CN103577701A - Method for computing control surface hinge moment coefficient when airplane incidence angle, sideslip angle and rudder deflection angle are all zero degree - Google Patents

Abstract

The invention belongs to the airplane aerodynamic computation technology, and relates to a method for computing a control surface hinge moment coefficient when an airplane incidence angle, a sideslip angle and a rudder deflection angle are all zero degree. The method for computing the control surface hinge moment coefficient is characterized by comprising the steps of determining computation conditions, computing the relative curvature degree of a mean camber line (3), computing the zero-lift incidence angle, computing a zero-lift hinge pitching moment, and computing the control surface hinge moment coefficient when the airplane incidence angle, the sideslip angle and the rudder deflection angle are all the zero degree. According to the method, the computation accuracy of the hinge moment coefficient obtained when the airplane incidence angle, the sideslip angle and the rudder deflection angle are all the zero degree is improved, and the airplane control performance and the air safety are ensured.

Description

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

Technical field

The invention belongs to aircraft aerodynamic force computing technique, control surface hinge moment coefficient computing method when relating to a kind of aircraft angle of attack, yaw angle and angle of rudder reflection and 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 the result that existing evaluation method calculates hinge moment is also incomplete, do not comprise null alpha, hinge moment coefficient when zero degree yaw angle and zero degree angle of rudder reflection, only has hinge moment with the derivative of the angle of attack with the derivative of angle of rudder reflection, this practical value that makes undoubtedly computing method is discount to some extent, because for yaw rudder and elevating rudder, its zero angle of attack, yaw angle and angle of rudder reflection hinge moment coefficient are zero substantially, do not affect use, but larger for this value of aileron, with zero, replace null alpha, hinge moment coefficient when zero degree yaw angle and zero degree angle of rudder reflection can cause the larger error of calculation, to exert an influence to aircraft handling performance, cause potential safety hazard.

Summary of the invention

The object of the invention is: control surface hinge moment coefficient computing method when proposing a kind of aircraft angle of attack, yaw angle and angle of rudder reflection and being zero degree, so that the computational accuracy of hinge moment coefficient while improving null alpha, zero degree yaw angle and zero degree angle of rudder reflection, guarantees aircraft handling performance and flight safety.

Technical scheme of the present invention is: control surface hinge moment coefficient computing method when aircraft angle of attack, yaw angle and angle of rudder reflection are zero degree, it is characterized in that, and the step of calculating control surface hinge moment coefficient is as follows:

1, determine design conditions: the aerofoil profile of the upper surface of stabilator 1 and lower surface is considered as to symmetrical profile, a tangential section of take on wing or empennage is reference section, this reference section was control surface hinge axis mid point B, and perpendicular to the section of wing or empennage 1/4 string of a musical instrument, the leading edge point of reference section is A, the trailing edge point of reference section is C, 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 is trailing edge point C of reference section, the control surface string of a musical instrument 4 is by control surface hinge axis mid point B, the mean camber line of control surface 2 is 3, the aft terminal of mean camber line 3 is trailing edge point C of reference section, the forward terminal of mean camber line 3 is intersection point D of the control surface string of a musical instrument 4 and control surface 2 leading edges, set up the two-dimensional coordinate system of reference section, the D point of take is initial point, take straight line DC as X-axis, and right-hand is positive dirction, take to cross D point and be Y-axis perpendicular to the straight line of X-axis, and top is positive dirction, the length of line segment DC is L,

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

2.1, calculate the unique point coordinate of control surface 2 coboundary curves: control surface 2 coboundary curves are divided into 14 coboundary unique point S _i, i=1,2 ..., 14, the 1 coboundary unique point S ₁to the 14th coboundary unique point S ₁₄horizontal ordinate 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; The control surface 2 coboundary curve calculation that provide according to drawing obtain the 1st coboundary unique point S ₁to the 14th coboundary unique point S ₁₄ordinate S _iy;

2.2, calculate the unique point coordinate of control surface 2 lower limb curves: control surface 2 lower limb curves are divided into 14 unique point M _i, i=1,2 ..., 14, the 1 lower limb unique point M ₁to the 14th lower limb unique point M ₁₄horizontal ordinate 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; The control surface 2 lower limb curve calculation that provide according to drawing obtain the 1st lower limb unique point M ₁to the 14th lower limb unique point M ₁₄ordinate M _iy;

2.3, the relative camber of calculating control surface 2 mean camber lines 3, is divided into 14 mean camber line unique point N by mean camber line _i, the 1st mean camber line unique point N ₁to the 14th mean camber line unique point N ₁₄horizontal ordinate 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; The 1st mean camber line unique point N ₁to the 14th mean camber line unique point N ₁₄ordinate respectively:

N _iY＝0.5×(S _iY+M _iY)/L…………………………………………………[1]

3, calculate zero and rise the angle of attack:

${\mathrm{\&alpha;}}_0=-\underset{1}{\overset{14}{\mathrm{\&Sigma;}}}{A}_i\&times;N_{\mathrm{iY}}\&CenterDot;\&CenterDot;\&CenterDot;\left[2\right]$

Wherein, calculating parameter A _ivalue is respectively: A ₁=2.9, A ₂=4.22, A ₃=3.12, A ₄=4.82, A ₅=5.88, A ₆=5.76, A ₇=6.26, A ₈=7.34, A ₉=9.83, A ₁₀=13.44, A ₁₁=23.5, A ₁₂=43.44, A ₁₃=119.7, A ₁₄=-329.8;

4, calculate zero and rise pitching moment:

$m_{z0}=\underset{1}{\overset{14}{\mathrm{\&Sigma;}}}{K}_i\&times;N_{\mathrm{iY}}\&CenterDot;\&CenterDot;\&CenterDot;\left[3\right]$

Wherein, calculating parameter K _ivalue is respectively: K ₁=0.238, K ₂=0.312, K ₃=0.208, K ₄=0.248, K ₅=0.148, K ₆=0.018, K ₇=-0.09, K ₈=-0.202, K ₉=-0.34, K ₁₀=-0.564, K ₁₁=-0.954, K ₁₂=-1.572, K ₁₃=-6.052, K ₁₄=-9.578;

Control surface hinge moment coefficient m when 5, calculating aircraft angle of attack, yaw angle and angle of rudder reflection are zero degree _j0:

$m_{j0}=({\mathrm{\&alpha;}}_1-{\mathrm{\&alpha;}}_0)\&times;m_j^{\mathrm{\&delta;}}+m_{z0}/(1-\mathrm{\&lambda;})\&CenterDot;\&CenterDot;\&CenterDot;\left[4\right]$

Wherein, λ is the tangential relative position of hinge axis, λ=DB/DC, α ₁for the control surface string of a musical instrument 4 angle with the reference section string of a musical instrument 5, with in control surface string of a musical instrument leading edge partially for just,

for the derivative of control surface hinge moment coefficient with angle of rudder reflection,

by drawing, provided.

Advantage of the present invention is: control surface hinge moment coefficient computing method when having proposed a kind of aircraft angle of attack, yaw angle and angle of rudder reflection and being zero degree, while having improved null alpha, zero degree yaw angle and zero degree angle of rudder reflection, the computational accuracy of hinge moment coefficient, has guaranteed aircraft handling performance and flight safety.

Accompanying drawing explanation

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

Embodiment

Below the present invention is described in further detail.Referring to Fig. 1, when aircraft angle of attack, yaw angle and angle of rudder reflection are zero degree, control surface hinge moment coefficient computing method, is characterized in that, the step of calculating control surface hinge moment coefficient is as follows:

1, determine design conditions: the aerofoil profile of the upper surface of stabilator 1 and lower surface is considered as to symmetrical profile, a tangential section of take on wing or empennage is reference section, this reference section was control surface hinge axis mid point B, and perpendicular to the section of wing or empennage 1/4 string of a musical instrument, the leading edge point of reference section is A, the trailing edge point of reference section is C, 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 is trailing edge point C of reference section, the control surface string of a musical instrument 4 is by control surface hinge axis mid point B, the mean camber line of control surface 2 is 3, the aft terminal of mean camber line 3 is trailing edge point C of reference section, the forward terminal of mean camber line 3 is intersection point D of the control surface string of a musical instrument 4 and control surface 2 leading edges, set up the two-dimensional coordinate system of reference section, the D point of take is initial point, take straight line DC as X-axis, and right-hand is positive dirction, take to cross D point and be Y-axis perpendicular to the straight line of X-axis, and top is positive dirction, the length of line segment DC is L,

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

2.1, calculate the unique point coordinate of control surface 2 coboundary curves: control surface 2 coboundary curves are divided into 14 coboundary unique point S _i, i=1,2 ..., 14, the 1 coboundary unique point S ₁to the 14th coboundary unique point S ₁₄horizontal ordinate 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; The control surface 2 coboundary curve calculation that provide according to drawing obtain the 1st coboundary unique point S ₁to the 14th coboundary unique point S ₁₄ordinate S _iy;

