Patent Description:
Some systems, such as unmanned aerial vehicles (UAV's) or the like often utilize electrical power for propulsion and operation of onboard systems. Some such systems, such as medium-sized UAV's that require power levels in the range of about <NUM> KW to <NUM> KW, have relatively short mission times because the energy density of batteries is far too low to effectively work in this power range, and conventional internal combustion engines and jet engines are very inefficient at these low power levels. One option that has been developed is a tethered UAV system in which the UAV is connected to a power source on the ground by a tether. Use of a tethered UAV allows for an increase in mission duration time, but reduces an operating height and distance in which the UAV may operate, due to the constraint of the tether. An untethered efficient power source that is lightweight with a high power density is greatly desired. <CIT> discloses a compressor impeller made with an additive manufacturing technique.

According to one embodiment, a compressor assembly is provided. The compressor assembly includes a compressor including a central shaft including an external surface. The compressor includes a shroud extending circumferentially around the central shaft. The shroud including a radially inward surface and a radially outward surface located opposite the radially inward surface. The external surface of the central shaft and the radially inward surface of the shroud are in a facing spaced relationship forming a core flow path therebetween. The compressor also includes a plurality of blades extending from the central shaft to the shroud. The compressor assembly also includes a compression ring extending circumferentially around the shroud, the compression ring being in an interference fit with the shroud. The compression ring is configured to apply a radially inward compressive force along one or more portions of the radially outward surface of the shroud, the radially inward compressive is configured to compress the shroud and the plurality of blades into the central shaft.

The compressor is formed via an additive manufacturing technique.

In addition to one or more of the features described above, or as an alternative, further embodiments may include that the compression ring is formed via subtractive machining.

In addition to one or more of the features described above, or as an alternative, further embodiments may include that the compressor has a first tensile strength. The compression ring has a second tensile strength that is greater than the first tensile strength.

In addition to one or more of the features described above, or as an alternative, further embodiments may include that the one or more portions include a first portion and a second portion located a first distance away from the first portion.

In addition to one or more of the features described above, or as an alternative, further embodiments may include that the radially outward surface of the shroud has a first outer diameter along the first portion of the shroud. The radially outward surface of the shroud has a second outer diameter along the second portion of the shroud. The second outer diameter being greater than the first outer diameter.

According to another embodiment, an electrical power generation system is provided. The electrical power generation system includes a micro-turbine alternator. The micro-turbine alternator includes a combustion chamber, at least one turbine driven by combustion gases from the combustion chamber, a compressor operably connected to the combustion chamber to provide a compressed airflow thereto, one or more shafts connecting the at least one turbine to the compressor such that rotation of the at least one turbine drives rotation of the first stage compressor and the second stage compressor, and an electric generator disposed along the one or more shafts such that electrical power is generated via rotation of the one or more shafts. The compressor includes a central shaft including an external surface and a shroud extending circumferentially around the central shaft. The shroud including a radially inward surface and a radially outward surface located opposite the radially inward surface. The external surface of the central shaft and the radially inward surface of the shroud are in a facing spaced relationship forming a core flow path therebetween. The compressor includes a plurality of blades extending from the central shaft to the shroud and a compression ring extending circumferentially around the shroud. The compression ring being in an interference fit with the shroud. The compression ring is configured to apply a radially inward compressive force along one or more portions of the radially outward surface of the shroud. The radially inward compressive is configured to compress the shroud and the plurality of blades into the central shaft.

According to another embodiment, a method of manufacturing a compressor assembly is provided. The method including: reducing a temperature of a compressor, the compressor including: a central shaft including an external surface; a shroud extending circumferentially around the central shaft, the shroud including a radially inward surface and a radially outward surface located opposite the radially inward surface. The external surface of the central shaft and the radially inward surface of the shroud are in a facing spaced relationship forming a core flow path therebetween. The compressor including a plurality of blades extending from the central shaft to the shroud. The method also including increasing a temperature of a compression ring; and sliding the compression ring onto the shroud of the compressor such that the compression ring extends circumferentially around the shroud. The compression ring being in an interference fit with the shroud once the temperature of compression ring and the temperature of the compressor reach an equilibrium. The compression ring is configured to apply a radially inward compressive force along one or more portions of the radially outward surface of the shroud. The radially inward compressive is configured to compress the shroud and the plurality of blades into the central shaft.

In addition to one or more of the features described above, or as an alternative, further embodiments may include forming the compression ring via subtractive machining.

In addition to one or more of the features described above, or as an alternative, further embodiments may include that the compressor has a first tensile strength, and wherein the compression ring has a second tensile strength that is greater than the first tensile strength.

In addition to one or more of the features described above, or as an alternative, further embodiments may include that the radially outward surface of the shroud has a first outer diameter along the first portion of the shroud, and wherein the radially outward surface of the shroud has a second outer diameter along the second portion of the shroud, the second outer diameter being greater than the first outer diameter.

