Patent Description:
A nacelle surrounds the engine. An inlet section of the nacelle is that portion of the nacelle that is forward of the fan section of the engine. One function of the inlet is to reduce noise. A minimum length of the inlet is typically required for noise reduction with high bypass ratio engines.

While longer inlets tend to improve noise reduction, that feature does not come without cost. A longer inlet is associated with increased weight and external drag. Additionally, the airflow at the inlet during takeoff typically creates a bending moment that is proportional to the length of the inlet. Longer inlets, therefore, tend to introduce additional load on the engine structure under such conditions.

A prior art gas turbine engine having the features of the preamble to claim <NUM> is disclosed in <CIT>. Prior art nacelle inlets for direct drive engines are also disclosed in <CIT>).

The present invention provides a gas turbine engine assembly according to claim <NUM>.

In a further non-limiting embodiment of the foregoing assembly, the dimensional relationship of L/D is between about <NUM> and about <NUM>.

In a further non-limiting embodiment of either of the foregoing assemblies, the dimensional relationship of L/D is between about <NUM> and about <NUM>.

In a further non-limiting embodiment of any of the foregoing assemblies, the dimensional relationship of L/D is about <NUM>.

In a further non-limiting embodiment of any of the foregoing assemblies, the dimension L is different at a plurality of locations on the fan case. A greatest value of L corresponds to a value of L/D that is at most about <NUM>, and a smallest value of L corresponds to a value of L/D that is at least about <NUM>.

In a further non-limiting embodiment of any of the foregoing assemblies, the dimension L varies and the dimensional relationship of L/D is based on an average value of L.

In a further non-limiting embodiment of any of the foregoing assemblies, the dimension L varies between a top of the inlet portion and a bottom of the inlet portion, and the dimensional relationship of L/D is based on a value of L near a midpoint between the top and the bottom of the inlet portion.

In a further non-limiting embodiment of any of the foregoing assemblies, the leading edges of the fan blades are in a reference plane, and the dimension L extends along a direction that is generally perpendicular to the reference plane.

In a further non-limiting embodiment of any of the foregoing assemblies, the engine has a central axis, the reference plane is generally perpendicular to the central axis, and the dimension L extends along a direction that is parallel to the central axis.

In a further non-limiting embodiment of any of the foregoing assemblies, the engine has a central axis, the forward edge on the fan case is in a reference plane, the leading edges of the fan blades are in a second reference plane, and the dimension L is measured between a first location where the central axis intersects the first reference plane and a second location where the central axis intersects the second reference plane.

In a further non-limiting embodiment of any of the foregoing assemblies, the fan is configured to deliver a portion of air into a compressor section and a portion of air into a bypass duct, a bypass ratio which is defined as a volume of air passing to the bypass duct compared to a volume of air passing into the compressor section being greater than or equal to about <NUM>, and the fan is configured to have a pressure ratio between about <NUM> and about <NUM> when operating at sea level.

In a further non-limiting embodiment of any of the foregoing assemblies, a fan blade tip speed of each of the fan blades is less than about <NUM> ft/second.

In a further non-limiting embodiment of any of the foregoing gas turbine engine assemblies, the geared architecture defines a gear reduction ratio greater than or equal to about <NUM>.

<FIG> schematically illustrates an example gas turbine engine <NUM> that includes a fan section <NUM>, a compressor section <NUM>, a combustor section <NUM> and a turbine section <NUM>. Alternative engines might include an augmenter section (not shown) among other systems or features. The fan section <NUM> drives air along a bypass flow path B while the compressor section <NUM> draws air in along a core flow path C where air is compressed and communicated to a combustor section <NUM>. In the combustor section <NUM>, air is mixed with fuel and ignited to generate a high pressure exhaust gas stream that expands through the turbine section <NUM> where energy is extracted and utilized to drive the fan section <NUM> and the compressor section <NUM>.

Although the disclosed non-limiting embodiment depicts a two-spool turbofan gas turbine engine, it should be understood that the concepts described herein are not limited to use with two-spool turbofans as the teachings may be applied to other types of turbine engines; for example a turbine engine including a three-spool architecture in which three spools concentrically rotate about a common axis and where a low spool enables a low pressure turbine to drive a fan via a gearbox, an intermediate spool that enables an intermediate pressure turbine to drive a first compressor of the compressor section, and a high spool that enables a high pressure turbine to drive a high pressure compressor of the compressor section.

