A method for evaluating the benefits of aviation hybrid electric propulsion systems

Abstract

The present application discloses a method for evaluating the benefits of an aviation hybrid electric propulsion system, including the following steps: S1, obtaining the input of the benefit evaluation from the aircraft manufacturer; S2, establishing a component coefficient table for hybrid electric, pure electric, and traditional propulsion systems; S3, calculating the weight of the gas turbine engine and the like according to the input of the benefit evaluation and the component coefficient table; S4, obtaining an iterative calculation formula for the cruise stage distance applicable to different propulsion systems, and determining the cruise stage distance through iterative calculation; S5, calculating the comprehensive total cost per kilometer of the unit commercial load, the maximum power capability index under high temperature conditions at sea level, and other related parameters; S6, taking the relevant parameters of the traditional propulsion system as a benchmark, and combining the relevant parameters of the propulsion system to be evaluated to calculate the relative proportion of the corresponding indicators of the propulsion system, and configuring the weight coefficient to calculate the comprehensive benefit coefficient of the propulsion system to be evaluated. The present application improves the accuracy of the evaluation results and provides a reliable basis for the design of the flight platform and the power selection.

Description

Technical Field

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

[0002] In the context of energy conservation and emission reduction, how to reduce the fuel consumption of the flight platform propulsion system has become a widely concerned issue. Since pure electric power is limited by issues such as battery energy density, the aviation turboshaft hybrid electric propulsion system has become an important development direction, especially the use of a series architecture, where the engine only generates electricity through a generator, rather than directly driving the propeller, thereby achieving mechanical decoupling of the engine and the propeller, and promoting distributed propulsion to become an important layout method for propeller or ducted fan aircraft. Compared with traditional propeller aircraft powered by turboprop engines, distributed propulsion can effectively improve the aerodynamic efficiency of the flight platform. As a power output gas turbine engine, the aviation turboshaft engine can directly drive the generator and has a natural advantage in building a hybrid electric propulsion system.

[0003] Although the goal of energy conservation and emission reduction can affect the selection of aircraft propulsion system solutions to a certain extent, in order to strongly promote the development, production and application of hybrid electric propulsion systems, it is necessary to ensure that they have comprehensive benefits relative to traditional and pure electric propulsion systems, including maximum range, unit cost, power margin and fuel consumption. At present, the benefit evaluation of hybrid electric propulsion systems only considers the aerodynamic benefits of aircraft, and rarely involves the impact of dedicated engines. The accuracy of the data cannot be guaranteed, and no effective technical indicators have been established, making it difficult to support aircraft developers in making scientific and reasonable choices between hybrid electric, pure electric and traditional propulsion systems. Summary of the invention

[0004] The present application provides a method for evaluating the benefits of an aviation hybrid electric propulsion system to solve the technical problem that the data accuracy of existing benefit evaluation methods cannot be guaranteed, making it difficult for aircraft developers to make scientific and reasonable choices among hybrid electric, pure electric and traditional propulsion systems.

[0005] The technical solutions adopted in this application are as follows:

[0006] A method for evaluating the benefits of an aviation hybrid electric propulsion system comprises the following steps:

[0007] S1. Obtaining inputs for benefit evaluation from the aircraft manufacturer, including at least effective load, commercial load, typical flight profile power spectra of different power schemes, and maximum power requirements of the aircraft for high temperature / high altitude flight, wherein the sum of the weights of the propulsion system, energy storage system, and fuel system can be obtained by subtracting the commercial load from the effective load, and the typical flight profile power spectrum includes the flight altitudes during takeoff, climb, cruise, landing, and landing go-around, the aircraft's power requirements for the propulsion system, working time, and flight speed;

[0008] S2. Establish component coefficient tables for hybrid electric, pure electric, and traditional propulsion systems to characterize the applicability or quantity of electrical components;

[0009] S3. Calculate the weight of the gas turbine engine, the weight of the generator, the weight of the generator controller, the weight of the drive motor controller, the weight of the drive motor, the weight of the power distribution system, the total weight of the cooling system, the total weight of the propulsion system, the total weight of the battery system and battery system parameters, and the fuel quality based on the input of the benefit evaluation and the component coefficient table;

[0010] S4. Establish an overall performance simulation model of a turboshaft engine for a hybrid electric propulsion system and an overall performance simulation model of a gas turbine engine for a conventional power scheme, calculate the engine fuel consumption rate at each stage according to the power specified by the typical flight profile power spectrum of different propulsion systems, and establish an iterative calculation formula for the cruise stage distance of the fuel part and an iterative calculation formula for the cruise stage distance of the battery part in parallel, thereby obtaining an iterative calculation formula for the cruise stage distance applicable to different propulsion systems, and determine the cruise stage distance through iterative calculation;

[0011] S5. Calculate the comprehensive total cost per kilometer of unit commercial load, the maximum power capability index under high temperature conditions at sea level, the maximum power capability index under high temperature conditions at plateau, the fuel consumption per kilometer of unit commercial load and the noise index of the propulsion system according to the maximum range, number of flights and relevant prices of the propulsion system to be evaluated based on the cruise stage distance;

[0012] S6. Taking the maximum range of the traditional propulsion system, the comprehensive total cost per kilometer of unit commercial load, the maximum power capability index under high temperature conditions at sea level, the maximum power capability index under high temperature conditions at plateau, the fuel consumption per kilometer of unit commercial load and the noise index of the propulsion system as the benchmark, the relative proportion of the corresponding indicators of the propulsion system is calculated in combination with the maximum range of the propulsion system to be evaluated, the comprehensive total cost per kilometer of unit commercial load, the maximum power capability index under high temperature conditions at sea level, the maximum power capability index under high temperature conditions at plateau, the fuel consumption per kilometer of unit commercial load and the noise index of the propulsion system, and the weight coefficient is configured to calculate the comprehensive benefit coefficient of the propulsion system to be evaluated.

[0013] Further, in step S3:

[0014] The weight of a gas turbine engine is calculated as follows:

[0015] m engine = max((1- H pto ) to ,(1- H pcl ) cl / α,P cr / α) / (η al d1 ×η alc d1 ×η sw d3- d1 ×η moc d2 ×η mo d2 ) / β / ρ en

[0016] Where: H pto , H pcl are the power mixing ratios in the take-off and climb phases, respectively. The traditional propulsion system takes a value of 0, the pure electric propulsion system takes a value of 1, and the hybrid electric propulsion system takes a value of 0 to 1; P to Pcl is the power requirement of the aircraft for the propulsion system during the take-off phase, in kW; Pcl is the power requirement of the aircraft for the propulsion system during the climb phase, in kW; cr is the power demand of the aircraft for the propulsion system during the cruise phase, in kW; α is the engine power altitude attenuation coefficient, with a value of 0.95~0.9 / km; η al is the generator efficiency; η alc is the generator controller efficiency; η sw is the efficiency of the power distribution switch; η moc is the efficiency of the drive motor controller; η mo is the efficiency of the driving motor; β is the engine power margin coefficient, which is 0.7 to 0.8; ρ en is the engine power density. The turboshaft engine dedicated to the hybrid electric propulsion system is 1.29 to 1.35 times that of the turboprop engine. d1 is the generator coefficient, d2 is the drive motor coefficient, and d3 is the power distribution system coefficient.

[0017] The weight of the generator is calculated by the following formula:

[0018] m al = d1×max((1- H pto ) to , (1- H pcl ) cl , P cr ) / (η al ×η alc ×η swd3- d1 ×η moc ×η mo ) / ρ al

[0019] Where: al is the generator power density;

[0020] The weight of the generator controller is calculated by the following formula:

[0021] m alc = d1×max((1- H pto ) to , (1- H pcl ) cl , P cr ) / (η alc ×η sw d3- d1 ×η moc ×η mo ) / ρ alc

[0022] Where: alc is the generator controller power density;

[0023] The weight of the drive motor controller is calculated by the following formula:

[0024] m moc = d2×max(P to ,P cl ,P cr ) / (η moc ×η mo ) / ρ moc

[0025] Where: moc is the power density of the drive motor controller;

[0026] The weight of the drive motor is calculated by the following formula:

[0027] m mo = d2×max(P to ,P cl ,P cr ) / η mo / ρ mo

[0028] Where: mo is the power density of the drive motor;

[0029] The weight of the power distribution system is calculated using the following formula:

[0030] m sw = d3×m sw0

[0031] Where: m sw0 The weight is for a single distribution switch.