As previously noted, an untethered, lightweight, high power density power source would allow systems like UAVs to have longer mission times without the height and distance limits of a tether. An approach to power generation involves micro-turbine alternator design utilizing an electric generator in combination with a compressor, turbine, and combustion chamber. The efficiency of the micro-turbine alternator is often largely dependent on the compressor design.

Some compressor designs utilize an open impeller. An open impeller may be defined as an impeller without a shroud. Open impellers may often be used to facilitate the manufacturing process, as it allows the impeller to be machined with standard cutting tools. A centrifugal compressor's efficiency is highly dependent on the tip leakage flow rate. The compressor's tip leakage flow rate is dependent on the distance between the impeller blade tips and the inside of the stationary housing. This is referred to as the tip gap. One method to eliminate a compressor's tip leakage is to build the impeller with a shroud, which may be referred to as a shrouded impeller. According to the invention, the shrouded impeller is built using an additive manufacturing method, such as, for example, 3D printing. Shrouded impellers built through additive manufacturing have a lower tensile strength than shrouded impellers built through a subtractive machining method. As the shrouded impeller rotates, the shrouded impeller experiences a large centrifugal force that causes the shrouded impeller to try to expand in the radial direction, which applies tensile stress to the impeller blades. This tensile stress that is applied to impeller blades may exceed the material strength of the additive manufactured material. Embodiments disclosed herein seek to significantly reduce the operational impeller blade stresses associated with implementing a shroud on a highspeed centrifugal compressor by installing a stress relieving compressor shroud compression ring over the outside of the compressor impeller encircling the impeller blades.

Referring to <FIG>, an isometric view of an unmanned aerial vehicle (UAV) <NUM> is illustrated in accordance with an embodiment of the present disclosure. The UAV <NUM> includes a propulsion/lift system <NUM>, for example a plurality of lift rotors <NUM>, operably connected to an electrical power generation system <NUM>, which includes a micro-turbine alternator system <NUM>. In an embodiment, the micro-turbine alternator system <NUM> is a high efficiency Brayton cycle micro-turbine alternator. The UAV <NUM> includes a propulsion system having electric motors <NUM> and lift rotors <NUM> associated with each electric motor <NUM>. Each lift rotor <NUM> is operably connected to the electric motor <NUM> that is configured to rotate the lift rotor <NUM> using electrical power generated by the micro-turbine alternator system <NUM> of the electrical power generation system <NUM>. The micro-turbine alternator system <NUM> is configured to convert fuel to electrical power to power at least the electric motors <NUM> of the lift rotors <NUM>. The fuel is provided from one or more fuel storage tanks <NUM> operably connected to the micro-turbine alternator system <NUM>. In some embodiments, the fuel utilized is JP-<NUM>. The micro-turbine alternator system <NUM> may utilize compressed air provided from a compressed air tank <NUM> at <NUM> psig (31MPa) and regulated to about <NUM> psig (<NUM> MPa). The compressed air from the compressed air tank <NUM> of <FIG> may be utilized to provide the motive pressure required to drive the liquid fuel through a turbine speed control valve (not shown) and into a combustion chamber. Alternatively, an electric driven pump may be used in place of the compressed air.

Referring now to <FIG>, with continued reference to <FIG>, an isometric view of an electrically-powered suit <NUM> is illustrated in accordance with an embodiment of the present disclosure. While in <FIG>, the micro-turbine alternator system <NUM> is described as utilized in a UAV <NUM>, the micro-turbine alternator system <NUM> disclosed herein may be readily applied to other systems, and may be utilized in, for example, an electrically-powered suit <NUM>, as shown in <FIG>.

The electrically-powered suit <NUM> is operably connected to an electrical power generation system <NUM>, which includes a micro-turbine alternator system <NUM>. The micro-turbine alternator system <NUM> is configured to convert fuel to electrical power to power the electrically-powered suit <NUM>. The fuel is provided from one or more fuel storage tanks <NUM> operably connected to the micro-turbine alternator system <NUM>. In some embodiments, the fuel utilized is JP-<NUM>. The fuel storage tanks <NUM> may be located on legs of the electrically-powered suit <NUM>, as illustrated in <FIG>.

It is understood that the micro-turbine alternator system <NUM> is not limited to a UAV <NUM> and an electrically-powered suit <NUM> application, and the micro-turbine alternator system <NUM> may be applied to other systems not disclosed herein.

Referring now to <FIG>, an isometric cut-away view of the micro-turbine alternator system <NUM> is illustrated, in accordance with an embodiment of the present disclosure. The micro-turbine alternator system <NUM> includes a first stage compressor <NUM>, a second stage compressor <NUM>, a third stage compressor <NUM>, a first stage turbine <NUM>, and a second stage turbine <NUM>. The first stage compressor <NUM>, the second stage compressor <NUM>, the third stage compressor <NUM>, the first stage turbine <NUM>, and the second stage turbine <NUM> are oriented along a central longitudinal axis A of the micro-turbine alternator system <NUM>. The micro-turbine alternator system <NUM> also includes an electric generator <NUM> located between the first stage compressor <NUM> and the second stage compressor <NUM> as measured along the central longitudinal axis A.

