RU2445599C1 - Method for determining turning angle of thrust vector of hypersonic once-through air jet engine with chamfer cut of nozzle as per results of flight tests on hypersonic flying laboratory - Google Patents

Abstract

FIELD: engines and pumps.

SUBSTANCE: in steady-state flight mode of flying laboratory (HFL), at which the sum of longitudinal aerodynamic moment and moment from engine traction force is equal to zero, pulse stepped deviation of altitude control through δ_pulse angle is set additionally, angles of deflections of altitude control and torque moments are measured in excited motion, and increment of aerodynamic moment, which is determined only with control deflection relative to its balancing value at pulse supply, is calculated. Then, as per measurement results obtained during flight tests, there calculated is derivative as per deviation angle of altitude control from aerodynamic moment coefficient; after supply of pulse there measured are current values of velocity pressure, deviation angle of altitude control and there determined is turning angle φ of thrust vector P_"двт".

EFFECT: improving the accuracy of determining the turning angle of thrust vector owing to additional measurements of parameters during short-period movement of flying laboratory around centre of mass.

3 dwg

Description

The field of technology.

The invention relates to aircraft engine building, and in particular to a new direction in it - hypersonic ramjet engines (SCJS), first of all, to determining the angle of rotation of the engine thrust force vector with oblique cut of the nozzle according to the results of scramjet flight tests at a hypersonic flying laboratory (HLL) .

The level of technology.

The modulus and angle of rotation of the thrust force vector of the scramjet engine with an oblique cut of the nozzle are the most important traction characteristics of the engines and flight performance characteristics of the HLL. A serious obstacle limiting the possibility of determining the traction characteristics of such engines on ground stands is the dependence of these characteristics on the conditions of flow around the HLL in field conditions, as well as the features of engine integration with the HLL glider: the entire lower surface of the airframe from the nose to the entrance to the scramjet is shaped as an air intake, and the aft from the exit from the combustion chamber to the bottom cut of the airframe is a nozzle of unilateral expansion. Of course, when installing the engine on another glider, not only the aerodynamic characteristics of the HLL, but also the traction characteristics of the scramjet engine change. Reproduction of real flight conditions on a stand at high speed heads and high temperatures is technically difficult to implement, requires large economic costs and is currently almost impossible.

A significant addition to the methodology for determining the thrust vector of an engine with an oblique nozzle cut introduces a change in the angle of rotation of the vector during engine operation; the deviation angle can reach 20 ÷ 25 °. The task of flight tests is to determine this angle as accurately as possible, especially at its low values.

Currently, studies are being conducted of the thrust characteristics of engines providing a deviation of the direction of the thrust force vector (I.A. Brailko, Yu.M. Klestov, S.Yu. Krasheninnikov, A.K. Mironov "Experimental and computational study of the aerodynamics of a flat rotary nozzle with a sharp a change in the contour in the throat. "TsIAM named after P.I. Baranov," Aeromechanics and Gas Dynamics ", No. 3). Nozzle studies are being conducted that could provide a rotation of the thrust force vector. One option is the geometric rotation of the nozzle, which is set by the control system and which can be accurately measured. For nozzles with flow deviation in the supersonic part, a dependence of the direction angle of the thrust force vector on the pressure drop is revealed. At small supersonic pressure drops on the nozzle, a significant excess of the effective angle of direction of the thrust force was found compared with the geometric angle of rotation of the supersonic nozzle flaps. Due to the oblique cut of the outlet nozzle and the interaction of the exhaust jet with the external flow, as well as with the lower part of the HLL housing, the jet deviates at the exit of the nozzle and, accordingly, the thrust force vector deviates from the axis oX ₁ . To date, the solution to the problem of determining the magnitude of this angle, including during flight tests of aircraft, has not been given due attention.

Close to the specified is the result set forth in patent RU No. 2122963 dated 10.12.1998, "The control system of a twin-engine aircraft by controlling the thrust vector." In this system, the signals to the stabilizers are coordinated depending on the mechanical rotation of the nozzle of the left and right engine in order to ensure stability in the longitudinal and directional control channels.

