The disclosure relates to operation of a multi-engine aircraft and, more particularly, to operation of engine air systems for such a multi-engine aircraft.
Bleed air produced by a gas turbine engine is compressed air from the compressor stage that is used for engine functions (such as cooling of turbines and to help seal bearing cavities, for example) and is also used for aircraft functions (such as engine starting, cabin pressure, pneumatic systems, pressurizing liquid tanks, etc.). Engine bleed air can be derived from the high pressure or the low pressure compressor stage, depending on the air pressure requirements and the engine operating condition. Low pressure stage air is typically used during high power engine operation, and high pressure stage air may be used during descent and other low power engine operations. Bleed air may be used within an engine for engine functions such as bearing cavity sealing and cooling of various engine components. Improvement is desirable to optimize operation of the efficiency of compressed air systems.
Failure to provide sufficient pressurized air for the engine demands during lower power operation states of an engine can however result in oil seal degradation, oil leakage contamination of cabin air and overheating of air cooled components. Improvement is desirable to monitor and control the delivery of sufficient compressed air for the air demands of engine components when the engine is running in low power operating regimes.
There is according provided, a method of operating a gas turbine engine having a bleed air system with a switching valve operable to switch between multiple air sources depending on an operating mode of the gas turbine engine, the operating mode including a motive powered mode and a standby mode, the method comprising: operating the switching valve to provide pressurized air from a selected one or more of the multiple air sources to a cavity of the gas turbine engine having seals, and pressurizing the seals of the cavity using the pressurized air; testing the switching valve to determine if the switching valve is functioning normally or abnormally, the switching valve functioning normally when operable to switch between the multiple air sources and functioning abnormally when switching between the multiple air sources is not possible; and if the switching valve is determined to be functioning abnormally, controlling the gas turbine engine to prevent a change of the operating mode thereof.
There is also provided, a gas turbine engine, comprising: a bleed air system including a switching valve having at least first and second air inlets and an outlet in communication with a cavity of the gas turbine engine having one or more seals therein, the first inlet of the air switching valve in communication with a low pressure air source, the second inlet of the air switching valve in communication with a high pressure air source; and a switching valve health monitor, the switching valve health monitor operable to determine if the switching valve is functioning normally or abnormally, the switching valve functioning normally when operable to switch between the first and second inlets, and the switching valve functioning abnormally when switching between the first and second inlets is not possible or the switching valve is otherwise not operating as expected.
There is further provided, a multi-engine aircraft comprising: a first gas turbine engine for providing a first engine power output; and a second engine for providing a second engine power output, the second engine having a bleed air system and an air switching system including: an switching valve having: a low pressure inlet in communication with a first bleed location in a compressor of the second engine; a high pressure inlet in communication with a second bleed location in the compressor of the second engine, the second bleed location being downstream of the first bleed location relative to main gas flow through the compressor of the second engine; and a switch outlet in communication with at least one cavity of the second engine having seals therein; wherein the switching valve is 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 switching valve health monitor, the switching valve health monitor including at least one of a pressure sensor and a position sensor operable to determine if the switching valve is functioning normally or abnormally, the switching valve functioning normally when operable to switch between the first and second positions and functioning abnormally when switching between the first and second positions is not possible.
There is also provided a method of operating a gas turbine engine having a bleed air system and a switching valve operable to switch between low and high pressure air source inlets and having an outlet in communication with the bleed air system, the method comprising: actuating the switching valve between a motive powered mode and a standby mode, wherein low pressure compressed air is conveyed to the bleed air system when the switching valve is in the motive powered mode, and wherein high pressure compressed air is conveyed to the bleed air system when the switching valve is in the standby mode; sensing a standby mode pressure using a bleed sensor located downstream from the switching valve; comparing the standby mode pressure to a reference pressure; and when the standby mode pressure is equal to or greater than the reference pressure, maintaining the switching valve in the standby mode; and when the standby mode pressure is less than the reference pressure, actuating the switching valve from the standby mode to the motive powered mode.
In the gas turbine engine as defined above, the control means are adapted, in at least one embodiment, for: before actuating the switching valve from the motive powered mode to the standby mode, sensing the reference pressure comprising an initiation pressure using the bleed sensor located downstream from the valve outlet of the switching valve; and after actuating the switching valve from the motive powered mode to the standby mode, comparing the standby mode pressure to the initiation pressure to determine an initial pressure differential; and when the initial pressure differential is equal to or greater than a downstream reference value, maintaining the switching valve in the standby mode; and when the initial pressure differential is less than the downstream reference value, actuating the switching valve from the standby mode to the motive powered mode
In the gas turbine engine as defined above, the control means are adapted, in at least one embodiment, for: comparing the initial pressure differential to the downstream reference value; and determining a switching valve status being one of: the switching valve is fully open; the switching valve is fully closed; and the switching valve is partially open.
