A variable mixture ratio design method for an aviation hybrid electric aircraft and a propulsion system

Abstract

The present application discloses a variable mixture ratio design method for an aviation hybrid electric aircraft and a propulsion system, including the following steps: S1, dividing the aircraft mission profile into take-off, climb and other stages, and establishing a linkage integrated design model for the hybrid electric aircraft and the propulsion system; S2, combining the various calculation formulas in the linkage integrated design model for the hybrid electric aircraft and the propulsion system to determine the aircraft and propulsion system parameters under different application scenarios; S3, by changing the mixture ratios in the take-off and climb stages under different application scenarios, optimizing and determining the mixture ratios corresponding to the minimum maximum take-off weight and the power cost per kilometer per unit commercial load under different application scenarios, achieving the optimal design of the aviation hybrid electric aircraft and the propulsion system, and providing a basis for the aircraft power selection. The present application can carry out hybrid electric system design based on existing turboshaft engines, achieve the optimal design of the aviation hybrid electric aircraft and the propulsion system, and provide a reliable basis for power selection.

Description

Technical Field

[0001] The present application relates to the field of aviation technology, and in particular, to a variable hybrid ratio design method for an aviation hybrid electric aircraft and a propulsion system. Background Art

[0002] Unlike traditional powered aircraft and pure electric powered aircraft, the range, weight, cost and other parameters of aviation turboshaft hybrid electric aircraft will also be affected by the distribution of the mixing ratio. The mixing ratio of the aviation turboshaft hybrid electric propulsion system refers to the proportion of the aircraft's power demand for the battery system in the total power demand, which should be determined according to the specific application scenario.

[0003] Compared with pure electric aircraft, hybrid electric aircraft have the characteristics of high fuel energy density and long range. However, in the face of special application scenarios, pure electric aircraft also have application value. In the competition with pure electric aircraft, the refined design of hybrid electric power system becomes the key, among which the power mixing ratio and engine selection are important parameters affecting the performance of hybrid electric power system.

[0004] In the existing hybrid electric design analysis, there is no practical method for determining the appropriate mixture ratio. Generally, the aircraft performance is only evaluated when the mixture ratio remains unchanged, and the variable design of the mixture ratio in different flight phases is not considered. At the same time, during the design, there is no linkage integration model between aircraft performance calculation and engine performance data. Generally, the propulsion system design is carried out by assuming the target engine, without considering how to quickly select a suitable engine from existing engines. Summary of the invention

[0005] The present application provides a variable mixture ratio design method for an aviation hybrid-electric aircraft and a propulsion system to solve the technical problem that the existing hybrid-electric design analysis does not provide a method for determining the optimal mixture ratio and cannot perform rapid engine selection.

[0006] This application is implemented through the following scheme:

[0007] A variable hybrid ratio design method for an aviation hybrid electric aircraft and a propulsion system, comprising the steps of:

[0008] S1. On the basis of dividing the aircraft mission profile into take-off, climb, cruise, descent, landing and go-around stages, a linkage integrated design model of the hybrid electric aircraft and the propulsion system is established, wherein the linkage integrated design model of the hybrid electric aircraft and the propulsion system includes a calculation formula for the maximum take-off weight of the aircraft, a calculation formula for the empty weight coefficient of the hybrid electric aircraft, a calculation formula for the fuel weight coefficient, a calculation formula for the weight coefficient of the energy storage system, a calculation formula for the power demand of the aircraft at each flight stage, a calculation formula for the weight of the engine used in the hybrid propulsion system, a calculation formula for the fuel consumption rate of the engine at each stage, a calculation formula for the weight of electrical components, and a calculation formula for the weight of the propulsion system;

[0009] S2. Combine the calculation formulas in the linkage integrated design model of the hybrid electric aircraft and the propulsion system, input the application scenario requirements, and determine the aircraft and propulsion system parameters under different application scenarios. The aircraft and propulsion system parameters include the maximum take-off weight, the maximum power capacity of the engine for the hybrid propulsion system, the weight of each electrical component, the fuel weight, and the weight of the energy storage system. The application scenario requirements include load, range, cruising altitude, flight speed at each stage, climb rate, and descent rate;

[0010] S3. By changing the mixture ratio in the take-off phase and the climb phase under different application scenarios, the mixture ratio corresponding to the lowest maximum take-off weight and the power cost per unit commercial load kilometer under different application scenarios is optimized and determined, which provides input for power cost evaluation and aircraft cost evaluation, realizes the optimal design of aviation hybrid-electric aircraft and propulsion system, and provides a basis for aircraft power selection.

[0011] Furthermore, in step S1, the maximum take-off weight calculation formula of the aircraft is specifically:

[0012] ; (1);

[0014] Where: W 0 is the maximum take-off weight of the aircraft; W e is the empty weight; W cp For load-bearing; W f is the fuel weight; W b is the weight of the energy storage system; W p is the weight of the propulsion system;

[0015] According to formula (1), the following formula can be obtained:

[0016] ; (2);

[0018] Where: W 0 is the maximum takeoff weight factor, is the empty machine weight factor; is the fuel weight factor; is the weight coefficient of the energy storage system; is the propulsion system weight factor.

[0019] Furthermore, the calculation formula of the empty weight coefficient of the hybrid electric aircraft is specifically:

[0020] ;

[0021] (3); where: a , c The values ​​are 1.4 and -0.1 respectively; K VS Related to aircraft structural design, revised based on existing aircraft data, W g，tra is the weight of a conventional engine.

