Aircraft thickness/camber control device for low sonic boom

An aircraft thickness/camber control device mounts to the lower surface of a airfoil configuration, for example on a fuselage, and extends along a longitudinal axis. The device, when deployed, generates expansions ahead of compressions generated by off-design conditions, inlet spillage for example, and enables maintenance of a low boom signature. The device, when positioned at appropriate locations, may also be used as a drag reduction device. The thickness/camber control device comprises a structural member capable of coupling to the airfoil at a position forward of the concentrated source of added compression and a control element. The control element is coupled to the structural member and controls the structural member to adjust thickness/camber of the configuration to cancel the far-field effect of the extra compression or concentrated pressure source.

BACKGROUND OF THE INVENTION

Air travelers have long sought the convenience and efficiency of widespread supersonic commercial aviation only to be denied by technological, economic, and political roadblocks. With operations spanning over a quarter of a century, the Concorde remains the only commercial aircraft that travels at supersonic speeds but struggles with technological obsolescence. Fuel consumption and maintenance requirements of the Concorde strain commercial feasibility in today's competitive environment. Possibly overshadowing 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 burden that restricts the Concorde to routes over oceans.

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.

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

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

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

Shaped sonic booms are only achieved deliberately. No existing aircraft creates a shaped sonic boom that persists for more than a fraction of the distance to the ground while flying at an efficient cruise altitude since non-shaped pressure distributions quickly coalesce into the fundamental N-wave shape. The N-wave form generates the largest possible shock magnitude from a particular disturbance. The N-wave shape results because the front of a supersonic aircraft generates an increase in ambient pressure while the rear generates a decrease in pressure. Variation in propagation speed stretches the disturbance during propagation to the ground. Shaped boom techniques typically attempt to prevent coalescing of the pressure disturbance by adding a large compression at the aircraft nose and an expansion at the tail with pressure in between constrained between the compression and expansion. The shaped boom stretches the ends of the signature faster than the in-between pressures, creating a non-N-wave sonic boom at the ground.

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

SUMMARY OF THE INVENTION

What is desired are external shapes and lift devices that facilitate sonic boom reduction.

According to various embodiments, an aircraft or aircraft control system utilizes a deployable and stowable structural element on the airfoil to counter the spillage at off-design conditions and obtain a lower amplitude sonic boom for a supersonic cruise aircraft.

In accordance with some embodiments of the disclosed aeronautical system, an aircraft fuselage thickness/camber control device create a local expansion that, when propagated to the far field, cancels the added compression generated by inlet spillage shock. The thickness/camber control device comprises a structural member capable of modifying the local camber and thickness of the fuselage at a position forward of the concentrated source of increase in compression, for example the spillage shock, and a control element. The control element is coupled to the structural member and controls the structural member to adjust thickness/camber of the airfoil to cancel the far-field effect of the concentrated source of increased compression.

In accordance with other embodiments, an aircraft comprises a bulge, bump, protrusion, or the like, that extends along a longitudinal axis forward and aft and that has a concentrated source of compression, a structural member capable of coupling to the bulge, bump, or protrusion at a position forward of the concentrated source of compression that results from an off design condition, and a control element. The control element is coupled to the structural member and controls the structural member to adjust thickness/camber of the fuselage to cancel the far-field effect of the concentrated source of pressure.

According to further embodiments, a method of reducing the sonic boom in a supersonic cruise aircraft comprises deploying a structural member on an lower surface of a fuselage that extends along a longitudinal axis forward and aft and that has a concentrated source of expansions at a position forward of the concentrated source of pressure. The method further comprises controlling the structural member to adjust thickness/camber of the airfoil to cancel the far-field effect of the concentrated source of compression.

DETAILED DESCRIPTION OF THE EMBODIMENTS

Referring toFIG. 1, a schematic pictorial and block diagram illustrates an example of an aircraft thickness/camber control device100extending along a longitudinal axis104forward and aft and that has a concentrated source of compression106for example that results from inlet shock or spillage. The thickness/camber control device100comprises a structural member108that mounts on the bottom of the outer mold line at a position forward of the concentrated source of compression106and a control element110. The control element110is connected to the structural member108and controls the structural member108to adjust thickness/camber of the airfoil102and generates expansions that cancel the far-field effect created by the concentrated source of compression106.