The example engine <NUM> generally includes a low speed spool <NUM> and a high speed spool <NUM> mounted for rotation about an engine central longitudinal axis A relative to an engine static structure <NUM> via several bearing systems <NUM>. It should be understood that various bearing systems <NUM> at various locations may alternatively or additionally be provided.

The low speed spool <NUM> generally includes an inner shaft <NUM> that connects a fan <NUM> and a low pressure (or first) compressor section <NUM> to a low pressure (or first) turbine section <NUM>. The inner shaft <NUM> drives the fan <NUM> through a speed change device, such as a geared architecture <NUM>, to drive the fan <NUM> at a lower speed than the low speed spool <NUM>. The high-speed spool <NUM> includes an outer shaft <NUM> that interconnects a high pressure (or second) compressor section <NUM> and a high pressure (or second) turbine section <NUM>. The inner shaft <NUM> and the outer shaft <NUM> are concentric and rotate via the bearing systems <NUM> about the engine central longitudinal axis X.

A combustor <NUM> is arranged between the high pressure compressor <NUM> and the high pressure turbine <NUM>. In one example, the high pressure turbine <NUM> includes at least two stages to provide a double stage high pressure turbine <NUM>. In another example, the high pressure turbine <NUM> includes only a single stage. As used herein, a "high pressure" compressor or turbine experiences a higher pressure than a corresponding "low pressure" compressor or turbine.

The example low pressure turbine <NUM> has a pressure ratio that is greater than about <NUM>. The pressure ratio of the example low pressure turbine <NUM> is measured prior to an inlet of the low pressure turbine <NUM> as related to the pressure measured at the outlet of the low pressure turbine <NUM> prior to an exhaust nozzle.

The mid-turbine frame <NUM> further supports bearing systems <NUM> in the turbine section <NUM> as well as setting airflow entering the low pressure turbine <NUM>.

The core airflow C is compressed by the low pressure compressor <NUM> then by the high pressure compressor <NUM> mixed with fuel and ignited in the combustor <NUM> to produce high speed exhaust gases that are then expanded through the high pressure turbine <NUM> and low pressure turbine <NUM>. The mid-turbine frame <NUM> includes vanes <NUM>, which are in the core airflow path and function as an inlet guide vane for the low pressure turbine <NUM>. Utilizing the vane <NUM> of the mid-turbine frame <NUM> as the inlet guide vane for low pressure turbine <NUM> decreases the length of the low pressure turbine <NUM> without increasing the axial length of the mid-turbine frame <NUM>. Reducing or eliminating the number of vanes in the low pressure turbine <NUM> shortens the axial length of the turbine section <NUM>. Thus, the compactness of the gas turbine engine <NUM> is increased and a higher power density may be achieved.

The disclosed gas turbine engine <NUM> in one example is a high-bypass geared aircraft engine. In a further example, the gas turbine engine <NUM> includes a bypass ratio greater than about six, with an example embodiment being greater than about ten. The example geared architecture <NUM> is an epicyclical gear train, such as a planetary gear system, star gear system or other known gear system, with a gear reduction ratio of greater than about <NUM>.

In one disclosed embodiment, the gas turbine engine <NUM> includes a bypass ratio greater than about ten (<NUM>:<NUM>) and the fan diameter is significantly larger than an outer diameter of the low pressure compressor <NUM>. It should be understood, however, that the above parameters are only exemplary of one embodiment of a gas turbine engine including a geared architecture and that the present disclosure is applicable to other gas turbine engines.

The fan section <NUM> of the engine <NUM> is designed for a particular flight condition -- typically cruise at about <NUM> Mach and about <NUM>,<NUM> feet (<NUM>). The flight condition of <NUM> Mach and <NUM>,<NUM> ft. (<NUM>), with the engine at its best fuel consumption - also known as "bucket cruise Thrust Specific Fuel Consumption ('TSFC')" - is the industry standard parameter of pound-mass (lbm) of fuel per hour being burned divided by pound-force (lbf) of thrust the engine produces at that minimum point.

In another non-limiting embodiment the low fan pressure ratio is less than about <NUM>.

