RU2773024C1 - Method for reproducing aerodynamic heating of aircraft elements - Google Patents

Abstract

FIELD: aircraft engineering.

SUBSTANCE: invention relates to ground tests of elements of aircraft, namely to methods for reproducing aerodynamic thermal effects on the surface of aircraft elements, for example, fairings of homing heads of aircraft missiles, antenna fairings, compartments with a rocket, in ground conditions. The proposed method for reproducing aerodynamic heating of aircraft elements by conducting a thermal experiment on a stand using a hot casing includes several cycles of heating aircraft elements with a heated air flow and radiant heat flow from the hot casing. Before conducting a thermal experiment, the air flow recovery temperature in flight and the heat transfer coefficient on the stand and in flight are calculated, provided that the heat transfer coefficient on the stand is less than the heat transfer coefficient in flight, then the first heating cycle of the aircraft elements is carried out with the stand heat transfer coefficient and the recovery temperature of the heated air flow equal to the recovery temperature in flight. The peculiarity of the proposed method is that the first heating cycle is carried out under the additional condition of a thermal experiment, according to which the temperature of the hot casing is equal to the temperature of the environment receiving thermal radiation from the aircraft elements in flight. Moreover, during the first heating cycle of the aircraft element, the surface temperature of the aircraft element is measured. Then, the second and subsequent heating cycles are carried out with a heated airflow and radiant heat flow from the hot casing at a constant stand heat transfer coefficient and recovery temperature equal to the recovery temperature in flight while maintaining the surface temperature of the hot casing, which is calculated before each current heating cycle of the aircraft element. Moreover, the thermal experiment is completed when the measured temperature of the aircraft element at the current heating cycle will differ from the measured temperature of the aircraft element at the previous heating cycle by no more than the value of the permissible measurement error or the temperatures are equal to each other.

EFFECT: increase in the accuracy and reliability of reproducing the thermal effect on the surface of the aircraft elements during aerodynamic heating while reducing the complexity of the experiment.

1 cl, 1 dwg

Description

The invention relates to ground testing of elements of aircraft (LA), and in particular to methods for reproducing aerodynamic thermal effects on the surface of aircraft elements, for example fairings for homing heads, aircraft missiles, antenna fairings, compartments with a rocket, in ground conditions.

To conduct thermal tests of aircraft elements on the ground, various installations are used: radiation heating stands, wind tunnels, thermal test stands based on fuel combustion in an air stream [Polezhaev Yu.V. etc. Materials and coatings in extreme conditions. In 3 vols. T. 3. Experimental studies. MSTU im. N.E. Bauman, 2002].

In the practice of ground tests, stands for radiation heating and / or heating with tape heaters are widely used, especially for the case of thermal tests at a given temperature field of the object being tested or at a given value of the heat flux entering the aircraft elements in flight [patent RU 2571442 "Method of thermal testing of rocket fairings from non-metallic materials”, patent RU 2676385 “Method of controlling heating during thermal testing of antenna radomes of missiles”].

A known method of thermal testing of the instrument compartment of the aircraft [RU 2526406 "Method of thermal testing of the instrument compartment of the aircraft"], in which the thermal experiment is carried out in two stages. At the first stage, a thermal test of a fragment of the full-scale heat-insulating package of the instrument compartment is carried out in a thermal chamber with a thermal load corresponding to the flight one, maintaining the calculated temperature values on the outer surface of the heat-insulation package, while creating boundary conditions for heat transfer on the inner surface of the heat-insulating package, simulating the conditions of heat removal from the housing shell into the instrument compartment. compartment. Then, according to the measured values of the temperatures of the inner surface of the heat-insulating package, a graph of the temperature dependence of the body of the instrument compartment is obtained from time to time. At the second stage, the body of the instrument compartment is heated without thermal insulation in accordance with the previously obtained temperature change schedule and at the same time measuring the temperatures of the gaseous medium and the instrument compartment equipment that produces heat in accordance with the flight sequence diagram.

