Air and ground consistency correction method for aerodynamic force of high-speed aircraft

Abstract

The invention relates to a heaven and earth consistency correction method for aerodynamic force of a high-speed aircraft. The method comprises the following steps: S1, calculating aerodynamic performance of the high-speed aircraft in different states in a to-be-corrected flight state and an incoming flow working condition; s2, the aerodynamic performance deviation rate of the high-speed aircraft in different states under the flight state to be corrected and the incoming flow working condition is calculated; s3, calculating a turbulence degree proportion, a thermal deformation degree proportion and a rudder deflection degree proportion of the aircraft; s4, solving a rudder deflection correction amount, a thermal deformation correction amount and a transition correction amount under the to-be-corrected flight state and the incoming flow working condition; and S5, correcting the aerodynamic performance of the aircraft obtained by ground simulation by adopting the rudder deflection correction, the thermal deformation correction and the transition correction, so that the aerodynamic performance of the aircraft is consistent with the aerodynamic performance of the aircraft in a real flight state.

Description

Technical Field

[0001] The invention belongs to the field of numerical simulation of aerodynamic performance of high-speed aircraft, and is used to solve the problem of aerodynamic performance of high-speed aircraft under real flight state and predicting aerodynamic performance of flight test aircraft using ground data. Background Art

[0002] With the rapid development of aircraft performance, the current theoretical calculation methods and ground test methods have certain limitations in the study of the aerodynamic performance of high-speed aircraft in real flight conditions. In order to make the theoretical calculation methods at the physical model level and the wind tunnel test environment better reflect the research status in the real flight environment, it is necessary to conduct a comprehensive study on the consistency of high-speed aircraft aerodynamic performance between the sky and the earth through a combination of various means. Summary of the invention

[0003] The technical problem solved by the present invention is: in view of the problem that the traditional theoretical calculation method of high-speed aircraft cannot well reflect the aerodynamic performance in the actual flight state, a method for correcting the sky-ground consistency of the aerodynamic force of high-speed aircraft is proposed. For the problem of the sky-ground difference in the aerodynamic performance of high-speed aircraft, based on the analysis and modeling of the sky-ground difference, a sky-ground consistency correction principle is established to realize the prediction of the aerodynamic performance of the aircraft under the actual flight state.

[0004] The technical solution of the present invention is: a method for correcting the sky-ground consistency of aerodynamic force of a high-speed aircraft, the method comprising the following steps:

[0005] S1. Calculate the aerodynamic performance of the high-speed aircraft under different states under the flight state to be corrected and the incoming flow conditions, including the aerodynamic performance of the reference state, no rudder deviation, and laminar flow, the aerodynamic performance of the reference state, no rudder deviation, and turbulent flow, the aerodynamic performance of the reference state, with rudder deviation, and laminar flow, the aerodynamic performance of the reference state, with rudder deviation, and turbulent flow, the aerodynamic performance of thermal deformation, no rudder deviation, and laminar flow, the aerodynamic performance of thermal deformation, no rudder deviation, and turbulent flow, the aerodynamic performance of the reference state, no rudder deviation, and transition, the aerodynamic performance of the reference state, maximum rudder deviation, and laminar flow, the aerodynamic performance of the reference state, minimum rudder deviation, and laminar flow, the aerodynamic performance of the reference state, maximum rudder deviation, and turbulent flow, and the aerodynamic performance of the reference state, minimum rudder deviation, and turbulent flow;

[0006] S2. Calculate the aerodynamic performance deviation rate of the high-speed aircraft under different states under the flight state to be corrected and the incoming flow conditions, including the aerodynamic performance deviation rate of the rudder-biased aircraft in the reference state and laminar flow, the aerodynamic performance deviation rate of the rudder-biased aircraft in the reference state and turbulent flow, the aerodynamic performance deviation rate of the thermal deformation aircraft in the thermal deformation and laminar flow, the aerodynamic performance deviation rate of the thermal deformation aircraft in the thermal deformation and turbulent flow, and the aerodynamic performance deviation rate of the transitional aircraft in the reference state and transition;

[0007] S3, calculating the proportion of turbulence degree, thermal deformation degree and rudder deflection degree of the aircraft;

[0008] S4, using the aerodynamic performance deviation rate of the aircraft in each rudder deviation state, the aerodynamic performance deviation rate of the aircraft in the thermal deformation state, and the aerodynamic performance deviation rate of the aircraft in the transition mode, as well as the aircraft turbulence degree ratio, thermal deformation degree ratio, and rudder deviation degree ratio, through interpolation fitting, calculate the rudder deviation correction amount, thermal deformation correction amount, and transition correction amount under the flight state to be corrected and the incoming flow condition;

[0009] S5. Use rudder deviation correction, thermal deformation correction and transition correction to correct the aerodynamic performance of the aircraft obtained by ground simulation to make it consistent with the aerodynamic performance of the aircraft under actual flight conditions.

[0010] Preferably, the step S1 is calculated using a CFD tool.