[0032] Furthermore, the total weight of the cooling system is calculated by the following steps:

[0033] Calculate the weight of the cooling system for the gas turbine engine and generator section:

[0034] m col,al =d1×max((1- H pto ) to , (1- H pcl ) cl , P cr )×(1-η al ×η sw ) / ρ co

[0035] Where: co is the power-to-weight ratio of the cooling system;

[0036] Calculate the weight of the drive motor cooling system:

[0037] m col,mo = d2×max(P to ,P cl ,P cr ) / (η mo ×η moc ×η sw ) ×(1-η mo ×η moc ×η sw ) / ρ co

[0038] Calculate the weight of the cooling system on the battery system side:

[0039] m col,bas =d2×max(H pto ×P to , H pcl ×P cl , (1-d1)×P cr )×P to ×(1-η sw d1 ) / ρ co

[0040] Calculate the total cooling system weight:

[0041] .

[0042] Furthermore, the total weight of the propulsion system is calculated by the following formula:

[0043] .

[0044] Furthermore, the total weight of the battery system and the battery system parameters are calculated as follows:

[0045] Calculate the weight of the energy storage system according to power requirements:

[0046] m bas,P= d2×max(H pto ×P to , H pcl ×P cl , (1- d1) × P cr ) / ( E bas ×C)

[0047] Where: E bas is the energy density of the battery system; C is the discharge rate, which represents the discharge capacity of the battery system;

[0048] Calculate the weight of the energy storage system according to energy demand:

[0049] m bas,E= d2×(1- d1) ×(m load -m pay -m pop )+ d1(H pto ×P to ×t+H pcl ×P cl ×t cl ) / E bas ×γ

[0050] Where: m load is the aircraft payload, m pay is the commercial load of the aircraft, γ is the battery residual power coefficient, which is 1.1 to 1.2, and t is the total time of the take-off, landing and go-around phase;

[0051] Calculate the total battery system weight:

[0052]

[0053] Different groups of battery system parameters are selected, including battery system energy density and discharge rate, and the total weight of the battery system corresponding to different battery system parameters is compared to determine the battery system parameters with the smallest total weight of the battery system.

[0054] Furthermore, the fuel quality is calculated by the following formula:

[0055] .

[0056] Furthermore, the step S4 specifically includes the steps of:

[0057] The overall performance simulation model of the turboshaft engine for the hybrid electric propulsion system and the overall performance simulation model of the gas turbine engine for the traditional power scheme are established respectively, and the engine fuel consumption rate at each stage is calculated according to the power specified by the typical flight profile power spectrum of different propulsion systems;

[0058] Calculate the energy conversion efficiency from the engine to the aircraft thruster input:

[0059]

[0060] Calculate the energy conversion efficiency from the energy storage system to the input of the aircraft thruster:

[0061]

[0062] Establish the iterative calculation formula for the cruise phase distance of the fuel part:

[0063] m load -m pay -m pop -m bas =(1- H pto ) × P to / η g2p ×SFC to ×t+((1- H pcl ) × P cl / η g2p ×SFC cl ×t cl +P cr / η g2p ×SFC cr ×(L cr1 / V cr )+d4∑(P i / η g2p ×SFC i ×t i ))×δ

[0064] Where: SFC to is the fuel consumption rate of the engine during takeoff; t is the working time during takeoff, landing and go-around; SFC cl is the fuel consumption rate of the engine during the climbing phase, P cl is the power in the climbing stage; t cl P is the working time of the climbing stage; cr is the cruise power; SFC cr is the fuel consumption rate of the engine during cruising stage; L cr1 The distance of the cruise phase of the fuel part needs to be determined by iterative calculation; V cr is the flight speed in the cruising phase; P i 、SFCi ,t i are the power, engine fuel consumption rate and working time of the remaining stages respectively; δ is the residual oil coefficient, ranging from 1.06 to 1.1, and d4 is the engine coefficient;

[0065] Establish the iterative calculation formula for the cruise phase distance of the battery part: d2×(1-d1) ×m bas =(d2×(1-d1)×H pto ×P to ×t+ d2×(1-d1) ×H pcl ×P cl ×t cl +P cr ×(L cr2 / V cr )+d2×(1-d1) ×∑(P i / t i )) / η b2p / E bas ×γ

[0066] Where, L cr2 For the cruise phase distance of the battery part, the iterative calculation formula for the cruise phase distance applicable to different propulsion systems is obtained. The cruise phase distance is determined by iterative calculation:

[0067] L cr =max(L cr1 , L cr2 )

[0068] Where: L cr is the final calculated cruise stage distance.

[0069] Furthermore, the step S5 specifically includes the steps of:

[0070] Establish the calculation formula of the maximum range L:

[0071] L=V cl ×t cl + L cr + V dec ×t dec

[0072] Where: V cl is the flight speed during the climb phase, V dec is the flight speed during the descent phase, t dec Working hours for the descent phase;

[0073] Calculate the purchase cost per kilometer of the battery system per commercial load by combining the maximum range and number of flights:

[0074] W bas = Ebas ×m bas ×U bas / (F×L×m pay )

[0075] Where: U bas is the price corresponding to the unit energy capacity of the battery system; F is the number of available discharges of the battery system, corresponding to the number of flights during the life of the battery system;

[0076] The purchase cost per kilometer of the unit commercial load of the aircraft engine is calculated based on the time life and cycle life:

[0077] W engine= MAX(F×t f / t engine , F / F engine ) ×MAX((1-H pto ) to , P cl / α, P cr / α) ×U eigine / (F×L×m pay )

[0078] Where: t f is the time of a single flight; t engine The time allowed for the engine to be used; F engine U is the number of cycles allowed for the engine; eigine is the price corresponding to the unit power of the engine;

[0079] Calculate the purchase cost per kilometer of the aircraft engine generator, generator controller, drive motor controller, drive motor, and cooling system:

[0080] W mo =F×t f / t mo ×m mo ×ρ mo ×U mo / (F×L×m pay )

[0081] W al = F × t f / t al ×m al ×ρ al ×U al / (F×L×m pay )

[0082] W alc = F × tf / t alc ×m alc ×ρ alc ×U alc / (F×L×m pay )

[0083] W moc = F × t f / t moc ×m moc ×ρ moc ×U moc / (F×L×m pay )

[0084] W cold = F × t f / t cold ×m cold ×ρ cold ×U cold / (F×L×m pay )

[0085] Where: t mo U is the time allowed for the drive motor to be used; mo is the price corresponding to the unit power of the drive motor;

[0086] The formula for calculating the purchasing cost per kilometer of unit commercial load for establishing a power distribution system is:

[0087] W sw =d3×U sw / (F×L×m pay )

[0088] Where: U sw is the unit price of the distribution switch;

[0089] Calculate the total purchase cost per kilometer of commercial load:

[0090] ;

[0091] Calculate the total operating cost per kilometer of commercial load:

[0092] W c= (m bas ×E bas / γ×U elc + m fule ×U fule / δ) / (L×m pay )

[0093] Where: U elc , U fule They are the unit mass cost of fuel and the unit cost of electricity;

[0094] The comprehensive total cost per kilometer of commercial load is calculated based on the total purchase cost per kilometer of commercial load and the total operating cost per kilometer of commercial load:

[0095] ;

[0096] Calculate the maximum power capability index at sea level under high temperature conditions:

[0097]

[0098] Where: P pop,hot P is the maximum power capability of the engine at sea level and high temperature; dem,hot The maximum power requirement of the aircraft under high temperature conditions at sea level;

[0099] Calculate the maximum power capability index under high temperature conditions in plateau:

[0100]

[0101] Where: P pop,pla P is the maximum power capability of the engine under high temperature conditions at plateau; dem,pla The maximum power requirement of the aircraft under high temperature conditions on the plateau;

[0102] Calculate the aircraft's fuel consumption per kilometer of commercial load:

[0103] M c= (m fule / δ) / (L×m pay )

[0104] According to the mixture ratio during takeoff and cruise, the propulsion system noise index is calculated:

[0105] N c= H pto ×H pcl .

