A method and system for determining a stall angle of attack of a statically unstable aircraft

Abstract

The application belongs to the field of aerodynamic design and stability and control design, and particularly relates to a static unstable aircraft stall angle of attack determination method and system, obtaining a current aircraft pitch-down angle acceleration, calculating a pitch moment coefficient requirement through the pitch-down angle acceleration, obtaining a control surface deflection meeting the pitch-down capability requirement, obtaining a control surface roll rudder efficiency, obtaining a control surface deflection meeting the roll capability requirement, subtracting the control surface deflection meeting the pitch-down capability requirement and the control surface deflection meeting the roll capability requirement from the maximum control surface deflection of the aircraft, obtaining a longitudinal control efficiency corresponding to the remaining control surface deflection as the balanced stall angle of attack. The static unstable aircraft stall angle of attack determination method based on aircraft trim, pitch-down and roll capability is proposed, which can reasonably and effectively determine the stall angle of attack of the aircraft, thereby providing input for formulating the aircraft stall speed, take-off and landing configuration characteristic speed and stall envelope.

Description

Technical Field

[0001] The present application belongs to the fields 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] An aircraft's stall angle of attack is closely linked to flight safety. It directly determines the aircraft's stall speed, characteristic speed during takeoff and landing, and flight envelope, ultimately impacting the aircraft's performance and quality parameters. Determining the aircraft's stall angle of attack is a primary issue in aircraft design.

[0003] Conventional aircraft are generally longitudinally statically stable. Near the stall angle of attack, airflow separates, the lift coefficient reaches its maximum value, and then begins to decrease. Simultaneously, longitudinal static stability weakens, and the pitching moment curve begins to show an inflection point. Therefore, the stall angle of attack is often determined by the angle corresponding to the maximum lift coefficient curve and the inflection point in the pitching moment curve.

[0004] For aircraft with unconventional layouts, pitching moment nonlinearity generally develops before lift nonlinearity. This is especially true for statically unstable aircraft, where the pitching moment characteristics exhibit static instability near the stall angle of attack, making it impossible to determine the stall angle of attack using traditional methods. Determining the stall angle of attack for statically unstable aircraft has become a key issue requiring urgent research. Summary of the Invention

[0005] The purpose of this 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 using 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 ; According to 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 based on the aircraft's roll capability and requirements need , according to the roll 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 roll 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 α s .

[0008] Preferably, the pitch angle acceleration is:

[0009]

[0010] Where 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 pitching 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] Where, is the roll damping derivative, l is the span, and V is the true airspeed.

[0017] Preferably, the pitch-down 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 downward pitch acceleration and calculate the pitch moment coefficient requirement C through the downward pitch acceleration. m_need , get the current control surface rudder effect C mδ ; According to 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 roll 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 roll 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] Where 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 pitching 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] Where, is the roll damping derivative, l is the span, and V is the true airspeed.

[0031] Preferably, the pitch-down 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 in this application propose a method for determining the stall angle of attack of a statically unstable aircraft based on the aircraft's balancing, pitching, and rolling capabilities. The method can reasonably and effectively determine the aircraft's stall angle of attack, thereby providing input for the formulation of the aircraft's stall speed, takeoff and landing configuration characteristic speed, and stall envelope. BRIEF DESCRIPTION OF THE DRAWINGS

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

[0034] Figure 1 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 clearly and completely describe the technical solutions in the embodiments of the present invention in conjunction with the accompanying drawings. 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 making creative efforts 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 is the boundary of safe flight. For conventional aircraft, near the stall angle of attack, due to airflow separation, the aircraft's lift coefficient reaches its maximum value and begins to decrease, and the pitching moment becomes nonlinear. Therefore, it is easy to use the angle of attack corresponding to the maximum value of the lift coefficient curve and the inflection point of the pitching 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, while trimmed, must possess a certain amount of pitch-down capability to keep the aircraft nose down, and a certain amount of roll capability to suppress the asymmetric roll caused by asymmetric airflow separation. Therefore, the aircraft's stall angle of attack can be determined based on these 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), pitching moment of inertia I y (kg*m2), true air speed V (m / s), velocity pressure qbar (N / m2), roll damping derivative Control surface pitch 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 required pitch moment coefficient C based on the pitch angle acceleration. m_need , get the current control surface rudder effect C mδ ; According to 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 flight quality specification GJB2874-97 "Flying quality of fly-by-wire aircraft" stipulates that in the event of a stall, the pitch control efficiency should be able to produce no less than 0.08 rad / s 2 Pitch angle acceleration capability, that is:

[0047]

[0048] To avoid over-design, determine the pitch-down 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] Where C m It is the pitch rudder effect.

