 118 in combination with control of center-of-gravity to reduce trim drag.

The control effector configuration, controlled by the Vehicle Management Computers 808, 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 808 implement processes for controlling the effectors, including the canards 118, in combination with center-of-gravity control 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 9, a schematic block diagram shows an embodiment of a suitable hydraulic power and distribution system architecture 900 for supplying actuating power to effectors and the center of gravity management system. For high reliability, the system 900 is highly redundant with a hydraulic system supplying three independent sources 902,904, 906 of hydraulic power to operate primary flight controls, landing gear 914, nose wheel steering 916, wheel brakes 918, and thrust reversers 920. The three independent systems 902,904, 906 give triple redundancy for continued safe flight and landing.

Hydraulic power for the systems is supplied by two engine driven pumps 922 and an AC motor pump 924 on system 1 902 and system 2 904. The engine driven pumps 922 can operate continuously while the AC motor pumps 924 operate on demand basis. Additionally, the AC motor pumps 924 are an extra source of hydraulic power that gives redundancy within each system. The AC motor pumps 924 can be operated on the ground for system checkout without running the engines or using a hydraulic ground carts.

System 3 906 has two air driven pumps 926 and an AC motor pump 924. One air driven pump 926 operates continuously while the other air driven pump 926 and the AC motor pump 924 operate on a demand basis. The AC motor pump 924 in system 3 906 can also be operated on the ground for system checkout without running the engines or using a hydraulic ground cart. System 3 906 also includes a ram air turbine 928 for emergency hydraulic and electrical power in the event of dual engine flameout. The ram air turbine 928 is sized to supply hydraulic and electrical power to essential equipment from the certified altitude to safe landing for level 3 handling quality.

Referring to FIGURE 10, a graph shows a center-of-gravity envelope for an aircraft embodiment. The center of gravity envelope shows the range of aircraft center of gravity from extreme forward to aft that can be attained by the aircraft, along with a nominal, median excursion. The graph shows subsonic and supersonic aerodynamic limits. The graph does not

show the expected excursion for typical mission or the rate at which the center of gravity can be moved using fuel management.

Referring to FIGURE 11, a graph shows an example of a control configuration that can be controlled by the Vehicle Management Computers 112 in an embodiment of the supersonic aircraft to attain longitudinal stability and control during takeoff and landing. Pitch axis static stability and controllability are assessed by determining the lift coefficient (CL) at a range of aircraft baseline pitch moment coefficients (CM) with all control surfaces at a null position as shown in the graph. The graph shows an example of a nominal center of gravity (CG) range of an aircraft embodiment.

Primary pitch control surfaces that can be controlled in combination with center-of- gravity include the canard and the ruddervators. Total pitch control power is supplied by full deflections of the canard and the ruddervators, shown in the CL vs. CM plot for the low speed takeoff 1100 and landing 1102 condition. In the example, full canard trailing edge down deflection is scheduled as a function of angle-of-attach alpha (a) to prevent canard stall.

Intersections of center of gravity (CG) lines with the CL-CM curves are trim controls. Trim control is appropriate for the nominal CG range of the aircraft in takeoff 1100 and landing 1102 configurations.

Referring to FIGURE 12, a graph shows an example of a control configuration that can be controlled by the Vehicle Management Computers 112 in an embodiment of the supersonic aircraft to attain longitudinal stability and control during supersonic cruise. The lift coefficient (CL) vs. pitch moment coefficient (CM) plot is depicted for a supersonic cruise condition of Mach 1.8 and includes flexible effects due to aircraft bending. The illustrative aircraft embodiment is stable in the pitch axis in the supersonic cruise condition. Moving the center-of-gravity (CG) aft reduces canard trim. In the center-of-gravity (CG) range from about 40% to approximately 50%, the aircraft has adequate control power for trim for the cruise angle-of-attack a of 2 to 3 degrees.

Referring to FIGURE 13, a schematic block diagram illustrates an embodiment of a fly- by-wire (FBW) flight control system 1300 that attains aircraft controllability via management of control effectors and center of gravity. The aircraft uses control laws to stabilize the aircraft and supply attain appropriate handling qualities to the pilot.

Control surface authority and actuator rate criteria can be predicted by evaluating closed- loop aircraft responses to various levels of atmospheric disturbance and to execute mission tasks for the entire flight envelope. The control laws are defined for three axes: pitch, roll and yaw.

For the pitch axis, angle-of-attack (or alpha) command system is used for the takeoff, approach

and landing conditions. The control system 1300 can attain precise alpha regulation for pitch control and is generally desired over manual handing by the pilot for satisfactory handling qualities. Calculation of control gains uses an automated algorithm based on the Dynamic Inversion theory. The desired closed-loop short-period eigenvalues are computed based on military standard (Mil-Std-1797) handling qualities criteria and are used for determinations of the feedback and feedforward gains. The gains are scheduled as a function of dynamic pressure. The handling qualities of the closed-loop vehicle are analyzed by using linear theory in the frequency- domain.

The inlet is a propulsion component that interfaces with airflow internal and external to an aircraft. The inlet functions to carry air required for combustion into the engine while modifying air stream conditions from free stream conditions to conditions suitable for usage at the entrance of a fan or compressor with minimum pressure loss and flow distortion. A fan or compressor commonly operates best with uniform air flow at a Mach number of approximately 0.5.

Installed engine performance is degraded by inlet installation losses including recovery, additive drag, cowl drag, bypass air, and boundary layer bleed drag. Various inlet designs can reduce installation losses.

Several factors operate in combination to affect inlet performance including flow uniformity and techniques for controlling flow matching, total pressure recovery, installation drag, starting and stability, and signatures such as acoustic, radar, and infrared. Inlet performance also includes considerations of weight, cost, durability, and reliability. Inlet design typically involves a balancing of these factors and considerations.

The function of an engine inlet for a supersonic aircraft is to supply a proper quantity and uniformity of air to the engine over a range of flight conditions from subsonic to supersonic.

Characteristics of supersonic flow complicate inlet design and integration of the inlet structure into the airframe. Various design considerations include considerations of inlet total pressure recovery and installed drag at cruise, location of the inlet on the aircraft, and the aircraft attitude envelope including angle of attack, angle of yaw, and crosswind takeoff. Further considerations include inlet total pressure recovery and distortion levels for engine operation, integration of the inlet exterior contour to the adjacent aircraft surface and to the fan or compressor. Further variables under consideration include boundary layer bleed air, bypass of excess inlet air, and suppression of engine fan noise.

Supersonic aircraft travel at a range of Mach number when accelerating to cruise after takeoff and deceleration at the end of a trip. What is desired is efficient operation over a range of Mach numbers from subsonic to supersonic to manage inlet total pressure recovery, drag, and distortion levels. Variable-geometry inlets can modify aerodynamic characteristics by adjusting to various flight speeds and conditions. Unfortunately, a variable-geometry inlet increases complexity, maintenance, and costs.

What is desired is an aircraft that is capable of efficient operation at supersonic speeds to achieve long-range travel at reasonable costs. What is also desired is a supersonic aircraft capable of suppression or management of sonic boom. What is further desired is an engine inlet that is suitable for usage in supersonic flight that avoids the maintenance and cost inherent in variable- geometry designs.

In some embodiments of the disclosed aeronautical system, an inlet for an aircraft engine comprises a fixed geometry axisymmetric inlet spike having a longitudinal axis and a curved exterior contour of varying height along the longitudinal axis, and an axisymmetric translating cowl. The cowl is mounted about the inlet spike and separated from the inlet spike by an annular duct. The inlet spike and translating cowl form an inlet with essentially isentropic external compression and no bleed.

The fixed geometry inlet avoids complex mechanical structures and actuator systems of variable geometry inlet systems, thereby reducing or eliminating cost, maintenance burden, and weight. The reduced weight of a fixed geometry system increases range and reduces fuel costs.

The fixed geometry inlet also can reduce operating cost and improve aircraft availability due to the reduced maintenance load. The fixed compression geometry inlet can also reduce or eliminate closed-loop, flight-critical control systems required by variable-geometry inlet designs. The fixed compression geometry design also avoids, reduces, or eliminates certification costs and risk.

The fixed compression geometry inlet design can also reduce or eliminate the difficulty in producing a suitable spike surface quality inherent in a variable-geometry inlet system. Suitable contours of the spike surface of a variable-geometry inlet can be difficult to attain due to the discontinuities and edges that can result from an articulating mechanical surface. In addition, the possibility of failure of the actuators or mechanical structures can lead to configurations with poor aerodynamics.

In other aircraft embodiments, an aircraft engine inlet comprises a cowl extending fore and aft along a longitudinal axis and encasing an interior duct and a fixed geometry center body.

The cowl includes a translating forward section and a stationary main body. The stationary main

body having a forward end with a rounded cowl lip. The fixed geometry center body extends fore and aft generally interior to the cowl and the interior duct. The forward portion of the center body has a contoured geometry that produces external compression with an initial oblique shock wave and isentropic compression focused forward and above the cowl forward end and no bleed.

In further embodiments, a supersonic aircraft comprises a fuselage, first and second wings coupled symmetrically to opposing lateral sides of the fuselage, and first and second engines respectively coupled beneath the first and second wings. The first and second engines further comprise fixed compression geometry axisymmetric external compression inlets with translating cowls and no bleed.