[0106] Furthermore, the step S6 specifically includes the steps of:

[0107] Taking the maximum range of the traditional propulsion system, the comprehensive total cost per kilometer of unit commercial load, the maximum power capability index under high temperature conditions at sea level, the maximum power capability index under high temperature conditions at plateau, the fuel consumption per kilometer of unit commercial load and the noise index of the propulsion system as the benchmark, the relative proportion of the corresponding indicators of the propulsion system is calculated in combination with the maximum range of the propulsion system to be evaluated, the comprehensive total cost per kilometer of unit commercial load, the maximum power capability index under high temperature conditions at sea level, the maximum power capability index under high temperature conditions at plateau, the fuel consumption per kilometer of unit commercial load and the noise index of the propulsion system to be evaluated, and the weight coefficient is configured to calculate the comprehensive benefit coefficient of the propulsion system to be evaluated:

[0108] Z = a1 × L / L tra + a2×W / W tra + a3×N p,hot / N p,hot,tra +a4×N p,pla / N p,pla,tra -a5×M c / M c,tra - a6×N c / N c,tra

[0109] Where: a1, a2, a3, a4, a5, a6 are weight coefficients, all of which are positive; L, W, N p,hot 、N p,pla 、M c 、N c They are the maximum range of the propulsion system to be evaluated, the comprehensive total cost per kilometer of unit commercial load, the maximum power capability index under high temperature conditions at sea level, the maximum power capability index under high temperature conditions at plateau, the fuel consumption per kilometer of unit commercial load and the noise index of the propulsion system; L tra , W tra 、N p,hot,tra 、N p,pla,tra 、M c,tra 、N c,tra They are the maximum range of the traditional propulsion system, the comprehensive total cost per unit commercial load kilometer, the maximum power capability index under high temperature conditions at sea level, the maximum power capability index under high temperature conditions at plateau, the fuel consumption per unit commercial load kilometer and the propulsion system noise index.

[0110] Furthermore, some weight coefficients satisfy the following conditions:

[0111] a1+a2+a3+a4=1;

[0112] The weight coefficients a5 and a6 are adjusted accordingly according to the requirements of the application scenario for fuel consumption or emissions and power noise.

[0113] Compared with the prior art, this application has the following beneficial effects:

[0114] The present application provides a method for evaluating the benefits of an aviation hybrid electric propulsion system. By using the benefit evaluation method of the present application, according to the power requirements of the aircraft and from the perspective of engine design, the comprehensive benefit coefficient of the propulsion system of the hybrid electric propulsion system scheme relative to the gas turbine engine propulsion system scheme is obtained, including the maximum power and power margin, and the maximum power and power margin benefits, fuel consumption benefits, weight benefits, cost benefits, etc. are obtained, so as to obtain more accurate information. The accuracy of the benefit evaluation result is high, which supports the aircraft developer to make a scientific and reasonable choice between the three propulsion systems of hybrid electric, pure electric and traditional, thereby providing a more reliable and important basis for the scheme design and power selection of the flight platform, ensuring that the power selection of the hybrid electric propulsion system scheme achieves a higher power output, and a power system with a more advanced technical level can be obtained, which can effectively save the development cost brought by the re-development or serial development of the engine; ensuring that the selected hybrid electric propulsion system scheme has both high thermal efficiency and low fuel consumption, and at the same time has the power requirements that meet different intake conditions, cruising, and emergency landing stages to ensure flight safety. BRIEF DESCRIPTION OF THE DRAWINGS

[0115] The drawings constituting a part of the present application are used to provide a further understanding of the present application. The illustrative embodiments and descriptions of the present application are used to explain the present application and do not constitute an improper limitation on the present application. In the drawings:

[0116] Figure 1 It is a flow chart of a method for evaluating the benefits of an aviation hybrid electric propulsion system according to a preferred embodiment of the present application;

[0117] Figure 2 This is a schematic diagram of the relationship between the maximum range and the technical parameters of the energy storage system under different power schemes for 50 fixed-wing aircraft;

[0118] Figure 3 This is a schematic diagram of the relationship between unit purchase cost and technical parameters of energy storage systems under different power solutions for 50 fixed-wing aircraft;

[0119] Figure 4 This is a schematic diagram of the relationship between unit operating cost and technical parameters of energy storage systems under different power schemes for 50 fixed-wing aircraft;

[0120] Figure 5 This is a schematic diagram of the relationship between the unit comprehensive cost and the technical parameters of the energy storage system under different power schemes for 50 fixed-wing aircraft;

[0121] Figure 6 This is a schematic diagram of the relationship between the fuel consumption per kilometer of unit commercial load and the technical parameters of the energy storage system under different power schemes for 50 fixed-wing aircraft;

[0122] Figure 7This is a schematic diagram of the changing relationship between the comprehensive benefit coefficient and the technical parameters of the energy storage system under different power schemes for 50 fixed-wing aircraft. DETAILED DESCRIPTION

[0123] 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.

[0124] like Figure 1 As shown, a preferred embodiment of the present application provides a method for evaluating the benefits of an aviation hybrid electric propulsion system, comprising the following steps:

[0125] S1. Obtain the inputs of benefit evaluation from the aircraft manufacturer, including at least the effective load, commercial load (including crew members), typical flight profile power spectra of different power schemes, and the maximum power demand of the aircraft for high temperature / high altitude flight, wherein the sum of the weights of the propulsion system, energy storage system, and fuel system can be obtained by subtracting the commercial load from the effective load. The typical flight profile power spectrum includes the flight altitude during takeoff, climb, cruise, landing, and landing go-around, the power demand of the aircraft for the propulsion system, working time, and flight speed. The power spectrum example is shown in Table 1, and the maximum power demand of the aircraft for high temperature flight at sea level / high altitude is shown in Table 2;

[0126] Table 1: Operating parameters of a 50-seat fixed-wing aircraft with conventional power at different stages (cruising altitude 6000m)

[0127]

[0128] Table 2: Maximum power requirements for aircraft flying at sea level or high altitude

[0129]

[0130] S2. Component coefficient tables for hybrid electric, pure electric, and traditional propulsion systems are established to characterize the applicability or quantity of electrical components (see Table 3). A unified weight summation formula can be established for propulsion systems of different schemes, thereby simplifying the calculation process and improving the efficiency of the calculation model. Among them, the generator coefficient d1 and the drive motor coefficient d2 represent whether the propulsion system includes a generator and a drive motor, respectively; the distribution system coefficient d3 represents the number of distribution switches; and the engine coefficient d4 represents the number of engines.

[0131] Table 3: Component coefficients of hybrid electric, pure electric and traditional propulsion systems

[0132]

[0133] S3. Calculate the weight of the gas turbine engine, the weight of the generator, the weight of the generator controller, the weight of the drive motor controller, the weight of the drive motor, the weight of the power distribution system, the total weight of the cooling system, the total weight of the propulsion system, the total weight of the battery system and battery system parameters, and the fuel quality based on the input of the benefit evaluation and the component coefficient table;

[0134] S4. Establish an overall performance simulation model of a turboshaft engine for a hybrid electric propulsion system and an overall performance simulation model of a gas turbine engine for a conventional power scheme, calculate the engine fuel consumption rate at each stage according to the power specified by the typical flight profile power spectrum of different propulsion systems, and establish an iterative calculation formula for the cruise stage distance of the fuel part and an iterative calculation formula for the cruise stage distance of the battery part in parallel, thereby obtaining an iterative calculation formula for the cruise stage distance applicable to different propulsion systems, and determine the cruise stage distance through iterative calculation;

[0135] S5. Calculate the comprehensive total cost per kilometer of unit commercial load, the maximum power capability index under high temperature conditions at sea level, the maximum power capability index under high temperature conditions at plateau, the fuel consumption per kilometer of unit commercial load and the noise index of the propulsion system according to the maximum range, number of flights and relevant prices of the propulsion system to be evaluated based on the cruise stage distance;

[0136] S6. Taking the maximum range of the traditional propulsion system, the comprehensive total cost per kilometer of unit commercial load, the maximum power capability index under high temperature conditions at sea level, the maximum power capability index under high temperature conditions at plateau, the fuel consumption per kilometer of unit commercial load and the noise index of the propulsion system as the benchmark, the relative proportion of the corresponding indicators of the propulsion system is calculated in combination with the maximum range of the propulsion system to be evaluated, the comprehensive total cost per kilometer of unit commercial load, the maximum power capability index under high temperature conditions at sea level, the maximum power capability index under high temperature conditions at plateau, the fuel consumption per kilometer of unit commercial load and the noise index of the propulsion system, and the weight coefficient is configured to calculate the comprehensive benefit coefficient of the propulsion system to be evaluated.