[0022] Furthermore, the fuel weight coefficient calculation formula is obtained by respectively determining the aircraft weight change coefficient caused by fuel consumption at different stages according to the flight altitude, speed mission parameters, combined with the aircraft power-to-weight ratio, lift-to-drag ratio and engine fuel consumption rate, and specifically includes the steps of:

[0023] Calculate the weight change coefficient during takeoff:

[0024] ; (4);

[0026] Where: W 1 is the weight of the aircraft at the end of the take-off phase; P m is the power-to-weight ratio of the aircraft during takeoff; sfc to is the fuel consumption rate of the aircraft during takeoff, which comes from the calculation results of the overall performance simulation model; η g2p is the energy conversion efficiency from engine to propeller; Hp to is the mixture ratio at takeoff stage;

[0027] Calculate the weight change coefficient during the climb phase:

[0028] ; (5);

[0030] Where: W 2 is the weight of the aircraft at the end of the take-off phase; H To help the aircraft climb to a certain altitude; V vcl is the aircraft climb rate; ld cl is the lift-to-drag ratio of the aircraft during the climbing phase; V cl It is the flight speed of the aircraft during the climbing phase; sfc cl is the fuel consumption rate of the aircraft during the climbing phase, which comes from the calculation results of the overall performance simulation model; η p is the propeller efficiency; Hpcl is the mixture ratio during the climbing phase;

[0031] Calculate the weight change coefficient during the cruise phase:

[0032] ; (6);

[0034] Where: W 3 is the weight of the aircraft at the end of the cruising phase; R cr For cruising voyage; V cr is the cruising speed; ld cr is the lift-to-drag ratio of the aircraft during cruising phase; sfc cr is the fuel consumption rate of the aircraft during the cruising phase, which comes from the calculation results of the overall performance simulation model;

[0035] Calculate the coefficient of weight change during the descent phase:

[0036] ; (7);

[0038] In the formula, W 4 is the weight of the aircraft at the end of the descent phase; V vdec is the aircraft descent rate; ld dec is the lift-to-drag ratio of the aircraft during the descent phase; V dec is the flight speed of the aircraft during the descent phase; sfc dec is the fuel consumption rate of the aircraft during the descent phase, which comes from the calculation results of the overall performance simulation model;

[0039] Calculate the weight change factor for landing / go-around phase:

[0040] The power requirement during the landing / go-around phase is the same as that during the take-off phase. The weight change factor during the landing / go-around phase is , determined with reference to the take-off phase;

[0041] Calculation formula for obtaining fuel weight coefficient:

[0042] According to the weight change coefficient of each stage, considering the residual fuel coefficient of 1.06, the aircraft fuel weight coefficient is determined as:

[0043] ; (8);

[0045] In the formula, It is the weight change coefficient during landing / go-around phase.

[0046] Furthermore, the energy storage system weight coefficient calculation formula is obtained by the following steps:

[0047] Calculate the energy storage system weight factor during takeoff:

[0048] ; (9);

[0050] Where: W bto The weight requirement of the energy storage system during the takeoff phase; t to The aircraft take-off phase time; η b2p is the energy conversion efficiency from the energy storage system to the propeller; E sb is the energy density of the energy storage system;

[0051] Calculate the energy storage system weight factor during the climbing phase:

[0052] ; (10);

[0054] Calculate the weight coefficient of the energy storage system during the cruise phase:

[0055] ; (11);

[0057] Calculate the weight coefficient of the energy storage system during the descent phase:

[0058] The engine does not stop during the descent phase, and only the engine provides power. The energy storage system does not work. The weight coefficient of the energy storage system during the descent phase =0 ;

[0059] Calculate the energy storage system weight factor during landing / go-around phase:

[0060] Energy storage system weight factor during landing / go-around phase , determined with reference to the take-off phase;

[0061] Calculate the energy storage system weight coefficient:

[0062] According to the energy storage system weight coefficient at each stage, taking into account 20% of unavailable electricity, the total value of the energy storage system weight coefficient is determined. At the same time, when calculating the energy storage system weight coefficient, the power requirements for the energy storage system are taken into account:

[0063] ; (12);

[0065] Where: is the weight coefficient of the energy storage system during the descent phase; is the weight coefficient of the energy storage system during the landing / go-around phase.

[0066] Furthermore, the calculation formula for the aircraft power demand at each flight stage includes:

[0067] Power calculation during takeoff:

[0068] Calculate the takeoff power based on the maximum takeoff weight and power-to-weight ratio of the aircraft:

[0069] ; (13);

[0071] Power calculation during the climbing phase:

[0072] ; (14);

[0074] Power calculation during cruise phase:

[0075] ; (15);

[0077] Power calculation during the descent phase:

[0078] ; (16);

[0080] Power calculation for landing / go-around phase:

[0081] The power requirement during the landing / go-around phase is the same as that during the take-off phase and is determined by referring to the power calculation for the take-off phase.

[0082] Furthermore, the weight calculation formula of the engine for the hybrid propulsion system is obtained by the following steps:

[0083] Establish a power-weight relationship table for existing turboshaft engines, where the power refers to the maximum power capability of the engine at sea level and standard atmospheric conditions;

[0084] Determine the weight of a conventional engine:

[0085] According to the power of each stage, the maximum power requirement of the traditional engine is calculated. The specific formula is as follows:

[0086] ; (17);

[0088] Where: α is the engine power altitude attenuation coefficient, which is 0.9~0.95 / km; βis the engine power margin coefficient, ranging from 0.7 to 0.8; η re is the efficiency of the turboprop engine reducer, ranging from 0.985 to 0.99;

[0089] According to the maximum power requirement and the power-weight relationship table of existing turboshaft engines, the engine with the smallest maximum power capacity in the relationship table is determined as the traditional engine under the condition that the maximum power capacity is not less than the power requirement, and the calculation formula for the weight of the traditional engine is obtained as follows:

[0090] ; (18);

[0092] Where: n For the number of traditional engines on the aircraft, the relationship between the number of engines and the aircraft takeoff weight is established based on the statistics of existing aircraft; W g，tra，0 is the weight of a single conventional engine; λ The weight increase coefficient of the reducer is 1.25 to 1.35;

[0093] Determine the weight of an engine for a hybrid propulsion system:

[0094] According to the power of the engine at each stage, the maximum power requirement of the engine used in the hybrid propulsion system is calculated. The specific formula is as follows:

[0095] ; (19);

[0097] According to the maximum power requirement and the power-weight relationship table of existing turboshaft engines, and under the condition that the maximum power capacity is not lower than the power requirement, the engine that meets the minimum maximum power capacity in the relationship table is determined as the engine for the hybrid propulsion system, and the weight of the engine for the hybrid propulsion system is obtained.

