GAS TURBINE ENGINE SYSTEM WITH MIXED-FLOW AUXILIARY POWER UNIT

Abstract

A scavenging air control system (16) is configured to vary the air pressure at the inlet of a gas turbine engine. The scavenging air control system (16) comprises a first gas turbine engine (12) configured to supply scavenging air, a second gas turbine engine (14) acting as an auxiliary power unit, and a scavenging air control system (16) configured to selectively supply scavenging air from the first gas turbine engine (12) to the second gas turbine engine (14). Fig. 2

Description

Description Title of the invention: GAS TURBINE ENGINE SYSTEM WITH AUXILIARY FLOW POWER UNIT BLEND SCOPE OF DISCLOSURE

[0001] — This disclosure relates generally to auxiliary groups of power installed in an aircraft, particularly for DC power management compressed air to these auxiliary power units. BACKGROUND

[0002] = Gas turbine engines are used to power aircraft, em- boats, power generators, and other similar items. Turbine engines Gas engines typically include a compressor, a combustion chamber, and a... turbine. The compressor compresses the air drawn into the engine and distributes air to High pressure in the combustion chamber. In the combustion chamber, Fuel is mixed with air at high pressure and ignited. The products of the combustion reaction in the combustion chamber is directed into the turbine where The work is extracted to drive the compressor and, sometimes, an output shaft. remaining combustion products are expelled outside the turbine and can to provide a boost in certain applications.

[0003] — Auxiliary power units are generally small gas turbines used for starting the main engine, supplying electrical power to the aircraft and the supply of compressed air for use in the systems of Environmental control. Modern aircraft are becoming increasingly electric, with electrically controlled compressors, in-flight entertainment electric and electromechanical actuators instead of hydraulic actuators. These loads can exceed the power generation capacity of the generators. generators mounted on the aircraft's primary propulsion engines, and can be supplemented by additional power generation from the groups power auxiliaries. SUMMARY

[0004] This disclosure may include one or more of the following features the following and their combinations.

