An iterative optimization design method for a hybrid power system of a more electric aircraft

Abstract

The application provides an iterative optimization design method of a hybrid power system of a more electric aircraft. The method is based on linear regression, part splitting and other methods, and realizes iterative optimization design of the more electric aircraft. The weight of the motor, battery and other equipment of the more electric aircraft is obtained through theoretical calculation, multiple parameters are introduced, and the proportion of the total energy occupied by the electric power supply is finally given from the aspects of emission, equipment cost, environmental benefit and the like.

Description

Technical Field

[0001] This invention relates to the field of aircraft design, specifically to an iterative optimization design method for a multi-electric aircraft hybrid power system. Background Technology

[0002] More electric aircraft have gradually become a hot topic in the aviation industry, especially in the civil aviation sector, in recent years. In the design process of more electric aircraft, the total weight of the aircraft and the proportion of energy supplied by electricity in a single flight are particularly important, and are the foundation and key to determining the success of the design.

[0003] In the literature "Hu Yu. Design and Optimization of Hybrid Electric Power System for General Aircraft", a propeller aircraft represented by the AMT-200 was used as the research object to design a hybrid electric power system for general aircraft. However, due to the significant differences in performance and values ​​between propeller engines and turbofan engines, and in the current era where large passenger aircraft powered by turbofan engines account for the vast majority, this method obviously cannot provide much reference for the modification and design of multi-electric aircraft.

[0004] In the literature "Chen Hongwei. Design and Research of Energy Management System for Hybrid Unmanned Aerial Vehicle", the optimization design is carried out using the unmanned aerial vehicle as the object. In the optimization design process, the final solution is determined by selecting a series of existing electric motors, batteries and other structures, but the overall allocation ratio is not involved. Summary of the Invention

[0005] To address the limitations and inadequacies of existing design and evaluation schemes for multi-electric large passenger aircraft, and the significant weight estimation errors during multi-electric aircraft design, this invention proposes an iterative optimization design method for the hybrid power system of multi-electric aircraft. This method calculates the aircraft's total weight and approximate electrical power supply values ​​using feasible and effective approaches, providing preliminary basis and ideas for the retrofit design of multi-electric aircraft.

[0006] The technical solution of this invention is as follows:

[0007] The iterative optimization design method for a multi-electric aircraft hybrid power system includes the following steps:

[0008] Step 1: Obtain flight data of the aircraft to be modified; based on the flight data, obtain the aircraft's lift coefficient C. L With drag coefficient C D Scatter plot of relationships;

[0009] Step 2: Obtain the lift coefficient C of the aircraft during the cruise phase based on the aircraft's cruise speed. L,cr and with C L,cr ±ΔC LTo determine the range of values, find the corresponding scatter plot of the lift-drag coefficient relationship during the cruise phase in the scatter plot obtained in step 1;

[0010] Step 3: Use the scatter plot of the lift-drag coefficient relationship obtained in Step 2 to perform linear regression to obtain the drag curve of the aircraft during the cruise phase.

[0011] Step 4: Obtain the flight speed during the cruise phase of the aircraft's flight plan. Using the drag curve for the cruise phase obtained in Step 3, determine the total power P during the cruise phase. T And total energy Q T ;

[0012] Step 5: Set the initial value τ for the proportion of electrical energy in the total energy of the aircraft after the power modification, and obtain the electrical energy expression Q. T τ and the expression for electrical energy output power P T τ; and based on the electrical energy output power, obtain the expression m for the corresponding motor mass with respect to τ. total (τ), based on electrical energy, the expression for the battery mass with respect to τ is obtained m. b (τ); combined with the set additional connecting component mass m s We obtain the expression for the maximum takeoff mass of the modified aircraft in terms of τ:

[0013] m MTO '(τ)=m MTO +m b (τ)+m total (τ)+m s

[0014] Where m MTO The maximum takeoff mass of the aircraft before the electrical modification;

[0015] Step 6: Based on the maximum takeoff mass m of the modified aircraft obtained in Step 5... MTO '(τ), assuming the flight speed remains constant during the cruise phase in the flight plan, yields the lift coefficient C of the modified aircraft during the cruise phase. L ′(τ), and based on the drag curve obtained in step 3, the drag coefficient C of the modified aircraft during the cruise phase is obtained. D ′(τ); based on the drag coefficient C D ′(τ) gives the drag power of the aircraft during the cruise phase, and thus the cruise power P of the engine. S,cr ′(τ); based on the engine's cruising power P S,cr ′(τ) and engine fuel flow rate Q cr The fuel consumption rate C′(τ) of the modified engine was obtained.