In some embodiments, an inlet for an aircraft engine comprises a cowl axisymmetric about a longitudinal axis and having a forward end and an aft end, and a center body. The cowl has a fixed portion and a translating portion coupled to a forwardmost end of the cowl fixed portion. The cowl encloses an axisymmetric interior duct and is translatable to enable additional air to enter the inlet at low speed flight Mach numbers. The cowl translation position is adjustable based on aircraft flight speed and conditions. The center body has a forward end and an aft end and extending parallel to the longitudinal axis within the interior duct. The center body has a fixed geometry. The inlet has isentropic compression and no bleed.

In additional embodiments, a supersonic aircraft comprises an aircraft main body, an engine coupled to the aircraft main body, and an inlet coupled to the engine. Air inflow into the engine is defined by the inlet and at least a portion of the aircraft main body. The inlet has a fixed geometry, a translating cowl, isentropic compression, and no bleed.

In another embodiment, an aircraft engine comprises a nacelle, a turbofan engine encased within the nacelle, an exhaust system coupled to the turbofan engine, and an inlet coupled to the turbofan engine. The inlet for an aircraft engine comprises a fixed geometry axisymmetric inlet spike having a longitudinal axis and a curved exterior contour of varying height along the longitudinal axis, and an axisymmetric translating cowl. The cowl is mounted about the inlet spike and separated from the inlet spike by an annular duct. The inlet spike and translating cowl forming an inlet with essentially isentropic external compression and no bleed.

Referring again to FIGUREs 1A, 1B, and 1C, the aircraft 100 is configured with an engine 116 and a fixed geometry compression inlet 119. The aircraft 100 is capable of supersonic flight and includes a fuselage 101, wings 104 coupled symmetrically to opposing lateral sides of the fuselage 101, and engines 116. In the illustrative configuration, the engines 116 are coupled

to lower surfaces of the wings 104. The engines 116 further have fixed compression geometry axisymmetric external compression inlets 119 with translating cowls 117 and no bleed.

In some embodiments, the engines 116 are twin non-afterburning turbofan engines. In some embodiments, the aircraft 100 has a comparatively low lift/drag ratio resulting from a configuration in which the wings 104 are thin, highly swept wings, the aircraft 100 has a plurality of distributed sharp edges, and the engines 116 have axisymmetric, external compression inlets 119.

In some embodiments, the engines 116 have an inlet design so that a Mach 1.8 cruise is attained in an overspeed condition and the engines 116 and inlets 119 configured to maximize range at Mach 1.6.

The illustrative supersonic aircraft 100 has one or more aspects that reduce the amplitude of sonic booms. In one aspect, the engines 116 are located in an aft position beneath the wings 104 in a configuration with a highly integrated wing/inlet geometry that produces a low-boom capability and low inlet/nacelle installation drag. Some aircraft embodiments can include an inverted V-tail 114 geometry that generates longitudinal trim in cruise in a manner that reduces sonic boom amplitude, and structurally supports an engine/nacelle system 115.

In one embodiment, the aircraft 100 has a design optimized for Mach 1.8 cruise with twin non-afterburning turbofan engines 116 set below and behind the wings 104. The engines 116 operate behind fixed compression geometry axisymmetric external compression inlets 119.

Engine cycle selection and engine sizing are determined based on considerations of reduced community and takeoff noise, operation at speeds in a transonic range, and suitable cruise thrust. The aircraft 100 has relatively large wings 104, for example with an area of approximately two thousand square feet, and a suitable engine takeoff thrust size to generate sufficient lift to attain suitable takeoff performance in even short airfields, for example with a runway lengths of about 6,000 feet, while having near-optimum lift-to-drag (L/D) performance during supersonic flight.

Aircraft length is set in conjunction with operating weight to achieve selected sonic boom overpressure levels. In a particular example, maximum takeoff weight is set at approximately 150,000 pounds and overall aircraft length set at approximately 130 feet. A variety of aircraft weight and length configurations can also suitably selected to attain a desired performance.

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 115 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.

In an illustrative embodiment, the aircraft 100 has an average lift/drag ratio (L/D) of 9.9 for a sustained cruise climb at 40-50 ft/min, compared to ratios higher than 10 for conventional business jets that are optimized for a subsonic mission. Design features of the present aircraft 100 that contribute to a comparatively low L/D include a thin, highly swept wing 104, sharp edges throughout the aircraft 100, and axisymmetric external compression engine inlets 119. The highly-swept, thin wings 104 use high energy vortices to create lift in takeoff and landing configurations. During second segment climb, aircraft performance is dominated by a capability to efficiently produce lift and thrust. Compared to conventional business jets, the illustrative aircraft generates more drag during second segment climb, resulting in a higher climb speed and a longer balanced field length.

Systems and subsystems of the aircraft can be composed of various suitable materials. In one specific embodiment, parametric weights are reduced by about ten percent in comparison to conventional aircraft by utilizing structural composite materials including aluminum-lithium skins for the wings, composite edges and control surfaces in the tail structures, and miscellaneous composite components in the body and in the nacelle/inlet structures.

Referring to FIGURE 14, a schematic pictorial diagram illustrates an embodiment of an inlet 1410 for an aircraft propulsion system 1400 that includes an axisymmetric translating cowl 1412 and a fixed geometry axisymmetric inlet spike 1414 or center body. The inlet spike 1414 extends along a longitudinal axis and has a curved exterior contour of varying height along the longitudinal axis. The axisymmetric translating cowl 1412 is mounted about the inlet spike 1414 and is separated from the inlet spike 1414 by an annular duct 1416. The annular duct 1416 has a size and form that create isentropic compression to an inlet throat 1418. The inlet spike 1414 and translating cowl 1412 form an inlet 1410 with characteristics of essentially isentropic external compression and no bleed.

The propulsion system 1400 is contained within a nacelle 1406 and has an external compression inlet 1410 configured for isentropic compression at a free stream Mach number up to 2.0, and an inlet throat Mach number nominally set in a range from 1.1 to 1.5. In a specific embodiment, the inlet 1410 is configured for isentropic compression at a free stream Mach number of approximately 1.8, and an inlet throat 1418 Mach number nominally set at approximately 1.3.

The axisymmetric translating cowl 1412 has a cowl lip 1420, an inner cowl wall 1422, and an outer cowl wall 1424. A translating cowl portion 1426 at the forward portion of the cowl 1412 is translated between extended and retracted with respect to a fixed cowl body 1428 using an actuation system 1440 to facilitate operation at different flight speeds and conditions. The actuation system 1440 is capable of translating the translating cowl portion 1426 fore and aft.

The inlet 1410 generates an initial oblique shock wave and focuses isentropic compression ahead of and above the cowl lip 1420 at the forward extremity of the translating cowl 1412.

The axisymmetric center body or spike 1414 includes a forward center body 1430, an exterior wall 1432, a cone 1434, a contoured body 1436, and an isentropic compression region 1438. The center body 1414 has a fixed and stationary position with respect to the propulsion system 1400.

The inlet 1410 is typically composed of a suitable metal, such as aluminum. Parts of the inlet and engine, such as the engine nacelle and inlet lip, can be composed of a more durable metal, for example titanium. In various embodiments, other metals and alloys may be used.

Referring to FIGURE 15, a schematic pictorial and airflow diagram illustrates airflow in an embodiment of an inlet 1410. The external compression inlet 1410, though suitable for either subsonic or supersonic applications, has an aerodynamic design that is best utilized for supersonic cruise applications. The airflow diagram shows the external compression inlet 1410 and flow areas. In various conditions, compression takes place through either one or a series of oblique shocks 1502 followed by a normal shock, or through one normal shock. Airflows are illustrated in free stream tube areas including inlet Ao ; engine Ao, and bypass Aobp areas. The external compression inlet 1410 has one or more oblique shocks and has an inlet throat 1418 at or near the leading edge 1504 of the cowl 1412. The inlet 1410 is configured so that the normal shock 1506 is positioned at or near the cowl lip 1420.

In some embodiments, the inlet spike 1414 includes a cone 1434 with an initial half angle in a range from 10° to 17° followed by an isentropic compression ramp 1438 with an angle increase in a range from 5'to 13'. The internal cowl 1422 has an angle in a range from 0° to 10° at the cowl leading edge, and the inlet throat to capture area ratio has a range from approximately 0. 70 to 0. 9.

In a particular embodiment, the inlet spike 1414 includes a cone 1434 with an initial half angle of approximately 15° followed by an isentropic compression ramp 1438 with an angle

increase of approximately 9°. The internal cowl 1422 has an angle of approximately 8°. The inlet throat to capture area ratio is approximately 0.73.

The inlet mass flow ratio is the ratio of the actual inlet mass flow rate to the mass flow rate of the geometric opening of the inlet for an undisturbed free stream flow. For free stream Mach number in a range from about 1.6 to 2.0, the external compression inlet 1410 mass flow ratio is approximately 1 so that the air spilled around the inlet is minimal or greatly reduced. In the illustrative inlet design, the external compression inlet 1410 accepts a mass flow rate that positions a terminal shock immediately outside the cowl lip 1420, the external compression inlet 1410 is matched to the engine, and the external compression inlet 1410 has no boundary layer bleed air. No bleed is incorporated into the inlet 1410 because the negative impact on installed specific fuel consumption and the effect on range of inlet bleed is two to three times the impact of pressure recovery loss. An inlet in a configuration that operates well with no bleed assures good supersonic cruise performance.