[0005] According to one illustrative aspect of the disclosure, a gas turbine engine system includes a first engine, a second engine and an air control system of sampling. The first engine includes a first compressor, a first a combustion chamber and a first turbine. The second engine includes a The second engine core and an airflow mixer are configured to receive ambient air and combine it with compressed air to form a combined airflow. The second engine core includes a second compressor, a second combustion chamber, and a second turbine. The airflow mixer is configured to receive ambient air and combine it with compressed air to form a combined airflow. It is further configured to direct this combined airflow into the second engine core. The bleed air control system is configured to control the inlet pressure of the second compressor. In some embodiments, the bleed air control system includes a duct, a flow control valve, and a control device. The duct is configured to direct compressed air from the first compressor to the airflow mixer. The flow control valve is in fluidic communication with the duct. The flow control valve is configured to selectively vary the flow of compressed air through the duct. The control device is programmed to adjust the flow control valve in response to a second compressor inlet pressure falling below a predetermined value, such that a combined flow pressure directed to the second engine core allows the second engine core to operate in a low ambient pressure environment. In some embodiments, the airflow mixer is a gas jet ejector. The gas jet ejector is configured to use compressed air from the first compressor as the motive fluid to create a vacuum and draw ambient air into the gas jet ejector, creating the combined flow directed into the core of the second motor. In some embodiments, the gas jet ejector comprises a nozzle and an ambient air inlet arranged circumferentially around the nozzle. The nozzle receives the compressed air stream and the ambient air inlet receives the ambient air. In some embodiments, the airflow mixer consists of an ambient air inlet and a check valve in fluidic communication with the duct. The check valve is configured to direct the compressed air flow into the ambient air inlet. In some embodiments, the second engine includes a backpressure regulator downstream of the second engine core. The backpressure regulator is configured to vary the exhaust air pressure of the second engine. In some embodiments, the sampling control system further includes a heat exchanger in fluidic communication with the duct. The heat exchanger is configured to remove heat from the compressed air. In some embodiments, the heat exchanger is configured to receive ambient air to remove heat from the compressed air. In some embodiments, the first motor further includes a blower. The heat exchanger is configured to receive air from the blower to remove heat from the compressed air. In some embodiments, the control device is programmed to vary the flow control valve and the back pressure regulator based on a difference between the pressure at an inlet of the second compressor and a reference pressure value on a predetermined pressure program of the second gas turbine engine. In some embodiments, the first compressor further comprises an intermediate pressure compressor section and a high-pressure compressor section. The intermediate pressure compressor section has a first port in fluidic communication with the duct. The high-pressure compressor section has a second port in fluidic communication with the duct. The control device is programmed to selectively open one of the first and second ports based on a pressure-to-flow ratio of the first compressor, which is stored in the control device. According to another illustrative aspect of the disclosure, a gas turbine engine system comprises a first engine, a second engine, and a bleed air control system. The first engine includes a first compressor, a first combustion chamber, and a first turbine. The second engine includes a second engine core and an airflow mixer. The second engine core includes a second compressor, a second combustion chamber, and a second turbine. The airflow mixer is positioned at the inlet of the second engine core. The sampling air control system comprises a duct, a flow control valve, and a control device. The duct connects the first compressor to the airflow mixer. The flow control valve is in fluidic communication with the duct. The control device is programmed to adjust the flow control valve in response to the second compressor's inlet pressure falling below a predetermined value. In some embodiments, the airflow mixer is a gas jet ejector. In some embodiments, the gas jet ejector further includes an ambient air inlet. Compressed air from the first compressor and ambient air enter the ejector coaxially, and the compressed air passes through a nozzle positioned radially towards the inside of the ambient air inlet. In some embodiments, the airflow mixer is an ambient air inlet and a check valve in fluidic communication with the duct. In some embodiments, the second motor includes a backpressure regulator downstream of the second motor core. In some embodiments, the sampling control system further includes a heat exchanger in fluidic communication with the duct. In some embodiments, the heat exchanger receives ambient air. In some embodiments, the first motor also includes a blower. The heat exchanger receives air from the blower. In some embodiments, the first compressor further comprises an intermediate pressure compressor section, an intermediate pressure compressor port in fluidic communication with the duct, a high-pressure compressor section, and a high-pressure compressor port in fluidic communication with the duct. The control device is programmed to selectively open the intermediate pressure compressor port and the high-pressure compressor port based on a pressure-to-flow ratio of the first compressor. According to an illustrative method of disclosure, a method of controlling a gas turbine engine system includes compressing a first air stream using a first compressor of a first gas turbine engine to provide a first compressed air stream, compressing a second air stream using a second compressor of a