Patent Description:
Air-breathing engine inlets have sometimes used internal diverters to remove or otherwise ameliorate boundary layers that may otherwise detrimentally affect engine performance. However internal diverters themselves have been associated with detrimental flow effects, for example flow disruptions such as shocks or expansion fans.

<CIT> discloses an air intake system suitable for a supersonic vehicle. The system includes a channel comprising an inlet and a side wall and a plenum coupled to the side wall. The plenum is configured to accept a flow of coolant. In certain embodiments, the coolant is the waste coolant from an on-board electronics cooling system. The system also includes a porous region in the side wall configured to allow a flow of bleed air from the channel through the porous region of the side wall into the plenum so as to aid the transition to supersonic flow. In certain embodiments, the flow of the bleed air is reduced at supersonic speeds by pressurization of the plenum with the coolant.

<CIT> discloses an air inlet for a ramjet, comprising a cowling defining a duct, a body mounted in said duct but partially extending forwardly therefrom, a streamlined member supported in front of and in spaced relation to said cowling and substantially forming an extension of said cowling which is separated therefrom by a slot, said streamlined member being substantially coextensive with the forwardly extending portion of said body and means for dividing said slot into a plurality of slots, whereby said <NUM> tains supersonic speeds and the separation of the boundary layer from the streamlined member is inhibited by said means.

<CIT> discloses, In a compressor having inner and outer walls of predetermined diameters respectively forming an annular passage, said passage having a longitudinal axis, a row of vanes fixed to said walls of airfoil shape and circurnferentially spaced transversely of the axis of said passage, each of said vanes substantailly spanning said annular passage in a radial direction and extending from one of said walls to the other, said vanes having a chordwise length extending along the longitudinal axis of said passage, and a protrusion extending into said passage from at least one of said walls and spanning the entire space between circumferentially adjacent vanes, said protrusion having a smoothly curved shape diverging at a predetermined rate from the diameter of said one wall and subsequently converging at a greater rate to the diameter-of said one wall in a direction downstream along said longitudinal axis, said protrusion having its point of maximum divergence located approximately within the last downstream quarter of the chordwise length of the vanes but upstream of the trailing edge thereof.

<NPL>, discloses designing an inlet by streamline tracing, comprising selecting a geometric shape of the projection of the desired air capture streamtube that lies within the circular capture streamline of a truncated Busemann flowfield.

<CIT> discloses a compact hypersonic modular inlet which divides a captured airstream and compresses it supersonically without the need for a variable geometry diffuser. According to the invention, an innerbody is provided with a plurality of perimetrically spaced inlet ducts which are defined by a single concave sidewall. The sidewalls intersect to form a center spike and radially directed, swept back leading edges. A cowling covers the inlets and is provided with a swept back leading edge adjacent to each duct, each leading edge having a profile conforming to a normal shock front.

According to an aspect of the invention, a method of configuring an inlet for a flight vehicle engine according to claim <NUM> is provided.

Preferred embodiments are provided in the dependent claims.

To the accomplishment of the foregoing and related ends, the invention comprises the features hereinafter fully described and particularly pointed out in the claims. The following description and the annexed drawings set forth in detail certain illustrative embodiments of the invention. These embodiments are indicative, however, of but a few of the various ways in which the principles of the invention may be employed. Other objects, advantages and novel features of the invention will become apparent from the following detailed description of the invention when considered in conjunction with the drawings.

An inlet for a flight vehicle engine, such as for a supersonic or hypersonic engine, includes an internal flow diverter to divert boundary layer flow. The flow diverter is configured to minimize disruption to flow outside the diverted boundary by being configured through use of a flow field that is also used to configure the walls of the inlet. The flow field that is used to configure an inlet-creating shape and a diverter-creating shape has the same flow generator, contraction ratio, compression ratio, mass capture ratio, pressure ratio between entrance and exit, and/or Mach number, for example. The internal diverter may be configured so as to allow selection of a leading edge shape for the internal diverter, for example to use a shape that helps avoid radar detection.