The gas turbine engine as defined above includes, in at least one embodiment: a source sensor located upstream from the high pressure inlet of the switching valve; and wherein the control means are adapted for: before actuating the switching valve from the motive powered mode to the standby mode, sensing the reference pressure comprising a source pressure using the source sensor; and after actuating the switching valve from the motive powered mode to the standby mode, comparing the standby mode pressure to the source pressure to determine a secondary pressure differential; and when the secondary pressure differential is less than an upstream reference value, maintaining the switching valve in the standby mode; and when the secondary pressure differential is equal to or greater than the upstream reference value, actuating the switching valve from the standby mode to the motive powered mode.
In the gas turbine engine as defined above, the control means are adapted, in at least one embodiment, for: comparing the secondary pressure differential to the upstream reference value; and determining a switching valve status being one of: the switching valve is fully open; the switching valve is fully closed; and the switching valve is partially open.
Pressure sensors upstream or downstream from the switching valve can be used to monitor the operation and determine the status of the switching valve. A change in static pressure or the measurement of a pressure differential across the switching valve will provide an indication of whether the switching valve is fully open, fully closed or partially open.
In accordance with embodiments of the present description, malfunction of the switching valve can be detected and can result in aborting of the standby mode operation. Aborting standby mode operation results in a return to the full power motive powered mode to provide sufficient compressed air to avoid engine damage, hazardous or undesirable conditions, such as oil seal degradation, oil contamination of the aircraft cabin air and overheating of air cooled components of the engine.
Further details of these and other aspects of the subject matter of this application will be apparent from the detailed description included below and the drawings.
More particularly, the multi-engine system 50 of this embodiment includes first and second turboshaft engines 10 each having a respective transmission 38 which are interconnected by a common output gearbox 40 to drive a common load 44. In one embodiment, the common load 44 may comprise a rotary wing of a rotary-wing aircraft. For example, the common load 44 may be a main rotor of the helicopter 1. Depending on the type of the common load 44 and on the operating speed thereof, each of turboshaft engines 10 may be drivingly coupled to the common load 44 via the output gearbox 40, which may be of the speed-reduction type.
For example, the gear box 40 may have a plurality of transmission shafts 42 to receive mechanical energy from respective output shafts 11 of respective turboshaft engines 10. The gear box 40 may be configured to direct at least some of the combined mechanical energy from the plurality of the turboshaft engines 10 toward a common output shaft 43 for driving the common load 44 at a suitable operating (e.g., rotational) speed. It is understood that the multi-engine system 50 may also be configured, for example, to drive accessories and/or other elements of an associated aircraft. As will be described, the gear box 40 may be configured to permit the common load 44 to be driven by either of the turboshaft engines 10 or, by a combination of both engines 10 together.
Referring now to
The engine 10 depicted in
Referring still to
As will be seen, the air bled from the engine 10 is, in at least a particular embodiment, to be re-used for engine secondary air demands for air only, but not the air demands of the aircraft fuselage or cabin. The present disclosure describes a method of monitoring the engine 10 when operating in a lower power, or “standby” mode and switching to an alternate bleed air source, while ensuring that the engine demands for pressurized air are satisfied, including pressurized oil seals, and turbine cooling systems. The methods and functions of the aircraft 1 described herein may be executed for example by any suitable controller(s), such as a full-authority digital controller (FADEC) (
The present description relates generally to operating a multi-engine aircraft 1, where a first engine of the gas turbine engines 10 is capable of providing motive power in flight to the aircraft 1 when operating in a “powered mode”, and a second engine of the gas turbine engines 10 can be switched between the powered mode, whereby the second engine also provides motive power to the aircraft, and a lower-power “standby mode”, whereby substantially no (or very little) motive power is provide to the aircraft. The second engine may therefore be switched into, and maintained in, this very low-power standby mode during flight. The standby mode of the second engine includes, but is not limited to a significantly reduced power mode, i.e. a so-called “sub-idle” regime or “idle cruise” regime. The term “standby mode” as used herein is therefore understood to be such a low power regime. By operating the second engine in this standby mode (sub-idle or idle cruise regime) during flight, fuel consumption of the second engine, and therefore over the overall powerplant, is reduced. The second engine however remains able to return to normal operational power (e.g. full power) if additional power needed by the aircraft.
Although various differential control sequences between the engines 10 are possible, in one particular embodiment the controller(s) (e.g. (FADEC)) may correspondingly control fuel flow rate to each engine 10 as follows. In the case of the standby engine, a fuel flow (and/or a fuel flow rate) provided to the standby engine may be controlled to be between 70% and 99.5% less than the fuel flow (and/or the fuel flow rate) provided to the active engine (i.e. the engine 10 operating in the powered mode). The standby engine may be maintained between 70% and 99.5% less than the fuel flow to the active engine. In some embodiments, the fuel flow rate difference between the active and standby engines may be controlled to be in a range of 70% and 90% of each other, with fuel flow to the standby engine being 70% to 90% less than the active engine. In some embodiments, the fuel flow rate difference may be controlled to be in a range of 80% and 90%, with fuel flow to the standby engine being 80% to 90% less than the active engine.
