SYSTEM AND METHOD FOR DIVERTING BOUNDARY LAYER AIR

Abstract: A diverterless engine inlet (10) system that integrates a "bump" (20) surface with a forward swept, aft-closing cowl (30) to divert substantially all of the boundary layer air from the inlet. The bump includes an isentropic compression surface raised outwardly from the body of the aircraft to form a portion of the inner surface of the inlet. The cowl couples to the body of the aircraft to form the remaining surfaces of the inlet and closes against the body of the aircraft at the aft-most points of the inlet opening. The bump and the cowl work together to divert low energy boundary layer air from the inlet during aircraft operation, thus eliminating the requirement for a boundary layer diverter. The diverterless inlet provides a lower cost, reduced complexity air induction system that can be used on a supersonic aircraft engine.

Description: SYSTEM AND METHOD FOR DIVERTING BOUNDARY LAYER AIR TECHNICAL FIELD OF THE INVENTION The present invention relates in general to engine inlet systems, and more particularly, to a system and method for diverting boundary layer air from the inlet of an supersonic aircraft engine . BACKGROUND OF THE INVENTION Many of today's fighter aircraft can cruise supersonic and are low observable aircraft with low radar cross-sections. The engine inlets on the aircraft must meet similar requirements. The inlet needs to work well supersonically and have a low radar cross-section. During high speed flight, a very low velocity, low pressure boundary layer of air builds up on the fuselage of a supersonic aircraft . Because this low energy air will cause poor engine performance, high speed aircraft have traditionally employ some type of boundary layer diverter system to prevent the boundary layer air from entering the inlet. Current advanced supersonic tactical aircraft utilize air induction systems that include boundary layer diverters, boundary layer bleed systems, and overboard bypass systems to divert this boundary layer air in order to provide higher engine inlet performance. A boundary layer diverter is essentially a gap between the side of the aircraft body and the inlet that diverts the low-pressure boundary layer air that builds up on the fuselage and prevents this boundary layer air from entering the engine. In addition to the diverter, traditional inlet systems also utilize a boundary layer bleed system on the edge of the inlet. The bleed system works in a manner similar to the diverter, except instead of skimming off the boundary layer air, the bleed system takes that air on-board, then vents it up through the airplane and exhausts it through a bleed exit on the aircraft. In addition to a diverter and a bleed system, some traditional inlet systems also use an overboard bypass system. The bypass system exists to correct a high speed problem with the aerodynamics of traditional inlet systems. The inlet has to flow a certain amount of air to work properly. At high speeds, the engine airflow demand cuts back to a level below that required for the inlet to work properly. The bypass system compensates by dumping the excess air overboard. Present air induction systems require these subsystems in order to make this low observable inlet design work properly at high speeds. These systems are highly complex and involve a variety of composite materials. These air induction systems increase the weight, the cost of production, mechanical complexity and the cost of maintenance of the aircraft. The "bump" concept of having a raised compression surface has been discussed in industry references, most recently in a book entitled "Intake Aerodynamics," edited by J. Seddon and E.L. Goldsmith, and published in 1985 by the American Institute of Aeronautics and Astronautics. Performance of an inlet with a compression surface was measured and documented in the 1956 report "Performance of External-Compression Bump Inlet at Mach Numbers of 1.5 to 2.0," by Simon, Brown, and Huff (NACA RM E56119) . However, the concept reported therein, unlike the present invention, utilized an unswept cowl and bump surface boundary layer bleed. Bump inlet technology was also examined in the Air Force Wright Laboratory program "Management of Advanced Inlet Boundary Layers" (Contract F33615-89-C-3000) . This program examined a bump inlet concept with a bleed system and a serrated cowl designed to optimize, rather than eliminate, boundary layer diverters and bleed systems. SUMMARY OF THE INVENTION In accordance with the present invention, a system and method for diverting boundary layer air from an aircraft engine is provided that substantially eliminates or reduces disadvantages and problems associated with previously developed engine inlet systems. The present invention provides a diverterless inlet system that uniquely integrates a "bump" including an isentropic compression surface with a forward swept, aft- closing cowl. The bump is a surface raised outwardly from the body of the aircraft. The cowl couples to the body of the aircraft and forms the outer surfaces of the inlet . The cowl closes against the body of the aircraft at the af -most points of the inlet opening. The bump and the cowl work together to divert boundary layer air and prevent substantially all of the lower energy boundary layer air from flowing through the inlet during aircraft operation. A technical advantage of the present invention is that the forward-swept, aft-closing cowl and the isentropic compression surface reduce the complexity of the aircraft's inlet system. In particular, the present invention does not require a boundary layer diverter, a side or splitter plate, a boundary layer bleed system, or an overboard bypass system. Furthermore, the present invention has no moving parts. This reduction in complexity reduces the tactical fighter aircraft's empty weight, production cost, and maintenance support requirements. These savings are estimated to be 250 pounds per aircraft, $225,000 per aircraft, and 0.03 maintenance man hours per flight hour, respectively. The present invention provides another technical advantage in that it is adaptable to virtually any supersonic aircraft. Because all inlet designs spill air from the inlet area, another technical advantage of the present invention is that it takes advantage of the fact that some air must be spilled outboard by spilling the lower quality boundary layer air, rather