second gas turbine engine, and determining an inlet pressure of an inlet of a second compressor, directing a portion of the first compressed air stream to the inlet of the second compressor in response to the fact that the inlet pressure of the second compressor is less than a predetermined value. These and other features of this disclosure will be more clearly illustrated by the following description of illustrative embodiments. Brief description of the drawings [Fig.1] is a schematic view of an aircraft having four propulsive gas turbine engines and an auxiliary power group supplied by a fifth gas turbine engine enclosed in a nacelle, one of the four propulsive gas turbine engines is connected to the gas turbine engine located in the nacelle as part of a gas turbine engine system shown in [Fig.2] to provide auxiliary power; [Fig.2] is a schematic view of a gas turbine engine system in accordance with this disclosure; the gas turbine engine system comprising a first gas turbine engine that produces power for an aircraft, a second gas turbine engine that provides auxiliary power to the first engine and the aircraft, a bleed air control system configured to selectively supply compressed bleed air from the first engine to an airflow mixer of the second engine to assist the second engine to operate at higher altitudes, and an exhaust backpressure regulator to maintain adequate exhaust backpressure for the second engine at higher altitudes; [Fig.3] is a perspective view of the second gas turbine engine included in the system of [Fig.2] showing that a container coupled to the aircraft is cut to reveal the second gas turbine engine within it as well as other auxiliary components such as batteries, a generator, etc.; [Fig.4] is a schematic view of an airflow mixer coupled to the inlet of a second gas turbine engine, where the airflow mixer is an ejector comprising an ambient air inlet and a compressor bleed air inlet; [Fig. 5] is a schematic view of an airflow mixer coupled to the inlet of a second gas turbine engine, where the airflow mixer is a fast-opening / closing valve comprising an ambient air inlet and a compressor bleed air valve; and |Fig.6] is a flowchart showing one embodiment of the logic that a control device of the [Fig.2] system uses to regulate the compressed bleed air flow from the first gas turbine engine to the airflow mixer of the second gas turbine engine to help the second engine operate at higher altitudes. DETAILED DESCRIPTION OF THE DRAWINGS In order to promote understanding of the principles of disclosure, reference will now be made to a number of illustrative embodiments shown in the drawings and specific language will be used to describe them. A gas turbine engine system 10 for producing propulsion and power for an aircraft 8 is schematically represented in Figures 1 and 2. The system 10 comprises a first gas turbine engine 12, a second gas turbine engine 14, and a bleed air control system 16. The first gas turbine engine 12 produces propulsion for the aircraft 8. The second gas turbine engine 14 is an auxiliary power unit and provides auxiliary power to the first gas turbine engine 12 and to the aircraft 8. The bleed air control system 16 controls the supply of bleed air. compressed 78 from the first gas turbine engine 12 to the second gas turbine engine 14 and to aircraft 8. In the illustrative embodiment, the aircraft 8 includes a number of gas turbine engines to propel the aircraft 8; however, only the first gas turbine engine 12 is configured to supply compressed bleed air 78 to the second gas turbine engine 14. In other embodiments, any number of the propulsion gas turbine engines of the aircraft 8 or other APUs not shown may be connected to and configured to supply compressed bleed air 78 to the auxiliary power unit / second gas turbine engine 14. In this embodiment shown in [Fig. 2], the first gas turbine engine 12 comprises a fan 22, a compressor 18, a combustion chamber 28, and a turbine 30. The compressor 18 comprises an intermediate pressure compressor 24 and a high-pressure compressor 26 in the illustrative embodiment. In some embodiments, the turbine 30 may include a high-pressure section, an intermediate pressure section, and a low-pressure section to drive the high-pressure compressor 26, the intermediate pressure compressor 24, and the fan 22, respectively. Air passes through the fan 22, which is driven by the turbine 30. Some of the air is diverted around the compressor 18 to provide thrust to the aircraft 8. Another portion of the air is compressed by the compressor 18 and mixed with fuel in the combustion chamber 28.The fuel-air mixture in the combustion chamber 28 is ignited, adding energy to the air, and some of this energy is extracted by the turbine 30. The second gas turbine engine 14 comprises a compressor 32, a combustion chamber 34, a turbine 36, a gearbox 38, an electric generator 40, and an airflow mixer 42, as shown in [Fig. 2]. In some embodiments, the turbine 36 may include a high-pressure section to drive the compressor 32 and a low-pressure section to drive the gearbox 38. Air is directed through the airflow mixer 42 and compressed as it enters the compressor 32 after passing through the airflow mixer 42. Fuel is added to the air in the combustion chamber 34 and ignited to add energy to the air. Energy is extracted from the burned air by the turbine 36, which powers the gearbox 38 and the compressor 32.The electric generator 40 is powered by the gearbox 38 and is configured to produce electricity for auxiliary use in the aircraft 8, the first gas turbine engine 12, and / or the second gas turbine engine 14 for example. The airflow mixer 42 is in fluidic communication with the first gas turbine engine 12 via the intake air control system 16, as shown in [Fig. 2]. At relatively low altitudes, the airflow mixer 42 does not receive compressed bleed air 78 from the first gas turbine engine and instead supplies ambient air to the compressor 32 of the second gas turbine engine 14. The second gas turbine engine 14 is configured to operate independently of the first gas turbine engine 12 when the inlet air pressure is sufficiently high. At relatively high altitudes, the low ambient pressure environment may cause the compressor 32 to operate below a desired