CN105819001A - Determination method for horizontal tail installation angle of horizontal tail fixed airplane - Google Patents

Abstract

The invention discloses a determination method for a horizontal tail installation angle of a horizontal tail fixed airplane. The method comprises the following steps: step 1: determining a first design optimization point; step 2: obtaining airplane parameters under the first design optimization point, and calculating a first horizontal tail installation angle of the first design optimization point through a formula; step 3: substituting the first horizontal tail installation angle into a flight process of the whole horizontal tail fixed airplane, and verifying whether an elevating rudder can guarantee normal flight of the horizontal tail fixed airplane in a self deflection angle range under the first horizontal tail installation angle or not; step 4: calculating a second horizontal tail installation angle according to a longitudinal maneuvering capability of the horizontal tail fixed airplane; and step 5: selecting one of the first horizontal tail installation angle and the second horizontal tail installation angle as an actual horizontal tail installation angle of the horizontal tail fixed airplane. The determination method for the horizontal tail installation angle of the horizontal tail fixed airplane, disclosed by the invention, can be used for determining a suitable horizontal tail installation angle.

Description

The horizontal tail setting angle of a kind of horizontal tail fastening aircraft determines method

Technical field

The present invention relates to aircraft controllability stability Design technical field, the horizontal tail setting angle particularly relating to a kind of horizontal tail fastening aircraft determines method.

Background technology

Horizontal tail and be positioned at the rudder face elevator of horizontal tail trailing edge the fore-and-aft control quality of aircraft is had important function.For immovable horizontal tail aircraft, horizontal tail fixes an established angle, and longitudinal trim and transshipping motor-driven completes by elevator.The established angle now fixing horizontal tail determines and important, and it relates to the manipulation with rudder scope and driver of elevator and experiences.

Prior art not yet has a kind of method quickly determining established angle.

Thus, it is desirable to have a kind of technical scheme overcomes or at least alleviates at least one drawbacks described above of prior art.

Summary of the invention

The horizontal tail setting angle that it is an object of the invention to provide a kind of horizontal tail fastening aircraft determines that method overcomes or at least alleviates at least one drawbacks described above of prior art.

For achieving the above object, the present invention provides the horizontal tail setting angle of a kind of horizontal tail fastening aircraft to determine method.The horizontal tail setting angle of described horizontal tail fastening aircraft determines that method comprises the steps:

Step 1: fly trim ability determine the first design optimization point according to the longitudinal direction of horizontal tail fastening aircraft is flat；

Step 2: obtain the aircraft parameter under the first design optimization point, and calculated the first horizontal tail setting angle of this first design optimization point by formula；

Step 3: this first horizontal tail setting angle is brought in whole horizontal tail fastening aircraft flight course, checking elevator under the first horizontal tail setting angle this described, the normal flight that whether ensure that described horizontal tail fastening aircraft in the range of deflection angles of himself of described elevator；The most then carry out next step；If it is not, then repeating said steps 1 is to described step 2, until the result in described step 3 is yes；

Step 4: according to the longitudinal maneuver ability of horizontal tail fastening aircraft and calculate the second horizontal tail setting angle by formula；

Step 5: select an actual horizontal tail setting angle as described horizontal tail fastening aircraft by formula between described first horizontal tail setting angle and described second horizontal tail setting angle.

Preferably, the aircraft parameter in described step 2 includes aircraft weight center of gravity, flying height, flight speed, engine condition, flap state.

Preferably, the formula in described step 2 is:

Wherein,

δ_eFor elevator drift angle, this formula takes O；G is aircraft weight, and S is wing area, c_AFor wing mean aerodynamic chord, P is motor power, y_PFor the motor power vertical distance away from the center of gravity of airplane, C_LFor lift coefficient, C_DFor resistance coefficient, Cm is longitudinal pitching moment coefficient, and α is the angle of attack,For engine installation angle, Phi is the first horizontal tail setting angle, and q is ram compression.

