Patent Application
20070181743
The present invention relates generally to engine inlet design, and more particularly, external compression inlet integration within aircraft design.
Low velocity, low pressure, boundary layer air (i.e., low energy air) builds up on the fuselage of a supersonic aircraft in front of the main engine inlets during high speed flight. This low energy air can cause poor engine performance. To address this problem, high-speed aircraft have traditionally employed boundary layer diverter systems to prevent low energy air from entering the air inlet. However, air induction systems often require complex subsystems in order to work properly at high speed. As such, these air inductions systems increase the weight, cost of production, mechanical complexity and cost of maintenance of the aircraft.
Air inlet systems for gas turbine powered supersonic aircraft decelerate the approaching flow to subsonic conditions prior to the engine face. Supersonically, this is accomplished through shock waves or isentropic compression generated externally, internally, or a “mixture” of both. Fixed geometry external compression inlets have typically been used for aircraft designed for short excursions to supersonic conditions, such as the F-16 and F-18, due to the relative simplicity and light weight of these designs. Higher speed capable aircraft, such as the F-15 and F-14, have employed variable geometry external compression inlets to obtain better engine/inlet airflow matching at low speeds and higher performance at supersonic speeds.
High supersonic cruise aircraft typically require maximum efficiency at the cruise point to obtain optimum range and payload. At design point cruise speeds above Mach 2, mixed compression inlet systems become more favorable due to reduced drag relative to an external compression system. Mixed compression inlets have been demonstrated in flight on 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 axisymmetric or two-dimensional compression schemes, in order to minimize shock interactions caused by complex 3D geometry. However, increasing demand for more integrated inlet/airframe concepts has resulted in the need for more exotic inlet aperture shapes.
Another solution provides a “bump” or raised external compression surface to divert low velocity, low energy air away from the aperture of the engine inlet before it can enter. However, such external compression surfaces usually focus solely on boundary layer diversion and not the cowl shape and inner duct walls associated with the cowl in order to create an integrated solution that may address the problems associated with advanced aircraft design.
Further limitations and disadvantages of conventional boundary layer diversion and modeling of and implementing compression surfaces to divert low energy air will become apparent to one of ordinary skill in the art to comparison with the present invention described herein.
The present invention provides for the use and design of Streamline Traced External Compression Inlets (STECI) that substantially eliminates or reduces disadvantages and problems associated with previously developed boundary layer diverting systems.
Embodiments of the present invention provide for a STECI and a methodology for designing such an inlet that represents a new solution to external compression inlet integration issues. The STECI concept may be derived using computational fluid dynamics (CFD). This allows an existing external compression, shock generating surface design to be used to obtain a CFD solution with slip wall boundaries at the inlet design point. This existing design serves as the “flow field generator” for the STECI. A high speed cruise inlet is typically designed such that the oblique shock produced by the external compression portion of the inlet lies on or very near the inlet cowl lip at the cruise point, minimizing the amount of air “spilled” by the inlet, but allowing for enough spilled air to maintain inlet airflow stability.
The design of a STECI inlet starts with a CFD solution produced by a shock-generating shape (the flow field generator), which could be as simple as a cone or wedge. The angle of the cone or wedge may be defined by a conceptual design trade study. The flow field generator may also be created by single or multiple complex surfaces. The aperture of the STECI inlet is defined according to shaping requirements. The desired cowl shape is projected onto a surface identical to, but downstream from, the aft-most external oblique shock produced by the flow field generator. The offset between the aft-most shock and the downstream aperture design surface allows for a small amount of air to be “spilled” by the inlet. The segments of the aperture that represent the leading edge of the STECI external compression surface are projected on the forward-most shock surface of the flow field generator. The cowl segments are then connected to the leading edge segments to form a full and continuous aperture. The resultant shape of the projected/offset cowl and the projected bump leading edge is the STECI aperture. Streamline seeds are next placed on the STECI aperture segments and are used to produce streamlines through the flow field generator CFD solution (which represents a physical flow field).
These streamlines provide the basis and tangency-reference for the surfaces that make up the portion of the STECI from the leading edge of the compression surface to the inlet throat. The surface tangencies feed into traditional methods to define and loft the subsonic diffuser from the inlet throat to the engine face.
Streamline tracing the cowl allows for alternate aperture shaping to be employed while producing the desired spillage and minimizing cowl drag, while maintaining the desired internal flow field and inlet performance. Without this technique, 3D shaped external compression inlets would require more extensive testing and design iterations to predict and optimize their performance. The techniques provided by embodiments of the present invention allow for the design of edge aligned, swept aperture shapes without incurring the aerodynamic penalties typically encountered with such designs.
By producing the design from the desired solution, the performance can easily be predicted using simple 1-D methods. The STECI has been computationally proven through CFD to produce uniform flow at the throat with low lip-loss and intended spillage characteristics.
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:
Preferred embodiments of the present invention are illustrated in the FIGs., like numerals being used to refer to like and corresponding parts of the various drawings.
Historic external compression air induction systems for gas turbine powered supersonic aircraft do not meet the current needs for inlet/airframe integration. These needs are driven by factors such as tailored aperture shapes, drag minimization, and propulsion performance. Embodiments of the present invention provide a method to create a Streamline Traced External Compression Inlet (STECI) that substantially addresses Inlet/Airframe integration issues. This method results in fewer design iterations than traditional historic external compression inlets.
