Statically unstable aircraft stall attack angle determination method and system

Abstract

The invention belongs to the field of aerodynamic design specialty and operation stability specialty design, and particularly relates to a statically unstable aircraft stall attack angle determination method and system.The method comprises the steps that the downward depression angle acceleration of a current aircraft is obtained, the pitching moment coefficient requirement is calculated through the downward depression angle acceleration, and the control surface skewness meeting the downward depression capacity requirement is obtained; the rolling steering effect of the control surface is obtained, and the skewness of the control surface meeting the rolling capacity requirement is obtained; and subtracting the control surface skewness meeting the pitching capability requirement and the control surface skewness meeting the rolling capability requirement from the maximum skewness of the control surface of the aircraft to obtain the longitudinal control efficiency corresponding to the remaining control surface skewness, and taking the longitudinal control efficiency as the balance stall attack angle. The method for determining the stall attack angle of the statically unstable airplane based on the balancing, pitching and rolling capacities of the airplane is provided, the stall attack angle of the airplane can be reasonably and effectively determined, and therefore input is provided for formulating the stall speed, the take-off and landing configuration characteristic speed, the stall envelope and the like of the airplane.

Description

Technical Field

[0001] The present application belongs to the field of aerodynamic design and control stability design, and particularly relates to a method and system for determining the stall angle of attack of a statically unstable aircraft. Background Art

[0002] The aircraft's stall angle of attack is closely related to flight safety, and directly determines the aircraft's stall speed, characteristic speed during takeoff and landing, and flight envelope, which in turn affects the aircraft's performance and quality parameters. Determining the aircraft's stall angle of attack is the primary issue that needs to be addressed in aircraft design.

[0003] Conventional aircraft are generally longitudinally statically stable. Near the stall angle of attack, the airflow separates, the lift coefficient reaches its maximum value and then begins to decrease. At the same time, the longitudinal static stability weakens, and the pitch moment curve begins to have an inflection point. Therefore, the angle of attack corresponding to the maximum value of the lift coefficient curve and the inflection point of the pitch moment curve is often used as the stall angle of attack.

[0004] For aircraft with unconventional layouts, generally speaking, the pitch moment nonlinearity occurs earlier than the lift nonlinearity. Especially for statically unstable aircraft, the pitch moment characteristics are statically unstable near the stall angle of attack, and the stall angle of attack of the aircraft cannot be determined by traditional methods. How to determine the stall angle of attack of statically unstable aircraft has become a key issue that needs to be studied urgently. Summary of the invention

[0005] The purpose of the present application is to provide a method and system for determining the stall angle of attack of a statically unstable aircraft, so as to solve the problem that the pitch moment characteristics of statically unstable aircraft are statically unstable near the stall angle of attack, and the stall angle of attack of the aircraft cannot be determined by traditional methods.

[0006] The technical solution of the present application is: a method for determining the stall angle of attack of a statically unstable aircraft, comprising: obtaining the downward pitch acceleration of the current aircraft, calculating the pitch moment coefficient requirement C through the downward pitch acceleration m_need , get the current control surface rudder effect C mδ ; Based on the obtained pitch moment coefficient requirement C m_need and control surface pitch rudder effect C mδ , the control surface deflection δ that meets the pitch-down capability requirement is obtained ea ; Determine the current required roll angular velocity p according to the aircraft's roll capability and requirements need , according to the rolling angular velocity p need Calculate the rolling moment coefficient C of the aircraft during steady-state rolling l_need ; Get the control surface roll rudder effect C lδ , according to the rolling moment coefficient C l_need and control surface roll rudder effect C lδ , the control surface deflection δ that meets the rolling capability requirements is obtained ea ;

[0007] Obtain the maximum deflection of the aircraft's control surface, subtract the deflection of the control surface that meets the pitch capability requirement and the deflection of the control surface that meets the rolling capability requirement from the maximum deflection of the aircraft's control surface, and obtain the longitudinal control effectiveness corresponding to the remaining control surface deflection as the equilibrium stall angle of attack α s .

[0008] Preferably, the pitch angle acceleration is:

[0009]

[0010] In the formula, C m is the pitch rudder effect; qbar is the velocity pressure; S is the wing reference area; C A is the average aerodynamic chord length; I y is the pitch moment of inertia.

