Patent Description:
One type of aircraft engine is a gas turbine engine that include various sections that are subject to high temperatures. Ensuring cooling of such components and sections is a goal of gas turbine engine systems. Cooling of cavities within such engines is important to ensure part life and efficient operation. However, using secondary flows of cooling air from a common source to cool multiple cavities and then recombining such flows in a common sink typically requires metering locations along such flow paths. These metering locations result in pressure drops. The pressure drops can result in insufficient cooling or may require boosted pressure at certain locations to ensure that downstream maintains sufficiently high pressure. As such, losses may be realized.

<CIT> discloses a cooling circuit for a gas turbine engine including a plurality of rotor discs each defining a cooling passage. A portion of air is extracted from the air passing through the compressor section of the engine and is directed to the cooling passages in order to cool the rotor discs and the rotor blades attached thereto.

According to a first aspect of the invention, a gas turbine engine according to claim <NUM> is provided.

Optionally, the at least one entrance flow path has a first cross-sectional area and the at least one exit aperture has a second cross-sectional area, wherein the first cross-sectional area is less than the second cross-sectional area.

Optionally, the at least one entrance flow path has a first cross-sectional area and the at least one exit aperture has a second cross-sectional area, wherein the second cross-sectional area is <NUM>%-<NUM>% of the second cross-sectional area.

Optionally, the at least one entrance flow path has a first cross-sectional area and the at least one exit aperture has a second cross-sectional area, wherein the second cross-sectional area is <NUM>%-<NUM>% of the first cross-sectional area.

Optionally, the first disk and the second disk are part of one of a compressor section and a turbine section of the gas turbine engine.

Optionally, the axial portion of the rotor arm is angled radially outward in a direction from the first end to the second end with respect to the engine axis.

Optionally, the at least one entrance flow path is defined by a plurality of slots formed in the rotor arm and the at least one exit flow aperture is a hole formed in the rotor arm.

Optionally, the first disk is configured to be rotationally driven by a shaft to rotationally drive rotation of the rotor arm and the second disk.

Optionally, the at least one entrance flow path is positioned upstream relative to the at least one exit flow aperture with respect to a direction of flow through the cavity.

The foregoing features and elements may be executed or utilized in various combinations without exclusivity, unless expressly indicated otherwise.

The exemplary gas turbine engine <NUM> is a two-spool turbofan engine that generally incorporates a fan section <NUM>, a compressor section <NUM>, a combustor section <NUM>, and a turbine section <NUM>. The fan section <NUM> drives air along a bypass flow path B, while the compressor section <NUM> drives air along a core flow path C for compression and communication into the combustor section <NUM>. Hot combustion gases generated in the combustor section <NUM> are expanded through the turbine section <NUM>. Although depicted as a turbofan gas turbine engine in the disclosed non-limiting embodiment, it should be understood that the concepts described herein are not limited to turbofan engines and these teachings could extend to other types of engines.

The gas turbine engine <NUM> generally includes a low-speed spool <NUM> and a high-speed spool <NUM> mounted for rotation about an engine centerline longitudinal axis A. The low-speed spool <NUM> and the high-speed spool <NUM> may be mounted relative to an engine static structure <NUM> via several bearing systems <NUM>. It should be understood that other bearing systems <NUM> may alternatively or additionally be provided.

The low-speed spool <NUM> generally includes an inner shaft <NUM> that interconnects a fan <NUM>, a low-pressure compressor <NUM> and a low-pressure turbine <NUM>. The inner shaft <NUM> can be connected to the fan <NUM> through 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 compressor <NUM> and a high-pressure turbine <NUM>. In this embodiment, the inner shaft <NUM> and the outer shaft <NUM> are supported at various axial locations by bearing systems <NUM> positioned within the engine static structure <NUM>.

A combustor <NUM> is arranged between the high-pressure compressor <NUM> and the high-pressure turbine <NUM>. A mid-turbine frame <NUM> may be arranged generally between the high-pressure turbine <NUM> and the low-pressure turbine <NUM>. The mid-turbine frame <NUM> can support one or more bearing systems <NUM> of the turbine section <NUM>. The mid-turbine frame <NUM> may include one or more airfoils <NUM> that extend within the core flow path C.

