An aircraft energy and thermal integrated management system and method based on intake bleed air

Abstract

The present application belongs to the field of aviation, and relates to an aircraft energy and heat comprehensive management system and method based on inlet air bleed, which solves the problems of large demand for aircraft energy supply and heat management capability, low efficiency of traditional air refrigeration and large influence of flight state on working state. The system comprises an inlet air bleed opening (1), a gas-liquid heat exchanger (3), a gas turbine power device (4) and an exhaust assembly (8). The inlet air bleed opening (1) is opened on the wall surface of the aircraft inlet, the inlet air bleed opening (1) is connected with the inlet of the gas-liquid heat exchanger (3), the outlet of the gas-liquid heat exchanger (3) is directly connected with the air inlet of the gas turbine power device (4), and the transmission shaft of the gas turbine power device (4) is connected with a load. The present application is suitable for various aircrafts.

Description

Technical Field

[0001] This invention belongs to the field of aviation and relates to an integrated management system and method for aircraft energy and heat based on air intake bleed air. Background Technology

[0002] Aircraft energy systems, including hydraulic and electrical systems, provide hydraulic and electrical energy for flight control, avionics, and hydraulic and environmental control systems. Thermal management systems, such as environmental control and fuel systems, collect waste heat generated by electronic equipment, hydraulic pumps, generators, and other electromechanical equipment, and dissipate it into the atmosphere via ram air or fuel. High-power loads, such as high-power electronic equipment and electric actuators, consume a large amount of energy and generate a significant amount of waste heat during operation. Therefore, improving the power capacity and heat dissipation capacity of aircraft energy systems is a crucial capability requirement for future aircraft.

[0003] However, the energy harvesting and heat dissipation capabilities of aircraft platforms cannot be infinitely expanded. On the one hand, the power extracted by the energy system is limited by the available shaft power of the engine; on the other hand, the air cooling and fuel cooling capabilities are limited by the volume of the engine's fuel inlet and the engine's fuel inlet temperature. In response to the increasingly severe challenges to aircraft energy supply and heat dissipation capabilities, the US military proposed the concept of integrated energy and heat management on the F-35 aircraft. This integrates components such as air cooling turbines, integrated starter-generator motors, compressors, and power turbines into a single shaft system. Its advantage lies in achieving energy and heat conversion simultaneously in a single device through the integration of energy generation and heat dissipation components, thereby improving system integration and power-to-weight ratio. However, this technology faces numerous technical challenges, such as multi-modal turbines and high-speed coaxial integrated starter-generator motors, making its implementation difficult. Currently, only the PTMS system equipped on the US military's F-35 aircraft achieves integrated energy and heat management. Summary of the Invention

[0004] The purpose of this invention is to provide an integrated aircraft energy and thermal management system and method based on intake bleed air, which solves the problems of high demand for current aircraft energy supply and thermal management capabilities, low efficiency of traditional air cooling, and significant influence of flight conditions on operating status.

[0005] The technical solution of this invention:

[0006] An integrated energy and thermal management system for aircraft based on inlet bleed air includes:

[0007] 1. Intake duct bleed-out opening; 3. Gas-liquid heat exchanger; 4. Gas turbine power unit; 8. Exhaust assembly;

[0008] An air intake opening 1 is made on the wall of the aircraft air intake. The air intake opening 1 is connected to the inlet of the gas-liquid heat exchanger 3. The outlet of the gas-liquid heat exchanger 3 is directly connected to the air intake of the gas turbine power unit 4. The drive shaft of the gas turbine power unit 4 is connected to the load.

[0009] The load includes an accessory housing 5 and a generator 7. The shaft power input interface of the accessory housing 5 is connected to the drive shaft of the gas turbine power unit 4, and the shaft power output interface of the accessory housing 5 is connected to the generator 7.

[0010] The load includes an accessory housing 5 and a hydraulic pump 6. The shaft power input interface of the accessory housing 5 is connected to the drive shaft of the gas turbine power unit 4, and the shaft power output interface of the accessory housing 5 is connected to the hydraulic pump 6.

