Abstract:
Hypersonic inlet systems and methods are disclosed. In one embodiment, an inlet for an airbreathing propulsion system includes an inboard surface at least partially shaped to conform to a plurality of streamline-traces of a design flowfield approaching an aperture, an outboard surface spaced apart from the inboard surface, an upper surface extending between the inboard and outboard surfaces, and a lower surface extending between the inboard and outboard surfaces, wherein leading edges of the inboard, outboard, upper, and lower surfaces cooperatively define the aperture.

Description:
CROSS-REFERENCE TO RELATED APPLICATIONS 
       [0001]    This application claims priority benefit from U.S. Provisional Patent Application No. 60/942,186, filed on Jun. 5, 2007, entitled “HYPERSONIC INLET SYSTEMS AND METHODS”. 
     
    
     STATEMENT OF GOVERNMENT INTEREST 
       [0002]    This invention was made with government support under subcontract number G1103-0002 to U.S. Government Contract NNL04AA01C awarded by the National Aeronautics and Space Administration. The U.S. Government has certain rights in this invention. 
     
    
     FIELD OF THE INVENTION 
       [0003]    The present disclosure relates to airbreathing propulsion systems, and more specifically, to hypersonic inlet systems and methods. 
       BACKGROUND OF THE INVENTION 
       [0004]    In order to operate efficiently, airbreathing hypersonic vehicles generally rely on at least two propulsion system types to complete their missions: one to propel the vehicle at relatively low speeds (Mach 0 to 3-4), and one to take over at higher speeds (Mach 3-4 to Mach 7-9 for hydrocarbon-fueled accelerator and cruise vehicles, and up to Mach 10-12 for hydrogen-fueled cruisers). The low-speed propulsion system is typically a turbine engine, designed to survive the thermal stresses of high-Mach operation and supply adequate thrust over the required speed range. High-speed thrust may be provided by a dual-mode ram/scramjet. 
         [0005]    Integration of these propulsion systems on a hypersonic vehicle may be enhanced by a common Multi-Role Air Induction System (MRAIS) to supply the needs of both propulsion system types, creating a so-called “turbine-base combined cycle” (TBCC) propulsion system. Requirements for an MRAIS include supplying the required amount of air with adequate pressure recovery and sufficient operability margin for each propulsion system independently, and also during propulsion system transition from low-speed to high-speed operating mode. MRAIS efficient operation and smooth mode transition rely on a well-designed, highly integrated system of inlet variable geometry and bleed. 
         [0006]    Prior art hypersonic inlet systems typically include variable geometry systems that are used to redirect and compress the incoming airflow during various portions of the vehicle&#39;s flight regime. Known hypersonic inlet systems include, for example, those systems disclosed in U.S. Pat. No. 5,054,288 issued to Salemann, and U.S. Pat. No. 5,337,975 issued to Peinemann. Typically, prior art TBCC inlets have relied upon a variable planar (or two-dimensional 2D) geometry integrated into an over/under arrangement, with the turbine flowpath being above the ramjet/scramjet flowpath, and having the turbine inlets external to, and forward of, the ramjet/scramjet inlet, while sharing a common external forebody. Typically, planar variable geometry features (e.g., flat flaps with effective sealing) have not been effectively integrated with inlets which are defined by axisymmetric flowfields. Axisymmetric flowfield inlets may offer benefits, including more efficient compression in converging (i.e., inward turning) flows, than 2-D flowfields which can have stronger shock waves and greater losses. Thus, although such prior art hypersonic inlet systems may provide desirable results, there is room for improvement. 
       SUMMARY OF THE INVENTION 
       [0007]    The present invention is directed to hypersonic inlet systems and methods. Embodiments of the present invention may advantageously provide desired inlet capabilities, including supplying the required amount of air with adequate pressure recovery and sufficient operability margin for each propulsion system independently, and also during propulsion system transition from low-speed to high-speed operating mode, without a great deal of mechanical complexity. 
         [0008]    In one embodiment, a method of designing an inlet for an airbreathing propulsion system includes establishing a design point flowfield, defining an aperture shape, tracing a plurality of streamlines around a perimeter of the aperture shape on the design flowfield to provide an intermediate shape, splitting the intermediate shape along a lateral centerline to provide a left lateral half and a right lateral half, and interchanging positions of the left lateral half and the right lateral half to form the inlet. In alternate embodiments, the design point flowfield may be an axisymmetric flowfield, and the aperture shape may be a two-dimensional (e.g. rectangular) shape. 