[0098] Furthermore, the engine fuel consumption rate at each stage is obtained by the following steps:

[0099] Establish a power-fuel consumption relationship table for existing turboshaft engines of various specifications at different altitudes;

[0100] According to the determined engine for the hybrid propulsion system, the power-fuel consumption relationship table of the corresponding engine is selected; according to the power-fuel consumption relationship table and the altitudes of different flight stages, the power-fuel consumption relationship table corresponding to the altitude of each flight stage is obtained through linear interpolation. When determining the fuel consumption rate of a certain power, the specific formula is:

[0101] ; (20);

[0103] Where: sfc i The engine power-fuel consumption rate relationship close to but lower than the current flight stage altitude in the original power-fuel consumption rate relationship table; sfc i+1 The engine power-fuel consumption rate relationship close to but higher than the current flight stage altitude in the original power-fuel consumption rate relationship table; H i It is the altitude in the original power-fuel consumption rate relationship table, close to but lower than the altitude of the current flight stage; H i+1 It is the altitude in the original power-fuel consumption rate relationship table, which is close to but higher than the altitude of the current flight stage; H fly is the current flight phase altitude;

[0104] According to the calculated aircraft power requirements at each stage, the power of the engine used in the hybrid propulsion system at each stage is determined. The calculation formula is as follows:

[0105] ; (twenty one);

[0107] Where: P j The power requirement of the aircraft at the current stage; Hp j is the mixing ratio of the current stage;

[0108] Using the power-fuel consumption rate relationship at the current stage altitude and the engine power of the hybrid propulsion system, the fuel consumption rate of the current flight stage is obtained through linear interpolation. The calculation formula is as follows:

[0109] ; (twenty two);

[0111] Where: sfc j,k The fuel consumption rate of the power-fuel consumption rate corresponding to the altitude of the current flight stage is close to but lower than the power of the current flight stage; sfc j,k+1 The fuel consumption rate of the power corresponding to the altitude of the current flight stage is close to but higher than the fuel consumption rate of the current flight stage; P j,k The power of the current flight phase in the power-fuel consumption rate relationship corresponding to the altitude is close to but lower than the power of the current flight phase; P j,k+1 The altitude close to but higher than the current flight stage altitude in the power-fuel consumption rate relationship corresponding to the current flight stage altitude;

[0112] After obtaining the fuel consumption rate of each stage, substitute it into the formula of each stage of the fuel weight coefficient calculation formula for the fuel weight coefficient calculation.

[0113] Furthermore, the electrical component weight calculation formula is obtained through the following steps:

[0114] The weight of each electrical component is related to its power level. The maximum power during takeoff, climb and cruise phases is selected, and the efficiency along the way is considered to obtain the nominal power of each electrical component. Finally, the weight of each electrical component is obtained according to the power-to-weight ratio of the electrical component. The weight of the generator is calculated as follows:

[0115] ; (twenty three);

[0117] Where: η al2p is the energy conversion efficiency from generator to propeller; ρ al is the generator power density.

[0118] Furthermore, the propulsion system weight calculation formula is as follows:

[0119] After determining the weight of the engine and the weight of each electrical component used in the hybrid propulsion system, the weight of the propulsion system is obtained:

[0120] ; (twenty four);

[0122] Where: W engine The weight of the engine for the hybrid propulsion system; W elec,l is the weight of different electrical components.

[0123] Compared with the prior art, this application has the following beneficial effects:

[0124] The present application provides a variable mixture ratio design method for an aviation hybrid electric aircraft and a propulsion system. The method divides the aircraft mission profile into take-off, climb, cruise, descent, landing and go-around stages, and configures different appropriate mixture ratios in high-power stages such as take-off and climb to achieve the optimal design of the aviation hybrid electric aircraft and the propulsion system. The present application establishes an integrated design model for the hybrid electric aircraft and the propulsion system, and can utilize existing turboshaft engine products, such as the power-fuel consumption characteristics of existing turboshaft engines, according to different usage scenarios, to achieve the purpose of designing a hybrid electric system based on existing turboshaft engines. The present application provides a strategy for designing hybrid electric aircraft according to different usage scenarios, to achieve the lowest propulsion system cost, and to reduce the weight of the aircraft as much as possible, while providing a scientific and accurate basis for cost evaluation and power selection of the aircraft.

[0125] In addition to the above-described purposes, features and advantages, the present application also has other purposes, features and advantages. The present application will be further described in detail with reference to the drawings. BRIEF DESCRIPTION OF THE DRAWINGS

[0126] The drawings constituting a part of the present application are used to provide a further understanding of the present application. The illustrative embodiments of the present application and their descriptions are used to explain the present application and do not constitute an improper limitation on the present application.

[0127] Figure 1 It is a flow chart of a variable hybrid ratio design method for an aviation hybrid-electric aircraft and a propulsion system according to a preferred embodiment of the present application.

[0128] Figure 2 It is a schematic diagram of the principle of a hybrid electric propulsion system of a preferred embodiment of the present application. DETAILED DESCRIPTION

[0129] The embodiments of the present application are described in detail below in conjunction with the accompanying drawings, but the present application can be implemented in a variety of different ways defined and covered below.

[0130] like Figure 1 and Figure 2 As shown, the preferred embodiment of the present application provides a variable hybrid ratio design method for an aviation hybrid electric aircraft and a propulsion system, comprising the steps of:

[0131] S1. On the basis of dividing the aircraft mission profile into take-off, climb, cruise, descent, landing and go-around stages, a linkage integrated design model of the hybrid electric aircraft and the propulsion system is established, wherein the linkage integrated design model of the hybrid electric aircraft and the propulsion system includes a calculation formula for the maximum take-off weight of the aircraft, a calculation formula for the empty weight coefficient of the hybrid electric aircraft, a calculation formula for the fuel weight coefficient, a calculation formula for the weight coefficient of the energy storage system, a calculation formula for the power demand of the aircraft at each flight stage, a calculation formula for the weight of the engine used in the hybrid propulsion system, a calculation formula for the fuel consumption rate of the engine at each stage, a calculation formula for the weight of electrical components, and a calculation formula for the weight of the propulsion system;

[0132] S2. Combine the calculation formulas in the linkage integrated design model of the hybrid electric aircraft and the propulsion system, input the application scenario requirements, and determine the aircraft and propulsion system parameters under different application scenarios. The aircraft and propulsion system parameters include the maximum take-off weight, the maximum power capacity of the engine for the hybrid propulsion system, the weight of each electrical component, the fuel weight, and the weight of the energy storage system. The application scenario requirements include load, range, cruising altitude, flight speed at each stage, climb rate, and descent rate;

[0133] S3. By changing the mixture ratio in the take-off phase and the climb phase under different application scenarios, the mixture ratio corresponding to the lowest maximum take-off weight and the power cost per unit commercial load kilometer under different application scenarios is optimized and determined, which provides input for power cost evaluation and aircraft cost evaluation, realizes the optimal design of aviation hybrid-electric aircraft and propulsion system, and provides a basis for aircraft power selection.