[0011] It also includes an air intake channel 2, and the air intake opening 1 of the air intake channel is connected to the inlet of the gas-liquid heat exchanger 3 through the air intake channel 2.

[0012] Air in the intake duct enters the gas-liquid heat exchanger 3 through the intake duct 1 to absorb heat, and then enters the gas turbine power unit 4 for combustion to do work. Exhaust gas is discharged through the exhaust assembly 8.

[0013] A method for integrated energy and thermal management of aircraft based on inlet bleed air includes:

[0014] After the gas turbine power unit 4 is started, under the suction of the internal compressor, the air in the intake duct enters the gas-liquid heat exchanger 3 through the intake duct 1 to absorb heat.

[0015] After absorbing heat, the air enters the gas turbine power unit 4 for combustion and work.

[0016] The output power of the gas turbine power unit 4 is supplied to the load connected to the gas turbine power unit 4.

[0017] The exhaust gas from the combustion of the gas turbine power unit 4 is discharged outside the machine through the exhaust assembly 8.

[0018] The load includes accessory housing 5, hydraulic pump 6 and generator 7;

[0019] The power output from the gas turbine power unit 4 is reduced in speed through the internal gear system of the accessory housing 5, which then drives the hydraulic pump 6 and the generator 7.

[0020] The beneficial effects of this invention are:

[0021] By adopting an air intake bleed scheme, the compression effect of the air intake during flight is utilized to increase the total air pressure at the system inlet, which is beneficial to improving the efficiency of the heat exchanger and the output power of the gas turbine power unit.

[0022] By utilizing the intake suction effect of the gas turbine power unit, the flow rate of the air-liquid heat exchanger is increased to improve the heat exchange power.

[0023] By utilizing the intake suction effect of a gas turbine power unit, the airflow to the cold side of the heat exchanger is controlled by adjusting the turbine speed at different flight airspeeds, thus solving the problem of unstable heat dissipation capacity of existing heat exchangers that directly use external ram air.

[0024] By utilizing the intake suction effect and shaft power output capability of the gas turbine power plant, it can simultaneously provide energy and thermal management capabilities for the aircraft during ground maintenance.

[0025] The turbine power unit is equipped with an accessory housing, which uses the gear system inside the housing to adjust the shaft power output speed to adapt to the speed requirements of different energy devices such as hydraulic pumps and generators.

[0026] Compared with foreign energy and thermal management schemes that design cooling turbines, generators, and gas turbine power units on the same axis, this scheme uses a gas-liquid heat exchanger instead of a cooling turbine. It takes advantage of the high temperature and large heat capacity of the liquid at the hot end of the heat exchanger to directly use the air in the intake duct for heat dissipation, eliminating the shaft connection between the cooling turbine and the gas turbine power unit and reducing technical difficulty. Attached Figure Description

[0027] Figure 1 This is a schematic diagram of the aircraft energy and thermal integrated management system based on intake bleed air, as per the present invention. Detailed Implementation

[0028] The technical solutions of the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are only some embodiments of the present invention, and not all embodiments. Based on the embodiments of the present invention, all other embodiments obtained by those of ordinary skill in the art without creative effort are within the scope of protection of the present invention.

