Patent Description:
Turbofan engines are widely employed for large commercial aircraft. As engines become larger and fans become wider, nacelles housing the fans must become shorter to achieve lower fuel burns. However, shorter nacelles, especially the resulting shorter inlets means that at adverse conditions such as high angles of attack or crosswind conditions the flow is more likely to separate behind the leading edge of the short inlet. The short inlet's smaller leading edge radius, and other features, makes it more difficult for flow to stay attached when airflow entering the engine must turn before heading in a direction approximately normal to the fan-face. If the flow separates at the leading-edge of the nacelle, the resulting distortion at the fan-face is undesirable. The separated flow may reduce performance, increase noise, and require heavier support structure to mitigate aerodynamically induced vibration. Existing solutions include simply making the inlet longer. Alternatively blow-in doors used in earlier nacelle designs may be employed. However, making the inlet longer is not viable option in many applications as it reduces effectiveness of the larger engine by creating excess drag and weight. Blow-in doors increase emitted noise from aircraft operations and are structurally complex. It is therefore desirable to provide alternative solutions for inlet flow control which overcome the constraints of prior art solutions and provide improved performance.

<CIT> discloses a nacelle which includes leading edge slats movable to modify inlet shape and area airflow into the nacelle. The articulating leading edge slats are movable for tailoring incoming airflow to current aircraft operating conditions. An actuator disposed within a thickness of the nacelle drives leading edge slats to a desired position to vary the inlet area and shape.

<CIT> discloses systems and methods for altering airflow to gas turbine engines. In this regard, a representative system includes a gas turbine engine inlet having a slat, the slat being movable between a retracted position and an extended position. In the extended position, the slat increases an effective diameter of the inlet compared to the diameter of the inlet when in the retracted position.

<CIT> discloses an aircraft engine air inlet noise suppression foil means comprising a plurality of segmented acoustically treated foil members stowed in recesses along the
wall of the inlet passageway during cruise flight. The foil members are movable to an operative position for takeoff and low speed flight where they collectively form a ringlike array which splits the inlet flow and increases the acoustically treated wetted surface area tending to reduce sound pressure levels. The preferred embodiment in this document provides stowed positions in auxiliary intake passageways which are opened by a translating cowl section, and foil members which automatically rotate to conform to local streamlines of the changing flow pattern in the inlet.

<CIT> discloses a gas turbine jet propulsion engine provided with flaps or the like whereby the intake cross sectional area may be selectively reduced to a value which causes the ingoing air to reach sonic velocity, thus effectively providing a block against the egress of noise from the engine compressor. The selective reduction of intake cross sectional area is accomplished by flaps or segment members movable in a generally axial direction with respect to the axis of the engine air intake.

<CIT> discloses a thrust reverser system capable of providing high efficiency within a tightly constrained nacelle. The thrust reverser system provides a displaceable internal door pivotally mounted within a transcowl. The displaceable internal door is rotatable about a pivot axis that is positioned aft of a front edge of the transcowl when the transcowl is in a deployed position.

And finally, <CIT> discloses a cascade set for creating sufficient drag to slow an aircraft. The cascade set includes one or more supporting vanes. A plurality of turning vanes are connected to the supporting vanes, and the turning vanes include forward and aft turning vanes. The forward turning vanes have a larger surface area than the aft turning vanes.

As disclosed herein a flow control system mountable on a leading edge of an aircraft engine inlet nacelle having the features of independent claim <NUM> is provided.

Embodiments of the flow control system are defined in dependent claims <NUM>-<NUM>.

The exemplary embodiments disclosed further provide a method for inlet flow control on an aircraft engine inlet nacelle comprising the steps listed in independent claim <NUM>.

Further embodiments of the method form the subject-matter of dependent claims <NUM>-<NUM>.

The features, functions, and advantages desired can be achieved independently in various exemplary 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.

The exemplary embodiments described herein provide translating multiple element turning vanes for adverse flow conditions in an ultra-short nacelle inlet to solve the problem of flow distortion on the fan face of the engine. The multiple element turning vanes are a deployable aerodynamic structure having a vane cascade with varying airfoil sections which are extended from the leading edge of the nacelle to decrease or eliminate flow separation from the inlet inner contour in off-nominal conditions such as crosswind and high angles of attack. The resulting variable geometry inlet deals with low speed and high angle of attack problems of separated flow, while still preserving the short nacelle by retracting into a recess in the nacelle leading edge to maintain cruise performance and the overall optimum performance of the larger engine.

