CN102901613A - Method for determining pressure center of reentry vehicle - Google Patents

Abstract

The invention discloses a method for determining the pressure center of a reentry vehicle. The method comprises the following steps of: acquiring a normal force coefficient CN=CN (alpha) and an axial force coefficient CA=CA (alpha), which are changed along with an angle of attack, of the reentry vehicle, and a pitching moment coefficient Mz=Mz (alpha) relative to the center of mass of the reentry vehicle by utilizing data of a wind tunnel experiment; then acquiring differential coefficients CN', CA' and Mz' of the CN, CA and Mz relative to the angle of attack alpha; and acquiring coordinates (x, y) of the pressure center under an elastic body shaft system by using the formula as shown in the specification. In the coordinates of the pressure center under a structural coordinate system, Xr is the distance between the vertex of the head of the reentry vehicle and the center of mass of the reentry vehicle, and Lr is the reference length of the reentry vehicle. By the method, the position of the pressure center of the small-slenderness-ratio reentry vehicle can be accurately obtained.

Description

Method is determined in a kind of reentry vehicle Center of Pressure

Technical field

The present invention relates to a kind of aircraft Center of Pressure and determine method, be used for wind tunnel test and accurately obtain little slenderness ratio reentry vehicle centre-of-pressure position.

Background technology

Aircraft is subject to the effect of aerodynamic force awing, and wherein normal force and axial force all can cause the pitching moment to the aircraft barycenter.For conventional high-fineness ratio profile aircraft such as guided missiles, axial force itself is just than the little magnitude of normal force or more, and the arm of force of axial force is more much smaller than the arm of force of normal force, axial force is less important to the contribution of pitching moment, so be defined as the intersection point that aircraft is subject to combined air force and the longitudinal axis awing for the Center of Pressure (the pressure heart) of the high-fineness ratio normal arrangement aircraft such as guided missile.During wind tunnel test, the normal arrangement aircraft centre of pressure coefficient

Can carry out approximate treatment by following formula.

Wherein, CN is the normal force coefficient of aircraft, and Mz is the pitching moment coefficient of the aircraft barycenter being got square, Xr is the distance that the aircraft barycenter is arrived on the Vehicle nose summit, and Lr is the aircraft reference length, δ _eElevating rudder drift angle, α for aircraft _tThe angle of attack for aircraft.

Can find out that from this computing formula the size of Center of Pressure depends on pitching moment coefficient Mz and normal force coefficient CN, when the axial force coefficient CA of aircraft to the contribution of pitching moment coefficient Mz hour, the method can be similar to the position that obtains its Center of Pressure.

For little slenderness ratio reentry vehicle, axial force is generally greater than normal force during Low Angle Of Attack, axial force is larger to the contribution of pitching moment, and the increase along with the angle of attack, axial force increases gradually to the contribution of pitching moment, and greater than the contribution of normal force to pitching moment, at this moment axial force is main to the contribution of moment, this situation is still calculated with said method and is pressed the heart, and then error is larger.Because lift-drag ratio is smaller, more than press heart define method obviously no longer applicable for the reentry vehicle profile, the aircraft pressure heart is the application point that aircraft is subject to combined air force awing strictly.And in conventional dynamometry wind tunnel test, the flight force and moment of balance measurement aircraft can only obtain the make a concerted effort size of size, direction and this arm of force d that makes a concerted effort of F of axial force and normal force, as shown in Figure 1, and can't accurately obtain the Center of Pressure of aircraft.

Summary of the invention

Technical matters to be solved by this invention provides a kind of reentry vehicle Center of Pressure and determines method, can accurately obtain the centre-of-pressure position of little slenderness ratio reentry vehicle.

The present invention includes following technical scheme:

Method is determined in a kind of reentry vehicle Center of Pressure, comprises the steps:

(1) utilize wind tunnel test data acquisition reentry vehicle with the normal force coefficient CN=CN (α) of angle of attack variation, axial force coefficient CA=CA (α), and pitching moment coefficient Mz=Mz (α);

(2) the normal force coefficient CN with angle of attack variation that obtains according to step (1), axial force coefficient CA, and pitching moment coefficient Mz obtains CN, CA, Mz about derivative CN ', CA ', the Mz ' of angle of attack;

(3) utilize following formula to obtain the pressure-acting point at the coordinate (x, y) of body axle system;

$\left\{\begin{array}{c}x=\frac{{\mathrm{Mz}}^{\′}}{{\mathrm{CN}}^{\′}-\frac{\mathrm{CN}}{\mathrm{CA}}{C}^{\′}}\\ y=\frac{\mathrm{Mz}}{\mathrm{CA}}-\frac{\mathrm{CN}}{\mathrm{CA}}\frac{{\mathrm{Mz}}^{\′}}{{\mathrm{CN}}^{\′}-\frac{\mathrm{CN}}{\mathrm{CA}}{\mathrm{CA}}^{\′}}\end{array}\right.$

(4) obtain the coordinate (x, y) of reentry vehicle Center of Pressure according to step (3); Structure coordinate is that the coordinate at downforce center is

Wherein

Xr is the distance that the aircraft barycenter is arrived on reentry vehicle head summit, and Lr is the reference length of aircraft.

