CN103577701B - Control surface hinge moment coefficient computational methods when aircraft angle of attack, yaw angle and angle of rudder reflection are zero degree - Google Patents

Abstract

The invention belongs to airplane aerodynamic computing technique, relate to control surface hinge moment coefficient computational methods when a kind of aircraft angle of attack, yaw angle and angle of rudder reflection are zero degree.It is characterized in that, the step calculating control surface hinge moment coefficient is as follows: determine design conditions；Calculate the relative camber of mean camber line (3)；Calculate zero liter of angle of attack；Calculate zero liter of pitching moment;Calculate control surface hinge moment coefficient when aircraft angle of attack, yaw angle and angle of rudder reflection are zero degree.The computational accuracy of hinge moment coefficient when the present invention improves null alpha, zero degree yaw angle and zero degree angle of rudder reflection, it is ensured that aircraft handling performance and flight safety.

Description

Control surface hinge moment coefficient computational methods when aircraft angle of attack, yaw angle and angle of rudder reflection are zero degree

Technical field

The invention belongs to airplane aerodynamic computing technique, relate to a kind of aircraft angle of attack, yaw angle and angle of rudder reflection Control surface hinge moment coefficient computational methods when being zero degree.

Background technology

Existing aircraft rudder surface hinge moment evaluation method system has " ESDU ", DATACOM, " Airplane Design ", " airplane design handbook ", " aviation aerodynamic force engineering calculation handbook " etc..But it is existing Evaluation method result that hinge moment is calculated not exclusively, the most do not include that null alpha, zero degree break away Hinge moment coefficient when angle and zero degree angle of rudder reflection, only hinge moment is with the derivative of the angle of attack with angle of rudder reflection Derivative, this makes the practical value discount of computational methods undoubtedly, because to rudder and elevator Speech, its zero angle of attack, yaw angle and angle of rudder reflection hinge moment coefficient are substantially zeroed, do not affect use, but right For aileron, this value is relatively big, with hinge when zero replacement null alpha, zero degree yaw angle and zero degree angle of rudder reflection Moment coefficient can cause bigger calculating error, aircraft handling performance will be produced impact, causes safety hidden Suffer from.

Summary of the invention

It is an object of the invention to: propose to handle when a kind of aircraft angle of attack, yaw angle and angle of rudder reflection are zero degree Face hinge moment coefficient computational methods, in order to when improving null alpha, zero degree yaw angle and zero degree angle of rudder reflection The computational accuracy of hinge moment coefficient, it is ensured that aircraft handling performance and flight safety.

The technical scheme is that control surface hinge when aircraft angle of attack, yaw angle and angle of rudder reflection are zero degree Chain moment coefficient computational methods, it is characterised in that the step calculating control surface hinge moment coefficient is as follows:

1, design conditions are determined: the aerofoil profile of the upper and lower surface of stabilization 1 is considered as symmetrical profile； With a tangential section on wing or empennage as reference section, this reference section was control surface hinge Axle midpoint B is also perpendicular to wing or the section of empennage 1/4 string of a musical instrument, and the leading edge point of reference section is A, The trailing edge point of reference section is C, and line segment AC is the reference section string of a musical instrument 5, the aft terminal of the control surface string of a musical instrument 4 Being the trailing edge point C of reference section, the control surface string of a musical instrument 4 passes through control surface hinge axis midpoint B, control surface 2 Mean camber line is 3, and the aft terminal of mean camber line 3 is the trailing edge point C of reference section, and the forward terminal of mean camber line 3 is The control surface string of a musical instrument 4 and the intersection point D of control surface 2 leading edge；Set up the two-dimensional coordinate system of reference section, with D point For initial point, with straight line DC as X-axis, right is positive direction, to cross D point and to be perpendicular to the straight line of X-axis and be Y-axis, is arranged above positive direction；The a length of L of line segment DC；

2, the relative camber of calculating mean camber line 3:

2.1, the characteristic point coordinate of control surface 2 top edge curve is calculated: divided by control surface 2 top edge curve It is 14 top edge characteristic points S_i, i=1,2 ..., the 14, the 1st top edge characteristic point S₁To the 14th top Edge characteristic point S₁₄Abscissa respectively: S_1X=0, S_2X=0.025L, S_3X=0.05L, S_4X=0.1L, S_5X=0.25L, S_6X=0.3L, S_7X=0.4L, S_8X=0.5L, S_9X=0.6L, S_10X=0.7L, S_11X=0.8L, S_12X=0.9L, S_13X=0.95L, S_14X=1L；It is calculated the 1st top according to the control surface 2 top edge curve that drawing is given Edge characteristic point S₁To the 14th top edge characteristic point S₁₄Vertical coordinate S_iy；