2.2, calculate the unique point coordinate of control surface 2 lower limb curves: control surface 2 lower limb curves are divided into 14 unique point M _i, i=1,2 ..., 14, the 1 lower limb unique point M ₁to the 14th lower limb unique point M ₁₄horizontal ordinate 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; The control surface 2 lower limb curve calculation that provide according to drawing obtain the 1st lower limb unique point M ₁to the 14th lower limb unique point M ₁₄ordinate M _iy;

2.3, the relative camber of calculating control surface 2 mean camber lines 3, is divided into 14 mean camber line unique point N by mean camber line _i, the 1st mean camber line unique point N ₁to the 14th mean camber line unique point N ₁₄horizontal ordinate 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; The 1st mean camber line unique point N ₁to the 14th mean camber line unique point N ₁₄ordinate respectively:

N _iY＝0.5×(S _iY+M _iY)/L…………………………………………………[1]

3, calculate zero and rise the angle of attack:

${\mathrm{\&alpha;}}_0=-\underset{1}{\overset{14}{\mathrm{\&Sigma;}}}{A}_i\&times;N_{\mathrm{iY}}\&CenterDot;\&CenterDot;\&CenterDot;\left[2\right]$

Wherein, calculating parameter A _ivalue is respectively: A ₁=2.9, A ₂=4.22, A ₃=3.12, A ₄=4.82, A ₅=5.88, A ₆=5.76, A ₇=6.26, A ₈=7.34, A ₉=9.83, A ₁₀=13.44, A ₁₁=23.5, A ₁₂=43.44, A ₁₃=119.7, A ₁₄=-329.8;

4, calculate zero and rise pitching moment:

$m_{z0}=\underset{1}{\overset{14}{\mathrm{\&Sigma;}}}{K}_i\&times;N_{\mathrm{iY}}\&CenterDot;\&CenterDot;\&CenterDot;\left[3\right]$

Wherein, calculating parameter K _ivalue is respectively: K ₁=0.238, K ₂=0.312, K ₃=0.208, K ₄=0.248, K ₅=0.148, K ₆=0.018, K ₇=-0.09, K ₈=-0.202, K ₉=-0.34, K ₁₀=-0.564, K ₁₁=-0.954, K ₁₂=-1.572, K ₁₃=-6.052, K ₁₄=-9.578;

Control surface hinge moment coefficient m when 5, calculating aircraft angle of attack, yaw angle and angle of rudder reflection are zero degree _j0:

$m_{j0}=({\mathrm{\&alpha;}}_1-{\mathrm{\&alpha;}}_0)\&times;m_j^{\mathrm{\&delta;}}+m_{z0}/(1-\mathrm{\&lambda;})\&CenterDot;\&CenterDot;\&CenterDot;\left[4\right]$

Wherein, λ is the tangential relative position of hinge axis, λ=DB/DC, α ₁for the control surface string of a musical instrument 4 angle with the reference section string of a musical instrument 5, with in control surface string of a musical instrument leading edge partially for just,

for the derivative of control surface hinge moment coefficient with angle of rudder reflection,

by drawing, provided.

Principle of work of the present invention is: control surface is considered as to an independently aerofoil, by calculating zero of control surface, rise the angle of attack and the zero pitching moment that rises, control surface deflect into zero rise the angle of attack in its suffered hinge moment be zero and rise pitching moment, the hinge moment being provided by drawing is again with angle of rudder reflection derivative, can calculate the hinge moment coefficient of aircraft angle of attack, yaw angle and angle of rudder reflection control surface while being zero degree.

Embodiment 1

Take certain model feeder liner aileron is example, and its aileron zero degree hinge moment coefficient test findings is m _j0=-0.118, by drawing, provided

calculating 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; Calculate zero and rise angle of attack α ₀=-0.16 and zero rises pitching moment m _z0=-0.04545; By drawing, obtain control surface with stabilator angle α ₁=6.63 °, hinge axis location λ=6.1% and formula $m_{j0}=({\mathrm{\&alpha;}}_1-{\mathrm{\&alpha;}}_0)\&times;m_j^{\mathrm{\&delta;}}+m_{z0}/(1-\mathrm{\&lambda;})$ 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

Take certain model transporter aileron is example, and its aileron zero degree hinge moment coefficient test findings is m _j0=-0.142, by drawing, provided

calculating 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; Calculate zero and rise angle of attack α ₀=-5.406 and zero rises pitching moment m _z0=-0.06983; By drawing, obtain control surface with stabilator angle α ₁=0.78, hinge axis location λ=4.784% and formula

obtain:

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 aircraft angle of attack, yaw angle and angle of rudder reflection are zero degree, control surface hinge moment coefficient computing method, is characterized in that, the step of calculating control surface hinge moment coefficient is as follows:

1.1, determine design conditions: the aerofoil profile of the upper surface of stabilator (1) and lower surface is considered as to symmetrical profile, a tangential section of take on wing or empennage is reference section, this reference section was control surface hinge axis mid point B, and perpendicular to the section of wing or empennage 1/4 string of a musical instrument, the leading edge point of reference section is A, the trailing edge point of reference section is C, 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) is the trailing edge point C of reference section, the control surface string of a musical instrument (4) is by control surface hinge axis mid point B, the mean camber line of control surface (2) is (3), the aft terminal of mean camber line (3) is the trailing edge point C of reference section, the forward terminal of mean camber line (3) is the intersection point D of the control surface string of a musical instrument (4) and control surface (2) leading edge, set up the two-dimensional coordinate system of reference section, the D point of take is initial point, take straight line DC as X-axis, and right-hand is positive dirction, take to cross D point and be Y-axis perpendicular to the straight line of X-axis, and top is positive dirction, the length of line segment DC is L,

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

1.2.1, calculate the unique point coordinate of control surface (2) coboundary curve: control surface (2) coboundary curve is divided into 14 coboundary unique point S _i, i=1,2 ..., 14, the 1 coboundary unique point S ₁to the 14th coboundary unique point S ₁₄horizontal ordinate 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; The control surface providing according to drawing (2) coboundary curve calculation obtains the 1st coboundary unique point S ₁to the 14th coboundary unique point S ₁₄ordinate S _iy;

1.2.2, calculate the unique point coordinate of control surface (2) lower limb curve: control surface (2) lower limb curve is divided into 14 unique point M _i, i=1,2 ..., 14, the 1 lower limb unique point M ₁to the 14th lower limb unique point M ₁₄horizontal ordinate 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; The control surface providing according to drawing (2) lower limb curve calculation obtains the 1st lower limb unique point M ₁to the 14th lower limb unique point M ₁₄ordinate M _iy;

1.2.3, calculate the relative camber of control surface (2) mean camber line (3), mean camber line is divided into 14 mean camber line unique point N _i, the 1st mean camber line unique point N ₁to the 14th mean camber line unique point N ₁₄horizontal ordinate 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; The 1st mean camber line unique point N ₁to the 14th mean camber line unique point N ₁₄ordinate respectively:

N _iY＝0.5×(S _iY+M _iY)/L………………………………………………[1]

1.3, calculate zero and rise the angle of attack:

${\mathrm{\&alpha;}}_0=-\underset{1}{\overset{14}{\mathrm{\&Sigma;}}}{A}_i\&times;N_{\mathrm{iY}}\&CenterDot;\&CenterDot;\&CenterDot;\left[2\right]$

Wherein, calculating parameter A _ivalue is respectively: A ₁=2.9, A ₂=4.22, A ₃=3.12, A ₄=4.82, A ₅=5.88, A ₆=5.76, A ₇=6.26, A ₈=7.34, A ₉=9.83, A ₁₀=13.44, A ₁₁=23.5, A ₁₂=43.44, A ₁₃=119.7, A ₁₄=-329.8;

1.4, calculate zero and rise pitching moment:

$m_{z0}=\underset{1}{\overset{14}{\mathrm{\&Sigma;}}}{K}_i\&times;N_{\mathrm{iY}}\&CenterDot;\&CenterDot;\&CenterDot;\left[3\right]$

Wherein, calculating parameter K _ivalue is respectively: K ₁=0.238, K ₂=0.312, K ₃=0.208, K ₄=0.248, K ₅=0.148, K ₆=0.018, K ₇=-0.09, K ₈=-0.202, K ₉=-0.34, K ₁₀=-0.564, K ₁₁=-0.954, K ₁₂=-1.572, K ₁₃=-6.052, K ₁₄=-9.578;

Control surface hinge moment coefficient m when 1.5, calculating aircraft angle of attack, yaw angle and angle of rudder reflection are zero degree _j0:

$m_{j0}=({\mathrm{\&alpha;}}_1-{\mathrm{\&alpha;}}_0)\&times;m_j^{\mathrm{\&delta;}}+m_{z0}/(1-\mathrm{\&lambda;})...\left[4\right]$

Wherein, λ is the tangential relative position of hinge axis, λ=DB/DC, α ₁for the angle of the control surface string of a musical instrument (4) and the reference section string of a musical instrument (5), with in control surface string of a musical instrument leading edge partially for just, for the derivative of control surface hinge moment coefficient with angle of rudder reflection,

by drawing, provided.

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