In wind tunnel test, allow the angle of attack of reentry vehicle change continuously, obtain flow parameter and the wind-tunnel balance output signal of the wind-tunnel corresponding with each angle of attack;

Obtain incoming flow dynamic pressure q according to flow parameter _∞, obtain the axial force X that aircraft is subject to according to the balance output signal, normal force Y, and to the pitching moment MZ of aircraft barycenter;

According to following formula reference axis to force coefficient CA, normal force coefficient CN with the aircraft barycenter is got the pitching moment coefficient Mz of square;

$\mathrm{CA}=\frac{X}{q_{\∞}\·S_r},$ $\mathrm{CN}=\frac{Y}{q_{\∞}\·S_r},$ $\mathrm{Mz}=\frac{\mathrm{MZ}}{q_{\∞}\·S_r\·L_r}.$

The present invention compared with prior art has following advantage:

(1) the present invention can accurately obtain the particular location of aircraft barycenter, and prior art can't obtain the particular location that aircraft is pressed the heart;

(2) the present invention has considered normal force, axial force to pressing the impact of heart position, and is more scientific and reasonable than prior art.

Description of drawings

Fig. 1 is that aircraft is subjected to making a concerted effort of aerodynamic force and confrontation mental and physical efforts arm synoptic diagram;

Fig. 2 is aircraft stressing conditions synoptic diagram under missile coordinate system and structure coordinate system;

Fig. 3 is typical angle of attack control curve;

Fig. 4 is that aircraft presses the heart with angle of attack variation position movement track synoptic diagram.

Embodiment

Below just by reference to the accompanying drawings the present invention is done further introduction.

In the wind tunnel test, metering system with static state, wind-tunnel balance is measured the flight force and moment of aircraft, can only obtain size that axial force and normal force make a concerted effort, direction and this with joint efforts to the size of the aircraft barycenter arm of force, and can't obtain the position that spot of resultant force is namely pressed the heart.The present invention is from dynamic viewpoint, and when aircraft changed continuously at the angle of attack, its suffered flight force and moment was continually varying, and then its Center of Pressure also is continually varying.Aircraft is when angle of attack variation, and the direction that its suffered aerodynamic force is made a concerted effort in axon system also can change, and namely aerodynamic force changes with angle of attack variation in the body axis coordinate system with joint efforts.

Aircraft is pressed the computing formula of the heart to derive by the following method and is drawn:

When if Aircraft Angle of Attack is α, incoming flow Mach 2 ship M, as shown in Figure 2, pressure-acting point (Center of Pressure) is (x, y) at the lower coordinate of body axis coordinate system (g-xy), has

Mz＝CN·x+CA·y (1)

Then when Aircraft Angle of Attack is (α+Δ α)

Mz+ΔMz＝(CN+ΔCN)·(x+Δx)+(CA+ΔCA)·(y+Δy)

Thereby

ΔMz＝CN·Δx+ΔCN·x+ΔCN·Δx+CA·Δy+ΔCA·y+ΔCA·Δy

The following formula both sides simultaneously just with Δ α, and are got the limit to Δ α → 0 and can be got

Mz′＝CN·x′+CN′·x+CA·y′+CA′·y (2)

Wherein, Mz ', CN ', CA ', x ', y ' are respectively Mz, CN, CA, x, y about the derivative of angle of attack a.

If y=kx+b, wherein k=-CN/CA

Y '=kx ' then

Bring (2) Shi Kede into

${\mathrm{Mz}}^{\′}=(C{N}^{\′}-\frac{\mathrm{CN}}{\mathrm{CA}}{\mathrm{CA}}^{\′})\·x$

Thereby

$\left\{\begin{array}{c}x=\frac{{\mathrm{Mz}}^{\′}}{{\mathrm{CN}}^{\′}-\frac{\mathrm{CN}}{\mathrm{CA}}{\mathrm{CA}}^{\′}}\\ y=\frac{\mathrm{Mz}}{\mathrm{CA}}-\frac{\mathrm{CN}}{\mathrm{CA}}\frac{{\mathrm{Mz}}^{\′}}{{\mathrm{CN}}^{\′}-\frac{\mathrm{CN}}{\mathrm{CA}}{\mathrm{CA}}^{\′}}\end{array}\right.---\left(3\right)$

Under the structure coordinate system (O-XY), the coordinate of Center of Pressure is in Fig. 2

Wherein Xr is the distance that the aircraft barycenter is arrived on reentry vehicle head summit, and Lr is reference length.