than the higher quality free stream air. BRIEF DESCRIPTION OF THE DRAWINGS For a more complete understanding of the present invention and the advantages thereof, reference is now made to the following description taken in conjunction with the accompanying drawings in which like reference numerals indicate like features and wherein: FIGURE IA shows one embodiment of the present invention side-mounted to the body of an aircraft; FIGURE IB illustrates a front view of a side-mounted embodiment of the present invention; FIGURE IC illustrates a top view of a side-mounted embodiment of the present invention; FIGURE ID illustrates a side view of a side-mounted embodiment of the present invention; FIGURE 2A shows another embodiment of the present invention mounted to the underside of a wing of an aircraft; FIGURE 2B illustrates a front view of a wing-mounted embodiment of the present invention; FIGURE 2C illustrates a side view of a wing-mounted embodiment of the present invention; FIGURE 2D illustrates a bottom view of a wing- mounted embodiment of the present invention; FIGURE 3 illustrates the one solution describing the boundary layer diversion associated with one embodiment of the present invention at Mach 1.6; FIGURE 4A illustrates an isometric view of the surface pressure coefficient of the same embodiment as shown in FIGURE 4A of the present invention at Mach 1.6; FIGURE 4B illustrates a bottom view of the boundary layer diversion of the particle traces of another embodiment of the present invention at Mach 1.6; FIGURE 5 illustrates one derivation of the compression surface from a conical flow field; FIGURE 6 illustrates an isentropic cone for deriving a compression surface; FIGURE 7 illustrates one derivation of the isentropic compression surface at Mach 1.6; FIGURE 8 illustrates an isentropic cone presented at an angle of attack to the air flow for deriving a compression surface; FIGURE 9A shows another embodiment of the present invention utilizing an isentropic cone; FIGURE 9B shows another embodiment of the present invention utilizing an isentropic cone at an angle of attack; FIGURE 10A is a graphical representation of the surface slope of the compression surface and the transition shoulder; FIGURE 10B is a graphical representation of another embodiment of the compression surface and transition shoulder; FIGURE 11 illustrates the shape of one embodiment of the transition shoulder of the present invention; FIGURE 12 shows one embodiment of the compression surface and transition shoulder of the present invention; and FIGURE 13 illustrates graphically the diversion of boundary layer air outboard by two embodiments of the present invention. DETAILED DESCRIPTION OF THE INVENTION The present invention provides a supersonic engine inlet concept for use in, for example, tactical aircraft applications, that utilizes a "bump" comprising a fixed three-dimensional, isentropic compression surface, a transition shoulder, and a diffuser fairing, combined with an aft-closing, forward-swept cowl which closes against the aircraft at the aft-most points of the inlet opening. The present invention eliminates the need for boundary layer diverters, overboard bypass systems, and boundary layer bleed systems currently used on conventional air induction systems for supersonic aircraft. Thus, the present invention reduces aircraft weight, cost, and complexity. These features are eliminated because the compression surface and cowl work synergistically to provide boundary layer diversion capability. The compression surface portion of the bump, can be designed to produce a conical flow field that can be equivalent to that of an axisymmetric body with a 12° semivertex angle isentropically blended to a 21° final turning angle. The span-wise static pressure on the surface can begin to divert boundary layer air outboard. The pressure differential between the inlet and the area surrounding the inlet further diverts the boundary layer air outboard. The compression surface can also serve to reduce terminal shock Mach number, thus reducing the tendency for shock-induced flow separation. Because the cowl closes against the forebody at the aft-most points of the inlet opening, low pressure boundary layer air, rather than free stream air, can be diverted out the side of the inlet. The inlet concept of the present invention can be utilized in any number of forebody/aperture integration schemes . FIGURE IA illustrates a side-mounted embodiment of the inlet 10 mounted to the body of an aircraft 90 with an opening 11 for receiving air into the inlet 10. The inlet 10 includes a bump 20 and a forward-swept, aft- closing cowl 30. The bump 20 is a raised surface formed outwardly from the aircraft towards the interior of the inlet 10. The bump 20 forms part of the inner surface of the inlet 10. The surface comprising the bump 20 begins to raise outwardly away from the body of the aircraft prior to the opening 11, so that boundary layer air will contact the bump 20 prior to arriving at the opening 11. The shape of the bump 20 can be varied depending on design parameters discussed more fully below. As shown in FIGURE IB, this side-mounted embodiment of the bump 20 extends outwardly from the body of the aircraft 90 with highest point of the bump 20 approximately at the center of the inlet and the bump gradually lessens in height as it approaches the cowl 30. FIGURE IC shows a side-mounted embodiment of the surface forming the bump 20 that gradually increase in height away from the body of the aircraft 90 near the opening 11, reaches a peak at a point inside the inlet 10, then gradually decreases in height near the rear of the inlet 10. The cowl 30 includes an aft-closing portion 38 and a forward-swept portion 39 shown generally in FIGURE IA. The embodiment of the af -closing portion 38 shown in FIGURES IB, IC, and ID includes a pair of aft-closing panel sections 31 and 33 that close against the body of the aircraft at the aft-most points 42 of the cowl. The embodiment of the forward-swept portion 39 shown in FIGURES IB, IC, and ID includes forward-swept panel section 32 that further includes tip 36. The forward- swept panel section 32 can include several sections joined together, or alternatively, can be a single piece. The embodiment of the cowl shown in FIGURE IA shows aft- closing panels 31 and 33 coupled to the body of the aircraft and extending outwardly from the aircraft. Aft- closing