Patent Description:
A gas turbine engine generally includes a compressor section, a combustion section, a turbine section, and an exhaust section. The compressor section progressively increases the pressure of a working fluid entering the gas turbine engine and supplies this compressed working fluid to the combustion section. The compressed working fluid and a fuel (e.g., natural gas) mix within the combustion section and burn in a combustion chamber to generate high pressure and high temperature combustion gases. The combustion gases flow from the combustion section into the turbine section where they expand to produce work. For example, expansion of the combustion gases in the turbine section may rotate a rotor shaft connected, e.g., to a generator to produce electricity. The combustion gases then exit the gas turbine via the exhaust section.

The turbine section generally includes a plurality of stator vanes and a corresponding plurality of rotor blades. Each stator vane and each rotor blade include an airfoil positioned within the flow of the combustion gases; thus, the airfoils are referred to as hot gas path components. The airfoils of each stator vane and each rotor blade typically extend radially outward from a platform, such as an inner platform in the case of a stator vane. The airfoil of each stator vane extends to an outer platform at a radially outer end of the stator vane airfoil. The airfoil of each rotor blade extends to a tip at a radially outer end of the rotor blade airfoil. Certain rotor blades may include a tip shroud coupled to the radially outer end of the airfoil. A fillet may be provided at each transition between the airfoil and the platform(s) and/or at the transition between the airfoil and the tip shroud.

The airfoil may extend from a leading edge to a trailing edge downstream of the leading edge and may define aerodynamic surfaces therebetween, such as a pressure side surface and a suction side surface. Because the airfoils are hot gas path components, the surfaces thereof, such as the aerodynamic surfaces, are typically treated to enhance their resistance to the high temperature environment of the hot gas path. One such surface treatment is a thermal barrier coating. In conventional airfoils, each layer of the thermal barrier coating is generally uniform, e.g., having a constant, uniform thickness, across the aerodynamic surface, both in the span direction (i.e., from the root to the tip) and the in the flow direction (i.e., from the leading edge to the trailing edge). However, the conditions present at various locations around the airfoil may differ, and the properties of the layers in the thermal barrier may also vary. For example, one layer may be more robust to physical impacts, while another layer may provide better temperature resistance.

Accordingly, an airfoil for a turbomachine having a thermal barrier coating that provides robust physical characteristics in selected areas or portions of the airfoil would be useful.

<CIT> relates to gas turbine airfoils. At least a portion of the airfoils are coated with a coating that provides for erosion and corrosion protection for the portion of the airfoils. <CIT> also shows a coating for protecting and repairing an airfoil surface.

Aspects and advantages of the systems in accordance with the present invention will be set forth in part in the following description, or may be obvious from the description, or mav be learned through practice of the technology.

In accordance with one embodiment, an airfoil for a turbomachine as defined in claim <NUM> is provided. The airfoil includes a root and a tip spaced radially outward from the root. A span of the airfoil is defined between the root and the tip. The airfoil also includes a leading edge extending across the span of the airfoil from the root to the tip and a trailing edge downstream of the leading edge along a flow direction. The trailing edge also extends across the span of the airfoil from the root to the tip. The airfoil further includes a pressure side surface extending between the root and the tip and extending between the leading edge and the trailing edge and a suction side surface extending between the root and the tip and extending between the leading edge and the trailing edge. The suction side surface opposes the pressure side surface. The airfoil also includes a thermal barrier coating on the pressure side surface and the suction side surface. The thermal barrier coating includes a base layer and a top coat. A thickness of the base layer varies across each of the pressure side surface and the suction side surface with a maximum thickness of the base layer at the leading edge.

In accordance with another embodiment, a turbomachine is provided. The turbomachine includes a compressor, a combustor disposed downstream from the compressor, and a turbine disposed downstream from the combustor. The turbine includes a rotor blade and a stator vane. At least one of the rotor blade and the stator vane includes an airfoil. The airfoil includes a root and a tip spaced radially outward from the root. A span of the airfoil is defined between the root and the tip. The airfoil also includes a leading edge extending across the span of the airfoil from the root to the tip and a trailing edge downstream of the leading edge along a flow direction. The trailing edge also extends across the span of the airfoil from the root to the tip. The airfoil further includes a pressure side surface extending between the root and the tip and extending between the leading edge and the trailing edge and a suction side surface extending between the root and the tip and extending between the leading edge and the trailing edge. The suction side surface opposes the pressure side surface. The airfoil also includes a thermal barrier coating on the pressure side surface and the suction side surface. The thermal barrier coating includes a base layer and a top coat. A thickness of the base layer varies across each of the pressure side surface and the suction side surface with a maximum thickness of the base layer at the leading edge.