A common disadvantage of these methods is the need to pre-calculate the temperature of the aircraft elements at the points under study before conducting a thermal experiment, which, in the absence of information about the thermophysical characteristics of the material, cannot be carried out with the required accuracy.

Also, to determine the actual temperature field of the aircraft elements, tests are carried out in supersonic wind tunnels with the provision of modes as close as possible to flight ones [Baranov A.N. et al. Static tests for the strength of supersonic aircraft. - M. Engineering, 1974.]. This type of reproduction of flight thermal conditions is laborious, requires large labor costs, financial costs and time to obtain results.

A known method of modeling the parameters of the environment during aerodynamic heating of the elements of an aircraft (LA), including heat-shielding materials, in ground conditions, while providing conditions for conducting a thermal experiment similar to those in flight. [CN 109029907 "Parametric similarity method for the conditions of an experiment on modeling an aerodynamic thermal environment"].

The method includes calculating the in-flight airflow recovery temperature and the in-flight heat transfer coefficient using known formulas. The heat flux density is then determined based on the calculated in-flight heat transfer coefficient. The surface temperature of an aircraft element in flight is determined according to the laws of heat transfer, then, in the process of an iterative process, due to the variable heat transfer coefficient on the bench, calculated before the start of the experiment based on the method of computational fluid dynamics, the flow recovery temperature on the bench is controlled, by which the heat flux density on the bench is determined.

The disadvantage of this method is that in order to calculate the surface temperature of the aircraft element, it is necessary to have information about the thermophysical characteristics of the material of the aircraft element. And also due to the fact that the airflow recovery temperature on the stand is lower than the airflow recovery temperature in flight, so it is impossible to simulate real flight conditions, and therefore the temperature field of the aircraft element will be unreliable.

, равного полетному

[Афанасьев В.А. Экспериментальная отработка космических летательных аппаратов М., МАИ, 1994].There is a known method of reproducing the thermal conditions of an aircraft at supersonic flight speeds with a subsonic flow of heated air with the placement of the object under study in a specially profiled channel (casing), which ensures that the heat flow enters the rocket

equal to the flight

[Afanasiev V.A. Experimental development of spacecraft M., MAI, 1994].

; (1)

; (one)

(2)

(2)

where

– коэффициент теплоотдачи на стенде и в полете;

is the heat transfer coefficient on the stand and in flight;

– температура восстановления потока на стенде и в полете;

is the flow recovery temperature on the stand and in flight;

– температура поверхности ракеты на стенде и в полете;

– temperature of the rocket surface on the stand and in flight;

– приведенная степень черноты поверхности ракеты на стенде и в полете;

- the reduced emissivity of the rocket surface on the stand and in flight;

σ is the Stefan-Boltzmann constant.

,

,

.The equality of temperatures at similar points of the aircraft on the stand and in flight is achieved during tests by this method at

,

,

.

,

, что фактически требует создания аэродинамической трубы, обеспечивающей достижение чисел Рейнольдса и Маха, равных полетным, что является чрезвычайно сложной технической задачей при высокой стоимости.The disadvantage of this known method of thermal testing is the need to ensure equality of heat transfer coefficients and recovery temperatures at

,

, which actually requires the creation of a wind tunnel that ensures the achievement of Reynolds and Mach numbers equal to the flight ones, which is an extremely complex technical task at a high cost.

A known method for determining the temperature field of aircraft elements during aerodynamic heating according to patent RU 2739524, which includes several cycles of heating aircraft elements with a heated air stream. Before conducting a thermal experiment, the airflow recovery temperature in flight and the heat transfer coefficients on the stand and in flight are calculated so that the heat transfer coefficient on the stand is less than the heat transfer coefficient in flight. The first heating cycle is carried out at a heat transfer coefficient equal to the heat transfer coefficient on the stand, and the temperature of the recovery of the heated air flow, equal to the temperature of the recovery of the air flow in flight, and subsequent heating cycles are carried out at a constant heat transfer coefficient and the calculated temperature of the recovery of the heated air flow. Moreover, the time of each cycle of heating the aircraft element is equal to the specified flight time, while the thermal experiment ends when the measured temperature of the surface of the aircraft element on the current cycle of heating the aircraft element will differ from the measured temperature of the surface of the aircraft element on the previous cycle of heating the aircraft element by no more than allowable error of the measurement system or temperature will be equal to each other.