Advantageously, by locating the electric generator <NUM> between the first stage compressor <NUM> and the second stage compressor <NUM>, the overall physical size of the micro-turbine alternator system <NUM> is reduced. As a result, the micro-turbine alternator system <NUM> according to one or more embodiments may be used in a UAV <NUM>, an electrically-powered suit <NUM>, or another system that benefits from untethered, lightweight power generation.

The micro-turbine alternator system <NUM> also includes an alternator stator cooling heat exchanger <NUM> configured to utilize airflow from the first stage compressor <NUM> to cool the electric generator <NUM>. The alternator stator cooling heat exchanger <NUM> may encircle or enclose the electric generator <NUM> and may be configured to pass airflow from the first stage compressor <NUM> through or around the electric generator <NUM>. Advantageously, by locating the electric generator <NUM> between the first stage compressor <NUM> and the second stage compressor <NUM>, moderately cool air in the core flow path C from the first stage compressor <NUM> is forced through the alternator stator cooling heat exchanger <NUM> and heat may be drawn out of the electric generator <NUM> and to the airflow within the alternator stator cooling heat exchanger <NUM>.

The electric generator <NUM> may be a permanent magnet alternator, an induction generator, a switched reluctance generator, a wound field generator, a hybrid generator, or any other type of alternator known to one of skill in the art. As illustrated in <FIG>, the electric generator <NUM> may be a permanent magnet alternator that includes a rotor element <NUM> and a stator element <NUM> radially outward from the rotor element. In other words, the rotor element <NUM> is located radially inward from the stator element <NUM> as measured relative to the central longitudinal axis A. It is understood that the embodiments disclosed herein may be applicable to a rotor element <NUM> that is located radially outward from the stator element <NUM>. The rotor element <NUM> may be rotated around the central longitudinal axis A to generate electricity.

The rotor element <NUM> includes an annular base member <NUM>, an annular array of permanent magnets <NUM> that are respectively coupled to an outer diameter of the annular base member <NUM>. The rotor element <NUM> may include a magnet retention band that fits over an outer diameter of the permanent magnet <NUM>, and keeps the permanent magnet <NUM> on the rotating annular base member <NUM>. In accordance with further embodiments, the stator element <NUM> includes a hub <NUM>, a plurality of spokes <NUM> extending radially inward from the hub <NUM> and conductive elements <NUM> that are wound around the spokes <NUM> to form windings. When the rotor element <NUM> is rotated around the central longitudinal axis A a rotating flux field is generated by the permanent magnets <NUM> and this rotating flux field generates an alternating current in the conductive elements <NUM> to generate electricity for use by the UAV <NUM> of <FIG> or the electrically-powered suit <NUM> of <FIG>.

The micro-turbine alternator system <NUM> includes a combustion chamber <NUM>, in which a fuel-air mixture is combusted, with the combustion products utilized to drive an electric generator <NUM>. In some embodiments, the fuel utilized in the combustion chamber <NUM> is JP-<NUM>. The micro-turbine alternator system <NUM> converts the energy of the combustion products into electrical power by urging the combustion products through the first stage turbine <NUM> and the second stage turbine <NUM>, which are operably connected to and configured to rotate the rotor element <NUM> of the electric generator <NUM>. The electrical energy generated by the electric generator <NUM> may then be rectified via a generator rectifier (not shown) and utilized by the propulsion/lift system <NUM> of <FIG> or the electrically-powered suit <NUM> of <FIG>. The compressed air from the compressed air tank <NUM> of <FIG> may be utilized to provide the motive pressure required to drive the liquid fuel through a turbine speed control valve (not shown) and into the combustion chamber <NUM>.

The first stage compressor <NUM> is located forward of the second stage compressor <NUM> and the third stage compressor <NUM> as measured along the central longitudinal axis A, and the second stage compressor <NUM> is located forward of the third stage compressor <NUM> as measured along the central longitudinal axis A. In other words, the second stage compressor <NUM> is located aft of the first stage compressor <NUM> and the third stage compressor <NUM> is located aft of the second stage compressor <NUM> as measured along the central longitudinal axis A. The forward direction D1 and the aft direction D2 are illustrated in <FIG>. The first stage turbine <NUM> is located forward of the second stage turbine <NUM> as measured along the central longitudinal axis A. In other words, the second stage turbine <NUM> is located aft of the first stage turbine <NUM> as measured along the central longitudinal axis A. The first stage compressor <NUM>, the second stage compressor <NUM>, and the third stage compressor <NUM> are located forward of first stage turbine <NUM> and the second stage turbine <NUM> as measured along the central longitudinal axis A.

