Patent Application
20240247612
This application relates to a method and apparatus for detecting a possible icing condition associated with an inlet to a gas turbine engine.
Gas turbine engines are known, and typically include a propulsor such as a fan or propeller, delivering air into a core engine. With a fan the air is also delivered into a bypass duct. With a propeller the air is also delivered outwardly of the core engine. In either case the propulsor provides propulsion to an associated aircraft.
In the core engine there is at least one compressor rotor, a combustor and a turbine section. Compressed air is delivered into the combustor where it is mixed with fuel and ignited. Products of this combustion pass downstream over at least one turbine rotor, driving it to rotate. The turbine rotor in turn rotates the compressor and propulsor rotors.
As known, gas turbine engines associated with aircraft may be exposed to extreme temperature conditions. Thus, icing conditions are often encountered by an upstream end of the gas turbine engine associated with the propulsor. Such icing conditions can raise challenges.
During flight an aircraft associated gas turbine engine may be exposed to icing and ice may accumulate or otherwise form on surfaces of an air inlet into the bypass duct or into the core engine. The icing conditions may reduce the performance of the air inlet, and could actually present a risk of damage to downstream engine components if accumulated ice became dislodged.
There are anti-icing systems associated with gas turbine engines on aircrafts. As an example, hot air may be tapped from the compressor to surfaces that may be subject to icing. Moreover, the wings of the aircraft may have anti-icing features that will also provide relief to icing on the gas turbine engines.
A gas turbine engine includes a propulsor for providing air into a core engine housing, and for providing air as propulsion air radially outwardly of the core engine housing. The core engine housing surrounds a compressor section. A combustor is positioned downstream of the compressor section and a turbine section is positioned downstream of the combustor. An aircraft fluid moves within an inlet to the gas turbine engine, and is exposed to inlet air at the inlet. There is a first temperature sensor for sensing a first temperature of the aircraft fluid at a first upstream point and a second temperature sensor for sensing a second temperature of the aircraft fluid at a second downstream point where the aircraft fluid has been exposed to the inlet air for a period of time. A control determines a temperature differential between first and second temperatures sensed by the first and second temperature sensors. The temperature differential is associated with a likelihood that an icing condition will occur, and the control is programmed to take a corrective action should an icing condition be identified.
These and other features of the present invention can be best understood from the following specification and drawings, the following of which is a brief description.
As mentioned above, an engine such as engine 20 may be subject to icing conditions during use on an aircraft. The icing conditions could be a function of a current atmospheric state. As examples, outside air temperature, altitude, air pressure, etc. all contribute to creating potential icing conditions. Also the operational states of the gas turbine engine and the associated aircraft impact conditions. Examples of such operational states include if the aircraft is at cruise, takeoff, landing, etc.
When icing conditions are anticipated an aircraft pilot may take steps to mitigate the risks involved with accumulated ice. As mentioned above, some gas turbine engines may include anti-icing or de-icing equipment which may be operated to reduce ice accumulation.
While mitigating actions and anti-deicing equipment may reduce the risks associated with icing conditions they may also present undesirable limitations on the gas turbine engine operation, as well as reducing gas turbine engine efficiency and performance.
As such, Applicant has recognized that it may be desirable to be able to accurately determine whether a gas turbine engine and air inlet is or is not exposed to icing condition. With an accurate determination the gas turbine engine can be optimally operated during icing conditions and non-icing conditions.
While the upstream location 56 is shown at the radially outer location, it should be understood that should the fluid be flowing radially outwardly the upstream location would be at the radially inner location. Moreover, while the temperature sensors are shown within a strut between a core engine outer housing and inner housing, the teachings of this disclosure could extend to fluids moving through the inlet air flow heading into the engine from any number of locations.
The strut 48 is positioned in the core inlet 32, but it could be positioned in other locations. Moreover, while oil is disclosed in this embodiment, other fluids may be utilized for the general teachings of this disclosure.