[0011] Preferably, the pitch moment coefficient requirement C m_need for:

[0012]

[0013] Preferably, the currently required roll angular velocity p need For: p need =30 / 3.9 / 57.3.

[0014] Preferably, the rolling moment coefficient requirement C l_need for:

[0015]

[0016] In the formula, is the roll damping derivative, l is the span, and V is the true airspeed.

[0017] Preferably, the pitch angle acceleration of the aircraft at the stall angle of attack is 0.08 rad / s 2 .

[0018] As a specific implementation, it includes a first control surface deflection calculation module, a second control surface deflection calculation module and a balanced stall angle of attack determination module;

[0019] The first control surface deflection calculation module is used to obtain the current aircraft's pitch angle acceleration and calculate the pitch moment coefficient requirement C through the pitch angle acceleration. m_need , get the current control surface rudder effect C mδ ; Based on the obtained pitch moment coefficient requirement C m_need and control surface pitch rudder effect C mδ , the control surface deflection δ that meets the pitch-down capability requirement is obtained ea ;

[0020] The second control surface deflection calculation module is used to determine the current required roll angular velocity p according to the aircraft's roll capability. need , according to the rolling angular velocity p need Calculate the rolling moment coefficient C of the aircraft during steady-state rolling l_need ; Get the control surface roll rudder effect C lδ , according to the rolling moment coefficient C l_need and control surface roll rudder effect C lδ , the control surface deflection δ that meets the rolling capability requirements is obtained ea ;

[0021] The balanced stall angle of attack determination module is used to obtain the maximum deflection of the aircraft's control surface. After subtracting the control surface deflection that meets the pitch capability requirement and the control surface deflection that meets the rolling capability requirement from the maximum deflection of the aircraft's control surface, the longitudinal control effectiveness corresponding to the remaining control surface deflection is obtained as the balanced stall angle of attack α s .

[0022] Preferably, the pitch angle acceleration is:

[0023]

[0024] In the formula, C m is the pitch rudder effect; qbar is the velocity pressure; S is the wing reference area; c A is the average aerodynamic chord length; I y is the pitch moment of inertia.

[0025] Preferably, the pitch moment coefficient requirement C m_need for:

[0026]

[0027] Preferably, the currently required roll angular velocity p need For: p need =30 / 3.9 / 57.3.

[0028] Preferably, the rolling moment coefficient requirement C l_need for:

[0029]

[0030] In the formula, is the roll damping derivative, l is the span, and V is the true airspeed.

[0031] Preferably, the pitch angle acceleration of the aircraft at the stall angle of attack is 0.08 rad / s 2 .

[0032] The method and system for determining the stall angle of attack of a statically unstable aircraft of the present application propose a method for determining the stall angle of attack of a statically unstable aircraft based on the aircraft's trim, pitch down and roll capabilities, which can reasonably and effectively determine the stall angle of attack of the aircraft, thereby providing input for the formulation of the aircraft's stall speed, characteristic speed of the take-off and landing configuration, and stall envelope. BRIEF DESCRIPTION OF THE DRAWINGS

[0033] In order to more clearly illustrate the technical solution provided by the present application, the following is a brief introduction to the accompanying drawings. Obviously, the accompanying drawings described below are only some embodiments of the present application.

[0034] Figure 1 A schematic diagram for determining longitudinal aerodynamic characteristics and stall angle of attack for this application;

[0035] Figure 2 This is a schematic diagram of the longitudinal aerodynamic characteristics of a statically unstable aircraft in this application;

[0036] Figure 3 This is a schematic diagram of the overall process of this application;

[0037] Figure 4 This is a schematic diagram for determining the stall angle of attack of the aircraft in this application example. DETAILED DESCRIPTION

[0038] The following will be combined with the drawings in the embodiments of the present invention to clearly and completely describe the technical solutions in the embodiments of the present invention. Obviously, the described embodiments are only part of the embodiments of the present invention, not all of the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by ordinary technicians in this field without creative work are within the scope of protection of the present invention.

[0039] A method for determining the stall angle of attack of a statically unstable aircraft. The stall angle of attack of an aircraft is the boundary of safe flight of an aircraft. For conventional aircraft, near the stall angle of attack, due to airflow separation, the aircraft lift coefficient reaches its maximum value and begins to decrease, and the pitch moment becomes nonlinear. Therefore, it is easier to use the angle of attack corresponding to the maximum value of the lift coefficient curve and the inflection point of the pitch moment curve as the stall angle of attack, see Figure 1 .