без учета влияния на температурное поле теплового излучения от окружающей среды с температурой

, что приводит к погрешности в найденных температурах объекта испытаний. Недостатком данного способа моделирования аэродинамического нагрева является также необходимость создания потока с температурой

, превышающей

, что недопустимо при ограничениях воздействующих температур.The implementation of this method is characterized by the fact that the test sample is affected by a heat flux

without taking into account the influence on the temperature field of thermal radiation from the environment with temperature

, which leads to errors in the found temperatures of the test object. The disadvantage of this method of modeling aerodynamic heating is also the need to create a flow with a temperature

exceeding

, which is unacceptable when the operating temperatures are limited.

The presented method in technical essence is the closest to the claimed invention and can act as a prototype.

The technical problem to be solved by this invention is the creation of a method for reproducing the aerodynamic heating of aircraft elements during flights at supersonic speeds, which ensures the determination in bench conditions with high accuracy of the temperature field of aircraft elements, for example, rocket fairings, with previously unknown thermophysical characteristics of the material (thermal conductivity , specific heat capacity, thermal diffusivity).

The technical result, to which the claimed invention is directed, is to increase the accuracy and reliability of the reproduction of the thermal effect on the surface of the aircraft elements during aerodynamic heating while reducing the complexity of the experiment.

The claimed technical result is achieved by implementing a method for reproducing the aerodynamic heating of aircraft elements by conducting a thermal experiment on a stand using a hot casing in which an aircraft element is placed, which includes several cycles of heating aircraft elements with an air flow and radiant heat flow from a hot casing, in which before conducting a thermal experiment, the air flow recovery temperature in flight and the heat transfer coefficient on the stand and in flight are calculated, provided that the heat transfer coefficient on the stand is less than the heat transfer coefficient in flight, then the first heating cycle of the aircraft elements is carried out at the stand heat transfer coefficient and the temperature of recovery of the heated air flow equal to the airflow recovery temperature in flight.

равна температуре окружающей среды, воспринимающей тепловое излучение от элементов ЛА в полете, причем в процессе первого цикла нагревания элемента ЛА измеряют температуру поверхности элемента ЛА, затем проводят второй и последующие циклы нагревания подогретым воздушным потоком и лучистым тепловым потоком от горячего кожуха при неизменном стендовом коэффициенте теплоотдачи и температуре восстановления, равной температуре восстановления в полете с поддержанием температуры поверхности горячего кожуха, которую рассчитывают перед каждым текущим циклом нагревания элемента ЛА по формуле: A feature of the proposed method is that the first heating cycle is carried out under the additional condition of conducting a thermal experiment, according to which the temperature of the hot casing

equal to the temperature of the environment that perceives thermal radiation from aircraft elements in flight, and during the first cycle of heating the aircraft element, the surface temperature of the aircraft element is measured, then the second and subsequent heating cycles are carried out with heated air flow and radiant heat flow from the hot casing at a constant bench heat transfer coefficient and a recovery temperature equal to the recovery temperature in flight while maintaining the surface temperature of the hot casing, which is calculated before each current cycle of heating the aircraft element according to the formula:

(3)

(3)

based on the surface temperature of the aircraft element, which is measured in the previous cycle of heating the aircraft element, taking into account the temperature of the hot casing obtained in the previous heating cycle,

where

n – aircraft element heating cycle by air flow and radiant heat flow from the hot casing, n=1…N;

– температура горячего кожуха на n-м цикле нагревания элемента ЛА;

is the temperature of the hot casing on the nth cycle of heating the aircraft element;

– температура горячего кожуха на (n-1) цикле нагревания элемента ЛА;

- temperature of the hot casing on (n-1) cycle of heating the aircraft element;

– температура восстановления в полете;