performance level and / or below a desired efficiency level. At these relatively high altitudes, the airflow mixer 42 receives both ambient air and compressed bleed air 78 from the first gas turbine engine 12.The airflow mixer 42 mixes ambient air and compressed air to provide an airflow to the compressor 32 with a higher pressure than that of ambient air alone, thereby improving the performance of the compressor 32 at high altitudes. The altitude at which the compressed bleed air 78 is used to supplement the ambient air may vary depending on the altitude at which the second gas turbine engine 14 is certified to operate and the magnitude of the electrical loads that the second gas turbine engine 14 must withstand. In some embodiments, the airflow mixer 42 may be used to mix the bleed air 78 with ambient air above 20,000 feet. The bleed air control system 16 is configured to selectively vary the flow of compressed bleed air 78 from the compressor 18 of the first engine 12 to the air flow mixer 42 and the compressor 32 of the second engine 14. The bleed air control system 16 includes an optional pre-cooler 62, a flow control valve 64, an optional heat exchanger 66, a backpressure regulator 68, a control device 70, and a conduit 80. Compressed bleed air 78 is drawn through the conduit 80 from the compressor 18 of the first engine 12 and passes through the pre-cooler 62, if present. The pre-cooler 62 uses ambient air or another heat sink to remove heat from the bleed air 78 before it enters the flow control valve 64.The pre-cooler 62 can already be present for the first gas turbine engine 12 and easily incorporated into the sampling control system 16. The compressed sampling air 78 can be drawn through one and / or the other of a first port 25 on the intermediate pressure compressor 24 and a second port 27 on the high pressure compressor 26 in the conduit 80. The flow control valve 64 is configured to open and close selectively to allow the compressed scavenging air 78 to flow through the duct 80 from the compressor 18 to the airflow mixer 42 or to block the compressed scavenging air 78 as suggested in [Fig. 2]. The control valve The flow control valve 64 can be partially opened to allow discretely or infinitely variable flow control of the compressed shunt air 78. The flow control valve 64 is located, for illustrative purposes, downstream of the pre-cooler 62. In other embodiments, the flow control valve 64 is located upstream of any pre-cooler 62. The heat exchanger 66, if present, uses ambient air or another heat sink such as blower air to remove heat from the compressed sampling air 78 flowing through the sampling air control system 16, as shown in [Fig. 2]. The heat exchanger 66 may include valves to vary the flow of ambient air through the heat exchanger 66. For illustrative purposes, the heat exchanger 66 is located downstream of the flow control valve 64. In other embodiments, the heat exchanger 66 is located upstream of the flow control valve 64. The backpressure regulating valve 68 is located downstream of the second gas turbine engine 14, as shown in [Fig. 2]. The backpressure regulator 68 is configured to open and close selectively when instructed by the control device 70 to maintain a desired pressure at the outlet of the second gas turbine engine 14. The air exiting the backpressure regulating valve 68 is vented to the ambient air. Controlling the backpressure with the backpressure regulating valve 68 allows the control device 70 to maintain a desired pressure change across the engine when the inlet pressure is varied, thus keeping the second gas turbine engine 14 operating at a predetermined pressure. In some embodiments, the second motor 14 is located in a nacelle 74 as shown in [Fig.3]. The nacelle 74 is coupled to the aircraft 8. By way of illustration, the nacelle 74 includes an ambient air inlet 46 and a compressed air inlet 72. The ambient air inlet 46 is connected to the airflow mixer 42 and is configured to receive ambient air from outside the nacelle 74 and direct the ambient air to the airflow mixer 42. The air inlet 72 fluidically connects the airflow mixer 42 to a duct 80 which supplies compressed bleed air 78 from the compressor 18 to the airflow mixer 42. An exhaust duct 76 is fluidly coupled to the turbine 36 of the second gas turbine engine 14 and is configured to direct the exhaust gases out of the second gas turbine engine 14 and out of the nacelle 74. As shown in [Fig.3], the electric generator 40 is located in the nacelle 74 and other components may be located in the nacelle 74 such as batteries, electrical circuits, cooling fluids, etc. A detailed view of the gas turbine engine 14 is shown in [Fig. 4]. As The airflow mixer 42, as shown, is an ejector and includes a compressed air inlet 44, an ambient air inlet 46, a convergent-divergent nozzle 48, a diffuser 50, and a pressure sensor 52. The compressed air inlet 44 is in fluidic communication with the duct 80 on one side and with the convergent-divergent nozzle 48 on the other. The ambient air inlet is arranged circumferentially around the compressed air inlet 44. Both the ambient air inlet 46 and the convergent-divergent nozzle 48 are in fluidic communication with the diffuser 50. The pressure sensor 52 is located where the diffuser 50 is in fluidic communication with the compressor 32. Compressed scavenging air 78, when supplied to the airflow mixer 42, enters the compressed air inlet 44 and passes through the nozzle 48, which converts some of the pressure in the compressed scavenging air 78 into velocity. The air exiting the nozzle 48 acts as a motive fluid and creates a vacuum that draws in or entrains the low-pressure ambient air coming through the ambient air inlet 46 to form a combined flow. This combined flow has a higher pressure than the ambient air alone and passes through the diffuser 50, increasing the pressure of the air entering the compressor 32 of the second gas turbine engine 14. The pressure sensor 52 measures the air pressure at the outlet of the diffuser 50 and at the inlet of the compressor 32.The pressure at compressor 32 can enable the second gas turbine engine 14 to operate at higher altitudes than it would be able to operate using ambient air alone in a low ambient pressure environment. Another embodiment