Referring to the drawings, <FIG> depicts a large commercial aircraft <NUM> employing high bypass ratio turbofan engines <NUM> having ultra-short nacelles <NUM>. A radial cascade <NUM> of translating turning vanes are configured to be deployed around a leading edge <NUM> of the nacelle <NUM> in translating turning vane segments <NUM> as seen in <FIG>. The leading edge <NUM> circumscribes an inlet <NUM> providing air flow into the nacelle <NUM> for the turbofan engine <NUM>. The cascade <NUM> of extendible turning vanes translates to be positionable over a range from a stowed or retracted position as seen in <FIG> to the fully deployed or extended position seen in <FIG> and in detail in <FIG> and <FIG>. The cascade <NUM> of extendible turning vanes provides a flow control system to reduce flow distortion in the inlet.

Each translating turning vane segment <NUM> has a nose vane 17a, and multiple trailing vanes 17b, 17c, 17d and 17e supported by longitudinal ribs <NUM>. The nose vane 17a has an outer contour <NUM> matching the nose contour of the leading edge <NUM>, whereby in the retracted position, laminar flow of the inlet may be substantially maintained. A closing vane 17f again has an outer contour <NUM> matching the contour of the leading edge <NUM> to provide a smooth aerodynamic transition with the translating turning vane segment <NUM> in the fully extended position aligning the closing vane 17f with the leading edge <NUM>.

While shown in the drawings as having a common shape, vanes 17b-17e may have differing airfoil shapes and chord for tailoring the aerodynamics of flow turning as required. As seen in <FIG>, the radial cascade <NUM> of extendable translating turning vane segments <NUM> has fully extended length <NUM> which is determined based on the number and spacing of vanes 17a-17e but is nominally <NUM>-<NUM>% of the nacelle length <NUM>. For the embodiment shown, the nose vane 17a and four additional vanes 17b-17e are used. In alternative embodiments between <NUM> and <NUM> vanes may be employed.

<FIG> show the translating turning vane segment <NUM> in alternate views for clarity. The entire radial cascade <NUM> may be a single element which is extended from the leading edge <NUM> as a unit or cartridge with a continuous nose vane 17a. Alternatively, the radial cascade <NUM> may be split into quadrants, individually operable to extend translating turning vane segments <NUM> or other desired segments spanning a portion of the circumference of the leading edge <NUM> for operational sequencing as will be described in greater detail subsequently. For simplicity in explanation, the description is provided herein in terms of an individual translating turning vane segment <NUM>.

The ribs <NUM> support the vanes 17a - 17e to provide spacing of flow channels or slots 26a - 26e between the vanes. Deployment of the translating turning vane segments <NUM> increases the effective chord of the nacelle.

As seen in <FIG> and <FIG>, the ribs <NUM> are supported by a plurality of bearings <NUM> for longitudinal extension from and retraction into the nacelle <NUM>. A plurality of actuators such as rollers <NUM>, which may be pinion gears engaging a gear rack in the ribs <NUM>, are engaged to extend or retract the translating turning vane segments <NUM> in the cascade <NUM>. Stepper motors or similar devices may be employed to drive the pinion gears for precise extension length for partial and full extension. In alternative embodiments, linear actuators may be employed to extend and retract the translating turning vane segments <NUM>. Depending on the selected configuration of the cascade <NUM> (a single cartridge, individual translating turning vane segments or partial circumferential segments) multiple vane segments <NUM> may be interconnected by mechanical or hydraulic linkages and the number of individual ribs or segments driven by separate actuators may vary. In certain embodiments a single actuator may be employed to extend the entire cascade <NUM>. In certain embodiments, one or more of the vanes 17a-17e in the translating turning vane segments <NUM> may be rotatable about an axis <NUM> (represented in <FIG>).

