A rapid calculation method for flight performance based on a prototype aircraft

Abstract

The present application provides a method for quickly calculating the flight performance based on a prototype aircraft. The method includes: taking the prototype aircraft as a platform, keeping the power plant of the target aircraft consistent with that of the prototype aircraft, so as to keep the power characteristics of the target aircraft unchanged, and making the overall design parameters of the target aircraft consistent with those of the prototype aircraft. The aerodynamic configuration of the target aircraft is changed or optimized according to the design requirements; by constructing a relationship model between the change of aerodynamic characteristics data and the change of flight performance data of the target aircraft, based on the flight performance data of the prototype aircraft, the flight performance data of the target aircraft under the aerodynamic configuration is obtained according to the change of the aerodynamic characteristics data of the prototype aircraft.

Description

Technical Field

[0001] This application belongs to the technical field of aircraft performance design, and particularly relates to a rapid calculation method for flight performance based on a prototype aircraft. Background Art

[0002] In aircraft design, improvement designs for lift augmentation and drag reduction are often carried out based on a prototype aircraft, or modification designs are carried out with the prototype aircraft as a platform. For example, taking a transport aircraft as the platform aircraft, adding equipment such as a refueling pod for modification design into a tanker, or adding radar equipment for modification design into an early warning aircraft. In these cases, the power plant of the aircraft basically remains unchanged, and the aerodynamic shape of the designed aircraft needs to be optimized or adaptively changed. At this time, it is necessary to recalculate or calculate the flight performance of the optimized-designed aircraft or the modified aircraft.

[0003] The aircraft after optimization design or modification based on the prototype aircraft is called the target aircraft. During the preliminary design stage of the target aircraft, the main flight performance design parameters of the target aircraft need to be calculated. During the aircraft flight test and operation stages, comprehensive operation data such as a flight performance manual needs to be provided for the target aircraft. Therefore, it is necessary to recalculate and update the flight performance data of the target aircraft. The conventional method of the existing technology is to recalculate the performance of each flight stage and the design mission profile according to the flight mechanics equations based on the overall parameters, aerodynamic data, and power characteristic data of the target aircraft. However, this method requires a very large amount of work, a lot of time and manpower, and has low work efficiency when determining and updating the flight performance data, which affects the iteration speed of the design scheme during the preliminary design stage. Summary of the Invention

[0004] The purpose of this application is to provide a rapid calculation method for flight performance based on a prototype aircraft to solve or alleviate at least one problem in the background art.

[0005] The technical solution of this application is: a rapid calculation method for flight performance based on a prototype aircraft, the method comprising:

[0006] Using the prototype aircraft as a platform, making the power plant of the target aircraft consistent with that of the prototype aircraft, so that the power characteristics of the target aircraft and the prototype aircraft remain unchanged, and making the overall design parameters of the target aircraft and the prototype aircraft consistent, and changing or optimizing the aerodynamic shape of the target aircraft according to the design requirements;

[0007] By constructing a relationship model between the change in aerodynamic characteristics data and the change in flight performance data of the target aircraft, based on the flight performance data of the prototype aircraft, the flight performance data of the target aircraft under the aerodynamic shape is obtained according to the change in the aerodynamic characteristics data of the prototype aircraft.

[0008] In this application, the relationship model between the changes in the aerodynamic characteristic data and the changes in the flight performance data of the target aircraft is constructed based on the longitudinal aerodynamic data of the target aircraft and the prototype aircraft.

[0009] In this application, the flight performance data of the prototype aircraft and the target aircraft include climb performance, climb limit weight, cruise or patrol performance, and descent performance. Among them, the climb performance includes climb rate, climb time, climb distance, and climb fuel consumption. The cruise or patrol performance includes range per unit fuel and hourly fuel consumption. The descent performance includes descent rate, descent time, descent distance, and descent fuel consumption.

