Patent Description:
A gas turbine engine typically includes a fan section, a compressor section, a combustor section and a turbine section. Air entering the compressor section is compressed and delivered into the combustor section where it is mixed with fuel and ignited to generate a high-speed exhaust gas flow. The high-speed exhaust gas flow expands through the turbine section to drive the compressor and the fan section. The compressor section typically includes low and high pressure compressors, and the turbine section includes low and high pressure turbines.

In one type of aircraft de-icing system, bleed air from the compressor section is provided to an aircraft structure susceptible to ice accretion, for example, aircraft wings or an engine fan nacelle inlet. The flow of hot fluid from the compressor section is regulating using a valve. Typically, actuation of the valve is controlled in response to a pressure sensor reading downstream from the valve. Improved robustness over such a system is desired.

<CIT> relates to a method of monitoring two temperature sensors. <CIT> relates to a fuzzy logic fault accommodation control system. <CIT> relates to a flow control arrangement for an intake anti-icing system for a nacelle of a gas turbine engine.

In one exemplary embodiment, an anti-icing system for an aircraft structure that includes a cavity that has an exterior surface subject to ice accretion and a bleed source that is configured to provide a fluid to the cavity via a duct. The system includes multiple temperature sensors, disposed at the aircraft structure, with each temperature sensor configured to detect a temperature associated with an aircraft structure location. A controller is in communication with the temperature sensors. The controller is programmed to compare outputs of the temperature sensors and determine a temperature sensor fault condition.

In a further embodiment of the above, the aircraft structure is a fan nacelle D-duct. The bleed source is a compressor section.

In a further embodiment of any of the above, the temperature sensors are inlet temperature sensors arranged within the D-duct.

In a further embodiment of any of the above, there are at least three inlet temperature sensors.

In a further embodiment of any of the above, the duct is arranged in the fan nacelle. A duct temperature sensor is arranged within the fan nacelle but outside the D-duct. The duct temperature sensor is in communication with the controller. The controller is programmed to determine a burst duct condition based upon an output from the duct temperature sensor.

In a further embodiment of any of the above, first and second valves are arranged in the duct between the bleed source and the cavity. The first and second valves are in communication with the controller. The controller is programmed to regulate at least one of a flow and a pressure of the fluid to the cavity with the first and second valves.

In a further embodiment of any of the above, the controller is programmed to command one of the first and second valves to a full open position and regulate the flow of the fluid with the other of the first and second valves.

In a further embodiment of any of the above, a pressure sensor is in communication with the duct and is arranged downstream from the first and second valves. The pressure sensor is in communication with the controller. The controller is programmed to determine a flow rate of the fluid through the duct in response to an output from the pressure sensor.

In another exemplary embodiment, an anti-icing system for an aircraft structure that includes a cavity that has an exterior surface subject to ice accretion and a bleed source that is configured to provide a fluid to the cavity via a duct. The system includes first and second valves that are arranged in the duct upstream from the temperature sensor. A controller is in communication with the first and second valves. The controller is programmed to command one of the first and second valves to a full open position and regulate the flow of the fluid with the other of the first and second valves in a first condition, and command the other of the first and second valves to the full open position and regulate the flow of the fluid with the one of the first and second valves in a second condition.

In a further embodiment of any of the above, a temperature sensor is disposed at the aircraft structure. The temperature sensor is configured to detect a temperature associated with an aircraft structure location. The controller is in communication with the temperature sensor. The controller is programmed to regulate the flow of fluid in response to an output from the temperature sensor.

In a further embodiment of any of the above, there are multiple temperature sensors at the aircraft structure. Each temperature sensor is configured to detect a temperature associated with an aircraft structure location. The controller is in communication with the temperature sensors. The controller is programmed to compare outputs of the temperature sensors and determine a temperature sensor fault condition.

In a further embodiment of any of the above, the aircraft structure is a fan nacelle D-duct. The bleed source is a compressor section. The temperature sensors are inlet temperature sensors arranged within the D-duct.

In a further embodiment of any of the above, the duct is arranged in the fan nacelle. A duct temperature sensor is arranged within the fan nacelle but outside the D-duct. The duct temperature sensor is in communication with the controller. The controller is programmed to determine a burst duct condition based upon an output from the duct temperature sensor. The controller is programmed to close at least one of the first and second valves in response to the burst duct condition.

