Abstract: A compact axial turbine configured to operate with high density working fluid is described. The turbine comprises an axial majority cantilevered turbomachinery shaft. Rotor assemblies and nozzle spacers communicate torque through turbine shaft splines, allowing them to be slid off the shaft for quick replacement in the field. The compact axial turbine houses turbomachinery within a separable inner casing encircled by a cartridge sleeve, thereby forming a cartridge which can itself be removed as a single component.

Abstract: A combustor structure of an embodiment is disposed to penetrate, from a direction perpendicular to an axial direction of a turbine rotor in a supercritical CO2 gas turbine which uses supercritical CO2 for a working fluid, a casing of the supercritical CO2 gas turbine. The combustor structure includes a plurality of combustors. Each of the combustors includes: a combustor liner in a cylindrical shape, which combusts fuel and an oxidant; a fuel supply part which is provided at an upstream end of the combustor liner and supplies the fuel into the combustor liner; and an oxidant supply part which is provided at the upstream end of the combustor liner and supplies the oxidant into the combustor liner.

Abstract: A containment structure for a rotor includes a shroud and a shroud reinforcement. The shroud is coaxial with and partially surrounds the rotor and includes a tubular section, a transition section, and a flange section. The tubular section extends axially past a first side of the rotor. The transition section connects to the tubular section and is adjacent to a curved side of the rotor. The flange section connects to the transition section opposite the tubular section. The flange section extends radially past a radially outer side of the rotor. The shroud reinforcement is connected to a radially outer surface of the transition section. The shroud reinforcement encloses the transition section and includes a support scaffold and a reinforcing material. The support scaffold includes a series of geometric retaining features encircling a radially outer surface of the transition section. The reinforcing material couples to the support scaffold and restricts shroud radial expansion.

Abstract: An assembly is provided for a gas turbine engine. This gas turbine engine assembly includes a stationary engine structure. The stationary engine structure includes a diffuser, a combustor, an engine case and a plenum. The combustor is disposed within the plenum. The engine case forms a peripheral boundary of the plenum. A gas path extends sequentially through the diffuser, the plenum and the combustor. A first section of the stationary engine structure is formed as a first monolithic body. The first section includes the diffuser and the combustor. A second section of the stationary structure is formed as a second monolithic body. The second section is configured as or otherwise includes the engine case.

Abstract: An engine system of an aircraft includes a gas turbine engine comprising at least one spool and at least one electric machine operably coupled with the at least one spool. A controller is configured to detect if the at least one spool of the gas turbine engine is in or is approaching an overspeed condition and apply a load to the at least one spool via the at least one electric machine.

Abstract: In a method for operating a combustor of an embodiment, before ignition in the combustor, a mixed gas containing oxygen is circulated through the combustor as a circulating gas. Then, in an operating time from the time of ignition in the combustor to the time of a rated load of a turbine, from the time of ignition until reaching stable combustion conditions allowing stable combustion, a combustion gas in which a controller controls a flow rate of a fuel supplied from a fuel supply part and a flow rate of an oxidant supplied from an oxidant supply part to maintain the same oxygen concentration as an oxygen concentration in the mixed gas is circulated as the circulating gas.

Abstract: Apparatus, systems, and methods use cryogenic liquids such as, for example, liquefied natural gas and liquefied air or liquefied air components to store thermal energy. The cryogenic liquids may be produced using electrically powered liquefaction methods, for example, using excess electric power during periods of over-generation on the electric grid.

Abstract: A gas turbine engine having: an engine core having, in serial flow communication, a low-pressure compressor, a high-pressure compressor, a high-pressure turbine drivingly connected to the high-pressure compressor, and a low-pressure turbine drivingly connected to an output shaft; and a clutch having a disengaged configuration in which the low-pressure turbine is drivingly disconnected from the low-pressure compressor such that, in the disengaged configuration, the clutch disengages the low-pressure turbine from the low-pressure compressor, and an engaged configuration in which the low-pressure turbine is drivingly connected to the low-pressure compressor, the low-pressure turbine drivingly connected to the output shaft in both of the engaged and disengaged configurations of the clutch.