2.2, the characteristic point coordinate of control surface 2 lower limb curve is calculated: divided by control surface 2 lower limb curve It is 14 characteristic points M_i, i=1,2 ..., the 14, the 1st lower limb characteristic point M₁To the 14th lower limb feature Point M₁₄Abscissa respectively: M_1X=0, M_2X=0.025L, M_3X=0.05L, M_4X=0.1L, M_5X=0.25L, M_6X=0.3L, M_7X=0.4L, M_8X=0.5L, M_9X=0.6L, M_10X=0.7L, M_11X=0.8L, M_12X=0.9L, M_13X=0.95L, M_14X=1L；It is following that the control surface 2 lower limb curve be given according to drawing is calculated the 1st Edge characteristic point M₁To the 14th lower limb characteristic point M₁₄Vertical coordinate M_iy；

2.3, calculate the relative camber of control surface 2 mean camber line 3, mean camber line is divided into 14 mean camber line spies Levy a N_i, the 1st mean camber line characteristic point N₁To the 14th mean camber line characteristic point N₁₄Abscissa respectively: N_1X=0, N_2X=0.025L, N_3X=0.05L, N_4X=0.1L, N_5X=0.25L, N_6X=0.3L, N_7X=0.4L, N_8X=0.5L, N_9X=0.6L, N_10X=0.7L, N_11X=0.8L, N_12X=0.9L, N_13X=0.95L, N_14X=1L；1st mean camber line is special Levy a N₁To the 14th mean camber line characteristic point N₁₄Vertical coordinate respectively:

N_iY=0.5 × (S_iY+M_iY)/L…………………………………………………[1]

3, zero liter of angle of attack is calculated:

${\mathrm{\α}}_0=-\underset{1}{\overset{14}{\mathrm{\Σ}}}{A}_i\×N_{\mathrm{iY}}\·\·\·\left[2\right]$

Wherein, parameter A is calculated_iValue is respectively as follows: A₁=2.9, A₂=4.22, A₃=3.12, A₄=4.82, A₅=5.88, A₆=5.76, A₇=6.26, A₈=7.34, A₉=9.83, A₁₀=13.44, A₁₁=23.5, A₁₂=43.44, A₁₃=119.7, A₁₄=-329.8；

4, zero liter of pitching moment is calculated:

$m_{z0}=\underset{1}{\overset{14}{\mathrm{\Σ}}}{K}_i\×N_{\mathrm{iY}}\·\·\·\left[3\right]$

Wherein, parameter K is calculated_iValue is respectively as follows: K₁=0.238, K₂=0.312, K₃=0.208, K₄=0.248, K₅=0.148, K₆=0.018, K₇=-0.09, K₈=-0.202, K₉=-0.34, K₁₀=-0.564, K₁₁=-0.954, K₁₂=-1.572, K₁₃=-6.052, K₁₄=-9.578；

5, control surface hinge moment coefficient m when aircraft angle of attack, yaw angle and angle of rudder reflection are zero degree is calculated_j0:

$m_{j0}=({\mathrm{\α}}_1-{\mathrm{\α}}_0)\×m_j^{\mathrm{\δ}}+m_{z0}/(1-\mathrm{\λ})\·\·\·\left[4\right]$

Wherein, λ is the tangential relative position of hinge axis, λ=DB/DC, α₁Cut open with calculating for the control surface string of a musical instrument 4 The angle of the face string of a musical instrument 5, partially for just in control surface string of a musical instrument leading edge,For control surface hinge moment coefficient with rudder The derivative of drift angle,Be given by drawing.

The invention have the advantage that and propose behaviour when a kind of aircraft angle of attack, yaw angle and angle of rudder reflection are zero degree Vertical face hinge moment coefficient computational methods, when improve null alpha, zero degree yaw angle and zero degree angle of rudder reflection The computational accuracy of hinge moment coefficient, it is ensured that aircraft handling performance and flight safety.

Accompanying drawing explanation

Fig. 1 is the Computing Principle schematic diagram of the present invention.