[0010] Further, preprocess the longitudinal aerodynamic data of the prototype aircraft. The preprocessing process includes:

[0011] Express the polar curve of the prototype aircraft as an nth-degree polynomial with the lift-to-drag ratio as the dependent variable and the lift coefficient as the independent variable:

[0012] K b (Ma,CL) = A Ma_n ×CL n +A Ma_n-1 ×CL n-1 +......+A Ma_0

[0013] In the formula, K b is the lift-to-drag ratio of the prototype aircraft, and A Ma_n , A Ma_n-1 , …, A Ma_0 are all polynomial coefficients, Ma is the Mach number, and CL is the lift coefficient.

[0014] Further, preprocess the longitudinal aerodynamic data of the target aircraft. The preprocessing process includes:

[0015] Express the polar curve of the target aircraft as an nth-degree polynomial with the lift-to-drag ratio as the dependent variable and the lift coefficient as the independent variable:

[0016] K x (Ma,CL) = B Ma_n ×CL n +B Ma_n-1 ×CL n-1 +......+B Ma_0

[0017] In the formula, K x is the lift-to-drag ratio of the target aircraft, and B Ma_n , B Ma_n-1 , ……, B Ma_0 are all polynomial coefficients, Ma is the Mach number, and CL is the lift coefficient;

[0018] At the same lift coefficient \(C_L\), the difference in lift-to-drag ratio \(\Delta K\) between the target aircraft and the prototype aircraft satisfies: \(\Delta K=(B Ma_n - A Ma_n )C_L n +(B Ma_n-1 - A Ma_n-1 )C_L n-1 +......+(B Ma_0 - A Ma_0 )。

[0019] Furthermore, under the same flight conditions, based on the climb performance data of the prototype aircraft, the climb performance of the target aircraft is determined according to the change in the lift-to-drag ratio. The process includes:

[0020] When the flight weight \(W\) is the same, according to the force balance relationship of the aircraft's steady climb, the lift coefficient can be expressed as:

[0021] Lift

[0022] Then the lift coefficient

[0023] The corresponding drag

[0024] The climb speed \(V = M_a\times a\);

[0025] The climb rate is then

[0026] In the formula, \(\rho\) is the air density, \(V\) is the cruise speed, \(S\) is the wing reference area, \(C_L\) is the lift coefficient, \(C_D\) is the drag coefficient, \(F\) is the installed thrust, \(Y\) is the lift, \(D\) is the drag, \(V_y\) is the climb rate, and \(K\) is the lift-to-drag ratio;

[0027] Under the same climb conditions, the relationship between the climb rates of the target aircraft and the prototype aircraft is as follows:

[0028]

[0029] In the formula, \(V_y x represents the climb rate of the target aircraft, \(V_y b represents the climb rate of the prototype aircraft, \(K b is the lift-to-drag ratio of the prototype aircraft, and \(K x is the lift-to-drag ratio of the target aircraft;

[0030] That is, the climb time, climb fuel consumption, and climb distance data of the target aircraft are obtained.

[0031] Furthermore, under the same climb conditions, based on the climb ability limit weight data of the prototype aircraft, the climb ability limit weight of the target aircraft is determined according to the change relationship between the climb ability limit weight and the lift-to-drag ratio. The process includes:

[0032] Climb capacity limit weight of the target aircraft

[0033] Wherein, CG0 is the required climb gradient value, and W b is the climb limit weight of the prototype aircraft, and K b is the lift-to-drag ratio of the prototype aircraft, and K x is the lift-to-drag ratio of the target aircraft;

[0034] Among them, the lift coefficient CL of the prototype aircraft b and the lift coefficient CL of the target aircraft x are determined by the following formula:

[0035] Wherein, CL is the lift coefficient, and W b is the climb capacity limit weight of the prototype aircraft, ρ is the atmospheric density, V is the cruise speed, and S is the wing reference area.