In a further embodiment of any of the above, a pressure sensor is in communication with the duct and the controller. The controller is programmed to command the first and second valves in response to an output from the pressure sensor.

In a further embodiment of any of the above, the first and second valves are at least one of a torque motor valve and a pulse width modulator solenoid.

In a further embodiment of any of the above, a bifurcation extends radially inward from the fan nacelle. The first and second valves are arranged in the bifurcation.

In a further embodiment of any of the above, the controller is programmed to determine an aircraft flight cycle. The controller is programmed to alternate between the first and second conditions in alternating aircraft flight cycles.

In a further embodiment of any of the above, the controller is programmed to determine an engine start condition. The controller is programmed to command the first and second valves to full open in the engine start condition.

In a further embodiment of any of the above, the controller is programmed to determine a compressor stall condition. The controller is programmed to command at least one of the first and second valves in the compressor stall condition to provide a desired stall margin to a compressor section.

Alternative engines might include an augmenter section (not shown) among other systems or features and/or may not include a gear reduction in the fan.

The exemplary engine <NUM> generally includes a low speed spool <NUM> and a high speed spool <NUM> mounted for rotation about an engine central longitudinal axis X relative to an engine static structure <NUM> via several bearing systems <NUM>.

The inner shaft <NUM> and the outer shaft <NUM> are concentric and rotate via bearing systems <NUM> about the engine central longitudinal axis X which is collinear with their longitudinal axes.

In a further example, the engine <NUM> bypass ratio is greater than about six (<NUM>:<NUM>), with an example embodiment being greater than about ten (<NUM>:<NUM>), the geared architecture <NUM> is an epicyclic gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about <NUM>:<NUM> and the low pressure turbine <NUM> has a pressure ratio that is greater than about five (<NUM>:<NUM>). In one disclosed embodiment, the engine <NUM> bypass ratio is greater than about ten (<NUM>:<NUM>), the fan diameter is significantly larger than that of the low pressure compressor <NUM>, and the low pressure turbine <NUM> has a pressure ratio that is greater than about five (<NUM>:<NUM>). The geared architecture <NUM> may be an epicycle gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about <NUM>:<NUM>.

The fan section <NUM> of the engine <NUM> is designed for a particular flight condition -- typically cruise at about <NUM> Mach (<NUM>/s) and about <NUM>,<NUM> feet (<NUM>,<NUM> meters). The flight condition of <NUM> Mach (<NUM>/s) and <NUM>,<NUM> ft (<NUM>,<NUM> meters), with the engine at its best fuel consumption - also known as "bucket cruise Thrust Specific Fuel Consumption ('TSFC')" - is the industry standard parameter of lbm of fuel being burned divided by lbf of thrust the engine produces at that minimum point. The low fan pressure ratio as disclosed herein according to one non-limiting embodiment is less than about <NUM>:<NUM>. "Low corrected fan tip speed" is the actual fan tip speed in ft/sec divided by an industry standard temperature correction of [(Tram °R) / (<NUM> °R)]<NUM> (<NUM> °R = <NUM>/<NUM>).

An example anti-icing system <NUM> is illustrated in <FIG>. The system <NUM> includes a bleed source <NUM>, such as a high pressure stage of the compressor section <NUM>, which supplies hot air to a cavity provided by a an inlet lip circumferential duct, sometimes configured with a "D" shape cross-section subsequently referred to as a "D" duct. Though shown half round and circumferential, it could be other shapes and be non-uniform as it traverses the inlet. It may also wrap up to <NUM>% of the inlet or a portion thereof. D-duct <NUM> of a fan nacelle <NUM>. The D-duct <NUM> provides an inlet into the fan section <NUM>. The inlet is subject to ice accretion during some environmental conditions during engine operation. Although the anti-icing system <NUM> is described with reference to a gas turbine engine fan nacelle, the system may be used to de-ice other aircraft structures, such as aircraft wings.