Typically, the largest concentrated source of compression106in an aircraft is a nacelle. An airfoil is generally designed for most aerodynamically-efficient performance at a particular Mach number or range of Mach numbers. In various circumstances and conditions, operation at off-design Mach numbers is desirable. For example, operating at off-design Mach numbers at selected times can increase aircraft range. During operations in off-design conditions, the nacelle106creates a spillage shock that is larger than a desired baseline cruise configuration for a low sonic boom signature. The control element110adjusts airfoil thickness/camber so that the far-field effects of the controlled structural member108and the nacelle106substantially cancel.

The thickness/camber control device100has particular utility in a supersonic cruise aircraft with an area distribution that matches a low sonic boom signature. In off-design conditions the concentrated source of compression106generates a spillage shock larger than the baseline cruise configuration for the low sonic boom signature. To reduce the sonic boom, the control element100adjusts the structural member108so that the far-field effect of the adjusted structural member108cancels the far-field effect resulting from the spillage shock. The control element110adjusts the structural member108in off-design conditions so that the expansion generated around the structural member108effectively reduces the equivalent area distribution ahead of the spillage shock.

The thickness/camber control device100utilizes the control element110to adjust the structural member108to effectively modify the camber of the airfoil102to minimize or reduce spillage shock resulting from the concentrated source. Controlled adjustments of the fuselage thickness and camber can operate either to block the nacelle spillage effects, thereby softening the sonic boom signature, or as a moveable active feedback device to adjust aircraft aerodynamics.

The thickness/camber control device100adjusts body area of the airfoil102and fuselage thickness/camber ratio to reduce sonic boom effects. The thickness/camber control device100does introduce some drag when deployed, typically a drag increase of approximately 5% for a particular embodiment. Accordingly, the thickness/camber control device302is stowed when unused so that drag is reduced when operating in compliance with design conditions. For example, the thickness/camber control device302can be deployed when flying over land to reduce or minimize sonic boom effects. When flying over oceans, the thickness/camber control device302can be stowed

The control element110includes any operating controls that can be manipulated by a pilot, either with or without supporting electronics, processors, computers, controllers, and the like. The control element110further includes any linkages or actuators connected to the structural element108to physically move the structural element108.

The airfoil102generally includes aircraft wings and also includes other aerodynamic shapes including a fuselage, tail, and other structures within the air stream.

FIGS. 2A through 2Fshow several examples of thickness/camber control devices that can be used to improve aerodynamics and reduce sonic boom amplitude in supersonic aircraft. Referring toFIG. 2A, some embodiments of the thickness/camber control device utilize a structural member in the form of a body flap200, such as control surface202hinged204to a lower section of the fuselage206forward of the wings208and nacelles.

FIG. 2Bshows an example of a thickness/camber control device embodiment that utilizes canards210as a structural member to effectively increase the thickness of the fuselage206forward of the nacelles. The canards210are stowable and deployable to cancel the far-field effect resulting from the spillage shock and tailor the sonic boom signature of the aircraft.

FIG. 2Cdepicts an embodiment of a thickness/camber control device that deploys and stows a pair of fairings220as structural members for controlling aerodynamics of an aircraft. The fairings220are stowed in the aircraft fuselage206within a compartment222covered by doors224. The fairings220are deployed by structural linkages that open the doors and extend the fairings220from the fuselage206.

FIG. 2Dshows an embodiment of a thickness/camber control device that deploys and stows a speed brake230to control aircraft aerodynamics and manage sonic boom signature.

FIGS. 2E and 2Frespectively show a cross-sectional view of an aircraft fuselage206with a thickness/camber control device200, and a side view of the aircraft. The thickness/camber control device240has the form of a protrusion240that can be rotated out of the fuselage206on deployment and into the fuselage206for stowage.

Other structures can be used such as protrusions, extensions, or wideners that are deployable and stowable to increase cross-sectional area of the fuselage forward of a concentrated source of compression, such as the nacelle or nacelles. The structures can be an active drag device, fairing, protrusion, or bank of actuators that extends from the fuselage, thereby enlarging the cross-sectional area of the airfoil, when deployed, and that forms a flush surface with the fuselage when stowed.

Referring toFIG. 3, a pictorial schematic diagram shows an example of a supersonic aircraft300that includes a thickness/camber control device302capable of improving flight aerodynamics and reducing sonic boom effects. The thickness/camber control device302comprises a structural member310coupled to the airfoil304at a position forward of the concentrated source of pressure308, and a control element312. The control element312is coupled to the structural member310and controls the structural member310to adjust thickness/camber of the airfoil304to cancel the far-field effect of the concentrated source of compression308.

