Abstract: A gas turbine engine which includes a combustion chamber adapted to be supplied with fuel by at least one atomizer nozzle. The gas turbine engine includes an air compressor for supplying compressed air to the combustion chamber, with the compressor including a secondary guide vane set having a plurality of ducts. At least one of the ducts of the secondary guide vane set is connected to the atomizer nozzle by means of a conduit. The at least one duct is equipped with a boundary layer bleed-off.

Abstract: A combustor for burning heavy residual fuel to drive a gas turbine engine has an elongated, imperforate vaporizing tube with an air blast nozzle at one end of the tube adjacent a cannister type preheater means arranged to swirl air blasted residual fuel in the tube to produce vaporization while preventing autoignition and wherein the vaporized fuel is processed through a primary crossed slot air mixer means for reducing the temperature of the air/fuel mixture passing from the vaporizing tube and to establish an equivalence ratio in the air/fuel mixture at a lean reaction zone combustion chamber for burning a lean air/fuel mixture from the primary air mixer means under controlled conditions and wherein a combustor torch igniter directs an igniting flame into the lean air/fuel mixture directed from the air mixer means so as to assure continuous combustion in the lean combustion zone so as to reduce hydrocarbon and carbon monoxide emissions from the combustion apparatus under varying conditions of gas turbine en

Abstract: The fuel flow for a turboprop or turboshaft engine is provided by a plurality of parallel circuits each of which supplies a portion of the fuel to the engine with different valves controlling the flow through each of the separate circuits. In the turboprop engine there is imposed on the parallel flow circuits a sensing device that determines the pressures in certain of the circuits and selects the appropriate pressure for signalling a pressure regulating valve whereby to vary the pressure drop through the fuel control.

Abstract: A gas turbine engine fuel burner comprises a convergent/divergent flow passage which receives a flow of swirling fuel and air upstream of the passage throat. The fuel and air mixture passes through the throat and is entrained into a further mass of swirling air which enters the passage through air swirlers in the wall of the divergent portion of the passage.The invention is intended to promote good fuel and air mixing, fine atomization and a reduction in carbon deposition on the burner.

Abstract: A gas turbine control for ameliorating the adverse affects of cyclic load applications to both single-shaft and two-shaft gas turbines is disclosed. During cyclic load conditions, air flow is governed to control the output of the gas turbine while fuel flow is controlled to hold the firing or exhaust temperatures substantially constant.

Abstract: A rotor assembly and methods for securing rotor blades within and removing rotor blades from rotor assemblies. The rotor assembly comprises a rotor disc defining a plurality of blade grooves, and including a plurality of tenons disposed between the blade grooves and defining a plurality of pin sockets radially extending inward from outside surfaces of the tenons; and a plurality of rotor blades, each blade including a root disposed within a blade groove to secure the blade against radial movement, and a blade platform overlaying a tenon and defining a radially extending pin aperture. The rotor assembly further comprises a plurality of locking pins radially extending through the pin apertures and into the pin sockets to secure the rotor blades against axial movement, each pin including a head and a base to limit radial movement of the pin.

Abstract: A solid propellant rocket motor devised so as to temporarily increase its thrust at a desired stage of its operation. The motor has a first propellant grain which is fixedly disposed in the combustion chamber and is formed with a hole that intersects an inital burning surface of the grain and a second propellant grain which is slidably and gastightly fitted in the hole of the first grain and is shorter in length than this hole. Initially, the second grain is prevented from moving relative to the first grain by a restraining mechanism including, for example, a plunger protruding into an aft end portion of the aforementioned hole to come into abutment with an end face of the second grain.

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 structure for and a method of mechanically locking an abrasive matrix material in honeycomb structure in a shroud for a turbine engine. The mechanical lock may be lanced portions of the honeycomb structure extending into adjacent honeycomb cells to engage the abrasive matrix material. In the method of the invention the lancing is accomplished during the manufacture of the honeycomb structure so that it engages the abrasive matrix material prior to setting thereof in the honeycomb material.

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 rotor blade, for example for a windmill, comprises a plurality of blade ctions between a blade root and a tip. The sections are interconnected by tensioning members such as steel cables or the like, preferably extending through the blade sections. Thus, each section may be manufactured individually and all components may be assembled at the construction site, whereby relatively large windmill rotors may be constructed.