The inner shaft <NUM> and the outer shaft <NUM> are concentric and rotate via the bearing systems <NUM> about the engine centerline longitudinal axis A, which is co-linear with their longitudinal axes. The core airflow is compressed by the low-pressure compressor <NUM> and the high-pressure compressor <NUM>, is mixed with fuel and burned in the combustor <NUM>, and is then expanded over the high-pressure turbine <NUM> and the low-pressure turbine <NUM>. The high-pressure turbine <NUM> and the low-pressure turbine <NUM> rotationally drive the respective high-speed spool <NUM> and the low-speed spool <NUM> in response to the expansion.

Each of the compressor section <NUM> and the turbine section <NUM> may include alternating rows of rotor assemblies and vane assemblies (shown schematically) that carry airfoils that extend into the core flow path C. For example, the rotor assemblies can carry a plurality of rotating blades <NUM>, while each vane assembly can carry a plurality of vanes <NUM> that extend into the core flow path C. The blades <NUM> of the rotor assemblies add or extract energy from the core airflow that is communicated through the gas turbine engine <NUM> along the core flow path C. The vanes <NUM> of the vane assemblies direct the core airflow to the blades <NUM> to either add or extract energy.

Various components of a gas turbine engine <NUM>, including but not limited to the airfoils of the blades <NUM> and the vanes <NUM> of the compressor section <NUM> and the turbine section <NUM>, may be subjected to repetitive thermal cycling under widely ranging temperatures and pressures. The hardware of the turbine section <NUM> is particularly subjected to relatively extreme operating conditions. Therefore, some components may require internal cooling circuits for cooling the parts during engine operation. Example cooling circuits that include features such as airflow bleed ports are discussed below.

Although a specific architecture for a gas turbine engine is depicted in the disclosed non-limiting example embodiment, it should be understood that the concepts described herein are not limited to use with the shown and described configuration. For example, the teachings provided herein may be applied to other types of engines. Some such example alternative engines may include, without limitation, turbojets, turboshafts, and other turbofan configurations (e.g., wherein an intermediate spool includes an intermediate pressure compressor ("IPC") between a low-pressure compressor ("LPC") and a high-pressure compressor ("HPC"), and an intermediate pressure turbine ("IPT") between the high-pressure turbine ("HPT") and the low-pressure turbine ("LPT")).

Turning now to <FIG>, a schematic illustration of a portion of a gas turbine engine <NUM> is shown. The illustration shown in <FIG> shows a portion of a compressor section, such as a high pressure compressor. It will be appreciated that the illustration of <FIG> may also represent a portion of a low pressure compressor, high pressure turbine, low pressure turbine, or other section of an aircraft engine, without departing from the scope of the present invention. The gas turbine engine <NUM> includes a rotor shaft <NUM>, a tie shaft <NUM>, a first disk <NUM>, a second disk <NUM>, and a rotor arm <NUM> connecting the first disk <NUM> and the second disk <NUM>. The rotor shaft <NUM> and the tie shaft <NUM> are configured to be rotationally driven by a turbine section (not shown) of the gas turbine engine <NUM>.

The first disk <NUM> includes one or more blades <NUM> connected thereto. It will be appreciated that the blades <NUM> may be fixedly attached or integrally formed with the first disk <NUM>. The second disk <NUM> may also include one or more blades (not shown), as will be appreciated by those of skill in the art. As shown, the second disk <NUM> is operably coupled to the rotor shaft <NUM>, and thus the second disk <NUM> may be rotationally driven by the rotor shaft <NUM>. The first disk <NUM> is coupled to the second disk <NUM> by the rotor arm <NUM> such that the first disk <NUM> may be rotationally driven by the second disk <NUM>. It will be appreciated that other connection mechanisms and arrangements may be possible. For example, the first disk <NUM> may be coupled to one or more of the shafts of the gas turbine engine to be rotationally driven thereby. Each of the disks <NUM>, <NUM> may include blades <NUM> that are rotationally driven by the respective disks <NUM>, <NUM>. The blades <NUM> may be arranged relative to vanes <NUM>, as will be appreciated by those of skill in the art.

Cooling may be provided throughout the gas turbine engine <NUM> to cool components thereof. For example, it may be desirable to cool cavities of the gas turbine engine <NUM>. As shown, the gas turbine engine <NUM>, in <FIG>, forms a plurality of cavities 216a-f. The cavities 216a-f are defined between components of the gas turbine engine <NUM>. Certain cavities (e.g., cavity 216e) may be relatively isolated due to structures within the gas turbine engine <NUM>. As shown, one cavity 216e is bounded by the second disk <NUM> and the rotor arm <NUM> (and, potentially, a portion of the first disk <NUM>). As such, providing cooling air therein may be difficult. Further, if air is supplied into such cavity 216e, removing the air from the cavity 216e and forming a cooling flow therethrough may also be difficult.