[0052] From formula (2), the pitching 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. need , according to the roll 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 second-level rolling capability specified in the flight quality standard GJB185-86 "Flight Quality of Manned Aircraft (Fixed Wing)" is a 30° roll 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 α s

[0061] The flight quality specification GJB185-86 "Flight quality of manned aircraft (fixed wing)" stipulates that for any available altitude, from the stall speed V s To maximum speed V max Within the level flight speed range between 0 and 1, the longitudinal control efficiency must be sufficient to enable the aircraft to reach any level flight speed within this speed range. Therefore, the aircraft must have the ability to maintain longitudinal balance 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 roll 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 α s The maximum capacity of the nose-up moment at the stall can be used to determine the stall angle α s .

[0063] In summary, this 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, takeoff 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, 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.

[0065] The first control surface deflection calculation module is used to obtain the current aircraft's downward pitch acceleration and calculate the pitch moment coefficient requirement C through the downward pitch acceleration. m_need , get the current control surface rudder effect C mδ ; According to 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 roll 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 roll 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] Where 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 pitching 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] Where, is the roll damping derivative, l is the span, and V is the true airspeed.

[0077] Preferably, the pitch-down 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 aircraft's maximum lift coefficient corresponds to an angle of attack of 16.7°, while the pitching moment coefficient is statically unstable within the entire angle of attack range and becomes nonlinear after an angle of attack of 15°, making it impossible to directly determine the aircraft's stall angle of attack.

[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, velocity pressure qbar=3376N / m 2 , C mδ =-0.0016, C lδ =-0.0015.

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

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

[0082] In step S300, after subtracting the control surface deflection that satisfies the pitch capability requirement and the control surface deflection that satisfies the roll capability requirement from the maximum control surface deflection of the aircraft, the remaining control surface deflection is 15.4°, and the corresponding pitch moment coefficient C is 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 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. Other structures can refer to common designs. In the absence of conflicts, 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 scope of protection 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 required pitch moment coefficient C based on the pitch angle acceleration m_need , get the current control surface rudder effect C mδ ; According to 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 ea1 ; Determine the current required roll angular velocity p according to the aircraft's roll capability and requirements need , according to the roll 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 ea2 ; 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 roll 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 α s .

2. The method for determining the stall angle of attack of a statically unstable aircraft according to claim 1, wherein: The pitch angle acceleration is: Where 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 pitching moment of inertia.

3. The method for determining the stall angle of attack of a statically unstable aircraft according to claim 2, wherein: Pitching 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, wherein: The current required roll angular velocity p need For: p need =30 / 3.9 / 57.3, where 3.9 and 30 respectively refer to a 30° roll in 3.9 seconds.

5. The method for determining the stall angle of attack of a statically unstable aircraft according to claim 4, wherein: Rolling moment coefficient requirement C 1_need for: Where, 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, wherein: The aircraft's pitch-down acceleration 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 downward pitch acceleration and calculate the pitch moment coefficient requirement C through the downward pitch acceleration. m_need , get the current control surface rudder effect C mδ ; According to 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 ea1 ; 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 and requirements. need , according to the roll angular velocity p need 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 l_need and control surface roll rudder effect C lδ , the control surface deflection δ that meets the rolling capability requirements is obtained ea2 ; 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 roll 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: Where 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 pitching moment of inertia.

9. The statically unstable aircraft stall angle of attack determination system according to claim 8, characterized in that: Pitching 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, where 3.9 and 30 respectively refer to a 30° roll in 3.9 seconds.

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: Where, 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 aircraft's pitch-down acceleration at the stall angle of attack 0.08 rad / s 2 .

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