For the embodiment shown, the outer extent of a slotted opening <NUM> in the leading edge <NUM> (best seen in <FIG> and <FIG>), through which the cascade <NUM> is extended and retracted, has an outer edge <NUM> relative to a nacelle inlet centerline <NUM> (seen in <FIG>) substantially aligned with a nominal cruise condition stagnation point <NUM> on the nacelle leading edge <NUM>. The cascade <NUM> is thus positioned radially inward from the stagnation point <NUM> to avoid interrupting flow on the exterior contour of the nacelle <NUM> for reduced drag at cruise. Matching of the outer contour <NUM> of the nose vane 17a to the contour of the leading edge of the nacelle reduces flow disruption on the internal surface of the inlet with the cascade <NUM> retracted for cruise conditions to minimize aerodynamic drag at cruise. The slotted opening <NUM> opens into a recess <NUM> in the nacelle <NUM> to house the cascade <NUM>.

Deployment of the cascade of translating turning vane segments <NUM> is demonstrated in the sequence of drawings in <FIG> (closed or retracted), <FIG> (partially extended) and <FIG> (fully extended). Exemplary embodiments will typically be operated at either the stowed/retracted or fully extended positions. However, in alternative embodiments, multiple partially extended positions may be employed, exposing differing numbers of vanes and slots to match aerodynamic operating requirements. As displayed in this sequence, extension of the entire radial cascade <NUM> of translating turning vane segments <NUM> is symmetrical about the nacelle inlet centerline <NUM>.

As seen in <FIG>, the partially extended translating turning vane segment <NUM> provides an intermediate chord extension for the nacelle <NUM>. Intermediate vane 17d has a contour comparable to the outer contour <NUM> of closing vane 17f to provide flow smoothing around and substantially close the slot <NUM> in the nacelle leading edge with the segment in the partially extended position.

<FIG> shows the symmetrical extended configuration of the radial cascade <NUM> of translating turning vane segments <NUM>. As annotated in <FIG>, quadrants 40a-40d around the nacelle may have differing aerodynamic conditions or effects created by angle of attack of the aircraft as a whole, cross winds, which may be partially shielded or mitigated by the fuselage <NUM> of the aircraft, or other aerodynamic phenomenon induced during flight, takeoff or landing of the aircraft. Each of the translating turning vane segments16 may be separately operable for extension and retraction. For high angle of attack operation of the aircraft, deployment of selectable groups of the translating turning vane segments <NUM> in at least lower outboard and lower inboard quadrants 40a and 40b would likely be desirable. For a strong outboard cross wind from the right, R, of the aircraft (left on the drawings as a front view of the aircraft), deployment of the translating turning vane segments grouped in lower and upper outboard quadrants 40a and 40d would be desirable. Similarly, for a strong inboard cross wind from the left, L, of the aircraft (right on the drawing) deployment of the translating turning vane segments grouped in lower and upper inboard quadrants 40b and 40c may be desirable. However, presence of the fuselage <NUM> may block left cross wind flow and deployment of the translating turning vane segments in upper inboard quadrant 40c may not be required. The descriptions herein are reversed for left and right designations for an engine mounted on the left side of the aircraft. Additionally, while shown in the drawings as equal quadrants, the "quadrants" may be interpreted as any selected arcuate segments of the circumference of the inlet.

For aircraft with certain operating conditions or engine mounting configurations, the cascade of translating turning vane segments may be altered to include only active devices in lower quadrants 40a and 40b, or those quadrants plus a lower portion of quadrants 40c and 40d which would be sufficient to accommodate all needed aerodynamic conditions.

Claim 1:
A flow control system mountable on a leading edge (<NUM>) of an aircraft engine inlet nacelle (<NUM>) housing a fan of a turbofan engine, the leading edge (<NUM>) circumscribing an inlet (<NUM>) providing air flow into the engine, and characterized in that the system comprising:
a cascade (<NUM>) of translating turning vanes (17a-17e) configured to extend from the leading edge (<NUM>) of the nacelle (<NUM>); and
at least one actuator (<NUM>) drivingly engaging the cascade of translating turning vanes to translate the cascade (<NUM>) from a retracted position to an extended position;
wherein the cascade (<NUM>) of translating turning vanes are engaged by a plurality of longitudinal ribs (<NUM>) forming translating turning vane segments (<NUM>); and
wherein each translating turning vane segment (<NUM>) has a nose vane (17a) and at least one trailing vane (17b, 17c, 17d, 17e) supported by the longitudinal ribs (<NUM>) to provide spacing of flow channels (26a - 26e) between the vanes.