CN102901613B - 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 is (X<->r-x,y), X<->r=Xr/Lr, 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, for wind tunnel test, 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 aircraft barycenter.For conventional high-fineness ratio profile aircraft such as guided missiles, axial force itself is just than the little magnitude or more of normal force, 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 for the Center of Pressure (the pressure heart) of the high-fineness ratio normal arrangement aircraft such as guided missile the intersection point that aircraft is subject to combined air force and the longitudinal axis awing.During wind tunnel test, the normal arrangement aircraft centre of pressure coefficient can carry out approximate treatment by formula below.

Wherein, CN is the normal force coefficient of aircraft, and Mz gets the pitching moment coefficient of square to aircraft barycenter, xrShi Vehicle nose summit is to the distance of aircraft barycenter, and Lr is aircraft reference length, δ _eelevating rudder drift angle, α for aircraft _tthe angle of attack for aircraft.

From this computing formula, can find out, 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, during Low Angle Of Attack, axial force is generally greater than normal force, 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 is greater than the contribution of normal force to pitching moment, and at this moment axial force is main to the contribution of moment, this situation is still calculated and is pressed the heart with said method, and error is larger.For reentry vehicle profile, because lift-drag ratio is smaller, above pressure heart define method is obviously no longer applicable, and 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 axial force and normal force the make a concerted effort size, direction of F and the size of the arm of force d that should make a concerted effort, as shown in Figure 1, and cannot accurately obtain the Center of Pressure of aircraft.

Summary of the invention

Technical matters to be solved by this invention is to provide 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:

A method is determined in 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 obtaining 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 pressure-acting point at the coordinate (x, y) of body axle system;

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

(4) according to step (3), obtain the coordinate (x, y) of reentry vehicle Center of Pressure; Structure coordinate is that the coordinate at downforce center is wherein xr is that reentry vehicle head summit is to the distance of aircraft barycenter, the reference length that Lr is 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;

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

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

$\mathrm{CA}=\frac{X}{q_{\&infin;}\&CenterDot;S_r},$ $\mathrm{CN}=\frac{Y}{q_{\&infin;}\&CenterDot;S_r},$ $\mathrm{Mz}=\frac{\mathrm{MZ}}{q_{\&infin;}\&CenterDot;S_r\&CenterDot;L_r}.$

The present invention compared with prior art tool has the following advantages:

(1) the present invention can accurately obtain the particular location of aircraft barycenter, and prior art cannot 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, more scientific and reasonable compared with prior art.

Accompanying drawing explanation

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

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

Fig. 3 is that the typical angle of attack is controlled curve;

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

Embodiment

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

In wind tunnel test, with static metering system, wind-tunnel balance is measured the flight force and moment of aircraft, can only obtain size, direction and this size of making a concerted effort to the aircraft barycenter arm of force that axial force and normal force are made a concerted effort, and the application point that cannot obtain making a concerted effort presses the position of the heart.The present invention is from dynamic viewpoint, and aircraft is when the angle of attack changes continuously, and its suffered flight force and moment is continually varying, and its Center of Pressure is also 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 aerodynamic force changes with angle of attack variation with joint efforts in body axis coordinate system.

Aircraft presses the computing formula of the heart to derive and draw by the following method:

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)

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

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

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 '

Bring (2) Shi Ke get into

${\mathrm{Mz}}^{\&prime;}=(C{N}^{\&prime;}-\frac{\mathrm{CN}}{\mathrm{CA}}{\mathrm{CA}}^{\&prime;})\&CenterDot;x$

Thereby

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

In Fig. 2, under structure coordinate system (O-XY), the coordinate of Center of Pressure is wherein xr is the distance that aircraft barycenter is arrived on reentry vehicle head summit, and Lr is reference length.

In testing by wind tunnel test, by measuring balance, can obtain aircraft with angle of attack continually varying aerodynamic force (moment) coefficient, as normal force coefficient CN=CN (α), axial force coefficient CA=CA (α), the pitching moment coefficient Mz=Mz (α) to aircraft barycenter.By supposing that known CN (α), CA (α) and Mz (α) are continuous function above, if it is smooth curve at a certain angle of attack minor function curve, corresponding coefficient exists the derivative of angle of attack, by the definition of derivative, is had

$f^{\&prime;}=\underset{\mathrm{\&Delta;\&alpha;}\&RightArrow;0}{\lim }\frac{f(\mathrm{\&alpha;}+\mathrm{\&Delta;\&alpha;})-f\left(\mathrm{\&alpha;}\right)}{\mathrm{\&Delta;\&alpha;}}$

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

Known according to Center of Pressure coordinate formula (3), except will measuring the aerodynamic force of aircraft 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, tries to achieve aerodynamic force larger with the error of angle of attack variation derivative with this, thereby it is also larger to obtain the error of Center of Pressure.Therefore need to solve aerodynamic force accurately measures problem with angle of attack variation derivative.

