A yaw rate over air minimum control speed influence correction method

Abstract

The application belongs to the field of aircraft flight mechanics design, and is a yaw angle rate influence correction method for the minimum control speed in the air. The slope angle of the actual state of the aircraft is first inversely deduced according to the slope angle required by the balanced yaw angle speed and the current slope angle of the flight test. Then the new slope angle is linearly fitted through the lateral component of the lift coefficient corresponding to different flight states of the aircraft, so as to correct the test curve. Finally, the minimum control speed in the air under a given flight state is inversely deduced through the corrected test curve. Finally, the difference between the minimum control speed in the air obtained by correction and the test minimum control speed in the air is obtained. The method is simple in analysis and use. Meanwhile, the method provides an accurate, effective and simple analysis means for the determination of the minimum safe take-off speed of the aircraft, and provides theoretical support and technical support for the safe flight in the engine failure state.

Description

Technical Field

[0001] This application belongs to the field of aircraft flight mechanics design, and specifically relates to a method for correcting the influence of yaw rate on minimum control speed in the air. Background Technology

[0002] Minimum airborne control speed is a core indicator in aircraft boundary flight dynamics design. It represents the minimum safe speed required to maintain directional control under a specific flight condition (single engine failure, full rudder deflection, fixed heading). Considering the difficulty in maintaining a steady heading during minimum airborne control speed testing, the aircraft may experience slight left or right yaws in real-time, inducing random yaw rates and causing deviations in the static minimum airborne control speed. Significant deviations can lead to heading divergence, resulting in loss of directional control. Accurately determining the minimum airborne control speed is crucial for ensuring flight safety under single engine failure conditions; however, currently, there is a lack of a rapid, reasonable, and simple method to express, characterize, and analyze the impact of yaw rates on the minimum airborne control speed. Summary of the Invention

[0003] The purpose of this application is to provide a method for correcting the influence of yaw rate on minimum control speed in the air, so as to solve the problems of complex calculation and poor calculation accuracy of minimum control speed in the air in the prior art.

[0004] The technical solution of this application is: a method for correcting the influence of yaw rate on minimum control speed in the air, comprising:

[0005] Given the current flight conditions, calculate the bank angle Δφ required to determine the equilibrium yaw rate;

[0006] Obtain the current flight pressure and calculate the lift coefficient C under the current flight conditions. L0 ;

[0007] Conduct aircraft test flights, and based on the current test flight results, including C L0 sinφ0 and C N0 Calculate the current slope angle φ0;

[0008] The actual bank angle of the aircraft is calculated based on the current bank angle φ0 and the bank angle Δφ required for the balance yaw rate: φ = φ0 + Δφ.

[0009] Determine the relationship between the lateral component of the aircraft's actual flight state and its actual lift coefficient, and calculate the lateral component C of the lift coefficient in the current actual flight state. Y =C L0 sinφ;

[0010] The lateral component of the lift coefficient corresponding to different flight states is calculated based on the relationship between the lateral component of the actual flight state and the bank angle of the actual flight state.

[0011] Linear fitting of data points was performed on the lateral component of the lift coefficient under different flight conditions to determine the curve C of the rear lateral force coefficient and the symmetrical yaw moment coefficient obtained from the aircraft test flight under the test flight conditions. L0 sinφ0~C N0 The curve C showing the rear lateral force coefficient under the current flight condition and the symmetrical yaw moment coefficient obtained from the aircraft test flight. L0 sinφ~C N0 ;

[0012] According to C L0 sinφ0~C N0 Curve and C L0 sinφ~C N0 The curve calculates the maximum rear lateral force coefficient C under standard atmospheric conditions, 0 altitude, and different flight speeds at the most severe bank angle in the current flight state. L *sin|φ| max The symmetrical yaw moment coefficient C under the current flight condition N ;

[0013] According to C L *sin|φ| max C in different states N The correspondence between them yields the curve C of the maximum rear lateral force coefficient and the symmetrical yaw moment coefficient under the current flight state. L *sin|φ| max ~C N ;

[0014] Analysis determines C L0 sinφ0~C N0 Curve, C L0 sinφ~C N0 Curve and C L *sin|φ| max ~C N Find the intersection points between the curves and identify two of them, one of which is the intersection point A0(x0,y0) obtained from the experimental curve, and the other intersection point A1(x1,y1) obtained from the modified curve;

[0015] Calculate the influence of yaw rate on minimum control speed in the air based on the coordinates of the two intersection points.