         [0009]    In another embodiment, an inlet for an airbreathing propulsion system includes an inboard surface at least partially shaped to conform to a plurality of streamline-traces of a design flowfield approaching an aperture, an outboard surface spaced apart from the inboard surface, an upper surface extending between the inboard and outboard surfaces, and a lower surface extending between the inboard and outboard surfaces, wherein leading edges of the inboard, outboard, upper, and lower surfaces cooperatively define the aperture. 
         [0010]    In a further embodiment, an airbreathing propulsion system includes an inlet assembly configured to receive an incoming airflow, the inlet assembly including a left inlet and a right inlet, each of the left and right inlets including: an inboard surface at least partially shaped to conform to a plurality of streamline-traces of a design flowfield downstream from an aperture; an outboard surface spaced apart from the inboard surface, the outboard surface comprising a movable high-speed flap; an upper surface extending between the inboard and outboard surfaces; a lower surface extending between the inboard and outboard surfaces, wherein leading edges of the inboard, outboard, upper, and lower surfaces cooperatively define the aperture; and a moveable low-speed flap spaced apart in an inboard direction from the high-speed flap. The system further including a low-speed diffuser duct operatively coupled to the inlet assembly, the moveable low-speed flap being configured to at least partially control a first airflow into the low-speed diffuser duct; a high-speed diffuser duct operatively coupled to the inlet assembly outboard from the low-speed diffuser duct, the moveable high-speed flap being configured to at least partially control a second airflow into the high-speed diffuser duct; a low-speed jet engine coupled to the low-speed diffuser duct; and a high-speed jet engine coupled to the high-speed diffuser duct. 
         [0011]    The features, functions, and advantages that have been discussed can be achieved independently in various embodiments of the present invention or may be combined in yet other embodiments further details of which can be seen with reference to the following description and drawings. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0012]    Embodiments of the present invention are described in detail below with reference to the following drawings. 
           [0013]      FIG. 1  is a schematic view of a hypersonic inlet design process in accordance with an embodiment of the invention; 
           [0014]      FIG. 2  is a front elevational view of a hypersonic vehicle having a hypersonic inlet in accordance with an embodiment of the invention; 
           [0015]      FIG. 3  is a top cross-sectional view of the hypersonic vehicle of  FIG. 2 ; 
           [0016]      FIG. 4  is an enlarged top cross-sectional view of the hypersonic vehicle of  FIG. 2 ; and 
           [0017]      FIGS. 5 and 6  are top cross-sectional views of the hypersonic vehicle having streamline-traced hypersonic inlets in accordance with alternate embodiments of the invention. 
       
    
    
     DETAILED DESCRIPTION 
       [0018]    The present invention relates to hypersonic inlet systems and methods. Many specific details of certain embodiments of the invention are set forth in the following description and in  FIGS. 1-6  to provide a thorough understanding of such embodiments. The present invention, however, may have additional embodiments, or may be practiced without one or more of the details described below. 
         [0019]    As described above, prior art turbine-based combined cycle (TBCC) inlets have traditionally relied upon a more purely planar (or two-dimensional 2D) geometry as integrated into an over/under arrangement, with the turbine flowpath being above the ramjet/scramjet flowpath. Also, the prior art approach generally had the turbine inlets external to, and forward of, the ramjet/scramjet inlet, while sharing a common external forebody, which facilitated incorporation of planar variable geometry. 
         [0020]    In general, embodiments of the present invention anchor to a potentially higher-performing streamline-traced inlet from an axisymmetric flowfield while combining this with a rectangular aperture to facilitate planar variable geometry (e.g., planar flaps) without increasing mechanical complexity. By movement of one or more flaps, each of the turbine inlets is moveable between a retracted position wherein there is relatively smaller or zero flow entering the turbine, and a deflected position wherein there is relatively greater flow from the ramjet/scramjet inlet entering the turbine. The resulting integration has the turbine and ramjet/scramjet flowpaths side by side instead of one over the other. 
         [0021]      FIG. 1  is a schematic view of a hypersonic inlet design process  100  in accordance with an embodiment of the invention. The design process  100  provides a single air induction system for the simultaneous or independent operation of a high-speed ram/scramjet and a separate low-speed turbine engine system. As shown in  FIG. 1 , in this embodiment, the design process  100  includes establishing a high-speed inward turning design point overall flowfield at a block  102 . The design point overall flowfield is preferably axisymmetric. As know to those skilled in the art, an axisymmetric flowfield is generally uniform and permits definition of flow properties by reference to a single dimension (e.g., a radial distance). The flowfield is defined by conducting a suitable and conventional flow analysis, such as using a computational fluid dynamics (CFD) code in a two-dimensional mode and with axisymetric boundary conditions. An idealized, converging inlet contour is modeled between a capture plane and a throat plane to define an initial pattern of shock waves in the flowfield. 
