Disk engine with circumferential swirl radial combustor

A disk engine and system configured to provide high power at a reduced axial length is disclosed herein. The disk engine includes a radial compressor, a compressor discharge manifold positioned circumferentially about compressor, a combustion chamber positioned within the discharge manifold and a radial turbine positioned radially inward of the combustion chamber.

TECHNICAL FIELD

The present disclosure generally relates to a disk engine and more particularly, but not exclusively to a disk engine with a radial combustor to define an axially compact power system.

BACKGROUND

Gas turbine engines are used to power aircraft, watercraft and land vehicles. While axial flow gas turbine engines are prevalent, some gas turbine engines have a radial flow compressor and a radial flow turbine which somewhat reduces the axial length of the engine. Some applications have requirements for an engine with an axial length that is smaller than prior art systems can provide. Accordingly, there remains a need for further contributions in this area of technology.

SUMMARY

One embodiment of the present disclosure includes a unique disk engine configured with a radial flow combustor to minimize axial space claim. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations wherein the disk engine includes a unique method and means for inducing circumferential combustor swirl to facilitate efficient combustion in a radial flow combustor. Further embodiments, forms, features, aspects, benefits, and advantages of the present application shall become apparent from the description and figures provided herewith.

DETAILED DESCRIPTION OF THE ILLUSTRATIVE EMBODIMENTS

The disk engine disclosed herein operates on a traditional Brayton cycle like other gas turbine engines, but the configuration is designed such that the axial length of the engine is significantly reduced over traditional engines. A circumferential swirl combustor positioned around the outside diameter of the engine is one feature that enables the compact engine architecture. The circumferential flow combustor enhances the mixing of combustion reactants and permits a reduction of the axial length of the combustor compared to other gas turbine engines. Swirling the reactants at a high centrifugal loading inside the combustion cavity has been shown to enhance Rayleigh-Taylor instabilities leading to enhanced mixing of the reactants. Greater mixing means that the reactions can happen faster in a smaller combustor volume. The outside diameter of the combustion cavity serves as the primary reaction zone, where the fuel is burned rich. The mixture moves radially inward, while simultaneously swirling circumferentially, where additional air is introduced and the remaining fuel is burned lean. This process follows the Rich Burn, Quick Quench, Lean Burn (RQL) cycle which is beneficial in reducing heat load to the engine hardware, and reducing NOx emissions.

The amount of centrifugal loading that can be applied to the reactants in the primary zone is limited to be able to sustain combustion. The target to achieve maximum mixing while ensuring combustion is sustained is approximately 2,500 g's (2,500 times the acceleration of gravity), with values above 3,500 g's often resulting in a flameout. The present design allows the fluid flow in the primary zone (cavity outer diameter) to be swirled in a helical pattern by using offset, horizontally opposed air injectors. The size of the air injector openings can be varied or adjusted in some embodiments so that the circumferential loading can be controlled at an optimum value across the entire engine operating range. The circumferential swirl of the combustion reactants (air and fuel) can be maintained throughout the entire combustor flowpath and continue directly into the radial turbine.

The length of an axial combustor in use in engines today is controlled by the amount of time it takes for the combustor to inject fuel and air, mix the fuel and air, and fully combust the fuel mixture prior to entering the turbine section. The method of injection and mixing in an axial combustor requires a relatively large minimum length to achieve these goals. The present invention provides a circumferential swirl to the combustor flow path so that the flow is redirected in the circumferential direction which in turn minimizes the axial length required to complete the combustion process prior to reaching the turbine inlet.

The combustion chamber is defined by a circumferential cavity with radial dilution inlet holes directing airflow into the chamber. The dilution inlet holes can be angled and horizontally opposed such that the fuel and air reactants inside are rotating at high centripetal acceleration in a spiral, cork-screw fashion. The inlet holes can include variable geometry features to limit the maximum circumferential acceleration in the cavity. The combustion chamber includes a primary combustion zone in the outer diameter region. A secondary combustion zone and dilution zone are located radially inward of the primary combustion zone. Additional variable geometry inlets can be used to control the air distribution throughout the combustion chamber. In the last stage of combustion proximate the dilution zone, the exhaust products are directed through a row of stator vanes to accelerate and direct the gases into a radial inflow turbine.

The combustion chamber of the disk engine as defined herein can achieve equal performance to a traditional gas turbine engine combustor, but with a significantly shorter axial length and corresponding weight reduction. Further reduction in axial length of the combustion chamber is achieved by positioning the turbine radially inward of the combustor. Additionally, this radial combustor design provides a relatively uniform temperature profile across the exit plane as combustion gases pass into the turbine rotor inlet. This is important, because if the temperature profile of the combustion products at the turbine rotor inlet includes hot spots or regions, the mechanical life and thermal efficiency of the turbine are reduced.