of an airflow mixer 242 is shown in [Fig.5]. The air flow mixer 242 is substantially similar to the air flow mixer 42 except where the following description differs from that of the air flow mixer 42. The air flow mixer 242 includes a check valve 244 and an ambient air inlet 246. The check valve 244 is in fluidic communication with the compressor 18 to receive compressed bleed air 78. The check valve 244 opens and closes as required to add compressed bleed air 78 to the ambient air in the air flow mixer 242 in order to maintain a desired pressure at the pressure sensor 52 for the compressor 32 of a second gas turbine engine 14. By way of illustration, the ambient air inlet 246 has a generally constant diameter and neither converges nor diverges. The control device 70 varies the amount of compressed sampling air 78 supplied to the airflow mixer 42 as shown in [Fig. 2]. The control device 70 includes a memory 71 in which instructions and data are stored, and a processor 73 configured to execute the instructions and access the data stored in the memory 71. The control device 70 is connected to the flow control valve 64 and backpressure regulating valve 68. The control device 70 reads the pressure from the pressure sensor 52 and compares this pressure to a reference pressure in a lookup table stored in memory 71. The control device 70 is configured to adjust the flow control valve 64 based on the reference pressure values ​​in a lookup table to vary the amount of compressed sampling air 78 supplied from the compressor 18 to the air flow mixer 42 in order to raise the pressure measured by the pressure sensor 52. The control device 70 is further configured to adjust the backpressure regulating valve 68 according to a pressure program stored in memory 71 to vary the exhaust air pressure at the exhaust duct 76 in order to maintain a desired pressure ratio across the second gas turbine engine 14. This can be useful because the low ambient pressure environment can cause the exhaust air pressure to be lower than expected at the exhaust duct 76 due to the fact that the intake air control system 16 raises the pressure at the inlet of the compressor 32. In configurations such as that shown in [Fig. 2], where the compressor 18 comprises an intermediate pressure compressor 24 and a high-pressure compressor 26, the control device 70 will selectively open an orifice on one or both of the intermediate pressure compressor 24 and the high-pressure compressor 26 based on a compressor map with a pressure-to-flow ratio stored in memory 71. This selective mixing of intermediate-pressure and high-pressure bleed air allows the control device 70 to control the pump margin of the first gas turbine engine 12. The intermediate-pressure and high-pressure bleed air can also be used to control the pressure of the air mixed with ambient air and supplied to the second gas turbine engine 14.In alternative embodiments, the engine control system of the first gas turbine engine 12 or of a fully mechanical system can control the ports of the intermediate pressure compressor 24 and the high pressure compressor 26. Maintaining these desired pressures allows the second gas turbine engine 12 to operate under more optimal conditions and to adapt to the load on the electric generator 40 when the aircraft 8 is at high altitudes. The second gas turbine engine 12 then functions as an electronic auxiliary power unit (e-APU). An e-APU operates similarly to a standard aircraft auxiliary power unit, except that it can be configured to produce only electricity and not supply compressed air to the aircraft or other engines. Figure 6 represents an embodiment of the logic 300 which is stored in memory 71. The control device 70 can execute the logic 300 via the processor 73 to maintain the desired input pressure at the pressure sensor 52 for the compressor 32. At step 302, the control device 70 checks the status of the second gas turbine engine 14 to determine whether the second gas turbine engine 14 is either running (operating) or in start-up mode. The control device 70 can perform this determination by querying the engine control device or receiving signals from the speed / pressure sensors. If the second gas turbine engine 14 is neither running nor in start-up mode, the control device 70 directs the flow control valve 64 to close as in step 304. If the control device 70 determines that the second gas turbine engine 14 is either running or in start-up mode, the control device will proceed to step 306. At step 306 shown in [Fig.[6], the control device 70 determines whether the second gas turbine engine 14 is running (operating) or in start-up mode (as opposed to neither being true in step 304). If the second gas turbine engine 14 is in start-up mode, the control device 70 proceeds to step 308. If the second gas turbine engine 14 is currently running, the control device 70 proceeds to step 314. In step 308 shown in [Fig. 6], the control device 70 checks the air pressure at the pressure sensor 52 at an inlet of the compressor 32 to determine whether it is within the start-up limits of the second gas turbine engine 14, which are found in a lookup table stored in memory 71.If the pressure at the pressure sensor 52 is within the starting limits of the second gas turbine engine 14, the control device 70 maintains the flow control valve 64 in its current position as in step 310. If the pressure at the pressure sensor 52 is below the starting limits of the second gas turbine engine 14, the control device 70 will direct the flow control valve 64 to open by a predetermined amount as in step 312, and then return to step 308. In step 314 shown in [Fig. 6], the control device 70 checks the air pressure at the pressure sensor 52 at an inlet of the compressor 32 to determine if it is higher than the reference pressure corresponding to the certified altitude of the second gas turbine engine 14, which is found in a lookup table stored in memory 71. If the pressure at the pressure sensor 52 is lower than the certified altitude of the second gas turbine engine 14, the control device 70 will direct the flow control valve 64 to open by a predetermined amount and will direct the back pressure regulator 68 to adjust itself according to the pressure program stored in memory 71 in order