Abstract: A surgical instrument includes an anvil and an elongate channel. The elongate channel includes a plurality of first electrical contacts and a plurality of electrical connectors comprising a plurality of second electrical contacts, wherein the electrical connectors are spring-biased such that a gap is maintained between the first electrical contacts and the second electrical contacts. The surgical instrument further includes a staple cartridge releasably attachable to the elongate channel, wherein the staple cartridge has a cartridge body comprising a plurality of staple cavities, a plurality of staples deployable from the staple cavities into the tissue, and a plurality of third electrical contacts, wherein the attachment of the staple cartridge to the elongate channel moves the electrical connectors causing the second electrical contacts to bridge the gap and become electrically coupled to the first electrical contacts.

Abstract: Disclosed herein are systems and methods for treating a magnesium or a magnesium alloy substrate. The system includes a cleaning composition having a pH greater than 8.5 or a solvent, and a pretreatment composition comprising an organophosphate compound and/or an organophosphonate compound. The method includes contacting a substrate surface with the cleaning composition or the solvent, and contacting the surface with the pretreatment composition. Also disclosed are substrates treated with one of the systems or methods. Also disclosed are magnesium or magnesium alloy substrates wherein, between the air/substrate interface and 500 nm below the air/substrate interface (a) oxygen and magnesium are present in a combined amount of at least 70 atomic %; (b) carbon is present in an amount of no more than 30 atomic %; and/or (c) fluoride is present in an amount of no more than 8 atomic %, wherein atomic % is measured by XPS depth profiling.

Abstract: A test rig that reproduces high temperature and high pressure conditions found in a gas turbine engine for testing materials under these conditions. The test rig includes a combustor surrounded by an upper plenum chamber and an exhaust plenum that receives a hot gas stream from the combustor. A transparent channeling vessel guides the hot gas stream from the combustor over a test specimen on which a material to be tested is mounted. The hot gas stream exits the clear channeling vessel and into the exhaust plenum where the stream is cooled by diluting the stream with cooling air. A viewing portal is located in the exhaust plenum so that the material can be observed through the clear vessel.

Abstract: A small gas turbine engine for use in an UAV such as a cruise missile, the gas turbine having a combustor forming a primary burn zone and a secondary burn zone, and in which fuel is injected into both the primary and the secondary burn zones by either a rotary cup injector or a plurality of fuel injector nozzles. The secondary burn zone with separate fuel injection allows for the diameter of the engine to be reduced in size but still allow for adequate power and efficiency to be reached for powering the vehicle. Air flow from the compressor is used to cool the combustor walls before being injected into the combustor, and to pass through and cool the guide nozzles and a main bearing located near the hot section of the combustor prior to being introduced into the combustor.

Abstract: A test rig that reproduces high temperature and high pressure conditions found in a gas turbine engine for testing materials under these conditions. The test rig includes a combustor surrounded by an upper plenum chamber and an exhaust plenum that receives a hot gas stream from the combustor. A transparent channeling vessel guides the hot gas stream from the combustor over a test specimen on which a material to be tested is mounted. The hot gas stream exits the clear channeling vessel and into the exhaust plenum where the stream is cooled by diluting the stream with cooling air. A viewing portal is located in the exhaust plenum so that the material can be observed through the clear vessel.

Abstract: A small gas turbine engine for use in an UAV such as a cruise missile, the gas turbine having a combustor forming a primary burn zone and a secondary burn zone, and in which fuel is injected into both the primary and the secondary burn zones by either a rotary cup injector or a plurality of fuel injector nozzles. The secondary burn zone with separate fuel injection allows for the diameter of the engine to be reduced in size but still allow for adequate power and efficiency to be reached for powering the vehicle. Air flow from the compressor is used to cool the combustor walls before being injected into the combustor, and to pass through and cool the guide nozzles and a main bearing located near the hot section of the combustor prior to being introduced into the combustor.

Abstract: A small gas turbine engine for use in an UAV such as a cruise missile, the gas turbine having a combustor forming a primary burn zone and a secondary burn zone, and in which fuel is injected into both the primary and the secondary burn zones by either a rotary cup injector or a plurality of fuel injector nozzles. The secondary burn zone with separate fuel injection allows for the diameter of the engine to be reduced in size but still allow for adequate power and efficiency to be reached for powering the vehicle. Air flow from the compressor is used to cool the combustor walls before being injected into the combustor, and to pass through and cool the guide nozzles and a main bearing located near the hot section of the combustor prior to being introduced into the combustor.

Abstract: A small gas turbine engine where the compressor and turbine are supported on a rotary shaft, and a main bearing is supported on the rotary shaft, the main bearing being located in a hot zone of the combustor. The main bearing includes cooling air passages within the races to provide cooling for the bearing. A cooling air is diverted from the compressor and passed through the bearing cooling passages for cooling the bearing, and then the cooling air is directed into the combustor. The cooling air is also passed through a guide nozzle before being passed through the bearing to cool both the guide nozzle and the bearing. A swirl cup injector is used to deliver the compressed air from the compressor and the cooling air from the bearing into the combustor, the swirl cup injector also acting to draw the cooling air through the bearing.

Abstract: A small gas turbine engine with a compressor, a combustor, and a turbine located downstream of the combustor. The compressor and turbine are supported on a rotary shaft, and a main bearing is support on the rotary shaft, the main bearing being located in a hot zone of the combustor. The main bearing includes cooling air passages within the races to provide cooling for the bearing. A cooling air is diverted from the compressor and passed through the bearing cooling passages for cooling the bearing, and then the cooling air is directed into the combustor. The cooling air is also passed through a guide nozzle before being passed through the bearing to cool both the guide nozzle and the bearing. A swirl cup injector is sued to deliver the compressed air from the compressor and the cooling air from the bearing into the combustor, the swirl cup injector also acting to draw the cooling air through the bearing.