A duct <NUM> extends from the bleed source <NUM> through a bifurcation <NUM> that extends radially inward from the fan nacelle <NUM>. The duct <NUM> is arranged within the fan nacelle <NUM> and carries fluid to a manifold and or nozzle combination <NUM> that distributes the hot air within the D-duct <NUM> to melt the ice from the exterior surface of the D-duct <NUM>. Spent air exits the D-duct <NUM> through a vent <NUM>, which may be provided by fixed or adjustable louvers.

First and second valves <NUM>, <NUM> are arranged in the duct <NUM> between the bleed source <NUM> and the cavity provided by the D-duct <NUM>. The valves are stepper servo valve (torque motors) controlled that can be fully opened or closed in approximately <NUM> seconds or less and moved to a locked open or other fixed position in the event of a valve failure. The first and second valves <NUM>, <NUM> are in communication with a controller <NUM> that is programmed to regulate a flow of the de-icing fluid, for example, bleed air, to the cavity using the first and second valves <NUM>, <NUM>. The controller <NUM> may have dual channels to provide system robustness. In one example, the controller <NUM> is provided by a full authority digital engine control (FADEC) located within the fan nacelle <NUM>. The controller <NUM> can be provided by a single unit or multiple units in communication with one another. Though torque motor controlled valves are illustrated, other electronic controls, or a combination thereof, can be used such as a pulse width modulator (PWM) solenoid. The valves may provid position feedback for back-up position control.

A pressure sensor <NUM> is in communication with the duct <NUM> and is arranged downstream from the first and second valves <NUM>, <NUM>. The pressure sensor <NUM> communicates with the controller <NUM>, which is programmed to determine a flow rate of the fluid through the duct <NUM> in response to an output from the pressure sensor <NUM> and control the first and second valves <NUM>, <NUM>.

A duct temperature sensor <NUM> is arranged downstream from the first and second valves <NUM>, <NUM> and within the fan nacelle <NUM> but outside the D-duct <NUM>. The duct temperature sensor <NUM> is in communication with the controller <NUM>. The controller <NUM> is programmed to determine a burst duct condition using diagnostics <NUM> based upon an output from the duct temperature sensor <NUM>. A temperature above a predetermined threshold is indicative of a burst duct leaking hot bleed air into the nacelle <NUM>. In response to a detected burst duct condition, the controller <NUM> includes a burst duct control mode <NUM> that is programmed to close at least one of the first and second valves <NUM>, <NUM> to cease flow of the bleed air. Temperature reduction after system shutdown indicates the issue was resolved. If the temperature does not drop, the duct temperature sensor <NUM> is likely faulty and the system <NUM> may resume operation.

In one example embodiment, multiple temperature sensors <NUM>, <NUM>, <NUM>, for example, at least three, are arranged within the D-duct <NUM> to measure the temperature of the bleed air at various temperature sensor locations. The temperature sensors are disposed at the aircraft structure, either directly in contact with the metal or adjacent fluid (such as air) temperature. The controller <NUM> is in communication with the temperature sensors <NUM>, <NUM>, <NUM>. The temperature sensors are used to provide feedback to determine if the inlet is sufficiently de-iced and to determine if the valves are faulty and providing too much hot air to the D-duct <NUM>, which could damage the fan nacelle <NUM>.

The controller <NUM> includes diagnostics <NUM> that are programmed to compare outputs of the temperature sensors and determine a temperature sensor fault condition <NUM>. Since the temperature sensors are spread out within the D-duct <NUM> at locations that are necessarily exposed to different bleed air temperatures, a normal operating temperature range or relationship for each sensor may be empirically determined. Temperature sensors closer to the duct <NUM> are expected to detect hotter air that more remote sensors. During a test cycle or during engine operation, the sensors referee one another in the sense that the controller <NUM> can determine if a temperature sensor is faulty based upon good readings from the other temperature sensors. The system <NUM> can continue operating with a bad temperature sensor by using the remaining inlet temperature sensors. Pressure sensor health <NUM> may also be determined.

During system operation, the controller <NUM> is programmed to command one of the first and second valves to a full open position and regulate the flow of the fluid with the other of the first and second valves in a first condition. The controller <NUM> command the other of the first and second valves to the full open position and regulate the flow of the fluid with the one of the first and second valves in a second condition. If one of the valves is determined to be faulty by the diagnostics <NUM>, a valve control mode <NUM> can be used to lock the bad valve open and the other valve can be used to regulate the flow of bleed air through the system.