[0016] Step 7: Establish the energy efficiency formula

[0017]

[0018] Where ω(τ) represents energy efficiency, k is the set fuel price, C is the engine fuel consumption rate before electrification, and Q... T ′(τ) represents the total energy during the cruise phase after the electrification modification, Q m (τ) represents the electrical energy during the cruise phase after the electrification modification, L lift To increase the mass of the aircraft, the energy consumption during takeoff is increased, ω solid The inherent cost of the electrification conversion is denoted by 'a', which is the set charging cost coefficient. Within the set range of energy ratio, the energy ratio τ corresponding to the maximum energy efficiency is obtained through iterative calculation.

[0019] Furthermore, in step 1, based on the flight data, using the formula...

[0020]

[0021] The lift coefficient C of the aircraft is obtained. L With drag coefficient C D The scatter plot of the relationship is given, where m is the aircraft mass, g is the weight acceleration, γ is the flight trajectory angle, ρ is the air density, V is the flight speed, S is the wing area, and E is the lift-to-drag ratio.

[0022] Furthermore, in step 2, based on the aircraft's cruising speed V... cr Through formula

[0023]

[0024] The lift coefficient C of the aircraft during the cruise phase is obtained. L,cr .

[0025] Furthermore, in step 2, ΔC L The value is 0.2.

[0026] Furthermore, in step 3, the resistance curve is represented as...

[0027] C D =C D,0 +kC L 2

[0028] With C L 2 C is the independent variable. D A scatter plot is created for the dependent variable, and linear regression is performed based on the scatter plot to obtain the induced drag factor k and the zero-lift drag coefficient C. D,0 This leads to the drag curve.

[0029] Furthermore, in step 4, based on the flight speed during the cruise phase in the aircraft's flight plan, and using the drag curve for the cruise phase obtained in step 3, the total power P for the cruise phase is determined. T And total energy Q T The process is as follows:

[0030] Based on the flight speed during the cruise phase of the aircraft's flight plan, according to the formula...

[0031]

[0032] The lift coefficient C of the aircraft during the cruise phase is obtained. L,cr Then, based on the drag curve during the aircraft's cruise phase, the corresponding drag coefficient C is obtained. D,cr According to the formula

[0033]

[0034] The total power P during the cruise phase is obtained. T In formula Q T =P T t, to obtain the total energy Q T , where t is the flight time.

[0035] Furthermore, in step 5, the expression for the motor mass m with respect to τ is obtained based on the electrical energy output power. total The process of (τ) is as follows:

[0036] The rated power, stator, rotor, and housing mass of multiple motor models were obtained. Regression analysis was performed with rated power as the independent variable and the weight of stator, rotor, and housing as the three dependent variables to obtain the relationship between rated power and the mass of stator, rotor, and housing, respectively.

[0037] The electrical energy output power P T Substituting τ into the expressions relating rated power to the mass of the stator, rotor, and housing, and summing them, we obtain the corresponding expression for the mass of the motor in terms of τ, m. total (τ).

[0038] Furthermore, in step 5, the expression for the battery mass with respect to τ is obtained based on electrical energy. b The process of (τ) is as follows: Based on the energy density ρ of the battery used, obtain m b (τ)=Q T τ / ρ.

[0039] Beneficial effects

[0040] The iterative optimization design method for multi-electric aircraft hybrid power systems proposed in this invention is based on linear regression, component breakdown, and other methods to achieve iterative optimization design of multi-electric aircraft. The weight of the multi-electric aircraft's electric motor, battery, and other equipment is calculated theoretically. Multiple parameters are introduced to comprehensively evaluate the system from multiple aspects, including emissions, equipment cost, and environmental benefits, ultimately determining the proportion of total energy supplied by electricity.

[0041] Additional aspects and advantages of the invention will be set forth in part in the description which follows, and in part will be obvious from the description, or may be learned by practice of the invention. Attached Figure Description

[0042] The above and / or additional aspects and advantages of the present invention will become apparent and readily understood from the description of the embodiments taken in conjunction with the following drawings, in which:

[0043] Figure 1 : A schematic diagram of the process of this invention. Detailed Implementation

[0044] The embodiments of the present invention are described in detail below. These embodiments are exemplary and intended to explain the present invention, and should not be construed as limiting the present invention.