Turning now to <FIG>, schematic illustrations of a portion of a gas turbine engine <NUM> are shown. The illustrations shown in <FIG> show a portion of a compressor section, such as a high pressure compressor. It will be appreciated that the illustrations of <FIG> may also represent a portion of a low pressure compressor, high pressure turbine, low pressure turbine, or other section of an aircraft engine, without departing from the scope of the present invention as defined by the appended claims. The gas turbine engine <NUM> includes a rotor shaft <NUM>, a tie shaft <NUM>, a first disk <NUM>, a second disk <NUM>, and a rotor arm <NUM> connecting the first disk <NUM> and the second disk <NUM>. The rotor shaft <NUM> and the tie shaft <NUM> are configured to be rotationally driven by a turbine section (not shown) of the gas turbine engine <NUM>. Similar to that shown and described above, the first disk <NUM> and second disk <NUM> may include one or more blades <NUM> arranged between vanes <NUM> of the gas turbine engine <NUM>.

To enable a cooling of components of the gas turbine engine <NUM>, one or more flow apertures may be formed in various components of the gas turbine engine <NUM> to form a cooling flow path. For example, the tie shaft <NUM> can include a flow aperture <NUM> configured to enable fluid flow (e.g., cooling air) from a first cavity 316a to a second cavity 316b, which in turn will flow axially forward (to the left in <FIG>) to a third cavity 316c and a fourth cavity 316d. According to the invention, the cooling flow is a forward flowing cooling flow (i.e., from aft to forward of the engine along an engine axis). A fifth cavity 316e may be supplied from other flow paths or cooling schemes, as will be appreciated by those of skill in the art. A primary cooling flow <NUM> is shown in <FIG>.

The flow aperture <NUM>, being significantly smaller in area than the cavities it connects, retards the cooling flow by causing a pressure drop therethrough. This pressure drop dictates that the second cavity 316b must operate at a lower pressure than the first cavity 316a, while serving to prevent or minimize any backflow from the second cavity 316b to the first cavity 316a. It then follows that the cooling flow feeding the first cavity 316a must come from a higher-pressure source than the sink to which the flow is routed (i.e., second, third, and fourth cavities 316b, 316c, 316d). It is noted that because the second, third, and fourth cavities 316b, 316c, 316d do not have any flow-retarding aperture(s) between them, the difference in operating pressures between these three cavities may be negligible.

However, a sixth cavity 316f in this embodiment, may be remote from any convenient or potential aperture(s) that could connect it from and to two cavities of different pressure, which would deprive it of cooling flow. As shown, the sixth cavity 316f is bounded by the first disk <NUM>, the second disk <NUM>, and the rotor arm <NUM>. To provide a cooling flow to the sixth cavity 316f, a cooling flow source is required. In this illustrative configuration, the preferred source for cooling flow is the primary cooling flow <NUM>. Splitting a portion of the primary cooling flow <NUM> to cool the sixth cavity 316f may result in pressure drops that can be detrimental for overall cooling effectiveness. For example, splitting secondary flow from a common source (primary cooling flow <NUM>) to cool multiple cavities (sixth cavity 316f) and then recombining in a common sink (back to primary cooling flow <NUM> in second cavity 316b) typically involves metering locations in each leg and measurable pressure drop. This pressure drop and/or change in pressure can detrimental to the achievable flow of the system. The desired flow direction through the sixth cavity 316f shown in <FIG> may not be achievable with metering apertures because the source and sink pressures (third cavity 316c) are the same/common.

Accordingly, embodiments of the present invention are directed to a cooling flow scheme that enables relatively low pressure cooling flow from the primary cooling flow <NUM> to be diverted into the sixth cavity 316f from the third cavity 316c and then causing the flow to pass through the sixth cavity 316f and back into the third cavity 316c to be recombined with the primary cooling flow <NUM>, as shown in <FIG>. To achieve this, an entrance flow path <NUM> and an exit flow aperture <NUM> may be formed in the structure of the rotor arm <NUM> to ensure a secondary cooling flow <NUM> to be generated and supply cool air into the sixth cavity 316f. Thus, the air in the sixth cavity 316f will have a higher pressure than the air in the cooling flow source cavity (e.g., third cavity 316c). As such, the cooling flow will require assistance to be directed radially outward and then a solution for having the cooling flow recombine with the primary cooling flow <NUM> is required.