[0011] Preferably, the aerodynamic performance deviation rate of the rudder-biased aircraft in the reference state and laminar flow and the aerodynamic performance deviation rate of the rudder-biased aircraft in the reference state and turbulent flow are calculated using the following formula:

[0012]

[0013] Where H is the altitude in km, Ma is the dimensionless parameter Mach number, α is the angle of attack, β is the sideslip angle, i is the aerodynamic performance term, i∈(x,y,z,mx,my,mz), x is the axial force, y is the normal force, z is the lateral force, mx is the rolling moment, my is the yaw moment, mz is the pitch moment, δ∈(δ1,δ2,…,δ n ),δi∈[δi min ,δi max ], δ is the rudder angle, and is selected from the preset rudder angle set (δ1, δ2, …, δ n ), δ i is the i-th element in the rudder angle set, δi min is the minimum rudder deflection of the aircraft, δi max is the maximum rudder deflection of the aircraft, bas represents the reference state, s is the incoming flow state, s∈(laminar, turbulence), when s=laminar represents laminar flow, s=turbulence represents turbulent flow.

[0014] Preferably, the aerodynamic performance deviation rate of thermal deformation aircraft with thermal deformation and laminar flow, and the aerodynamic performance deviation rate of thermal deformation aircraft with thermal deformation and turbulent flow are calculated using the following formula:

[0015]

[0016] Among them, H is the altitude in km, Ma is the dimensionless parameter Mach number, α is the angle of attack, β is the sideslip angle, i is the aerodynamic performance, hd represents thermal deformation, i∈(x,y,z,mx,my,mz), x is the axial force, normal force, and lateral force, mx is the rolling moment, my is the yaw moment, mz is the pitch moment, s is the incoming flow state, s∈(laminar,turbulence), when s=laminar, it means laminar flow, and s=turbulence, it means turbulent flow.

[0017] Preferably, the aerodynamic performance deviation rate of the transitional aircraft in the reference state and the transitional state;

[0018]

[0019] Among them, H is the altitude in km, Ma is the dimensionless parameter Mach number, α is the angle of attack, β is the sideslip angle, i is the aerodynamic performance, bas represents the reference state, i∈(x,y,z,mx,my,mz), x is the axial force, normal force, and lateral force, mx is the rolling moment, my is the yaw moment, mz is the pitch moment, s is the incoming flow state, s∈(laminar,turbulence), when s=laminar represents laminar flow, s=turbulence represents turbulent flow.

[0020] Preferably, the turbulence degree accounts for γ H,Ma,α,β The calculation is done using the following formula:

[0021]

[0022] Wherein, H is the altitude in km, Ma is the dimensionless parameter Mach number, α is the angle of attack, β is the sideslip angle, bas represents the reference state, i is the aerodynamic performance term, i∈(x,y,z,mx,my,mz), x is the axial force, y is the normal force, z is the lateral force, mx is the rolling moment, my is the yaw moment, mz is the pitch moment, s is the incoming flow state, s∈(laminar,turbulence), when s=laminar represents laminar flow, s=turbulence represents turbulent flow; Represents the average value of all aerodynamic performance indicators.

[0023] Preferably, the proportion of aircraft thermal deformation degree τ is calculated H,Ma,α,β The calculation formula is as follows:

[0024]

[0025] Among them, H is the altitude in km, Ma is the dimensionless parameter Mach number, α is the angle of attack, β is the sideslip angle, bas represents the reference state, hd represents the thermal deformation, i represents the aerodynamic performance, i∈(x,y,z,mx,my,mz), x represents the axial force, y represents the normal force, z represents the lateral force, mx represents the rolling moment, my represents the yaw moment, mz represents the pitching moment, s represents the incoming flow state, s∈(laminar,turbulence), when s=laminar represents laminar flow, s=turbulence represents turbulent flow, δ represents the rudder deflection angle, It means to find the average value of each aerodynamic performance. It is the aerodynamic performance in the state of maximum thermal deformation.

[0026] Preferably, the proportion of the aircraft rudder deviation σ is calculated H,Ma,α,β The calculation formula is as follows:

[0027]

[0028] Among them, H is the height in km, Ma is the dimensionless parameter Mach number, α is the angle of attack, β is the sideslip angle, i is the aerodynamic performance, i∈(x,y,z,mx,my,mz), x is the axial force, normal force, lateral force, mx is roll, my is yaw, mz is pitch, δ∈(δ1,δ2,…,δ n ),δi∈[δi min ,δi max ], δ is the rudder angle, and is selected from the preset rudder angle set (δ1, δ2, …, δ n ), δ i is the i-th element in the rudder angle set, δi min is the minimum rudder deflection of the aircraft, δi max is the maximum rudder deflection of the aircraft, bas represents the reference state, s represents the incoming flow state, s∈(laminar, turbulence), when s=laminar represents laminar flow, s=turbulence represents turbulent flow;

[0029] C_max H,Ma,α,β,bas,i,s,δ is the maximum rudder deflection aerodynamic performance, C_max H,Ma,α,β,bas,i,s,δi =max(C H,Ma,α,β,bas,i,s,δi ,δi∈[δi min ,δi max ]);

[0030] C_min H,Ma,α,β,bas,i,s,δ The minimum rudder deflection aerodynamic performance: C_min H,Ma,α,β,bas,i,s,δi =min(C H,Ma,α,β,bas,i,s,δi ,δi∈[δi min ,δi max ]).