These and other features, aspects, and advantages of the present systems will become better understood with reference to the following description and appended claims.

A full and enabling disclosure of the present systems, including the best mode of making and using the present systems and methods, directed to one of ordinary skill in the art, is set forth in the specification, which makes reference to the appended figures, in which:.

Reference will now be made in detail to embodiments of the present systems, one or more examples of which are illustrated in the accompanying drawings. Each example is provided by way of explanation, rather than limitation of, the technology. In fact, it will be apparent to those skilled in the art that various modifications and variations can be made in the present technology without departing from the scope of the claims. Thus, it is intended that the present disclosure covers such modifications and variations as come within the scope of the appended claims.

Like or similar designations in the drawings and description have been used to refer to like or similar parts of the technology. As used herein, the terms "first," "second," and "third" may be used interchangeably to distinguish one component from another and are not intended to signify location or importance of the individual components.

As used herein, the terms "upstream" (or "forward") and "downstream" (or "aft") refer to the relative direction with respect to fluid flow in a fluid pathway. The term "radially" refers to the relative direction that is substantially perpendicular to an axial centerline of a particular component; the term "axially" refers to the relative direction that is substantially parallel and/or coaxially aligned to an axial centerline of a particular component; and the term "circumferentially" refers to the relative direction that extends around the axial centerline of a particular component.

Terms of approximation, such as "generally" or "about" include values within ten percent greater or less than the stated value. When used in the context of an angle or direction, such terms include within ten degrees greater or less than the stated angle or direction. For example, "generally vertical" includes directions within ten degrees of vertical in any direction, e.g., clockwise or counter-clockwise.

Although an industrial or land-based gas turbine is shown and described herein, the present systems as shown and described herein are not limited to a land-based and/or industrial gas turbine, unless otherwise specified in the claims. For example, the technology as described herein may be used in any type of turbomachine including, but not limited to, a steam turbine, an aircraft gas turbine or a marine gas turbine.

Referring now to the drawings, wherein identical numerals indicate the same elements throughout the figures, <FIG> illustrates a schematic diagram of one embodiment of a turbomachine, which in the illustrated embodiment is a gas turbine <NUM>. It should be understood that the gas turbine <NUM> of the present invention need not be a gas turbine engine, but rather may be any suitable turbomachine, such as a steam turbine engine or other suitable engine.

As shown, the gas turbine <NUM> generally includes an inlet section <NUM>, a compressor section <NUM> disposed downstream of the inlet section <NUM>, a plurality of combustors (not shown) within a combustor section <NUM> disposed downstream of the compressor section <NUM>, a turbine section <NUM> disposed downstream of the combustor section <NUM>, and an exhaust section <NUM> disposed downstream of the turbine section <NUM>. Additionally, the gas turbine <NUM> may include one or more shafts <NUM> coupled between the compressor section <NUM> and the turbine section <NUM>.

The compressor section <NUM> may generally include a plurality of rotor disks <NUM> (one of which is shown) and a plurality of rotor blades <NUM> extending radially outwardly from and connected to each rotor disk <NUM>. Each rotor disk <NUM> in turn may be coupled to or form a portion of the shaft <NUM> that extends through the compressor section <NUM>.

The turbine section <NUM> may generally include a plurality of rotor disks <NUM> (one of which is shown) and a plurality of rotor blades <NUM> extending radially outwardly from and being interconnected to each rotor disk <NUM>. Each rotor disk <NUM> in turn may be coupled to or form a portion of the shaft <NUM> that extends through the turbine section <NUM>. The turbine section <NUM> further includes an outer casing <NUM> that circumferentially surrounds the portion of the shaft <NUM> and the rotor blades <NUM>, thereby at least partially defining a hot gas path <NUM> through the turbine section <NUM>. The turbine section may also include a plurality of stator vanes <NUM>, which are mounted to the casing <NUM> within the hot gas path <NUM>.