Preferably, the verification method in described step 3 specifically, be divided into the section of departing, Climb Enroute section, cruise section, normal decline section and approach segment by the flight of whole horizontal tail fastening aircraft, and wherein, the section of departing uses equation below to verify:

$\left\{\begin{array}{c}G-{\mathrm{qSC}}_L(\mathrm{\α},{\mathrm{\δ}}_e,Phi)-Psin(\mathrm{\α}+{\mathrm{\φ}}_p)=0\\ Pcos(\mathrm{\α}+{\mathrm{\φ}}_p)-{\mathrm{qSC}}_D(\mathrm{\α},{\mathrm{\δ}}_e,Phi)=0\\ {\mathrm{qSc}}_A{C}_m(\mathrm{\α},{\mathrm{\δ}}_e,Phi)-{\mathrm{Py}}_p=0\end{array}\right.;$

Climb Enroute section, cruise section, normal decline section and approach segment verified by equation below:

$\left\{\begin{array}{c}G-{\mathrm{qSC}}_L(\mathrm{\α},{\mathrm{\δ}}_e,Phi)-P\sin (\mathrm{\α}+{\mathrm{\φ}}_p)-R=0\\ {\mathrm{qSc}}_A{C}_m(\mathrm{\α},{\mathrm{\δ}}_e,Phi)-R(x+{\mathrm{fy}}_g)-{\mathrm{Py}}_p=0\end{array}\right.$

Verify the δ in above-mentioned formula_eMeet following condition:

δe_zmax≤δ_e≤δe_fmax；Wherein,

δ_eFor elevator drift angle；G is aircraft weight, and S is wing area, c_AFor wing mean aerodynamic chord, P is motor power, y_PFor the motor power vertical distance away from the center of gravity of airplane, C_LFor lift coefficient, C_DFor resistance coefficient, Cm is longitudinal pitching moment coefficient, and α is the angle of attack,For engine installation angle, Phi is the first horizontal tail setting angle, and q is ram compression；, μ is the relative density of aircraft in lengthwise movement；δe_zmaxFor elevator positively biased maximum, δ e_fmaxFor elevator negative bias maximum.

Preferably, described step 4 particularly as follows:

Pass through formula δ_e=δ_e1+δ_e2Obtain the elevator degree of bias, δ_e2Obtained by equation below:

${\mathrm{\δ}}_{e2}=(N-1)\×(-\frac{C_{\mathrm{Lpf}}}{C_{m{C}_L=c}^{{\mathrm{\δ}}_e}}(C_m^{C_L}+\frac{C_m^{\stackrel{\‾}{q}}}{\mathrm{\μ}}));$

Wherein, δ_e1Pass through formula:Obtain；

Wherein, for phi assignment, and in the range of desin speed, choose multiple speed point, thus obtain different δ_e, by each δ_eIt is fitted by statistical method, thus chooses and meet pre-conditioned δ_e, meet described pre-conditioned δ_eCorresponding phi is the second horizontal tail setting angle；

In above-mentioned formula: δ_eFor elevator drift angle；G is aircraft weight, and S is wing area, c_AFor wing mean aerodynamic chord, P is motor power, y_PFor the motor power vertical distance away from the center of gravity of airplane, C_LFor lift coefficient, C_DFor resistance coefficient, Cm is longitudinal pitching moment coefficient, and α is the angle of attack,For engine installation angle, Phi is the first horizontal tail setting angle, and q is ram compression；, μ is the relative density of aircraft in lengthwise movement；δe_zmaxFor elevator positively biased maximum, δ e_fmaxFor elevator negative bias maximum；-the pitching moment coefficient derivative to the angle of attack,For slope of lift curve,C_LpfFor putting down lift coefficient when flying,It is elevator efficiency during constant for lift coefficient；δ_e1For the balance rudder face degree of bias；δ_e2On the basis of trim, rudder face degree of bias increment when making motor-driven.

Preferably, described pre-conditioned in described step 4 is:

Preferably, the formula in described step 5 particularly as follows:

Phi=k1* the first horizontal tail setting angle+k2* the second horizontal tail setting angle；

K1 flies weight coefficient for flat, and k2 is overload weight coefficient, and span is:

0≤k1≤10≤k2≤1k1+k2=1, wherein, K1 accounting is 0-0.8, and described K2 accounting is 0-0.2.

