Abstract:
A system, method, and apparatus for throat corner scoop offtake for mixed compression inlets for high speed aircraft engine applications is disclosed. The throat corner scoops are small air intakes located inside the large mixed compression inlet. They are positioned in a region otherwise prone to generate low pressure airflow. The throat corner scoops capture and remove the low pressure airflow from the bulk stream that is passed on to the engine. This location also provides inlet stability enhancement, and the airflow is used on the auxiliary systems.

Description:
BACKGROUND OF THE INVENTION 
   1. Technical Field 
   The present invention relates in general to inlet design for aircraft engines and, in particular, to an improved system, method, and apparatus for throat corner scoop offtake for mixed compression inlets for high speed aircraft engine applications. 
   2. Description of the Related Art 
   Air inlet systems for gas turbine powered supersonic aircraft are required to decelerate the approaching flow to subsonic conditions before it reaches the engine face. Supersonically, this can be done through shock waves or isentropic compression generated externally, internally, or by a mixture of both. Fixed geometry external compression inlets have typically been used for aircraft (e.g., the F-16 and F-18) designed for short excursions to supersonic conditions, due to the relative simplicity and light weight of these designs. Aircraft capable of higher speeds, such as the F-14 and F-15, have employed variable geometry external compression inlets to obtain better engine and inlet airflow matching at low speeds, and higher performance at supersonic speeds. 
   High altitude supersonic cruise aircraft typically require maximum efficiency at the cruise point to obtain optimum range and payload. At speeds above Mach 2, mixed compression inlet systems become favorable over external compression systems due to reduced drag. Mixed compression inlets have been demonstrated in flight on aircraft such as the A-12, SR-71, D-21, and XB-70. Several other designs have been tested over the past 50 years. All of these mixed compression designs were based on either axi-symmetric or two-dimensional compression schemes in order to minimize shock interactions caused by complex, three-dimensional geometry. 
   As shown in  FIG. 1 , axi-symmetric mixed compression inlet designs  11  typically include a throat bleed system that removes the low pressure boundary layer from the main duct  13 . This provides terminal normal shock stability and reduces shock/boundary layer interaction, which reduces overall pressure recovery and increases distortion. In this example, the throat bleed system includes both a centerbody shock trap  15  and a cowl slot  17 . The low energy air captured in the shock trap  15  would likely be exhausted overboard as it typically does not have enough energy to be used as utility flow. Having more energy due to a larger dynamic pressure component, the cowl slot  17  could possibly be used for utility flow. Various approaches to these bleed systems have been implemented in the industry for axi-symmetric and two-dimensional mixed compression inlets. Increasing demand for more integrated inlet and airframe concepts has resulted in the need for more exotic inlet aperture shapes. These exotic shapes impose additional geometric constraints that require novel approaches to bleed system design and integration. 
   Exotically-shaped, high speed engine inlets can suffer from several diverse performance losses. First, mixed compression inlets with duct wall interfaces that form acute angles (such as streamline traced inlets) can develop vorticity and a thick boundary layer (e.g., corner flow) in these regions which can cause separation and flowfield distortion that reduces engine performance. Second, a mixed compression inlet can undergo a process called “unstart” in which terminal shock stability is lost and airflow to the engine is drastically reduced, which consequently reduces engine performance. Third, airflow from the engine inlet is required for nacelle ventilation, environmental control systems, and various other utility and subsystems. While current state-of-the-art bleed system designs and integration approaches are workable for axi-symmetric and two-dimensional mixed compression inlets, an improved solution would be desirable for advanced shaped mixed compression inlet concepts that impose additional requirements based on geometrical constraints. 
   SUMMARY OF THE INVENTION 
   One embodiment of a system, method, and apparatus for an inlet throat bleed system for exotically-shaped mixed compression inlets comprises a throat corner scoop offtake. The invention addresses all three problems described above in the background. Throat corner scoops are small air intakes located inside the large mixed compression inlet. They are positioned in a region otherwise prone to generate low pressure airflow. The throat scoops capture and remove the low pressure airflow from the bulk stream that is passed on to the engine. This location also provides inlet stability enhancement, and the airflow is used on the auxiliary systems. 