panels 31 and 33 include leading edge 37 that extends from the closure point 42 toward the front of the aircraft. Forward-swept panel section 32 joins aft- closing panel section 31 to aft-closing panel section 33 to form the cowl 30 that provides the outer frame of the inlet 10. Coupling the cowl 30 to the body of the aircraft forms the opening 11 of the inlet 10. The forward-swept panel section 32 of cowl 30 includes a triangular shaped tip 36 coming to a point at the forward-most point of the forward-swept panel section 32 to form the forward-swept portion 39 of the cowl. This side-mounted embodiment of the cowl 30 is also shown in FIGURES IC and ID. FIGURES IC and ID show the forward- swept portion 39 of the cowl 30 extending towards the front of the aircraft. FIGURE ID illustrates that the tip 36 of the forward-swept panel section 32 can be triangular in shape with an apex of the triangle approximately at the centerline of the inlet 10. As shown in FIGURE IC, the aft-closing panel sections 31 and 33 can close against the body of the aircraft at closure points 42. As illustrated in FIGURES IC and ID, cowl closure points 42 can be located at the aft-most points of the opening 11. As illustrated in FIGURE IB, forward- swept panel section 32 can join aft-closing panel sections 31 and 33 to create a cowl 30 that forms an inlet opening 11 with an approximately trapezoid shape as the cowl 30 closes against the body of the aircraft. FIGURE 2A illustrates a wing-mounted embodiment of the inlet 10 with an opening 11 for receiving air into the inlet 10. In this alternative embodiment, the raised surface of the bump 20 is formed outwardly from the bottom of the wing 98 of the aircraft 90. As before, the inlet 10 includes a bump 20 and a cowl 30. FIGURE 2B, 2C and 2D show the front, side and bottom views respectively of the side-mounted embodiment of FIGURE 2B. FIGURES 2A- D again illustrate an embodiment of the inlet 10 with an approximately trapezoidal shaped opening 11, an aft- closing, forward-swept cowl 30 with an approximately triangular tip 36 at the forward most part of the forward-swept panel section 32, and a bump 20 raised outwardly in a generally curved manner with a peak approximately on the centerline of the inlet 10. FIGURES 1A-D and 2A-D are by way of illustration and not limitation. For example, the panel sections 31, 32, and 33 that form the outer portion of the cowl 34 could be formed from a single piece. Furthermore, panel sections 31, 32, and 33 could be formed such that, when coupled to the body of the aircraft, the shape of the opening 11 was approximately elliptical. For further example, the tip 36 could be formed with a curved outer portion rather than the shape of a triangle. In operation, the bump 20 and cowl 30 work together to divert substantially all of the boundary layer air from the inlet 10. As the aircraft moves, the boundary layer air flows toward the inlet 10 while remaining approximately near the aircraft surface. Prior to reaching the inlet opening 11, the boundary layer air contacts the bump 20 which alters boundary layer air's path of motion to begin diverting this air away from the inlet opening 11. The shape of the cowl 30 helps to establish a pressure differential so that the pressure near the inlet opening 11 and inside the inlet 10 is higher than the pressure outside the opening 11. The shape of the cowl 30 creates a significantly lower pressure at the closure points 42. Thus, once the boundary layer air begins moving outboard due to the bump 20, the boundary layer air moves to the lower pressure regions near the closure points 42 and outside the inlet 10, rather than the higher pressure regions near the opening 11 of the inlet 10. This pressure gradient continues to divert boundary layer air during operation of the aircraft. Computer programs generate an image of an aircraft using computational fluid dynamics (CFD) analysis and aerodynamic simulation to model this inlet design. The computer program illustrates what the pressure field looks like around the inlet 10 and where the air flows in and around the inlet 10. FIGURE 3 represents a solution of the present invention at Mach 1.6 for a side-mounted inlet 10. As indicated above, the interior of the inlet 10 and near the opening 11 are at a higher pressure region relative to the exterior of the cowl 30 and the closure points 42. The white lines represent the paths 50 that particles or elements of boundary layer air travel. The paths 50 show how the bump 20 and cowl 30 work together to divert this boundary layer overboard. As the particles in the paths 50 approach the inlet 10, the bump 20 starts subtly moving the boundary layer air outboard. As shown by the paths 50, the boundary layer air particles then take a more extreme turn and go toward the outer edges of the cowl 30 due to the high pressure in the inlet 11 area that forces the boundary layer air to spill out to a lower pressure region. The higher pressure diverts the lower pressure, lower velocity boundary layer out around the cowl 30 and prevents substantially all of this lower energy boundary layer air from entering the inlet 10. The inlet concept of the present invention will divert boundary layer air just by using the cowl 30 concept. However, the present invention without the bump 20 will only divert boundary layer to a certain Mach level. In order to divert substantially all of the boundary layer air at Mach 1.6 or higher, the present invention must utilize both the bump 20 and the cowl 30. FIGURE 4A shows a CFD analysis on another embodiment of the present invention. In FIGURE 4, the surface pressure of the inlet 10 area is coded in terms of pressure contour where the dotted areas indicate lower pressure regions and cross-hatched areas represent regions of higher pressure. The areas marked 44 near the closure points 42 represent the lowest pressure region areas. FIGURE 4A demonstrates that, in operation, the interior of the inlet 10 is at a higher pressure relative to the exterior of the cowl 30. FIGURE 4B shows that as the particle trace paths 50 of the boundary layer air approach the bump 20, a subtle deflection of the particles outboard occurs as shown at point A. Due to the higher pressure region in the inlet 10, the boundary layer air particles then take a more extreme turn at point B and are diverted toward the outer edges and the closure points 42 of the cowl 30. As illustrated in