The well-known "Stand for determining the thrust vector of an engine with an oblique nozzle", patent C2 No. 2274764 dated April 20, 2006. In the proposed methodology, the thrust force vector module is determined by the readings of the force meters in the directions of two orthogonal axes. The direction of the thrust force vector is determined by geometric constructions of the actual engine test circuit (by superimposing the engine projection on the calculated thrust vector).

A significant drawback of the proposed method is that it can only be used when working on the stand. In addition, the rotation angle is determined by complex geometric constructions, which will not allow to determine it with high accuracy.

Known "The method of measuring the thrust of a jet engine in real time, patent EP No. 0342970. A2 of May 19, 1988. In this method, the total thrust of the engine is defined as the difference in total force, including the lifting force of the aircraft and the force of aerodynamic drag, taking into account angles of attack and slip To determine these components in the calculation algorithm, a large number of measured parameters are used, including overloads measured using accelerometers.

A significant drawback of this method is the large error due to the measurement of a large number of parameters in flight, including the thermodynamic parameters of a gas mixture: temperatures, isentropic parameters (k), gas constant (R), taking into account the real gas composition. In this method, the task is to determine only the modulus of the thrust force vector.

The well-known “Method of measuring traction in flight of a hypersonic ramjet engine (SCJP) of an unmanned hypersonic flying laboratory (HLL)”, patent RU No. 2242736 of December 20, 2004

In this method, the engine thrust is determined by the increment of longitudinal acceleration (along the axis oX ₁ ) created due to the action of the thrust in flight. To determine engine thrust, aerodynamic drag forces are separated along the longitudinal axis of the GLL, Earth's gravity, and engine thrust. For this purpose, the fuel supply to the combustion chamber is turned off and on at short adjacent intervals of time t ₁ and t ₂ , which do not exceed 1 second. The engine traction force is determined by the formula

R _DW = m _l · ω _R ,

where m _LL - the mass of GLL,

ω _R - GLL acceleration under the action of traction.

GLL acceleration is calculated by the formula

ω _R = (n _xt1 -n _xt2 ) g,

where n _xt1 and n _xt2 are the values of the longitudinal overloads at time t ₁ and t ₂ ,

g is the acceleration of gravity.

A significant drawback of such a measurement system with one accelerometer is the inability to determine the angle of rotation of the thrust force vector.

The closest technical solution adopted for the prototype is “A method for determining the thrust vector of a hypersonic ramjet engine (SCJP) with an oblique nozzle cut according to the results of its flight tests at a hypersonic flying laboratory (HLL)”; Patent No. 2314503 of July 18, 2004.

The technical result when using this method is achieved by measuring the overload created by the action of the engine traction force in flight in a strictly defined direction, and external trajectory measurements of relative speed. With a certain rotation of the measuring axis of the accelerometer, the component from the action of the total aerodynamic forces is excluded in its readings and only the component from the traction force remains.

In this case, a formula is given for calculating the rotation angle of the measuring axis (sensitivity axis) depending on the aerodynamic characteristics of the GLL; The specified aerodynamic characteristics of the GLL, taken as nominal, are used. The formula is obtained from the condition that the projection of the vector of the total aerodynamic force, which includes the drag force Q and the lift force Y, is equal to zero.

Calculations of the angle of rotation of the thrust force vector at small values and in the presence of large errors of accelerometers showed that it is impossible to further improve the accuracy of determining the angle of rotation on the basis of the measuring instruments indicated in the prototype. Therefore, it is necessary to use other characteristic parameters that are most sensitive to changes in the angle of rotation of the thrust force vector, and means of measuring these parameters. These parameters can be the moments of GLL rotation relative to the center of mass during rotation of the traction force vector and the elevation deflection angles to eliminate these moments.

Disclosure of the invention.

The technical result to which the invention is directed is to increase the accuracy of determining the angle of rotation of the thrust vector of a hypersonic ramjet with an oblique nozzle cut by attracting additional parameter measurements in a short-period GLL motion around the center of mass.