[0031] Preferably, according to the rudder deviation correction amount Thermal deformation correction and transition correction And the reference aircraft aerodynamic performance C H,Ma,α,β,bas,i,s＝turb,δ The formula for calculating the aerodynamic performance of the aircraft in flight is as follows:

[0032] The advantages of the present invention compared with the prior art are:

[0033] (1) The present invention implements a discrete partial derivative method for each influencing factor of aerodynamic performance, thereby achieving a separate correction for each factor. Compared with the prior art method of using a transition model and a dynamic grid of an aerodynamic rudder to perform a mechanical-thermal coupling numerical calculation, the numerical calculation process is simplified while meeting the prediction accuracy.

[0034] (2) The comprehensive correction module of the present invention ensures the accuracy of aerodynamic performance prediction under flight conditions by compensating for the effects of rudder deflection, thermal deformation, and transition on aerodynamic performance. BRIEF DESCRIPTION OF THE DRAWINGS

[0035] Figure 1 This is a flow chart of the method for correcting the consistency between heaven and earth;

[0036] Figure 2 The following is a comparison chart of the results before and after correction and the flight test results. DETAILED DESCRIPTION

[0037] The present invention is described in detail below with reference to the accompanying drawings and specific embodiments.

[0038] like Figure 1 As shown, the method for correcting the sky-ground consistency of the aerodynamic force of a high-speed aircraft includes the following steps:

[0039] S1. Calculate the aerodynamic performance of the high-speed aircraft under different states under the flight state to be corrected and the incoming flow conditions, including the aerodynamic performance of the reference state, no rudder deviation, and laminar flow, the aerodynamic performance of the reference state, no rudder deviation, and turbulent flow, the aerodynamic performance of the reference state, with rudder deviation, and laminar flow, the aerodynamic performance of the reference state, with rudder deviation, and turbulent flow, the aerodynamic performance of thermal deformation, no rudder deviation, and laminar flow, the aerodynamic performance of the reference state, no rudder deviation, and transition, the aerodynamic performance of the reference state, maximum rudder deviation, and laminar flow, the aerodynamic performance of the reference state, minimum rudder deviation, and laminar flow, the aerodynamic performance of the reference state, maximum rudder deviation, and turbulent flow, and the aerodynamic performance of the reference state, minimum rudder deviation, and turbulent flow;

[0040] This step is calculated using CFD tools, and the specific calculation method is:

[0041] (1.1) For the aircraft shape in the reference state and without rudder deviation, according to the input flight state corresponding to the altitude, Mach number, angle of attack, sideslip angle and other incoming flow conditions, select "Laminar flow" in the "Viscous model option" of the CFD tool, and use the CFD tool to calculate the "reference state, no rudder deviation, laminar flow aerodynamic performance";

[0042] (1.2) For the aircraft shape in the reference state and without rudder deviation, according to the input flight state corresponding to the altitude, Mach number, angle of attack, sideslip angle and other incoming flow conditions, select "turbulence model" in the "viscous model option" of the CFD tool, and use the CFD tool to calculate the "reference state, no rudder deviation, turbulent aerodynamic performance";

[0043] (1.3) For the aircraft shape in the reference state and with rudder deflection, according to the input flight state corresponding to the altitude, Mach number, angle of attack, sideslip angle and other incoming flow conditions, select "Laminar flow" in the "Viscous model option" of the CFD tool, and use the CFD tool to calculate the "reference state, with rudder deflection, laminar flow aerodynamic performance";

[0044] (1.4) For the aircraft shape in the reference state and with rudder deflection, according to the input flight state corresponding to the altitude, Mach number, angle of attack, sideslip angle and other incoming flow conditions, select "turbulence model" in the "viscous model option" of the CFD tool, and use the CFD tool to calculate the "reference state, with rudder deflection, turbulent aerodynamic performance";

[0045] (1.5) For the aircraft shape with thermal deformation and no rudder deflection, according to the input flight state corresponding to the altitude, Mach number, angle of attack, sideslip angle and other incoming flow conditions, select "Laminar flow" in the "Viscous model option" of the CFD tool, and use the CFD tool to calculate the "thermal deformation, no rudder deflection, laminar flow aerodynamic performance";

[0046] (1.6) For the aircraft shape with thermal deformation and no rudder deflection, according to the input flight state corresponding to the altitude, Mach number, angle of attack, sideslip angle and other incoming flow conditions, select "turbulence model" in the "viscous model option" of the CFD tool, and use the CFD tool to calculate the "thermal deformation, no rudder deflection, turbulent aerodynamic performance";

