The present disclosure relates generally to aircraft air systems, and more specifically to an air circuit for providing air to an environmental control system.
Aircraft, such as commercial airliners, typically include multiple gas turbine engines configured to generate thrust. The gas turbine engines include a compressor section that compresses air, a combustor section that mixes the air with a fuel and ignites the mixture, and a turbine section across which the resultant combustion products are expanded.
As the compressor section draws in atmospheric air and compresses it, the air from the compressor section is suitable for provision to the environmental control system (ECS) of the aircraft. The ECS system supplies air for various applications on the aircraft. As an example, it supplies conditioned cabin air for the passenger cabin and the cockpit. As such, it must be a particular pressure and temperature.
In existing ECS configurations, air is bled from the compressor section at a temperature and a pressure in excess of the temperature and pressure required by the ECS and is conditioned to reduced temperature using a pre-cooler and to reduced pressure using a pressure regulating valve. In this manner, pressure in excess of that required by the ECS is dumped across the pressure regulating valve. The excess pressure dump results in an overall efficiency loss to the engine.
As part of the typical ECS system, the air downstream of the gas turbine engine is passed through an air cycle machine. In an air cycle machine, the compressed air tapped from the engine compressor is typically cooled by ram air, compressed to high pressure and temperature, cooled to low temperature and high pressure again by ram air, and expanded across a turbine (which drives the air cycle compressor) to create cold air at cabin pressure. Thus the air cycle machine requires only a moderate pressure to operate (as air is further compressed in the machine), while the engine bleed pressure can be well in excess of moderate pressure in many operation points.
Of course, during operation of a gas turbine engine during flight condition, the power supplied by the engine changes dramatically. In prior art systems, operating at high power, the air is at an undesirably high pressure. On the other hand, at certain low power conditions, the air from any one tap might be at an undesirably low pressure.
As such, the prior art has not been efficient.
It is well known that fuel efficiency is a driving force in modern aircraft engine design. The increase of even a small percent of fuel burn efficiency is a very valuable goal.
In a featured embodiment, an aircraft has a gas turbine engine including a compressor section that includes at least one compressor bleed. An environmental control system has an air input configured to receive pressurized cabin air. An intercooler has an input and an output. A selection valve is configured to selectively connect the bleeds to an intercooler input. At least one auxiliary compressor is connected to the intercooler output. An output of at least one auxiliary compressors is connected to an ECS air input. A controller is configured to receive contemporaneous operational data, calculate minimum configuration requirements to satisfy environmental demands, and transmit calculated configuration requirements to at least the selection valve to achieve a desired pressure and temperature for the air downstream of the auxiliary compressor.
In another embodiment according to the previous embodiment, at least one auxiliary compressor comprises a plurality of auxiliary compressors.
In another embodiment according to any of the previous embodiments, the selection valve selectively connects at least one of the bleeds to provide air to the intercooler at a pressure below a desired pressure of the ECS air input.
In another embodiment according to any of the previous embodiments, at least one of the compressor bleeds is a compressor bleed positioned between a low pressure compressor and a high pressure compressor.
In another embodiment according to any of the previous embodiments, the intercooler is an air to air heat exchanger, and cooled by fan air from a fan in an associated gas turbine engine.
In another embodiment according to any of the previous embodiments, the intercooler is mounted in an outer fan case.
In another embodiment according to any of the previous embodiments, the intercooler is mounted within a bypass duct.
In another embodiment according to any of the previous embodiments, the intercooler is mounted within a core engine housing.
In another embodiment according to any of the previous embodiments, the intercooler is located in an upper bifurcation.
In another embodiment according to any of the previous embodiments, the intercooler is located in a lower bifurcation.
In another embodiment according to any of the previous embodiments, a second valve is between the intercooler and the at least one auxiliary compressor, the controller also controls the second valve such that the controller controls a state of the selection valve and a state of the second valve.
In another embodiment according to any of the previous embodiments, the controller includes memory storing instructions configured to cause the controller to connect at least one of the bleeds having an ECS operating requirement.
In another embodiment according to any of the previous embodiments, at least one auxiliary compressor comprises a plurality of auxiliary compressors and wherein the controller includes a memory storing instructions configured to cause the control to alternate auxiliary compressors operating as a primary compressor on a per flight basis.
In another embodiment according to any of the previous embodiments, at least one of the at least one auxiliary compressors includes an electric motor, and wherein the electric motor is configured to drive rotation of the corresponding auxiliary compressor.
In another embodiment according to any of the previous embodiments, at least one of the at least one auxiliary compressor includes a mechanical motor, and wherein the mechanical motor is configured to drive rotation of the corresponding auxiliary compressor.