The micro-turbine alternator system <NUM> includes a compressor shaft <NUM> oriented along and co-axial to the central longitudinal axis A. In an embodiment, the compressor shaft <NUM> is a tie bolt and is used to compress a rotating group of components including the first stage compressor <NUM>, compressor transfer tube <NUM>, the compressor shaft <NUM>, and a second journal bearing <NUM> in the axial direction, causing the multi-segment shaft to act as a single stiff shaft. The compressor shaft <NUM> may be attached or operably connected to the first stage compressor <NUM>. The micro-turbine alternator system <NUM> includes a turbine shaft <NUM> oriented along and co-axial to the central longitudinal axis A. The turbine shaft <NUM> may be attached or operably connected to the first stage turbine <NUM> and the second stage turbine <NUM>.

The micro-turbine alternator system <NUM> includes a coupling assembly <NUM> configured to operably connect the turbine shaft <NUM> to the compressor shaft <NUM>. The coupling assembly <NUM> may be attached or operably connected to the second stage compressor <NUM>. The compressor shaft <NUM> extends in the aft direction D2 away from the first stage compressor <NUM> and through the electric generator <NUM> to operably connect to the coupling assembly <NUM>. In an embodiment, the compressor shaft <NUM> is located radially inward of the rotor element <NUM>.

Advantageously, locating the electric generator <NUM> between the first stage compressor <NUM> and the second stage compressor <NUM> allows the first stage compressor <NUM> to have a reduced inlet hub diameter that is smaller than a diameter of the rotor element <NUM>. Having a reduced inlet hub diameter DIA1 reduces the inlet flow relative velocity, increasing the aerodynamic performance of the first stage compressor <NUM> and increasing the swallowing capacity of the first stage compressor <NUM>. If the electric generator <NUM> was located forward of the first stage compressor <NUM>, then the compressor shaft <NUM> would have to extend forward of the first stage compressor <NUM> and thus the inlet hub diameter DIA1 would have to be increased to a diameter of the compressor shaft <NUM>, thus decreasing the aerodynamic performance of the first stage compressor <NUM> and decreasing the swallowing capacity of the first stage compressor <NUM>.

The turbine shaft <NUM> extends in the forward direction D1 away from the first stage turbine <NUM> to operably connect to the coupling assembly <NUM>. The turbine shaft <NUM>, the coupling assembly <NUM>, and the compressor shaft <NUM> are configured to rotate in unison. Thus, when exhaust <NUM> from the combustion chamber <NUM> drives rotation of the first stage turbine <NUM> and the second stage turbine <NUM>, the rotation of the first stage turbine <NUM> and the second stage turbine <NUM> drives rotation of the turbine shaft <NUM>, which drives rotation of the coupling assembly <NUM> and the compressor shaft <NUM>. The rotation of the compressor shaft <NUM> drives rotation of the first stage compressor <NUM>. The rotation of the coupling assembly <NUM> drives rotation of the second stage compressor <NUM>. The third stage compressor <NUM> is operably connected to the second stage compressor <NUM> and the turbine shaft <NUM>, and thus rotation of the second stage compressor <NUM> and the turbine shaft <NUM> drives rotation of the third stage compressor <NUM>.

It is understood that while the compressor shaft <NUM>, the turbine shaft <NUM>, and the coupling assembly <NUM> are described as three different shafts, the embodiments disclosed herein may be applicable to micro-turbine alternator system <NUM> having one or more shafts. In an embodiment, the electric generator <NUM> is disposed along the one or more shafts between the first stage compressor <NUM> and the second stage compressor <NUM>. In another embodiment, the electric generator <NUM> is disposed along the compressor shaft <NUM> between the first stage compressor <NUM> and the second stage compressor <NUM>. The electric generator <NUM> is located aft of the first stage compressor <NUM> and forward of the second stage compressor <NUM>. In another embodiment, at least one of the one or more drive shafts passes through the electric generator <NUM>. In another embodiment, the compressor shaft <NUM> passes through the electric generator <NUM>.

The compressor shaft <NUM>, the turbine shaft <NUM>, and the coupling assembly <NUM> are coaxial and rotate via the bearing systems about the central longitudinal axis A, which is colinear with their longitudinal axes. The bearing system includes a first journal bearing <NUM> located between the compressor transfer tube <NUM> and the frame <NUM> of the micro-turbine alternator system <NUM>. The bearing system includes a second journal bearing <NUM> located between the coupling assembly <NUM> and the frame <NUM> of the micro-turbine alternator system <NUM>. The bearing system includes a third journal bearing <NUM> located between the turbine shaft <NUM> and the frame <NUM> of the micro-turbine alternator system <NUM>.

Advantageously, locating the electric generator <NUM> between the first stage compressor <NUM> and the second stage compressor <NUM> provides for very effective bearing placement around the compressor shaft <NUM>, which increases the stiffness of the compressor shaft <NUM>. The increased stiffness of the compressor shaft <NUM> allows for an increase in the critical speed of the compressor shaft <NUM>.