to maintain the desired pressure ratio through the second gas turbine engine 14 as in a step 316, and then will return to step 314. If the pressure at the pressure sensor 52 is higher than the certified altitude of the second gas turbine engine 14, the control device 70 will proceed to a step 318. In step 318 shown in [Fig. 6], the control device 70 checks the air pressure at the pressure sensor 52 at an inlet of the compressor 32 to determine if it is higher than the reference pressure required to match the load on the electric generator 40, which is found in a lookup table stored in memory 71. If the pressure at the pressure sensor 52 is higher than the pressure required to match the load on the electric generator 40, the control device 70 will direct the flow control valve 64 to close by a predetermined amount as in step 320, and then return to step 314. If the pressure at the pressure sensor 52 is lower than the pressure required to match the load on the electric generator 40, the control device 70 will proceed to step 322. In step 322 shown in [Fig.6], the control device 70 checks the air pressure at the pressure sensor 52 at an inlet of the compressor 32 to determine if it is equal to the reference pressure required to adapt to the load on the electric generator 40, which is found in a lookup table stored in memory 71. If the pressure at the pressure sensor 52 is equal to the pressure required to adapt to the load on the electric generator 40, the control device 70 will direct the flow control valve 64 to remain in its current position as in step 324.If the pressure at the pressure sensor 52 is not equal to the pressure required to adapt to the load on the electric generator 40, the control device 70 will direct the flow control valve 64 to open by a predetermined amount and will direct the back pressure regulator 68 to adjust according to the pressure program stored in memory 71 in order to maintain the desired pressure ratio through the second gas turbine engine 14 as in a step 326, and then return to step 314. Modern aircraft are becoming increasingly electric, with electrically controlled compressors, electric in-flight entertainment, and electromechanical actuators instead of hydraulic ones. These loads can exceed the power generation capacity of the generators mounted on the aircraft's primary propulsion engines. Therefore, it may be desirable to generate additional electricity from auxiliary power units, which When they generate only electrical power, they can be called e-APUs. Auxiliary power units are typically much smaller gas turbines used for starting the main engine and generating some additional power when the main propulsion engines are shut down. An electric APU, or e-APU, would only provide electricity rather than intake air. To reduce costs, it may be desirable to use existing certified gas turbine engines for this purpose. This presents a challenge, however, as the most suitable gas turbines in terms of size are typically helicopter or general aviation engines, which are not certified for or capable of operating at the high altitudes of modern commercial airliners. Furthermore, these engines can experience significant vertical gradients in their power output capacity as altitude increases, limiting the amount of electrical power that can be supplied. One proposed solution to this challenge, as disclosed herein, may involve combining high-altitude, low-pressure, low-temperature ambient air with a small amount of compressor bleed air from the main propulsion engines to supply the e-APU's inlet stream. A gas jet ejector may be used to blend the low- and high-pressure streams to create a medium-pressure stream acceptable for the e-APU gas turbine engine. Other methods for combining the streams may also be feasible, such as a fast-opening, closing check valve on the compressed bleed air stream. One advantage of this arrangement is that lower altitude conditions can be simulated at the e-APU inlet. Using compressor bleed air from the main propulsion engine can reduce the operating line, increasing the pump margin for the main engine depending on the bleed location, whereas the more conventional method of directly extracting power from the shaft decreases the pump margin for the main engine. In the case of a three-spool engine with intermediate pressure (IP) compressor outlet bleed, this can increase the pump margin on the IP compressor but decrease the pump margin on the high pressure (HP) compressor.By varying the mixing ratio of compressor bleed air and ambient air to the e-APU, it is possible to allow the e-APU to operate without compressor bleed air up to its maximum envelope and then, as altitude increases, additional compressor bleed air can be routed to the e-APU. Possibly, given that the power output of the e-APU may decrease to As altitude increases, compressor bleed air can be selectively added even at the e-APU's nominal operating altitude to increase power capacity. A theoretical planned control algorithm for the e-APU could be as shown in the flowchart in [Fig. 6]. Aircraft main engines are typically designed to provide more high-pressure bleed air than the aircraft requires to account for failure scenarios. This disclosure utilizes excess bleed air from the main propulsion engines to drive an ejector that feeds the inlet of a second gas turbine engine used for power generation. At lower altitudes, no bleed air is drawn in, and subsequently, the programmed bleed air increases with altitude. The warm bleed air is mixed with cool ambient air to provide an acceptable inlet temperature for the power-generating gas turbine engine. The power-generating gas turbine engine is able to remain within its certified operating envelope by utilizing the high-pressure air produced by the ejector.This disclosure also includes an engine exhaust backpressure regulator to maintain acceptable exhaust conditions for engine operation. Although the disclosure has been illustrated and described in detail in the preceding drawings and description, it is to be regarded as exemplary and not restrictive, it being understood that only illustrative embodiments thereof have been represented and described and that all changes and modifications which fall within the spirit of the disclosure must be protected.

