Abstract: A circumferentially extending structure 30 for containing particles having an axial component of velocity and a radial component of velocity is disclosed. The structure includes an inner casing 48 and an outer casing 50. The outer casing is formed of a plurality of layers of flexible fabric and includes layers having differing masses per unit length in the fabric. The layers of fabric extend about the inner casing. A means for applying a load to the fabric engages the fabric to place the fabric in tension. One such means is a plurality of bars 80, 82 and turnbuckles 74. In one embodiment, the fabric has a greater mass per unit length in the leading edge region and a greater tension is applied to the leading edge region than other regions of the fabric. A method of forming the structures is disclosed.

Abstract: A rotor telemetry apparatus 60 for providing a data communication link between a rotating member 68 and a stationary member 66 of a rotary machine is disclosed. The rotor telemetry apparatus includes a pair of coils 72 each mounted on an associated member for transferring electrical energy between the stationary member 66 and rotating member 68 of the turbomachine. At least one coil of the pair of coils is held by a first layer 88 of bonding agent and is attached to one of the members by a second layer 90 of bonding agent which is filled with a magnetizable material having a magnetic permeability under operating conditions that is greater than the magnetic permeability of the bonding agent. A method of assembling the coils to the stationary and rotating members is also disclosed.

Abstract: A rotor stage assembly 20 of the type adapted for use in an axial flow gas turbine engine is disclosed. The assembly includes a rotor disk 36 having blade attachment slots 42 configured to receive a plurality of rotor blades 40. Each rotor blade is insertable into the disk in a generally axial direction. Each rotor blade is retained against fore and aft movement by a lock pin 52. The disk is adapted by a groove 44 and the blade is adapted by a groove 48 to receive the lock pin. During assembly the lock pin is slidable in the grooves 44,48 to bring the lock pin into engagement with the corresponding groove in a rotor blade in a corresponding groove in the disk.

Abstract: A detector-transducer 26 for sensing temperatures in an engine is disclosed. A temperature signal generated by a temperature sensing device 28 is biased by a supplementary thermocouple 30. A resistor 52 is in series with the temperature sensing device. The supplementary thermocouple is in parallel with the resistor 52. In one embodiment the temperature sensing device is the junction of a second thermocouple.

Abstract: A pressure probe 26 for sensing total pressure in a duct 16, 22 bounding a flow path 12 for a working medium fluid is disclosed. The probe has an alignment portion 32, a sensor portion 34 spaced from the alignment portion leaving a plenum 36 therebetween and passageways 41 for ducting the flow away from the plenum. In one particular embodiment a transition passageway 54 extends between the plenum and the passageway 41. The alignment portion has a first passageway 42 directed towards a second passageway 40 in the sensor portion. The sensor portion has a wall 38 disposed about the second passageway 40.

Abstract: A nonrotating flow directing assembly 14 for a gas turbine engine is disclosed. The flow directing assembly is formed of a circumferentially segmented inner case 24 supported by an annular sleeve 22. The inner case 24 is formed of a plurality of arcuate segments 26 extending axially continuously through the engine. A method of assembling the circumferentially continuous annular sleeve about an axially continuous rotor is also disclosed. In an alternate embodiment the inner case 124 is formed of several pluralities of arcuate segments 126.

Abstract: A flow directing assembly 14 having an airfoil section or shape 28 of the type adapted for use in an axial flow gas turbine engine is disclosed. The cambered meanline MCL of the airfoil shape is formed of a front circular arc FA and a rear circular arc RA. A thickness distribution TD is applied to the meanline to form the convex suction surface 20 and the concave pressure surface 22. The airfoil section exhibits good aerodynamic performance as compared with an equivalent circular arc airfoil in a transonic flow field. A method for making the airfoil shape is disclosed. The method includes the steps of: forming a cambered meanline of two circular arcs; forming a thickness distribution about the conical chord line B.sub.t ; and applying thickness distribution to the cambered meanline such that a portion of the suction surface is stretched and a portion of the pressure surface is compressed.

Abstract: A nonrotating flow directing assembly 14 for a gas turbine engine is disclosed. The flow directing assembly is formed of a circumferentially segmented inner case 24 supported by an annular sleeve 22. The inner case 24 is formed of a plurality of arcuate segments 26 extending axially continuously through the engine. A method of assembling the circumferentially continuous annular sleeve about an axially continuous rotor is also disclosed. In an alternate embodiment the inner case 124 is formed of several pluralities of arcuate segments 126.

Abstract: An annular ring 22 for an axial flow gas turbine engine disposed inwardly of an array of stator vanes 14 and between an annular flow path 18 for hot working medium gases and an annular flow path 20 for cooling air is disclosed. The ring abuttingly engages the array of stator vanes in the radial direction to provide radial support and abuttingly engages the array of stator vanes and an outer air seal block the intermixing of gases from the working medium flow path and gases from the cooling air flow path. The annular ring includes a first annular section 28 and a second annular section 30. The first annular section has an axial width to radial height ratio greater than one and the second annular section has an axial width to radial height ratio less than one to adapt the ring for axial flexibility and radial stiffness and to enable the ring to accommodate growth and axial growth between the outer air seal and the array of stator vanes while remaining within the elastic limit.