[0036] Furthermore, under the same flight conditions, based on the unit fuel range and hourly fuel consumption of the prototype aircraft, the cruise performance data of the target aircraft are determined according to the changes in the unit fuel range, hourly fuel consumption of the target aircraft, the unit fuel range and hourly fuel consumption of the prototype aircraft, and the lift-to-drag ratio. The process is as follows:

[0037] Unit fuel range of the target aircraft

[0038] Hourly fuel consumption of the target aircraft

[0039] Wherein, r b is the unit fuel range of the prototype aircraft, r x is the unit fuel range of the target aircraft, q b is the hourly fuel consumption of the prototype aircraft, q x is the hourly fuel consumption of the target aircraft, K b is the cruise lift-to-drag ratio of the prototype aircraft, and K x is the cruise lift-to-drag ratio of the target aircraft.

[0040] The fast calculation method of flight performance based on the prototype aircraft of the present application has the following advantages:

[0041] 1) The method of the present application can be used for the estimation of typical flight performance parameters in the design of the target aircraft scheme, flight test and rapid update of flight performance data during use. The estimation content mainly includes the climb / descent performance, climb capacity limit weight, and cruise / patrol performance of the target aircraft;

[0042] 2) When calculating relevant flight performance data for the optimized design based on the prototype aircraft or the modified aircraft, the method of this application does not require a large number of new data iterative calculations using the conventional method of solving flight mechanics equations. The calculation formula is simple and easy to implement. There is no need for iterative calculations during the implementation process, which greatly reduces the repetitive labor of aircraft designers and improves work efficiency; in the design of modified or improved aircraft, calculations can be quickly performed to obtain results that meet the engineering accuracy requirements, which can quickly promote the iteration of design schemes and shorten the development cycle;

[0043] 3) For limited optimization and improvement on the prototype aircraft, in the compilation of relevant flight performance data such as the aircraft flight performance manual, data can be updated quickly and in a timely manner, effectively improving the maintenance work efficiency of user materials and the technical support efficiency for users. BRIEF DESCRIPTION OF THE DRAWINGS

[0044] To more clearly illustrate the technical solutions provided in this application, the accompanying drawings will be briefly introduced below. Obviously, the drawings described below are only some embodiments of this application.

[0045] Figure 1 It is a flowchart of the flight performance rapid calculation method based on the prototype aircraft in this application.

[0046] Figure 2 It is a schematic diagram of the function curve of the lift-drag ratio K of the prototype aircraft in an embodiment of this application b expressed as a function of the lift coefficient CL b schematic diagram.

[0047] Figure 3 It is a schematic diagram of the function curve of the lift-drag ratio K of the target aircraft in an embodiment of this application x expressed as a function of the lift coefficient CL x schematic diagram. DETAILED DESCRIPTION OF THE EMBODIMENTS

[0048] To make the purpose, technical solutions, and advantages of the implementation of this application clearer, the technical solutions in the embodiments of this application will be described in more detail below with reference to the accompanying drawings in the embodiments of this application.

[0049] As Figure 1 shown, this application provides a rapid flight performance calculation method based on a prototype aircraft, which specifically includes the following steps:

[0050] 1) Taking the prototype aircraft as a platform, keeping the power plant of the target aircraft consistent with that of the prototype aircraft, so that the power characteristics of the target aircraft and the prototype aircraft remain unchanged, and the overall design parameters of the target aircraft and the prototype aircraft are consistent. The aerodynamic configuration of the target aircraft is changed or optimized according to design requirements.

[0051] Among them, the input data are the longitudinal aerodynamic force data of the target aircraft and the prototype aircraft, the climb performance of the prototype aircraft (including climb rate, climb time, climb distance, climb fuel consumption), the climb limit weight, the cruise / patrol performance (range per unit fuel, fuel consumption per hour), and the descent performance data (including descent rate, descent time, descent distance, descent fuel consumption).

[0052] 2) By constructing a relationship model between the change in the aerodynamic characteristics data and the change in the flight performance data of the target aircraft, based on the flight performance data of the prototype aircraft, the flight performance data of the target aircraft under the aerodynamic configuration are obtained according to the change in the aerodynamic characteristics data of the prototype aircraft.