[0134] The present embodiment provides a variable mixture ratio design method for an aviation hybrid electric aircraft and a propulsion system. The method divides the aircraft mission profile into take-off, climb, cruise, descent, landing and go-around stages, and configures different appropriate mixture ratios in high-power stages such as take-off and climb to achieve the optimal design of the aviation hybrid electric aircraft and the propulsion system. The present embodiment establishes a linkage integrated design model for the hybrid electric aircraft and the propulsion system, and can utilize existing turboshaft engine products, such as the power-fuel consumption characteristics of existing turboshaft engines, according to different usage scenarios, to achieve the purpose of designing a hybrid electric system based on existing turboshaft engines. The present embodiment provides a strategy for designing hybrid electric aircraft according to different usage scenarios, to achieve the lowest propulsion system cost, and to reduce the weight of the aircraft as much as possible, while providing a scientific and accurate basis for cost evaluation and power selection of the aircraft.

[0135] Preferably, in step S1, the maximum take-off weight calculation formula of the aircraft is specifically:

[0136] ; (1);

[0138] Where: W0 is the maximum take-off weight of the aircraft; W e is the empty weight; W cp For load-bearing; W f is the fuel weight; W b is the weight of the energy storage system; W p is the weight of the propulsion system;

[0139] According to formula (1), the following formula can be obtained:

[0140] ; (2);

[0142] Where: W 0 is the maximum takeoff weight factor, is the empty machine weight factor; is the fuel weight factor; is the weight coefficient of the energy storage system; is the propulsion system weight factor.

[0143] Preferably, the calculation formula for the empty weight coefficient of the hybrid electric aircraft is specifically:

[0144] ; (3);

[0146] Where: a , c The values ​​are 1.4 and -0.1 respectively; K VS Related to aircraft structural design, revised based on existing aircraft data, W g，tra is the weight of a conventional engine.

[0147] Generally, the empty weight of an aircraft includes the weight of the engine, and the calculation formula for the empty weight coefficient of a fixed-wing aircraft is calculated using the following empirical formula. However, the empty weight here refers to the empty weight without the traditional engine, because the mass of the hybrid electric propulsion system includes the weight of the engine. Therefore, when calculating the empty weight coefficient of the hybrid electric aircraft, this embodiment needs to subtract the traditional engine weight coefficient from the existing empirical formula for the empty weight coefficient of the fixed-wing aircraft.

[0148] Preferably, the fuel weight coefficient calculation formula is obtained by determining the aircraft weight change coefficient caused by fuel consumption at different stages according to the flight altitude, speed mission parameters, combined with the aircraft power-to-weight ratio, lift-to-drag ratio and engine fuel consumption rate, and specifically includes the steps of:

[0149] Calculate the weight change coefficient during takeoff:

[0150] ; (4);

[0152] Where: W 1 is the weight of the aircraft at the end of the take-off phase; P m is the power-to-weight ratio of the aircraft during takeoff; sfc to is the fuel consumption rate of the aircraft during takeoff, which comes from the calculation results of the overall performance simulation model; η g2p is the energy conversion efficiency from engine to propeller; Hp to is the mixture ratio at takeoff stage;

[0153] Calculate the weight change coefficient during the climb phase:

[0154] ; (5);

[0156] Where: W 2 is the weight of the aircraft at the end of the take-off phase; H To help the aircraft climb to a certain altitude; V vcl is the aircraft climb rate; ld cl is the lift-to-drag ratio of the aircraft during the climbing phase; V cl It is the flight speed of the aircraft during the climbing phase; sfc cl is the fuel consumption rate of the aircraft during the climbing phase, which comes from the calculation results of the overall performance simulation model; η p is the propeller efficiency; Hp cl is the mixture ratio during the climbing phase;

[0157] Calculate the weight change coefficient during the cruise phase:

[0158] ; (6);

[0160] Where: W 3 is the weight of the aircraft at the end of the cruising phase; R cr For cruising voyage; V cr is the cruising speed; ld cr is the lift-to-drag ratio of the aircraft during cruising phase; sfccr is the fuel consumption rate of the aircraft during the cruising phase, which comes from the calculation results of the overall performance simulation model;

[0161] Calculate the coefficient of weight change during the descent phase:

[0162] ; (7);

[0164] In the formula, W 4 is the weight of the aircraft at the end of the descent phase; V vdec is the aircraft descent rate; ld dec is the lift-to-drag ratio of the aircraft during the descent phase; V dec is the flight speed of the aircraft during the descent phase; sfc dec is the fuel consumption rate of the aircraft during the descent phase, which comes from the calculation results of the overall performance simulation model;

[0165] Calculate the weight change factor for landing / go-around phase:

[0166] The power requirement during the landing / go-around phase is the same as that during the take-off phase. The weight change factor during the landing / go-around phase is , determined with reference to the take-off phase;

[0167] Calculation formula for obtaining fuel weight coefficient:

[0168] According to the weight change coefficient of each stage, considering the residual fuel coefficient of 1.06, the aircraft fuel weight coefficient is determined as:

[0169] ; (8);

[0171] In the formula, It is the weight change coefficient during landing / go-around phase.