[0029] This invention discloses an integrated energy and thermal management system and method for aircraft based on inlet bleed air. The system includes a liquid-air heat exchanger integrated into the inlet duct, a turbine power unit, a transmission casing, a hydraulic pump, a generator, and a turbine power exhaust assembly. Its operating principle is as follows: after the gas turbine power unit starts, air in the inlet duct is drawn in by the gas turbine power unit, passes through the heat exchanger in the bleed air duct, and enters the gas turbine power unit for combustion. Similar to conventional gas turbine power units, the shaft power generated by the turbine drives the hydraulic pump and generator through the transmission casing, providing electro-hydraulic energy to the aircraft. Simultaneously, the air in the bleed air duct, under the suction effect of the compressor, provides ram air energy to the heat exchanger, cooling the high-temperature liquid within it. The method for achieving integrated power and thermal management is as follows: when the gas turbine power unit is running, by controlling the speed of the gas turbine power unit, the airflow and heat exchange capacity into the heat exchanger are adjusted. Alternatively, the speed of the gas turbine power unit can be kept constant while altering the flow rate and heat exchange capacity of the hot-side fluid in the heat exchanger, thus achieving thermal management capabilities. By adjusting the fuel injection quantity in the internal combustion chamber of the gas turbine power unit according to the inlet air temperature and flow rate, the turbine output torque and output power can be changed, thereby achieving energy management capabilities.

[0030] This invention discloses an integrated aircraft energy and thermal management system based on inlet bleed air, such as... Figure 1 As shown, it includes:

[0031] 1. Intake duct bleed-out opening; 3. Gas-liquid heat exchanger; 4. Gas turbine power unit; 8. Exhaust assembly;

[0032] An air intake opening 1 is made on the wall of the aircraft air intake. The air intake opening 1 is connected to the inlet of the gas-liquid heat exchanger 3. The outlet of the gas-liquid heat exchanger 3 is directly connected to the air intake of the gas turbine power unit 4. The drive shaft of the gas turbine power unit 4 is connected to the load.

[0033] The load includes an accessory housing 5 and a generator 7. The shaft power input interface of the accessory housing 5 is connected to the drive shaft of the gas turbine power unit 4, and the shaft power output interface of the accessory housing 5 is connected to the generator 7.

[0034] The load includes an accessory housing 5 and a hydraulic pump 6. The shaft power input interface of the accessory housing 5 is connected to the drive shaft of the gas turbine power unit 4, and the shaft power output interface of the accessory housing 5 is connected to the hydraulic pump 6.

[0035] It also includes an air intake channel 2, and the air intake opening 1 of the air intake channel is connected to the inlet of the gas-liquid heat exchanger 3 through the air intake channel 2.

[0036] Air in the intake duct enters the gas-liquid heat exchanger 3 through the intake duct 1 to absorb heat, and then enters the gas turbine power unit 4 for combustion to do work. Exhaust gas is discharged through the exhaust assembly 8.

[0037] The operating principle of this system is as follows: After the gas turbine power unit starts, air in the intake duct is drawn in by the gas turbine power unit, passes through the heat exchanger in the duct, and enters the gas turbine power unit for combustion. Similar to traditional gas turbine power units, the shaft power generated by the turbine drives the hydraulic pump and generator through the transmission casing, providing electro-hydraulic energy to the aircraft. Simultaneously, the air in the duct, under the suction of the compressor, provides ram air energy to the heat exchanger, cooling the high-temperature liquid within. The system achieves integrated power and thermal management by: controlling the gas turbine power unit's speed to adjust the airflow and heat exchange capacity into the heat exchanger; or maintaining a constant gas turbine power unit speed while altering the flow rate and heat exchange capacity of the hot-side fluid in the heat exchanger, thus achieving thermal management. Based on the inlet air temperature and flow rate of the gas turbine power unit, the fuel injection quantity in its internal combustion chamber is adjusted, changing the turbine output torque and output power, thereby achieving energy management.

[0038] This invention also provides a method for integrated aircraft energy and thermal management based on inlet bleed air, comprising:

[0039] After the gas turbine power unit 4 is started, under the suction of the internal compressor, the air in the intake duct enters the gas-liquid heat exchanger 3 through the intake duct 1 to absorb heat.

[0040] After absorbing heat, the air enters the gas turbine power unit 4 for combustion and work.

[0041] The output power of the gas turbine power unit 4 is supplied to the load connected to the gas turbine power unit 4.

[0042] The exhaust gas from the combustion of the gas turbine power unit 4 is discharged outside the machine through the exhaust assembly 8.