Claims

Demands

1. DC gas turbine motor system (10) comprising a first engine (12) which includes a first compressor, a first combustion chamber and first turbine, a second engine (14) which includes a second engine core and an airflow mixer (42), the core of the second engine including a second compressor, a second chamber combustion and a second turbine, the airflow mixer (42) being configured to receive an ambient airflow and combine the flow ambient air is mixed with compressed air to form a combined airflow, and further configured to direct the combined airflow into the core of second engine, and a sampling air control system (16) configured for to control the inlet pressure of the second compressor, the system air sampling control (16) comprising a duct, a flow control valve (64) and a control device (70), the conduit being configured to direct the compressed air from the first compressor to airflow mixer, control valve flow (64) being in fluidic communication with the conduit, the flow control valve (64) being configured to do selectively vary the flow of compressed air through the duct, and the control device (70) being programmed to adjust the valve flow control (64) in response to the fact that a pressure the inlet of the second compressor is less than a prede- completed so that a pressure of the combined flow which is directed into the second engine core allows the second engine core to operate in a low ambient pressure environment.

2. Gas turbine engine system of claim 1, wherein the airflow mixer (42) is a gas jet ejector, the jet ejector of gas being configured to use compressed air from the first Compressor as a working fluid to create a vacuum and draw in air ambient in the gas jet ejector and create the combined flow directed into the core of the second engine.

3. Gas turbine engine system of claim 2, wherein The gas jet ejector includes a nozzle (48) and an air inlet ambient arranged circumferentially around the nozzle (48), the nozzle (48) receives the compressed air flow and the ambient air inlet receives the ambient air.

