Abstract: A fuel control system for a gas turbine engine having a staged combustor includes apparatus and methods to operate the combustor during pilot operation, staged operation, and transition between the two modes of operation. Various construction details are developed which provide a fuel control having a switch with built in hysteresis for determining the transition points between operating modes. In one embodiment, the combustor is switched from pilot to staged operation at a thrust level greater than the thrust level at which the combustor is switched from staged to pilot operation. During transitions between operating modes, the changes in fuel flows are limited by a plurality of rate limits to prevent combustor fuel spikes and blow-outs from occurring. During engine starts, the fuel lines to the main combustor are pre-filled to provide responsiveness during the transition from pilot to staged operation.

Abstract: A gas turbine engine combustor can operate on either liquid fuel or gaseous fuel while maintaining low levels of nitrous oxide emissions. An upstream end of each of a plurality of mixing tubes includes an external manifold that is supplied with gaseous fuel. The gaseous fuel is injected into the mixing tubes through a plurality of orifices with sufficient momentum to be uniformly distributed across the flow path of each mixing tube. The optimum fuel/air ratio for low nitrous oxide emissions does not exceed seventy five percent (75%) of the stoichiometric ratio for a given fuel.

Abstract: A linkage arrangement (20) joins a nacelle (12) enclosing a gas turbine engine (10) and a pylon (18) suspended from an aircraft wing. The linkage arrangement (20) counteracts the deflections of the nacelle (12) with respect to the free air stream due to the aerodynamic load. The linkage arrangement (20) includes a plurality of links (22, 24, 32, 34) that substantially duplicate the motion of a four bar linkage mechanism.

Abstract: A brush seal includes an outer retainer ring and a plurality of layers of wire fibers extending inward at a skewed angle from the outer retainer ring and forming a center opening to accommodate an igniter plug therein. The brush seal is mounted onto a combustor wall to allow movement of the igniter plug relative to the combustor wall without damaging the igniter plug and to prevent leakage of air into the combustor.

Abstract: A cleaning method for gas turbine engine airfoils includes a step of immersing an ultrasonic agitator, such as a welding horn, into a tank with a cleaning solution and a step of directing the ultrasonic agitator onto the portion of the airfoil having the crust layer. A subsequent step of high pressure water jet spray removes the crust debris. The cleaning method of the present invention significantly increases the power density of the ultrasonic cleaning.

Abstract: A method of minimizing nitrous oxide emissions in a gas turbine engine includes a step of monitoring pressure fluctuations within a gas turbine engine combustor and a step of adjusting fuel flow to the combustor if the pressure fluctuations therein exceed a preestablished threshold. The combustor can be operated at the lowest levels of nitrous oxide emissions since the onset of a combustor instability and/or blowout is identified early to allow for corrective action so that such undesirable events are prevented.

Abstract: A system for building, testing, and repairing gas turbine engines and for transporting gas turbine engines from one location to another includes a transporter and a build cell. The build cell is a frame structure supported by a plurality of pedestals for mounting the gas turbine engine thereon and is transportable by the transporter to allow assembly and disassembly of the gas turbine engine at any location. The transporter removes the gas turbine engine from either a build cell or another type of a structure at a first location and transports the gas turbine engine either with or without the build cell to a second location and loads the gas turbine engine onto either a build cell or a second structure.

Abstract: An engine case of a gas turbine engine is selectively coated with a thermal barrier coating to control axial clearance between rotating and stationary airfoils. The coating is applied to the thinner portions of the engine case to retard thermal expansion of these portions of the engine case during transient conditions of the gas turbine engine operation. The selectively coated engine case responds substantially uniformly to heating and thermal expansion during transient conditions, thereby reducing axial vane lean in gas turbine engines.

Abstract: A system for repairing damaged airfoils for gas turbine engines includes a plurality of rigid attachment tools. Each attachment tool is individually shaped to access a leading edge of the damaged airfoil of a particular stage of the gas turbine engine. The attachment tools enable repair of the damaged airfoils without disassembling of the gas turbine engine.