[0172] During the flight, fuel consumption will cause the aircraft weight to change at each stage. This embodiment can determine the aircraft weight change coefficient caused by fuel consumption at different stages based on mission parameters such as flight altitude and speed, combined with technical parameters such as aircraft power-to-weight ratio, lift-to-drag ratio, and engine fuel consumption rate.

[0173] Preferably, the amount of electricity in the energy storage system will gradually decrease during the process of providing electricity. Assuming that the energy density remains unchanged, the demand for the weight of the energy storage system can be converted. Therefore, the energy storage system weight coefficient calculation formula is obtained through the following steps:

[0174] Calculate the energy storage system weight factor during takeoff:

[0175] ; (9);

[0177] Where: W bto The weight requirement of the energy storage system during the takeoff phase; t to The aircraft take-off phase time; η b2p is the energy conversion efficiency from the energy storage system to the propeller; E sb is the energy density of the energy storage system;

[0178] Calculate the energy storage system weight factor during the climbing phase:

[0179] ; (10);

[0181] Calculate the weight coefficient of the energy storage system during the cruise phase:

[0182] ; (11);

[0184] Calculate the weight coefficient of the energy storage system during the descent phase:

[0185] The engine does not stop during the descent phase, and only the engine provides power. The energy storage system does not work. The weight coefficient of the energy storage system during the descent phase ;

[0186] Calculate the energy storage system weight factor for the landing / go-around phase: Energy storage system weight factor for the landing / go-around phase , determined with reference to the take-off phase;

[0187] Calculate the weight coefficient of the energy storage system: According to the weight coefficient of the energy storage system at each stage, consider 20% of the unavailable electricity and determine the total value of the weight coefficient of the energy storage system. At the same time, when calculating the weight coefficient of the energy storage system, take into account the power requirements of the energy storage system:

[0188] ; (12);

[0190] Where: is the weight coefficient of the energy storage system during the descent phase; is the weight coefficient of the energy storage system during the landing / go-around phase.

[0191] Preferably, the calculation formula for aircraft power demand at each flight stage includes:

[0192] Power calculation during takeoff: Calculate the power during takeoff based on the maximum takeoff weight and power-to-weight ratio of the aircraft:

[0193] ; (13);

[0195] Power calculation during the climbing phase:

[0196] ; (14);

[0198] Power calculation during cruise phase:

[0199] ; (15);

[0201] Power calculation during the descent phase:

[0202] ; (16);

[0204] Power calculation for landing / go-around phase: The power requirement for landing / go-around phase is the same as that for take-off phase, and is determined by referring to the power calculation for take-off phase.

[0205] Preferably, the weight calculation formula for the engine used in the hybrid propulsion system is obtained through the following steps: establish a power-weight relationship table of existing turboshaft engines, where the power refers to the maximum power capacity of the engine at sea level and standard atmospheric conditions. Examples are shown in Table 1. There are five turboshaft engines with different parameters.

[0206] Table 1: Power-weight relationship of existing turboshaft engines

[0207] ;

[0208] Determine the weight of a conventional engine:

[0209] According to the power of each stage, the maximum power requirement of the traditional engine is calculated. The specific formula is as follows:

[0210] ; (17);

[0212] Where: α is the engine power altitude attenuation coefficient, which takes into account the characteristic that the available engine power decreases when the altitude increases, and the value is 0.9~0.95 / km; β is the engine power margin factor, which takes into account the atmospheric temperature change and the margin required for safety, and its value is 0.7 to 0.8; η re is the efficiency of the turboprop engine reducer, which is mainly based on the structural difference between the turboprop engine and the turboshaft engine, and the value is 0.985~0.99;

[0213] According to the maximum power requirement and the power-weight relationship table of existing turboshaft engines, the engine with the smallest maximum power capacity in the relationship table is determined as the traditional engine under the condition that the maximum power capacity is not less than the power requirement, and the calculation formula for the weight of the traditional engine is obtained as follows:

[0214] ; (18);

[0216] Where: n For the number of traditional engines on the aircraft, the relationship between the number of engines and the aircraft takeoff weight is established based on the statistics of existing aircraft; W g，tra，0 is the weight of a single conventional engine; λ The weight increase coefficient of the reducer is mainly due to the structural difference between turboprop engines and turboshaft engines, and the value is 1.25 to 1.35;

[0217] Determine the weight of an engine for a hybrid propulsion system:

[0218] According to the power of the engine at each stage, the maximum power requirement of the engine used in the hybrid propulsion system is calculated. The specific formula is as follows:

[0219] ; (19);

[0221] According to the maximum power requirement and the power-weight relationship table of existing turboshaft engines, and under the condition that the maximum power capacity is not lower than the power requirement, the engine that meets the minimum maximum power capacity in the relationship table is determined as the engine for the hybrid propulsion system, and the weight of the engine for the hybrid propulsion system is obtained.

[0222] Preferably, the engine fuel consumption rate at each stage is obtained by the following steps:

[0223] Taking into account the different altitudes at different flight stages, a power-fuel consumption rate relationship table of existing turboshaft engines of various specifications at different altitudes is established, as shown in Table 2:

[0224] Table 2: Power-fuel consumption relationship of 1100kW engine

[0225] ;

[0226] According to the determined engine for the hybrid propulsion system, the power-fuel consumption relationship table of the corresponding engine is selected; according to the power-fuel consumption relationship table and the altitude of different flight stages, the power-fuel consumption relationship table corresponding to the altitude of each flight stage is obtained through linear interpolation, as shown in Table 3:

[0227] Table 3: Power-fuel consumption rate relationship table for different flight phases and corresponding altitudes

[0228] ;

[0229] Therefore, when determining the fuel consumption rate of a certain power, the specific formula is:

[0230] ; (20);

[0232] Where: sfc i The engine power-fuel consumption rate relationship close to but lower than the current flight stage altitude in the original power-fuel consumption rate relationship table; sfc i+1 The engine power-fuel consumption rate relationship close to but higher than the current flight stage altitude in the original power-fuel consumption rate relationship table; H i It is the altitude in the original power-fuel consumption rate relationship table, close to but lower than the altitude of the current flight stage; H i+1 It is the altitude in the original power-fuel consumption rate relationship table, which is close to but higher than the altitude of the current flight stage; H fly is the current flight phase altitude;