[0053] Specifically, first, preprocess the aerodynamic force data required for estimating the target aircraft and the prototype aircraft. The preprocessing process is as follows:

[0054] Express the polar curve of the prototype aircraft as an nth-degree polynomial with the lift-to-drag ratio K b as the dependent variable and the lift coefficient CL as the independent variable:

[0055] K b (Ma,CL) = A Ma_n ×CL n + A Ma_n-1 ×CL n-1 +……+ A Ma_0 (1)

[0056] At the same time, express the polar curve of the target aircraft as an nth-degree polynomial with the lift-to-drag ratio K x as the dependent variable and the lift coefficient CL as the independent variable:

[0057] K x (Ma,CL) = B Ma_n ×CL n + B Ma_n-1 ×CL n-1 +……+ B Ma_0 (2)

[0058] Therefore, under the same lift coefficient CL, the lift-to-drag ratio difference ΔK between the target aircraft and the prototype aircraft is:

[0059] ΔK = (A Ma_n - B Ma_n )×CL n +(A Ma_n-1 - B Ma_n-1 )×CL n-1 +…+(A Ma_0 - B Ma_0 ) (3)

[0060] In the above three equations, K is the lift-to-drag ratio, the subscript b represents the prototype aircraft, the subscript x represents the target aircraft, CL is the lift coefficient, Ma is the Mach number, A Ma_n 、A Ma_n-1 、…、A Ma_0 and B Ma_n 、B Ma_n-1 、…、B Ma_0 are all coefficients of the polynomial, n is the number of polynomial terms and also the exponent.

[0061] After that, the climb / descent performance data of the target aircraft are calculated based on the performance data of the prototype aircraft under the same flight conditions according to the change of the lift-to-drag ratio. The calculation process is as follows:

[0062] According to the force balance relationship and kinematic relationship of the aircraft in steady climb, the climb rate Vy is related to the lift coefficient CL and the corresponding lift-to-drag ratio K:

[0063] When the aircraft is in steady climb, the lift Y satisfies

[0064] Then there is

[0065] The corresponding drag satisfies

[0066] The climb speed V is V = Ma × a; (5)

[0067] According to the above equation, the expression of the climb rate can be obtained:

[0068]

[0069] In the formula, CL is the lift coefficient, CD is the drag coefficient, F is the installed thrust, Y is the lift, D is the drag, and Vy is the climb rate.

[0070] Take the climb performance data of the prototype aircraft, which are the known quantities in Table 1. According to Table 1, the climb height h is a known quantity, and the height step is Δh = h i+1 -h i-1 、W b_0 、W x_0 The initial climb weights of the prototype aircraft and the modified aircraft are known quantities, and Vy b_i is the climb rate of the prototype aircraft.

[0071] Under the same atmospheric conditions, when the initial climb height, initial climb weight, and climb speed are the same, the climb performance data of the prototype aircraft can be used to find: the climb weight W x_i 、the climb rate Vy x_1 、the climb time t x_i 、the climb fuel consumption Wf x_i, Climb distance Dis x_i , which is the unknown in Table 1.

[0072] Table 1 Climb performance data of prototype aircraft and target aircraft

[0073]

[0074]

[0075] According to the relationships in the above steps, the climb performance data of the target aircraft can be calculated as follows:

[0076] According to the relationships in Equations (4) to (8), under the conditions of known climb weight and climb speed, first determine the lift coefficient CL, and then on the CL-K curve corresponding to the climb speed, determine the corresponding lift-to-drag ratio K according to Equations (1) and (2) b , K x .

[0077] Specifically, first determine the climb rate of the target aircraft and the prototype aircraft under the same climb conditions. The formula is as follows:

[0078]

[0079] The time t from altitude h0 to h1 x_1 can be expressed as

[0080] The fuel consumption Dis from altitude h0 to h1 x_1 can be expressed as:

[0081]

[0082] The climb distance Wf from altitude h0 to h1 x_1 can be expressed as:

[0083]

[0084] At altitude h1, the flight weight of the target aircraft is: W x_1 = W x_0 - Wf x_1 ; (13)

[0085] At altitude h i , the climb data of the target aircraft are as follows in Equations (14), (15), (16), (17), where i > 1 and i is an integer:

[0086]

[0087]

[0088]

[0089] W x_i = W x_i-1 - Wf x_i (17)

[0090] It should be noted that the above method for estimating the climb performance data of the target aircraft is also applicable to the estimation of the descent performance of the target aircraft.