[0137] This embodiment provides a method for evaluating the benefits of an aviation hybrid electric propulsion system. By using the benefit evaluation method of the present application, according to the aircraft power demand, from the perspective of engine design, the comprehensive benefit coefficient of the propulsion system of the hybrid electric propulsion system scheme relative to the gas turbine engine propulsion system scheme is obtained, including the maximum power and power margin, and the maximum power and power margin benefits, fuel consumption benefits, weight benefits, cost benefits, etc. are obtained, so as to obtain more accurate information. The accuracy of the benefit evaluation result is high, which supports the aircraft developer to make a scientific and reasonable choice between the three propulsion systems of hybrid electric, pure electric and traditional, thereby providing a more reliable and important basis for the scheme design and power selection of the flight platform, ensuring that the power selection of the hybrid electric propulsion system scheme achieves a higher power output, and a power system with a more advanced technical level can be obtained, which can effectively save the development cost brought by the re-development or serial development of the engine; ensuring that the selected hybrid electric propulsion system scheme has both high thermal efficiency and low fuel consumption, and at the same time has the power requirements that meet different intake conditions, cruising, and emergency landing stages to ensure flight safety.

[0138] Preferably, the present embodiment adopts a series hybrid electric propulsion system, and utilizes a battery system to provide partial power during the high-power stage of takeoff, so as to reduce the power demand of the turboshaft engine dedicated to the hybrid electric propulsion system, thereby reducing the number of engines or reducing the weight. Since the demand for engine power during the cruise stage is relatively small, cruise flight can still be achieved after the number of engines is reduced or the weight is reduced. For a turboshaft engine, the design point generally selects a high-power state, so the higher the engine power, the higher the thermal efficiency; for the same low-power level, small turboshaft engines benefit from targeted design, and the thermal efficiency will also be relatively improved. Therefore, the series hybrid electric propulsion system of the present invention can effectively improve the thermal efficiency during the cruise stage.

[0139] When calculating the weight of a gas turbine engine, a power mix ratio is established to characterize the proportion of power provided by the energy storage system in the total power demand. When the aircraft is climbing or cruising, the power capacity provided by the engine will decrease. By setting the engine power altitude attenuation coefficient, the power demand equivalent to the ground during the climb and cruise stages can be obtained. The nominal power of the engine can be obtained by selecting the maximum equivalent power during takeoff, climb, and cruise stages, considering the efficiency along the way from the gas turbine engine to the propeller input, as well as the atmospheric temperature changes and the power margin statistics required for safety. Finally, the weight of the gas turbine engine can be obtained based on the engine power-to-weight ratio. Among them, when calculating the power loss along the way, the component coefficients d1, d2, and d3 are used to establish a weight calculation formula for a gas turbine engine that is applicable to both traditional propulsion systems and hybrid electric propulsion systems. The specific details are as follows:

[0140] In step S3: the weight of the gas turbine engine is calculated by the following formula:

[0141] m engine = max((1- H pto ) to ,(1- H pcl ) cl / α,P cr / α) / (η al d1 ×η alc d1 ×η sw d3- d1 ×η moc d2 ×η mo d2 ) / β / ρ en

[0142] Where: H pto , H pcl are the power mixing ratios in the take-off and climb phases, respectively. The traditional propulsion system takes a value of 0, the pure electric propulsion system takes a value of 1, and the hybrid electric propulsion system takes a value of 0 to 1; P to Pcl is the power requirement of the aircraft for the propulsion system during the take-off phase, in kW; Pcl is the power requirement of the aircraft for the propulsion system during the climb phase, in kW; cr is the power demand of the aircraft for the propulsion system during the cruise phase, in kW; α is the engine power altitude attenuation coefficient, which takes into account the characteristic that the available power of the engine decreases when the altitude increases, and is taken as 0.95~0.9 / km; η al is the generator efficiency; η alc is the generator controller efficiency; η sw is the efficiency of the power distribution switch; η moc is the efficiency of the drive motor controller; η mo is the efficiency of the drive motor; β is the engine power margin coefficient, which takes into account the atmospheric temperature change and the margin required for safety, and is set to 0.7-0.8; ρ en is the engine power density, which depends on the level of technological development. Considering the influence of the turboprop engine reducer, the turboshaft engine dedicated to the hybrid electric propulsion system is 1.29 to 1.35 times that of the turboprop engine. d1 is the generator coefficient, d2 is the drive motor coefficient, and d3 is the power distribution system coefficient.

[0143] The weight of the generator is mainly related to the power level of the generator. The maximum power during takeoff, climb, and cruise phases is selected, and the efficiency along the generator to the input end of the propeller is considered to obtain the nominal power of the generator. Finally, the weight of the generator can be obtained according to the power-to-weight ratio of the generator. Among them, the component coefficient d1 is used to make the calculated value of the generator weight of the pure electric and traditional propulsion system 0; d3-d1 is used to calculate the number of distribution switches along the way, and then consider its impact on efficiency loss. Therefore, the weight of the generator is calculated by the following formula:

[0144] m al = d1×max((1- H pto ) to , (1- H pcl ) cl , P cr ) / (η al ×η alc ×η sw d3- d1 ×η moc ×η mo ) / ρ al

[0145] Where: al is the generator power density;

[0146] The weight of the generator controller is calculated by the following formula:

[0147] m alc = d1×max((1- H pto ) to , (1- H pcl ) cl , P cr ) / (η alc ×η sw d3- d1 ×η moc ×η mo ) / ρ alc

[0148] Where: alc is the generator controller power density;

[0149] The weight of the drive motor controller is calculated by the following formula:

[0150] m moc = d2×max(P to ,P cl ,P cr ) / (η moc ×η mo ) / ρ moc

[0151] Where: moc is the power density of the drive motor controller;

[0152] The weight of the drive motor is calculated by the following formula:

[0153] m mo = d2×max(P to ,P cl ,P cr ) / η mo / ρ mo

[0154] Where: mo is the power density of the drive motor;

[0155] The weight of the power distribution system is calculated using the following formula:

[0156] m sw = d3×m sw0

[0157] Where: m sw0 The weight is for a single distribution switch.

[0158] This embodiment calculates the weight of the gas turbine engine, the weight of the generator, the weight of the generator controller, the weight of the drive motor controller, the weight of the drive motor, the weight of the power distribution system, the total weight of the cooling system, the total weight of the propulsion system, the total weight of the battery system and battery system parameters, and the fuel quality based on the input of the benefit evaluation and the component coefficient table. The purpose and benefit of this embodiment is to establish a calculation model that is applicable to hybrid electric, pure electric and traditional propulsion systems at the same time, avoiding repeated modeling processes and improving the efficiency of the calculation process.

[0159] Preferably, the total weight of the cooling system is calculated by the following steps:

[0160] The weight of the cooling system is mainly related to the power loss along the way. By selecting the maximum power during takeoff, climb, and cruise, and considering the efficiency along the way from the gas turbine engine to the input end of the drive motor, the heat dissipation power requirement can be obtained. Finally, the weight of the cooling system can be obtained based on the power-to-weight ratio of the cooling system. Among them, the component coefficient d1 is used to make the calculation result of this part of the cooling system of the pure electric and traditional propulsion system take a value of 0. Therefore, the weight of the cooling system of the gas turbine engine and the generator is calculated as follows:

[0161] m col,al =d1×max((1- H pto ) to , (1- H pcl ) cl , P cr )×(1-ηal ×η sw ) / ρ co

[0162] Where: co is the power-to-weight ratio of the cooling system;

[0163] Calculate the weight of the drive motor cooling system:

[0164] m col,mo = d2×max(P to ,P cl ,P cr ) / (η mo ×η moc ×η sw ) ×(1-η mo ×η moc ×η sw ) / ρ co ;

[0165] Using the component coefficient d2, the calculation result of the cooling system of the traditional propulsion system is set to 0; using (1-d1), it is indicated that the calculation of the heat dissipation of the energy storage system of the hybrid electric propulsion system does not consider the cooling requirements of the climbing and cruising stages; using d1 to represent the number of distribution switches to be considered, therefore, the weight of the cooling system at the battery system end is calculated as follows:

[0166] m col,bas =d2×max(H pto ×P to , H pcl ×P cl , (1-d1)×P cr )×P to ×(1-η sw d1 ) / ρ co ;

[0167] Calculate the total cooling system weight:

[0168] .

[0169] Compared with traditional propulsion systems, the total weight of the cooling system of the hybrid electric propulsion system includes the weight of the cooling system of the gas turbine engine and generator, the weight of the cooling system of the drive motor end, and the weight of the cooling system of the battery system end. Therefore, the total weight of the cooling system of the hybrid electric propulsion system is required and obtained.

[0170] Preferably, the total weight of the propulsion system is calculated by the following formula:

[0171] .