By obtaining aircraft with angle of attack continually varying aerodynamic force (moment) coefficient by measuring balance in the wind tunnel test test, such as normal force coefficient CN=CN (α), axial force coefficient CA=CA (α) is to the pitching moment coefficient Mz=Mz (α) of aircraft barycenter.By top hypothesis as can be known CN (α), CA (α) and Mz (α) be continuous function, if it is smooth curve at a certain angle of attack minor function curve, then corresponding coefficient exists the derivative of angle of attack, is had by the definition of derivative

$f^{\′}=\underset{\mathrm{\Δ\α}\→0}{\lim }\frac{f(\mathrm{\α}+\mathrm{\Δ\α})-f\left(\mathrm{\α}\right)}{\mathrm{\Δ\α}}$

CN then, CA, MZ, x, y about the derivative of α all can in the hope of.

According to Center of Pressure coordinate formula (3) as can be known, except measuring aircraft the aerodynamic force under each attitude, also need to obtain aircraft under each attitude aerodynamic force with the derivative of angle of attack variation.In conventional test, attack angle mechanism adopts ladder erect-position version, and the angle of attack number of measurement is less, and it is larger with the error of angle of attack variation derivative to try to achieve aerodynamic force with this, thereby the error that obtains the Center of Pressure is also larger.Therefore need to solve aerodynamic force and accurately measure problem with the angle of attack variation derivative.

Raising by the wind tunnel test measuring technology, when carrying out permanent dynamometer check, adopt the angle of attack to change continuously and the real time data acquisition experimental technique, this technology can increase considerably the quantity of test figure, accurately reflect studied object aerodynamic characteristic with the rule of angle of attack variation, thereby measure accurate derivative.This experimental technique comprises that mainly Flow Field in Wind Tunnel control, the angle of attack change control and three aspects such as real time data acquisition and processing continuously.During test, system's control wind-tunnel starts, after the flow field was set up and be stable, angle of attack control system began to control attack angle mechanism deflection, when mechanism arrives given angle of attack starting point and begins at the uniform velocity to change, data acquisition system (DAS) begins flow parameter, the angle of attack and balance output by the collection of corresponding sequential real-time synchronization, until attack angle mechanism arrives angle of attack scope terminal point, mechanism's fly back, wind-tunnel cut-offs, data acquisition system (DAS) gathers end zero, off-test.For the supersonic wind tunnel test, when the wind-tunnel blockage ratio meets the demands, as long as angle of attack variation speed is fit to, just can obtain stable flow field.Therefore, the research of experimental technique mainly concentrates on the angle of attack and changes continuously control and real time data acquisition and processing two aspects.

1. the continuous change control method of the angle of attack

Because test remains static wind tunnel test, requires angle of attack variation speed suitable, it is can be because the angle of attack changes fast inconsistent during with angle of attack stepped change to guarantee that model streams not, and test findings can not brought unsteady aerodynamic characteristic into.Simultaneously, the too fast meeting of angle of attack variation makes the disturbance of model and supporting mechanism stream field can not get recovering, and can not obtain stable flow field.Therefore, obtaining suitable angle of attack variation speed is the key of experimental technique success.In addition, for avoiding inertial force that the model acceleration and deceleration cause on the impact of measurement result, require that angle of attack variation is at the uniform velocity in the test angle of attack scope.

For this reason, respectively added one section acceleration (deceleration) section in the front and back of the required conversion range of angle of attack was of test, to guarantee that angle of attack pace of change is at the uniform velocity in the required angle of attack scope.Common angle of attack variation mode is: attack angle mechanism rapidly moves to maximum angle of attack point after starting, and deceleration stops, then oppositely accelerate, and before required angle of attack starting point, reach required speed, then begin uniform motion, until required end of extent (EOE) point beginning mechanism returns zero.Fig. 3 is typical angle of attack control curve.

2. real time data acquisition and treatment technology

The angle of attack changes experimental technique continuously needs data acquisition system (DAS) to gather the tunnel airstream parameter by real-time synchronization, the mass data such as angle of attack value and the output of balance signal, and this needs system hardware and software that enough response speeds and sample frequency are arranged.

