Abstract: A small gas turbine engine, such as is used to power a UAV, that includes at least one centrifugal compressor having an impeller with blades that form a gap between the blade tips and stationary shroud of the gas turbine engine, and where a resistance heating element, such as an impingement air manifold, is secured to or bonded to a compressor casing of the gas turbine engine in order to use heat to control the gap between the impeller blades and the stationary shroud. The resistance heating element is activated at cruise mode to move the shroud toward the impeller. Additionally or alternatively, the compressor casing is heated with bled-off compressed air to move the shroud toward the impeller. A capacitance tip clearance sensor can be mounted on the impeller shroud to monitor and control tip clearance in real time.

Abstract: A two-dimensional variable area plug (2D VAP) nozzle assembly for a high-speed flight vehicle. In one embodiment, a 2D VAP nozzle assembly comprises a nozzle including a plurality of sidewalls; a plug body within the nozzle, the plug body abutting at least two of the plurality of sidewalls; a first convergent flap hingedly connected to at least one of the plurality of sidewalls; and a second convergent flap hingedly connected to at least one of the plurality of sidewalls. In one embodiment, the nozzle assembly includes only a first convergent flap and a second convergent flap, without diverging flaps. The 2D VAP nozzle assembly has a simplified design with reduced sidewall length, which results in reduced manufacturing and maintenance costs.

Abstract: An aero vehicle with a dual mode propulsion system that can operate in a first mode below a ramjet speed and a second mode at a ramjet speed. The first mode is operated using a compressor driven by an electric motor supplied with electrical power from a power pack. When the vehicle has reached a speed where the ramjet engine can operate, the power pack is dropped from the vehicle to decrease the weight of the vehicle.

Abstract: A geared turbofan engine with at least one compression stage and at least one turbine stage on a high speed shaft, the high speed shaft coupled through a speed reduction gear box to a low speed shaft with a fan and a starter/generator. The low speed shaft is collinear with the high speed shaft but does not rotate within the high speed shaft. The speed reduction gear box is positioned between and mechanically couples the high speed shaft and the low speed shaft, which allows the fan and the integrated starter/generator on the low speed shaft to operate at a lower speed than the high speed shaft.

Abstract: A lay shaft assembly for a speed reduction gearbox in a geared turbofan gas turbine engine, the lay shaft assembly having an output gear formed on a lay shaft as a single piece, and an input welded gear formed as a separate piece that is welded to the shaft in close proximity to the output gear. The output gear is machined and precision-ground prior to welding the input welded gear to the lay shaft so that the two gears can be located in close proximity. The input welded gear is precision-ground after the gear has been welded to the shaft.

Abstract: A geared turbofan engine with at least one compression stage and at least one turbine stage on a high speed shaft, the high speed shaft coupled through a speed reduction gear box to a low speed shaft with a fan and a starter/generator. The low speed shaft is collinear with the high speed shaft but does not rotate within the high speed shaft. The speed reduction gear box is positioned between and mechanically couples the high speed shaft and the low speed shaft, which allows the fan and the integrated starter/generator on the low speed shaft to operate at a lower speed than the high speed shaft.

Abstract: A small gas turbine engine, such as is used to power a UAV, that includes at least one centrifugal compressor having an impeller with blades that form a gap between the blade tips and stationary shroud of the gas turbine engine, and where a resistance heating element is secured to or bonded to a compressor casing of the gas turbine engine in order to use heat to control the gap between the impeller blades and the stationary shroud. The resistance heating element is activated at cruise mode to move the shroud toward the impeller. Additionally or alternatively, the compressor casing is heated with bled-off compressed air to move the shroud toward the impeller. A capacitance tip clearance sensor can be mounted on the impeller shroud to monitor and control tip clearance in real time.

Abstract: A gas turbine engine with a bearing that is supplied with fuel and compressed air for both cooling and lubrication of the bearing. The bearing housing contains a serpentine flow cooling circuit with each leg formed from a plurality of parallel flow channels, and where fuel is supplied to the cooling circuit and then discharged into a combustor. Some of the fuel is also bled off and passed into the bearing compartment along with compressed air bled off from the compressor to both cool and lubricate the bearing.

Abstract: An aero gas turbine engine with a fan stage and a bearing rotatably supporting the fan stage, where the bearing is cooled by passing cooling air through the bearing, and the spent bearing cooling air is re-injected into the air flow path at a location where the effect is minimalized. The location is just downstream from a leading edge of a stator vane and along a pressure side surface of the vane. A fine mist of oil can be discharged into the bearing cooling air to also lubricate the bearing, where both the oil and the spent cooling air is reintroduced into the air flow path.

Abstract: A turbocharged gas turbine engine with an electric generator to provide electrical power for an aircraft (e.g., UAV) with multiple propulsor fans each driven by an electric motor, where the engine includes a low spool that drives a main fan and a high spool that drives a high speed electric generator. The low pressure compressor supplies low pressure air to an inlet of the high pressure compressor. A row of stator vanes in the high pressure turbine is cooled using cooling air bled off from the low pressure compressor outlet that is passed through an intercooler and a boost compressor, where the spent vane cooling air is discharged into the combustor. The low pressure turbine and the two compressors each include a variable inlet guide vane to control the power level of the engine. Bypass flow from the main fan is used to cool hot parts of the engine.