<FIG> schematically shows an air vehicle <NUM> that is powered by a propulsion system <NUM> that is mechanically coupled to a fuselage <NUM>. The air vehicle <NUM> may be a missile, projectile, an unmanned aircraft (an unmanned aerial vehicle or UAV), manned aircraft or an access-to-space vehicle. The air vehicle may have any of a variety of sizes, and any of a variety of operating conditions. In much of the description below the air vehicle <NUM> is described in terms of a high supersonic to hypersonic air vehicle, with a Mach number ranging from <NUM> to <NUM>, or more broadly with a Mach number of <NUM> to <NUM>. However the air vehicle <NUM> may operate at lower supersonic speeds (Mach number greater than one), or even at subsonic speeds.

The propulsion system <NUM> may be coupled to the fuselage <NUM> in any of a variety ways, including parts of the propulsion system <NUM> being integrally formed with parts of the fuselage <NUM>. The fuselage <NUM> may have any of a variety of suitable shapes, and may include additional components for carrying out one or more operations of the air vehicle <NUM>. Such additional components, to give a few nonlimiting examples, may include control systems (such as for steering), lift-producing and/or control surfaces (such as wings, fins, or canards, either fixed in position or movable in whole or in part), communication systems, cooling systems, sensors or other data-collecting systems, and/or any of a variety of payloads.

With reference in addition to <FIG>, the propulsion system <NUM> includes an air inlet <NUM>, an isolator or diffuser <NUM>, and a combustor or engine combustor <NUM>. Various shocks <NUM> occur upstream of and in the various parts of the propulsion system <NUM>. The air inlet <NUM> takes in air from the freestream and compresses the air, with one or more shocks perhaps occurring as the flow is compressed. The flow captured by the inlet <NUM> is also decelerated in the inlet <NUM>. The compressed air then exits the air inlet <NUM> to enter into the isolator <NUM>. There may be a throat <NUM>, a minimum area location at the boundary between the air inlet <NUM> and the isolator <NUM>.

The isolator <NUM> functions to keep the shocks stable, isolates dynamic flow fluctuations between the inlet and engine, provides demanded pressure rises, and/or provides desired flow patterns at its downstream end, where the air passes from the isolator <NUM> to the combustor <NUM>. There may be a shock train at lower flight speeds that further decelerates flow from supersonic at the throat <NUM>, to subsonic at the entrance to the combustor <NUM>.

In the combustor <NUM> fuel is added to the air flow, mixed, combustion occurs, and the combusted flow is passed through a nozzle <NUM>, producing thrust from the propulsion system <NUM>, which is used to propel the air vehicle <NUM>. Combustion products are exhausted from a downstream end of the combustor <NUM> through the nozzle <NUM>. The propulsion system <NUM> thus defines a flow path or propulsion flow path through the inlet <NUM>, the isolator <NUM>, the engine combustor <NUM>, and the nozzle <NUM>.

The combustor <NUM> may be any of variety of suitable devices for burning a fuel-air or fuel-oxidizer mixture and producing thrust. For example the combustor <NUM> (and/or the engine <NUM>) may be a ramjet, a scramjet, a dual-mode ramjet/scramjet, constant-volume combustion device, or perhaps a turbojet. In <FIG> the combustor <NUM> is shown as having a turbine <NUM>, but in many embodiments the combustor <NUM> has no turbine (or other moving parts).

In general the inlet <NUM> may have any of a variety of suitable shapes, for example being round, elliptical, or rectangular. The isolator <NUM> may have a general shape that makes the transition between a square, rectangular, trapezoidal or elliptical shape of the inlet <NUM> (to give a few examples) to a round or other-shaped combustor <NUM>. The inlet <NUM> and the combustor <NUM> may be in line with each other, or may be offset from one another and at different angular orientations. Many variations are possible for the configuration of the isolator <NUM>, and the examples given herein should not be considered as limiting to the invention.

<FIG> shows further details of one embodiment of the inlet <NUM>. The inlet <NUM> includes walls <NUM> that define an interior space <NUM> within the walls <NUM>. Air enters the interior space <NUM> and is compressed and directed downstream to the isolator <NUM> (<FIG>).