[0045] like Figure 1 The flowchart shown is for the optimized design of a more-electric aircraft. This method first analyzes and calculates the drag curve during flight, and then optimizes the design based on actual data. Regression analysis is used to establish a regression model based on the ideal drag curve and actual flight data, yielding the actual drag curve during flight, which facilitates subsequent optimization calculations.

[0046] Regression analysis is a commonly used statistical method for analyzing data. Its purpose is to understand whether there is a correlation between two or more variables, and to estimate the conditional expectation of the dependent variable from the given independent variables through the establishment of a mathematical model. Regression with only one independent variable is called univariate regression, while regression with more than one independent variable is called multivariate regression. If the independent and dependent variables show a linear correlation, it is called linear regression; otherwise, it is called nonlinear regression. In practical engineering problems, the specific function should be set according to different situations. In the regression calculations involved in this method, because the resistance curve exhibits a significant nonlinear relationship, nonlinear regression and iterative calculations are used for the solution. In the process of estimating the quality of electrical equipment, a scatter plot is first created, and then the model to be used is determined based on a rough judgment.

[0047] The specific steps are as follows:

[0048] Step 1: Obtain flight data of the aircraft to be modified; based on the flight data, obtain the aircraft's lift coefficient C.L With drag coefficient C D Scatter plot of relationships.

[0049] Here, we can obtain the aircraft's speed, engine thrust, and a series of parameters during flight from the flight manual of a known aircraft model. Based on these parameters, we can use regression analysis to calculate the lift-drag curve of the aircraft near the cruise point, which will facilitate subsequent calculations.

[0050] Specifically, based on flight data, through formulas

[0051]

[0052] The lift coefficient C of the aircraft is obtained. L With drag coefficient C D The scatter plot of the relationship is given, where m is the aircraft mass, g is the weight acceleration, γ is the flight trajectory angle, ρ is the air density, V is the flight speed, S is the wing area, and E is the lift-to-drag ratio.

[0053] Step 2: Determining the lift-drag curve through linear regression over the entire flight envelope would introduce significant errors. Since energy management is most critical for civil aircraft during the cruise phase, more accurate determination of the lift and drag coefficients during the cruise phase is essential for providing more precise optimization solutions. Therefore, this invention primarily studies the aircraft's cruise state and calculates the lift-drag curve near the cruise point to avoid generating further errors.

[0054] Based on the aircraft's cruising speed V cr Through formula

[0055]

[0056] The lift coefficient C of the aircraft during the cruise phase is obtained. L,cr and with C L,cr The range is ±0.2. Find the corresponding scatter plot of the lift-drag coefficient relationship during the cruise phase in the scatter plot obtained in step 1.

[0057] Step 3: Using the scatter plot of the lift-drag coefficient relationship obtained in Step 2, perform linear regression to obtain the drag pole curve for the aircraft's cruise phase; represent the drag pole curve as follows:

[0058] C D =C D,0 +kC L 2

[0059] With C L 2 C is the independent variable. DA scatter plot is created for the dependent variable, and linear regression is performed based on the scatter plot to obtain the induced drag factor k and the zero-lift drag coefficient C. D,0 This leads to the drag curve.

[0060] In addition, based on the engine's cruising power P S,cr and engine fuel flow rate Q cr The engine fuel consumption rate C is obtained.

[0061] Step 4: Obtain the flight speed during the cruise phase of the aircraft's flight plan. Using the drag curve for the cruise phase obtained in Step 3, determine the total power P during the cruise phase. T And total energy Q T .

[0062] Based on the flight speed during the cruise phase of the aircraft's flight plan, according to the formula...

[0063]

[0064] The lift coefficient C of the aircraft during the cruise phase is obtained. L,cr Then, based on the drag curve during the aircraft's cruise phase, the corresponding drag coefficient C is obtained. D,cr According to the formula

[0065]

[0066] The total power P during the cruise phase is obtained. T In formula Q T =P T t, to obtain the total energy Q T , where t is the flight time.

[0067] Step 5: Set the initial value τ for the proportion of electrical energy in the total energy of the aircraft after the power modification, and obtain the electrical energy expression Q. T τ and the expression for electrical energy output power P T τ.