FIGURE 4B, the bump 20 causes the boundary layer air to start moving outboard and the higher pressure forces the boundary layer air to take a more extreme turn towards the outer edges of the cowl 30. As shown in FIGURE 9A, the bump 20 comprises a compression surface 72, a transition shoulder 73, and a diffuser fairing 74. The shape of the compression surface 72 of the bump 20 can be determined by CFD computerized analysis. As illustrated in FIGURE 5, in a computerized model, placing a virtual right circular cone 60 with a constant semi-vertex angle in a supersonic CFD derived flow field 61 will create a conical shock 62 field that is axisymmetric around the virtual cone 60 emanating from the apex 65 of the virtual cone 60. A plane of particles 63 is released and sent through the axisymmetric conical shock 62 at a distance above the apex of the cone known as k. The width of the cone formed by the conical shock 62 at the intersections of the conical shock cone with the plane of particles 63 is denoted as w. The ratio of k to w (k/w) is approximately 0.1 for the embodiment shown in FIGURE 5. The particles 64 in the plane of particle released 63 propagate into the flow field and will alter course as the particles 64 pass through the conical shock 62. The shape formed above the original plane of release 63 after passing through the conical shock 62 defines the shape of a compression surface 72 that can be applied to the surface of the aircraft. If a compression surface 72 formed from this procedure is placed in the same flow field, the particles will flow as shown in FIGURE 5. A computer program could store this compression surface 72 in its memory and apply the necessary contour of the compression surface 72 to the given shape of the aircraft . The computer must contain a definition of the flow field generated by the virtual cone and a definition of the particle release plane 63, then the computer program can generate the compression surface. In an alternative embodiment, a parabolic shape, an elliptical shape, or a wedge could be used, rather than a right circular cone, to produce a flow field that could generate a compression surface 72. FIGURE 6 illustrates that an isentropic cone 70, rather than a right circular virtual cone 60 with a constant semi-vertex angle, can be used to create an isentropic compression surface 72. The isentropic cone 70 is used in the same manner described above to produce a flow field through which a plane of particles is released to determine the shape of the compression surface 72. The isentropic cone 70 has a cross section defined by an initial semi-vertex angle δ. relative to the centerline of the isentropic cone 70 and a final vertex angle δ ; relative to the centerline of the isentropic cone 70. As shown in FIGURE 6, the isentropic cone 70 has a smaller initial semi-vertex angle that gradually increases to a larger final vertex angle. The embodiment of FIGURE 7 shows an the isentropic cone 70 with a lower semi-vertex angle defined by a deflection angle of 12° that gradually increases to reach a final vertex angle with a final deflection angle of 21°. These initial and final deflection angles are illustrative and can be modified to achieve a total turning angle of approximately 21 degrees. FIGURE 8 shows yet another embodiment of the virtual cone used to define the shape of the bump 20 where the isentropic cone 70 has been positioned relative to the release of particles 63 so that the centerline of the isentropic cone 70 is offset from the airflow release plane of particles 63 by an angle of attack . Positioning a virtual cone at an angle of attack relative to a computer generated supersonic CFD derived flow field 61 will create a non-axisymmetric (or distorted) conical shock 66 wave emanating from the apex of the virtual cone. Whereas an axisymmetric conical shock wave 62 has a circular cross section, the non- axisymmetric shock wave 66 will have an elliptical cross section. The angle of attack illustrated in FIGURE 8 is approximately 7 degrees, though the angle of attack could be less than or greater than that illustrative value. The angle of attack can be introduced for either a right circular or an isentropic cone 70. Positioning the cone at an angle of attack α with respect to the plane of particle release 63 results in a compression surface 72 with increased ability to diver boundary layer air as compared to a compression surface 72 formed from a cone not positioned at an angle of attack a with respect to the flow field. FIGURE 9A shows an isentropic cone 70 CFD solution for a compression surface 72. Through the procedures described above, the isentropic cone 70 will allow a CFD solution of the shape of an isentropic compression surface 72 of a different shape than the compression surface created by a circular cone of a constant semi- vertex angle. When the CFD derived flow field contacts the isentropic cone 70 shown in FIGURE 9A, it produces a conical shock 62 and an infinite number of weak shocks 90 that form an isentropic compression fan 91. The isentropic compression fan 91 compresses to a focal point 92, after which the conical shock 62 returns to a single shock state. An isentropic compression surface 72 provides technical advantages over a compression surface created from a constant vertex right circular cone. The isentropic compression surface 72 allows the utilization of a smaller inlet capture area to reduce weight and aid integration onto the aircraft. Furthermore, designing the isentropic compression surface 72 from an isentropic cone 70 that produces a focal point 92 outside the cowl 30 will allow for better performance because the air flows through a series of weak shock waves 90 rather than merely one strong conical shock wave and a terminal shock wave. A bump 20 with a compression surface 72 shaped based on an isentropic cone 70 with an initial deflection angle of 12° and a final deflection angle of 21°, for example, will produce an isentropic compression fan 91 with a focal point 92 outside the cowl 30. FIGURE 9B shows the integration of one embodiment of an isentropic bump 20 with the forward-swept aft-closing cowl 30 that includes an isentropic compression surface 72 created from an isentropic cone 70 positioned at an angle of attack . The compression surface 72 in this embodiment has all of the advantages of a compression surface 72 formed without an angle of attack, with the additional