, измеряют в возмущенном движении углы отклонений рулей высоты и вращающие моменты, при таком испытании измеренные моменты будут определять отклонение только аэродинамического момента M_Zаэр. от балансировочного M_{Zаэр.бал.}, который уравновешен моментом от тяги двигателя. Определяют отклонение момента как разность ΔM_Zаэр.=M_Zаэр.-M_{Zаэр.бал.}, вносят коррекцию в значение ΔM_Zаэр. на величину

гдеThe specified technical result is achieved by the fact that in the method for determining the thrust vector of a hypersonic ramjet with oblique cut of the nozzle according to the results of its flight tests at a hypersonic flying laboratory, including the operation of measuring the overload on the HLL from the thrust of the engine, provided that it is equal to zero in the readings accelerometer projection of the full aerodynamic force vector, calculation of the module of the thrust force vector P _∂Вt current time, in the steady state flight mode of the HLL, at the torus the sum of the longitudinal aerodynamic moment and moment from the engine thrust force is zero, an additional impulse stepwise deviation of the elevator by an angle δ _{imp is set} , the value of which is selected based on the results of preliminary mathematical modeling so that the difference Δα = α-α ₀ between the current value of the angle attack α in the oscillation mode and its value α ₀ at the time of the pulse in modulus did not exceed the value required during the experiment

, the angles of deviations of elevators and torques are measured in a perturbed motion, in such a test the measured moments will determine the deviation of only the aerodynamic moment M _Zaer. from balancing M _Z.air.bal. which is balanced by the moment from the engine thrust. The moment deviation is determined as the difference ΔM _Zaer. = M _Zaer. -M _Zaer.bal. , make a correction in the value of ΔM _Zaer. by the amount

Where

- производная по углу атаки от коэффициента аэродинамического момента. Определяют предварительно расчетным путем и по результатам наземных стендовых испытаний координаты l и h точки приложения вектора силы тяги на срезе сопла относительно центра масс ГЛЛ, далее вычисляют приращение аэродинамического момента, обусловленное только отклонением руля относительно его балансировочного значения при подаче импульса и равное разности:

- derivative of the angle of attack from the coefficient of aerodynamic moment. The l and h coordinates of the point of application of the thrust force vector at the nozzle exit relative to the center of mass of the HLL are determined preliminarily by calculation and according to the results of ground bench tests, then the aerodynamic moment increment is calculated, which is caused only by the deviation of the steering wheel relative to its balancing value when applying a pulse and equal to the difference:

ΔM _Z ≈ΔM _Zaer. -ΔM _Zα , where

ΔM _Zaer. measured value of the aerodynamic moment in disturbed motion.

Then, according to the results of measurements obtained in flight tests, the derivative of the angle of deviation of the elevator from the coefficient of aerodynamic moment is calculated according to the formula

where Δδ = δ-δ ₀ is the difference between the measured values of the angle of deviation of the elevator δ in the perturbed movement and its balancing value δ ₀ measured when applying the pulse,

q is the measured value of the pressure head,

s, b - characteristic area and length of the product.

In the steady state, after applying the pulse, the current values of the pressure head, the elevator deflection angle are measured, and the rotation angle φ of the thrust force vector - Pt is _found as a result of solving the trigonometric equation:

where R _dt - the value of the modulus of the thrust force vector at the current time t,

which is determined by the readings of accelerometers,

δ _t , q _t - current elevator and velocity head in steady state after applying a pulse,

l and h are the coordinates of the point of application of the thrust force vector at the nozzle exit.

Thus, a technical result is achieved, the invention is aimed at: increasing the accuracy of determining the angle of rotation of the thrust force vector of the scramjet engine with an oblique cut of the nozzle.