[0040] For statically unstable aircraft, the aircraft is statically unstable near the stall angle of attack, and generally speaking, the torque nonlinearity occurs earlier than the angle of attack nonlinearity, see Figure 2 Therefore, the stall angle of attack of a statically unstable aircraft cannot be determined using the above method.

[0041] When recovering from a stall angle of attack, the aircraft, on the basis of trim, needs to have a certain pitch-down capability to make the aircraft lower its head, and at the same time, needs to have a certain rolling capability to suppress the asymmetric rolling caused by the asymmetric separation of airflow. Therefore, the stall angle of attack of the aircraft can be determined based on the above requirements.

[0042] The parameters required for this application are as follows:

[0043] Wing reference area S (m2), average aerodynamic chord length c A (m), span l(m), pitch moment of inertia I y (kg*m2), vacuum speed V(m / s), speed pressure qbar(N / m2), roll damping derivative Control surface pitch rudder effect C mδ and control surface roll rudder effect C lδ wait.

[0044] like Figure 3 , specifically including the following steps:

[0045] Step S100, obtain the current aircraft's pitch angle acceleration, and calculate the pitch moment coefficient requirement C through the pitch angle acceleration m_need , get the current control surface rudder effect C mδ ; Based on the obtained pitch moment coefficient requirement C m_need and control surface pitch rudder effect C mδ , the control surface deflection δ that meets the pitch-down capability requirement is obtained ea .

[0046] The flying quality standard GJB2874-97 "Flying quality of fly-by-wire control system aircraft" stipulates that when stalling, the pitch control efficiency should be able to produce no less than 0.08rad / s 2 Pitch angle acceleration capability, that is:

[0047]

[0048] To avoid over-design, determine the downward pitch acceleration of the aircraft at the stall angle of attack. 0.08 rad / s 2 .

[0049] The pitch angle acceleration can be calculated using the following formula:

[0050]

[0051] In the formula, C m It is the pitch rudder effect.

[0052] From formula (2), the pitch moment coefficient requirement C can be derived: m_need :

[0053]

[0054] Step S200, determine the current required roll angular velocity p according to the aircraft's roll capability and requirements need , according to the rolling angular velocity pneed Calculate the rolling moment coefficient C of the aircraft during steady-state rolling 1_need ; Get the control surface roll rudder effect C lδ , according to the rolling moment coefficient C 1_need and control surface roll rudder effect C lδ , the control surface deflection δ corresponding to the rolling capability requirement is obtained ea .

[0055] The secondary rolling capability specified in the flight quality specification GJB185-86 "Flight Quality of Manned Aircraft (Fixed Wing)" is a roll of 30° within 3.9 seconds.

[0056] According to the above requirements, determine the required roll angular velocity p need for:

[0057] p need =30 / 3.9 / 57.3 (4)

[0058] When the aircraft is in steady-state rolling, the rolling moment generated by the elevon and the rolling damping moment are balanced, and the rolling moment coefficient requirement C can be obtained. l_need :

[0059]

[0060] Step S300, determining the stall angle of attack α s

[0061] The flight quality standard GJB185-86 "Flight quality of manned aircraft (fixed wing)" stipulates: For any available altitude, from the stall speed V s To maximum speed V max In the level flight speed range between , the longitudinal control efficiency must be sufficient to enable the aircraft to reach any level flight speed within this speed range. Therefore, the aircraft needs to have longitudinal balance capability at the stall speed, that is, the stall angle of attack.

[0062] Obtain the maximum deflection of the aircraft's control surface, subtract the deflection of the control surface that meets the pitch capability requirement and the deflection of the control surface that meets the rolling capability requirement from the maximum deflection of the aircraft's control surface, and obtain the longitudinal control effectiveness corresponding to the remaining control surface deflection as the equilibrium stall angle of attack α s The maximum capacity of the nose-up moment at the stalling position can be used to determine the stalling angle α s .

[0063] In summary, the present application proposes a method for determining the stall angle of attack of a statically unstable aircraft based on the aircraft's trim, pitch and roll capabilities, which can reasonably and effectively determine the aircraft's stall angle of attack, thereby providing input for the formulation of the aircraft's stall speed, take-off and landing configuration characteristic speed, and stall envelope.