Abstract: An airfoil (20) for a gas turbine engine (10) includes cooling passages (40), (50) extending radially within the airfoil to circulate cooling air therethrough. Pluralities of small crossover holes (48), (66), (72) are formed within the walls (50), (68), (74), respectively, to allow cooling air to flow between the cooling passages. Optimum stiffness parameters are provided to improve producability of the airfoils with small geometric features, such as the crossover holes, as well as to improve the overall cooling scheme without jeopardizing manufacturability of airfoils.

Abstract: A hollow airfoil (10) for a gas turbine engine having a leading edge (12), a trailing edge (14), a pressure side (20), and a suction side (22) includes a plurality of internal spanwise stiffening ribs (31-35) that are arranged in a logarithmic pattern. The particular arrangement of internal fibs (31-35) optimizes stiffness of the airfoil (10) without significantly increasing the weight thereof.

Abstract: A thrust reverser (30) of a gas turbine engine (10) includes a blocker door (36) and a plurality of cascades (38). The thrust reverser (30) and the blocker door (36) have a stowed position and a deployed position. In the deployed position the blocker door (36) "leaks" airflow therethrough without generating substantial forward thrust. The leaked airflow reduces the amount of total airflow that must be accommodated by the cascades (38) of the thrust reverser (30), thereby decreasing the overall size of the cascades (38) and of the associated thrust reverser hardware and subsequently reducing the overall weight thereof.

Abstract: A rotor blade for a gas turbine engine includes a bowed surface on a tip region of the suction side thereof. The curvature of the bowed surface progressively increases toward the tip of the blade. The bowed surface results in a reduction of tip leakage through a tip clearance from the pressure side to the suction side of the blade and reduces mixing loss due to tip leakage.

Abstract: A friction damper (50) provides friction damping for gas turbine engine airfoils to reduce vibrations therein. The friction damper (50) includes a plate (52) underlying radially outer shrouds (30) of two adjacent airfoils (20). Friction generated between an outer surface (53) of the plate (52) and undersides (42) of the adjacent shrouds (30) reduces undesirable vibrations in the airfoils (20).

Abstract: Cooling of a turbine airfoil such as a turbine stator vane in a gas turbine engine is enhanced by the provision of additional cooling air through a secondary air inlet communicating with a medial portion of an internal, serpentine cooling air passage.

Abstract: An airfoil for a gas turbine engine, having an airfoil portion with a leading and a trailing edge, includes dual pressure source cooling. The trailing edge of the airfoil includes an internal trailing edge passage and the leading edge includes an internal leading edge passage. Compressor bleed air at higher pressure is channeled through the leading edge passage whereas compressor bleed air at lower pressure is channeled through the trailing edge passage. The higher internal pressure in the leading edge ensures that the inward flow of products of combustion does not occur.

Abstract: A gas turbine engine (10) includes a plurality of rotor blades (24) having a root (33) and a tip (34) and a pressure side (32) and a suction side (30). The rotor blades (24) are enclosed in an engine case (11). A variable position engine case liner (50) is disposed radially inward of the engine case (11) and radially outward of the rotor blade tips (34). Biasing mechanism (60), disposed between the liner (50) and the engine case (11), urge the liner (50) radially inward from the engine case (11). The tips (34) of the rotor blades (24) are angled with respect to the direction of rotation of the blades (24) on the pressure side (32) thereof to deflect pressurized air against the liner (50). The liner (50) adjusts its position to maintain a minimum tip clearance (70) between the liner (50) and the tips (34) of the rotor blades (24) during different operating conditions of the gas turbine engine (10).

Abstract: A gas turbine engine (10) includes a multi-stage combustion chamber (19) having a pilot stage (22) and a main stage (24), wherein during the low power setting only the pilot stage (22) is operating. During the non-operation of the main stage (24) a problem of coking occurs in a fuel supply system in general, and in a check valve (30) distributing fuel to the main stage, in particular. The check valve (30) is cooled by pilot fuel flowing through an auxiliary chamber (52) and is insulated by an air jacket (60) that is formed between the outer housing and the inner housing of the check valve.

Abstract: A gas turbine engine fan blade having an airfoil portion and a dovetail root portion is enhanced by having spaced apart supporting ribs extend into the dovetail root portion. The supporting ribs define a plurality of hollow cavities extending into the dovetail root portion of the fan blade. The extended cavities reduce the weight of each blade and also improve the integrity of each blade.