[0016] Preferably, the formula for calculating the slope angle Δφ required to balance the yaw rate is:

[0017] g = 9.8 * 6356766 2 / (6356766+H) 2

[0018] △φ=-V / 3.6*△r / g

[0019] In the formula, g is the acceleration due to gravity, H is the flight altitude, V is the flight speed, and Δr is the yaw rate.

[0020] Preferably, the lift coefficient C in the current flight state L0 The calculation formula is:

[0021] ρ=P / (287.05*(288.15-0.0065*H))

[0022] C L0 =2*G*g / (ρ*V 2 / 3.6 2 *S)

[0023] In the formula, ρ is the atmospheric density, P is the atmospheric pressure, G is the aircraft weight, H is the flight altitude, g is the gravitational acceleration, V is the flight speed, and S is the wing area.

[0024] Preferably, the formula for calculating the current slope angle φ0 is:

[0025] SN=C L0 sinφ0 / C L0

[0026] φ0=asin(SN)

[0027] In the formula, a is a constant, and SN is the sine value of the slope angle.

[0028] Preferably, the symmetrical yaw moment coefficient C under the current flight state N The calculation formula is:

[0029] C N =△T*L E / (0.5*ρ*V 2 ) / S / b A

[0030] In the formula, S is the wing area, b A Where ρ is the wingspan, ρ is the atmospheric density, V is the flight speed, ΔT is the engine asymmetric thrust, and L is the air velocity. E This is the lateral distance from the engine centerline to the plane of symmetry of the fuselage.

[0031] Preferably, the formula for determining the coordinates of the two intersection points is:

[0032] x0=(C L sinφ)0

[0033] x1=(C L sinφ)1

[0034] y0=(C N )0

[0035] y1=(C N )1

[0036] In the formula, the subscripts outside the parentheses represent the rear lateral force coefficient value and the symmetrical yaw moment coefficient value obtained from the test flight of the aircraft, respectively, under the flight state corresponding to the intersection point.

[0037] Preferably, the formula for calculating the influence of the yaw rate on the minimum control speed in the air, ΔV, is as follows:

[0038]

[0039] In the formula, ρ is the atmospheric density, G is the weight of the aircraft, g is the gravitational acceleration, and S is the wing area.

[0040] The method for correcting the influence of yaw rate on minimum control speed in this application first derives the actual bank angle of the aircraft based on the bank angle required to balance the yaw rate and the current bank angle during the test flight. Then, it linearly fits the new bank angle to the data points using the lateral component of the lift coefficient corresponding to different flight states, thereby correcting the test curve. Finally, it calculates the minimum control speed in the air under a given flight state by using the corrected test curve. Finally, it compares the corrected minimum control speed in the air with the test minimum control speed in the air to obtain the difference between the two. This method is clear, simple to analyze, and easy to use. At the same time, it provides an accurate, effective, and simple analytical means for determining the minimum safe takeoff speed of the aircraft, and provides theoretical support and technical guarantee for ensuring safe flight under engine failure conditions. Attached Figure Description

[0041] To more clearly illustrate the technical solutions provided in this application, the accompanying drawings will be briefly described below. Obviously, the drawings described below are merely some embodiments of this application.

[0042] Figure 1 This is a schematic diagram of the overall process of this application;

[0043] Figure 2 For this application C L0 sinφ~C N0 Schematic diagram of the curve;

[0044] Figure 3 For this application C L sinφ~C N Schematic diagram of the curve;

[0045] Figure 4 For this application CL sinφ~C N Schematic diagram of the curve. Detailed Implementation

[0046] The technical solutions of the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are only some embodiments of the present invention, and not all embodiments. Based on the embodiments of the present invention, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of the present invention.