At any rate, a location receiving an aircraft fluid that is exposed to the inlet airflow is preferably utilized. Whatever temperature the air flow (here core air C, but again the temperature may be sensed elsewhere) may be at, it will generally cool the fluid within the conduit 50 between locations 56 and 54. The temperature sensors 54 and 56 thus sense a temperature difference which is communicated to a control 100.
The control 100 may be a standalone controller, or may be a full authority digital electronic controller for the overall engine (FADEC). In the exemplary example, the control 100 also controls a deicing system such as ports 102 which receive heated air as from a source 106 which taps air from the compressor section 33 of the gas turbine engine 20.
The control 100 controls a valve 104 to selectively supply this air to ports 102. While this particular deicing system is shown deicing the radially inner surface of the fan case 22, there are a number of other deicing system which could be utilized within the broad teachings of this disclosure. This deicing system is illustrated merely for example.
Applicant has recognized the temperature differential between the sensors 54 and 56 will be a function of various atmospheric and gas turbine engine conditions including outside air temperature, air speed, fluid flow rate, fluid composition (e.g. oil, fuel, pressurized bleed air, etc.) and what condition the gas turbine engine may be at including idle, take off, cruise, etc. By accounting for the various atmospheric and gas turbine engine conditions a differential temperature threshold, say −20° C., may be determined for the measured differential temperature of the fluid. The differential temperature threshold may correspond to a probability that icing conditions exist for the inlet at the current atmospheric and operational conditions.
The differential temperature threshold may be determined by analysis or testing of the gas turbine engine to determine differential temperature threshold. Look up tables may be developed for various atmospheric and gas turbine engine conditions. After determining differential temperature thresholds by test or analysis, the control 100 may be effectively programmed with information as shown in graphs like
As shown in
The main oil temperature is a temperature which is utilized by the engine control and provided to various control systems, and to the pilot, to provide an indication of the current temperature of the oil. While the location of the identified main oil temperature is not particularly controlling, for purposes of this disclosure, it is simply a reference temperature which may then be utilized to determine the likelihood of icing conditions.
If an icing condition is identified, and exists for a predetermined minimum time, then an icing condition may be identified. If an icing condition is identified, the engine control may take one or more actions such as executing a pilot notification. Exemplary pilot notifications might include a warning light, an audible alarm, etc. Alternatively, or in combination, the anti-icing system might be actuated by the control.
Step 200 asks if the ambient temperature is below a set maximum temperature that is determined to be indicative of when icing conditions could occur. If so, at step 202 the calibrated air speed at an ambient temperature and MOT is determined. If the calibrated air speed is to the right of the slope (see
At step 204 the control system will confirm icing conditions if the delta temperature is above a minimum threshold (
At step 206 a cockpit notification might then occur. Alternatively, or in combination with step 206 an anti-icing measure could be automatically actuated at step 208.
The controller is programmed to associate a plurality of main oil temperatures, and relative to an ambient temperature and a calibrated air speed, and identify the icing condition should the current calibrated air speed at a current main oil temperature be greater than an identified calibrated air speed that might indicative potential icing.
The control is also programmed to identify a particular temperature differential with a particular reference fluid temperature to identify whether the icing condition is occurring.
In a featured embodiment, a gas turbine engine 20 includes a propulsor 24/26 for providing air into a core engine housing, and for providing air as propulsion air radially outwardly of the core engine housing. The core engine housing surrounds a compressor section. A combustor is positioned downstream of the compressor section and a turbine section is positioned downstream of the combustor. An aircraft fluid moves within an inlet to the gas turbine engine, and is exposed to inlet air at the inlet. There is a first temperature sensor 56 for sensing a first temperature of the aircraft fluid at a first upstream point and a second temperature sensor 54 for sensing a second temperature of the aircraft fluid at a second downstream point where the aircraft fluid has been exposed to the inlet air for a period of time. A control 100 determines a temperature differential between first and second temperatures sensed by the first and second temperature sensors. The temperature differential is associated with a likelihood that an icing condition will occur, and the control is programmed to take a corrective action should an icing condition be identified.