[0091] Then, calculate the weight limit for the climb ability of the target aircraft.

[0092] Under the same climb conditions, based on the weight limit data for the climb ability of the prototype aircraft, calculate the weight limit for the climb ability of the target aircraft according to the variation relationship between the weight limit for the climb ability and the lift-to-drag ratio. The process is as follows:

[0093] The target aircraft and the prototype aircraft use the same configuration, with the same atmospheric temperature and altitude during flight, and the climb speed V b ≈ V x , according to the force balance relationship of the aircraft's steady climb, when the climb gradient requirement of the prototype aircraft is not less than CG0, the maximum allowable flight weight W0 of the prototype aircraft, the engine thrust F, and the corresponding lift-to-drag ratio K b , and the relationship with the climb gradient requirement value CG0 is as follows:

[0094]

[0095] When the climb gradient requirement of the target aircraft is not less than CG0, the maximum allowable flight weight W x and the engine thrust F, the corresponding lift-to-drag ratio K x , and the relationship with the climb gradient requirement value CG0 is as follows:

[0096]

[0097] From equations (18) and (19), the calculation formula for the climb limit weight of the target aircraft can be obtained as follows:

[0098]

[0099] In equation (20), the climb gradient requirement value CG0 is a known quantity, W0 is the climb limit weight of the prototype aircraft, which is a known quantity, K b and K x are calculated from equations (1) and (2). The process is as follows:

[0100] The lift coefficient CL b of the prototype aircraft at the climb limit weight W b is calculated by the following formula:

[0101]

[0102] From the above formula, using the relationship curve (Equation 1) between the lift coefficient CL and the lift-to-drag ratio K of the prototype aircraft, the corresponding lift-to-drag ratio K b value can be obtained.

[0103] Let the lift coefficient CL of the target aircraft at the climb limit weight W x be approximately CL x , and using the relationship curve (Equation 2) between the lift coefficient CL and the lift-to-drag ratio K of the target aircraft, the corresponding lift-to-drag ratio K b value can be obtained. x value.

[0104] Finally, calculate the cruise performance data of the target aircraft.

[0105] Based on the cruise performance data of the prototype aircraft under the same flight conditions, calculate the cruise performance data of the target aircraft through the variation relationship between the fuel range per unit, fuel consumption per hour, and lift-to-drag ratio under the same flight conditions. The calculation process is as follows:

[0106] The fuel range per unit r can be expressed as the ratio of the flight distance per unit time (speed V) to the fuel consumption per unit time (Wf). The fuel consumption per unit time is equal to the product of the net thrust of the engine and the specific fuel consumption rate.

[0107] When the aircraft is flying at a constant horizontal speed, the thrust is balanced with the drag, and the lift is balanced with the gravity. Therefore, the fuel range per unit of the prototype aircraft at the flight weight W and flight speed V is expressed as:

[0108]

[0109] The calculation formula for the fuel consumption per hour of the prototype aircraft is:

[0110] The cruise lift coefficient is:

[0111] where Wf cru is the fuel consumption per hour in the cruise state, sfc is the specific fuel consumption rate, F cru is the required thrust during cruise, L is the lift during cruise, D is the drag during cruise, and K b is the lift-to-drag ratio of the prototype aircraft corresponding to the cruise state.

[0112] Under the same atmospheric temperature, flight altitude, flight weight, and flight speed as the prototype aircraft, respectively according to Equation (22) and Equation (23), the expression (25) for the fuel range per unit r x of the target aircraft and the expression (26) for the cruise fuel consumption per hour of the target aircraft can be derived:

[0113] Unit fuel range of the target aircraft

[0114] Cruising hourly fuel consumption of the target aircraft

[0115] Wherein, r b , q b are known quantities, and K b , K x are calculated from Equation 24, Equation 1, and Equation 2.