Abstract: A small gas turbine engine with a bearing cooling and lubricating passage arrangement for a high speed rotor shaft. The engine includes a bypass fan and a compressor. The rotor shaft includes a central passage extending through the entire shaft, and where the rotor shaft is supported by a forward bearing and a rearward bearing. Cooling air for the bearings is diverted from the bypass air and is channeled through the bearings. Fuel is added to the cooling air at a location upstream of the bearing to provide lubrication. The cooling air and lubricating fuel passes through the bearings and into the rotating central shaft, and is then forced to flow toward a radial passage located adjacent to the combustor. The fuel is collected on the central shaft surface and forced out the radial passage and into the combustor. The cooling air continuous out from the central shaft to be mixed with the engine exhaust.

Abstract: A small gas turbine engine for use in an UAV such as a cruise missile, the gas turbine having a combustor forming a primary burn zone and a secondary burn zone, and in which fuel is injected into both the primary and the secondary burn zones by either a rotary cup injector or a plurality of fuel injector nozzles. The secondary burn zone with separate fuel injection allows for the diameter of the engine to be reduced in size but still allow for adequate power and efficiency to be reached for powering the vehicle. Air flow from the compressor is used to cool the combustor walls before being injected into the combustor, and to pass through and cool the guide nozzles and a main bearing located near the hot section of the combustor prior to being introduced into the combustor.

Abstract: A rotary cup fuel injector for use in a gas turbine engine, the rotary cup injector being rotatably secured to the rotor shaft and positioned radially inward of the combustor, the rotary cup injector having a front face with a parabolic shape and a rear face with a slanted and flat shape. The front and rear faces of the injector form a thin film of fuel on the surfaces and—because of the high rotational speed—produce fine droplet's of fuel to be injected into a combustion chamber. The two faces inject fuel into two combustion zones of the combustor.

Abstract: A small gas turbine engine for use in an UAV such as a cruise missile, the gas turbine having a combustor forming a primary burn zone and a secondary burn zone, and in which fuel is injected into both the primary and the secondary burn zones by either a rotary cup injector or a plurality of fuel injector nozzles. The secondary burn zone with separate fuel injection allows for the diameter of the engine to be reduced in size but still allow for adequate power and efficiency to be reached for powering the vehicle. Air flow from the compressor is used to cool the combustor walls before being injected into the combustor, and to pass through and cool the guide nozzles and a main bearing located near the hot section of the combustor prior to being introduced into the combustor.

Abstract: A small gas turbine engine with a compressor, a combustor, and a turbine located downstream of the combustor. The compressor and turbine are supported on a rotary shaft, and a main bearing is support on the rotary shaft, the main bearing being located in a hot zone of the combustor. The main bearing includes cooling air passages within the races to provide cooling for the bearing. A cooling air is diverted from the compressor and passed through the bearing cooling passages for cooling the bearing, and then the cooling air is directed into the combustor. The cooling air is also passed through a guide nozzle before being passed through the bearing to cool both the guide nozzle and the bearing. A swirl cup injector is sued to deliver the compressed air from the compressor and the cooling air from the bearing into the combustor, the swirl cup injector also acting to draw the cooling air through the bearing.

Abstract: A combustor system for a miniature gas turbine engine includes fuel injection via orifices which direct fuel into fuel-air injection tubes which feed a fuel-rich mixture of fuel and air into a leading end of the combustor liner to form a primary burning region. Fuel system pressures are kept low and controlled by control of the fuel injection port size and number. Fuel breakup is via airblast and tube wall impingement. The fuel-air injection tubes are directed circumferentially, radially outward and toward the front end of the combustor. Air is fed into the combustor such that two distinct burning regions are created. Each region approximates a “well-stirred reactor” and the combination of the two regions results in an efficient use of combustion volume.

Abstract: A combustor system for a miniature gas turbine engine includes fuel injection via orifices which direct fuel into fuel-air injection tubes which feed a fuel-rich mixture of fuel and air into a leading end of the combustor liner to form a primary burning region. Fuel system pressures are kept low and controlled by control of the fuel injection port size and number. Fuel breakup is via airblast and tube wall impingement. The fuel-air injection tubes are directed circumferentially, radially outward and toward the front end of the combustor. Air is fed into the combustor such that two distinct burning regions are created. Each region approximates a “well-stirred reactor” and the combination of the two regions results in an efficient use of combustion volume.

Abstract: Accuracy and predictability are provided in a fuel control system for an engine system including a fuel supply and a combustion chamber. The fuel control system includes at least one fuel injector having a calibrated nozzle to direct a fuel flow from the fuel supply into the combustion chamber of the engine, a servo valve connected between the fuel supply and the calibrated nozzle to meter the fuel flow from the fuel supply to the calibrated nozzle, a pressure transducer operably associated with the calibrated nozzle to measure a pressure drop in the fuel flow across the calibrated nozzle, and a controller operably associated with the pressure transducer and the servo valve to control the metering of the fuel flow by the servo valve. The calibrated nozzle has a known fuel flow vs. pressure drop characteristic. The controller controls the metering of the fuel flow by the servo valve based on the pressure drop measured by the pressure transducer and the known fuel flow vs.

Abstract: A reverse flow annular combustor is provided having downstream wall comprised of a shroud covering the outer surface of the outer transition liner. The shroud abuts the transition liner at a plurality of points to define a plurality of passages between the shroud and the liner. The passages receiving cooling through the shroud at its radial inner end and expel the cooling air through the liner at its outer radial portion and into the hot gas in the combustor. Thus, the air used to cool the liner is also used for dilution mixing with the combustion gas.