Referring now in addition to <FIG>, the inlet <NUM> includes a flow diverter <NUM> that is used to skim off a boundary layer of the flow from a top wall <NUM> of the inlet <NUM>. The internal diverter <NUM> is spaced away from the top wall <NUM> at a front end <NUM> of the internal diverter <NUM>. The internal diverter <NUM> is angled back in toward the top wall <NUM>, and in the illustrated embodiment the internal diverter <NUM> makes contact with top wall <NUM> toward a downstream end <NUM> of the internal diverter <NUM>.

The flow diverter <NUM> is attached to side walls <NUM> and <NUM>, near the tops of the side walls <NUM> and <NUM>. The boundary layer flow captured by the internal diverter <NUM> is directed through slots at the sides of the inlet <NUM>, out of the inlet <NUM>.

Removing the boundary layer flow using the internal diverter <NUM> may improve performance of the engine combustor <NUM> (<FIG>), for example by providing more uniform momentum in the intake air provided to the combustor <NUM>.

It is advantageous that the internal diverter <NUM> not unnecessarily impact the flow of air through the inlet <NUM>. Toward that end, the internal diverter <NUM> may be configured such that the flow qualities within the inlet <NUM> maintain the same sort of flow used in configuring the inlet walls <NUM>. The inlet walls <NUM> are configured using streamline traces in a defined flow field, and the internal diverter <NUM> is configured using the same flow field. This helps in maintaining the flow qualities of the original inlet shape. The internal diverter <NUM> has an aft-swept leading edge <NUM>, which is one of a variety of possible leading edge shapes, as is discussed further below.

The illustrated inlet <NUM> has a rectangular cross-section shape. Alternatively the inlet <NUM> could have other suitable shapes, such as trapezoidal, rounded, or circular. More broadly, the inlet could have any of a variety of further shapes, such as any shapes used to conform to a vehicle body shape.

<FIG> shows a high-level flow chart of a method or process <NUM> of configuring the inlet. In step <NUM> the inlet walls <NUM> (<FIG>) are streamlined traced using a flow field. The flow field is a theoretical (ideal) flow field past a surface or shape, such as a Busemann inlet. The flow field may model a supersonic flow field with characteristics, such as Mach number, contraction ratio (the ratio of the upstream and downstream areas of the inlet), compression ratio, pressure ratio between entrance and exit, mass capture ratio, pressure ratio between entrance and exit, and air pressure, corresponding to operating conditions for the engine. Shock locations within the flow field may be determined based on the flow characteristics. These factors may be taken into account in configuring the walls <NUM> of the inlet <NUM>.

In step <NUM> a shape is used for configuring the internal diverter <NUM> (<FIG>). The diverter-configuring shape is a shape created using the same flow field as was used in step <NUM> to configure the inlet walls <NUM>. The similarity includes (for example) using the same original generating body for streamline tracing, a geometrically similar leading edge shape (of higher aspect ratio), the same Mach number and contraction ratio, and the same throat shape. The identity of the throat shape may be used to allow the internal diverter <NUM> to blend into the shape of the inlet walls <NUM>. As illustrated in <FIG>, the diverter-creating shape <NUM> has a shape similar to that of an inlet-creating shape <NUM> used for configuring the inlet walls <NUM> (<FIG>). As illustrated, the diverter-creating shape <NUM> is shorter and wider than the inlet-creating shape <NUM>, giving the diverter-creating shape <NUM> a larger aspect ratio than the inlet-creating shape <NUM>. That the diverter-creating shape <NUM> is shorter than the inlet-creating shape <NUM>, with a vertical separation at a forward part of the shapes <NUM> and <NUM>. The amount of this separation (a vertical separation in the illustrated embodiment) corresponds to a desired height of the boundary layer to be removed at the axial station of interest. The shapes <NUM> and <NUM> are configured to converge with each other downstream of their leading edges, because the shapes <NUM> and <NUM> blend to the same throat shape.

The flow diverter <NUM>, or most of the flow diverter <NUM> (such as the upstream-most part of the flow diverter <NUM>) may be selected at any of a variety of locations within the diverter-creating shape <NUM>. One example is shown as reference number <NUM> in <FIG>. Thus the diverter-creating shape <NUM> is acting as a template for configuring the internal diverter <NUM>.