[0172] Preferably, the total weight of the battery system and the battery system parameters are calculated by the following steps:

[0173] Due to the high power demand for takeoff power, the battery is required to have a short-term high-rate discharge, which puts forward requirements for its power density. The method for calculating power density is the product of the energy density of the energy storage system and the discharge rate, and then divided by 1h. Therefore, the weight of the energy storage system calculated according to the power demand is:

[0174] m bas,P= d2×max(H pto ×P to , H pcl ×P cl , (1- d1) × P cr ) / ( E bas ×C)

[0175] Where: E bas is the energy density of the battery system; C is the discharge rate, which represents the discharge capacity of the battery system. The hybrid electric propulsion system selects a battery system with a discharge rate of 9 to 11;

[0176] For pure electric propulsion systems, the weight of the energy storage system can be directly calculated based on the weight of the payload, commercial load and propulsion system; for hybrid electric propulsion systems, the weight of the energy storage system is determined based on the energy demand during takeoff. Using the component coefficients d1, d2 and (1-d1), the calculation formula is applicable to different propulsion systems. Therefore, the specific weight of the energy storage system calculated according to the energy demand is:

[0177] m bas,E= d2×(1- d1) ×(m load -m pay -m pop )+ d1(H pto ×P to ×t+H pcl ×P cl ×t cl ) / E bas ×γ

[0178] Where: m load is the aircraft payload, m pay is the commercial load of the aircraft, γ is the battery residual power coefficient, which is 1.1 to 1.2, and t is the total time of the take-off, landing and go-around phase;

[0179] Calculate the total weight of the battery system and determine the battery system parameters:

[0180]

[0181] Different groups of battery system parameters are selected, including battery system energy density and discharge rate, and the weights of the energy storage systems calculated according to different requirements are compared to determine the battery system parameters with the smallest total battery system weight.

[0182] Specifically, the fuel quality is calculated by the following formula:

[0183] .

[0184] Preferably, the step S4 specifically comprises the steps of:

[0185] The overall performance simulation model of the turboshaft engine for the hybrid electric propulsion system and the overall performance simulation model of the gas turbine engine for the traditional power scheme are established respectively, and the engine fuel consumption rate at each stage is calculated according to the power specified by the typical flight profile power spectrum of different propulsion systems;

[0186] Calculate the energy conversion efficiency from the engine to the aircraft thruster input:

[0187]

[0188] Calculate the energy conversion efficiency from the energy storage system to the input of the aircraft thruster:

[0189]

[0190] According to the weight of the payload, commercial load and propulsion system, the sum of the weight of the battery system and the energy storage system can be obtained. The fuel mass can be obtained by using the fuel consumption of each stage and considering the residual fuel coefficient. Among them, the fuel consumption in the take-off stage needs to consider the influence of the mixing ratio, and the fuel consumption in the cruise stage is related to the given range. The cruise stage range needs to be determined through iteration. Using the component coefficient d4, the calculation result of the pure electric propulsion system is made 0. Therefore, the iterative calculation formula for the cruise stage distance of the fuel part is established as follows:

[0191] m load -m pay -m pop -m bas =(1- H pto ) × P to / η g2p ×SFC to ×t+((1- H pcl ) × P cl / η g2p ×SFC cl ×t cl +P cr / η g2p ×SFC cr ×(L cr1 / V cr)+d4∑(P i / η g2p ×SFC i ×t i ))×δ

[0192] Where: SFC to is the fuel consumption rate of the engine during takeoff; t is the working time during takeoff, landing and go-around; SFC cl is the fuel consumption rate of the engine during the climbing phase, P cl is the power in the climbing stage; t cl P is the working time of the climbing stage; cr is the cruise power; SFC cr is the fuel consumption rate of the engine during cruising stage; L cr is the cruise stage distance, which needs to be determined by iterative calculation; V cr is the flight speed in the cruising phase; P i 、SFC i ,t i are the power, engine fuel consumption rate and working time of the remaining stages respectively; δ is the residual oil coefficient, ranging from 1.06 to 1.1, and d4 is the engine coefficient;

[0193] The quality of the battery system can be obtained by using the power consumption of each stage, combining the battery energy density, and considering the battery residual power coefficient. Among them, the power consumption in the take-off stage needs to consider the influence of the mixing ratio, the power consumption in the cruise stage is related to the given range, and the cruise stage range needs to be determined through iteration. The applicability of different propulsion systems is characterized by d2 and (1-d1) in the remaining stages. Therefore, the iterative calculation formula for the cruise stage distance of the battery part is established as follows: d2×(1-d1) ×m bas =(d2×(1-d1) ×H pto ×P to ×t+ d2×(1-d1) ×H pcl ×P cl ×t cl +P cr ×(L cr2 / V cr )+d2×(1-d1) ×∑(P i / t i ) ) / η b2p / E bas ×γ

[0194] The iterative calculation formula for the cruise phase distance applicable to different propulsion systems is obtained, and the cruise phase distance is determined through iterative calculation.

[0195] L cr =max(L cr1 , L cr2 )

[0196] Where: L cr is the final calculated cruise stage distance.

[0197] After obtaining the iterative calculation formula for the cruise phase distance applicable to different propulsion systems, this embodiment determines the cruise phase distance through iterative calculation. Its purposes and benefits include: the established iterative formula can be simultaneously applicable to the range calculation of three propulsion systems: hybrid electric, pure electric and traditional, avoiding repeated modeling processes and improving the efficiency of the calculation process.

[0198] Preferably, the step S5 specifically comprises the steps of:

[0199] Establish the calculation formula of the maximum range L:

[0200] L=V cl ×t cl + L cr + V dec ×t dec

[0201] Where: V cl is the flight speed during the climb phase, V dec is the flight speed during the descent phase, t dec Working hours for the descent phase;

[0202] The price of the battery system is calculated based on the battery capacity, while the battery life is related to the number of available discharges. Based on the battery mass and energy density, the battery capacity and price can be obtained. Therefore, the purchase cost per kilometer of the battery system per unit commercial load is calculated based on the maximum range and the number of flights:

[0203] W bas = E bas ×m bas ×U bas / (F×L×m pay )

[0204] Where: U bas is the price corresponding to the unit energy capacity of the battery system; F is the number of available discharges of the battery system, corresponding to the number of flights during the life of the battery system;

[0205] The price of aircraft engines is related to nominal power. During the battery life (i.e., the number of flights), the number of times the engine is replaced must take into account both the time life and the cycle life, whichever comes first. Therefore, the specific purchase cost per kilometer of commercial load of aircraft engines calculated based on the time life and cycle life is:

[0206] W engine= MAX(F×t f / t engine , F / F engine ) ×MAX((1-H pto ) to , P cl / α, P cr / α) ×U eigine / (F×L×m pay )

[0207] Where: t f is the time of a single flight; t engine The time allowed for the engine to be used; F engine U is the number of cycles allowed for the engine; eigine is the price corresponding to the unit power of the engine;

[0208] Calculate the purchase cost per kilometer of the aircraft engine generator, generator controller, drive motor controller, drive motor, and cooling system:

[0209] W mo =F×t f / t mo ×m mo ×ρ mo ×U mo / (F×L×m pay )

[0210] W al = F × t f / t al ×m al ×ρ al ×U al / (F×L×m pay )

[0211] W alc = F × t f / t alc ×m alc ×ρ alc ×U alc / (F×L×m pay )

[0212] W moc = F × t f / t moc ×m moc ×ρ moc ×U moc / (F×L×m pay )

[0213] W cold = F × t f / tcold ×m cold ×ρ cold ×U cold / (F×L×m pay )

[0214] The price of each component is related to the power level, and the number of replacements of each component is mainly considered by the life span, where: t mo U is the time allowed for the drive motor to be used; mo is the price corresponding to the unit power of the drive motor;

[0215] The formula for calculating the purchasing cost per kilometer of unit commercial load for establishing a power distribution system is:

[0216] W sw =d3×U sw / (F×L×m pay )

[0217] Where: U sw is the unit price of the distribution switch;

[0218] Calculate the total purchase cost per kilometer of commercial load:

[0219] ;

[0220] Calculate the total operating cost per kilometer of commercial load:

[0221] W c= (m bas ×E bas / γ×U elc + m fule ×U fule / δ) / (L×m pay )

[0222] Where: U elc , U fule They are the unit mass cost of fuel and the unit cost of electricity;

[0223] The comprehensive total cost per kilometer of commercial load is calculated based on the total purchase cost per kilometer of commercial load and the total operating cost per kilometer of commercial load:

[0224] ;

[0225] Calculate the maximum power capability index at sea level under high temperature conditions:

[0226]

[0227] Where: P pop,hot P is the maximum power capability of the engine at sea level and high temperature; dem,hotis the maximum power requirement of the aircraft under high temperature conditions at sea level;

[0228] Calculate the maximum power capability index under high temperature conditions in plateau:

[0229]

[0230] Where: P pop,pla P is the maximum power capability of the engine under high temperature conditions at plateau; dem,pla The maximum power requirement of the aircraft under high temperature conditions on the plateau;

[0231] The maximum power capability index is used to compare the maximum power and power margin of the three propulsion systems, hybrid electric, pure electric, and traditional, at different altitudes and temperatures to obtain the comprehensive power margin. Due to the problem of insufficient power margin of the propulsion system, it mainly occurs in flights under high temperature conditions at sea level and high temperature at plateau. Therefore, the evaluation of power margin mainly considers these two conditions;

[0232] Calculate the aircraft's fuel consumption per kilometer of commercial load:

[0233] M c= (m fule / δ) / (L×m pay )

[0234] The takeoff and climb phases are closest to the ground. Therefore, the propulsion system noise index is calculated based on the mixture ratio of the takeoff and cruise phases:

[0235] N c= H pto ×H pcl .