During operation, a working fluid such as air flows through the inlet section <NUM> and into the compressor section <NUM> where the air is progressively compressed, thus providing pressurized air to the combustors of the combustor section <NUM>. The pressurized air is mixed with fuel and burned within each combustor to produce combustion gases <NUM>. The combustion gases <NUM> flow through the hot gas path <NUM> from the combustor section <NUM> into the turbine section <NUM>, wherein energy (kinetic and/or thermal) is transferred from the combustion gases <NUM> to the rotor blades <NUM>, causing the shaft <NUM> to rotate. The mechanical rotational energy may then be used to power the compressor section <NUM> and/or to generate electricity. The combustion gases <NUM> exiting the turbine section <NUM> may then be exhausted from the gas turbine <NUM> via the exhaust section <NUM>.

<FIG> illustrates a perspective view of an exemplary airfoil <NUM>, which may be incorporated into the rotor blade <NUM> and/or stator vane <NUM> of the turbine section <NUM> of the gas turbine <NUM>. As illustrated in <FIG>, the airfoil <NUM> may extend radially outward from a root <NUM> to a tip <NUM>. The airfoil <NUM> includes a pressure side surface <NUM> and an opposing suction side surface <NUM> (<FIG>). The pressure side surface <NUM> and the suction side surface <NUM> are joined together or interconnected at a leading edge <NUM> of the airfoil <NUM>, which is oriented into the flow of combustion gases <NUM> (<FIG>). The pressure side surface <NUM> and the suction side surface <NUM> are also joined together or interconnected at a trailing edge <NUM> of the airfoil <NUM> spaced downstream from the leading edge <NUM>. The pressure side surface <NUM> is generally concave, and the suction side surface <NUM> is generally convex.

Referring particularly to <FIG>, the airfoil <NUM> defines a span <NUM> extending from the root <NUM> to the tip <NUM>. In particular, the root <NUM> is positioned at zero percent (<NUM>%) of the span <NUM>, and the tip <NUM> is positioned at one hundred percent (<NUM>%) of the span <NUM>.

<FIG> provide various cross-sectional views of an exemplary airfoil <NUM>. It should be noted that each of the sectional views in <FIG> is a constant-span section. For example, <FIG> may be taken at about fifty percent (<NUM>%) of the span <NUM>, and the entirety of the section through the airfoil <NUM> as shown in <FIG> lies at the same position along the span <NUM>, e.g., at about fifty percent (<NUM>%) of the span <NUM>.

As may be seen, e.g., in <FIG>, the airfoil <NUM> defines a camber line <NUM>. More specifically, the camber line <NUM> extends from the leading edge <NUM> to the trailing edge <NUM>. The camber line <NUM> is also positioned between and equidistant from the pressure side surface <NUM> and the suction side surface <NUM>. Also, as may generally be seen in <FIG>, a thermal barrier coating <NUM> may be provided on the outermost surface of the airfoil <NUM>. For example, as illustrated in <FIG>, the thermal barrier coating <NUM> may be provided on each of the pressure side surface <NUM> and the suction side surface <NUM>.

<FIG> provides an enlarged view of a leading edge portion of the airfoil <NUM>, e.g., a portion of the airfoil <NUM> including the leading edge <NUM> and parts of the airfoil <NUM> proximate thereto. <FIG> provides an enlarged view of a mid-forward portion of the airfoil <NUM>, e.g., a portion of the airfoil <NUM> that is aft of the leading edge <NUM> and forward of a midpoint of the airfoil <NUM>, the midpoint being defined along the direction of flow of combustion gases. <FIG> provides an enlarged view of a mid-aft portion of the airfoil <NUM>, e.g., a portion of the airfoil <NUM> that is aft of the midpoint of the airfoil <NUM> and forward of the trailing edge <NUM>. <FIG> provides an enlarged view of a trailing edge portion of the airfoil <NUM>, e.g., a portion of the airfoil <NUM> including the trailing edge <NUM> and parts of the airfoil <NUM> proximate thereto, according to one or more embodiments. <FIG> and <FIG> provide further views of the trailing edge portion of the airfoil according to one or more additional embodiments.