Abstract: A gas turbine engine has a first spool having a low pressure compressor section in fluid communication with an air inlet, the low pressure compressor section including a first plurality of variable guide vanes therein, and a low pressure turbine section drivingly engaged to the low pressure compressor section. A second spool has a high pressure compressor section in fluid communication with the low pressure compressor section to receive pressurized air therefrom, the high pressure compressor section including a second plurality of variable guide vanes at an entry thereof, and a high pressure turbine section drivingly engaged to the high pressure compressor section, the high pressure turbine section disposed upstream of the low pressure turbine section and in fluid communication therewith. An output drive shaft drivingly engages the low pressure turbine section and is adapted to drivingly engage a rotatable load of the gas turbine engine.

Abstract: An apparatus and method for diffusing an airflow in a turbine engine can include a turbine center frame positioned between a high pressure turbine and a low pressure turbine of a turbine section of the engine. The turbine center frame can include two or more diffusion sections for diffusing the airflow. A stabilization section can be provided between the two or more diffusion sections to stabilize the airflow.

Abstract: An electric compressor system for a vehicle includes: an electric motor having a rotor and a motor shaft which selectively rotate in a first rotation direction or a second rotation direction; an external rotation shaft extending from the motor shaft of the electric motor; a first compressor unit connected to the external rotation shaft and selectively compressing a first fluid according to the rotation direction of the external rotation shaft; and a second compressor unit connected to the external rotation shaft and selectively compressing a second fluid according to the rotation direction of the external rotation shaft, wherein the first compressor unit and the second compressor unit are sequentially arranged on the external rotation shaft, the first compressor unit is fluidly connected to a first fluid system, and the second compressor unit is fluidly connected to a second fluid system.

Abstract: An axial-direction passage, a forced vortex passage, a first blade array passage, and a second blade array passage are formed in a rotor shaft of a turbine. Cooling air from an air extraction port of a compressor flows through the axial-direction passage which extends in an axial direction. The forced vortex passage is connected to the axial-direction passage and extends outwards in the radial direction relative to an axial line from a connecting portion between the forced vortex passage and the axial-direction passage. The first blade array passage is connected to an end portion on the outer side in the radial direction of the forced vortex passage and guides cooling air to a first blade row among a plurality of blade rows. The second blade array passage is connected to an end portion of the forced vortex passage and guides cooling air to a second blade row.

Abstract: A fan assembly that has a fan hub having a plurality of blades extending radially therefrom and having a variable pitch and a controller in communication with the fan assembly to reposition a pitch of the plurality of blades. The controller establishes a variable maximum pitch of the plurality of blades based on an ambient temperature.

Abstract: An apparatus is provided that includes a rotor bore and a bore basket. The bore basket extends axially between a first and second axial ends. The bore basket includes an engagement panel, and cylindrical inner and outer radial panels. The inner and outer radial panels extend substantially between the first and second axial ends. The inner radial panel is disposed radially inside of and separated from the outer radial panel, defining an annular passage disposed there between. The bore basket includes a plurality of inlet apertures in fluid communication with the annular passage at a first axial position, and a plurality of exit apertures in fluid communication with the annular passage at a second axial position. The second axial end of the bore basket is attached to an inner radial hub of the rotor bore, and the rotor bore and bore basket are a unitary structure.

Abstract: The present invention discloses embodiments for a power augmentation system of a gas turbine engine resulting in performance improvements while also improving efficiency. The invention provides systems and methods for generating a heated air supply by way of mixing compressed air from an electrically-driven process with air drawn from the engine compressor discharge plenum.

Abstract: A cooling system for a turbine engine is provided. The turbine engine includes a compressor, a compressor discharge chamber (CDC), a combustor assembly, and a turbine coupled in a serial flow relationship such that a first portion of air from the CDC is channeled to the combustor assembly. The turbine is coupled to the compressor via a rotor. The cooling system includes an air duct configured to channel a second portion of air from the CDC to a mid-rotor region of the rotor, and a fluid supply system coupled to the air duct at a coupling. The fluid supply system is configured to channel a flow of fluid to the coupling. The coupling is configured to cool the second portion of CDC air via absorption of heat by the fluid from the second portion of CDC air.

Abstract: A tap is connected to a location upstream of a downstream most location in a compressor to a heat exchanger. Downstream of the heat exchanger is a shut off valve and a cooling compressor. The cooling compressor is connected to a chamber provided with a check valve configured to selectively allow flow directly from a downstream location in the compressor. There is a system for stopping operation of the cooling compressor, and a control for closing the shut off valve. The cooling compressor is configured to compress air to a greater pressure than the higher pressure, such that the check valve is configured to maintain a closed position, but when said cooling compressor is not providing compressed air, the at least one check valve is configured to allow said higher pressure flow into the chamber. A method is also disclosed.