Abstract: A thrust nozzle for a reaction engine having a combustion chamber with a sustained flight discharge nozzle defined therein for the passage of thrust gases and a sustained flight thrust arrangement connected to the combustion chamber for supplying thrust gases through the combustion chamber, the thrust nozzle comprising a starting nozzle formed on an interior wall of the combustion chamber radially inwardly of the sustained flight discharge nozzle made of a plurality of layers of ablatable material which are ablated by the passing thrust gases to discharge vaporized or fragmentary parts of the layers of ablatable material.

Abstract: The object of the invention is to provide more effective film cooling to protect a wall surface from a hot fluid which impinges on or flows along the surface. A film of cooling fluid having increased area is provided by changing the direction of a stream of cooling fluid through an angle of from 135.degree. to 165.degree. before injecting it through the wall into the hot flowing gas.As shown in FIG. 1, cooling fluid is injected from an orifice (16) through a wall (10) into a hot flowing gas (11) at an angle (20) to form a cooling fluid film (12). Cooling fluid is supplied to the orifice (16) from a cooling fluid source (13) via a turbulence control passageway (14) having a curved portion (18) between two straight portions (17 and 19).The angle (24) through which the direction of the cooling fluid is turned results in less mixing of the cooling fluid with the hot gas (11), thereby substantially increasing the length of the film (12) in a downstream direction.FIGS.

Abstract: An annular combustor for a jet turbine engine of generally rectangular cr-section having an axially-projecting aft exhaust outlet is lined by a plurality of circumferentially-extending heat-insulating liner segments, preferably ceramic, spaced from the metal shell of the combustor by spring-like spacers. The liner segments have offset tabs at their edges to slidably mate with each other and with retaining clips adjacent to the exhaust outlet, thereby being slidably secured to accept expansion of the metal shell without placing appreciable stress on the liner segments. Arcuate, generally planer liner segments along the forward and aft walls of the combustor are keyed in place by a radially-inward flanged liner segment along the outer perimetrical wall of the shell.

Abstract: In order to cool the annular upstream wall of a gas turbine engine combustion chamber and thereby to prevent the deposition of carbon particles, the wall is formed of upstream and downstream portions which between them define a chamber arranged to receive a flow of cooling air. The downstream portion has a central opening for an airspray fuel burner and on each side of the opening the downstream portion has a pair of facets each facet having apertures for the throughflow of cooling air over the face of the downstream portion.

Abstract: A fuel control system for a dual-shaft gas turbine engine having a gasifier turbine and a power turbine comprises a device for supplying fuel to a combustor of the engine by an amount necessary for engine ignition during engine starting. Under normal operating conditions, fuel is supplied to the combustor by a closed loop controller. An accelerating device increases the fuel supply to the combustor as time progresses from a first time when engine ignition takes place to a second time when the closed loop control of fuel supply takes place, thereby shortening the time period required for the engine to reach normal operation.

Abstract: A variable turbine inlet nozzle assembly is provided for use in high temperature environments. The nozzle assembly comprises a plurality of vanes and each vane includes a shaft extending outwardly from it. The shafts in turn are rotatably mounted to an annular mounting member so that the vanes are circumferentially spaced from each other and positioned within a gas stream passageway formed within the turbine housing between the combuster and turbine stages. An actuating assembly includes a ring rotatably mounted to the turbine housing concentrically with the mounting member while actuating arms extending between the ring and each shaft selectively rotates or pivots the vanes in unison with each other. Each vane further includes a hollow interior which is fluidly connected with the turbine engine compressor output to provide cooling to and prevent thermal distortion of the vanes.

Abstract: Disclosed is a high-pressure centrifugal blower, which comprises an impeller with a back plate and pressure leveling blades secured to the back side of that back plate. The blades extend from the periphery of the impeller over some distance radially inwards. In order to regulate the pressure control clearance between the blades and the back wall of the blower case, control members have been installed between the blades and the back wall of the case.

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.

Abstract: This disclosure relates to a bearing assembly for a high speed rotating shaft, such as a turbocharger shaft. The bearing assembly includes a sleeve bearing and a ball bearing. The sleeve bearing carries the radial load adjacent one end of the shaft, and the ball bearing carries the radial load adjacent the other end of the shaft. The shaft is subject to thrust loads in either axial direction, and the ball bearing also carries the thrust load in at least one direction.