The fixed geometry external compression inlet 1410 has several characteristics that are beneficial for supersonic, low-boom performance. The fixed geometry axi-symmetric inlet 1410 is simple with no mechanical components for changing inlet geometry, thereby minimizing maintenance and costs. The fixed geometry inlet 1410 generates a normal shock. The fixed geometry inlet 1410 enables the engine to be moved to a more forward location with respect to the aircraft, alleviating problems created by flutter.

The fixed geometry inlet 1410 enables improved performance in comparison to alternative variable geometry approaches. For example, a variable geometry axi-symmetric inlet inherently and greatly increases weight, cost, and reduces reliability. Furthermore, a variable- geometry inlet requires actuators and other moving components that create maintenance and troubleshooting difficulties that are inconsistent with the economics of business jet operations.

Although a two-dimensional (2D) variable geometry induction system avoids some difficulties of an axi-symmetric system, the 2D variable geometry system has additional drawbacks including increased weight, and wave drag resulting from increased volume requirements, than comparable axi-symmetric designs. A 2D inlet also impacts low sonic boom performance.

Referring to FIGUREs 16A and 16B, two schematic pictorial diagrams illustrate the engine inlet 1410 with translating cowl retracted and extended, respectively. The cowl 1412 is retracted for high speed/high flow conditions. The cowl 1412 is extended for takeoff and low speed operations. The cowl 1412 is also extended during high speed bypass conditions to form an

auxiliary slot 1610 that functions as an emergency bypass system that prevents buzz during supersonic operation. Buzz is a low-frequency, high-amplitude pressure oscillation that is linked to shock/boundary layer and/or shock/shock interaction at a relatively low inlet flow ratio. In some conditions, for example subcritical operation of the external compression inlet 1410, the terminal normal shock and the boundary layer formed on the cone wall can cause the boundary layer to separate. The separated boundary layer can cause variations in the inlet mass flow rate that repetitively move the normal shock forward and backward along the ramp, creating the buzz.

The translating cowl 1412 enables the engine to operate efficiently in a wide range of flight conditions, for example for Mach numbers ranging from 0 to 2.0.

Referring to FIGURE 17, a schematic pictorial diagram depicts aerodynamic characteristics of the external compression, fixed compression geometry inlet 1410. The propulsion system 1400 is positioned in a free stream supersonic air stream at Mach=1. 8. The compression region of the inlet spike 1414 generates one or more oblique shock waves 1710 extending from the inlet spike 1414 to the cowl lip 1420. The air stream aft of the oblique shock wave 1710 is slowed, for example to approximately a Mach=1.3 range. The CFD analysis diagram shows a small or negligible amount of spillage at the cowl lip 1420. The external compression inlet 1410 also produces a normal shock wave 1712 aft of the oblique shock wave 1710 and extending from the cowl lip 1420 to a position near the maximum width of the inlet spike 1414. Aft of the normal shock wave 1712, the air stream slows to subsonic levels, for example Mach 0.9, with additional slowing of the air stream, for example to the range around Mach 0.4, at the engine intake.

In some embodiments, the inlet 1410 is configured for isentropic compression at a free stream Mach number in a range up to 2.0, and an inlet throat Mach number in a range from approximately 1.1 to 1. 5. The inlet 1410 has an isentropic region 1716 between the initial oblique shock wave 1710 and the normal shock wave 1712.

The inlet 1410 includes a cowl 1412 that is axisymmetric about a longitudinal axis 1714.

The cowl 1412 has a fixed portion 1428 and a translating portion 1426. The cowl translating portion 1426 is translatable to control the inlet engine air flow match which forms an air shock 1712 at the forwardmost end of the cowl translatable portion 1426. The cowl translation position is adjustable based on aircraft flight speed and/or conditions. The cowl 1412 encloses an axisymmetric interior duct 1416. The inlet 1410 also includes a center body 1414 extending parallel to the longitudinal axis 1714 within the interior duct 1416. The center body 1414 has a fixed geometry in an inlet 1410 with isentropic compression and no bleed.

Referring to FIGURE 18, a schematic pictorial diagram shows a side view of a nacelle 1822. The nacelle 1822 is designed for Mach 1.8 free stream performance. The nacelle 1822 is installed under the aircraft wing 1802. In an illustrative embodiment, the nacelle 1822 comprises an inlet system 1810 that is axi-symmetric with external compression, no bleed or bypass. The inlet system 1810 has a translating cowl 1812 and is acoustically treated to reduce sonic boom.

The inlet system 1810 has an inlet spike 1814 held by three spike struts 1818. The inlet 1810 is mounted under a boundary layer diverter 1820. Aft of the inlet 1810 and contained within an engine nacelle 1822 is a turbofan engine 1800. In an illustrative embodiment, the turbofan engine 1800 is 33, 000-lbf thrust class, non-augmented or dry, with medium bypass, and Full Authority Digital Engine Control (FADEC) control. The engine 1800 has an exhaust system 1824 that is integral with the engine 1800 with a variable throat or exit 1828. The exhaust system 1824 is low-noise and has thrust reversing.

The nacelle system 1822 encloses fire detection, fire extinguishing, and ventilation systems. The nacelle 1822 contains three actuators 1830 that open and translate forward the inlet cowling 1826.

The inlet compression system 1810 is a light-weight, fixed compression geometry design that is configured for usage with engines from several manufacturers. Engine and inlet characteristics are configured to coordinate engine airflow schedules and flight Mach number for the fixed geometry inlet, matching airflow at all pertinent Mach numbers so that no bypass or excessive subcritical spillage occurs under nominal conditions. Airflows at off-nominal conditions are matched using engine trim.

In one embodiment, an inlet/engine configuration is based on a General Electric (GE) supersonic aircraft engine that maintains a status range of 3600 nautical miles (nmi) at Mach 1.8.

The fixed compression geometry engine inlet is optimized for Mach 1.8. A maximum Mach 1.8 capable design represents performance of the Mach 1.8 designed engine cruising at Mach 1.6. A Mach 1.8 capable engine flying at Mach 1.6 increases range and engine life, and potentially improves performance on hot-temperature days. For General Electric, Rolls-Royce, and Pratt & Whitney engines designed with fixed compression geometry external compression inlets optimized for Mach 1.8 cruise, maximum range occurs at Mach 1.6 due to an optimal wing/engine match at Mach 1.6.

In an alternative embodiment, an engine is configured with a fixed compression geometry inlet optimized for Mach 1.6, increasing range to approximately 4250 nmi due to a lift/drag ratio increased by a full percentage point, and a lower engine weight enabling more fuel to bum in cruise.

Referring to FIGURE 19, a schematic pictorial diagram shows an example of an engine 1900 integrated into an aircraft 1902 beneath a wing 1904 and partially supported by a stabilizer 1906. More specifically, a left engine 1900 is coupled to a left outboard wing 1904 connected to an aircraft fuselage 1908 by a left wing rib 1910. In the illustrative embodiment, a left inverted stabilizer 1906 is coupled to the left outboard wing 1904 at the position of the left engine 1900.

The illustration shows a left engine inlet cowling 1912 in an opened position, extended from a left engine inlet bypass lip 1930. One of three inlet cowling actuators 1916 is shown that function to extend and retract the translatable cowling 1912.

The engine 1900 has an inlet 1918 including the inlet cowling 1912 that extends fore and aft along a longitudinal axis 1920 and encases an interior duct 1922. The cowling 1912 includes a translating forward section 1924 and a stationary main body 1928. The stationary main body 1928 has a forward end with a rounded cowl lip 1930. The inlet 1918 also includes a fixed geometry center body 1932 that extends fore and aft generally interior to the cowling 1912 and the interior duct 1922. The forward portion of the center body 1932 has a contoured geometry that produces external compression with an initial oblique shock wave and isentropic compression focused forward and above the cowling forward end, and has no bleed.

In some embodiments, the cowling 1912, the interior duct 1922, and the center body 1932 are axially symmetric about a common longitudinal center line 1920.

In some embodiments, the inlet 1918 is configured for isentropic compression at a free stream Mach number up to 2.0, a throat Mach number nominally set in a range from 1. 1 to 1.5. In some embodiments, the center body 1932 includes a cone 1934 with an initial half angle in a range from 13° to 17° followed by an isentropic compression ramp with an angle increase in a range from 7° to 13°, an internal cowl angle in a range from 0° to 10°, and a throat to capture area ratio in a range from approximately 0.70 to 0.9.

Referring to FIGUREs 20A and 20B, pictorial diagrams respectively show frontal and bottom views of a wing/nacelle/inlet geometry 2000 in an embodiment of an aircraft including engine inlets 2010 with a fixed geometry, a translating cowl, isentropic compression, and no bleed.

FIGUREs 21A, 21B, and 1 show alternative embodiments of suitable engine inlets with a fixed geometry, a translating cowl, isentropic compression, and no bleed.

FIGURE 22 is a perspective pictorial diagram showing an example of an inlet 2210 mounted beneath a wing segment 2212 for wind tunnel testing of a configuration of the inlet 2210, nacelle 2214, and throttle plug 2216.

Referring to a graph shown in FIGURE 23A, a chart represents a typical Mach 1.8 mission output with an engine with fixed compression geometry external compression inlet optimized for Mach 1.8. The aircraft lifts-off from the ground at a weight of approximately 150, 0001bs and burns about 60, 0001bs of fuel during cruise, yielding a supersonic range of about 3600 nmi at Mach 1.8. Maximum aircraft altitude is about 61,000 feet at the end of cruise.