The horizontal tail setting angle of the horizontal tail fastening aircraft in the present invention determines that method is capable of determining that suitable horizontal tail established angle, it is thus possible to guaranteeing that elevator is on the premise of full mission profile meets design requirement, alleviate driver's burden in design optimization section, and improve fore-and-aft control quality.This method is short and sweet, and application is good, workable.

Accompanying drawing explanation

Fig. 1 is the schematic flow sheet that the horizontal tail setting angle of horizontal tail fastening aircraft according to an embodiment of the invention determines method.

Fig. 2 is pre-conditioned middle positive skewness curve values and the trendgram of negative bias degree curve values in embodiment illustrated in fig. 1.

Detailed description of the invention

Clearer for the purpose making the present invention implement, technical scheme and advantage, below in conjunction with the accompanying drawing in the embodiment of the present invention, the technical scheme in the embodiment of the present invention is further described in more detail.In the accompanying drawings, the most same or similar label represents same or similar element or has the element of same or like function.Described embodiment is a part of embodiment of the present invention rather than whole embodiments.The embodiment described below with reference to accompanying drawing is exemplary, it is intended to is used for explaining the present invention, and is not considered as limiting the invention.Based on the embodiment in the present invention, the every other embodiment that those of ordinary skill in the art are obtained under not making creative work premise, broadly fall into the scope of protection of the invention.Below in conjunction with the accompanying drawings embodiments of the invention are described in detail.

In describing the invention; it will be appreciated that; term " orientation or the position relationship of the instruction such as " center ", " longitudinally ", " laterally ", "front", "rear", "left", "right", " vertically ", " level ", " top ", " end " " interior ", " outward " they be based on orientation shown in the drawings or position relationship; be for only for ease of the description present invention and simplifying and describe; rather than instruction or imply the device of indication or element must have specific orientation, with specific azimuth configuration and operation, therefore it is not intended that limiting the scope of the invention.

Fig. 1 is the schematic flow sheet that the horizontal tail setting angle of horizontal tail fastening aircraft according to an embodiment of the invention determines method.Fig. 2 is pre-conditioned middle positive skewness curve values and the trendgram of negative bias degree curve values in embodiment illustrated in fig. 1.

The horizontal tail setting angle of horizontal tail fastening aircraft as shown in Figure 1 determines that method comprises the steps:

Step 1: fly trim ability determine the first design optimization point according to the longitudinal direction of horizontal tail fastening aircraft is flat；

Step 2: obtain the aircraft parameter under the first design optimization point, and calculated the first horizontal tail setting angle of this first design optimization point by formula；

Step 3: this first horizontal tail setting angle is brought in whole horizontal tail fastening aircraft flight course, checking elevator under the first horizontal tail setting angle this described, the normal flight that whether ensure that described horizontal tail fastening aircraft in the range of deflection angles of himself of described elevator；The most then carry out next step；If it is not, then repeating said steps 1 is to described step 2, until the result in described step 3 is yes；Specifically, optimizing first and take increment near point, change the first optimization point, repetition step 1 is to step 2, until the result in step 3 is yes；

Step 4: according to the longitudinal maneuver ability of horizontal tail fastening aircraft and calculate the second horizontal tail setting angle by formula；

Step 5: select an actual horizontal tail setting angle as described horizontal tail fastening aircraft by formula between described first horizontal tail setting angle and described second horizontal tail setting angle.

In the present embodiment, the aircraft parameter in described step 2 includes aircraft weight center of gravity, flying height, flight speed, engine condition, flap state.

In the present embodiment, the formula in described step 2 is:

Wherein,

δ_eFor elevator drift angle, this formula takes O；G is aircraft weight, and S is wing area, c_AFor wing mean aerodynamic chord, P is motor power, y_PFor the motor power vertical distance away from the center of gravity of airplane, C_LFor lift coefficient, C_DFor resistance coefficient, Cm is longitudinal pitching moment coefficient, and α is the angle of attack,For engine installation angle, Phi is the first horizontal tail setting angle, and q is ram compression.

