A system includes an engine manufacturer database communicatively coupled to a blockchain database through a network and a ground station configured to wirelessly communicate with a communication adapter of a gas turbine engine of an aircraft. The communication adapter includes a communication interface configured to communicate with an engine control of a gas turbine engine. The system is further configured to monitor the blockchain database for a configuration update associated with the aircraft and update the engine manufacturer database based on the configuration update. The system is further configured to command a synchronization of the configuration update from the engine manufacturer database to a communication adapter of the gas turbine engine tracked by the engine manufacturer database and transmit the configuration update wirelessly to the communication adapter through the communication interface to update a data storage unit of the gas turbine engine with the configuration update.

A bumpless transfer fault tolerant control method for aero-engine under actuator fault is disclosed. For an aero-engine actuator fault, by adopting an undesired oscillation problem produced by an active fault tolerant control method based on a virtual actuator, in order to solve the shortage of the existing control method, a bumpless transfer active fault tolerant control design method for the aero-engine actuator fault is provided, which can guarantee that a control system of the reconfigured aero-engine not only has the same state and output as an original fault-free system without changing the structure and parameters of a controller, to achieve a desired control objective, and that a reconfigured system has a smooth transient state, that is, output parameters such as rotational speed, temperature and pressure do not produce the undesired transient characteristics such as overshoot or oscillation.

Variable geometry turbines having new configurations of vanes are disclosed. New methods for designing new configurations of vanes for geometry turbines having are also disclosed.

A system for determining and/or controlling motor thrust and engine thrust in a parallel hybrid aircraft. One or more sensors may be configured to monitor one or more flight parameters to generate sensor information. User input including one or more pilot estimates may be received. The sensor information may be obtained. A performance thrust ratio may be calculated based on the user input, the sensor information, an aerodynamic model, a propeller model, and a battery model. The performance thrust ratio may be used to control the motor thrust and engine thrust to improve utilization of electric energy throughout a flight. A first thrust setting for the motor and/or a second thrust setting for the engine may be determined based on the performance thrust ratio.

In an assembly method for a turbine, measured shape data is acquired by measuring a shape for each of a plurality of casing components in a state in which the plurality of casing components are not fastened to each other. self-weighted state shape data, which is shape data when self-weight is applied, is created for each of the plurality of casing components. A reference shape model is corrected based on a difference between the measured shape data of a target measurement part and the self-weighted state shape data of the target measurement part. By using the corrected shape model, fastened state shape data, which is shape data in a state in which the plurality of casing components are fastened to each other, is estimated for each of the plurality of casing components.

A control system for turbine systems configured to utilize an intelligent model of particulate presence and accumulation within turbine systems to address engine maintenance, erosion, corrosion, and parts failure mitigation is disclosed. The control system may build an intelligent model of fluid flow based on the data value measured by at least one sensor and based on a database of known data values to provide an estimation of amount of ingress of air intake particles into the turbine system, fouling within the turbine system, erosion of at least a portion of the turbine system, and performance degradation rate of the turbine system.

A dispatch advisor for operating a power plant having at least one gas turbine with flexibility is described. The dispatch advisor can generate a representation of a flexible base load map for operating the power plant. The representation can include an aggregation of a primary base load operating space and an expanded portion of the base load operating space. The representation offers a range of operating values for operational parameters of the power plant during base load at various base load settings at predetermined ambient conditions and corresponding power output and efficiency values that are attained while operating the power plant at the range of operating values. This offers an operator of the power plant with flexibility in controlling the plant during base load.

A system includes a gas turbine system including a compressor, a combustor, a turbine, and an exhaust section. The system also includes multiple sensors coupled to components of the gas turbine system. The system further includes a controller communicatively coupled to the gas turbine system and the multiple sensors and configured to control operations of the gas turbine system, wherein the controller is configured to calculate a surrogate value for turbine exit Mach number based on the feedback from the multiple sensors and to utilize the surrogate value to derive a control action for the gas turbine system.

The method of assembling the rotor assembly can include obtaining geometrical reference values about the individual rotor components, based on the geometrical reference values, determining a combination of relative circumferential positions of the individual rotor components associated to a bow shape configuration of the centers of mass along the axially-extending sequence; and assembling the rotor components to one another in said determined combination of relative circumferential positions, into the rotor assembly.

A method for balancing a set of blades intended to be arranged on a bare disc of an aircraft engine, the bare disc comprising a defined number of numbered cells (ai) intended to receive the same defined number of blades, which can have a spread of mass, the method comprising the following steps:—sorting the blades by monotonic order of their mass (mi) to form an ordered set of blades,—separating the ordered set of blades in a balanced manner into four lobes constituted by a first large lobe, by a second large lobe, by a first small lobe and by a second small lobe, the blades being classified into each lobe according to a current placement order, and—arranging the four lobes on the bare disc by making the current placement order of the blades correspond to the numbered cells of the bare disc.