Also, advantageously, by locating the electric generator <NUM> between the first stage compressor <NUM> and the second stage compressor <NUM>, the alternator stator cooling heat exchanger <NUM> helps reduce the operating temperature of the electric generator <NUM>, while the airflow through the alternator stator cooling heat exchanger <NUM> also experiences a pressure drop. This pressure drop through the alternator stator cooling heat exchanger <NUM> forces some of the airflow from the first stage compressor <NUM> through the rotor element <NUM> and to a stator gap between the rotor element <NUM> and the stator element <NUM>, which provides cooling air to the rotor element <NUM>, the first journal bearing <NUM>, and the second journal bearing <NUM>.

The compressor transfer tube <NUM> extends from the first stage compressor <NUM> to the second stage compressor <NUM> through the electric generator <NUM>. The compressor transfer tube <NUM> is co-axial with the electric generator <NUM>. The rotor element <NUM> with the annular base member <NUM> and the annular array of permanent magnets <NUM> are located radially inward of the compressor transfer tube <NUM> measured relative to the central longitudinal axis A. The stator element <NUM> with the hub <NUM>, the conductive elements <NUM>, and the spokes <NUM> are located radially outward of the compressor transfer tube <NUM> measured relative to the central longitudinal axis A.

The first stage compressor <NUM>, the second stage compressor <NUM>, and the third stage compressor <NUM> drive air along a core flow path C for compression and communication in the combustion chamber <NUM>. The airflow in the core flow path C is compressed by the first stage compressor <NUM>, the second stage compressor <NUM>, and the third stage compressor <NUM>, is mixed with fuel and burned in the combustion chamber <NUM>, and is then expanded over the first stage turbine <NUM> and the second stage turbine <NUM>. The first stage turbine <NUM> and the second stage turbine <NUM> rotationally drive the turbine shaft <NUM> in response to the expansion. The combustion products are exhausted from the second stage turbine <NUM> through a turbine exit <NUM>.

Each of the first stage compressor <NUM>, the second stage compressor <NUM>, the third stage compressor <NUM>, the first stage turbine <NUM>, and the second stage turbine <NUM> may include rows of rotor 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>. The blades <NUM> of the rotor assemblies create or extract energy (in the form of pressure) from the core airflow that is communicated through the micro-turbine alternator system <NUM> along the core flow path C.

The micro-turbine alternator system <NUM> may include an auxiliary turbo charger <NUM> to pre-compress the airflow <NUM> prior to entering the core flow path C. The auxiliary turbo charger <NUM> includes a turbo compressor <NUM> and a turbine <NUM> operably connected to the turbo compressor <NUM> through a turbo compressor drive shaft <NUM>. The turbo compressor <NUM> is configured to rotate when the turbine <NUM> rotates.

The turbo compressor <NUM> is configured to pull external airflow <NUM> through one or more air inlets <NUM> in the frame <NUM> into a compressor flow path C1. The turbo compressor <NUM> is configured to compress the external airflow <NUM> in the compressor flow path C1 and deliver the airflow <NUM> to the first stage compressor <NUM> in the core airflow path C.

Each of the turbine <NUM> and the turbo compressor <NUM> may include rows of rotor assemblies (shown schematically) that carry airfoils that extend into the compressor flow path C1. For example, the rotor assemblies can carry a plurality of rotating blades <NUM>. The blades <NUM> of the rotor assemblies for the turbine <NUM> extract energy (in the form of pressure and temperature) from the exhaust <NUM> that is communicated through the micro-turbine alternator system <NUM> along the core flow path C. The blades <NUM> of the rotor assemblies for the turbo compressor <NUM> create energy (in the form of pressure and temperature) from the airflow <NUM> that is communicated through the micro-turbine alternator system <NUM> along the compressor flow path C1.

Combustor exhaust <NUM> exiting the turbine exit <NUM> is directed to the turbine <NUM> of the auxiliary turbo charger <NUM>. The exhaust <NUM> is then expanded over the turbine <NUM> of the auxiliary turbo charger <NUM>. The turbine <NUM> rotationally drives the turbo compressor drive shaft <NUM> in response to the expansion. Rotation of the turbo compressor drive shaft <NUM> causes the turbo compressor <NUM> to rotate and compress the airflow <NUM> within the compressor flow path C1.

Some embodiments further include a thermal electric energy recovery system <NUM>, configured to recover additional energy from exhaust <NUM> of the micro-turbine alternator system <NUM> before the exhaust <NUM> has flowed through the turbine <NUM> of the auxiliary turbo charger <NUM>.

Referring now to <FIG>, with continued reference to <FIG>, an isometric view of the second stage compressor <NUM> is illustrated in <FIG> and an isometric cutaway view of the second stage compressor <NUM> is illustrated in <FIG>, in accordance with an embodiment of the present disclosure. It is understood that while <FIG> and the associated description discuss the embodiments disclosed in relation with the second stage compressor <NUM>, the embodiments disclosed herein are not limited to the second stage compressor <NUM> and may be applicable to other compressors within the micro-turbine alternator system <NUM> or any other system where compressors or pumps are required.