Wind-tunnel force balance commonly used has higher frequency response characteristic usually, can satisfy the test demand.For improving the data acquisition system (DAS) response characteristic, adopt the HBM MG Cplus of company data acquisition system (DAS) to carry out data acquisition.This system can satisfy the collection of high-precision analog signal and high speed, high number of divisions word signal, and the digital measurement rate is up to 19200 value/second/passages.

For satisfying test mass data acquisition and processing needs, Data Acquisition ﹠ Processing Software needs to gather and to record continuously, in real time flow parameter, angle of attack value and the balance output signal of corresponding sequential in required angle of attack scope, the mass data under energy fast processing, demonstration and the output angle of attack change continuously.

Contrast and analysis through calibration model test and a large amount of model tests, angle of attack test findings and ladder angle of attack test findings consistance are fine continuously, satisfy the requirement of test precision, show the test figure reliable results that continuous angle of attack variation gathers, solved simultaneously the Measurement accuracy problem of aerodynamic force with the derivative of angle of attack variation.

The data that gather are handled as follows: obtain incoming flow dynamic pressure q according to flow parameter _∞, obtain the axial force X that aircraft is subject to, normal force Y, and pitching moment MZ according to the balance output signal; According to following formula reference axis to force coefficient CA, normal force coefficient CN and pitching moment coefficient Mz;

S wherein _rBe area of reference, L _rBe reference length.

The present invention uses the angle of attack to change continuously and the real time data acquisition experimental technique, can obtain abundanter test figure, presses scheming to calculate the actual characteristic that the result more meets aircraft.

Get reentry vehicle for little slenderness ratio, can obtain the particular location of aircraft pressure-acting point by the present invention, thereby solved the Accurate Prediction problem of reentry body Center of Pressure, obtained the Center of Pressure accurate location and have good reference significance with the angle of attack variation rule for mass center of reentry vehicle Position Design and Design of Attitude Control System.Be that certain aircraft presses the heart to change synoptic diagram with angle of attack as shown in Figure 4, the A point is α=-2.The time centre-of-pressure position, the B point is α=-60.The time centre-of-pressure position.

The unspecified part of the present invention belongs to general knowledge as well known to those skilled in the art.

Claims (2)

1. method is determined in a reentry vehicle Center of Pressure, comprises the steps:

(1) utilize wind tunnel test data acquisition reentry vehicle with the normal force coefficient CN=CN (α) of angle of attack variation, axial force coefficient CA=CA (α) is with the pitching moment coefficient Mz=Mz (α) of relative aircraft barycenter;

(2) the normal force coefficient CN with angle of attack variation that obtains according to step (1), axial force coefficient CA, and pitching moment coefficient Mz obtains CN, CA, Mz about derivative CN ', CA ', the Mz ' of angle of attack;

(3) utilize following formula to obtain the coordinate (x, y) of pressure-acting point under body axle system;

$\left\{\begin{array}{c}x=\frac{{\mathrm{Mz}}^{\′}}{{\mathrm{CN}}^{\′}-\frac{\mathrm{CN}}{\mathrm{CA}}{\mathrm{CA}}^{\′}}\\ y=\frac{\mathrm{Mz}}{\mathrm{CA}}-\frac{\mathrm{CN}}{\mathrm{CA}}\frac{{\mathrm{Mz}}^{\′}}{{\mathrm{CN}}^{\′}-\frac{\mathrm{CN}}{\mathrm{CA}}{\mathrm{CA}}^{\′}}\end{array}\right.$

(4) coordinate (x, y) that obtains according to step (3) obtains the coordinate of reentry vehicle Center of Pressure; The coordinate of Center of Pressure under structure coordinate system is

Wherein

Xr is the distance that the aircraft barycenter is arrived on reentry vehicle head summit, and Lr is the reference length of aircraft.

2. method is determined in reentry vehicle as claimed in claim 1 Center of Pressure, and it is characterized in that: described step (1) specifically comprises the steps:

In wind tunnel test, allow the angle of attack of reentry vehicle change continuously, obtain tunnel airstream parameter and the wind-tunnel balance output signal corresponding with each angle of attack;

Obtain incoming flow dynamic pressure q according to flow parameter _∞, obtain the axial force X that aircraft is subject to according to the balance output signal, normal force Y, and barycenter is got the pitching moment MZ of square;

According to following formula reference axis to force coefficient CA, normal force coefficient CN and pitching moment coefficient Mz; $\mathrm{CA}=\frac{X}{q_{\∞}\·S_r},$ $\mathrm{CN}=\frac{Y}{q_{\∞}\·S_r},$ $\mathrm{Mz}=\frac{\mathrm{MZ}}{q_{\∞}\·S_r\·L_r}.$

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