Patent Application
20140039704
The present invention relates generally to systems and methods for the protection of a gas turbine hot gas path and rotor parts from thermal distress.
Improved gas turbine performance may be achieved by running successively higher levels of turbine hot gas path inlet temperature. A common method of achieving increased turbine inlet gas temperature without deteriorating or otherwise distressing hot gas path parts involves applying cooling to turbine component parts. Turbine cooling flow may be bled from the gas turbine compressor and routed directly to the turbine or may be routed through a heat exchanger, or may be taken from an external source. The flow is routed through the secondary flow system to turbine blades and vanes, their platforms, seals, turbine disks, spacers, flanges and other turbine parts. Cooling flow that is routed through the gas turbine rotor may, under normal operation, be metered across seals and/or balanced across multiple seals in a design that ensures that all turbine parts are properly supplied with cooling to maintain metal temperatures within acceptable limits. Abnormal hardware events may result in damage to rotor seals which, in turn, may result in thermal distress of hot gas path parts.
One scenario which may disrupt flow is blockage of the flow paths by a foreign object. This may disrupt cooling flow, thus starving some parts and oversupplying others. With no intervention, this may cause increased metal temperature or inappropriate cooling distribution causing thermal damage, erosion, cracking, etc.
Traditionally, the secondary air flow region of the gas turbine has been monitored through the use of temperature instrumentation. Increased secondary air flow temperatures indicate insufficient cooling, thus a severe protective action may be taken by the turbine controller or the operator to prevent imminent and costly damage. The protective action may be an automatic shut-down or trip or a manual shut-down.
The invention is a model-based approach using the engine control effectors to maintain the gas turbine parts within acceptable limits and allow continued operation at a reduced level of performance. This is achieved by modeling the metal temperatures and supply pressures in substantially real time, and using a controller to adjust the effectors to maintain the desired margin.
According to one exemplary embodiment of the invention, a system for operating a gas turbine, comprises a controller configured to: receive input from a plurality of sensors that sense parameters of the gas turbine during operation; run a first model of the operation of the gas turbine from one or more of the sensed parameters; determine one or more unmeasured variables of the operation from the first model; run a second model of process variables from one or more of the sensed parameters and one or more of the unmeasured variables; determine differences between the process variables and associated boundaries; and adjust one or more effectors of the gas turbine to maintain a predetermined margin between the process variables and hardware physical limits.
According to another exemplary embodiment of the invention, a method of operating a gas turbine comprises receiving input from a plurality of sensors that sense parameters of the gas turbine during operation; running a first model of the operation of the gas turbine from one or more of the sensed parameters; determining one or more unmeasured variables of the operation from the first model; running a second model of process variables from one or more of the sensed parameters and one or more of the unmeasured variables; determining differences between the process variables and associated boundaries; and adjusting one or more effectors of the gas turbine to maintain a predetermined margin between the process variables and hardware physical limits.
Referring to
As used herein, the term “process variables” will be understood to mean hot gas path metal temperatures, gas path and secondary flow path pressures and temperatures, as well as secondary flows and backflow margins. Similarly, the term “unmeasured variables” will be understood to include pressures and temperatures as well as component efficiencies, backflow margins, thrust and airflows.
In addition, it will be understood that that the term “substantially real time” contemplates implementations based on modeled variables that, for example, lead the part temperature(s). In other words, it may be more straightforward and substantially as accurate to control the variables that are faster such as gas path variables rather than certain slower, metal temperature variables that take time to develop.
The control system 14 may be, for example a computer configured to run software programs for performing the required calculations and creating the real time models. It should also be appreciated that the control system 14 may use existing circuits, known programming methods, structures and controllers well within the skill of the art.
The PI controller 30 may be configured to, for example, determine an effect or position 32 to protect the hot gas path and rotor parts of the gas turbine 12 from damage due to engine-to-engine variation, deterioration, mechanical faults, failures or damage to the engine or any of the engine components, etc. and mechanical faults, failures or damage relating to the control system 14 or its components. Effector position selection logic 34 may be provided to select an effector(s) that may be adjusted. The positions(s) 36 of the selected effector(s) is (are) adjusted to maintain the gas turbine parts within acceptable boundaries and allow continued operation at a reduced level of performance without exceeding hardware limits. The substantially real-time modeling of the operation of the gas turbine and the metal temperatures and cooling supply pressures allows the control system 14 to adjust the position of a selected effector to maintain a desired margin between the process variable(s) of the model(s) and the process variable boundaries.
The effectors may be, for example, actuators in the gas turbine that include fuel metering valves, inlet guide vanes, variable stator vanes, variable geometry, bleed valves, starter valves, clearance control valves, inlet bleed heat, and/or variable exhaust valves.
The hot gas path metal temperatures 22, 24, backflow margins, pressures, rotor speeds, actuator or effector positions, and/or flows or other process variables 38 representing the state of the turbine parts are modeled in substantially real time within the control system 14. The model 20 may be physics-based, neural net, or regression-based, or may use variable outputs from a physics-based model. A boundary is defined for each process variable. If a process variable increases and impinges on the boundary, changes to control effector positions 36 are applied.
The gas turbine model 20 may be a model of any physical system, control system, or property of the turbine or turbine subsystem, including but not limited to, the turbine itself, the gas path and gas path dynamics, actuators, effectors, or other controlling devices that modify or change any turbine behavior, sensors, monitors, or sensing systems, the fuel metering system, the fuel delivery system, the lubrication system and/or the hydraulic system. The model 20 may represent each of the main components of the gas turbine engine 12 at a system level, including for example the inlet, fan, compressor, combustor, high pressure turbine, low pressure turbine, afterburner, and variable area exhaust nozzle.
The invention has applications in both the power and aviation industries. In the commercial power industry the ability to continue operating with engine deterioration or damage is advantageous in giving the operator the opportunity to make an informed decision about when to shut down for repairs and, under certain circumstances, will permit the operator to continue selling power to customers. In the aviation industry, a high level of importance is placed on the ability to continue running gas turbines while the aircraft is in flight regardless of their condition. Every opportunity is provided in gas turbine operation to ensure the ability to land at the nearest airport in civilian applications, or to “get home” in military applications with all gas turbines running. Even in multi-engined aircraft, shut down of an engine can significantly impact aircraft flight characteristics. Continued operation with hot gas path part damage can lead to engine failure, and thus is a safety critical issue.
The invention methodology introduces flexibility into the control methodology, including the ability to maintain a higher turbine power level. The invention also facilitates reduction of power to the degree necessary to protect the gas turbine parts, thus reducing the impact on operation of the gas turbine. The invention may also prevent an unnecessary outage of the gas turbine and associated loss of revenue by allowing continued operation at a reduced output until the next scheduled maintenance period.
While the invention has been described in connection with what is presently considered to be the most practical and preferred embodiments, it is to be understood that the invention is not to be limited to the disclosed embodiments, but on the contrary, is intended to cover various modifications and equivalent arrangements included within the spirit and scope of the appended claims.