The invention is illustrated by drawings, in which:

figure 1 shows a diagram of the distribution of forces and moments relative to the center of mass of the GLL:

1 - HLL,

2 - experimental scramjet,

3 - elevator,

4 - nozzle cut;

oX ₁ Y ₁ - the associated coordinate system,

oXY - speed coordinate system, the oX axis is rotated by the angle of attack α relative to the oX axis ₁ , the drag force Q and the lifting force Y are directed along the oX and oY axes, the relative velocity vector V is directed along the oX axis;

R _dv - engine thrust reduced to the center of mass,

R _p - reactive engine thrust;

φ is the angle of rotation of the thrust force vector;

M _aer. , M _dv - moments of rotation relative to the center of mass from the action of aerodynamic forces and engine thrust;

figure 2 shows the distribution of thrust of the engine at the nozzle exit:

1 - engine path,

2 - engine nozzle,

P ₁ / P ₂ - thrust of the engine on the upper inner and outer sections of the nozzle;

P _B - the total thrust of the engine on the upper cut of the nozzle (geometric sum of the force vectors P ₁ and P ₂ ,);

P _H - thrust of the engine at the bottom of the nozzle,

P _C is the total thrust of the engine at the nozzle exit (geometric sum of the force vectors P _B and P _H at the point of intersection),

R _p - reactive thrust;

P _DW - the total thrust of the engine, reduced to the nozzle exit (geometric sum of the force vectors P _P and P _C );

figure 3 presents as an example the results of mathematical modeling; the time variation of the parameters and characteristics in a short-period GLL motion around the center of mass is shown; the stepwise deviation of the elevator by an angle Δδ = 5 ° at t = 60 sec is set, the deviation of the traction force vector by an angle of φ = 5 °.

M _zaer. and M _zdv. (Mzar and Mzdv) - moments from aerodynamic forces and from engine traction,

ΔM _z , (dMz) is the sum of these moments,

- частная производная коэффициента момента тангажа (заданная - mzdl, вычисленная по результатам измерений (результаты математического моделирования - mzdlu),

is the partial derivative of the pitch moment coefficient (given is mzdl calculated from the measurement results (mathematical modeling results are mzdlu),

δ (delgr, in degrees) is the angle of deviation of the elevator relative to the balancing value,

α (adgr) - angle of attack,

φ (firgr) is the angle of deviation of the thrust force vector.

The proposed method for determining the thrust vector of a hypersonic ramjet engine (SCJP) (2) with an oblique cut of the nozzle according to the results of flight tests at a hypersonic flying laboratory (HLL) (1) is carried out in the following sequence:

set on the axes of rotation of the rudders of height sensors of moments and angles of deviation of the rudders;

install pressure sensors (pressure head);

along the oX ₁ and oY ₁ axes, the l and h coordinates are measured from the center of mass of the GLL to the calculated point of application of the thrust force at the nozzle exit;

according to the simulation results, approximately the time for applying the stepwise disturbance to the elevator is selected;

during the flight, a stepwise disturbance is set on the elevator;

measure the values of moments, angles of deviation of the elevators, pressure head;

в измеренное значение аэродинамического момента при отклонении угла атаки от заданного номинального;make a correction based on a given partial derivative

to the measured value of the aerodynamic moment when the angle of attack deviates from the specified nominal;

;calculate the value of the partial derivative

;

calculate the value of the angle of rotation of the thrust vector of the engine as a result of solving the trigonometric equation:

The results of calculating the angle are presented in figure 3.

The principle of separation in the readings of accelerometer projections of the overload from aerodynamic forces and engine traction and measuring instruments (accelerometers and external trajectory measurements) used in the method for determining the angle of rotation presented in the prototype does not allow to achieve a further significant increase in the accuracy of determining the angle of rotation of the vector of thrust force. Therefore, the task is to use other measured parameters that are more sensitive to changes in the angle of rotation of the thrust force vector. These parameters can be the moments of rotation around the center of mass of the aircraft and the corresponding angles of deviation of the elevators.

The dynamics of the spatial motion of an aircraft is determined by two basic laws: changes in the momentum and moment of momentum. The first of them characterizes the long-period motion of the center of mass under the action of the resultant, passing through the center of mass of the aircraft, of all external forces, the second - the short-period rotational motion around the center of mass under the action of the main moment relative to the center of mass of all external forces.