[0068] Based on the electrical energy output power, the expression for the mass of the motor with respect to τ is obtained. total (τ):

[0069] Unlike most studies that use robust methods to estimate the overall weight of an electric motor, this invention uses a method that breaks down the core components of the motor for estimation, establishes the relationship between each component and the motor's power, and finally adds them together to obtain a more accurate result.

[0070] This embodiment uses a relatively advanced three-phase asynchronous motor as an example, dividing the motor into three parts: stator, rotor, and housing. Rated power, stator, rotor, and housing mass data for multiple motor models were obtained from the market and motor manufacturers. Using rated power as the independent variable and the weight of the stator, rotor, and housing as the three dependent variables, scatter plots were created, and a suitable model was established for regression analysis. The results were then corrected based on the Multiple R-value, and three regression analyses were performed to obtain the relationship between rated power and the mass of the stator, rotor, and housing, respectively.

[0071] The electrical energy output power P T Substituting τ into the expressions relating rated power to the mass of the stator, rotor, and housing, and summing them, we obtain the corresponding expression for the mass of the motor in terms of τ, m. total (τ).

[0072] Furthermore, based on electrical energy, an expression for the battery mass m with respect to τ is derived. b (τ):

[0073] Here, based on the energy density ρ of the battery used, m is obtained. b (τ)=Q T τ / ρ.

[0074] After determining the mass of each major component, the mass of the additional connecting parts of each component remains roughly the same. Here, based on the actual situation of the aircraft, we set it to m. s .

[0075] Thus, we obtain the expression for the maximum takeoff mass of the modified aircraft in terms of τ:

[0076] m MTO '(τ)=m MTO +m b (τ)+m total (τ)+m s

[0077] Where m MTO The maximum takeoff mass of the aircraft before the electrical modification;

[0078] Step 6: Based on the maximum takeoff mass m of the modified aircraft obtained in Step 5... MTO '(τ), assuming the flight speed remains constant during the cruise phase of the flight plan, substitute...

[0079]

[0080] The lift coefficient C of the modified aircraft during the cruise phase was obtained. L ′(τ), and based on the drag curve obtained in step 3, the drag coefficient C of the modified aircraft during the cruise phase is obtained. D ′(τ);

[0081] Based on the drag coefficient C D ′(τ), using the formula

[0082]

[0083] Obtain the drag power P of the aircraft during the cruise phase. D (τ), and then combined with engine efficiency, the engine's cruising power P is obtained. S,cr ′(τ); based on the engine's cruising power P S,cr ′(τ) and engine fuel flow rate Q cr The fuel consumption rate of the modified engine was obtained:

[0084]

[0085] Step 7: Establish the energy efficiency formula

[0086]

[0087] Where ω(τ) represents energy efficiency, k is the set fuel price, C is the engine fuel consumption rate before electrification, and Q... T ′(τ) represents the total energy during the cruise phase after the electrification modification, Q m (τ) represents the electrical energy during the cruise phase after the electrification modification, L lift To increase the mass of the aircraft, the energy consumption during takeoff is increased, ω solid The inherent cost of the electric vehicle modification is represented by 'a', where 'a' is the set charging cost coefficient.

[0088] Within the set range of electricity consumption percentages, the electricity consumption percentage τ corresponding to the maximum energy efficiency is obtained through iterative calculation.

[0089] In summary, a clear, concise, and simple iterative optimization scheme for multi-electric aircraft can be obtained. Although this method is relatively simple, it can quickly help researchers determine a suitable energy allocation ratio.

[0090] Although embodiments of the present invention have been shown and described above, it is understood that the above embodiments are exemplary and should not be construed as limiting the present invention. Those skilled in the art can make changes, modifications, substitutions and variations to the above embodiments within the scope of the present invention without departing from the principles and spirit of the present invention.