The second stage compressor <NUM> includes central shaft <NUM>. The central shaft <NUM> is coaxial to a compressor longitudinal axis B. The central shaft <NUM> rotates about the compressor longitudinal axis B. When the second stage compressor <NUM> is installed within the micro-turbine alternator system <NUM> of <FIG>, the compressor longitudinal axis B is colinear with the central longitudinal axis A. In other words, the compressor longitudinal axis B and the central longitudinal axis A are the same axis when the second stage compressor <NUM> is installed within the micro-turbine alternator system <NUM> of <FIG>.

The central shaft <NUM> includes an external surface <NUM> and an internal surface <NUM>. The central shaft <NUM> includes a passageway <NUM> formed therein. The internal surface <NUM> defines the passageway <NUM>. The passageway <NUM> is coaxial with the compressor longitudinal axis B. The passageway <NUM> may be tubular in shape and configured to fit the turbine shaft <NUM> (See <FIG>). In other words, the turbine shaft <NUM> is configured to fit within the passageway <NUM>.

The second stage compressor <NUM> includes a shroud <NUM> extending circumferentially around the central shaft <NUM>. The shroud <NUM> is separated from the central shaft <NUM> by a gap G1. The gap G1 extends circumferentially around the compressor longitudinal axis B and may vary in size moving from a forward end <NUM> of the shroud <NUM> to an aft end <NUM> of the shroud <NUM>. The shroud <NUM> encircles the central shaft <NUM>. The shroud <NUM> includes a radially outward surface <NUM> and a radially inward surface <NUM> located opposite the radially outward surface <NUM>. The core flow path C is defined between the external surface <NUM> of the central shaft <NUM> and the radially inward surface <NUM> of the shroud <NUM>. In other words, the external surface <NUM> of the central shaft <NUM> and the radially inward surface <NUM> of the shroud <NUM> are in a facing spaced relationship forming the core flow path C therebetween.

The radially outward surface <NUM> of the shroud <NUM> may have a first outer diameter OD1 along a first portion <NUM> of the shroud <NUM>. Alternatively, the first outer diameter OD1 may be slightly raised in the first portion <NUM> with an undercut aft of the first portion <NUM> in the radially outward surface <NUM>. The undercut may facilitate grinding operations. The first portion <NUM> may be located at the forward end <NUM> of the shroud <NUM>. The radially outward surface <NUM> of the shroud <NUM> may have a second outer diameter OD2 along a second portion <NUM> of the shroud <NUM>. The second outer diameter OD2 is greater than the first outer diameter OD1. The second portion <NUM> of the shroud <NUM> is located at a first distance DIS1 away from the first portion <NUM> as measured along the compressor longitudinal axis B. The second portion <NUM> may be closer to the aft end <NUM> of the shroud <NUM> than to the forward end <NUM>.

The second stage compressor <NUM> includes a plurality of blades <NUM> circumferentially encircling the central shaft <NUM>. Each of the plurality of blades <NUM> extend from the external surface <NUM> of the central shaft <NUM> to the radially inward surface <NUM> of the shroud <NUM>. The blades <NUM> of the second stage compressor <NUM> transfer mechanical energy of the rotating shaft into pneumatic energy in the fluid stream (in the form of dynamic pressure) by compressing and accelerating the airflow in the core airflow path C. The blades <NUM> may be contoured between the external surface <NUM> of the central shaft <NUM> and the radially inward surface <NUM> of the shroud <NUM> to appropriately compress and accelerate the airflow in the core airflow path C as required.

The second stage compressor <NUM> is a monolithic structure rather than being assembled from separate individually formed components that are then assembled. The term monolithic may be defined as an object that is cast or formed as single piece without joints or seams. In other words, the second stage compressor <NUM> is formed as a single piece comprising a unitary structure. In an embodiment, the second stage compressor <NUM> has no joints or seams. The second stage compressor <NUM> is manufactured or formed via additive manufacturing. Additive manufacturing may include, but is not limited to 3D printing, laser powder bed fusion (L-PBF) additive manufacturing, investment casting (using the rapid prototype method) or any other additive manufacturing technique known to one of skill in the art.

Referring now to <FIG>, with continued reference to <FIG>, an isometric view of a compressor assembly <NUM> is illustrated in <FIG> and an isometric cutaway view of the compressor assembly <NUM> is illustrated in <FIG>, in accordance with an embodiment of the present disclosure.

The compressor assembly <NUM> includes the second stage compressor <NUM> and the compression ring <NUM> extending circumferentially around the shroud <NUM> of the second stage compressor <NUM>. It is understood that while <FIG> and the associated description discuss the embodiments disclosed in relation with the second stage compressor <NUM>, the embodiments disclosed herein are not limited to the second stage compressor <NUM> and may be applicable to other compressors within the micro-turbine alternator system <NUM> or any other system where compressors are required.