[0047] (1.7) For the aircraft shape in the reference state and without rudder deflection, according to the input flight state corresponding to the altitude, Mach number, angle of attack, sideslip angle and other incoming flow conditions, select "transition model" in the "viscous model option" of the CFD tool, and use the CFD tool to calculate the "reference state, no rudder deflection, transition aerodynamic performance";

[0048] (1.8) For the aircraft shape in the reference state and maximum rudder deflection, according to the input flight state corresponding to the altitude, Mach number, angle of attack, sideslip angle and other incoming flow conditions, select "Laminar flow" in the "Viscous model option" of the CFD tool, and use the CFD tool to calculate the "reference state, maximum rudder deflection, laminar flow aerodynamic performance";

[0049] (1.9) For the aircraft shape in the reference state and minimum rudder deflection, according to the input flight state corresponding to the altitude, Mach number, angle of attack, sideslip angle and other incoming flow conditions, select "Laminar flow" in the "Viscous model option" of the CFD tool, and use the CFD tool to calculate the "reference state, minimum rudder deflection, laminar flow aerodynamic performance";

[0050] (1.10) For the aircraft shape in the reference state and maximum rudder deflection, according to the input flight state corresponding to the altitude, Mach number, angle of attack, sideslip angle and other incoming flow conditions, select "turbulence model" in the "viscous model option" of the CFD tool, and use the CFD tool to calculate the "reference state, maximum rudder deflection, turbulent aerodynamic performance";

[0051] (1.11) For the aircraft shape in the reference state and minimum rudder deflection, according to the input flight state corresponding to the altitude, Mach number, angle of attack, sideslip angle and other incoming flow conditions, select "Turbulence Model" in the "Viscous Model Option" of the CFD tool, and use the CFD tool to calculate the "reference state, minimum rudder deflection, turbulent aerodynamic performance".

[0052] S2. Calculate the aerodynamic performance deviation rate of the high-speed aircraft under different states under the flight state to be corrected and the incoming flow conditions, including the aerodynamic performance deviation rate of the rudder-biased aircraft in the reference state and laminar flow, the aerodynamic performance deviation rate of the rudder-biased aircraft in the reference state and turbulent flow, the aerodynamic performance deviation rate of the thermal deformation aircraft in the thermal deformation and laminar flow, the aerodynamic performance deviation rate of the thermal deformation aircraft in the thermal deformation and turbulent flow, and the aerodynamic performance deviation rate of the transitional aircraft in the reference state and transition;

[0053] Specific calculation method:

[0054] Using the “baseline state, aerodynamic performance without rudder deflection, laminar flow” and “baseline state, aerodynamic performance with rudder deflection, laminar flow”, calculate the “baseline state, laminar flow aerodynamic performance deviation rate of rudder deflection aircraft”;

[0055] Using the “baseline state, aerodynamic performance without rudder deflection and turbulence” and “baseline state, aerodynamic performance with rudder deflection and turbulence”, calculate the “baseline state, turbulence, rudder deflection aircraft aerodynamic performance deviation rate”;

[0056] The deviation rate of the aerodynamic performance of the aircraft with rudder deviation in the reference state and laminar flow and the deviation rate of the aerodynamic performance of the aircraft with rudder deviation in the reference state and turbulent flow are calculated using the following formula:

[0057]

[0058] Where H is the altitude in km, Ma is the dimensionless parameter Mach number, α is the angle of attack, β is the sideslip angle, i is the aerodynamic performance, i∈(x,y,z,mx,my,mz), x is the axial force, y is the normal force, z is the lateral force, mx is the rolling moment, my is the yaw moment, mz is the pitch moment, δ∈(δ1,δ2,…,δ n ),δi∈[δi min ,δi max ], δ is the rudder angle, and is selected from the preset rudder angle set (δ1, δ2, …, δ n ), δ i is the i-th element in the rudder angle set, δi min is the minimum rudder deflection of the aircraft, δi max is the maximum rudder deflection of the aircraft, bas represents the reference state, s is the incoming flow state, s∈(laminar, turbulence), when s=laminar represents laminar flow, s=turbulence represents turbulent flow.

[0059] Using the “baseline state, no rudder deviation, laminar flow aerodynamic performance” and “thermal deformation, no rudder deviation, laminar flow aerodynamic performance”, calculate the “thermal deformation, laminar flow thermal deformation aircraft aerodynamic performance deviation rate”;

[0060] Using the “baseline state, no rudder deviation, turbulent aerodynamic performance” and “thermal deformation, no rudder deviation, turbulent aerodynamic performance”, calculate the “thermal deformation, turbulent thermal deformation aircraft aerodynamic performance deviation rate”;

[0061] The aerodynamic performance deviation rate of thermal deformation aircraft with thermal deformation and laminar flow, and the aerodynamic performance deviation rate of thermal deformation aircraft with thermal deformation and turbulent flow are calculated using the following formula:

[0062]

[0063] Among them, H is the altitude in km, Ma is the dimensionless parameter Mach number, α is the angle of attack, β is the sideslip angle, i is the aerodynamic performance, hd represents thermal deformation, i∈(x,y,z,mx,my,mz), x is the axial force, y is the normal force, z is the lateral force, mx is the rolling moment, my is the yaw moment, mz is the pitch moment, s is the incoming flow state, s∈(laminar,turbulence), when s=laminar, it means laminar flow, s=turbulence, it means turbulent flow.