Abstract: An inlet air cooling system used in a gas turbine for supplying power to a refrigerant compressor for compressing refrigerant in a natural gas liquefaction plant includes: an inlet air cooler for cooling inlet air of the gas turbine; chiller motors used for a chiller for cooling coolant supplied to the inlet air cooler; a first variable speed driver for supplying electric power to each of the one or more chiller motors; and an electric generator driven by the gas turbine, wherein the electric generator is electrically connected to the first variable speed driver, and electric power generated by the electric generator can be supplied to each of the chiller motors from the first variable speed driver without using a main power line of an electric power system, which enables efficient electric power supply to the motors via the variable speed driver.

Abstract: The present invention relates to a gas turbine which includes a cooling system provided with cooling air supply paths bypassed through an outside casing and, more specifically, to a gas turbine including a cooling system and a cooling method, wherein, in order to supply cooling air to a plurality of turbine blades and other devices provided to the inside of the gas turbine, cooling air supply paths are not provided to the rotor center shaft of the gas turbine but provided around the outer casing of the gas turbine so as to achieve the increased effect in the aerodynamic efficiency of the compressor and the turbine. According to the above structure and method, it may be possible to provide the cooling air supply paths not to the rotor center shaft of the gas turbine but around the outer casing of the gas turbine, thereby finally achieving the increased effect in the aerodynamic efficiency of the compressor and the turbine.

Abstract: Apparatus, systems, and methods store energy by liquefying a gas such as air, for example, and then recover the energy by regasifying the liquid and combusting or otherwise reacting the gas with a fuel to drive a heat engine. The process of liquefying the gas may be powered with electric power from the grid, for example, and the heat engine may be used to generate electricity. Hence, in effect these apparatus, systems, and methods may provide for storing electric power from the grid and then subsequently delivering it back to the grid.

Abstract: A turbine exhaust case frame (100) comprises an inner ring (104), an outer ring (102), and a plurality of load-bearing struts (106). The inner ring is configured to carry load from inner bearings. The outer ring features a multi-function boss (116) with a service line aperture (124) and a mounting point for the turbine exhaust case. The load-bearing struts connect the inner ring to the outer ring, and have a service line passage (128) extending from the service line aperture to the inner ring.

Abstract: Gas turbine engines and related systems involving air-cooled vanes are provided. In this regard, a representative vane for a gas turbine engine includes: an airfoil having a leading edge, a pressure surface, a trailing edge and a suction surface; and a cooling air channel; the suction surface being formed by an exterior surface of a first wall portion and an exterior surface of a second wall portion, the first wall portion spanning a length of the suction surface between the second wall portion and the trailing edge; the cooling air channel being defined, at least in part, by an interior surface of the first wall portion, the first wall portion exhibiting a thickness that is thinner than a thickness exhibited by the second wall portion.

Abstract: A mid-turbine frame is provided. The mid-turbine frame may comprise an inner case having an annular surface with an interface section disposed on the annular surface. The interface section may include an interface feature. A balancing section may be disposed on the annular surface defining an opening and disposed circumferentially adjacent the interface section.

Abstract: Provided is a gas turbine apparatus including: a turbine unit comprising an output shaft; a cooling gas generation unit, comprising a rotation shaft, which receives power from the output shaft through the rotation shaft and generates a compressed cooling gas; a first duct unit which transfers the generated compressed cooling gas to the turbine unit; a clutch unit which controls a power transfer connection between the output shaft and the rotation shaft; and a control unit which controls the transferring of the generated compressed cooling gas.

Abstract: A thermal management system for a gas turbine engine includes a high pressure compressor rotor having a plurality of disks and a shaft; at least one forward cavity between a rim of at least one of the disks and the shaft and at least one aft cavity formed between a rim of at least one other of the disks and the shaft; a first flow of cooling air in the at least one forward cavity; and a second flow of cooling air in the at least one aft cavity.