Throttle is held at 100% for the entire cruise portion of the mission to attain the highest possible L/D. Cruise L/D can be higher if the aircraft operates at higher altitude, although the weight increase associated with a thrust increase and increasing SFC with altitude yields lower performance. Cruise thrust decreases steadily as the aircraft climbs due to a relatively flat engine lapse rate at supersonic speeds.

Referring to the graph shown in FIGURE 23B, a graph shows flight profile selection for subsonic and supersonic performance of a selected engine and inlet configuration. A subsonic cruise Mach number of 0.9 is selected based on performance of the engine with the illustrative engine and inlet configuration.

Several design techniques can be used to configure an aircraft for a range capability that is greater than a baseline Mach 1.8 point design approach, yet supply a greater speed capability than a Mach 1.6 point design method. One technique is to design a Mach 1.6 inlet and engine and cruise off-design at Mach 1.8 to improve range over a Mach 1.8 design inlet, for example attaining a 150-250 nmi improvement in range. A second technique involves designing the aircraft as a Mach 1.6 point design for maximum range and accepting any overspeed capability that happens to occur, resulting in a small speed increase for a fully optimized Mach 1.6 engine design that is further limited by engine life reduction as well as degradation of inlet stability and distortion. In a slight variation to the second approach, the engine can be configured as a Mach 1.6 point design with the engine and subsystem design Mach numbers tailored to any speed a Mach 1.6 inlet is capable of attaining in an overspeed condition. The range benefit is even smaller than the effect of a pure Mach 1.6 aircraft but the overspeed capability can be improved although not to the level of a Mach 1.8 design. A third approach incorporates a variable geometry inlet into an otherwise Mach 1.8 configuration so that efficient on-design inlet performance can be obtained from a range from Mach 1.6 to Mach 1.8, resulting in a small range penalty due to higher weight and higher losses inherent to the variable geometry inlet. Mach 1.6 performance of the third approach is further hindered due to increased inlet weight.

In a fourth approach, the inlet design Mach number is set such that a Mach 1.8 cruise can be attained in an overspeed condition with engine, subsystem, and aerodynamic design configured to maximize range at Mach 1.6. The illustrative concept does not operate on-design in a purest

sense, although enabling the largest range of a fixed compression geometry inlet capable of cruising at Mach 1.8. Potentially, flight at a lower than design Mach number using the fixed geometry external compression engine can increase spillage drag and integrate the inlet and propulsion system in a manner that results in a higher drag.

The illustrative aircraft has inlet, engine, and airframe generally designed for Mach 1.8 performance, and further includes optimizations to improve various performance aspects. The configuration enables cruising at a slightly lower Mach number than 1.8 to attain a higher range performance. In an illustrative embodiment, the wings are sized slightly larger than normal for a Mach 1.8 design to improve takeoff and landing performance. For example, the wing area can be set at approximately 2000 square feet to attain desired takeoff and landing performance. The optimum wing area for Mach 1.8 cruise is about 1800ft2 and about 2000 fe for Mach 1.6 cruise.

The slightly larger wing size is nearly optimum for a Mach 1.6 design, improving cruise performance below Mach 1.8. The wing leading edge sweep is set as low as possible for a Mach 1.8 subsonic leading edge design such that the wing design point is on the"knee"of a notional drag-sweep curve. At Mach 1.6, the wing is off the knee and has a slightly lower drag. The wing leading and trailing edge sweeps are reduced to improve stability and control, improve low speed performance, reduce structural weight, and improve performance at slightly lower cruise Mach number.

Other mission-related characteristics associated with cruising at lower Mach numbers include a tendency to cruise at lower altitudes at lower Mach numbers, a consequence of an optimum lift coefficient occurring at lower altitude as a consequence of lower speed.

Furthermore, suitable engines for the desired Mach performance typically produce lower specific fuel consumption at the lower altitudes. Also, lower cruise altitudes yield excess thrust at cruise, enabling a reduction is engine cruise thrust requirement and reduced engine weight. Additionally, lower cruise altitudes allow cruise to begin earlier and end later in a mission so that the aircraft spends proportionately more of a mission in a cruise condition. Also, lower cruise Mach numbers yield lower total air temperatures, benefit engine and subsystem life. Lower cruise Mach numbers can also reduce emissions.

Referring again to FIGUREs 1A, 1B, and 1C, schematic pictorial diagrams respectively showing side, front, and top views of a supersonic aircraft 100 with an engine structural integration system 103 that, in various embodiments, is capable of improving aircraft performance by facilitating positive aerodynamic effects associated with deflecting wing, inverted-V tail rudder surfaces, and canards. The engine structural integration system 103 is positioned in an aft location beneath a wing 104 and has a highly integrated wing/inlet geometry to produce low-

boom compatibility and low inlet/nacelle installation drag. The aircraft has an inverted V-tail geometry 114 that generates low-sonic-boom longitudinal trim in cruise and structural support for the engine structural integration system 103.

In the illustrative embodiment, the aircraft 100 has an elongated nose 111 with a conical tip 113 and an inverted V-tail surface 114 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 audible. The supersonic aircraft 100 creates an N-shaped pressure wave caused by overpressure at the nose 111 and under pressure at the tail 114.

Pressure rises rapidly at the nose 111, declines to an under pressure condition at the tail 114, 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 113 of the nose 111 can 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 114 can generate additional lift near the tail to improve aerodynamics and reduce boom.

The illustrative aircraft arrangement 100 has twin non-afterburning turbofan engines 116 set below and behind the wing 104. The non-afterburning turbofan engines 116 operate behind simple fixed-geometry axisymmetric external compression inlets 119. Considerations of community noise and takeoff, transonic, and cruise thrust specifications determine engine cycle selection and engine sizing.

The shaping of the supersonic aircraft 100 including aspects of the wing 104, the tail assembly or empennage 114, and the engine structural integration system 103 are adapted according to sonic boom signature and supersonic cruise drag considerations. The empennage or tail system 114 includes stabilizers, elevators, and rudders in the inverted V-tail geometry 114.

The inverted V-tail geometry 114 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 114 support the nacelles 122 and non- afterburning turbofan engines 116 in combination with support of the wing 104 to handle flutter.

Inverted V-tail control surfaces, termed ruddervators 124, adjust aircraft longitudinal lift distribution throughout the flight envelope to maintain a low boom, low drag trim condition.

Shapes of the fuselage 101, the wing 104, and empennage 114 are integrated with the entire aircraft configuration so as to be conducive to attaining a low-boom signature and supersonic cruise drag levels. The wing 104 and/or fuselage 101 form an airfoil having aerodynamic characteristics appropriate for low-boom supersonic and transonic flight.

Referring to FIGUREs 24A and 24B, a pictorial frontal view and a three-dimensional view from the rear of the aircraft show structural integration of a propulsion system 2402 into an aircraft. Propulsion system 2402 includes an engine 2404, and an inlet 2406 including a translating cowl 2408. Propulsion system structural integration elements 2402 integrate the propulsion system into an aircraft wing 2412 that is further coupled to an inverted V-tail panel 2414. Propulsion system structural integration elements 2402 include a structural nacelle 2416, an engine support torque box 2418, and a wing structural fairing 2420. The wing 2412 has a leading edge 2422, a trailing edge 2424, upper 2426 and lower 2428 aerodynamic surfaces, and inboard 2430 and outboard 2432 ends. The engine 2404 is mounted onto the wing lower aerodynamic surface 2428 behind the wing trailing edge 2424.

The engine support torque box 2418 is a weight-bearing support structure that couples the structural nacelle 2416 to the wing 2412, and supports the weight of the engine 2404. The torque box 2418 includes a top panel 2434, a bottom panel 2436, and a median honeycomb layer 2438 including a plurality cells 2440 separated by structural cross-members 2442. The cells 2440 typically have different sizes and wall thicknesses although cells with common size and thickness can be used in some embodiments. The median honeycomb layer 2438 is disposed between the top 2434 and bottom 2436 panels. The torque box 2418 has an aerodynamic cross-section that integrates into the wing 2412.

The structural cross-members 2442 in the torque box 2418 are bonded to spars and beams 2444 in the wing 2412 to form shear connections that prevent deformation of the torque box 2418 under loading stresses capable of occurring during various aspects of flight including take-off, landing, cruise, other flight conditions, and emergency conditions. The torque box 2418 transfers bending moments associated with weight and other loadings into the support structure of the wing 2412 and inverted V-tail panel 2414. The spars and beams 2444 are load-bearing and support stresses including weight, torque, air drag, and engine thrust, and isolates vibration produced by the engine 2404 from the wing 2412 and inverted V-tail panel 2414.

The wing structural fairing 2420 covers the torque box 2418 and functions as an auxiliary structure or external surface to improve aerodynamic characteristics and reduce drag. The structural fairing 2420 can be constructed of any suitable material such as aluminum, alloy, or other metal, fiberglass, or the like. The wing structural fairing 2420 attaches to forward and aft

edges of the torque box 2418 and supplies aerodynamic leading and trailing edges to the torque box assembly.