[0064] The “aerodynamic performance of the reference state, no rudder deviation, and turbulence” and the “aerodynamic performance of the reference state, no rudder deviation, and transition” are used to calculate the “aerodynamic performance deviation rate of the transitional aircraft from the reference state to the transition”.

[0065] Deviation rate of aerodynamic performance of transitional aircraft in reference state and transitional state;

[0066]

[0067] Among them, H is the altitude in km, Ma is the dimensionless parameter Mach number, α is the angle of attack, β is the sideslip angle, i is the aerodynamic performance, bas represents the reference state, i∈(x,y,z,mx,my,mz), x is the axial force, y is the normal force, z is the lateral force, mx is the rolling moment, my is the yaw moment, mz is the pitch moment, s is the incoming flow state, s∈(laminar,turbulence), when s=laminar represents laminar flow, s=turbulence represents turbulent flow.

[0068] S3, calculating the proportion of turbulence degree, thermal deformation degree and rudder deflection degree of the aircraft;

[0069] The turbulence degree accounts for γ H,Ma,α,β The calculation is done using the following formula:

[0070]

[0071] Among them, H is the altitude in km, Ma is the dimensionless parameter Mach number, α is the angle of attack, β is the sideslip angle, bas represents the reference state, i is the aerodynamic performance, i∈(x,y,z,mx,my,mz), x is the axial force, normal force, and lateral force, mx is roll, my is yaw, mz is pitch, s is the incoming flow state, s∈(laminar,turbulence), when s=laminar, it means laminar flow, and s=turbulence, it means turbulent flow.

[0072] It means to find the average value of each aerodynamic performance.

[0073] γ H,Ma,α,β = 0 is laminar flow state, γ H,Ma,α,β =1 is turbulent state,

[0074] Calculate the proportion of thermal deformation of the aircraft τ H,Ma,α,β The calculation formula is as follows:

[0075]

[0076] Among them, H is the altitude in km, Ma is the dimensionless parameter Mach number, α is the angle of attack, β is the sideslip angle, bas represents the reference state, hd represents, i is the aerodynamic performance, i∈(x,y,z,mx,my,mz), x is the axial force, normal force, lateral force, mx is roll, my is yaw, mz is pitch, s is the incoming flow state, s∈(laminar,turbulence), when s=laminar represents laminar flow, s=turbulence represents turbulent flow, δ is the rudder deflection angle, It means to find the average value of each aerodynamic performance. It is the aerodynamic performance in the state of maximum thermal deformation.

[0077] where τ H,Ma,α,β = 0 means no thermal deformation state, τ H,Ma,α,β =1 is the complete thermal deformation state, and the thermal deformation degree of the aircraft accounts for τ H,Ma,α,β

[0078] Calculate the proportion of the aircraft's rudder deviation σ H,Ma,α,β The calculation formula is as follows:

[0079]

[0080] Among them, H is the height in km, Ma is the dimensionless parameter Mach number, α is the angle of attack, β is the sideslip angle, i is the aerodynamic performance, i∈(x,y,z,mx,my,mz), x is the axial force, normal force, lateral force, mx is roll, my is yaw, mz is pitch, δ∈(δ1,δ2,…,δ n ),δi∈[δi min ,δi max ], δ is the rudder angle, and is selected from the preset rudder angle set (δ1, δ2, …, δ n ), δ i is the i-th element in the rudder angle set, δi min is the minimum rudder deflection of the aircraft, δi max is the maximum rudder deflection of the aircraft, bas represents the reference state, s is the incoming flow state, s∈(laminar, turbulence), when s=laminar represents laminar flow, s=turbulence represents turbulent flow.

[0081] C_max H,Ma,α,β,bas,i,s,δ is the maximum rudder deflection aerodynamic performance, C_max H,Ma,α,β,bas,i,s,δi =max(C H,Ma,α,β,bas,i,s,δi ,δi∈[δi min ,δi max ]);

[0082] C_min H,Ma,α,β,bas,i,s,δ The minimum rudder deflection aerodynamic performance: C_minH,Ma,α,β,bas,i,s,δi =min(C H,Ma,α,β,bas,i,s,δi ,δi∈[δi min ,δi max ]);

[0083] where σ H,Ma,α,β = 0 means no rudder bias state, σ H,Ma,α,β =1 means full rudder bias state.