The aircraft300further comprises a fuselage314, an aircraft wing316coupled to the fuselage, and an engine nacelle318. Commonly, the nacelle318is the largest concentrated source of compression308in an aircraft. In the illustrative aircraft300, the engine nacelle318is coupled to the aircraft wing316. The structural member310is coupled to the fuselage314forward of the engine nacelle318and is controllably deployed and stowed to adjust thickness/camber of the fuselage314to cancel the far-field effect of the engine nacelle318.

Referring toFIG. 4, a graph and schematic pictorial view of an aircraft show an example of a technique for minimizing or reducing sonic boom effects using the thickness/camber control device302. Jones-George-Seebass-Darden sonic boom minimization theory states a ground signature will have minimum shock strength (ramp signature) by following a calculated equivalent area distribution400, defined by a program SEEB, which becomes a design goal. To attain the goal signature defined by the SEEB curve400for predetermined flight conditions of aircraft weight, altitude, and Mach number, a control procedure either deducts or adds to the configuration equivalent areas. If Mach angle cross-sectional areas402are configured to approximate the SEEB curve400, a control procedure is termed “area boom-ruling,” as distinguished from “lift boom-ruling” if the lift distribution on the aircraft were modified to match the SEEB curve400.

In off-design conditions, the concentrated source of pressure created by the nacelle308generates the spillage shock406that exceeds the baseline cruise level configuration for a low sonic boom signature400. The thickness/camber control device302adjusts airfoil aerodynamics in the off-design conditions to generate an expansion404around the structural member310to effectively reduce the equivalent area distribution ahead of the shock406as shown in the controlled equivalent area plot402that is beneath the SEEB curve400. The control element312adjusts the structural member310to improve aerodynamic flow fields for flight at Mach numbers different from the Mach number to which the aircraft design is optimized.

The equivalent area curve402of the aircraft300shows the increase in equivalent area resulting from the spillage shock406. The thickness/camber control device302corrects for the increase in equivalent area by creating the expansion404that pulls the aircraft equivalent area curve402below the SEEB curve400at all positions relative to the aircraft300. To attain the reduced sonic boom goal, the aircraft equivalent area curve402can fall below but not above the SEEB curve400.

The thickness/camber control device302can remain stowed while the aircraft300is operating with optimal aerodynamics at a predetermined design condition. However, the aircraft300can be alternatively operated at other conditions, for example to increase range or other purposes, with aerodynamics controlled by deploying the thickness/camber control device302to reduce sonic boom effects.

Referring toFIG. 5, a flow chart depicts an embodiment of a method500for reducing the sonic boom in a supersonic cruise aircraft. The method500comprises deploying a structural member502on an airfoil that extends along a longitudinal axis forward and aft and that has a concentrated source of compression at a position forward of the concentrated source of compression. The method500further comprises controlling the structural member504to adjust thickness/camber of the airfoil506thereby canceling the far-field effect508of the concentrated source of compression.

One aspect of the method is that the aircraft is designed so that the structural member is positioned510at a location on the airfoil that facilitates cancellation of the far-field effect of the concentrated source of pressure.

The aircraft can be flown at off-design conditions512causing the concentrated compression to create a spillage shock larger than a baseline cruise configuration for a low sonic boom signature514. In some embodiments, the method includes adjusting the structural member to vary airfoil thickness/camber so that the far-field effects of the controlled structural member and the concentrated source of compression substantially cancel516.

Referring toFIGS. 6A,6B, and6C, schematic pictorial diagrams respectively showing side, front, and top views of a supersonic aircraft600with a thickness/camber control device601that, in various embodiments, is capable of improving aircraft performance by facilitating positive aerodynamic effects including adjustment of flow fields to improve aerodynamics at a range of air speeds and maintaining a low sonic boom signature. The aircraft600comprises a fuselage604and an aircraft wing608mounted on the fuselage604. The wing608has a leading edge606. The wing608extends from an inboard edge at the fuselage604to an outboard edge at the wing tip. The aircraft lift device601further comprises a strake602capable of coupling to the fuselage604and extending to the leading edge606of the wing608. The aircraft600further comprises a Krueger flap614coupled to the leading edge606of an inboard portion of the wing608adjacent the strake602, and a leading edge flap612coupled to the leading edge606of the wing608and extending from a junction at the Krueger flap614to an outboard portion of the wing608.