[0047] A method for correcting the influence of yaw rate on minimum control speed in the air, where the aircraft's balance bank angle Δφ (°) and the lateral component of the aircraft's lift coefficient C are considered. Y Minimum air control speed correction ΔV (km / h).

[0048] like Figure 1 As shown, it includes the following steps:

[0049] Step S100: Given the current flight state (G, H, V, Δr), calculate the bank angle Δφ required to determine the balance yaw rate.

[0050] The formula for calculating the bank angle Δφ required to balance the yaw rate is:

[0051] g = 9.8 * 6356766 2 / (6356766+H) 2

[0052] △φ=-V / 3.6*△r / g

[0053] In the formula, g is the acceleration due to gravity, H is the flight altitude, V is the flight speed, and Δr is the yaw rate.

[0054] Step S200: Obtain the current flight pressure and calculate the lift coefficient C under the current flight condition. L0 .

[0055] Lift coefficient C under current flight conditions L0 The calculation formula is:

[0056] ρ=P / (287.05*(288.15-0.0065*H))

[0057] C L0 =2*G*g / (ρ*V 2 / 3.6 2 *S)

[0058] In the formula, ρ is the atmospheric density, P is the atmospheric pressure, G is the aircraft weight, V is the flight speed, and S is the wing area.

[0059] Step S300: Conduct an aircraft test flight, based on the current test flight results (C) L0 sinφ0,C N0 ), including C L0 sinφ0 and C N0 Calculate the current slope angle φ0.

[0060] The formula for calculating the current slope angle φ0 is:

[0061] SN=C L0 sinφ0 / C L0

[0062] φ0=asin(SN)

[0063] In the formula, a is a constant, and SN is the sine value of the slope angle.

[0064] Since the bank angle of the aircraft in its actual state cannot be directly obtained, this application compares and analyzes the bank angle required for the balance yaw rate in the actual flight state with the current bank angle in the test flight state in order to accurately obtain the bank angle of the aircraft in its actual state.

[0065] Step S400: Calculate the actual bank angle of the aircraft, φ = φ0 + Δφ, based on the current bank angle φ0 and the bank angle Δφ required to balance the yaw rate.

[0066] The above steps deduce the equilibrium bank angle from the yaw rate and the current test flight speed. The equilibrium bank angle and the test bank angle are then superimposed to calculate the actual aircraft bank angle in real time.

[0067] Step S500: Determine the relationship between the lateral component of the actual flight state and the lift coefficient of the actual aircraft state, and calculate the lateral component C of the lift coefficient of the current actual aircraft state. Y =C L0 sinφ.

[0068] Step S600: Calculate the aircraft's different flight states (G) based on the relationship between the lateral component of the actual flight state and the bank angle of the actual aircraft state. i H i V i ,△r i The lateral component of the lift coefficient C corresponding to i = 1, 2, ... n Yi .

[0069] Step S700: Perform linear fitting of data points for the lateral component of the lift coefficient corresponding to different flight states to determine the curve C of the rear lateral force coefficient and the symmetrical yaw moment coefficient obtained from the aircraft test flight under the test flight state. L0 sinφ0~CN0 The curve C showing the rear lateral force coefficient under the current flight condition and the symmetrical yaw moment coefficient obtained from the aircraft test flight. L0 sinφ~C N0 By linearly fitting the data points, a new corrected test curve that meets the requirements of the flight test state is generated.

[0070] Step S800, according to C L0 sinφ0~C N0 Curve and C L0 sinφ~C N0 Curve calculations are performed under standard atmospheric conditions, at 0 altitude, and at different flight speeds at the most severe bank angle (|φ|). max Maximum rear lateral force coefficient C under the current flight state at ≤5°) L *sin|φ| max The symmetrical yaw moment coefficient C under the current flight condition N .