advantage that the isentropic bump 20 of FIGURE 9B can have an increased performance in diverting boundary layer air relative to an isentropic bump 20 with a compression surface 72 formed from an isentropic cone 70 without an angle of attack . FIGURE 9B shows a cut-away side view of the bump 20 comprising a compression surface 72, a transition shoulder 73, and a diffuser fairing 74. The compression surface-72 terminates at the k/w point within the inlet. The transition shoulder 73 begins at the termination of the compression surface 72. In the embodiment described in FIGURE 9B, the surface of the transition shoulder 73 initially has an angle in relation to the body of the aircraft approximately equal to the total turning angle of the compression surface 72. This angle gradually decreases until the angle of the surface of the transition shoulder 73 in relation to the body of the aircraft approaches approximately zero so the surface of the transition shoulder 73 is approximately parallel to the body of the aircraft, at which point the transition shoulder 73 terminates. The termination point of the transition shoulder 73 (when the angle of the surface of the transition shoulder 73 in relation to the aircraft body is approximately zero) defines the minimum throat area 76, or minimum flow area, of the inlet. A diffuser fairing 74 begins at the termination of the transition shoulder 73. The diffuser fairing 74 can gradually decrease in the height the surface of the bump 20 is raised away from the body of the aircraft body as the diffuser fairing 74 continues toward the aft of the inlet. FIGURES 10A and 10B show graphical representations of two embodiments of the compression surface 72 and the transition shoulder 73. The graphs in FIGURES 10A and 10B plot the surface slope in degrees on the y axis versus the bump station in inches (moving forward to aft) on the x axis. The different lines of plotted data on each graph indicate the surface slope versus the bump station at various radial locations on the surface of the bump 20. The inset indicates how to determine the radial location, described by the angle φ with relation to the bump centerline. For example, when the angle φ is zero degrees, this radial location represents approximately the plane of symmetry for the inlet (for example, the [red] line represents the plot for a radial location described when φ is approximately zero degrees) . The angle φ increases as the radial location on the surface of the bump 20 moves outboard. The solid vertical line in the graph represents the point where the compression surface terminates and the transition shoulder begins; thus, the portion of the graph to the left of the solid line represents the compression surface 72 and the portion to the right represents the transition shoulder 73. The compression surface 72 geometry is identical on both FIGURES 10A and 10B and is defined by the CFD analysis described earlier. In both FIGURES, the dashed line, shown in the legend representing a φ of approximately 0°, shows a surface slope of approximately 19.5° at the termination of the compression surface 72; the dash, two dot, dash line, representing a φ of approximately 40°, shows a surface slope of approximately 14° at the termination of the compression surface 72. These FIGURES illustrate that as the radial location moves outboard, the peak angle at that bump station declines. In FIGURE 10A, the slope of the surface reaches a maximum for each radial location of the bump at the point where the compression surface terminates and the transition shoulder begins, or at approximately forty-one inches for the embodiment of FIGURE 10A. The reduction in slope from the beginning of the transition shoulder at approximately forty-one inches to the end of the transition shoulder at approximately sixty-seven inches is constant. Each of the curves is a straight line, representing a pure radius of curvature having a constant turn rate, from the maximum surface slope at the termination of the compression surface 72 to a zero surface slope at the termination of the transition shoulder 73. Thus, for the transition shoulder described in FIGURE 10A, the change in degrees per inch is approximately a constant. In FIGURE 10B, represents a different shape of the transition shoulder 74. In FIGURE 10B, the slope of the surface reaches a maximum at a different bump station location for each different radial location. For example, while the dashed line for a φ of approximately 0° still shows a peak surface slope (of approximately 19.5°) at the termination of the compression surface 72, the dash-two dot-dash line for o of approximately 40° shows a peak surface slope of approximately 19.5° at a bump station well into the transition shoulder 73 portion of the bump 20 (at approximately 50 inches) . Thus, rather than having reduction in surface slope for all radial locations beginning at the termination of the compression surface, the peak surface slope occurs increasing further into the transition shoulder 73 until a radial location of approximately 60°. Rather than a straight line from the point where each radial cut intersects the termination of the compression surface to the termination of the transition shoulder (as in FIGURE 10A) , a significant number of the radial cut plots rise to a nearly equivalent peak surface slope before falling back to zero. For example, the dash-two dot-dash line for the radial location of approximately 40° intersects the vertical separation line in FIGURE 10B at approximately 14°. However, instead of dropping immediately down to 0°, the dash-two dot-dash line increases again from 14° to approximately 19°, where it then drops down to 0° at a bump station of approximately sixty-six inches. These FIGURE illustrate that while the slope of the dashed line for φ of approximately 0° in FIGURE 10A has the highest slope, the slope of the dashed line for φ of approximately 0° in FIGURE 10B has a lower slope than a significant number of the other plotted lines. Controlling the slopes of the lines after the termination of the compression surface 72 determines the shape of the transition shoulder 73. In operation, the embodiment of the transition shoulder 73 in FIGURE 10B provides a technical advantage over the embodiment of the transition shoulder 73 of FIGURE 10A by increasing the amount of boundary layer air that will be diverted outboard. The amount of boundary layer air