In accordance with embodiments of the present invention, the rotor arm <NUM> is modified to define or form a pressure gain feature. As the rotor arm <NUM> is rotated, the rotor arm <NUM> and radially distending the features of the entrance flow path <NUM> will operate as a centrifugal pump to increase the pressure of the air traversing through the entrance flow path <NUM> by raising the angular momentum of the air as it moves outboard through the rotating passages, as described herein. This high pressure air will then cool within the sixth cavity 316f. The cooling air will then exit the sixth cavity 316f through the exit flow aperture <NUM> due to the air pressure being relatively high within the sixth cavity 316f as compared to the pressure within the third cavity 316c, in which the air will rejoin the primary cooling flow <NUM>. This flow path is referred to as the secondary cooling flow <NUM> and is illustratively shown in <FIG>. The result is a leg of the split primary cooling flow <NUM> is routed through a pressure gain feature such as slot(s) in a rotating flange joint or tube(s) attached to the rotor arm <NUM> to flow from radially inward to radially outward. This will centrifugally pump up the pressure of the air so that it can flow through the sixth cavity 316f and be recombined with the primary cooling flow <NUM> at or near their common source pressure within the third cavity 316c.

The features of each of the entrance flow path(s) <NUM> and the exit flow aperture(s) <NUM> may be formed within the material of the rotor arm <NUM>. For example, the channels, apertures, holes, slots, and other flow path features may be machined, cast, or otherwise formed within the rotor arm <NUM> during or after manufacture of the rotor arm <NUM>. In some configurations, the entrance flow path <NUM> may be formed by one or more flow channels or flow slots. Multiple entrance flow paths <NUM> and exit flow apertures <NUM> may be formed at different circumferential positions about the rotor arm <NUM> to provide cooling about an entire circumference (at a given radial position) of the gas turbine engine <NUM>. The number, shape, size, and length of the entrance flow paths <NUM> and exit flow apertures <NUM> may be selected to achieve a desired cooling of the sixth cavity 316f and the material of the components that define such sixth cavity 316f.

Turning now to <FIG>, a schematic illustration of a portion of a gas turbine engine <NUM> is shown. The illustration shown in <FIG> is an enlarged illustration having a configuration substantially similar to that shown and described with respect to <FIG>, and thus description of similar features may be omitted herein. The illustration of <FIG> illustrates a first disk <NUM>, a second disk <NUM>, and a rotor arm <NUM> connecting the first disk <NUM> and the second disk <NUM>.

The rotor arm <NUM> includes one or more entrance flow paths <NUM> and one or more exit flow apertures <NUM>, as described herein. The entrance flow paths <NUM> are radial flow paths that extend substantially radially through a portion of the rotor arm <NUM>. For example, as shown, the rotor arm <NUM> includes a radial portion <NUM> and an axial portion <NUM>. In this configuration, the radial portion <NUM> is at an aft end of the rotor arm <NUM>, when installed in the gas turbine engine <NUM> and connects to the first disk <NUM>. The axial portion <NUM>, in this embodiment, extends axially forward from the axial portion <NUM> to connect with the second disk <NUM>. The radial portion <NUM> extends radially from an inner diameter end <NUM> to an outer diameter end <NUM>. As shown, the axial portion <NUM> does not extend parallel to an engine axis but rather is angled radially outward from a first end <NUM> to a second end <NUM>. In this embodiment, the first end <NUM> is aft of the second end <NUM>. The outer diameter end <NUM> of the radial portion <NUM> joins with or defines the intersection with the first end <NUM> of the axial portion <NUM>.

In operation without a pressure gain device, the air pressure at the inner diameter end <NUM> is not greater than the air pressure at the outer diameter end <NUM>. As such, without aid, the cooling flow will not travel outboard. However, an outboard radial air flow in slots (or holes) through the entrance flow paths <NUM> increases total pressure, with such entrance flow paths <NUM> operating like a centrifugal compressor stage, by raising angular momentum. This causes the air at the inner diameter end <NUM> to flow through the entrance flow paths <NUM> to a cavity <NUM>. The geometry, number, and length of the entrance flow paths <NUM> combined with the rotation of the rotor arm <NUM> causes a pressure differential to operate as a centrifugal pump to direct a portion of a primary cooling flow <NUM> into a secondary cooling flow <NUM> and into the cavity <NUM>.