The structural nacelle 2416 encases the engine 2404 and defines the inlet 2406 including an inlet duct 2446 that receives airflow into the engine 2404 and extends along the engine 2404 fore to aft surrounding the engine casing. The illustrative structural nacelle 2416 has a generally cylindrical-type form with approximately annular walls for encasing an axisymmetric engine 2404 and inlet 2406 with some variation from a precise cylindrical geometry. The illustrative structural nacelle 2416 has an annular cross-section separated into an upper portion 2448 and a lower portion 2450 with an approximately or roughly semicircular division. The upper portion 2448 is load-bearing, structurally supporting the engine 2404, and supplies additional stiffness to the engine support structure. The lower portion 2450 is nonload-bearing and generally includes a door for accessing the engine 2404.

The structural nacelle 2416 can have various different geometries in different aircraft embodiments. For example, the structural nacelle 2416 can have different shapes and sizes to accommodate different engine models and engines from different manufacturers. A particular nacelle may be smaller and consequently have lower weight, skin friction, and aeroelastic impact on the wing 2412. A smaller engine can also have additional tip back angle capability at nonzero roll angles due to smaller diameter and a shorter length. Wave drag for a structural nacelle 2416 depends not only on engine size, and thus nacelle volume, but also on inlet capture area relative to the maximum nacelle cross-sectional area. Accordingly, even larger engines can have a suitable wave drag if the inlet is appropriately configured.

Referring to FIGURE 25, a schematic pictorial structural diagram illustrates an example of a supersonic aircraft 2500 that includes the engine structural integration system 103. The supersonic aircraft 2500 includes a left wing section 2502 attached to the left side 2504 of a center body/inboard wing section 2506 and a lower part 2508 of a left inverted stabilizer 2510. A left leading edge flap 2512 and left aileron 2514 are attached to the left wing section 2502 forward spar 2516. Wing skins 2518 have integral stiffeners 2520 machined in a panel 2522 that runs between the wing spars 2516.

A tail structure 2524 includes three sections, a tail structure section 2526, a vertical stabilizer to inverted stabilizer joint section 2528, and inverted stabilizer to nacelle joint section 2530. A vertical stabilizer 2532 is attached to the top of the center body and aft body section 2534. The top of the vertical stabilizer 2532 is attached to the top of left 2510 and right 2536 inverted stabilizers. The lower end of left inverted stabilizer 2510 is attached to the surface of a

left torque box 2538. A left ruddervator 2540 is attached to the aft of the left inverted stabilizer 2510. A rudder 2542 is attached to the end of the vertical stabilizer 2532.

Referring to FIGURE 26, a schematic pictorial diagram depicts the integration of a left nacelle 2600, left wing 2602, and left inverted V-tail stabilizer 2604. The top of the right and left 2604 inverted stabilizers are attached to the top of the vertical stabilizer.

Referring to FIGUREs 27A, 27B, and 27C, front, bottom, side pictorial structural views show an example of a nacelle, wing, and tail configuration. A nacelle structure 2700 includes a right nacelle structure 2702 and left nacelle structure 2704. The right nacelle structure 2702 is attached to the right wing section 2706 and the lower right inverted stabilizer 2708. The left nacelle structure 2704 is attached to the left wing section 2710 and the lower left inverted stabilizer 2712. A left torque box 2716 is attached to the top of the left wing surface 2710 and engine inboard 2718 and outboard 2720 diverters. The left engine outboard diverter 2720 attached to the lower surface of the left wing 2710 and the top of the engine nacelle 2704. The left engine outboard diverter 2720 attaches the frames of the left engine nacelle 2704.

Referring to FIGUREs 28A and 28B, side and front schematic views respectively show an embodiment of an engine attachment/nacelle concept. An engine 2802 is supported and contained within a load-bearing nacelle 2804. The top portion of the nacelle 2804 is load bearing and supplies additional stiffness to the engine support structure. The lower portion of the nacelle 2804 is a non-load-bearing access door 2806. An inlet 2807 is coupled to the forward section of the engine 2802. A nozzle 2808 is mounted directly to the engine 2802.

A V-tail 2810 forms a structural coupling between the wing 2812 and fuselage. The inverted V-tail 2810 forms a highly stable structure that supports the significant axial loads induced on the V-tail 2810 by wind bending. An engine support torque box 2816 distributes stresses throughout an extended area of the wing 2812. The nacelle 2804 is coupled to the wing 2812 using a forward thrust mount 2818, forward torque mounts 2820, and an aft mount 2822.

Various aircraft embodiments can have one or more unique aspects including the inverted V-tail, aft engine placement on the lower wing surface, a slender fuselage and thin wing, distinctive airfoil geometry, and large rudder. In light of these aspects, different configurations may implement structures and operating techniques to address aerodynamic characteristics such as body freedom flutter, V-tail flutter and divergence, and wing flutter arising from thin wing and angle of attack aspects.

Referring to FIGUREs 29A, 29B, 29C, and 29D, multiple views show nacelle integration into an aircraft. Referring to FIGURE 29A, a schematic pictorial diagram shows an example of

an engine 2900 integrated into an aircraft 2902 beneath a wing 2904 and partially supported by a stabilizer 2906. More specifically, a left engine 2900 is coupled to a left outboard wing 2904 connected to an aircraft fuselage 2908 by a left wing rib 2910. In the illustrative embodiment, a left inverted stabilizer 2906 is coupled to the left outboard wing 2904 at the position of the left engine 2900. The illustration shows a left engine inlet cowling 2912 in an opened position, extended from a left engine inlet bypass lip 2914. One of three inlet cowling actuators 2916 is shown that function to extend forward and retract aft the translatable cowling 2912.

The engine 2900 has an inlet 2918 including the inlet cowling 2912 that extends fore and aft along a longitudinal axis 2920 and encases an interior duct 2922. The cowling 2912 includes a translating forward section 2924 and a stationary main body 2928. The stationary main body 2928 has a forward end with a rounded cowl lip 2930. The inlet 2918 also includes a fixed geometry center body 2932 that extends fore and aft generally interior to the cowling 2912 and the interior duct 2922. The forward portion of the center body 2932 has a contoured geometry that produces external compression with an initial oblique shock wave and isentropic compression focused forward and above the cowling forward end, and has no bleed.

In some embodiments, the cowling 2912, the interior duct 2922, and the center body 2932 are axially symmetric about a common longitudinal center line 2920.

In some embodiments, the inlet 2918 is configured for isentropic compression at a free stream Mach number in a range from 1.6 to 2.0, a throat Mach number nominally set in a range from 1. 1 to 1.5. In some embodiments, the center body 2932 includes a cone 2934 with an initial half angle in a range from 13° to 17° followed by an isentropic compression ramp with an angle in a range from 7° to 13°, an internal cowl angle in a range from 6° to 10°, and a throat to capture area ratio in a range from approximately 0.70 to 0.76.

FIGURE 29B shows a left engine nacelle integration and multiple engine mounts.

Engine 2900 is secured to the nacelle 2901 using a left engine mount forward fitting 2940 and left engine mount aft fitting 2942. The engine mount forward fitting 2940 is connected with links (not shown) to inboard and outboard engine mount lugs 2944. A spherical type bearing on the center of the engine mount forward fitting 2940 connects to a pin on the top of the engine 2900.

The engine mount aft fitting 2942 connects by two bolts loaded in tension on top of the engine 2900.

FIGURE 29C is a view aft at the nacelle integration to the wing 2904. The top of torque box is removed for clarity. An accessory access panel is shown on the bottom of nacelle 2901 and the nacelle skin is removed for clarity. The torque box 2950 includes a left inboard torque

box channel 2952, left torque box support 2954, and left outboard torque box channel 2956. The torque box 2950 also includes ducts for carrying fluids for the aircraft environmental control system. A diverter 2958 is positioned between the torque box 2950 and the left outboard wing 2904 FIGURE 29D is a view with an accessory access door removed and lower nacelle doors open for engine removal and/or installation.. The engine 2900 and nacelle 2901 are shown with the inlet cowling 2912 in an open position. One of the three actuators 2916 is shown. The actuator 2916 moves the cowling 2912 through the operation of actuator push rod 2960 in response to movement of a flexible synchronization shaft 2962.

The engine 2900 is secured in the nacelle 2901 by an engine mount forward fitting 2940 and the engine mount aft fitting 2942. The nacelle 2901 has an accessory access panel frame 2964 containing elements including an accessory drive gear box 2966 and pre-cooler 2968..

Referring to FIGUREs 30A and 30B, computational fluid dynamic (CFD) images respectively show aerodynamic characteristics of two nacelle configurations. FIGURE 30A shows a nacelle 3000 with a reflexed plate 3002, diverter 3004, and inlet 3006 with terminating center body 3008 designed for free stream Mach number 1.75 and engine face plate Mach number 0.686 in a configuration that produces a reduced drag. The nacelle 3000 is designed with an upper surface dip 3010 that reduces shock at the diverter 3004.

FIGURE 30B shows a nacelle 3020 with a reflexed plate 3022 that is also designed for a free stream Mach number of 1.75. The surface dip 3010 of nacelle 3000 results in an approximately 3 count drag reduction.

In the past decades, travel by aircraft has become commonplace, population has greatly increased particularly in urban and metropolitan areas, and the number of daily flights has expanded proportionately. Population density enlargement in the vicinity of airports in combination with a high frequency of takeoffs and landings has expectedly resulted in public criticism of nuisance, inconvenience, and damage created by noise inherent to the air travel industry. National and international agencies, aircraft manufacturers, engine manufacturers, and others have responded to the criticism by establishing noise emission standards for aircraft and aircraft traffic limitations for particular communities. The United States Federal Aviation Administration (FAA) has imposed noise limits on takeoff and landing.