The compression ring <NUM> is a stress relieving compressor shroud compression ring and is configured to relieve stress on the second stage compressor <NUM> during operation by compressing the second stage compressor <NUM>. The compression ring <NUM> is configured to relieve stress on the second stage compressor <NUM> by compressing the shroud <NUM>. The compression ring <NUM> is configured to apply an approximately equal pressure circumferentially around the radially outward surface <NUM> of the shroud <NUM> towards central shaft <NUM> and the compressor longitudinal axis B.

As previously noted, since the second stage compressor <NUM> is manufactured utilizing additive manufacturing techniques it may have a reduces tensile strength in comparison to a subtractive manufactured impeller. In an embodiment, the second stage compressor <NUM> may be composed of titanium. The material strength capability or tensile strength for additive manufactured titanium may be about <NUM> ksi. However, due to the high rotational operating speed of the second stage compressor <NUM>, the second stage compressor <NUM> may experience a tensile stress of about <NUM> ksi. Embodiments disclosed herein seek to utilize a compression ring <NUM> that is installed via an interference fit around the shroud <NUM> of the second stage compressor <NUM>. The compression ring <NUM> bridges the gap between the material strength capability of the additively manufactured second stage compressor <NUM> and the operational tensile stress experienced during operation by compressing the shroud <NUM> and the blades <NUM> into the central shaft <NUM>.

The compression ring <NUM> includes a radially inner surface <NUM> and a radially outer surface <NUM> opposite the radially inner surface <NUM>. The radially inner surface <NUM> of the compression ring <NUM> is configured to mate flush with one or more portions <NUM>, <NUM> of the radially outward surface <NUM> of the shroud <NUM>. The radially inner surface <NUM> of the compression ring <NUM> is configured to apply a radially inward compressive force F1 along the one or more portions <NUM>, <NUM> of the radially outward surface <NUM> of the shroud <NUM>. The radially inward compressive force F1 is configured to compress the shroud <NUM> and the blades <NUM> into the central shaft <NUM>, which helps relieve operational tensile stress on the shroud <NUM> and the blades <NUM> when rotating at operational speeds. As shown in <FIG>, the radially inward compressive force F1 is directed towards the compressor longitudinal axis B.

The compression ring <NUM> utilizes an interference fit with the shroud <NUM> to place the shroud <NUM> in compression when the second stage compressor <NUM> is at rest. More specifically, the radially inner surface <NUM> of the compression ring <NUM> utilizes an interference fit with the radially outward surface <NUM> of the shroud <NUM> to place the shroud <NUM> in compression when the second stage compressor <NUM> is at rest. The radially inner surface <NUM> of the compression ring <NUM> utilizes an interference fit with one or more portions <NUM>, <NUM> of the radially outward surface <NUM> of the shroud <NUM> to place the shroud <NUM> in compression when the second stage compressor <NUM> is at rest.

As the rotational speed of the compressor assembly <NUM> increases, the compression stress from the compression ring <NUM> decreases, until the micro-turbine alternator system <NUM> reaches about <NUM>% speed. At this speed, the blades <NUM> of the second stage compressor <NUM> may not be subject to any stress. As the speed continues to increase, the blade <NUM> stress starts to increase in the tensile direction. By full speed, the tensile stress in the blades <NUM> may be about <NUM>% of the tensile stress that would be present without the compression ring <NUM> helping to support the mass of the shroud <NUM>.

An inner diameter ID1, ID2 of the radially inner surface <NUM> of the compression ring <NUM> may vary in size to mate properly with the first portion <NUM> and the second portion <NUM> of the radially outward surface <NUM> of the shroud <NUM>. The radially inner surface <NUM> of the compression ring <NUM> includes a first area <NUM> and a second area <NUM>.

The second area <NUM> is located at a first distance DIS1 away from the first area <NUM> as measured along the compressor longitudinal axis B. The second area <NUM> may be closer to an aft end <NUM> of the compression ring <NUM> than to a forward end <NUM>.

The first area <NUM> of the radially inner surface <NUM> of the compression ring <NUM> is configured to mate flush with the first portion <NUM> of the radially outward surface <NUM> of the shroud <NUM>. The second area <NUM> of the radially inner surface <NUM> of the compression ring <NUM> is configured to mate flush with the second portion <NUM> of the radially outward surface <NUM> of the shroud <NUM>.

The radially inner surface <NUM> of the compression ring <NUM> has a first inner diameter ID1 along the first area <NUM> of the radially inner surface <NUM> of the compression ring <NUM>. The radially inner surface <NUM> of the compression ring <NUM> has a second inner diameter ID2 along the second area <NUM> of the radially inner surface <NUM> of the compression ring <NUM>. The second inner diameter ID2 is greater than the first inner diameter ID2.