Referring now toFIG. 1, a perspective view of a disk engine10according to one illustrative embodiment is disclosed herein. The disk engine10includes a compressor side20and a turbine side30. Air represented by arrow39enters the disk engine10through an inlet40. A plurality of fuel inlet ports or injectors50are operably connected to the disk engine10through an outer circumferential perimeter wall52. Fuel is combusted with the air and the resulting exhaust products represented by arrow59exits the disk engine10through the exhaust nozzle outlet60.

Referring toFIG. 2, a cutaway view of the disk engine10is shown. A radial compressor70receives air from an inlet40. The radial compressor70compresses the air and discharges the pressurized air into a circumferential manifold80extending around the disk engine10radially inward of the outer perimeter wall52. A circumferential combustion chamber or combustor90is positioned within the manifold80. The combustor90includes a plurality of inlet holes or apertures92to permit the pressurized air to transfer from the manifold into the combustor90. The inlet apertures92can be offset from side to side of the combustor90. Furthermore, the inlet apertures92can be angled relative to one another. The offset positions and angle variation provides a swirling motion within the combustor to promote mixing of the air and fuel to speed up the combustion event.

A radial turbine120receives high temperature pressurized combustion flow exiting the combustor90, which drives the turbine at a high rotational speed. The rotational speed is dependent on physical size and operating conditions, but can range anywhere from a few thousand RPM to over one hundred thousand RPM. The turbine120drives the compressor70and can be connected to other shaft driven devices such as an electric power generator (not shown) or the like. The exhaust flow is accelerated through the exhaust nozzle60and can generate a thrust sufficient to power an aircraft or a missile.

Referring toFIG. 3, an enlarged view of a portion of the disk engine10depicts working fluid flow paths represented by arrows140through the engine. The compressor inlet airflow39enters the compressor70and is compressed to a design pressure. The compressed air140exits through the compressor outlet and is directed into the manifold80. The compressed air then enters the combustor90at an outer primary combustion zone150, at a secondary zone160radially inward of the primary zone and at a dilution zone170radially inward of the secondary zone160. The airflow from each of the zones150,160and170can be combusted with fuel or used to dilute and cool the combustion products or other hot regions of the engine10. After exiting the combustor90, the combustion product working fluid is directed to the turbine inlet180. The working fluid provides energy to the turbine and exits as a turbine outlet flow190prior to discharging from the engine10as an exhaust outlet flow59.

Referring toFIG. 4, a side view of the engine10shows a plurality of dilution holes94positioned internal to each of a plurality of turbine inlet guide vanes or stators122positioned around the turbine120. Each stator122include an upper hook portion123that wraps around a dilution hole94to provide some directional control of the airflow passing therethrough.

Referring toFIG. 5, a cutaway side view of the engine10shows a plurality of compressor outlet guide vanes72positioned around the compressor70. The arrows represent compressed airflow exiting though the outlet guide vanes72as is known to those skilled in the art.

Referring toFIG. 6, another side view of the engine10shows a bearing mechanism220structured to rotatably support the compressor70and the turbine120as they rotate at high speeds.

Referring toFIGS. 7A and 7B, a schematic illustration of a sliding tabbed plate for use as a variable area air inlet device is shown. A combustor wall250for a combustion chamber can include an aperture260to permit compressed air to flow therethrough. A sliding cover270can be moved relative to the aperture260from a completely open position to a closed position. The sliding cover can be moved via electric or mechanical actuation as would be known to those skilled in the art.FIG. 7Ashows the variable area inlet280completely open andFIG. 7Bshows the variable are inlet280partially closed with a sliding cover270. In the disclosed embodiment, the aperture200is rectangular in shape, however other shapes are contemplated herein.

Referring toFIGS. 8A and 8B, a schematic illustration of a rotatable plate for use as a variable area air inlet device according to another illustrative embodiment is shown. A combustor wall300for a combustion chamber can include a variable aperture320to permit compressed air to flow therethrough. A rotatable plate330includes a control aperture340that when aligned with the aperture320of the combustor wall300, airflow350can pass through both apertures320and340without impedance or restriction. The rotatable plate330can be rotated as illustrated by double arrow360, such that the aperture320can be partially or completely closed.FIG. 8Ashows the variable inlet320completely open with unrestricted flow represented by arrows350.FIG. 8Bshows the variable aperture320completely closed as the control aperture340is rotated away from the variable aperture resulting in the flow370being prevented from passing through the aperture320.