Noise rules generally limit aircraft noise that can be emitted during takeoff and during approach to landing. Aircraft noise has many constituent parts including engine fan noise, engine combustion noise, airframe noise, and jet noise caused by shearing of airflow. Engine noise

during takeoff is usually the largest noise component because the engine is then at the highest power setting. Jet noise is a prevalent engine noise component at high engine thrust conditions.

Many techniques for suppressing engine noise have been developed. In one example, engine secondary-to-primary mass flow bypass ratios are increased to values of five to eight to decrease peak jet velocities, shear layer velocity gradients, and turbulence, thereby reducing noise. Some aircraft use"hush kits"such as ejectors or free mixers, and forced mixers to mix high velocity hot engine streams with cooler low velocity free stream air to decrease peak jet velocity and shift from low frequency to more absorbable high frequency noise. In other examples, some conventional engine noise reduction systems use suction devices. Alternatively, suction devices have been used to reduce aerodynamic drag.

Various techniques have been developed to reduce airframe noise, defined as objectionable audible noises during departure and approach conditions from an aircraft and induced by airflow, not related to the engine during operation. Airframe noise can reach or exceed engine noise levels during aircraft landing. Conventional techniques typically address airframe noise by thickening the shear layer adjacent the end of the aircraft body by positioning protuberances adjacent the end of the body. Unfortunately, the devices attached to an aircraft may introduce new noise sources even while reducing some airflow-related noise.

Known techniques successfully reduce noise levels, at least to some degree. However, further reductions are always desirable. Furthermore, the conventional techniques impact performance by one or more of adding weight to the aircraft, reducing engine performance, reducing aerodynamic performance, increasing fuel consumption, reducing range, and/or increasing engine complexity in ways that can compromise engine performance and reliability, increasing the possibility of breakdown and increasing cost.

Noise abatement flight procedures are constantly evolving with advances in technology, improved aircraft design, and implementation of airspace management procedures. Many efforts to address aircraft noise have been targeted to reduction of noise at the source. Aircraft are required to meet government noise certification standards. Compliance with these standards must be considered in the design of new aircraft.

An aircraft includes an Automatic Takeoff Thrust Management System (ATTMS) to reduce or minimize takeoff noise in a limited takeoff field length. In particular embodiments, the automatic takeoff thrust management system can include a programmed lapse rate function that automatically reduces thrust in appropriate conditions after takeoff.

The automatic takeoff thrust management system not only increases power in the event of an engine failure but also modulates thrust as soon as the aircraft establishes takeoff climb conditions. Previous systems have increased thrust only in the event of engine failure, as specified according to Federal Aviation Regulations (FAR) 25.904 and Appendix I25. 5, and required that reduced thrust be set during takeoff roll. The automatic takeoff thrust management system described herein reduces takeoff sound levels while supplying additional thrust for heightened climb performance if an engine failure event occurs during takeoff.

Aircraft typically employ normal takeoff thrust from all engines during take-off roll phase to achieve desired takeoff field lengths. After liftoff, less thrust is needed for the takeoff-climb phase with all engines operating. Usage of full takeoff thrust after liftoff results in noise levels greater than desired for the aircraft. If all engines are operating, thrust is automatically reduced by a selected amount after liftoff once the takeoff-climb phase is securely established.

The automatic takeoff thrust management system performs analysis to determine the point in the flight path that the reduction takes place. The analysis includes logical determination that climb is established based on one or more parameters. Suitable parameters include, but are not limited to, weight-on-wheels sensing, main landing gear position, airspeed, angle-of-attack, rate- of-climb, and others.

The automatic takeoff thrust management system controls thrust reduction to approximate thrust lapse-rate effects that occur naturally due to increasing altitude and airspeed. A selected thrust reduction characteristic can be programmed into control schedules that respond to signals received from sensors and/or control actuators in the aircraft.

An aircraft thrust management system, if all engines are operating thrust is automatically reduced by a selected amount after liftoff once take-off climb is safely established. In a particular embodiment, thrust is automatically reduced by approximately ten percent once climb is established.

An automatic takeoff thrust management system can be used in an aircraft with at least two engines. The management system comprises an aircraft status sensor or set of sensors capable of detecting establishment of takeoff climb conditions, and engine failure detectors respectively coupled to the at least two engines and capable of detecting engine failure. The management system further comprises thrust management modules respectively coupled to the at least two engines and capable of controlling the thrust of the engines, and a controller coupled to the aircraft status sensors, the engine failure detectors, and the thrust control modules. The

controller reduces thrust by a selected amount upon detecting establishment of takeoff climb conditions and, if engine failure is detected, restores thrust to the initial or a higher schedule.

In a particular embodiment, takeoff distance in compliance with Federal Aviation Regulations (FAR) Part 25 is based at least in part on One-Engine-Inoperative (OEI) acceleration that occurs if an engine fails after reaching a takeoff decision speed. The automatic takeoff thrust management system is programmed to sense an engine failure event after reaching takeoff decision speed and responding by increasing available thrust. In the illustrative embodiment, the automatic takeoff thrust management system increases thrust approximately ten percent for OEI acceleration. The automatic takeoff thrust management system is programmed to sense engine failure and responds by increasing thrust on the operating engine to a maximum OEI thrust rating.

The one-engine-operative (OEI) thrust rating can be imposed temporarily, for example for several minutes, in the event of an engine failure during takeoff. The automatic takeoff thrust management system detects the OEI condition based on information and/or absence of information in engine control system sensors and/or dedicated sensors on the failed engine.

If an engine failure occurs during take-off climb after the climb established point, the automatic takeoff thrust management system automatically increases thrust. In some embodiments, the automatic takeoff thrust management system increases takeoff thrust to maximum OEI rating.

The maximum landing weight of the aircraft is often such that unmodified takeoff thrust or go-around thrust is sufficient to meet all Federal Aviation Administration (FAA) climb requirements for landing-climb (go-around) without using automatic thrust management or control. Therefore, in some embodiments the automatic takeoff thrust management system can be enabled to function only during takeoff.

An aircraft comprises a fuselage, wings coupled to opposing sides of the fuselage, at least two engines mounted on the paired wings symmetrically with respect to the fuselage, and an automatic takeoff thrust management system. The automatic takeoff thrust management system comprises an aircraft status sensor or set of sensors capable of detecting establishment of takeoff climb conditions, and engine failure detectors respectively coupled to the at least two engines and capable of detecting engine failure. The automatic takeoff thrust management system further comprises thrust control modules respectively coupled to the at least two engines and capable of controlling the thrust of the engines, and a controller coupled to the aircraft status sensors, the engine failure detectors, and the thrust control modules. The thrust management system reduces thrust by a selected amount upon detecting establishment of takeoff climb conditions and, if engine failure is detected, restoring thrust to the initial or a higher schedule.

A method of automatically controlling takeoff thrust in an aircraft comprises detecting establishment of takeoff climb conditions, detecting engine failure if engine failure occurs, reducing thrust by a selected amount upon detection of established takeoff climb conditions. The method further comprises restoring thrust to the initial or a higher schedule if engine failure is detected.

An article of manufacture comprises a computer usable medium having computer readable program code means embodied therein for detecting establishment of takeoff climb conditions, and a computer readable program code means for detecting engine failure if engine failure occurs. The article of manufacture further comprises a computer readable program code means for reducing thrust by a selected amount upon detecting establishment of takeoff climb conditions, and a computer readable program code means operational when engine failure is detected for restoring thrust to the initial or a higher schedule.

A high performance aircraft has design characteristics that specify excess thrust for all- engine takeoff climb, full thrust for takeoff roll, and full thrust or higher for engine-out takeoff climb. In some embodiments, the high performance aircraft is a supersonic transport aircraft.

Use of full takeoff thrust for all-engine takeoff climb can result in higher than desired noise levels.

Current Federal Aviation Administration (FAA) airworthiness standards define takeoff procedures where no change in thrust that requires action by the pilot may be made until the airplane is 400 feet above the takeoff surface. An automatic takeoff thrust control system satisfies FAA requirements and reduces community noise exposure. The automatic takeoff thrust management system reduces thrust after climb conditions are established, but, in the event of engine failure, increases thrust to meet FAA engine-out climb specifications.

Referring to FIGURE 31, a schematic block diagram depicts an automatic thrust management system 3100 that can be used in an aircraft. In the illustrative embodiment, the automatic thrust management system 3100 comprises a first engine 3120 and a second engine 3130, a computer 3110, sensors 3112, and a thrust level control 3114. The individual engines 3120 and 3130 each include failure detectors 3122 and 3132, respectively, and thrust control modules 3124 and 3134, respectively.

The automatic takeoff thrust management system 3100 comprises one or more sensors 3112 that supply signals to a vehicle management computer or controller 3110. The computer 3110 responds to the sensed signals by sending signals that control engine thrust level. In some embodiments, for normal operation the computer reduces thrust by approximately ten percent when aircraft status sensors indicate the aircraft has established takeoff climb conditions. If

engine failure occurs after the thrust decrease, engine status sensors detect the engine failure and restore thrust to the initial or a higher schedule.

The automatic takeoff thrust management system 3100 reduces takeoff sound levels while supplying safe climb performance in the event of engine failure during takeoff. The automatic takeoff thrust management system increases power upon detection of engine failure, but also reduces thrust under normal conditions once takeoff-climb is established.