The spatial system of forces, the main vector of which is not equal to zero and not perpendicular to the main moment, can be reduced to dynamo - the totality of the force and pair, the plane of which is perpendicular to the line of action of the force. The force vector can be carried along the line of action of the force. The forces in the longitudinal plane are considered. It is advisable to choose the reduction center in the center of mass of the aircraft, in accordance with the accepted distribution of forces: aerodynamic forces can be reduced to one resultant - total aerodynamic force, and the resulting moment - the total aerodynamic moment acting relative to the center of mass of the aircraft. When the traction force vector is rotated, the system of forces can be reduced to the same indicated system. Additionally, the moment arising when the force vector is deflected by an angle φ must be compensated by the aerodynamic moment due to the deviation of the elevator in order to fulfill the balancing condition at a given angle of attack: M _dw = M _aer .

The specified system of forces and moments is the basis of the algorithm for determining the angle of rotation of the thrust force vector scramjet, see figure 1.

The main essence of the problem lies in the fact that it is necessary to find a way to separate the components in the sensor’s readings, depending on the aerodynamic forces and on the rotation of the engine’s traction force.

This result is achieved, first of all, due to the fact that in the steady state flight mode of the HLL the sum of the longitudinal aerodynamic moment and the moment of the engine thrust is zero. When the elevator deviates from this equilibrium balancing value, a moment arises which in such a test will contain only the aerodynamic component of the moment. From the measured values of the moment and the angle of deviation of the rudders, you can calculate the aerodynamic moment coefficient and then the total aerodynamic moment, which in steady state was balanced by the moment from the engine.

As can be seen from figure 1, in the case of selecting the point of supply of forces in the center of mass in the absence of rotation of the vector of traction force, a moment is created (in Fig. 1, it is converging), which is balanced by the aerodynamic moment at the nominal balancing value of the elevator. If the thrust force vector deviates up from the zero position, a diving moment will be created.

The system of equations of angular motion has the form:

,где

,Where

- момент от аэродинамических сил;

- moment from aerodynamic forces;

- приведенный момент от аэродинамических сил;

- reduced moment from aerodynamic forces;

- приведенный момент от силы тяги двигателя;

- reduced moment of engine traction;

ω _x , ω _y , ω _z - the angular velocity of the GLL rotation relative to the associated axes;

I _x , I _y , I _z - moments of inertia of the GRL relative to the associated axes;

υ is the pitch angle;

- производные коэффициента аэродинамического момента соответственно по углу атаки, углу отклонения руля и по угловой скорости;

- derivatives of the aerodynamic moment coefficient, respectively, with respect to the angle of attack, steering angle and angular velocity;

taken b = 4.39 m ² , I _z = 1400 kg m ² .

The equation of moment from aerodynamic forces can be represented as:

где

Where

α _δ - balancing angle of attack.

From this formula it follows that in the steady state with α = α _{δ the} increment of the aerodynamic moment will correspond to such a deviation of the steering wheel with respect to the balancing value, which will balance the moment from the engine traction force due to the rotation of the traction force vector relative to the calculated nominal value. In the magnitude of the moment from the engine thrust, the moment component is considered, due to the rotation of the thrust force vector, and the moment component _{ΔM zdv.} = -P _dv. h, equal to the magnitude of the converting moment in the absence of rotation of the thrust force vector. The direction in the absence of rotation of the vector is taken as the nominal design. It may not coincide with the longitudinal axis of the engine and will depend on the distribution of forces at the nozzle exit. In the general case, all forces can be reduced to a dynamo in the center of mass of the GLL. The distribution of forces on the oblique section of the nozzle is shown in figure 2.

This figure shows the calculated values of the forces with the number M = 7 and the angle of attack α = 4 ° for one of the traction characteristics;

here R _N and R _In - the total forces on the lower and upper surfaces of the nozzle,

P _R - reactive force.