         [0022]    In one particular embodiment, the freestream Mach number of the design point overall flowfield is Mach 7. At a block  104 , a high-speed aperture shape and aspect ratio (AR) are defined. The aperture shape may be rectangular, and preferably provides adequate width and height to integrate low-speed flaps. In one particular embodiment, the aspect ratio AR is 4.3. A rectangular shape facilitates an effective integration of planar flaps for variable geometry, although other shapes (e.g., trapezoids, non-uniform shapes) may be used. Other design Mach numbers and aperture dimension ratios do not depart from the scope of the invention. 
         [0023]    The results of establishing the high-speed inward turning design point overall flowfield (block  102 ) and defining the high-speed aperture shape and aspect ratio AR (block  104 ) are used to trace streamlines around a perimeter of the aperture on the design flowfield at a block  106 . As known to those skilled in the art, the perimeter of the aperture is projected longitudinally to intersect with the flowfield definition, particularly with the initial shock wave. Streamlines are then traced downstream from the points of intersection as they would flow in the established flowfield. The streamlines collectively define a streamtube forming a favorable inlet contour (the inlet of one embodiment being shown in block  106  of  FIG. 1 ). 
         [0024]    Next, the resultant streamline traced inlet is split along its lateral centerline, and the left and right halves are interchanged, at a block  108 . The interchanging of the halves may provide these advantages: 1) a larger average width for incorporating the turbine inlet and duct without degrading performance; 2) improved structural efficiency and avoidance of relatively thin, cantilevered forward portions; and 3) a more open inlet which may facilitate free bleed of flow and avoid trapping shock waves for improved operability. As further shown in  FIG. 1 , the high-speed and low-speed inlet flap surfaces are defined at a block  110 . Finally, the high-speed and low-speed flap rotation schedules are established at a block  112 . The flap schedules may be determined based on a variety of factors, including freestream Mach number, capture area to throat area contraction ratio AOA, and airflow requirements of the turbojet. 
         [0025]      FIG. 2  is a partial, front elevational view of a hypersonic vehicle  200  having a pair of inlets  210  in accordance with an embodiment of the invention.  FIG. 3  is a top cross-sectional view of the hypersonic vehicle  200  of  FIG. 2 . As best shown in  FIG. 2 , the aperture of each inlet  210  is partially defined by an upper surface leading edge  212  and a lower surface leading edge  214 . A centerbody  216  is positioned between the inlets  210 . The shape of the surfaces of the centerbody  216  is defined by the interchanging of the positions of the left and right halves of the resultant shape defined by the streamline-tracing described above with respect to  FIG. 1  (block  108 ). 
         [0026]    Each inlet  210  further includes a low-speed inlet flap  218  configured to selectively open and close a low-speed diffuser duct  220  leading to a turbojet engine  222 . More specifically, in an open position (see lower flap  218  of  FIG. 3 ), the low-speed inlet flap  218  is rotated outwardly so that a portion of the captured airflow may pass to the turbojet engine  222  via the low-speed diffuser duct  220 , while in a closed position (see upper flap  218  of  FIG. 3 ), the low-speed inlet flap  218  is rotated inwardly so that the entrance to the low-speed diffuser duct  220  is blocked. Thus, the low-speed inlet flap  218  provides airflow capture and inlet operability control for the turbojet engine  222  during low-speed operation and transition. 
         [0027]    As further shown in  FIG. 3 , the centerbody  216  may include one or more variable geometry regions  230 . Preferably, the variable geometry regions  230  may be positioned proximate a throat region  232  leading into the low-speed diffuser duct  220  and may be equipped with one or more bleed regions  215  to enable proper control of the boundary layer, particularly at locations proximate to normal or oblique shock wave boundary layer interactions. The variable geometry regions  230  and bleed regions  215  may improve controllability and performance of the inlet  210 , such as by permitting improved control of an expansion rate of the low-speed (e.g. subsonic) flow entering the low-speed diffuser duct  226 . Other arrangements of components and variable geometry regions do not depart from the scope of this invention. 