[0233] According to the calculated aircraft power requirements at each stage, the power of the engine used in the hybrid propulsion system at each stage is determined. The calculation formula is as follows:

[0234] ; (twenty one);

[0236] Where: P j The power requirement of the aircraft at the current stage; Hp j is the mixing ratio of the current stage;

[0237] Using the power-fuel consumption rate relationship at the current stage altitude and the engine power of the hybrid propulsion system, the fuel consumption rate of the current flight stage is obtained through linear interpolation. The calculation formula is as follows:

[0238] ; (twenty two);

[0240] Where: sfc j,k The fuel consumption rate of the power-fuel consumption rate corresponding to the altitude of the current flight stage is close to but lower than the power of the current flight stage; sfc j,k+1 The fuel consumption rate of the power corresponding to the altitude of the current flight stage is close to but higher than the fuel consumption rate of the current flight stage; Pj,k The power of the current flight phase in the power-fuel consumption rate relationship corresponding to the altitude of the current flight phase, which is close to but lower than the power of the current flight phase; P j,k+1 The altitude close to but higher than the current flight stage altitude in the power-fuel consumption rate relationship corresponding to the current flight stage altitude;

[0241] After obtaining the fuel consumption rate of each stage, substitute it into the formula of each stage of the fuel weight coefficient calculation formula for the fuel weight coefficient calculation.

[0242] Preferably, the electrical component weight calculation formula is obtained by the following steps:

[0243] The weight of each electrical component is related to its power level. The maximum power of takeoff, climb and cruise phases is selected, and the efficiency along the way is considered to obtain the nominal power of each electrical component. Finally, the weight of each electrical component is obtained according to the power-to-weight ratio of the electrical component. The weight of the generator is calculated as follows. Other components can refer to:

[0244] ; (twenty three);

[0246] Where: η al2p is the energy conversion efficiency from generator to propeller; ρ al is the generator power density.

[0247] Preferably, the propulsion system weight calculation formula is specifically:

[0248] After determining the weight of the engine and the weight of each electrical component used in the hybrid propulsion system, the weight of the propulsion system is obtained:

[0249] ; (twenty four);

[0251] Where: W engine The weight of the engine for the hybrid propulsion system; W elec,l is the weight of different electrical components.

[0252] According to the method provided by the present invention, the hybrid electric propulsion system applicable to the five application scenarios in Table 4 is calculated to determine the appropriate mixing ratio and the maximum take-off weight and other parameters.

[0253] Table 4: Parameters of different application scenarios

[0254] ;

[0255] Taking application scenario 3 as an example, by using the method of the present invention, the influence of the mixing ratio on the maximum take-off weight can be obtained. At the same time, combined with the calculation formula of the cost application, the influence of the mixing ratio on the power cost per unit commercial load kilometer can also be obtained.

[0256] It can be seen from Tables 5 and 6 that when the climb mixture ratio is 0 and the takeoff mixture ratio is 0.5, the maximum takeoff weight and the power cost per unit commercial load kilometer are both near the lowest level.

[0257] Table 5: Maximum takeoff weights corresponding to different mixing ratios under application scenario 3 (unit: kg)

[0258] ;

[0259] Table 6: Power cost per kilometer of unit commercial load corresponding to different mixing ratios under application scenario 3 (unit: yuan / (ton·km))

[0260] ;

[0261] According to the calculation results of different application scenarios, the optimal take-off phase mixture ratio and cruise phase mixture ratio were obtained, and the maximum take-off weight, the maximum power capacity of the engine for the hybrid propulsion system, the weight of each electrical component, the fuel weight and the energy storage system weight were determined. See Table 7 for details.

[0262] Table 7: Aircraft and propulsion system parameters in different application scenarios

[0263] ;

[0264] By using the data obtained by the method of the present invention and combining it with the aircraft cost calculation method, the aircraft unit commercial load per kilometer cost can be obtained, and then compared with the pure electric propulsion system in terms of takeoff weight and cost. As can be seen from Table 8, in application scenario 1, pure electric and hybrid electric are relatively close, but with the increase of range, speed and altitude, the advantages of hybrid electric are further highlighted.

[0265] Table 8

[0266]

[0267] The above embodiments of the present application provide the following features:

[0268] 1) A variable mixture ratio design method for aviation hybrid electric aircraft and propulsion system is proposed. By configuring different mixture ratios in high-power stages such as takeoff and climb, the optimal design of aviation hybrid electric aircraft and propulsion system is achieved;

[0269] 2) An integrated design model of the hybrid electric aircraft and propulsion system was established, and the power-fuel consumption characteristics of the existing turboshaft engine were used to achieve the purpose of designing a hybrid electric system based on the existing turboshaft engine;

[0270] 3) It provides strategies for designing hybrid electric aircraft according to different usage scenarios, and provides a basis for the selection of aircraft power.

[0271] In summary, this application establishes an integrated design model of hybrid-electric aircraft and propulsion system, utilizes the power-fuel consumption characteristics of existing turboshaft engines, and realizes the rapid design of hybrid electric systems based on existing turboshaft engines; sets the mixture ratio in takeoff and cruise phases respectively, and iteratively optimizes and determines the appropriate mixture ratio, and provides a strategy for designing hybrid-electric aircraft according to different usage scenarios, providing a basis for cost evaluation and power selection of aircraft.

[0272] Although the preferred embodiments of the present application have been described, those skilled in the art may make other changes and modifications to these embodiments once they have learned the basic creative concept. Therefore, the appended claims are intended to be interpreted as including the preferred embodiments and all changes and modifications falling within the scope of the present application.

[0273] Obviously, those skilled in the art can make various changes and modifications to the present application without departing from the spirit and scope of the present application. Thus, if these modifications and variations of the present application fall within the scope of the claims of the present application and their equivalents, the present application is also intended to include these modifications and variations.