[0116] If the flight weights in Equation (25) and Equation (26) are the real-time cruising weights of the aircraft at a certain altitude and speed, these two calculation formulas can be used to calculate the cruising performance data at a single point. If the flight weight is the average flight weight, these two calculation formulas can be used to calculate the range change of the target aircraft compared with the prototype aircraft.

[0117] The fast estimation method of flight performance based on the prototype aircraft provided by this application is different from the conventional method in that:

[0118] 1) The input data includes the main overall parameters of the aircraft, the flight performance data of the prototype aircraft, the longitudinal trim aerodynamic data of the prototype aircraft, and the longitudinal trim aerodynamic data of the target aircraft, without using the engine power characteristic data;

[0119] 2) Based on the flight performance data of the prototype aircraft, the changes in flight performance parameters are correlated with the aerodynamic forces (mainly the lift-to-drag ratio K), and the mechanical equations of aircraft flight do not need to be solved during the calculation process;

[0120] 3) The content of the flight performance data estimated quickly includes the climb / descent performance, cruise / patrol performance, and climb limit weight of the target aircraft.

[0121] Therefore, the fast calculation method of flight performance based on the prototype aircraft of this application has the following advantages:

[0122] 1) The method of this application can be used for the estimation of typical flight performance parameters in the design of the target aircraft scheme, flight test, and the rapid update of flight performance data during use. The estimation content mainly includes the climb / descent performance, climb capacity limit weight, and cruise / patrol performance of the target aircraft;

[0123] 2) When calculating relevant flight performance data for the optimized design based on the prototype aircraft or the modified aircraft using the method of this application, there is no need to adopt the conventional method of solving flight mechanics equations for a large number of new data iterative calculations. The calculation formula is simple and easy to implement. There is no need for iterative calculations during the implementation process, which greatly reduces the repetitive labor of aircraft designers and improves work efficiency; in the design of modified or improved aircraft, calculations can be quickly carried out to obtain results that meet the engineering accuracy requirements, which can quickly promote the iteration of design schemes and shorten the development cycle;

[0124] 3) For limited optimization and improvement on the prototype aircraft, in the compilation of relevant flight performance data such as the aircraft flight performance manual, data can be updated quickly and in a timely manner, effectively improving the work efficiency of user data maintenance and the technical support efficiency for users.

[0125] As mentioned above, the above are only specific implementation manners of this application, but the protection scope of this application is not limited thereto. Any changes or substitutions that can be easily thought of by those skilled in the art within the technical scope disclosed in this application should be covered by the protection scope of this application. Therefore, the protection scope of this application should be subject to the protection scope of the claims.