The automatic thrust management system 3100 receives control signals from a thrust level control 3114 that may be set manually by a pilot to select a desired total output thrust by the engines 3120 and 3130. The thrust level control 3114 generates manual thrust signals that may be overridden by automatic controls. Sensors 3112 are included for detecting various control parameters such as engine speed or Mach number, engine inlet temperature, engine revolutions per minute, engine inlet pressure, weight on wheels, and others. The computer 3110 is connected to receive signals from the thrust level control 3114 and sensor signals from the sensors 3112.

The computer 3110 comprises processing, storage, and logic elements capable of executing programs, methods, and processes for monitoring and analyzing the thrust level and sensor signals and, based on control requests and signal analysis, generating a command value. The command value controls the thrust control modules 3124 and 3134 of the respective engines 3120 and 3130.

The thrust control modules 3124 and 3134 respond to the command value by continuously regulating fuel flow and other control effectors for the engines 3120 and 3130.

Engine failure detectors 3122 and 3132 typically monitor selected operational parameters and indicate failure in response to selected parameters or combinations of parameters exceeding predetermined limit values. If one or more of the engine failure detectors 3122 and 3132 detect engine failure or a condition indicating the risk of failure, signals indicative of the condition pass to the computer 3110. Typically the signals are also supplied by visual or audio warning to the pilot.

In response to signals from the failure detectors 3122 and 3132, the computer 3110 can automatically respond to complete or partial engine failure using various techniques to control operating engine power levels. In a particular embodiment, the engine failure detectors 3122 and 3132 monitor engine revolutions per minute (RPM) with an RPM loss being indicative of possible engine failure. A desired failure detector works in the expected manner and avoids false indications. Some embodiments may use other or additional sensors for redundancy and self- checking. Multiple or dual sensors may improve reliability by avoiding false positive and false negative indications.

In various embodiments, the computer 3110 can manipulate engine limits, rotor speeds, and turbine temperature to supply sufficient engine power for safe aircraft operation. The computer 3110 can address partial or intermittent engine failure that results in power loss, for example by increasing the power in remaining operating engines.

The engine failure detectors 3122 and 3132 respectively receive signals from engine sensors 3126 and 3136 and can use various techniques to detect partial or complete power loss or engine failure. The failure detectors 3122 and 3132 assess engine performance by monitoring one or more engine performance parameters selected from among engine rotational speeds, engine pressure ratios, and exhaust gas temperatures. In some embodiments, oil supply to critical parts such as bearings and fuel supply can be monitored for indicators of quantity, pressure, and temperature. In some embodiments, vibration can be monitored during engine operation to detect improper balance from failure of rotating parts or other mechanical distress. The illustrative parameters may be monitored to detect early indications of total or partial engine failure.

In various embodiments, particular temperature parameters for monitoring may include inlet, external air, compressor, turbine, bleed air, and exhaust temperatures. In some embodiments, particular pressure parameters that may be monitored include inlet, compressor discharge, lubrication oil, and bleed air pressures. Oil system measurements may be selected from among air quantity, filter status, oil consumption, contamination, and debris. Vibration sensors may detect vibration in afterburners, rotors, shafts, bearings, reduction gears, and others.

Miscellaneous monitored parameters may include life usage such as hours of operation, start times, fatigue, stresses, and cracks. The engine failure detectors 3122 and 3132 may monitor speeds, throttle position, nozzle position, stator position, and fuel flow.

In another example, the engine failure detectors 3122 and 3132 can monitor engine performance including engine pressure ratio, fuel flow, rotational speed, exhaust gas temperature, and throttle position. The engine failure detectors 3122 and 3132 also monitor mechanical performance including oil consumption and vibration amplitude.

The engine failure detectors 3122 and 3132 can send information to the computer 3110 to analyze the monitored parameters, in some embodiments comparing the monitored parameters to stored reference levels, evaluating shifts through time trending. In other examples, the computer 3110 can perform expert analysis based on a library of faults determined by field experience and manufacturer data, neural network based diagnosis using nonlinear modeling techniques, artificial intelligence diagnostic techniques, and the like.

Sensors 3112 indicate status of the aircraft, for example whether the aircraft has established takeoff climb conditions. Status sensors 3112 are used in the illustrative embodiment to automatically control takeoff thrust. The sensors 3112 can also be used to detect status for purposes of stability augmentation for pitch handling during flight, and for proper elevator adjustment.

In various embodiments, different sensors 3112 and sensor combinations can be used. A highly useful status sensor is a weight-on-wheels detector. A typical weight-on-wheels detector uses one or more strain gages mounted on landing gear posts to determine whether a force is above or below a predetermined threshold force indicative of aircraft contact with the ground.

Other typical weight-on-wheels detectors use position sensors or switches to detect compression of the landing gear as it supports the aircraft weight.

Other embodiments may supplement or replace the weight-on-wheels detector with other sensors 3112. A landing gear truck attitude sensor can detect truck angle relative to landing gear posts, monitoring for an upward swinging tilt of the gear indicative of takeoff or a downward shift of the gear truck at touchdown.

Referring to FIGURE 32, a graph shows a takeoff-climb profile definition from Federal Aviation Regulations Part 36 used for acoustic noise determination and depicts a measurement configuration and operations that occur at various positions in three-dimensional space. In the test configuration example, lateral microphones 3210 are positioned 1476 feet from the centerline of the takeoff approach. A flyover microphone 3212 is positioned along the takeoff approach centerline at a distance 21,325 feet from the start position. From the start of takeoff roll (A), throttles are maintained at fixed takeoff position until transition to a manual throttle cutback point (D). Normal takeoff power is applied from the start of takeoff roll (A), through a position (VI) at which takeoff decision speed is attained, through liftoff (B), start of first constant climb (C), and establishment of climb (C*) at an altitude greater than 35 feet from the takeoff surface. The automatic takeoff thrust management system begins a programmed lapse rate (PLR) downtrim when aircraft flight and system parameters indicate that climb is safely established (C*). Manual throttle modulation for cutback begins (D) at an altitude determined by FAA specifications based on the aircraft design configuration. The climb profile transitions from the start of manual power cutback (D) to start of a second constant climb (E). The aircraft passes over the flyover microphone 3212 at the end of the noise certification flight path (F).

Referring to FIGURE 33, a graph illustrates a notational time history of engine thrust for multiple scenarios that may occur during the take-off profile climb. The graph shows net thrust (FN) as a function of time. A natural thrust lapse due to increasing airspeed and altitude causes a

general downward thrust trend with increasing time. During a normal takeoff, a programmed lapse rate (PLR) downtrim reduces available thrust to an operating schedule 3306 generally parallel to the initial 3304 schedule or a maximum one-engine-inoperative (OEI) schedule 3302.

In the illustrative example, at takeoff (A) the maximum OEI takeoff rating is approximately ten percent higher than the normal takeoff rating. The PLR schedule is approximately ten percent less than the initial schedule.

The aircraft typically begins the start of takeoff roll (A) at the normal takeoff rating and progresses on that initial schedule 3304. The aircraft reaches the takeoff decision speed (vol). If the Automatic Takeoff Thrust Management System detects engine failure or low thrust after the aircraft reaches the takeoff decision speed (VI), the system selects the one-engine-operative (OEI) rating schedule 3302, boosting thrust, illustratively by approximately ten percent, and lifting-off (B) using elevated thrust. Otherwise, in normal conditions the engines continue on the initial schedule 3304 through lift-off (B), start of first constant climb (C), and determination that climb is established (C*) at an altitude at least 35 feet above the takeoff surface. When climb is established (C*), the automatic takeoff thrust management system reduces thrust to the programmed lapse rate (PLR) downtrim and the takeoff climb proceeds according to the PLR schedule 3306. The Automatic Takeoff Thrust Management System continues to monitor for opposite engine failure or low thrust and, if detected, responds by removing PLR downtrim and increasing available thrust to the OEI rating schedule 3302. In normal conditions, in absence of engine failure or low thrust, the Automatic Takeoff Thrust Management System maintains the PLR downtrim, for example reducing thrust by approximately ten percent. At or above a selected altitude above the takeoff surface, as determined by FAA regulations, manual throttle reductions may be made during climb for further noise abatement. A pilot may make throttle lever movements to modulate thrust between normal idle levels 3304 and the PLR schedule 3306.

In the event of an engine failure, available thrust from the operating engine is reset to the OEI schedule 3302 and, if the pilot moves the manual throttle, thrust is modulated between normal idle levels and the increased OEI schedule.

Referring to FIGUREs 34 and 35, schematic block diagrams respectively depict examples of automatic takeoff thrust management system embodiments 3400 and 3500. In an illustrative example, an automatic takeoff thrust management system can be implemented with logical elements distributed among the thrust controller 3114 and engine thrust control modules 3124 and 3134 shown in FIGURE 31. The automatic takeoff thrust management system embodiments 3400 and 3500 are shown in dual-engine implementations with each engine comprising an engine controller element 3420 and 3520, respectively. In some systems, each

engine controller stores rating schedule information including maximum continuous, programmed lapse rate (PLR), One-Engine-Inoperative (OEI), flight idle, and ground idle rating schedules.

The illustrative automatic takeoff thrust management system embodiments 3400 and 3500 each also comprise an aircraft Vehicle Management System (VMS) 3410 and 3510, respectively.

Decision logic that selects from among the rating schedules may reside in the engine controllers, the aircraft Vehicle Management System (VMS), or may be distributed among the VMS and engine controllers.

