A zero running clearance system for gas turbine engines includes a gas turbine engine compressor (14) having one or more first rotor assemblies (30) and one or more first stator assemblies (28) disposed adjacent the one or more first rotor assemblies (30); a gas turbine engine turbine (18) having one or more second rotor assemblies (104) and one or more second stator assemblies (102) disposed adjacent the one or more second rotor assemblies (104); one or more shroud segments (80) having blade seal surfaces (120) disposed radially outside of the tips of blades of both the first and second rotor assemblies (30;104) and a clearance distance extending between the blade tips and blade seal surfaces; and, an actuator (58) selectively operable to axially move the first and said second rotor assemblies (30;104) to alter the clearance distance.

Gas turbine engine blade tip clearance apparatus and method

A shroud (16) includes suspension apparatus (50), a control ring (52) and a blade outer air seal (54). The suspension apparatus is compliant and is disposed between the control ring 52 and the case 32. The control ring (52) carries the seal (54). The suspension apparatus comprises a hoop (56) from which extend, circumferentially, a number of cantilevered leaves (64) which are attached to the control ring (52).

Rotor blade outer tip seal apparatus

A turbine blade shroud assembly for a gas turbine engine includes a metal substrate ring on the engine, a continous ceramic barrier ring inside the substrate ring and exposed to hot gas in a hot gas flow path of the engine, and a wire mesh compliant ring between the barrier and substrate rings. The temperature of the barrier ring increases faster than the temperature of the substrate ring as the temperature in the hot gas flow path increases. The coefficient of thermal expansion of the substrate ring is less than the coefficient of thermal expansion of the barrier ring so that the barrier ring expands relative to the substrate ring with increasing temperature in the hot gas flow path and development of tensile hoop stress in the ceramic barrier ring is minimized.

Turbine blade shroud assembly

Described is reinforcement of a fan case (1002) in a gas turbine jet engine. In one embodiment, a containment ring (1102) and a heat resistance ring (1112) are shrink interference fit on the inside diameter of the fan case, the containment ring where the large fan blades turn, and the heat resistance ring where heated air from backfiring heats up the fan case. In one example, the containment ring is made of a super alloy to provide added strength to the fan case should a fan blade break, containing the fan blade within the fan case. The heat resistance ring is made of titanium or other suitable material. Additionally, one or more stiff ener rings (1104,1106) may be shrink interference fit on the outside diameter of the fan case. The containment ring and stiffener rings can reduce the flight weight of the fan case and lower the material costs, while increasing the containment strength of the fan case. Other embodiments are described and claimed.

Fan case reinforcement in a gas turbine jet engine

In one embodiment, a low or high pressure turbine case (204) is machined on its outside surface to form circumferential notches (302). The notches may coincide with internal locations of seals (210) for the blades (208), or with 'hot spots' that have been identified, for example. A stiffener ring (304) may be shrunk with an interference fit into each notch (302) through inducing temperature differentials between the ring and the case. The radially compressive circumferential force exerted by each ring can inhibit the low or high pressure turbine case from expanding as much as it would otherwise. In some applications, a stiffener ring can improve blade tip clearance or counterbalance 'hot spots', stiffen the case, improve case cooling, or other benefits, depending upon the particular application. In one embodiment, notches may be avoided. In an alternate embodiment, C-rings, or multiple segmented rings, may be coupled together by hydraulic, electrical, or other means and actuated by a controller to exert adjustable radially compressive circumferential force. Other embodiments are described and claimed.

Turbine case reinforcement in a gas turbine jet engine

A gas turbine engine blade tip seal consists of a sealing ring which is controlled by means of two annular control members The members are arranged such that one control member has a relatively rapid thermal response rate and the other control member has a relatively slow thermal response rate, the sealing ring being controlled such that a preferred tip clearance is maintained under varying engine operating conditions

Roy Talbot Hirst

Papers

Gas turbine engines