As may be seen in <FIG>, the base or core material of the airfoil <NUM> may comprise a substrate <NUM> onto which the thermal barrier coating <NUM> is applied. For example, the substrate <NUM> may be or include a metallic material, such as an alloy including iron and nickel or cobalt, e.g., high-temperature steel, superalloys, and/or other suitable metal alloys. The thermal barrier coating <NUM> may include, for example, ceramic material. The thermal barrier coating <NUM> is generally formed on an exterior surface <NUM> of the substrate <NUM>. In various embodiments, the airfoil <NUM> may include the thermal barrier coating <NUM> directly on the exterior surface <NUM> or may include a bond coat <NUM> formed directly on the exterior surface <NUM> of the substrate <NUM>. In embodiments wherein the bond coat <NUM> is provided, the thermal barrier coating <NUM> may be formed directly on the bond coat <NUM>.

As may be seen in <FIG>, the thermal barrier coating <NUM> may include a plurality of layers. For example, the thermal barrier coating <NUM> may include a base layer <NUM> closer to the substrate <NUM> and a top coat <NUM> outward and/or on top of the base layer <NUM>.

The thickness of the base layer <NUM> may vary across each of the pressure side surface <NUM> and the suction side surface <NUM>. For example, the thickness of the base layer <NUM> may decrease across each of the pressure side surface <NUM> and the suction side surface <NUM>. In at least some embodiments, the thickness of the base layer <NUM> may decrease from a maximum at the leading edge <NUM> to a minimum at the trailing edge <NUM>. Providing a maximum thickness of the base layer <NUM> at and around the leading edge <NUM> may advantageously provide improved physical resistance at the leading edge portion of the airfoil <NUM>, which may be advantageous when the leading edge portion of the airfoil <NUM> is more likely than downstream portions of the airfoil <NUM> to experience a physical impact and/or erosion.

Thus, in various embodiments, the thickness of the base layer <NUM> at the leading edge <NUM> may be greater than a thickness of the top coat <NUM> at the leading edge <NUM>. In some embodiments, the thickness of the base layer <NUM> may taper continuously from the leading edge <NUM> to the trailing edge <NUM> across each of the pressure side surface <NUM> and the suction side surface <NUM>. In some embodiments, the thickness of the base layer <NUM> at the trailing edge <NUM> may be less than the thickness of the top coat <NUM> at the trailing edge <NUM>.

In some embodiments, the thermal barrier coating <NUM> may consist of only the base layer <NUM> at the leading edge <NUM>, e.g., the ratio of base layer <NUM> to top coat <NUM> may be <NUM>:<NUM> in the leading edge portion, and there may be no top coat <NUM> at the leading edge <NUM>. In alternative embodiments, the thermal barrier coating <NUM> may include mostly base layer <NUM> at and around the leading edge, such as a ratio of base layer <NUM> to top coat <NUM> of about <NUM>:<NUM>, such as about <NUM>:<NUM>.

In some embodiments, e.g., as illustrated in <FIG>, no thermal barrier coating <NUM> may be provided at the trailing edge <NUM>. For example, in such embodiments, the base coat <NUM> and the top coat <NUM> may both taper down to zero thickness at or proximate to the trailing edge <NUM>, as illustrated in <FIG>.

In other embodiments, the thermal barrier coating <NUM> may continue across the trailing edge <NUM>. For example, in some embodiments, the proportion of the layers of the thermal barrier coating <NUM> at the trailing edge <NUM> may be the same as or similar to the proportion at the leading edge <NUM> (e.g., as illustrated in <FIG>). In such embodiments, e.g., as illustrated in <FIG>, the top coat <NUM> may taper down in thickness at or approaching the trailing edge <NUM>, such that the ratio of base layer <NUM> to top coat <NUM> may be <NUM>:<NUM> in the trailing edge portion, and there may be no top coat <NUM> at the trailing edge <NUM>. Additionally, and similar to the leading edge <NUM> as described above, the thermal barrier coating <NUM> may include mostly base layer <NUM> at and around the trailing edge <NUM>, such as a ratio of base layer <NUM> to top coat <NUM> of about <NUM>:<NUM>, such as about <NUM>:<NUM>.

In some embodiments, e.g., as illustrated in <FIG>, both layers of the thermal barrier coating <NUM>, e.g., the base layer <NUM> and the top coat <NUM>, may wrap around the trailing edge <NUM> with a generally constant thickness around the trailing edge <NUM>.