   Mixed compression inlets traditionally have been designed as axi-symmetric or two-dimensional configurations. Such designs typically include a throat bleed system that removes the low pressure boundary layer from the main duct via a shock trap or bleed holes and slots, which are designed to remove boundary layer and provide terminal shock stability. The throat corner scoops of the present invention are designed to do this for non-axi-symmetric mixed compression inlet configurations such as, for example, advanced diverterless streamline-traced inlets. In contrast to current and future architectures, the previous designs did not have acute corners and therefore did not need to address the issue of corner flow affecting engine operation. Thus, the present solution has the added advantage of enhancing the viability of advanced streamline traced mixed compression inlets. 
   For example, the corner throat scoop is well suited for advanced streamline traced mixed compression inlets that incorporate acute corners and angles into the forward diffuser geometry. The scoop walls allow for a favorable aerodynamic transition as the main duct flow passes through a streamline-traced forward diffuser and moves toward a circular engine face or a bifurcated dual engine configuration. The invention also addresses several diverse design issues with a single solution. 
   The foregoing and other objects and advantages of the present invention will be apparent to those skilled in the art, in view of the following detailed description of the present invention, taken in conjunction with the appended claims and the accompanying drawings. 

   
     BRIEF DESCRIPTION OF THE DRAWINGS 
     So that the manner in which the features and advantages of the present invention, which will become apparent, are attained and can be understood in more detail, more particular description of the invention briefly summarized above may be had by reference to the embodiments thereof that are illustrated in the appended drawings which form a part of this specification. It is to be noted, however, that the drawings illustrate only some embodiments of the invention and therefore are not to be considered limiting of its scope as the invention may admit to other equally effective embodiments. 
       FIG. 1  is a half-sectional side view of a conventional, axi-symmetric aircraft engine inlet equipped with a conventional shock trap and cowl slot; 
       FIG. 2  is an isometric view of one embodiment of non-axi-symmetric, aircraft engine inlet constructed in accordance with the present invention; 
       FIG. 3  is a schematic side view of the inlet of  FIG. 2  and is constructed in accordance with the present invention; 
       FIG. 4  is a sectional end view of one embodiment of the inlet of  FIG. 2  taken along the line  4 - 4  of  FIG. 3  and is constructed in accordance with the present invention; 
       FIG. 5  is a sectional end view of another embodiment of the inlet of  FIG. 2  taken along the line  4 - 4  of  FIG. 3  and is constructed in accordance with the present invention; 
       FIG. 6  is an isometric view of an interior of the inlet of  FIG. 2  showing sectional images of air flow therethrough without the improvement of the present invention; and 
       FIG. 7  is a high level flow diagram of one embodiment of method constructed in accordance with the present invention. 
   

   DETAILED DESCRIPTION OF THE INVENTION 
   Referring to  FIGS. 2-7 , embodiments of a system, method and apparatus for manipulating airflow into a high speed aircraft engine is disclosed. As shown in  FIGS. 2-4 , one embodiment of the invention comprises a non-axi-symmetric, mixed compression inlet  31 . The mixed compression inlet  31  comprises a forward supersonic diffuser  33 , a minimum area throat region or throat  35 , and a subsonic diffuser section downstream of the throat  35  to provide subsonic airflow to a high speed aircraft engine  39 . The airflow  37  ( FIG. 3 ) is captured by the inlet  31 , compressed in the supersonic section  33  to the terminal normal shock just downstream of the throat  35 , and is further decelerated in the subsonic diffuser to the desired Mach number at the engine  39 . The engine  39  may be a single engine or may be bifurcated to feed multiple engines. 
   A small air intake or scoop  41  is located inside the mixed compression inlet  31 . In one embodiment, the scoop  41  is a throat corner scoop  41  that protrudes into an interior volume region of the mixed compression inlet  31  that is prone to generate a boundary layer of low pressure airflow  43  ( FIG. 6 ) as a component of the bulk airflow stream  37 . For ease of reference, only “sections” of the airflow  37  are illustrated in  FIG. 6 , including general gradations between high pressure flow  44  and low pressure flow  43 . The low pressure airflow  43  results from the acute angle formed by the shape of the supersonic diffuser  33 . The scoop  41  is not flush with the interior surfaces of the mixed compression inlet  31 , and it is isolated in location rather than integrated into the perimeter or circumferential geometry of the mixed compression inlet  31 . 