In another embodiment according to the previous embodiment, a strut 48 extends between an outer engine core housing 30 and an inner engine core housing 46 and the aircraft fluid passes through the strut.
In another embodiment according to any of the previous embodiments, the aircraft fluid is oil being routed through the strut to a heat exchanger 53.
In another embodiment according to any of the previous embodiments, a fan case 22 surrounds the propulsor, and the propulsor is a fan rotor. The fan rotor delivers air into a bypass duct defined between the fan case and the outer core engine housing, and the heat exchanger is positioned in the bypass duct.
In another embodiment according to any of the previous embodiments, the control 100 is programmed to identify a particular temperature differential with a particular reference fluid temperature to identify whether the icing condition is occurring.
In another embodiment according to any of the previous embodiments, the control 100 is programmed to identify the icing condition should the temperature differential continue to suggest an icing condition for a minimum period of time.
In another embodiment according to any of the previous embodiments, the aircraft fluid is oil, and the reference fluid temperature is a main oil temperature utilized by the control 100.
In another embodiment according to any of the previous embodiments, the controller 100 is also programmed to associate a plurality of main oil temperatures, and relative to an ambient temperature and a calibrated air speed, and identify the icing condition should the current calibrated air speed at a current main oil temperature be greater than an identified calibrated air speed that might indicative potential icing.
In another embodiment according to any of the previous embodiments, the aircraft fluid is oil being routed through the strut to a heat exchanger.
In another embodiment according to any of the previous embodiments, the control is programmed to identify the icing condition should the temperature differential continue to suggest an icing condition for a minimum period of time.
In another featured embodiment, a method of operating a gas turbine engine includes the steps of providing air from a propulsor into a core engine housing, and as propulsion air radially outwardly of the core engine housing. An aircraft fluid is moved within an inlet to the gas turbine engine, and exposed to inlet air 32. The method then senses a first upstream temperature 56 of the aircraft fluid and a second downstream temperature 54 after the aircraft fluid has been exposed to the inlet air for a period of time. A temperature differential is determined between the first and second temperatures. The temperature differential is associated with a likelihood that an icing condition will occur, and a corrective action is taken should an icing condition be identified.
In another embodiment according to any of the previous embodiments, a strut 48 extends between an outer engine core housing 30 and an inner engine core housing 46 and the aircraft fluid passing through the strut.
In another embodiment according to any of the previous embodiments, the aircraft fluid is oil routed through the strut to a heat exchanger 53.
In another embodiment according to any of the previous embodiments, a fan case 22 surrounds a fan rotor, and the propulsion air delivered by the fan rotor is delivered into a bypass duct 128 defined between the fan case and an outer core engine housing, and the heat exchanger 53 being positioned in the bypass duct.
In another embodiment according to any of the previous embodiments, a control 100 is programmed to identify a particular temperature differential with a particular reference fluid temperature to identify whether the icing condition is occurring.
In another embodiment according to any of the previous embodiments, the control 100 identifying the icing condition should the temperature differential continue to suggest an icing condition for a minimum period of time.
In another embodiment according to any of the previous embodiments, the aircraft fluid is oil, and the reference fluid temperature is a main oil temperature utilized by the control 100.
In another embodiment according to any of the previous embodiments, the controller 100 is also programmed to associate a plurality of main oil temperatures, and relative to an ambient temperature and a calibrated air speed, and identify the icing condition should the current calibrated air speed at a current main oil temperature be greater than an identified calibrated air speed that might indicative potential icing.
In another embodiment according to any of the previous embodiments, the aircraft fluid is oil routed through the strut to a heat exchanger 53.
In another embodiment according to any of the previous embodiments, the control identifying the icing condition should the temperature differential continue to suggest an icing condition for a minimum period of time.
Although embodiments have been disclosed, a worker of ordinary skill in this art would recognize that modifications would come within the scope of this disclosure. For that reason, the following claims should be studied to determine the true scope and content of this disclosure.