The air within the cavity <NUM> will then exit the cavity <NUM> through one or more exit flow apertures <NUM>. The flow exits the cavity <NUM> through the exit flow apertures <NUM> which is driven by a pressure differential created in the exit flow apertures. That is, a relatively high pressure in cavity <NUM> and relatively low pressure along the primary cooling flow <NUM> causes the air within the cavity <NUM> to flow through the exit flow apertures <NUM> back to the primary cooling flow <NUM>. The exit flow apertures <NUM> may be sized larger than the inlet slots or holes of the entrance flow path <NUM> (e.g., <NUM>%-<NUM>% of inlet slot area) to not significantly restrict the flow, while achieving high enough velocity to prevent reverse flow.

Turning now to <FIG>, schematic illustrations of a portion of a gas turbine engine <NUM> having a rotor arm <NUM> in accordance with an embodiment of the present invention are shown. <FIG> illustrates the rotor arm <NUM> as installed in the gas turbine engine <NUM> and <FIG> illustrates a perspective illustration of the rotor arm <NUM>. The gas turbine engine of <FIG> is an enlarged illustration having a configuration substantially similar to that shown and described above, and thus description of similar features may be omitted herein. The gas turbine engine <NUM> includes a first disk <NUM>, a second disk <NUM>, and the rotor arm <NUM> connecting the first disk <NUM> and the second disk <NUM>.

The components <NUM>, <NUM>, <NUM> of the gas turbine engine <NUM> are annular or circumferential structures arranged about an engine axis <NUM>. The components <NUM>, <NUM>, <NUM> may be rotationally driven by one or more shafts, as will be appreciated by those of skill in the art. The rotor arm <NUM> has a substantially conical or annular shape, as shown in <FIG>. The rotor arm <NUM>, when installed between and connecting the first disk <NUM> and the second disk <NUM>, defines a cavity <NUM> which may require cooling air during operation. The rotor arm <NUM> thus include one or more entrance flow paths <NUM> to supply cooling air into the cavity <NUM> and one or more exit flow apertures <NUM> to expel such air back to a primary cooling flow, as described above.

As shown in <FIG>, the rotor arm <NUM> has a radial portion <NUM> and an axial portion <NUM>. The radial portion <NUM> extends substantially radially from an inner diameter end <NUM> to an outer diameter end <NUM> of the radial portion <NUM>. The radial potion <NUM> joins with or connects to the axial portion <NUM> at the outer diameter end <NUM>. The axial portion <NUM> extends from the radial portion <NUM> at a first end <NUM> to a second end <NUM>. In this embodiment, the first end <NUM> is aft of the second end <NUM>, although the opposite may be true for other engine configurations and depending on a direction of a primary cooling flow. The entrance flow paths <NUM> are formed within and define channels within the radial portion <NUM> and have apertures or openings formed in the axial portion <NUM> at the first end of the axial portion <NUM>. The exit flow apertures <NUM> are arranged proximate the second end <NUM> of the axial portion <NUM>.

The axial flow path portion of the entrance flow paths <NUM> may be defined by conduits, channels, or slots. The cross-sectional area (in a flow direction through the feature) of such entrance flow paths <NUM> may be selected to ensure a desired air flow and/or pressure differential to be generated during rotation of the rotor arm <NUM> within the gas turbine engine <NUM>. At a component level, for example and in accordance with some embodiments, the size and shape of the entrance flow paths <NUM> may be selected such that between <NUM>%-<NUM>% of a primary cooling flow may be extracted and conveyed into the cavity <NUM>. The exit flow apertures <NUM> may be selected to be larger in cross-sectional area (in a flow direction through the feature) than the cross-sectional area of the entrance flow paths <NUM>. For example, in some non-limiting embodiments, the cross-sectional area of the exit flow apertures <NUM> may be <NUM>%-<NUM>% larger than the cross-sectional area of the entrance flow paths <NUM>.

Although shown and described with a forward-flowing cooling flow (with the first disk aft of the second disk), it will be appreciated that embodiments of the present invention will operate similarly to an aft-flowing cooling flow. Embodiments of the present invention are directed to a component that in-part defines a remote cavity to incorporate flow passages that operate as a pump during rotation to direct relatively cool (and initially relatively low pressure) air into a remote cavity at higher pressure. This is achieved through the entrance flow paths having channels or passages that, when rotated, operate as a centrifugal pump to increase the pressure of the air within the passage and thus cause such air to flow outboard into the cavity. Once in the cavity, exit flow apertures provide a low pressure differential path for the flow to rejoin the primary cooling flow.