In the present embodiment,

Verification method in described step 3 specifically, be divided into the section of departing, Climb Enroute section, cruise section, normal decline section and approach segment by the flight of whole horizontal tail fastening aircraft, and wherein, the section of departing uses equation below to verify:

$\left\{\begin{array}{c}G-{\mathrm{qSC}}_L(\mathrm{\α},{\mathrm{\δ}}_e,Phi)-Psin(\mathrm{\α}+{\mathrm{\φ}}_p)=0\\ Pcos(\mathrm{\α}+{\mathrm{\φ}}_p)-{\mathrm{qSC}}_D(\mathrm{\α},{\mathrm{\δ}}_e,Phi)=0\\ {\mathrm{qSc}}_A{C}_m(\mathrm{\α},{\mathrm{\δ}}_e,Phi)-{\mathrm{Py}}_p=0\end{array}\right.;$

Climb Enroute section, cruise section, normal decline section and approach segment verified by equation below:

$\left\{\begin{array}{c}G-{\mathrm{qSC}}_L(\mathrm{\α},{\mathrm{\δ}}_e,Phi)-P\sin (\mathrm{\α}+{\mathrm{\φ}}_p)-R=0\\ {\mathrm{qSc}}_A{C}_m(\mathrm{\α},{\mathrm{\δ}}_e,Phi)-R(x+{\mathrm{fy}}_g)-{\mathrm{Py}}_p=0\end{array}\right.$

Verify the δ in above-mentioned formula_eMeet following condition:

δe_zmax≤δ_e≤δe_fmax；Wherein,

δ_eFor elevator drift angle；G is aircraft weight, and S is wing area, c_AFor wing mean aerodynamic chord, P is motor power, y_PFor the motor power vertical distance away from the center of gravity of airplane, C_LFor lift coefficient, C_DFor resistance coefficient, Cm is longitudinal pitching moment coefficient, and α is the angle of attack,For engine installation angle, Phi is the first horizontal tail setting angle, and q is ram compression；, μ is the relative density of aircraft in lengthwise movement；δe_zmaxFor elevator positively biased maximum, δ e_fmaxFor elevator negative bias maximum.

In this embodiment, described step 4 particularly as follows:

Pass through formula δ_e=δ_e1+δ_e2Obtain the elevator degree of bias, δ_e2Obtained by equation below:

${\mathrm{\δ}}_{e2}=(N-1)\×(-\frac{C_{\mathrm{Lpf}}}{C_{m{C}_L=c}^{{\mathrm{\δ}}_e}}(C_m^{C_L}+\frac{C_m^{\stackrel{\‾}{q}}}{\mathrm{\μ}}));$

Wherein, δ_e1Pass through formula:Obtain；

Wherein, for phi assignment, and in the range of desin speed, choose multiple speed point, thus obtain different δ_e, by each δ_eIt is fitted by statistical method, thus chooses and meet pre-conditioned δ_e, meet described pre-conditioned δ_eCorresponding phi is the second horizontal tail setting angle；

In above-mentioned formula: δ_eFor elevator drift angle；G is aircraft weight, and S is wing area, c_AFor wing mean aerodynamic chord, P is motor power, y_PFor the motor power vertical distance away from the center of gravity of airplane, C_LFor lift coefficient, C_DFor resistance coefficient, Cm is longitudinal pitching moment coefficient, and α is the angle of attack,For engine installation angle, Phi is the first horizontal tail setting angle, and q is ram compression；, μ is the relative density of aircraft in lengthwise movement；δe_zmaxFor elevator positively biased maximum, δ e_fmaxFor elevator negative bias maximum；-the pitching moment coefficient derivative to the angle of attack,For slope of lift curve,C_LpfFor putting down lift coefficient when flying,It is elevator efficiency during constant for lift coefficient；δ_e1For the balance rudder face degree of bias；δ_e2On the basis of trim, rudder face degree of bias increment when making motor-driven.

Seeing Fig. 2, described pre-conditioned in described step 4 is:

Specifically, the formula left sideAs in figure 2 it is shown, on the right of formulaFor given value.