The aircraft600further comprises a control element, such as the control element310shown inFIG. 3and at least one structural member such as the structural members shown inFIGS. 2A–2F. The control element is used to adjust the position of deployment of the thickness/camber control device601to improve aerodynamic flow fields for flight at Mach numbers different from the Mach number to which the aircraft design is optimized.

In illustrative embodiments, the aircraft600has engines616positioned in aft locations beneath the wings608and have a highly integrated wing/inlet geometry626to produce low-boom compatibility and low inlet/nacelle installation drag. The aircraft600can have an inverted V-tail geometry632that enhances low-sonic-boom longitudinal trim in cruise and allows better structural support for the engines616.

In the illustrative embodiment, the aircraft600has an blunted nose628with a conical tip630and an inverted V-tail surface632that overlaps the wing608, 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 aircraft600creates an N-shaped pressure wave caused by overpressure at the nose628and under pressure at the tail634. Pressure rises rapidly at the nose628, declines to an under pressure condition at the tail634, 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 tip630of the nose628can 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 aircraft600has a sharply swept arrow wing configuration608that reduces peak overpressure in the wave by spreading wing lift along the aircraft length. The wing configuration608has reduced wing leading and trailing edge sweeps. The inverted V-tail632can generate additional lift near the tail to improve aerodynamics and reduce boom.

The illustrative aircraft arrangement600has twin non-afterburning turbofan engines616set below and behind the wing608. The non-afterburning turbofan engines616operate behind simple fixed-geometry axisymmetric external compression inlets618. Considerations of community noise and takeoff, transonic, and cruise thrust specifications determine engine cycle selection and engine sizing.

The shaping of the supersonic aircraft600including aspects of the wing608, the tail assembly or empennage620, and the engine616structural integration are adapted according to sonic boom signature and supersonic cruise drag considerations. The empennage or tail system620includes stabilizers, elevators, and rudders in the inverted V-tail geometry632. The inverted V-tail geometry632supports nacelles622in highly suitable positions relative to the wing608to suppress boom, and trims the supersonic aircraft600in cruise to attain an improved low-boom lift distribution. Panels of the inverted V-tail632support the nacelles622and non-afterburning turbofan engines616in combination with support of the wing608to handle flutter. Inverted V-tail control surfaces, termed ruddervators624, adjust aircraft longitudinal lift distribution throughout the flight envelope to maintain a low boom, low drag trim condition.

The shape of the fuselage604, the wing608, and empennage620are integrated with the entire aircraft configuration so as to be conducive to attaining a low-boom signature and supersonic cruise drag levels. The wing608and/or fuselage604are integrated to achieve low-boom supersonic flight.

The wings608can have a substantial dihedral, or “gulling” incorporated into the wings608inboard of the engines616. The dihedral geometry is most pronounced at the wing trailing edge. The gull or dihedral results from twisting and cambering the wing608for 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 of the wing608can be configured to integrate with the nacelle622and a diverter formed between the nacelle622and the wing608to follow the contour of a low-sonic-boom fuselage604with as close a normal intersection as possible to attain low interference drag. In some embodiments, an inboard flap hinge line is fully contained within the wing contour with the wing upper and lower surfaces held as planar as possible to facilitate seal design.

With the resulting wing configuration, the wing gull raises the engines616to increase available tip back angle and reduce thrust-induced pitching moments. The gull enhances low-boom signature by vertically staggering the wing longitudinal lift distribution and lowers the aircraft body or fuselage604to reduce the height of the cabin door638above the ground, thereby reducing entry stair length. The low fuselage604assists in maintaining a low aircraft center of gravity, reducing tip over angle and promoting ground stability. The wing gull forms a wrapping of the wing608around the nacelle622that enhances favorable interference between the nacelles618and the wing608, resulting in a wing/body/nacelle geometry conducive to successful ditching and gear-up landings.

The leading edge surfaces of the wing608, including the leading-edge flap of the strake602, the Krueger flap614, and the leading edge flap612are controlled by one or more control elements to adjust aerodynamic flow fields, thereby improving aerodynamic performance in operation at various airspeeds. In addition, the leading edge surfaces can be controlled to adjust the leading-edge surface to maintain a low sonic boom signature. In some conditions, the control elements can deflect the strake602to reduce lift ahead of spillage at an off-design condition and maintain a low sonic boom signature.