[0084] S4, using the aerodynamic performance deviation rate of the aircraft in each rudder deviation state, the aerodynamic performance deviation rate of the aircraft in the thermal deformation state, and the aerodynamic performance deviation rate of the aircraft in the transition mode, as well as the aircraft turbulence degree ratio, thermal deformation degree ratio, and rudder deviation degree ratio, through interpolation fitting, calculate the rudder deviation correction amount, thermal deformation correction amount, and transition correction amount under the flight state to be corrected and the incoming flow condition;

[0085] The proportion of aircraft turbulence level γ H,Ma,α,β , and the aerodynamic performance deviation rate ΔC of the transitional aircraft under the corresponding (H, Ma, α, β) state H,Ma,α,β,bas,i,s , the number is n+1, calculate the Lagrange interpolation function:

[0086]

[0087] The transition correction can be calculated from this interpolation function and the turbulence degree ratio:

[0088]

[0089] The proportion of the aircraft rudder deviation σ H,Ma,α,β , corresponding to the deviation rate of aerodynamic performance of the rudder-biased aircraft under the (H, Ma, α, β) state ΔC H,Ma,α,β,bas,i,s＝turb,δ , the number is l+1, calculate the Lagrange interpolation function:

[0090]

[0091] The rudder deviation correction is calculated using this interpolation function and the rudder deviation ratio:

[0092]

[0093] The proportion of thermal deformation degree of aircraft τ H,Ma,α,β , and the aerodynamic performance deviation rate ΔC of the thermal deformation aircraft under the corresponding (H, Ma, α, β) state H,Ma,α,β,hd,i,s＝turb , the number is m+1, calculate the Lagrange interpolation function:

[0094]

[0095] The thermal deformation correction can be calculated from this interpolation function and the thermal deformation ratio:

[0096]

[0097] S5. Use rudder deviation correction, thermal deformation correction and transition correction to correct the aerodynamic performance of the aircraft obtained by ground simulation to make it consistent with the aerodynamic performance under actual flight conditions.

[0098] According to the rudder deviation correction Thermal deformation correction and transition correction And the reference aircraft aerodynamic performance C H,Ma,α,β,bas,i,s＝turb,δ The formula for calculating the aerodynamic performance of the aircraft in flight is as follows:

[0099] Example:

[0100] Taking a certain aircraft as an example, try to obtain the axial force, normal force coefficient and lift-to-drag ratio of the aircraft under a certain flight state U.

[0101] 1. Calculate the aerodynamic performance of the reference aircraft:

[0102] C U,bas,x,turb,δ＝0 =-0.0938, C U,bas,y,turb,δ＝0 =0.7146;

[0103] Rudder deviation aircraft aerodynamic performance deviation rate:

[0104] ΔC U,bas,x,lami,δ =1.26%, ΔC U,bas,x,turb,δ =0.78%;

[0105] ΔC U,bas,y,lami,δ =2.21%, ΔC U,bas,y,turb,δ =1.97%;

[0106] The aerodynamic performance deviation rate of the aircraft in the thermal deformation state is calculated:

[0107] ΔC U,hd,x,lami =10.29%, ΔC U,hd,x,turb =-12.66%;

[0108] ΔC U,hd,y,lami =8.03%, ΔC U,hd,y,turb =6.39%;

[0109] The aerodynamic performance deviation rate of the aircraft in the transition mode is calculated:

[0110] ΔC U,bas,x,lami =11.61%, ΔC U,bas,x,turb =0.76%;

[0111] ΔC U,bas,y,lami=-15.93%, ΔC U,bas,y,turb =-1.34%.

[0112] 2. Calculate the proportion of the aircraft's rudder deviation: σ U =0.0599;

[0113] Aircraft thermal deformation ratio: τ U =0.2086;

[0114] Aircraft turbulence degree ratio: γ U =0.63.

[0115] The rudder deviation correction value is obtained by fitting:

[0116]

[0117] Thermal deformation correction:

[0118]

[0119] Transition correction:

[0120]

[0121] 3. Calculate the aerodynamic performance of the aircraft in flight:

[0122] C 飞行_U,x =-0.0828, C 飞行_U,y =-0.7762;

[0123] In this flight state, the lift-to-drag ratio of the aircraft is approximately 4.81.

[0124] The flight test results are: C 飞试_U,x =-0.0816, C 飞试_U,y =-0.7707;

[0125] The lift-to-drag ratio in this state during the flight test was 4.83.

[0126] The present invention uses conventional aerodynamic and thermal performance parameter solution methods to obtain the aerodynamic performance of the aircraft in each state and the deviation rate of the aerodynamic performance of the aircraft. The correction amount is obtained by interpolation fitting and other methods to calculate the aerodynamic performance of the aircraft in the flight state. This method can significantly correct the difference between the earth and the sky in the aerodynamic performance of the aircraft, and achieve the consistency of the aerodynamic performance of the aircraft in the sky and the sky. It is a powerful tool for evaluating the aerodynamic performance of high-speed aircraft in flight. The present invention can better calculate the aerodynamic performance of the aircraft with consistency between the sky and the sky under the actual flight state, and at the same time provides a method for using ground data to predict the aerodynamic performance of the flight test aircraft, such as Figure 2 shown.