"Low corrected fan tip speed" is the actual fan tip speed in ft/sec divided by an industry standard temperature correction of [(Tram °R) / (<NUM> °R)]<NUM> (where °R=K x <NUM>/<NUM>). The "Low corrected fan tip speed", as disclosed herein according to one non-limiting embodiment, is less than about <NUM> ft/second (<NUM>/s).

<FIG> illustrates an example embodiment of the engine <NUM> with a nacelle or cowling <NUM>, that surrounds the entire engine. In this disclosure, like reference numerals designate like elements where appropriate and reference numerals with the addition of one-hundred or multiples thereof designate modified elements that are understood to incorporate the same features and benefits of the corresponding original elements. An inlet portion <NUM> is situated forward of the fan <NUM>. In this example, the inlet portion <NUM> has a leading edge <NUM>, which may be defined by the inlet side cut on the cowling <NUM>. The leading edge <NUM> is generally within a first reference plane <NUM>.

The nacelle <NUM> in some examples includes a flange <NUM> that is received against a leading edge on a fan case <NUM>. The inlet portion <NUM> has a length L between a selected location corresponding to the leading edge <NUM>, such as a location within the reference plane <NUM>, and a forward most portion <NUM> on leading edges on the fan blades <NUM> of the fan <NUM>. In this example, the length L may be considered an axial length of the inlet portion <NUM> because the length L is taken along a direction parallel to the central longitudinal axis A of the engine <NUM>. In the illustrated example, the inlet section of the nacelle <NUM> and the section of the fan case <NUM> that is forward of the blades <NUM> collectively establish the overall effective length L. In other words, in this example the length L of the inlet portion <NUM> includes the length of the inlet section of the nacelle <NUM> and some of the fan case <NUM>.

The fan blades <NUM> are three-dimensional swept fan blades (each having a similar side profile shown in <FIG>). In some examples, the fan blades <NUM> are forward-swept fan blades 92A (shown in <FIG>). In other examples, the fan blades <NUM> are rearward-swept fan blades 92B (shown in <FIG>). In further examples, the fan blades <NUM> include both forward-swept and rearward-swept portions (shown in <FIG>). A forward-swept fan blade is configured to have a radial portion of a leading edge of the fan blade forward of other portions of the leading edge. A rearward-swept fan blade is configured to have a radial portion of a leading edge of the fan blade rearward of other portions of the leading edge. A three-dimensional swept fan blade is twisted about an axis R extending in a radial direction between a tip or outermost edge <NUM> and a root <NUM> of the fan blade (shown in an axial view in <FIG>).

The fan blades <NUM> establish a diameter between circumferentially outermost edges <NUM>. The fan diameter D is shown in <FIG> as a dimension extending between the edges <NUM> of two of the fan blades <NUM> that are parallel to each other and extending in opposite directions away from the central axis A. In the illustration, the forward most portions <NUM> on the fan blades <NUM> are within a second reference plane <NUM>. In this example, the second reference plane <NUM> is oriented generally perpendicular to the central axis A of the engine <NUM>. The first reference plane <NUM> in this example is oriented at an oblique angle relative to the second reference plane <NUM> and the central axis A. In the illustrated example the oblique angle of orientation of the first reference plane <NUM> is approximately <NUM>°.

The length L is selected to establish a desired dimensional relationship between L and D. The dimensional relationship of L/D (e.g., the ratio of L/D) is between about <NUM> and about <NUM>. In some example embodiments, the dimensional relationship of L/D is between about <NUM> and about <NUM>. In some examples L/D is between about <NUM> and about <NUM>. In some example embodiments, the dimensional relationship of L/D is about <NUM>.

As can be appreciated from <FIG>, the length L of the inlet portion <NUM> (i.e., the combined length of the nacelle inlet and the forward section of the fan case) is different at different locations along a perimeter of the fan case <NUM>. The leading edge <NUM> is further from the second reference plane <NUM> near the top (according to the drawing) of the engine assembly than it is near the bottom (according to the drawing) of the engine assembly. The greatest length L in this example corresponds to a value for L/D that is no more than about <NUM>. The smallest length L in the illustrated example corresponds to a value for L/D that is at least about <NUM>. The value of L/D may vary between those two limits at different locations on the leading edge <NUM>.