The automatic takeoff thrust management system embodiments 3400 and 3500 illustrate two example architectures for functional partitioning. Depicted signal and data flows do not necessarily represent separate physical signal connections since many signals may also or otherwise be combined onto data buses with appropriate redundancy. In various embodiments, physical interfaces may be digital data buses, analog wires, wireless controls, and other communication elements, or may be combinations of interfaces.

Referring to FIGURE 34, a schematic block diagram illustrates an"aircraft-centric" automatic takeoff thrust management system 3400, a configuration that predominantly positions logical elements for selecting thrust schedules in the aircraft Vehicle Management System (VMS) 3410. The automatic takeoff thrust management system 3400 further comprises right and left engine controllers 3420, each of which further comprises thrust scheduling algorithms 3422.

Functional interfaces from the VMS 3410 to the right and left engine controllers 3420 are discrete on/off signals that manage direct thrust schedule selection. Signals from cockpit throttle levers 3408 are routed to the VMS 3410, which responds by communicating commands to the engine controllers 3420. VMS 3410 can pass the commands through without change or can modify the commands. The discrete signals from VMS 3410 are generated from among an"ON GROUND" logic element 3430, a"CLIMB ESTABLISHED"logic element 3432, and a"ONE ENGINE OPERATIVE (OEI) "logic element 3434.

Referring to FIGURE 35, a schematic block diagram illustrates an"engine-centric" automatic takeoff thrust management system 3500, a configuration that predominantly positions logical elements for selecting thrust schedules in right and left engine controller elements 3520.

The aircraft Vehicle Management System (VMS) 3510 predominantly supplies the engine controllers 3520 with data used to execute schedule selection logic contained within each of the right and left engine controllers 3520. Each of the right engine and left engine controllers 3520 comprise thrust scheduling algorithms 3522 that further comprise schedule selection logic elements, including an"ON GROUND"logic element 3530, a"CLIMB ESTABLISHED"logic <BR> <BR> element 3532, and a"ONE ENGINE OPERATIVE (OEI) "logic element 3534. Data from the

VMS 3510 supplied to the logic elements includes weight-on-wheels (WOW) sensor information, main landing gear (MLG) attitude and configuration information, airspeed, altitude, and other information. Each of the right and left cockpit throttle levers 3508 is connected directly to the corresponding right and left engine controllers 3520 to supply engine throttle control signals.

Signals from the throttle levers 3508 in combination with other engine parameter information pass through the right and left engine controllers 3520 and are supplied to the VMS 3510 via separate data paths from each engine controller 3520.

Referring to FIGURE 36, a schematic block diagram shows an example of a thrust command logic structure 3600 for implementing engine controller thrust scheduling algorithms.

In the illustrative embodiment, the thrust command logic structure 3600 include logical structures that are shared among interrelated logic elements including an"ON GROUND"logic element 3630, a"CLIMB ESTABLISHED"logic element 3632, and a"ONE ENGINE OPERATIVE <BR> <BR> (OEI) "logic element 3634. Engine controller thrust scheduling algorithms 3602 combine the logical elements 3630-3634 and a plurality of logical functions including a Programmed Lapse Rate (PLR) Cutback function 3610, a"One Engine Inoperative" (OEI) rating adder function 3612, a maximum continuous rating schedule function 3614, a flight idle function 3616, and a ground idle schedule function 3618. The illustrative example shows schedule selection logic elements 3630-3634 contained within the engine controller thrust scheduling algorithms 3602. In some other embodiments, the schedule selection logic elements 3630-3634 can be located elsewhere, for example in the Vehicle Management System (VMS).

Air data and signals from the aircraft Vehicle Management System (VMS) activate rating schedules and adjustments for PLR cutback and OEI supplementation via communication with functions 3610-3618. The"CLIMB ESTABLISHED"logic element 3632 activates the Programmed Lapse Rate (PLR) Cutback function 3610 to invoke the programmed lapse rate. The OEI logic element 3634, as directed by opposite engine controller data or a VMS signal 3622, activates the OEI rating adder function 3612 to add thrust when an engine is inoperative or thrust is otherwise low. When opposite engine controller data or VMS signals 3622 activate OEI logic 3634, the programmed lapse rate (PLR) logic path is interrupted so that full OEI thrust becomes available. The maximum continuous rating schedule function 3614 determines an upper thrust bound level subject to correction by the OEI rating adder function 3612 or PLR cutback function 3610. Thrust reduction by the PLR cutback function 3610, thrust increase by the OEI rating adder function 3612, and form of the continuous rating schedule function 3614 can vary in amount and form based on ambient, engine, and aircraft conditions including temperatures, pressures, and other parameters.

In conditions determined by air data signals from VMS 3620, the"ON GROUND"logic 3630 selects from among the flight idle schedules 3616 and ground idle schedules 3618 to determine a lower thrust bound. Form of the flight idle schedules 3616 and the ground idle schedules 3618 vary according to various ambient, engine, and aircraft conditions including temperatures, pressures, and other parameters.

Idle schedules 3616,3618 or high power schedules 3614 can define the lower and upper thrust bounds. In either case, cockpit throttle lever command signals 3624 are mapped between the bounds to generate continuous and essentially linear thrust response between idle and takeoff stops on the cockpit throttle quadrant.

Referring to FIGURE 1A in combination with FIGURE 31, a schematic pictorial view depicts an aircraft that comprises the described automatic takeoff thrust management system. In a particular embodiment, the aircraft is a Quiet Supersonic Transport (QSST) aircraft. The aircraft 3100 comprises an airframe 102 or fuselage, wings 104, engines 116 attached to the wings 104, and tail 3114. Interior to the aircraft 3100 and shown in block diagram form is the computer 3110, aircraft status sensors 3112, engine failure detectors 3122 and 3132, and engine thrust control modules 3124 and 3134.

In an illustrative embodiment, the aircraft 3100 is a high performance aircraft that has design characteristics specifying excess thrust for all-engine takeoff climb, full thrust for takeoff roll, and full thrust for engine-out takeoff climb. In some embodiments, the high performance aircraft 3100 is a supersonic transport aircraft. The high performance aircraft 3100 is designed for high performance cruise capability and accordingly has wings, aerodynamics and engines to attain best cruise performance. Increased maximum thrust level are required to achieve design takeoff field lengths but result in excess all-engine climb capability.

A disadvantage of increasing thrust to decrease the required takeoff field is that all-engine climb is greater than required to meet FAA requirements. Use of full takeoff thrust for all-engine takeoff climb can result in higher than desired noise levels. Current Federal Aviation Administration (FAA) airworthiness standards define takeoff procedures where no change in thrust that requires action by the pilot may be made until the airplane is 3400 feet above the takeoff surface. The Automatic Takeoff Thrust Management System enables automatic thrust management to reduce noise.

The computer 3110, aircraft sensors 3112, engine failure detectors 3122 and 3132, and engine thrust control modules 3124 and 3134 operate in combination as the Automatic Takeoff Thrust Management System 3100. The Automatic Takeoff Thrust Management System 3100

operates to satisfy Federal Aviation Administration (FAA) requirements and reduce community noise exposure. The automatic takeoff thrust management system 3100 reduces thrust after takeoff climb is established, but, in the event of engine failure, increases thrust to meet FAA engine-out climb specifications.

The automatic takeoff thrust management system 3100 comprises one or more aircraft status sensors 3112 that supply signals to the computer 3110, executing as a vehicle management computer or controller. The computer 3110 responds to the sensed signals by sending signals to the engine thrust control modules 3124 and 3134 that control engine thrust level. In some embodiments, for normal operation the computer 3110 reduces thrust by approximately ten percent when aircraft status sensors 3112 indicate the aircraft 3100 has established takeoff climb conditions. If either of engines 3120 and 3130 fails after the thrust decrease, engine status sensors in the engine failure detectors 3122 and 3132 detect the engine failure and restore thrust to the initial or a higher schedule..

The automatic takeoff thrust management system 3100 reduces takeoff sound levels while supplying safe climb performance in the event of engine failure during takeoff.

The automatic takeoff thrust management system 3100 increases power upon detection of engine failure, but also reduces thrust under normal conditions after takeoff climb has been established.

Although the specification describes the"aircraft-centric"and"engine-centric" architectures in detail, other suitable architectures may be used including combined approaches that allocate some control elements to a central management system and other control elements to systems more associated with the engines. Various implementations may take into account the particular sensors and information for making control decisions. Some implementations may centralize control decisions; other implementations may distribute control. Logic elements distributed into engines may implement control functionality tailored to the particular engine characteristics, and design practices.

Also, although thrust is described as reduced by approximately ten percent during takeoff climb, other percentage reductions may be used depending on safety considerations, aircraft performance, engine capabilities, and the like.

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 center-of-gravity control system and techniques for controlling center-of-gravity in combination with control effectors, such as flaps, rudders, elevators, ruddervators, canards, and the like 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 illustrative fuel tank and pump configuration includes a specific number of tanks and pumps in a particular geometric configuration. Other suitable embodiments may have more or fewer tanks, more or fewer pumps, and different types of tanks and pumps. In addition, the tanks and pumps may be arranged in different configurations and placed in different locations. The proportion of tank volume arranged in the wings and fuselage may be different. In various embodiments, the number of computers and centralization or decentralization of computing resources may be varied. For example, control functionality may be dispersed throughout the aircraft or confined to the aircraft central control system.