When these formula both sides are unequal, change phi institute assigned value, thus finally make this formula restrain.This value is the second horizontal tail setting angle.

In the present embodiment, the formula in described step 5 particularly as follows:

Phi=k1* the first horizontal tail setting angle+k2* the second horizontal tail setting angle；

K1 flies weight coefficient for flat, and k2 is overload weight coefficient, and span is:

0≤k1≤10≤k2≤1k1+k2=1, wherein, K1 accounting is 0-0.8, and described K2 accounting is 0-0.2.

The horizontal tail setting angle of the horizontal tail fastening aircraft in the present invention determines that method is capable of determining that suitable horizontal tail established angle, it is thus possible to guaranteeing that elevator is on the premise of full mission profile meets design requirement, alleviate driver's burden in design optimization section, and improve fore-and-aft control quality.This method is short and sweet, and application is good, workable.

Last it is noted that above example is only in order to illustrate technical scheme, it is not intended to limit.Although the present invention being described in detail with reference to previous embodiment, it will be understood by those within the art that: the technical scheme described in foregoing embodiments still can be modified by it, or wherein portion of techniques feature is carried out equivalent；And these amendments or replacement, do not make the essence of appropriate technical solution depart from the spirit and scope of various embodiments of the present invention technical scheme.

Claims (7)

1. the horizontal tail setting angle of a horizontal tail fastening aircraft determines method, it is characterised in that the horizontal tail setting angle of described horizontal tail fastening aircraft determines that method comprises the steps:

Step 1: fly trim ability determine the first design optimization point according to the longitudinal direction of horizontal tail fastening aircraft is flat；

Step 2: obtain the aircraft parameter under the first design optimization point, and calculated the first horizontal tail setting angle of this first design optimization point by formula；

Step 3: this first horizontal tail setting angle is brought in whole horizontal tail fastening aircraft flight course, checking elevator under the first horizontal tail setting angle this described, the normal flight that whether ensure that described horizontal tail fastening aircraft in the range of deflection angles of himself of described elevator；The most then carry out next step；If it is not, then repeating said steps 1 is to described step 2, until the result in described step 3 is yes；

Step 4: according to the longitudinal maneuver ability of horizontal tail fastening aircraft and calculate the second horizontal tail setting angle by formula；

Step 5: select an actual horizontal tail setting angle as described horizontal tail fastening aircraft by formula between described first horizontal tail setting angle and described second horizontal tail setting angle.

2. the horizontal tail setting angle of horizontal tail fastening aircraft as claimed in claim 1 determines method, it is characterised in that the aircraft parameter in described step 2 includes aircraft weight center of gravity, flying height, flight speed, engine condition, flap state.

3. the horizontal tail setting angle of horizontal tail fastening aircraft as claimed in claim 2 determines method, it is characterised in that the formula in described step 2 is:

Wherein,

δ_eFor elevator drift angle, this formula takes O；G is aircraft weight, and S is wing area, c_AFor wing mean aerodynamic chord, P is motor power, y_PFor the motor power vertical distance away from the center of gravity of airplane, C_LFor lift coefficient, C_DFor resistance coefficient, Cm is longitudinal pitching moment coefficient, and α is the angle of attack,For engine installation angle, Phi is the first horizontal tail setting angle, and q is ram compression.

4. the horizontal tail setting angle of horizontal tail fastening aircraft as claimed in claim 3 determines method, it is characterised in that

Verification method in described step 3 specifically, be divided into the section of departing, Climb Enroute section, cruise section, normal decline section and approach segment by the flight of whole horizontal tail fastening aircraft, and wherein, the section of departing uses equation below to verify:

$\left\{\begin{array}{c}G-{\mathrm{qSC}}_L(\mathrm{\α},{\mathrm{\δ}}_e,Phi)-Psin(\mathrm{\α}+{\mathrm{\φ}}_p)=0\\ Pcos(\mathrm{\α}+{\mathrm{\φ}}_p)-{\mathrm{qSC}}_D(\mathrm{\α},{\mathrm{\δ}}_e,Phi)=0\\ {\mathrm{qSc}}_A{C}_m(\mathrm{\α},{\mathrm{\δ}}_e,Phi)-{\mathrm{Py}}_p=0\end{array}\right.;$