Abstract: The invention provides a gas turbine cooling system that efficiently cools a first-stage turbine wheel and the attachment portions of first-stage turbine blades. A gas turbine comprises: a compressor 1 for compressing air; a combustor 2 for combusting the air compressed by the compressor 1 with fuel; and a turbine 3 driven by the combustion gas generated by the combustor 2, the turbine 3 including at least one turbine wheel 12a having turbine blades 11a on an outer circumferential section thereof. A gas turbine cooling system comprises a group of impingement cooling holes 32, provided in a partition wall 16a that separates the exit space 42 of the compressor 1 and the wheel space 43a located upstream of the turbine wheel 12a, for ejecting the compressed air 101 in the exit space 42 toward the turbine wheel 12a and the attachment portions of the turbine blades 11a.

Abstract: A converging-diverging nozzle that has particular application for providing a cooling air flow to ring segments in a gas turbine engine. The engine includes a turbine section that receives a hot working gas. The turbine section includes at least one row of vanes, at least one row of blades and a plurality of ring segments forming at least one ring. The ring segments and the vanes are mounted to a vane carrier, where the vane carrier includes a cooling flow channel for each of the ring segments that receives an air flow to cool the ring segments. A plug is provided in each channel and has an internal bore shaped to define the converging-diverging nozzle through which the air flows so as to create a supersonic flow that reduces the temperature of the air and thus provides more cooling for the same amount of air flow.

Abstract: A method for operating a hydrogen-fueled gas turbine is provided wherein a supply of fuel is passed to a gas turbine combustor, and a supply of nitrogen and sufficient air to provide at least sufficient compressed air to the gas turbine for fuel combustion is passed to a compressor. A sufficient portion of the compressor discharge flow is passed to a combustor for fuel rich combustion of the fuel flow to the combustor and the fuel is combusted to produce hot combustion gases that are, in turn, passed to a turbine.

Abstract: An interface assembly for a combustor includes an interface housing having a channel defined by a forward wall and at least one aft wall segment, the aft wall segment operatively coupled to an aft flange of a flow sleeve. Also included is a piston ring fittingly disposed in the channel.

Abstract: The present invention relates to an isothermal compression type heat engine using air as a heat working medium. It is technically characterized as follows. An air compressor is used to replace an adiabatic air compressor. Two engineering courses being air compression and expansion work are performed separately in different engine members, so as to significantly increase a working pressure of the heat engine. Temperature gradient type heat preservation composite tubes are disposed to recycle remaining heat of exhaust gas with high efficiency, so as to distinctly improve the heat efficiency of the heat engine.

Abstract: A rotational machine such as a turbocompressor has a fluid recovery system for recovering leaked working fluid such as gaseous helium in a helium circuit which has leaked past a shaft seal, a purifier being provided for removing contaminants from the working fluid, and turbocompressor may have one fluid such as helium or hydrogen working through one turbo component such as a turbine thereof and a second working fluid such as air or helium working through a second turbo component such as a compressor thereof, the rotational machine being installable in an engine of a flying machine.

Abstract: A fan blade for a gas turbine engine includes an airfoil that includes leading and trailing edges joined by pressure and suction sides to provide an exterior airfoil surface that extends in a radial direction to a tip. The external airfoil surface is formed in substantial conformance with multiple cross-sectional profiles of the airfoil described by a set of Cartesian coordinates set forth in Table 1. The Cartesian coordinates are provided by an axial coordinate scaled by a local axial chord. A circumferential coordinate is scaled by the local axial chord, and a span location. The local axial chord corresponds to a width of the airfoil between the leading and trailing edges at the span location.

Abstract: Provided is a gas turbine system capable of dealing with a request for output increase even when high-pressure hot water generated using solar thermal energy cannot be used according to the operating state of the gas turbine system. A gas turbine system which sucks in intake air from an air intake duct by a compressor and drives a gas turbine by combustion gas obtained by burning air and fuel by a combustor, said gas turbine system being provided with pipes for generating high-pressure hot water by providing a solar collecting tube that utilizes solar heat and spraying the high-pressure hot water into the intake air sucked in by the compressor, and pipes for spraying normal temperature water into the intake air sucked in by the compressor.

Abstract: A method for positioning sensors about a sensor ring includes the steps of assigning each sensor in a plurality of sensors a sensor number selected from a set of sensor numbers, where the set of sensor numbers is a whole number in the range of 0 to N, and where N is the total number of sensors in said plurality of sensors minus one, disposing a first sensor at a circumferential angular position zero on the sensor ring, and disposing each sensor in the plurality of sensors at a circumferential angular position about the sensor ring, wherein the circumferential angular position is defined by an offset from a circumferential angular position zero and the offset is equal to a base arc length between sensors multiplied by the sensor number of the sensor plus a base offset arc length multiplied by the sensor number of the sensor.