Detailed description of the invention

Below the present invention is described in further detail.See Fig. 1, aircraft angle of attack, yaw angle and angle of rudder reflection Control surface hinge moment coefficient computational methods when being zero degree, it is characterised in that calculate control surface hinge power The step of moment coefficient is as follows:

1, design conditions are determined: the aerofoil profile of the upper and lower surface of stabilization 1 is considered as symmetrical profile； With a tangential section on wing or empennage as reference section, this reference section was control surface hinge Axle midpoint B is also perpendicular to wing or the section of empennage 1/4 string of a musical instrument, and the leading edge point of reference section is A, The trailing edge point of reference section is C, and line segment AC is the reference section string of a musical instrument 5, the aft terminal of the control surface string of a musical instrument 4 Being the trailing edge point C of reference section, the control surface string of a musical instrument 4 passes through control surface hinge axis midpoint B, control surface 2 Mean camber line is 3, and the aft terminal of mean camber line 3 is the trailing edge point C of reference section, and the forward terminal of mean camber line 3 is The control surface string of a musical instrument 4 and the intersection point D of control surface 2 leading edge；Set up the two-dimensional coordinate system of reference section, with D point For initial point, with straight line DC as X-axis, right is positive direction, to cross D point and to be perpendicular to the straight line of X-axis and be Y-axis, is arranged above positive direction；The a length of L of line segment DC；

2, the relative camber of calculating mean camber line 3:

2.1, the characteristic point coordinate of control surface 2 top edge curve is calculated: divided by control surface 2 top edge curve It is 14 top edge characteristic points S_i, i=1,2 ..., the 14, the 1st top edge characteristic point S₁To the 14th top Edge characteristic point S₁₄Abscissa respectively: S_1X=0, S_2X=0.025L, S_3X=0.05L, S_4X=0.1L, S_5X=0.25L, S_6X=0.3L, S_7X=0.4L, S_8X=0.5L, S_9X=0.6L, S_10X=0.7L, S_11X=0.8L, S_12X=0.9L, S_13X=0.95L, S_14X=1L；It is calculated the 1st top according to the control surface 2 top edge curve that drawing is given Edge characteristic point S₁To the 14th top edge characteristic point S₁₄Vertical coordinate S_iy；

2.2, the characteristic point coordinate of control surface 2 lower limb curve is calculated: divided by control surface 2 lower limb curve It is 14 characteristic points M_i, i=1,2 ..., the 14, the 1st lower limb characteristic point M₁To the 14th lower limb feature Point M₁₄Abscissa respectively: M_1X=0, M_2X=0.025L, M_3X=0.05L, M_4X=0.1L, M_5X=0.25L, M_6X=0.3L, M_7X=0.4L, M_8X=0.5L, M_9X=0.6L, M_10X=0.7L, M_11X=0.8L, M_12X=0.9L, M_13X=0.95L, M_14X=1L；It is following that the control surface 2 lower limb curve be given according to drawing is calculated the 1st Edge characteristic point M₁To the 14th lower limb characteristic point M₁₄Vertical coordinate M_iy；

2.3, calculate the relative camber of control surface 2 mean camber line 3, mean camber line is divided into 14 mean camber line spies Levy a N_i, the 1st mean camber line characteristic point N₁To the 14th mean camber line characteristic point N₁₄Abscissa respectively: N_1X=0, N_2X=0.025L, N_3X=0.05L, N_4X=0.1L, N_5X=0.25L, N_6X=0.3L, N_7X=0.4L, N_8X=0.5L, N_9X=0.6L, N_10X=0.7L, N_11X=0.8L, N_12X=0.9L, N_13X=0.95L, N_14X=1L；1st mean camber line is special Levy a N₁To the 14th mean camber line characteristic point N₁₄Vertical coordinate respectively:

N_iY=0.5 × (S_iY+M_iY)/L…………………………………………………[1]

3, zero liter of angle of attack is calculated:

${\mathrm{\α}}_0=-\underset{1}{\overset{14}{\mathrm{\Σ}}}{A}_i\×N_{\mathrm{iY}}\·\·\·\left[2\right]$

Wherein, parameter A is calculated_iValue is respectively as follows: A₁=2.9, A₂=4.22, A₃=3.12, A₄=4.82, A₅=5.88, A₆=5.76, A₇=6.26, A₈=7.34, A₉=9.83, A₁₀=13.44, A₁₁=23.5, A₁₂=43.44, A₁₃=119.7, A₁₄=-329.8；

4, zero liter of pitching moment is calculated:

$m_{z0}=\underset{1}{\overset{14}{\mathrm{\Σ}}}{K}_i\×N_{\mathrm{iY}}\·\·\·\left[3\right]$

Wherein, parameter K is calculated_iValue is respectively as follows: K₁=0.238, K₂=0.312, K₃=0.208, K₄=0.248, K₅=0.148, K₆=0.018, K₇=-0.09, K₈=-0.202, K₉=-0.34, K₁₀=-0.564, K₁₁=-0.954, K₁₂=-1.572, K₁₃=-6.052, K₁₄=-9.578；