By the raising of wind tunnel test measuring technology, when carrying out permanent dynamometer check, adopt the angle of attack to change continuously and real time data acquisition experimental technique, this technology can increase considerably the quantity of test figure, accurately reflect that studied object aerodynamic characteristic is with the rule of angle of attack variation, thereby measure accurate derivative.This experimental technique mainly comprises that Flow Field in Wind Tunnel is controlled, the angle of attack changes three aspects such as control and real time data acquisition and processing continuously.During test, system is controlled wind-tunnel and is started, after flow field is set up and stablized, angle of attack control system starts to control attack angle mechanism deflection, when mechanism arrives given angle of attack starting point and starts at the uniform velocity to change, data acquisition system (DAS) starts 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, mechanism's fly back, wind-tunnel cut-offs, data acquisition system (DAS) gathers end zero, off-test.For supersonic wind tunnel test, when wind-tunnel blockage ratio meets the demands, as long as angle of attack variation speed is applicable 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, require angle of attack variation speed suitable, guarantee that model streams not can be because the angle of attack changes fast inconsistent during with angle of attack stepped change, test findings can not brought unsteady aerodynamic characteristic into.Meanwhile, 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 that experimental technique is successfully crucial.In addition, for avoiding the impact on measurement result of inertial force that model acceleration and deceleration cause, requiring angle of attack variation within the scope of the test angle of attack is at the uniform velocity.

For this reason, in the front and back of the required conversion range of angle of attack was of test, respectively added one section of acceleration (deceleration) section, to guarantee that within the scope of the required angle of attack, angle of attack pace of change is at the uniform velocity.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 reach required speed before required angle of attack starting point, then start uniform motion, until required end of extent (EOE) point starts mechanism's back to zero.Fig. 3 is that the typical angle of attack is controlled curve.

2. real time data acquisition and treatment technology

The angle of attack changes experimental technique continuously needs data acquisition system (DAS) to gather 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 to have enough response speeds and sample frequency.

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

For meeting test mass data acquisition and processing needs, Data Acquisition & Processing Software needs within the scope of the required angle of attack, to gather and to record continuously, in real time flow parameter, angle of attack value and the balance output signal of corresponding sequential, 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, meet the requirement of test precision, show the test figure reliable results that continuous angle of attack variation gathers, solved the Measurement accuracy problem of aerodynamic force with the derivative of angle of attack variation simultaneously.

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

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

For little slenderness ratio, obtain reentry vehicle, by the present invention, can obtain the particular location of aircraft pressure-acting point, thereby solved the Accurate Prediction problem of reentry body Center of Pressure, obtained Center of Pressure accurate location and for mass center of reentry vehicle Position Design and Design of Attitude Control System, there is good reference significance with angle of attack variation rule.Be that certain aircraft presses the heart to change schematic diagram with angle of attack as shown in Figure 4, A point is α=-2.Time centre-of-pressure position, B point is α=-60.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. a method is determined in 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), with the pitching moment coefficient Mz=Mz (α) of relative aircraft barycenter;

(2) the normal force coefficient CN with angle of attack variation obtaining 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;

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

(4) coordinate (x, y) obtaining 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 that reentry vehicle head summit is to the distance of aircraft barycenter, the reference length that Lr is aircraft.

2. method is determined in reentry vehicle as claimed in claim 1 Center of Pressure, 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;

According to tunnel airstream gain of parameter incoming flow dynamic pressure q _∞, according to wind-tunnel balance output signal, obtain the axial force X that aircraft is subject to, normal force Y and barycenter is got to 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; $\mathrm{CA}=\frac{X}{q_{\&infin;}\&CenterDot;S_r},\mathrm{CN}=\frac{Y}{q_{\&infin;}\&CenterDot;S_r},\mathrm{Mz}=\frac{\mathrm{MZ}}{q_{\&infin;}\&CenterDot;S_r\&CenterDot;L_r},$ S _rfor area of reference.