[0064] As a specific implementation, a statically unstable aircraft stall angle of attack determination system is also included. Based on the above design, the system includes a first control surface deflection calculation module, a second control surface deflection calculation module and a balanced stall angle of attack determination module.

[0065] The first control surface deflection calculation module is used to obtain the current aircraft's pitch angle acceleration and calculate the pitch moment coefficient requirement C through the pitch angle acceleration. m_need , get the current control surface rudder effect C mδ ; Based on the obtained pitch moment coefficient requirement C m_need and control surface pitch rudder effect C mδ , the control surface deflection δ that meets the pitch-down capability requirement is obtained ea ;

[0066] The second control surface deflection calculation module is used to determine the current required roll angular velocity p according to the aircraft's roll capability. need , according to the rolling angular velocity p need Calculate the rolling moment coefficient C of the aircraft during steady-state rolling l_need ; Get the control surface roll rudder effect C lδ , according to the rolling moment coefficient C l_need and control surface roll rudder effect C lδ , the control surface deflection δ that meets the rolling capability requirements is obtained ea ;

[0067] The balanced stall angle of attack determination module is used to obtain the maximum deflection of the aircraft's control surface. After subtracting the control surface deflection that meets the pitch capability requirement and the control surface deflection that meets the rolling capability requirement from the maximum deflection of the aircraft's control surface, the longitudinal control effectiveness corresponding to the remaining control surface deflection is obtained as the balanced stall angle of attack α s .

[0068] Preferably, the pitch angle acceleration is:

[0069]

[0070] In the formula, C m is the pitch rudder effect; qbar is the velocity pressure; S is the wing reference area; c A is the average aerodynamic chord length; I y is the pitch moment of inertia.

[0071] Preferably, the pitch moment coefficient requirement C m_need for:

[0072]

[0073] Preferably, the currently required roll angular velocity p need For: p need =30 / 3.9 / 57.3.

[0074] Preferably, the rolling moment coefficient requirement C l_need for:

[0075]

[0076] In the formula, is the roll damping derivative, l is the span, and V is the true airspeed.

[0077] Preferably, the pitch angle acceleration of the aircraft at the stall angle of attack is 0.08 rad / s 2 .

[0078] As a specific implementation method, the following is described with a specific example: Figure 2 It can be seen that the maximum lift coefficient of the aircraft corresponds to an angle of attack of 16.7°, while the pitch moment coefficient is statically unstable in the entire angle of attack range and becomes nonlinear after an angle of attack of 15°, making it impossible to directly determine the stall angle of attack of the aircraft.

[0079] The wing reference area of ​​the example aircraft is S = 485m 2 , mean aerodynamic chord length c A =13m, span l = 55(m), pitching moment of inertia I y =1900000kg*m 2 , vacuum speed V = 267m / s, speed pressure qbar = 3376N / m 2 , C mδ =-0.0016, C lδ =-0.0015.

[0080] In step S100, the pitch moment coefficient requirement C that satisfies the pitch capability requirement is obtained. m_need =-0.0071, corresponding control surface deflection δ ea =4.5°.

[0081] In step S200, the rolling moment coefficient requirement C that meets the secondary rolling capability requirement is obtained. 1_Need =-0.0153, corresponding control surface deflection δ ea =10.1°.

[0082] In step S300, after the deflection of the control surface that meets the pitch capability requirement and the deflection of the control surface that meets the rolling capability requirement are subtracted from the maximum deflection of the control surface of the aircraft, the remaining control surface deflection is 15.4°, and the corresponding pitch moment coefficient C m =-0.0244 (negative value represents pitch down), the maximum pitch moment coefficient C that can be balanced m is 0.0244, thus determining the stall angle of attack αs =12.5°, see Figure 3 .

[0083] Finally, it should be noted that: the drawings of the embodiments disclosed in the present invention only involve structures related to the embodiments disclosed in the present invention, and other structures can refer to the general design. In the absence of conflict, the same embodiment and different embodiments of the present invention can be combined with each other;

[0084] Finally: 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 in the protection scope of the present invention.