Claims

1. A variable hybrid ratio design method for an aviation hybrid electric aircraft and a propulsion system, characterized in that: Includes steps: S1. On the basis of dividing the aircraft mission profile into take-off, climb, cruise, descent, landing and go-around stages, a linkage integrated design model of the hybrid electric aircraft and the propulsion system is established, wherein the linkage integrated design model of the hybrid electric aircraft and the propulsion system includes a calculation formula for the maximum take-off weight of the aircraft, a calculation formula for the empty weight coefficient of the hybrid electric aircraft, a calculation formula for the fuel weight coefficient, a calculation formula for the weight coefficient of the energy storage system, a calculation formula for the power demand of the aircraft at each flight stage, a calculation formula for the weight of the engine used in the hybrid propulsion system, a calculation formula for the fuel consumption rate of the engine at each stage, a calculation formula for the weight of electrical components, and a calculation formula for the weight of the propulsion system; S2. Combine the calculation formulas in the linkage integrated design model of the hybrid electric aircraft and the propulsion system, input the application scenario requirements, and determine the aircraft and propulsion system parameters under different application scenarios. The aircraft and propulsion system parameters include the maximum take-off weight, the maximum power capacity of the engine for the hybrid propulsion system, the weight of each electrical component, the fuel weight, and the weight of the energy storage system. The application scenario requirements include load, range, cruising altitude, flight speed at each stage, climb rate, and descent rate; S3. By changing the mixture ratio in the take-off phase and the climb phase under different application scenarios, the mixture ratio corresponding to the lowest maximum take-off weight and the power cost per unit commercial load kilometer under different application scenarios is optimized and determined, which provides input for power cost evaluation and aircraft cost evaluation, realizes the optimal design of aviation hybrid-electric aircraft and propulsion system, and provides a basis for aircraft power selection.

2. The variable mixture ratio design method of the aviation hybrid electric aircraft and propulsion system according to claim 1, characterized in that: In step S1, the maximum take-off weight calculation formula of the aircraft is specifically: ； Where: W 0 is the maximum take-off weight of the aircraft; W e is the empty weight; W cp For load-bearing; W f is the fuel weight; W b is the weight of the energy storage system; W p is the weight of the propulsion system; According to the calculation formula of the maximum take-off weight of the aircraft, the following formula can be obtained: ； Where: W 0 is the maximum takeoff weight factor, is the empty machine weight factor; is the fuel weight factor; is the weight coefficient of the energy storage system; is the propulsion system weight factor.

3. The variable mixture ratio design method of the aviation hybrid electric aircraft and propulsion system according to claim 2 is characterized in that: The calculation formula for the empty weight coefficient of the hybrid electric aircraft is specifically: ； Where: a , c The values ​​are 1.4 and -0.1 respectively; K VS Related to aircraft structural design, revised based on existing aircraft data, W g，tra is the weight of a conventional engine.

4. The variable mixture ratio design method of the aviation hybrid electric aircraft and propulsion system according to claim 3 is characterized in that: The fuel weight coefficient calculation formula is obtained by determining the aircraft weight change coefficient caused by fuel consumption at different stages based on the flight altitude, speed mission parameters, combined with the aircraft power-to-weight ratio, lift-to-drag ratio and engine fuel consumption rate, and specifically includes the following steps: Calculate the weight change coefficient during takeoff: ； Where: W 1 is the weight of the aircraft at the end of the take-off phase; P m is the power-to-weight ratio of the aircraft during takeoff; sfc to is the fuel consumption rate of the aircraft during takeoff, which comes from the calculation results of the overall performance simulation model; η g2p is the energy conversion efficiency from engine to propeller; Hp to is the mixture ratio at takeoff stage; Calculate the weight change coefficient during the climb phase: ； Where: W 2 is the weight of the aircraft at the end of the take-off phase; H To help the aircraft climb to a higher altitude; V vcl is the aircraft climb rate; ld cl is the lift-to-drag ratio of the aircraft during the climbing phase; V cl It is the flight speed of the aircraft during the climbing phase; sfc cl is the fuel consumption rate during the climbing phase, which comes from the calculation results of the overall performance simulation model; η p is the propeller efficiency; Hp cl is the mixture ratio during the climbing phase; Calculate the weight change coefficient during the cruise phase: ； Where: W 3 is the weight of the aircraft at the end of the cruising phase; R cr For cruising voyage; V cr is the cruising speed; ld cr is the lift-to-drag ratio of the aircraft during cruising phase; sfc cr is the fuel consumption rate of the aircraft during the cruising phase, which comes from the calculation results of the overall performance simulation model; Calculate the coefficient of weight change during the descent phase: ； In the formula, W 4 is the weight of the aircraft at the end of the descent phase; V vdec is the aircraft descent rate; ld dec is the lift-to-drag ratio of the aircraft during the descent phase; V dec is the flight speed of the aircraft during the descent phase; sfc dec is the fuel consumption rate of the aircraft during the descent phase, which comes from the calculation results of the overall performance simulation model; Calculate the weight change factor for landing / go-around phase: The power requirement during the landing / go-around phase is the same as that during the take-off phase. The weight change factor during the landing / go-around phase is , determined with reference to the take-off phase; Calculation formula for obtaining fuel weight coefficient: According to the weight change coefficient of each stage, considering the residual fuel coefficient of 1.06, the aircraft fuel weight coefficient is determined as: ； In the formula, It is the weight change coefficient during landing / go-around phase.

5. The variable mixture ratio design method of the aviation hybrid electric aircraft and propulsion system according to claim 4, characterized in that: The energy storage system weight coefficient calculation formula is obtained by the following steps: Calculate the energy storage system weight factor during takeoff: ； Where: W bto The weight requirement of the energy storage system during the takeoff phase; t to The aircraft take-off phase time; η b2p is the energy conversion efficiency from the energy storage system to the propeller; E sb is the energy density of the energy storage system; Calculate the energy storage system weight factor during the climbing phase: ； Calculate the weight coefficient of the energy storage system during the cruise phase: ； Calculate the weight coefficient of the energy storage system during the descent phase: The engine does not stop during the descent phase, and only the engine provides power. The energy storage system does not work. The weight coefficient of the energy storage system during the descent phase ; Calculate the energy storage system weight factor during landing / go-around phase: Energy storage system weight factor during landing / go-around phase , determined with reference to the take-off phase; Calculate the weight coefficient of the energy storage system: According to the energy storage system weight coefficient at each stage, taking into account 20% of unavailable electricity, the total value of the energy storage system weight coefficient is determined. At the same time, when calculating the energy storage system weight coefficient, the power requirements for the energy storage system are taken into account: ； Where: is the weight coefficient of the energy storage system during the descent phase; is the weight coefficient of the energy storage system during the landing / go-around phase.