Referring again to <FIG>, in some embodiments, the leading edge portion and/or the area or portion of the airfoil <NUM> where the base layer <NUM> thickness is at a maximum may encompass a high-impact zone <NUM> on the airfoil <NUM>. It should be noted that the high-impact zone <NUM> refers to a relatively high probability of physical impact and/or erosion within the delineated area <NUM> relative to the remainder of the airfoil <NUM>, in particular, the remainder of the aerodynamic surfaces <NUM> and <NUM> thereof. Additionally, it should be understood that the high-impact zone <NUM> is generally symmetrical about the leading edge <NUM> and extends over about the same distance along the length of the airfoil <NUM> on both the pressure side <NUM> (as shown in <FIG>) and the suction side <NUM>.

As noted above, the length of the airfoil <NUM> is defined along the flow direction. In at least some embodiments, the leading edge portion may be coextensive with the high-impact zone <NUM>, e.g., the maximum thickness of the base layer <NUM> may be provided at the leading edge <NUM> and throughout the high-impact zone <NUM>. Moving from the leading edge <NUM> to the trailing edge <NUM>, the thickness of the base layer <NUM> may be at a maximum in the area <NUM>, may decrease in the first intermediate zone <NUM>, and may decrease again in the second intermediate zone <NUM>. As a result, the thickness of the base layer <NUM> may be less in the second intermediate zone <NUM> than in the first intermediate zone <NUM>. Also, the thickness of the base layer <NUM> may decrease again in a downstream or aft zone <NUM>. In some embodiments, the thickness of the base layer <NUM> may be at a minimum at the trailing edge <NUM> and/or within the area of the aft zone, as indicated by <NUM> in <FIG>.

In some embodiments, the minimum thickness of the base layer <NUM> may account for about seventy percent (<NUM>%) of the total thermal barrier coating <NUM> or less, such as about sixty percent (<NUM>%) or less, such as about fifty percent (<NUM>%) or less, such as about forty percent (<NUM>%) or less, or such as about thirty percent (<NUM>%) or less.

Still with reference to <FIG>, in some embodiments, the thickness of the base layer <NUM> may also vary across the span <NUM> of the airfoil <NUM>. For example, the ratio of the base layer <NUM> and top coat <NUM> within the thermal barrier coating <NUM> may be at the base ratio, e.g., minimum thickness of the base layer <NUM>, across the full length of the airfoil <NUM> at the root <NUM> and the thickness of the base layer <NUM> may increase at about ten percent (<NUM>%) to about twenty percent (<NUM>%) of the span <NUM>. That is, the thermal barrier coating <NUM> may transition from the minimum thickness of the base layer <NUM> in the aft zone <NUM> to one of the intermediate zones <NUM> or <NUM> at about five percent (<NUM>%) of the span <NUM> or greater, such as at about ten percent (<NUM>%) of the span <NUM>, such as at about twenty percent (<NUM>%) of the span <NUM>, such as at about thirty percent (<NUM>%) of the span <NUM>, and/or within a zone extending from about five percent (<NUM>%) to about thirty percent (<NUM>%) of the span <NUM>. In various embodiments, there may be a second spanwise increase in the thickness of the base layer <NUM>, e.g., a transition from the second intermediate zone <NUM> to the first intermediate zone <NUM>, which may occur at about ten percent (<NUM>%) of the span <NUM> or greater, such as at about twenty percent (<NUM>%) of the span <NUM>, such as at about thirty percent (<NUM>%) of the span <NUM>, such as about forty percent (<NUM>%) of the span <NUM>, and/or within a zone extending from about ten percent (<NUM>%) of the span <NUM> to about forty percent (<NUM>%) of the span <NUM>.

In some embodiments, the thickness of the base layer <NUM> at the leading edge <NUM> may vary across the span <NUM> of the airfoil <NUM>, while also being constant across the span <NUM> at other portions of the airfoil <NUM>, such as at the trailing edge <NUM>. In such embodiments, the maximum thickness of the base layer <NUM> may be defined at the leading edge <NUM>, and in particular at about a mid-span point (e.g., about fifty percent span) on the leading edge <NUM>. For example, the maximum thickness of the base layer <NUM> may be provided from about forty percent of the span <NUM> outward, e.g., starting at forty percent span and continuing to one hundred percent span, such as from about fifty percent span outward, or such as from about sixty percent span outward.