Claims

1. A method for determining the stall angle of attack of a statically unstable aircraft, characterized in that: include: Get the current aircraft's pitch angle acceleration, and calculate the pitch moment coefficient requirement C through the pitch angle acceleration m_need , get the current control surface rudder effect C mδ ; Based on the obtained pitch moment coefficient requirement C m_need and control surface pitch rudder effect C mδ , the control surface deflection δ that meets the pitch-down capability requirement is obtained ea ; Determine the current required roll angular velocity p according to the aircraft's roll capability and requirements need , according to the rolling angular velocity p need Calculate the rolling moment coefficient C of the aircraft during steady-state rolling l_need ; Get the rolling rudder effect of the control surface Clδ , according to the rolling moment coefficient C l_need and control surface roll rudder effect C lδ , the control surface deflection δ that meets the rolling capability requirements is obtained ea ; Obtain the maximum deflection of the aircraft's control surface, subtract the deflection of the control surface that meets the pitch capability requirement and the deflection of the control surface that meets the rolling capability requirement from the maximum deflection of the aircraft's control surface, and obtain the longitudinal control effectiveness corresponding to the remaining control surface deflection as the equilibrium stall angle of attack α s .

2. The method for determining the stall angle of attack of a statically unstable aircraft according to claim 1, characterized in that: The pitch angle acceleration is: In the formula, C m is the pitch rudder effect; qbar is the velocity pressure; S is the wing reference area; c A is the average aerodynamic chord length; I y is the pitch moment of inertia.

3. The method for determining the stall angle of attack of a statically unstable aircraft according to claim 2, characterized in that: Pitch moment coefficient requirement C m_need for:

4. The method for determining the stall angle of attack of a statically unstable aircraft according to claim 1, characterized in that: The current required roll angular velocity p need For: n eed =30 / 3.9 / 57.

3.

5. The method for determining the stall angle of attack of a statically unstable aircraft according to claim 4, characterized in that: Rolling moment coefficient requirement C 1_need for: In the formula, is the roll damping derivative, l is the span, and V is the true airspeed.

6. The method for determining the stall angle of attack of a statically unstable aircraft according to claim 1, characterized in that: The pitch-down acceleration of the aircraft at the stall angle of attack 0.08 rad / s 2 .

7. A system for determining the stall angle of attack of a statically unstable aircraft, using the method according to any one of claims 1 to 6, characterized in that: It includes a first control surface deflection calculation module, a second control surface deflection calculation module and a balanced stall angle of attack determination module; The first control surface deflection calculation module is used to obtain the current aircraft's pitch angle acceleration and calculate the pitch moment coefficient requirement C through the pitch angle acceleration. m_need , get the current control surface rudder effect C mδ ; Based on the obtained pitch moment coefficient requirement C m_need and control surface pitch rudder effect C mδ , the control surface deflection δ that meets the pitch-down capability requirement is obtained ea ; The second control surface deflection calculation module is used to determine the current required roll angular velocity p according to the aircraft's roll capability. need , according to the rolling angular velocity p need Calculate the rolling moment coefficient C of the aircraft during steady-state rolling l_need ; Get the control surface roll rudder effect C lδ , according to the rolling moment coefficient C l_need and control surface roll rudder effect C lδ , the control surface deflection δ that meets the rolling capability requirements is obtained ea ; The balanced stall angle of attack determination module is used to obtain the maximum deflection of the aircraft's control surface. After subtracting the control surface deflection that meets the pitch capability requirement and the control surface deflection that meets the rolling capability requirement from the maximum deflection of the aircraft's control surface, the longitudinal control effectiveness corresponding to the remaining control surface deflection is obtained as the balanced stall angle of attack α s .

8. The statically unstable aircraft stall angle of attack determination system according to claim 7, characterized in that: The pitch angle acceleration is: In the formula, C m is the pitch rudder effect; qbar is the velocity pressure; S is the wing reference area; c A is the average aerodynamic chord length; I y is the pitch moment of inertia.

9. The statically unstable aircraft stall angle of attack determination system according to claim 8, characterized in that: Pitch moment coefficient requirement C m_need for:

10. The statically unstable aircraft stall angle of attack determination system according to claim 7, characterized in that: The current required roll angular velocity p need For: p need =30 / 3.9 / 57.

3.

11. The statically unstable aircraft stall angle of attack determination system according to claim 10, characterized in that: Rolling moment coefficient requirement C l_need for: In the formula, is the roll damping derivative, l is the span, and V is the true airspeed.

12. The statically unstable aircraft stall angle of attack determination system according to claim 7, characterized in that: The pitch-down acceleration of the aircraft at the stall angle of attack 0.08 rad / s 2 .

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