Abstract: The present application and the resultant patent provide a dual fuel combustor for a gas turbine engine. The combustor may include a primary premixer positioned within a head end plenum of the combustor, and a dual fuel, injection system positioned within the head end plenum and upstream of the premixer. The injection system may be configured to inject a gas fuel about an inlet end of the premixer when the combustor operates on the gas fuel. The injection system also may be configured to vaporize and inject a liquid fuel about the inlet end of the premixer when the combustor operates on the liquid fuel. The present application and the resultant patent also provide a related method of operating a dual fuel combustor.

Abstract: A method of recovering heat energy from a cooling medium used to cool hot gas path components in a turbine engine includes cooling one or more hot gas path components with the cooling medium; supplying spent cooling medium used to cool the one or more hot gas path components to a heat exchanger; supplying air (e.g., compressor discharge air) to the heat exchanger so as to be in heat exchange relationship with the spent cooling medium and thereby add heat to the compressor discharge air; and supplying the air heated in the heat exchanger to at least one combustor.

Abstract: A component formed by an additive manufacturing process includes a body and a first vibration damper. The body is formed from an additive manufacturing material, and defines at least a first cavity completely enclosed within the body. The first vibration damper is disposed within the first cavity. The first vibration damper includes a flowable medium and a first solidified element formed from the additive manufacturing material. The flowable medium surrounds the first solidified element.

Abstract: A gas turbine engine section has a rotor carrying a plurality of blades. The blades have airfoils which define a radially outer tip. A blade outer air seal is positioned radially outwardly of the tips of the blades. The blade outer air seal is provided by at least a plurality of circumferentially spaced segments, which slide circumferentially relative to each other to adjust an inner diameter of an inner surface of the blade outer air seal segments. An actuator actuates the blade outer air seal segments to slide towards each other to control a clearance between the inner periphery of the blade outer air seal segments and the radially outer tip of the blade airfoils. A gas turbine engine is also disclosed.

Abstract: An inner barrel member for a gas turbomachine includes a body having an outer surface, an inner surface and one or more radial flow splitters provided on the outer surface. The one or more radial flow splitters are configured and disposed to be arranged along a combustor centerline at a combustion flow outlet radially inwardly of a transition piece.

Abstract: This invention concerns a system for cooling components in a gas turbine engine, the gas turbine engine including a compressor for driving a primary gas flow to a combustor and a turbine arranged to be driven by combustion gases from the combustor, wherein the system includes: an annular cooling flow passage arranged for fluid communication between the compressor and the turbine, the flow passage having a first inlet arranged to receive gas from the primary gas flow downstream of compressor, and a second inlet located upstream of the first inlet, wherein the annular cooling flow passage has at least one internal wall for guiding airflow from the first inlet towards the airflow from the second inlet, the airflow from the first and second inlets coalesce within the annular flow passage prior to passing along the passage in a direction from the compressor to the turbine.

Abstract: An improved combustor for a gas turbine is operable to provide high combustion efficiency in a compact combustion chamber. The combustor includes a counter swirl doublet for improved fuel/air mixing. The enhanced combustor assembly and method of operation improves operation of the turbine.

Abstract: A system includes an oxidant compressor and a gas turbine engine. The gas turbine engine includes a combustor section having a turbine combustor, a turbine driven by combustion products from the turbine combustor, and an exhaust gas compressor driven by the turbine. The exhaust gas compressor is configured to compress and route an exhaust flow to the turbine combustor and the oxidant compressor is configured to compress and route an oxidant flow to the turbine combustor. The gas turbine engine also includes an inlet oxidant heating system configured to route at least one of a first portion of the combustion products, or a second portion of the exhaust flow, or any combination thereof, to an inlet of the oxidant compressor.

Abstract: The invention relates to a method for operating a gas turbine which includes a compressor with annular inlet area, at least two burners, a combustion chamber and a turbine. According to the method, at least one first partial intake flow, consisting of oxygen-reduced gas which has an oxygen concentration which is lower than the average oxygen concentration of the compressor intake flow, and at least one second partial intake flow, consisting of fresh air, are fed to the compressor in an alternating manner in the circumferential direction of the inlet area. In addition, the invention relates to a gas turbine power plant with a gas turbine, the compressor inlet of which includes at least one first segment and at least one second segment which are arranged in an alternating manner around a compressor inlet in the circumferential direction, wherein a feed for an oxygen-reduced gas is connected to the first segment and a fresh air feed is connected to the second segment of the compressor inlet.