The intersection of the inlet walls <NUM> and the internal diverter <NUM> form the lateral extent of the surface <NUM>. The intersection of the original inlet top wall <NUM> and the diverter wall <NUM> form the aft extent of the internal diverter surface <NUM>.

One advantage of configuring the flow diverter <NUM> to be on the shape <NUM> is that this minimizes disruption of the flow through the inlet <NUM>. This is so, at least in part, because the shape <NUM> is on streamlines similar to the same flow as that used for creating the inlet walls <NUM>. Because the diverter shape is along streamlines of the original inlet, the leading edge of the diverter surface may exist at any location on the diverter, and may be of arbitrary (or at least somewhat arbitrary) shape. Another advantage is that this process allows variability in the shape of the leading edge of the internal diverter <NUM>, without causing significant impact or disruption on the flow. The trailing edge of the internal diverter <NUM> is where the internal diverter <NUM> merges with the walls <NUM>, may correspond to a location where the shapes <NUM> and <NUM> converge, such as at or near a downstream throat of the inlet <NUM>.

The method <NUM> allows the internal diverter <NUM> to be configured with its leading edge away from an upstream end of the inlet <NUM>, which is in contrast to how diverters are usually configured. The placement of the diverter edge downstream of the upstream end of the inlet <NUM> allows more flexibility in the configuration of diverters, for example to achieve desired characteristics in radar visibility.

Streamline-traced inlets tend to be long and have a boundary layer concentrated along the center of the body of the inlet. An inlet diverter such as the internal diverter <NUM> helps reduce the opportunity for boundary-layer separation along the inlet centerline. The internal diverter <NUM> may also be useful in aiding the inlet starting process.

Many different types of flow fields may be used in configuring the shapes <NUM> and <NUM>, for example modeling flow over wedges, cones, osculating cones, or in a Busemann inlet. For example, in a supersonic flow over a sharp-edged wedge, the streamlines downstream of a shock anchored on the front edge are identical, regardless of position on the shock. The streamlines are also tangent to the generating wedge, with no curvature, and flow field properties downstream of the leading shock are uniform.

The shapes <NUM> and <NUM> may be shaped to follow any of a variety of Busemann-based inlet shapes, for example by use of computer tools for generating blended, streamline-traced Busemann inlets with arbitrary leading edges and throats. The leading edge can be configured with a shape selected for ease of manufacture, such as a straight leading edge, or a leading edge made up of line segments. As an alternative to Busemann inlet flows, any three-dimensional flowfield may be used as a generator.

<FIG> shows various possible configurations of the internal diverter, with various leading edge shapes and/or positions within the inlet <NUM>. It will be appreciated that the illustrated embodiments are only examples, and that a variety of other leading edge shapes are possible. As an example, the leading edges may be curvilinear.

<FIG> shows an internal diverter 42b with a single-notch leading edge <NUM>. The internal diverter 42b is in a relatively close to a leading edge (upstream end) <NUM> of the inlet <NUM>, while still being downstream from the inlet upstream end <NUM>. <FIG> shows an internal diverter 42c that has the same shape as the internal diverter 42b, while being further downstream within the inlet <NUM>. <FIG> shows an internal diverter 42d with a W-shape leading edge <NUM>, with multiple line segments angled relative to one another. <FIG> shows another example, an internal diverter 42e with a W-shape leading edge <NUM>.

Claim 1:
A method of configuring an inlet (<NUM>) for a flight vehicle engine (<NUM>), the method comprising:
configuring (<NUM>) streamline-traced inlet walls (<NUM>) of the inlet, the streamline-traced inlet walls defining an interior space (<NUM>) within the streamline-traced inlet walls;
configuring (<NUM>) a streamline-traced internal diverter (<NUM>) within the inlet, for removing, from the inlet, boundary layer flow along one of the streamline-traced inlet walls; and
forming the inlet walls to have one or more slots through which flow captured by the streamline-traced internal diverter is directed, out of the inlet;
wherein the streamline-traced inlet walls are configured using an inlet-creating shape along streamlines of a flow field; and
wherein the streamline-traced internal diverter is configured using a diverter-creating shape along streamlines of the flow field used for the inlet-creating shape.