Advantageously, embodiments described herein provide for improved mechanisms for cooling cavities within gas turbine engines. Specifically, for example, embodiments of the present invention may provide for cooling of otherwise dead air cavities without adding pressure drops within the system. Further, advantageously, embodiments of the present invention can result in a net-zero pressure change even while cooling a previously "dead air" cavity. By increasing the angular velocity of the air, the pressure increases, this can result in cooling airflow that flows outboard despite coming from a lower inboard source pressure. By having a radial span within and along the entrance flow paths the work done on the air increases due to increased angular momentum at the greater radial positions. This causes the air to flow into the cavity, and then the air can exit such cavity through one or more exit flow apertures.

Advantageously, embodiments of the present invention provide for a purging flow that is extracted from and a portion of a primary cooling flow within a gas turbine engine. The purging flow will flow through the entrance flow paths due to the angular momentum increase at larger radial distances. The air will enter the cavity and purge through the exit flow apertures to rejoin the primary cooling flow with negligible net pressure drop. The size and number of entrance flow paths may be selected to meter the flow to a portion or fraction of the total available primary cooling flow. The size and number of entrance flow paths may be selected to ensure that the main flow does not reference into a feedback loop. Between <NUM>%-<NUM>% of the total airflow of the primary cooling flow at the location of the entrance flow paths is extracted. Similarly, the size and number of exit apertures may be selected to not significantly restrict the flow while achieving high enough velocity to prevent reverse flow therethrough.

It will be appreciated that in some embodiments of the present invention, the number of entrance flow paths and the number of exit flow apertures may be the same. However, in other embodiments, the number of entrance flow paths may exceed the number of exit flow apertures. In other embodiments, the number of entrance flow paths may be less than the number of exit flow apertures. As such, the specific number and relationship of number of features relative to each other is not to be limiting and may be selected to achieve a desired flow through a remote cavity, as described herein. Further, although illustratively shown with the entrance flow paths and the exit flow apertures being aligned in a circumferential direction, such arrangement is not to be limiting. For example, in some embodiments, the position of the exit flow apertures may be offset in a circumferential direction from the locations of the entrance flow paths on the rotor arm. Accordingly, the above description and accompanying illustrations are provided to be illustrative and for explanatory purposes only, and are not intended to be limiting to the specific configurations and embodiments described herein.

As used herein, the terms "about" and "substantially" are intended to include the degree of error associated with measurement of the particular quantity based upon the equipment available at the time of filing the application. For example, the terms may include a range of ± <NUM>%, or <NUM>%, or <NUM>% of a given value or other percentage change as will be appreciated by those of skill in the art for the particular measurement and/or dimensions referred to herein.

Claim 1:
A gas turbine engine (<NUM>) comprising:
a first disk (<NUM>);
a second disk (<NUM>), wherein the first disk is arranged aft of the second disk along an axis of the gas turbine engine; and
a rotor arm (<NUM>) arranged between and connecting the first disk (<NUM>) to the second disk (<NUM>), wherein a cavity (<NUM>) is defined at least between the rotor arm (<NUM>) and the first disk (<NUM>), and wherein the first disk (<NUM>), the second disk (<NUM>) and the rotor arm (<NUM>) are arranged about an engine axis (<NUM>),
characterised in that the rotor arm comprises:
a radial portion (<NUM>) having an inner diameter end (<NUM>) and an outer diameter end (<NUM>);
an axial portion (<NUM>) having a first end (<NUM>) and a second end (<NUM>), wherein the first end (<NUM>) of the axial portion (<NUM>) is connected to the outer diameter end (<NUM>) of the radial portion (<NUM>);
at least one entrance flow path (<NUM>) defined within the radial portion (<NUM>) extending from the inner diameter end (<NUM>) to the outer diameter end (<NUM>); and
at least one exit aperture (<NUM>) arranged proximate the second end (<NUM>) of the axial portion (<NUM>),
wherein a primary cooling flow (<NUM>) is passed through the engine (<NUM>) as a forward flowing cooling flow from aft to forward along the engine axis (<NUM>),
wherein the at least one entrance flow path (<NUM>) is configured to extract a portion of the primary cooling flow (<NUM>) and direct said extracted flow into the cavity (<NUM>) as a secondary cooling flow (<NUM>),
wherein the secondary cooling flow (<NUM>) is directed back to the primary cooling flow (<NUM>) through the at least one exit aperture (<NUM>), and
wherein the portion of the primary cooling flow (<NUM>) that is directed to the cavity (<NUM>) is <NUM>- <NUM>% of a total flow of the primary cooling flow (<NUM>).