[0127] The present invention is not limited to the correction calculation of the sky-earth consistency of the aerodynamic performance of a high-speed aircraft in flight state, but can be extended to the calculation of aerodynamic performance of any aircraft with sky-earth difference or sky-earth consistency.

[0128] The above is a detailed introduction to the method for correcting the sky-to-ground consistency of the aerodynamic performance of a supersonic aircraft of the present invention. This article uses specific cases to illustrate the principles and implementation methods of the present invention. The above specific case description is only used to help understand the method and core idea of ​​the present invention. It should be pointed out that any simple modification made to the above examples without departing from the technical essence of the present invention belongs to the technical scope of the present invention.

[0129] The contents not described in detail in the specification of the present invention belong to the common knowledge of those skilled in the art.

Claims

1. A method for correcting the sky-ground consistency of aerodynamic forces of a high-speed aircraft, characterized in that The steps include: S1. Calculate the aerodynamic performance of the high-speed aircraft under different states under the flight state to be corrected and the incoming flow conditions, including the aerodynamic performance of the reference state, no rudder deviation, and laminar flow, the aerodynamic performance of the reference state, no rudder deviation, and turbulent flow, the aerodynamic performance of the reference state, with rudder deviation, and laminar flow, the aerodynamic performance of the reference state, with rudder deviation, and turbulent flow, the aerodynamic performance of thermal deformation, no rudder deviation, and laminar flow, the aerodynamic performance of thermal deformation, no rudder deviation, and turbulent flow, the aerodynamic performance of the reference state, no rudder deviation, and transition, the aerodynamic performance of the reference state, maximum rudder deviation, and laminar flow, the aerodynamic performance of the reference state, minimum rudder deviation, and laminar flow, the aerodynamic performance of the reference state, maximum rudder deviation, and turbulent flow, and the aerodynamic performance of the reference state, minimum rudder deviation, and turbulent flow; S2. Calculate the aerodynamic performance deviation rate of the high-speed aircraft under different states under the flight state to be corrected and the incoming flow conditions, including the aerodynamic performance deviation rate of the rudder-biased aircraft in the reference state and laminar flow, the aerodynamic performance deviation rate of the rudder-biased aircraft in the reference state and turbulent flow, the aerodynamic performance deviation rate of the thermal deformation aircraft in the thermal deformation and laminar flow, the aerodynamic performance deviation rate of the thermal deformation aircraft in the thermal deformation and turbulent flow, and the aerodynamic performance deviation rate of the transitional aircraft in the reference state and transition; S3, calculating the proportion of turbulence degree, thermal deformation degree and rudder deflection degree of the aircraft; S4, using the aerodynamic performance deviation rate of the aircraft in each rudder deviation state, the aerodynamic performance deviation rate of the aircraft in the thermal deformation state, and the aerodynamic performance deviation rate of the aircraft in the transition mode, as well as the aircraft turbulence degree ratio, thermal deformation degree ratio, and rudder deviation degree ratio, through interpolation fitting, calculate the rudder deviation correction amount, thermal deformation correction amount, and transition correction amount under the flight state to be corrected and the incoming flow condition; S5. Use rudder deviation correction, thermal deformation correction and transition correction to correct the aerodynamic performance of the aircraft obtained by ground simulation to make it consistent with the aerodynamic performance of the aircraft under actual flight conditions.

2. The method for correcting the sky-ground consistency of aerodynamic force of a high-speed aircraft according to claim 1, characterized in that: The step S1 is calculated using a CFD tool.

3. The method for correcting the sky-ground consistency of aerodynamic force of a high-speed aircraft according to claim 1, characterized in that: The deviation rate of the aerodynamic performance of the aircraft with rudder deviation in the reference state and laminar flow and the deviation rate of the aerodynamic performance of the aircraft with rudder deviation in the reference state and turbulent flow are calculated using the following formula: Where H is the altitude in km, Ma is the dimensionless parameter Mach number, α is the angle of attack, β is the sideslip angle, i is the aerodynamic performance term, i∈(x,y,z,mx,my,mz), x is the axial force, y is the normal force, z is the lateral force, mx is the rolling moment, my is the yaw moment, mz is the pitch moment, δ∈(δ1,δ2,…,δ n ),δi∈[δi min ,δi max ], δ is the rudder angle, and is selected from the preset rudder angle set (δ1, δ2, …, δ n ), δ i is the i-th element in the rudder angle set, δi min is the minimum rudder deflection of the aircraft, δi max is the maximum rudder deflection of the aircraft, bas represents the reference state, s is the incoming flow state, s∈(laminar, turbulence), when s=laminar represents laminar flow, s=turbulence represents turbulent flow.