diverted outboard increases because this embodiment of the transition shoulder 73 of FIGURE 10B creates a greater spanwise static pressure gradient than the transition shoulder 73 embodiment of FIGURE 10A. FIGURE 11 shows the design parameters for determining a transition shoulder with a varying rate of surface slope change relative to radial location. On FIGURE 11 the bump 20 has a transition shoulder 73 beginning at the Bstop station 110 and terminating at the throat station 111. The surface 112 of the transition shoulder 73 shown in FIGURE 11 consists of two surface portions along surface 112, the first defined by the length L-_ from the Bstop station 110 to the intersection of the transition shoulder 113, and the second defined by L 2 from the intersection 113 to the throat station 111. The shape of the surface portion 112 of the transition shoulder 73 is defined by the radius R : for the length L : and by the radius R ; over the length h 2 , where the radii R-_ and R, are defined by the following formulas: L./2 R, θ.-θ, θ,+θ. Sin( ——- ) Cos ( ——-) R 2 = L 2 /Sin{Q 2 ) where θ-_ = Surface Slope Angle at Bstop Station θ = Maximum Surface Slope Angle L 2 = F - S - , roa F.S.) Bstop ≤ A < F.S.) τ ^ -F.S.) A ≡ θ 2 /1.5 @ φ = 55' 97/35105 PO7US97/04438 20 F • s • ) r.-. ca and represent fuselage stations on the surface of the aircraft . For a typical aircraft application, bump station of zero inches can correspond a fuselage station of approximately 262.3 inches. F.S.) 3ε- ~ c represents the point at which the compression surface 72 ends and transition shoulder 73 begins. F.S.) : . roa- represents the point at which the transition shoulder 73 terminates. The throat station point is a function of the length of the aircraft and the Bstop station is a user defined point. L + L 2 is a constant distance defined by the difference between F.S.) rrr . i- and F.S.) 3E = -.. The maximum surface slope angle (θ ; ) is a constant (for the embodiment illustrated in FIGURE 10B, θ 2 equals approximately 19.5°) that is determined at the Bstop station for a radial location of zero degrees. The surface slope angle at Bstop Station (Θ-) changes based on the radial location. In FIGURE 10B, a radial location of 0° the surface slope angle is approximately 19.5°; for a radial location of 10° the surface slope angle is approximately 18°; and so forth. At the Bstop station, L 2 is defined to be zero. L 2 is then determined for a radial location φ of 55° according to the above equation (L 2 =θ 2 /1.5 @ φ = 55°) . The 1.5 factor was determined through analysis as a constant whose use would result in an improved transition shoulder. Other factors could possibly be used in this equation. By assuming a linear regression between the points defining L 2 @ φ=0° and L 2 @ Φ=55° with a straight line, L 2 can be determined for each of the radial locations of 10, 20, 30, 40, and 50 degrees. Other regression analyses could also be used to determine L 2 . The embodiment described by FIGURES 10B and 11 use a straight line regression to determine L 2 for radial locations up to 55°. For radial locations ranging exceeding 55°, the user defines L 2 . The FIGURE 10B shows one example of how to define θ : and θ 2 at radial locations in excess of 55°. Once θ 2 , θ_, L : , and L ; . have been determined for each radial location φ, R and R 2 can be calculated for each radial location φ using the equations for R : and R 2 stated earlier. As shown in FIGURE 11, the slope of the surface of the transition shoulder 73 will increase from the Bstop station 110 over the length L to a maximum slope at the intersection 113, and will then decrease over the length L 2 to a slope of zero at the throat station 111. Thus, the graphical representation of the transition- shoulder surface slope of FIGURE 10B shows that the intersection 113 for the transition shoulder represented in FIGURE 10B occurs at a bump station of approximately 50 inches for a radial location of 40 degrees. As FIGURE 4B illustrates, the maximum surface slope of the transition shoulder 73 will occur at different bump stations for different radial locations φ. FIGURE 12 shows an isometric view of the compression surface 72 and transition shoulder 73 graphically represented in FIGURE 10B. As shown in FIGURE 12, the maximum or peak surface slope occurs at varying distances aft of the initiation of the compression surface 120. This varying aft distance is illustrated by the line depicting the locus of peak surface slope 120. This also indicates that h λ and L 2 will vary as the radial location φ varies. FIGURE 13 illustrates improved boundary layer diversion of the present invention using a compression surface designed from an isentropic cone with a seven degree angle of attack and a transition shoulder with a varying rate of decrease in surface slope (illustrated as Design B) as compared to the boundary layer diversion of an embodiment of the present invention using a compression surface designed with no angle of attack and with a transition shoulder with a constant rate of change in surface slope (illustrated as Design A) . Design B of FIGURE 13 shows the particle trace paths 50 of the boundary layer air pushed further outboard as compared to the particle trace paths 50 of the boundary layer air associated with the embodiment of Design A. The CFD analysis performed on these two embodiments indicate approximately 1% improvement in total performance for Design B over Design A. In summary, the present invention provides a diverterless engine inlet system that utilizes a bump with an isentropic compression surface, a transition shoulder and a diffuser fairing in combination with an aft-closing forward-swept cowl to divert boundary layer air from an aircraft engine inlet. The present invention eliminates the need for a boundary layer diverter, an overboard bypass system, and a boundary layer bleed system currently used on conventional air induction systems for supersonic aircraft. The present invention reduces aircraft weight, cost, and complexity. These features are eliminated because the compression surface and cowl work synergistically to provide a passive boundary layer diversion capability. Although the present invention has been described in detail, it should be understood that various changes, substitutions and alterations can be made hereto without departing from the spirit and scope of the invention as described by the appended claims.

