Abstract: A control for a hydraulic propeller pitch actuator includes a low power pilot valve responsive to digital electronic input signals to control movement of a hydraulic servomotor driving a higher power servovalve. Negative feedback motion is delivered from the servomotor to both the pilot valve and the servovalve.

Abstract: A fluidic generator (38) employing a fluidic feedback oscillator (12) in combination with a magnetogenerator (41) is described. Oscillational changes in pressure produced within the feedback oscillator are transmitted to opposite sides of a pole cap (74) which reciprocates a coil (70) in a magnetic field.

Abstract: A metallic regenerator seal is provided having multi-layer coating comprising a NiCrAlY bond layer, a yttria stabilized zirconia (YSZ) intermediate layer, and a ceramic high temperature solid lubricant surface layer comprising zinc oxide, calcium fluoride, and tin oxide. An array of discontinuous grooves is laser machined into the outer surface of the solid lubricant surface layer making the coating strain tolerant.

Abstract: An actuating system effects rotation of wing slat and flap panels of an aircraft and includes actuators having arms rigidly mounted to the panel and allowing some wing-to-panel rotation during wing/panel bending, and actuators having arms mounted to the panel to allow angular displacement of the arms relative to the rigidly mounted arms, whereby structural deflection of the frame of the aircraft is inhibited from imposing unanticipated destructive loads on either the actuating system components or the aircraft frame itself.

Abstract: A multifunction secondary power system (10) for use aboard an aircraft or other vehicle provides the functions both of an auxiliary power unit and of an emergency power unit. The power system (10) includes two turbine engines (12, 22) operable under differing conditions, a gearbox (34) receiving power from the two engines (12, 22) selectively through overrunning clutches (74,84), and a plurality of accessory devices (36, 38, 40) receiving shaft power from either of the two engines (12, 22) through the gearbox (34). The gear box (34) includes a third clutch (64) operably interposed between the clutches (74, 84) which selectively delivers shaft power from the engine (22) to the engine (12) to affect the start thereof.

Abstract: An annular gas turbine engine combustor includes additional axial primary airflow openings, formed as arcuately segmented openings on opposite sides of each of the airblast fuel nozzles of the combustor, for reducing overall combustor diameter while maintaining flame stabilization and avoiding unacceptable thermal variations in the combustion process.

Abstract: A gas turbine engine has a low cost combustor liner cooling system combining the benefits of high internal heat removal with improved film cooling by employing a large number of strategically positioned, laser-drilled cooling passages. Cooling air flows through these specially tailored passages to absorb heat from the liner prior to injection as a protective film on the interior surface. The passages are set in staggered rows on a thickened portion of the liner and have a rough internal heat transfer surface and an exit with a steep injection angle to evenly distribute the cooling film along the interior surface of the liner.

Abstract: To prevent propeller overshoot, a propeller governor control generates a signal which modifies a speed command issued to a propeller governor. The signal causes the propeller governor to change blade pitch before the propeller reaches its commanded speed. The propeller governor control also compensates for small calibration errors in the propeller governor.

Abstract: A power management system for turbine engines includes a plurality of switches, each allowing for engine speed and engine torque to be controlled together. The switches correspond to various modes of operation of the turbine engine. When a switch is selected, a mode signal is supplied to a FADEC which, in response, issues speed and torque commands to an associated engine. The FADEC includes function generators that issue the torque commands as a function of outside ambient conditions. The torque commands are continuously updated for changing ambient conditions. Further, takeoff torque can be corrected for heating conditions. The power management system can also provide torque commands as a linear function of power lever angle.

Abstract: A turbine or compressor blade designed for deliberate, controlled failure in a gas turbine engine during testing of the engines, includes a lower-strength tubular insert in an axial through bore in the root portion of the blade, which insert fails at preselected conditions to induce blade failure.

Abstract: A turbine or compressor blade designed for deliberate, controlled failure in a gas turbine engine during testing of the engines, includes a lower-strength tubular insert in an axial through bore in the root portion of the blade, which insert fails at preselected conditions to induce blade failure.