при расчетном балансировочном положении руля высоты, т.е. при отсутствии поворота вектора силы тяги сумма момента аэродинамических сил и указанного кабрирующего момента от силы тяги двигателя будет равна нулю. В установившемся режиме дополнительное отклонение руля высоты от указанного балансировочного значения будет зависеть практически только от угла поворота вектора силы тяги.As a result of geometrical addition of the indicated forces P _Н and P _В, we obtain the total value of the force at the nozzle exit - Р _С , and the total value of the thrust force P _dv. - adding the result of forces F _C and F _R. Along the line of action, the total thrust is reduced to the nozzle exit. Through the obtained point in the direction parallel to the axis ox ₁ passes the traction force vector at zero rotation angle. Component of the moment ΔM _zdv. = -P _dv. h, due to the location of the engine with an oblique cut of the nozzle under the fuselage, below the axle ox ₁ , in the absence of rotation of the thrust force vector is balanced by the aerodynamic moment

with the calculated balancing position of the elevator, i.e. in the absence of a rotation of the thrust force vector, the sum of the moment of aerodynamic forces and the indicated convertible moment of the thrust of the engine will be zero. In the steady state, the additional deviation of the elevator from the specified balancing value will depend almost exclusively on the angle of rotation of the thrust force vector.

In accordance with _figure 1 and the accepted conditions, the formula for calculating the moment M _zdv. created by the thrust of the engine on the shoulder l relative to the center of mass, has the form:

For one of the considered HLL assemblies, it was obtained: l = 1.561 m, h = 0.24 m. When the thrust force vector is rotated, the moment equilibrium under balancing conditions will be satisfied even with a different elevator. The difference between this value and the balancing one will characterize the rotation of the thrust force vector. Since at small angles of rotation the change in projection onto the axis ox _{1 is} insignificant, the moment of force will depend mainly on the magnitude of the projection of force onto the axis o ₁ .

The specified form of writing equations of moments allows us to calculate the aerodynamic moment M _zaer. when the elevator deviates from zero. Thus, the problem of determining the angle of rotation of the thrust force vector in a linear approximation is solved. Relative to the indicated initial scheme, the moments below are considered when the thrust force vector is rotated. This approach to solving the problem is legitimate and allows us to simplify its solution, primarily at small angles φ.

The ability to improve the accuracy of determining rotation angles in a small range of values will compensate for the shortcomings of the algorithm described in the prototype. It is assumed that the module of the thrust force vector is determined by the readings of the accelerometer at the current time in the entire range of possible values of the angle φ.

In addition to the above, when compiling the algorithm for determining the angle of rotation φ, a large velocity head, relations between the coefficients of aerodynamic moments and the requirements for the stabilization of the angle of attack and slip are taken into account: to ensure flight at almost constant angles of attack and slip, i.e. at small angular velocities ω _x , ω _y , ω _z.

Тогда второе из уравнений (3) можно представить в виде:At high pressure heads q and small angular velocities ω _x , ω _y , ω _z, we can neglect the product ω _x · ω _y and, therefore, the moment of inertial forces. Also accepted

Then the second of equations (3) can be represented as:

.

.

Figure 3 shows the changes in parameters and characteristics in a short-period GLL motion around the center of mass; a stepped deviation of the elevator by an angle Δδ = 5 ° at t = 60 sec; deviation of the thrust force vector by an angle φ = 5 °. At the same time, the mode of continuous stabilization of the angle of attack is maintained. Such a flight regime with a slight disturbance practically does not violate the required conditions for maintaining the values of the angle of attack and the pressure head, i.e. required conditions at the inlet to the air intake. At the same time, it provides additional information on the effect of high-frequency oscillations of the angle of attack on the magnitude of the traction force. Such information can be obtained based on the readings of accelerometers at a certain orientation of the measuring axes; the rationale for this possibility was given in the prototype.

As can be seen from figure 3, under nominal conditions in steady state at t> 72 sec. the calculated value of the angle φ≈5 ° (4.85 ° ... 5.15 °), i.e. close to the specified over a significant time interval (Δt≈20 sec).