Climb Enroute section, cruise section, normal decline section and approach segment verified by equation below:

$\left\{\begin{array}{c}G-{\mathrm{qSC}}_L\left(\mathrm{\α},{\mathrm{\δ}}_e,Phi\right)-P\sin \left(\mathrm{\α}+{\mathrm{\φ}}_p\right)-R=0\\ {\mathrm{qSc}}_A{C}_m\left(\mathrm{\α},{\mathrm{\δ}}_e,Phi\right)-R\left(x+{\mathrm{fy}}_g\right)-{\mathrm{Py}}_p=0\end{array}\right.$

Verify the δ in above-mentioned formula_eMeet following condition:

δe_zmax≤δ_e≤δe_fmax；Wherein,

δ_eFor elevator drift angle；G is aircraft weight, and S is wing area, c_AFor wing mean aerodynamic chord, P is motor power, y_PFor the motor power vertical distance away from the center of gravity of airplane, C_LFor lift coefficient, C_DFor resistance coefficient, Cm is longitudinal pitching moment coefficient, and α is the angle of attack,For engine installation angle, Phi is the first horizontal tail setting angle, and q is ram compression；, μ is the relative density of aircraft in lengthwise movement；δe_zmaxFor elevator positively biased maximum, δ e_fmaxFor elevator negative bias maximum.

5. the horizontal tail setting angle of horizontal tail fastening aircraft as claimed in claim 4 determines method, it is characterised in that described step 4 particularly as follows:

Pass through formula δ_e=δ_e1+δ_e2Obtain the elevator degree of bias, δ_e2Obtained by equation below:

${\mathrm{\δ}}_{e2}=(N-1)\×(-\frac{C_{Lpf}}{C_{{\mathrm{mC}}_L=c}^{{\mathrm{\δ}}_e}}(C_m^{C_L}+\frac{C_m^{\stackrel{\‾}{q}}}{\mathrm{\μ}}\left)\right);$

Wherein, δ_e1Pass through formula:Obtain；

Wherein, for phi assignment, and in the range of desin speed, choose multiple speed point, thus obtain different δ_e, by each δ_eIt is fitted by statistical method, thus chooses and meet pre-conditioned δ_e, meet described pre-conditioned δ_eCorresponding phi is the second horizontal tail setting angle；

In above-mentioned formula: δ_eFor elevator drift angle；G is aircraft weight, and S is wing area, c_AFor wing mean aerodynamic chord, P is motor power, y_PFor the motor power vertical distance away from the center of gravity of airplane, C_LFor lift coefficient, C_DFor resistance coefficient, Cm is longitudinal pitching moment coefficient, and α is the angle of attack,For engine installation angle, Phi is the first horizontal tail setting angle, and q is ram compression；, μ is the relative density of aircraft in lengthwise movement；δe_zmaxFor elevator positively biased maximum, δ e_fmaxFor elevator negative bias maximum；-the pitching moment coefficient derivative to the angle of attack,For slope of lift curve,C_LpfFor putting down lift coefficient when flying,It is elevator efficiency during constant for lift coefficient；δ_e1For the balance rudder face degree of bias；δ_e2On the basis of trim, rudder face degree of bias increment when making motor-driven.

6. the horizontal tail setting angle of horizontal tail fastening aircraft as claimed in claim 5 determines method, it is characterised in that described pre-conditioned in described step 4 is:

7. the horizontal tail setting angle of horizontal tail fastening aircraft as claimed in claim 6 determines method, it is characterised in that formula in described step 5 particularly as follows:

Phi=k1* the first horizontal tail setting angle+k2* the second horizontal tail setting angle；

K1 flies weight coefficient for flat, and k2 is overload weight coefficient, and span is:

0≤k1≤10≤k2≤1k1+k2=1, wherein, K1 accounting is 0-0.8, and described K2 accounting is 0-0.2.