Patent Application
20130183136
The present disclosure relates to gas turbine engines, and more particularly to a start system and method therefor.
Gas turbine engines are typically started through the high spool with an air turbine starter or electric starter/generator which drives a towershaft through a gearbox. In flight, accessories and the generator are driven from this high-spool-mounted gearbox. In such an arrangement, horsepower is extracted from the high spool, which may debit high compressor stability. Thus, one challenge in start systems for gas turbine engines concerns the design of an operable system, within the available space in the gas turbine engine, to improve compressor stability.
A gas turbine engine according to an exemplary aspect of the present disclosure includes a high spool along an engine axis, the high spool operable to communicate a core airflow through a core flow path. A fan section is operable to communicate an airflow into a secondary flow path and the core flow path. A low spool is located along the engine axis to drive the fan section. A flow control mechanism is operable to direct at least a portion of the airflow from the secondary flow path into the core flow path.
In a further embodiment of the foregoing gas turbine engine, the flow control mechanism is movable between an open position and a closed position.
In a further embodiment of the foregoing gas turbine engine, the flow control mechanism is a fan duct blocker ring.
In a further embodiment of the foregoing gas turbine engine, the flow control mechanism is a nozzle.
A further embodiment of the foregoing gas turbine engine includes a starter/generator connected to the low spool.
In a further embodiment of the foregoing gas turbine engine, the starter/generator is forward of the low spool.
In a further embodiment of the foregoing gas turbine engine, the starter/generator is aft of the low spool.
In a further embodiment of the foregoing gas turbine engine, the starter/generator is driven by the low spool through a towershaft.
In a further embodiment of the foregoing gas turbine engine, the high spool along the engine axis with a high pressure compressor section and a high pressure turbine section, the high pressure compressor section includes a variable guide vane.
In a further embodiment of the foregoing gas turbine engine, the high spool along the engine axis with a high pressure compressor section and a high pressure turbine section, the high pressure turbine section includes a variable guide vane.
In a further embodiment of the foregoing gas turbine engine, the fan section includes a variable pitch fan blade.
In a further embodiment of the foregoing gas turbine engine, the fan section includes a variable pitch vane.
In a further embodiment of the foregoing gas turbine engine, the high spool along the engine axis with a high pressure compressor section and a high pressure turbine section, a start bleed in selective communication with the high pressure compressor section.
In a further embodiment of the foregoing gas turbine engine, the low spool drives the fan section through a geared architecture.
A method of operating a gas turbine engine gas turbine engine according to a further exemplary aspect of the present disclosure includes driving a low spool to generate a fan airflow from a fan section connected to the low spool and modulating a flow control mechanism to direct at least a portion of the fan airflow from a secondary flow path into a core flow path to drive a high spool.
In a further embodiment of the foregoing method, no accessory gearbox is driven by the high spool.
A further embodiment of the foregoing method includes rotating the high spool in a direction opposite that of the low spool.
A further embodiment of the foregoing method includes driving the fan section through a geared architecture driven by the low spool.
A further embodiment of the foregoing method includes driving the low spool with a starter/generator.
A further embodiment of the foregoing method includes driving a constant speed transmission with the low spool.
Various features will become apparent to those skilled in the art from the following detailed description of the disclosed non-limiting embodiment. The drawings that accompany the detailed description can be briefly described as follows:
The fan section 22 and the LPT 30 are coupled by a first shaft 36 to define a low spool 38. The HPC 24, the combustor section 26 and the HPT 28 are generally referred to as an engine core. The HPC 24 and the HPT 28 are coupled by a second shaft 40 to define a high spool 42. In the disclosed, non-limiting embodiment the low spool 38 and the high spool 42 counter-rotate about axis X.
The low spool 38 drives the fan section 22 either directly or through a geared architecture 38G (illustrated schematically) to drive a fan 44 such as a multi-stage fan with a first fan stage 44A and a second fan stage 44B, at a lower speed than the low spool 38. It should be appreciated that a low pressure compressor may alternatively or additionally be coupled to the first shaft 36 of the low spool 38.
An engine case structure 48 generally includes an outer case structure 50, an intermediate case structure 52 and an inner case structure 54. It should be understood that various structures individual or collectively within the engine may define the case structures 50, 52, 54 to essentially define an exoskeleton that supports the spools 38, 42 for rotation therein. The outer case structure 50 and the intermediate case structure 52 define a generally annular secondary flow path 56. The intermediate case structure 52 and the inner case structure 54 define a core flow path 58 radially within the secondary flow path 56.
A starter/generator 70, which, for example, may be in the Megawatt class for military weapon systems, or in the kilowatt class for civilian and military “all-electric” aircraft as well as others, is connected to the low spool 38. In one disclosed non-limiting embodiment, the starter/generator 70, is located aft of the LPT 30 within a tail cone T. Alternatively, the starter/generator 70, may be located within a fan hub 22H (
Air which enters the fan section 22 is divided between a core airflow and a fan airflow. The core airflow passes through the HPC 24, the combustor section 26, the HPT 28, the LPT 30 then the augmentor section 34 where fuel may be selectively injected and burned to generate additional thrust through the nozzle section 32. The fan section 22 communicates the fan airflow into the secondary flow path 56 as well as the core flow path 58 as the fan 44 is upstream of a split 52S between the secondary flow path 56 and the core flow path 58. The fan airflow may be utilized for a multiple of purposes to include, for example, cooling, pressurization and VSTOL operations. The fan airflow passes through an annulus defined by the outer case structure 50 and the intermediate case structure 52 then may be at least partially injected into the core flow adjacent the nozzle section 32.