может быть использовано при послеполетной обработке в широком промежутке времени как до подачи импульса, так и в установившемся режиме после возмущения. Это позволит определять угол поворота вектора силы тяги на достаточно продолжительном интервале времени.With continuous measurement of rudder torques and angles, the calculated value

It can be used in post-flight processing over a wide period of time both before the pulse is applied and in the steady state after perturbation. This will make it possible to determine the angle of rotation of the thrust force vector over a sufficiently long time interval.

To solve the problem in a wider time range, an additional oscillatory mode in the initial section from the moment the engine is turned on can also be used. Therefore, it is advisable to start the scramjet with a delay of several seconds after separation from the accelerator, from the moment the required balancing angle of attack is established. Similarly, after the engine is finished, maintain stabilization of the set values of the pressure head and angle of attack.

As can be seen from figure 3, in the steady state ΔM _z ≈ 0. In the transition process with a constant value of the traction force, we obtain ΔM _z ≈ΔM _zaer. ; deviations of the angle of attack from the nominal modulo do not exceed 0.3 °. With small deviations of the angle of attack from the nominal value, the increment of the moment can be represented as:

, где

where

Δδ = δ _t -δ _to is the difference between the current value of the rudder and its value before applying a pulse.

The moment can be measured quite accurately using a torque sensor (in particular, DM-18, DM-20, VT-191); measurement error of moments ≈1%.

, используя расчетные значения моментных характеристик и измеренные значения скоростного напора, угла атаки и угловой скорости. При ω_z≈0 правомерно ранее указанное принятое условие

=0 и можно пренебречь составляющей момента демпфирования, в сравнении со статической составляющей от поворота руля высоты и отклонения угла атаки от расчетного балансировочного.If necessary, in case of large fluctuations, you can make a correction for the moment

using calculated values of moment characteristics and measured values of pressure head, angle of attack and angular velocity. When ω _z ≈ 0, the previously accepted accepted condition is valid

= 0 and we can neglect the component of the damping moment, in comparison with the static component from the rotation of the elevator and the deviation of the angle of attack from the calculated balancing one.

As a result, we obtain the balance of moments - aerodynamic, only from the magnitude of the angle of deviation of the steering wheel relative to the balancing value, and from the traction force of the engine when the force vector is rotated through an angle φ. In order to increase the accuracy of calculations when processing the measurement results, it is necessary to take for calculation the values of the measured parameters at the moments of maximum deviations of the rudder in the oscillation mode after disturbances.

Therefore, using the values ΔM _z , δ, q measured in a flight experiment, we can calculate the partial derivative of the longitudinal moment coefficient with respect to the elevator deflection angle as follows:

на достаточно продолжительном участке полета можно принять

≈const и использовать полученные соотношения на всем этом участке полета.Note that here the quantity Δδ is the deviation of the elevator relative to the balancing value measured at the time of the pulse. Then, in the steady state, the increment of the moment due to the deviation of the thrust force vector must be balanced by the increment of the aerodynamic moment, which can be calculated using this formula. Subject to a slow change in the derivative

on a sufficiently long flight section, you can take

≈const and use the obtained ratios throughout this portion of the flight.

It is possible to shift the point of application of the thrust force relative to the calculated point at the nozzle exit, i.e. deviation from the accepted pattern of distribution of forces. Then the converting moment will change by the value of P _dvt · Δh and, in accordance with this, the balancing value of the elevator, relative to which the increment Δδ is calculated. The change in the shoulder l, affecting the value of the angle φ, will change slightly. In particular, when the nozzle is cut at 45 °, the change in shoulder l will not exceed 2%. Consequently, the adopted force distribution system remains valid when the point of application of the traction force is displaced in a rather wide range.

The results of statistical modeling of the measuring system confirm its operability under the action of random disturbances. Errors in determining the angle of rotation of the thrust force vector for given random perturbations are characterized by the following parameters: the mean square deviation does not exceed σ _φ = 0.2 ° with the mathematical expectation m _φ = 0.2 °.

