Abstract: A fastening system for an aircraft includes first and second parts of the aircraft, a bolt hole including a bolt countersink and being defined by a bolt hole surface of the first part, a nut hole including a nut countersink and being defined by a nut hole surface of the second part, a bolt having a shank including threads, and a bolt head with an undersurface complementarily shaped to the bolt countersink, and a nut having threads and having a nut chamfer complementarily shaped to the nut countersink, the shank being dimensioned relative to the first and second parts, the bolt hole and the nut hole such that the shank is spaced from one of or both of the bolt hole surface and the nut hole surface. A method of fastening a first aircraft part with a second aircraft part using a bolt and a nut is also described.

Abstract: A fastening system for an aircraft includes first and second parts of the aircraft, a bolt hole including a bolt countersink and being defined by a bolt hole surface of the first part, a nut hole including a nut countersink and being defined by a nut hole surface of the second part, a bolt having a shank including threads, and a bolt head with an undersurface complementarily shaped to the bolt countersink, and a nut having threads and having a nut chamfer complementarily shaped to the nut countersink, the shank being dimensioned relative to the first and second parts, the bolt hole and the nut hole such that the shank is spaced from one of or both of the bolt hole surface and the nut hole surface. A method of fastening a first aircraft part with a second aircraft part using a bolt and a nut is also described.

Abstract: The gas turbine compressor for an aircraft gas turbine engine includes a compressor rotor having a plurality of compressor blades circumferentially distributed around a hub. Each of the compressor blades has an airfoil extending radially outward from the hub to a blade tip. A circumferential row of the compressor blades includes two or more different blade types, at least one modified blade of the two or more different blade types having means for generating different shock patterns between adjacent ones of the two or more different blade types when the gas turbine compressor operates in supersonic flow regimes. The means for generating different shock patterns on the modified blade aerodynamically mistune the two or more different blade types.

Abstract: The gas turbine compressor for an aircraft gas turbine engine includes a compressor rotor having a plurality of compressor blades circumferentially distributed around a hub. Each of the compressor blades has an airfoil extending radially outward from the hub to a blade tip. A circumferential row of the compressor blades includes two or more different blade types, at least one modified blade of the two or more different blade types having means for generating different shock patterns between adjacent ones of the two or more different blade types when the gas turbine compressor operates in supersonic flow regimes. The means for generating different shock patterns on the modified blade aerodynamically mistune the two or more different blade types.

Abstract: A compressor rotor for a gas turbine engine is described which includes sets of blades having different airfoil thickness distributions, each including a frequency modifier forming a thickness differential relative to a baseline blade thickness. The frequency modifiers provide different natural vibration frequencies for each of the blades, and facilitate modifying natural vibration frequency separation between adjacent blades of the compressor rotor.

Abstract: A compressor rotor, such as a fan, for a gas turbine engine is described which includes alternating at least first and second blade types. The leading edge of the second blade types includes a leading edge tip cutback extending to the blade tip thereof. The leading edge tip cutback of the second blade type defines a chord length at the blade tip of the second blade types that is less than that of the first blades types. The first and second blade types generate different shock patterns when the fan or compressor rotor operates in supersonic flow regimes.

Abstract: A compressor rotor for a gas turbine engine is described which includes sets of blades having different airfoil thickness distributions, each including a frequency modifier forming a thickness differential relative to a baseline blade thickness. The frequency modifiers provide different natural vibration frequencies for each of the blades, and facilitate modifying natural vibration frequency separation between adjacent blades of the compressor rotor.

Abstract: A compressor rotor for a gas turbine engine is described which includes sets of blades having different airfoil thickness distributions, each including a frequency modifier forming a thickness differential relative to a baseline blade thickness. The frequency modifiers provide different natural vibration frequencies for each of the blades, and facilitate modifying natural vibration frequency separation between adjacent blades of the compressor rotor.

Abstract: A compressor rotor, such as a fan, for a gas turbine engine is described which includes alternating at least first and second blade types. The leading edge of the second blade types includes a leading edge tip cutback extending to the blade tip thereof. The leading edge tip cutback of the second blade type defines a chord length at the blade tip of the second blade types that is less than that of the first blades types. The first and second blade types generate different shock patterns when the fan or compressor rotor operates in supersonic flow regimes.

Abstract: A compressor rotor for a gas turbine engine is described which includes sets of blades having different airfoil thickness distributions, each including a frequency modifier forming a thickness differential relative to a baseline blade thickness. The frequency modifiers provide different natural vibration frequencies for each of the blades, and facilitate modifying natural vibration frequency separation between adjacent blades of the compressor rotor.

Abstract: A multiplexer selects an input signal from one of plural sensors. An A/D converter converts the selected input signal to a digital input value. A processor then determines, based on the digital input value, an output resistance to which the resistance of the selected sensor is to be translated. Finally, the processor controls a digital potentiometer corresponding to the selected sensor to the determined output resistance. In one embodiment, a conditioning circuit adjusts the gain of the input signal according to calibration data, effectively performing at least a part of the translation in the analog domain.