In order to accomplish the interference fit, when disassembled, the first inner diameter ID1 of the radially inner surface <NUM> of the compression ring <NUM> is less than the first outer diameter OD1 of the radially outward surface <NUM> of the shroud <NUM> and the second inner diameter ID2 of the radially inner surface <NUM> of the compression ring <NUM> is less than the second outer diameter OD2 of the radially outward surface <NUM> of the shroud <NUM>. To assemble, the compression ring <NUM> is expanded by a heat source, the second stage compressor <NUM> is shrunk by a cold source, and then the compression ring <NUM> is slid onto the shroud <NUM>. Once assembled and the temperature of compression ring <NUM> and the second stage compressor <NUM> reach an equilibrium, the first inner diameter ID of the radially inner surface <NUM> of the compression ring <NUM> is about equal to the first outer diameter OD1 of the radially outward surface <NUM> of the shroud <NUM> and the second inner diameter ID2 of the radially inner surface <NUM> of the compression ring <NUM> is about equal to the second outer diameter OD2 of the radially outward surface <NUM> of the shroud <NUM>.

Once the second stage compressor <NUM> starts to spin, the pre-loaded blades <NUM> (in compression while at rest) relax as centrifugal force causes the shroud <NUM> and the compression ring <NUM> to expand. Advantageously, the compression ring <NUM> is configured to allow the transfer of the centrifugal load from the shroud <NUM> to the compression ring <NUM>.

In an embodiment, the compression ring <NUM> is formed via subtractive machining and thus has an increased tensile strength in comparison to the second stage compressor <NUM> that was additively manufactured. In an embodiment, the compression ring <NUM> may be machined from a titanium allow billet with a tensile strength of about <NUM> ksi. In another embodiment, the second stage compressor <NUM> has a first tensile strength and the compression ring <NUM> has a second tensile strength that is greater than the first tensile strength.

In an embodiment, the second stage compressor <NUM> is composed of additive manufactured titanium with a tensile strength of about <NUM> ksi and the compression ring <NUM> may be machined from a titanium alloy billet with a tensile strength of about <NUM> ksi, which would advantageously reduce the maximum tensile stresses experienced in the additively manufactured second stage compressor <NUM> and shroud <NUM> to less than <NUM> ksi during rotational operation.

Referring now to <FIG>, with continued reference to <FIG>, a flow chart of a method <NUM> of manufacturing a compressor assembly <NUM> is illustrated, in accordance with an embodiment of the disclosure. At block <NUM>, a temperature of a compressor <NUM> is reduced. At block <NUM>, a temperature of a compression ring <NUM> is increased. Block <NUM> may occur prior to block <NUM>, after block <NUM>, or simultaneous to block <NUM>. Block <NUM> occurs after both block <NUM> and <NUM>. At block <NUM>, the compression ring <NUM> is slid onto the shroud <NUM> of the compressor <NUM> such that the compression ring <NUM> extends circumferentially around the shroud <NUM>. The compression ring <NUM> being in an interference fit with the shroud <NUM> once the temperature of compression ring <NUM> and the temperature of the compressor <NUM> reach an equilibrium.

The compression ring <NUM> is configured to apply a radially inward compressive force F1 along one or more portions <NUM>, <NUM> of the radially outward surface <NUM> of the shroud <NUM>. The radially inward compressive force F1 is configured to compress the shroud <NUM> and the plurality of blades <NUM> into the central shaft <NUM>.

The method <NUM> includes forming the compressor <NUM> via an additive manufacturing technique. The method <NUM> may further include forming the compression ring <NUM> via subtractive machining. In an embodiment, the compressor <NUM> has a first tensile strength and the compression ring <NUM> has a second tensile strength that is greater than the first tensile strength.

While the above description has described the flow process of <FIG> in a particular order, it should be appreciated that unless otherwise specifically required in the attached claims that the ordering of the steps may be varied, and the order of the steps may occur simultaneously or near simultaneously.

Technical effects and benefits of the features described herein include utilizing a compression ring in an interference fit with a shroud of a compressor to increase the tensile strength of the compressor.

Claim 1:
A compressor assembly (<NUM>), comprising:
a compressor (<NUM>), comprising:
a central shaft (<NUM>) comprising an external surface (<NUM>);
a shroud (<NUM>) extending circumferentially around the central shaft, the shroud comprising a radially inward surface (<NUM>) and a radially outward surface (<NUM>) located opposite the radially inward surface, wherein the external surface of the central shaft and the radially inward surface of the shroud are in a facing spaced relationship forming a core flow path (C) therebetween; and
a plurality of blades (<NUM>) extending from the central shaft to the shroud;
wherein the compressor (<NUM>) is formed via an additive manufacturing technique; characterised in that the assembly further comprises
a compression ring (<NUM>) extending circumferentially around the shroud, the compression ring being in an interference fit with the shroud,
wherein the compression ring is configured to apply a radially inward compressive force (F1) along one or more portions of the radially outward surface of the shroud, the radially inward compressive is configured to compress the shroud and the plurality of blades into the central shaft.