Abstract: A combustion system for a gas turbine engine provides a reaction limiting means for preventing the combustion process within the combustion chamber of the engine from proceeding towards completion by limiting the reaction rate or duration of the reaction process sufficiently such that a clean combustion gas is produced having substantially less carbon particles than for a combustor effective for essentially complete combustion. The invention also provides an efficient method of removing carbon deposits which do form by controlling the combustor shutdown technique in such a way as to cause the carbon to be oxidized. The preferred embodiment provides an aircraft emergency power unit having a high pressure air storage tank, an aviation fuel storage tank, and a combustor with a reaction limiting means which combusts pressurized air and aviation fuel in a fuel rich ratio to produce a motive combustion gas in a reaction limiting means.

Abstract: A fuel control system for accurately controlling the fuel flow rate to a gas turbine engine during the initial stages of startup is provided. The fuel control system has a novel, closed control loop that measures the actual fuel flow rate to the primary nozzles of the engine, and uses this measurement to trim the commanded fuel flow rate signal generated by preprogrammed schedules in the engine's electronic control unit. By the novel arrangement of proportional and integral controllers, the fuel control system maintains a forward path gain of unity while permitting the independent adjustment of the gain of the controllers to achieve the system stability and response necessary for quick correction of fuel flow rate errors.

Abstract: An aircraft is maintained at a constant glide slope and air speed by trimming delta torque commands provided by an autopilot. Delta torque commands that change by more than twenty percent are trimmed automatically to maintain the constant glide angle and to assist the pilot in maintaining air speed within .+-.5 knots on the approach glide slope.

Abstract: A turbine rotor is formed from a turbine disk having a rim with a circumferential direction, and a plurality of turbine blade segments each fixed to the turbine disk around a circumference of the rim of the turbine disk. The turbine blade segments are oriented such that the elastic modulus of the turbine blade segments parallel to the circumferential direction is less than that of the rim of the turbine disk parallel to the circumferential direction. This arrangement is preferably achieved using single crystal turbine blade segments made of an an alloy having a cubic crystal structure, such as a nickel-based superalloy, with a <010> direction oriented radially and a <001> direction oriented circumferentially.

Abstract: A bicast vane and shroud rings is provided in which thermal stresses are minimized. The vane has a flange portion along its inner and outer edge. These flange portions extend from the mid-chord of the vane part way towards the leading and trailing edges and follow the camber of the vane. Each of the flange portions has a width greater than that of the vane so has to form an overhang or lip about which conventional shroud rings are cast. As the flange portions approach the leading and trailing edges their width decreases and the overhang blends into the inner and outer edge to form rounded tops and bottoms at the leading and trailing edge. These rounded tops and bottoms reduce stress concentrations in the shroud rings and increase shroud material near these edges. Additionally, a stress relieving member is provided that embraces the top and bottom edges of the thin trailing edge, and inhibits the melting of the trailing edge during bicasting.

Abstract: A gas turbine engine comprised of a gas generator in flow series arrangement with a free turbine is provided, that includes a bleed port disposed between the gas generator and the free turbine. The flow of hot gas through the free turbine is controlled by a valve operably disposed in a duct downstream of the bleed port. The gas turbine engine can rapidly change its power output by first bleeding a portion of the hot gas exiting the gas generator overboard, and then increasing or decreasing the amount of bleed flow, thereby producing rapid changes in the free turbine's output.

Abstract: A high temperature ceramic/metallic turbine engine includes a metallic housing which journals a rotor member of the turbine engine. A ceramic disk-like shroud portion of the engine is supported on the metallic housing portion and maintains a close running clearance with the rotor member. A ceramic spacer assembly maintains the close running clearance of the shroud portion and rotor member despite differential thermal movements between the shroud portion and metallic housing portion.

Abstract: A centrifugal compressor impeller for gas turbine engines designed to efficiently operate at both part-speed and full speed design points includes slanted vanes in a secondary air inlet to the impeller. The vanes are tangentially slanted in the direction of rotation of the impeller to preferentially augment airflow into the impeller at maximum speed, while at part-speed discouraging airflow out of its impeller through the secondary inlet.