Claims

1. An iterative optimization design method for a multi-electric aircraft hybrid power system, characterized in that, Includes the following steps: Step 1: Obtain flight data of the aircraft to be modified; based on the flight data, obtain the aircraft's lift coefficient C. L With drag coefficient C D Scatter plot of relationships; Step 2: Obtain the lift coefficient of the aircraft during the cruise phase based on the aircraft's cruise speed. and with ±Δ To determine the range of values, find the corresponding scatter plot of the lift-drag coefficient relationship during the cruise phase in the scatter plot obtained in step 1; Step 3: Use the scatter plot of the lift-drag coefficient relationship obtained in Step 2 to perform linear regression to obtain the drag curve of the aircraft during the cruise phase. Step 4: Obtain the flight speed during the cruise phase of the aircraft's flight plan. Using the drag curve for the cruise phase obtained in Step 3, determine the total power during the cruise phase. and total energy ; Step 5: Set the initial value for the proportion of electrical energy in the total energy of the aircraft after the power modification. The electrical energy expression is obtained. and the expression for electrical energy output power ; And based on the power output, the corresponding mass of the motor is obtained. expression The battery mass is obtained based on electrical energy. expression Combined with the set mass of additional connecting components The maximum takeoff mass of the modified aircraft was obtained. The expression: in The maximum takeoff mass of the aircraft before the electrical modification; Step 6: Based on the maximum takeoff mass of the modified aircraft obtained in Step 5. Under the condition that the flight speed remains constant during the cruise phase of the flight plan, the lift coefficient of the modified aircraft during the cruise phase is obtained. Based on the drag curve obtained in step 3, the drag coefficient of the modified aircraft during the cruise phase is obtained. According to the drag coefficient The drag power of the aircraft during the cruise phase is obtained, and thus the cruise power of the engine is obtained. Based on the engine's cruising power and engine fuel flow The modified engine's fuel consumption rate was obtained. ; Step 7: Establish the energy efficiency formula in For energy efficiency, k is the set fuel price, and C is the engine fuel consumption rate before the electrification conversion. This refers to the total energy consumed during the cruise phase after the electric modification. For the electric power needed for the cruise phase after the electric modification, Adding to the aircraft's mass increases energy consumption during takeoff. Due to the inherent costs of electric vehicle retrofitting, The set charging cost coefficient is used; within the set range of energy consumption ratio, the energy consumption ratio corresponding to the maximum energy efficiency is obtained through iterative calculation. .

2. The iterative optimization design method for a multi-electric aircraft hybrid power system as described in claim 1, characterized in that, In step 1, based on the flight data, using the formula... ， The lift coefficient C of the aircraft is obtained. L With drag coefficient C D A scatter plot of the relationships, where For aircraft quality, For weight acceleration, For the flight trajectory angle, air density, Where S is the flight speed, S is the wing area, and E is the lift-to-drag ratio.

3. The iterative optimization design method for a multi-electric aircraft hybrid power system as described in claim 2, characterized in that, In step 2, based on the aircraft's cruising speed Through formula Obtain the lift coefficient of the aircraft during the cruise phase. .

4. The iterative optimization design method for a multi-electric aircraft hybrid power system as described in claim 1, characterized in that, In step 2, Δ The value is 0.

2.

5. The iterative optimization design method for a multi-electric aircraft hybrid power system as described in claim 1, characterized in that, In step 3, the drag curve is represented as follows: by As the independent variable, Plot a scatter plot for the dependent variable, and perform linear regression based on the scatter plot to obtain the induced drag factor k and the zero-lift drag coefficient. This leads to the drag curve.

6. The iterative optimization design method for a multi-electric aircraft hybrid power system as described in claim 2, characterized in that, In step 4, based on the flight speed during the cruise phase in the aircraft's flight plan, and using the drag curve for the cruise phase obtained in step 3, the total power during the cruise phase is determined. and total energy The process is as follows: According to the flight speed during the cruise phase of the aircraft's flight plan According to the formula Obtain the lift coefficient of the aircraft during the cruise phase. Then, based on the drag curve during the aircraft's cruise phase, the corresponding drag coefficient is obtained. According to the formula Obtain the total power during the cruise phase In the formula To obtain the total energy , where t is the flight time.

7. The iterative optimization design method for a multi-electric aircraft hybrid power system as described in claim 1, characterized in that, In step 5, the mass of the motor is obtained based on the electrical energy output power. expression The process is as follows: The rated power, stator, rotor, and housing mass of multiple motor models were obtained. Regression analysis was performed with rated power as the independent variable and the weight of stator, rotor, and housing as the three dependent variables to obtain the relationship between rated power and the mass of stator, rotor, and housing, respectively. Power output of electrical energy Substituting the values ​​of rated power with respect to the mass of the stator, rotor, and housing into the given expressions, and summing them, we obtain the corresponding motor mass with respect to... expression .

8. The iterative optimization design method for a multi-electric aircraft hybrid power system as described in claim 1, characterized in that, In step 5, the battery mass is obtained based on the electrical energy. expression The process is as follows: based on the energy density of the battery used... ,get .

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