Therefore, when using the presented method, it is possible to significantly increase the accuracy of determining the angle of rotation of the thrust force vector. This, first of all, is a consequence of the fact that the rotation of the force vector even at a small angle affects primarily the magnitude of the moment. The change in the modulus of the thrust force vector is negligible. With increasing distance from the center of mass of the aircraft to the nozzle exit (arm shoulder), the magnitude of the moment and, therefore, the accuracy of calculating the angle will increase. To clarify the obtained values of the module and the angle of rotation of the thrust force vector, post-flight mathematical modeling of flight paths should be performed with the use of the obtained processing results of the on-board measurement system and the results of external trajectory measurements.

The claimed solution makes it possible to determine the angle of rotation in a flight experiment and to clarify the value of the module of the thrust force vector (from the prototype). This method can be used in testing all engines, in particular engines, in which the dependence of the direction angle of the thrust force vector on the magnitude of the pressure drop in the engine path is revealed.

The proposed method can be used to determine the vector of thrust by the results of measurements of moments in a short-period motion, deviation angles of elevators and high-speed pressure when flying aircraft with ramjet engines.

With a more accurate determination of the angle of rotation of the thrust force vector, the characteristics of the gas flow from the nozzle, the controllability reserves of the HLL in longitudinal motion can be refined, which will improve the accuracy of stabilization of the angle of attack and ultimately give recommendations when developing the engine.

Claims (1)

A method for determining the angle of rotation of the thrust force vector of a hypersonic ramjet with an oblique cut of the nozzle according to the results of its flight tests at a hypersonic flying laboratory, including the operation of measuring the overload on the HLL from the force of the engine's thrust provided that the projection of the aerodynamic force vector projection accelerometer is equal to zero , calculation of the modulus of the thrust force vector P _dt at the current time, characterized in that in the steady state flight mode GLL, in which the sum is longitudinal aerodynamic moment and moment from the engine thrust force is zero, an impulse stepwise deviation of the elevator by an angle δ _{imp is} additionally set, the value of which is selected on the basis of preliminary mathematical modeling so that the difference Δα = α-α ₀ between the current value of the angle of attack α mode oscillations and its value α ₀ at time modulo the pulse does not exceed the desired value for the experiment | Δα | <0.5 °, measured in the perturbed motion angles of the elevators and rotational deviations Suitable times for such testing points will be measured to determine the deviation of only aerodynamic torque M M _Zaer by balancing _Zaer.bal which moment is balanced by the thrust, the time deviation is determined as the difference _Zaep ΔM = M -M _{Zaep Zaep.bal,} making correction value ΔM _Zaer by

where

- derivative of the angle of attack from the coefficient of aerodynamic moment; preliminarily determined by calculation and according to the results of ground-based bench tests, the coordinates l and h of the point of application of the thrust force vector at the nozzle exit relative to the center of mass of the HLL, then the aerodynamic moment increment is calculated, which is caused only by the deviation of the steering wheel relative to its balancing value when applying a pulse and equal to the difference:
ΔM _Z ≈ΔM _Zaer -ΔM _Zα , where
ΔM _Zaer - the measured value of the aerodynamic moment in perturbed motion, then, according to the results of measurements obtained in flight tests, the derivative is calculated by the angle of deviation of the elevator from the coefficient of aerodynamic moment according to the formula:

Δδ = δ-δ ₀ is the difference between the measured values of the angle of deviation of the elevator δ in the perturbed movement and its balancing value δ ₀ measured when applying the pulse,
q is the measured value of the pressure head,
s, b - characteristic area and length of the product,
in the steady state, after applying the pulse, measure the current values of the velocity head, the elevator deflection angle and find the rotation angle φ of the thrust force vector P _dt as a result of solving the trigonometric equation:

P _dt is the value of the module of the thrust force vector at the current time t, which is determined by the readings of accelerometers,
δ _t , q _t - current elevator and velocity head in steady state after applying a pulse,
l and h are the coordinates of the point of application of the thrust force vector at the nozzle exit.

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