Historical supersonic external compression inlets are based on simple 2D ramp or Axisymmetric configurations, or have been designed using Caret shaping or Diverterless Supersonic Inlet methodologies. None of these historical configurations or methods incorporates the cowl shape and inner duct walls associated with the cowl into a streamlined traced inlet solution. As a result, extra design iterations were needed to address lip loss, 3D effects, spillage, and other undesirable performance traits that became apparent once Computational Fluid Dynamics (CFD) and/or wind tunnel tests were used to quantify performance.
Embodiments of the present invention provide for the design and implementation of exotic engine inlets which may be used in high performance vehicles such as tactical aircraft. These exotic aperture shapes can be obtained by tracing particle streamlines from existing CFD (Computational Fluid Dynamics) solutions. The STECI is a 3D external compression inlet design concept derived from CFD by streamline tracing the supersonic section from a shock-generating geometry or “flow field generator”.
The STECI includes an external inlet compression surface combined with a cowl aperture shape. This inlet eliminates the need for traditional boundary layer diverters, overboard bypass systems, and boundary layer bleed systems which have been used in conventional air induction systems. Thus, the STECI may reduce vehicle (aircraft) weight, cost and complexity. These features are eliminated because of the compression surface and cowl work synergistically to provide boundary layer diversion capability.
The compression surface is designed to produce a suitable inlet-compatible flow field. The span wise static pressure of the surface can begin to divert boundary layer air outward. The pressure differential between the inlet and the area surrounding the inlet further diverts the boundary layer air outboard. The compression surface serves to reduce the Mach number just upstream of the terminal shock, thus reducing the tendency for shock-induced flow separation. The cowl is positioned to minimize the intake of low energy air and maximize the intake of free stream air within the inlet.
Compression surface 14 and cowl 18 work together to divert substantially all of the low energy air from inlet 10. As low energy, boundary layer air flows towards inlet 10, the low energy air remains approximately near the vehicle surface. As shown in
Previous techniques of developing and defining compression surfaces have often relied on expensive wind tunnel tests of actual models of proposed compression surfaces and cowl configurations on an aircraft. To properly model the actual performance of this geometry the overall aircraft design needs to be taken into account as well. Most design process focus on the inlet design but not the overall integrated design of the inlet within a given vehicle. The present invention provides a methodology for determining the compression surface and cowl geometry integrated within a vehicle that eliminates much of the modeling associated with prior solutions.
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Embodiments of the present invention provide an advanced inlet concept that represents a new solution to external compression inlet integration issues. The STECI starts with a CFD solution produced by a shock-generating shape which may be defined by a conceptual design trade study. The shock generating shape may be created using single or multiple surfaces depending on the number of upstream oblique shocks required to optimize the flow field at the throat of the inlet. An aperture is then defined according to shaping requirements. This desired aperture is projected onto surfaces identical to or similar to the external shock system produced by the shock generating shape. The projected cowl portion of the aperture is drawn on a surface duplicated from the aft-most oblique shock, and offset downstream from the shock to allow for air to be “spilled” by the inlet. The resultant shape of the projected/offset cowl and the projected bump leading edge is the STECI aperture. Streamline seeds are next placed along the STECI aperture and are used to produce streamlines through the CFD solution (which represents a physical flow field). These streamlines provide the basis for the surfaces that make up the portion of the STECI from the leading edge of the bump surface to the inlet throat and the tangencies for surfaces that will exist downstream of the throat. Traditional methods are used to define and loft the subsonic diffuser from the inlet aperture to the engine face. The STECI method can be used with more complex shock generating shapes with multiple sections in developing integrated geometry operable to supersonic cruise Mach numbers.
Embodiments of the present invention overcome problems associated with traditional supersonic external compression inlets driven by propulsion performance that did not address airframe/inlet integration issues. The geometries provided by the present invention support integration driven by total system performance, including the need for more exotic inlet aperture shapes and boundary layer management methods. STECI integration concerns during the initial design cycle instead of an afterthought. This allows for rapid development of highly integrated external compression inlets which provide excellent propulsion performance without compromising airframe/inlet integration.
As one of average skill in the art will appreciate, the term “communicatively coupled”, as may be used herein, includes wireless and wired, direct coupling and indirect coupling via another component, element, circuit, or module. As one of average skill in the art will also appreciate, inferred coupling (i.e., where one element is coupled to another element by inference) includes wireless and wired, direct and indirect coupling between two elements in the same manner as “communicatively coupled”.
The present invention has also been described above with the aid of method steps illustrating the performance of specified functions and relationships thereof. The boundaries and sequence of these functional building blocks and method steps have been arbitrarily defined herein for convenience of description. Alternate boundaries and sequences can be defined so long as the specified functions and relationships are appropriately performed. Any such alternate boundaries or sequences are thus within the scope and spirit of the claimed invention.
The present invention has been described above with the aid of functional building blocks illustrating the performance of certain significant functions. The boundaries of these functional building blocks have been arbitrarily defined for convenience of description. Alternate boundaries could be defined as long as the certain significant functions are appropriately performed. Similarly, flow diagram blocks may also have been arbitrarily defined herein to illustrate certain significant functionality. To the extent used, the flow diagram block boundaries and sequence could have been defined otherwise and still perform the certain significant functionality. Such alternate definitions of both functional building blocks and flow diagram blocks and sequences are thus within the scope and spirit of the claimed invention.
One of average skill in the art will also recognize that the functional building blocks, and other illustrative blocks, modules and components herein, can be implemented as illustrated or by discrete components, application specific integrated circuits, processors executing appropriate software and the like or any combination thereof.
Moreover, although described in detail for purposes of clarity and understanding by way of the aforementioned embodiments, the present invention is not limited to such embodiments. It will be obvious to one of average skill in the art that various changes and modifications may be practiced within the spirit and scope of the invention, as limited only by the scope of the appended claims.