[0236] The purposes and benefits of this embodiment include:

[0237] The maximum range is used to characterize the range capability of an aircraft and is an important parameter for determining whether the aircraft can meet the requirements of application scenarios.

[0238] The number of flights is related to the allowable number of cycles of the battery and engine, which directly affects the life and cost of the propulsion system.

[0239] The comprehensive total cost per unit commercial load-kilometer represents the cost corresponding to transporting one ton of commercial load per kilometer, which directly affects the economic benefits of the aircraft.

[0240] The maximum power capability index under high temperature conditions at sea level represents the ability of the propulsion system to meet the power requirements of the aircraft under high temperature conditions at sea level, and directly affects the commercial load capacity of the aircraft under high temperature conditions at sea level.

[0241] The maximum power capability index under high-altitude and high-temperature conditions characterizes the ability of the propulsion system to meet the power requirements of the aircraft under high-altitude and high-temperature conditions, and directly affects the commercial carrying capacity of the aircraft under high-altitude and high-temperature conditions.

[0242] Fuel consumption per unit of commercial load-kilometer represents the fuel consumption corresponding to one ton of commercial load transported per kilometer, which directly affects the aircraft's range, emissions and economic benefits.

[0243] The propulsion system noise index represents the power noise level of the aircraft during operation, and directly affects whether the aircraft can operate for a long time in densely populated areas or quiet areas.

[0244] Preferably, the step S6 specifically comprises the steps of:

[0245] Taking the maximum range of the traditional propulsion system, the comprehensive total cost per kilometer of unit commercial load, the maximum power capability index under high temperature conditions at sea level, the maximum power capability index under high temperature conditions at plateau, the fuel consumption per kilometer of unit commercial load and the noise index of the propulsion system as the benchmark, the relative proportion of the corresponding indicators of the propulsion system is calculated in combination with the maximum range of the propulsion system to be evaluated, the comprehensive total cost per kilometer of unit commercial load, the maximum power capability index under high temperature conditions at sea level, the maximum power capability index under high temperature conditions at plateau, the fuel consumption per kilometer of unit commercial load and the noise index of the propulsion system to be evaluated, and the weight coefficient is configured to calculate the comprehensive benefit coefficient of the propulsion system to be evaluated:

[0246] Z = a1 × L / L tra + a2×W / W tra + a3×N p,hot / N p,hot,tra +a4×N p,pla / N p,pla,tra - a5×M c / M c,tra - a6×N c / N c,tra

[0247] Where: a1, a2, a3, a4, a5, a6 are weight coefficients, all of which are positive; L, W, N p,hot 、N p,pla 、M c 、N c They are the maximum range of the propulsion system to be evaluated, the comprehensive total cost per kilometer of unit commercial load, the maximum power capability index under high temperature conditions at sea level, the maximum power capability index under high temperature conditions at plateau, the fuel consumption per kilometer of unit commercial load and the noise index of the propulsion system; L tra , W tra 、N p,hot,tra 、N p,pla,tra 、Mc,tra 、N c,tra They are the maximum range of the traditional propulsion system, the comprehensive total cost per unit commercial load kilometer, the maximum power capability index under high temperature conditions at sea level, the maximum power capability index under high temperature conditions at plateau, the fuel consumption per unit commercial load kilometer and the propulsion system noise index.

[0248] This embodiment takes the maximum range of the traditional propulsion system, the comprehensive total cost per unit commercial load kilometer, the maximum power capability index under high temperature conditions at sea level, the maximum power capability index under high temperature conditions at plateau, the fuel consumption per unit commercial load kilometer and the propulsion system noise index as benchmarks, and calculates the relative proportions of the corresponding indicators of the propulsion system based on the maximum range of the propulsion system to be evaluated, the comprehensive total cost per unit commercial load kilometer, the maximum power capability index under high temperature conditions at sea level, the maximum power capability index under high temperature conditions at plateau, the fuel consumption per unit commercial load kilometer and the propulsion system noise index, and configures the weight coefficient to calculate the comprehensive benefit coefficient of the propulsion system to be evaluated, which has the following benefits and purposes: it can comprehensively evaluate the advantages and disadvantages of the three propulsion systems of hybrid electric, pure electric and traditional, thereby providing a more reliable and important basis for the design of the flight platform and the power selection.

[0249] Specifically, some weight coefficients satisfy the following conditions:

[0250] a1+a2+a3+a4=1;

[0251] The weight coefficients a5 and a6 are adjusted accordingly according to the requirements of the application scenario for fuel consumption or emissions and power noise. For example, for areas with higher environmental protection requirements, the value of a5 should be increased, such as 1; for areas with strict control of noise levels, the value of a6 should be increased.

[0252] In this embodiment, the purpose and benefit of setting the above-mentioned weight coefficients is to evaluate the benefits of the propulsion system according to different application scenarios, and then determine a suitable propulsion system solution.

[0253] In view of the future development trend of battery systems, the benefit evaluation method of this application was used to conduct a benefit evaluation of the hybrid electric power system of a 50-seat fixed-wing aircraft, and compared the maximum range and comprehensive cost with traditional power and pure electric power. The maximum range and comprehensive cost were compared with traditional power and pure electric power. When designing and evaluating the benefits of hybrid electric power solutions, the following factors are mainly considered:

[0254] 1) The 50-seat fixed-wing aircraft is based on the Cyber ​​2000 configuration;

[0255] 2) The flight mission mainly includes five stages: take-off, climb, cruise, landing and landing go-around;

[0256] 3) The hybrid electric power adopts a series configuration, that is, the engine generates electricity through the generator instead of directly driving the propeller, which supports the aircraft to realize the design of distributed propulsion layout, so that the aircraft can obtain better aerodynamic benefits and thus improve the lift-to-drag ratio;

[0257] 4) Hybrid electric power includes three configurations:

[0258] Hybrid electric-silent configuration, takeoff, climb phase (at least below 2 km), landing, and go-around all use energy storage (i.e. battery) system to provide power to achieve silent takeoff and landing;

[0259] Hybrid electric-semi-silent configuration, the power mixing ratio during takeoff and climb is 50%, that is, the energy storage system and vortex electric system each provide half of the power for the flight platform to reduce the noise level during takeoff and landing;

[0260] In the hybrid electric-long-range configuration, the energy storage system only provides partial power during takeoff and go-around, and fuel is used in the remaining stages to achieve a longer range;

[0261] 5) Taking into account factors such as improvements in fuselage materials and the integration of energy storage systems and fuselage designs, it is estimated that the weight of aircraft fuselage structures can be reduced by 10% by 2030;

[0262] 6) In addition to meeting the energy needs of the aircraft, the energy storage system should also meet the high-power discharge requirements during the take-off phase and the compensation requirements during power changes.

[0263] from Figure 2 It can be seen that when the energy storage system develops to the level of 700-1000 Wh / kg, the range of a pure electric aircraft is only 383-638 km, while the range of a hybrid electric system can reach 1733-2296 km.

[0264] like Figure 3 , 4 As shown in Figure 5, from the perspective of cost, due to the increase in energy density, the total purchase cost per commercial load kilometer, the total operating cost per commercial load kilometer, and the comprehensive cost per commercial load kilometer of pure electric vehicles will all decrease. However, even when the energy storage system develops to 1000 Wh / kg, the unit comprehensive cost of pure electric power is still about 37.5% higher than that of traditional power, and nearly 2.5 times that of hybrid electric power.