Claims: WHAT IS CLAIMED IS: 1. a system for diverting boundary layer air from an inlet for an aircraft engine comprising: a bump having a surface raised outwardly from the body of the aircraft to begin diverting said boundary layer air prior to said boundary layer air entering said inlet, said bump positioned prior to an opening of the inlet and extending toward the rear of the inlet to form at least a portion of an inner surface of the inlet; and a cowl coupled to the body of the aircraft to define the opening of the inlet comprising; an aft-closing portion closed against the body of the aircraft at the aft-most points of the opening of the inlet; and a forward-swept portion, coupled to the aft- closing portion, that extends toward the front of the aircraft; said cowl working in conjunction with said bump to further divert said boundary layer air and prevent substantially all of said boundary layer air from entering said inlet during operation. 2. The system of Claim 1 wherein said bump further comprises a compression surface, wherein said compression surface is defined by deriving a supersonic flow field through use of computational fluid dynamics, placing a virtual cone in the derived supersonic flow field to create a conical flow field and a conical shock around the cone emanating from the apex of the virtual cone, releasing a plane of particles through the cone at a defined distance away from the apex of the cone to alter the course of the plane of particles as the particle plane pass through the conical shock to form a shape above the plane of release that defines the shape of the compression surface. 3. The system of Claim 2 wherein the centerline of the virtual cone is positioned at an angle relative to the plane of release of the particles. 4. The system of Claim 2 wherein the cone is an isentropic cone with a lower semi-vertex angle that gradually increases to a final vertex angle greater than the lower vertex angle. 5. The system of Claim 4 wherein the lower semi- vertex angle has a deflection angle of approximately 12 degrees and the final vertex angle has a deflection angle of approximately 21 degrees. 6. The system of Claim 2 wherein the cone is a right circular cone of a constant semi-vertex angle. 7. The system of Claim 2 wherein the distance from the apex of said virtual cone to the released plane of particles is defined by a k/w ratio. 8. The system of Claim 7 wherein the k/w ratio is approximately 0.1. 9. The system of Claim 2 wherein the compression surface terminates within the inlet at the k/w location. 10. The system of Claim 1 wherein air passing into the inlet must pass through a series of shock waves created by the bump and the cowl. 11. The system of Claim 10 wherein the series of shock waves comprise an initial conical shock, a plurality of weaker shocks, and a terminal shock. 97/35105 25 12. The system of Claim 2 wherein the bump further comprises; a transition shoulder beginning at the termination of the compression surface, wherein said transition shoulder having an angle in relation to the body of the aircraft, said angle gradually decreasing, said transition shoulder terminating when said angle in relation to the aircraft body is approximately zero, wherein the point at which said angle is approximately zero defines the minimum flow area of the inlet; and a diffuser fairing beginning at the termination of the transition shoulder and gradually reducing in height raised away from the body of the aircraft body as the bump continues toward the aft of the inlet. 13. The system of Claim 1 wherein the aft-closing portion further comprises a pair of opposing aft-closing panel sections, and wherein the forward-swept portion further comprises at least one forward-swept panel section located between the pair of aft-closing panel sections, and coupled to the pair of aft-closing panel sections to form said cowl. 14. The system of Claim 1 wherein the forward-swept portion and the aft-closing portion constitute a single, continuous structure that couples to the body of the aircraft to form the cowl. 15. The system of Claim 13 wherein the forward- swept panel section and the aft-closing panel sections constitute a single formed structure that couples to the body of the aircraft to form the cowl. 16. The system of Claim 13 wherein the aft-closing panel sections further comprise a leading edge, said leading edge extending from the point at which the aft- closing panel sections couple to the aircraft toward the front of the aircraft. 17. The system of Claim 13 wherein the aft-closing panel sections extend away from the body of the aircraft toward each other and are coupled to said forward-swept panel section to form an inlet opening that is approximately trapezoidal in shape. 18. The system of Claim 13 wherein the forward- swept panel section includes a tip extending from the intersection of the forward-swept panel section with the aft-closing panel sections toward the front of the aircraft. 19. The system of Claim of 18 wherein the tip of the forward-swept panel section is approximately triangular in shape. 20. The system of Claim 19 wherein the tip of the forward-swept panel section takes approximately the shape of an isosceles triangle. 21. a method for diverting boundary layer air from an inlet for an engine on an aircraft in flight, comprising; altering the path of the boundary layer air flowing toward the inlet by placing a bump, having a surface raised outwardly from the body of the aircraft, in the path the boundary layer air will flow to enter the inlet, positioning said bump to cause said boundary layer air to contact the bump prior to said boundary layer air entering the inlet; creating a pressure differential in the interior of the inlet by coupling a cowl to the body of the aircraft, wherein at least a portion of the bump is enclosed within the inlet by the cowl; and diverting substantially all of said boundary layer air to prevent said boundary layer from entering said inlet . 22. The method of Claim 21 wherein placing a bump further comprises forming a compression surface at the forward portion of the bump, said forming a compression surface further comprising: deriving a supersonic flow field through use of computational fluid dynamics; placing a virtual cone in the derived supersonic flow field to create a conical flow field and a conical shock around the cone emanating from the apex of the virtual cone; releasing a plane of particles through the cone at a defined distance away from the apex of the cone to alter the course of the plane of particles as the particle plane pass through the conical shock; and forming a shape above the plane of release that defines the shape of the compression surface bump. 23. The method of Claim 22 wherein placing a virtual cone in the derived supersonic flow field further comprises positioning the virtual cone at an angle relative to the plane of release of the particles. 