Claims

1. A rapid flight performance calculation method based on a prototype aircraft, characterized in that, The method includes: Using the prototype aircraft as a platform, making the power plant of the target aircraft consistent with that of the prototype aircraft, so that the power characteristics of the target aircraft remain unchanged compared with the prototype aircraft, and making the overall design parameters of the target aircraft consistent with those of the prototype aircraft. The aerodynamic configuration of the target aircraft is changed or optimized according to design requirements. Among them, the relationship model between the changes in aerodynamic characteristic data and flight performance data of the target aircraft is constructed based on the longitudinal aerodynamic force data of the target aircraft and the prototype aircraft. The longitudinal aerodynamic forces of the prototype aircraft and the target aircraft include climb performance, climb limit weight, cruise or patrol performance, and descent performance. Among them, the climb performance includes climb rate, climb time, climb distance, and climb fuel consumption. The cruise or patrol performance includes specific fuel range and fuel consumption per hour. The descent performance includes descent rate, descent time, descent distance, and descent fuel consumption; By constructing the relationship model between the changes in aerodynamic characteristic data and flight performance data of the target aircraft, based on the flight performance data of the prototype aircraft, the flight performance data of the target aircraft under the aerodynamic configuration is obtained according to the changes in the aerodynamic characteristic data of the prototype aircraft, including: Preprocessing the longitudinal aerodynamic force data of the prototype aircraft. The process is as follows: Express the polar curve of the prototype aircraft as an nth-degree polynomial with the lift-to-drag ratio as the dependent variable and the lift coefficient as the independent variable: K b (Ma,CL) = A Ma_n ×CL n +A Ma_n-1 ×CL n-1 +......+A Ma_0 where K b is the lift-drag ratio of the prototype aircraft, A Ma_n , A Ma_n-1 , …, A Ma_0 are all polynomial coefficients, Ma is the Mach number, and CL is the lift coefficient; Preprocessing the longitudinal aerodynamic force data of the target aircraft. The process is as follows: Express the polar curve of the target aircraft as an nth-degree polynomial with the lift-to-drag ratio as the dependent variable and the lift coefficient as the independent variable: K x (Ma,CL) = B Ma_n ×CL n +B Ma_n-1 ×CL n-1 +......+B Ma_0 Where K x is the lift-drag ratio of the target aircraft, B Ma_n , B Ma_n-1 , ……, B Ma_0 are all polynomial coefficients, Ma is the Mach number, and CL is the lift coefficient; Under the same lift coefficient \(C_L\), the lift-drag ratio difference \(\Delta K\) between the target aircraft and the prototype aircraft satisfies: \(\Delta K=(B Ma_n -A Ma_n )C_L n +(B Ma_n-1 -A Ma_n-1 )C_L n-1 +\cdots+(B Ma_0 -A Ma_0 ) Under the same flight conditions, based on the climb performance data of the prototype aircraft, the climb performance of the target aircraft is determined according to the change in the lift-to-drag ratio. The process includes: When the flight weight W is the same, according to the force balance relationship of the aircraft in steady climb, the lift coefficient can be expressed as: Lift The lift coefficient Corresponding resistance The climb speed V = Ma × a; The rate of climb is then In the formula, ρ is the air density, V is the climb speed, S is the wing reference area, CL is the lift coefficient, CD is the drag coefficient, F is the installed thrust, Y is the lift, D is the drag, and K is the lift-to-drag ratio; Under the same climb conditions, the following relationship exists between the climb rates of the target aircraft and the prototype aircraft: where Vy x represents the climb rate of the target aircraft, and Vy b represents the climb rate of the prototype aircraft, and K b is the lift-to-drag ratio of the prototype aircraft, and K x is the lift-to-drag ratio of the target aircraft; That is, the climb time, climb fuel consumption, and climb distance data of the target aircraft are obtained; Under the same climb conditions, based on the climb capability limit weight data of the prototype aircraft, the climb capability limit weight of the target aircraft is determined according to the change relationship between the climb capability limit weight and the lift-to-drag ratio. The process includes: Climbing ability limited weight of the target aircraft where CG0 is the required climb gradient value, W b is the climb limit weight of the prototype aircraft, K b is the lift-to-drag ratio of the prototype aircraft, K x is the lift-to-drag ratio of the target aircraft; Among them, the lift coefficient CL of the prototype aircraft b and the lift coefficient CL of the target aircraft x are determined by the following formula: CL x ≈CL b Where, W b is the climb limit weight of the prototype aircraft, ρ is the atmospheric density, V is the cruise speed, and S is the wing reference area; Finally, under the same flight conditions, based on the specific fuel range and fuel consumption per hour of the prototype aircraft, the cruise performance data of the target aircraft is determined according to the specific fuel range and fuel consumption per hour of the target aircraft, the specific fuel range and fuel consumption per hour of the prototype aircraft, and the change in the lift-to-drag ratio. The process is as follows: Unit fuel range of the target aircraft Fuel consumption per hour of the target aircraft where r b is the fuel range per unit fuel of the prototype aircraft, r x is the fuel range per unit fuel of the target aircraft, q b is the fuel consumption per hour of the prototype aircraft, q x is the fuel consumption per hour of the target aircraft, K b is the lift-drag ratio of the prototype aircraft, K x is the lift-drag ratio of the target aircraft.

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