4. Gas turbine engine system of claim 1, wherein the airflow mixer (42) is an inlet for ambient air and a non-return valve (244) in fluidic communication with the conduit, the non-return valve (244) being configured to direct the airflow compressed in the ambient air intake.

5. Gas turbine engine system of claim 1, wherein the second motor (14) includes a backpressure regulator in downstream of the second engine core, the backpressure regulator being configured to vary the pressure of an exhaust air from the second engine (14).

6. Gas turbine engine system of claim 1, wherein the The sampling control system also includes a heat exchanger. of heat (66) in fluidic communication with the duct, the exchanger heat (66) being configured to remove heat from the air compressed.

7. Gas turbine engine system of claim 6, wherein The heat exchanger (66) is configured to receive ambient air. to dissipate heat from the compressed air.

8. Gas turbine engine system of claim 6, wherein the The first engine (12) also includes a blower, and the heat exchanger of heat (66) is configured to receive air from the blower for to remove heat from the compressed air.

9. Gas turbine engine system of claim 5, wherein the control device (70) is programmed to vary the valve flow control (64) and backpressure regulator (68) based on a difference between the pressure at an inlet of the second compressor and a reference pressure value on a predetermined pressure program of the second turbine engine (14) gas-powered.

10. Gas turbine engine system of claim 1, wherein the The first compressor also includes an intermediate pressure compressor section having a first orifice in fluidic communication with the duct a high-pressure compressor section having a second port in fluidic communication with the conduit, and in which the control device (70) is programmed to open selectively one of the first and second orifices on the base of a pressure ratio relative to the predetermined flow rate of the first compressor stored on the control device (70).

11. Gas turbine engine system comprising a first engine (12) which includes a first compressor, a first combustion chamber and first turbine, a second engine (14) which includes a second engine core and an airflow mixer, the core of the second engine comprising a second compressor, a second combustion chamber and a second turbine, and the airflow mixer (42) being positioned at second engine core input level, and a sampling air control system (16) comprising a conduit, a flow control valve (64) and a device control (70), the conduit connecting the first compressor to air flow mixer, flow control valve (64) being in fluidic communication with the conduit, and in which the control device (70) is programmed to adjust the valve of flow control (64) in response to the fact that a pressure the inlet of the second compressor is less than a prede- Done.

12. Gas turbine engine system of claim 11, wherein the airflow mixer (42) is a gas jet ejector.

13. Gas turbine engine system of claim 12, wherein The gas jet ejector further includes an ambient air inlet and in which the compressed air from the first compressor and the air ambient air enters the ejector coaxially, the compressed air passing to through a nozzle positioned radially towards the inside of the air intake ambient.

14. Gas turbine engine system of claim 11, wherein the airflow mixer (42) is an inlet for ambient air and a non-return valve (244) in fluidic communication with the conduit.

15. Gas turbine engine system of claim 11, in which the second motor (14) includes a back pressure regulator (68) downstream of the second engine core.

16. | Gas turbine engine system of claim 11, wherein The sampling control system also includes a heat exchanger (66) in fluidic communication with the duct.

17. | Gas turbine engine system of claim 16, in which The heat exchanger (66) receives ambient air.

18. Gas turbine engine system of claim 16, wherein the first engine (12) also includes a blower, and the heat exchanger heat (66) receives air from the blower.

19. Gas turbine engine system of claim 11, wherein the first compressor also includes an intermediate pressure compressor section, an orifice of the intermediate pressure compressor in communication fluidic with the conduit a high-pressure compressor section, and an orifice of the high-pressure compressor in fluid communication with the conduit, and in which the control device (70) is programmed to selectively open the compressor pressure orifice intermediate and the high-pressure compressor orifice on the base of a pressure ratio relative to the predetermined flow rate of the first com- presser.

20. Method for controlling a gas turbine engine system, the process comprising the compression of a first airflow using a first com- pressurizer of a first gas turbine engine (12) to provide a first stream of compressed air, the compression of a second air stream using a second compressor of a second gas turbine engine (14), determining the inlet pressure of a second inlet compressor, and the direction of part of the first stream of compressed air towards the inlet of the second compressor in response to the fact that the inlet pressure of the the second compressor is below a predetermined value.

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