[0043] The load includes accessory housing 5, hydraulic pump 6 and generator 7;

[0044] The power output from the gas turbine power unit 4 is reduced in speed through the internal gear system of the accessory housing 5, which then drives the hydraulic pump 6 and the generator 7.

[0045] In this method, the gas turbine power unit includes an intake duct structure whose function is to introduce air from the intake duct; the hot-side fluid of the liquid air heat exchanger is high-temperature fuel oil or coolant, and the cold-side fluid is air drawn into the gas turbine power unit from the intake duct; it consists of an intake casing, compressor, combustion chamber, and power turbine, and its main function is to burn fuel oil to obtain shaft power; the transmission casing functions to reduce the shaft power generated by the gas turbine unit and transmit it to energy devices such as hydraulic pumps and generators; the exhaust assembly functions to exhaust the high-temperature gas generated by the gas turbine unit to the outside of the machine.

[0046] Example

[0047] The system of this invention is arranged in a compartment near the aircraft's air intake. An air intake opening 1 is located on the air intake wall. An air intake duct 2 is installed between the air intake opening 1 and the inlet of a gas-liquid heat exchanger 3. The outlet of the gas-liquid heat exchanger 3 is directly connected to the air intake of a gas turbine power unit 4. An exhaust assembly 8 is connected to the exhaust port of the gas turbine power unit 4, discharging the exhaust gases after combustion to the outside of the aircraft. The accessory housing has three shaft power interfaces. One shaft power input interface is connected to the power output shaft of the gas turbine power unit 4, and the two functional output interfaces are connected to a hydraulic pump 6 and a generator 7, respectively.

[0048] After the gas turbine power unit starts, under the suction of its internal compressor, the air in the intake duct passes sequentially through the bleed air opening 1, the bleed air passage 2, the gas-liquid heat exchanger 3, the gas turbine power unit 4, and finally is discharged to the outside of the machine through the exhaust assembly 8. The air will lose heat from the hot side liquid in the gas-liquid heat exchanger 3 through the convection heat exchanger, and generate shaft power through compression, combustion and expansion processes in the gas turbine power unit.

[0049] After the gas turbine power unit is started, the shaft power generated by its power turbine unit 3 not only maintains the operation of its internal compressor, but also transmits the shaft power to the accessory housing 5 through a mechanical interface. After being reduced in speed by the gear system inside the accessory housing 5, it drives the hydraulic pump 6 and the generator 7, which can provide hydraulic and electrical energy to the aircraft platform.

[0050] The main feature of this embodiment is:

[0051] The system mainly consists of an exhaust duct, a liquid-air heat exchanger, a gas turbine power unit, an accessory casing, and an exhaust assembly.

[0052] Air is drawn out from the intake manifold through the airflow channel, and then passes through the liquid-air heat exchanger to exchange heat with the high-temperature liquid. Fuel enters the gas turbine power unit and is burned to generate shaft power output to the casing. Exhaust gas is discharged to the outside of the engine through the exhaust assembly.

[0053] Under the suction of the gas turbine unit, air and high-temperature liquid undergo forced convection heat exchange in the heat exchanger;

[0054] The power obtained by the gas turbine unit from burning fuel is partly converted into shaft power to supply the casing, and partly converted into the suction and compression power of the compressor.

[0055] To achieve continuous heat dissipation and shaft power, the gas turbine power unit should be in ignition operation at all times.

[0056] This patent addresses the current challenges of high energy supply and thermal management requirements in aircraft, the low efficiency of traditional air cooling systems, and the significant influence of flight platform on operational status. It proposes a comprehensive aircraft energy and thermal management system and method based on intake bleed air. This system utilizes an air-liquid heat exchanger, which is simpler and more reliable than traditional turbine coolers. The heat exchanger is positioned within the turbine power unit's intake duct, leveraging the turbine power unit's suction effect to increase the cold-side airflow of the heat exchanger and improve its efficiency. While the turbine power unit drives a hydraulic pump and generator via the casing to provide energy to the aircraft, it also utilizes the compressor's suction effect to provide active airflow control for the air heat exchanger. This system offers higher heat exchange efficiency than existing air heat exchangers and avoids the uncontrolled cold-side condition issues of existing air heat exchangers that directly use ram air. Furthermore, this system distributes turbine power between the suction air and shaft power extraction, resulting in enhanced energy distribution and management capabilities. This system utilizes air energy instead of mechanical energy as the transmission medium for energy and heat management, and has lower performance requirements for components such as turbines, heat exchangers, and generators. It is a solution for realizing the concept of advanced energy and integrated management technology in the current state of China's industrial capabilities.