In another embodiment, the controller(s) may operate one engine in a standby mode at a power substantially lower than a rated cruise power level of that engine, and in some embodiments at zero output power and in other embodiments less than 10% output power relative to a reference power (provided at a reference fuel flow). Alternately still, in some embodiments, the controller(s) may control the standby engine to operate at a power in a range of 0% to 1% of a rated full-power of the standby engine (i.e. the power output of the second engine to the common gearbox remains between 0% to 1% of a rated full-power of the second engine when the second engine is operating in the standby mode).
The first engine as described herein may be referred to as a “main” or “primary” engine, and the second engine may be referred to herein as a “reserve” engine. However, it is to be understood that both are capable of operating at full power to propel the aircraft 1 if desired/necessary.
Referring now to
Accordingly, a switching valve health monitor (SVHM) (
With reference to
With reference to
As shown in
With reference to
Returning to
As indicated by return arrow 26 in
As indicated by return arrow 27 in
The means by which the control system interrogates pressure sensor 30 or 31 and determines the status of the switching valve 14, is described below with reference to
The air source input switching valve 14 has: a low pressure inlet 16 in communication with a low pressure air source 17; a high pressure inlet 20 in communication with a high pressure air source 21; and at least one valve outlet 19 in communication with the bleed air system 18 of the engine. In the example illustrated, the bleed air system 18 includes conduits conveying compressed air through a stationary strut 28 to an engine shaft 29 which can be directed to cool the turbines 11 (
When the switching valve 14 is in the motive powered mode, as shown with dashed arrows in
The aircraft pilot requests that the engine enter the idle mode by engaging the control system to actuate the switching valve 14 from the motive powered mode to the standby mode. The control system commences a system check by sensing a standby mode pressure using a bleed sensor 30 located downstream from the switching valve 14. The standby mode pressure is compared to a reference pressure, which could be an absolute static pressure value, an allowable range of pressure values or an expected pressure differential.
When the standby mode pressure measured by the bleed sensor 30 is equal to or greater than the reference pressure, the switching valve 14 is maintained in the standby mode (see box 24,
When the standby mode pressure is less than the reference pressure, the control system actuates the switching valve 14 from the standby mode to the motive powered mode, thereby inhibiting the operation of the engine in standby mode. This result is because the air pressure delivered is less than the pressure required for engine demands.
The bleed sensor 30 can also be interrogated to obtain a pressure reading in the bleed air system 18 when the switching valve 14 is in the motive powered mode, shown in
After actuating the switching valve 14 from the motive powered mode (as in
By comparing the initial pressure differential to the downstream reference value, a switching valve status can be determined namely: the switching valve 14 is fully open (change in pressure is adequate); the switching valve 14 is fully closed (virtually no change in pressure detected); or the switching valve 14 is partially open (inadequate pressure change detected).
An optional method includes additional pressure sensors to determine the air pressure upstream from the switching valve 14. Before actuating the switching valve 14 from the motive powered mode (
After actuating the switching valve 14 from the motive powered mode to the standby mode, the standby mode pressure (from bleed sensor 30) is compared to the source pressure (from source sensor 31) to determine a secondary pressure differential. When the secondary pressure differential is less than an upstream reference value, the switching valve 14 is maintained in the standby mode (i.e. sufficient air pressure detected). When the secondary pressure differential is equal to or greater than the upstream reference value, the switching valve 14 is actuated from the standby mode to the motive powered mode (i.e. not sufficient air pressure detected).
The secondary pressure differential when compared to the upstream reference value can also be used to determine a switching valve status namely: the switching valve is fully open (bleed sensor 30 and source sensor 31 measure equal pressures approximately); the switching valve is fully closed (bleed sensor 30 measures a pressure significantly lower than source sensor 31); and the switching valve is partially open (bleed sensor 30 measures a pressure less than source sensor 31 and insufficient pressure differential is detected).
The above description is meant to be exemplary only, and one skilled in the relevant arts will recognize that changes may be made to the embodiments described without departing from the scope of the invention disclosed. The present disclosure may be embodied in other specific forms without departing from the subject matter of the claims. The present disclosure is intended to cover and embrace all suitable changes in technology. Modifications which fall within the scope of the present invention will be apparent to those skilled in the art, in light of a review of this disclosure, and such modifications are intended to fall within the appended claims. Also, the scope of the claims should not be limited by the preferred embodiments set forth in the examples, but should be given the broadest interpretation consistent with the description as a whole.
The present application claims the benefit of U.S. Patent Application No. 62/835,806, filed Apr. 18, 2019, entitled “HEALTH MONITOR FOR AIR SWITCHING SYSTEM”, the entirety of which is hereby incorporated herein by reference in its entirety.
Number | Date | Country | |
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62835806 | Apr 2019 | US |