In another embodiment according to any of the previous embodiments, a low pressure spool including a fan drive turbine, for driving a fan rotor, provides power to drive the at least one auxiliary compressor.
In another embodiment according to any of the previous embodiments, the rotation of the low pressure spool creates electricity in an associated generator which is then provided to drive a motor the at least one auxiliary compressor.
In another embodiment according to any of the previous embodiments, the low pressure spool includes a takeoff shaft for mechanically driving a drive to the at least one auxiliary compressor.
In another featured embodiment, a method for supplying engine air to an environmental control system (ECS) includes selecting a compressor bleed from a plurality of compressor bleeds, the selected compressor bleed providing air at a higher temperature than a required ECS inlet air temperature maximum and at a lower pressure than a required ECS inlet air pressure. The bleed air is cooled from the selected bleed using an intercooler such that the bleed air is below the required ECS inlet air temperature maximum. The bleed air uses at least one auxiliary compressor such that the bleed air is at least the required ECS inlet air pressure. The cooled compressed bleed air is provided to an ECS air inlet.
In another featured embodiment, a system for use on a turbine engine powered aircraft to provide conditioned air to an environmental control system has a heat exchanger having an input and an output. A selection valve is configured to selectively connect the heat exchanger input to at least one bleed of a compressor of the turbine engine. At least one auxiliary compressor is connected to the heat exchanger output, wherein an output of the at least one auxiliary compressor is connected to an input of the environmental control system. A controller is configured to receive contemporaneous operational data, calculate minimum configuration requirements to satisfy environmental demands, and transmit calculated configuration requirements to at least the selection valve to achieve a desired pressure and temperature for the air downstream of the auxiliary compressor.
These and other features can be best understood from the following specification and drawings, of which the following is a brief description.
The exemplary engine 20 generally includes a low speed spool 30 and a high speed spool 32 mounted for rotation about an engine central longitudinal axis A relative to an engine static structure 36 via several bearing systems 38. It should be understood that various bearing systems 38 at various locations may alternatively or additionally be provided, and the location of bearing systems 38 may be varied as appropriate to the application.
The low speed spool 30 generally includes an inner shaft 40 that interconnects a fan 42, a first (or low) pressure compressor 44 and a first (or low) pressure turbine 46. The inner shaft 40 is connected to the fan 42 through a speed change mechanism, which in exemplary gas turbine engine 20 is illustrated as a geared architecture 48 to drive the fan 42 at a lower speed than the low speed spool 30. The high speed spool 32 includes an outer shaft 50 that interconnects a second (or high) pressure compressor 52 and a second (or high) pressure turbine 54. A combustor 56 is arranged in exemplary gas turbine 20 between the high pressure compressor 52 and the high pressure turbine 54. A mid-turbine frame 57 of the engine static structure 36 is arranged generally between the high pressure turbine 54 and the low pressure turbine 46. The mid-turbine frame 57 further supports bearing systems 38 in the turbine section 28. The inner shaft 40 and the outer shaft 50 are concentric and rotate via bearing systems 38 about the engine central longitudinal axis A which is collinear with their longitudinal axes.
The core airflow is compressed by the low pressure compressor 44 then the high pressure compressor 52, mixed and burned with fuel in the combustor 56, then expanded over the high pressure turbine 54 and low pressure turbine 46. The mid-turbine frame 57 includes airfoils 59 which are in the core airflow path C. The turbines 46, 54 rotationally drive the respective low speed spool 30 and high speed spool 32 in response to the expansion. It will be appreciated that each of the positions of the fan section 22, compressor section 24, combustor section 26, turbine section 28, and fan drive gear system 48 may be varied. For example, gear system 48 may be located aft of combustor section 26 or even aft of turbine section 28, and fan section 22 may be positioned forward or aft of the location of gear system 48.
The engine 20 in one example is a high-bypass geared aircraft engine. In a further example, the engine 20 bypass ratio is greater than about six (6), with an example embodiment being greater than about ten (10), the geared architecture 48 is an epicyclic gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3 and the low pressure turbine 46 has a pressure ratio that is greater than about five. In one disclosed embodiment, the engine 20 bypass ratio is greater than about ten (10:1), the fan diameter is significantly larger than that of the low pressure compressor 44, and the low pressure turbine 46 has a pressure ratio that is greater than about five (5:1). Low pressure turbine 46 pressure ratio is pressure measured prior to inlet of low pressure turbine 46 as related to the pressure at the outlet of the low pressure turbine 46 prior to an exhaust nozzle. The geared architecture 48 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 2.3:1. It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present invention is applicable to other gas turbine engines including direct drive turbofans.