5, control surface hinge moment coefficient m when aircraft angle of attack, yaw angle and angle of rudder reflection are zero degree is calculated_j0:

$m_{j0}=({\mathrm{\α}}_1-{\mathrm{\α}}_0)\×m_j^{\mathrm{\δ}}+m_{z0}/(1-\mathrm{\λ})\·\·\·\left[4\right]$

Wherein, λ is the tangential relative position of hinge axis, λ=DB/DC, α₁Cut open with calculating for the control surface string of a musical instrument 4 The angle of the face string of a musical instrument 5, partially for just in control surface string of a musical instrument leading edge,For control surface hinge moment coefficient with rudder The derivative of drift angle,Be given by drawing.

The operation principle of the present invention is: control surface is considered as an independent aerofoil, by being calculated behaviour Zero liter of angle of attack in vertical face and zero liter of pitching moment, then control surface is deflecting into zero liter of angle of attack when suffered by it Hinge moment be zero liter of pitching moment, then the hinge moment be given by drawing is with angle of rudder reflection derivative, can Calculate the hinge moment coefficient of control surface when aircraft angle of attack, yaw angle and angle of rudder reflection are zero degree.

Embodiment 1

As a example by certain model feeder liner aileron, its aileron zero degree hinge moment coefficient result of the test is m_j0=-0.118, is given by drawingCalculating its relative camber value result is: N_1Y=0.14968, N_2Y=-0.05276, N_3Y=-0.05418, N_4Y=-0.05094, N_5Y=-0.03726, N_6Y=-0.02481, N_7Y=-0.01387, N_8Y=-0.00507, N_9Y=0.001519, N_10Y=0.005813, N_11Y=0.006988, N_12Y=0.005266, N_13Y=0.003038, N_14Y=0；It is calculated zero liter of angle of attack α₀=-0.16 and zero liter of pitching Moment m_z0=-0.04545；Control surface is obtained with stabilization angle α by drawing₁=6.63 °, hinge axis location λ=6.1% and formula $m_{j0}=({\mathrm{\α}}_1-{\mathrm{\α}}_0)\×m_j^{\mathrm{\δ}}+m_{z0}/(1-\mathrm{\λ})$ Obtain:

m_j0=(6.63-(-0.16))×(-0.00849)+(-0.04545)/(1-0.061)=-0.106；

Accuracy in computation is 0.106/0.118=90%；

Embodiment 2

As a example by certain model transporter aileron, its aileron zero degree hinge moment coefficient result of the test is m_j0=-0.142, is given by drawingCalculating its relative camber value result is: N_1Y=0.075025, N_2Y=0.06378, N_3Y=0.057848, N_4Y=0.05802, N_5Y=0.05852, N_6Y=0.05756, N_7Y=0.05454, N_8Y=0.049807, N_9Y=0.045766, N_10Y=0.03731, N_11Y=0.02746, N_12Y=0.0148, N_13Y=0.007162, N_14Y=0；It is calculated zero liter of angle of attack α₀=-5.406 With zero liter of pitching moment m_z0=-0.06983；Control surface is obtained with stabilization angle α by drawing₁=0.78, hinge Shaft position λ=4.784% and formulaObtain:

m_j0=(0.78-(-5.406))×(-0.009276)+(-0.06983)/(1-0.04784)=-0.131；

Accuracy in computation is 0.137/0.142=97%.

Claims (1)

1. when an aircraft angle of attack, yaw angle and angle of rudder reflection are zero degree, control surface hinge moment coefficient calculates Method, it is characterised in that the step calculating control surface hinge moment coefficient is as follows:

1.1, design conditions are determined: the aerofoil profile of the upper and lower surface of stabilization (1) is considered as symmetric form Face；With a tangential section on wing or empennage as reference section, this reference section was control surface Hinge axis midpoint B is also perpendicular to wing or the section of empennage 1/4 string of a musical instrument, and the leading edge point of reference section is A, the trailing edge point of reference section is C, and line segment AC is the reference section string of a musical instrument (5), after the control surface string of a musical instrument (4) End points is the trailing edge point C of reference section, and the control surface string of a musical instrument (4) passes through control surface hinge axis midpoint B, handles The mean camber line in face (2) is (3), and the aft terminal of mean camber line (3) is the trailing edge point C of reference section, mean camber line (3) Forward terminal be the intersection point D of the control surface string of a musical instrument (4) and control surface (2) leading edge；Set up the two dimension seat of reference section Mark system, with D point as initial point, with straight line DC as X-axis, right is positive direction, to cross D point and to be perpendicular to The straight line of X-axis is Y-axis, is arranged above positive direction；The a length of L of line segment DC；