The entire system <NUM>, or portions of the system, can be checked at engine start up to determine if the valves and sensors are functioning properly. In one example, the controller <NUM> is programmed to determine an aircraft flight cycle, which is one take-off and landing. The controller <NUM> is programmed to alternate between the first and second conditions in alternating aircraft flight cycles. In this manner the use and wear to of the first and second valves <NUM>, <NUM> are more evenly distributed.

The de-icing system <NUM> can be actuated manually by the pilot using a switch <NUM>. Alternatively, the system <NUM> is actuated automatically using an algorithm that predicts ice accretion based upon inputs from sensors <NUM>, such as outside ambient temperature, low pressure compressor inlet air temperature, high spool speed, environmental control system status and wing anti-ice status. The system may be programmed to achieve maximum inlet temperature at take-off.

The de-icing system <NUM> provides other functionality during engine operation. The controller <NUM> is programmed to determine a compressor stall condition <NUM> in the compressor section <NUM>. For example, the controller <NUM> programmed to command at least one of the first and second valves <NUM>, <NUM>, typically by opening the valves, in the compressor stall condition to provide a desired stall margin to a compressor section <NUM>. The controller <NUM> is also programmed to enable an engine start condition <NUM>. The controller <NUM> is programmed to command the first and second valves <NUM>, <NUM> to full open in the engine start condition, which enables the compressor section <NUM> to rotate more rapidly making the engine easier to start.

Primary control of the system is based on closed loop control satisfying the outputs of the temperature sensors, inferring or directly measuring metal temperature. With this scheme, it is possible to limit the rate of change of metal temperature to reduce transient stress, and or preheat the inlet prior to takeoff. Additionally, this could also be used to shock the ice. Backup alternate control can be accomplished based on closing the loop on pressure sensor <NUM>. This pressure combined with known downstream pressure losses and inferred or measured duct temperature can be used to calculate flow. The system can also be controlled to known valve positions based on position or angle transducers.

It should also be understood that although a particular component arrangement is disclosed in the illustrated embodiment, other arrangements will benefit herefrom. For example, more or fewer pressure and/or temperature sensors than disclosed may be used. 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 invention.

Although the different examples have specific components shown in the illustrations, embodiments of this invention are not limited to those particular combinations.

Claim 1:
An anti-icing system (<NUM>) comprising:
an aircraft structure that includes a cavity having an exterior surface subject to ice accretion and a bleed source (<NUM>) configured to provide a fluid to the cavity via a duct (<NUM>), wherein the duct delivers fluid to the aircraft structure;
multiple temperature sensors (<NUM>, <NUM>, <NUM>), disposed at the aircraft structure, each temperature sensor (<NUM>, <NUM>, <NUM>) configured to detect a temperature associated with an aircraft structure location;
a controller (<NUM>) in communication with the temperature sensors (<NUM>, <NUM>, <NUM>), the controller (<NUM>) programmed to compare outputs of the temperature sensors (<NUM>, <NUM>, <NUM>) and determine a temperature sensor fault condition (<NUM>);
at least two valves (<NUM>, <NUM>) arranged within the duct and configured to regulate fluid to the aircraft structure;
wherein the temperature sensors (<NUM>, <NUM>, <NUM>) are disposed at the aircraft structure downstream from the duct; and
wherein the controller (<NUM>) is configured to receive feedback from the multiple temperature sensors (<NUM>, <NUM>, <NUM>) to determine at least one of a de-iced condition and a faulty valve condition, wherein the at least two valves comprise first and second valves (<NUM>, <NUM>), wherein the controller (<NUM>) is programmed to command one of the first and second valves (<NUM>,<NUM>) to a full open position and regulate the flow of the fluid with the other of the first and second valves (<NUM>, <NUM>) in a first condition, and to command the other of the first and second valves (<NUM>, <NUM>) to the full open position and regulate the flow of the fluid with the one of the first and second valves (<NUM>, <NUM>) in a second condition; and
if one of the first and second valves (<NUM>, <NUM>) is determined to be faulty by the controller (<NUM>), a valve control mode (<NUM>) is configured to lock the faulty valve open.