A significant amount of thrust is provided by the bypass flow B due to the high bypass ratio. The fan section 22 of the engine 20 is designed for a particular flight condition—typically cruise at about 0.8 Mach and about 35,000 feet (10668 meters). The flight condition of 0.8 Mach and 35,000 ft (10668 m), with the engine at its best fuel consumption—also known as “bucket cruise Thrust Specific Fuel Consumption (‘TSFCT’)”—is the industry standard parameter of lbm of fuel being burned divided by lbf of thrust the engine produces at that minimum point. “Low fan pressure ratio” is the pressure ratio across the fan blade alone, without a Fan Exit Guide Vane (“FEGV”) system. The low fan pressure ratio as disclosed herein according to one non-limiting embodiment is less than about 1.45. “Low corrected fan tip speed” is the actual fan tip speed in ft/sec divided by an industry standard temperature correction of [(Tram °R)/(518.7° R)]̂0.5. The “Low corrected fan tip speed” as disclosed herein according to one non-limiting embodiment is less than about 1150 ft/second (350.5 m/s).
In order to provide air from the compressor section 24 to the aircraft environmental control system (ECS), multiple bleeds are incorporated in the compressor section 24 (illustrated schematically in
An aircraft controller 101 controls a selection valve 140 such that, at any given time, air is provided from a bleed 102, 104, 106, 108 having the appropriate flow requirements of the ECS at the current operating conditions of the aircraft. While the bleed 102, 104, 106, 108 selected by the controller 101 provides air at acceptable flow levels, the bleed 102, 104, 106, 108 is selected to provide air that is under pressured. In other words, the pressure of the air provided by the selected bleed 102, 104, 106, 108 is below the pressure required by the ECS. Further, the air selected generally exceeds the temperature requirements of the ECS.
After passing through the selection valve 140, the air is passed to the intercooler 130. The intercooler 130 is a heat exchanger that cools the bleed air prior to providing the air to the ECS. The exemplary intercooler 130 utilizes fan air, provided from the bypass flowpath of the engine 120, to cool the air in an air to air heat exchanger format.
Cooled air from the intercooler 130 is provided to a second valve 150. The second valve 150 is controlled by the aircraft controller 101 and provides air to a first auxiliary compressor 160, a second auxiliary compressor 162, or both the first and second auxiliary compressor 160, 162. Each of the auxiliary compressors 160, 162 is driven by a corresponding electric motor 164, 166 and raises the pressure of the air to a required pressure level for provision to the ECS. Also, the controller 101 is shown controlling both motors 164/166. In alternative examples, one or both of the electric motors 164, 166 can be replaced or supplemented by a mechanical motor.
Once pressurized via the auxiliary compressors 160, 162 the air is provided to the ECS. In alternative examples, a single auxiliary compressor 160 can be used in place of the first and second auxiliary compressors 160, 162. In yet further alternative examples, three or more auxiliary compressors can be included, with the controller 101 rotating between the auxiliary compressors as necessary.
By cooling the bleed air prior to providing the bleed air to auxiliary compressors 160, 162, the amount of work required to compress the air at the auxiliary compressor 160, 162 is reduced, thereby achieving a fuel efficiency savings relative to not cooling the air before the auxiliary compressors.
Controller 101 may include memory storing instructions configured to cause the controller to connect a bleed having a required flowrate and pressure for an ECS operating requirement, and wherein the connected bleed has a pressure requirement below a pressure requirement of the ECS inlet.
It should be understood the controller is programmed to achieve a desired pressure and temperature which may vary with the demand for the ECS system, and which may also vary based upon the operating condition of the gas turbine engine. The controller is operable to control valve 140, valve 150, and the compressor 160/162. In addition, an optional variable speed transmission 165 may be provided such that the speed of the auxiliary compressor 160 can be controlled to achieve the desired conditions.
The controller is scheduled to understand the pressure required at the ECS inlet for various flight conditions. It selects the appropriate bleed port that supplies the highest pressure that is below the ECS required pressure. Power is supplied to the compressor such that the resultant pressure meets requirements. In some operating modes (descent or failure modes), the controller may select a pressure in excess of the ECS demand and use a regulating valve to reduce the pressure. This would not be a normal operating mode as it is contrary to the system goal (not to waste energy), but may be used to provide fail safe operation for failure conditions, or conditions like descent where very little fuel is burned (and thus there is not much of an opportunity for fuel burn reduction). Such operation will simplify the system design without compromising typical mission fuel burn reduction.