[0071] C L *sin|φ| max The calculation method is described in step S200, where the symmetrical yaw moment coefficient C is under the current flight condition. N The calculation formula is:

[0072] C N =△T*L E / (0.5*ρ*V 2 ) / S / b A

[0073] In the formula, S is the wing area, b A Where ρ is the wingspan, ρ is the atmospheric density, V is the flight speed, ΔT is the engine asymmetric thrust, and L is the air velocity. E This is the lateral distance from the engine centerline to the plane of symmetry of the fuselage.

[0074] Step S900, according to C L *sin|φ| max C in different states N The correspondence between them yields the curve C of the maximum rear lateral force coefficient and the symmetrical yaw moment coefficient under the current flight state. L *sin|φ| max ~C N .

[0075] Step S1000, analyze and determine C L0 sinφ0~C N0 Curve, C L0 sinφ~C N0 Curve and C L *sin|φ| max ~CN Find the intersection points between the curves and identify two of them, one of which is the intersection point A0(x0,y0) obtained from the experimental curve, and the other intersection point A1(x1,y1) obtained from the modified curve.

[0076] The formula for determining the coordinates of two intersection points is:

[0077] x0=(C L sinφ)0

[0078] x1=(C L sinφ)1

[0079] y0=(C N )0

[0080] y1=(C N )1

[0081] In the formula, the subscripts outside the parentheses represent the rear lateral force coefficient value and the symmetrical yaw moment coefficient value obtained from the test flight of the aircraft, respectively, under the flight state corresponding to the intersection point.

[0082] Step S1100: Calculate the influence of yaw angular velocity on minimum control speed in the air based on the coordinates of the two intersection points.

[0083] The formula for calculating the effect of yaw rate on minimum control speed in the air, ΔV, is as follows:

[0084]

[0085] In the formula, ρ is the atmospheric density, G is the weight of the aircraft, g is the gravitational acceleration, and S is the wing area.

[0086] This application first derives the actual bank angle of the aircraft based on the bank angle required to balance the yaw rate and the current bank angle during the test flight. Then, it linearly fits the new bank angle to the data points using the lateral component of the lift coefficient corresponding to different flight states, thereby correcting the test curve. Finally, it calculates the minimum control speed in the air under a given flight state by using the corrected test curve. Finally, it compares the corrected minimum control speed in the air with the test minimum control speed in the air to obtain the difference between the two. This method is clear, simple to analyze, and easy to use. At the same time, it provides an accurate, effective, and simple analytical means for determining the minimum safe takeoff speed of the aircraft, and provides theoretical support and technical guarantee for ensuring safe flight under engine failure conditions.

[0087] This method can effectively correct the influence of random yaw rate during flight testing. Based on the correction results, a more accurate minimum control speed in the air can be obtained, ensuring safe and controllable flight of the aircraft.

[0088] As a specific implementation method, the following example illustrates the process:

[0089] Given standard atmospheric conditions, a certain type of aircraft weighs 120,000 kg; flight altitude H = 1500 m; flight speed 220 km / h; the asymmetric thrust ΔT of the engine is interpolated with respect to speed, temperature, and altitude; the lateral distance of the failed engine from the aircraft's center of gravity is 10 m; the above conditions are only applicable to steps S100 to S500).

[0090] Step S100: Given the yaw rate Δr = -0.2°, calculate the bank angle required to balance the yaw rate;

[0091] g = 9.8 * 6356766 2 / (6356766+1500) 2 =9.795

[0092] △φ=-V / 3.6*△r / g=-(220 / 3.6)*-0.2 / 9.795=1.25°

[0093] Step S200: Calculate the lift coefficient (C) under the current flight condition. L0 )

[0094] P=101325*((288.15-0.0065*1500) / 288.15)(-0.034 / -0.0065)

[0095] =84629.1

[0096] ρ=84629.1 / (287.05*(288.15-0.0065*1500))=1.059

[0097] C L0 =2*G*g / (ρ*V 2 / 3.6 2 *S)

[0098] =2*120000*9.795 / (1.059*(220 / 3.6)^2*288) = 2.064

[0099] Step S300, the current test flight results of the aircraft are known (C L0 sinφ0,C N0 ), Calculate the current slope angle (φ0):

[0100] C L0 sinφ0=0.07

[0101] C N0 =0.03

[0102] SN=CL0 sinφ0 / C L0 =0.07 / 2.064=0.0339

[0103] φ0=asin(SN)=asin(0.0339)*57.3=1.94°

[0104] Step S400: Calculate the actual bank angle of the current flight state.

