Aircraft bleed air systems and methods

Abstract

A method of determining a state of a switching valve of a bleed air system in a gas turbine engine of an aircraft includes determining a temperature in a main gas path of the gas turbine engine at a location at or downstream of the combustor and then actuating the switching valve of the bleed air system to bleed air from a different bleed location. The temperature is determined after actuating the switching valve. In response to determining that a difference between the temperatures before and after the actuation of the switching valve meets a threshold, executing at least one of a first action and a second action with respect to the gas turbine engine and/or the switching valve.

What is claimed is: 
     
         1 . A method of determining a state of a switching valve of a bleed air system in a gas turbine engine of an aircraft, the method comprising:
 when the gas turbine engine is operating, using the bleed air system to provide compressed air to a plurality of parts of the gas turbine engine, by bleeding the compressed air from one of:
 a first bleed location of a compressor section of the gas turbine engine, and 
 a second bleed location that is downstream of the first bleed location relative to air flow through the compressor section; 
   determining a temperature in a main gas path of the gas turbine engine at a location at or downstream of a combustor of the gas turbine engine and then actuating the switching valve of the bleed air system to bleed air from the other one of the first and second bleed locations;   determining the temperature of the gas turbine engine after the step of actuating the switching valve; and   in response to determining that a difference between the temperatures before and after the step of actuating the switching valve meets a threshold, executing at least one of a first action with respect to at least one of the gas turbine engine and the switching valve, and else executing a second action with respect to the at least one of the gas turbine engine and the switching valve, the second action being different from the first action.   
     
     
         2 . The method of  claim 1 , wherein the operating the gas turbine engine includes operating the gas turbine engine in one of a powered mode to a standby mode, when the gas turbine engine is operating in the powered mode, the step of bleeding air is from the first bleed location, and when the gas turbine engine is operating in the standby mode, the step of bleeding air is from the second bleed location. 
     
     
         3 . The method of  claim 2 , wherein when the gas turbine engine is operating in the powered mode, the first action includes one of:
 switching the gas turbine engine to the standby mode in response to a command from an operator of the aircraft to switch the gas turbine engine to the standby mode, and   allowing a switch of the gas turbine engine to the standby mode in response to the command.   
     
     
         4 . The method of  claim 3 , wherein when the gas turbine engine is operating in the powered mode, the second action includes preventing the gas turbine engine from switching to the standby mode in response to the command. 
     
     
         5 . The method of  claim 1 , wherein the first action includes generating an indication that the switching valve is in a functional state. 
     
     
         6 . The method of  claim 1 , wherein the second action includes generating an indication that the switching valve is in a malfunction state. 
     
     
         7 . The method of  claim 6 , wherein the temperature is an inter-turbine temperature of the gas turbine engine. 
     
     
         8 . A multi-engine aircraft comprising:
 a first gas turbine engine for providing a first engine power output; and   a second gas turbine engine for providing a second engine power output, the second gas turbine engine having a bleed air system and a bleed air system including: an switching valve having: a low pressure inlet in communication with a first bleed location in a compressor section of the second gas turbine engine; a high pressure inlet in communication with a second bleed location in the compressor section, the second bleed location being downstream of the first bleed location relative to main gas flow through the compressor section; and a switch outlet in communication with the bleed air system, the switching valve displaceable between at least first and second positions, the first position interconnecting the lower pressure inlet and the switch outlet, and the second position interconnecting the high pressure inlet and the switch outlet; and   a control system configured to:
 actuate the switching valve between the first and second positions, 
 determine a temperature change in a main gas path of the second gas turbine engine resulting from the switching valve being displaced from one of the first and second positions to the other one of the first and second positions, determine whether the temperature change meets a threshold, and 
 in response to determining that the temperature change meets the threshold, execute an action with respect to the first gas turbine engine, the second gas turbine engine, and/or the switching valve. 
   
     
     
         9 . The multi-engine aircraft according to  claim 8 , wherein the action includes switching the second engine between a powered mode and a standby mode. 
     
     
         10 . The multi-engine aircraft according to  claim 9 , wherein the control system is configured to:
 receive a command for from an operator of the multi-engine aircraft, the command requesting the switching the second engine between the powered mode and the standby mode;   determine whether the temperature change meets the threshold in response to receiving the command; and   execute the action in response to determining that the temperature change meets the threshold.   
     
     
         11 . The multi-engine aircraft according to  claim 10 , wherein the temperature is an inter-turbine temperature of the gas turbine engine. 
     
     
         12 . The multi-engine aircraft according to  claim 11 , wherein the control system is configured to prevent the switching the second engine between the powered mode and the standby mode in response to determining that the temperature change is below the threshold.