6. The variable mixture ratio design method of the aviation hybrid electric aircraft and propulsion system according to claim 5, characterized in that: The calculation formula for aircraft power requirements at each flight stage includes: Power calculation during takeoff: Calculate the takeoff power based on the maximum takeoff weight and power-to-weight ratio of the aircraft: ; Power calculation during the climbing phase: ； Power calculation during cruise phase: ； Power calculation during the descent phase: ； Power calculation for landing / go-around phase: The power requirement during the landing / go-around phase is the same as that during the take-off phase and is determined by referring to the power calculation for the take-off phase.

7. The variable mixture ratio design method of the aviation hybrid electric aircraft and propulsion system according to claim 6, characterized in that: The weight calculation formula for the hybrid propulsion system engine is obtained by the following steps: Establish a power-weight relationship table for existing turboshaft engines, where the power refers to the maximum power capability of the engine at sea level and standard atmospheric conditions; Determine the weight of a conventional engine: According to the power of each stage, the maximum power requirement of the traditional engine is calculated. The specific formula is as follows: ； Where: α is the engine power altitude attenuation coefficient, which is 0.9~0.95 / km; β is the engine power margin coefficient, ranging from 0.7 to 0.8; η re is the efficiency of the turboprop engine reducer, ranging from 0.985 to 0.99; According to the maximum power requirement and the power-weight relationship table of existing turboshaft engines, the engine with the smallest maximum power capacity in the relationship table is determined as the traditional engine under the condition that the maximum power capacity is not less than the power requirement, and the calculation formula for the weight of the traditional engine is obtained as follows: ； Where: n For the number of traditional engines on the aircraft, the relationship between the number of engines and the aircraft takeoff weight is established based on the statistics of existing aircraft; W g，tra，0 is the weight of a single conventional engine; λ The weight increase coefficient of the reducer is 1.25 to 1.35; Determine the weight of an engine for a hybrid propulsion system: According to the power of the engine at each stage, the maximum power requirement of the engine used in the hybrid propulsion system is calculated. The specific formula is as follows: ； According to the maximum power requirement and the power-weight relationship table of existing turboshaft engines, and under the condition that the maximum power capacity is not lower than the power requirement, the engine that meets the minimum maximum power capacity in the relationship table is determined as the engine for the hybrid propulsion system, and the weight of the engine for the hybrid propulsion system is obtained.

8. The variable mixture ratio design method of the aviation hybrid electric aircraft and propulsion system according to claim 7, characterized in that: The engine fuel consumption rate at each stage is obtained through the following steps: Establish a power-fuel consumption relationship table for existing turboshaft engines of various specifications at different altitudes; According to the determined engine for the hybrid propulsion system, the power-fuel consumption relationship table of the corresponding engine is selected; according to the power-fuel consumption relationship table and the altitudes of different flight stages, the power-fuel consumption relationship table corresponding to the altitude of each flight stage is obtained through linear interpolation. When determining the fuel consumption rate of a certain power, the specific formula is: ； Where: sfc i The engine power-fuel consumption rate relationship close to but lower than the current flight stage altitude in the original power-fuel consumption rate relationship table; sfc i+1 The engine power-fuel consumption rate relationship close to but higher than the current flight stage altitude in the original power-fuel consumption rate relationship table; H i It is the altitude in the original power-fuel consumption rate relationship table, close to but lower than the altitude of the current flight stage; H i+1 It is the altitude in the original power-fuel consumption rate relationship table, which is close to but higher than the altitude of the current flight stage; H fly is the altitude of the current flight phase; According to the calculated aircraft power requirements at each stage, the power of the engine used in the hybrid propulsion system at each stage is determined. The calculation formula is as follows: ； Where: P j The power requirement of the aircraft at the current stage; Hp j is the mixing ratio of the current stage; Using the power-fuel consumption rate relationship at the current stage altitude and the engine power of the hybrid propulsion system, the fuel consumption rate of the current flight stage is obtained through linear interpolation. The calculation formula is as follows: ； Where: sfc j,k The fuel consumption rate of the power-fuel consumption rate corresponding to the altitude of the current flight stage is close to but lower than the power of the current flight stage; sfc j,k+1 The fuel consumption rate of the power corresponding to the altitude of the current flight stage is close to but higher than the fuel consumption rate of the current flight stage; P j,k The power of the current flight phase in the power-fuel consumption rate relationship corresponding to the altitude is close to but lower than the power of the current flight phase; P j,k+1 The altitude close to but higher than the current flight stage altitude in the power-fuel consumption rate relationship corresponding to the current flight stage altitude; After obtaining the fuel consumption rate of each stage, substitute it into the formula of each stage of the fuel weight coefficient calculation formula for the fuel weight coefficient calculation.

9. The variable mixture ratio design method of the aviation hybrid electric aircraft and propulsion system according to claim 8, characterized in that: The electrical component weight calculation formula is obtained through the following steps: The weight of each electrical component is related to its power level. The maximum power during takeoff, climb and cruise phases is selected, and the efficiency along the way is considered to obtain the nominal power of each electrical component. Finally, the weight of each electrical component is obtained according to the power-to-weight ratio of the electrical component. The weight of the generator is calculated as follows: ； Where: η al2p is the energy conversion efficiency from generator to propeller; ρ al is the generator power density.

10. The variable mixture ratio design method of the aviation hybrid electric aircraft and propulsion system according to claim 9, characterized in that: The specific formula for calculating the weight of the propulsion system is: After determining the weight of the engine and the weight of each electrical component used in the hybrid propulsion system, the weight of the propulsion system is obtained: ； Where: W engine The weight of the engine for the hybrid propulsion system; W elec,l is the weight of different electrical components.

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