         [0028]    Each inlet  210  further includes a high-speed inlet flap  224  positioned along an outboard edge of the inlet  210 . The high-speed inlet flap  224  is configured to selectively deflect toward the centerbody  216  to control airflow entering a high-speed diffuser duct  226  leading to a dual mode ramjet/scramjet engine  228 . More specifically, in a deflected position (see upper flap  224  of  FIG. 3 ), the high-speed inlet flap  224  is rotated inwardly, while in a retracted position (see lower flap  218  of  FIG. 3 ), the high-speed inlet flap  224  is rotated outwardly. In this way, the high-speed inlet flap  224  controls inlet contraction ratio and operability during operation of the ramjet/scramjet engines  228 . 
         [0029]      FIG. 4  shows additional details of the high-speed inlet flaps  224 . In this embodiment, the high-speed inlet flap  224  is a relatively longer flap that is positioned adjacent a fixed cowl  225  in the retracted position (see upper flap  224  of  FIG. 4 ). In the deflected position (see lower flap  224  of  FIG. 4 ), the high-speed inlet flap  224  is pivoted (i.e. inwardly turned) away from the fixed cowl  225  to control an internal area distribution within a compression zone  227  situated between the high-speed inlet flap  224  and the centerbody  216 /low-speed inlet flap  218  surfaces. 
         [0030]    Embodiments of the present invention may advantageously provide a streamline-traced hypersonic inlet which provides the necessary variable geometry without compromising overall system performance. Furthermore, embodiments of the present invention may provide the desired inlet capabilities, including supplying the required amount of air with adequate pressure recovery and sufficient operability margin for each propulsion system independently, and also during propulsion system transition from low-speed to high-speed operating mode, without a great deal of mechanical complexity. 
         [0031]    It will be appreciated that a variety of alternate embodiments of the invention may be conceived, and that the invention is not limited to the particular embodiments described above. In the following discussion of alternate embodiments, components which remain unchanged from the previously described embodiments are designated with like reference numerals. For the sake of brevity, only substantial structural and operational differences from the previously-discussed embodiments will be described. 
         [0032]      FIG. 5  is a top cross-sectional view of a hypersonic vehicle  500  having streamline-traced hypersonic inlets  510  in accordance with an alternate embodiment of the invention. In this embodiment, each inlet  510  includes a high-speed inlet flap  524  having a first portion  540  and a second portion  542 . In a first mode of operation, the first and second portions  540 ,  542  of the high-speed inlet flap  524  remain coupled together (see upper flap  524  of  FIG. 5 ), and the high-speed inlet flap  524  performs as described above with respect to the inlet  210  shown in  FIG. 4 . 
         [0033]    In a second mode of operation, however, the first portion  540  remains fixed relative to the fixed cowl  225  (see lower flap  524  of  FIG. 5 ), and the second portion  542  pivots (i.e. inwardly turns) independently of the first portion  540 . As shown in  FIG. 5 , the second portion  542  does not extend to the leading edge of the fixed cowl  225 . Consequently, in the deflected position (see lower flap  524  of  FIG. 5 ), the second portion  542  of the high-speed inlet flap  524  controls an internal area distribution within a relatively-smaller compression zone  527  which, in this embodiment, is situated between the second portion  542  and the low-speed inlet flap  218 . 
         [0034]      FIG. 6  is a top cross-sectional view of a hypersonic vehicle  600  having streamline-traced hypersonic inlets  610  in accordance with yet another embodiment of the invention. In this embodiment, each inlet  610  includes a high-speed inlet flap  624  having a first portion  640  and one or more second portions  642 . In a first mode of operation, the first and second portions  640 ,  642  of the high-speed inlet flap  624  remain coupled together (see upper flap  624  of  FIG. 6 ), and the high-speed inlet flap  624  performs as described above with respect to the inlet  210  shown in  FIG. 4 . 
         [0035]    In a second mode of operation, however, the first portion  640  remains fixed relative to the fixed cowl  225  (see lower flap  624  of  FIG. 6 ), and the second portion(s)  642  are actuated (e.g. by retracting fore or aft) independently of the first portion  640  to open up a bypass channel  644 . Thus, in the second mode of operation, the second portion(s)  642  allow some of the captured flow to be bypassed, providing additional control of the flow entering the high-speed diffuser duct  226 . 
         [0036]    While preferred and alternate embodiments of the invention have been illustrated and described, as noted above, many changes can be made without departing from the spirit and scope of the invention. Accordingly, the scope of the invention is not limited by the disclosure of the preferred embodiment. Instead, the invention should be determined entirely by reference to the claims that follow.