Abstract: The invention is directed to a process to obtain a compressed gas starting from a starting gas having a lower pressure by performing the following steps: (i) increasing the pressure and temperature of a gas having an intermediate pressure by means of indirect heat exchange against a fluid having a higher temperature to obtain a gas high in pressure and temperature, (ii) obtaining part of the gas high in temperature and pressure as the compressed gas, (iii) using another part of the gas high in temperature and pressure as a driving gas to increase the pressure of the starting gas in one or more stages to obtain the gas having an intermediate pressure for use in step (i). The invention is also directed to a configuration wherein the process can be performed and directed to a process to generate energy using the process.

Abstract: A hybrid electric rotary engine is provided having a pair of rotors separated by a divider and configured for rotation in opposite directions, a timing gear engaged between the inner faces of the rotors, and at least one pair of slanted rotor openings in the rim of each rotor for alignment with at least one pair of slanted divider openings to form a pair of combustion chambers in communication with at least one pair of exhaust chambers for venting of exhaust and to provide rotational thrust.

Abstract: A gas turbine engine system includes a compressor, a combustor, and a turbine. The combustor is coupled to the compressor and disposed downstream of the compressor. The combustor includes a secondary combustor section coupled to a primary combustor section and disposed downstream of the primary combustor section. The combustor also includes a transition nozzle coupled to the secondary combustor section and disposed downstream of the secondary combustor section. The combustor further includes an injector coupled to the secondary combustor section, for injecting an air-fuel mixture to the secondary combustor section. The turbine is coupled to the combustor and disposed downstream of the transition nozzle; wherein the transition nozzle is oriented substantially tangential to the turbine.

Abstract: A thermal/electrical power converter includes a gas turbine with an input couplable to an output of an inert gas thermal power source, a compressor including an output couplable to an input of the inert gas thermal power source, and a generator coupled to the gas turbine. The thermal/electrical power converter also includes a heat exchanger with an input coupled to an output of the gas turbine and an output coupled to an input of the compressor. The heat exchanger includes a series-coupled super-heater heat exchanger, a boiler heat exchanger and a water preheater heat exchanger. The thermal/electrical power converter also includes a reservoir tank and reservoir tank control valves configured to regulate a power output of the thermal/electrical power converter.

Abstract: Embodiments of a Gas Turbine Engine (“GTE”) are provided, as are embodiments of a plasma flow-controlled intake system for deployment on a GTE. In one embodiment, the GTE includes a turbine section, a combustion section upstream of the turbine section, a compressor section upstream of the combustion section, and intake section upstream of the compressor section. The intake section includes a plenum, a first inlet fluidly coupled to the plenum, and a flow-obstructing structure projecting into the plenum and having an outer surface impinged by the airflow directed into the plenum through the first inlet during operation of the GTE. A first array of plasma actuators is disposed on flow-obstructing structure and, when activated, suppresses vortex shedding of the air flowing over the outer surface of the flow-obstructing structure.

Abstract: One embodiment of the present invention is a unique aircraft propulsion gas turbine engine. Another embodiment is a unique gas turbine engine. Another embodiment is a unique gas turbine engine. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for gas turbine engines with heat exchange systems. Further embodiments, forms, features, aspects, benefits, and advantages of the present application will become apparent from the description and figures provided herewith.

Abstract: The invention relates to rotary units and rotary mechanisms that are suitable for use in numerous applications. Rotary units typically include rotational components that are configured to rotate. In some embodiments, for example, multiple rotary units are assembled in rotary mechanisms such that neighboring pairs of rotational components counter-rotate or contra-rotate relative to one another during operation of the rotary mechanisms. Rotational components generally include one or more implements that are structured to perform or effect one or more types of work as the rotational components rotate relative to one another in a given rotary mechanism. In certain embodiments, implements are configured to rotate and/or to effect the movement of other components as rotational components rotate. In some embodiments, engines include rotary mechanisms and are used in, for example, ground vehicles, marine vehicles, aircraft, or devices.