The aircraft 10 could be said to include a gas turbine engine having a compressor section including at least one compressor bleed. An environment control system (ECS) has an air input configured to receive pressurized cabin air. An intercooler has an input and an output. A selection valve is configured to selectively connect the bleed to the intercooler input. At least one auxiliary compressor is connected to the intercooler output. An output of the at least one auxiliary compressor is connected to the ECS air input. A controller is configured to receive contemporaneous operational data, calculate minimum configuration requirements to satisfy environmental demands, and transmit calculated configuration requirements to at least the selection valve to achieve a desired pressure and temperature for the air downstream of the auxiliary compressor.
The operational data may include engine performance demands and/or atmospheric data, as well as other conditions as appropriate. The environmental demands may be as known in the aircraft art.
While the circuit 100 is illustrated in
In the exemplary circuit 100 only one of the auxiliary compressors 160, 162 may be required to provide sufficient pressurization to the ECS during standard operating conditions. As such, only a single auxiliary compressor 160, 162 is typically operated during a flight. In order to even out wear between the auxiliary compressors 160, 162 the primary operating auxiliary compressor 160, 162 is alternated between flights on a per flight basis. Alternating between auxiliary compressors 160, 162 further allows earlier detection, and correction, of a damaged or inoperable second auxiliary compressor 162.
The controller may also include memory storing instructions configured to cause the controller to alternate auxiliary compressors operating as a primary compressor on a per flight basis.
During flight, should one engine 120 shut down, either due to mechanical failure, or for any other reason, the air provided from the bleeds 102, 104, 106, 108, is reduced. By way of example, if there are two engines 120, and one shuts down, the air provided to the auxiliary compressors 160, 162 is cut by ⅔ of the normal flow. In order to remedy this, in the exemplary system when one engine 120 shuts down, the currently inactive auxiliary compressor 160, 162 begins operating simultaneously with the currently operating auxiliary compressor 160, 162. The simultaneous operations ensure that any lost pressure due to the loss of an engine is compensated for using air from the operating engine or engines. In aircraft having more than two auxiliary compressors 160, 162, the controller 101 can apply a proportional control to one or more of the auxiliary compressors to ensure that adequate pressure is maintained at the ECS in proportion to the pressure lost due to the lack of operation of the engine. Again, the controller 101 is programmed to achieve this control.
A disclosed system for use on a turbine engine powered aircraft provides conditioned air to an environmental control system. A heat exchanger has an input and an output. A selection valve is configured to selectively connect the heat exchanger input to at least one bleed of a compressor of the turbine engine. At least one auxiliary compressor is connected to the heat exchanger output. An output of the at least one auxiliary compressor is connected to an input of the environmental control system. A controller is configured to receive contemporaneous operational data, calculate minimum configuration requirements to satisfy environmental demands, and transmit calculated configuration requirements to at least the selection valve to achieve a desired pressure and temperature for the air downstream of the auxiliary compressor.
A method for supplying engine air to an environmental control system (ECS) includes selecting a compressor bleed from a plurality of compressor bleeds, the selected compressor bleed providing air at a higher temperature than a required ECS inlet air temperature maximum and at a lower pressure than a required ECS inlet air pressure, cooling the bleed air from the selected bleed using an intercooler such that the bleed air is below the required ECS inlet air temperature maximum, including the bleed air using at least one auxiliary compressor such that the bleed air is at least the required ECS inlet air pressure, and providing the cooled compressed bleed air to an ECS air inlet.
While certain drive connections are disclosed above for the auxiliary compressor 160/162, it should also be understood that a hydraulic drive may be utilized.
It should be understood that the embodiments shown in
One example auxiliary compressor 200 is illustrated in
A high pressure compressor 220 is driven by a high pressure turbine 224. A combustor 226 is intermediate the two.
A bypass duct 228 is shown. As known, the fan rotor delivers air into the bypass duct 228 as bypass air B and also into a core engine as core air flow C. The core engine is defined by a core engine housing 230. A fan case 232 sits outwardly of the fan rotor 212.
In the embodiment of
In an alternative embodiment shown in
Finally, the heat exchanger may be mounted externally from the engine, closer to the location of the auxiliary compressors, using either cooling air bled from the engine fan duct, or ram air.
It is further understood that any of the above described concepts can be used alone or in combination with any or all of the other above described concepts. Although an embodiment has been disclosed, a worker of ordinary skill in this art would recognize that certain modifications would come within the scope of this invention. For that reason, the following claims should be studied to determine the true scope and content of this disclosure.
This application claims priority to U.S. Provisional Application No. 62/432,110, filed Dec. 9, 2017.
Number | Date | Country | |
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62432110 | Dec 2016 | US |