4. The method for correcting the sky-ground consistency of the aerodynamic force of a supersonic aircraft according to claim 3, characterized in that: The aerodynamic performance deviation rate of thermal deformation aircraft for thermal deformation and laminar flow, and the aerodynamic performance deviation rate of thermal deformation aircraft for thermal deformation and turbulent flow are calculated using the following formula: Among them, H is the altitude in km, Ma is the dimensionless parameter Mach number, α is the angle of attack, β is the sideslip angle, i is the aerodynamic performance, hd represents thermal deformation, i∈(x,y,z,mx,my,mz), x is the axial force, normal force, and lateral force, mx is the rolling moment, my is the yaw moment, mz is the pitch moment, s is the incoming flow state, s∈(laminar,turbulence), when s=laminar, it means laminar flow, and s=turbulence, it means turbulent flow.

5. The method for correcting the sky-ground consistency of aerodynamic force of a supersonic aircraft according to claim 1, characterized in that: Deviation rate of aerodynamic performance of transitional aircraft in reference state and transition; Among them, H is the altitude in km, Ma is the dimensionless parameter Mach number, α is the angle of attack, β is the sideslip angle, i is the aerodynamic performance, bas represents the reference state, i∈(x,y,z,mx,my,mz), x is the axial force, normal force, and lateral force, mx is the rolling moment, my is the yaw moment, mz is the pitch moment, s is the incoming flow state, s∈(laminar,turbulence), when s=laminar represents laminar flow, s=turbulence represents turbulent flow.

6. The method for correcting the sky-ground consistency of the aerodynamic force of a supersonic aircraft according to claim 1, characterized in that: The turbulence degree accounts for γ H,Ma,α,β The calculation is done using the following formula: Wherein, H is the altitude in km, Ma is the dimensionless parameter Mach number, α is the angle of attack, β is the sideslip angle, bas represents the reference state, i is the aerodynamic performance term, i∈(x,y,z,mx,my,mz), x is the axial force, y is the normal force, z is the lateral force, mx is the rolling moment, my is the yaw moment, mz is the pitch moment, s is the incoming flow state, s∈(laminar,turbulence), when s=laminar represents laminar flow, s=turbulence represents turbulent flow; Represents the average value of all aerodynamic performance indicators.

7. The method for correcting the sky-ground consistency of the aerodynamic force of a supersonic aircraft according to claim 1, characterized in that: Calculate the proportion of thermal deformation of the aircraft τ H,Ma,α,β The calculation formula is as follows: Among them, H is the altitude in km, Ma is the dimensionless parameter Mach number, α is the angle of attack, β is the sideslip angle, bas represents the reference state, hd represents the thermal deformation, i represents the aerodynamic performance, i∈(x,y,z,mx,my,mz), x represents the axial force, y represents the normal force, z represents the lateral force, mx represents the rolling moment, my represents the yaw moment, mz represents the pitching moment, s represents the incoming flow state, s∈(laminar,turbulence), when s=laminar represents laminar flow, s=turbulence represents turbulent flow, δ represents the rudder deflection angle, It means to find the average value of each aerodynamic performance. It is the aerodynamic performance in the state of maximum thermal deformation.

8. The method for correcting the sky-ground consistency of aerodynamic force of a supersonic aircraft according to claim 1, characterized in that: Calculate the proportion of the aircraft's rudder deviation σ H,Ma,α,β The calculation formula is as follows: Among them, H is the height in km, Ma is the dimensionless parameter Mach number, α is the angle of attack, β is the sideslip angle, i is the aerodynamic performance, i∈(x,y,z,mx,my,mz), x is the axial force, normal force, lateral force, mx is roll, my is yaw, mz is pitch, δ∈(δ1,δ2,…,δ n ),δi∈[δi min ,δi max ], δ is the rudder angle, and is selected from the preset rudder angle set (δ1, δ2, …, δ n ), δ i is the i-th element in the rudder angle set, δi min is the minimum rudder deflection of the aircraft, δi max is the maximum rudder deflection of the aircraft, bas represents the reference state, s represents the incoming flow state, s∈(laminar, turbulence), when s=laminar represents laminar flow, s=turbulence represents turbulent flow; C_max H,Ma,α,β,bas,i,s,δ is the maximum rudder deflection aerodynamic performance, C_max H,Ma,α,β,bas,i,s,δi =max(C H,Ma,α,β,bas,i,s,δi ,δi∈[δi min ,δi max ]); C_min H,Ma,α,β,bas,i,s,δ The minimum rudder deflection aerodynamic performance: C_min H,Ma,α,β,bas,i,s,δi =min(C H,Ma,α,β,bas,i,s,δi ,δi∈[δi min ,δi max ]).

9. The method for correcting the sky-ground consistency of aerodynamic force of a supersonic aircraft according to claim 1, characterized in that: According to the rudder deviation correction Thermal deformation correction and transition correction And the reference aircraft aerodynamic performance C H,Ma,α,β,bas,i,s＝turb,δ The formula for calculating the aerodynamic performance of the aircraft in flight is as follows:

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