– in-flight recovery temperature;

– разница между коэффициентами теплоотдачи в полете и на стенде;

is the difference between the heat transfer coefficients in flight and on the stand;

– температура наружной поверхности элемента ЛА, измеренная на (n-1) цикле нагревания элемента ЛА;

- temperature of the outer surface of the aircraft element, measured on (n-1) cycle of heating the aircraft element;

σ - Stefan-Boltzmann constant (5.67 * 10 ^-8 W / (m ² * K ⁴ );

– приведенная степень черноты,

- reduced degree of emissivity,

moreover, the thermal experiment is completed when the measured temperature of the aircraft element in the current heating cycle differs from the measured temperature of the aircraft element in the previous heating cycle by no more than the allowable measurement error or the temperatures are equal to each other.

Thus, when testing according to the proposed method, the equality of temperatures on the bench and in flight is achieved by the method of successive approximations by compensating for the reduced intensity of the convective heat flux (the heat transfer coefficient on the bench is less than the flight one) by increasing the radiant heat flux from the heated casing, while the equality of the heat flux entering on an aircraft element, for example, a rocket fairing on the stand and in flight. The calculation of the required jacket temperature is based on the measured surface temperature of the rocket fairing in the previous heating cycle at a known recovery temperature, flight and bench heat transfer coefficients, and the reduced emissivity of the surface of the aircraft element.

Figure 1 shows a diagram of the installation for reproducing the aerodynamic heating of the tested aircraft sample, which is represented by a rocket fairing, where 1 is a hot casing; 2 – rocket fairing; 3 - air blower with heater; 4 – temperature sensors (resistance thermocouples, thermocouples); 5 – system of registration of parameters; 6 - infrared lamps; 7 - cooling system; 8 – heating control system; 9 - base; 10 - power floor.

The reproduction of the aerodynamic heating of the aircraft elements according to this invention is carried out as follows (Fig. 1).

The thermal experiment is carried out on a thermal bench, so that a test sample is placed in the hot casing (1), which can be a rocket fairing (2), which is subjected to several cycles of heating by convection from air using an air blower with a heater (3) and radiant heat flow from a hot casing (1), the temperature of which is changed by infrared lamps (6) with a cooling system (7). Based on the data taken from the temperature sensors (4) at the control points, the parameter recording system (5) sends a control action to the heating control system (8).

The air blower (3) can be turbojet engines, gas-plasma heaters, liquid jet engines, electric arc heaters, ramjet engines, etc.

The rationale for the proposed method of thermal testing is based on the following.

It should be noted that the rocket fairing is considered as a test sample.

и стендовое (экспериментальное)

свяжем с помощью ряда Тейлора:Flight (desired) temperature value of the rocket fairing

and bench (experimental)

connect using the Taylor series:

(4)

(four)

where

- температура обтекателя ракеты при нагреве его потоком с стендовым коэффициентом теплообмена

, при

,

, первое приближение к

(первый цикл нагревания),

- температура окружающей среды;

is the temperature of the rocket fairing when it is heated by a flow with a bench heat transfer coefficient

, at

,

, the first approximation to

(first heating cycle),

- ambient temperature;

;

;

и

и лучистым тепловым потоком с

и

- первое приближение к

, описывается системой уравнений:The temperature field of the rocket fairing on the stand when it is heated by an air flow with

and

and radiant heat flux with

and

- first approach

, is described by the system of equations:

; (5)

; (5)

; (6)

; (6)

; (7)

; (7)

; (8)

; (eight)

where

c, ρ(r) are the heat capacity and thermal conductivity of the rocket fairing material;

R is the outer radius of the rocket element (rocket fairing);

r is the transverse coordinate;

- температура поверхности обтекателя ракеты.

- surface temperature of the rocket fairing.