Abstract: A hollow rotor shaft 34 having a bore 36 sealed by an end cap 40 is disclosed. The end cap abuttingly engages a first side 42 and second side 44 on the shaft which extend radially inwardly into the bore of the shaft. The sides face in opposite directions and are oriented in a generally axial direction. The end cap includes at least one dog 48, a cover 46 pressed tightly against the first side by the rotatable dog and a lock means such as the bolt 50. The dogs are rotatable to a first position during insertion and rotatable to a second position during assembly by the lock means which extends through the cover to permit assembly and disassembly of the end cap from the exterior of the rotor shaft. In one embodiment a gasket 72 is trapped between the side of the shaft and the cover by the side of the shaft and the cover.

Abstract: A fan case 30 which extends circumferentially about a fan rotor assembly 26 in a turbofan gas turbine engine 10 is disclosed. The fan case has a shell 66 outwardly of the rotor assembly for blocking the rotor blade 38 during a blade failure from penetrating the portion of the shell immediately outwardly of the rotor blade. The material of the shell increases in thickness in the axially rearward direction and has a maximum thickness rearward of a reference plane passing through the mid-chord point of the rotor blade.

Abstract: A fixture 34 for positioning a shroud 22 of a worn rotor blade 14 with respect to an apparatus for removing material is disclosed. The fixture is formed of a base 42, an axial locator 44 which is adapted to engage the root 18 of the rotor blade, and a tangential locator 46 which engages the root of the rotor blade at a hinge joint 58 and which engages the airfoil 20 of the rotor blade. A method of restoring a shroud of the rotor blade axially with respect to the root and tangentially with respect to the airfoil is also disclosed. In an alternate embodiment, a second locator 160 for indexing the locator from an airfoil edge is included.

Abstract: A shroud ring (22) for a gas turbine engine having a stator structure (14) and an array of inwardly extending stator vanes (20) is disclosed. The shroud ring is circumferentially segmented to accommodate inward and outward movement of the vanes and the stator structure. The shroud ring has a segmented outer ring (28) and a segmented inner ring (30). The outer ring engages the vanes to support the vanes and to position the inner ring and an integral seal member (24) about a rotor structure. In one embodiment the vanes (20) are rotatable about a spanwise axis S. In a second embodiment, the vanes (120) are integrally attached to the shroud ring (122).

Abstract: A rotor stage assembly 20 of the type adapted for use in an axial flow gas turbine engine is disclosed. The assembly includes a rotor disk 36, at least two rotor blades 42 and a scalloped lock pin 52. The scalloped lock pin has lugs 54. The disk is adapted by a groove 40 and each blade is adapted by a groove 48 to receive the scalloped lock pin. During assembly the scalloped lock pin is alignable with blade attachment slots 44 in the disk to enable insertion of the blades and is slidable in the grooves 40, 48 to bring each lug 54 on the scalloped pin into engagement with a corresponding groove in a rotor blade.

Abstract: A combined powerplant 10 employing a gas turbine engine cycle 12 and a steam turbine engine cycle 14 is disclosed. The powerplant includes a partial gasifier 22 and an auxiliary combustor 20. The powerplant utilizes the heating value of coal in a two-stage combustion process by burning the gases derived from the coal within the gas turbine engine and by burning the char derived from the coal externally of the engine in the auxiliary combustor. Heat is transferred from the auxiliary combustor to the respective working mediums of both engines.

Abstract: An apparatus 10 for consuming coal in powering a gas turbine engine 12 is disclosed. The apparatus includes an auxiliary combustor 16 and a partial gasifier 18. The gasifier 18 produces combustible char and combustible fluids. The combustible fluids are burned in the gas turbine engine. The combustible char is burned in the auxiliary combustor to form high temperature gases. Heat is transferred from the high temperature gases to the working medium flow path of the gas turbine engine. The exhaust from the gas turbine engine provides combustion air to the fluidized bed combustor.

Abstract: A sound absorbing structure for use in gas turbine engines is disclosed. Various construction details are developed for mounting the liner of the structure to enable the damping of acoustically caused vibrations and to isolate panels of the liner from mechanically caused vibrations. A viscoelastic material suspends the sound absorbing panels of the liner from the case.

Abstract: A coolable rotor blade assembly 10 of the type adapted for use in an axial flow gas turbine engine is disclosed. The assembly includes a rotor blade 12 and an impingement tube 14. A fiber metal seal 56 is disposed between the base of the rotor blade and the impingement tube. The seal contacts a seat 52 on the rotor blade and a face 54 on the impingement tube. The fiber metal seal has sufficient durability, flexibility, compressibility and density to enable sealing between the rotor blade and the impingement tube in a high temperature environment.

Abstract: A case assembly for supporting stator vanes of a gas turbine engine is disclosed. Various construction details and techniques which enable support of the vanes across the flow path of the engine are discussed. The construction details and techniques which provide damping to the vanes and accommodate differential thermal expansion between the vanes and the case assembly are developed. The vane mounting system disclosed is built around the concept of enabling sliding friction between components of the case assembly to dampen vane vibrations.

Abstract: A compression section of a gas turbine engine having an annular flow path is disclosed. Various construction details which increase the efficiency of an array of rotor blades in the compression section are developed. The annular flow path is contoured to cause the streamlines of the flow path to follow a pattern of varying radial curvature. In one embodiment, a conical surface extending between the base of each airfoil on the inner wall causes a flow path contraction and a cylindrical surface on the outer wall facing the tip of each airfoil enables close clearances.