[0265] According to the maximum power demand of the aircraft during high temperature / high altitude flight, the maximum power capability index under high temperature conditions at sea level and the maximum power capability index under high temperature conditions at high altitude are calculated. As can be seen from Tables 4 and 5, the hybrid silent and pure electric propulsion systems are less affected by high temperature and high altitude due to the use of battery system energy during takeoff and climb phases, with the support of battery system temperature control technology, and the maximum power capability index under high temperature conditions at sea level and the maximum power capability index under high temperature conditions at high altitude are both large.

[0266] Table 4: Maximum power capability index under high temperature conditions at sea level

[0267]

[0268] Table 5: Maximum power capability index under high temperature conditions in plateau

[0269]

[0270] Based on the maximum range, fuel consumption and payload, calculate the fuel consumption per kilometer of payload for different propulsion systems (see Figure 6 ).

[0271] According to the mixing ratio in the take-off and climb phases, the noise index of different propulsion systems is obtained, as shown in Table 6.

[0272] Table 6: Noise index

[0273]

[0274] For the scenario of short-distance commuter flights in suburban areas, the weight coefficients are set as shown in Table 7.

[0275] Table 7:

[0276]

[0277] According to the ratio of the technical indicators of different propulsion systems to the traditional propulsion system and the set weight coefficient, the change trend of the comprehensive benefit coefficient of the propulsion system is obtained. Figure 7 It can be seen that hybrid silent and hybrid long-range are more suitable.

[0278] The parts not described in detail in this application are all known technologies to those skilled in the art. Finally, it should be noted that the above embodiments are only used to illustrate the technical solutions of this application and are not limiting. Although this application is described in detail with reference to the preferred embodiments, those skilled in the art should understand that the technical solutions of this application can be modified or replaced by equivalents without departing from the purpose and scope of the technical solutions of this application, which should be included in the scope of the claims of this application.

Claims

1. A method for evaluating the benefits of an aviation hybrid electric propulsion system, characterized in that: The following steps are involved: S1. Obtaining inputs for benefit evaluation from the aircraft manufacturer, including at least effective load, commercial load, typical flight profile power spectra of different power schemes, and maximum power requirements of the aircraft for high temperature / high altitude flight, wherein the sum of the weights of the propulsion system, energy storage system, and fuel system can be obtained by subtracting the commercial load from the effective load, and the typical flight profile power spectrum includes the flight altitudes during takeoff, climb, cruise, landing, and landing go-around, the aircraft's power requirements for the propulsion system, working time, and flight speed; S2. Establish component coefficient tables for hybrid electric, pure electric, and traditional propulsion systems to characterize the applicability or quantity of electrical components; S3. Calculate the weight of the gas turbine engine, the weight of the generator, the weight of the generator controller, the weight of the drive motor controller, the weight of the drive motor, the weight of the power distribution system, the total weight of the cooling system, the total weight of the propulsion system, the total weight of the battery system and battery system parameters, and the fuel quality based on the input of the benefit evaluation and the component coefficient table; S4. Establish an overall performance simulation model of a turboshaft engine for a hybrid electric propulsion system and an overall performance simulation model of a gas turbine engine for a conventional power scheme, calculate the engine fuel consumption rate at each stage according to the power specified by the typical flight profile power spectrum of different propulsion systems, and establish an iterative calculation formula for the cruise stage distance of the fuel part and an iterative calculation formula for the cruise stage distance of the battery part in parallel, thereby obtaining an iterative calculation formula for the cruise stage distance applicable to different propulsion systems, and determine the cruise stage distance through iterative calculation; S5. Calculate the comprehensive total cost per kilometer of unit commercial load, the maximum power capability index under high temperature conditions at sea level, the maximum power capability index under high temperature conditions at plateau, the fuel consumption per kilometer of unit commercial load and the noise index of the propulsion system according to the maximum range, number of flights and relevant prices of the propulsion system to be evaluated based on the cruise stage distance; S6. Taking the maximum range of the traditional propulsion system, the comprehensive total cost per kilometer of unit commercial load, the maximum power capability index under high temperature conditions at sea level, the maximum power capability index under high temperature conditions at plateau, the fuel consumption per kilometer of unit commercial load and the noise index of the propulsion system as the benchmark, the relative proportion of the corresponding indicators of the propulsion system is calculated in combination with the maximum range of the propulsion system to be evaluated, the comprehensive total cost per kilometer of unit commercial load, the maximum power capability index under high temperature conditions at sea level, the maximum power capability index under high temperature conditions at plateau, the fuel consumption per kilometer of unit commercial load and the noise index of the propulsion system, and the weight coefficient is configured to calculate the comprehensive benefit coefficient of the propulsion system to be evaluated.

2. The method for evaluating the benefits of an aviation hybrid electric propulsion system according to claim 1, characterized in that: In step S3: The weight of a gas turbine engine is calculated as follows: m engine = max((1- H pto ) P to ,(1- H pcl ) P cl / a,P cr / a) / (h al d1 ×h alc d1 ×h sw d3- d1 ×h moc d2 ×h mo d2 ) / b / r en Where: H pto , H pcl are the power mixing ratios in the take-off and climb phases, respectively. The traditional propulsion system takes a value of 0, the pure electric propulsion system takes a value of 1, and the hybrid electric propulsion system takes a value of 0 to 1; P to Pcl is the power requirement of the aircraft for the propulsion system during the take-off phase, in kW; Pcl is the power requirement of the aircraft for the propulsion system during the climb phase, in kW; cr is the power demand of the aircraft for the propulsion system during the cruise phase, in kW; α is the engine power altitude attenuation coefficient, with a value of 0.95~0.9 / km; η al is the generator efficiency; η alc is the generator controller efficiency; η sw for the efficiency of power distribution switches; η moc To drive the motor controller efficiency; η mo is the efficiency of the driving motor; β is the engine power margin coefficient, which is 0.7 to 0.8; ρ en is the engine power density. The turboshaft engine dedicated to the hybrid electric propulsion system is 1.29 to 1.35 times that of the turboprop engine. d1 is the generator coefficient, d2 is the drive motor coefficient, and d3 is the power distribution system coefficient. The weight of the generator is calculated by the following formula: m al = d1×max((1- H pto ) P to ,(1- H pcl ) P cl ,P cr ) / (the al ×h alc ×h sw d3- d1 ×h moc ×h mo ) / p al Where: al is the generator power density; The weight of the generator controller is calculated by the following formula: m alc = d1×max((1- H pto ) P to ,(1- H pcl ) P cl ,P cr ) / (the alc ×h sw d3- d1 ×h moc ×h mo ) / p alc Where: alc is the generator controller power density; The weight of the drive motor controller is calculated by the following formula: m moc = d2×max(P to ,P cl ,P cr ) / (the moc ×h mo ) / p moc Where: moc is the power density of the drive motor controller; The weight of the drive motor is calculated by the following formula: m mo = d2×max(P to ,P cl ,P cr ) / h mo / r mo Where: mo is the power density of the drive motor; The weight of the power distribution system is calculated using the following formula: m sw = d3×m sw0 Where: m sw0 The weight is for a single distribution switch.

3. The method for evaluating the benefits of an aviation hybrid electric propulsion system according to claim 2, characterized in that: The total weight of the cooling system is calculated as follows: Calculate the weight of the cooling system for the gas turbine engine and generator section: m col,al =d1×max((1- H pto ) P to ,(1- H pcl ) P cl ,P cr ) ×(1-n al ×h sw ) / p co Where: co is the power-to-weight ratio of the cooling system; Calculate the weight of the drive motor cooling system: m col,mo = d2×max(P to ,P cl ,P cr ) / (the mo ×h moc ×h sw ) ×(1-n mo ×h moc ×h sw ) / p co Calculate the weight of the cooling system on the battery system side: m col,bas =d2×max(H pto ×P to ,H pcl ×P cl ,(1-d1)×P cr )×P to ×(1-η sw d1 ) / ρ co Calculate the total cooling system weight: 。 4. The method for evaluating the benefits of an aviation hybrid electric propulsion system according to claim 3, characterized in that: The total weight of the propulsion system is calculated as follows: 。 5. The method for evaluating the benefits of an aviation hybrid electric propulsion system according to claim 4, characterized in that: The total weight of the battery system and the battery system parameters are calculated as follows: Calculate the weight of the energy storage system according to power requirements: m bas,P= d2×max(H pto ×P to ,H pcl ×P cl ,(1- d1) ×P cr ) / ( E bas ×C) Where: E bas is the energy density of the battery system; C is the discharge rate, which represents the discharge capacity of the battery system; Calculate the weight of the energy storage system according to energy demand: m bas,E= d2×(1- d1) ×(m load -m pay -m pop )+ d1(H pto ×P to ×t+H pcl ×P cl ×t cl ) / E bas ×γ Where: m load is the aircraft payload, m pay is the commercial load of the aircraft, γ is the battery residual power coefficient, which is 1.1 to 1.2, and t is the total time of the take-off, landing and go-around phase; Calculate the total battery system weight: ； Different groups of battery system parameters are selected, including battery system energy density and discharge rate, and the total weight of the battery system corresponding to different battery system parameters is compared to determine the battery system parameters with the smallest total weight of the battery system.