In one example where the leading edge <NUM> has a variable distance from the second reference plane <NUM>, the dimensional relationship L/D is taken based upon a measurement of L that corresponds to an average measurement of the dimension between the leading edge <NUM> of the inlet portion <NUM> and the average location of the leading edge on the fan blades <NUM>. Stated another way, L/D in such an embodiment is based on a measurement of the average distance between the reference planes <NUM> and <NUM>. In another example where the dimension between the first reference plane <NUM> and the second reference plane <NUM> varies, the dimension L used for the dimensional relationship L/D is taken at a midpoint between a portion of the leading edge <NUM> that is most forward and another portion of the leading edge <NUM> that is most aft.

In another example, the dimension L is measured between a first location where the central longitudinal axis A of the engine intersects the first reference plane <NUM> and a second location where the axis A intersects the second reference plane <NUM>.

The dimensional relationship of L/D is smaller than that found on typical gas turbine engines. The corresponding dimensional relationship on most gas turbine engines is greater than <NUM>. Providing a shorter inlet portion length L facilitates reducing the weight of the engine assembly. A shorter inlet portion length also reduces the overall length of the nacelle and reduces external drag. Additionally, having a shorter inlet portion <NUM> reduces the bending moment and corresponding load on the engine structure during flight conditions, such as takeoff. A shorter inlet portion <NUM> also can contribute to providing more clearance with respect to cargo doors and other mechanical components in the vicinity of the engine.

The example engine <NUM> is a high bypass ratio engine having a larger fan with respect to the engine core components and lower exhaust stream velocities compared to engines with lower bypass ratios. Higher bypass ratio engines tend to have fan noise as a more significant source of noise compared to other sources. The illustrated example includes a shorter inlet yet does not have an associated effective perceived noise level that is noticeably greater than other configurations with longer inlets. One reason for this is that the example engine <NUM> includes a low pressure ratio fan that operates at a slower fan speed, which is associated with less fan noise. In one example, the fan <NUM> has a pressure ratio between about <NUM> and about <NUM>. A pressure ratio within that range corresponds to the engine operating at a cruise design point in some example implementations and/or at sea level in other example implementations. The shorter length L of the inlet portion <NUM> combined with the low pressure ratio of the fan <NUM>, which has a slower fan speed enabled by the geared architecture <NUM> of the engine <NUM>, results in an acceptable perceived engine noise level. Additionally, the geared architecture <NUM> enables the fan <NUM> to rotate at a slower speed and a lower fan tip relative Mach number which is associated with a reduced fan noise signature. The geared architecture <NUM> reduces the fan tip relative Mach number below <NUM> at the critical condition for noise attenuation, such as at full-takeoff, and in some instances, into the sub-sonic range at Mach <NUM> and below. Less acoustic liner material is necessary to maintain acceptable noise attenuation control because of the reduced fan source noise.

Utilizing a dimensional relationship as described above allows for realizing a relatively shorter inlet on a gas turbine engine while maintaining sufficient noise attenuation control. Additionally, the short inlet portion <NUM> combined with the low pressure ratio fan <NUM> provides improved propulsive efficiency and lower installed fuel burn compared to conventional gas turbine engine propulsion systems.

Claim 1:
A gas turbine engine assembly, comprising:
a fan (<NUM>) including a plurality of fan blades (<NUM>), a diameter of the fan (<NUM>) having a dimension D that is based on a dimension of the fan blades (<NUM>), each fan blade (<NUM>) having a leading edge;
an inlet portion (<NUM>) forward of the fan (<NUM>), a length of the inlet portion (<NUM>) having a dimension L between a forward most portion of the leading edges of the fan blades (<NUM>) and a forward edge (<NUM>) on the inlet portion (<NUM>);
a compressor section (<NUM>);
a turbine section (<NUM>); and
a geared architecture (<NUM>) configured to drive the fan (<NUM>) at a speed that is less than an input speed in the geared architecture (<NUM>), wherein the turbine section (<NUM>) is configured to drive the compressor section (<NUM>) and the geared architecture (<NUM>);
characterized in that:
a dimensional relationship of L/D is between <NUM> and <NUM>; and in that
each of the fan blades (<NUM>) is a three-dimensional swept fan blade (<NUM>) twisted about an axis (R) extending in a radial direction between a tip (<NUM>) and a root (<NUM>) of the fan blade (<NUM>).