   The throat corner scoop  41  captures and removes the low pressure airflow  43  from the bulk airflow stream  37 . The throat corner scoop  41  also provides inlet shock stability enhancement by inherently increasing the rate of airflow as the normal shock moves forward of the scoop in the event of an imminent unstart. In addition, the throat corner scoop provides a source of additional airflow via duct  45  and the like for an auxiliary system  47  of the aircraft, such as secondary ventilation (e.g., nacelle ventilation), cooling, and/or airflow for the environmental control system (ECS), i.e., air conditioning for the pilot, avionics, and other temperature sensitive equipment. 
   In one embodiment, the mixed compression inlet  31  comprises a geometry that includes acute corners  49  (see, e.g.,  FIG. 4 ) that can create voracity and accumulate boundary layer. Subsonic diffuser  33  has an upper wall  51  and a lower wall  53  that converge toward each other as they approach each corner  49 . An inboard or sidewall  55  joins upper and lower walls  51 ,  53  inboard from each corner  49 . In the embodiments shown, the throat corner scoops  41  are located in at least one of the acute corners (two shown). The acute corners may be located at the throat  35  of the forward supersonic diffuser  33  of the mixed compression inlet  31 . The throat corner scoop  41  allows for a favorable aerodynamic transition as the bulk airflow stream  37  passes through the throat  35  of the forward diffuser  33 . The arrangement in  FIG. 5  is similar to  FIG. 4 , but corners  49 ′ are rounded. Upper and lower walls  51 ′,  53 ′ of subsonic diffuser  33 ′ are joined by sidewalls  55 ′ to define scoops  41 ′. 
   As shown in the illustrated embodiments, the throat corner scoop  41  comprises a plurality of discrete throat corner scoops, each of which is located in a low pressure airflow region inside the mixed compression inlet. A leading edge of the throat corner scoop  41  may be located at the throat  35 . The scoop  41  may comprise a small air intake located inside the mixed compression inlet and positioned in a region prone to generate low pressure airflow as a component of the bulk airflow stream, such that the scoop captures and removing the low pressure airflow from the bulk airflow stream. In  FIG. 2 , an aperture  57  for subsonic diffuser  33  is the farthest upstream location where subsonic diffuser  33  first completely surrounds and constrains the main bulk air flow. The air intake to scoop  41  is a considerable distance rearward from aperture  57 . 
   Referring now to  FIG. 7 , one embodiment of a method of manipulating airflow for a high speed aircraft engine is disclosed. The method begins as indicated at step  71 , and comprises providing a mixed compression inlet for a bulk airflow stream having a forward supersonic diffuser that transitions through a minimum area throat region into a subsonic diffuser for delivering subsonic airflow to the high speed aircraft engine (step  73 ); locating a scoop comprising a small air intake inside the mixed compression inlet and positioning the scoop in a region prone to generate low pressure airflow as a component of the bulk airflow stream (step  75 ); capturing and removing the low pressure airflow from the bulk airflow stream with the scoop (step  77 ); before ending as indicated at step  79 . 
   In other embodiments, the method may comprise providing inlet stability enhancement and a source of additional airflow for an auxiliary system selected from the group consisting of nacelle ventilation and an environmental control system. The method also may comprise providing the mixed compression inlet as a non-axi-symmetric design comprising a geometry that includes at least one corner formed at an acute angle, and the scoop is located in that acute angled corner. 
   The invention has many advantages, including providing boundary layer removal for shock and boundary layer interaction reduction, which is important for all high speed inlets. It also provides mixed compression stability margin for mixed compression inlets. In addition, the invention provides a high pressure source for utility flow needs, which is important for highly integrated inlet systems. Furthermore, this design removes the corner vortex that enables the viability of the streamline traced inlet for advanced inlets. All of these advantages are provided simultaneously with a single device. 
   While the invention has been shown or described in only some of its forms, it should be apparent to those skilled in the art that it is not so limited, but is susceptible to various changes without departing from the scope of the invention.