The nozzle section 32 receives flow from both the secondary flow path 56 and the core flow path 58. It should be understood that various fixed, variable, convergent/divergent, two-dimensional and three-dimensional nozzle systems as well as Variable Area Fan Nacelle (VAFN) may be utilized herewith.
The fan airflow as defined herein is any airflow different from the primary combustion gas core airflow through the engine core and may otherwise be described as secondary airflow or bypass air flow. Although depicted as an augmented low bypass gas turbine engine with a bypass ratio of approximately 0.7 in the disclosed non-limiting embodiment, it should be appreciated that the concepts described herein are applicable to other gas turbine engines such as high bypass engines with a VAFN.
The fan airflow is selectively communicated through a flow control mechanism 60 into the secondary flow path 56. The flow control mechanism 60 may include various structures such as rotationally variable splitter vanes, pneumatic or mechanical operated blocker doors, or other structures that operate as a choke point to define a variable area throat and selectively control flow into the secondary flow path 56 such that a selective percentage of flow from the fan section 22 is divided between the secondary flow path 56 and the core flow path 58.
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A module 202 may be utilized to execute the start mode algorithm 200. In one non-limiting embodiment, the module 202 may be an engine FADEC, a portion of a flight control computer, a portion of a central vehicle control, an interactive vehicle dynamics simulator unit or other system. The module typically includes a processor; a memory and an interface. The processor may be any type of known microprocessor having desired performance characteristics. The memory may be computer readable medium which stores the data and control algorithms described herein. The interface facilitates communication with the engine 20 as well as other avionics and vehicle systems.
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With the flow control mechanism 60 closed, approximately two-times (2×) or other multiple of the torque is provided to the high spool 42 from the fan airflow which reduces the work required to spin the high spool 42. This essentially simulates an in-air windmill restart as the flow through the core drives the HPT 28 to thereby drive the HPC 24 and pressurize the core flow.
The flow control mechanism 60 directs the fan airflow to accelerate the high spool 42 to a defined light off window to await ignition. When the proper high spool 42 air flow (driven by fuel air ratio) and high spool 42 mechanical speed (driven by stability requirements) reach the defined light-off window, fuel is communicated to the combustor section 26 and ignited. The high spool 42 is then accelerated to a self-sustaining speed.
The flow control mechanism 60 may remain closed until the high spool 42 reaches the self-sustaining speed and generates enough torque to continue independent acceleration. Once this occurs, the flow control mechanism 60 is opened and the starter/generator 70 is shifted into generate mode.
As the low spool 38 has a substantially larger inertia than the high spool 42, the engine 20 may be more tolerant to power extraction. This architecture obviates the need for a starter/generator on the high spool 42 and enables a completely extraction-less high spool 42, which reduces the required surge margin. Hence, the HPC 24 can be designed more aggressively as the high spool 42 need not be designed with extra surge margin to accommodate power extraction associated with an attached motor/generator, to facilitate a more efficient design and potentially fewer stages. This will positively affect Thrust Specific Fuel Consumption (TSFC) and reduce overall engine weight.
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The variable turbine vanes 100, 102 facilitate low spool start modes as—like the variable EGV on the HPC 24—they are modulated to increase airflow during start modes. The variable turbine vanes 100, 102 also may be utilized to change the incidence angle of the flow on the turbine blades 104A, 104B (of the illuminated two stage HPT) to provide for more efficient operation at this off-design condition. That is, the HPT 28 would generate more torque and therefore accelerate the HPC 24 more rapidly for the same airflow.
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In the start mode, the fan section 22 operates at an off-design condition, and may be relatively inefficient such that a selective change in pitch via the variable pitch mechanism 110 adjusts the fan section 22 to increase airflow for same start torque.
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Alternatively, or in addition, a towershaft-driven constant speed transmission 130 mechanically driven by the low spool 38. That is, no power extraction need be obtained from the high spool 42. The constant speed transmission 130 is driven by the low spool 38 such that the speed varies linearly with the speed of the low spool 38 which may operate at speed excursions of up to 80% between idle to max take-off conditions. The constant speed transmission accessory gearbox 130 maintains constant output speed despite speed excursions of the low spool 38. That is, the constant speed transmission 130 provides, for example, a 5:1 continuously variable gear ratio capability which automatically selects the most optimum gear ratio to maintain the constant output speed in response to the low spool 38 speed excursions.
It should be understood that relative positional terms such as “forward,” “aft,” “upper,” “lower,” “above,” “below,” and the like are with reference to the normal operational attitude of the vehicle and should not be considered otherwise limiting.
It should also be understood that like reference numerals identify corresponding or similar elements throughout the several drawings. It should also be understood that although a particular component arrangement is disclosed in the illustrated embodiment, other arrangements will benefit herefrom.
Although the different non-limiting embodiments have specific illustrated components, the embodiments of this invention are not limited to those particular combinations. It is possible to use some of the components or features from any of the non-limiting embodiments in combination with features or components from any of the other non-limiting embodiments.
Although particular step sequences are shown, described, and claimed, it should be understood that steps may be performed in any order, separated or combined unless otherwise indicated and will still benefit from the present disclosure.
The foregoing description is exemplary rather than defined by the limitations within. Various non-limiting embodiments are disclosed herein, however, one of ordinary skill in the art would recognize that various modifications and variations in light of the above teachings will fall within the scope of the appended claims. It is therefore to be understood that within the scope of the appended claims, the disclosure may be practiced other than as specifically described. For that reason the appended claims should be studied to determine true scope and content.