6. The method for evaluating the benefits of an aviation hybrid electric propulsion system according to claim 5, characterized in that: The fuel quality is calculated by the following formula: 。 7. The method for evaluating the benefits of an aviation hybrid electric propulsion system according to claim 6, characterized in that: The step S4 specifically comprises the following steps: The overall performance simulation model of the turboshaft engine for the hybrid electric propulsion system and the overall performance simulation model of the gas turbine engine for the traditional power scheme are established respectively, and the engine fuel consumption rate at each stage is calculated according to the power specified by the typical flight profile power spectrum of different propulsion systems; Calculate the energy conversion efficiency from the engine to the aircraft thruster input: ； Calculate the energy conversion efficiency from the energy storage system to the input of the aircraft thruster: ； Establish the iterative calculation formula for the cruise phase distance of the fuel part: m load -m pay -m pop -m bas =(1- H pto ) ×P to / or g2p ×SFC to ×t+((1- H pcl ) ×P cl / or g2p ×SFC cl ×t cl +P cr / or g2p ×SFC cr ×(L cr1 / V cr )+d4∑(P i / or g2p ×SFC i ×t i ))×d Where: SFC to is the fuel consumption rate of the engine during takeoff; t is the working time during takeoff, landing and go-around; SFC cl is the fuel consumption rate of the engine during the climbing phase, P cl is the power in the climbing stage; t cl P is the working time of the climbing stage; cr is the cruise power; SFC cr is the fuel consumption rate of the engine during cruising stage; L cr1 The distance of the cruise phase of the fuel part needs to be determined by iterative calculation; V cr is the flight speed in the cruising phase; P i 、SFC i ,t i are the power, engine fuel consumption and working time in the remaining stages respectively; δ is the residual oil coefficient, ranging from 1.06 to 1.1, and d4 is the engine coefficient; Establish the iterative calculation formula for the cruise phase distance of the battery part: d2×(1-d1)×m bas =(d2×(1-d1)×H pto ×P to ×t+ d2×(1-d1) ×H pcl ×P cl ×t cl +P cr ×(L cr2 / V cr )+d2×(1-d1) ×∑(P i / t i ) ) / η b2p / E bas ×γ Where, L cr2 For the cruise phase distance of the battery part, the iterative calculation formula for the cruise phase distance applicable to different propulsion systems is obtained. The cruise phase distance is determined by iterative calculation: L cr =max(L cr1 ,L cr2 ) Where: L cr is the final calculated cruise stage distance.

8. The method for evaluating the benefits of an aviation hybrid electric propulsion system according to claim 7, characterized in that: The step S5 specifically comprises the following steps: Establish the calculation formula of the maximum range L: L=V cl ×t cl + L cr + V dec ×t dec Where: V cl is the flight speed during the climb phase, V dec is the flight speed during the descent phase, t dec Working hours for the descent phase; Calculate the purchase cost per kilometer of the battery system per commercial load by combining the maximum range and number of flights: W bas = E bas ×m bas ×U bas / (F×L×m pay ) Where: U bas is the price corresponding to the unit energy capacity of the battery system; F is the number of available discharges of the battery system, corresponding to the number of flights during the life of the battery system; The purchase cost per kilometer of the unit commercial load of the aircraft engine is calculated based on the time life and cycle life: W engine= MAX(F×t f / t engine ,F / F engine ) ×MAX((1-H pto ) P to , P cl / α, P cr / α) ×U eigine / ( F×L×m pay ) Where: t f is the time of a single flight; t engine The time allowed for the engine to be used; F engine U is the number of cycles allowed for the engine; eigine is the price corresponding to the unit power of the engine; Calculate the purchase cost per kilometer of the aircraft engine generator, generator controller, drive motor controller, drive motor, and cooling system: W mo =F×t f / t mo ×m mo ×ρ mo ×U mo / (F×L×m pay ) W al = F×t f / t al ×m al ×ρ al ×U al / (F×L×m pay ) W alc = F×t f / t alc ×m alc ×ρ alc ×U alc / (F×L×m pay ) W moc = F×t f / t moc ×m moc ×ρ moc ×U moc / (F×L×m pay ) W cold = F×t f / t cold ×m cold ×ρ cold ×U cold / (F×L×m pay ) Where: t mo U is the time allowed for the drive motor to be used; mo is the price corresponding to the unit power of the drive motor; The formula for calculating the purchasing cost per kilometer of unit commercial load for establishing a power distribution system is: IN sw =d3×U sw / (F×L×m pay ) Where: U sw is the unit price of the distribution switch; Calculate the total purchase cost per kilometer of unit commercial load: ； Calculate the total operating cost per kilometer of commercial load: W c= (m bas ×E bas / γ×U elc + m fule ×U fule / δ) / ( L×m pay ) Where: U elc , U fule They are the unit mass cost of fuel and the unit cost of electricity; The comprehensive total cost per kilometer of commercial load is calculated based on the total purchase cost per kilometer of commercial load and the total operating cost per kilometer of commercial load: ； Calculate the maximum power capability index at sea level under high temperature conditions: ； Where: P pop,hot P is the maximum power capability of the engine at sea level and high temperature; dem,hot The maximum power requirement of the aircraft under high temperature conditions at sea level; Calculate the maximum power capability index under high temperature conditions in plateau: ； Where: P pop,pla P is the maximum power capability of the engine under high temperature conditions at plateau; dem,pla The maximum power requirement of the aircraft under high temperature conditions on the plateau; Calculate the aircraft's fuel consumption per kilometer of commercial load: M c= (m fule / δ) / ( L×m pay ) According to the mixture ratio during takeoff and cruise, the propulsion system noise index is calculated: N c= H pto ×H pcl 。 9. The method for evaluating the benefits of an aviation hybrid electric propulsion system according to claim 8, characterized in that: The step S6 specifically comprises the following steps: Taking the maximum range of the traditional propulsion system, the comprehensive total cost per kilometer of unit commercial load, the maximum power capability index under high temperature conditions at sea level, the maximum power capability index under high temperature conditions at plateau, the fuel consumption per kilometer of unit commercial load and the noise index of the propulsion system as the benchmark, the relative proportion of the corresponding indicators of the propulsion system is calculated in combination with the maximum range of the propulsion system to be evaluated, the comprehensive total cost per kilometer of unit commercial load, the maximum power capability index under high temperature conditions at sea level, the maximum power capability index under high temperature conditions at plateau, the fuel consumption per kilometer of unit commercial load and the noise index of the propulsion system to be evaluated, and the weight coefficient is configured to calculate the comprehensive benefit coefficient of the propulsion system to be evaluated: Z= a1×L / L tra + a2×W / W tra + a3×N p,hot / N p,hot,tra +a4×N p,pla / N p,pla,tra - a5×M c / M c,tra - a6×N c / N c,tra Where: a1, a2, a3, a4, a5, a6 are weight coefficients, all of which are positive; L, W, N p,hot 、N p,pla 、M c 、N c They are the maximum range of the propulsion system to be evaluated, the comprehensive total cost per kilometer of unit commercial load, the maximum power capability index under high temperature conditions at sea level, the maximum power capability index under high temperature conditions at plateau, the fuel consumption per kilometer of unit commercial load and the noise index of the propulsion system; L tra , W tra 、N p,hot,tra 、N p,pla,tra 、M c,tra 、N c,tra They are the maximum range of the traditional propulsion system, the comprehensive total cost per unit commercial load kilometer, the maximum power capability index under high temperature conditions at sea level, the maximum power capability index under high temperature conditions at plateau, the fuel consumption per unit commercial load kilometer and the propulsion system noise index.

10. The method for evaluating the benefits of an aviation hybrid electric propulsion system according to claim 9, characterized in that: Some weight coefficients meet the following conditions: a1+a2+a3+a4=1; The weight coefficients a5 and a6 are adjusted accordingly according to the requirements of the application scenario for fuel consumption or emissions and power noise.

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