Moreover, it should be understood that additional embodiments may include more or fewer transitions in one or both of the span direction and the flow direction, e.g., along the length of the airfoil <NUM>. For example, some embodiments may include only one intermediate zone, or three intermediate zones, four intermediate zones, or more. The transition or transitions in thickness of the base layer <NUM> are generally gradual and tapering such that the lines indicating the various zones in <FIG> are to be understood as for illustrative purposes only and not as sharp boundaries between areas of varying thicknesses of the base layer <NUM>. Additionally, some embodiments may include a continuous variation of the thickness of the base layer <NUM>, such that there are effectively infinite or constant transitions in the thickness of the base layer <NUM>. For example, when the thermal barrier coating <NUM> is continuous around the airfoil <NUM>, e.g., as illustrated in <FIG>, the varying thickness of the base layer <NUM> may thus form an effectively infinite loop around the perimeter of the airfoil <NUM>.

In at least some embodiments, the total thickness of the thermal barrier coating <NUM> may be the same across a majority of the airfoil <NUM>, such as at least from the leading edge <NUM> to at least about seventy-five percent (<NUM>%) of the length of the airfoil <NUM>, e.g., where the length of the airfoil <NUM> is defined along the flow direction. For example, the overall thickness of the thermal barrier coating <NUM> may be constant over at least about eighty-five percent (<NUM>%) of the length of the airfoil <NUM>, such as at least about ninety percent (<NUM>%) of the length of the airfoil <NUM>, such as at least about ninety five percent (<NUM>%) of the length of the airfoil <NUM>, such as about ninety eight percent (<NUM>%) of the length of the airfoil <NUM> or more. In some embodiments, the total thickness of the thermal barrier coating <NUM> may be constant over the entire length of the airfoil <NUM> from the leading edge <NUM> to the trailing edge <NUM>. Providing a uniform, constant overall thickness to the thermal barrier coating <NUM> across all or most of the length of the airfoil <NUM> may advantageously improve the balance, e.g., mass distribution, of the airfoil <NUM>, while also allowing improved resistance to physical impacts and/or erosion at the leading edge <NUM> and/or leading edge portion with the relatively thicker base layer <NUM>, as described above.

Claim 1:
An airfoil (<NUM>) for a turbomachine, the airfoil (<NUM>) comprising:
a root (<NUM>);
a tip (<NUM>) spaced radially outward from the root (<NUM>), the root (<NUM>) and the tip (<NUM>) defining a span (<NUM>) of the airfoil (<NUM>) therebetween;
a leading edge (<NUM>) extending across the span (<NUM>) of the airfoil (<NUM>) from the root (<NUM>) to the tip (<NUM>);
a trailing edge (<NUM>) downstream of the leading edge (<NUM>) along a flow direction, the trailing edge (<NUM>) extending across the span (<NUM>) of the airfoil (<NUM>) from the root (<NUM>) to the tip (<NUM>);
a pressure side surface (<NUM>) extending between the root (<NUM>) and the tip (<NUM>) and extending between the leading edge (<NUM>) and the trailing edge (<NUM>);
a suction side surface (<NUM>) extending between the root (<NUM>) and the tip (<NUM>) and extending between the leading edge (<NUM>) and the trailing edge (<NUM>), the suction side surface (<NUM>) opposing the pressure side surface (<NUM>); and
a thermal barrier coating (<NUM>) on the pressure side surface (<NUM>) and the suction side surface (<NUM>), the thermal barrier coating (<NUM>) comprising a base layer (<NUM>) and a top coat (<NUM>), wherein a thickness of the base layer (<NUM>) varies across each of the pressure side surface (<NUM>) and the suction side surface (<NUM>) with a maximum thickness of the base layer (<NUM>) at the leading edge (<NUM>),
characterized in that the airfoil comprises a leading edge portion (<NUM>), an aft zone (<NUM>) and an intermediate zone (<NUM>) between the leading edge portion and the aft zone,
wherein the thickness of the base layer (<NUM>) is at a maximum in the leading edge portion (<NUM>), the thickness of the base layer (<NUM>) in the intermediate zone (<NUM>) is less than the thickness of the base layer (<NUM>) in the leading edge portion (<NUM>), and wherein the thickness of the base layer (<NUM>) is at a minimum at the trailing edge (<NUM>) and in the aft zone (<NUM>).