Abstract: An example system may include a computing device including a finite element analysis module solving for a finite element model representing a component including a metal or an alloy and including a plurality of respective elements. The finite element analysis module may solve a respective stress St and a respective temperature Tt at each respective element during the predetermined cooling operating. The finite element analysis module may determine a respective impact energy ET based on the temperature Tt and cooling rate, using a predetermined cooling rate-dependent energy relationship that relates a temperature of the metal or the alloy to an impact energy, determine a respective weakness index Wt=[A×ET/St]n (A being a predetermined constant, n being a predetermined real number greater than or equal to 1), and identify a respective element having a minimum weakness index less than a predetermined weakness index threshold as a cracking-prone element.

Abstract: A variable pitch fan assembly for an aero gas turbine engine that includes a row of fixed pitch blades in a core stream flow and a row of variable pitch blades in a bypass stream flow, and where a sync ring with a number of hydraulic actuators within a flow splitter of the fan stage is used to vary a pitch of each of the bypass stream flow blades. Hydraulic fluid lines pass through the shaft and through the fixed pitch fan blades into the hydraulic actuators to move the sync ring and vary a pitch of the variable pitch bypass flow fan blades.

Abstract: A gas turbine engine for a small aircraft such as a UAV having a bypass flow with a variable area bypass nozzle located at an outlet of the bypass channel, the nozzle having one position with a maximum flow area and a second position with a minimal flow area. The compressor is a twin stream compressor with an inner flow path for compressed air to the combustor and an outer flow path for the bypass channel. A fan stage can be used in front of the compressor.

Abstract: A turbomachine includes a vane, a rotation support coupled to an end of the vane, and a spindle coupled to the rotation support. The spindle, the vane, and the rotation support are rotationally aligned. An annular sleeve defines the spindle. The annular sleeve contacts the rotation support at a radially inward extent and contacts a turbine casing at a radially outward extent. A first rolling element engages the annular sleeve substantially near the radially outward extent. The first rolling element is coupled to the turbine casing. A second rolling element engages the annular sleeve substantially near the radially inward extent. The second rolling element is coupled to an outer endwall ring. A center of mass of the annular sleeve is positioned between the first and second rolling elements.

Abstract: A small aircraft such as a UAV having a gas turbine engine with the outer bypass elements removed from the engine, and where the airframe is modified to form the bypass duct and exhaust nozzle for the fan produced bypass flow of the engine in order to reduce weight, size and cost of the small aircraft. The outer casing and the exhaust nozzle is removed from the engine and optimally integrated into the airframe structure.

Abstract: An example system may include a computing device including a finite element analysis module solving for a finite element model representing a component including a metal or an alloy and including a plurality of respective elements. The finite element analysis module may solve a respective stress St and a respective temperature Tt at each respective element during the predetermined cooling operating. The finite element analysis module may determine a respective impact energy ET based on the temperature Tt and cooling rate, using a predetermined cooling rate-dependent energy relationship that relates a temperature of the metal or the alloy to an impact energy, determine a respective weakness index Wt=[A×ET/St]n (A being a predetermined constant, n being a predetermined real number greater than or equal to 1), and identify a respective element having a minimum weakness index less than a predetermined weakness index threshold as a cracking-prone element.

Abstract: A gas turbine engine includes a compressor assembly that is rotationally coupled to a shaft. The engine includes a radial diffuser assembly coupled to a shroud of the compressor assembly and positioned to receive compressed air from a centrifugal impeller. The radial diffuser assembly includes a first arcuate wall and a second wall, and a near axial diffuser coupled to the radial diffuser assembly and positioned to receive the compressed air from the radial diffuser assembly. The gas turbine includes a gas seal coupled between the second wall and a wall of the near axial diffuser, the gas seal configured to prevent the compressed air from passing through the seal while allowing relative motion between the radial diffuser assembly and the near axial diffuser.

Abstract: A gas turbine engine includes a compressor assembly that is rotationally coupled to a shaft, the compressor assembly having a centrifugal impeller and a shroud covering a bladed portion of the centrifugal impeller. The compressor assembly includes a diffuser that is attached to the shroud via a pair of flanges, the diffuser including a strut that is mounted through an aft-extending leg to a base of an intercase. A sealing assembly is attached to the diffuser and is attachable to a transition duct that is positioned to receive air from the diffuser. The sealing assembly is configured to prevent air from passing through the sealing assembly while allowing relative motion to occur between the transition duct and the diffuser.

Abstract: An adaptive fan system for a variable cycle turbofan engine having at least one turbine, includes a shaft structured to receive mechanical power from a turbine in the variable cycle turbofan engine and a first fan stage that is coupled directly to the shaft and is driven directly by the shaft. A transmission system is coupled to the shaft and a second fan stage is coupled to the shaft via the transmission system and is driven by the shaft via the transmission system. The transmission system is structured to selectively vary a speed between high and low speeds at which power is supplied from the shaft to the second fan stage relative to at least one of the shaft and the first fan stage, wherein the first fan stage and the second fan stage are configured to rotate in the same direction and at the same high speed to increase the mechanical power of the turbine.

Abstract: An exemplary gas turbine engine may include a compressor, a turbine, and a combustor disposed between the compressor and the turbine. The combustor generally may have an inner casing. The exemplary gas turbine may further include a pre-swirl nozzle configured to receive a cooling stream. The cooling stream may be supplied from a cooled cooling air stream. The pre-swirl nozzle may further be configured to direct at least a portion of the cooling stream to the turbine. The pre-swirl nozzle may be flexibly mounted to the inner casing of the combustor.