[0105] φ=φ0+△φ=1.94°+1.25°=3.19°

[0106] Step S500: Calculate the lateral component of the actual lift coefficient in the current flight state.

[0107] △C Y =C L0 sinφ=2.064*sin(3.19 / 57.3)=0.115

[0108] Step S600: Calculate the different flight state points (G) according to steps (1) to (5) above. i H i V i ,△r i ) i＝1,2,…n C Yi The calculation state points are shown in Table 1; the calculation results are shown in Table 2.

[0109] Table 1 Calculation of state points

[0110]

[0111]

[0112]

[0113] Table 2 C Y(i＝1,2,…6) Calculation results

[0114] Step S700: Use linear fitting of data points to determine C. L0 sinφ0~C N0 Curve and C L0 sinφ~C N0 Curve, results see Figure 2 , Figure 2 The Δr in the middle ranges from -0.3° to 0.3° from left to right, with a step size of 0.1°.

[0115] Step S800: Calculate the flight speeds under standard atmospheric conditions, at 0 altitude, and at the most severe bank angle (|φ|). max C at ≤5° L *sin|φ| maxand C N The calculation results are shown in Table 3;

[0116] Table 3 Calculation Results

[0117]

[0118]

[0119] Step S900, determine C L *sin|φ| max ~C N Curve, see curve Figure 3 ;

[0120] Step S1000: Determine the two intersection points A0(x0,y0) and A1(x1,y1). Here, we take Δr = -0.2° as an example for calculation. The results are shown below. Figure 4 ;

[0121] The intersection points A0 and A1 are obtained by solving the system of two linear equations in two variables.

[0122] That is, A1 = (-0.189, 0.044)

[0123] That is, A0 = (-0.208, 0.049)

[0124] Step S1100: Determine the influence of yaw rate on minimum control speed in the air, ΔV.

[0125]

[0126] Finally, the following points should be noted: First, in the description of this application, it should be noted that, unless otherwise specified and limited, the terms "installation", "connection", and "linkage" should be interpreted broadly, and can be mechanical or electrical connections, or internal connections between two components, or direct connections. "Up", "down", "left", "right", etc. are only used to indicate relative positional relationships. When the absolute position of the described object changes, the relative positional relationship may change.

[0127] Secondly: The accompanying drawings of the embodiments disclosed in this invention only involve the structures involved in the embodiments disclosed in this invention. Other structures can refer to the general design. In the absence of conflict, the same embodiment and different embodiments of this invention can be combined with each other.

[0128] In conclusion, the above description is only a preferred embodiment of the present invention and is not intended to limit the present invention. Any modifications, equivalent substitutions, improvements, etc., made within the spirit and principles of the present invention should be included within the protection scope of the present invention.