Referring toFIGS. 7A,7B,7C, and7D, a series of graphs illustrate theory upon which a low sonic boom signature is attained by controlling deployment of the thickness/camber control device601, reducing sonic boom loudness while maintaining long supersonic range. The leading edge control elements reduce sonic boom loudness by shaping the sonic boom for low shock strengths.FIG. 7Ais a graph showing the pressure distribution from a conventional supersonic aircraft. The pressure distribution coalesces into an N-wave at the ground, a shape corresponding to the largest shock strength and thus the greatest loudness. One technique for reducing sonic boom amplitude at the ground involves a minimization theory in which a pressure distribution caused by a low boom aircraft follows an inversely calculated distribution to generate low shock strength at the ground. Contrary to intuition, a low boom distribution occurs when a strong leading edge compression quickly reduces in magnitude, followed by a gradually increasing weak compression that rapidly inverts into a weak expansion, followed by a stronger trailing edge expansion that gradually recompresses to ambient. Boom minimization occurs when an aircraft produces an inversely calculated pressure distribution. The pressure distribution produced by an aircraft results from a Mach angle, cross-sectional area distribution, for example as shown inFIG. 7B, and a Mach angle lift distribution, as shown inFIG. 7C. The thickness/camber control device operates to generate a local expansion on the airfoil to counteract spillage shock from the nacelles, thereby shaping the active area distribution to reduce sonic boom amplitude at the ground. A minimized pressure distribution, shown inFIG. 7D, occurs when the sum of the area pressure distribution and the lift pressure disturbance equal the minimized pressure distribution. The leading edge devices described herein can be used to shape the pressure distribution.

Referring toFIG. 8, a graph further illustrates theory of equivalent area minimization to reduce sonic boom signature, showing effective area against axial location along the longitudinal axis of the aircraft. When equivalent area due to geometric area and lift sum to the minimized distribution, a minimized ground sonic boom occurs. The thickness/camber control device is controlled to modify the airflow, counteracting the spillage shock generated by the nacelles, and possibly stretching the lifting length to move the active area distribution closer to the distribution that shapes the sonic boom signature.

Referring again toFIGS. 6A through 6C, the illustrative aircraft600utilizes control of the thickness/camber control device601, in accordance with an equivalent area technique to reduce sonic boom signature. Equivalent area is the Mach angle area distribution of an axisymmetric body that generates the same disturbance as a given geometric area or given lift distribution. The equivalent area due to geometric area can be approximated as equal to the Mach angle area distribution. The equivalent area due to lift is equal to the integral of the Mach lift per unit of stream wise length times atmospheric constants.

In the illustrative embodiment, the leading edge control surfaces are controlled to reduce or minimize sonic boom by deflecting the air flow to reduce lift ahead of the spillage due to nacelles622. For example, if the aircraft600is flying in an off-design condition in which the nacelles622are spilling air and are thus generating stronger shocks and stronger compressions, the leading edge control surfaces and be actuated to compensate by creating an expansion of air flow that blocks the spillage from coalescing into an N-wave.

The wings and engine are generally designed for selected for usage at various air speeds. Engine616and inlet626characteristics are configured to coordinate engine airflow schedules and flight Mach number. In a particular embodiment, a fixed geometry inlet626can be utilized, for example to reduce propulsion system weight and complexity, and thereby improve efficiency and performance. In particular fixed-geometry inlet configurations, airflow is matched 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 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. The Mach 1.8-capable engine flying at Mach 1.6 increases range and engine life, and potentially improves performance on hot-temperature days.

In an alternative embodiment, an engine616is configured with a fixed compression geometry inlet optimized for Mach 1.6, increasing range to approximately 4250 nmi by increasing lift/drag ratio by a full point, and a lower engine weight enabling more fuel to burn in cruise.

Various 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. Translating cowls can also be used to enhance subsonic performance.

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.

An illustrative aircraft600can have inlet626, engine616, 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.

The control elements operating the thickness/camber control device601can be controlled to further facilitate operation of the aircraft600at off-design Mach numbers.

Other mission-related characteristics facilitated by control of the leading edge surfaces include a capability to cruise at lower Mach numbers, and a tendency to cruise at lower altitudes and lower Mach numbers, resulting from 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.

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, the structural member or body flap can be used as a sonic boom reduction device in conjunction with or separately from other boom devices such as a leading edge flap. Sonic boom reduction can be affected undertrack and off-track.