24. The method of Claim 22 wherein placing a virtual cone further comprises placing an isentropic virtual cone with a lower semi-vertex angle that gradually increases to a final vertex angle greater than the lower vertex angle. 25. The method of Claim 24 wherein placing an isentropic virtual cone further comprises defining the isentropic virtual cone as having a lower semi-vertex angle with a deflection angle of approximately 12 degrees and a final vertex angle with a deflection angle of approximately 21 degrees. 26. The method of Claim 22 wherein placing the virtual cone further comprises placing a right circular cone of a constant semi-vertex angle. 27. The method of Claim 22 wherein releasing a plane of particles through the cone at a defined distance further comprises releasing the plane of particles through the cone at a distance defined by a k/w ratio. 28. The method of Claim 27 wherein releasing the plane of particles through the cone at a distance defined by a k/w ratio further comprises establishing a k/w ratio of approximately 0.1. 29. The method of Claim 22 wherein forming a compression surface further comprises terminating the compression surface within the inlet at the k/w location. 30. The method of Claim 22 wherein placing a bump further comprises: forming a transition shoulder beginning at the termination of the formed compression surface, wherein said forming a transition shoulder with an angle in relation to the body of the aircraft that gradually decreases; terminating the transition shoulder when said angle in relation to the aircraft body is approximately zero; and forming a diffuser fairing from the termination point of transition shoulder that gradually reduces in height raised away from the body of the aircraft body toward the aft of the inlet . 31. The method of Claim 21 wherein coupling a cowl to the body of the aircraft further comprises coupling a forward-swept portion between a pair of opposing aft- closing panel sections. 32. The method of Claim 21 wherein coupling a cowl to the body of the aircraft further comprises forming the forward swept portion from at least one forward-swept panel section. 33. The method of Claim 21 wherein coupling a cowl to the body of the aircraft further comprises forming the cowl from a single, continuous structure. 34. The method of Claim 31 wherein coupling a cowl to the body of the aircraft further comprises forming the aft-closing panel sections to include a leading edge and extending the leading edge from the point at which the aft-closing panel sections couple to the aircraft toward the front of the aircraft. 35. The method of Claim 31 wherein coupling a cowl to the body of the aircraft further comprises forming an inlet opening with an approximately trapezoidal shape by extending the aft-closing panel sections away from the body of the aircraft toward each other and coupling the aft-closing panel sections to the forward-swept panel section. 36. The method of Claim 31 wherein coupling a cowl to the body of the aircraft further comprises forming the forward-swept panel section to include an approximately triangular tip extending from the intersection of the forward-swept panel section with the aft-closing panel sections toward the front of the aircraft. 37. The system of Claim 36 wherein forming the forward-swept panel section further comprises forming the tip of the forward-swept panel section to have approximately the shape of an isosceles triangle. 38. a method for determining, through the use of computerized modeling, the shape of a compression surface to use in a system for diverting boundary layer air from an inlet comprising the steps of: deriving a supersonic flow field through use of computational fluid dynamics; placing a virtual cone in the derived supersonic flow field; creating a conical flow field, and an accompanying conical shock, around the cone emanating from the apex of the virtual cone; releasing a plane of particles through the cone at a defined distance above the apex of the cone; altering the course of the plane of particles as the particle plane pass through the conical shock to form a shape above the plane of particle release, said shape defining the shape of the compression surface. 39. The method of Claim 38 further comprising applying the compression surface to the surface of an aircraft as part of a boundary layer diversion system. 40. The method of Claim 38 wherein placing a virtual cone in the derived supersonic flow field further comprises positioning the virtual cone at an angle relative to the plane of release of the particles. 41. The method of Claim 38 wherein placing a virtual cone further comprises placing an isentropic virtual cone with a lower semi-vertex angle that gradually increases to a final vertex angle greater than the lower vertex angle. 42. The method of Claim 24 wherein placing an isentropic virtual cone further comprises defining the isentropic virtual cone as having a lower semi-vertex angle with a deflection angle of approximately 12 degrees and a final vertex angle with a deflection angle of approximately 21 degrees. 43. The system of Claim 38 wherein placing a virtual cone further comprises placing a right circular virtual cone of a constant semi-vertex angle. 44. The method of Claim 38 wherein releasing a plane of particles through the cone at a defined distance further comprises releasing the plane of particles through the cone at a distance defined by a k/w ratio. 45. The method of Claim 38 wherein releasing the plane of particles through the cone at a distance defined by a k/w ratio further comprises establishing a k/w ratio of approximately 0.1. 46. The method of Claim 38 wherein forming a compression surface further comprises terminating the compression surface within the inlet at the k/w location.

Inventors: HAMSTRA JEFFREY WILLIAM, SYLVESTER THOMAS GERARD, HAMSTRA, JEFFREY, WILLIAM, SYLVESTER, THOMAS, GERARD
Applicants: LOCKHEED CORP, LOCKHEED MARTIN CORPORATION
Application year: 1997
Main classification: F02C7
IPC codes: B64C7, B64C21, B64D33