Referring now toFIGS. 9-13, an alternate embodiment of a disk engine1010is illustrated in the various views. Many of the features of the disk engine1010may be the same or similar to the disk engine10and, therefore will not be described again. With specific reference toFIGS. 9-11, the disk engine1010includes a plurality of fuel injection inlet ports or fuel injectors1050engaged through the compressor side wall20around the radial compressor70. Each fuel injector1050extends through a corresponding air swirler1060that is operable to generate additional turbulence into the compressed air discharged from the radial compressor70. The swirlers1060facilitate mixing of the fuel and air so that the reactants can combust quickly in a circumferential combustor1065. A combustion liner1070is positioned within a circumferential manifold1072such that a combustion liner passageway1080is formed therebetween to transport compressor discharge air therethrough. The combustor or combustion chamber1072is located radially inward of the combustion liner1070.

A plurality of secondary air injection apertures1090are formed in an outer perimeter wall of the combustion liner1070. Air can be directed from the combustion liner passageway1080through the secondary air injection apertures1090and into a primary combustion zone1093and a secondary combustion zone2015of the combustor1065. A plurality of dilution cooling holes2010are formed in sidewalls of the combustion liner1070. The dilution cooling holes2010transport cooling air from the combustion liner passageway1080into the secondary zone2015of the combustor1065. The air passing through the dilution cooling holes2010can be used to cool certain portions of the disk engine1010and provide addition air to support complete lean combustion of any remaining unburned reactants.

A centerbody2012extends into the combustor1065and includes a plurality of cooling pathways2016that receive and transport compressor discharge air to cool various components internal to the disk engine1010. A turbine stator2020includes a stator cooling passageway2022formed therein that is in fluid communication with one or more of the cooling pathways2016of the centerbody2012.

Referring more particularly toFIG. 11, the compressor discharge flow paths are illustrated with arrows1095. A substantial majority of the compressor discharge flows through the plurality of swirlers1060where fuel is injected through the fuel injection ports1050to provide a combustion mixture with a desired air/fuel ratio to the combustor1065. A portion of the compressor discharge flow is diverted through the combustion liner passageway1080and other cooling passageways as described inFIG. 10.

Referring now toFIG. 12, a compressor side cutaway view of the disk engine1010is illustrated with an enlarged view of a swirler1060. A plurality of swirlers1060are positioned around the compressor70radially outward of the compressor outlet stator vanes72. Each swirler1060includes an angled inlet scoop2040that is proportional to the angle of the compressor outlet vanes72. The scoop angle may vary, but in one form can be approximately 45 degrees relative to a radial centerline. The swirler1060includes a plurality of swirler vanes2050configured to generate turbulence and direct the compressor discharge flow through a swirling flow path represented by arrows2060.

Referring now toFIG. 13, a turbine side cutaway view of the disk engine1010is illustrated with an enlarged view of a turbine inlet guide vane122. The turbine inlet guide vane122includes an upper hook portion123wrapped around a stator coolant hole94. The turbine rotates in the direction of arrow2080and rotationally drives the compressor as is well understood.

In one aspect, the present disclosure includes a disk engine comprising: a radial compressor; a compressor discharge manifold positioned circumferentially around the compressor; a combustion chamber positioned within the discharge manifold; and a radial turbine positioned radially inward of the combustion chamber.

In refining aspects the present disclosure includes a plurality of fuel inlet ports positioned through an outer perimeter wall of the discharge manifold; wherein the combustion chamber includes a primary combustion zone in a radially outer region; wherein the combustion chamber includes a secondary combustion zone radially inward of the primary combustion zone; wherein the combustion chamber includes a dilution combustion zone radially inward of the secondary combustion zone; a plurality of offset air inlet holes positioned on opposing sides of the combustion chamber; wherein opposing pairs of the plurality of offset air inlet holes are angled relative to one another; wherein the air inlet holes are positioned proximate a primary combustion zone, a secondary combustion zone and a dilution zone of the combustion chamber; a variable geometry inlet flow control device in fluid communication with one or more of the air inlet holes; wherein the variable geometry inlet flow control device includes a sliding plate operable to selectively block a flow aperture; wherein the sliding plate is movable between fully open position and a closed position; and wherein the variable geometry inlet flow control device includes a rotating plate with one or more though apertures configured to rotate and selectively block a portion of one or more of the air inlet holes.

In another aspect, the present disclosure includes a disk engine comprising: a radial compressor enclosed by a compressor sidewall; a compressor discharge manifold positioned circumferentially around the compressor; a combustion chamber positioned within the discharge manifold; a radial turbine positioned radially inward of the combustion chamber; and a plurality of fuel injectors extending through the compressor sidewall around a compressor discharge region of the radial compressor.

In refining aspects the present disclosure includes a plurality of air swirler devices positioned downstream of the radial compressor; wherein each fuel injector extends through one of the air swirler devices; a combustion liner formed within the discharge manifold; a combustion liner cooling passageway formed between the combustion liner and the discharge manifold; a plurality of cooling air holes formed through the combustion liner; a center body having a plurality of cooling passageways extending into the discharge manifold and a turbine stator having a cooling passageway in fluid communication with at least one of the cooling passageways in the center body.