1.2, the relative camber of calculating mean camber line (3):

1.2.1 the characteristic point coordinate of control surface (2) top edge curve, is calculated: control surface (2) top edge is bent Line is divided into 14 top edge characteristic points S_i, i=1,2 ..., the 14, the 1st top edge characteristic point S₁To the 14th Top edge characteristic point S₁₄Abscissa respectively: S_1X=0, S_2X=0.025L, S_3X=0.05L, S_4X=0.1L, S_5X=0.25L, S_6X=0.3L, S_7X=0.4L, S_8X=0.5L, S_9X=0.6L, S_10X=0.7L, S_11X=0.8L, S_12X=0.9L, S_13X=0.95L, S_14X=1L；It is calculated the 1st top according to control surface (2) the top edge curve that drawing is given Edge characteristic point S₁To the 14th top edge characteristic point S₁₄Vertical coordinate S_iY；

1.2.2 the characteristic point coordinate of control surface (2) lower limb curve, is calculated: control surface (2) lower limb is bent Line is divided into 14 characteristic points M_i, i=1,2 ..., the 14, the 1st lower limb characteristic point M₁To the 14th lower limb Characteristic point M₁₄Abscissa respectively: M_1X=0, M_2X=0.025L, M_3X=0.05L, M_4X=0.1L, M_5X=0.25L, M_6X=0.3L, M_7X=0.4L, M_8X=0.5L, M_9X=0.6L, M_10X=0.7L, M_11X=0.8L, M_12X=0.9L, M_13X=0.95L, M_14X=1L；It is following that control surface (2) the lower limb curve be given according to drawing is calculated the 1st Edge characteristic point M₁To the 14th lower limb characteristic point M₁₄Vertical coordinate M_iY；

1.2.3, calculate control surface (2) mean camber line (3) relative camber, mean camber line is divided into 14 middle arcs Line feature point N_i, the 1st mean camber line characteristic point N₁To the 14th mean camber line characteristic point N₁₄Abscissa respectively: N_1X=0, N_2X=0.025L, N_3X=0.05L, N_4X=0.1L, N_5X=0.25L, N_6X=0.3L, N_7X=0.4L, N_8X=0.5L, N_9X=0.6L, N_10X=0.7L, N_11X=0.8L, N_12X=0.9L, N_13X=0.95L, N_14X=1L；1st mean camber line is special Levy a N₁To the 14th mean camber line characteristic point N₁₄Vertical coordinate respectively:

N_iY=0.5 × (S_iY+M_iY)/L………………………………………………[1]

1.3, zero liter of angle of attack is calculated:

${\mathrm{\α}}_0=-\underset{1}{\overset{14}{\Σ}}{A}_i\×N_{iY}...\[2\]$

Wherein, parameter A is calculated_iValue is respectively as follows: A₁=2.9, A₂=4.22, A₃=3.12, A₄=4.82, A₅=5.88, A₆=5.76, A₇=6.26, A₈=7.34, A₉=9.83, A₁₀=13.44, A₁₁=23.5, A₁₂=43.44, A₁₃=119.7, A₁₄=-329.8；

1.4, zero liter of pitching moment is calculated:

$m_{z0}=\underset{1}{\overset{14}{\Σ}}{K}_i\×N_{iY}...\[3\]$

Wherein, parameter K is calculated_iValue is respectively as follows: K₁=0.238, K₂=0.312, K₃=0.208, K₄=0.248, K₅=0.148, K₆=0.018, K₇=-0.09, K₈=-0.202, K₉=-0.34, K₁₀=-0.564, K₁₁=-0.954, K₁₂=-1.572, K₁₃=-6.052, K₁₄=-9.578；

1.5, control surface hinge moment coefficient when aircraft angle of attack, yaw angle and angle of rudder reflection are zero degree is calculated m_j0:

$m_{j0}=({\mathrm{\α}}_1-{\mathrm{\α}}_0)\×m_j^{\mathrm{\δ}}+m_{z0}/(1-\mathrm{\λ})...\[4\]$

Wherein, λ is the tangential relative position of hinge axis, λ=DB/DC, α₁For the control surface string of a musical instrument (4) and calculating The angle of section chord line (5), partially for just in control surface string of a musical instrument leading edge,For control surface hinge moment coefficient With the derivative of angle of rudder reflection,Be given by drawing.