[0057] The above description is merely a specific embodiment of the present invention, providing a detailed description of the invention. Parts not covered herein are conventional techniques. However, the scope of protection of the present invention is not limited thereto. Any variations or substitutions that can be easily conceived by those skilled in the art within the scope of the technology disclosed in the present invention should be included within the scope of protection of the present invention. The scope of protection of the present invention should be determined by the scope of the claims.

Claims

1. An integrated energy and thermal management system for aircraft based on inlet bleed air, characterized in that, include: Air intake opening (1), gas-liquid heat exchanger (3), gas turbine power unit (4), exhaust assembly (8); An air intake opening (1) is made on the wall of the aircraft air intake. The air intake opening (1) is connected to the inlet of the gas-liquid heat exchanger (3). The outlet of the gas-liquid heat exchanger (3) is directly connected to the air intake of the gas turbine power unit (4). The drive shaft of the gas turbine power unit (4) is connected to the load.

2. The aircraft energy and thermal integrated management system based on inlet bleed air as described in claim 1, characterized in that, The load includes an accessory housing (5) and a generator (7). The shaft power input interface of the accessory housing (5) is connected to the drive shaft of the gas turbine power unit (4), and the shaft power output interface of the accessory housing (5) is connected to the generator (7).

3. The aircraft energy and thermal integrated management system based on inlet bleed air as described in claim 1, characterized in that, The load includes an accessory housing (5) and a hydraulic pump (6). The shaft power input interface of the accessory housing (5) is connected to the drive shaft of the gas turbine power unit (4), and the shaft power output interface of the accessory housing (5) is connected to the hydraulic pump (6).

4. The aircraft energy and thermal integrated management system based on inlet bleed air as described in claim 1, characterized in that, It also includes an air intake channel (2), wherein the air intake opening (1) of the air intake channel is connected to the inlet of the gas-liquid heat exchanger (3) through the air intake channel (2).

5. The aircraft energy and thermal integrated management system based on inlet bleed air as described in claim 1, characterized in that, Air in the intake duct enters the gas-liquid heat exchanger (3) through the intake duct sump (1) to absorb heat, enters the gas turbine power unit (4) to burn and do work, and exhaust gas is discharged through the exhaust assembly (8).

6. A method for integrated energy and thermal management of aircraft based on inlet bleed air, characterized in that, The method for the aircraft energy and thermal integrated management system based on inlet bleed air as described in any one of claims 1 to 5 includes: After the gas turbine power unit (4) is started, under the suction action of the internal compressor, the air in the intake duct enters the gas-liquid heat exchanger (3) through the intake duct slurry opening (1) to absorb heat. After absorbing heat, the air enters the gas turbine power unit (4) for combustion and work. The output power of the gas turbine power unit (4) is supplied to the load connected to the gas turbine power unit (4).

7. The aircraft energy and thermal integrated management method based on inlet bleed air as described in claim 6, characterized in that, The exhaust gas from the combustion of the gas turbine power unit (4) is discharged outside the machine through the exhaust assembly (8).

8. The aircraft energy and thermal integrated management method based on inlet bleed air as described in claim 6, characterized in that, The load includes an accessory housing (5), a hydraulic pump (6), and a generator (7); The power output of the gas turbine power unit (4) is reduced by the gear system inside the accessory housing (5) and then drives the hydraulic pump (6) and generator (7) to work.

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