Claims

1. A method for correcting the influence of yaw rate on minimum control speed in the air, characterized in that, include: Given the current flight conditions, calculate the bank angle Δφ required to determine the equilibrium yaw rate; Obtain the current flight pressure and calculate the lift coefficient C under the current flight conditions. L0 ; Conduct aircraft test flights, and based on the current test flight results, including C L0 sinφ0 and C N0 Calculate the current slope angle φ0; The actual bank angle of the aircraft is calculated based on the current bank angle φ0 and the bank angle Δφ required for the balance yaw rate: φ = φ0 + Δφ. Determine the relationship between the lateral component of the aircraft's actual flight state and its actual lift coefficient, and calculate the lateral component C of the lift coefficient in the current actual flight state. Y =C L0 sinφ; The lateral component of the lift coefficient corresponding to different flight states is calculated based on the relationship between the lateral component of the actual flight state and the bank angle of the actual flight state. Linear fitting of data points was performed on the lateral component of the lift coefficient under different flight conditions to determine the curve C of the rear lateral force coefficient and the symmetrical yaw moment coefficient obtained from the aircraft test flight under the test flight conditions. L0 sinφ0~C N0 The curve C showing the rear lateral force coefficient under the current flight condition and the symmetrical yaw moment coefficient obtained from the aircraft test flight. L0 sinφ~C N0 ; According to C L0 sinφ0~C N0 Curve and C L0 sinφ~C N0 The curve calculates the maximum rear lateral force coefficient C under standard atmospheric conditions, 0 altitude, and different flight speeds at the most severe bank angle in the current flight state. L *sin|φ| max The symmetrical yaw moment coefficient C under the current flight condition N ; According to C L *sin|φ| max C in different states N The correspondence between them yields the curve C of the maximum rear lateral force coefficient and the symmetrical yaw moment coefficient under the current flight state. L *sin|φ| max ~C N ; Analysis determines C L0 sinφ0~C N0 Curve, C L0 sinφ~C N0 Curve and C L *sin|φ| max ~C N Find the intersection points between the curves and identify two of them, one of which is the intersection point A0(x0,y0) obtained from the experimental curve, and the other intersection point A1(x1,y1) obtained from the modified curve; Calculate the influence of yaw rate on minimum control speed in the air based on the coordinates of the two intersection points.

2. The method for correcting the influence of yaw rate on minimum control speed in air as described in claim 1, characterized in that, The formula for calculating the slope angle Δφ required to balance the yaw rate is: g=9.8*6356766 2 / (6356766+H) 2 △φ=-V / 3.6*△r / g In the formula, g is the acceleration due to gravity, H is the flight altitude, V is the flight speed, and Δr is the yaw rate.

3. The method for correcting the influence of yaw rate on minimum control speed in air as described in claim 1, characterized in that, The lift coefficient C under the current flight condition L0 The calculation formula is: ρ=P / (287.05*(288.15-0.0065*H)) C L0 =2*G*g / (ρ*V 2 / 3.6 2 *S) In the formula, ρ is the atmospheric density, P is the atmospheric pressure, G is the aircraft weight, H is the flight altitude, g is the gravitational acceleration, V is the flight speed, and S is the wing area.

4. The method for correcting the influence of yaw rate on minimum control speed in air as described in claim 1, characterized in that, The formula for calculating the current slope angle φ0 is: SN=C L0 sinφ0 / C L0 φ0=asin(SN) In the formula, a is a constant, and SN is the sine value of the slope angle.

5. The method for correcting the influence of yaw rate on minimum control speed in air as described in claim 1, characterized in that, The symmetrical yaw moment coefficient C under the current flight state N The calculation formula is: C N =△T*L E / (0.5*ρ*V 2 ) / S / b A In the formula, S is the wing area, b A Where ρ is the wingspan, ρ is the atmospheric density, V is the flight speed, ΔT is the engine asymmetric thrust, and L is the air velocity. E This is the lateral distance from the engine centerline to the plane of symmetry of the fuselage.

6. The method for correcting the influence of yaw rate on minimum control speed in air as described in claim 1, characterized in that, The formula for determining the coordinates of the two intersection points is: x0=(C L sinφ)0 x1=(C L sinφ)1 y0=(C N )0 y1=(C N )1 In the formula, the subscripts outside the parentheses represent the rear lateral force coefficient value and the symmetrical yaw moment coefficient value obtained from the test flight of the aircraft, respectively, under the flight state corresponding to the intersection point.

7. The method for correcting the influence of yaw rate on minimum control speed in air as described in claim 1, characterized in that, The formula for calculating the influence of the yaw rate on the minimum control speed in the air, ΔV, is as follows: In the formula, ρ is the atmospheric density, G is the weight of the aircraft, g is the gravitational acceleration, and S is the wing area.

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