и выполнив ряд преобразований с учетом, что вторым приближением

к

будет:Differentiating the system (4-7) with respect to

and performing a number of transformations, taking into account that the second approximation

to

will be:

(9)

(9)

we obtain a system of equations in the form:

(10)

(ten)

(11)

(eleven)

; (12)

; (12)

; (13)

; (13)

where

(14)

(fourteen)

It follows from equality (9) and system (10-14) that this system describes the temperature field of the rocket fairing in the second approximation (the second heating cycle).

и температурой

, температурой кожуха

(14), получим температурное поле

(в том числе и

) во втором приближении.Therefore, heating the test object with a flow with

and temperature

, casing temperature

(14), we obtain the temperature field

(including

) in the second approximation.

In the general case, repeating the above methodology n times, we arrive at a system of equations:

(15)

(fifteen)

(16)

(16)

; (17)

; (17)

; (18)

; (eighteen)

where

(19)

(19)

The system of equations (15-19) determines the natural temperature field of the rocket fairing in the (n-1)-th approximation.

Thus, the natural temperature field of the rocket fairing during testing is determined as follows:

,

,

, в результате эксперимента определяется температура обтекателя ракеты в первом приближении

; - the first heating cycle (out of n cycles, n=1…N) with

,

,

, as a result of the experiment, the temperature of the rocket fairing is determined in the first approximation

;

,

и

, в результате эксперимента определяется температура обтекателя ракеты во втором приближении

;- second heating cycle with

,

and

, as a result of the experiment, the temperature of the rocket fairing is determined in the second approximation

;

,

и

, в результате эксперимента находится температура обтекателя ракеты в n-ом приближении к

.- nth heating cycle with

,

and

, as a result of the experiment, the temperature of the rocket fairing is found in the nth approximation to

.

The number of rocket fairing heating cycles is determined from the condition that the measured temperature of the rocket fairing surface in the current rocket fairing heating cycle will differ from the measured rocket fairing surface temperature in the previous heating cycle by no more than the allowable error of the measurement system or the temperatures will be equal to each other.

Claims (12)

A method for reproducing aerodynamic heating of aircraft (AC) elements by conducting a thermal experiment on a test bench, which includes several cycles of heating the aircraft elements with a heated air flow, in which, before conducting a thermal experiment, the air flow recovery temperature in flight and the heat transfer coefficient on the test bench and in flight are calculated provided that the heat transfer coefficient on the bench is less than the heat transfer coefficient in flight, then the first cycle of heating the aircraft elements is carried out at the bench heat transfer coefficient and the recovery temperature of the heated air flow equal to the air flow recovery temperature in flight, characterized in that the thermal experiment is carried out using a hot casing, in which the aircraft element is placed, for additional heating of the aircraft elements by radiant heat flux, provided that the first heating cycle is carried out at a temperature of the hot casing equal to the ambient temperature, receiving thermal radiation from aircraft elements in flight, and during the first cycle of heating the aircraft element, the surface temperature of the aircraft element is measured, then the second and subsequent heating cycles are carried out with heated air flow and radiant heat flow from the hot casing at a constant bench heat transfer coefficient and a recovery temperature equal to in-flight recovery temperature while maintaining the temperature of the hot casing surface, which is calculated before each current cycle of heating the aircraft element according to the formula:

based on the surface temperature of the aircraft element, which is measured in the previous cycle of heating the aircraft element, taking into account the temperature of the hot casing obtained in the previous heating cycle, where n is the cycle of heating the aircraft element with air flow and radiant heat flow from the hot casing, n=1…N;

- temperature of the hot casing on the n-th cycle of heating the aircraft element;

- temperature of the hot casing on (n-1) cycle of heating the aircraft element;

- in-flight recovery temperature;

- the difference between the heat transfer coefficients in flight and on the stand;

- temperature of the outer surface of the aircraft element, measured on (n-1) cycle of heating the aircraft element; σ - Stefan-Boltzmann constant (5.67 * 10 ^-8 W / (m ² * K ⁴ );

- reduced degree of emissivity, moreover, the thermal experiment is completed when the measured temperature of the aircraft element in the current heating cycle differs from the measured temperature of the aircraft element in the previous heating cycle by no more than the allowable measurement error or the temperatures are equal to each other.

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